A pre-diffuser fairing for a gas turbine engine is disclosed. In various embodiments, the pre-diffuser fairing includes a first side wall, a first radially inward portion and a first radially outward portion and a second side wall, a second radially inward portion and a second radially outward portion. The first side wall and the second side wall are spaced apart to form a cavity configured to receive a strut.

The disclosure provides a blade with an airfoil including: a root region at a first radial end; a tip region at a second radial end opposite the first radial end; and a midspan region between the root region and the tip region and at least one endwall connected with the root region or the tip region of the airfoil along the suction side, the pressure side, the trailing edge and the leading edge, wherein the midspan region includes a reduced axial width relative to an axial width of the root region and an axial width of the tip region, and a reduced opening-to-pitch ratio at the midspan region relative to the root and tip regions.

A method of assembling and disassembling a gas turbine carries out various disassembly and reassembly processes depending on circumstances. In one process, a first-stage blade assembly and a first-stage vane assembly of a turbine section are disassembled from a gas turbine by sequential steps of disassembling a combustor assembly; disassembling a first-stage vane assembly; and disassembling a first-stage blade assembly. In another process, a fourth-stage blade assembly of a turbine section is disassembled from a gas turbine by sequential steps of disassembling a diffuser loading slot from a rear diffuser; and disassembling a fourth-stage blade assembly from a turbine disk. In another process, a rear bearing assembly of a turbine section is disassembled from a gas turbine by sequential steps of disassembling a rear diffuser cover from a rear diffuser; and supporting one end of a rotor shaft and disassembling a rear bearing from a rotor shaft support.

A gaspath component for a gas turbine engine includes a platform having at least one internal cooling passage. The at least one internal cooling passage has a plurality of trip strips extending into the cooling passage from at least one internal surface of the cooling passage. Each of the trip strips is defined by a z-shaped configuration.

A turbine rotor blade having an airfoil that includes a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of a pressure side as set forth in Table I. The Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along the Z axis. The X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value. The pressure side sections may be joined smoothly with one another to form the pressure side portion.

A turbine engine that includes an engine casing including a fluid supply plenum, a mating surface, and a nozzle supply passage and a cavity flow passage that both extend between the fluid supply plenum and the mating surface. The turbine engine further includes a turbine nozzle assembly including a mating band. The mating band includes an inlet scoop in flow communication with the nozzle supply passage. An interface is defined between the mating band and a first portion of the mating surface, and a band cavity is defined between the mating band and a second portion of the mating surface. The cavity flow passage couples the fluid supply plenum in flow communication with the band cavity.

A gas turbine, including: a combustion chamber; a high-pressure turbine with a first turbine guide vane ring that is arranged downstream of the combustion chamber, wherein the first turbine guide vane ring has a plurality of turbine nozzle guide vane segments that respectively include at least one guide vane, an outer platform, and an inner platform; and an outer housing. Provision is made that the turbine nozzle guide vane segments are fixed in the radial direction at the outer housing, wherein occurring radial loads are transferred into the outer housing. The invention further relates to a method for attaching a turbine nozzle guide vane segment of a gas turbine.

Disclosed herein is a gas turbine blade. The gas turbine blade includes a turbine blade (33) provided in a turbine, and film cooling elements (100), each including a cooling channel (110) for cooling of the turbine blade (33), an outlet (120) through which cooling air is discharged, and a plurality of ribs (130), wherein the outlet (120) extends from a longitudinally extended end of the cooling channel (110) to an outer surface of the turbine blade (33) and has a width increased from one end of the cooling channel (110) to the outer surface of the turbine blade (33), and the ribs (130) face each other on inner walls of the outlet (120).

A multi-piece fan spacer for a gas turbine engine includes at least one lug comprising a platform portion, a connection portion radially inward of the platform portion, and a support connecting the platform portion to the connection portion. The multi-piece fan spacer includes a plurality of platforms. Each of the platforms is connected to at least one axially adjacent platform portion.

A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting them. The engine includes a fan, with a plurality of fan blades, located upstream of the core and a gearbox receiving an input from the core shaft and outputting drive so the fan is at a lower rotational speed than the core shaft. The turbine includes a plurality of stages of axially spaced rotor blades mounted on a rotor, which are surrounded by a turbine casing. The turbine has an inlet defined at an upstream end of a first stage of blades and an outlet defined at a downstream end of a last stage of blades and a ratio of the area of the outlet to the inlet is at between 2.5 and 3.5. This increases the pressure ratio of and power extracted from the turbine and the engine.