1565results about "Jet-propulsion engine testing" patented technology

The invention relates to the technical field of combustion monitoring, and discloses a combustion instability boundary prediction system based on complex environment condition analysis, and the system comprises a multi-source sensing module, a data fusion module, an anti-interference processing module, an instability feature library, an intelligent decision engine and a digital twinborn body. A dynamic environment interference compensation mechanism is constructed, spectrum correction parameters are configured according to different environment working conditions during real-time prediction of a combustion instability boundary, the problem of spectrum distortion caused by environment disturbance on combustion dynamic parameters is eliminated in real time, spectrum interference deviation of a combustion measurement position can be automatically sensed and corrected, and real-time prediction of the combustion instability boundary is realized. The accuracy of combustion instability sensitive feature extraction is guaranteed, the instability early warning error is reduced, a flame pose real-time tracking system is deployed, fuel control parameter self-adaptive adjustment is triggered immediately when it is detected that the pose is abnormal, autonomous correction of the combustion instability feature recognition position is achieved, and it is guaranteed that the feature capture space is accurately positioned.

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 discloses a two-degree-of-freedom clamp capable of adjusting the bending moment and direction of a turbine blade and a test method. The two-degree-of-freedom clamp comprises a blade clamping component, an upper adjustable U-shaped head, a lower adjustable U-shaped head, an upper clamping rod and a lower clamping rod. The blade clamping component is in threaded connection with the upper adjustable U-shaped head and the lower adjustable U-shaped head, the other end of the upper adjustable U-shaped head and the other end of the lower adjustable U-shaped head are connected with the upper clamping rod and the lower clamping rod respectively, and the other end of the upper clamping rod and the other end of the lower clamping rod are connected with an upper clamping head and a lower clamping head of the fatigue testing machine respectively. The upper clamping rod and the upper adjustable U-shaped head are adjustably connected in pairs in the transverse direction intersecting with the vertical axis so as to adjust the eccentricity of the turbine blade load. The eccentric angle of the blade load can be adjusted by adjusting the rotation cycle of the blade clamping component. According to the clamp disclosed by the invention, by adjusting the eccentricity and the eccentricity angle during a turbine blade test, different bending moment directions and relative sizes of different sections during the turbine blade test can be realized, so that the aim of simulating a real load is fulfilled.

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.

An inspection method is provided during which a head of an inspection scope is inserted into an interior of a powerplant. The head of the inspection scope includes an actuator. The powerplant includes a component located within the interior of the powerplant. The head of the inspection scope is abutted against a surface of the component within the interior of the powerplant. The head of the inspection scope is preloaded against the surface of the component using a preload device located outside of the interior of the powerplant. Vibrations are induced in the component using the actuator. A vibratory response in the component excited by the vibrations is measured using a sensor to provide sensor data.

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 method, a device and equipment for monitoring the degradation state of a turboshaft engine and a medium, and aims at overcoming the defects of an existing method for sensing the degradation state of the turboshaft engine based on limited measurement parameters. A degeneration factor range linear solving method based on a degeneration factor fingerprint map theory is introduced on the basis of a traditional particle swarm optimization algorithm, and a more accurate and reasonable particle optimization limitation range is provided for the particle swarm optimization algorithm under the condition that available system input parameters are insufficient. For any sensor actual acquisition main performance parameter degradation situation combination, a corresponding specific optimization range can be provided, and the situation that degradation factor optimization falls into local optimization in a plurality of possible optimization spaces due to insufficient data complexity is avoided. Therefore, degradation state monitoring of the turboshaft engine is optimized, and the overall performance degradation trend of the engine is accurately obtained.

The invention relates to the technical field of aero-engine rotor vibration tests, in particular to an aero-engine rotor unsteady-state thermally induced vibration test system which comprises a rotor experiment table, an air inlet and exhaust system, a lubricating oil system, a protection system and a heating system. The rotor experiment table comprises a first supporting system, a second supporting system, a rotor outer protection casing, a gas compressor disc, a gas compressor sealing disc, a turbine disc, a turbine front sealing disc, a front shaft diameter, an inter-disc drum, a rear shaft diameter, a first squirrel-cage elastic support, a ball bearing, a second squirrel-cage elastic support and a rolling rod bearing. The first supporting system is composed of a first squirrel-cage type elastic support and a ball bearing, meanwhile, an air inlet is formed in the first supporting system, and the ball bearing in the first supporting system is connected with the front shaft diameter in an interference fit mode. Rotor unsteady-state characteristics considering heating conditions are finally created, and vibration response and durability of a wheel disc structure under the unsteady-state thermal load condition are studied in combination with test conditions such as temperature vibration and the like.

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 state monitoring method and device and a computer readable storage medium, and relates to the technical field of aero-engine state monitoring. The method comprises the following steps: acquiring a magnetic flux variable quantity, an optical signal, a thermal radiation signal and a charge variable quantity through a composite sensor, and converting the magnetic flux variable quantity, the optical signal, the thermal radiation signal and the charge variable quantity into coil output voltage and optical signal voltage, and converting into thermal radiation exitance voltage signals and variable voltage under different wavelengths; obtaining a blade tip gap measured value according to the coil output voltage and the distance value between the probe and the blade tip; based on the blade tip timing principle, blade vibration displacement is obtained according to the optical signal voltage; performing colorimetric processing on the thermal radiation exitance voltage signals under different wavelengths to obtain a radiation exitance ratio at different wavelengths, and further obtaining the temperature of a turbine blade temperature field; and calculating the casing acceleration according to the changing voltage. In conclusion, state parameters in multiple aspects are obtained at the same measuring point through one composite sensor, and the engineering testing steps are simplified.

The invention relates to the technical field of aero-engines, and particularly provides a variable cycle engine duct experiment system and experiment method. A variable cycle engine duct experiment system comprises: a first combined experiment member comprising a main flow valve and a shunt ring; the second combined experiment piece comprises a first outer culvert, a second outer culvert and an ejector; the inlet of the variable capacitor second external culvert experiment piece is connected with the external culvert outlet of the shunt ring, and the outlet is connected with the second external culvert; an inlet of the first gas collection tank is communicated with an inner culvert outlet of the shunting ring; the bypass ratio adjusting valve is used for adjusting the bypass ratio at the shunting ring; an inlet of the second gas collection tank is connected with an outlet of the first gas collection tank through an inner culvert pressurization simulator, and an outlet is communicated with the first outer culvert; an inlet of the variable capacitor confluence duct experiment piece is communicated with the second outer duct; and the outer culvert throttle valve is connected with an outlet of the variable capacitor confluence culvert experiment piece. According to the technical scheme, engine duct experiments in multiple modes can be achieved, and the experiment efficiency is improved.

The invention relates to the field of aero-engine testing, in particular to an aero-engine rotating speed signal generator and a control method, and the system comprises the steps that a main control module receives a test demand input by a user, selects and determines the model, matching frequency, voltage and phase parameters of a tachometer sensor needing to be simulated based on the test demand, and transmits the parameters to a DSP processor; the DSP processor receives parameters issued by the main control module and generates an SPWM signal through a timer; the timer is configured to be in a circulation mode, and the period of the PWM signal is changed by changing the initial value of the timer; the inversion module converts the isolated SPWM waveform into a three-phase alternating current signal and outputs the three-phase alternating current signal; and the detection module measures data of the three-phase alternating current signal output by the inversion module in real time and feeds back the data to the main control module to form closed-loop control. Test errors caused by performance errors of the sensor are avoided, and the actual performance of the tachometer indicator can be evaluated more accurately.

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 floating ring seal test device and a detection system, the device comprises a shell and a test part, the shell is internally provided with an accommodating cavity, the test part comprises a rotating shaft, two bearing parts and a test part, the rotating shaft is sleeved with the two bearing parts at intervals in the axial direction, and the test part is arranged between the two bearing parts; the rotating shaft penetrates through the containing cavity in the horizontal direction and is installed on the shell through two bearing pieces. The test piece is arranged in the containing cavity and comprises a sealing runway arranged on the rotating shaft in a sleeving mode and two floating ring pieces installed on the shell, the sealing runway comprises two annular protruding plates arranged at the two axial ends of the test piece respectively, and the two floating ring pieces and the two annular protruding plates are arranged in a one-to-one correspondence mode. The floating ring piece is used for performing non-contact sealing with the outer wall of the annular convex plate; and the two floating ring pieces, the sealing runway and the shell are enclosed to form a test cavity for introducing a test medium. Two floating ring pieces can be tested at the same time under the same working condition, the test efficiency is remarkably improved, and the failure risk of a test part is reduced.

The invention discloses an aerospace equipment performance evaluation method based on digital test data fusion, and belongs to the field of electronic engineering and computer science, and the method comprises the steps: designing a physical test and a digital test, selecting related indexes of aerospace equipment, and carrying out corresponding physical and digital tests around the related indexes of the aerospace equipment; performing preprocessing and multi-source data integration on the physical test data; performing data quality evaluation and virtual data division on the digital test data; and performing double fusion on the result through an adaptive fusion algorithm to obtain a performance evaluation result. The advantages of the digital test and the physical test can be fully exerted, and comprehensive and accurate evaluation of the aerospace equipment performance is realized through deep fusion of the digital test data and the physical test data.

The invention provides an impeller high-temperature over-rotation test device. The impeller high-temperature over-rotation test device comprises a long shaft, an impeller, a press-fitting assembly, a buffer assembly and a fixing piece. The long shaft comprises a first shaft body and a second shaft body which are concentric and different in diameter, the impeller is sleeved on the first shaft body, and press-fitting parts are formed at two ends. The first press-fitting assembly and the second press-fitting assembly are pressed on the press-fitting parts on the two sides of the impeller correspondingly, relative rotation is prevented through the rectangular keys, the double press-fitting assemblies are symmetrically arranged to balance axial force, and rotation stability is improved. The left / right clamping sleeve is matched with the conical surface of the left / right taper sleeve, and a tensioning wire groove is formed to adapt to impellers of different sizes and compensate size changes. The buffering assemblies are arranged in a mirror symmetry mode, the horn mouth design optimizes stress distribution, long-axis thermal expansion is absorbed, and vibration impact is resisted. The fixing piece provides axial fixing force, and the axial rigidity of the system is optimized. The device can ensure the reliable positioning of the impeller at high temperature and ultrahigh speed, and improves the test precision and reliability.

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 discloses a multichannel thermocouple temperature compensation method for a liquid rocket engine test, and relates to the technical field of liquid rocket engine tests, and the method comprises the steps: determining the type of a target thermocouple temperature sensor according to the characteristics of a tested object in a to-be-tested liquid rocket engine, and determining a corresponding temperature compensation wire; a target thermocouple temperature sensor is connected with a temperature compensation wire to form an acquisition device; a thermocouple voltage signal and a thermistor voltage signal are acquired in real time through an acquisition device, and the signals are transmitted to a processor based on a wiring terminal in the temperature compensation cabinet; and the thermocouple temperature corresponding to the input data is determined through a preset cold end compensation algorithm in the processor. According to the invention, the precision and reliability of temperature measurement can be improved, the system structure is simplified, the system cost is reduced, and the system is suitable for temperature measurement requirements of multi-channel, high-precision and complex environments in a liquid rocket engine test.

An inspection method is provided during which a distal end of an inspection probe is inserted into an interior of a powerplant. The inspection probe includes a body and a head pivotally connected to the body. The head includes an actuator, and the head is disposed at the distal end of the inspection probe. The powerplant includes a component within the interior of the powerplant. The head is arranged with the component. The arranging includes pivoting the head relative to the body and abutting the head against a surface of the component. Vibrations in the component are induced using the actuator. A vibratory response in the component excited by the vibrations is measured using a sensor to provide sensor data.

The invention relates to the technical field of aero-engine state monitoring and fault diagnosis, in particular to a fault monitoring method and device for an aero-engine and a storage medium. The method comprises the steps that vibration signals of all parts in the working process of the aero-engine are obtained; dividing the vibration signal into a steady-state signal, a periodic impact signal or an unsteady-state transient signal based on the short-time energy and the zero-crossing rate; when the vibration signal is a steady-state signal, determining whether a fault exists or not by adopting a time domain feature extraction and classification model; when the vibration signal is a periodic impact signal, Fourier transform and envelope spectrum analysis are adopted to determine whether a fault exists; and when the vibration signal is an unstable transient signal, determining whether a fault exists or not by adopting time-frequency analysis. Compared with a single analysis algorithm, the method has the advantages that different algorithms are matched for fault monitoring based on the characteristics of the vibration signals, the monitoring accuracy is improved, and the false alarm rate is reduced.

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 relates to the technical field of aero-engine combustion chamber performance testing, in particular to a three-dimensional chromatography testing system and method for the ignition process of an aero-engine combustion chamber. The device is characterized by comprising a guide rail arranged on the periphery of an igniter of a combustion chamber, a plurality of imaging units are arranged on the guide rail, each imaging unit is composed of an imaging lens, a flexible optical fiber bundle and a four-degree-of-freedom sliding block, a high-speed camera is arranged corresponding to the flexible optical fiber bundle, and the high-speed camera is connected with a synchronous controller through a signal line. And the synchronous controller is connected with the igniter through a signal line. By introducing the flexible optical fiber bundle and the synchronous controller, the number of the high-speed cameras is reduced to one, so that the test cost is reduced, the problem of synchronous control of the high-speed cameras and the igniter is solved, and the synchronism of the high-speed chromatographic measurement test bed to the ignition process is guaranteed. Meanwhile, the imaging lens can be placed on the guide rail at any angle through the sliding block, and test operation is facilitated.

Disclosed is a cart (2) for transporting an aircraft engine, comprising a base frame (4) equipped with wheels (6) for riding on a floor; at least two engine arms (14, 16) extending horizontally and movable vertically relative to the base frame, structured and designed for supporting the aircraft engine; and at least two adapter arms (20) extending horizontally at a higher level than the at least two engine arms, structured and designed for supporting an adapter coupled at the top of the aircraft engine. Disclosed is also a method of transporting an aircraft engine to a test cell.

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 relates to a self-adaptive engine double-duct flow test device, which comprises an air supply system A used for providing air flow in a simulated engine duct and passing through a front fan, and a duct test section C used for simulating an inner duct, an outer duct and an adjusting structure of a double-duct fan engine and carrying out a flow test between the ducts, the total pressure control system B is arranged between the air supply system A and the duct test section C; a flow matching and adjusting simulation method between the double ducts is provided, the motion law of the valve is continuously adjusted and dispersed into a plurality of steady-state working conditions for test measurement, or the motion process of the valve is dynamically measured, so that various test requirements are met. According to the invention, the flow area of the duct throat part is continuously adjustable, and the duct throat part is cooperatively matched with a duct inlet total pressure control system, so that the stability of the inlet condition of a test section is ensured, and meanwhile, the flow characteristics among multiple ducts under various different matching and adjusting working conditions are successfully reproduced; and important support is provided for deeply exploring a coupling mechanism and a flow evolution rule between the ducts.