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 thermal management system for a gas turbine engine includes a heat exchanger including first and second sides, with the first side in contact with flow path air flowing through a flow path of the engine. Furthermore, the system includes a housing positioned relative to the heat exchanger such that the housing and the second side of the heat exchanger define a plenum configured to receive bleed air from the engine. Moreover, the system includes and at least one of a plurality of fins extending outward from the second side of the heat exchanger in a radial direction into the plenum and along the second surface of the heat exchanger in the circumferential direction or an impingement plate defining a plurality of impingement apertures, with each impingement aperture configured to direct an impingement jet of the bleed air within the plenum onto the second side of the heat exchanger.

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: This disclosure describes a power system that includes a first compressor with an air inlet and a compressed air outlet; a nitrogen separator coupled to the compressed air outlet, the nitrogen separator comprising a nitrogen concentrate outlet and a byproduct outlet; a second compressor coupled to the nitrogen concentrate outlet, the second compressor having a high pressure outlet for supplying high pressure concentrated nitrogen to an underground storage; and a turbine generator with an inlet for high pressure concentrated nitrogen for coupling to an underground storage.

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: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

Abstract: A fuel injector includes a fuel nozzle configured to issue a spray of fuel from a fuel outlet in a downstream direction along an injection axis. The fuel nozzle includes a nozzle body that defines a main flow passage therethrough. An injection fuel line is in fluid communication with the fuel nozzle to supply fuel to the fuel nozzle. A torch ignitor with a flame outlet opens into the main flow passage of the fuel nozzle for issuing flame into the main flow passage. The flame outlet meets the main flow passage at a position that is downstream of the fuel outlet with respect to the downstream direction along the injection axis.

Abstract: An assembly for a turbomachine can include a fuel nozzle and a combustion liner. The fuel nozzle and the combustion liner can be attached to each using a plurality of clip joints such that the combustion liner is longitudinally fixed relative to the fuel nozzle but such that the combustion liner can radially move relative to the fuel nozzle.

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: An air cap is configured to be affixed to a nozzle assembly as a downstream facing component of the nozzle assembly. The air cap includes a plurality of air wipe passages with outlets defined in a circumferential direction around a downstream facing surface of the air cap. A nozzle assembly can include an outer wall wherein the air cap is connected to the nozzle assembly outboard of the outer wall. An inner wall can be connected to the outer wall by a plurality of air swirl vanes. One or more fuel circuit components can be connected to the nozzle assembly inboard of the inner wall.

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 turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter (D). The heat exchanger module comprises a plurality of heat transfer elements. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct. The inlet duct has a fluid path length along a central axis of the inlet duct between a downstream-most face of the heat transfer elements and an upstream-most face of the fan assembly. The fluid path length is less than 10.0*D.

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: A fuel supply device includes a first electric pump which discharges a fuel toward an engine at a first flow rate, a second electric pump which discharges the fuel at a second flow rate and merges the fuel discharged at the second flow rate with the fuel discharged at the first flow rate, a first control device which is configured to control the first electric pump, and a second control device which is configured to control the second electric pump, in which the first control device and the second control device control the first electric pump and the second electric pump such that a total flow rate of the first flow rate and the second flow rate is a fuel flow rate target value.

Abstract: Disclosed is a mechanism to exhaust refrigerant vapor resulting from operation of a thermal management system that is used to cool a thermal load by a vehicle, such as an airborne vehicle. The thermal management system includes an open circuit refrigeration system featuring a receiver configured to store a liquid refrigerant fluid, an evaporator configured to extract heat from the thermal load that contacts the evaporator, and an exhaust line, where the receiver, the evaporator, and the exhaust line are connected to provide an open refrigerant fluid flow path. Other implementations of open circuit refrigeration systems include the use of a gas receiver, a pump and an ejector are also described, as are other mechanisms to exhaust refrigerant vapor resulting from operation of the thermal management system.

Abstract: A control system for an aircraft hybrid propulsion system comprising a gas turbine engine coupled to an electric generator, a propulsor coupled to an electric motor, and an electrical storage device coupled to the motor and the generator. The control system configured to operate the propulsion system in a first descent mode and a second descent mode. In the first descent mode, the gas turbine engine is operated at a first engine power level and the generator is operated at a first generator power level. In the second descent mode, the gas turbine engine is operated at a second engine power level, higher than the first engine power level, and the generator is operated at a second generator power level, higher than the first generator power level. Electric power generated by the electric generator during operation in the second descent mode is stored in the electrical storage device.

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 system for removing heat from an aircraft component includes: a heat exchanger in proximity to an aircraft component and having an inlet, an outlet, and an internal surface coated with a catalyst; a source of hydrocarbon fuel in fluid communication with the inlet of the heat exchanger; a source of oxygen in fluid communication with the inlet of the heat exchanger; and a distribution system. A method of removing heat from an aircraft component includes providing a heat exchanger in proximity to an aircraft component, the heat exchanger being in fluid communication with a source of hydrocarbon fuel and a source of water and having an internal surface coated with a catalyst; introducing a hydrocarbon fuel into the heat exchanger; introducing oxygen into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel.

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: The upstream-side wall portion 54 includes, in the circumferential direction thereof, a first region 31 where air inlets 30 are formed at a lower density, and a second region 32 which is disposed at a position offset from the first region 31 in the circumferential direction, and in which the air inlets 30 are formed at a higher density than in the first region 31.