Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a compressor section, a combustor, and a turbine section arranged in serial flow order, the turbomachine further including an outer casing and, during operation of the gas turbine engine, an exhaust gas flow from the turbine section; and a fuel cell assembly positioned within the outer casing of the turbomachine, the fuel cell assembly having a fuel cell defining an inlet positioned to receive a portion of the exhaust gas flow, an outlet positioned to provide output products directly to the exhaust gas flow, or both.

Abstract: A gas turbine engine includes a compressor section and a turbine section. A heat exchanger is operable to cool a fluid utilized by a component of the gas turbine engine. The heat exchanger has an inlet and an outlet for the accessory fluid. A strain gauge is associated with the fluid inlet and operable to produce a signal indicative of a temperature of the fluid at the fluid inlet and a strain gauge is associated with the fluid outlet and operable to produce a signal indicative of a temperature of the fluid at the fluid outlet. A controller determines a temperature differential across the heat exchanger by utilizing the signal from the strain gauge at the fluid inlet and the signal from the strain gauge at the fluid outlet. A method and a heat exchanger are also disclosed.

Abstract: A method for testing a hybrid power plant equipping a rotary-wing aircraft comprising at least one lift rotor, the hybrid power plant being configured to rotate the at least one lift rotor, the hybrid power plant comprising at least one heat engine configured to rotate the at least one lift rotor; and at least one electric motor supplied with electrical energy by at least one electrical power source, the at least one electric motor being configured to rotate the at least one lift rotor at least when at least one of the at least one heat engine fails.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a compressor section, a combustor, and a turbine section arranged in serial flow order, the turbomachine further including an outer casing; a fuel cell assembly positioned within the outer casing of the turbomachine, the fuel cell assembly including a fuel cell and an output products line in fluid communication with an outlet of the fuel cell for receiving output products from the fuel cell; a fuel delivery assembly including a fuel line; and a reformer in communication with the fuel line and the output products from the fuel cell during operation of the gas turbine engine to convert a fuel flow through the fuel line to a hydrogen rich gas.

Abstract: A hybrid aircraft power plant includes an air mover for propelling an aircraft, a gas turbine engine operable to drive the air mover, a bifurcated exhaust duct conveying combustion gas from the gas turbine engine, and an electric motor mounted to the exhaust duct and operable to drive the air mover. The turbine engine includes a turbine extracting energy from an annular stream of the combustion gas and being rotatable about a rotation axis. The exhaust duct includes a first branch conveying a first portion of the combustion gas toward a first direction away from a rotation axis of the turbine, and a second branch conveying a second portion of the combustion gas toward a second direction away from the rotation axis.

Abstract: An assembly has: a gas turbine engine including: a compressor including a rotor including blades rotatable about a central axis and disposed in a gas path of the gas turbine engine, a stator including one or more vanes disposed in the gas path, and a compressor case disposed around the stator; a combustor downstream of the compressor along the gas path; and a combustor case surrounding the combustor and secured to the compressor case; and a retainer for interlocking the stator with the compressor case when disassembling the combustor case from the compressor case, the retainer including: a base secured to the compressor case; and a pin extending from the base and radially inwardly through the compressor case via an aperture defined in the compressor case, the pin engaging with the stator to prevent axial movement of the stator relative to the compressor case.

Abstract: An anti-ice swirl system for an aircraft engine includes a nacelle including a first nacelle section exposed to ambient environment and facing an oncoming airstream when an aircraft powered by the engine is in flight. The first nacelle section defines a circumferential D-duct. The anti-ice swirl system also includes a nozzle configured to inject into and circulate through the D-duct a controlled unsteady flow of heated fluid. The controlled unsteady flow mixes newly injected heated fluid with air inside the D-duct and mitigates icing of the first nacelle section when the aircraft is in flight.

Abstract: An engine system includes an engine core assembly, a fuel system and a cooling circuit. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The combustor section includes a combustor adjacent a plenum. The fuel system includes a fuel flowpath, a fuel reservoir, a fuel-cooling air heat exchanger and a fuel injector. The fuel flowpath is configured to direct hydrogen fuel, received from the fuel reservoir, through the fuel-cooling air heat exchanger to the fuel injector. The fuel injector is configured to introduce the hydrogen fuel into the combustor. The cooling circuit is configured to direct cooling air, bled from the plenum, through the fuel-cooling air heat exchanger to the turbine section. The fuel-cooling air heat exchanger is configured to exchange heat energy between the hydrogen fuel flowing within the fuel flowpath and the cooling air flowing within the cooling circuit.

Abstract: An ignition and combustion assist system and method comprising a plasma igniter and electronic driver unit for use with gas turbine engines operating under low air densities, reduced voltage conditions and overall pressure ratios of 3:1 to 7:1. The plasma igniter has an inner chamber housing a centrally positioned and electrically isolated electrode attached to an electrical lead, driver unit, and AC or DC power supply. The electrode features a corner positioned near an outlet end of the igniter, where a plasma arc ignites a fuel-air mixture creating a flame extending into a primary burn region of a combustor of the gas turbine. The driver unit is in two embodiments and configured with low-cost microsecond voltage wave time periods or energy-efficient nano-second pulses. The method uses the plasma igniter and the electronic driver units described herein separately with other components or together.

Abstract: A bleed system including control circuitry and a variable jet pump. The control circuitry is configured to receive a signal indicative of a fluid parameter in the bleed system and cause the jet pump to alter a mixing ratio of a higher pressure gas and a lower pressure gas based on the signal. The jet pump is configured to combine the lower pressure gas and the higher pressure gas in the mixing ratio to generate a mixed gas. The jet pump is configured to supply the mixed gas to one or more gas loads in the bleed system. In examples, the control circuitry is configured to establish a system setpoint for the fluid parameter based on an operating status of the one or more gas loads.

Abstract: A gas turbine facility includes: a gas turbine; an ammonia supply device that supplies ammonia to a combustor of the gas turbine; a flow path forming frame forming an exhaust gas flow path through which exhaust gas from the gas turbine flows; a sprinkling device having a sprinkler that sprinkles water into the exhaust gas flow path; and a sprinkling controller that controls an operation of the sprinkling device. The sprinkling controller instructs the sprinkling device to start sprinkling the water on condition that an instruction to start supply of the ammonia to the combustor from the ammonia supply device or an instruction to stop the supply of the ammonia is received.

Abstract: Even if the fuel pressure to a torch for igniting an air-fuel mixture in a gas turbine engine is set to an ideal value, the actual fuel pressure supplied to the torch may differ due to modifications, drift, changes in ambient temperature or pressure, fuel composition, and/or the like. This may result in failure of the gas turbine engine to ignite. Accordingly, disclosed embodiments introduce a torch active pressure control (TAPC) system into the flow path of fuel to the torch. The TAPC system is controlled to modulate the fuel pressure of the fuel supplied to the torch, according to a modulation profile. The modulation profile may traverse a range of fuel pressures in each of one or a plurality of sinusoidal, V-shaped, stepped, or other-shaped cycles, to increase the probability that the ideal fuel pressure, and therefore, flame height, is achieved, thereby increasing the likelihood of a successful lightoff.

Abstract: Example systems and methods for integrating an air-cooled oil cooler into an inlet frame of an engine are provided. An example gas turbine engine comprises an inlet frame including an inlet plenum at a first end of an inlet frame flow path, the inlet frame defining an airflow entrance to a compressor; the compressor to provide an airflow through the inlet frame; a pump to circulate a fluid; a plurality of struts supporting the inlet frame flow path; and a fluid core connected to the pump and built at least partially into the inlet frame of the gas turbine engine, the fluid core to route the fluid through the inlet plenum.

Abstract: A compressor has: a rotor and a shroud having a internal and external sides and an air outlet; an annular case mounted to the external side and enclosing a plenum, the annular case having inner and outer sides and defining an aperture; a valve mounted to the annular case at the aperture and having: a support secured to the annular case and defining a seat surrounding a bleed passage, the seat located on the outer side, the support defining a guiding aperture, and a valve member having a head and a stem slidably received within the guiding aperture, the valve member having a closed position and an open position, the head abutting the valve seat in the closed position and offset from the valve seat in the open position; and an actuator engaged to the valve member for moving the valve member between the closed position and the open position.

Abstract: A system includes an engine case for a gas turbine engine defined around a longitudinal axis. An access opening is defined through the engine case for access from outside the engine case to a space inside the engine case. A combustor is housed in the space inside the engine case. The combustor includes an inner annular wall and an outer annular wall radially outboard from the inner annular wall. The inner annular wall includes a first rail on an upstream end thereof. The outer annular wall includes a second rail on an upstream end thereof. The first and second rails each include a respective access portion configured to receive line replaceable injector components from through the access opening of the engine case into the first and second rails to form a combustor dome of the combustor.

Abstract: A turbomachine turbine assembly having a longitudinal axis and including an exhaust casing, an ejection cone arranged downstream of the exhaust casing, the ejection cone having an outer annular wall for the flow of a primary air stream and an acoustic box arranged radially inside said outer annular wall. The acoustic box can include a radially inner annular wall and a connection member interposed longitudinally between the exhaust casing and the inner annular wall of the ejection cone. The connection member can be fastened to the exhaust casing and to the inner annular wall. The upstream end of the outer annular wall can be connected to the connection member with a degree of freedom in the radial direction and a degree of freedom in the longitudinal direction.

Abstract: A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor to combust fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; a bearing supporting the fan shaft; an oil loop system supplying oil to the bearing; and a heat exchange system including: an air-oil heat exchanger through which oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow to transfer heat between the oil and the fuel; and a bypass pipe to allow a proportion of the oil to flow past one of the air-oil and the fuel-oil heat exchanger; and a bypass valve to allow the proportion of the oil sent through the bypass pipe to be varied.

Abstract: An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

Abstract: An example system includes a first gas-turbine engine configured to propel an aircraft, the first gas-turbine engine comprising a first electric starter, the first electric starter configured to rotate a spool of the first gas-turbine engine; and one or more controllers collectively configured to: cause, following operation of the first gas-turbine engine, the first electric starter to perform barring of the first gas-turbine engine; measure, during the barring of the first gas-turbine engine, values of one or more parameters of the first gas-turbine engine; and determine, based on the values of the one or more parameters, whether the first electric starter is available for use in performing mid-air restart of the first gas-turbine engine.

Abstract: The present disclosure provides for improving flame propagation in a combustor, particularly in a combustor for use in a power production system and method. At least one stream being passed into the combustor (e.g., a fuel, an oxidant, a diluent, a coolant, a working fluid, water, or steam) can be independently heated to a temperature that is about the autoignition temperature of the fuel or greater. Further, flame stabilization, including promoting ignition, can be improved as described herein through controlled addition of one or more NOx species into the combustor or combustion chamber.