Abstract: A fuel nozzle guide assembly for a gas turbine engine is provided. The fuel nozzle guide assembly includes a housing; a tube arranged in the housing and defining a portion of a first fluid passage therein, the first fluid passage configured to include a first fluid, wherein a second fluid passage is defined, in part, between an exterior surface of the tube and an interior surface of the housing, the second fluid passage configured to include a second fluid; an air inflow tube, the air inflow tube defining a central air passage and configured to include a third fluid; an air inflow assembly defining a third fluid passage and configured to include a fourth fluid; and a nozzle outlet configured to receive each of the first fluid, the second fluid, the third fluid, and the fourth fluid to cause mixing thereof.

Abstract: A coupling includes first and second housings, a valve element, and a guide link. The first housing is rotatable relative to the second housing. The first housing has helical cam slots and the second housing has linear cam tracks. The guide link is in the first housing and is fixed with the valve element. The guide link carries cam rollers. The cam rollers ride in the helical slots and the linear cam tracks. Rotation of the first housing causes rotation of the helical slots. The rotation of the helical slots induces the cam rollers to travel along the helical slots and along the linear cam tracks such that the guide link translates linearly. The valve element moves with the guide link between open and closed positions.

Abstract: A combustor for a gas turbine system includes a combustion liner including a primary combustion zone and a secondary combustion zone. The combustor body is additively manufactured and made of a first material. The combustor also includes axial fuel stage (AFS) immersed injectors extending radially into the combustion liner in the secondary combustion zone. Each AFS immersed injector extends through an opening in the combustion liner and is additively manufactured with a second material different than the first material. A coupler fixes each AFS immersed injector in a respective opening in the combustion liner. The additive manufacturing results in as much as a 70% reduction in parts within a given combustor and allows use of high temperature and high oxidation tolerant, hot gas path (HGP) materials for the AFS immersed injectors.

Abstract: A fuel injection assembly for a combustor of a gas turbine engine includes a body with a head section that defines an air intake chamber. A mixing chamber is disposed below the air intake chamber. A deflection member is disposed in the air intake chamber and includes an upper deflection surface and sides spaced apart from inner wall surfaces of the air intake chamber such that peripheral air passages are defined for air to pass from the air intake chamber into the mixing chamber. A plurality of radially extending and longitudinally spaced-apart ribs is disposed along longitudinal side walls of the mixing chamber. A fuel plenum is disposed along the longitudinal side walls of the mixing chamber, and a plurality of fuel injection holes is defined from the fuel plenum through the longitudinal side walls of the mixing chamber and at least some of the ribs.

Abstract: A system includes a hydrogen fuel delivery system for an engine. The hydrogen fuel delivery system includes a pump configured to pump hydrogen from a tank storing the hydrogen in a liquid state through the hydrogen fuel delivery system, and an evaporator configured to convert at least some of the hydrogen in the liquid state to a gaseous state. The system also includes a heat source thermally coupled with the hydrogen fuel delivery system and fluidly uncoupled from the hydrogen fuel delivery system. The hydrogen fuel delivery system is configured to supply the hydrogen to the engine for combustion and cool the heat source using the hydrogen.

Abstract: A rocket engine system including a thrust chamber including walls that define an interior surface and a combustion section which is fluidically coupled to an output section. A coolant source containing a coolant. A means of heating the coolant. At least one port configured to apply the coolant to the interior surface to achieve a film cooling of the interior surface and wherein the coolant source is fluidically coupled to the means of heating and the at least one port.

Abstract: This invention exploits the fundamental physics of turbulent mixing to improve bi-propellant chemical rocket engines. Instead of the many axial reactant injectors in conventional engines, only two transverse injectors are required to generate a pair of large, counter-rotating, streamwise vortices and to mix the reactants for the case of a circular combustion chamber. The vortex pair controls the rate of mixing and the rate of heat release, insensitive to pressure waves and relatively immune to combustion instabilities. The injector of a liquid propellant generates droplets, some of which may contact the opposite wall of the combustion chamber to aid in wall cooling. With the dramatic reduction in the number of injectors, the invention lowers engine mass and manufacturing costs while improving engine reliability.

Abstract: A thrust reverser for an aircraft propulsion system includes a fixed thrust reverser structure including a wall, a sleeve configured to translate between a first stowed position and a second position, and a blocker door assembly including a plurality of blocker doors. An upstream end of the sleeve is disposed at the wall with the translating sleeve in the first stowed position. A locking assembly includes a locking body and a receiver. Each blocker door includes a body pivotably mounted to the sleeve by a hinge and the locking body. Translation of the sleeve from the first to the second position effects pivoting of the blocker door body from a first stowed position to a second position. In a locked position, the locking body is engaged with the wall and disposed at the receiver. In an unlocked position, the locking body is disengaged with the wall and the receiver.

Abstract: An air turbine starter for starting an engine includes a starter housing having an inlet, an outlet, and a flow path extending between the inlet and the outlet. The air turbine starter includes a rotatable turbine member having a central hub defining a platform and a set of blades extending radially from the platform. The air turbine starter also includes a damping member.

Abstract: A turbine engine assembly generates an exhaust gas flow that is communicated through a core flow path. The exhaust gas flow is split into a first exhaust gas flow and a second exhaust gas flow. Water is extracted in a condenser from the second exhaust gas flow. The extracted water is transformed into a steam flow in an evaporator system utilizing thermal energy from at least the second exhaust gas flow. An exit flow from the condenser is communicated through an exhaust compressor and compressed to a higher pressure exit flow.

Abstract: A gas turbine cogeneration system includes: a gas turbine; a first fuel supply facility for supplying a first fuel to the combustor; a second fuel supply facility for supplying a second fuel with a calorific value per mass lower than the first fuel; a heat recovery steam generator for generating steam using exhaust gas discharged from the turbine; a steam supply line for supplying steam discharged from the heat recovery steam generator to a steam consumer; a steam extraction line for supplying steam extracted from the steam supply line to both a head end side of the combustor and a turbine side of the combustor; and a fuel control part for controlling the first and second fuel supply facilities so that the amount of the first fuel decreases and the amount of the second fuel increases with a decrease in extraction amount of steam by the steam extraction line.

Abstract: A method for monitoring a bleed valve of an aircraft engine includes obtaining operational data of the aircraft engine, the operational data indicative of operational performance of the aircraft engine. The method further includes utilizing one or more neural networks to determine a health status of a bleed valve of the aircraft engine based on the operational data, the bleed valve configured to bleed air from a compressor section of the aircraft engine, wherein the one or more neural networks have been trained with root cause data relating to a plurality of potential failure modes of the bleed valve.

Abstract: A casing system mounted between an upstream solenoid valve of an upstream pipe and a downstream solenoid valve of a downstream pipe and which has a casing divided by a separation wall into a first volume which is placed under reduced pressure and a second volume, a heat exchanger which is fixed inside the casing with a first inlet which is arranged in the first volume and which is connected to the upstream pipe and a first outlet which is arranged in the second volume and which is connected to the downstream pipe, and a control unit which closes the solenoid valves if the deviation between two successive values of the pressure in the first volume or in the second volume is greater than a threshold.

Abstract: An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

Abstract: A gas turbine engine includes a turbine for driving a shaft to in turn drive a rotor. A combustor receives compressed air, and combusts the compressed air. The shaft is supported by at least one bearing. An oil delivery system is associated with the at least one bearing including an oil scoop rotating with the at least one shaft and a stationary oil jet provided with a source of lubricant. A mixed fluid chamber is within the shaft. The mixed fluid chamber is connected to an air exit. There further is an oil chamber radially outward of the mixed fluid chamber. The oil chamber is connected to the at least one bearing.

Abstract: A system includes a burn flow line and a flow divider valve in fluid communication with the burn flow line to split flow from the burn flow line to a primary line and to a secondary line for supplying fuel to a set of primary engine nozzles and to a set of secondary engine nozzles, respectively. A sensor operatively connected to the flow divider valve to produce a feedback signal indicative of flow split between the primary line and the secondary line. A controller is operatively connected to receive the feedback signal from the senor, compare the feedback signal to a stored flow split value for a mismatch, and output an alert message upon detecting the mismatch.

Abstract: A turbine engine assembly includes a first case structure with a first flange and a second case structure that includes a second flange. The second flange is configured for securement to the first flange at a connection interface. At least one steam conduit is in thermal communication with the connection interface and configured to receive a portion of a flow of steam to heat the connection interface. Heating the connection interface provides for control of a thermal gradient generated by a difference in temperature in temperatures on either side of the connection interface.

Abstract: A combined gas turbine engine and hydrogen fuel cell system includes a hydrogen fuelled gas turbine engine, a cryogenic liquid hydrogen fuel tank, a first fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, a second fuel offtake arranged to divert a portion of hydrogen fuel from the main fuel conduit downstream of the heat exchanger, and a hydrogen fuel cell configured and arranged to produce electric power using hydrogen fuel diverted from the second fuel offtake.

Abstract: An aircraft propulsion system includes a compressor section where an inlet airflow is compressed, a combustor section where the compressed inlet airflow is mixed with fuel and ignited to generate an exhaust gas flow that is communicated through a core flow path, a turbine section through which the exhaust gas flow expands to generate a mechanical power output, the exhaust gas flow is split into at least a first exhaust gas flow and a second exhaust gas flow, and a condenser where water is extracted from the second exhaust gas flow, and a bottoming cycle where extracted water is transformed into a steam flow and injected into the core flow path.

Abstract: A combustion section of a gas turbine includes an annular combustion chamber, a burner arrangement upstream of the combustion chamber, an annular compressor diffusor, and an annular combustion casing. The combustion casing has an outer section wall, an upstream section wall, and an inner section wall. The inner section wall extends along the rotor axis and is arranged between the burner arrangement and the compressor diffusor. A shielding is attached to the inner section wall at the side facing the rotor axis with a free space in-between at least. An outer diffusor wall of the compressor diffusor extends in axial direction up to the upstream half of the shielding.