21results about "Turbine/propulsion engine ignition" patented technology

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.

Disclosed is a propulsion assembly (100) for a hybrid aircraft, comprising a first engine (1) and a second engine (2) each having a gas generator (12, 22) and a free turbine (11, 21), a main rotor (62) coupled to the free turbines (11, 21) of the first and second engines (1, 2), the first engine (1) comprising a first electric machine (30) and a second electric machine (32) that is less powerful than the first electric machine (30), one of the first and second electric machines (30, 32) being able to be coupled to the gas generator (12) and to rotate the gas generator during a starting phase of the engine, and also being able to be coupled to the free turbine (11) in order to generated electric energy after the starting phase, the other of the first and second electric machines being coupled to the gas generator (12) only.

This invention discloses a staged fuel supply system and starting method based on natural gas, specifically relating to the field of gas turbine technology. The system includes a fuel supply shut-off valve, a fuel filter, an online fuel composition monitoring unit, a metering system, an electric heater, and parallel standby and main fuel branches. The gas turbine electronic controller calculates the fuel activity coefficient based on the fuel composition and presets the ignition flow rate and heating temperature. The starting method includes dry-running, ignition, acceleration, and target state stages: during the ignition stage, the fuel supply pattern is corrected according to the fuel activity coefficient; during the acceleration stage, the main fuel rate is reduced based on vibration spectrum feedforward control; and a backfire treatment is included. This invention can adapt to fuel composition, actively suppress backfire, and improve ignition reliability, operational safety, and component lifespan.

A propulsion system (20) includes an open propulsor rotor (34), an open guide vane structure (36), a turbine engine (32) and a blade actuation system (48). The open propulsor rotor (34) includes a plurality of open propulsor blades (42). The turbine engine (32) includes a compressor section (65) and a rotating structure (86). The rotating structure (86) is operatively coupled to the open propulsor rotor (34). The rotating structure (86) includes a compressor rotor in the compressor section (65). The blade actuation system (48) is configured to adjust pitch of at least one of the open propulsor blades (42). A controller (110) is configured to control operation of the blade actuation system (48) during aircraft flight to tailor the pitch. The turbine engine (32) is configured to drive rotation of the open propulsor rotor (34) about the axis during a first mode of operation. The open propulsor rotor (34) is configured to windmill about the axis to drive rotation of the rotating structure (86) during a second mode of operation.

The present invention relates to air turbine starters and systems for gas turbine engines. Specifically, a system includes an air turbine starter having a housing defining an air flow path between an inlet and an outlet. A turbine shaft is located in the housing and a plurality of blades are coupled to the turbine shaft and disposed in the main air flow path. An electric motor is within the housing and an outer surface of the electric motor defines a portion of the main air flow path, the electric motor having a drive shaft. A clutch assembly is within the housing to selectively couple the electric motor drive shaft and the turbine shaft. Further, a valve is to fluidly couple an ambient air supply to the inlet when open and a controller is to open the valve when the electric motor is primed to enable ambient air to flow through the main air flow path and over the outer surface of the electric motor.

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 application provides a lubricating oil electromagnetic valve capable of reducing the starting load of an aviation auxiliary power device, comprising a blocking block, a driving mechanism and a mounting shell; the mounting shell is detachably connected to the auxiliary power device; the mounting shell has a communication chamber, a first through hole and a second through hole which are in communication with each other; the communication chamber is in communication with the lubricating oil pump inlet of the auxiliary power device through the first through hole; the communication chamber is in communication with the air inlet of the auxiliary power device through the second through hole; the blocking block is arranged in the communication chamber and has a blocking state and a communication state; the driving mechanism is arranged on the mounting shell and is used to drive the blocking block to move along the length direction of the mounting shell, so that the blocking block is switched from the blocking state to the communication state or from the communication state to the blocking state. In this way, the state of the blocking block is changed through the driving mechanism, the switching of the lubricating oil path and the air path is realized, and the working reliability and service life of the auxiliary power device are improved.

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

Methods and apparatuses for a heat source of a hydrogen fuel supply system for a turbine engine are disclosed. An example apparatus includes a heat source configured to combust gaseous hydrogen to produce heat energy and combustion byproducts, a line to transport gaseous hydrogen from a hydrogen supply to the heat source, a vent coupled to the heat source to vent excess gaseous hydrogen and combustion byproducts, and a heat exchanger coupled to the heat source to heat liquid hydrogen in the line using heat energy of the heat source.

The application discloses a voltage establishment compensation control method and system for an aviation starting generator, and belongs to the technical field of motor control. The application analyzes parameters that determine the dynamic characteristics of voltage establishment, and designs a voltage establishment control strategy considering the parameters. The control strategy firstly calculates the maximum available power in the voltage establishment process of the starting generator under the premise of considering the influence of copper loss; then, based on dynamic safety constraints and the maximum available power, the optimal instantaneous voltage establishment rate of the starting generator is calculated, and the maximum voltage change of the starting generator is determined; finally, based on the maximum voltage change, the reference voltage value of the current starting generator is calculated, so that the voltage rate compensation of the voltage establishment process is realized. The application can maximize the voltage establishment rate under the premise of ensuring the safety margin, realize fast and stable voltage establishment, and is suitable for multi-working-condition change scenes in a wide speed range.

This invention provides a bypass system for an intercooled cycle gas turbine, comprising a three-shaft intercooled cycle gas turbine, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a power turbine, and an intercooler. An intercooler inlet transition section is provided between the low-pressure compressor and the intercooler, and an intercooler outlet transition section is provided between the intercooler and the high-pressure compressor. By adding a bypass system with a control valve to the intercooler inlet transition section and the intercooler outlet transition section of the three-shaft intercooled cycle gas turbine, before the intercooled cycle gas turbine starts, an electric actuator first opens the control valve, placing the control valve on the bypass flow path in the open position. After successful and stable combustion in the combustion chamber, once the high-pressure rotor speed reaches a certain value, the control valve is gradually closed, eventually reaching the closed position. This technology enables the rapid start-up of the three-shaft intercooled cycle gas turbine.

Aspects of the invention relate to a starting method (100) for an aircraft engine (1), in which the engine (1) is connected to a lubrication circuit (6) comprising in particular an oil pump system (7), the lubrication circuit (6) being configured and arranged to circulate oil in the engine (1), and in which the operating modes of the engine (1) comprise a stop mode and a standby mode, the starting method (100) being characterized in that, during a starting phase, it comprises the following steps: - measuring (101) the oil temperature, the step (101) of measuring being performed by a temperature detection device; - selecting (102), as a function of the measured temperature compared to a threshold temperature and as a function of the operating mode of the engine, a starting oil flow profile to be applied in the engine, the step (102) of selecting being performed by a computer; - applying the selected starting oil flow profile by means of the oil pump system (103), the oil pump system being controlled by the computer.

A fuel system includes a reservoir, a heat exchanger, a heating device, a cracker and a fuel injector. The heat exchanger is configured to vaporize a first flow of liquid ammonia received from the reservoir into ammonia gas using heat energy. The heat exchanger is configured to receive the heat energy from the heating device during a first operating mode. The heat exchanger is configured to receive the heat energy from a turbine engine a second operating mode. The cracker configured to at least partially crack the ammonia gas into hydrogen gas and nitrogen gas. The fuel system is configured to deliver the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas to a fuel injector for injection into a combustion chamber.

This gas turbine control device controls a gas turbine equipped with: a combustion chamber capable of mixing and burning a first fuel and a second fuel with combustion air; a bypass channel capable of diverting and supplying at least a portion of the combustion air to a downstream chamber of the combustion chamber; and a bypass valve provided in the bypass channel. The gas turbine control device performs control such that the mixed combustion rate of the second fuel increases when the gas turbine load increases, and it controls the opening of the bypass valve based on the mixed combustion rate.The opening control of the bypass valve is carried out in such a way that the opening of the bypass valve is controlled so that the bypass valve is completely closed when the opening is reduced inversely proportional to the load of the gas turbine at a second load that is smaller than a first load at which the bypass valve is completely closed.

A cross-start bleed air regulation system and method are disclosed for regulating bleed air pressure during cross-starting of an engine to be started using bleed air supplied by a gas-supplying engine. The system includes: a starter branch that branches off from the environmental control line connecting the gas-supplying engine and the engine to be started, and connects to the starter of the engine to be started; a pressure sensor located in the starter branch for detecting the actual pipe pressure of the bleed air flowing into the starter from the gas-supplying engine; a starter control unit that receives the actual pipe pressure detected by the pressure sensor and generates a control command based on a comparison between the actual pipe pressure and a target pipe pressure; and a throttle control unit that receives the control command from the starter control unit and drives the throttle lever of the gas-supplying engine to move based on the control command. This enables fully automated and precise control of the aircraft engine cross-starting process.

A combustion device 10 includes: a liner 13b having an opening 13f at an end and a combustion chamber 13c formed inside the liner 13b; a first ammonia flow passage 24 inserted through the opening 13f of the liner 13b, the first ammonia flow passage 24 communicating with the combustion chamber 13c via an injection valve 14c and supplied with liquid ammonia; a second ammonia flow passage 25 covering an outer periphery of the first ammonia flow passage 24; and an air flow passage 23 covering an outer periphery of the second ammonia flow passage 25.

A propulsion system includes an open propulsor rotor, an open guide vane structure, a turbine engine and a blade actuation system. The open propulsor rotor includes a plurality of open propulsor blades. The turbine engine includes a compressor section and a rotating structure. The rotating structure is operatively coupled to the open propulsor rotor. The rotating structure includes a compressor rotor in the compressor section. The blade actuation system is configured to adjust pitch of at least one of the open propulsor blades. A controller is configured to control operation of the blade actuation system during aircraft flight to tailor the pitch. The turbine engine is configured to drive rotation of the open propulsor rotor about the axis during a first mode of operation. The open propulsor rotor is configured to windmill about the axis to drive rotation of the rotating structure during a second mode of operation.

The present invention relates to a cold start method for a thermodynamic system comprising two turbochargers (TC1, TC2), two catalytic combustion chambers (CCC1, CCC2), the second catalytic combustion chamber (CCC2) being equipped with electric heating means (RE), two electric machines (MEL1, MEL2) each connected to one of the turbochargers (TC1, TC2), fuel injection means (FI1, FI2), a heat recovery unit (REC) characterized in that it successively comprises a step (10) of activating the electric heater (RE) until reaching a first determined temperature of the second catalytic combustion chamber (CCC2), a step of rotating the two turbochargers (TC1, TC2) by their electric machine (MEL1, MEL2) at minimum rotation speeds, a step (20) of starting fuel injection in the second catalytic combustion chamber (CCC2),a step (30) of starting fuel injection into the first combustion chamber (CCC1) when it reaches a second determined temperature. Figure 2,