Gas turbine, gas turbine blade used therefor and manufacturing method for gas turbine blade

Description

[0001] The invention relates to a gas turbine blade of a heavy-duty gas turbine according to the first part of claim 1, to a method for producing such a gas turbine blade and to the use of such blades in a heavy-duty gas turbine.

[0002] Primarily Ni-base superalloys have hitherto been used as materials for the rotor blades of electricity generating gas turbines. To improve the thermal efficiency of gas turbines, the temperature of gas has been increased year after year. To cope with such an increase in the gas temperature, conventional casting blades having complicated cooling holes therein have been employed.

[0003] Single-crystal wings have already been used as rotor blades of aircraft jet engines. Alloys for casting the single-crystal wing are developed on the assumption that they do not have grain boundaries, and therefore they do not contain grain boundary strengthening elements such as B, Zr and Hf. For this reason, the grain boundaries of single-crystal alloys are weak. At least a portion of a casting must be single-crystallized before the casting can be used. In order to use the single-crystal wing as a gas turbine rotor blade, it is indispensable for the entire casting to be single-crystallized.

[0004] Most single-crystal castings are manufactured by a unidirectional solidification process disclosed in JP-A-51-41851 and JP-A-1-26796. This process is a process in which a casting is withdrawn downwardly from a heated furnace and is solidified gradually from the lower end to the upper end thereof.

[0005] The rotor blade for the aircraft jet engine has a length of approximately 10 cm, and the cross-section area of a shaft is 10 cm² at the largest. The size of a platform extending horizontally from the main body of the rotor blade is small. Because the entire rotor blade is such a small component, a single-crystal wing can be manufactured by solidifying a wing-shaped casting through the above unidirectional solidification process.

[0006] In the document Revue Technique Sulzer, vol. 70, No. 3, 1988, pages 11 to 16; Staub: "An alternative process for the manufacture of single crystal gas turbine blades" there is described the directional and single crystal solidification of a cast turbine blade by a controlled unidirectional heat flow. For creating and maintaining such unidirectional heat flow a heat source, a watercooled sink and a heat reservoir are provided. The heat reservoir is encased in a modified mold and is capable of storing and dispensing the heat capacity necessary for the solidification. By the virtue of its geometrical design and of suitable configuration of insulating materials, the reservoir concentrates the heat content of the gating and feeding system to a controlled heat flow. The monocrystal blades produced by this process have a relatively small length of maximally 100 mm and are tested and used in aeroplane gas turbines.

[0007] However, rotor blades in electricity generating gas turbines are larger than those in aircraft jet engines. The former have a length of 14-16 cm at the shortest or more and shanks having a cross-section area of 15 cm² or more. It is therefore difficult to manufacture the former in a single-crystal structure. There are portions, such as the platform and sealing portions extending from the side of the shank, protruding horizontally from the direction in which the casting is solidified. Even when the casting is solidified by the conventional unidirectional solidification process, the entire casting cannot be single-crystallized. The following reason may be attributed to the non-single crystallization. When the casting is solidified, the horizontally protruding portion begins to solidify from the outer periphery of the casting. Since the horizontally protruding portion has no relationship with the other portion of the casting, it will have crystal orientation different from that of the other portion. When this portion and the other portion of the casting are further solidified and the crystals ot both come into contact with each other, the contacting surface is formed into a grain boundary, thus preventing a single crystal from growing.

[0008] It is thus impossible to form an entire large turbine blade for use in an electricity generating gas turbine in a single-crystal structure.

[0009] The JP-A-63-118037 discloses a Ni-base single-crystal heat-resisting alloy with improved strength at high temperature and structural stability, consisting of, by weight, 5 - 10 % Co, 5 - 10 % Cr, 9 - 15 % W, 1 - 5 % Mo, 4,5 - 6 % Al, 5 - 10 % Ta, balance mainly Ni, (W+Mo+Ta) being 17 - 24 %, which alloy is subjected to unidirectional solidification, thus forming a single crystal, further to solution treatment, rapid cooling and age precipitation treatment. By said constitution, said alloy can allow a raised gas inlet temperature when used for moving blades of a high pressure turbine.

[0010] An object of the invention is to provide a large single-crystal turbine blade excellent in tensile and creep strength and in thermal fatigue performance at heat and stress. Another object of the invention is to provide a manufacturing method for such a turbine blade. A further object is to provide the use of said turbine blade in a heavy-duty gas turbine having high thermal efficiency.

[0011] To achieve the above objects, this invention provides a gas turbine blade comprising the features of claim 1.

[0012] The protrusions provided in the shank of the turbine blade may be sealing portions, in a single stage or multistages, provided on both surfaces along a surface where the wing rotates. The edge of the sealing portion bends towards the wing. The protrusion provided in the shank is one platform provided on both surfaces intersecting with the surface where the wing rotates. The shank, in which the protrusions are provided, has a cross-section area of not less than 15 cm². The shank and the wing including the dovetail and the protrusions are made of the Ni-base alloy in which the γ' phase is precipitated in a single-crystal base of the γ phase. The gas turbine blade has an overall length of not less than 180 mm in a longer direction thereof. The wing weighs not more than 30%, particularly 20-30%, of the overall weight of the gas turbine blade.

[0013] This invention also provides a manufacturing method for a gas turbine blade as claimed in claim 6.

[0014] The invention further provides the use of a gas turbine blade in a heavy-duty gas turbine as claimed in claim 9.

[0015] In order for the gas turbine blade to solidify in one direction, the mold having the by-pass formed in the protrusion is employed separately from the other mold used for the dovetail, the shank and the wing. The manufacturing method for the gas turbine blade, according to this invention, is capable of manufacturing a large gas turbine blade having a complicated configuration and the single-crystal structure.

[0016] Although the turbine blade of the invention is a large blade having the protrusion formed where the cross-section area of the blade is 15 cm² or more, it has more strength than a blade made of a polycrystal having grain boundaries because it is made in the single-crystal structure.

[0017] Desirably, Ni-base alloys should be used for the turbine blade in this invention, each alloy containing by weight 0.15% or less C or preferably 0.02% as an impurity; 0.03% or less Si; more preferably an impurity; 2.0% or less Mn; 5-14% Cr; 1-7% Aℓ; 1-5% Ti; 2.0% or less Nb; 2-15% W; 5% or less Mo; 12% or less Ta, more preferably 2-10%; 10% or less Co; 0.2% or less Hf; 3.0% or less Re; and 0.02% or less B. Table 1 shows the above Ni-base alloys, indicating weight percent of the elements in the alloys.

[0018] Desirably, Co-based alloys may be used in this invention, each alloy containing by weight 0.2-0.6% C; 0.5% or less Si; 2% or less Mn; 20-30% Cr; 20% or less Ni; 5% or less Mo; 2-15% W; 5% or less Nb; 0.5% or less Ti; 0.5% or less Aℓ; 5% or less Fe) 0.02% or less B; 0.5% or less Zr; 5% or less Ta; and the remaining weight percent constitutes Co. Table 2 shows the above Co-based alloys, used for a turbine nozzle serving as a stator blade, indicating weight percent of the elements in the alloys.

[0019] The gas turbine of this invention is more efficient because it is large and permits an operating gas temperature to increase to 1400°C or more at an early stage of the operation.

[0020] Crystal orientation in the horizontally protruding portion with respect to the direction in which solidification advances is oriented so that it may be in the same crystal orientation as the casting. It is thus possible to efficiently manufacture the large single-crystal rotor blade.

[0021] Because the characteristics of the single-crystal rotor blade of the invention are excellent at high temperatures, the service life of the blade is extended,

and the thermal efficiency of the gas turbine caused by an increase in the fuel gas temperature is increased to 34%.

BRIEF DESCRIPTION OF THE DRAWINGS

[0022] 

Fig. 1 is a perspective view of a turbine rotor blade in accordance with an embodiment of the present invention;

Fig. 2 is a vertical cross-sectional view of a mold, illustrating a manufacturing method for the turbine rotor blade shown in Fig. 1;

Fig. 3 is a front view showing a turbine rotor blade of another embodiment of this invention;

Fig. 4 is a vertical cross-sectional view of a mold, illustrating another manufacturing method for the turbine rotor blade shown in Fig. 3;

Fig. 5 is a plan view of the mold shown in Fig. 4;

Fig. 6 is a plan view of a mold in comparison with the mold shown in Fig. 4; and

Fig. 7 is a cross-sectional view showing the rotary portion of a gas turbine in accordance with this invention.

First Embodiment

[0023] Fig. 1 is a perspective view of the rotor blade, according to the present invention, of an electricity generating gas turbine. Fig. 2 is a vertical cross-sectional view showing a manufacturing method for the rotor blade. This method employs a mold of this invention to manufacture the rotor blade.

[0024] As shown in Fig. 2, first, a shell mold 2, made of alumina, is secured to a water-cooled chill 1, and is placed in a mold heating heater 3 in which it is heated to not less than the melting temperature of a Ni-base alloy. Next, a dissolved alloy is poured into the mold 2, and then the water-cooled chill 1 is withdrawn downwardly to solidify the alloy by a unidirectional solidification process. When the alloy is thus solidified, many crystals are first formed in a starter 4 at the lower end of the mold 2, and are then formed into one single crystal in a selector 5, capable of rotating 360°, while the alloy is still being solidified. The single crystal becomes larger in an enlarged section 6. The alloy is solidified and formed into a casting 7, which is composed of a wing 8 having cooling holes formed therein, a shank 9 on the wing 8, and a Christmas tree-shaped dovetail 10 on the shank 9. (These components 8, 9 and 10 are shown upside down in Fig. 1.) Sealing portions or protrusions 11, the end of each bending toward the wing 8, protrude from the shank 9. As shown in Fig. 2, the turbine blade is cast from the wing 8 of the turbine rotor blade to the shank 9 and the dovetail 10 shown in Fig. 1.

[0025] In this embodiment, a by-pass mold 12 different from the casting 7 is provided from the point of section enlargement 6 to the sealing portions or protrusions 11. The provision of the by-pass mold 12 permits the entire rotor blade of the turbine to be single-crystallized. The turbine rotor blade shown in Fig. 1 measures approximately 180 mm high x 40 mm wide x 100 mm long, as denoted by numerals 13, 14 and 15, respectively. The wing 8 measures approximately 90 mm high, and weighs approximately 30% of the weight of the entire turbine rotor blade. The cross-section area of the shank 9, where the sealing portions or protrusions 11 are formed, is 40 cm². The sealing portions 11 each extend approximately 15 mm.

[0026] The casting heater 3 is maintained at high temperatures until the casting 7 is withdrawn and solidified completely. The casting process mentioned above is performed in a vacuum. After the turbine rotor blade made from the single crystal has been cast, it is subjected to a solution heat treatment in a vacuum at temperatures of 1300-1350°C for 2-10 hours. A eutectic γ' phase formed by solidifying the alloy is changed into a γ phase. The turbine rotor blade is then subjected to an aging treatment at temperatures of 980-1080°C for 4-15 hours and at temperatures of 800-900°c for 10-25 hours. Horn-shaped γ' phases, each having an average size of 3-5 µm, are precipitated in the γ phase.

[0027] Table 3 shows conditions for casting the single-crystal wing.

Table 3

Mold heating temperature	1560°C
Pouring temperature	1550°C
Withdrawal velocity	10 cm/h
Mold material	ceramic
Degree of vacuum	2.7 x 10-3 h Pa (2 x 10-3 Torr)or less
Alloys	Nos. 2 and 10

Table 4 shows the comparison between the yield of single-crystal wings manufactured by the method of this invention and the yield of such wings manufactured by the conventional method.

Table 4

	Yields
Alloys	This invention	Conventional method
No. 2	75%	0%
No. 10	83%	0%

[0028] The turbine rotor blade is shrunk at the upper portion of a platform, and the secondary growth of a long, thin dendrite is found at the lower portion of the platform.

[0029] As shown in Table 4, this invention makes it possible to manufacture a large single-crystal wing which cannot be manufactured by the conventional method. In this embodiment, since the wing of the turbine rotor blade, which requires the highest strength and ductility, is first solidified, the time during which the rotor blade is in contact with the molten mold is shortened. It is possible to obtain a turbine rotor blade made of an alloy containing elements which vary little and have few defects; as a result, a turbine rotor blade having the required characteristics can be manufactured. It takes approximately one hour for the wing to solidify, and approximately two hours for the other components and the dovetail to solidify finally. The elements in an alloy vary, and particularly Cr varies greatly. As described in this embodiment, however, if a large amount of Cr, 8.5 wt% and particularly 10 wt% or more, is contained in an alloy, it varies little and is very effecive in being used for turbine rotor blades. On the contrary, 8.5 wt% or less Cr varies greatly.

[0030] The by-pass mold 12, different from the mold used for forming the turbine rotor blade, may be provided in a position which is above the selector 5 in a selector method or above a seed in a seed method, and which is anywhere below the sealing portions or protrusions 11. However, after the single-crystal has been cast, the by-pass mold 12 must be removed; therefore desirably, the by-pass mold 12 should be provided in the enlarged section 6, shown in Fig. 2, which is above the selector 5 or the seed and is below the wing 8.

[0031] The rotor blade is solidified from the wing 8 to the dovetail 10 for the following reasons. The wing 8 of the gas turbine rotor blade is the essential part of the rotor blade, and is subjected to high temperatures and stress. It therefore must possess fewer defects and be of a higher-quality than any other components. The wing 8 is first solidified, so that the time during which it is held at high temperatures is shortened. In order to make the elements vary little, such casting is suitable for manufacturing the rotor blade of the gas turbine. A plurality of cooling holes are provided leading from the wing 8 to the dovetail 10, and are used for cooling the components by a refrigerant. A core for the cooling holes is used as the mold. The speed at which the alloy is solidified varies from 1 to 50 cm/h according to the size of the casting to be solidified. The wing 8 can be solidified faster than the shank 9 and the dovetail 10.

[0032] Although the manufacturing method for the rotor blade of a gas turbine has been described, it is possible to allow a single crystal to grow for stator blades by the same method as described above.

Second Embodiment

[0033] A rotor blade having substantially the same configuration as that of the rotor blade in the first embodiment is cast using the alloy of No. 2. The same casting conditions and the unidirectional solidification process as those in the first embodiment are employed in the second embodiment. The blade measures 160 mm high; a wing measures 70 mm high; and a shank and a dovetail each measure 90 mm high.

[0034] Fig. 3 shows the front view of this rotor blade. Because the rotor blade has a wide platform 17, when it is solidified by the unidirectional solidification process, a new crystal is formed at the platform 17, thus preventing a single crystal from growing. To solve this problem, the present invention is applied to the method of manufacturing the rotor blade. As shown in Fig. 4, a portion near the edge of the platform 17 is connected to a portion above a selector 5 by means of a by-pass mold 12, different from the mold for forming a casting 7. Such connection makes it possible for a single crystal to grow. The by-pass mold 12 has a thickness of 1 mm and a width of 20 mm. Fig. 4 shows the cross-sectional configuration of the rotor blade; Fig. 6 shows how the new crystal grows in the conventional method, as seen from the upper portion of the wing 8; and Fig. 5 shows how the new crystal does not grow in this invention, as seen also from the upper portion of the wing 8. In Fig. 6 numeral 18 denotes a grain boundary, and numeral 19 denotes the new crystal. This invention makes it possible for the single crystal to grow, instead of a new crystal growing.

Third Embodiment

[0035] Fig. 7 is a partial cross-sectional view showing the rotary portion of a gas turbine. In the drawing, the Ni-base alloy of No. 2 made of the single crystal, obtained in the first embodiment of this invention, is used for first turbine blades 20a. In this embodiment, the gas turbine has two turbine disks 21a and 21b. The first turbine disk 21a having a central hole 22 therein is disposed upstream of a gas flow, and the second turbine disk 21b, is disposed downstream of the gas flow. A martensitic heat resisting steel containing 12% Cr is used for the final stage compressor disk 23, a distant piece 24, a turbine spacer 25, a turbine stacking bolt 26 and a compressor stacking bolt 27. The turbine blades 20b in the second turbine disk 21b, a turbine nozzle 28b, a liner 30 of a combustor 29, compressor blades 31, a compressor nozzle 32, a diaphragm 33 and a shroud 34 are made of alloys. The elements contained in these alloys are shown in Table 5. The turbine nozzle 28 in a first stage and the turbine blade 20 are made of a single-crystal casting.

[0036] The turbine nozzle 28a in the first stage is made of an alloy of No. 13, and is composed of one segment for each wing in the same manner as in the turbine blade. The turbine nozzle 28a is disposed on a circumference, and has a diaphragm and a length which is substantially equal to the wing of the blade. Numeral 35 denotes a turbine stab shaft, and numeral 36 denotes a compressor stab shaft. A compressor used in this embodiment has 17 stages. The turbine blades 20a, the turbine nozzle 28a, a shroud segment 34a (1) and the diaphragm, all shown in Table 5, are used in the first stage upstream of the gas flow, whereas a shroud segment 34b (2) is used in the second stage.

[0037] In this embodiment, a layer made of a highly-concentrated alloy containing Aℓ, Cr and other elements, or made of a mixture containing oxides, may be used as a coating layer which is resistant to oxidation and corrosion at temperatures higher than those at which an alloy serving as a base material is resistant to oxidation and corrosion.

[0038] The crystal may be formed so that its orientation becomes (001) in the direction in which a centrifugal force is applied. A blade having high strength is obtainable by forming the crystal in this way.

[0039] According to the gas turbine thus constructed, when electricity on the order of 50 Mw is generated, the gas temperature at the entrance of the turbine nozzle in the first stage is capable of rising as high as 1500°C, and the metal temperature at the blade in the first stage is capable of rising as high as 1000°C. Thirty four percent thermal efficiency is obtainable. As mentioned above, the heat resisting steel having higher creep rupture strength and fewer defects caused by heat is used for the turbine disk, the distant piece, the spacer, the final stage of the compressor disk, and the stacking bolt. The alloy having strength at high temperatures is used for the turbine blade; the alloy having strength and ductility at high temperatures is used for the turbine nozzle; and the alloy having high fatigue performance and strength at high temperatures is used for the liner of the combustor. It is thus possible to obtain a gas turbine which is more reliable in various aspects than the conventional art.

Claims

1. Gasturbine blade of a Ni- or Co-based alloy for heavy-duty gas turbines, comprising

- a dovetail (10) for securing in a turbine disc (21),

- a shank (9) integrally formed to said dovetail (10) and provided with one or more integrally formed sealing protrusions (11) on at least one side surface and with an integrally formed platform (17) on its upper side,

- a wing (8) integrally formed on the platform (17) of said shank (9) and provided with cooling holes leading to the dovetail (10),

characterized in that

- in the whole turbine blade the γ phase of the alloy is in a single-crystal structure and

- the crystal orientation in the sealing protrusions (11) and in the platform (17) has the same direction as the crystal orientation in the other parts (8, 9, 10) of the turbine blade.

2. Gasturbine blade according to claim 1,
characterized in that the sealing protrusions (11) are provided on both sides of said shank (9) along the rotating surface of the turbine blade.

3. Gasturbine blade according to claim 2,
characterized in that
the edges of said sealing protrusions (11) are bent toward the wing (8).

4. Gasturbine blade according to one of the claims 1 to 3,
characterized in that
the cross-section area of the shank (9) in which the sealing protrusions (11) are provided is not less than 15 cm².

5. Gasturbine blade according to one of the claims 1 to 4,
characterized in that
the overall height of the turbine blade is 160 mm or not less than 180 mm.

6. Method for producing a gasturbine blade according to one of the claims 1 to 5,
   characterized by the steps of

- connecting a by-pass mold (12) for forming the protrusions (11) and the platform (17) with the shank forming part of a main mold (2) and with a mold portion (6) below the wing forming part of the main mold (2),

- pouring a dissolved Ni- or Co-based alloy melt simultaneously into said main mold (2) and into said by-pass mold (12), said molds (2, 12) are heated by a mold heater (3) to a temperature not less than the melting temperature of the alloy,

- unidirectionally solidifying the alloy in the main mold (2) and in the by-pass mold (12) by out-drawing the molds (2, 12) from the heater (3) with a predetermined speed for generating a single-crystal structure of the whole cast turbine blade,

- heat-treating the cast turbine blade for precipitating the eutectic γ' phase formed by the solidifying of the alloy into the γ phase and

- removing the by-pass mold (12) after casting.

7. Method according to claim 6,
characterized in that
the wing (8) of the turbine blade has been solidified first.

8. Method according to claim 6 or 7,
characterized by

a solution heat treatment of the cast turbine blade in a vacuum at temperatures of 1300 - 1350 °C for 2 to 10 hours, and

an aging heat treatment of the turbine blade at temperatures of 980 - 1080 °C for 4 to 15 hours and at temperatures of 800 to 900 °C for 10 to 25 hours.

9. Use of a gas turbine blade according to any of claims 1 to 5 in a heavy-duty gas turbine, having the features of

- a multistage compressor (31, 32), a distant piece (24), a combustion liner (29, 30), and a gasturbine with at least one nozzle diaphragm (28, 33) and at least one turbine disk (21a),

- the blades (20a, 28) of at least one turbine disk (20a) and of at least one nozzle diaphragm (33) are made of a Ni- or Co-based alloy

- the final stage compressor disk (23) and the distant piece (24) are made of a martensitic heat resisting steel, the liner (29) of the combustor (30) is made of an alloy of high fatigue performance and strength at high temperatures,

- said gas turbine is operated at a temperature which is not less than 1400 °C and the temperature at the turbine blades is not less than 1000 °C under working stress.

Ansprüche

1. Gasturbinenschaufel aus einer Legierung auf Ni- oder Co-Basis für Hochleistungs-Gasturbinen, mit

- einem Schwalbenschwanzteil (10) zur Befestigung in einem Turbinenrad (21),

- einem mit dem Schwalbenschwanzteil (10) einteilig ausgebildeten Schaftteil (9), das einen oder mehrere angeformte Dichtungsansätze (11) und an seiner Oberseite eine angeformte Plattform (17) aufweist,

- einem auf der Plattform (17) des Schaftteils (9) angeformten Flügel (8), der mit zum Schwalbenschwanzteil (10) führenden Kühlkanälen versehen ist,

dadurch gekennzeichnet , daß

- in der gesamten Turbinenschaufel die γ-Phase der Legierung in einer Einkristall-Struktur vorliegt und

- die Kristallorientierung in den Dichtungsansätzen (11) und in der Plattform (17) die gleiche Richtung wie die Kristallorientierung in den anderen Teilen (8, 9, 10) der Turbinenschaufel hat.

2. Gasturbinenschaufel nach Anspruch 1,
dadurch gekennzeichnet, daß die Dichtungsansätze (11) an beiden Seiten des Schaftteils (9) längs der Drehfläche der Turbinenschaufel vorgesehen sind.

3. Gasturbinenschaufel nach Anspruch 2,
dadurch gekennzeichnet, daß die Ränder der Dichtungsansätze (11) gegen den Flügel (8) gebogen sind.

4. Gasturbinenschaufel nach einem der Ansprüche 1 bis 3,
dadurch gekennzeichnet, daß der mit den Dichtungsansätzen (11) versehene Querschnittsbereich des Schaftteils (9) nicht kleiner als 15 cm² ist.

5. Gasturbinenschaufel nach einem der Ansprüche 1 bis 4,
dadurch gekennzeichnet, daß die Gesamthöhe der Turbinenschaufel 160 mm oder nicht weniger als 180 mm beträgt.

6. Verfahren zum Herstellen einer Gasturbinenschaufel nach einem der Ansprüche 1 bis 5,
gekennzeichnet durch die Schritte

- Verbinden einer Bypass-Form (12) zum Formen der Dichtungsansätze (11) und der Plattform (17) mit dem Schaftformteil einer Grundform (2) und mit einem Formteil (6) unter dem Flügelformteil der Grundform (2),

- Gießen einer gelösten auf Ni oder Co basierenden Legierungsschmelze gleichzeitig in die Grundform (2) und in die Bypass-Form (12), wobei die Formen (2 und 12) von einer Formheizung (3) auf eine Temperatur nicht unter der Schmelztemperatur der Legierung beheizt werden,

- unidirektionales Verfestigen der Legierung in der Grundform (2) und in der Bypass-Form (12) durch Herausziehen der Formen (2, 12) aus der Heizung mit einer vorgegebenen Geschwindigkeit zur Erzeugung einer Mono-Kristall-Struktur der gesamten Turbinenschaufel,

- einer Wärmebehandlung der Turbinenschaufel zum Ausfällen der bei der Verfestigung der Legierung gebildeten eutektischen γ'-Phase in die γ-Phase und

- Entfernen der Bypass-Form (12) nach dem Gießvorgang.

7. Verfahren nach Anspruch 6,
   dadurch gekennzeichnet, daß der Flügel (8) der Turbinenschaufel zuerst verfestigt wird.

8. Verfahren nach Anspruch 6 oder 7, gekennzeichnet durch eine Lösungswärmebehandlung des Turbinenschaufel-Gußteils von 2 bis 10 h im Vakuum bei Temperaturen von 1300 - 1350 °C und eine Alterungswärmebehandlung der Turbinenschaufel von 4 bis 15 h bei Temperaturen von 980 - 1080 °C und von 10 bis 25 h bei Temperaturen von 800 bis 900 °C.

9. Verwendung einer Gasturbinenschaufel nach einem der Ansprüche 1 bis 5 in einer Hochleistungsgasturbine mit den Merkmalen

- einem mehrstufigen Kompressor (31, 32), einem Abstandshalter (24), einer Brennerauskleidung (29; 30) und einer Gasturbine mit mindestens einem Leitrad (28, 33) und mindestens einem Turbinenrad (21a),

- die Schaufeln (20a, 28) des zumindest einen Turbinenrads (20a) und des zumindest einen Leitrads (33) bestehend aus einer Legierung auf Ni- oder Co-Basis,

- das endstufige Kompressorrad (23) und der Abstandshalter (24) bestehen aus einem warmfesten Martensit-Stahl, die Auskleidung (29) der Brenner (30) bestehen aus einer Legierung von hoher Dauerwechselfestigkeit und Zähigkeit bei hohen Temperaturen,

- wobei die Turbine bei einer Temperatur von nicht kleiner als 1400 °C betrieben wird und die Temperatur an den Turbinenschaufeln bei Betriebsbeanspruchungen nicht unter 1000 °C liegt.

Revendications

1. Aube de turbine à gaz réalisée en un alliage à base de Ni ou de Co pour des turbines à gaz à forte capacité, comprenant :

une queue d'aronde (10) pour la fixation dans un disque de turbine (21),

une tige (9) formée d'une seule pièce avec ladite queue d'aronde (10) et comportant une ou plusieurs saillies d'étanchéité formées d'une seule pièce (11) sur au moins une surface latérale, et une plate-forme formée d'une seule pièce (17) sur sa face supérieure,

une aile (8) formée d'une seule pièce sur la plate-forme (17) de ladite tige (9) et comportant des trous de refroidissement menant à la queue d'aronde (10),

   caractérisée en ce que :

dans toute l'aube de turbine, la phase γ de l'alliage est une structure mono-cristalline, et

l'orientation cristalline dans les saillies d'étanchéité (11) et dans la plate-forme (17) a la même direction que l'orientation cristalline dans les autres parties (8, 9, 10) de l'aube de turbine.

2. Aube de turbine à gaz selon la revendication 1, caractérisée en ce que les saillies d'étanchéité (11) sont disposées sur les deux côtés de ladite tige (9) le long de la surface de rotation de l'aube de turbine.

3. Aube de turbine à gaz selon la revendication 2, caractérisée en que les bords desdites saillies d'étanchéité (11) sont incurvés vers l'aile (8).

4. Aube de turbine à gaz selon l'une des revendications 1 à 3, caractérisée en ce que la surface de section transversale de la tige (9) dans laquelle sont disposées les saillies d'étanchéité (11) n'est pas inférieure à 15 cm².

5. Aube de turbine à gaz selon l'une des revendications 1 à 4, caractérisée en ce que la hauteur totale de l'aube de turbine est de 160 mm ou n'est pas inférieure à 180 mm.

6. Procédé pour produire une aube de turbine à gaz selon l'une des revendications 1 à 5, caractérisée par les étapes suivantes :

la connexion d'un moule de dérivation (12) pour former les saillies (11) et la plate-forme (17) avec la tige faisant partie d'un moule principal (2) et avec une partie de moule (6) en-dessous de l'aile faisant partie du moule principal (2),

le versage d'un bain de fusion d'alliage à base de Ni ou de Co dissous simultanément dans ledit moule principal (2) et dans ledit moule de dérivation (12), lesdits moules (2, 12) étant chauffés par un dispositif de chauffage de moule (3) à une température qui n'est pas inférieure à la température de fusion de l'alliage,

la solidification unidirectionnelle de l'alliage dans le moule principal (2) et dans le moule de dérivation (12) en retirant les moules (2, 12) de l'élément chauffant (3) à une vitesse prédéterminée pour générer une structure mono-cristalline de la totalité de l'aube de turbine moulée,

le traitement à chaud de l'aube de turbine moulée pour précipiter la phase eutectique γ' formée par la solidification de l'alliage sous la forme de la phase γ, et

le retrait du moule de dérivation (12) après le moulage.

7. Procédé selon la revendication 6, caractérisé en ce que l'aile (8) de l'aube de turbine a été solidifiée en premier.

8. Procédé selon la revendication 6 ou 7, caractérisé par :

un traitement thermique en solution de l'aube de turbine moulée dans le vide à des températures comprises entre 1.300 et 1.350°C pendant 2 à 10 heures, et

un traitement thermique de vieillissement de l'aube de turbine à des températures comprises entre 980 et 1.080°C pendant 4 à 15 heures et à des températures comprises entre 800 et 900°C pendant 10 à 25 heures.

9. Utilisation d'une aube de turbine à gaz selon l'une quelconque des revendications 1 à 5 dans une turbine à gaz de forte capacité, ayant les caractéristiques suivantes :

un compresseur à étages multiples (31, 32), une pièce distante (24), une chemise de combustion (29, 30), et une turbine à gaz avec au moins un distributeur annulaire (28, 33) et au moins un disque de turbine (21a),

les aubes (20a, 28) d'au moins un disque de turbine (20a) et d'au moins un distributeur annulaire (33) étant faites en un alliage à base de Ni ou de Co,

le disque de compresseur d'étage final (23) et la pièce distante (24) étant faits en un acier martensitique résistant à la chaleur, la chemise (29) de la chambre de combustion (30) étant faite en un alliage ayant des hautes performances vis-à-vis de la fatigue et une haute résistance aux températures élevées,

ladite turbine à gaz fonctionnant à une température qui n'est pas inférieure à 1.400°C, et la température des aubes de turbine n'étant pas inférieure à 1.000°C sous des contraintes de travail.

Drawing