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.

Abstract: A vapor-based system and method for treating one or more components of a gas turbine engine. The system includes a treatment compound contained in a storage vessel. The storage vessel being operably coupled to a delivery module. The delivery module delivering the treatment compound at one or more locations of the gas turbine engine such that the treatment compound is a vapor when exposed to an engine air-path.

Abstract: A self-adapting gas turbine firebox with variable geometry includes at least one system with variable opening for automatically adjusting a combustion configuration inside a combustion chamber according to an air temperature, particularly of air from a compressor. The adjustment is performed by a thermosensitive member that controls at least one cross-sectional area of a passage to the combustion chamber, for air which participates in the combustion of fuel or participates in a dilution of gases effective inside the combustion chamber.

Abstract: A gas turbine engine has a compressor section with a downstream most end and a cooling air tap at a location spaced upstream from the downstream most end. The cooling air tap is passed through at least one boost compressor and at least one heat exchanger, and then passed to a turbine section to cool the turbine section. The boost compressor is driven by a driveshaft which is driven by the turbine section. A boost turbine selectively drives the boost compressor.

Abstract: Methods and systems for operating a gas turbine engine having variable geometry mechanisms are described. The method comprises detecting a failure event associated with the gas turbine engine, identifying a location of the failure event, selecting an aerodynamic load modulation schedule for the variable geometry mechanisms of the gas turbine engine as a function of the location of the failure event, and applying the aerodynamic load modulation schedule as selected to the variable geometry mechanisms during the failure event.

Abstract: A control system operable includes a modified primary control parameter for an aircraft, the control system includes: a primary control parameter leg configured to output a demand in an aircraft primary control parameter; a primary control parameter compensation leg configured to receive a change in absolute levels and/or spatial distributions of swirl angle and/or fan pressure at a primary control parameter relative to a reference and convert the change into the primary control parameter; a processor adapted to receive the demand in the primary control parameter output from the primary control parameter leg and the change to the primary control parameter output from the primary control parameter compensation leg; compare the demand in the primary control parameter output from the primary control parameter leg and the change to the primary control parameter output from the primary control parameter compensation leg; and generate a modified primary control parameter for the aircraft.

Abstract: A gas turbine engine swirler that includes a tubular body having a forward face, an aft end, and a throat. A plurality of primary swirl vanes that is positioned between the aft end and the forward face. A plurality of secondary swirl vanes that is positioned between the primary swirl vanes and the aft end. The plurality of primary swirl vanes and the plurality of secondary swirl vanes are configured such that the throat is fluidly connected to a plenum that is positioned outside of the tubular body. A tubular ferrule is positioned such that it joins the body at the forward face thereof. Each of the primary swirl vanes extend radially inwardly to a vane lip. The secondary swirl vanes extend radially inwardly for swirling air therefrom. The body also includes a tubular Venturi that extends aft from between the primary swirler vanes and the secondary swirler vanes for radially separating air swirled therefrom.

Abstract: Supercritical fluid systems and aircraft power systems are described. The systems include a compressor, a turbine operably coupled to the compressor, a generator operably coupled to the turbine and configured to generate power, a primary working fluid flow path having a primary working fluid configured to pass through the compressor, a separator, the turbine, and back to the compressor, and a secondary working fluid flow path passing through the generator, the compressor, the separator, and back to the generator. The primary working fluid is supercritical carbon dioxide (sCO2) and the secondary working fluid is a fluid having at least one of a density less than the primary working fluid and a molecular size smaller than the primary working fluid.

Abstract: A combustor assembly comprising a deflector wall in which a plurality of openings is defined through the deflector wall and around the fuel nozzle opening. The plurality of openings defines a first set of openings at a first radius, a second set of openings at or greater than a second radius greater than the first radius, and a third set of openings at one or more of a third radius between the first radius and the second radius. The first set of openings defines one or more of a first angle relative to the radial direction between approximately 60 degrees and approximately 100 degrees. The second set of openings defines one or more of a second angle between approximately zero and approximately 30 degrees. The third set of openings defines one or more of a third angle between the first angle and the second angle.

Abstract: A combustor includes: a first cylindrical body which is configured to hold a fuel nozzle extending in an axial line direction and through which air flows toward a downstream side thereof; a second cylindrical body that is connected to a downstream side of the first cylindrical body; and an outer shell that has an inner peripheral surface configured to define an air introduction channel through which air is introduced such that the air reverses course at an upstream end of the first cylindrical body and flows toward the downstream side together with an outer peripheral surface of the first cylindrical body. The inner peripheral surface has an outside narrowing surface that is formed to extend inward in a radial direction toward the upstream end of the first cylindrical body.

Abstract: Disclosed herein is a combustor capable of improving impingement cooling performance by reducing the influence of a cross-flow on a jet flow. The combustor includes a combustion liner, in which fuel sprayed from fuel nozzles of a gas turbine is mixed with compressed air and the mixture is combusted, a sleeve surrounding the outer surface of the combustion liner while being spaced apart therefrom in order to form a flow passage for compressed air, a cooling hole formed in the sleeve to introduce compressed air discharged from a compressor into the flow passage, and a first auxiliary hole formed upstream of the cooling hole in the airflow direction. Compressed air discharged from the compressor is introduced into the flow passage through the first auxiliary hole, impinges with compressed air flowing through the flow passage, and forms an air column.

Abstract: A combustor assembly for a gas turbine engine defining a radial direction and a circumferential direction includes a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal member including a body, a flange, and a radial element, the body defining a body surface extending along the radial direction and positioned adjacent the interface surface of the at least one liner, the flange extending forward from the body, and the radial element coupled to the flange and extending into the radial opening defined by the at least one liner.

Abstract: A system effective for dual utilization of cooling fluid in a gas turbine engine is provided. A cooling annulus is subject to a hot-temperature combustion flow received from a combustor basket and includes a liner including a feed channel to receive cooling fluid. A feed manifold is in fluid communication with feed channel to feed cooling fluid to a plurality of conduits in fluid communication with a plurality of exit orifices that is in fluid communication with a plurality of resonators. A distributor manifold includes a plurality of manifold sectors in fluid communication with a plurality of conduits arranged to convey cooling fluid. Some of the plurality of resonators operates with different amounts of cooling fluid. A group of the plurality of exit orifices is configured to supply an amount of cooling fluid appropriate for a resonator in fluid communication with the group of the plurality of exit orifices.

Abstract: An air intake system includes an exterior housing for a vehicle, the exterior housing including an outer surface including a recessed portion defined therein. The recessed portion includes an angled bottom member having a first end and a second end that is coupled to the outer surface. The recessed portion further includes a first sidewall, a second sidewall opposing the first sidewall, and an inlet opening defined within the recessed portion. The inlet opening is bounded by the first sidewall, the second sidewall, and the second end, and the inlet opening is configured to receive a fluid stream therethrough. The air intake system further includes an actuation component coupled to the angled bottom member. The actuation component includes a shape memory alloy, and the actuation component is responsive to a change in a thermal condition and configured to move the second end, thereby regulating the inlet opening.