A conventional gas turbine blade using a steam cooling system, in which high-pressure steam is supplied to an inner portion of the blade to cool the blade therewith and recover thermal energy, has been proposed but has a problem to be solved concerning its strength. A gas turbine blade capable of solving this problem free from the strength problem and sufficiently resistant to a flow of the high-pressure steam, wherein a cooling medium flow passage extending in the lengthwise direction of the blade is formed in the interior thereof, reinforcing ribs connecting front and rear walls of the blade together and extending in the direction of a flow of the cooling medium being provided in the cooling medium flow passage, whereby the strength of the blade is secured without hampering the flow of the cooling medium.

Gas turbine blade

An improved, highly stable blade-to-blade vibration damper (22) configuration is disclosed which provides substantially continuous blade vibration damping and sealing of an interplatform gap due to positional stability thereof. Disposed in a subplatform cavity, the damper (22) is comprised of a generally triangular shaped body having a vertex (42) thereof aligned with the interplatform gap (64). A primary damper load face (38) abuts a first inclined platform load face (66) and a secondary damper load (40) face abuts a second inclined platformload face (68). Maintenance of sliding planar contact between primary (38) and first load (66) faces and sliding linear contact between secondary (40) and second load (68) faces is afforded by orienting inclined platform faces to have an included angle greater than that of the damper vertex (42) and offsetting a damper center of gravity (58) toward the primary load face (38). The damper (22) may include one or more legs (52,54) to orient the damper in the cavity and one or more extending tabs (56) to discourage hot gas flow thereby. In a preferred embodiment, the damper is utilized as an as-cast member, without the need for additional processing or machining, and the vertex (42) included angle is greater than about 90 degrees.

Stable blade vibration damper for gas turbine engine

A particle separator (10) for use with a turbine engine (12) is arranged to divide a flow of air into a radially inner air flow and a radially outer air flow. The particle separator (10) is further arranged to separate inner particles from the inner air flow to allow air from the inner air flow to enter the turbine engine without the inner particles and arranged to separate outer particles from the outer air flow to allow air from the outer air flow to enter the turbine engine without the outer particles.

Particle separator for a turbine engine

An end seal for cooling air manifolds below the bases of a stage of air cooled turbine blades on a turbine wheel, the end seal including a seal ring having an annular planar portion abutting an end face of the turbine wheel inboard of the air manifolds and a right cylindrical flange perpendicular to the planar portion and underlying a corresponding cylindrical shoulder on the wheel and on the blade bases outboard of the air manifolds, and a plurality of end plates on the wheel having lugs underlying the cylindrical flange. The planar portion is axially clamped to the wheel to define a 360° inboard seal while centrifugal forces developed by the end plates during rotation of the wheel cause the lugs to radially clamp the cylindrical flange against the cylindrical shoulder on the wheel and blade bases to define a 360° outboard seal.

End seal for turbine blade bases

A shield is disposed in spaced relationship to the exterior surface of the combustor liner of a gas turbine engine and extends over a plurality of very small holes formed in the combustor liner so as to block flow of air to these holes. Further, the shield provides a space between the shield and the combustor liner for receiving reverse flow of air. A portion of the air flowing along the exterior surface of the combustor liner is diverted for reverse flow in the path provided between the shield and the combustor liner. A lesser plurality of substantially larger holes are provided in the combustor liner at the point of reversal of the air so that the dirt particles in the air, because of the centrifugal force acting thereon, tend to flow through these larger holes. A second group of larger holes are provided in the combustor liner approximately at the forward end of the space between the shield and the combustor liner. In flowing to the very small holes in the combustor liner the air must make a second reversal and substantially all of any remaining dirt particles are caused to flow through this second group of larger holes during this second reversal of air flow. Accordingly, the air reaching the very small holes in the combustor liner and being directed to the interior surface of the combustor liner for cooling thereof is substantially devoid of dirt particles and the possibility of plugging any of these very small holes is minimized.

Combustor liner cooling arrangement

The invention is an improved inlet particle separator for removing extraneous matter from a stream of air directed into the engine's core section. The improved separator utilizes two stages of separation. The first stage is an axial flow separator for initially separating engine inlet air into a first flow of relatively contaminated air and a second flow of relatively clean air. This first stage of separation is accomplished by sharply turning the air flow radially inwardly so that the relatively dense extraneous matter continues in its original direction into a scavenge system. A second stage of separation is accomplished in the scavenge system and comprises a centrifugal separator that separates the first flow of air into third and fourth flows of relatively contaminated air for the purpose of protecting a blower that powers the flow through the scavenge system and which is in flow communication with only the relatively less contaminated fourth flow of air.

Axial flow inlet particle separator

Disclosed is a bladed disk assembly having a new and improved rotor blade. The blade includes a shank and dovetail extending from an airfoil portion thereof. The shank includes a pocket extending therein from an end surface thereof. The shank pocket allows for reduced weight of the blade while maintaining acceptable bending stiffness thereof. In an exemplary embodiment of the invention, the bladed disk assembly is effective for being interchangeable with a blisk assembly having an inlet hub radius ratio of about less than about 0.5 and a root solidity of greater than about 2.2.

Gas turbine bladed disk assembly

In an axial flow apparatus including stator and rotor elements, means are included for constraining axial movement of individual blade members relative to the rotor axis when the rotor is subjected to high centrifugal loading. The constraining means include a ring which is circumferentially expandable under centrifugal loading. Part of the outward radial force causing the expansion of the ring is directed axially against the blade members to provide an axially constraining force which is proportional to the centrifugal loading on the rotor. The invention herein described was made in the course of or under a contract or subcontract thereunder, (or grant) with the Department of the Army.

Blade constraining structure

To reduce weight while maintaining acceptable bending stiffness, a gas turbine blade comprising a shank (38), a dovetail (40) and an airfoil portion (22) includes a pocket (58) extending from an end surface of the shank. In one embodiment a disk assembly is interchangeable with one having an inlet hub radius ratio of about less than 0.5. and a root solidity of greater than about 2.2.

Gas turbine blades

ADVANCED VANE FOR A DISC AND ROTOR VANE ASSEMBLY WITH LOW WEIGHT AND THE SAME STRENGTH CHARACTERISTICS AS CONVENTIONAL ROTOR ASSEMBLIES. IT INCLUDES: -A BLADE 22 HAVING A SUBBASE 26; -A FOOT 38 EXTENDING FROM THE BASE 26 OF THE BLADE AND; -A DOVETAIL 40 EXTENDING FROM FOOT 38. FOOT 38 HAS TWO OPPOSED SIDE SURFACES 48, AND A FIRST AND SECOND END SURFACES 50, 52 OPPOSED, SIDE SURFACES 48 BEING LONGER THAN THE SURFACES END 52 AND THE SECOND END SURFACE OF THE FOOT INCLUDING A SHELF 58 EXTENDING INNER TOWARD THE FIRST END SURFACE 52 OF THE FOOT AND BETWEEN THE DOVETAIL 40 AND THE BASE 26 OF THE BLADE 22. APPLICATION TO GAS TURBINE ENGINES.

David Dubble Klassen

Papers

Axial flow inlet particle separator

Gas turbine bladed disk assembly

Blade constraining structure

Gas turbine blades

Aube de rotor