High-temperature alloy and articles made therefrom

Details High-temperature alloy and articles made therefrom High-temperature alloy and articles made therefrom

2001-10-24

2003-06-24

King, Roy (Department: 1742)

Metal treatment

Stock

Noble metal base

C420S456000, C420S580000, C420S462000

Reexamination Certificate

active

06582534

ABSTRACT:

BACKGROUND OF INVENTION
The present invention relates to materials designed to withstand high temperatures. More particularly, this invention relates to heat-resistant alloys for high-temperature applications, such as, for instance, gas turbine engine components of aircraft engines and power generation equipment.
There is a continuing demand in many industries, notably in the aircraft engine and power generation industries where efficiency directly relates to operating temperature, for alloys that exhibit sufficient levels of strength and oxidation resistance at increasingly higher temperatures. Gas turbine airfoils on such components as vanes and blades are usually made of materials known in the art as “superalloys.” The term “superalloy” is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more additional elements to enhance high temperature performance, including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The term “based” as used in, for example, “nickel-based superalloy” is widely accepted in the art to mean that the element upon which the alloy is “based” is the single largest elemental component by weight in the alloy composition. Generally recognized to have service capabilities limited to a temperature of about 1100° C., conventional superalloys used in gas turbine airfoils often operate at the upper limits of their practical service temperature range. In typical jet engines, for example, bulk average airfoil temperatures range from about 900° C. to about 1000° C., while airfoil leading and trailing edge and tip temperatures can reach about 1150° C. or more. At such elevated temperatures, the oxidation process consumes conventional superalloy parts, forming a weak, brittle metal oxide that is prone to chip or spall away from the part. Maximum temperatures are expected in future applications to be over about 1300° C., at which point many conventional superalloys begin to melt. Clearly, new materials must be developed if the efficiency enhancements available at higher operating temperatures are to be exploited.
The so-called “refractory superalloys,” as described in Koizumi et al., U.S. Pat. No. 6,071,470, represent a class of alloys designed to operate at higher temperatures than those of conventional superalloys. According to Koizumi et al., refractory superalloys consist essentially of a primary constituent selected from the group consisting of iridium (Ir), rhodium (Rh), and a mixture thereof, and one or more additive elements selected from the group consisting of niobium (Nb), tantalum (Ta), hafnium (Hf), zirconium (Zr), uranium (U), vanadium (V), titanium (Ti), and aluminum (Al). The refractory superalloys have a microstructure containing an FCC (face-centered cubic)-type crystalline structure phase and an L12 type crystalline structure phase, and the one or more additive elements are present in a total amount within the range of from 2 atom % to 22 atom %.
SUMMARY OF INVENTION
Although the refractory superalloys have shown potential to become replacements for conventional superalloys in present and future gas turbine engine designs, it has been shown that many alloys of this class do not meet all of the desired performance criteria for high-temperature applications. Therefore, the need persists for alloys with improved high-temperature properties.
The present invention provides several embodiments that address this need. One embodiment provides an alloy comprising rhodium, aluminum, and chromium, wherein a microstructure of the alloy comprises a face-centered-cubic phase and a B2-structured phase and is essentially free of an L1
2
-structured phase at temperatures greater than about 1000° C.
A second embodiment provides a gas turbine engine component comprising an alloy comprising rhodium, aluminum, and chromium; wherein a microstructure of the alloy of the engine component comprises a face-centered-cubic phase and a B2-structured phase and is essentially free of an L1
2
-structured phase at temperatures greater than about 1000° C.


REFERENCES:
patent: 3640705 (1972-02-01), Selman et al.
patent: 3709667 (1973-01-01), Selman et al.
patent: 3904404 (1975-09-01), Suits
patent: 3918139 (1975-11-01), Felten
patent: 4737205 (1988-04-01), Selman et al.
patent: 6071470 (2000-06-01), Koizumi et al.
patent: 2269393 (1994-02-01), None

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