Patent Application
20100043929
The present invention relates to a single crystal component, in particular a single crystal turbine blade and a method of heat treating a single crystal turbine blade, in particular a method of heat treating a single crystal turbine blade.
Nickel base single crystal superalloy components, e.g. turbine blades and turbine vanes, are used in the hottest regions of an engine, e.g. a gas turbine engine. In a gas turbine engine hot combustion gases from the combustor impinge on turbine blades within the turbine section and the turbine blades convert the kinetic energy of the hot combustion gases into useful work energy to drive the compressor section. The high-pressure turbine blades are actively cooled and are often provided with thermal barrier coatings to reduce metal temperatures and minimise creep deformation of the turbine blades. However, nickel base single crystal superalloy turbine blades are subject to a wide range of stresses and temperatures that dictate the dominant mode of deformation in these materials.
Resistance to creep deformation is traditionally considered to be one of the most important design factors for a nickel base single crystal superalloy turbine blade. Nickel base single crystal superalloy turbine blades are carefully processed, heat treated, to optimise the resulting grain and precipitate structure for creep resistance. A nickel base single crystal superalloy is cast into an investment casting mould, which has a single crystal selector or a single crystal seed, and the nickel base single crystal superalloy is directionally solidified and the orientation of the single crystal is carefully controlled such that the <001> direction of the single crystal is closely aligned with the primary stress axis of the turbine blade.
Following directional solidification, multiple heat treatments, as shown in
An initial solution heat treatment S removes the residual gamma/gamma prime eutectic within the as cast microstructure of the nickel base single crystal superalloy turbine blade and assists homogenisation of the dendritic microstructure. After the solution heat treatment S the nickel base single crystal superalloy turbine blades are cooled at a specific cooling rate and then primary aged PA to nucleate a single distribution of gamma prime phase precipitates. Upon subsequent cooling to room temperature nano-sized secondary gamma prime phase precipitates often form within the gamma channels to relieve super-saturation of elements contained within the gamma phase matrix, as shown in
Nickel base single crystal superalloy turbine blades are intentionally processed, heat treated, to produce a comparatively homogeneous gamma prime precipitate size, shape and distribution contained throughout the microstructure.
Due to the large variation in operating temperatures, stresses and loading conditions experienced by a nickel base single crystal turbine blade during operation, the turbine blade is subjected to a wide range of deformation mechanisms. The aerofoil of the turbine blade is subjected to elevated temperatures, greater than 750° C., and undergoes creep deformation during service. The root of the turbine blade is subjected to lower temperatures, less than 750° C., and does not experience significant amounts of creep. The root of the turbine blade is subjected to large stresses during service that ultimately lead to deformation and failure due to low cycle fatigue.
Currently the gamma prime phase precipitates within the entire nickel base single crystal superalloy turbine blade are arranged to provide high degree of creep resistance in the aerofoil of the turbine blade without considering the effect on the low cycle fatigue life of the root of the turbine blade.
It is known from U.S. Pat. No. 4,921,405 and U.S. Pat. No. 5,451,142 to provide a layer/coating of polycrystalline superalloy bonded to the root of a nickel base single crystal superalloy turbine blade. The polycrystalline superalloy is plasma sprayed onto the root of the turbine blade. The layer of polycrystalline superalloy is intended to increase the low cycle fatigue life of the turbine blade.
U.S. Pat. No. 5,106,266 discloses providing a core of polycrystalline superalloy within and bonded in a cavity within the root of a nickel base single crystal superalloy turbine blade. The polycrystalline superalloy is plasma sprayed into the cavity in the root of the turbine blade. The core of polycrystalline superalloy is intended to increase the low cycle fatigue life of the turbine blade.
A problem with the prior art is that it requires a layer of polycrystalline superalloy on the root of the turbine blade or a core of polycrystalline superalloy within the root of the turbine blade to increase the fatigue life of a nickel base single crystal superalloy turbine blade.
Accordingly the present invention seeks to provide a novel single crystal turbine blade which reduces, preferably overcomes, the above mentioned problem.
Accordingly the present invention provides a single crystal component comprising a first region and a second region, the component comprising a nickel base single crystal superalloy having a gamma phase matrix and gamma prime phase precipitates distributed in the gamma phase matrix, wherein the first region of the component comprising a bi-modal distribution of gamma prime phase precipitates in the gamma phase matrix and the second region of the component comprising a uni-modal distribution of gamma prime phase precipitates in the gamma phase matrix, the uni-modal distribution of gamma prime phase precipitates consisting of primary cuboidal gamma prime phase precipitates and the bi-modal distribution of gamma prime phase precipitates consisting of primary cuboidal gamma prime phase precipitates and secondary spherical gamma prime phase precipitates and/or cuboidal gamma prime phase precipitates whereby the first region of the component has enhanced resistance to low cycle fatigue and the second region of the component has enhanced resistance to creep deformation.
Preferably the primary cuboidal gamma prime phase precipitates in the second region of the component have edge lengths within the range 200 to 700 nm.
Preferably the primary cuboidal gamma prime phase precipitates in the second region of the component have edge lengths within the range 250 to 500 nm.
Preferably the volume fraction of primary cuboidal gamma prime phase precipitates in the second region of the component is at least 50 vol %.
Preferably the volume fraction of primary cuboidal gamma prime phase precipitates in the second region of the component is at least 60 vol %.
Preferably the volume fraction of primary cuboidal gamma prime phase precipitates in the second region of the component is 70 vol %.
Preferably the gamma channel widths between the gamma prime phase precipitates in the second region of the component is within the range 10 to 100 nm.
Preferably the gamma channel widths between the gamma prime phase precipitates in the second region of the component is within the range 30 to 60 nm.
Preferably the primary cuboidal gamma prime phase precipitates in the first region of the component have edge lengths within the range of 200 to 700 nm.
Preferably the primary cuboidal gamma prime phase precipitates in the first region of the component have edge lengths within the range of 250 to 500 nm.
Preferably the secondary spheroidal and/or cuboidal gamma prime phase precipitates in the first region of the component have edge lengths within the range of 1 to 50 nm.
Preferably the secondary spheroidal and/or cuboidal gamma prime phase precipitates in the first region of the component have edge lengths within the range of 5 to 20 nm.
Preferably the secondary spheroidal and/or cuboidal gamma prime phase precipitates in the first region of the component are contained within the gamma channels between the primary cuboidal gamma prime phase precipitates.
Preferably the component is a turbine blade comprising a root, a shank, a platform and an aerofoil. Preferably the first region comprises the root and the shank and the second region comprises the platform and the aerofoil.
The single crystal turbine blade may comprise a shroud and the second region comprises the platform, the aerofoil and the shroud.
Preferably the nickel base single crystal superalloy comprises a 3rd, 4th or a 5th generation single crystal superalloy.
The present invention also provides a method of heat treating a single crystal component comprising a first region and a second region, the method comprising heat treating the component to produce primary cuboidal gamma prime precipitates in a gamma phase matrix in the first region and the second region, heat treating the component to produce secondary gamma prime phase precipitates in the first region.
Preferably the method comprises solution heat treating the component, primary ageing the component and secondary ageing the component wherein the primary ageing of the component comprises cooling the first region of the component at a lower cooling rate than the cooling rate of the second region of the component during the cooling from the primary ageing temperature of the primary ageing heat treatment to room temperature.
The method may comprise providing insulation on the first region of the component during the primary ageing heat treatment or providing insulation on the first region of the component during the cooling part of the primary ageing heat treatment to reduce the cooling rate of the first region of the component after the primary ageing heat treatment.
The method may comprise heating the first region of the component to further reduce the cooling rate of the first region of the component after the primary ageing heat treatment.
The method may comprise submerging the first region of the component in a molten salt bath maintained at an elevated temperature to reduce the cooling rate of the first region of the component.
The method may comprise directing a cooling fluid onto the surface of the second region of the component to reduce the cooling rate of the first region of the component relative to the cooling rate of the second region of the component.
Alternatively the method comprises solution heat treating the component, primary ageing the component and secondary ageing the component wherein the secondary ageing of the component comprises heating the second region of the component to a temperature above the solvus temperature of the fine secondary gamma prime precipitates while heating the first region of the component to a temperature below the solvus temperature of the fine secondary gamma prime precipitates during the secondary ageing heat treatment.
The method may comprise heating the second region of the component to the temperature above the solvus temperature of the fine secondary gamma prime precipitates to dissolve them.
The method may comprise providing insulation on the first region of the component and locally heat treating the first region of the component to precipitate secondary gamma prime phase precipitates while the insulation maintains the second region of the component at a temperature below the secondary gamma prime solvus so that secondary gamma prime phase precipitates are not formed in the second region of the component.
The method may comprise heating the whole of the component during part of the primary heat treatment, secondary heat treatment and coating heat treatment step and controlling the heating such that different temperature gradients are produced in the first region and the second region of the component.
The present invention will be more fully described by way of example with reference to the accompanying drawings in which:—
A single crystal turbine blade 10, as shown in
The uni-modal distribution of primary gamma prime phase precipitates consists of cuboidal gamma prime phase precipitates. The bi-modal distribution of gamma prime phase precipitates consists of primary cuboidal gamma prime phase precipitates and secondary spherical gamma prime phase precipitates and/or cuboidal gamma prime phase precipitates.
The root 12 and the shank 14 of the turbine blade 10 have enhanced resistance to low cycle fatigue and the platform 16 and the aerofoil of the turbine blade 10 have enhanced resistance to creep deformation.
The primary cuboidal gamma prime phase precipitates in the platform and the aerofoil of the turbine blade have edge lengths within the range 200 to 700 nm, more preferably within the range 250 to 500 nm. The volume fraction of primary cuboidal gamma prime phase precipitates in the platform and the aerofoil of the turbine blade is at least 50 vol %, more preferably at least 60 vol %, for example 70 vol %. The gamma channel widths between the gamma prime phase precipitates in the platform 16 and aerofoil 18 of the turbine blade 10 is within the range 10 to 100 nm, more preferably within the range 30 to 60 nm. Due to the high anti-phase boundary (APB) energy associated with dislocation shearing of the gamma prime phase precipitates, dislocations are forced to bow in-between the gamma prime phase precipitates within the gamma phase channels during deformation. The combination of a large volume fraction of gamma prime phase precipitates and narrow gamma phase channels maximises the Orowan resistance to dislocation bowing thereby minimising creep deformation at elevated temperature, above 750° C.
The primary cuboidal gamma prime phase precipitates in the root 12 and shank 14 of the turbine blade 10 have edge lengths within the range of 200 to 700 nm, more preferably within the range of 250 to 500 nm. The secondary spheroidal and/or cuboidal gamma prime phase precipitates in the root 12 and shank 14 of the turbine blade 10 have edge lengths within the range of 1 to 50 nm, more preferably within the range of 5 to 20 nm. The secondary spheroidal and/or cuboidal gamma prime phase precipitates in the root 12 and the shank 14 of the turbine blade 10 are contained within the gamma channels between the primary cuboidal gamma prime phase precipitates. Low cycle fatigue is the dominant failure mechanism at lower temperatures and occurs at temperatures lower than those for creep deformation and the fine dispersion of secondary gamma prime phase precipitates effectively serve as obstacles to restrict dislocation mobility. The root 12 and shank 14 of the turbine blade 10 are not subjected to elevated temperatures, above 750° C., and do not suffer dislocation climb mechanisms or dissolution of the small secondary gamma prime precipitates and hence the low cycle fatigue response and low cycle fatigue life of the root 10 of the turbine blade 10 is improved.
The uni-modal distribution of gamma prime phase precipitates in the aerofoil 18 and the platform 16 and the narrow gamma channels associated with the microstructure offer maximum resistance against tensile and creep deformation at temperatures above 800° C. in the aerofoil 18 of the turbine blade 10. The b-modal distribution of gamma prime phase precipitates in the root 12 and shank 14 improves the low cycle fatigue properties of the turbine blade 10. The turbine blade 10 is able to sustain a greater number of engine operating cycles before inspection or overhaul of the turbine blades 10 is required or the root 12 of the turbine blade 10 is able to withstand higher stresses.
The present invention is applicable to any nickel base single crystal superalloy strengthened by gamma prime phase precipitates. The present invention is particularly applicable to nickel base single crystal superalloys with relatively high content of refractory elements, e.g. tantalum, tungsten, molybdenum, niobium and/or rhenium, and in particular to 3rd, 4th or 5th generation nickel base single crystal superalloys. This is due to the high concentration of refractory elements supersaturating the gamma phase, modifying the kinetics of the superalloy system and stabilising the secondary gamma prime phase precipitates to temperatures above 750° C.
The 3rd generation nickel base single crystal superalloys have greater than 3 wt % rhenium, the 4th generation nickel base single crystal superalloys have less than 4 wt % ruthenium and the 5th generation of nickel base single crystal superalloys have more than 4 wt % ruthenium. An example of a 3rd generation nickel base single crystal superalloy is CMSX10, an example of a 4th generation nickel base single crystal superalloy is TMS138A and an example of a 5th generation nickel base single crystal superalloy is TMS196.
CMSX10 consists of 5.7 wt % Al, 2.0 wt % Cr, 3.0 wt % Co, 0.4 wt % Mo, 0.2 wt % Ti, 8.0 wt % Ta, 5.0 wt % W, 6.0 wt % Re, 0.03 wt % Hf and the balance Ni plus incidental impurities.
TMS138A consists of 5.7 wt % Al, 3.2 wt % Cr, 5.8 wt % Co, 2.9 wt % Mo, 5.6 wt % Ta, 5.6 wt % W, 5.8 wt % Re, 3.6 wt % Ru, 0.01 wt % Hf and the balance Ni plus incidental impurities.
TMS196 consists of 5.6 wt % Al, 4.6 wt % Cr, 5.6 wt % Co, 2.4 wt % Mo, 5.6 wt % Ta, 5.0 wt % W, 6.4 wt % Re, 5.0 wt % Ru, 0.1 wt % Hf and the balance Ni plus incidental impurities.
A further single crystal turbine blade 110, as shown in
A nickel base single crystal superalloy turbine blade according to the present invention has a non-uniform gamma prime phase precipitate microstructure which possesses mechanical properties that match more closely the actual operating conditions of the turbine blade and provides performance advantages over a turbine blade with a uniform gamma prime phase precipitate microstructure. This extends turbine blade life, particularly turbine blades prone to low cycle fatigue failure below the platform of the turbine blade in the root and/or shank of the turbine blade. Alternatively the turbine rotor/turbine disc and turbine blades may be operated at higher rotational speeds to increase the efficiency of the gas turbine engine with the same turbine blade life.
The nickel base single crystal turbine blade is produced by pouring the molten nickel base superalloy in an investment mould and then directionally solidifying the nickel base superalloy within the investment mould to produce a nickel base superalloy turbine blade 10. The investment mould either contains a single crystal seed at the base of the investment mould and contacting a chill plate or the investment mould comprises a single crystal selector, e.g. a spiral, to ensure that a single crystal is formed in the investment mould during directional solidification.
The dual microstructure is produced in the nickel base single crystal superalloy turbine blade 10 by heat treating the turbine blade 10 to produce primary cuboidal gamma prime precipitates in a gamma phase matrix in the root 12, the shank 14, the platform 16 and the aerofoil 18 and by heat treating the turbine blade 10 to produce secondary gamma prime phase precipitates in the root 12 and the shank 14.
In particular the dual microstructure is produced by cooling the root 12 and the shank 14 of the turbine blade 10 at a lower cooling rate B than the cooling rate A of the platform 16 and the aerofoil 18 of the turbine blade 10 during the cooling from the primary ageing temperature of the primary ageing heat treatment PA to room temperature as shown in
In a first method this is achieved by applying insulation on the root 12 and shank 14 of the turbine blade 10 during the primary ageing heat treatment or alternatively only during the cooling part of the primary ageing heat treatment. The insulation may be a metallic, refractory or other suitable insulation material located on the root 12 and shank 14 of the turbine blade 10. The increased thermal mass around the root 12 and the shank 14 reduces the cooling rate for the root 12 and the shank 14 relative to the un-insulated platform 16 and aerofoil 18 and therefore promotes the precipitation of the fine secondary gamma prime precipitates. The insulating material is selected to give the desired cooling rate in conjunction with a process model of the cooling of the turbine blade after the primary ageing heat treatment PA.
In a second method this is achieved by applying insulation on the root 12 and shank 14 of the turbine blade 10 and providing a heater in the insulation to heat the root 12 and the shank 14 of the turbine blade 14 to further reduce the cooling rate of the root 12 and the shank 14 of the turbine blade 10 after the primary ageing heat treatment PA. It may be possible to simply provide a heater to heat the root 12 and the shank 14 without the insulation to reduce the cooling rate of the root 12 and the shank 14 of the turbine blade 10 after the primary ageing heat treatment PA. The heater may be a localised induction heater, direct or indirect.
In a third method the root 12 and shank 14 of the turbine blade 10 are submerged in a molten salt bath maintained at an elevated temperature to reduce the cooling rate of the root 12 and shank 14 of the turbine blade 10.
In a fourth method the dimensions of the cast turbine blade 10 in the region of the root 12 and shank 14 are oversized so that the mass in the region of the root 12 and shank 14 is increased and hence reduces the cooling rate of the root 12 and shank 14 of the turbine blade 10. However, the additional mass in the region of the root 12 and shank 14 of the turbine blade 10 results in the removal of more material by machining to the required finished dimensions.
In a fifth method a cooling fluid, liquid or gas, is directed onto the internal surface or external surface of the platform 16 and the aerofoil 18 of the turbine blade 10 to reduce the cooling rate of the root 12 and the shank 14 of the turbine blade 10 relative to the cooling rate of the platform 16 and the aerofoil 18 of the turbine blade 10. This method may be used in conjunction with the first, second, third and fourth methods discussed above.
Alternatively the dual microstructure is produced by heating the platform 16 and the aerofoil 18 of the turbine blade 10 to a temperature C above the solvus temperature of the fine secondary gamma prime precipitates while heating the root 12 and the shank 14 of the turbine blade 10 to a temperature D below the solvus temperature of the fine secondary gamma prime precipitates during the secondary ageing heat treatment SA, as shown in
In a sixth method this is achieved by providing a heater to heat the platform 16 and the aerofoil 18 to the temperature C above the solvus temperature of the fine secondary gamma prime precipitates to dissolve them. The heater may be a localised induction heater, direct or indirect.
In a seventh method insulation is provided on the platform 16 and aerofoil 18 of the turbine blade 10 and the root 12 and the shank 14 of the turbine blade 10 are given a local heat treatment to precipitate secondary gamma prime phase precipitates while the insulation maintains the platform 16 and the aerofoil 18 at a temperature below the secondary gamma prime solvus so that secondary gamma prime phase precipitates are not formed in the platform 16 and the aerofoil 18. The insulation may be cooled by a cooling fluid, liquid or gas.
In an eighth method induction heaters are provided to heat to the whole of the turbine blade 10 as part of the primary heat treatment, secondary heat treatment and coating heat treatment step. The current supplied to the induction heaters/induction coils is controlled such that different temperature gradients are produced in different sections of the root 12, shank 14, platform 16 and aerofoil 18.
In general the cooling rate B is lower, or slower, than cooling rate A. The rate chosen affects the solvus temperature Tsolv of the fine gamma prime precipitates and therefore temperatures C and D depend on the previous cooling rate A or B. Each nickel base superalloy will have its own cooling rates A and B and temperatures C and D.
Although the present invention has been described with reference to a nickel base superalloy turbine blade it is equally applicable to other nickel base superalloy components where it is desired to have different mechanical properties in different regions of the nickel base single crystal component.
| Number | Date | Country | |
|---|---|---|---|
| 61136271 | Aug 2008 | US |
