158 results about "Power turbine" patented technology

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

The invention discloses a power turbine optimization method and system based on data driving, and the method comprises the steps: selecting a plurality of optimization working condition points, and determining turbine operation working condition parameters under each optimization working condition point; obtaining a blade profile optimization geometric parameter, and determining the dimension and the variation range of the blade profile optimization geometric parameter; forming a plurality of groups of geometric parameter schemes based on the blade profile optimization geometric parameters, and establishing a turbine three-dimensional model and dividing computational grids; cFD simulation is carried out on the turbine under each group of geometric parameter schemes, aerodynamic performance parameters of each geometric parameter scheme under each working condition are obtained, and a sample library is formed based on the blade profile optimization geometric parameters, turbine operation working condition parameters and the aerodynamic performance parameters so as to establish a neural network prediction model; and performing multi-objective optimization on the blade profile optimization geometric parameters by adopting a neural network prediction model to determine optimal blade profile geometric parameters. By means of the scheme, the optimal blade profile geometric parameter combination meeting the multiple performance targets at the same time can be efficiently searched in the design space.

A fully-submerged multi-stage Horizontally-oriented Conical-helical Hydrokinetic Turbine that captures water stream flows (river, tidal, deep ocean currents) and funnels this kinetic-rich concentrated flow into a conical turbine system consisting of a rotating arrangement of helical hydrofoil blades wherein the internal stream is forced to exit out through the frustum turbine blade assembly. With a higher Capacity Factor and Power Coefficient than conventional renewables, this turbine is designed to be an environmentally friendly baseload electricity generator to meet the ever-increasing electricity demands.

The invention discloses a power turbine rotor structure design method, which belongs to the technical field of aero-engines, and comprises the following steps: arranging an oil film damper according to strain energy distribution and vibration mode; performing structure optimization on the power turbine rotor according to the radial deformation condition of the connection interface; designing the sectional wall thickness of the power turbine shaft; a balance boss is arranged on the outer wall of the maximum displacement position in each order of vibration mode power turbine shaft; and the gap between the central pull rod of the power turbine rotor and the power turbine disc is designed according to the maximum vibration displacement of the power turbine disc. According to the invention, the oil film dampers are arranged at reasonable positions, so that the rotor has good vibration reduction; connection interface deformation coordination calculation is carried out, a connection structure is optimized, and the connection interface is not damaged under long-term work; and designing and optimizing the segmented shaft wall thickness, optimizing and designing the critical rotating speed margin, considering the influence of the limit unbalance on the gap, and realizing a good dynamic balance characteristic.

The invention relates to the technical field of aero-engines, and discloses a starting oil supply rule control method and device for an unmanned aerial engine, and the method comprises the steps: obtaining the rotating speed and acceleration at a target moment and a historical moment, and judging whether a first switching condition is met or not; if yes, the inlet temperature of the power turbine at the target moment and the historical moment is obtained, and whether a second switching condition is met or not is judged; if yes, the fuel given quantity and the fuel feedback quantity are obtained, and whether a third switching condition is met or not is judged; if yes, judging whether a fourth switching condition is met or not based on the first time consumption, the second time consumption and the third time consumption; if yes, whether a fifth switching condition is met or not is judged based on the first time consumption, the second time consumption and the fourth time consumption; if yes, switching to a third starting stage in advance, and executing the oil supply rule of the third starting stage. Through multi-stage judgment, oil supply rules are timely and effectively switched, suspension protection is prevented from being triggered, and the starting success rate of the engine is increased.

An environmental control system (ECS) of an aircraft, having a heat exchanger; a compressor that receives a first airflow from a first air source, compresses the first airflow, and directs the first airflow to the heat exchanger; an axial flow turbine that receives a second airflow from a second air source, extracts energy from the second airflow, and directs the second airflow to the heat exchanger, wherein the heat exchanger directs the first airflow to a cabin of the aircraft and directs the second airflow overboard; and a shaft connected between the turbine and the compressor.

The invention relates to the technical field of gas turbine control, in particular to a gas turbine multi-parameter cooperative control and dynamic safety adjusting system and method. According to the system, PID outputs of high-pressure compressor rotating speed tracking, power turbine rotating speed tracking, active power tracking and parameter limit control loops are generated in parallel through a multi-parameter cooperative control module, and a minimum fuel demand signal is selected as a control reference based on competitive logic; through a dynamic safety adjusting module, a dynamic threshold curve linked with the inlet air temperature is preset, action allowance is superposed, and active intervention before parameter overrun is achieved; and through the fault tolerance module, the operation is maintained by adopting historical data fitting and an online compensation strategy when the sensor has a fault. According to the method, the problems of single parameter control response delay, insufficient limit protection, poor fault tolerance and the like are solved, and the control precision, response speed and operation reliability of the gas turbine under variable load, extreme environment and sensor abnormal working conditions are improved.

A lubricating system is disclosed. The lubricating system includes: at least one first to-be-lubricated component, wherein an inlet of each of the at least one first to-be-lubricated component is connected with a first lubrication oil inlet pipe, and an outlet of the each of the at least one first to-be-lubricated component is connected with a first lubrication oil outlet pipe; and at least one second to-be-lubricated component, wherein an inlet of each of the second to-be-lubricated component is connected with a second lubrication oil inlet pipe, and an outlet of the each of the at least one second to-be-lubricated component is connected with a second lubrication oil outlet pipe. An operating pressure of the each of the at least one first to-be-lubricated component is different from a working pressure of the each of the at least one second to-be-lubricated component.

The invention discloses a hybrid power turboshaft engine and a working method, the engine comprises a core machine system and a fuel cell power supply system, and the core machine system comprises an air inlet channel, an air compressor, a main combustion chamber, a gas turbine and a power turbine; the fuel cell power supply system comprises a liquid ammonia storage tank, a fuel pump, a liquid ammonia decomposer, a pressure reducing valve, a heat exchanger and a proton exchange membrane fuel cell. According to the invention, a hybrid power form of combining a turbine power system and a fuel cell power system is adopted, so that the output power of the turboshaft engine is ensured, and meanwhile, the proton exchange membrane fuel cell provides electric energy for airborne electric equipment of the engine; the problems that the electric power requirement of an existing turboshaft engine is gradually improved, power of storage batteries such as lithium batteries is insufficient, and electric power is difficult to extract from a power turbine are solved. Meanwhile, the ammonia working medium serves as power generation fuel, the energy conversion efficiency of the aviation turboshaft engine is improved, and energy consumption is reduced.

This invention discloses a one-dimensional design method, system, device, and storage medium for a variable-speed power turbine. The one-dimensional design method for the variable-speed power turbine iteratively optimizes the design speed point by calculating the comprehensive performance of the variable-speed power turbine under all operating conditions. It takes into account the characteristics of the wide operating range of the variable-speed power turbine and establishes a comprehensive performance evaluation index and system applicable to the operating characteristics of the variable-speed power turbine. It accurately reflects the comprehensive performance of the variable-speed power turbine throughout the entire mission cycle and accurately evaluates and predicts the full-mission performance of the variable-speed power turbine. It can achieve the goal of efficient aerodynamic design of the variable-speed power turbine under a wide range of operating conditions and meet the design requirements of stable and efficient operation of the variable-speed power turbine within a wide range of operating conditions.

The application discloses a turbo-shaft engine ground bench performance debugging method based on turbine guide vane area, comprising the following steps: obtaining different gas turbine guide vane areas A GT and power turbine guide vane areas A PT under the condition of same gas generator physical rotating speed t4.5 , corresponding gas temperature change data ΔT, power change data ΔP, and forming gas temperature change and power change tables respectively; obtaining different A GT and A PT under the condition of same take-off power g , corresponding specific fuel consumption change data Δsfc and rotating speed change data Δn, and forming specific fuel consumption change table and rotating speed change table respectively; obtaining A GT and A PT corresponding to the change data which simultaneously satisfy the related constraint conditions from the two-dimensional linear interpolation of each table, and taking A GT and A PT as candidate A GT and A PT ; and finding A GT and A PT with the minimum adjustment amount for performance debugging. The application can improve the debugging efficiency and reduce the debugging cost.

The invention relates to the technical field of gas turbine control, and discloses an intelligent fuzzy self-tuning PID optimization control method for the rotating speed of a gas turbine, which adopts the intelligent fuzzy self-tuning PID optimization control method for the rotating speed of the gas turbine based on an improved cloud drift optimization algorithm, is used for adaptively tuning the control parameters of fuzzy PID, can improve the fuzzy PID performance, and can improve the stability of the rotating speed of the gas turbine. The control effect of the gas turbine under complex conditions is improved, and the control precision and robustness of the gas turbine are improved. Compared with an existing classical intelligent optimization algorithm, the method is high in convergence speed and good in optimization effect, the precision of gas turbine power turbine rotating speed control is effectively improved, the adjusting time is shortened, and the overshoot is reduced.

The invention discloses a vertical torque measuring tooth adjusting device, the vertical torque measuring tooth adjusting device abuts against a power turbine shaft, a through cavity is arranged in the power turbine shaft along the length direction, a torque measuring shaft is arranged in the through cavity and is in clearance fit with the power turbine shaft, and the vertical torque measuring tooth adjusting device comprises a supporting assembly, a fixing assembly and an ultrasonic vibration assembly, the fixing assembly comprises a fixing base and a fixing piece, the fixing base is connected with the supporting assembly, the fixing piece is connected with the fixing base, the power turbine shaft is fixed to the fixing base through the fixing piece, and the power turbine shaft is in the vertical direction; the ultrasonic vibration assembly comprises a first ultrasonic vibration structure, the first ultrasonic vibration structure is installed on the supporting assembly, and the first ultrasonic vibration structure abuts against the periphery of the power turbine shaft. Ultrasonic waves emitted by the ultrasonic vibration assembly are applied to the power turbine shaft, so that micro motion is generated between the torsion measuring shaft and the power turbine shaft, the clamping stagnation or hysteresis phenomenon in the torsion measuring shaft adjusting process is reduced or eliminated, and adjustment is more accurate.

This application discloses a method, apparatus, equipment, and medium for calibrating the torque of an aero-gas turbine shaft engine based on a torque meter. The method includes the following steps: S1, engine steady-state performance testing: starting the engine to idle, then slowly increasing the engine to several takeoff torque states, recording the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter at each takeoff torque state; S2, torque calibration: selecting N=2 takeoff torque states for the engine, the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter, and obtaining the calibration function relationship through linear fitting; S3, post-calibration check: calculating the calibrated torque values ​​of the power turbine shaft torque sensor at several takeoff torque states based on the calibration function relationship. If the differences between each calibrated torque value and the torque value corresponding to the torque meter all meet the requirements, then the torque calibration meets the requirements. This application improves the accuracy of engine torque calibration.

The invention discloses a power turbine over-rotation protection method and device, electronic equipment and a storage medium. The power turbine over-rotation protection method comprises the steps that S1, the rotating speed of a speed reducer unit paddle shaft and the rotating speed of the rear end of a power turbine shaft are detected through a control system; s2, judging whether the difference between the rotating speed of a paddle shaft of a speed reducer unit and the rotating speed of the rear end of a power turbine shaft exceeds a preset limit value or not through a control system, if yes, turning to step S4, and if not, turning to step S3; s3, judging whether the rotating speed of the aero-engine reaches an over-rotating condition or not through the control system, if yes, turning to the step S4, and if not, turning to the step S1; and S4, over-rotation protection is conducted on the power turbine rotor, and fuel supply is cut off. According to the power turbine over-rotation protection method, fuel supply of the engine is cut off in a short time after the shaft of the power turbine is broken, and the effect of protecting the rotor of the engine is achieved.

A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft / sec)1 / 2 and three hundred (ft / sec)1 / 2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft / sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:LCORE=[m  (20+m)*n  (10+n)](1100)*DCORE+CIS.

The invention relates to a hydrogen-electricity energy turbine hybrid propulsion engine system, which relates to the technical field of engines, and is characterized in that a motor driven by a hydrogen fuel cell and a hydrogen fuel gas turbine engine are connected in parallel by using a gear transmission unit, and the gear transmission unit has three switchable transmission modes; the propulsion structure establishes a transmission relationship with the motor and a power turbine of the hydrogen fuel gas turbine engine at the same time through a first transmission mode; the propelling structure only establishes a transmission relation with the motor through the second transmission mode; the propulsion structure establishes a transmission relationship only with the power turbine of the hydrogen fuel gas turbine engine through a third transmission mode. At least one of the hydrogen fuel gas turbine engine and the hydrogen fuel cell can be selectively used, so that the response speed of the engine system is increased, and infrared characteristics and noise of the engine system are reduced.

This invention belongs to the field of aero-engine turboshaft technology and discloses an automatic monitoring method and system for the performance degradation of aero-engine turboshaft engines. The method includes: determining the number of monitoring states and the altitude range corresponding to each monitoring state; determining the converted power value for each monitoring state; determining the turboshaft engine power; determining the converted gas generator speed and converted turbine inlet temperature corresponding to the converted power value for each monitoring state; determining the variation trend of the converted gas generator speed and converted turbine inlet temperature with the number of sorties for each monitoring state, and characterizing the degradation trend of the turboshaft engine through these trends. This invention does not require baseline performance, can monitor the engine performance variation trend with the number of sorties, and can monitor the performance degradation of the engine over operating time, without adding extra maintenance workload.

This invention relates to a dynamic balancing method for a flexible rotor system in an engine, belonging to the technical field of mechanical rotor dynamic balancing. It is applicable to the dynamic balancing of low-pressure rotors or power turbine rotors in turbofan engines when using flexible rotors. The turbofan engine includes a compressor blade disk and a hollow connecting long shaft. The dynamic balancing method includes: S1, inspection of the manufacturing process; S2, dynamic balancing tests on the connecting long shaft and blade disk respectively; S3, assembly and testing of the connecting long shaft and blade disk bearing the first identification on the test bench of the dynamic balancing device according to design requirements; S4, checking whether the dynamic indicators meet the standards; S5, full-range speed vibration testing of the connecting long shaft and blade disk as a whole on the platform of a rotor dynamics experimental apparatus; and S6, verification of the repeatability of assembly parameters and dynamic balancing, thus improving the efficiency of dynamic balancing.

The present application relates to the technical field of aviation system, more particularly to an aviation engine integrated with an inerting system and an aviation aircraft. It solves the problem that the traditional on-board nitrogen generation system technology needs to use the engine bleed air, thereby affecting the engine performance. The aviation engine comprises a compressor, a rotating shaft, a gas turbine, a power turbine, a generator, a compressor air splitter, an anode tail gas splitter, a combustion chamber, an anode tail gas blower, a reforming reactor, a fuel inerting system, a fuel pump, an electric propulsion system and a cathode tail gas splitter, wherein the fuel inerting system comprises a solid oxide fuel cell, a reformer heat exchange channel, a first heat exchanger, a second heat exchanger and an oil tank. It reduces the influence of the inerting system on the weight of the aircraft, avoids the additional demand of the nitrogen generation system on the engine bleed air, and also meets the demand for electric energy under the trend of aircraft electrification, thereby ensuring the normal working performance of the engine.

The application discloses a power turbine rotor combination blade root X-ray detection method. The power turbine rotor combination comprises a long rod and a wheel disc located on two sides of a blade respectively, and the wheel disc has a bottom recess at the center. The X-ray directional penetration is implemented on all blade roots of the power turbine rotor combination in the circumferential direction, and a film is tightly attached to the circular arc surface of the annular structure during the X-ray directional penetration. The detection method creatively regards the blade root as an annular structure, reduces the penetration thickness, can significantly reduce energy loss and improve detection sensitivity, realizes complete image of the detected part, improves the contrast of the film, and guarantees stable and reliable detection quality of the blade root. The directional penetration device is used for auxiliary detection, can quickly clamp and stabilize the power turbine rotor combination, guarantees that the ray beam is perpendicular to the blade root, and tightly attaches the film and the effective penetration area, and improves the image contrast and the detection sensitivity.

The invention discloses a novel turboshaft engine double-rotor coaxial nested transmission structure which comprises a power turbine shaft, a power turbine disc is fixedly arranged at one end of the power turbine shaft, a spline shaft is arranged on the side wall of the other end of the power turbine shaft, and a gas compressor shaft is arranged on the power turbine shaft between the spline shaft and the power turbine disc in a sleeving mode. The gas compressor shaft is coaxial with the axis of the power turbine shaft, a fastening seat is arranged on the gas compressor shaft, a centrifugal impeller is sleeved on the gas compressor shaft on one side of the fastening seat, the centrifugal impeller is fixedly connected with the gas compressor shaft, a gas turbine blade is sleeved on one side of the centrifugal impeller, and the gas turbine blade is fixedly connected with the centrifugal impeller. A hold-down ring is arranged on one side of the gas turbine blade. The coaxial nesting design is adopted for the power turbine shaft and the gas compressor shaft, the radial size of the whole structure can be obviously reduced, a transmission system is more compact, and the turbine engine is matched with the limited installation space of a turboshaft engine; the power turbine disc is directly fixed to one end of the power turbine shaft, and power can be efficiently transmitted.

The invention relates to the technical field of aviation gas turbine engines, and provides a power turbine test piece rotor connecting structure which comprises a connecting shaft, a turbine shaft and multiple stages of blade discs, the connecting shaft is matched with the turbine shaft, parts in the multiple stages of blade discs are arranged on the connecting shaft, gaps are formed between the multiple stages of blade discs on the connecting shaft, and the multiple stages of blade discs are arranged in the gaps. And the multi-stage blade disc far away from the connecting shaft is connected with the multi-stage blade disc on the connecting shaft. The connecting shaft is connected with the blade disc and the turbine shaft, the connecting shaft can lift the hub bottom hole to be larger than the position height of a traditional hub bottom hole, the length of the center plate is shortened, the weight of the single blade disc is reduced, and therefore the overall weight of the rotor is reduced.

The application relates to the technical field of aero-engines and discloses a starting fuel supply law control method and device for an aero-unmanned engine, which comprises the following steps: obtaining the rotating speed and acceleration of a target moment and a historical moment, judging whether a first switching condition is met, obtaining the power turbine inlet temperature of the target moment and the historical moment if the first switching condition is met, judging whether a second switching condition is met, obtaining the fuel given quantity and the fuel feedback quantity if the second switching condition is met, judging whether a third switching condition is met, judging whether a fourth switching condition is met based on a first time consumption, a second time consumption and a third time consumption if the third switching condition is met, judging whether a fifth switching condition is met based on the first time consumption, the second time consumption and a fourth time consumption if the fourth switching condition is met, turning into a starting third stage in advance if the fifth switching condition is met, and executing the fuel supply law of the starting third stage. Through multi-stage judgment, the application realizes timely and effective switching of the fuel supply law, avoids triggering the suspension protection, and improves the engine starting success rate.

Systems and methods to pump fracturing fluid into a wellhead may include a gas turbine engine including a compressor turbine shaft connected to a compressor, and a power turbine output shaft connected to a power turbine. The compressor turbine shaft and the power turbine output shaft may be rotatable at different rotational speeds. The systems may also include a transmission including a transmission input shaft connected to the power turbine output shaft and a transmission output shaft connected to a hydraulic fracturing pump. The systems may also include a fracturing unit controller configured to control one or more of the rotational speeds of the compressor turbine shaft, the power turbine output shaft, or the transmission output shaft based at least in part on target signals and fluid flow signals indicative of one or more of pressure or flow rate associated with fracturing fluid pumped into the wellhead.

The invention discloses an aviation turbine generator, which belongs to the field of aviation turbines and comprises a main shaft and a driver, the first rotor of the driver is sleeved with the air compressor, and the main shaft is sleeved with the air compressor; the gas turbine is sleeved outside the main shaft; the power turbine shaft is rotationally sleeved outside the axial middle part of the main shaft; the power turbine shaft is sleeved with the power turbine; the generator is connected with the power turbine shaft; the compressor, the cold end of the heat exchanger, the combustion chamber, the gas turbine, the power turbine and the hot end of the heat exchanger are sequentially communicated, and the combustion chamber is communicated with the oil injection assembly. The invention has the technical effects that the utilization rate of waste heat is high, the dynamic balance of the power turbine shaft is easy, and the overall length of the power turbine shaft is not easy to increase.

The invention discloses a compressed air energy storage system started by a low-power turbine and a control method, the system comprises a main body module and a starting control module, and a compression unit of the main body module adopts a plurality of sections of independent gas compressors and corresponding driving motors to form a gas compression section group; the driving motor can be controlled to be switched among cut-out, low-torque boosting and rated driving. The starting control module is provided with a starting gas storage container and a low-power starting turbine, an output shaft of the starting turbine applies pull-back torque to the target gas compression section set through an auxiliary power interface, and when the pull-back torque reaches a threshold value, the pull-back torque is seamlessly switched to be driven by the driving motor. The control method comprises the steps of starting condition detection, motor switching-out, turbine driving starting, cooperative acceleration transition, driving mode switching and the like, and soft starting and starting torque peak clipping are achieved. The system has the advantages of being low in starting current, small in mechanical impact, high in reliability, good in expandability and the like, and an efficient solution is provided for engineering application of a high-power compressed air energy storage system.

This invention relates to the field of aero-engine technology, specifically to an engine gearbox and an aero-engine having the same. The engine gearbox includes a gearbox body, within which a power output shaft assembly is integrated. The input end of the power output shaft assembly is adapted for transmission with the engine body, and the output end is adapted for transmission with a load mechanism. A central gear is mounted on the power output shaft assembly. An accessory drive gear assembly meshes with the central gear and is adapted for transmission with an accessory mechanism. This allows the load mechanism to rotate synchronously at high speed with the engine's power turbine rotor. Simultaneously, a portion of the power is transmitted to the accessory mechanism via the accessory drive gear assembly through the central gear, enabling the power turbine rotor's rotational speed to drive the accessory mechanisms at their respective suitable speeds after gear shifting.

The invention relates to a turbine engine (1) of an aircraft (2), comprising: - a gas generator (3) of longitudinal axis (X1) which comprises a power turbine (4); - an unducted fan (5) which is movable about a longitudinal axis (X2) which is parallel to the axis (X1) and offset radially with respect to the axis (X1), the fan (5) being driven by the power turbine (4) via a speed reducer (6) which is arranged radially above an air inlet (7) of the gas generator (3); - an unshrouded stator (8) which is rotationally fixed and arranged axially downstream of the unducted fan (5); characterized in that the stator (8) comprises a ring (9) which supports vanes (10, 11) and which jointly surrounds a casing (12) of the speed reducer (6) and the air inlet (7), the ring (9) being fastened to the casing (12) of the speed reducer (6).