Abstract: A fuel injection nozzle for a combustion turbine engine has thermal stress-relief vanes, which accommodate and relieve localized thermal stresses within its monolithic, three-dimensional nozzle structure, imparted by heat transfer during engine combustion. At least one first vane is coupled to opposing, spaced nozzle sleeves at both ends. At least one cantilever-like second vane is coupled to one of the opposing sleeves on one end, while the other free or floating end is spaced by a second vane gap from the other opposing sleeve. Some embodiments include a plurality of second vanes, which have locally varying orientation, and/or structure, and/or second vane gaps, for normalizing spatially and/or temporally thermal stresses within the nozzle structure. The monolithic structure is fabricated, in some nozzle embodiments, by additive manufacturing.

Abstract: A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; a combustor; an electrical energy storage unit; and an engine controller configured to: in response to receipt of a request for an increased electrical power supply by the engine (dPD) within a requested timeframe (dt), evaluating a current HP compressor surge margin (dRH/RH) and a current LP compressor surge margin (dRL/RL) based on the operating points thereof; in response to evaluating that one or more of the HP compressor and the LP compressor has insufficient surge margin to facilitate a change of electrical power supply at a requested rate (dPD/dt), meeting the requested electrical power demand using the electrical energy storage unit whilst increasing fuel flow to the combustor to accelerate the engine.

Abstract: A system for starting a turbine engine. The system may comprise a gearbox, a first starter, and a second starter. The gearbox may have a gearbox input shaft. The gearbox may be coupled to the turbine engine. The gearbox input shaft may be rotatively coupled to a spool of the turbine engine. The first starter may be coupled to the gearbox input shaft. The second starter may have a second-starter output shaft. The second-starter output shaft may be coaxial with the gearbox input shaft. The second starter may be coupled to the gearbox input shaft through the first starter.

Abstract: A gas turbine engine includes a rotatable first shaft, a first disk connected to the first shaft, a second disk connected to the first shaft, a combustor radially outward from the first disk and the second disk, and a heat exchanger connected to the combustor aft of the second disk. The first disk includes a row of low pressure compressor blades and a row of high pressure turbine blades connected to a radially outer end of the row of low pressure compressor blades. The second disk includes a row of high pressure compressor blades and a row of low pressure turbine blades connected to a radially outer end of the row of high pressure compressor blades.

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 power extraction system for a gas turbine engine may comprise a low spool transmission and a low spool accessory gearbox. The low spool accessory gearbox may comprise a generator and a dual clutch transmission. The dual clutch transmission may be coupled between the low spool transmission and the generator.

Abstract: In an embodiment, a system includes a gas turbine controller. The gas turbine controller is configured to receive a plurality of sensor signals from a fuel composition sensor, a pressure sensor, a temperature sensor, a flow sensor, or a combination thereof, included in a gas turbine engine system. The controller is further configured to execute a gas turbine model by applying the plurality of sensor signals as input to derive a plurality of estimated gas turbine engine parameters. The controller is also configured to execute a flame holding model by applying the plurality of sensor signals and the plurality of estimated gas turbine engine parameters as input to derive a steam flow to fuel flow ratio that minimizes or eliminates flame holding in a fuel nozzle of the gas turbine engine system.

Abstract: An aircraft turbine engine assembly including a nacelle structure with a strut attachment zone, a first cover forming an air intake lip, a hinged cover arranged on the nacelle upper portion, between the first cover and the attachment zone, a hinge system of the cover comprising, on either side of a mid-plane of the nacelle, a front rod and a rear rod that are one in front of the other. The two ends of each rod are co-planar, parallel to the mid-plane. A first end of each rod is hinged on the cover. A second end of each rod is hinged on the structure. A locking system is actuatable from the outside of the nacelle between a locked and an unlocked positon. The nacelle comprises at least one cylinder with a first end mounted hinged on the structure and the other end mounted hinged on the cover.

Abstract: A system may be provided that includes a hot section component of a gas turbine engine. The hot section component includes a dual wall, which includes a first wall and a second wall. The first wall includes multiple impingement cooling holes extending through the first wall. The second wall is positioned adjacent the first wall. The first wall and the second wall together define a cooling passage between the first wall and the second wall. Multiple pockets are in a surface of the second wall. Each of the pockets is positioned opposite a respective one of the impingement cooling holes. Each of the pockets is configured to receive a cooling fluid from the respective one of the impingement cooling holes and direct the cooling fluid into the cooling passage. The cooling passage includes a single cooling passage into which the pockets are configured to direct the cooling fluid.

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 turbomachine assembly for an aircraft comprising a nacelle with a nacelle structure, cowls attached to the structure to create an aerodynamic surface, an articulated cowl, comprising four edges, in the upper portion of the nacelle, an articulation at a first edge of the articulated cowl attached between the articulated cowl and the nacelle structure allowing the articulated cowl to move between a closed and an open position, a locking system at a second edge of the articulated cowl, opposite the first edge. The locking system comprises a handle actuable from the exterior of the nacelle, between a first position and a second position, a locking device movable between a locking position and an unlocking position, and a transmission system that moves the locking device from the locking position to the unlocking position when the handle passes from the first position to the second position, and vice versa.

Abstract: A combustor provided with: a nozzle having formed therein an air jet section that causes air to be jetted from a nozzle section; a cylindrical part covering the nozzle from the outer peripheral side thereof and forming an air flow path between the cylindrical part and the nozzle; and a pressure loss section provided to the air flow path. The pressure loss section causes a loss of pressure in the air that flows through the air flow path. The nozzle is provided with: at least one air inlet section that takes in air from an outer peripheral surface that is an upstream side from the pressure loss section; and a flow path-forming section forming a discharge air flow path that guides air that is taken in from the at least one air inlet section to the air jet section.

Abstract: The invention aims to shorten the time required for start-up and prevent excessive increases in the heat loads on turbine blades.

Abstract: A hybrid-electric propulsion system for an aircraft includes a turbomachine, the turbomachine including a first spool and a second spool. A method for operating the hybrid electric propulsion system includes operating, by one or more computing devices, the turbomachine such that the first spool mechanically drives a prime propulsor of the hybrid-electric propulsion system; and modifying, by the one or more computing devices, a speed relationship parameter defined between the first spool and second spool by providing electrical power to, or drawing electrical power from, an electric machine mechanically coupled to the first spool, the second spool, or both.

Abstract: The present disclosure relates to systems and methods that provide a low pressure liquid CO2 stream. In particular, the present disclosure provides systems and methods wherein a high pressure CO2 stream, such as a recycle CO2 stream from a power production process using predominately CO2 as a working fluid, can be divided such that a portion thereof can be expanded and used as a cooling stream in a heat exchanger to cool the remaining portion of the high pressure CO2 stream, which can then be expanded to form a low pressure CO2 stream, which may be in a mixed form with CO2 vapor. The systems and methods can be utilized to provide net CO2 from combustion in a liquid form that is easily transportable.

Abstract: A turbomachine comprising an engine body and an annular nacelle. The nacelle has an air intake comprising an outer skin, an inner skin, an air intake lip connecting the outer and inner skins together, and an annular frame connecting the outer and inner skins together. The turbomachine comprises a system for de-icing the air intake having an air bleeding port for bleeding a flow of hot air from the engine body, and two interconnected heating systems. An inner skin heating system receives the hot air flow to heat, by means of a gas-liquid heat exchanger, a heat transfer liquid contained in a hydraulic circuit arranged in part in the heat exchanger and in the inner skin. An air intake lip heating system diffuses the flow of hot air, once it has passed through the inner skin heating system, through the annular frame to heat the air intake lip.

Abstract: An anti-icing system is disclosed. In various embodiments, the anti-icing system includes an inner duct having a first end configured to deliver heated gas to a plenum and a second end spaced from the first end; an outer duct circumferentially encompassing at least a portion of the inner duct; and a seal system disposed proximate the second end, the seal system including a first annular seal having a radially inner end positioned proximate a flange disposed on the inner duct.

Abstract: A softwall containment system for a machine includes an inner wall circumscribing a bladed rotatable member of the machine and extending from a forward end to an aft end. The softwall containment system also includes an outer wall circumscribing the inner wall and extending from a forward end to an aft end. The outer wall is spaced radially outwardly from the inner wall. The outer wall extends axially between a first joint coupling the forward end of the inner wall to the forward end of the outer wall and a second joint coupling the aft end of the inner wall to the aft end of the outer wall. The softwall containment system also includes an anti-ballistic material wrap covering the outer wall, the anti-ballistic material wrap including at least one of a first extension extending forward of the first joint and a second extension extending aft of the second joint.

Abstract: A bypass turbojet nacelle comprising a fixed structure, a fixed cowl and a cowl movable in translation between advanced and withdrawn positions by a ram with a reverser flap movable thereon, and a deployment system comprising a first slider integral with the movable cowl, a first groove comprising a first part and a curved, second part integral with the fixed structure, a second slider movable in translation in the second groove part, a first connecting rod rotatable on the first slider and on the second slider, and a second connecting rod articulated on both the first connecting rod first end and the reverser flap. The ram rod is articulated on the first connecting rod. Moving the first slider from the advanced to the withdrawn position moves the second slider along the first part and, when the first slider reaches the withdrawn position, the second slider moves along the second part.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.