836results about "Jet-propulsion engine testing" patented technology

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.

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 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 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.

The present application relates to the field of gas turbine technology, specifically to a sand supply device for gas turbine engine sand dust deposition experiment; including box, air intake system, air outlet system, stirring system and dust concentration measuring instrument; air intake system is arranged in the lower part of the box; the air outlet system is arranged in the upper part of the box, and the air outlet system contains height-adjustable air outlet; the input end of the air outlet is provided with an air outlet inlet flow diverter; the stirring system is installed at the bottom of the box, including servo motor, motor support and stirring impeller. According to the target particle concentration, the height of the air outlet inlet flow diverter is adjusted as the main means, and the air inlet flow and speed of the air inlet and the stirring impeller speed are fine-tuned, and through the coupling effect of the three, the high, medium and low concentration levels can be continuously adjusted, covering the engine full working condition sand dust simulation demand, solving the problem of single adjustment mode and insufficient precision of traditional device.

The invention provides an aero-engine turbine interstage casing thermal-mechanical coupling loading test device, which comprises a mechanical loading test tool clamp, a temperature field control system and a test assembly, and is characterized in that the test assembly is used for simulating a turbine interstage casing assembly and is installed on the mechanical loading test tool clamp; the mechanical loading test tool clamp is used for composite loading of radial force and axial force of the test assembly. The temperature field control system comprises a cooling system and a heating system, and the cooling system and the heating system enable the test assembly to reach an expected target temperature field by controlling the temperature to adjust the current of the heater and the flow of cold air. The device has the advantages of being simple in structure, flexible to install, high in control precision, easy to operate and the like, and can be used for controlling the deformation discordance of the turbine interstage casing supporting plate and applying the mechanical load of the casing, so that an interstage casing outer ring stress-strain assessment point reaches the working state of an engine, and the working efficiency of the engine is improved. And checking whether the strength design of the inter-stage casing meets the expectation or not.

The systems and methods of the invention pertain to analyzing steam generator tube data for the detection of wear. Further, the invention is capable of performing a comparison of current tube signal data to baseline or historic tube signal data, e.g., from previous and / or the first, in-service inspection of the steam generator. The systems and methods are automated and can generate results to show potential tube-to-tube contact wear areas as well as the progression of tube-to-tube gap reduction within a steam generator tube bundle. In certain embodiments, the invention is capable of comparing current and historical eddy current data to determine the difference that may be related to degradation or other interested phenomena, and of processing and trending historical comparison results to establish normal variance and detect abnormal variances.

The application discloses a fuel step rapid injection device and injection method, and belongs to the field of aero-engine whole machine test, and aims at solving the problems of complex system, high cost, low injection precision and slow response. The device comprises a servo motor, a screw pair, an oil injector, first and second remote switch valves, first and second check valves and a safety valve. The servo motor drives the oil injector piston rod through the screw pair, the oil injector rod cavity is connected with the low-pressure fuel system through the second remote switch valve, and is connected with the engine fuel manifold through the first remote switch valve and the first check valve. A bypass connected with the fuel control device through the second check valve is arranged between the first check valve and the fuel manifold. In normal operation, the fuel control device supplies fuel to the fuel manifold through the second check valve; in step injection, the second remote switch valve is opened, the piston rod is extended to suck oil, then the piston rod is closed, the first remote switch valve is opened, and the piston rod is quickly retracted to inject fuel into the manifold. The application can be used in aero-engine surge test.

The application provides a device for testing a rotating part in a high-temperature environment, comprising: a test section for providing a simulated high-temperature environment for the rotating part, and provided with a first through hole and a second through hole; a first tension assembly for pulling an end of the rotating part extending out of the first through hole; a second tension assembly for pulling an end of the rotating part extending out of the second through hole; a first sleeve connected outside the first through hole and sleeved outside the first tension assembly, and provided with at least one first air inlet channel; a second sleeve connected outside the second through hole and sleeved outside the second tension assembly, and provided with at least one second air inlet channel; and an air pressure source connected to the at least one first air inlet channel and / or the at least one second air inlet channel and providing an air pressure greater than or equal to the environmental pressure in the test section. The application can accurately simulate the thermal-mechanical coupling loading environment in which the rotating part is in actual work, and effectively improve the reliability and accuracy of the test results.

Device (1) for characterizing wear on a gas turbine engine comprising: - an abradable material (8) having a wear end (80) adapted to be in contact with a gaseous fluid flow (F) flowing from upstream to downstream of the gas turbine engine when the gas turbine engine is running, the abradable material (8) being electrically conductive; - a magnetic field generator (7) adapted to emit a magnetic field in the abradable material (8); - an electronic board (10) configured to calculate information enabling the characterization of the wear of the gas turbine engine from data relating to a physical response of said abradable material (8) to said magnetic field. Figure for the abbreviation: Figure 2.

This invention provides an attitude-adjustable aero-engine test bench, relating to the technical field of aero-engine test benches. The test bench includes: a mounting plate; two support frames symmetrically arranged on the mounting plate, the tops of which are rotatably connected to the main mounting joints on both sides of the aero-engine; and at least one first telescopic assembly disposed between the two support frames, one end of which is hinged to the mounting plate, and the other end of which is rotatably connected to a secondary mounting joint above the aero-engine's exhaust nozzle. When it is necessary to adjust the angle of attack or attitude angle of the aero-engine, the length of the first telescopic assembly is adjusted to allow the aero-engine to rotate around the main mounting joint, changing the engine's tilt angle and simulating the engine's state during UAV flight to meet experimental requirements.

This invention relates to the field of engine technology and discloses a noise reduction method for engine turbochargers. The method involves obtaining an initial value for the backflow gap between the guide ring and the compressor housing; setting n improved gap values ​​based on the initial value; adjusting the height of the guide ring according to the improved gap values ​​to fabricate the nth guide ring; using the nth guide ring to fabricate a turbocharger and testing it; acquiring operating data of the turbocharger under different operating conditions; determining whether the operating data meets noise requirements; if it meets the noise requirements, the nth guide ring is the optimal guide ring; if it does not meet the noise requirements, the test is repeated. Therefore, this invention's engine turbocharger noise reduction method, based on the initial turbocharger structure, plans an improved structure and conducts noise tests on the improved turbocharger to determine the structure that meets noise requirements, thereby solving the problem of turbocharger noise exceeding limits. Furthermore, this invention has the advantages of simple operation, low cost, and ease of implementation.

The invention provides a thermal shock test device and a test method, and belongs to the technical field of aero-engine hot end component detection. The test device comprises a test rack body, a heating assembly, a displacement control assembly and a cooling assembly, the test rack body is provided with a placing space for the heating assembly, the displacement control assembly and the cooling assembly; the heating assembly is provided with a channel opening; the displacement control assembly is arranged on the test rack body and is used for moving a test piece into the heating assembly through the channel opening or moving the test piece out of the heating assembly; and the cooling assembly is arranged on the test rack body and is used for cooling the test piece pulled out of the heating assembly. The test device can be used for carrying out thermal shock test on a test piece.

The present invention provides a method for detecting the current state of damage of a machine (1), comprising acquiring a plurality of first pressure signals caused by a combustion process and at least one second pressure signal caused by the combustion and a blade rotation process, transforming the plurality of first pressure signals and the at least one second pressure signal from the time domain into the frequency domain to achieve a plurality of transformed first pressure signals and at least one transformed second pressure signal, subtracting at least one of the plurality of the first averaged pressure signal from the at least one second averaged pressure signal to obtain a subtracted pressure signal and determining the current status of the machine based on a cepstrum analysis of the subtracted pressure signal by comparison to values obtained from an undamaged machine. The present invention further comprises a corresponding system and a computer readable storage medium.

The application discloses an aero-engine parameter-oriented comprehensive early warning system and method, relates to the technical field of aero-engine early warning, and comprises the following steps: acquiring multiple-channel operation parameters of an aero-engine, performing uniform time reference resampling, dividing a sliding window, establishing a stage environment bin identification, calculating a sliding window representative quantity of a consistent parameter pair, and establishing a fleet reference table; calculating a logarithmic ratio residual, generating a sliding window residual list, calculating a robust deviation quantity, and generating a local sliding window evidence record; acquiring a local comprehensive evidence strength according to the local sliding window evidence record, calculating a persistence index, and generating a single-machine decision record; gathering a same-bin fleet snapshot table, calculating a mechanism commonality proportion, generating a cooperative conclusion record, and acquiring a batch disposal list. The application improves the accuracy and environmental adaptability of parameter anomaly detection and improves the fine management level and risk prevention and control capability of aero-engine groups.

The utility model provides a test installation device and test platform for a gas turbine test piece, which comprises an installation platform, a test transfer box and a lifting assembly, and is characterized in that the installation platform is provided with an accommodating cavity with a top opening; the test transfer box is provided with a first position and a second position, the test transfer box is located in the containing cavity at the first position and located outside the installation platform at the second position, the test transfer box is slidably connected with a top opening of the containing cavity, the lifting assembly is connected with the containing cavity, and the lifting assembly is connected with the installation platform. And the lifting assembly is in transmission connection with the test transfer box, so that the test transfer box is driven by the lifting assembly to be switched between the first position and the second position. According to the utility model, high-altitude installation operation of a test interface and nearby switching requirements of a test point are considered, installation and maintenance are convenient, and occupied space is small.

An apparatus, comprising an optical probe having optical components therein. A gas path router receives an airflow from a gas path of a gas turbine engine. The gas path router is configured to route the airflow from the gas path past the optical components of the optical probe to purge debris from the optical components.

Plant monitoring method that uses a Mahalanobis distance calculated from data of several variables indicating the state of a plant, the method comprising: a step (S2) of capturing initial data, which are data from a past first period (T1) up to a current point in time; a prediction step (S6) of predicting second data, which are data from a second period (T2) after the current time; and a unit space generation step (S8) of generating a unit space which is a basis for calculating the Mahalanobis distance, based on the first data and the second data, in the prediction step (S6) the second data are predicted on the basis of third data, which are the data of a third period (T3) obtained by shifting the first period (T1) a prescribed time into the past, fourth data, which are the data of a fourth period (T4) obtained by shifting the second period (T2) a prescribed time into the past, and the first data, where the prescribed time is longer than the second period (T2).

This invention discloses a foreign object impact detection device for aero-engine fan blades, relating to the field of aero-engine performance testing technology. The technical solution includes a detection unit, a foreign object launching unit, and a linkage unit. The detection unit includes an electric conductor connector, a camera connector, a high-speed detection camera, and a controller connector. The foreign object launching unit includes a feeding mechanism, a low-speed launching mechanism, a high-speed launching mechanism, and a cylinder. The linkage unit can switch the connection between any launching mechanism and the cylinder. The cylinder can store energy for the connected launching mechanism. After storing energy, the low-speed and high-speed launching mechanisms can launch the material from the feeding mechanism. The linkage unit can synchronously drive the controller connector to reciprocate. Through its overall structural design, this invention can not only detect the impact of different impact velocities on fan blades but also provide more accurate detection results.

The application discloses a kind of gas analyzer's calibration device and gas analysis system, calibration device includes several groups of calibration module, each calibration module includes controller, and three-way regulating valve equipment and calibration unit are connected with controller communication;Different calibration module is used to calibrate different categories of standard gas;Calibration unit, three-way regulating valve equipment and gas analyzer are sequentially communicated by pipeline;Controller is used to generate first control instruction when calibration, based on first control instruction control the valve opening of corresponding calibration unit;Controller is also used to generate second control instruction when calibration, based on second control instruction drive corresponding three-way regulating valve equipment switches to target working state, to make standard gas that calibration unit exports pass three-way regulating valve equipment and be delivered to corresponding gas analyzer, using standard gas to calibrate gas analyzer and carry out calibration operation.The device ensures the measurement accuracy of analyzer and the accuracy of test result.

The invention provides a convection heat exchange device for casing heating, which comprises a lower rotation stage, a lower air inlet stage, a test casing section, an upper exhaust section and an upper rotation stage, the test casing section is arranged between the upper exhaust section and the lower air inlet section, the upper rotation stage is connected to the upper part of the upper exhaust section, and the lower rotation stage is connected to the lower part of the lower air inlet section; a heat convection inner sleeve is further mounted on the inner wall between the lower rotating stage and the upper rotating stage, the heat convection inner sleeve comprises an inner layer, an outer layer and a flow guide plate, a heat preservation cavity is formed between the inner layer and the outer layer, a heat convection cavity with a spiral ascending channel is formed between the inner layer and the test casing section, and the flow guide plate is arranged between the outer layer and the heat convection cavity. The flow guide plate is fixed on the outer surface of the inner layer in a spiral rising mode, by the adoption of the convection heat exchange device for heating the casing, the purpose of evenly heating the casing can be achieved, and the problem that the casing is not evenly heated by hot air is solved.

This application belongs to the field of aero-engine technology, and specifically relates to a method and apparatus for simulating aircraft bleed air during sustained test of a turbofan engine, wherein the inlet temperature T1 and inlet pressure P1 of the engine are preset to a range; Step S2: Calculate the aircraft bleed air volume W fy The range, the total bleed air temperature T of the aircraft fy The range and total bleed air pressure P fy Range; Step S3: Calculate the cold diameter D of the bleed air nozzle. fy Range; Step S4: Select the cold-state diameter D of the bleed air nozzle fy An aircraft bleed air volume verification test was conducted to obtain the bleed air volume W for the test. fyr When the induced draft volume W in the experiment fyr The requirements are met, and the cold-state diameter D of the expiratory air nozzle is adopted. fy When the induced draft volume W in the experiment fyr If the requirements are not met, return to step S4; ensure that the calculation error of the aircraft bleed air volume is within the required range, and obtain the actual bleed air volume of the engine in various states to the maximum extent through test program design. This test method has the advantages of simple test method, low test cost and high replacement efficiency.

The invention belongs to the field of aviation lubricating oil system accessory structure design and test verification, and discloses a lubricating oil heat dissipation ejector with an improved ejection ratio and a test method and device.The test device comprises a front air duct, a lubricating oil radiator, a to-be-verified lubricating oil heat dissipation ejector and a rear air duct; a front air duct and a rear air duct are respectively arranged in front of and behind a lubricating oil radiator, a lubricating oil radiating ejector is arranged on the lubricating oil radiator, a sealing gasket is covered outside the lubricating oil radiating ejector, and a hot edge inlet pipeline is arranged at an inlet of the lubricating oil radiating ejector; driving fluid enters through the front air duct, and mixed air flows out through the rear air duct; and the hot side inlet pipeline and the rear air duct are provided with flowmeters. According to the optimization and improvement design and test verification method for the lubricating oil heat dissipation ejector, a process route combining integrally-cast titanium alloy forming with high-smoothness inner wall machining is adopted, 3D printing special-shaped air duct test verification and a heat-flow coupling simulation analysis technology are matched, and it is ensured that the product can meet the requirement of an aero-engine.