Shrouded turbine blade with contoured platform and axial dovetail

Abstract

A turbine blade includes an airfoil having a root, a tip, a concave pressure side, and a laterally opposite convex suction side, the pressure and suction sides extending in chord between opposite leading and trailing edges; an outer platform disposed at the tip of the airfoil, the outer platform having spaced-apart lateral edges which each define an interlocking element; an inner platform with two spaced-apart curved lateral edges disposed at the root of the airfoil, the inner platform having a hot side facing the airfoil which is contoured in a non-axisymmetric shape; and a dovetail extending radially inward from the opposite side of the inner platform, wherein the dovetail is axially straight.

Description

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH AND DEVELOPMENT[0001]The U.S. Government may have certain rights in this invention pursuant to contract number W911W6-07-2-0002 awarded by the Department of the Army.BACKGROUND OF THE INVENTION[0002]The present invention relates generally to gas turbine engines, and more specifically, to turbines therein.[0003]In a gas turbine engine, air is pressurized in a compressor and subsequently mixed with fuel and burned in a combustor to generate combustion gases. One or more turbines downstream of the combustor extract energy from the combustion gases to drive the compressor, as well as a fan, shaft, propeller, or other mechanical load. Each turbine comprises one or more rotors each comprising a disk carrying an array of turbine blades or buckets. A stationary nozzle comprising an array of stator vanes having radially outer and inner endwalls in the form of annular bands is disposed upstream of each rotor, and serves to optimally direct the ...

AI Technical Summary

Benefits of technology

"The present invention provides a turbine blade with a 3D-countoured inner band surface and an axially straight dovetail. The blade has an airfoil with a root, tip, concave pressure side, and convex suction side, and an outer platform with interlocking elements. The blade also has an inner platform with curved lateral edges and a hot side facing the airfoil. The blade is disposed in an annular side-by-side array to define a plurality of flow passages each containing two inner platforms, two outer platforms, and adjacent airfoils. The inner platforms have a non-axisymmetric shape including a peak and trough that define an arcuate channel extending axially along the inner platform. The technical effects of the invention include improved cooling efficiency, reduced pressure drop, and improved airflow control."

Problems solved by technology

Undesirable pressure losses in the combustion gas flowpaths therefore correspond with undesirable reduction in overall turbine efficiency.

One common source of turbine pressure losses is the formation of horseshoe and passage vortices generated as the combustion gases are split in their travel around the airfoil leading edges.

The interaction of the pressure and suction side vortices occurs near the mid-span region of the airfoils and creates total pressure loss and a corresponding reduction in turbine efficiency.

These vortices also create turbulence and increase undesirable heating of the endwalls.

Since the horseshoe and passage vortices are formed at the junctions of turbine rotor blades and their integral root platforms, as well at the junctions of nozzle stator vanes and their outer and inner bands, corresponding losses in turbine efficiency are created, as well as additional heating of the corresponding endwall components.

Because of the manufacturing and assembly tolerances, the trailing edge ridge will be interrupted and may see a forward facing step that adversely affects performance.

However, this has only been possible with a blade that does not have a tip shroud and can be individually assembled into a rotor disk.

For a shrouded turbine blade, the interlocking tip shroud, the curved platform and a conventional curved dovetail make rotor assembly impossible.

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