Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Having wear liner – sheathing or insert
Reexamination Certificate
2001-09-27
2003-04-29
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Having wear liner, sheathing or insert
C416S09700R
Reexamination Certificate
active
06554575
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom. A high pressure turbine powers the compressor, and a low pressure turbine powers an upstream fan in a turbofan aircraft engine embodiment.
The first stage turbine blades first receive the hot combustion gases from the combustor and are typically air cooled by using air bled from the compressor. Turbine blade cooling is quite esoteric and the art is well crowded in view of the complex nature of blade cooling.
A typical turbine blade includes a generally concave pressure side and an opposite, generally convex suction side extending axially or chordally between leading and trailing edges which extend radially in span from root to tip of the blade. The airfoil portion of the blade is hollow and extends radially outwardly from a supporting dovetail which mounts the blade in a supporting rotor disk.
Cooling air is channeled to each blade through the dovetail and various internal passages are formed inside the airfoil for tailoring cooling thereof to mitigate the various heat loads experienced around the outer surface of the airfoil.
The radially outer end or tip of the airfoil is particularly difficult to cool since it is exposed to hot combustion gases along both the pressure and suction sides of the airfoil as well as in the radial clearance or gap formed with the surrounding stator casing or shroud. Since turbine blades are subject to occasional tip rubs, the airfoil tip is typically formed by squealer rib extensions of the pressure and suction sides which join together at the leading and trailing edges and define an open tip plenum therebetween having a floor which encloses the internal passages of the airfoil.
A significant advancement in blade tip cooling is U.S. Pat. No. 5,261,789 which discloses the use of a tip shelf along the pressure side of the turbine blade. The tip shelf is fed with cooling air through holes formed therethrough and interrupts the flow of combustion gases along the pressure side of the blade tip. Improved cooling of the blade tip including the pressure side tip rib is obtained.
The typical profile of a turbine blade is a crescent in which the blade increases in width aft from the leading edge and then converges and decreases in width to the narrow trailing edge thereof. The desired aerodynamic performance of the blade controls the aerodynamic profile of the pressure and suction sides and typically results in relatively thin trailing edge regions of the blade.
Since the tip shelf described above is recessed into the pressure side of the turbine blade, the blade must have sufficient thickness near the trailing edge in which the shelf may be formed.
However, in turbine blades having relatively thin trailing edge regions, there is insufficient width in which the tip shelf may be formed while having sufficient strength for a long useful life of the blade. In such configurations, the tip shelf may be terminated forward of the trailing edge and introduced in the blade where space permits. Terminating the tip shelf in this manner will form a vertical face where the tip shelf blends with the pressure side, with the boundary of the tip shelf having horizontal and vertical lines intersecting at an aft corner.
Since the combustion gas flow can be tripped or interrupted by the aft shelf corner during operation, such tripping of the gas flow can increase the combustion gas heat transfer so that as the combustion gases reattach to the pressure side downstream of the terminated tip shelf, local heating of the tip can result in rapid oxidation of the blade tip which adversely affects the useful life of the blade.
Accordingly, it is desired to provide a turbine blade having improved tip cooling notwithstanding termination of the tip shelf upstream of the blade trailing edge.
BRIEF SUMMARY OF THE INVENTION
A gas turbine engine blade includes pressure and suction sides extending between leading and trailing edges and root to tip. The pressure side includes a tip rib recessed therein to define a tip shelf terminating in an inclined ramp. The ramp may be aligned with streamlines of combustion gas flow for preventing interruption thereof and the increase of heat transfer therefrom.
REFERENCES:
patent: 5261789 (1993-11-01), Butts et al.
patent: 6422821 (2002-07-01), Lee et al.
GE Aircraft Engine, “Blade Tip,” in commercial use in USA more than one year.
Frey David Alan
Harris Daniel John
Keith Sean Robert
Leeke Leslie Eugene
Lenahan Dean Thomas
Andes William Scott
Conte Francis L.
General Electric Company
Look Edward K.
McCoy Kimya N
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