Stock material or miscellaneous articles – Composite – Of metal
Reexamination Certificate
1999-09-10
2001-09-25
Copenheaver, Blaine (Department: 1775)
Stock material or miscellaneous articles
Composite
Of metal
C428S471000, C428S472000, C428S472200, C428S937000, C416S24100B
Reexamination Certificate
active
06294260
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to the field of thermal barrier coatings, and more particularly to a thermal barrier coating for a very high temperature application such as a combustion turbine engine. In particular, this invention relates to the field of multiphase ceramic thermal barrier coatings for high temperature application for coating superalloy components of a combustion turbine.
2. Background Information
The demand for continued improvement in the efficiency of combustion turbine and combined cycle power plants has driven the designers of these systems to specify increasingly higher firing temperatures in the combustion portions of these systems. Although nickel and cobalt based “superalloy” materials are now used for components in the hot gas flow path, such as combustor transition pieces and turbine rotating and stationary blades, even these superalloy materials are not capable of surviving long term operation at temperatures sometimes exceeding 1,200° C.
Examples of cobalt or nickel based superalloys are, for example, Cr.Al.Co.Ta.Mo.W which has been used for making SC turbine blades and vanes for gas turbines, as taught, for example, in U.S. Pat. No. 5,716,720 (Murphy).
These turbine components are generally protected by a basecoat of MCrA
1
Y, where M is selected from the group of Fe, Co, Ni, and their mixtures, as taught for example, by U.S. Pat. Nos. 5,763,107 and 5,846,605 (both Rickerby et al.) and by U.S. Pat. Nos. 4,916,022; 5,238,752; 5,562,998; and 5,683,825 (Solfest et al.; Duderstadt et al.; Strangman; and Bruce et al., respectively). These basecoats are usually covered by an aluminum oxide layer and a final thermal barrier coating (“TBC”). The standard thermal barrier coating, however, is made from yttria-stabilized zirconia, ceria-stabilized zirconia, scandia-stabilized zirconia or non-stabilized zirconia, as taught, for example, by U. S. Pat. No. 5,780,110 (Schaeffer et al.). A particularly useful state of the art TBC is 8 wt. % yttria stabilized zirconia (“8YSZ”).
Many of the ceramic thermal barrier layers are deposited as a columnar structure in the direction of the coating layer thickness, as taught in Murphy U.S. Pat. No. 5,716,720. This structure can be formed by plasma assisted physical vapor deposition or electron beam physical vapor deposition, as taught, for example, in Rickerby et al. U.S. Pat. No. 5,846,605 and in Murphy U.S. Pat. No. 5,716,720, respectively.
Another type of ceramic thermal barrier is deposited as a flat layer parallel to the surface of the superalloy structure, and applied, for example, by plasma spraying techniques, as taught in U.S. Pat. 5,824,423 (Maxwell et al.). The Rickerby et al., U.S. Pat. No. 5,846,605 characterizes these type of coatings as having a “brick wall” microstructure which has poor erosion resistance but a lower thermal conductivity than columnar structures.
Much of the development in this field of technology has been driven by the aircraft engine industry, where turbine engines are required to operate at high temperatures, and are also subjected to frequent temperature transients as the power level of the engine is varied. A combustion turbine engine installed in a land-based power generating plant is also subjected to high operating temperatures and temperature transients, but it may also be required to operate at full power and at its highest temperatures for very long periods of time, such as for days or even weeks at a time. Prior art insulating systems are susceptible to degradation under such conditions at the elevated temperatures demanded in the most modern combustion turbine systems.
Modern gas turbine engines can achieve higher efficiencies by increasing the turbine inlet temperatures. This subjects the TBCs to high temperatures. TBC materials that are phase stable at high temperatures upon long term exposure will be required. The current state-of-the-art air plasma sprayed (“APS”) 8YSZ coatings destabilize above approximately 1200° C. In addition, the long term high temperature exposure leads to potential sintering and loss of strain compliance, and possible premature TBC failure. 8YSZ coatings are also susceptible to corrosion upon exposure to contaminants in the fuel and erosion due to foreign object damage. Therefore, some of the key requirements for new TBC candidates for high temperature applications are high temperature phase stability, a reduced tendency to sinter, good corrosion and erosion resistance, all of them to be maintained upon long term exposure. These requirements are in addition to the primary needs of a TBC, such as, a low thermal conductivity with minimal coefficient of thermal expansion mismatch with the superalloy substrate. Further advances in gas turbine operating temperatures therefore require a plasma sprayed ceramic thermal barrier coating capable of surface temperature in excess of 1200° C.
SUMMARY OF THE INVENTION
Therefore, it is a main object of this invention to provide an improved thermal barrier coating for use on underlayers, such as alumina and MCrAlY, protecting turbine components, such as superalloy turbine blade assemblies that can operate over 1200° C.
These and other objects of the invention are accomplished by providing a turbine component comprising a metal alloy substrate and a flat thermal barrier coating on and parallel to the substrate surface, the thermal barrier coating comprising (a) a planar-grained ceramic oxide structural base material layer with open porosity and microcracks generally parallel to the surface of the substance and (b) a heat resistant ceramic oxide overlay covering the top surface of the planar-grained base material and infiltrating the open porosity and microcracks of the base, where the ceramic oxide overlay comprises the reaction product of a ceramic oxide overlay precursor material and the ceramic oxide base structure. The ceramic oxide overlay precursor material consists essentially of the composition C
z
O
w
. The ceramic oxide base structure consists essentially of the composition (A,B)
x
O
y
, where A and B are preferably selected from the group consisting of Al, Ca, Mg, Zr, Y, Sc and rare earth elements where rare earth elements are selected from La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Dy, Ho, Er, Tm and Yb, which will react as oxides with C
z
O
w
. C also is selected from the group consisting of Al, Ca, Mg, Zr, Y, Sc and rare earth elements where rare earth elements are selected from La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Dy, Ho, Er, Tm and Yb, that will react as oxides with (A,B)
x
O
y
. Preferably, C can be any A or B compound not used as A or B, but C is preferably Al, that is if A═Ca and B═Mg then C should not be Ca or Mg.
The preferred precursor material is a thin coating of alumina, Al
2
O
3
and the preferred base material is yttria stabilized zirconia where yttria, Y
2
O
3
, content can range from dopant amounts of 10 wt %-20 wt % of the total up to 60 wt % of the total with zirconia ZrO
2
. The base structure is one of a series of microfissures, microcracks, and splat boundaries all generally parallel to the super alloy top surface and disposed throughout the coating. The reaction product can be prompted upon heating to about 1200° C. to 1500° C. and has the composition, in this preferred case, of a material comprising Y
3
Al
5
O
12
. Another preferred material is the use of rare earth oxide stabilized zirconia, and an oxide such as Al
2
O
3
to form a reaction product containing ReO and Al
2
O
3
, for example, Re
3
Al
5
O
12
.
The invention also resides in a method of making a turbine component having a coated, adherent flat thermal barrier coating on its surface comprising the steps of: (a) providing a nickel or cobalt based superalloy substrate having a flat top surface; (b) depositing a planar-grained ceramic oxide base thermal barrier layer comprising stabilized zirconia, where the thermal barrier layer comprises discrete flat sections with microcrack volumes between the sections, all generally parallel to the flat top surface of the superalloy sub
Copenheaver Blaine
Siemens Westinghouse Power Corporation
Young Bryant
LandOfFree
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