US20140099209A1

US20140099209A1 - Gas turbine component, method for its production and casting mold for use of this method

Info

Publication number: US20140099209A1
Application number: US14/033,589
Authority: US
Inventors: Fathi Ahmad; Wayne J. McDonald; Uwe Paul
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2012-10-08
Filing date: 2013-09-23
Publication date: 2014-04-10
Also published as: EP2716386A1

Abstract

A method is provided for casting a gas turbine component having at least a first section and a second section. As per the method, a liquid metal is cast into a casting mold and solidification of the metal is induced by means of controlled cooling. The cooling of the liquid metal in the first section is controlled in such a way that the metal solidifies in a directionally solidified crystal structure or a single-crystal structure. In the second section the cooling is carried out in such a way that it solidifies in a multidirectional crystal structure.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority of European Patent Office application No. 12187594.2 EP filed Oct. 8, 2013. All of the applications are incorporated by reference herein in their entirety.

FIELD OF INVENTION

The present invention relates to a gas turbine component and to a method for its production. The invention furthermore relates to a casting mold which may be used in the method for producing the gas turbine component.

BACKGROUND OF INVENTION

Gas turbine components, for example guide vanes and rotor blades, of gas turbines are exposed to harsh conditions during the operation of a gas turbine. Particularly in the case of rotor blades, centrifugal forces act during operation, and these necessitate a high creep strength of the material. The high creep strength must also exist at the temperatures prevailing during the operation of a gas turbine. Large rotor blades, in particular, experience high centrifugal forces which require high creep strength at high temperatures.
At the high temperatures prevailing in the gas turbines, particularly in the front stages, above all grain boundaries extending transversely to the main stress direction in the material structure weaken the material, so that a higher creep strength can be achieved by eliminating grain boundaries. In particular, large gas turbine components are therefore increasingly being made of material having directionally solidified structure or single-crystal structure. In this case, a directionally solidified structure is intended in particular to mean columnar crystalline forms, in which there are no grain boundaries in a particular spatial direction.
The production of directionally solidified or single-crystal gas turbine components is carried out by filling a casting mold, heated to above the melting temperature of the material to be cast, with the material, the melt of the material at the bottom of the casting mold being at a cooler position and forming a directional grain structure there. The mold with the material is then moved slowly from the hot part of the furnace into a cooler part of the furnace, so that the cooling material propagates the directionally directed structure of the already solidified material and therefore solidifies in a directionally solidified microstructure or a single-crystal microstructure. A corresponding method is described, for example, in US 2010/0193150 A1.
Particularly in the case of large components, however, the slow movement of the casting mold from the hot region to the cool region of the furnace requires a corresponding length of time which reduces the throughput of the production of the components. The reason for the long length of time, however, is not only the size of the component, which determines the length of time required for extraction of the component from the furnace with a given movement speed, but also the larger mass of the large components, which reduces the cooling rate, and this itself slows the solidification of the component.

SUMMARY OF INVENTION

It is an object of the present invention to provide an advantageous method for producing a gas turbine component. It is another object of the present invention to provide a casting mold which may be used advantageously in the method according to the invention. It is yet another object of the present invention to provide an advantageous gas turbine component.
The above objects are achieved by the features of the independent claim(s).
In the method according to the invention for casting a gas turbine component having at least a first section and a second section, a liquid metal is cast into a casting mold and solidification of the metal is induced by means of controlled cooling. The cooling of the liquid metal in the first section of the gas turbine component is controlled in such a way that the metal solidifies in a directionally solidified crystal structure or in a single-crystal structure. In the second section the cooling is carried out in such a way that it solidifies in a multidirectional crystal structure (globulitic or polycrystalline crystal structure).
The invention is based on the discovery that, for example in the case of a gas turbine rotor blade, the creep strength at high temperature is weakened primarily in the region of the main blade by a globulitic or polycrystalline crystal structure. At the blade platform and at the blade root, conversely, lower temperatures prevail during operation of the blade, at which a globulitic or polycrystalline crystal structure supports a good creep strength.
It is at this point that the method according to the invention is relevant. By subjecting merely one section of the gas turbine component, for example the main blade of a turbine blade, to directionally solidified or single-crystal solidification, but conversely subjecting the second section to multidirectional solidification, the length of time for the production of the component can be reduced since globulitic solidification can take place more rapidly than directionally solidified or single-crystal solidification. That section in which the multidirectional solidification takes place can be moved rapidly from the hot furnace region into the cooler furnace region, so that the length of time required for the directionally solidified solidification, or the single-crystal solidification, is no longer determined by the dimension of the entire component and the mass of the entire component, but then only by that part of the component which must have the high creep strength at high temperature. In the turbine blade mentioned by way of example, the section with the blade root and the blade platform can therefore be moved rapidly from the hot furnace region into the cooler furnace region, so that the lengths of time required for inducing the directional solidification or the single-crystal solidification are then determined only by the amount and mass of the main blade.
By the method according to the invention, overall, the production time for producing a gas turbine component, in particular a turbine blade, can be reduced without those sections of the component which are exposed to high centrifugal forces at high temperatures being weakened. Owing to the reduced production time, the production costs for the gas turbine components can be reduced.
If rapid removal of the second section from the furnace is carried out after the directionally solidified solidification or the single-crystal solidification has taken place in the first section, there may be a tendency for the remaining liquid metal to solidify first in the narrow low-mass regions before it solidifies in the wider and higher-mass regions. This can lead to regions in which liquid metal is surrounded by already solidified metal. When the final cooling and solidification of the metal then takes place, volume changes in the cooling metal lead to there being a higher overall porosity in the wide and high-mass regions than in the previously solidified narrow and low-mass regions.
In order to avoid the occurrence of an increased overall porosity, the method according to the invention may be refined in such a way that, when the second section of the gas turbine component comprises a low-mass region and a high-mass region, additional liquid metal is supplied to the high-mass region at least when the directionally solidified crystal structure or the single-crystal structure is being formed. In other words, at least during the rapid cooling during the production of the multidirectionally solidified structure, additional liquid metal is supplied to the high-mass region of the second section in order to compensate for the volume change during cooling and thus avoid the occurrence of an increased porosity.
The additional liquid metal may, in particular, be stored in the form of a reservoir of liquid metal which adjoins the high-mass region. It is furthermore advantageous for the additional liquid metal to be kept liquid for a long time by thermal insulation of the reservoir. In this way, it is possible for the reservoir of liquid metal to be available long enough in order to be able to compensate for the volume change in the high-mass region of the second section until solidification. As an alternative or in addition to the insulation, the additional liquid metal may also be kept liquid by adding an exothermic material to the free surface of the additional liquid metal in the reservoir.
When the gas turbine component is a rotor blade or guide vane, the first section of the gas turbine component is the main blade of the rotor blade or guide vane. In particular when the gas turbine component is a rotor blade, the second section may include the blade root and the blade platform, the high-mass region and the low-mass region both being part of the blade root and the low-mass region lying between the high-mass region and the platform. The supply of the additional liquid metal may in this case be carried out to side surfaces of the high-mass region or to the end surface of the high-mass region, which constitutes the end of the rotor blade facing away from the main blade.
With the method according to the invention, it is possible to produce a gas turbine component according to the invention having a first section and a second section, which is formed integrally with the first section. The first section has a directionally solidified crystal structure or a single-crystal structure, while the second section has a multidirectionally solidified crystal structure. The gas turbine component may, for example, be formed as a guide vane or, in particular, as a rotor blade, the first section being formed by the main blade.
The gas turbine component according to the invention offers the advantage that it can be produced rapidly, since only those sections which are exposed to the high temperatures during operation, and which therefore need to have high creep strength at high temperatures, solidify in a directionally solidified or single-crystal fashion in a time-consuming method. The remaining sections, on the other hand, can solidify more rapidly in a multidirectionally solidified fashion.
A casting mold according to the invention for casting a gas turbine component by means of the method according to the invention comprises at least a first casting mold section for molding the first section of the gas turbine component and a second casting mold section for molding the second section of the gas turbine component. The casting mold furthermore comprises a supply channel which communicates with the second casting mold section in order to permit supply of additional liquid metal into the second casting mold section. Owing to the fact that the supply of additional liquid metal into the second casting mold section is made possible, it is possible to avoid the aforementioned volume changes leading to an increased porosity of the solidified metal in the second casting mold section during the rapid cooling of the second casting mold section.
In a refinement of the casting mold according to the invention, the supply channel comprises a volume for receiving a reservoir of liquid metal. The casting mold may be thermally insulated where the supply channel is located, in order to keep the reservoir of liquid metal liquid for as long as possible.
In the casting mold according to the invention, the second casting mold section may comprise a mold region with a narrow intermediate space for molding the low-mass region of the gas turbine component and a casting mold region with a wide intermediate space for molding the high-mass region of the gas turbine component. The supply channel then leads to the casting mold region with the wide intermediate space. The above-described problem of increased porosity in the solidified material due to the volume change during the solidification is most serious in the region of the wide intermediate space. The supply of the additional liquid material is therefore preferably carried out into the casting mold region with the wide intermediate space.
If the gas turbine component is a rotor blade or guide vane, the first casting mold section may in particular have a contour for molding the main blade. In general, the main blade constitutes the section of the turbine blade most stressed at high temperatures, so that this section is produced in a single-crystal structure or directionally solidified crystal structure.
In particular when the gas turbine component is a rotor blade, the second casting mold section may include a blade root casting mold section for molding the blade root and a platform casting mold section for molding the blade platform. The casting mold region with the narrow intermediate space and the casting mold region with the wide intermediate space are then both part of the blade root casting mold section, the casting mold region with the narrow intermediate space lying between the casting mold region with the wide intermediate space and the platform casting mold section. This configuration of the casting mold leads to the additional liquid metal being supplied to those regions of the rotor blade in which the susceptibility to the formation of an increased porosity is greatest, namely the high-mass part of the blade root. The supply openings for supplying the additional liquid metal may, in particular, lead to casting mold side surfaces of the casting mold region with the wide intermediate space or to a casting mold end surface of the casting mold region with the wide intermediate space, the end surface defining the end surface of the blade root facing away from the main blade.

BRIEF DESCRIPTION OF THE DRAWINGS

Other features, properties and advantages of the present invention may be found in the following description of exemplary embodiments with reference to the appended figures.

FIG. 1 shows a first exemplary embodiment of a casting mold according to the invention in a sectional view.

FIG. 2 shows the casting mold of FIG. 1 in a side view.

FIG. 3 shows a flow chart of the method according to the invention.

FIG. 4 shows a gas turbine blade as an exemplary embodiment of the gas turbine component according to the invention.

FIG. 5 shows a second exemplary embodiment of the casting mold according to the invention in a sectional view.

FIG. 6 shows the casting mold of FIG. 5 in a side view.

FIG. 7 shows by way of example a gas turbine in a partial longitudinal section.

FIG. 8 shows in perspective view a rotor blade or guide vane of a turbomachine.

FIG. 9 shows a combustion chamber of a gas turbine.

DETAILED DESCRIPTION OF INVENTION

A first exemplary embodiment of a casting mold according to the invention will be described below with reference to FIGS. 1 and 2, before the way in which the casting method according to the invention is carried out by using the casting mold according to the invention is then described with reference to FIG. 3.
The exemplary embodiment of a casting mold according to the invention, as shown in FIGS. 1 and 2, shows a casting mold for casting a gas turbine rotor blade. The casting mold 1 comprises a first casting mold section 3 for molding the main blade 4 of the turbine blade to be cast, and a second casting mold section 5 for molding the blade platform 6 and the blade root 7 of the turbine blade to be cast. The main blade in this case constitutes a first section of the gas turbine component, which is intended to be cast in a directionally solidified structure, also abbreviated to DS structure. Instead of with a directionally solidified crystal structure, the main blade may also be produced with a single-crystal structure. The blade platform 6 and the blade root 7 constitute a second section of the gas turbine component, in which solidification of the liquid metal in a directionally solidified structure is not necessary. This section therefore also constitutes a section in which the liquid metal can solidify in a multidirectional structure.
The blade root 7 has a wide high-mass region 8 and a narrow low-mass region 9, which lies between the high-mass region 8 of the blade root 7 and the blade platform 6. Correspondingly, the second casting mold section 5 has a casting mold region 10 with a wide intermediate space for molding the wide high-mass region 8 of the blade root 7, and a casting mold region 11 with a narrow intermediate space for molding the narrow low-mass region 9 of the blade root 7.
The casting mold 1 furthermore comprises at least one supply channel 12, which leads to the casting mold region 10 with the wide intermediate space and is used to supply additional liquid metal into the casting mold region 10 with the wide intermediate space during the casting process. In the present exemplary embodiment, there are two supply channels 12, which lead to the side surfaces 8 a, 8 b of the wide high-mass region 8 of the blade root 7. In other words, the supply channels 12 lead to casting mold side surfaces 10 of the casting mold region with a wide intermediate space. These supply channels 12 are configured in terms of their volume in such a way that they can act as a reservoir for the liquid metal. Furthermore, the casting mold 1 is provided in the region of the reservoir with a thermal insulation 13, which in the present exemplary embodiment is applied on the outer side of the casting mold 1. As an alternative, however, the insulation may also be applied on the inner side of the casting mold, or the material of the casting mold may have a higher insulating capacity in the region of the supply channels 12 than in other regions of the casting mold.
The casting method according to the invention will be explained in more detail below with reference to FIG. 3, which represents a flow chart of the casting method.
In a first step S1 of the method according to the invention, the casting mold 1 is filled with liquid metal. The casting mold is in this case located in a special furnace which, in a vertical arrangement, has a hot zone and a cooler zone arranged underneath, and in which the casting mold can be moved continuously with a controlled speed from the hot zone into the cooler zone. In this case, it is merely the relative movement between the casting mold and the hot/cool boundary of the furnace which is important, i.e. the movement of the filled casting mold 1 from the hot region into the cooler region of the furnace may be carried out by moving the casting mold itself, by moving the hot/cool boundary of the furnace with a stationary casting mold, or by moving both the casting mold and the hot/cool boundary of the furnace.
In the present exemplary embodiment, the casting mold is filled with metal in such a way that not only the casting mold 1 itself, but also the reservoir formed by the supply channels 12 is filled with liquid metal.
After the casting mold 1 has been filled with the liquid metal, in step S2 a slow movement of the casting mold section 3 formed in order to form the main blade 2 takes place from the hot region into the cooler region of the furnace. For example, it may be moved out in such a way that first the main blade tip is moved out of the hot region into the cooler region of the furnace, so that solidification of the liquid metal is induced in the region of the blade tip. The movement speed, the temperature distribution in the furnace and the casting mold structure in the blade tip region are in this case selected in such a way that this solidification takes place in a directionally solidified crystal structure or a single-crystal structure. By further controlled slow movement of the first casting mold section out of the hot region into the cooler region of the furnace, it is possible for the solidification front in the liquid metal to be moved in the direction of the blade platform 6 and in this case to adopt and propagate the crystal structure of the already solidified metal part. In this way, by slow movement of the first casting mold section 3 out of the hot region into the cooler region of the furnace, it is possible for the entire main blade 4 to consist of a directionally solidified or single-crystal structure, when the metal has solidified in the main blade region.
After the directional solidification of the metal in the first casting mold section 3 has taken place, in a next step S3 the second casting mold section 5, which is used to form the blade platform 6 and blade root 7, can be moved out of the hot region into the cooler region of the furnace more rapidly than was the case for the first casting mold section 3. As a consequence of the more rapid outward movement, controlled slow solidification then no longer takes place in a directionally solidified or single-crystal structure, but instead in a multidirectional crystal structure. The grain boundaries occurring in the case of such a multidirectional crystal structure are not in this case problematic, since the multidirectional crystal structure supports the creep strength of the blade platform and of the blade root at the temperatures prevailing at the blade platform and the blade root during the operation of the turbine blade. Owing to the more rapid movement from the hot region into the cooler region of the furnace, and the resulting more rapid solidification of the liquid metal in the second casting mold region, the production time of the gas turbine blade, and therefore the production costs, can be reduced.
However, the rapid movement of the second casting mold section 5 from the hot region into the cooler region of the furnace in step S3 also entails the tendency that the metal solidifies in the casting mold region 11 with the narrow intermediate space, in which there is less mass of liquid metal, more rapidly than in the casting mold region 10 with the wide intermediate space, in which there is more mass of liquid metal. This can lead to enclosed volumes of liquid metal which are bounded by the casting mold or already solidified metal. When these volumes ultimately solidify last, they may have a higher overall porosity than the other regions of blade root or the blade platform owing to the volume reduction of the metal associated with the solidification. In order to avoid this undesired effect, additional liquid metal, which is supplied to the casting mold region 10 with the wide intermediate space, lies in the reservoir formed by the supply channels 12 at least during the cooling of the liquid metal in the second casting mold section 5, in order to compensate by additional material for the volume reduction occurring during cooling and therefore counteract the formation of an increased overall porosity (step S4).
With the aid of the thermal insulation 13 of the casting mold 1, where the supply channels 12 are located, it is possible for the liquid metal in the reservoir to be kept liquid for long enough until the high-mass region 8 of the blade root 7 is fully solidified. In addition or as an alternative, the heat required to keep the metal liquid in the reservoir may be supplied by using an exothermic material on the free surface of the liquid metal in the reservoir.
After the cooling and solidification of the metal located in the second casting mold section 5, in step S5 the casting mold may be removed.
After the removal of the casting mold in step S5, finishing of the blade root 7 may then be carried out in step S6 in order to produce the intended contour of the blade root.
In the described method, it is sufficient for the additional liquid metal to lie in the reservoir during the cooling of the metal in the second casting mold section 5, i.e. in step S4. Typically, however, the reservoir is already filled with it when the hot casting mold located in the furnace is filled, in order to avoid subsequent filling of the reservoir and therefore a second filling process.
The result of the method according to the invention is a turbine blade 15 as schematically represented in FIG. 4. It comprises a main blade 4, which has a directionally solidified crystal structure, that is to say a so-called columnar crystal structure having grains which extend over the entire length of the workpiece, or a single-crystal structure. The blade platform 6 and the blade root 7, on the other hand, have a multidirectional crystal structure, i.e. a polycrystalline or globulitic crystal structure. Where particular requirements are to be placed on the creep strength at high temperatures, namely in the region of the main blade 7, the turbine blade therefore has a high creep strength at the temperatures prevailing there during operation, while in the region of the blade platform 6 and the blade root 4, the requirements for the creep strength at the lower temperatures prevailing there during operation in comparison with the main blade are supported by the polycrystalline crystal structure.
A second exemplary embodiment of a casting mold according to the invention, which may be used in a method as described with reference to FIG. 3, will be described below with reference to FIGS. 5 and 6.
Elements which correspond to the casting mold from FIGS. 1 and 2 are denoted by the same references as in FIGS. 1 and 2, and will not be explained again in order to avoid repetition.
The casting mold shown in FIGS. 5 and 6 differs from the casting mold shown in FIGS. 1 and 2 in the arrangement of the supply channel for the additional liquid metal. While in the first exemplary embodiment of the casting mold there are two supply channels, which open from the side into the casting mold region 10 for molding the high-mass blade root section 8, in the exemplary embodiment of FIGS. 5 and 6 there is merely one supply channel 16, which leads to the end surface of the region of the casting mold 1 forming the blade root. In the finished turbine blade, this end surface forms the end surface 8 c of the blade root 7 facing away from the main blade 5. In comparison with the solution shown in FIGS. 1 and 2, in which the additional liquid metal was supplied from the reservoir to the side surfaces 8 a, 8 b of the blade root, the solution shown in FIGS. 5 and 6 offers the advantage that the end surface 8 c has a simpler contour than the side surfaces 8 a, 8 b, and the final contour can therefore be produced with little outlay after removal of the casting mold.
The present invention has been described in detail with the aid of specific exemplary embodiments for explanatory purposes. Differences from the described exemplary embodiments are however possible, for which reason the invention is not intended to be limited to features of the exemplary embodiments, but merely by the appended claims. For example, it is possible for the liquid metal of the reservoir to be supplied to the high-mass region of the blade root 7 not surface-wide but in the form of a grid. To this end, for example, the reservoir 16 of FIG. 5 may be separated from the end surface 17 by a ceramic layer, the ceramic layer having openings arranged in the form of a grid in order to allow the liquid metal to flow through into the high-mass region 8 of the blade root. A similar solution is, of course, also possible in the exemplary embodiment represented in FIGS. 1 and 2. Arrangements of the supply channels other than those shown in the figures are also possible, so long as they do not hinder the slow movement of the first casting mold section out of the furnace during the production of the directionally solidified or single-crystal structure.
A gas turbine having turbine components, which contains components that can be produced by the method according to the invention, will be described below.
FIG. 7 shows by way of example a gas turbine 100 in a partial longitudinal section.
The gas turbine 100 internally comprises a rotor 103, which will also be referred to as the turbine rotor, mounted so as to rotate about a rotation axis 102 and having a shaft 101.
Successively along the rotor 103, there are an intake manifold 104, a compressor 105, an e.g. toroidal combustion chamber 110, in particular a ring combustion chamber, having a plurality of burners 107 arranged coaxially, a turbine 108 and the exhaust manifold 109.
The ring combustion chamber 110 communicates with an e.g. annular hot gas channel 111. There, for example, four successively connected turbine stages 112 form the turbine 108.
Each turbine stage 112 is formed for example by two blade rings. As seen in the flow direction of a working medium 113, a guide vane row 115 is followed in the hot gas channel 111 by a row 125 formed by rotor blades 120.
The guide vanes 130 are in this case fastened on an inner housing 138 of a stator 143, while the rotor blades 120 of a row 125 are fitted on the rotor 103, for example by means of a turbine disk 133.
Coupled to the rotor 103, there is a generator or a work engine (not shown).
During operation of the gas turbine 100, air 135 is taken in and compressed by the compressor 105 through the intake manifold 104. The compressed air provided at the turbine-side end of the compressor 105 is delivered to the burners 107 and mixed there with a fuel. The mixture is then burnt to form the working medium 113 in the combustion chamber 110. From there, the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working medium 113 expands by imparting momentum, so that the rotor blades 120 drive the rotor 103 and the latter drives the work engine coupled to it.
The components exposed to the hot working medium 113 experience thermal loads during operation of the gas turbine 100. Apart from the heat shield elements lining the ring combustion chamber 110, the guide vanes 130 and rotor blades 120 of the first turbine stage 112, as seen in the flow direction of the working medium 113, are thermally loaded the most.
In order to withstand the temperatures prevailing there, they may be cooled by means of a coolant.
Substrates of the components may likewise comprise a directional structure, i.e. they are single-crystal (SX structure) or comprise only longitudinally directed grains (DS structure).
Iron-, nickel- or cobalt-based superalloys are for example used as the material for the components, in particular for the turbine blades 120, 130 and components of the combustion chamber 110.
Such superalloys are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.
The blades 120, 130 may also have coatings against corrosion (MCrAlX: M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon, scandium (Sc) and/or at least one rare earth element, or hafnium). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
On the MCrAlX, there may furthermore be a thermal barrier layer, which consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam physical vapor deposition (EB-PVD).
The guide vanes 130 comprise a guide vane root (not shown here) facing the inner housing 138 of the turbine 108, and a guide vane head lying opposite the guide vane root. The guide vane head faces the rotor 103 and is fixed on a fastening ring 140 of the stator 143.
FIG. 8 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121.
The turbomachine may be a gas turbine of an aircraft or of a power plant for electricity generation, a steam turbine or a compressor.
The blade 120, 130 comprises, successively along the longitudinal axis 121, a fastening zone 400, a blade platform 403 adjacent thereto as well as a main blade 406 and a blade tip 415.
As a guide vane 130, the vane 130 may have a further platform (not shown) at its vane tip 415.
A blade root 183 which is used to fasten the rotor blades 120, 130 on a shaft or a disk (not shown) is formed in the fastening zone 400.
The blade root 183 is configured, for example, as a hammerhead. Other configurations as a firtree or dovetail root are possible.
The blade 120, 130 comprises a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade 406.
In conventional blades 120, 130, for example solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade 120, 130.
Such superalloys are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.
The blade 120, 130 may in this case be manufactured by a casting method, also by means of directional solidification, by a forging method, by a machining method or combinations thereof.
Workpieces with a single-crystal structure or single-crystal structures are used as components for machines which are exposed to heavy mechanical, thermal and/or chemical loads during operation.
Such single-crystal workpieces are manufactured, for example, by directional solidification from the melt. These are casting methods in which the liquid metal alloy is solidified to form a single-crystal structure, i.e. to form the single-crystal workpiece, or is directionally solidified.
Dendritic crystals are in this case aligned along the heat flux and form either a rod crystalline grain structure (columnar, i.e. grains which extend over the entire length of the workpiece and in this case, according to general terminology usage, are referred to as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of a single crystal. It is necessary to avoid the transition to globulitic (polycrystalline) solidification in these methods, since nondirectional growth will necessarily form transverse and longitudinal grain boundaries which negate the beneficial properties of the directionally solidified or single-crystal component.
When directionally solidified structures are referred to in general, this is intended to mean both single crystals which have no grain boundaries or at most small-angle grain boundaries, and also rod crystal structures which, although they do have grain boundaries extending in the longitudinal direction, do not have any transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures.
Such methods are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1.
The blades 120, 130 may also have coatings against corrosion or oxidation, for example MCrAlX (M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf)). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
The density is preferably 95% of the theoretical density.
A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an interlayer or as the outermost layer).
The layer composition preferably comprises Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. Besides these cobalt-based protective coatings, it is also preferable to use nickel-based protective layers such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re.
On the MCrAlX, there may furthermore be a thermal barrier layer, which is preferably the outermost layer and consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
The thermal barrier layer covers the entire MCrAlX layer.
Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam physical vapor deposition (EB-PVD).
Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance. The thermal barrier layer is thus preferably more porous than the MCrAlX layer.
Refurbishment means that components 120, 130 may need to be stripped of protective layers (for example by sandblasting) after their use. The corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the component 120, 130 are also repaired. The component 120, 130 is then recoated and the component 120, 130 is used again.
The blade 120, 130 may be designed to be a hollow or solid. If the blade 120, 130 is intended to be cooled, it will be hollow and optionally also comprise film cooling holes 418 (indicated by dashes).
FIG. 9 shows a combustion chamber 110 of a gas turbine. The combustion chamber 110 is designed for example as a so-called ring combustion chamber in which a multiplicity of burners 107, arranged in the circumferential direction around a rotation axis 102, open into a common combustion chamber space 154 and generates the flames 156. To this end, the combustion chamber 110 as a whole is designed as an annular structure which is positioned around the rotation axis 102.
In order to achieve a comparatively high efficiency, the combustion chamber 110 is designed for a relatively high temperature of the working medium M of about 1000° C. to 1600° C. In order to permit a comparatively long operating time even under these operating parameters which are unfavorable for the materials, the combustion chamber wall 153 is provided with an inner lining formed by heat shield elements 155 on its side facing the working medium M.
Each heat shield element 155 consisting of an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) on the working medium side, or is made of refractory material (solid ceramic blocks).
These protective layers may be similar to the turbine blades, i.e. for example MCrAlX means: M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
On the MCrAlX, there may furthermore be a, for example, ceramic thermal barrier layer which consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam physical vapor deposition (EB-PVD).
Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance.
Refurbishment means that heat shield elements 155 may need to be stripped of protective layers (for example by sandblasting) after their use. The corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the heat shield element 155 are also repaired. The heat shield elements 155 are then recoated and the heat shield elements 155 are used again.
Owing to the high temperatures inside the combustion chamber 110, a cooling system may also be provided for the heat shield elements 155 or for their retaining elements. The heat shield elements 155 are then, for example, hollow and optionally also have cooling holes (not shown) opening into the combustion chamber space 154.

Claims

1. A method for casting a gas turbine component having at least a first section and a second section, the method comprising:

casting a liquid metal into a casting mold, and

inducing solidification of the metal by means of controlled cooling,

wherein the cooling of the liquid metal in the first section is controlled in such a way that the metal solidifies in a directionally solidified crystal structure or a single-crystal structure, and

wherein in the second section the cooling is carried out in such a way that it solidifies in a multidirectional crystal structure.

2. The method as claimed in claim 1, wherein the second section of the gas turbine component comprises a low-mass region and a high-mass region, wherein the method further comprises supplying additional liquid metal to the high-mass region at least when the directionally solidified or single-crystal structure is being formed.

3. The method as claimed in claim 2, comprising storing the additional liquid metal in the form of a reservoir of liquid metal which adjoins the wide region.

4. The method as claimed in claim 3, wherein the additional liquid metal is kept liquid by thermal insulation of the reservoir.

5. The method as claimed in claim 3, wherein the additional liquid metal is kept liquid by adding an exothermic material to the free surface of the additional liquid metal in the reservoir.

6. The method as claimed in claim 1, wherein the gas turbine component is a rotor blade or guide vane, and the first section is the main blade of the rotor blade or guide vane.

7. The method as claimed in claim 6, wherein

the gas turbine component is a rotor blade,

the second section includes the blade root and the blade platform,

the high-mass region and the low-mass region are both part of the blade root, and

the low-mass region lies between the high-mass region and the platform.

8. The method as claimed in claim 7, wherein the high-mass region comprises side surfaces, and the additional liquid metal is supplied to the side surfaces of the high-mass region.

9. The method as claimed in claim 7, wherein the high-mass region comprises an end surface which constitutes the end of the rotor blade facing away from the main blade, and wherein the additional liquid metal is supplied to the end surface.

10. A gas turbine component, comprising:

a first section, and

a second section, which is formed integrally with the first section,

wherein the first section has a directionally solidified crystal structure or a single-crystal structure and the second section has a multidirectionally solidified crystal structure.

11. The gas turbine component as claimed in claim 10, wherein the gas turbine component is formed as a rotor blade or guide vane, wherein the first section is formed by the main blade.

12. A casting mold for casting a gas turbine component by means of the method as claimed in one of claims 1 to 9, the casting mold comprising:

at least a first casting mold section for molding the first section of the gas turbine component,

a second casting mold section for molding the second section of the gas turbine component, and

a supply channel which communicates with the second casting mold section in order to permit supply of additional liquid metal into the second casting mold section.

13. The casting mold as claimed in claim 12, wherein the supply channel comprises a volume for receiving a reservoir of liquid metal.

14. The casting mold as claimed in claim 12, wherein the casting mold has a thermal insulation where the supply channel is located.

15. The casting mold as claimed in claim 12, wherein the second casting mold section comprises a casting mold region with a narrow intermediate space for molding the low-mass region of the gas turbine component and a casting mold region with a wide intermediate space for molding the high-mass region of the gas turbine component, and wherein the supply channel leads to the casting mold region with the wide intermediate space.

16. The casting mold as claimed in claim 12, wherein

the gas turbine component is a rotor blade or guide vane, and

the first casting mold section has a contour for molding the main blade.

17. The casting mold as claimed in claim, wherein

the gas turbine component is a rotor blade,

the second casting mold section includes a blade root casting mold section for molding the blade root and a platform casting mold section for molding the blade platform,

the casting mold region with the narrow intermediate space and the casting mold region with the wide intermediate space are both part of the blade root casting mold section, and

the casting mold region with the narrow intermediate space lies between the casting mold region with the wide intermediate space and the platform casting mold section.

18. The casting mold as claimed in claim 17, wherein the casting mold region with the wide intermediate space comprises casting mold side surfaces, and wherein supply channels lead to the casting mold side surfaces of the casting mold region with the wide intermediate space.

19. The casting mold as claimed in claim 17, wherein the casting mold region with the wide intermediate space comprises a casting mold end surface for defining the end surface of the blade root facing away from the main blade, and the at least one supply channel leads to the casting mold end surface.