390results 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.

This disclosure generally relates to power generation methods and systems based on gas turbine engines, and particularly to mobile and adaptive power generation systems and methods based on gas turbine engine for supplying mechanical and / or electrical power for fracturing operations at an oil wellsite. Various systems, platforms, components, devices, and methods are provided for flexibly and adaptively configure one of more gas turbines, hydraulic pumps, and electric generators to support both fracturing and electric demands at a well site. The disclosed implementations enable and facilitate a mobile, adaptive, and reconfigurable power system to provide both mechanical and electric power for hydraulic fracturing operation.

A system and process for restarting a turbomachine includes a shutdown cooldown protection process implemented by a plant level control system or direct control system of the turbomachine. The system and process for restarting ensure rotating components are cooled as expected prior to a unit restart. This system and process for restarting will lockout an ability to restart if an improper cooldown of rotating components is detected. If this lockout is enabled, delaying restart for the rotating components to cool naturally is needed, or the operator could decide to force cool the components.

An engine system for an aircraft includes a gas turbine engine and a control system. The control system is configured to motor the gas turbine engine, absent fuel burn, during a taxi mode of the aircraft. The control system is further configured to accelerate a motoring speed of the gas turbine engine, absent fuel burn, above an idle speed of the gas turbine engine to provide propulsion during the taxi mode. The control system is configured to decrease the motoring speed of the gas turbine engine, absent fuel burn, based on a change in a starting mode of the gas turbine engine or the aircraft reaching a targeted new position.

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.

An apparatus can be arranged and configured so that after a gas turbine is shut down due to a detected condition, a higher autoignition temperature fuel can be utilized to purge a fuel delivery system and facilitate a quicker restart of the turbine. Embodiments can utilize conduits, valves, and a control system for detection of a fault condition warranting a shut down, purging of a lower autoignition temperature fuel via venting and use of a higher autoignition temperature fuel for purging to facilitate subsequent restarting of the turbine. Embodiments can be configured so an inert gas purge is not needed for venting and purging of the fuel delivery system.

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.

An apparatus can be arranged and configured so that after a gas turbine is shut down due to a detected condition, a higher autoignition temperature fuel can be utilized to purge a fuel delivery system and facilitate a quicker restart of the turbine. Embodiments can utilize conduits, valves, and a control system for detection of a fault condition warranting a shut down, purging of a lower autoignition temperature fuel via venting and use of a higher autoignition temperature fuel for purging to facilitate subsequent restarting of the turbine. Embodiments can be configured so an inert gas purge is not needed for venting and purging of the fuel delivery system.

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.

Traditionally, during ignition of a gas turbine engine, the fuel control valves are controlled to be nearly closed, in order to provide the small amount of fuel that is necessary to ignite the gas turbine engine. The error that is inherent in providing fuel flow through small openings results in a higher likelihood of start-up failures. Accordingly, a lightoff fuel pressure reduction system is disclosed that uses a pressure reducing regulator on a bypass flow path to temporarily reduce the pressure of fuel supplied to the fuel control valves. In this case, the fuel control valves may be maintained in a more open position during the ignition phase, which reduces the likelihood of ignition failures.

The utility model relates to a ground starting system of an aero-engine. Comprising a control module used for controlling starting of the aero-engine, a power module used for transmitting power to the aero-engine and providing starting force for the aero-engine, a measuring module used for measuring operation data of the aero-engine and feeding back the operation data to the control module, and an oil supply module used for providing fuel oil for the aero-engine. The separation opportunity of the starting device and the engine is accurately controlled in a mode of combining the PLC and the electric control clutch, abrasion to equipment in the starting process is reduced, and the starting device can be universally used for engines with different separation rotating speeds.

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.

A method of starting a liquid hydrogen fuelled gas turbine engine (103) of an aircraft propulsion system. The aircraft propulsion system comprises a hydrogen storage tank (104) configured to store liquid hydrogen, a liquid hydrogen pump (205) configured to pump hydrogen in at least a liquid state, a core combustor (203) configured to receive hydrogen fuel from the hydrogen fuel pump (205), and a hydrogen fuel vent (214) provided downstream of the liquid hydrogen pump (205), and configured to selectively vent hydrogen fuel. The method comprises, in a liquid priming step, flowing hydrogen from the hydrogen storage tank (104) through the liquid hydrogen pump (205) and venting hydrogen through the hydrogen fuel vent (214) until the hydrogen pump (205) is primed with liquid hydrogen, then, in a liquid pumping step, operating the liquid hydrogen pump (205) to pump liquid hydrogen to the core combustor (203) at a required flow rate and pressure for engine ignition in an engine ignition step.

A power generation system (100) includes a combustion system (108), a liquid supply system (104), and a vapor supply system (106). The combustion system (108) is configured to generate power by combusting an alternative fuel. The liquid supply system (104) is configured to channel a liquid alternative fuel to the combustion system (108). The vapor supply system (106) is configured to channel a vapor alternative fuel to the combustion system (108). The combustion system (108) is ignited by combusting the liquid alternative fuel from the liquid supply system (104) and is operated by combusting the vapor alternative fuel from the vapor supply system (106).

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 gas turbine engine of a plurality of gas turbine engines of an aircraft. The apparatus includes a fuel supply system, a combustor, and a controller. The controller is configured to cause fuel to be introduced to the combustor of the gas turbine engine at a first threshold rotational speed of the gas-turbine engine during a normal starting operation in which the aircraft is on the ground. The controller is configured to cause fuel to be introduced to the combustor at a second threshold rotational speed of the gas-turbine engine during an emergency in-flight restarting operation in which the aircraft is in-flight. The second threshold rotational speed is lower than the first threshold rotational speed. Introducing fuel at the second threshold rotational speed results in a higher temperature in a turbine of the gas turbine engine than introducing fuel at the first threshold rotational speed.

A GPU output from a ground power unit (GPU) can be converted to power suitable for starting a jet engine of an aircraft by an auto-transformer rectifier unit (ATRU) of the aircraft. Multiple ATRUs, each connected to a corresponding GPU, may output multiple ATRU outputs, and combining these multiple ATRU outputs can generate a combined power, which can satisfy the power requirements for starting a jet engine when they exceed what can be supplied by a single GPU. The combining is possible when the outputs of the ATRUs are synchronized, which can require generating and sharing a sync signal between the ATRUs that is synchronous with an ATRU designated as the master (i.e., reference) for the other ATRUs.

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.

To prevent erroneous detection of flashback.SOLUTION: A gas turbine control device is targeted for control of a gas turbine that comprises a combustor capable of mixing and burning a first fuel and a second fuel with combustion air, a bypass passage capable of bypassing and supplying at least a part of combustion air to a downstream side space of the combustor, and a bypass valve provided in the bypass passage. The gas turbine control device performs control so as to increase a mixed combustion rate of the second fuel as the load of a gas turbine increases, and controls an opening of the bypass valve on the basis of the mixed combustion rate. The opening control of the bypass valve is performed so as to control the opening of the bypass valve so that the bypass valve is fully closed in a second load smaller than a first load in which the bypass valve is fully closed when the opening is reduced in inverse proportion to the load of the gas turbine.SELECTED DRAWING: Figure 3