Metal working – Method of mechanical manufacture – Impeller making
Reexamination Certificate
2002-03-21
2004-09-14
Cuda-Rosenbaum, Irene (Department: 3726)
Metal working
Method of mechanical manufacture
Impeller making
C029S889100
Reexamination Certificate
active
06789315
ABSTRACT:
This invention relates to gas turbine engines and, more particularly, to the stationary nozzle vane structure of the gas turbine engine.
BACKGROUND OF THE INVENTION
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the resulting hot combustion gases are passed through a gas turbine mounted on the same shaft. The flow of combustion gas turns the gas turbine by contacting an airfoil portion of the turbine blade, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. There may additionally be a bypass fan that forces air around the center core of the engine, driven by a shaft extending from the turbine section.
The combustion gas flowing from the combustor of the gas turbine engine enters a stationary (that is, not rotating) nozzle structure having a plurality of stationary nozzle vanes that extend radially around the circumference of the combustion gas flow path of the generally cylindrical gas turbine engine. The gas turbine nozzle vanes act as a nozzle to accelerate and redirect the combustion gas flow slightly so that it enters the turbine at the proper velocity and angle. The redirected combustion gas flow impinges upon turbine blades extending radially from a turbine wheel that is rotatable about the turbine shaft, causing the turbine wheel to turn and thence to drive the compressor.
One of the important operating parameters of the gas turbine engine is the area of the stationary nozzle vane structure through which the combustion gas flow passes, termed the throat area. The total throat area, a design parameter of the gas turbine engine, determines the turbine flow function, which in turn sets the compressor pressure ratio and operating line, and is an important parameter in determining the compressor stall margin. Desirably, the pairwise throat area between each pair of gas turbine nozzle vanes is maintained within a selected pairwise throat area target range, and also the total throat area for all of the pairs of gas turbine nozzle vanes is maintained within a selected total throat area target range.
The nozzle vanes are manufactured to close dimensional tolerances, but the variations within the tolerances and the assembly tolerances may be sufficiently great that, upon assembly of the gas turbine engine, adjacent pairs of gas turbine nozzle vanes do not meet the pairwise throat area target range, or that the total set of nozzle vanes do not meet the total throat area target range. The problem of deviation from the throat area target ranges becomes even more significant after the gas turbine engine has been in service and the dimensions of the nozzle vanes have changed by nonuniform amounts from nozzle vane to nozzle vane due to erosion, burning of the vane material, and other effects. In that case, the overhauling of the gas turbine engine includes replacing or repairing the individual nozzle vanes, and then selecting the proper pairings of nozzle vanes to meet the pairwise and total throat area target ranges.
Even though the individual nozzle vanes are expensive to produce, it is often necessary during the overhaul of the gas turbine engine to scrap many of the used nozzle vanes and replace them with new-make nozzle vanes in order to meet the throat area target ranges. Repair techniques have been proposed for the nozzle vanes, but the available techniques are not fully acceptable because they may result in undesirable alterations to the performance of the nozzle vanes. There is accordingly a need for an improved approach to the repair of nozzle vanes so that their performance is acceptable and also so that the throat area target ranges are met. The present invention fulfills this need, and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
The present invention provides a technique for establishing the throat area of the gas turbine nozzle, and a repair procedure for the individual nozzle vanes. The present approach allows both the pairwise throat area target range and the total throat area target range to be met. The repair procedure is utilized to provide the individual nozzle vanes with the proper dimensions to meet the pairwise throat area target range, and then the pairs of nozzle vanes are combined to meet the total throat area target range. It may also be used to alter the dimensions of the nozzle vanes for other reasons. The repair procedure substantially maintains the airfoil shape of the nozzle vanes.
A method for establishing a final throat area of a gas turbine nozzle comprises the steps of providing a final pairwise throat area target range, providing at least two gas turbine nozzle vanes, and thereafter determining a pairwise initial throat area between each pair of gas turbine nozzle vanes. Thereafter, for each pair of gas turbine nozzle vanes whose pairwise initial throat area is not within the final pairwise throat area target range, a trailing edge of one of the gas turbine nozzle vanes is extended responsive to the step of determining the initial pairwise throat area, so that the final pairwise throat area is within the final pairwise throat area target range.
The method is typically extended to a nozzle plurality of pairs of gas turbine nozzle vanes together comprising a gas turbine nozzle circumferential structure. A final total throat area target range is provided, and the nozzle plurality of pairs of gas turbine nozzle vanes is selected so that a sum of their final pairwise throat areas is within the final total throat area target range.
The step of extending the trailing edge preferably includes the steps of providing one of the gas turbine nozzle vanes having a rounded trailing edge extending lengthwise between a root and a tip of the gas turbine nozzle vane, affixing an extension wire to the trailing edge extending lengthwise along the trailing edge, applying a braze material overlying the extension wire and the trailing edge, and brazing the braze material to the extension wire and to the trailing edge. The brazing is desirably accomplished by heating the braze material to a brazing temperature to melt at least a portion of the braze material and, upon cooling, to bond the extension wire and the braze material to the trailing edge. The extending of the trailing edge may be accomplished in conjunction with the establishing of the final throat areas as discussed above, or independently of any considerations of the throat area and for other reasons.
In a typical case, the gas turbine nozzle vane is made of a nozzle-vane nickel-base superalloy, and the extension wire is made of an extension-wire nickel-base superalloy. For example, the nozzle-vane nickel-base superalloy may be directionally solidified Rene
R
142 or Alloy X-40, and the extension wire nickel-base superalloy may be Rene
R
142 nickel-base superalloy.
In a preferred approach, the step of applying the braze material includes the step of applying a first layer of a high-melt filler alloy into a gap region between the extension wire and the trailing edge, and applying a second layer of a braze composition overlying the first layer. The first layer is normally a first-layer nickel-base superalloy having a first-layer-nickel-base-superalloy melting point greater than the brazing temperature, such as Rene
R
N4 or Rene
R
142 nickel-base superalloys. The second layer is a second-layer nickel-base superalloy braze material having a second-layer-nickel-base-superalloy melting point less than the brazing temperature, such as a mixture of 40 percent by weight Rene
R
142 superalloy and 60 percent by weight of Rene
R
80 nickel-base superalloy modified by the addition of silicon and possibly boron to reduce its melting point. The brazing temperature for this preferred braze material is typically from about 2190° F. to about 2240° F.
The present approach alters the nozzle vanes by extending their trailing edges. The shape of the airfoil of the nozzle vanes is
Budinger David Edwin
de Rezende Tathiana Carneiro
Marques Marcio Lins Tinoco
Messelling William G.
von Montfort Andreas Martin
Cuda-Rosenbaum Irene
Garmong Gregory O.
General Electric Company
Maria Carmen Santa
McNees Wallace & Nurick LLC
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