Abstract: A heat engine including a liner extended along a first dimension and a second dimension, wherein the liner includes a plurality of cavities formed by a honeycomb structure. Each cavity forms a volume extended from an opening at a fluid contact side of the liner and along a third dimension substantially orthogonal to the first dimension and the second dimension. The opening at the fluid contact side includes a cross sectional area of each respective volume of the plurality of cavities. The plurality of cavities is non-uniform with respect to the cross sectional area at the opening along the first dimension, the second dimension, or both.

Abstract: Nozzle of a spacecraft engine, comprising a nozzle body extending along a main direction, from a proximal end secured to the engine of the spacecraft, and a free distal end, said nozzle body having a plurality of stiffeners extending from an outer surface of the nozzle body, radially with respect to the main direction, said nozzle comprising a thermal insulation system, comprising at least one strip of insulating material disposed so as to surround the outer surface of the nozzle body over at least part of the height of said nozzle body, said thermal insulation system comprising a plurality of holding elements, each holding element being positioned so as to surround said strips of insulating material, and being disposed between two stiffeners of the nozzle body.

Abstract: A system for bleeding air from a core flow path of a gas turbine engine is disclosed. In various embodiments, the system includes a bleed valve having a bleed valve inlet configured to receive a bleed air from a first access point to the core flow path and a bleed valve outlet; and an air motor having a first air motor inlet configured to receive the bleed air from the bleed valve outlet and a first air motor outlet configured to exhaust the bleed air, the air motor configured to pump the bleed air from the core flow path of the gas turbine engine.

Abstract: A turbine for a turbomachine has a first rotor and a second rotor configured to pivot about an axis in two opposite directions of rotation. The first rotor includes a radially outer drum from which blades extend radially inwards. The turbine has cooling means attached relative to the outer drum. The cooling means include a support plate with at least one first orifice and a calibration plate that is attached to the support plate and located radially inside the support plate. The calibration plate has at least one second calibration orifice facing the outer drum to allow the passage of cooling air radially from the outside to the inside through the first and second orifices.

Abstract: Steam production apparatuses are provided. The apparatuses can include at least two compartments that are mechanically engaged. Processes for the production of steam are also provided. The processes can include providing liquid water to a reactive material within a first compartment to generate steam within the first compartment; transferring at least some of the steam to a second compartment that is mechanically engaged with the first compartment; and exposing the steam from the first compartment to material within the second compartment that extends when exposed to the steam, the extending of the material reducing the volume of the first compartment.

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: A liner for a combustor in a gas turbine engine and a related method. 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 plurality of dilution inserts. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage 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 is provided for alternating starter use during multi-engine motoring in an aircraft. The system includes a first engine starting system of a first engine. A first controller is in communication with a second controller that controls a second engine starting system of a second engine, the first controller being configured to intermittently direct power to the first engine starting system to alternately accelerate and decelerate the first engine during motoring with respect to the second engine.

Abstract: A gas turbine engine includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with an ammonia based fuel and ignited to generate a high energy combusted gas flow that is expanded through a turbine section. The turbine section is mechanically coupled to drive the compressor section. An ammonia flow path communicates an ammonia flow to the combustor section. A cracking device is disposed in the ammonia flow path. The cracking device is configured to decompose the ammonia flow into a fuel flow containing hydrogen (H2). At least one heat exchanger is upstream of the cracking device that provides thermal communication between the ammonia flow and a working fluid flow such that the ammonia fluid flow accepts thermal energy from the working fluid flow.

Abstract: A gas turbine engine assembly includes a gas turbine engine with a combustion section, a fuel delivery system, and a thermal energy management system. The fuel delivery assembly provides a fuel to the combustion section of the gas turbine engine. The thermal energy management system includes a thermal transport bus, a heat source heat exchanger, and a heat sink heat exchanger. The thermal transport bus has a portion of the fuel configured to flow therethrough. The fuel is disposed as a heat exchange fluid of the thermal energy management system. The heat source heat exchanger is in thermal communication with the flow of fuel through the transport bus. The heat sink heat exchanger is in thermal communication with the flow of fuel through the transport bus.

Abstract: An aircraft propulsion system includes a gas turbine engine; a generator; a storage battery; a motor which drives a rotor, using at least one of the electric power which is output from the generator and the electric power which is output from the storage battery; a detection unit which detects the number of revolutions of the engine shaft; an engine control unit which controls at least a fuel flow rate of the gas turbine engine; and a generator control unit which controls the operation of the generator. When the number of revolutions satisfies a predetermined condition, at least the generator control unit executes a control for reducing a sudden change in the number of revolutions.

Abstract: Electric aircraft power plants and associated methods are provided. One power plant includes an emergency power unit (EPU) for providing electric power in the event of a malfunction of a battery pack of an electric aircraft to permit the electric aircraft to make an emergency maneuver. The EPU includes a rocket engine for generating a stream of exhaust fluid using a rocket propellant, a turbine operatively connected to extract energy from the stream of exhaust fluid generated by the rocket engine, and an electric generator operatively connected to be driven by the turbine and to supply electric power to an electric motor propelling the electric aircraft.

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 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: Systems and methods for preventing fuel leakage in a gas turbine engine are provided. A fuel accumulation system includes a control valve section fluidly coupled to a fuel manifold passage and an accumulator valve section fluidly coupled at a first side to the control valve section. The control valve section is configured to control expansion of a fluid flowing in the fuel manifold passage. The accumulator valve section is configured to receive fluid expanded in the fuel manifold passage via the control valve section.

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: 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.