Abstract: A method of assembling a mobile electric power generation system comprises providing a generator trailer having an electric generator, a gas turbine trailer having a gas turbine, an exhaust trailer having an exhaust silencer, and an air filter and ventilation trailer having an air inlet device. Each of the trailers further comprise a trailer alignment system. The method includes arranging the gas turbine trailer at a work site such that the generator trailer and the gas turbine trailer are in an end-to-end configuration. The method includes operably coupling each of the exhaust trailer and the air filter and ventilation trailer to the gas turbine trailer, and correcting a position of the at least one of the generator trailer, the gas turbine trailer, the exhaust trailer, and the air filter and ventilation trailer using the trailer alignment system thereof.

Abstract: A system includes a combustor. The combustor has a combustor wall with a combustor dome at an upstream end of the combustor wall, and an outlet at a downstream end of the combustor wall opposite the upstream end. The combustor wall includes an inner wall portion and an outer wall portion defining an interior of the combustor therebetween. Each of the inner wall portion and outer wall portion extends from the combustor dome to the downstream end of the combustor wall. The combustor wall includes an air cooling passage embedded inside at least one of the inner wall portion and the outer wall portion. The air cooling passage extends from the upstream end of the combustor wall to the downstream end of the combustor wall.

Abstract: A hybrid electric propulsion system including: a gas turbine engine comprising a low speed spool, a high speed spool, and a combustor; a lubrication circuit comprising a bearing compartment, a supply pump, and a scavenger pump; an electric motor configured to augment rotational power of the low speed spool or the high speed spool; and a controller operable to: control the electric motor based upon a pressure differential between an interior of the bearing compartment and an exterior of the bearing compartment and to drive rotation of the low speed spool and/or the high speed spool via the electric motor responsive to a thrust command while fuel flow to the combustor is inhibited.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber; and a fence disposed between the forward and aft liner segments, wherein the fence extends in the circumferential direction, and wherein the fence extends into the combustion chamber along the radial direction.

Abstract: A powerplant system is provided that includes a gas turbine engine system and an actuation system. The gas turbine engine system includes a rotating structure, a stationary structure and one or more bearings rotatably mounting the rotating structure to the stationary structure. The actuation system is configured to passively rotate the rotating structure about a rotational axis during transportation of the gas turbine engine system.

Abstract: A method for assembling a fuel distribution system for a turbomachine fuel injector includes inserting a liquid fuel distributor into an interior cavity of a shroud to create a liquid fuel distribution circuit between the liquid fuel distributor and the shroud and inserting a gas fuel distributor into the interior cavity of the shroud and into an interior cavity of the liquid fuel distributor to create a gas fuel distribution circuit between the gas fuel distributor and the liquid fuel distributor. The method includes inserting a fuel transfer tube into an outer diameter of the shroud. The method includes brazing or shrink fitting at least one of the fuel transfer tube, the gas fuel distributor, or the liquid fuel distributor to the shroud.

Abstract: An embodiment of a combustor for a gas turbine engine includes a combustor case, a combustor liner disposed within the combustor case, a fuel nozzle at an upstream end of the combustor liner, a torch igniter within the combustor case, and a removable fuel injector. The torch igniter includes a combustion chamber, a cap configured to receive a removable fuel injector and a surface igniter, a tip, an annular igniter wall extending from the cap to the tip and defining a radial extent of the combustion chamber, a structural wall coaxial with and surrounding the igniter wall, and an outlet passage within the tip which fluidly connects the combustion chamber to the combustor. The removable fuel injector extends through a fuel injector opening of the combustor case. The diameter of the fuel injector opening is wider than a fuel injector diameter of the removable fuel injector.

Abstract: A liner for a combustor in a gas turbine engine. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution passage having a concatenated geometry extending through the liner body. The concatenated geometry has a plurality of discrete dilution holes, an annular slot, and a fence concatenated with the plurality of discrete dilution holes. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the plurality of discrete dilution holes from the cold side to the hot side and a second dilution air flow flowing through the annular slot from the cold side to the hot side into an integrated dilution air flow, and (ii) to inject the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor.

Abstract: A generator assembly includes a stator assembly coupled to an engine stator component of a propulsion engine, the stator assembly including: a stator support structure fixedly attached to the engine stator component; a stator disposed on a supporting surface of the stator support; a manifold coupled to the stator support, the manifold defining a connection volume and including at least one coolant opening at a connection end of the manifold; and an electrical connector extending between the stator and a connection device disposed on the connection end. The generator assembly also includes a rotor assembly comprising a rotor support structure connected to a shaft of the propulsion engine and a rotor attached to the rotor support structure, wherein the rotor rotates in conjunction with the shaft to generate a power signal that travels through the electrical connector to the connection device.

Abstract: A cooling system for cooling hot components of a radial or axial gas turbine engine, which includes a recuperator heat exchanger, provides engine cooling without loss of thermal efficiency. Air flow leaving a compressor is split between a recuperator flow path and a bleed flow path. Air in the bleed flow path flows through the hot parts of the engine, thereby cooling the engine and heating the air. The air in the bleed flow path is combined with the output flow from a combustor and directed into a turbine inlet. A reduction of air flow in the recuperator flow path increases the thermal effectiveness of the recuperator heat exchanger by increasing a ratio of hot and cold flows inside the heat exchanger. The increase in thermal effectiveness of the heat exchanger compensates for energy losses incurred by diverting a portion of the compressor air flow for cooling.

Abstract: A gas turbine engine and combustor assembly are provided, the combustor assembly including a first liner and a second liner together defining at least in part a combustion chamber, wherein the first liner and the second liner are separated by a gap along the longitudinal direction, and wherein the first liner is forward of the second liner relative to a flow of fluid through the combustion chamber along the longitudinal direction, and wherein the gap is extended along the circumferential direction.

Abstract: In a method for operating a combustor of an embodiment, before ignition in the combustor, a mixed gas containing oxygen is circulated through the combustor as a circulating gas. Then, in an operating time from the time of ignition in the combustor to the time of a rated load of a turbine, from the time of ignition until reaching stable combustion conditions allowing stable combustion, a combustion gas in which a controller controls a flow rate of a fuel supplied from a fuel supply part and a flow rate of an oxidant supplied from an oxidant supply part to maintain the same oxygen concentration as an oxygen concentration in the mixed gas is circulated as the circulating gas.

Abstract: A particle separator includes a separator body in a primary fluid passageway of a machine. The primary fluid passageway includes one or more bleed holes through which a diverted portion of the fluid flowing in the primary fluid passageway toward a volume of the machine is diverted into an auxiliary flow passageway that bypasses the volume and directs the diverted portion of the fluid toward one or more other components of the machine. The separator body is coupled with the inner wall and/or outer wall of the primary fluid passageway. The separator body includes an upstream edge positioned to separate at least some particles carried by the fluid from the fluid as the diverted portion of the fluid bends around and flows over the at least one upstream edge of the separator body and into the auxiliary flow passageway.

Abstract: There are provided a throttle mechanism and the like that are capable of easily changing a cross-sectional area of a flow path according to an operating state. The throttle mechanism in an embodiment is a throttle mechanism that controls a flow rate of a fluid flowing through a flow path, and is configured to make a cross-sectional area of the flow path change autonomously according to temperature.

Abstract: An aircraft power system includes an auxiliary power unit (APU), a generator, and a constant speed drive (CSD). The APU drives an APU drive shaft at a first rotational speed during a first condition and a second rotational speed during a second condition. The generator is rotatably coupled to a generator shaft and produces an AC voltage having a target frequency in response to rotation of the generator shaft at a target rotational speed. The CSD unit receives the first rotational speed from the APU drive shaft and rotates the generator shaft at the target rotational speed based on the first rotational speed. The CSD further receives the second rotational speed from the APU drive shaft and rotates the generator shaft at the target rotational speed based on the second rotational speed.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inner structure that extends axially along and circumferentially about an axial centerline. The nacelle inner structure includes an internal compartment and a cowl. The internal compartment is configured to house a core of a gas turbine engine. The cowl is configured to form an outer radial periphery of the internal compartment. An aft end portion of the cowl is also configured to form an outer radial periphery of a compartment exhaust to the internal compartment. The aft end portion of the cowl includes a plurality of axial fingers arranged circumferentially about the axial centerline in an array.

Abstract: A ducted heat exchanger system for a gas turbine engine includes an additive manufactured heat exchanger core with a contoured external and/or internal geometry. A method of additively manufacturing a heat exchanger for a gas turbine engine includes additively manufacturing a core of a heat exchanger to set a ratio of local surface area to flow area to control a pressure drop per unit length along the core.

Abstract: A combustor of an aircraft engine comprises a liner defining a primary and a dilution zone having a hot surface exposed to a flow of combustion gases traveling from the primary zone downstream to the dilution zone and a cold surface. Dilution holes extending through the liner from the cold to the hot surface delimit the primary from the dilution zone. Effusion holes extending through the liner from the cold to the hot surface direct cooling air into the dilution zone. Two or more rows of effusion holes positioned within three dilution hole diameters downstream of the dilution holes are oriented relative to the liner to direct the cooling air in a cooling direction that is at least one of normal to the direction of the flow of gases passing adjacent the effusion holes, and against the direction of the flow of gases passing adjacent the effusion holes.

Abstract: A power plant is configured to output power to a grid power system and comprises a hydrogen generation system configured to produce hydrogen, a gas turbine combined cycle power plant comprising a gas turbine engine configured to combust hydrogen from the hydrogen generation system to generate a gas stream that can be used to rotate a turbine shaft and a heat recovery steam generator (HRSG) configured to generate steam with the gas stream of the gas turbine engine to rotate a steam turbine, a storage system configured to store hydrogen produced by the hydrogen generation system, and a controller configured to operate the hydrogen generation system with electricity from the grid power system when the grid power system has excess energy and balance active and reactive loads on the grid power system using at least one of the hydrogen generation system and the gas turbine combined cycle power plant.

Abstract: A method of operating a combustor of a gas turbine, the combustor including a combustor liner that defines a total combustion chamber volume, and has a primary combustion zone defining a primary volume. The combustor liner includes a movable portion that is arranged to be actuated to adjust a percentage of the primary volume with respect to the total combustion chamber volume. The method includes, at a first operating state of the gas turbine, adjusting a size of the primary volume to a first percentage of the total combustion chamber volume by actuating the movable portion to adjust the size of the primary volume, and at a second operating state of the gas turbine different from the first operating state, adjusting the size of the primary volume to a second percentage of the total combustion chamber volume by actuating the movable portion to adjust the size of the primary volume.