US20080063533A1

US20080063533A1 - Turbine blade for a gas turbine engine

Info

Publication number: US20080063533A1
Application number: US11/798,589
Authority: US
Inventors: Christopher Boyle; Philip Jennings; Julian Mason-Flucke; Michael Wardle; Paul Withey
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2006-06-07
Filing date: 2007-05-15
Publication date: 2008-03-13
Also published as: GB0611385D0; GB2440127A; GB2440127B

Abstract

A turbine blade is manufactured as a single-crystal casting from a metal alloy, without a solution heat treatment step.

Description

This invention relates to a turbine blade for a gas turbine engine, the use of such a turbine blade in a gas turbine engine, and to a method of manufacturing a turbine blade.
Turbine blades in gas turbine engines operate at the limits of their material properties. They may be exposed to temperatures in excess of 2000° C. and are subjected to severe stress, both from gas flow past the blades and from centrifugal forces.
It is known to form turbine blades as single-crystal castings from specialized metal alloys, thereby providing high strength required to avoid failure under operational loads. The alloys used tend to be nickel/aluminum alloys, with various other components selected to enhance the properties of the alloy. Typical alloys used for manufacturing turbine blades for gas turbine engines are disclosed, by way of example, in U.S. Pat. No. 4,222,794, U.S. Pat. No. 4,582,548, U.S. Pat. No. 4,643,782 and U.S. Pat. No. 5,540,790. Some of the alloys disclosed in these patent specifications are commercially available, for example under the designations CMSX-3, CMSX-4 (available from Cannon Muskegon Corporation of Muskegon, Mich., USA) and PW 1484.
It has been considered essential for turbine blades manufactured as single-crystal castings to be heat treated before use to relieve residual stresses, thereby optimizing the mechanical properties of the alloy. Early single crystal alloys were not heat treatable because the temperature at which the necessary strengthening changes occurred was above the melting point of the material. Hence such alloys were not used to produce turbine blades because their poor microstructure inherently meant their mechanical integrity was insufficient for such applications. Improved single crystal materials are now available which enable solution heat treatment, thus delivering optimal mechanical properties.
Residual stresses in single-crystal castings arise as a result of differential contraction of different parts of the casting as it cools. Solution heat treatment relieves these stresses but a disadvantage is that re-crystallization of the material may occur, which will weaken the structure.
Re-crystallization can be minimized by appropriate design of the turbine blade. In particular, it appears that re-crystallization is inhibited if internal webs within the turbine blade extend perpendicular to, or close to perpendicular to, the external walls of the turbine blade, if the webs are relatively thick, and if the spacing between adjacent cooling holes is relatively large. However, a turbine blade designed within these constraints may not have optimum performance. For example, thicker webs increase the weight of the blade, while the angles of the webs relative to the outer walls of the blade and the spacing of cooling holes can affect the cooling efficiency of the blade, in terms of the quantity of cooling air required to maintain a desired temperature.
Nevertheless, it has until now been believed that solution heat treatment of single-crystal turbine blades provides the only route by which an acceptable operational life can be achieved, and solution heat treatment has therefore been regarded as an essential step in the manufacture of such turbine blades.
According to one aspect of the present invention, there is provided a finished turbine blade for a gas turbine engine, comprising a single-crystal casting of a metal alloy having a solvus temperature which is less than its incipient melting point, characterised in that the turbine blade has not been subjected to solution heat treatment after casting.
Another aspect of the present invention provides the use in a gas turbine engine of a turbine blade comprising a single-crystal casting of a metal alloy having a solvus temperature which is less than its incipient melting point, characterised in that the turbine blade has not been subjected to solution heat treatment after casting.
A further aspect of the present invention provides a gas turbine engine, characterised in that the engine includes a turbine blade comprising a single-crystal casting of metal alloy having a solvus temperature which is less than its incipient melting point, characterised in that the turbine blade has not been subjected to solution heat treatment after casting.
Thus, while it was previously considered to be essential to heat treat a turbine blade formed as a single crystal casting in order to achieve desired physical properties to ensure the required operational life of the component, the present invention arises from the realization that the additional design flexibility which arises if solution heat treatment is avoided can compensate for any resulting deficiencies in the physical properties of the material which would otherwise lead to a reduced operational life.
The metal alloy from which the turbine blade is made is preferably a nickel-based alloy, such as a nickel aluminum alloy, including SRR99, CMSX-3, CMSX-4 and PWA 1484. The alloy may contain other alloying components, such as hafnium, rhenium, titanium, chromium or gallium. The solvus temperature of the metal alloy should be less than the incipient melting point of the alloy.
The turbine blade may be provided with internal cooling passages in the form of cavities extending through the blade. By virtue of the design freedom which results from the omission of any solution heat treatment step, internal walls within the turbine blade, which separate adjacent cavities from one another, may be thinner than in a turbine blade which is subjected to a solution heat treatment step. For example, the thickness ratio between internal walls which separate adjacent cavities from one another and external walls which separate the cavities from the exterior of the turbine blade, may be less than 1.5:1 and preferably less than 1.25:1. Also, the angle at which an internal wall meets the external wall may be smaller than in a turbine blade which has been subjected to a solution heat treatment step. For example, an internal wall may meet an external wall at an angle less than 60° and possibly less than 50° or 45°.
The internal walls may be provided with through holes, for example for the passage of cooling air, and these holes may be more closely spaced than in a turbine blade that has been subjected to solution heat treatment. For example, the centreline spacing between adjacent holes may be less than 6 times the hole diameter, or even less than 5 times or 4 times the hole diameter.
Another aspect of the present invention provides a method of manufacturing a turbine blade for use in a gas turbine engine, the method comprising casting the turbine blade as a single crystal from a metal alloy, without a subsequent solution heat treatment step, said metal alloy having a solvus temperature less than its incipient melting point.
For a better understanding of the present invention, and to show more clearly how it may be carried into effect, reference will now be made, by way of example, to the accompanying drawings, in which:
FIG. 1 (PRIOR ART) is a sectional view of a turbine blade for a gas turbine engine, the manufacture of which includes a solution heat treatment step;
FIG. 2 corresponds to FIG. 1, but shows a turbine blade configuration suitable for a turbine blade which is not subjected to solution heat treatment after casting;
FIG. 3 (PRIOR ART) is a sectional view on the line A-A in FIG. 1; and
FIG. 4 is a sectional view taken on the line B-B in FIG. 2.
The turbine blade shown in FIG. 1 is hollow, comprising a plurality of cavities 2 through which, in use, cooling air may flow in order to cool the turbine blade. Thus, the cavities 2 may be connected at one or both ends to cooling air supply chambers or ducts situated externally of the turbine blade itself.
The cavities 2 are separated from each other by internal walls 4. Cooling holes 6 are provided which extend through the walls 4 to allow the flow of cooling air between adjacent cavities 2. The cavities 2 are also bounded by external walls 6 which separate the cavities 2 from the outside of the turbine blade.
The turbine blade shown in FIG. 1 is cast as a single crystal from a suitable single-crystal alloy, for example CMSX-3, CMSX-4 or PW1484. It has been considered to be essential for the turbine blade, after casting, to be subjected to a solution heat treatment process in order to relieve residual stresses in the cast turbine blade. Solution heat treatment processes have been required in order to enhance the strength of the cast turbine blade, and in particular to enhance the fatigue strength and creep strength. Consequently, alloys for use in the manufacture of single-crystal turbine blades have been formulated so as to have a solvus temperature (ie the temperature at which the necessary strengthening changes will occur) which is below the incipient melting point of the alloy. Consequently, the beneficial changes achieved by solution heat treatment occur before the cast turbine blade begins to melt.
However, the solution heat treatment process is known to cause recrystallization of the alloy, which weakens the structure in the regions at which recrystallization occurs. The configuration shown in FIG. 1 is designed to minimize such recrystallization. Thus, the internal walls 4 are relatively thick, the angles at which they meet the external walls 6 are relatively large, and the spacing between adjacent cooling holes 8 is also relatively large.
By way of example, in the configuration shown in FIG. 1, the internal walls 4 have a thickness T_iwhich is significantly larger than the thickness T_eof the external walls 6. For example, the ratio T_i/T_eis greater than 1:5:1. In the embodiment shown, it is approximately 3:1.
Also, the angles at which the internal walls 4 meet the external walls 6, as measured between the general central axis of the internal wall 4 in the section shown in FIG. 1 and the tangent to the external surface of the external wall 6 at the point of intersection with the central axis of the internal wall 4, as designated by way of example by α in FIG. 1, is preferably as close to 90° as possible. Although this cannot be achieved if the internal walls 4 are to be generally straight, owing to the curved and tapering nature of the turbine blade, it is preferable for the angle α to be no less than 60°.
As shown in FIG. 3, the cooling holes 8 of the turbine blade configuration of FIG. 1 are disposed in a single line at a centreline spacing S which is relatively large. Expressed as a multiple of the diameter D of each cooling hole, the spacing S is preferably at least 5 times the diameter D and, in the embodiment shown in FIG. 3, is approximately 7 times the diameter D.
The configuration shown in FIGS. 1 and 3, established with a view to a minimizing recrystallization of the material of the turbine blade during solution heat treatment, has disadvantages. In particular, the relatively thick internal walls 4 increase the volume of alloy in the turbine blade, and consequently also increase the weight of the turbine blade. Also, the relatively thick internal walls 4 reduce the sizes of the cavities 2, so reducing the flow passage size for cooling air. Cooling air flow is also restricted by the relative spacing S between adjacent cooling holes 8, since this restricts the number of cooling 8 that can be provided. The need for a relatively large angle α makes it impossible to angle the internal walls 4 relatively to the external walls 6 by narrow angles, which could allow the cooling holes 8 to be directed at the external walls 6, so as provide impingement cooling.
These disadvantages are overcome in the turbine blade configuration shown in FIGS. 2 and 4, in which the features of the turbine blade are designated by the same reference numbers as in FIGS. 1 and 3.
In the configuration shown in FIGS. 2 and 4, the internal walls 4 are significantly thinner than those of the configuration shown in FIG. 1. In the case of some of the walls 4, the thickness T_iis comparable to the thickness T_eof the external walls 6, so that the ratio T_i/T_eis close to 1. Generally, it is preferred for the ratio T_i/T_eto be no greater than 1.5:1, or more preferably, no greater than 1.25:1.
Furthermore, some of the internal walls 4 meet the external walls 6 at angles α significantly less than the corresponding angles α of the configuration shown in FIG. 1. For example, the angle α for at least some of the internal walls 4 in the configuration of FIG. 2 may be less than 60° or even less than 50°, and in some cases may be 45° or smaller.
As shown in FIG. 4, the spacing between adjacent cooling holes 8 is significantly smaller than that of the configuration of FIG. 3. Thus, the spacing S may be less than 6 times the diameter D of each cooling hole, or even less than 5 times the diameter D. In the embodiment shown in FIG. 4, the spacing S is only about 3 times the diameter D. This enables the cooling holes to be arranged in two rows, which would not be possible in the configuration shown in FIG. 3 if the spacing S is to be sufficiently large to avoid recrystallization of the alloy during solution heat treatment.
As a result of the greater design freedom applicable to the configuration shown in FIGS. 2 and 4, the reduced thickness T_iof the internal walls 4 results in a weight reduction and an increase in the flow capability of the cavities 2. The ability to use a relatively acute angle α between the internal walls 4 and the external walls 6 means that the cooling holes 8 can be oriented so that they can direct cooling air onto a nearby external wall 6, so providing impingement cooling. Also, the ability to decrease the spacing S between cooling holes, means that the number of holes 8 can be increased, which not only increases the flow of cooling air, but also increases the surface area available for the transfer of heat from the alloy of the turbine blade to the cooling air.
Thus, although the process of manufacturing the turbine blade shown in FIGS. 2 and 4 without solution heat treatment means that, in some respects, the strength of the turbine blade may not be fully optimized, the advantages arising from weight reduction, increased cooling air flow capability and effective orientation of the cooling holes 8 means that the loss of fatigue and creep strength are outweighed. As a result, contrary to expectations, a turbine blade manufactured to the configuration shown in FIGS. 2 and 4 can have an operational life similar to, or exceeding, that of a turbine blade having the configuration of FIGS. 1 and 3, despite the fact that the turbine blade of FIGS. 2 and 4 is manufactured without a solution heat treatment step as used in the turbine blade of FIGS. 1 and 3.
The omission of the solution heat treatment step has the additional advantage that the overall manufacturing time and cost is reduced. Furthermore, the rate of rejection of turbine blades manufactured as single-crystal cast alloy components can be reduced, since many turbine blades are rejected largely as a result of an unacceptable degree of recrystallization during the solution heat treatment process.

Claims

1-6. (canceled)

7. A gas turbine engine, characterised in that the engine includes a turbine blade comprising a single-crystal casting of metal alloy having a solvus temperature which is less than its incipient melting point, characterised in that the turbine blade has not been subjected to solution heat treatment after casting.

8. A gas turbine engine as claimed in claim 7 wherein the metal alloy is a Ni ALX alloy in which X is one or more of hafnium, rhenium, titanium, chromium or gallium.

9. A gas turbine engine as claimed in claim 7 wherein the turbine blade is provided with cavities which are separated from adjacent cavities by internal walls and which are separated from the exterior of the turbine blade by external walls.

10. A gas turbine engine as claimed in claim 9 wherein the thickness ratio Ti/Te of the internal and external walls is not greater than 1.5:1.

11. A gas turbine engine as claimed in claim 9 wherein at least one of the internal walls is provided with through holes, at least some of which are spaced apart by a spacing(s) which is not greater than 6 times the transverse dimension of the holes.

12. A gas turbine engine as claimed in claim 9 wherein at least one of the internal walls meets an adjacent external wall at an angle α which is not greater than 60°.

13. A finished turbine blade for a gas turbine engine, comprising a single-crystal casting of a metal alloy having a solvus temperature which is less than its incipient melting point, characterised in that the turbine blade has not been subjected to solution heat treatment after casting.

14. A turbine blade as claimed in claim 13 characterised in that the metal alloy is a NiAlX alloy in which X is one or more of hafnium, rhenium, titanium, chromium or gallium.

15. A turbine blade as claimed in claim 13 characterised in that the turbine blade is provided with cavities which are separated from adjacent cavities by internal walls and which are separated from the exterior of the turbine blade by external walls.

16. A turbine blade as claimed in claim 15 characterised in that the thickness of ratio Ti/Te of the internal and external walls is not greater than 1.5:1.

17. A turbine blade as claimed in claim 15 characterised in that at least one of the internal walls is provided with through holes, at least some of which are spaced apart by a spacing(s) which is not greater than 6 times the transverse dimension of the holes.

18. A turbine blade as claimed in claim 15 characterised in that at least one of the internal walls meets an adjacent external wall at an angle α which is not greater than 60°.

19. A method of manufacturing a turbine blade for use in a gas turbine engine, the method comprising casting the turbine blade as a single crystal from a metal alloy, without a subsequent solution heat treatment step, said metal alloy having a solvus temperature less than its incipient melting point.