10838results about "Jet propulsion plants" patented technology

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.

An engine for aeromechanical instability abatement includes a sensor configured to capture data from rotating blades of the engine system, a airflow effector device, and an engine controller. The engine controller is configured to control the airflow effector device according to a nominal schedule, detect, based on a signal from the sensor indicating a vibration amplitude of the rotating blades within a frequency band, an incipient instability condition, in response to the incipient instability condition being present, determine a modified control parameter for at least one of the airflow effector device, and control the airflow effector device according to the modified control parameter, deviating from the nominal schedule.

The invention belongs to the field of data processing, and particularly relates to a dynamic cooperative control system and method for a gas turbine and a micro-grid, and the method comprises the steps: constructing a micro-grid real-time monitoring module, continuously collecting distributed energy real-time output, controllable load demands, bus voltage frequency and equipment state parameters, and carrying out the filtering and noise reduction through a preprocessing unit, thereby guaranteeing the data precision; calculating a real-time power difference value based on the preprocessed data, calling an adaptive neural fuzzy inference system, taking the power difference value, the bus voltage deviation and the frequency deviation as input, and judging whether the power difference value, the bus voltage deviation and the frequency deviation exceed a preset threshold value by means of a fuzzy rule base and a neural network model; if the threshold values are not exceeded, the current states of the gas turbine and the energy storage system are maintained; if any one exceeds the threshold value, a dynamic cooperative control instruction is triggered, precise cooperative control of the gas turbine and the micro-grid is achieved, and the operation stability, the operation efficiency and the reliability of the micro-grid are improved.

A method of operating a gas turbine engine including: an engine core with a turbine, compressor, a fuel combustor, and a core shaft connecting the turbine to the compressor; a fan; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; and a heat exchange system including: an air-oil heat exchanger through which the oil flows; a fuel-oil heat exchanger through which the oil and the fuel flow; a temperature sensor downstream of the fuel-oil heat exchanger; and a valve arranged to allow oil or air flow rate through at least one heat exchanger to be varied, the method including: determining if the fuel temperature has increased above a set threshold at cruise conditions; and in response, controlling the valve so as to change the flow rate through the heat exchanger.

The invention provides a one-key start-stop operation monitoring method and system for a gas-steam combined cycle unit, and relates to the technical field of equipment state monitoring, and the method comprises the steps: obtaining operation data, carrying out the data cleaning, noise reduction and standardization, and extracting the performance characteristics of the unit; constructing a component correlation model based on the spatial-temporal characteristics, and optimizing operation efficiency parameters; and performing fault diagnosis by using a multi-modal feature enhancement network, constructing a layered optimization control system, and performing control strategy optimization to obtain an optimal control strategy. According to the invention, intelligent monitoring, fault diagnosis and optimal control of the gas-steam combined cycle unit can be realized, the operation efficiency and reliability of the unit are improved, the fault risk is reduced, and the service life of equipment is prolonged.

A system for supplying power to a portable battery pack including a battery enclosed by a wearable and replaceable pouch or skin using at least one solar panel is disclosed, wherein the pouch or skin can be provided in different colors and / or patterns. Further, the pouch or skin can be MOLLE-compatible. The battery comprises a battery element housed between a battery cover and a back plate, wherein the battery element, battery cover, and back plate have a slight curvature or contour. Further, the battery comprises flexible leads.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by: VG / VGB. VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly. The gas turbine engine includes an electric power system including at least one electric machine. The electric power system includes a plurality of power converters and a plurality of power distribution management units. At least two of the plurality of power converters or the plurality of power distribution management units are integrated together in a single housing.

A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3. The lubricant extraction volume ratio defined by:VGVGB.VGis a gutter volume of the gutter and VGB is a gearbox volume. The gas turbine engine includes a lubrication system that includes a lubricant tank that stores lubricant, one or more primary gearbox lubricant supply lines, one or more secondary gearbox lubricant supply lines, and a lubricant pump for supplying the lubricant to the gearbox assembly through the primary gearbox lubricant supply lines and the secondary gearbox lubricant supply lines. The lubrication system modulates a mass flow rate of the lubricant to the gearbox assembly through the primary gearbox lubricant supply lines or the secondary gearbox lubricant supply lines.

A propulsion system for an aircraft as disclosed herein may include a nacelle, a shaft positioned centrally within a cylindrical passageway of the nacelle, a fan coupled to one end of the shaft, a turbine coupled to an opposite end of the shaft, an electric motor coupled to the shaft, a compressor positioned within the cylindrical passageway, and a solid oxide fuel cell positioned with a hollow ring-shaped interior of the nacelle. The hollow ring-shaped interior may surround and be isolated from the cylindrical passageway. The turbine may be configured to provide primary torque to the shaft while the electric motor may be configured to provide additional torque to the shaft. The electric motor may be powered an electric output of the solid oxide fuel cell while the turbine may be powered at least in part by output gases from the solid oxide fuel cell.

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined byVGVGB.VG is a gutter volume of the gutter and VGB is a gearbox volume of the gearbox. A gas turbine engine includes the gearbox assembly and a lubrication system. The lubrication system includes a tank that stores a lubricant therein, one or more lubricant pumps for generating a flow of lubricant from the tank to the gearbox assembly, and a flow rate control valve. The flow rate control valve has a flow inlet and a flow outlet and defines a variable throughput from the flow inlet to the flow outlet for controlling a flowrate of the lubricant to the gearbox assembly.

An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio isEIcruise⁡(lean),SAFEIcruise⁡(lean),FF,where EIcruise(lean),SAF isEImaxTO,SAF+EIclimb,SAF2and EIcruise(lean),FF isEImaxTO,FF+EIclimb,FF2.EImaxTO,SAF is nvPM emissions index operating at 100% available thrust and EIclimb,SAF is the nvPM emissions index operating at 85% available thrust if fuel includes sustainable aviation fuel. EImaxTO,FF is the nvPM emissions index operating at 100% available thrust and EIclimb,FF is the nvPM emissions index at 85% available thrust if fuel is a fossil-based hydrocarbon. The lean cruise nvPM emissions index ratio is less than 1.

Method for braking an aircraft turbine engine including a fan or a propeller connected to a turbine by a shaft and an electric generator connected to the shaft, the method comprising the following steps:a request to brake the turbine engine by thrust reversal;a calculation of a turbine braking setpoint by a control system; andin response to the braking setpoint, an adjustment by the control system of a resistance of a dissipative load to draw electric power from the electric generator to the dissipative load.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:EImaxTO / FmaxTOEIidle / Fidlewhere: EIidle is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN for the given operating conditions; and Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN for the given operating conditions. The thrust nvPM emissions index ratio is greater than 0.001. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

An aeronautical thruster of longitudinal axis has a hub, an annular row of unducted upstream rotor blades and an annular row of unducted downstream stator blades. Each downstream stator blade is of variable pitch, and at least one of the downstream stator blades is in a closed-pitch configuration relative to another of the downstream stator blades in that it has a pitch angle smaller than the pitch angle of the other downstream stator blade.

The invention provides a thermal protection coating as well as a preparation method and application thereof. The thermal protection coating comprises a NiCoCrAlY first bonding layer, a NiCoCrAlY second bonding layer, a CeYSZ first ceramic layer and a GdYbYSZ second ceramic layer which are sequentially arranged on the surface of a base material, and the second NiCoCrAlY bonding layer is made of CeO2 (CeO2) and Al2O3 (Al2O3). According to the invention, a thermal protection coating structure with double bonding layers and double ceramic layers is constructed on the surface of the base material, and CeO2 and Al2O3 nano-particle reinforced phases are introduced, so that the use temperature of the thermal protection coating is increased, and the thermal strain tolerance, the crack resistance and the oxidation resistance are greatly improved.

The invention relates to the technical field of energy, in particular to a dynamic self-adaptive control method and system for a gas-steam combined cycle unit, and the method comprises the steps: constructing a multi-modal model library comprising a plurality of operation modes; operating parameters and environment parameters of the unit are collected in real time, and multi-source data fusion and system state evaluation are carried out; carrying out multi-time scale prediction on the load demand, fuel characteristics and environmental parameters of the unit based on a deep learning model; dynamically selecting model precision according to a prediction result, and implementing a multi-mode collaborative optimization decision to generate a control instruction; and executing the control instruction and monitoring an execution effect, and carrying out adaptive adjustment on model parameters and strategies according to feedback. According to the method, the unit can maintain efficient and stable operation under the scenes of power grid peak regulation, rapid load fluctuation, complex environment change and the like, meanwhile, pollutant emission is reduced, and the service life of key equipment is prolonged.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

Disclosed is an assembly for a turbomachine comprising: ⋅—a second connection element (8) having a third end (81) which is connected to a turbine shaft (2) and is mounted on a second bearing (5), and a fourth end (82) connected to a mechanical transmission device (6), the second connection element (8) having a second radial compliance (SR2); and ⋅—a third connection element (9) having a fifth end (91) fixedly mounted on a casing (1), and a sixth end (92) connected to the mechanical transmission device (6), the third connection element (9) having a third radial compliance (SR3), wherein a ratio of the first radial compliance (SR1) to the third radial compliance (SR3) is strictly lower than 10% and / or a ratio of the first radial compliance (SR1) to the second radial compliance (SR2) is strictly lower than 4%.

A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and Wf,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine in kg / s is less than 2. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio isEIidle×Wf,idleRImaxTO×Wf,maxTO.EIidle and EImaxTO are the system loss corrected nvPM emissions index in mg / kg operating at around 7% and 100% available thrust for the given operating conditions, respectively. Wf,idle and Wf,maxTO are the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% and 100% available thrust for the given operating conditions, respectively. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having 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; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT / (AHPCExit2×1000).

A hybrid-electric propulsion (HEP) system is provided that includes a gas turbine engine, an electrical power motive system, a system controller, and a propulsor. The gas turbine engine has a free turbine configuration and a compressor. The electrical power motive system has first and second electric motors and first and second inverters. The gas turbine engine provides motive force to the propulsor. The first electric motor is configurable in a drive mode or in generator mode. The second electric motor is in communication with the compressor. The system controller is in communication with the gas turbine engine, the first and second inverters, and a non-transitory memory storing instructions, which instructions cause the system controller to control the second inverter to operate the second electric motor to provide a motive force to the compressor of the gas turbine engine during a low power setting of the gas turbine engine.

A turbomachine for an aircraft comprising a stator, a turbine shaft and an electric machine arranged to the rear of the turbine shaft, and comprising an electric machine stator and an electric machine rotor, a rotationally fixed coupling carriage capable of moving in axial translation relative to the electric machine rotor between coupled and uncoupled positions, coupling members which are designed to engage with one another after the coupling carriage has moved from the uncoupled position to the coupled position and to transmit a torque from the turbine shaft to the electric machine rotor when they are engaged, and an elastic return member designed to elastically return the coupling carriage in axial translation to the coupled position.

A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.

A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the fuel spray nozzles. Also disclosed is a method of operating a gas turbine engine.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF and EImaxTO,FF are respectively the nvPM emissions index in mg / kg when operating around 100% available thrust for given operating conditions if fuel provided to fuel spray nozzles includes either sustainable aviation fuel (SAF) or fossil-based hydrocarbon fuel; Wf,maxTO is mass flow rate of fuel provided to fuel spray nozzles in kg / s when gas turbine engine is operating at around 100% available thrust for given operating conditions. MTO nvPM emissions index is less than 2. The gas turbine engine is configured to provide fuel including SAF to fuel spray nozzles. Also disclosed is method of operating gas turbine engine.