Composite gas turbine blade and method of manufacturing same

• US 4,869,645 A
• Filed: 03/11/1988
• Issued: 09/26/1989
• Est. Priority Date: 03/19/1987
• Status: Expired due to Fees

• Current Assignee: Brown Boveri & CIE AG (ABB Limited)
• Original Assignee: Brown Boveri & CIE AG (ABB Limited)
• Inventors: Verpoort, Clemens
• Primary Examiner(s): Garrett, Robert E.
• Assistant Examiner(s): Pitko, Joseph M.
• Application Number:
  US07/167,015
• Time in Patent Office: 564 Days
• Field of Search: 416/213 R, 416/241 R, 416/189, 29/156.8 H, 29/156.8 B, 29/156.8 R, 29/527.1, 29/527.5, 164/98, 164/103, 164/105, 415/212 R
• US Class Current: 416/241.00R
• CPC Class Codes: B22D 19/00 Casting in, on, or around objects which form part of the product B22D23/04 takes precedence; alumino-thermic welding B23K23/00; coating by casting molten material on the substrate C23C6/00
• CPC Class Codes: Y10T 29/49337 Composite blade
• CPC Class Codes: Y10T 29/49988 Metal casting

Abstract

A composite gas turbine blade consists of an airfoil (1) in an oxide-dispersion-hardened nickel-based superalloy, in the condition of longitudinally directed coarse columnar crystals, and a shroud plate (6) or a shroud and a root (7), the latter items in a non-dispersion-hardened nickel-based superalloy (cast alloy). The gas turbine blade is manufactured by casting in and casting round, using the non-dispersion-hardened superalloy mentioned, the tip end (2) and root end (3)--provided with depressions (4) and/or protrusions (5)--of the airfoil (1), after preheating the latter to a temperature of between 50° and 300° C. below the solidus temperature of the lowest melting phase of the airfoil material. The casting temperature for this should be a maximum of 100° C. above the liquidus temperature of the highest melting phase of this non-dispersion-hardened alloy. Any melting onto the airfoil (1) and any metallurgical connection is to be avoided. It is advantageous to provide a thermally insulating, mechanically damping intermediate layer (16) of an oxide of at least one of the elements Cr, Al, Si, Ti and Zr with a thickness of 5 to 200 μm between the airfoil (1), on the one hand, and the shroud plate (6) and the root (7), on the other.

Claims

1. A method for manufacturing a composite gas turbine blade including a root, an airfoil and a shroud, the airfoil including an oxide-dispersion-hardened nickel-based superalloy in the condition of longitudinally directed coarse columnar crystals, wherein the outside surface of both the tip end and the root end of the airfoil are provided with depressions and/or protrusions, comprising the steps of:

inserting the airfoil in a mold having the negative shape of the shroud and the root in such a way that the tip end and the root end protrude into the hollow space of the mold;

preheating the airfoil to a temperature that is between 50° and

300°

C. below the solidus temperature of the lowest melting phase of the airfoil material;

filling the hollow space of the mold with the melt of a non-dispersion-hardened nickel-based superalloy intended for the shroud and the root at a casting temperature which is at the most 100°

C. above the liquidus temperature of the highest melting phase of the non-dispersion-hardened nickel-based superalloy, in such a way that the tip end and the root end of the airfoil are completely cast around and cast in;

controlling the temperature of the melt, after the conclusion of the casting procedure and during solidification, and that of the airfoil so that any melting onto the airfoil and any metallurgical connection between the material of the airfoil and that of the shroud and the root is avoided; and

cooling the whole workpiece to room temperature.

2. A method as claimed in claim 1, wherein the airfoil is machined out of semi-finished product which has previously been subjected to a heat treatment to increase the ductility at right angles to the longitudinal direction of the columnar crystals, or wherein the airfoil is subjected to a corresponding heat treatment after its manufacture, which consists of a heat treatment at or immediately above the lowest possible heat treatment temperature for solution of the γ

'"'"'-phase in the γ

-matrix of the airfoil material, followed by slow cooling at a maximum cooling rate of 5°

C./ min.

3. A method as claimed in claim 1, wherein, before it is cast around and cast in, the airfoil is preheated to a temperature which at least reaches a value of 50°

C. below the lowest possible heat treatment temperature for solution of the γ

'"'"'-phase in the γ

-matrix of the airfoil material, and wherein the airfoil, after it is cast around and cast in, is cooled at a maximum cooling rate of 5°

C./min. at least down to a temperature of 600°

C., while the solidified melt forming the shroud and/or the root is cooled at an arbitrary cooling rate.

4. A method as claimed in claim 1, wherein the airfoil is provided with an intermediate layer of an oxide of at least one of the elements Cr, Al, Si, Ti and Zr of between 5 μ

m and 200 μ

m thickness at least at the tip end and at the root end before it is placed in the mold.

5. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the air-foil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 15.0 % by weight Al = 4.5 % by weight Ti = 2.5 % by weight Mo = 2.0 % by weight W = 4.0 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is pre-heated to a temperature of between 1140° and

1180°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 16.0 % by weight Co = 8.5 % by weight Mo = 1.75 % by weight W = 2.6 % by weight Ta = 1.75 % by weight Nb = 0.9 % by weight Al = 3.4 % by weight Ti = 3.4 % by weight Zr = 0.1 % by weight B = 0.01 % by weight C = 0.11 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1380°

C.

6. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the airfoil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 15.0 % by weight Al = 4.5 % by weight Ti = 2.5 % by weight Mo = 2.0 % by weight W = 4.0 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is preheated to a temperature of between 1160° and

1200°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 22.4 % by weight Co = 19.0 % by weight Ta = 1.4 % by weight Nb = 1.0 % by weight Al = 1.9 % by weight Ti = 3.7 % by weight Zr = 0.1 % by weight C = 0.15 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1400°

C.

7. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the air-foil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 20.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Zr = 0.19 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is preheated to a temperature of between 1120° and

1160°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 16.0 % by weight Co = 8.5 % by weight Mo = 1.75 % by weight W = 2.6 % by weight Ta = 1.75 % by weight Nb = 0.9 % by weight Al = 3.4 % by weight Ti = 3.4 % by weight Zr = 0.1 % by weight B = 0.01 % by weight C = 0.11 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1380°

C.

8. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the airfoil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 20.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Zr = 0.19 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is preheated to a temperature of between 1120° and

1160°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 22.4 % by weight Co = 19.0 % by weight W = 2.0 % by weight Ta = 1.4 % by weight Nb = 1.0 % by weight Al = 1.9 % by weight Ti = 3.7 % by weight Zr = 0.1 % by weight C = 0.15 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1400°

C.

9. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the airfoil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 17.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is preheated to a temperature of between 1130° and

1170°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 16.0 % by weight Co = 8.5 % by weight Mo = 1.75 % by weight W = 2.6 % by weight Ta = 1.75 % by weight Nb = 0.9 % by weight Al = 3.4 % by weight Ti = 3.4 % by weight Zr = 0.1 % by weight B = 0.01 % by weight C = 0.11 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1380°

C.

10. A method as claimed in claim 1, wherein the oxide-dispersion-hardened nickel-based superalloy of the airfoil has the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 17.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest ______________________________________

and wherein the airfoil is preheated to a temperature of between 1130° and

1170°

C., wherein furthermore the nickel-based superalloy of the root and the shroud has the following composition;

space="preserve" listing-type="tabular">______________________________________ Cr = 22.4 % by weight Co = 19.0 % by weight W = 2.0 % by weight Ta = 1.4 % by weight Nb = 1.0 % by weight Al = 1.9 % by weight Ti = 3.7 % by weight Zr = 0.1 % by weight C = 0.15 % by weight Ni = rest ______________________________________

and wherein the maximum casting temperature of the melt of the abovementioned composition is 1400°

C.

11. A method as claimed in claim 1, wherein the complete workpiece, after it has been cooled to room temperature, is again heated to a temperature of between 1050°

and 1200°

C. and at least the root and the shroud are subjected to postcompression by hot isostatic pressing, in such a way that the workpiece is first heated to a temperature which is at least 100°

C. and a maximum of 150°

C. lower than the recrystallization temperature of the material of both the airfoil and the shroud and the root and is, after this, placed under a pressure of between 1000 and 3000 bar at this temperature for between 2 and 24 hours and is then cooled at a maximum rate of 5°

C./min at least down to a temperature of 600°

C.

12. A composite gas turbine blade, comprising:

a root;

an airfoil having a root end adjacent the root and a tip end at the opposite end of the airfoil, said airfoil having depressions and/or protrusions on the root end and the tip end; and

a shroud;

said airfoil comprising an oxide-dispersion-hardened nickel-based superalloy having longitudinally directed coarse columnar crystals;

the root and the shroud comprising a non-dispersion-hardened nickel-based cast superalloy;

the root and the shroud being connected purely mechanically to said airfoil by virtue of said root and said shroud including portions cast into the depressions and/or the protrusions on the root end and the tip end of the outside surface of the airfoil, while maintaining a metallic discontinuity between the airfoil and the shroud and root, and without any metallurgical connection therebetween.

13. A gas turbine blade as claimed in claim 12, wherein the metallic discontinuity between the airfoil and the shroud and/or between the airfoil and the root includes of a maximum width of 5 μ

m formed partially from a natural oxide layer and partially from hollow spaces.

14. A gas turbine blade as claimed in claim 12, wherein an intermediate layer of an oxide of at least one of the elements Cr, Al, Si, Ti and Zr of a thickness of between 5 μ

m and 200 μ

m is present on the surface of the airfoil in the metallic discontinuity between the airfoil and the root and/or between the airfoil and the shroud.

15. A gas turbine blade as claimed in claim 14, wherein the intermediate layer is designed as a firmly adhering layer of at least of 100 μ

m thickness on the surface of the airfoil, acting as thermal insulation in service, and consisting mainly of Al₂ O₃ or of ZrO₂ stabilized with Y₂ O₃.

16. A gas turbine blade as claimed in claim 12, wherein the airfoil is comprised of an oxide-dispersion-hardened, but not precipitation-hardened, nickel-based superalloy with increased ductility at right angles to the longitudinal direction of the columnar crystals.

17. A gas turbine blade or vane as claimed in claim 12, wherein the airfoil is comprised of an alloy with the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 15.0 % by weight Al = 4.5 % by weight Ti = 2.5 % by weight Mo = 2.0 % by weight W = 4.0 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest. ______________________________________

18. A gas turbine blade as claimed in claim 12, wherein the airfoil is comprised of an alloy with the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 20.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Zr = 0.19 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest. ______________________________________

19. A gas turbine blade as claimed in claim 12, wherein the airfoil is comprised of an alloy with the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 17.0 % by weight Al = 6.0 % by weight Mo = 2.0 % by weight W = 3.5 % by weight Ta = 2.0 % by weight Zr = 0.15 % by weight B = 0.01 % by weight C = 0.05 % by weight Y.sub.2 O.sub.3 = 1.1 % by weight Ni = rest. ______________________________________

20. A gas turbine blade as claimed in claim 12, wherein the root (7) and the shroud are comprised of an alloy with the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 16.0 % by weight Co = 8.5 % by weight Mo = 1.75 % by weight W = 2.6 % by weight Ta = 1.75 % by weight Nb = 0.9 % by weight Al = 3.4 % by weight Ti = 3.4 % by weight Zr = 0.1 % by weight B = 0.01 % by weight C = 0.11 % by weight Ni = rest. ______________________________________

21. A gas turbine blade as claimed in claim 12 wherein the root and the shroud are comprised of an alloy with the following composition:

space="preserve" listing-type="tabular">______________________________________ Cr = 22.4 % by weight Co = 19.0 % by weight W = 2.0 % by weight Ta = 1.4 % by weight Nb = 1.0 % by weight Al = 1.9 % by weight Ti = 3.7 % by weight Zr = 0.1 % by weight C = 0.15 % by weight Ni = rest. ______________________________________