283results about "Turbine/propulsion engine ignition" patented technology

A method is provided for operating a powerplant. During this method, a hydrocarbon fuel is delivered to a turbine engine for combustion within a combustion chamber of the turbine engine during initial startup operation of the turbine engine. Ammonia is at least partially cracked into hydrogen gas and nitrogen gas. A non-hydrocarbon fuel is delivered to the turbine engine for combustion within the combustion chamber of the turbine engine during post-startup operation of the turbine engine. The non-hydrocarbon fuel is or otherwise includes the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas.

A propulsion assembly for a hybrid aircraft including a first and a second engine each having a gas generator and a free turbine, a main rotor coupled to the free turbine of the first and second engines, the first engine including a first electric machine and a second electric machine of lower power than the first electric machine, one of the first or of the second electric machine being able to be coupled to the gas generator and to set the gas generator in rotation during a start phase of the engine, and being further able to be coupled to the free turbine in order to generate electrical energy after the start phase, the other of the first or of the second electric machine being coupled to the gas generator only.

The invention particularly relates to a hypersonic aerospace engine propelled by electric energy and chemical energy cooperatively. The hypersonic aerospace engine comprises an electric energy propelling assembly and a chemical energy propelling assembly. The electric energy propulsion assembly comprises a first medium-voltage magnetic suspension motor and a second medium-voltage magnetic suspension motor which are of an outer rotor motor structure and an inner rotor motor structure respectively. An air pressurization channel is arranged between the first motor rotor and the second motor rotor; the chemical energy propelling assembly comprises a combustion chamber, a fuel nozzle and an igniter; the air ionization electrification device is used for electrifying the air after ionizing the air; the combustion chamber is connected to the rear side of the electric energy propelling assembly and communicates with the air pressurization channel. The fuel nozzle and the igniter are both arranged at an outlet of the air pressurization channel. The electric energy and chemical energy collaborative propulsion mode has better controllability and operation stability; the zero-speed starting can be directly realized after electrification, an additional starter is not needed, the structural complexity is reduced, the flying distance is higher, and the voyage is farther.

A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.

An auxiliary power generator system for an aircraft including an auxiliary power unit (APU) with dual and fully redundant starter and electrical power generation systems that connects independently to a normal electric network or an emergency electric network of an aircraft. A method for supplying electric power to an aircraft by means of the system.

The invention relates to the field of start-stop control, and discloses a one-key start-stop control system of a single-shaft gas turbine, which is used for solving the technical bottlenecks of the traditional control method in the aspects of state perception, service life management and the like. Comprising the steps that a gas turbine state parameter set containing rotor thermal stress, combustion chamber temperature and exhaust temperature distribution is generated; and generating a life loss evaluation result and an operation efficiency evaluation result by adopting the life evaluation model and the efficiency calculation model, and generating a start-stop control prediction sequence through a dynamic programming algorithm. The service life loss and the operation efficiency are quantitatively evaluated in real time, so that the operation cost problem that the service life and the operation efficiency of the equipment are difficult to consider in the existing start-stop technology is solved, and the full-life-cycle operation cost is remarkably reduced while the long-term reliability of the equipment is guaranteed; maintenance cost can be effectively reduced, fuel consumption is saved, and power grid response economy is improved.

A starting apparatus for a first gas turbine engine of a plurality of gas turbine engines of an aircraft. The apparatus includes an air turbine starter, an electric machine, and a controller. The controller is configured to receive an emergency restart command for the first gas turbine engine while the aircraft is in-flight, determine whether the first gas turbine engine is in operation, determine whether at least a second gas turbine engine of the plurality of gas turbine engines is in operation, and, responsive to receiving the emergency restart command and determining that at least the second gas turbine engine of the plurality of gas turbine engines is in operation and that the first gas turbine engine is not in operation, perform an emergency restart of the first gas turbine engine.

A starting apparatus for a first gas turbine engine of a plurality of gas turbine engines of an aircraft. The apparatus includes an air turbine starter, an electric machine, and a controller. The controller is configured to receive an emergency restart command for the first gas turbine engine while the aircraft is in-flight, determine whether the first gas turbine engine is in operation, determine whether at least a second gas turbine engine of the plurality of gas turbine engines is in operation, and, responsive to receiving the emergency restart command and determining that at least the second gas turbine engine of the plurality of gas turbine engines is in operation and that the first gas turbine engine is not in operation, perform an emergency restart of the first gas turbine engine.

The application provides a high-power low-voltage starting rectifying device, which comprises a three-phase input power distribution and filter unit, a three-phase 12-pulse rectifying unit, a DC / DC unit I, a DC / DC unit II and a control module; the three-phase input power distribution and filter unit is connected with the three-phase 12-pulse rectifying unit, the three-phase 12-pulse rectifying unit adopts a differential connection triangle autotransformer 12-pulse rectifying technology and is connected with the DC / DC unit I and the DC / DC unit II respectively; the DC / DC unit I adopts a three-channel staggered parallel phase-shift full-bridge technology, and the DC / DC unit II adopts a single-channel phase-shift full-bridge technology; the control module is connected with the DC / DC unit I and the DC / DC unit II respectively; the control module is used for controlling the DC / DC unit I and the DC / DC unit II to convert direct current voltage and supply a hovercraft gas turbine. The hovercraft gas turbine can be started.

An aspect includes a hybrid gas turbine engine system (201; 301) of a hybrid electric aircraft (200; 300). The hybrid gas turbine engine system includes a gas turbine engine (20), an electric motor (312) operable to perform an electric taxiing of the hybrid electric aircraft, and a controller (216). The controller is operable to prevent fuel flow to the gas turbine engine (20) during at least a portion of the electric taxiing and monitor for a powered warm-up request during the electric taxiing. A powered warm-up state (405) of the gas turbine engine (20) is initiated based on detecting the powered warm-up request. The powered warm-up state (405) adds heat to one or more components of the gas turbine engine prior to transitioning to a takeoff power state. The gas turbine engine transitions from the powered warm-up state to the takeoff power state after reaching a target temperature (408) of the one or more components in the powered warm-up state.

This gas turbine control device uses a gas turbine as a control target, the gas turbine being provided with: a combustor capable of mixed combustion of a first fuel and a second fuel with combustion air; a bypass path capable of bypass-supplying at least a portion of the combustion air to the downstream-side space of the combustor; and a bypass valve provided in the bypass path. The gas turbine control device performs control so as to increase the mixed combustion rate of the second fuel as the load of the gas turbine increases, and controls the degree of opening of the bypass valve in accordance with the mixed combustion rate. The opening control of the bypass valve is performed such that, when the opening is inversely proportional to the load of the gas turbine, the opening of the bypass valve is controlled such that the bypass valve is in a fully closed state at a second load that is smaller than a first load in which the bypass valve is in a fully closed state.

To provide a fuel supply system for driving a gas turbine cogeneration system using off-gas generated in a process in which bio liquid fuel is generated from biomass as fuel, and a fuel supply method for the gas turbine cogeneration system.SOLUTION: A fuel supply system includes: a fuel gas supply line for supplying fuel gas to a combustor of a gas turbine; a first off-gas supply device for supplying first off-gas generated in a fuel refining plant to the combustor; a second off-gas supply device for supplying second off-gas that is generated in a bio liquid fuel generation plant and has a heating value per unit mass smaller than that of the fuel gas to the combustor; a gas mixing device for mixing the fuel gas supplied by the fuel gas supply line, the first off-gas supplied by the first off-gas supply device and the second off-gas supplied by the second off-gas supply device; and a mixed gas fuel supply line for supplying the mixed gas fuel generated in the gas mixing device to the combustor.SELECTED DRAWING: Figure 4

Based on a state of the gas turbine, a normal start mode or a hot start mode is selected as the start mode. When start-up control is started based on the start mode, the speed of the gas turbine gradually increases. In a low speed range, an opening degree of an inlet guide vane provided by a compressor of the gas turbine is maintained at a first opening degree, and an exhaust valve is maintained at a second opening degree. In a high speed range, the inlet guide vane is controlled at a first intermediate opening degree larger than the first opening degree, and the exhaust valve is controlled at a second intermediate opening degree smaller than the second opening degree. The first intermediate opening degree is set larger in the hot start mode than in the normal start mode. The second intermediate opening degree is set larger in the hot start mode than in the normal start mode.

An air turbine starter system of an aircraft includes an air turbine starter having a data acquisition system disposed thereon. The data acquisition system includes a data acquisition controller, a memory device, and a set of starter sensors.

A propulsion assembly (100) comprising: first and second engines (1, 2) each having a gas generator (12, 22) and a free turbine (11, 21); a main rotor (62) coupled to the free turbine (11, 21); the engines (1, 2) each comprise: a first electric machine (30, 40) coupled only to the gas generator (12, 22); and a second electric machine (32, 42) coupled to the gas generator (12, 22) by means of first coupling means (34, 44) when rotating in a first direction and coupled to the free turbine (11, 21) by means of second coupling means (36, 46) when rotating in a second direction, the first electric machine (30, 40) selectively operating in a motor mode or a generator mode, the second electric machine (32, 42) being coupled to the gas generator (12, 22) by means of first coupling means (34, 44) when rotating in a first direction and to the free turbine (11, 21) by means of second coupling means (36, 46) when rotating in a second direction. The second electric machine (32, 42) operates in a motor mode when rotated in a first direction and selectively operates in a motor mode or a generator mode when rotated in a second direction.

There is disclosed an aircraft turbine engine comprising a restart motor (ME2) configured to drive rotation of a high-voltage shaft (22) using an alternating restart current (CA2) delivered by a power supply system (3) comprising: a generator (ME1) for generating an alternating supply current (CA1) by extracting torque from a low-voltage shaft, an electrical rectifier (31) connected to the generator (ME1), an electrical converter (32) connected to the restart motor (ME2), an electrical switching device (33) configured to prevent current from flowing between the electrical rectifier (31) and the electrical converter (32) in an open state and to allow it to flow in a closed state (F1), control means (4) configured to command the closed state (F1) and to activate, during the in-flight restart, the means (D) for modulating the electrical power of the AC restart current (CA2).

Provided is a gas turbine system in which electricity consumption of a battery when starting or stopping a gas turbine engine is reduced and sufficient battery power during flight can be secured. The present invention includes a plurality of gas turbine engines 21, 22, and 23, generators 41, 42, and 43, a battery 5, and a controller 7. The controller 7 decides at least one main starting gas turbine engine 21 from the plurality of gas turbine engines before the plurality of gas turbine engines are started, starts the main starting gas turbine engine 21 using electricity from the battery 5, and starts the secondary starting gas turbine engines 22 and 23 of the plurality of gas turbine engines other than the main starting gas turbine engine 21 using electricity from the generator 41 connected to the main starting gas turbine engine 21 after the main starting gas turbine engine 21 reaches a steady state.

An aircraft (400) comprises an engine system (300) which includes a gas turbine engine (100) and a fuel system (376). The engine has combustion apparatus having an annular array of alternating primary and secondary fuel-injectors. The fuel system comprises a fuel store (370A, 370B) and a controller (374). The controller is arranged to receive one or more signals indicative of one or more of (i) starting of the gas turbine engine; (ii) flame-out of the combustion apparatus; and (ii) a manoeuvring of the aircraft associated with a risk of flame-out of the combustion apparatus or preparation for such manoeuvring; and in response thereto to control the fuel system to commence supply of fuel from the fuel store to the secondary fuel-injectors of the combustion apparatus. Compared to known aircraft, engine lighting and relighting are achieved for more rapidly and reliably and the aircraft is less susceptible to flame-out, especially where hydrogen fuel is used.

This aircraft propulsion system (12) comprises a turboshaft engine (20), a control system (24), a device (72) for determining the current engine speed (Rc) of the turboshaft engine (20), and a rotating machine (22) coupled to the turboshaft engine's (20) drive shaft for generating electrical power by drawing mechanical power from the drive shaft. The control system (24) is configured to limit the current electrical power (Pc) generated by the rotating machine (22), at least when the engine is accelerating, to a maximum permissible power (Pmax) that is a function of the current engine speed (Rc) before acceleration. Figure for the abbreviation: Fig. 2

The invention relates to an adaptive starting method for a turboshaft engine (10) comprising a gas generator (11) to which an electric machine (21) is mechanically connected and which is electrically connected to one or more electrical energy sources (26-28). According to at least one characteristic of each electrical energy source (26-28), a maximum electrical strength of the current that can be delivered by the one or more electrical energy sources (26-28) is determined, and a maximum engine torque supplied by the electric machine (21) powered by each electrical energy source (26-28) at said maximum electrical strength is determined. The maximum engine torque is compared with a specified engine torque envelope (70) permitted by the gas generator (11). The maximum electrical strength of the current supplied by the one or more electrical energy sources (26-28) is adjusted to comply with the prescribed envelope (70) in order to start the gas generator (11).

A method for operating a hybrid-electric propulsion system of an aircraft is provided. The hybrid-electric propulsion system includes a gas turbine engine having a high pressure system, a low pressure system, and an electric machine coupled to one of the high pressure system or low pressure system. The method includes receiving data indicative of an actual or anticipated in-flight shutdown of the gas turbine engine; and adding power to the gas turbine engine through the electric machine in response to receiving data indicative of the actual or anticipated in-flight shutdown of the gas turbine engine.

An air turbine starter system (200) of an aircraft (100) includes an air turbine starter (210) having a data acquisition system (330) disposed thereon. The data acquisition system (330) includes a data acquisition controller (335), a memory device (340), and a set of starter sensors (310).

The aircraft propulsion system comprises: - a turboshaft engine; - a starter; - a freewheel clutch system; - a control system designed to, with the turboshaft engine at idle: - receive a standby request (Rv); and - in response to the standby request (Rv), shut down the turboshaft engine. The control system is designed to, in response to the standby request (Rv): - before the turboshaft engine shuts down, command the starter so that the starter speed (ND) becomes non-zero; and - after shutdown, command the starter to maintain the starter speed (ND) non-zero while the gas generator speed (NGG) decreases to the standby speed (NGGv). Figure for the abbreviation: Fig. 3

The present invention provides a rapid response control design method for the starting process of an aeroengine, belonging to the field of engine control. Using the N-dot control plan, an intelligent optimization algorithm is utilized to optimize the acceleration target curve, and then the controller is designed based on the active disturbance rejection control theory. Finally, the starting process time of the engine is minimized to achieve rapid response. The present invention can solve the problems that the time consistency and reliability of the transition state of an aeroengine are difficult to guarantee under traditional control methods. It is a control design method for the rapid response of the transition state of an aeroengine based on the N-dot control plan and active disturbance rejection control (ADRC), which can shorten the transition state adjustment time on the premise of ensuring the stable operation of the aeroengine.

The utility model provides a coaxial hybrid power system engine which comprises an engine shell, an air inlet channel and a combustion chamber which are communicated with each other are arranged in the engine shell, and an installation chamber is arranged between the air inlet channel and the combustion chamber. The gas compressor is arranged in the gas inlet channel and is close to the gas inlet of the gas inlet channel; the starting integrated motor is arranged in the mounting chamber and located on the rear side of the gas compressor; the turbine is arranged in the combustion chamber; and the main rotating shaft is rotationally arranged in the engine shell, the main rotating shaft sequentially penetrates through an air compression rotor of the air compressor, a motor rotor of the starting and generating integrated motor and a turbine rotor of the turbine motor from front to back in the axial direction, and the air compression rotor, the motor rotor and the turbine rotor are all coaxially and fixedly connected with the main rotating shaft to form a power transmission system. The engine improves air inlet efficiency, reduces vibration, improves stability, and has the advantages of being compact in structure, light in weight, low in installation accuracy requirement, high in air inlet efficiency, small in vibration and good in stability.

An adaptive starting method for a turboshaft engine comprising a gas generator, an electric machine connected mechanically to the gas generator and electrically to one or more electrical energy sources. A maximum electrical intensity of an electric current that can be delivered by the electrical energy hgsource or sources is determined as a function of at least one characteristic of each electrical energy source, a maximum engine torque supplied by the electric machine powered by each electrical energy source with the maximum electrical intensity is determined. The maximum engine torque is compared with a required engine torque envelope permitted by the gas generator. The maximum electrical intensity of the electric current supplied by the electrical energy source or sources is adapted in order to comply with the required envelope in order to start the gas generator.

An alternative fuel quick start system for a gas turbine engine having a combustor and a compressor is provided. The alternative fuel quick start system may include an alternative liquid fuel, a thermal energy storage tank, and a vaporizer. The alternative liquid fuel is pressure-bearing stored in the thermal energy storage tank such that release of the alternative liquid fuel from the thermal energy storage tank vaporizes the alternative fuel. When the gas turbine engine is started, the thermal energy storage tank supplies vaporized alternative fuel to the combustor until the vaporizer can operate. The alternative liquid fuel may be propane, ethanol, or methanol.

An aspect includes a hybrid gas turbine engine system (201; 301) of a hybrid electric aircraft (200; 300). The hybrid gas turbine engine system includes a gas turbine engine (20), an electric motor (312) operable to perform an electric taxiing of the hybrid electric aircraft, and a controller (216). The controller is operable to prevent fuel flow to the gas turbine engine (20) during at least a portion of the electric taxiing and monitor for a powered warm-up request during the electric taxiing. A powered warm-up state (405) of the gas turbine engine (20) is initiated based on detecting the powered warm-up request. The powered warm-up state (405) adds heat to one or more components of the gas turbine engine prior to transitioning to a takeoff power state. The gas turbine engine transitions from the powered warm-up state to the takeoff power state after reaching a target temperature (408) of the one or more components in the powered warm-up state.