591 results about "Gas turbines" patented technology

A system for preventing wear during a wash of a gas turbine engine is provided. The system includes an accessory gearbox, a transfer gearbox drivingly coupled to the accessory gearbox, and a lubrication system fluidly coupled to the accessory gearbox and the transfer gearbox. The lubrication system includes a supply circuit for supplying lubricant from a lubricant reservoir to the accessory gearbox and the transfer gearbox. The lubrication system further includes a return circuit for returning the lubricant from the accessory gearbox and the transfer gearbox to the lubricant reservoir. The lubrication system further includes a wash bypass line extending from the return circuit to the supply circuit.

This invention relates to planetary damping spring dampers for gearboxes. Specifically, a planetary gear arrangement includes planetary gears rotatable on planetary bearings mounted to a bracket of the planetary gear arrangement via support pins. A damping spring is disposed between the outer surface of the outer ring of the bearing and the opposing inner surface of the planetary gear, and includes an annular gap. A spring diaphragm damper is disposed between the cylindrical outer surface of the support pin and the opposing inner surface of the inner ring of the planetary bearing, and also includes an annular gap. The support pin (96) and / or the outer ring includes an oil supply hole leading to the corresponding annular gap. Gas turbine engines include planetary gear arrangements having such planetary bearings.

A gas turbine unit includes a compressor (2), a turbine (5), a combustor assembly (3), and a fuel supply assembly (4); the combustor assembly (3) includes at least one combustor unit (7); the combustor unit (7) includes a first-stage combustor (8) and a second-stage combustor (9); the first-stage combustor (8) includes a first-stage incinerator (11) and a first-stage combustion chamber (12); the first-stage fuel is an amino fuel; wherein the second-stage combustor (9) includes a second-stage incinerator (13) and a second-stage combustion chamber (14); the fuel supply assembly (4) is configured to supply at least one first-stage fuel to the first-stage incinerator (11) such that the combustion stoichiometry of the first-stage incinerator (11) is unbalanced and there is an excess of first-stage fuel in the first-stage incinerator (11) to obtain unburned first-stage fuel reaching the second-stage incinerator (13).

This disclosure relates to a gas turbine and a thermal power generation system. The gas turbine includes a main compressor assembly, an auxiliary compressor, a cooling assembly, and a turbine assembly. The main compressor assembly includes a primary compression unit, a final compression unit, and multiple intermediate compression units located between the primary and final compression units. Each intermediate compression unit has an exhaust port communicating with its internal cavity. The intake port of the auxiliary compressor is connected to the exhaust port of one of the intermediate compression units via the cooling assembly. The exhaust port of the auxiliary compressor is connected to the internal chamber of the turbine assembly. The cooling assembly is disposed between the intermediate compression unit connected to the intake port of the auxiliary compressor and the turbine assembly, and is used to cool the gas before it enters the internal chamber of the turbine assembly.

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.

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.

A combustor assembly (3) for a gas turbine unit (1) comprises at least one combustor unit (7) provided with a liner (16) extending substantially along a combustor axis (B); the combustor unit (7) comprising a first-stage combustor (8) and a second-stage combustor (9), which is arranged downstream the first-stage combustor (8) along the gas flow direction (G); the first-stage combustor (8) comprising a first-stage burner (11), fed with first-stage fuel and first-stage air, and a first-stage combustion chamber (12), where the first-stage fuel is burned; wherein the second-stage combustor (9) comprises a second-stage burner (13) and a second-stage combustion chamber (14); the second-stage burner (13) being arranged downstream of the first-stage combustion chamber (12); the second-stage combustion chamber (14) being fed with hot gas leaving the first-stage combustion chamber (12) and passing through the second-stage burner (13), with dilution air and with a second-stage fuel; wherein the dilution air and the second-stage fuel are injected simultaneously within the second-stage burner (13).

The application discloses a gas turbine, a gas turbine blade and a leaf-shaped film cooling hole structure of the gas turbine blade. The existing traditional cylindrical film cooling hole is reshaped into a leaf-shaped structure including a left half circular arc, a right upper half elliptical arc, a right middle half elliptical arc and a right lower half elliptical arc, so that the cooling gas can be divided into three paths when being sprayed out of the hole, and is sprayed out along the passage secondary flow same direction, the film hole center line direction and the passage secondary flow opposite direction respectively, thereby the blowing-off effect of the passage secondary flow on the cooling film is inhibited, the wall-attaching ability and the along-flow development ability of the cooling film are enhanced, the covering uniformity of the cooling film on the base surface is improved, and the cooling efficiency and the service life of the blade are improved. The leaf-shaped film cooling hole structure has the characteristics of strong functionality and wide applicability, and can be widely applied to the cooling technical field of the gas turbine blade of an aero-engine or a ground gas turbine.

This invention discloses a protection and control method and system for a high-pressure feedwater pump system during the start-up and shutdown process of a gas turbine unit, belonging to the field of intelligent control of gas turbine units. It includes a first subtraction module, a first less than comparison module, a second less than comparison module, a third less than comparison module, a first quality judgment module, a second quality judgment module, a first NOT module, a second NOT module, a third NOT module, a first AND module, a second AND module, a third AND module, a fourth AND module, a first delay module, a second delay module, a third delay module, and a first OR module. The main functions of this invention include: achieving adaptive protection throughout the start-up and shutdown process, multi-system collaborative control, improving the safety and reliability of pump unit operation, reducing the intensity of manual intervention, and adapting to the operational needs of frequent start-up and shutdown and rapid load changes of the unit.

This invention discloses a method and device for auxiliary diagnosis of gas turbine faults, belonging to the field of industrial equipment fault diagnosis technology. Utilizing intelligent retrieval algorithms and efficient database support, it can quickly match relevant cases from a massive database of historical cases and provide diagnostic suggestions when a gas turbine experiences a warning or anomaly, shortening the fault diagnosis time. By characterizing historical fault cases of the gas turbine and combining them with current operating background parameters for multi-dimensional, dynamically weighted similarity matching, it can more accurately identify historical cases most similar to the current abnormal operating conditions, thereby improving the accuracy of fault diagnosis. In the specific application scenario of gas turbine power plants, it demonstrates unique technical contributions and significant practical value, which is of great significance for ensuring the safe and stable operation of gas turbines, improving power generation efficiency, and reducing operation and maintenance costs.

This invention relates to a design method and system for multi-principal element alloys with enhanced L12 phase dissolution temperature based on multi-scale features. The method obtains the optimal alloy composition by constructing a neural network model. This model uses the alloy composition-atomic scale descriptor L12 phase performance characteristics as input features, and the L12 phase interface fracture energy, L12 phase local thermal vibration entropy, and L12 phase interface phonon density of states overlap integral as output targets. The initial weights and thresholds of the neural network model are optimized using a genetic algorithm. The optimized weights and thresholds are then used to train the neural network model, resulting in a regression model for predicting L12 phase performance. This invention comprehensively demonstrates the enhanced thermal stability of the L12 phase from three dimensions—interface fracture resistance, atomic thermal vibration stability, and interface phonon matching—by outputting three unique physical parameters. This provides a new approach for the design of high-performance heat-resistant alloys and can be applied to the material development of hot-end components such as aero-engine turbine blades and gas turbine combustors.

A ring segment cooling structure with enhanced cooling performance, a ring segment (1600), and a gas turbine (1300) including the same. The ring segment cooling structure includes a cooling channel (1630) and a rib turbulator (1637). The cooling channel is formed inside a shield wall (1611) of a ring segment mounted to a turbine casing that accommodates blades (1400) and is formed in a wavy shape so as to allow cooling air to flow through the cooling channel. The rib turbulator is installed in the wavy-shaped cooling channel.

This prevents leakage of lubricating oil sealed in the bearings located in the exhaust chamber during a sudden load reduction operation of a gas turbine. [Solution] The gas turbine system according to the present invention comprises a gas turbine, a heat recovery device for recovering waste heat from the exhaust gas of the gas turbine, a blower for supplying the exhaust gas of the heat recovery device, and a carbon dioxide recovery device for recovering carbon dioxide from the exhaust gas. The exhaust chamber of a gas turbine houses a bearing section that rotatably supports the gas turbine rotor. Sealing air is supplied to the bearing section by a sealing air supply system. If the gas turbine's operating state is determined to be a sudden load reduction operation, the sealing air supply system is controlled to increase the flow rate of sealing air compared to the gas turbine's normal operation.

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine.

Whether the co-firing ratio is low or high, interlock control is appropriately implemented when the co-firing ratio fluctuates due to the influence of external disturbances. [Solution] This application relates to a gas turbine control device for controlling a gas turbine equipped with a fuel supply system that supplies a first mixed fuel of a first fuel and a second fuel to a first fuel nozzle. The gas turbine control device determines whether or not interlock control of the gas turbine is necessary by comparing the deviation between a target mixing ratio, which is a target value for the mixing ratio of the second fuel in the first mixed fuel, and the mixing ratio of the second fuel in the first mixed fuel, with a reference value. The reference value for determining whether interlock control is necessary differs between the first mixing ratio region, where the combustibility index of the first mixed fuel decreases with increasing mixing ratio, and the second mixing ratio region, where the combustibility index of the first mixed fuel increases with increasing mixing ratio.

This disclosure provides embodiments of a speed control method, apparatus, device, and medium applicable to pure hydrogen gas turbines. One specific implementation of the method includes: determining the grid connection status of a target pure hydrogen gas turbine; in response to determining that the grid connection status meets preset grid connection conditions, acquiring the current unit speed and current unit power; correcting the current unit speed based on the acquired ambient temperature information from the speed sensor to obtain a corrected current unit speed; generating the unit speed to be adjusted based on the current unit power and a preset grid load power; generating a speed fuel reference value based on the current unit speed and the unit speed to be adjusted; determining a unit fuel reference value based on the speed fuel reference value, the acquired acceleration fuel reference value, and the temperature fuel reference value; and controlling the various fuel valves included in the target pure hydrogen gas turbine based on the unit fuel reference value. This implementation can improve the stability of gas turbine operation.

A variable pitch fan for a gas turbine engine includes a hub, a fan blade supported by the hub, the fan blade including a blade root, and a retention system including a bearing and a retainer, the retainer securing the bearing to the blade root. The hub is rotatably supported by the bearing.

A combustor (100) for a gas turbine comprising a combustion chamber (102) having an inlet (104). The combustion chamber inlet (104) is defined by a burner (30). The burner comprises a pilot burner section (200) centred on the centre axis (Y); a swirler section (300) comprising vanes (302) which extend radially outwards from the pilot burner section (200); and a main burner section (400) which extends radially outwards from, and surrounds, the swirler section (300).

The application discloses a welding seam processing device for machining of a high-pressure outer cylinder of a gas turbine, and belongs to the field of machining, which comprises a six-axis mechanical arm and a mounting seat mounted on the six-axis mechanical arm, wherein the end of the mounting seat away from the six-axis mechanical arm is provided with a floating seat, the center of the end of the floating seat away from the elastic connecting piece is fixedly connected with a connecting block, the connecting block is provided with a polishing mechanism, and the floating seat is provided with a guide mechanism for guiding a polishing path. The application has the effects of adapting to complex curved surface polishing, reducing the problem of uneven polishing of the welding seam surface, and further improving product quality.

The present invention belongs to the field of gas turbines, and aims to provide a gas turbine disk overspeed pre-rotation test method for both structural integrity assessment and service life extension. The method comprises the following steps: sequentially performing rotor assembly and dynamic balancing, rotor installation, an ambient-temperature trial run, temperature field calibration, a cold-state pre-rotation test and a hot-state pre-rotation test; and then performing test result validity analysis, and if criteria are met, stopping the process, and if the criteria are not met, repeating the process until the criteria are met, during which dynamic balancing criterion determination, installation criterion determination, vibration amplitude criterion determination, temperature field criterion determination, etc. are required to be performed respectively. The present invention can effectively extend the service life of disks, can reduce the production cost of the disks, and can also complete the structural integrity assessment of the disks.

The application discloses a compressor, which comprises a cylinder, a moving blade, a variable-diameter ring and an adjusting assembly. An inner wall surface of the cylinder is provided with an annular groove. The moving blade is rotatably arranged in the cylinder. The variable-diameter ring is arranged in the annular groove. The variable-diameter ring comprises a plurality of adjusting members and an elastic sealing member. The adjusting members are arranged at intervals along the circumference of the cylinder. One end of the elastic sealing member is connected with one of the two adjacent adjusting members, and the other end of the elastic sealing member is connected with the other of the two adjacent adjusting members. The variable-diameter ring is arranged opposite to the blade tip of the moving blade and has a gap. The adjusting assembly is used for adjusting the position of the adjusting members in the radial direction of the cylinder, so as to change the gap size between the variable-diameter ring and the blade tip of the moving blade. The application guarantees the wear compensation during the operation of the compressor, avoids the defect that the traditional abradable coating cannot be recovered, and improves the overall efficiency and reliability of the gas turbine.

The application discloses an NH3-carbon-hydrogen fuel blending flame soot generation model and a prediction method based on a nitridation mechanism, and belongs to the technical field of combustion pollution control and numerical simulation. The application constructs a surface nitridation reaction population balance model containing NH3, NH2 and HCN double reaction paths on the basis of a traditional HACA soot model, determines reaction rate constants through molecular dynamics and quantum chemistry calculation, and clearly defines the core hypothesis that inertization of nitrogen-containing functional groups permanently blocks HACA growth. The model is solved by using a shadow projection method and is coupled with an OPPDIF flame solver to realize accurate prediction of soot generation in an ammonia-doped flame. The application can accurately distinguish between chemical soot suppression and dilution effect of ammonia, significantly improves the prediction accuracy of soot volume fraction and particle size, reveals the double-phase soot suppression mechanism of the gas phase surface, and is suitable for soot emission prediction and low-emission design of ammonia-doped combustion equipment such as gas turbines, internal combustion engines and ship propulsion.

The application discloses a nano-structure thermal protection coating for hydrogen fuel gas turbine and a preparation method thereof. The nano-structure thermal protection coating takes high-temperature alloy as a substrate, and sequentially deposits a bonding bottom layer, a yttrium stabilized zirconium oxide or zirconate intermediate layer and a silicate top layer on the surface of the substrate. The application starts from the design of the coating structure, and aims at the problems of low service temperature and poor water-oxygen corrosion resistance of the existing thermal protection coating for gas turbine. The water-oxygen corrosion resistance function of the silicate material is superimposed on the thermal protection coating for gas turbine, and the coating structure is optimized to form a new thermal protection coating structure resistant to high-temperature corrosion. The new thermal protection coating structure can not only meet more extreme service environments, but also resist water-oxygen corrosion at high temperature. The new thermal protection coating structure is suitable for hydrogen fuel power generation gas turbine, and is also suitable for hydrogen fuel or hydrogen mixed marine engines, hydrogen fuel rocket engines and other hydrogen fuel gas turbines.

Modular engine (1) for missiles, comprising: an engine inlet (10); a gas turbine (12); a payload turbine (14); at least one propulsor (20) and a nozzle assembly (16), wherein an outlet of the engine inlet (10) is connected to an inlet of the gas turbine (12), and wherein an outlet of the gas turbine (12) is selectively connectable to an inlet of the nozzle assembly (16) and / or to an inlet of the payload turbine (14) for activating and deactivating the payload turbine (14) for driving the at least one propulsor (20), characterized in that the engine (1) is designed to deactivate the payload turbine by jettisoning the payload turbine (14) and / or the at least one propulsor (20).

This invention provides a novel vibration fatigue testing method for gas turbine blades. Addressing the problems of large specimen wear, long testing time, and limited results associated with traditional step-and-step stress loading methods for obtaining blade fatigue limits, this invention excites blade resonance on an electromagnetic vibration table using novel step-and-step stress loading techniques on identical blade specimens. While retaining the step-and-step matching standard, when limited test data cannot meet the matching requirements, a stress "concession" method is rationally applied based on the turbine blade structure and fatigue strength level to eliminate test data with significant stress deviations, thereby obtaining the median fatigue strength of blades from the same batch. This invention makes stress matching faster and more flexible, requiring only 2-3 stress levels to achieve a test fitting result with a confidence level of over 95%, demonstrating high accuracy. This method is applicable to fatigue limit tests of all gas turbine blades.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUTADTExit)2*EGTAHPCExit*1⁢0-1⁢1;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

The present application relates to the technical field of starting control of large load rotor system, and particularly relates to a composite driving method of large load rotor system, which comprises the following steps: after receiving a starting instruction, a control unit drives a gas turbine unit to work, gas output by the gas turbine is guided to be sprayed from a rotor nozzle at the end of a rotor, so that the rotor starts to rotate; when the rotating speed of the gas turbine of the gas turbine unit reaches a self-sustaining rotating speed, the gas turbine provides gas for a piston engine unit, drives the piston engine to start, and the mechanical torque output by the piston engine is transmitted to the rotor through a main transmission shaft, and the gas output by the piston engine is guided to be sprayed from the rotor nozzle; the control unit adjusts the engagement degree of a clutch of the piston engine unit according to the rotating speed of the rotor, and when the rotating speed of the rotor reaches a preset working rotating speed, the clutch is completely tightened, and the starting of the large load rotor system is completed; the present application can realize reliable starting of the large load rotor system.