CN1034828C - Gas turbine blade, method of manufacture and use thereof - Google Patents

Abstract

A gas turbine blade comprising a dovetail for fixing to a disc; a root connected to said dovetail and forming one or more integral protrusions on the side of the dovetail; a root connected to said dovetail The airfoil at the root is characterized in that the gas turbine blade is made of a nickel-based alloy in which the γ' phase is substantially condensed in the formed γ phase of a single crystal structure. Has good high temperature performance. A method for manufacturing the gas turbine blade and its use for a gas turbine are also disclosed.

Description

Gas turbine blade, method of manufacture and use thereof

The invention relates to a gas turbine blade, its manufacturing method and use.

Early nickel-based heat-resistant alloys are still used as materials for power-generating gas turbine rotor blades. In order to improve the thermal efficiency of the turbine, the gas temperature is increased year by year.

To cope with the increase in gas temperature, conventional cast blades with complex cooling holes have been used.

Single crystal blades have been used as rotor blades for aircraft jet engines. Alloys for cast single crystal airfoils were developed on the assumption that the alloys have no grain boundaries and therefore do not contain grain boundary strengthening elements such as B, Zr and Hf. Therefore, the grain boundaries of single crystal alloys are insufficient. At least a portion of the casting must be monocrystalline before the casting is used. In order to use such monocrystalline airfoils as rotor blades for gas turbines, a casting that is entirely monocrystalline is essential. Most single crystal castings are produced by the directional solidification method already disclosed in Japanese Patent Nos. 51-4185551 and 51-26796. This method is to pull the casting down from the heated furnace body and gradually solidify from the bottom to the top from the bottom.

The rotor blade used by aircraft jet engine is about 10 centimeters long, and the maximum area of the cross-section of the shaft is about 10 centimeters ² , and the size of the platform extending horizontally from the blade main body is very small. Because the size of the entire rotor blade is very small, the single-crystal airfoil can be produced by solidifying the airfoil-shaped casting by using the above-mentioned directional solidification method.

However, the rotor blades in gas turbines for power generation are larger than those in aircraft jet engines. The shortest rotor blades of gas turbines for power generation are 14-16 cm, and the root cross-sectional area is at least 15 cm ² . Therefore, it is difficult to manufacture rotor blades for power generating gas turbines with a single crystal structure. Some parts, such as the boss and the sealing part projecting from the side of the root, project horizontally from the direction of solidification of the casting. Even when such castings are solidified by conventional directional solidification methods, it is impossible for the entire casting to be completely single crystal. The reason for the non-single crystal is as follows. As the casting solidifies, the horizontal bulge begins to solidify from the outer edge of the casting. Since the horizontal bulge is not related to the rest of the casting, its crystallization direction is different from that of the rest of the casting. When this part and other parts of the casting are further solidified and the crystals of the two contact each other, a grain boundary is formed at the interface, which prevents the growth of single crystals.

Thus, it is not possible to form a complete large turbine blade in a monocrystalline structure for use in a gas turbine for power generation.

It is an object of the present invention to provide a large single crystal turbine blade which exhibits excellent tensile and creep strength and thermal fatigue properties under thermal and stress conditions. Another object of the present invention is to provide a method of manufacturing such a turbine blade. A further object of the present invention is to provide a heavy duty gas turbine with high thermal efficiency.

In order to achieve the above object, the present invention provides a gas turbine blade, which includes: a dovetail for fastening on the wheel disk; a root, which is connected with the dovetail and has one or more integrally formed on the dovetail A protrusion on the side; an airfoil attached to the root; where the gas turbine blade is made of a nickel-based alloy whose r' phase is precipitated in the r phase forming a single crystal structure.

The protruding portion at the root of the turbine blade can be a single or multi-stage sealing portion, which is located on both faces along the rotating surface of the blade. The edge of the sealing portion is bent towards the airfoil. On the protruding portion at the root is a boss on the two faces intersecting the rotating face of the blade. Roots with protruding parts have a cross-sectional area of not less than 15 cm ² . The root and airfoil, including the dovetail and protrusions, are made of a nickel-based alloy in which the r' phase is precipitated in a single crystal matrix of the r-phase. The overall length of the gas turbine blade in its length direction is not less than 180mm. The weight of the airfoil is not more than 30% of the total weight of the gas turbine blade, especially between 20-30%.

The present invention also provides a method of manufacturing a gas turbine blade; the blade comprising a dovetail fastened to a disk, a root connected to the dovetail and having a protrusion integrally formed on the side of the dovetail; A blade airfoil connected to the root, the manufacturing method comprising the steps of: connecting a bypass mold corresponding to the protruding part to a main mold corresponding to the dovetail, the root, and the blade airfoil; The molten nickel-based alloy in the main and bypass molds gradually solidifies and casts a single crystal structure.

The present invention further provides a gas turbine blade, which includes: a dovetail as a fastening wheel; a root connected to the dovetail and integrally formed with one or more protrusions on the side of the dovetail; The airfoil at the root; where the gas turbine blade is unidirectionally solidified from the airfoil tip to the dovetail, made of a single crystal nickel-based alloy in r-phase.

The present invention provides a heavy-duty gas turbine, which includes: a compressor; a combustion sleeve; a single-stage or multi-stage turbine blade, which has a dovetail fastened on the turbine disk, and its total length is not less than 180 mm, and it is made of single crystal Made of nickel-based alloy, its r-phase is a single crystal; the turbine nozzle is set corresponding to the turbine blade; the working temperature of the gas is not lower than 1400°C, and the metal temperature of the first-stage blade is not lower than 1000°C under the working stress state .

To solidify the gas turbine blade in one direction, a die with bypass formed in the bulge is used, separate from the other dies for the dovetail, root and airfoil. According to the method of manufacturing a gas turbine blade of the present invention, a large gas turbine blade having a complex shape and a single crystal structure can be manufactured.

Although the turbine blade of the present invention is a large blade having the shape of a protruding part, and the cross-sectional area of the blade at the protruding part is set to be 15 ^mm2 or more, it is stronger than a blade with multiple crystal grains having multiple grain boundaries, Because it is a single crystal structure.

Preferably adopt nickel base alloy as turbine blade of the present invention, the weight composition of every kind of alloy is: carbon C is equal to or less than 0.15%, is preferably 0.02% as impurity; Silicon Si is equal to or less than 0.03%; Preferably as a kind of Impurities: Manganese Mn is equal to or less than 2.0%; Chromium Cr is equal to 5-14%; Aluminum Al is equal to 1-7%; Titanium Ti is equal to 1-5%; Niobium Nb is equal to or less than 2.0%; Tungsten W is equal to 2-15%; Molybdenum Mo is equal to or less than 5%; Tantalum Ta is equal to or less than 12%, preferably 2-10%; Cobalt Co is equal to or less than 10%; Hafnium Hf is equal to or less than 0.2%; Rhenium Re is equal to or less than 3.0%; Boron B Equal to or less than 0.02%. Table 1 shows the above-mentioned nickel-based alloys, and the percentages in the alloys are by weight.

It is hoped that cobalt-based alloys can be used in the present invention, and the weight ratio contained in each alloy is: carbon C=0.2-0.6%; silicon Si≤0.5%; manganese Mn≤2%; chromium Cr=20-30%; nickel Ni≤20% %; molybdenum Mo≤5%; tungsten W=2-15%; niobium Nb≤5%; titanium Ti≤0.5%; aluminum Al≤0.5%; iron Fe≤5%; %; tantalum Ta≤5%; the rest is cobalt Co. Table 2 presents cobalt-based alloys for turbine nozzles for stator blades, giving the weight percentages of the alloying elements.

The gas turbine of the present invention has higher efficiency because it is large and allows gas operating temperatures up to 1400°C or more during the initial stages of operation.

In the horizontal bulges, the crystallographic direction corresponds to the direction of solidification advance, so the same crystallographic orientation is possible in the casting. Therefore, it is possible to manufacture large single crystal rotor blades efficiently.

Due to the excellent high temperature characteristics of the single crystal rotor blade of the present invention, the service life of the blade is extended, and the thermal efficiency of the gas turbine is increased to 34% due to the increase of the gas temperature.

Figure 1 is a perspective view of a turbine rotor blade according to one embodiment of the present invention;

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

Fig. 3 is a front view of a rotor blade according to another embodiment of the present invention;

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

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

Figure 6 is a plan view of a mold compared with the mold shown in Figure 4;

Fig. 7 shows a cross-sectional view of the rotor part of the gas turbine of the present invention.

First example:

Figure 1 is a perspective view of a rotor blade of a power generating gas turbine according to the present invention. Fig. 2 is a vertical cross-sectional view showing a method of manufacturing a rotor blade for a power generating gas turbine of the present invention, which method uses a mold of the present invention to manufacture the rotor blade.

As shown in Figure 2, first a shell mold 2 made of alumina is fixed on a cold iron 1 cooled by water, and it is placed on the shell mold heater 3 for heating, and it is heated to not low temperature inside the heater. at the melting temperature of nickel-based alloys. Then see that the melted alloy is poured into the shell mold 2, and then the water-cooled cold iron 1 is pulled down to solidify it with the directional solidification method. In this way, when the alloy is solidified, many crystal grains are first formed in the starter 4 at the lower part of the shell mold 2, and then a single crystal is formed in the selector 5, which can turn 360° while the alloy is still solidifying. The single crystal becomes larger at the enlarged section 6 . The alloy solidifies and forms a casting 7 comprising an airfoil 8 with cooling holes, a root 9 on the airfoil 8 and a dovetail 10 in the shape of a Christmas tree on the root 9 . (The three parts 8 , 9 , 10 are shown from top to bottom in FIG. 1 ) A sealing portion or projection 11 , the end of which is bent towards the airfoil 8 , protrudes from the dovetail 10 . As shown in Figure 2, the turbine blade is cast from the airfoil 8 of the turbine rotor blade to the root 9 and dovetail, as shown in Figure 1.

In this embodiment, a bypass die 12 different from the casting 7 is arranged from the enlarged section 6 to the sealing or projection 11 . The configured bypass mold 12 makes the entire turbine rotor blade a single crystal. The height × width × length of the turbine blade shown in Fig. 1 is 180 millimeters × 40 millimeters × 100 millimeters, represented by serial number 13, 14 and 15 respectively among the figure. The airfoil 8 is about 90 mm high, and its weight accounts for about 30% of the weight of the entire turbine rotor blade. The cross-sectional area of the root 9 forming the sealing part or the protruding part ¹¹ is 40 cm2, and each end of the sealing part 11 protrudes about 15 mm. .

The casting heater 3 keeps high temperature until the casting 7 is pulled out completely and solidified completely, and the above-mentioned casting process is completed in a vacuum. After the turbine rotor blade made of single crystal is cast, the next step is to keep it in vacuum at a temperature of 1300-1350°C for 2-10 hours for solution heat treatment. The r' phase of the eutectic formed by the solidified alloy dissolves into the r phase. Then, the temperature is kept at 980-1080°C for 4-15 hours, and the temperature is 800-900°C for 10-25 hours, and the aging treatment of the turbine rotor blades is carried out. Angular r' phases with an average size of 3-5 microns are precipitated in the r phase.

Table 3 gives the conditions for casting single crystal airfoils.

Table 4 shows the comparison between the single crystal airfoil manufactured by the method of the present invention and the airfoil manufactured by the general method.

The turbine rotor blade shrinks at the upper part of the boss, and secondary elongated dendritic crystals grow at the lower part of the boss.

As shown in Table 2, the present invention enables the casting of large single-crystal airfoils that cannot be produced by conventional methods. In this embodiment, since the airfoil of the turbine rotor blade, which requires high strength and toughness, solidifies first, the contact time with the mold is shortened. In this way, a turbine rotor blade with little change in alloying elements and almost no defects can be obtained, so that a turbine rotor blade with desired characteristics can be manufactured. The leaf body takes about one hour to set, and the other parts and dovetails take two hours to set. However, there are changes in the elements in the alloy, especially the change of chromium Cr is very large. As described in this example, if the content of chromium Cr in the alloy is as large as 8.5%, especially to 10% or more, it changes little and is very effective for the turbine rotor blade. On the contrary, if the Cr content is less than or equal to 8.5%, it varies greatly.

The bypass mold 12, which is different from the shell mold for forming turbine rotor blades, can be arranged above the selector 5 when using the selector method, and can be arranged above the seed crystal when using the seed crystal method, but they are always located at Any position below the sealing portion or the protruding portion 11. However, after the single crystal is cast, the bypass mold 12 must be removed. According to the requirements, the bypass mold should be set at the enlarged section 6 shown in Figure 2, that is, above the selector or seed crystal, at the blade body 8 the following.

The rotor blade is solidified from the airfoil 8 to the dovetail for the following reasons. The airfoil 8 of the gas turbine rotor blade is the basic part of the rotor blade, which is subject to high temperature and high stress. Therefore, it must have few defects and be of higher quality than other parts. The airfoil 8 is solidified first, so its time in the high temperature zone is shorter. Such castings are suitable for the manufacture of gas turbine rotor blades due to the low elemental variation. A set of cooling holes leads from the airfoil 8 to the dovetail 10 for cooling the parts with coolant. A core with cooling holes is used as the mould. The rate at which the alloy solidifies varies from 1 to 50 cm/hour, depending on the size of the casting being solidified. The airfoil 8 solidifies faster than the root 9 and the dovetail 10 .

Although the method of manufacturing a gas turbine rotor blade has been described, the growth of a single crystal of a stator blade can also be performed using the same method as described above.

Second embodiment. A rotor blade having substantially the same shape as the rotor blade of the first embodiment was cast from Alloy No. 2. The same casting conditions and directional solidification method as in the first embodiment were used in the second embodiment. The blade height is 160 mm, the blade height is 70 mm, and the root and dovetail height are 90 mm.

Figure 3 shows a front view of such a blade. Since the rotor blade has a wide platform 17, when it is solidified by directional solidification, a new crystal is formed on the platform 17, thus preventing the growth of single crystals. In order to solve this problem, the present invention is applied to a method of manufacturing a rotor blade. As shown in FIG. 4 , the portion near the edge of the platform 17 is connected to the upper part of the selector 5 through a bypass die 12 which is different from the die from which the casting 7 is formed. Such connections allow single crystals to grow. The bypass mold 12 has a thickness of 1 millimeter and a width of 20 millimeters. Fig. 4 shows the shape of the cross-section of the rotor blade. Figure 5 shows how new crystals grow in a conventional method, as seen from the upper part of the airfoil 8. Fig. 6 shows how the new crystals of the inventive method do not grow, which can also be seen from the blade body 8 top. In Fig. 6, numeral 18 denotes a grain boundary, and numeral 19 denotes a new crystal. The present invention makes it possible to replace new crystal growth with single crystal growth.

third embodiment.

Fig. 7 is a cross-sectional view of the rotating part of a gas turbine. In the figure, the nickel-based No. 2 alloy made of single crystal obtained by the method of the first embodiment of the present invention is used on the first-stage turbine blade 20 . In this embodiment, the turbine wheel 21 has two stages. The first stage is arranged upstream of the gas flow and the second stage having a center 22 is arranged downstream of the gas flow. A martensitic heat-resistant steel containing 12% chromium Cr is used on the last stage disc 23 of the compressor, the spacer 24 , the turbine diaphragm 25 , the turbine combination bolt 26 and the compressor combination bolt 27 . Second stage turbine blades 20, turbine nozzles 28, combustor 29 sleeve 30, compressor blades 31, compressor nozzles 32, baffles 33 and shrouds 34 are all made of alloy. The elements contained in these alloys are shown in Table 5. The first stage turbine nozzle 28 and turbine blades 20 are single crystal cast.

The first stage turbine nozzle 28 is fabricated from Alloy 13 and includes airfoil sections of the same type as turbine blades. The turbine nozzle 28 is mounted on a circumference and has a partition having a length substantially equal to the airfoil of the blade. The serial number 35 represents the turbine main shaft, and the serial number 36 represents the compressor main shaft. The compressor in the present embodiment has 17 stages in total, and the turbine blades shown in Table 5, the turbine nozzle, the shroud (1) and the baffle are all used in the upstream of the gas flow of the first stage, and the shroud (2) is used for the second level.

In this embodiment, a coating made of a highly concentrated alloy including aluminum Al, chromium Cr and other elements, or a mixture containing oxides, can be used to resist high temperature oxidation and corrosion, which is more resistant to high temperature oxidation and corrosion than a coating made of a Alloys of base materials resist oxidation and corrosion at higher temperatures.

A crystal can be shaped so that its orientation can become [001] in the direction of centrifugal force. With crystals formed in this way, high-strength blades can be obtained.

According to the structure of this gas turbine, when the power generation is about 50MW, the gas temperature at the inlet of the first-stage nozzle can reach 1500°C, the metal temperature of the first-stage blade can reach 1000°C, and the thermal efficiency can reach 34%. As mentioned above, heat-resistant steels with high creep rupture strength and few defects caused by heat can be suitable for manufacturing turbine disks, separators, diaphragms, last stages of compressor disks, composite bolts. Alloys with high temperature strength are used in the manufacture of turbine blades. Alloy with high temperature strength and ductility, used in the manufacture of turbine nozzles. An alloy with high fatigue properties and high temperature strength, used in the manufacture of combustor sleeves. In this way it is possible to obtain a gas turbine which is in all respects more reliable than the prior art.

Table 1 Chromium Cr Molybdenum Mo Tungsten W Rhenium Re Aluminum Al Titanium Ti Tantalum Ta Cobalt Co Hafnium Hf Niobium Nb Nickel Ni 1 10.0 - 4.0 - 5.0 1.5 12.0 5.0 - - the remaining 2 9.0 1.0 10.5 - 5.8 1.2 3.3 - - - the remaining 3 9.0 1.5 6.0 - 3.7 4.2 4.0 7.5 - 0.5 the remaining 4 6.6 0.6 6.4 3.0 5.6 1.0 6.5 9.6 0.1 - the remaining 5 5.6 1.9 10.9 - 5.1 - 7.7 8.2 - - the remaining 6 10.0 0.7 6 0.1 5.4 2 5.4 4.5 - - the remaining 7 18.4 3.0 1.5 - 2.5 5.0 - 15.0 - 0.02 the remaining 8 8.5 - 9.5 - 5.5 2.2 2.8 5.0 - - the remaining 9 10.0 0.7 2.0 0.25 12.0 1.2 2.6 - - - the remaining 10 6.6 - 12.8 - 5.2 - 7.7 - - - the remaining

Table 2 Carbon C Chromium Cr Nickel Ni Cobalt Co Molybdenum Mo Tungsten W Niobium Nb Titanium Ti Aluminum Al Iron Fe Boron B Zirconium Zr Tantalum Ta 11 0.38 20.0 20.0 the remaining 4.0 4.0 4.0 - - 4.0 - - - 12 0.45 21.0 ≤1.0 the remaining - 11.0 2.0 - - 2.0 - - - 13 0.25 29.5 10.5 the remaining - 7.0 - - - 2.0 0.01 - - 14 0.60 24.0 10.0 the remaining - 7.0 - 0.2 - - - 0.5 3.5 15 0.60 24.0 10.0 the remaining - 7.0 - 0.25 0.18 - - - 3.5

Table 3 Table 4 mold heating temperature 1560°C Burning temperature 1550°C pull down speed 10cm/hour model material ceramics Vacuum ≤2×10 Torr alloy No. 2 and No. 3 alloy product this invention common method number 2 75% 0% No. 10 83% 0%

Claims (11)

1. a gas-turbine blade comprises as the dovetail that is fixed on the wheel disc; A root, it is connected in described dovetail, and forms one or more complete projections in the dovetail side; A blade that is connected in described root is characterized in that described gas-turbine blade made by nickel-base alloy, wherein γ ' phase condense in formed a kind of monoclinic crystal structure basically γ mutually in.

2. according to the described gas-turbine blade of claim 1, it is characterized in that the projection that is configured in described root is single-stage or multi-stage sealed part, be configured in along on two surfaces of described blade rotation face.

3. according to the described gas-turbine blade of claim 2, it is characterized in that a kind of structure is arranged, wherein the edge of each hermetic unit is towards described blade bending, and slides with respect to nozzle, so that the gas flow of seal flow.

4. according to the described gas-turbine blade of claim 1, it is characterized in that projection, be one and be positioned at the two lip-deep platforms that intersect with the blade rotation surface in described root configuration.

5. according to each described gas-turbine blade in the claim 1 to 4, it is characterized in that the root of tool, have and be no less than 15 centimetres at projection ²Cross sectional area.

6. according to each described gas-turbine blade in the claim 1 to 4, it is characterized in that the described root and the blade that comprise described dovetail and projection make with nickel-base alloy, wherein γ ' separates out in the monocrystalline matrix of γ sun.

7. according to each described gas-turbine blade in the claim 1 to 4, it is characterized in that its length at length direction is no less than 180 millimeters.

8. according to each described gas-turbine blade in the claim 1 to 4, it is characterized in that the weight of described blade is not quite less than 30% of described gas-turbine blade gross weight.

9. a gas-turbine blade comprises as the dovetail that is fixed on the wheel disc; It is connected in described dovetail root, and forms an a plurality of complete projection in the dovetail side; A blade that is connected in described root is characterized in that edge to the described dovetail from described blade is solidified described gas-turbine blade by a unidirectional solidification process, and a γ is that the monocrystal nickel-base alloy is made mutually.

10. method of making gas-turbine blade, this blade comprise one as the dovetail that is fixed to the part on the wheel disc; A root that is connected in described dovetail, and at the whole formation in the side of dovetail projection; With a blade that is connected in described root, described manufacture method comprises the steps:

To be connected to main mould corresponding to the bypass mould of projection, on root and the blade corresponding to dovetail; With

Solidify gradually along a direction with identical speed by the deposite metal that makes the nickel-base alloy in main mould and the bypass mould and to cast a monoclinic crystal structure.

11. gas turbine, its work fuel gas temperature is not less than 1400 ℃, it comprises a gas compressor, a combustion chamber sleeve and a plurality of turbine blade, described turbine blade be single-stage or multistage turbine vanes fixed on the turbine disk, nozzle arrangement with described turbine blade relevant position on, the metal temperature of first order turbine blade in working order the time is not less than 1000 ℃, the blade length of turbine blade is not less than 180 millimeters, and each turbine blade comprises as the dovetail that is fixed on the wheel disc; A root, it is connected in described dovetail, and forms one or more complete projections in the dovetail side; Platform on the described root with the blade that is connected with the platform of described root, is characterized in that described gas-turbine blade made by nickel-base alloy, wherein γ ' phase condense in a kind of monoclinic crystal structure that forms basically γ mutually in.

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