JPH08121104A - Turbine blade and gas turbine with highly corrosion-resistant coating layer - Google Patents

Abstract

(57) [Summary] [PROBLEMS] To provide a turbine blade provided with a coating layer which is superior to a MCrAlY alloy and has excellent corrosion resistance even at a metal temperature of 900 ° C or higher, and a gas turbine using the blade. In a turbine blade having a coating layer on a part of the surface of a heat-resistant alloy base material containing Ni as a main component, at least one of the coating layers has a thickness of 0.5 μm or more.
A metal layer 10 made of Ir having a purity of 30% or more and a thickness of 10 μm or less
2 and a ZrO ₂ -based columnar ceramic TBC layer 104 having a width of 20 to 600 μm is coated on the metal layer 102. [Effect] It is possible to supply a turbine blade provided with a coating layer having excellent corrosion resistance, and by using the blade, a gas turbine having a high combustion temperature can be provided, and the efficiency of the turbine can be improved.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine turbine blade provided with a highly corrosion resistant coating layer, and more particularly to a high efficiency gas turbine provided with the turbine blade.

[0002]

2. Description of the Related Art Gas turbines for power generation are being increased in operating temperature for the purpose of improving efficiency. As a result,
There is a strong demand for improvement in high-temperature durability of turbine blades, which are heat-resistant parts. Against this background, heat resistant alloys having high temperature strength and excellent reliability are being developed, but the heat resistant temperature is limited. As a method of reducing the temperature of parts used under high temperature conditions,
Thermal barrier coating that covers the surface of parts with ceramics with low thermal conductivity (eg, ZrO ₂ -based ceramics)
(Thermal Barrier Coating: abbreviated as TBC) is known. In TBC, the base metal temperature of parts is 50-100.
C can be reduced.

On the other hand, a technique of coating MCrAlY alloy on the surface of a component has been widely used in combination with the above TBC for the purpose of high temperature corrosion resistance and high temperature oxidation resistance. Normally,
The structure has a structure in which the surface of the component is first coated with an MCrAlY layer, and a TBC layer of ceramics is coated on top of it. MCr
The AlY alloy has Cr of 15 to 30% and Al of 5 to 15 as M in which one or more kinds of bulk selected from Fe, Co and Ni are used.
%, Y 0.1 to 1% is a high corrosion resistant alloy.
For example, in Japanese Patent Laid-Open No. 4-266889, an MCrAlY alloy (Co-32% Ni-21% Cr) is formed on the surface of a Ni-base superalloy substrate.
-8% Al-0.5% Y) layer is provided to ensure corrosion resistance, and a ceramic-coated turbine blade having a double-layered ZrO ₂ -6% Y ₂ O ₃ layer TBC is disclosed.

[0004]

The above-mentioned prior art is intended for the base metal temperature of the gas turbine blade up to about 850 ° C., and is considered for a high temperature gas turbine that requires a base metal temperature higher than that. Not not. No particular consideration is given to high temperature corrosion and oxidation.

As described above, the operating temperature of the gas turbine for power generation is being increased to improve the efficiency, but when the temperature of the combustion gas rises, the high temperature members typified by turbine blades are as follows. It is required to take appropriate measures.

(1) The cooling effect of the blade and the heat shield effect by TBC are strengthened to suppress the rise of the metal temperature to the minimum.

(2) Corresponding to the improvement of the operating temperature, the service temperature of the base alloy is increased and the metal temperature is higher than before.

Here, strengthening the blade cooling is not desirable because the pressure loss increases due to the increased cooling air and the power generation efficiency of the gas turbine decreases. In addition, since there is a limit to the heat shielding effect of TBC, it is essential to deal with (2) of (1) and (2) above. However, the MCrAlY alloy layer, which has been improved in high temperature corrosion resistance and oxidation resistance by the conventional technique, does not exhibit excellent corrosion resistance and oxidation resistance at a temperature significantly higher than 900 ° C. Therefore, even if a heat resistant alloy excellent in terms of service temperature, that is, high temperature strength is developed,
The simple application of conventional heat-resistant and corrosion-resistant coating technology cannot cope with the harsh gas turbine combustion gas atmosphere at even higher temperatures.

An object of the present invention is to provide a turbine blade provided with a coating layer which is superior to the MCrAlY alloy and has excellent corrosion resistance even at a metal temperature of 900 ° C. or higher, and a gas turbine using the blade.

[0010]

In order to achieve the above object, according to the present invention, in a turbine blade having a coating layer on a part of the surface of a heat-resistant alloy base material containing Ni as a main component, At least one layer, by weight, has a purity of 30
% Or more Ir is used as the metal layer. Ir
Is a noble metal and has a high melting point of 2400 ° C. or higher.
Excellent high temperature corrosion resistance and high temperature oxidation resistance, 1500 ° C
It is stable even at ultra-high temperatures exceeding 0. If this Ir is used as the metal layer of the coating layer, necessary and sufficient high temperature corrosion resistance and oxidation resistance can be obtained even when the metal temperature of the turbine blade is 900 ° C. or higher.

To achieve the above object, in the present invention, in a turbine blade having a coating layer on a part of the surface of a heat-resistant alloy base material containing Ni as a main component, at least one of the coating layers is The metal layer is made of an alloy containing 30% to 90% by weight of Ir. These metal layers are particularly effective in the thickness range of 0.5 μm or more and 300 μm or less.

Further, in the present invention, on the upper part of the metal layer,
A stabilized ZrO ₂ -based ceramic layer containing ZrO ₂ as a main component and one or more stabilizers selected from Y ₂ O ₃ , MgO, and CaO is coated in a width of 20 to 600 μm. This ceramic layer exerts a heat shielding effect as TBC and reduces the metal temperature.

Further, in order to achieve the above object, in the present invention, in a turbine blade having a coating layer on a part of the surface of a heat-resistant alloy base material containing Ni as a main component, at least one of the coating layers is One or more selected from Fe, Co, and Ni are used as a bulk, and Cr15 to 30 in a weight ratio.
%, Al 5 to 15%, Y 0.1 to 1%, Ir 0.1 to 10
% Is a metal layer composed of an alloy of. This is based on a MCrAlY alloy that has been conventionally used as a metal layer for improving corrosion resistance, and has Ir added thereto to improve high temperature corrosion resistance and high temperature oxidation characteristics.

When such a metal layer is used,
It is desirable to provide a diffusion prevention layer containing an intermetallic compound or ceramics as a main component between the metal layer and the heat resistant alloy substrate.

In the present invention, the base metal constituting these turbine blades is a unidirectionally solidified material containing Ni as a main component, a single crystal material, or an oxide dispersion strengthening material, and the alloy is 14 kgf / mm ² , it is desirable 10 ⁵ hours service temperature of 900 ° C. or higher.

Further, in order to achieve the above object, at least one of the gas turbines for power generation uses the turbine blade.

[0017]

In the currently used high temperature gas turbine, the metal temperature of the first-stage turbine blade is 800 ° C to 900 ° C, and the material damage in this temperature range is dominated by the high temperature corrosion from the sulfur component contained in the fuel rather than the high temperature oxidation. Become. On the other hand, the metal temperature will rise from 900 ° C to 1 with future operating temperature rise.
Above 000 ° C, material damage is dominated by high temperature oxidation rather than high temperature corrosion. MCrAlY shows sufficient corrosion resistance in the working temperature range, but 900 ° C to 1000 ° C.
Above ℃ is not enough. Particularly in a high temperature oxidizing atmosphere of 1000 ° C. or higher, the MCrAlY layer rapidly loses the function of blocking oxygen.

In the present invention, Ir or an Ir alloy is used as the highly corrosion resistant coating layer of the gas turbine blade. Ir
Is a noble metal and has a high melting point of 2400 ° C. or higher.
Excellent high temperature corrosion resistance and high temperature oxidation resistance, 1500 ℃
It is stable even at ultra-high temperatures exceeding 0. If this Ir is used as the metal layer of the coating layer, necessary and sufficient high temperature corrosion resistance and oxidation resistance can be obtained even if the metal temperature of the turbine blade is 900 to 1000 ° C. or higher.

The component of the metal layer containing Ir as a main component is effective when Ir is contained in an amount of 30 to 100%. Ir is a scarce resource, and its specific gravity is
It is a large and heavy element of 22.4. Therefore, alloying with another element which is cheap and light is effective in terms of economical weight and weight reduction of the device. In that case, it is a condition that the high temperature corrosion resistance and the oxidation resistance thereof are not deteriorated even if they are alloyed. When alloying Ir at a weight ratio of 30% or less, the selection range of alloying elements becomes extremely small, and it becomes difficult to obtain the effects of high temperature corrosion resistance and oxidation resistance.

The thickness of the metal layer containing Ir as a main component is preferably 0.5 μm or more and 300 μm or less. The film thickness is 0.
If the thickness is less than 5 μm, the components of the metal layer disappear and deteriorate due to diffusion into the base material or diffusion of alloy components from the base material, and sufficient effects of high temperature corrosion resistance and oxidation resistance cannot be obtained. On the other hand, when the film thickness is 300 μm or more, problems such as peeling and cracking are likely to occur, which is not suitable.

Further, in the present invention, one or more kinds selected from Fe, Co and Ni is used as a bulk, and Cr 15 to 3 in a weight ratio.
0%, Al 5 to 15%, Y 0.1 to 1%, Ir 0.1 to 1
You may use the alloy which consists of 0%. This is the conventional MC
It is based on the rAlY alloy and has Ir added thereto to improve the high temperature corrosion resistance and the high temperature oxidation characteristics, and particularly to the high temperature oxidation which becomes remarkable at 1000 ° C. or higher.
In this case, the amount of Ir is 10% or less, and a sufficient effect is exhibited.
To further assure use at 1000 ° C. and above, a layer of superior oxidation resistance may be coated on top of this modified MCrAlY layer. Furthermore, this improved MCrAlY
When using, the metal temperature is 900-1000 ℃
In the above case, MCrAlY may diffuse into the base material and disappear after long-term use. Therefore, it is preferable to provide a diffusion prevention layer between the base material and the metal layer. This diffusion prevention layer is N
Intermetallic compounds such as iAl are preferred.

In the present invention, ZrO stabilized on cubic zirconia is provided on the metal layer by ZrO ₂ as a main component and at least one stabilizer selected from Y ₂ O ₃ , MgO and CaO. It is also effective to coat the ₂ type ceramic layer with a width of 20 to 600 μm. Depending on the type of stabilizer,
It is important to adjust the addition amount to obtain a completely stabilized zirconia. This ceramic layer exerts a heat shielding effect as TBC and reduces the metal temperature. That is, it becomes possible to use it at a higher operating temperature.

The base metal constituting these turbine blades is a unidirectionally solidified material containing Ni as a main component, a single crystal material, or an oxide dispersion strengthening material. These have higher high-temperature strength than conventional polycrystalline Ni-base superalloys, and can withstand a rise in metal temperature accompanying a rise in operating temperature of a gas turbine. Especially 14 kgf / mm ^2, desirably 10 ⁵ hours service temperature of 900 ° C. or higher. By using such an alloy as a base material and using a turbine blade having an Ir or Ir alloy metal layer as described above as a coating layer, particularly as a first stage blade,
For the first time, a high-efficiency gas turbine with a high operating temperature becomes possible.

[0024]

EXAMPLES The present invention will be specifically described below with reference to examples.

Example 1 FIG. 1 shows a schematic sectional view of a coating layer of a turbine blade according to the present invention, and FIG. 2 shows a perspective view of the produced turbine blade. The base alloy is Ni
The base superalloy single crystal alloy CMSX-4 was used to cast a single crystal by the unidirectional solidification method. The coating layer 90 of the present invention was provided on the wing portion 200 and the platform portion 202, which are the portions exposed to the combustion gas. The method for producing a turbine blade according to the present invention is as follows. First, a single crystal blade is obtained by the unidirectional solidification casting method. First, a metal layer 102 having a thickness of 50 μm is provided on the coating portion of the blade by using a powder of Ir alone by plasma spraying in a reduced pressure atmosphere. The conditions are such that the above alloy powder is charged into a plasma jet (50 kW) formed by using a mixed gas of Ar-7% H ₂ and thermal spraying is performed, and the atmospheric pressure during the thermal spraying is about 50 Torr. As the pretreatment, degreasing and washing of the test piece base material and blasting with Al ₂ O ₃ grit are performed. After that, a ZrO ₂ system columnar ceramic TBC layer 104 is formed on the surface of the blade provided with the metal layer 102 by using a film forming apparatus equipped with an evaporation source and an ion beam source.
Was produced. ZrO ₂ -6% Y ₂ O ₃ was used as the vapor deposition material, and oxygen ions were used as the ion beam. Ion beam irradiation is used to improve the adhesion of the ceramic coating. ZrO ₂ system columnar ceramics TBC having a columnar structure with high adhesion by electron beam evaporation source and oxygen ion irradiation
The layer 104 was provided to 300 μm.

Example 2 FIG. 3 shows a schematic sectional view of a coating portion of a turbine blade according to the present invention having another coating layer according to the present invention. The base material was a Ni-based unidirectionally solidified material. The coating layer 90 was provided on the wing surface and the platform that are exposed to the combustion gas. The manufacturing method according to this example is as follows. First, a unidirectionally solidified structure is obtained by a unidirectionally solidified casting method. Ir-30% Fe alloy powder was used for the coating portion of this blade, and the Ir alloy layer 1 having a thickness of 100 μm was formed by plasma spraying in a reduced pressure atmosphere.
03 is provided. The conditions are the same as in Example 1. Then, a ZrO ₂ -based ceramic TB was formed on the surface of the blade provided with the Ir alloy layer 103 by using a film forming apparatus equipped with an evaporation source.
The C layer 105 was formed to a thickness of 300 μm. Zr as vapor deposition material
The O ₂ -6% Y ₂ O ₃ was used.

In this embodiment, the alloy composition of the Ir alloy layer 103 is Ir-30% Fe, but Ir is 30% to 90%.
If so, any alloy component may be used.

<Embodiment 3> FIG. 4 shows a schematic sectional view of a coating portion of a turbine blade according to the present invention having another coating layer according to the present invention. The base material was a Ni-based single crystal material. The coating layer 90 was provided on the wing surface and the platform that are exposed to the combustion gas. The manufacturing method according to this example is as follows. First, a single crystal blade is obtained by the unidirectional solidification casting method. Ir is applied to the coated part of this blade.
An Ir alloy layer 103 having a thickness of 200 μm is provided by plasma spraying in a reduced pressure atmosphere using -30% Fe alloy powder.
The conditions are the same as in Example 1. In the case of this embodiment,
No TBC of the ceramic layer is provided.

<Embodiment 4> FIG. 5 shows a schematic sectional view of a coating portion of a turbine blade according to the present invention having another coating layer according to the present invention. The base material was a Ni-based single crystal material. The coating layer 90 was provided on the wing surface and the platform that are exposed to the combustion gas. The manufacturing method according to this example is as follows. First, a single crystal blade is obtained by the unidirectional solidification casting method. M on the coating of this blade
CrAlY alloy (Co-32% Ni-21% Cr-8%
An MCrAlY layer 108 having a thickness of 60 μm is provided by plasma spraying in a reduced pressure atmosphere using Al-0.5% Y) powder. Similarly, a powder of Ir alone is used to form a metal layer 102 having a thickness of 10 μm by plasma spraying in a reduced pressure atmosphere. The thermal spraying conditions are the same as in Example 1. Then MCr
A ZrO ₂ system columnar ceramic T similar to that of Example 1 was formed by using a film forming apparatus equipped with an evaporation source and an ion beam source on the surface of the blade provided with the AlY layers (metal layers) 108 and 102.
The BC layer 104 is produced.

Example 5 FIG. 6 shows a schematic sectional view of a coating portion of a turbine blade according to the present invention having another coating layer according to the present invention. The base material was a Ni-based unidirectionally solidified material. The coating layer 90 was provided on the wing surface and the platform that are exposed to the combustion gas. The manufacturing method according to this example is as follows. First, a unidirectionally solidifying casting method is used to obtain a unidirectionally solidifying blade. A metal layer 102 of Ir having a thickness of 20 μm is provided on the coating portion of the blade by using a powder of Ir alone by plasma spraying in a reduced pressure atmosphere.
Similarly, an Ir alloy layer 103 having a thickness of 60 μm is formed by plasma spraying in a reduced pressure atmosphere using Ir-30% Fe powder.
To provide. The thermal spraying conditions are the same as in Example 1. afterwards,
The same ZrO ₂ system columnar ceramic TBC layer 104 as in Example 1 was formed on the surface of the blade provided with the metal layers 102 and 103 by using a film forming apparatus equipped with a vapor deposition source and an ion beam source.
Is prepared.

Example 6 FIG. 7 shows a schematic sectional view of a coating portion of a turbine blade according to the present invention having another coating layer according to the present invention. The base material was a Ni-based oxide dispersion strengthened alloy. An intermetallic compound NiAl alloy powder (Ni-50 at.% Al) is used to coat the blade with a diffusion prevention layer of 20 μm in a reduced pressure atmosphere plasma spraying method. Next, Co-32% Ni-21% Cr-8
A coating layer of 100 μm is formed in the same manner using% Al-0.5% Y-5% Ir alloy powder. The thermal spraying conditions are the same as in Example 1. Although the ceramic TBC layer is not provided in this embodiment, the TBC layer shown in Embodiment 1 may be provided in order to reduce the metal temperature.

Example 7 In this example, the high temperature oxidation resistance characteristics of MCrAlY and Ir were compared in order to explain the effect of the present invention. Figure 8 shows a typical MCrAlY, CoNiCrAl.
Y8426 (Co-32% Ni-21% Cr-8% Al-0.
It is the result of the oxidation test of the powder of 5% Y). The test is carried out with a thermobalance (TG-DTA), the atmosphere is an atmosphere with a flow rate of 100 cc, and the temperature condition is 20 ° C./min.
The weight change due to oxidation was examined by holding for a time. Sample powder 0.
17 g was placed in an alumina crucible, and the standard sample was alumina powder. CoNiCrAlY8426 started to increase in weight due to oxidation at around 910 ° C, which is a temperature rising process up to 1300 ° C, and showed a large increase in weight thereafter. In the end, it showed a weight gain of about 27 mg, which was about 16% of the total, and 9
It was found that sufficient oxidation resistance was not exhibited in the temperature range of 00 ° C to 1000 ° C or higher.

On the other hand, FIG. 9 shows the result of the oxidation test of Ir alone, but it was found that it was hardly oxidized even in the atmosphere at 1500 ° C., and it became clear that it exhibited sufficient high temperature oxidation resistance.

<Embodiment 8> FIG. 10 is a sectional view of a rotating portion of a gas turbine for power generation according to the present invention in which the gas turbine blades described in Embodiments 1 to 6 are used as a first stage blade of a gas turbine.

3 is a first stage blade, 5 is a second stage blade,
7 is a third stage blade, 20 is a first stage nozzle, 25 is a second stage nozzle, 27 is a third stage nozzle, 4 is a turbine disk,
10 is a turbine stub shaft, 13 is a turbine stacking bolt, 18 is a turbine spacer, 19 is a distant piece, 6 is a compressor disk, 8 is a compressor stacking bolt, 9 is a compressor stub shaft, and 11 is a hole. The gas turbine of the present invention has three turbine blades and three turbine nozzles.

The gas turbine in this embodiment is mainly composed of a heavy tuty type, a single shaft type, a horizontal split casing, a stacking type rotor, a compressor is a 17-stage axial flow type, a turbine is a 3-stage impulse type, and a combustion type. The vessel has a verse flow type, 16 cans, and a slot cool type.

The first-stage blade 3 of the gas turbine in this embodiment has the coating layer shown in the first or second embodiment. The base material part is a single crystal cast of Ni-base superalloy, Cr 6-8% by weight, Mo 0.5-1.
%, W6-8%, Re1-4%, Al4-6%, Ta6
-9%, Co 0.5-10%, Hf 0.03-0.13
%, And the balance being an unavoidable impurity and an alloy composed of Ni. The total length of the blade of the first stage blade is about 220 mm. This single crystal casting has a useful temperature of 105 kg / mm ² for 105 hours of 930 ° C. to 940 ° C., and complicated air cooling holes are provided inside, and it is cooled by compressed air during operation.
The cooling method is a closed method, and the cooling structure is a staggered rib method.

Second stage blade 5 and third stage blade 7
Is by weight, Cr 12 to 16%, Mo 0.5 to 2%, W2
~ 5%, Al 2.5-5%, Ti 3-5%, Ta 1.5-
3%, Co8-10%, C0.05-0.15%, B0.0
It is composed of a Ni-base superalloy consisting of Ni and 0.05 to 0.02% and the balance unavoidable impurities. These blades have an equiaxed crystal structure obtained by ordinary casting. Each has an internal cooling hole and is cooled by compressed air. The service temperature of these materials for 105 hours at 14 kgf / mm ² is 8
It is 40 ° C to 860 ° C. The blade surface was coated with a diffusion coating of Cr or Al to improve corrosion resistance.

The weight of the first stage nozzle 20 is Cr24-3.
0%, Ni8-12%, W6-9%, Ti0.1-0.4
%, Co8-10%, C0.2-0.4%, B0.005
% Or less, Fe 1.0% or less, Z 1.0% or less, Nb 0.3
% Or less, Hf of 1.0% or less, Ta of 2.0% or less, and a Co-based superalloy ordinary cast material (equiaxed crystal structure) consisting of Co and the balance unavoidable impurities are used. The service temperature of this alloy for 105 hours at 6 kgf / mm ² is 900 ° C to 910 ° C.
The cooling is a closed system impingement cooling. A thermal barrier coating layer is provided on the outer surface of the first stage nozzle in contact with the flame. It consists of a Y ₂ O ₃ -stabilized zirconia layer having a columnar crystal structure and a bonding layer between the base metal and the zirconia layer. The tie layer comprises Al2-5% by weight,
It is made of an alloy containing Cr of 20 to 30% and Y of 0.1 to 1% and the balance of Ni or Ni + Co, and also has an effect of improving corrosion resistance.

Second stage nozzle 25 and third stage nozzle 27
By weight, Cr 21 to 24%, Co 18 to 23%, C
0.05 to 0.20%, W1 to 8%, Al1 to 2%, Ti
2-3%, Ta 0.5-1.5%, B 0.05-0.15
%, And an alloy Ni consisting of Ni and Ni
It is composed of a base superalloy. These nozzles have an equiaxed crystal structure obtained by ordinary casting. Although it is not particularly necessary to provide a thermal barrier coating layer, the second stage nozzle is provided with a diffusion coating of Cr or Al in order to enhance corrosion resistance.
Each has an internal cooling hole and is cooled by compressed air. The service temperature of these materials for 105 hours at 6 kgf / mm ² is 840 ° C to 860 ° C.

In this embodiment, the turbine disk 4 has a weight of C0.05 to 0.2%, Si0.5% or less, Mn1% or less, Cr8 to 13%, Ni3% or less, Mo1.5 to 3%.
%, V 0.05% to 0.3%, Nb 0.02 to 0.2%, N
Ni with 0.02 to 0.1% and the balance substantially Ni
A base forged alloy is used. The Ni-based forged alloy has a temperature of 450 ° C,
10h creep rupture strength of 50kg / mm ² or more, 20 ℃ V
The Notch Charpy impact value is 7 kg-m / cm or more, which sufficiently satisfies the strength required as a material for a high temperature gas turbine.

The compressor blade has 17 stages and the obtained air compression ratio is 18.

As the fuel used, natural gas and light oil are used.

With the above structure, a more reliable and highly balanced gas turbine can be obtained. The gas inlet temperature to the first-stage turbine nozzle is 1500 ° C, the metal temperature of the first-stage turbine blade is 920 ° C, and the exhaust gas of the gas turbine is exhausted. The temperature is 650 ° C., and it is possible to achieve a power generation gas turbine with a power generation efficiency of 37% or more in LHV display.

[0045]

According to the present invention, it is possible to supply a turbine blade having a coating layer that exhibits excellent corrosion resistance even at a metal temperature of 900 ° C. or higher, and by using the blade, a gas having a high combustion temperature can be supplied. A turbine is possible and the efficiency of the turbine can be increased.

[Brief description of drawings]

FIG. 1 is a schematic cross-sectional view of a high corrosion resistant coating layer.

FIG. 2 is a perspective view of a gas turbine blade.

FIG. 3 is a schematic sectional view of a high corrosion resistant coating layer.

FIG. 4 is a schematic cross-sectional view of a high corrosion resistant coating layer.

FIG. 5 is a schematic cross-sectional view of a high corrosion resistant coating layer.

FIG. 6 is a schematic sectional view of a high corrosion resistant coating layer.

FIG. 7 is a schematic cross-sectional view of a high corrosion resistant coating layer.

FIG. 8 is an oxidation test result of MCrAlY (CoNICrAlY8426).

FIG. 9: Ir oxidation test result.

FIG. 10 is a sectional view of a rotating portion of the gas turbine.

[Explanation of symbols]

3 ... First stage blade, 4 ... Turbine disc, 5 ... Second stage blade, 6 ... Compressor disc, 7 ... Third stage blade, 8 ... Compressor stacking bolt, 9 ... Compressor stub shaft, 10 ... Turbine stub shaft, 11 ... Hole , 13 ... Turbine stacking bolts, 1
7 ... Compressor blade, 18 ... Turbine spacer,
19 ... Distant piece, 20 ... First stage nozzle, 25 ...
2nd stage nozzle, 27 ... 3rd stage nozzle, 90 ... Coating layer, 9
9 ... Ni-based oxide dispersion strengthened blade, 100 ... Ni-based superalloy unidirectionally solidifying blade, 101 ... Ni-based superalloy single crystal blade, 102 ... Ir metal layer, 103 ... Ir alloy layer, 104 ... Columnar ceramic TBC layer , 105 ... Ceramic TBC layer, 107 ... MCrAlY + Ir layer, 10
8 ... MCrAlY layer, 109 ... Diffusion prevention layer, 200 ... Wing portion, 202 ... Platform portion.

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Inventor Hiroshi Fukui 7-1-1 Omika-cho, Hitachi-shi, Ibaraki Hitachi Ltd. Hitachi Research Laboratory

Claims (8)

[Claims] 1. A turbine blade comprising a metal layer containing Ir having a purity of 30% or more in a weight ratio on at least a part of a surface of a base material made of a heat-resistant alloy containing Ni as a main component. 2. A turbine blade comprising a metal layer containing Ir having a purity of 90% or more in a weight ratio on at least a part of a surface of a base material made of a heat-resistant alloy containing Ni as a main component. 3. The turbine blade according to claim 1, wherein the metal layer has a thickness of 0.5 μm or more and 300 μm or more.
A turbine blade characterized by having a diameter of not more than μm. 4. The turbine blade according to claim 1, wherein a stabilizing Zr is provided on the metal layer.
A turbine blade having a ceramic layer of O ₂ with a thickness of 20 to 600 μm. 5. A weight ratio of Cr 15 to 30% and Al 5 to 1
5%, Y 0.1 to 1%, Ir 0.1 to 10% and the balance F
A turbine blade having a metal layer composed of an alloy composed of one or more selected from e, Co and Ni on at least a part of the surface of a base material composed of a heat resistant alloy containing Ni as a main component. . 6. The turbine blade according to claim 5, wherein a diffusion prevention layer containing an intermetallic compound or ceramics as a main component is provided between the metal layer and the base material made of the heat resistant alloy. . 7. The turbine blade according to claim 1, wherein the base material made of the heat-resistant alloy is a unidirectionally solidified material, a single crystal material, or an oxide dispersion strengthening material.
The alloy has a 14 kgf / mm ² , 10 ⁵ hour service temperature of 900
A turbine blade having a temperature of ℃ or above. 8. A gas turbine for rotating a combustion gas combusted by air compressed by a compressor through a nozzle to collide with blades respectively implanted in a plurality of disks to rotate the blades. Is a turbine blade according to any one of claims 1 to 7.