Specialized metallurgical processes – compositions for use therei – Compositions – Consolidated metal powder compositions
Reexamination Certificate
2000-09-15
2003-04-22
Jenkins, Daniel J. (Department: 1742)
Specialized metallurgical processes, compositions for use therei
Compositions
Consolidated metal powder compositions
C419S028000, C419S029000, C419S049000, C419S048000
Reexamination Certificate
active
06551372
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to powder metallurgy superalloy gas turbine engine components. More particularly, one embodiment of the present invention defines a wrought powder metallurgy nickel base superalloy gas turbine engine disk having enhanced fatigue crack growth resistance and a superior balance of mechanical properties. Although the present invention was developed for gas turbine engine components, certain applications may be in other fields.
The performance of a gas turbine engine is generally limited by the high temperature performance of the gas turbine engine's compressor and turbine disks, blades and vanes. In a typical gas turbine engine fuel and air are mixed and burned, and the hot combustion gas flow is directed against the vanes, which turn the gas flow against the turbine blades. The blades are mounted on the turbine disk, and the rotation of the turbine disk generates power which can be used to draw more air into the engine and drive a propulsion device such as a fan or propeller. The gas turbine engine disks, blades and vanes must therefore operate in an extremely hostile environment, of high temperature, high loading, fatigue, oxidation and corrosion. Gas turbine engine designers have focused much effort to improving the performance of materials that are used to fabricate the gas turbine engine's disks, blades and vanes.
For more than thirty years there has been continuing development on materials to enable engine components, such as compressor and turbine disks, to be operated under more rigorous conditions. A nickel base superalloy known as Waspaloy was introduced in 1967, and is still used in many applications today despite it's limitations of strength and maximum temperature of use. A cast/wrought-nickel base superalloy UDIMET 720 (UDIMET is a registered trademark of Special Metals Corporation) was introduced by Special Metals Corporation for selected components such as turbine blades used in industrial gas turbine engines. However, early applications of cast/wrought UDIMET 720 to aircraft gas turbine engine disk applications were hampered because the compositions used for disk forgings were susceptible to chemical segregation that can lead to low yields and a wide variability in grain size and heat treatment response. Further, problems were also encountered related to the formation of boride and carbide stringers that can act as nucleation sites that lead to early fatigue cracking and premature component failure.
In the 1980's several changes were made to the processing methods and the chemistry used for a disk component formed from cast/wrought UDIMET 720 to address the prior limitations. As an example, melt practices were changed from a double melt (vacuum induction melt plus vacuum arc remelt) to a triple melt (vacuum induction melt plus electro-slag remelt plus vacuum arc remelt) to minimize contamination and improve structure. Elements that can lead to the formation of stringers were analyzed and adjusted. More specifically, the carbon and boron levels were reduced from the levels in the earlier material. Further, the chromium levels were also reduced to prevent deleterious sigma phase formation.
The continued advancement in gas turbine engine designs requires the freedom to utilize significantly larger disks having enhanced micro-structural control and quality levels. Further, many modern design parameters require a disk material having defect tolerance, while maintaining resistance to burst yielding and fatigue crack initiation. Defect tolerance generally means that disks must have the capability to operate with either manufacturing defects that might escape non-destructive inspection during processing or post manufacturing defects that might arise from handling or service induced distress.
Heretofore, there has been a need for a high strength and defect tolerant disk for a gas turbine engine. The present invention satisfies this and other needs in a novel and unobvious way.
SUMMARY OF THE INVENTION
One form of the present invention contemplates a wrought powder metallurgy nickel based superalloy gas turbine engine disk.
Another form of the present invention contemplates a process of making a wrought powder metallurgy nickel based superalloy gas turbine engine disk. The disk is substantially free of chemical segregation and has grain sizes that promote a unique balance of fatigue crack growth resistance, low cycle fatigue capability, creep rupture strength and tensile properties.
Another form of the present invention contemplates a dual microstructure wrought powder metallurgy disk having a coarse grained rim and a fine grained bore.
Yet another form of the present invention contemplates a gas turbine engine disk, comprising: a main body member formed of a gamma prime strengthened wrought powder metallurgy composition consisting essentially of, in weight percent, 0.015%-0.035% carbon, 15.5%-16.5% chromium, 14%-15.5% cobalt, 2.75%-3.25% molybdenum, 4.75%-5.25% titanium, 2.25%-2.75% aluminum, 1%-1.5% tungsten, 0.030%-0.090% zirconium, 0.020%-0.050% boron, up to 0.90% hafnium, and the balance nickel plus incidental impurities; and the main body member has a substantially segregation free homogenous microstructure having an average grain size within a range of ASTM
5
(25 microns) to ASTM
14
(3 microns).
One aspect of the present invention contemplates a process of preparing a nickel base powder metal superalloy gas turbine engine disk, comprising: furnishing a composition consisting essentially of, in weight percent, 0.015%-0.035% carbon, 15.5%-16.5% chromium, 14%-15.5% cobalt, 2.75%-3.25% molybdenum, 4.75%-5.25% titanium, 2.25%-2.75% aluminum, 1%-1.5% tungsten, 0.030%-0.090% zirconium, 0.020%-0.050% boron, up to 0.90% hafnium, and the balance nickel plus incidental impurities; consolidating the composition to produce a preform member; thermomechanically working the preform to produce a wrought member; and, heat treating the wrought member.
One object of the present invention is to provide a unique gas turbine engine disk.
Related objects and advantages of the present invention will be apparent from the following description.
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The Homogenisation of Single-Crystal Superalloys.
Microstructural Analysis.
An Oxidation-Resistant Coating Alloy For Gamma Titanium Aluminides.
Creep Life Extension of a Single Crystal Superalloy by Re-Heat-Treatment.
Ewing Bruce A.
Green Kenneth A.
Helmink Randolph C.
Jain Sushil K.
James Allister
Jenkins Daniel J.
Rolls-Royce Corporation
Woodard, Emhardt, Naughton Moriarty & McNett LLP
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