Abstract: Hybrid propulsion is obtained by arranging, around a conventional turbomachine, a reversible electric machine linked to the low-pressure shaft via a mechanical movement transmission, a one-way clutch being arranged on the low-pressure shaft between the fan and a low-pressure compressor. The low-pressure shaft includes two aligned portions which can be separated and reconnected. A casing includes a first portion able to tilt around a horizontal transverse axis of the aircraft relative to a second portions of the casing. One portion of the low-pressure shaft, the fan and the reversible electric machine are mounted on the first portion of the casing, and the other portion of the low-pressure shaft is mounted on the second portion of the casing. This device facilitates electric taxiing of the aircraft, using the machine as a motor, whereas the rear of the low-pressure shaft remains stationary.

Abstract: A combustor includes an inner liner forming a combustion chamber; an outer liner surrounding the inner liner to form a cooling passage in which compressed air flows; and a plurality of cooling guides installed around an inner circumferential surface of the outer liner to surround the combustion chamber, each of the cooling guides protruding from the inner circumferential surface to create an impinging jet from the compressed air flowing in the cooling passage. The plurality of cooling guides surrounding the combustion chamber are installed at regular intervals in a flow direction of the compressed air, and are arranged in staggered axial rows. Each cooling guide includes an air guiding surface facing the flow of the compressed air to guide the compressed air toward the inner liner. Accordingly, liner cooling efficiency can be enhanced by more effectively guiding the impinging jet toward the inner liner.

Abstract: An aircraft turbomachine, including a fan that is able to rotate inside a casing, and an electric machine including a rotor secured to the fan and a stator secured to the casing, wherein the rotor of the electric machine includes a ring that is able to rotate inside the stator, which is linked by arms to a cone mounted upstream from the fan.

Abstract: A method for operating a power plant having a gas turbine, a heat recovery steam generator, a steam turbine, an auxiliary heat source, and a control system, wherein the method includes controlling the power plant such that the heat recovery steam generator receives an input of heat from the gas turbine; determining the gas turbine is operating at its maximum capacity or at an upper end of its control range and the power plant is operating at less than a target value for a power plant capacity; determining a target pressure value immediately upstream of the steam turbine, wherein the target pressure value is derived from a primary pressure for the steam turbine and a steam turbine capacity for the steam turbine; based upon the target pressure value, controlling the heat store to release heat into the heat recovery steam generator to achieve the predefined power plant capacity.

Abstract: A gearbox assembly for a gas turbine engine includes a planetary gear set comprising a sun gear, a planet gear, and a ring gear, the sun gear configured for connection to a first drive shaft of the gas turbine engine and the ring gear configured for connection to a second drive shaft of the gas turbine engine, and an electric machine assembly comprising an input and an electric machine. The electric machine comprising a rotor coupled to the input and a stator fixedly positioned within the gearbox assembly, the input of the electric machine assembly coupled to one of the sun gear, the ring gear, or the planet gear of the planetary gear set.

Abstract: Described herein are embodiments of systems and methods for the removal of a direct drive unit (DDU) housed in an enclosure, such as a direct drive turbine (DDT) connected to a gearbox for driving a driveshaft connected to a pump for use in hydraulic fracturing operations.

Abstract: Fuel nozzle assemblies are provided. For example, a fuel nozzle assembly for a combustor system comprises a fuel nozzle having a pilot swirler and an outlet defined in an outlet end, as well as a main mixer attached to the outlet end and extending about the outlet. A total combustor airflow through the combustor system comprises a pilot swirler airflow that is greater than about 14% and a main mixer airflow that is less than about 50% of the total combustor airflow. In further embodiments, the fuel nozzle also comprises main and pilot fuel injectors that each are configured to receive a portion of a fuel flow to the fuel nozzle. The fuel nozzle provides less than about 80% of the fuel flow to the main fuel injector at a high power operating condition of a gas turbine engine in which the fuel nozzle assembly is installed.

Abstract: A combustor having bundled tube fuel nozzles is provided. At least one of the fuel nozzles has a dynamics-mitigating adapter removably coupled thereto. The adapter includes a mounting body defining at least one flow passage aligned with an inlet of at least one tube of the at least one fuel nozzle. The at least one flow passage extends an axial length of the at least one tube. The adapter may include extenders aligned with each tube of the fuel nozzle, and the extenders may have identical or different lengths. Adapters may be used for each fuel nozzle of the combustor. The mounting body may be a monolithic unit through which the flow passages are defined or may include a plurality of extenders affixed to and extending upstream of the mounting body.

Abstract: Systems and methods are disclosed for controlling surge margin in the compressor section of a gas turbine engine. A first compressor section and a second compressor section are in fluid communication with a bypass conduit. An auxiliary turbine and discharge conduit are positioned in the bypass conduit. Fluid flow from the compressor sections into the bypass conduit is controlled by bypass control valves.

Abstract: Described herein are embodiments of systems and methods for the removal of a direct drive unit (DDU) housed in an enclosure, such as a direct drive turbine (DDT) connected to a gearbox for driving a driveshaft connected to a pump for use in hydraulic fracturing operations.

Abstract: The electrical power generation system including a micro-turbine including a combustor, a first stage turbine configured to be driven by a combustor exhaust from the combustor, at least one compressor operably connected to the combustor to provide a compressed airflow to the combustor, a catalytic converter configured convert the combustor exhaust to a catalytic exhaust that includes at least exothermic heat, a second stage turbine configured to be driven by the catalytic exhaust from the catalytic converter, and one or more shafts connecting the first stage turbine and the second stage turbine to the at least one compressor such that rotation of the first stage turbine and the second stage turbine drives rotation of the at least one compressor.

Abstract: A gas turbine engine maintenance stand for a gas turbine engine that comprises modules including an engine core module that houses a high pressure compressor and a turbine module that houses a low pressure turbine. The gas turbine engine maintenance stand has a pair of base beams, each base beam having a first end, a midsection and a second end. The stand has an engine core module support that straddles the pair of base beams at their midsection, the engine core module support having engine core module support arms that are configured to engage opposing sides of the engine core module of the gas turbine engine. The stand also has a turbine module support that straddles the pair of beams adjacent their second ends, the turbine module support having turbine module support arms that are configured to engage opposing sides of the turbine module of the gas turbine engine.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a low speed spool and a high speed spool; a rotor assembly coupled to the low speed spool; an electric machine rotatable with the low speed spool for extracting power from the low speed spool, for adding power to the low speed spool, or both; and an inter-spool clutch positioned between the low speed spool and the high speed spool for selectively coupling the low speed spool to the high speed spool.

Abstract: A diffuser pipe includes a tubular body defining a pipe center axis extending therethrough. The tubular body includes a first portion extending in a first direction from an inlet of the tubular body, a second portion extending in a second direction transverse to the first direction and terminating at a pipe outlet, and a bend portion fluidly linking the first portion and the second portion. A stiffening rib extends outwardly from an outer surface of the first portion of the tubular body.

Abstract: Combustor systems are provided. For example, a combustor system comprises a combustor having forward and aft ends and including annular inner and outer liners that each extend generally along an axial direction and define a combustion chamber therebetween. The combustor system also comprises a fuel nozzle having an outlet defined in an outlet end of the fuel nozzle and including a pilot swirler. The outlet is positioned at the forward end of the combustor to direct a fuel-air mixture into the combustion chamber. The combustor system further comprises a main mixer attached to the outlet end of the fuel nozzle and extending about the outlet. A total combustor airflow through the combustor comprises a pilot swirler airflow that is greater than about 14% of the total combustor airflow and a main mixer airflow that is less than about 50% of the total combustor airflow.

Abstract: Described herein are embodiments of systems and methods for the removal of a direct drive unit (DDU) housed in an enclosure, such as a direct drive turbine (DDT) connected to a gearbox for driving a driveshaft connected to a pump for use in hydraulic fracturing operations.

Abstract: A ceramic matrix composite (CMC) component including a plurality of layers of a CMC and a directionally controllable CMC insert. The directionally controllable CMC insert is disposed in the plurality of layers of a ceramic matrix composite. The directionally controllable CMC insert includes an optimized architecture to strengthen a high stress region of the CMC component. The directionally controllable CMC insert is geometrically configured and disposed within the plurality of layers of the CMC to redirect a crack in the CMC component toward a region of low crack growth driving force. A turbomachine and method of forming a turbomachine member including a plurality of layers of a CMC and having the directionally controllable CMC insert disposed in a shaped void are additionally disclosed.

Abstract: A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.

Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

Abstract: A system for energy conversion is provided including a propulsion system, a fuel circuit, a combustion device and turbine separate from the propulsion system, and a load device. A fuel flow control device separates a flow of fuel from the fuel tank into a first portion of fuel and a second portion of fuel. A fuel circuit is configured to provide the first portion of fuel to a heat addition system at the propulsion system. A combustion device receives a flow of oxidizer from a compressor section of the propulsion system via a fluid circuit. The fuel circuit provides the second portion of fuel to the combustion device. The fluid circuit flows combustion gases to the turbine and the propulsion system. The load device is operably coupled to the turbine to receive an output torque via expansion of the combustion gases through the turbine.