GB2440127A

GB2440127A - Non-solution treated gas turbine blades

Info

Publication number: GB2440127A
Application number: GB0611385A
Authority: GB
Inventors: Christopher James George Boyle; Philip Andrew Jennings; Julian Charles Mason-Flucke; Michael Edwin Wardle; Paul Anthony Withey
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2006-06-07
Filing date: 2006-06-07
Publication date: 2008-01-23
Anticipated expiration: 2026-06-07
Also published as: US20080063533A1; GB0611385D0; GB2440127B

Abstract

Use in a gas turbine engine of a turbine blade which comprises a single-crystal metal casting which has a solvus temperature which is less than its incipient melting point (i.e. it is heat-treatable), wherein the blade has not been subjected to a solution heat treating after casting. The blade has cavities 2 separated from each other by internal walls 4 and from the exterior by external walls 6, where the thickness ration of internal : external walls (Ti : Te) is not greater than 1.5:1. At least one of the internal walls 4 meets an adjacent external wall 6 at an angle a not greater than 60{. At least one of the internal walls 4 has through holes 8 separated by a spacing which is not greater than 6 times the traverse diameter of the holes.

Description

A TURBINE BLADE FOR A GAS TURBINE ENGINE

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 specialised metal alloys, thereby providing high strength required to avoid failure under operational loads.

The alloys used tend to be nickel/aluminium 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 US 4222794, US 4582548, US 4643782 and US 5540790. 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, Michigan, 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 optimising 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-crystallisation of the material may occur, which will weaken the structure.

Re-crystallisation can be minimised by appropriate design of the turbine blade. In particular, it appears that re-crystallisation 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.

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 realisation 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 aluminium 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 600 and possibly less than 50 or 450* 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 times or 4 times the hole diameter.

Another aspect of the present invention provides a method of manufacturing a turbine blade for 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: Figure 1 is a sectional view of a turbine blade for a gas turbine engine, the manufacture of which includes a solution heat treatment step; Figure 2 corresponds to Figure 1, but shows a turbine blade configuration suitable for a turbine blade which is not subjected to solution heat treatment after casting; Figure 3 is a sectional view on the line A-A in Figure 1; and Figure 4 is a sectional view taken on the line B-B in Figure 2.

The turbine blade shown in Figure 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 Figure 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 recrystallisation of the alloy, which weakens the structure in the regions at which recrystallisation occurs. The configuration shown in Figure 1 is designed to minimise such recrystallisation. 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 Figure 1, the internal walls 4 have a thickness T which is significantly larger than the thickness Te of the external walls 6.

For example, the ratio Ti/Te is greater than 1:5:1. In the embodiment shown, it is approximately 3:1.

Also, the angles at which the internal waIls 4 meet the external walls 6, as measured between the general central axis of the internal wall 4 in the section shown in Figure 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 a in Figure 1, is preferably as close to 900 as possible. Although this cannot be achieved if the internal waIls 4 are to be generally straight, owing to the curved and tapering nature of the turbine blade, it is preferable for the angle a to be no less than 60 .

As shown in Figure 3, the cooling holes 8 of the turbine blade configuration of Figure 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 Figure 3, is approximately 7 times the diameter D. The configuration shown in Figures 1 and 3, established with a view to a minimising recrystallisation 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 a makes it impossible to angle the internal waIls 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 Figures 2 and 4, in which the features of the turbine blade are designated by the same reference numbers as in Figures 1 and 3.

In the configuration shown in Figures 2 and 4, the internal walls 4 are significantly thinner than those of the configuration shown in Figure 1. In the case of some of the walls 4, the thickness T is comparable to the thickness Te of the external walls 6, SO that the ratio Ti/Te is close to 1. Generally, it is preferred for the ratio Ti/Te to 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 a significantly less than the corresponding angles a of the configuration shown in Figure 1. For example, the angle a for at least some of the internal walls 4 in the configuration of Figure 2 may be less than 600 or even less than 500, and in some cases may be 45 or smaller.

As shown in Figure 4, the spacing between adjacent cooling holes 8 is significantly smaller than that of the configuration of Figure 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 Figure 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 Figure 3 if the spacing S is to be sufficiently large to avoid recrystallisation of the alloy during solution heat treatment.

As a result of the greater design freedom applicable to the configuration shown in Figures 2 and 4, the reduced thickness T of 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 cx between the internal walls 4 and the external waIls 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 Figures 2 and 4 without solution heat treatment means that, in some respects, the strength of the turbine blade may not be fully optimised, 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 Figures 2 and 4 can have an operational life similar to, or exceeding, that of a turbine blade having the configuration of Figures 1 and 3, despite the fact that the turbine blade of Figures 2 and 4 is manufactured without a solution heat treatment step as used in the turbine blade of Figures 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 recrystallisation during the solution heat treatment process.

Claims

CLAIMS

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.

2 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.

3 A turbine blade as claimed in claim 1, or use of a turbine blade as claimed in claim 2, characterised in that the metal alloy is a NiAIX alloy in which X is one or more of hafnium, rhenium, titanium, chromium or gallium.

4 A turbine blade or use of a turbine blade in accordance with any one of the preceding claims, characterised in that the turbine blade is provided with cavities (2) which are separated from adjacent cavities (2) by internal walls (4) and which are separated from the exterior of the turbine blade by external walls (6).

A turbine blade or use of a turbine blade in accordance with claim 4, characterised in that the thickness ratio Ti/Te of the internal and external walls (4, 6) is not greaterthan 1.5:1.

6 A turbine blade or use of a turbine blade in accordance with claim 4 or 5, characterised in that at least one of the internal walls (4) is provided with through holes (8), at least some of which are spaced apart by a spacing(s) which is not greater than 6 times the transverse dimension (D) of the holes (8).

7 A turbine blade or use of a turbine blade in accordance with any one of claims 4 to 6, characterised in that at least one of the internal walls (4) meets an adjacent external wall (6) at an angle ci which is not greater than 60 .

8 A gas turbine engine, characterised in that the engine includes a turbine blade in accordance with claim 1, or in accordance with any one of claims 3 to 7 when appendant to claim 1.

9 A method of manufacturing a turbine blade for 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.

A turbine blade or use of a turbine blade substantially as hereinbefore described and/or as shown in Figures 2 to 4.

11 A gas turbine engine substantially as hereinbefore described and/or as shown in Figures 2 to 4.

12 A method of manufacturing a turbine blade for a gas turbine engine substantially as hereinbefore described and/or as shown in Figures 2 to 4.

Amendments to the claims have been filed as follows

CLAIMS

I 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.

2 Use of a turbine blade as claimed in claim 1, characterised in that the metal alloy is a NiAIX alloy in which X is one or more of hafnium, rhenium, titanium, chromium or gallium.

3 Use of a turbine blade in accordance with claim 1 or claim 2, characterised in that the turbine blade is provided with cavities (2) which are separated from adjacent cavities (2) by internal walls (4) and which are separated from the exterior of the turbine blade by external walls (6).

4 Use of a turbine blade in accordance with claim 3, characterised in that the thickness ratio Ti'Te of the internal and external walls (4, 6) is not greater than 1.5:1.

Use of a turbine blade in accordance with claim 3 or 4, characterised in that at

V

 least one of the internal walls (4) is provided with through holes (8), at least some of which are spaced apart by a spacing(s) which is not greater than 6 times the transverse dimension (D) of the holes (8). J0 it

6 Use of a turbine blade in accordance with any one of claims 3 to 5, characterised in that at least one of the internal walls (4) meets an adjacent external wall (6) at an angle a which is not greater than 600.

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 or claim 8 wherein the turbine blade is provided with cavities (2) which are separated from adjacent cavities (2) by internal walls (4) and which are separated from the exterior of the turbine blade by external walls (6).

A gas turbine engine as claimed in claim 9 wherein the thickness ratio Ti/Te of the internal and external walls (4, 6) is not greater than 1.5:1.

11 A gas turbine engine as claimed in claim 9 or claim 10 wherein at least one of the internal walls (4) is provided with through holes (8), at least some of which are spaced apart by a spacing(s) which is not greater than 6 times the transverse dimension (D) of the holes (8).

12 A gas turbine engine as claimed in any one of claims 9 to 11 wherein at least one of the internal walls (4) meets an adjacent external wall (6) at an angle a which is not greater than 60 .

13 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.

* 14 Use of a turbine blade substantially as hereinbefore described and/or as shown in Figures 2 and/or 4.

A gas turbine engine substantially as hereinbefore described and/or as shown in Figures 2 and/or 4.

16 A method of manufacturing a turbine blade for a gas turbine engine substantially as herein before described and/or as shown in Figures 2 and/or 4.