Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2000-02-18
2001-12-04
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S09700R
Reexamination Certificate
active
06325593
ABSTRACT:
BACKGROUND OF THE INVENTION
The Government has rights to this invention pursuant to a contract by the United States Air Force.
FIELD OF THE INVENTION
This invention relates to cooling of gas turbine engine turbine vanes and blades and, more particularly, to film cooling of airfoil trailing edges of ceramic vanes and blades.
DISCUSSION OF THE BACKGROUND ART
A gas turbine engine includes a compressor that compresses air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases flow downstream through one or more stages of turbines which extract energy therefrom for powering the compressor and producing additional output power for driving a fan for powering an aircraft in flight for example. A turbine stage includes a row of turbine rotor blades secured to the outer perimeter of a rotor disk, with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk.
The blades and vanes often include airfoils with hollow interiors extending between leading and trailing edges of the airfoil. Cooling air is flowed into the hollow interior for internal cooling of the airfoil and flowed out through film cooling holes for external cooling of the airfoil. The temperatures within gas turbines may exceed 2500 degrees Fahrenheit, and cooling of turbine vane and blade airfoils is very important in terms of vane and blade longevity. Without cooling, turbine vane and blade airfoils would rapidly deteriorate. Improved cooling for turbine airfoils is very desirable and much effort has been devoted by those skilled in the blade cooling arts to devise improved cooling designs in order to enhance cooling. The turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air necessarily decreases the overall efficiency of the engine. Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Different cooling passage configurations may be used within the hollow interior of the airfoil. Straight pass through of cooling air or impingement cooling air using impingement baffles are two types of cooling configurations used within the hollow interior to cool the airfoil. Film cooling air is discharged from the hollow interior through various film cooling holes disposed around the leading edge and mid-chord areas of the outer surface of the airfoil.
Since the overall efficiency of the gas turbine engine is directly related to the temperature of the combustion gases, engine efficiency is limited by the ability to operate the turbine nozzle at high temperature. Conventional turbine vane and blade materials are superalloys, such as single crystal nickel based materials, which have allowed engines to be operated at relatively high thermal efficiency. Further advances in engine efficiency by increasing combustion gas temperature has led to development of a class of high temperature ceramic materials one particular useful class of which is referred to as ceramic matrix composite (CMC) which have substantially higher temperature capabilities than conventional superalloys. CMC materials also maintain strength at relatively higher temperatures than that of conventional superalloys.
However, ceramic materials are relatively brittle when compared to conventional superalloys with a corresponding loss of toughness inherent therein. These materials also have substantially different coefficients of thermal expansion, or different thermal conductivities, compared to conventional superalloys. Advanced military engines are being designed to operate under conditions that require the trailing edge to operate with greater pressure loading, about twice that of modern commercial engines. High internal pressures required to cool turbine airfoils cause ballooning stresses at the trailing edge and high pressure loading across the turbine vane causes spanwise and chordwise bending loads for which a nearly solid trailing edge provides adequate structural strength. A problem is having enough structural integrity at the trailing edge while still being able to deliver cooling flow out the trailing edge as well as bleed flow to maintain film cooling. Low strain to failure ratio materials that are brittle such as ceramics offer very good thermal characteristics. SiC—SiC, a ceramic matrix composite (CMC) material, is being considered for use for turbine vanes because of its high temperature capability. The SiC—SiC CMC is a ceramic and has a very low failure strain level and consequently, a relatively low design stress when compared to typical Ni based super alloys.
An improved turbine airfoil and nozzle vane configuration is required that provides sufficient strength and cooling characteristics to meet the stress and temperature requirements in a high temperature and pressure environment and use ceramic materials that are characterized by a low thermal gradient capability.
SUMMARY OF THE INVENTION
A gas turbine engine hollow turbine airfoil suitable for use in a turbine vane includes an outer wall surrounding a hollow interior. The airfoil has pressure and suction sides extending aftwardly in a chordwise direction from a leading edge to a trailing edge. The outer wall extends radially outwardly in a spanwise direction from an airfoil base to an airfoil tip and widthwise spaced apart pressure and suction side portions extending chordwise between the leading edge and a trailing edge block. The trailing edge block extends aftwardly from the hollow interior and terminates at the trailing edge, a plurality of trailing edge cooling air ducts extend chordwise through the trailing edge block aftwardly from the hollow interior, and a plurality of trailing edge film cooling holes extend from the ducts through the trailing edge block. The cooling air ducts are preferably centered along or near a neutral axis of chordwise and spanwise bending of the trailing edge block. The invention is particularly suitable for turbine airfoils and vanes made with a ceramic matrix composite material such as one having a SiC matrix and, more particularly, with an SiC—SiC ceramic matrix composite. The trailing edge cooling air ducts converge aftwardly from the hollow interior. In a more particular embodiment, the trailing edge cooling air ducts converge in width and have a substantially constant spanwise height.
The invention includes, but is not limited to, three embodiments with means for terminating the trailing edge cooling air ducts. In the first embodiment, the trailing edge cooling air ducts extend aftwardly from the hollow interior completely through the trailing edge block. In the second embodiment, the trailing edge cooling air ducts terminate within the trailing edge block forward of the trailing edge. In the third embodiment, throttling holes extend from the trailing edge cooling air ducts, which terminate within the block, aftwardly completely through the trailing edge block. The trailing edge cooling air ducts have substantially constant spanwise heights and converging cross-sectional widths perpendicular to a span of the airfoil that converge in an aftwardly chordwise direction from the hollow interior.
In one more particular embodiment, the trailing edge block has a block spanwise bow wherein the trailing edge block is bowed outwardly in a normal direction to the pressure side of the airfoil. The block spanwise bow is preferably graduated in the chordwise direction to minimize bending of the airfoil trailing edge block near the trailing edge. The trailing edge is preferably bowed outwardly in a generally aftwardly chordwise direction.
The airfoil of the present invention is particularly suitable in a vane for a gas turbine engine such as the first stage of a high pressure turbine. The vane includes the hollow airfoil disposed between radially inner and outer segmented platforms that are
Darkins, Jr. Toby G.
Doughty Roger L.
Noe Mark E.
Shelton Monty L.
General Electric Company
Hess Andrew C.
Look Edward K.
Nguyen Ninh
Young Rodney M.
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