207 results about "Tip clearance" patented technology

The invention discloses an engine fault detection method, device and equipment based on a single sensor and a medium, and relates to the technical field of fault detection. The method comprises the following steps: acquiring an eddy current original signal and an optical fiber pulse signal of the optical fiber eddy current composite sensor; the optical fiber eddy current composite sensor is installed at the position of an aero-engine casing. Calculating a blade radial displacement signal according to the eddy current original signal, and calculating a blade circumferential displacement signal according to the optical fiber pulse signal; filtering the blade circumferential displacement signal to obtain a target circumferential displacement signal; resampling the radial displacement signal of the blade to obtain a time sequence blade tip clearance signal; and engine blade fault information is determined according to the target circumferential displacement signal and the time sequence blade tip gap signal. Thus, the eddy current original signal and the optical fiber pulse signal are directly measured through the single sensor to detect faults, the transmission path is shortened, the diagnosis sensitivity and accuracy and the measuring point utilization rate are improved, and the sensor arrangement complexity is reduced.

Disclosed in the present invention is a turbine transition state blade tip clearance estimation method based on a long short-term memory neural network. The method comprises: first, using data concatenation and introduction of new parameters to construct multi-transition-state process transient data; then using sensitivity analysis to screen for feature parameters having a great influence on a blade tip clearance, and carrying out dimension reduction on the influence parameters; and on the basis of a long short-term memory neural network, constructing a mapping model between each feature parameter and the blade tip clearance, and building an efficient and high-precision turbine transition state blade tip clearance estimation model. By means of the turbine transition state blade tip clearance estimation model, turbine blade tip clearances of multiple transition state processes can be efficiently and accurately estimated.

The invention belongs to the technical field of design of a cold-state blade tip gap between an aero-engine rotor blade and a cartridge receiver, and particularly relates to a method for determining the cold-state blade tip gap between the aero-engine rotor blade and the cartridge receiver. According to the method, the steady-state point is calculated, including the overall flexural deflection of the rotor, the local variation caused by rotor and stator machining runout, the rotor bearing clearance and the like, and transition state calculation is added while steady-state point calculation is performed, so that the actual use scene of the aero-engine can be better considered, and the long-term use requirement is met.

An operating method is provided during which a command is received to increase thrust generated by a propulsor rotor from a first thrust level to a second thrust level. The propulsor rotor is operatively coupled to an engine core and an electric machine. The engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The engine core is operated in a transient state to increase power output from the engine core to the propulsor rotor from a first power level to a second power level in response to the command. The electric machine is operated to boost the power output from the engine core to the propulsor rotor while the engine core is operating in the transient state. A clearance control system for the engine core is operated based on the operation of the electric machine.

The invention discloses a calculation method of a fan blade tip gap accelerated flow analytical model, and belongs to the technical field of fan speed field calculation. By analyzing the aerodynamic characteristics of a fan wake flow gap acceleration area, the correlation between the power increase of the side-by-side fan and the fan wake flow gap acceleration area is disclosed, and an analytical model is constructed according to the self-similarity of the gap acceleration area and an eddy current equation. According to the model, local accelerated flow caused by rotation of adjacent rotor blade tips under the same plane is considered, by adopting the method, the used model has high precision in the aspect of predicting the flow field characteristics of the blade tip wake flow gap acceleration area, the prediction effect is close to that of a high-fidelity numerical simulation method, but the calculation cost is far lower than that of traditional CFD simulation; the method is well applicable to processing and simulating wind power plant structures which are closely arranged along the y-z same plane, such as a double-head floating fan and a fan wall, and shows good engineering application value.

The invention discloses a cross-scale multi-physics-field coupling aero-engine warming-up process analysis method. The method comprises the steps that firstly, an engine air system one-dimensional network model, a fuel and lubricating oil system one-dimensional network model, a wheel disc and cartridge receiver two-dimensional axial symmetry model, a blade parameterization model and a main runner quasi-one-dimensional model are established; then calculating an air system boundary and a fuel and lubricating oil system boundary by using a main runner cavity method, then calculating the fuel and lubricating oil system boundary and the air system boundary by using a one-dimensional transient fluid network method to obtain a heat exchange boundary of a wheel disc and a cartridge receiver, and considering a thermal load and a centrifugal load in an unsteady state heat conduction process; calculating the radial displacement of the wheel disc and the casing by using a two-dimensional axisymmetric finite element method, and calculating the temperature and the radial displacement of the blades by using an engineering algorithm; and finally, integrating the radial displacement of the wheel disc, the blades and the casing and the initial blade tip clearance value to obtain the blade tip clearance value of each stage in the transient process, and feeding back the blade tip clearance value to the mainstream model again to realize iterative calculation. The analysis precision of each physical field is improved, and the calculation time consumption is reduced.

An operating method is provided during which a command is received to decrease thrust generated by a propulsor rotor from a first thrust level to a second thrust level. The propulsor rotor is operatively coupled to an engine core and an electric machine. The engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The engine core is operated in a transient state to decrease total power of the engine core from a first power level to a second power level in response to the command. The electric machine is operated as a generator to reduce power output from the engine core to the propulsor rotor while the engine core is operating in the transient state. A clearance control system for the engine core is operated based on the operation of the electric machine.

An operating method is provided during which a command is received to decrease thrust generated by a propulsor rotor from a first thrust level to a second thrust level. The propulsor rotor is operatively coupled to an engine core and an electric machine. The engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The engine core is operated in a transient state to decrease total power of the engine core from a first power level to a second power level in response to the command. The electric machine is operated as a generator to reduce power output from the engine core to the propulsor rotor while the engine core is operating in the transient state. A clearance control system for the engine core is operated based on the operation of the electric machine.

The invention discloses a method for calculating blade tip clearance and vibration of a turbomachine blade based on continuous Gaussian fitting, and belongs to the technical field of fault diagnosis. A continuous Gaussian fitting (CGF) method is proposed to be used for obtaining a data signal which can be used for more accurately calculating a blade tip gap and a blade tip vibration characteristic. Compared with a conventional threshold value recognition method (TVM), the CGF method has the advantages that Gaussian fitting calibration is performed on blade tip state signals collected by the capacitive sensor, interference of environmental noise is greatly reduced, and the accuracy of blade tip clearance and vibration calculation is remarkably improved. Experimental results show that the CGF method shows better calculation stability under different rotating speed working conditions, especially at a low rotating speed, the calculation stability of blade tip vibration is 4.15 times that of a TVM method, and blade resonance can be recognized more accurately. According to the method, real-time monitoring can be realized, and more accurate blade tip clearance and vibration data can be provided.

The invention discloses an aero-engine blade data storage method, device and equipment and a medium, and relates to the technical field of data processing. The method comprises the steps that signals of all blades of the aero-engine are collected through a plurality of sensors, and blade vibration signals and blade tip gap signals corresponding to the sensors are obtained; generating health characteristic parameters corresponding to the sensors according to the blade vibration signals and the blade tip gap signals corresponding to the sensors, and generating signal auxiliary information corresponding to the sensors according to the health characteristic parameters; determining a time cache granularity corresponding to each sensor according to each piece of signal subsidiary information, and allocating a continuous address cache region corresponding to each sensor; and storing a blade vibration signal, a blade tip gap signal, a health characteristic parameter and signal affiliated information corresponding to each sensor according to each time cache granularity and each continuous address cache region. In this way, efficient and low-redundancy storage can be achieved, and high real-time performance and flexible expansion are achieved.

The invention belongs to the technical field of airplane two-power systems, and discloses a rotor blade tip clearance measuring method and device, and the method comprises the steps: firstly, measuring the distance a between the outer ring surface of an air inlet channel outer ring support and the end surface of an air inlet channel outer ring, and measuring the height of the end surface of the air inlet channel outer ring on a plane; then directly measuring the distance between the lower end surface of the air inlet bearing end cover of the product and the surface of the air inlet bearing end cover, which is used for assembling the L-shaped sealing gasket, and measuring the distance between the bottom of the air inlet outer ring and the lower end surface of the air inlet bearing end cover; calculating to obtain a distance A from a gap design point on the air inlet channel outer ring to a surface, used for assembling an L-shaped sealing gasket, of the air inlet bearing end cover; measuring the bearing axial clearance f of the product bearing under the pretightening force; measuring the thickness h of the L-shaped sealing gasket, and measuring the distance g from a gap design point on the centrifugal impeller to the bottom of the bearing inner ring; the distance B between the clearance design point on the impeller and the end face of the L-shaped sealing gasket is obtained, wherein B = g + the theoretical distance between the end faces of the inner ring and the outer ring of the bearing-f / 2 + h; and calculating the blade tip clearance C = B-A at the clearance design point on the centrifugal impeller. Compared with an existing measuring method, the method has the advantages that the blade tip clearance is measured in the product assembling process, the whole product does not need to be disassembled and assembled repeatedly, meanwhile, the product does not need to be transferred back and forth between measuring equipment, the product assembling efficiency is improved, and the method is more suitable for batch operation.

Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes a first annular substrate extending along an axial direction; a second annular substrate positioned radially inward relative to the first annular substrate, the second annular substrate movably coupled to the first annular substrate; and an actuator coupled to the second annular substrate such that a force applied by the actuator moves the second annular substrate relative to the first annular substrate to adjust a tip clearance.

A turbine assembly includes a bladed rotor mounted for rotation about an axis of the gas turbine engine, a case assembly, and an actively-cooled tip clearance system. The tip clearance system includes a tip clearance sensor located radially outward of an inner case of the case assembly. The actively-cooled tip clearance system includes a sensor is configured to monitor a tip clearance formed between the bladed rotor and the case assembly during operation of the gas turbine engine.

The invention relates to the technical field of fluid machine design and artificial intelligence crossing, in particular to an optimization design method and system oriented to turbomachinery, a storage medium and a computer program. According to the method, a standardized data bus of a parameter layer, a configuration layer, a geometric layer and a result layer penetrates through an average line, parameterized geometry and multi-objective optimization are carried out, thermophysical property calibers are unified, and sample-level archiving and playback are achieved. Through LHS / Sobol sampling, Kriging proxy, NSGA-II and other algorithms, isentropic efficiency, pressure ratio and structure / noise indexes are collaboratively optimized under the condition that constraints of mass flow, minimum thickness, tip clearance and the like are met; the system is coupled with CFD / FEM through an adapter, and supports GUI and batch processing. Compared with the prior art, the design efficiency, the cross-module consistency and the result traceability are improved, and a manufacturable Pareto optimal blade scheme is obtained.

The invention relates to the technical field of aero-engine state monitoring, and discloses an aero-engine blade tip gap measuring probe and system based on an eddy current method.The aero-engine blade tip gap measuring probe adopts a coplanar Archimedes spiral double-coil array, and in the measuring process, the blade tip gap can be measured by applying anti-phase excitation; the common-mode noise suppression in the blade tip clearance measurement process is realized by combining a differential amplification circuit; signal extraction and temperature drift compensation are realized through a phase-locked amplification technology and an impedance-resistance orthogonal decoupling architecture, and finally micron-order gap analysis is realized. The aero-engine blade tip clearance measurement system provided by the invention can still keep excellent reliability under extreme working conditions of high temperature, high rotating speed and the like of a turbine, and meanwhile, the measurement precision of the blade tip clearance can be remarkably improved.

The invention provides a high-pressure turbine outer ring connecting and positioning structure capable of being used for adjusting blade tip gaps, and belongs to the field of aero-engines. The outer ring fixing ring is fixed on the turbine case; the high-pressure turbine outer ring is mounted on the outer ring fixing ring through a connecting pin; the connecting pins are provided with pin heads with circumferential positioning surfaces, large-diameter matching sections and small-diameter matching sections, the circumferential positioning surfaces are matched with planes or arc surfaces on the outer ring fixing ring to realize circumferential positioning, the large-diameter matching sections are matched with mounting holes in the outer ring fixing ring, and the small-diameter matching sections are matched with pin holes in the outer ring of the high-pressure turbine; the connecting pin comprises a non-eccentric structure and an eccentric structure with a small-diameter matching section facing downwards or upwards at different degrees; when the outer ring of the high-pressure turbine is assembled on the outer ring fixing ring, the radial height of the outer ring of the high-pressure turbine is adjusted by selecting and matching the connecting pins with different eccentricity as required, so that the blade tip clearance between the blade and the outer ring of the high-pressure turbine is adjusted.

Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes a first annular substrate extending along an axial direction, the first annular substrate defining a cavity at a radially inward surface of the first annular substrate, a second annular substrate positioned at least partially within the cavity of the first annular substrate, and a tension belt extending circumferentially around a periphery of the second annular substrate.

The invention discloses a turbine blade tip clearance passive matching method based on engine requirements, and belongs to the technical field of aero-engines, and the turbine blade tip clearance passive matching method comprises the steps that the minimum allowable working clearance is determined; determining a typical tip clearance analysis load spectrum; performing thermal analysis and deformation analysis on the rotor and the stator; according to the deformation results of the rotor and the stator, a blade tip gap change rule is analyzed, and a blade tip gap change rule is obtained; according to different engine requirements, blade tip clearance matching is optimized according to the material, the structure and the cooling form of the cartridge receiver; determining an initial cold-state gap according to the narrow-point gap and the minimum allowable working gap; and carrying out iterative calculation, and determining a final blade tip clearance matching scheme. Turbine rotor deformation and turbine stator deformation can be matched so as to ensure that the turbine blade tip clearance is as small as possible in the focused operation state, that is, the turbine blade tip clearance is reduced as far as possible, the turbine efficiency is improved, and the oil consumption rate is reduced on the premise of ensuring the safety of an engine.

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

The invention discloses an aero-engine three-dimensional blade tip gap signal undersampling fractal dimension reconstruction method and system, and belongs to the field of blade non-contact testing, and the method comprises the steps: building a sampling model of a measurement signal for a three-dimensional blade tip gap undersampling signal; the method comprises the following steps: constructing a Fourier sparse matrix, performing sparse representation on an original signal in a frequency domain, deducing an observation matrix, and establishing a three-dimensional blade tip gap signal compressed sensing model; selecting the number of three-dimensional blade tip clearance sensors through a correlation coefficient minimization theory, and optimizing the mounting positions of the sensors; the method provides theoretical guidance for quantitative diagnosis of the blade crack of the aero-engine based on the three-dimensional blade tip gap.

The invention discloses a rotor blade tip gap design method based on cold and hot state transformation and an aero-engine. The design method comprises the steps that the design working condition of cold and hot state transformation of a fan rotor is determined; performing pneumatic calculation on fan components; designing the structure of the hot-state blade S1; reversely solving the hot-state blade S1 to obtain a cold-state blade S2; designing a cold-state blade structure; performing forward solution on the cold-state blade to obtain a hot-state blade S3; forward solving is carried out on the fan rotor cold-state blade S2 to obtain a corresponding hot-state blade S3; performing error analysis on the thermal-state blade S3 and the thermal-state blade S1; blade tip models of the cold-state blade S2 and the hot-state blade S1 are compared and analyzed, and the front edge radial deformation delta q and the tail edge radial deformation delta h of the fan rotor are obtained; and according to the designed thermal state gap value delta r, the front edge gap Sq and the tail edge gap Sh of the blade tip of the fan rotor blade are obtained.

The invention discloses a synchronous decoupling measurement method for vibration displacement and blade top clearance of a shrouded blade, and relates to the field of energy power engineering, and the method comprises the following steps: 1, calibrating the axial vibration displacement and blade top clearance of the shrouded blade; step 2, installing an eddy current sensor; 3, the rotor is made to operate normally, the vibration state of the shrouded blade is measured in the rotating process of the shrouded blade, and a measurement signal is obtained; step 4, carrying out preprocessing on the measurement signal; 5, the sensing area variation of the sensor and the measurement distance variation from the sensor to the shroud are measured; and 6, calculating the vibration displacement and the blade top clearance of the shrouded blade. According to the method, the defects of the prior art can be overcome, synchronous and decoupling measurement on the vibration displacement and the blade top clearance of the shrouded blade can be realized, the measurement accuracy and credibility can be effectively improved, and the measurement requirement can be effectively met.

The invention discloses a cooling structure of a blade top of a high-pressure turbine blade, and belongs to the technical field of cooling structures, the cooling structure comprises a main flow channel, the bottom of the main flow channel is connected with a cold air channel, a double-layer wall is arranged in the cold air channel, air film holes are formed in the top of the double-layer wall in an array mode, and turbulent flow columns are arranged in the double-layer wall in an array mode. Impact holes are formed in the turbulent flow column in an array manner; according to the cooling structure for the blade top of the high-pressure turbine blade, the cold air coverage area of the blade top can be effectively increased, the blade top is prevented from being directly scoured by blade top gap leakage flow, separation bubbles formed by attachment of the blade top gap leakage flow and separation are weakened or even eliminated, the flow loss of the blade top area is reduced, and the temperature distribution of the blade top is improved; the cooling effect of the blade top close to the trailing edge area is effectively improved, and the service life of the turbine blade is prolonged.

Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes an annular substrate extending along an axial direction, the annular substrate including a first surface and a second surface that is radially inward relative to the first surface; an actuator structure coupled to the second surface of the annular substrate; and a smart structure cantilevered from the second surface of the annular substrate, the smart structure including: a support structure, a first region of the support structure coupled to the second surface of the annular substrate, a second region of the support structure coupled to the actuator structure; and a radially inward surface defining a variable surface, the support structure to move the variable surface in a radial direction.

The invention relates to the technical field of gas compressor test design, in particular to a test method for verifying the influence of gas compressor blade tip clearance on performance, which comprises the following steps: constructing an input data sequence formed by inlet temperature and rotating speed; a theoretical model of the inlet temperature and the blade tip thermal-state gap is constructed; acquiring an actual blade tip thermal clearance corresponding to the rotating speed of one group of input data of the gas compressor, and acquiring a correction model; obtaining a final inlet temperature and a final rotating speed corresponding to each to-be-verified blade tip thermal-state gap according to the corrected correction model; an actual performance curve of the gas compressor at the final rotating speed is obtained; and evaluating the influence value of the blade tip thermal-state clearance on the performance of the gas compressor. By obtaining the performance characteristics of the gas compressor under different blade tip gaps, the problems of repeated disassembly, assembly, coping and loading and unloading caused by a traditional method are avoided, so that the test efficiency is improved, and the development period is shortened.

A turbine blade and related method are disclosed. The blade includes an airfoil body having a radial outermost surface relative to a turbine rotor, a bond coat over the radial outermost surface, and an outermost abrasive layer over the bond coat. The outermost abrasive layer may include a graded oxide ceramic layer or a graded metal ceramic layer. In any event, the outermost abrasive layer is vacuum heat treated to harden the layer, resulting in increased rub and improved clearance between the blade tip, creating a smaller tip gap and lower leakage compared to conventional systems. The heat-treated abrasive layer is also resistant to oxidation at higher temperatures, e.g., greater than 1090° C. (˜2000°F.), caused by the increased use of higher temperature fuels, such as hydrogen. The vacuum heat treatment also results in thermal cycle testing improvements.

Radiation emitted by a radiation source (208, 408) along a projected path P0 can be diverted by a divergence angle by a lens (234, 434) in the projected path P0 with a first surface (236, 436) and a second surface (240, 440) spaced apart along the projected path P0 with the second surface (240, 440) farther from the radiation source (208, 408) than the first surface (236, 436). The first surface (236, 436) can be curved and have a normal point perpendicular to a line parallel to the projected path P0 and offset from the projected path P0 by an eccentricity. Radiation striking the first surface (236, 436) can be diverted by a first angle relative to the projected path P0 and can continue to strike the second surface (240, 440), which diverts the radiation by a second angle. The sum of the first and second angles is the divergence angle. The apparatus can be used in a probe assembly (200) to determine tip clearance of a turbine blade (120) during operation of a gas turbine engine (50).

The application discloses a kind of magnetic fluid sealing rim impeller structures for inhibiting blade tip clearance leakage, including pump shell, be provided with blade in pump shell, a circle of blade cover plate is equipped in the periphery of blade, blade cover plate and pump shell have sealing gap in radial direction, one or more magnetic fluid sealing rings are arranged in sealing gap along the axial direction;Magnetic force structure is arranged in the outer wall of pump shell in annular, magnetic concentration structure is arranged in the inner wall of pump shell in annular, closed magnetic circuit is formed between magnetic concentration structure and magnetic force structure, so that gradient magnetic field is formed at sealing gap;Magnetic fluid is adsorbed and positioned in the corresponding area of sealing gap under the action of gradient magnetic field, and forms a magnetic fluid sealing ring which is consistent with the shape of sealing gap, continuous and stable.The application utilizes the self-positioning characteristics of magnetic fluid under magnetic field, forms dynamic, self-adapting sealing interface between blade cover plate and pump shell, realizes zero leakage operation, improves the hydraulic efficiency and stability of axial flow pump unit.

An operating method is provided during which a command is received to decrease thrust generated by a propulsor rotor (26) from a first thrust level to a second thrust level. The propulsor rotor (26) is operatively coupled to an engine core (14) and an electric machine (66). The engine core (14) includes a flowpath (32), a compressor section (28), a combustor section (29) and a turbine section (30). The engine core (14) is operated in a transient state to decrease total power of the engine core (14) from a first power level to a second power level in response to the command. The electric machine (66) is operated as a generator to reduce power output from the engine core (14) to the propulsor rotor (26) while the engine core (14) is operating in the transient state. A clearance control system (20) for the engine core (14) is operated based on the operation of the electric machine (66).

The invention discloses a gas turbine blade tip radial clearance adjusting device and method and a mounting method.The device comprises a fixing ring (1), an adjusting part (2) and a cylinder jacking device (3), the adjusting part (2) surrounds the fixing ring (1) in the radial direction and is fastened, the cylinder jacking device (3) is detachably connected to the adjusting part (2), and the cylinder jacking device (3) exerts force in the radial direction of the fixing ring (1), so that the gas turbine blade tip radial clearance is adjusted. The device is used for controlling deformation of an air cylinder and adjusting blade tip radial clearances. The invention further comprises a mounting method and a gap adjusting method of the using device, the adjusting method comprises the steps that blade tip gap values are collected in real time after a test is started, if the blade tip gap values are smaller than a safety value, the cylinder deformation quantity needing to be adjusted is calculated according to the blade tip gap values in all directions of a cylinder body, and the cylinder deformation quantity is adjusted. The force application position and the force application value are calculated according to the cylinder deformation quantity, and the cylinder ejector (3) is controlled to adjust the blade tip clearance value according to the force application position and the force application value.