Abstract: To achieve a reduction in waste heat loss of a fuel with improved efficiency in arrangement of a plurality of cavities, or manifold spaces. A multi-stage combustor includes a combustor liner configured to define a combustion chamber therein, a plurality of fuel nozzles configured to inject a fuel, and a manifold configured to distribute the fuel to the plurality of fuel nozzles. The manifold is disposed on a central axis of the combustor liner in a central axis extending direction. The manifold includes a casing and a plurality of partitions inserted in the casing so as to be arranged in a central axis direction of the combustor liner to define a plurality of cavities divided by the partitions. The plurality of cavities are layered in the central axis direction of the combustor liner inside of the casing, and are connected to the corresponding fuel nozzles.

Abstract: A combustor for a turbomachine includes a head end, a liner at least partially defining a hot gas path, and a flow sleeve circumferentially surrounding at least a portion of the liner. The flow sleeve is spaced from the liner to form a cooling flow annulus therebetween. The cooling flow annulus is in direct fluid communication with a high pressure plenum, whereby air from the high pressure plenum flows into the cooling flow annulus and from the cooling flow annulus to the head end. The combustor further includes a first combustion zone and a second combustion zone downstream of the first combustion zone along the hot gas path. A plurality of fuel injectors are in fluid communication with the second combustion zone and are not in direct fluid communication with the high pressure plenum.

Abstract: A method of controlling a gas turbine engine capable of operating in at least a high output power range, a medium-high output power range, a medium-low output power range and a low output power range. The method includes during the medium-high output power range bleeding a gas from a downstream part of the compressor to an upstream part of the compressor so that a first predetermined temperature of the combustor is maintained, during the medium-low output power range bleeding a gas from a downstream part of the compressor to an upstream part of the compressor and bleeding a gas from the downstream part of the compressor to the exhaust so that a second predetermined temperature of the combustor is maintained.

Abstract: A fuel system for a gas turbine engine includes a secondary flow lockout valve. The secondary flow lockout valve includes a valve body having a first end that defines an inlet and a second end. The valve body includes at least one primary outlet bore and at least one secondary outlet bore. The valve body defines a channel in fluid communication with the primary outlet bore. The secondary flow lockout valve includes a cover that cooperates with the second end of the valve body to define a chamber. The chamber is in fluid communication with the channel such that the valve body is movable between at least a first position in which the primary outlet bore is open and a second position in which both the primary outlet bore and the secondary outlet bore are open based on a pressure in the chamber.

Abstract: A control device for a gas turbine include: a target value calculation part configured to calculate a control target value being a target value of an output of the gas turbine; and a command value calculation part configured to calculate a fuel command value on the basis of a deviation between the control target value and an actual output value of the gas turbine. The target value calculation part is configured to: set the control target value to a value which is greater than an output demand value of the gas turbine immediately before a difference between the output demand value and the actual output value becomes not greater than a threshold; and subtract the control target value from the value after the difference becomes not greater than the threshold.

Abstract: A method and system for cleaning a fuel nozzle during engine operation is provided. Operations include operating the compressor section to provide the flow of oxidizer at a first oxidizer flow condition to the combustion chamber, wherein the first oxidizer flow condition comprises an environmental parameter; operating the fuel system at a first fuel flow condition to produce a fuel-oxidizer ratio at the combustion chamber; comparing the environmental parameter to a first environmental parameter threshold; and transitioning the fuel system to a second fuel flow condition corresponding to a cleaning condition at the fuel nozzle if the environmental parameter is equal to or greater than the first environmental threshold.

Abstract: An accessory system for a gas turbine engine having a driveshaft is provided. The accessory system includes a towershaft coupled to the driveshaft and driven by the driveshaft. The accessory system also includes a shaft including a first shaft bevel gear coupled to a towershaft bevel gear. The shaft is rotatable by the towershaft. The accessory system includes a first accessory drive shaft having a first accessory bevel gear driven by the shaft. The accessory system also includes a second accessory drive shaft having a second accessory bevel gear driven by the shaft.

Abstract: Intercooling systems and methods for an aircraft engine are provided. An intercooling system includes: a first inlet configured to receive a first air flow of ambient air into the aircraft engine; a second inlet separate from the first inlet and configured to receive a second air flow of ambient air into the aircraft engine separately from the first air flow of ambient air; and a heat exchanger configured to facilitate heat transfer between at least a portion of the first air flow compressed by a compressor section of the aircraft engine and the second air flow.

Abstract: A combustor adapted for use in a gas turbine engine a combustor shell, a heat shield, and a heat shield retainer. The combustor shell is made from metallic materials and is formed to define an internal cavity. The heat shield is formed from ceramic matrix composite materials and is coupled to the dome panel. The heat shield retainer is configured to retain the heat shield to the combustor shell.

Abstract: An accessory system for a gas turbine engine having a driveshaft is provided. The accessory system includes a towershaft coupled to the driveshaft and driven by the driveshaft. The accessory system also includes a shaft including a first shaft bevel gear coupled to a towershaft bevel gear. The shaft is rotatable by the towershaft. The accessory system includes a first accessory drive shaft having a first accessory bevel gear driven by the shaft. The accessory system also includes a second accessory drive shaft having a second accessory bevel gear driven by the shaft.

Abstract: A nozzle having a forward portion skewed downwards and an aft portion translated downwards provides sufficient clearance between the nozzle and the heat shield structure to prevent contact in the event of large deflections (e.g., as associated with a fan blade Out (FBO) condition). Such large deflections must be accounted for to meet federal aviation regulations and gain airplane CFR 14 Part 25 Certification.

Abstract: A system and method for diagnosing load compressor health state for an auxiliary power unit that includes a power compressor, a combustor, a power turbine, and a load compressor is provided. The auxiliary power unit is operated and bleed air is discharged from the load compressor at a bleed air pressure. Using a pressure sensor, the bleed air pressure discharged from the load compressor is sensed and supplied to a processor. In the processor, power compressor health state is diagnosed based solely on the sensed bleed air pressure.

Abstract: An engine system for an aircraft includes a first gas turbine engine, a second gas turbine engine, and a control system. The control system is configured to operate the first gas turbine engine with an idle fuel burn schedule in a taxi mode of the aircraft and operate the second gas turbine engine with a sub-idle fuel burn schedule in the taxi mode of the aircraft. The sub-idle fuel burn schedule includes a reduction of the idle fuel burn schedule. A fuel flow of the first gas turbine engine and the second gas turbine engine is increased above the idle fuel burn schedule prior to takeoff of the aircraft.

Abstract: A method of actively controlling torsional resonance of a rotating shaft of an engine is provided. The shaft has a rotational velocity characterised by a low frequency, rotational velocity term and a high frequency, oscillatory term superimposed on the low frequency term, the oscillatory term being caused by torsional resonance. The method including: measuring the rotational velocity of the shaft; extracting the oscillatory term from the measured rotational velocity; and on the basis of the extracted oscillatory term, applying a torque component to the shaft, the torque component being modulated at the same frequency as the torsional resonance to counteract the torsional resonance.

Abstract: An exhaust duct of a gas turbine engine comprises a hub defining a radially-inner surface of a substantially annular exhaust gas path, and a strut extending into the exhaust gas path. The strut is attached to the hub via a first fastener at a forward fastening location closer to a leading edge of the strut than to a trailing edge of the strut, and via a second fastener at an aft fastening location closer to the trailing edge than to the leading edge. The second fastener is engaged with an appendage of the strut. The appendage is received into a receptacle formed in the hub open to the radially-inner surface.

Abstract: Auxiliary power systems, aircraft including the same, and related methods. An auxiliary power system comprises an auxiliary power unit (APU) controller and an APU with an air intake, a powerhead, and a load compressor stage. The load compressor stage includes a flow regulator assembly, a load compressor, and a bleed air temperature (BAT) sensor for generating a BAT signal. The APU controller regulates a flow rate of a load compressor airflow through the load compressor based on the BAT signal. A method of utilizing an auxiliary power system includes compressing a load compressor airflow to generate a bleed air flow, measuring the BAT with a BAT sensor, generating a BAT signal based on the BAT, transmitting the BAT signal to an APU controller, generating a flow regulator command with the APU controller, transmitting the flow regulator command to a flow regulator assembly, and controlling a flow regulator assembly.

Abstract: A gaspath component for a gas turbine engine includes a platform. A body extends outward from the platform and includes at least one internal cooling passage. The internal cooling passage includes at least one particulate redirection feature defined at an end of the body opposite the platform. The at least one particulate redirection feature includes a first face oblique to and facing an expected flow of fluid through the internal cooling passage and defines at least one corresponding opening.

Abstract: A method of controlling a gas turbine engine capable of operating in at least a high, medium-high, medium, medium-low, and low output power ranges. The method includes during the medium-high output power range varying the angle of the variable guide vanes so that a third predetermined temperature of the combustor is maintained, during the medium output power range the variable guide vanes are closed and bleeding a gas from a downstream part of the compressor to an upstream part of the compressor so that a first predetermined temperature of the combustor is maintained, during the medium-low output power range the variable guide vanes are closed and bleeding a gas from a downstream part of the compressor to an upstream part of the compressor and bleeding a gas from the downstream part of the compressor to the exhaust so that a second predetermined temperature of the combustor is maintained.

Abstract: In a gas turbine, a plurality of blade ring parts each include a plurality of first cooling flow passages, a plurality of second cooling flow passages, and a return flow passage. The first cooling flow passages are disposed on the outer side in a radial direction centering on a rotation axis, extend in an axial direction, and are disposed aligned in an axial rotation direction. The second cooling flow passages are disposed on an inner side in the radial direction with respect to the first cooling flow passages, extend in the axial direction, and are disposed aligned in the axial rotation direction. The return flow passage connects end parts of each of the first cooling flow passages and the second cooling flow passages on the same side in the axial direction with each other.

Abstract: A method and system for fuel nozzle cleaning during engine operation is provided. The steps or operations include operating the compressor section to provide the flow of oxidizer to the combustion chamber, operating the fuel system at a first fuel flow condition, in which the first fuel flow condition provides a fuel-oxidizer ratio at the combustion chamber, operating the fuel system at a second fuel flow condition, in which the second fuel flow condition provides the fuel-oxidizer ratio at the combustion chamber equal at the first fuel flow condition and the second fuel flow condition, and operating the fuel system at a third fuel flow condition after operating the fuel system at the second fuel flow condition.