The proposed solution relates to a nozzle assembly for a combustion chamber of an engine, having at least one nozzle, which includes a nozzle head that extends along a nozzle longitudinal axis and has at least one nozzle exit opening for injecting fuel into a combustion space of the combustion chamber, and at least one air-guiding duct, by way of which a swirl-affected air flow can be created in the direction of the combustion space along an outer lateral surface of the nozzle head.

At least one guide element which protrudes on the outer lateral surface of the nozzle head is in a flow path for the air flow of the at least one air-guiding duct and is configured to guide at least a part of the air flow radially inwards in relation to the nozzle longitudinal axis into a central injection region downstream of the nozzle exit opening.

A turbofan engine includes a propulsor section that has a propulsor shaft in driving engagement with a propulsor. An epicyclic gear system has a gear mesh lateral stiffness and a gear mesh transverse stiffness. A gear system input defines a gear system input lateral stiffness and a gear system input transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness. A first turbine section rotates at a first speed, and a second turbine rotates at a second speed that is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area of the first turbine, a second performance quantity is defined as the product of the second speed squared and the second area of the second turbine, and a performance quantity ratio is between 0.5 and 1.5.

A rotor blade is provided. The rotor blade includes a main body having a shank, an airfoil extends radially outwardly from the shank, and a platform. The main body includes a pressure side slash face and a suction side slash face. A slot is defined within each of the pressure side slash face and the suction side slash face. The slot of the pressure side slash face and the slot of the suction side slash face each include an upstream end portion that defines an end and a main body portion extending from the upstream end portion. The upstream end portion tapers from the end to the main body portion. The main body portion further includes a retention wall that covers a portion of the end and that defines an opening. The retention wall further includes an inner retention surface. The retention wall defines an offset from the opening.

A multipoint fuel injection system comprises an injection system segment including a circumferentially extending outer support defining a fuel manifold with a plurality of manifold passages extending circumferentially therethrough. A first connector is included at a first circumferential end of the outer support and a second connector is included at a second circumferential end of the outer support opposite the first circumferential end. The first and second connectors are each configured to connect each manifold passage with a manifold passages of a respective outer support of a circumferentially adjacent injection system segment. The system includes a circumferentially extending inner support and a plurality of circumferentially spaced apart feed arms extending radially between the inner support and the outer support. A plurality of outlet openings extend in an axial direction from each feed arm for feeding respective injection nozzles.

A gas turbine engine includes a turbine section that has a plurality of turbine vanes. Each of the turbine vanes includes inner and outer platforms and an airfoil section that extends there between. The airfoil section is hollow and rib-less and has a first end at the outer platform and a second end at the inner platform. The airfoil section is tangentially bowed from the first end to the second end with a radius of curvature that is from 17 centimeters to 130 centimeters.

A method of disassembling a rotor module of a gas turbine engine. The gas turbine engine having a rotor output shaft. The rotor module having a centre-bolt, a sleeve, at least one rotor stage, at least one stator stage, a casing and an axis. The method having the steps: attaching a fixture to the at least one rotor stage, attaching the fixture to the casing, detaching the centre-bolt from the at least one rotor stage, detaching the sleeve from the output shaft, attaching the fixture to the sleeve, and removing the rotor module and fixture from the rotor output shaft. There is also presented a method of assembling the rotor module to the gas turbine engine and the apparatus used for disassembly and assembly.

This blade repair method has: a first welding step in which overlay welding in which a first welding material is used is performed to form a notched part and a bury a first region positioned on a blade-body side with a first welding material; and a second welding step in which, after the first welding step, overlay welding in which a second welding material is used is performed to form a notched part and bury a second region positioned on a front-surface side of a platform with the second welding material. The high-temperature strength of the second welding material is higher than the high-temperature strength of the first welding material, the weldability of the first welding material is higher than the weldability of the second welding material, and the second region is located in a range from 1.0 mm to 3.0 mm (inclusive) from the front surface of the platform toward the blade body.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.

A gear reduction drives a fan rotor at a speed slower than a fan drive turbine. The turbine section further includes a high pressure turbine driving a high pressure compressor. The fan drive turbine and low pressure compressor are connected by a shaft and the low pressure turbine. The shaft and the low pressure compressor define a low pressure spool, the low pressure spool has a torque at maximum takeoff defined in ft-lbs and also having a low pressure spool power defined in horsepower and at maximum takeoff, and a ratio of the low pressure spool torque to the low pressure spool power being defined, with the low pressure spool power being defined in horsepower, and the ratio of the low pressure spool torque to the low pressure spool power being greater than or equal to 0.6 ft-lb/hp and less than or equal to 1.2 ft-lb/hp.

A component for a gas turbine engine. The component includes a continuous fiber blade including an airfoil extending radially between a root and a tip and a blade attachment feature positioned at or adjacent to the root. The component further includes a platform coupled to the root of the continuous fiber blade. The platform includes a plurality of chopped fibers. Additionally, the component includes a thermoplastic polymer contained in both the continuous fiber blade and the platform. Moreover, the continuous fiber blade and platform are coupled together such that the continuous fiber blade and platform form a monolithic composite body.