144results 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.

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.

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.

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.

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 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 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.

The present invention provides a gas turbine control device that can shorten the gas turbine startup time regardless of changes in atmospheric temperature. [Solution] The gas turbine control device includes a startup fuel command value output unit that outputs a startup fuel command value, a fuel control unit that calculates a fuel limit value based on the rotational speed and acceleration rate setting value of the gas turbine, and a selection unit that selects the larger of the startup fuel command value and the fuel limit value as the fuel command value. The fuel control unit includes an acceleration rate limit unit that sets the acceleration rate limit value to a first acceleration rate limit value when the value obtained by subtracting the startup fuel command value from the fuel limit value is less than a predetermined value, and switches the acceleration rate limit value to a second acceleration rate limit value that is lower than the first acceleration rate limit value when the value obtained by subtracting the startup fuel command value from the fuel limit value is greater than a predetermined value, and an acceleration rate setting unit that sets the acceleration rate setting value based on the acceleration rate limit value.

A device for monitoring a life condition of a spark igniter in a gas turbine ignition system. The device includes an evaluation circuit having circuit components that include a hold capacitor, a transistor, and an operational amplifier arranged to form a sample-and- hold circuit, wherein an igniter spark impulse signal is applied to an input node of the evaluation circuit causing the transistor to turn on and the hold capacitor to discharge for a duration of the igniter spark impulse signal, and wherein a discharged voltage at the hold capacitor is maintained and output by the operational amplifier, the discharged voltage representing the duration of the igniter spark impulse and indicating the life condition of the spark igniter.

Propulsion system (1) for an aircraft, comprising - at least one gas turbine (2), - at least one fluid system (4) for providing a heated fluid (6), - comprising an external auxiliary power unit (40) - wherein the at least one gas turbine (2) and the at least one fluid system (4) are coupled and arranged in such a fluid-conducting manner that at least one section (30, 34, 36) of the gas turbine (2) is supplied with the heated fluid (6), - wherein the fluid system (4) comprises a combustion unit (12) designed separately from the auxiliary engine (40), which is configured to produce the heated fluid (6) by burning fuel, - characterized in that a fluid supplied to the combustion unit (12) is bleed air from the auxiliary engine (40).

The application discloses a kind of information visualization and intelligent gas turbine one-key starting method, by clicking the start button of display screen, to sequentially start the step sequence process of gas turbine and receive the feedback of each step sequence;During the starting process of gas turbine, information interaction can be carried out with multiple systems;When each step sequence process signal and feedback signal are all displayed normally, the gas turbine is started normally;When one of step sequence process signal or feedback signal is abnormal, the signal displayed in the display screen is abnormal, and an alarm signal is sent out.The application discloses a kind of gas turbine one-key starting system, the application can monitor the starting step sequence process in the starting process of gas turbine in real time, when the step sequence is not carried out normally, timely alarm information is sent out, so that personnel can take timely treatment measures, which can not only guarantee the safety of unit, but also improve the work efficiency of power plant.In the starting process, multiple systems can carry out multidimensional information interaction.It provides effective reference for maintenance personnel.

The invention relates to an aircraft propulsion system (100), which comprises a turbine engine (102A, 102B) provided with a rotary portion (104A, 104B), a starter (122A, 122B), a freewheel clutch system (124A, 124B) and a control system (126) designed, when the turboshaft engine is operating at idling speed, to receive a standby request (Rv), and in response to the standby request (Rv), to turn off the turbine engine (102A, 102B). The control system (126) is designed, in response to the standby request (Rv), before the turbine engine is switched off, to control the starter so that the speed (ND) of the starter becomes non-zero, and after the turbine engine is switched off, to control the starter to maintain the non-zero speed (ND) of the starter during the decrease in the speed (NGG) of the rotary portion to the standby speed (NGGv).

A fuel system is configured to deliver a first fuel and / or a second fuel to a combustor section. The fuel system includes a first fuel source, a second fuel source, a first control valve, a second control valve, a metering valve and a fuel injector. The first fuel source is configured to crack ammonia into hydrogen gas and nitrogen gas and subsequently output the first fuel. The second fuel source is configured to output the second fuel. The metering valve is configured to meter a flow of the first fuel received from the first fuel source through the first control valve to the fuel injector during a first operating mode. The metering valve is configured to meter a flow of the second fuel received from the second fuel source through the second control valve to the fuel injector during a second operating mode.

The invention relates to a method for controlling a two-stage thermodynamic system comprising a first turbocharger (TC1), a second turbocharger (TC2), first (CC1) and second (CC2) combustion chambers, and at least one fuel injector between the second turbocharger (TC2) and the second combustion chamber (CC2). The first turbocharger (TC1) comprises a first turbine (T2), the second turbocharger (TC2) comprises a second turbine (T1), and the second combustion chamber comprises an electric heating element. The method comprises the steps of: a) Activating the electric heating element of the second combustion chamber; b) Reversing the operation of the first turbine (T2) to switch the second turbine to compressor mode. (Fig. 1)

The present application provides a rapid start system for alternative fuels for a gas turbine engine comprising a combustion chamber and a compressor. The rapid start system for alternative fuels can include an alternative liquid fuel, a thermal energy storage tank, and a vaporizer. The alternative liquid fuel is stored under pressure in the thermal energy storage tank, so that releasing the alternative liquid fuel from the thermal energy storage tank causes it to vaporize. Upon starting the gas turbine engine, the thermal energy storage tank delivers the vaporized alternative fuel to the combustion chamber until the vaporizer is ready for operation. The alternative liquid fuel can be propane, ethanol, or methanol.

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.

The invention discloses a self-adaptive adjustment gas turbine purging linkage control method, and relates to the technical field of voltage control, a multi-region and three-dimensional gas monitoring network structure is constructed, the precision and coverage dimension of combustible gas concentration monitoring are improved, the distribution characteristics and temperature gradient changes of gases with different densities can be comprehensively captured, and the gas concentration monitoring accuracy is improved. According to the method, an efficient and accurate purging decision is achieved, the starting safety and response speed of the gas turbine are improved, the starting time in the peak shaving process of a power grid is remarkably shortened, a threat sub-equipment recognition mechanism is constructed, damage of thermal shock to equipment is effectively prevented, the needed purging duration can be accurately calculated according to actual monitoring data, and the method is suitable for popularization and application. Energy waste and time delay caused by excessive purging are avoided, meanwhile, potential safety hazards caused by insufficient purging are prevented, the service life of equipment is effectively prolonged, the maintenance cost is reduced, and the economic benefit and technical value of the gas turbine serving as a power grid peak regulation resource are improved.

A combustion device (10) is provided with: a liner (13b) having an opening (13f) provided at an end thereof and having a combustion chamber (13c) formed therein; a first ammonia flow path (24) that is inserted through an opening (13f) of the liner (13b), communicates with the combustion chamber (13c) via an injection valve (14c), and is supplied with liquid ammonia; a second ammonia flow path (25) that covers the outer periphery of the first ammonia flow path (24); and an air flow path (23) that covers the outer periphery of the second ammonia flow path (25).

The invention relates to a fuel supply system (15) for a combustion chamber of an aircraft turbomachine, the system comprising a first fuel supply manifold (17) comprising a start injector (21), a cartridge (23) configured to contain fuel and a first line (25) configured to deliver fuel from the cartridge (23) to the start injector (21), and a second fuel supply manifold (19) comprising a second line (25) configured to deliver fuel from the cartridge (23) to the start injector (21), the second manifold (19) includes a nominal operating injector (27), a second line (29), and a pump (31) configured to send fuel from the fuel tank to the nominal operating injector (27) via the second line (29).

Aircraft power and propulsion systems 400 and methods 600, 700 of starting gas turbine engines 101 of aircraft power and propulsion systems 400 are disclosed. In one example, an aircraft power and propulsion system 400 comprises a gas turbine engine 101 having a first and second spools, the first spool being a designated starting spool; first and second DC electrical networks 431, 423 having first and second operating voltages; a first electrical machine 411 mechanically coupled with the first spool, the first electrical machine 411 having a first winding set 4111 and a second winding set 4112; a second electrical machine 413 mechanically coupled with the second spool; a first AC:DC converter, an AC-side of which is connected to the first winding set 4111 of the first electrical machine 411, and a DC-side of which is connected to the first DC electrical network 431; a second AC:DC converter 4212, an AC-side of which is connected to the second winding set 4112 of the first electrical machine 411, and a DC-side of which is connected to the second DC electrical network 432; a third AC:DC converter, a DC-side of which is connected to the first DC electrical network 431; and a controllable switching arrangement 441-443 configured so that a connection to an AC-side of the third AC:DC converter 4231 is switchable between: a winding 4131 of the second electrical machine; or the second DC electrical network 432 via an inductor 445, 446.

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.

According to the present invention, if a first battery, which is a power source of a high voltage system, is normal when starting a turbine engine, a turbine control device for controlling the turbine generator assists with starting the turbine engine by driving a power generation inverter by using the power of the first battery, and if the first battery is abnormal, the turbine control device electrically disconnects the first battery and drive inverter and assists with starting the turbine engine by driving the power generation inverter by using power from a second battery, which is a power source of a low voltage system. The amount of power consumption when assisting with starting the turbine engine by using the power of the second battery is made smaller than the amount of power consumption when assisting with starting the turbine engine by using the power of the first battery.