1091results about "Gas-turbine engine testing" patented technology

The invention relates to a multi-dimensional high-resolution turbine blade heat-flow coupling field synchronous measurement method and system, and the method comprises the steps: obtaining a target spot image of known space distribution in a measurement space, and building a three-dimensional space mapping function; carrying out particle matching on the velocity field measurement original image based on a three-dimensional space mapping function, and reconstructing a three-dimensional particle trajectory by adopting a self-adaptive grid algorithm combined with machine learning to obtain a high-resolution three-dimensional velocity vector field; acquiring a temperature field measurement original image, performing three-dimensional space registration according to a three-dimensional space mapping function, further calculating voxel-level temperature distribution, and inverting a three-dimensional temperature field; and carrying out coupling analysis on the three-dimensional velocity vector field and the three-dimensional temperature field, and outputting a three-dimensional heat-flow coupling cloud picture. According to the invention, multi-dimensional and high-precision synchronous measurement of the temperature field and the velocity field in the complex flow field of the turbine blade is realized, the spatial resolution and dynamic response capabilities are remarkably improved, and a reliable experimental basis is provided for analysis and optimization of the cooling performance of the blade.

The invention discloses a creep test system of a turbine rotor blade film hole simulation piece and a test method of the creep test system. Comprising a clamping assembly used for fixing a test piece, a loading assembly used for applying a load to the test piece, a strain measuring assembly used for measuring the strain rate of the test piece, a heating assembly used for covering the test piece to form a heating channel, and a cooling assembly used for manufacturing and conveying cooling airflow to flow through the interior of the test piece, the strain measurement assembly is used for measuring the external temperature of the test piece and the internal temperature of the test piece, the control assembly is connected with the strain measurement assembly, the heating assembly, the cooling assembly and the temperature measurement assembly, and the heating assembly is used for manufacturing and conveying high-temperature airflow to flow through the heating channel so as to heat the outside of the test piece to a first set temperature; the cooling assembly is used for cooling the interior of the test piece to a second set temperature, and the heating assembly and the cooling assembly are used for forming a dynamic convection environment around the test piece.

The invention belongs to the technical field of aero-engine high-temperature part test, and particularly relates to a thermal mechanical composite fatigue test device in a turbine blade rotation environment, which comprises a high-temperature gas system, a stator system, a rotor system, a support system, a test system, an auxiliary system and a turbine rotor blade test piece, the high-temperature fuel gas system is used for providing a fuel gas environment for a turbine rotor blade test piece and controlling the high-low temperature circulation state of fuel gas, a combustion chamber is arranged at the front end of the high-temperature fuel gas system and provided with a double-oil-way nozzle, a fuel gas flowing system is arranged at the rear end of the high-temperature fuel gas system, and a blade grid structure is designed for the fuel gas flowing system according to the geometrical shape of a turbine rotor blade. And a spray cooling section is arranged at the rear part of the test section. The device is composed of six core systems including a high-temperature fuel gas system, a stator system, a rotor system, a supporting system, a testing system and an auxiliary system, all the systems work cooperatively, and integrated simulation of the high-temperature fuel gas environment, rotating centrifugal loads and heat-force synchronous circulation is achieved.

The invention discloses an engine fault diagnosis method, device and equipment and a storage medium, and relates to the technical field of fault diagnosis. The method comprises the following steps: synchronously acquiring a sound signal and a vibration signal of an engine in a normal operation state, and performing feature extraction on the sound signal and the vibration signal to obtain corresponding target features; the target features comprise time domain features, frequency domain features and time-frequency domain features; determining a statistical index of each target feature through statistical characteristic analysis, and performing feature fusion on all the target features according to a target fusion mode based on the statistical indexes to obtain a fusion index; and an early warning threshold value is determined according to the data statistical characteristics of the fusion index, so that fault monitoring is carried out on the engine by using the early warning threshold value. Fault diagnosis is carried out according to the early warning threshold value determined according to the sound vibration data of the engine in the normal operation state, high dependence on fault samples is avoided, and the comprehensiveness of fault analysis is improved.

The invention provides an aero-engine test bed. The aero-engine test bed comprises a rack and an auxiliary support, the auxiliary support comprises a mounting platform, a mounting seat, a connecting piece, a supporting arm, a field force limiting mechanism and a controller; the mounting platform is arranged on the rack; the two mounting seats are oppositely arranged on the mounting platform; the connecting piece is arranged between the two mounting seats and is opposite to the mounting platform; the connecting piece is configured to move in the vertical direction; the supporting arm corresponds to the mounting seat; the two ends of each supporting arm are hinged to the corresponding mounting base and the corresponding connecting piece correspondingly. The supporting arm is provided with an installation joint used for being connected to an aero-engine. The field force limiting mechanism comprises a first component and a second component which are oppositely arranged; the first component is arranged on the mounting platform, and the second component is arranged on the connecting piece; and the controller is electrically connected with the field force limiting mechanism. Through cooperative work of the field force limiting mechanism and the controller, the supporting rigidity of the auxiliary support can be dynamically adjusted according to the real-time rotating speed of the aero-engine in the test run process.

An inspection system detects defects within a component using eddy current thermography. An infrared camera detects heat from a defect within the component generated by the eddy current interfering with the defect. The eddy current is generated in the component by a magnetic field applied to the component by an inductor assembly. The inductor assembly includes a magnetic core having magnetic core material and two arms, each arm including a coil to generate the magnetic field. The inductor assembly is configured to apply the magnetic field at an angle offset from a horizontal axis of the magnetic core such that the eddy current flows at an angle within the component relative to the offset angle. Further, the magnetic core material can be shaped to apply the magnetic field at the offset angle or to accommodate complex component shapes.

A design method for a blade with orientation structures, a blade and a performance test method for a blade are provided, which relate to the technical field of blade design, and are universal to blades whose parameters are completely or partially known. In the design method, orientation awls are provided at a trailing edge of a blade from a variable-pitch propeller; orientation awls are provided at a trailing edge of a blade from a fixed-pitch propeller or a fan blade from a jet engine, and orientation grooves are provided on a back of the blade. Meanwhile, during providing orientation awls for either blade, the blade is stretched in a chord direction to ensure that an area of a pressure surface of a modified blade is equal to an area of a pressure surface of its original.

The invention discloses a micro-scale turbine blade boundary layer heat-flow double-field synchronous test method and system. The test method comprises the steps that temperature field initial data and velocity field initial data are acquired; based on the initial data of the temperature field and the initial data of the velocity field, the distribution state of the microparticles in the boundary layer is diagnosed, and according to the diagnosis result, the optimal distribution of the microparticles in the boundary layer is obtained by feeding back and controlling the operation parameters of the reverse cyclone separation system; based on the optimal distribution, an improved PIV algorithm is adopted to calculate a boundary layer velocity field, and the improved PIV algorithm is obtained by introducing multistage grid adaptive encryption, physical constraint conditions and a transition area recognition mechanism into the PIV algorithm; a boundary layer temperature field reconstruction model is established, depth phase coupling is carried out on a temperature field reconstruction result and a boundary layer velocity field, synchronous analysis of heat-flow coupling behaviors in a turbine blade boundary layer is achieved, and finally a boundary layer coupling cloud picture reflecting heat-flow interaction characteristics is obtained.

The invention relates to an aero-engine vibration fault early warning method based on working condition self-adaption. The method comprises the steps that vibration signals and rotating speed data in the operation process of an aero-engine are synchronously collected through a vibration sensor and a rotating speed sensor; calculating an instantaneous rotating speed change rate based on the collected rotating speed data, and determining the working condition of the aero-engine according to the instantaneous rotating speed change rate; in the transient working condition freezing fault early warning mechanism, when the working condition of the aero-engine is a steady-state working condition, rotating speed data are discretized into N partitions, and vibration time sequence characteristics are predicted by adopting trend prediction models of the corresponding partitions to obtain predicted values; and calculating an absolute residual error between the characteristics of the collected vibration signal of the aero-engine and the predicted value, and carrying out double-window dynamic baseline management and graded early warning based on the absolute residual error. The aero-engine fault detection method can improve the aero-engine fault detection accuracy.

The invention discloses an aero-engine compressor outlet static pressure sensor fault diagnosis method, which comprises the steps of constructing an aero-engine general model, correcting parameters in the aero-engine general model according to test run data of an aero-engine, and obtaining a special model of an engine of a specified model; according to the special model, calculating a residual value of a gas compressor outlet pressure measured value and a gas compressor outlet pressure simulation value of the engine of the specified model; according to the technical requirements of the engine of the specified model, a residual error threshold value of the outlet pressure of a gas compressor is set; and when the absolute value of the residual error value is greater than the residual error threshold value, diagnosing that the static pressure sensor at the outlet of the gas compressor is faulty, and triggering an alarm signal. According to the method, a special model is constructed through an engine mechanism, diagnosis is performed in combination with actually measured data, the model precision is improved, rapid identification of early faults is realized, dynamic changes of the engine under different working conditions are adapted, and the robustness of a diagnosis result is improved.

The invention discloses a device and a method for simulating heat exchange and friction resistance measurement of an aviation high-speed motor under different working conditions, and relates to the field of heat management of high-speed rotating equipment. An additional heating sheet is attached to the outer surface of the inner rotor, and different surface structures are detachably assembled on the outer rotor to form a stator and rotor slit flow channel boundary; cooling liquid forms one-inlet and one-outlet forced convection through the radial through hole of the end cover. The rotating speed larger than or equal to 20,000 rpm, runner pressure drop and temperature data are synchronously collected through an axle load type sensor, and the heat exchange coefficient, the flow resistance coefficient and the friction resistance coefficient are calculated in real time based on energy conservation, a Darcy formula and a torque empirical formula. A real rotor surface structure is reproduced under the condition of high-speed rotation, working medium leakage is restrained through mechanical sealing, and runner blockage caused by centrifugal deformation of the surface structure is prevented through the outer rotor designed in an integrally-formed mode; the obtained convective heat transfer-resistance coefficient-friction resistance coefficient all-parameter curved surface database can guide optimization of the cooling flow channel of the aviation motor.

The invention relates to the technical field of aeroengine turbine blade thermodynamic performance testing, in particular to a two-dimensional multi-mode turbine blade heat-flow double-field synchronous testing method and system.The method comprises the steps that phosphorescent particles are evenly scattered in an experimental flow field of a target turbine blade, and after the phosphorescent particles are excited through double lasers, the target turbine blade is subjected to heat-flow double-field synchronous testing; collecting a flow field image; inverting a temperature field by using a phosphorescence intensity ratio method based on the flow field image, and calculating a velocity field through an improved PIV algorithm; and after time-space registration is carried out on the temperature field data and the velocity field data, fusion is carried out to generate a composite heat-flow field, dynamic coupling characteristics are analyzed through a vortex center tracking algorithm, and the heat-flow double-field synchronous test of the target turbine blade is completed. Through ultraviolet / PIV dual-laser excitation, multispectral separation acquisition and high-precision data fusion, time-space synchronous measurement of the temperature field and the velocity field of the boundary layer of the turbine blade is realized.

The invention provides a diffuser inlet speed distribution simulation device and full-ring combustion chamber test equipment. The diffuser inlet speed distribution simulation device comprises a current limiting ring; the flow limiting ring is provided with a plurality of openings evenly distributed in the circumferential direction of the flow limiting ring. The opening is symmetrical relative to the radial direction of the flow limiting ring, and on the section perpendicular to the axial direction of the flow limiting ring, the opening is provided with an opening outer edge, an opening inner edge and two opening side edges connected with the opening outer edge and the opening inner edge; the outer edge of the opening is a straight line perpendicular to the radial direction of the flow limiting ring or an arc line with the center of the flow limiting ring as the center, and the distance between the two end points of the outer edge of the opening in the circumferential direction of the flow limiting ring is a first distance a. The inner edge of the opening is a straight line perpendicular to the radial direction of the flow limiting ring or an arc line with the center of the flow limiting ring as the center, and the distance between the two end points of the inner edge of the opening in the circumferential direction of the flow limiting ring is a second distance b. The maximum distance between the edges of the two opening sides in the direction perpendicular to the radial direction of the flow limiting ring is a third distance c; and cgt; a, and cgt; b.

The invention discloses a single-sensing double-blade tip timing vibration speed measurement method, system, medium and equipment, and the method comprises the steps: only installing a blade tip timing sensor in the circumferential direction of a rotor blade case, and obtaining a blade arrival time sequence generated when rotor blades sequentially pass through the blade tip timing sensor; calculating the instantaneous rotating speed of the rotor based on the arrival time intervals of all adjacent rotor blades; the blade tip linear velocity between the two rotor blades is extracted by combining the circumferential included angle and the measuring point radius of any two rotor blades in the blade disc; removing a trend term related to the rotating speed from the blade tip linear velocity to obtain a velocity quantity reflecting the real vibration characteristics of any two rotor blades; a two-degree-of-freedom dynamic model is constructed by combining prior information of a blade Campbell diagram, and double-blade vibration parameter identification is realized by adopting a nonlinear least square fitting method.

The invention discloses a crack strong-robustness time-frequency compressed sensing monitoring method, system, medium and equipment based on inter-blade tip timing, and the method comprises the steps: determining the installation number and installation positions of blade tip timing sensors, and carrying out the circumferential installation of the blade tip timing sensors on a rotor blade case; the blade tip timing sensors are used for obtaining a time sequence when the rotor blades reach the positions of the blade tip timing sensors; based on each time sequence and the rotor blade included angle, obtaining an inter-blade tip coupling vibration speed measured by each blade tip timing sensor; an undersampling vibration speed sparse representation model is constructed based on the inter-blade tip coupling vibration speed, dynamic frequencies of adjacent rotor blades are decoupled and extracted by using a sparse reconstruction algorithm, and non-baseline monitoring is performed on a crack propagation process according to the evolution of a vibration frequency difference between the blades.

The invention discloses a fault diagnosis system for an aviation turboshaft engine, and relates to the technical field of fault diagnosis. An acquisition module acquires horizontal, vertical and axial vibration signals; the processing module is used for generating real-time feature vectors by introducing dynamic short-time Fourier transform, variational mode decomposition and singular value decomposition of an unfixed-width Gaussian window; the incremental diagnosis module calculates the matching degree between a real-time feature vector or a prediction feature vector and a known class through a dynamic expansion network, selects to belong to the known class or triggers an expansion mechanism according to an adaptive resonance theory, constructs joint loss by using a Bayesian learning framework after expansion, and performs Gaussian process modeling; loss is determined through Bayesian optimization, and network parameters are updated; the prediction module generates a prediction feature vector when there is no fault; the maintenance module sends or stores a maintenance scheme according to needs, self-adaptive recognition of known classes and new fault classes is achieved, the state can be pre-judged in advance, the operation safety of the engine is effectively guaranteed, and the maintenance cost is reduced.

The invention discloses a high-altitude test bed ejection system and an adjusting and parking control method thereof, relates to the technical field of aero-engines, and solves the problems that in the prior art, the control strategy is single, the abrupt change of cabin pressure during parking is easily caused, and all parts of the ejection system bear huge pressure impact. The micro-cabin pressure regulating device comprises a high-altitude cabin unit, an exhaust silencing unit and a multi-stage ejector unit, and accurate micro-cabin pressure regulating can be achieved through mutual cooperation of the three units. In addition, according to the adjusting method, the priority sequence of the ejectors at all levels is strictly followed, the next level is started after the previous level reaches the limit, the stability of the adjusting process is ensured, physical limit and a stable waiting period are further set in the process, and the adjusting feasibility under the limiting working condition is ensured; according to the parking method provided by the invention, a real-time pressure rate detection-pause-air supply buffer closed loop mechanism is introduced, so that accurate control and self-adaptive adjustment of the pressure in the high-altitude cabin and stable parking of the high-altitude cabin are realized.

The invention relates to the technical field of aero-engine vibration control, and discloses an on-line automatic balance system for a turboprop engine propeller rotor, which comprises a signal acquisition and analysis module, an automatic balance state monitoring and control module and an automatic balance actuator, the signal acquisition and analysis module is used for monitoring and acquiring an original vibration signal of the rotor system in real time to obtain vibration monitoring information, and transmitting the vibration monitoring information to the automatic balance state monitoring and control module; the automatic balance state monitoring and control module is used for receiving the vibration monitoring information and outputting a driving control instruction to the automatic balance actuator according to the vibration monitoring information; the driving control instruction comprises the moving direction and the stepping amount of the balancing weight; the automatic balance actuator comprises two balance weight discs, and the balance weight discs are used for moving according to the driving control instruction so as to carry out vibration suppression. According to the invention, the supporting stability and the assembling economy are improved.

The invention provides a five-hole probe correction method, which is used for correcting a five-hole probe in a cascade test, a left side pressure measurement hole of the corrected five-hole probe is a first hole, a central pressure measurement hole is a second hole, a right side pressure measurement hole is a third hole, a lower pressure measurement hole is a fourth hole, and an upper pressure measurement hole is a fifth hole. The correction method comprises the following steps: S1, acquiring a calibration curve of the five-hole probe, and acquiring total pressure distribution of a cascade outlet; s2, calculating to obtain a corrected pressure measurement value of the first hole and a corrected pressure measurement value of the third hole; s3, obtaining a corrected yaw angle and a corrected pitch angle by using the calibration curve; and S4, obtaining a corrected total pressure value and a corrected static pressure value by using the calibration curve. According to the invention, the problem of measurement distortion of angle measurement and pressure measurement of the probe calibration curve obtained under the condition of uniform incoming flow in wake measurement with non-uniform actual circumferential total pressure is solved, and the airflow angle in the wake of the cascade can be measured more accurately.

The invention belongs to the technical field of aero-engine health management, and particularly relates to an intelligent diagnosis system and method for an aero-engine. Comprising an engine, a multi-physical field sensor network installed on the engine, a real-time feature extraction module deployed at an edge computing node, a central diagnosis server integrated with a machine learning engine, and a historical knowledge base for storing structured fault cases and health management. Through the diagnosis mechanism, the health condition of the engine is monitored in real time, and the fault diagnosis efficiency is improved by more than 50%.

The invention discloses a circumferential pull rod rotor test bench and method. The circumferential pull rod rotor test bench comprises a grooved platform, a driving motor, a load device and a circumferential pull rod rotor, wherein the driving motor and the load device are mounted on the grooved platform; the circumferential pull rod rotor comprises a plurality of gas compressor wheel discs, a plurality of turbine wheel discs and an intermediate shaft; a plurality of gas compressor pull rod bolts penetrate through the gas compressor wheel disc, and the gas compressor wheel disc is connected in a pre-tightening manner through the gas compressor pull rod nuts; a plurality of turbine pull rod bolts penetrate through the turbine wheel disc, and the turbine wheel disc is connected in a pre-tightening mode through turbine pull rod nuts. The test bed further comprises a gas compressor casing and a turbine casing, and the gas compressor casing and the turbine casing are both installed on the grooved platform and are adjustable in position. Dynamic characteristics and various rotor faults of the circumferential pull rod rotor of the gas turbine can be effectively simulated, and the test bench is reliable in structure and stable in operation.

The present disclosure discloses a method for detecting anomalies in acroengine variable working conditions based on a multi-band wavelet network, including the steps of: collecting monitoring signals of an acroengine in a healthy state under a plurality of working conditions and corresponding working condition information; constructing and training a multi-band wavelet network using the signals and working conditions; calculating a state monitoring index of the signal using the trained network; determining a monitoring threshold according to the state monitoring index; and determining whether or not the engine is abnormal by comparing the state monitoring indexes of the monitoring signals of the aeroengine under the plurality of working conditions to the monitoring threshold. The present disclosure fully considers the advantages and disadvantages of the signal processing technique with the deep learning network for fault diagnosis.

The invention discloses an aircraft engine journal multi-type load composite fatigue test system. The invention discloses an aircraft engine journal multi-type load composite fatigue test system which comprises a supporting frame. The outer ring surface of the torque disc is provided with two symmetrical torque actuators, and the center of the torque disc is axially connected with an axial force actuator; the bending moment disc is mounted between the torque disc and the upper adapting section of the shaft neck, and two symmetrical bending moment actuators are mounted at the two ends of the bending moment disc; the heater is arranged outside the shaft neck, the bottom is connected with the bottom frame through the heater lower support, and the upper end is connected with the heater upper cover plate; the lower end of the journal is connected with the horizontal base through a lower switching section, and the upper end is connected with the bending moment disc through an upper switching section; the shaft neck is filled with an upper filling section and a lower filling section. By adopting an up-and-down and inside-and-outside laminated structure, combined load loading of axial force, dynamic bending moment, main torque, dynamic torque, high temperature and high pressure (replacing centrifugal load) can be realized, so that a test system is high in integration degree and high in index capability, and can be suitable for multi-type load combined fatigue tests of similar parts.

The invention belongs to the field of aerospace sensors, and discloses an intelligent wide-range pressure measurement system and method for an aero-engine. The system comprises a pressure acquisition module which comprises at least two pressure sensors with different measuring ranges and overlapped measuring intervals, and an electromagnetic valve group used for switching gas circuits of the pressure sensors; the intelligent control module communicates with an engine FADEC to obtain real-time working condition parameters, predicts the pressure change trend in combination with a pre-stored pneumatic model, switches gas circuits in advance before the pressure reaches a range critical point, triggers a special collection mode to lock an adaptive sensor under a transient working condition, and collects data at the highest sampling rate. The measurement method comprises the steps of working condition parameter acquisition, pressure trend prediction, advanced range switching and self-diagnosis calibration, and through multi-sensor data comparison under environment pressure or stable working conditions, temperature drift errors are dynamically compensated, and faults of the sensors and transmission pipelines are accurately diagnosed. According to the scheme, the problems of switching delay, data distortion, sensor precision drift and difficulty in fault discovery of a traditional system are effectively solved, high precision and stability of full-range pressure measurement are guaranteed, and reliable support is provided for precise regulation and control and flight safety of an aero-engine.

A method (100) and system (10) for evaluating a component includes receiving (102) a component definition that includes component geometry and one or more of a boundary condition set and a load set. A comparator model (34) compares (106) component geometry to one or more design envelopes (32A), and compares load sets to one or more prior load sets, each from a database of stored component definitions (32F, 32G). Based on the comparison, an analysis model (36) forms (108) a detailed analysis or an iterative analysis to produce a result set. A results module (22) generates (110) an evaluation disposition of the component definition based on the result set.

The application provides an aero-engine high-altitude test bed intake and exhaust system, and belongs to the technical field of intake and exhaust systems, and comprises the following steps: obtaining a flowchart of preset test conditions and an intake and exhaust system, and selecting a suitable test cabin as a first test cabin based on the preset test conditions; performing data analysis based on the preset test conditions to obtain a target working condition, and calculating the operation of each valve based on the target working condition; monitoring the actual operation working condition of the test cabin in real time, and obtaining feedback information based on the real-time detection result; inputting the feedback information into a feedback-regulation scheme mapping table, determining different regulation schemes corresponding to different feedback information, and adjusting the operation working condition based on the regulation scheme. The test cabin provides different temperature and flow rate intake conditions for the engine or engine components in the test cabin, so that different test working conditions are simulated, and the simulation demand is further met.

This invention relates to the field of spectroscopic measurement technology, and particularly to a measurement system and method for measuring the concentration of high-temperature combustion gases in aero-engines. The measurement system includes a supercontinuum laser source, a first collimating mirror, a second collimating mirror, a spectrometer, and a computer. The measurement method includes acquiring the actual transmittance spectrum of the high-temperature combustion gases; calculating the theoretical transmittance spectrum of the high-temperature combustion gases; constructing constraints; fitting the actual transmittance spectrum and the theoretical transmittance spectrum into the constraints; and calculating the concentration of the high-temperature combustion gases. The measurement system has a lightweight structure, requiring only holes for mounting the laser emission and absorption devices on the inner wall of the aero-engine casing. The measurement method uses a supercontinuum laser source, which, compared to ordinary lasers used in active gas detection, offers better spatial coherence, a wider spectral adjustment range, and higher intensity, enabling the detection of more types of gases. The method uses a multi-band combustion gas absorption model to solve for the temperature of the high-temperature combustion gases, improving the accuracy of the concentration calculation.

Variable pitch fan blade (8a) for an unfaired turbomachine fan, the blade (8a) comprising an aerodynamic blade (9) and a base (10), the blade (9) comprising a recess opening simultaneously onto the upper and lower surfaces (14) and a leading edge (15) of the blade (9), the blade (8a) further comprising a measuring device (26) housed in the recess and clipped onto the blade (9), the measuring device (26) comprising a streamlined body (28) whose outer surface (29) is complementary to the outer surface (30) of the blade (9) to ensure aerodynamic continuity, the body (28) carrying at least one measuring instrument (31) configured to measure at least one parameter of an airflow upstream of the blade (9). Figure for the abstract: 2

The invention discloses a static detection method of an aero-engine pulsating pressure testing system and a related device, and belongs to the technical field of aero-engine pulsating pressure testing. The method comprises the following steps: carrying out integrity and electrical characteristic detection on a sensing unit; checking the operation state of the signal processing unit; a background signal is collected in a static environment, and whether the overall performance of the system meets a preset requirement is judged according to the background signal. The system comprises a sensing unit detection module, a signal processing unit detection module and a background signal detection module. By comprehensively detecting the sensor state, circuit connection and system background noise under the condition of no dynamic excitation, potential faults can be found in advance, parking troubleshooting caused by data abnormity in the test run process is avoided, and the stability and operation and maintenance efficiency of the test system are remarkably improved.

A kind of turbine component transition state test import flow detection method, design utilizes the steady total pressure of turbine import before transition state starts, dynamic total pressure, the steady static pressure of turbine import before transition state starts, dynamic static pressure, and the dynamic total pressure of turbine import at time under transition state, dynamic static pressure, the corrected total pressure of turbine import at time under transition state, corrected static pressure are calculated to obtain, and then the Mach number of turbine import at time under transition state, flow coefficient are calculated, combined with the steady flow of turbine import before transition state starts, the steady flow of turbine import after transition state terminates, the conversion flow coefficient of turbine import before transition state starts, after transition state terminates is calculated, the flow coefficient of turbine import at time is calculated by steady dynamic mapping interpolation, finally the flow of turbine import at time is calculated.