A moving blade 21d is disposed in a last stage closest to a diffuser 10 among a plurality of stages of a turbine 9 including a turbine rotor 12 and a stationary body 14. The diffuser 10 is connected to an outlet side of the stationary body 14. A distal end of the moving blade 21d is opposed to a seal fin 38 provided in the stationary body 14. The moving blade 21d includes a blade section 26, a cover 27 and a guide 32 provided on a moving blade distal end face 31, which is a surface of the cover 27. The moving blade distal end face 31 extends in a rotation axis direction of the turbine rotor 12, and the guide 32 includes a guide surface 41 located on a side close to the diffuser 10 with respect to the seal fin 38 and formed to incline upward in a direction from the seal fin 38 toward the diffuser 10.

An apparatus and a process for converting a twin spool aero gas turbine engine to an industrial gas turbine engine, where the fan of the aero engine is removed and replaced with an electric generator, a power turbine is added that drives a low pressure compressor that is removed from the aero engine, variable guide vanes are positioned between the high pressure turbine and the power turbine, and a low pressure compressed air line is connected between the outlet of the low pressure compressor and an inlet to the high pressure compressor, where a hot gas flow produced in the combustor first flows through the high pressure turbine, then through the low pressure turbine, and then through the power turbine.

The present invention relates to a rotor blade arrangement for a turbomachine, with a rotor blade which has a sealing tip radially on the outside, and with a seal arrangement, wherein the seal arrangement forms a radially inwardly open cavity, in which the sealing tip is arranged, to which end the seal arrangement has a first sealing element, namely a first seal carrier with a first run-in coating, and a second sealing element, wherein the first run-in coating delimits the cavity radially on the outside, and the second sealing element delimits the cavity in an axial direction, and wherein the first and the second sealing element are assembled.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A turbine for a gas turbine engine includes, among other things, a shaft rotatable about a longitudinal axis, a turbine rotor including one or more rows of turbine blades and a disk assembly coupled to the shaft. The disk assembly includes one or more disks each having an attachment region extending radially between an inner boundary and an outer boundary, the outer boundary is established by an outer periphery of the respective disk, the attachment region defines an array of slots distributed about the outer periphery, each of the slots extends radially inwardly from the outer boundary to the inner boundary, and each of the slots is dimensioned to receive a root section of a respective one of the turbine blades to mount the turbine blades to the disk assembly.

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is between 0.50 and 0.55, or is greater than 0.55 and less than or equal to 0.65.

To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

A turbine includes a rotor including a rotation shaft that rotates around an axis and a blade row formed on an outer surface of the rotation shaft; a casing, which covers the rotor, has a casing inner surface being expanded radially outward approaching a downstream side of the casing in a direction of the axis; and an inner member body formed to line the casing inner surface of the casing such that an extraction port is formed between an upstream side end of the inner member body and the casing inner surface. A discharge port is formed between a downstream side end of the inner peripheral member body and the casing inner peripheral surface. The extraction port and the discharge port are formed in an annular shape centered on the axis. A flow path cross-sectional area of the discharge port is smaller than that of the extraction port.

A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

A compressor section is in fluid communication with a fan, which includes a first compressor section and a second compressor section. A turbine section includes a first turbine section driving the fan and the first compressor section and a second turbine section driving the second compressor section and the second compressor rotor. A first performance quantity is defined as a product of the first speed squared and the first area. A second performance quantity is defined as a product of the second speed squared and the second exit area. A performance ratio of the first performance quantity to the second performance quantity is between about 0.2 and about 0.8. A gear reduction is included between the first turbine section and the first compressor section.