阅读说明：本技术 航空器发动机壳体用Ni基超耐热合金及由其制成的航空器发动机壳体 (Ni-based superalloy for aircraft engine case and aircraft engine case made of same ) 是由 松井孝宪 福田正 于 2019-09-26 设计创作，主要内容包括：提供一种高温区域的拉伸特性、低循环疲劳特性等高温特性优异、加工性也优异的航空器发动机壳体用Ni基超耐热合金及由其制成的航空器发动机壳体。适合用于航空器发动机壳体的Ni基超耐热合金具有如下组成：以质量％计含有Co：4.0～11.0％、Cr：12.0～17.0％、Al：2.0～4.0％、Ti：2.0～4.0％、Al+Ti：4.6～6.7％、Mo：大于5.5％且为10.0％以下、W：大于0％且为4.0％以下、B：0.001～0.040％、C：0.02～0.06％、Zr：0～0.05％以下、Mg：0～0.005％以下、P：0～0.01％以下、Nb：0～1.0％以下、Ta：0～1.0％以下、Fe：0～2.0％以下,余量为Ni和不可避免的杂质。(Provided are a Ni-based superalloy for an aircraft engine casing, which has excellent high-temperature characteristics such as tensile characteristics and low cycle fatigue characteristics in a high-temperature region, and also has excellent workability, and an aircraft engine casing made of the Ni-based superalloy. A Ni-based superalloy suitable for use in aircraft engine casings has the following composition: contains Co in mass%: 4.0-11.0%, Cr: 12.0 to 17.0%, Al: 2.0-4.0%, Ti: 2.0-4.0%, Al + Ti: 4.6-6.7%, Mo: greater than 5.5% and 10.0% or less, W: greater than 0% and 4.0% or less, B: 0.001 to 0.040%, C: 0.02 to 0.06%, Zr: 0-0.05% or less of Mg: 0 to 0.005% or less, P: 0 to 0.01% or less, Nb: 0 to 1.0% or less, Ta: 0 to 1.0% or less, Fe: 0 to 2.0% or less, and the balance of Ni and unavoidable impurities.)

1. A Ni-based superalloy for aircraft engine housings, characterized by the following composition:

contains in mass%

Co：4.0～11.0％、

Cr：12.0～17.0％、

Al：2.0～4.0％、

Ti：2.0～4.0％、

Al+Ti：4.6～6.7％、

Mo: more than 5.5% and less than 10.0%,

W: more than 0% and not more than 4.0%,

B：0.001～0.040％、

C：0.02～0.06％、

Zr: less than 0.05% (including 0%)

Mg: less than 0.005% (including 0%)

P: less than 0.01% (including 0%)

Nb: less than 1.0 percent (including 0 percent),

Ta: less than 1.0 percent (including 0 percent),

Fe: less than 2.0% (including 0%),

the balance being Ni and unavoidable impurities.

2. The Ni-based superalloy for aircraft engine housings of claim 1, wherein M is calculated from thermodynamic calculations based on the composition and composition of the Ni-based superalloy for aircraft engine housings of claim 1₆The difference (T1-T2) between the solid solution temperature T1 of the C-type carbide and the solid solution temperature T2 of the gamma' phase is equal to or higher than-30 ℃ and equal to or lower than (T1-T2) and equal to or lower than +40 ℃.

3. The Ni-based superalloy for aircraft engine casings as claimed in claim 1 or 2, wherein the number of failure cycles in a low cycle fatigue test with a total strain range of 0.6% at 750 ℃ is 1.0 x 10⁵The above.

4. The Ni-based superalloy for an aircraft engine casing according to any of claims 1 to 3, having a tensile strength of 1000MPa or more at 750 ℃.

5. An aircraft engine casing, characterized in that it is made of the Ni-based super heat resistant alloy for aircraft engine casings according to any one of claims 1 to 4.

Technical Field

The present invention relates to a Ni-based superalloy for an aircraft engine case and an aircraft engine case made of the Ni-based superalloy, and more particularly, to a Ni-based superalloy for an aircraft engine case having excellent high-temperature material properties such as tensile properties and low cycle fatigue properties in a high-temperature region and having excellent workability, and an aircraft engine case made of the Ni-based superalloy.

Background

Since Ni-based superalloy has excellent strength, toughness, corrosion resistance, oxidation resistance, and the like, it is used in various technical fields requiring heat resistance, including aircraft engine components, power generation gas turbine components, and the like.

In particular, in structural members for aircraft engines, in order to improve the fuel economy, it is necessary to increase the combustion temperature, and therefore, Ni-based superalloys having more excellent high-temperature characteristics are required.

In order to meet this demand, a large number of Ni-based superalloys have been developed.

For example, patent document 1 describes "a nickel-based alloy containing, in weight%, about 0.10% or less of carbon, about 12 to about 20% of chromium, about 4% or less of molybdenum, about 6% or less of tungsten (the total of molybdenum and tungsten is about 2% or more and about 8% or less), about 5 to about 12% of cobalt, about 14% or less of iron, about 4 to about 8% of niobium, about 0.6 to about 2.6% of aluminum, about 0.4 to about 1.4% of titanium, about 0.003 to about 0.03% of phosphorus, about 0.003 to about 0.015% of boron, and nickel and inevitable impurities, wherein the total atomic% of aluminum and titanium is about 2 to about 6%, the atomic% ratio of aluminum and titanium is about 1.5 or more, and the value of the total atomic% of aluminum and titanium divided by the atomic% of niobium is about 0.8 to about 1.8" (see claim 1.8), the nickel-based alloys described above are also described as being useful in gas turbine engine components (e.g., disks, blades, fasteners, housings, shafts).

In addition, in the above patent document 1, Alloy718 (corresponding to UNS N07718) and a Waspaloy equivalent Alloy (corresponding to UNS N07001), which are commercially available Ni-based super heat-resistant alloys, are used as comparative conventional alloys, and it is considered that since the tensile strength and temperature stability of the Ni Alloy described in the above patent document 1 are extremely close to the values of the Waspaloy equivalent Alloy and are more excellent than Alloy718, the Ni Alloy described in patent document 1 is more excellent in stress rupture and creep life than both Alloy718 and Waspaloy equivalent Alloy, the time-dependent stress rupture and temperature stability associated with creep behavior are comparable to the Waspaloy equivalent alloy, therefore, the Ni Alloy described in patent document 1 has higher tensile strength, stress cracking property, creep life and long-term temperature stability than Alloy718 and Waspaloy equivalent alloys, and maintain good hot workability, weldability, and advantageous cost as compared with these alloys.

Here, "UNS" is "Unified numbering System" specified in SAE HS-1086 and ASTM DS-566, and the above-mentioned N07718, N07001 denote the alloy-specific numbers registered herein.

In patent document 1, a typical composition of Alloy718 is considered to be "in mass%, C: 0.08% or less, Mn: 0.35% or less, P: 0.015% or less, S: 0.015% or less, Si: 0.35% or less, Cr: 17-21%, Ni: 50-55%, Mo: 2.8 to 3.3%, Nb and Ta: 4.75-5.5%, Ti: 0.65 to 1.15%, Al: 0.2-0.8%, Co: 1% or less, B: 0.006% or less, Cu: 0.3% or less, with the balance being Fe and unavoidable impurities ", and a typical composition of the Waspaloy equivalent alloy is considered to be" C: 0.02 to 0.10%, Mn: 0.1% or less, P: 0.015% or less, S: 0.015% or less, Si: 0.15% or less, Cr: 18-21%, Fe: 2% or less, Mo: 3.5-5.0%, Ti: 2.75-3.25%, Al: 1.2-1.6%, Co: 12-15%, B: 0.003 to 0.01%, Cu: 0.1% or less, Zr: 0.02 to 0.08%, and the balance of Ni and inevitable impurities.

Disclosure of Invention

Problems to be solved by the invention

In recent years, since fuel costs account for a high percentage of the use cost of an aircraft, in order to improve the fuel consumption rate, as one of the countermeasures, the combustion temperature has been considered to be high, and the use temperature of the material tends to increase as the combustion temperature increases.

However, including conventional Ni-based superalloys such as those described in patent document 1, Ni-based superalloys achieve high strength by precipitation strengthening based on intermetallic compounds, and the higher the use temperature, the more remarkable the phase change, and therefore the strength characteristics are likely to change during use, and it is difficult to maintain the initial characteristics in design.

In particular, an aircraft engine casing is a member whose use environment is high temperature. For example, in the case of a burner, it becomes high because it is used in an environment exposed to combustion gas. According to the engine member, the use temperature and the use time are limited, and the use is realized without causing practical problems. A representative component is a disk located between the blades and the shaft in the rotor section. The temperature on the inner diameter side in contact with the shaft is relatively low, and the temperature on the outer diameter side in contact with the blades is high. In the disk, which is located at the highest temperature in front of the high-pressure turbine, temperatures of around 700 ℃ are sometimes reached. Therefore, a molten/forged Ni-based superalloy for disks which can withstand temperatures of about 700 ℃ under time constraints has been proposed. The housing does not produce as much stress as the disk, but is sometimes used at a higher temperature than the disk as in the above-described burner. Therefore, if the previously proposed Ni-based superalloy is to be used directly in an engine case, the alloy is used in a region where the mechanical properties are not optimal. Specifically, not only the conventional Ni-based superalloy which is widely used and has high versatility, but also the Ni-based superalloy for disks which is proposed to exhibit excellent strength characteristics in a temperature range of about 700 ℃ if it is used with time limitation is not suitable for a combustor case of an engine which requires a further high combustion temperature, and a problem of insufficient characteristics begins to occur.

However, although this engine case focuses on the necessity of high-temperature creep characteristics or high-temperature rupture characteristics, which are deformation or breakage due to a continuous stress load lower than the conditional yield strength occurring in a series of cycles from the start to the stop of the engine, the high-temperature low-cycle fatigue characteristics, which are breakage due to the stress load and stress relief occurring in the series of cycles, are also important characteristics in terms of the life, no proposal has been made to focus on the low-cycle fatigue strength at high temperatures.

The purpose of the present invention is to provide a Ni-based superalloy for an aircraft engine casing, which has sufficient high-temperature strength and excellent low-cycle fatigue strength even when the engine casing is used in a high temperature environment of 750 ℃ or higher, and which has excellent hot workability, and an aircraft engine casing using the Ni-based superalloy.

The term "engine casing" as used herein refers to, for example, a combustor casing, a high-pressure turbine casing, a high-pressure compressor casing, a low-pressure turbine casing, and the like of an aircraft jet engine, and the aircraft engine casing is a structural member required in all regions from the front to the rear of the engine.

The present invention is a Ni-based superalloy suitable for use in an aircraft engine casing, which exhibits its characteristics in a high-temperature region, and is mainly applicable to a combustor casing and a high-pressure turbine casing disposed behind the combustor casing. These are cylindrical blanks, and vary in size depending on the size (thrust force) of the engine, for example, 0.3 to 1.5m in outer diameter and 1m or less in height (length in the axial direction of the engine) which are greatly different depending on the engine design.

In addition, although the use temperature of the high-pressure compressor casing located in front of the combustor and the low-pressure turbine casing located behind the high-pressure turbine casing is lower than that of the combustor casing and the high-pressure turbine casing, the Ni-based superalloy of the present invention is still suitable.

The low-pressure turbine housing has a tapered shape with a larger diameter toward the rear, and the maximum diameter (toward the rear of the engine) can be as large as more than 2m, but may be 2.5m or less. In addition, the high-pressure compressor casing is instead shaped to have a larger diameter at the front face, and there is no difference in diameter as in the low-pressure turbine casing.

Means for solving the problems

The present inventors have conducted extensive studies on the alloy components and the composition ranges thereof for a Ni-based superalloy that is a material for aircraft engine housings and has excellent high-temperature strength and low-cycle fatigue properties in a high-temperature environment of 750 ℃ or higher and excellent workability, and as a result, have obtained the following findings.

First, while the Ni-based superalloy generally forms a structure in which a γ 'phase is precipitated in a γ phase (matrix) or a structure in which a γ "phase is precipitated in a γ phase for the purpose of high-temperature and high-strength, a conventional γ - γ' type Ni-based superalloy undergoes a rapid strength decrease with an increase in temperature in a temperature range exceeding 650 ℃, and a γ - γ" type Ni-based superalloy undergoes a more rapid strength decrease.

However, it has been found that by controlling the alloy composition and composition of the Ni-based superalloy within an appropriate range, excellent material properties of the Ni-based superalloy in a high temperature range can be ensured by improving the solid solution strengthening ability in addition to precipitation strengthening even in a temperature range (for example, 750 ℃ or higher) in which the strengthening contribution of the γ' phase is small.

Next, in terms of exhibiting Low Cycle Fatigue (LCF) characteristics equivalent to those at the start of use even during use, phase stability of Ni-based superalloy is generally regarded as important, but in precipitation-strengthened alloys, phase change inevitably occurs when used at high temperatures for a long period of time.

Accordingly, it has been found that by controlling the alloy composition and composition of the Ni-based superalloy within an appropriate range, rather than suppressing the phase change and placing importance on the phase stability, as a result of the decrease in strength and improvement in ductility due to the phase change, a Ni-based superalloy exhibiting excellent low cycle fatigue characteristics in a high temperature region (e.g., 750 ℃) can be obtained, and a Ni-based superalloy having low cycle fatigue characteristics after long-term use in a high temperature region that are almost the same as (without significant deterioration) the initial low cycle fatigue characteristics can be obtained.

Further, the aircraft engine case is generally manufactured by hot forging an ingot obtained by melting to prepare a forged bar blank, and then subjecting the forged bar blank to the steps of hot forging, hot rolling, solution treatment, aging treatment, and the like, and by controlling the alloy composition and composition of the Ni-based superalloy to an appropriate range, the M is reduced₆The difference between the solution temperature T1 of the C-type carbide and the solution temperature T2 of the gamma' phase is within a predetermined range, and M can be used₆The dispersion of C-type carbide suppresses the rapid growth of crystal grains in a temperature region slightly lower or slightly higher than the solid solution temperature T2 of the gamma' phase, while suppressing the rapid growth of crystal grains by mixing M₆The C-type carbide serves as a recrystallization site, and a Ni-based superalloy having a uniform metallurgical structure (grain size, precipitated phase) and excellent workability can be obtained, and as a result, an aircraft engine casing having excellent high-temperature performance can be produced.

The present invention has been made based on the above-mentioned findings, and an Ni-based superalloy for an aircraft engine case of the present invention and an aircraft engine case made of the Ni-based superalloy are characterized in that:

(1) a Ni-based superalloy for aircraft engine casings having the following composition: contains Co in mass%: 4.0-11.0%, Cr: 12.0 to 17.0%, Al: 2.0-4.0%, Ti: 2.0-4.0%, Al + Ti: 4.6-6.7%, Mo: greater than 5.5% and 10.0% or less, W: greater than 0% and 4.0% or less, B: 0.001 to 0.040%, C: 0.02 to 0.06%, Zr: 0.05% or less, Mg: 0.005% or less, P: 0.01% or less, Nb: 1.0% or less, Ta: 1.0% or less, Fe: 2.0% or less, the balance being Ni and unavoidable impurities;

(2) the Ni-based superalloy for an aircraft engine casing according to the item (1), wherein M is calculated by thermodynamic calculations based on the composition and composition of the Ni-based superalloy for an aircraft engine casing according to the item (1)₆The difference (T1-T2) between the solid solubility temperature T1 of the C-type carbide and the solid solubility temperature T2 of the gamma' phase is-30 ℃ or higher (T1-T2) or lower +40 ℃, preferably-20 ℃ or lower (T1-T2) or lower +20 ℃;

(3) the Ni-based superalloy for aircraft engine housings of (1) or (2), wherein the number of failure cycles in a low cycle fatigue test in which a total strain range at 750 ℃ is 0.6% is 1.0X 10⁵The above;

(4) the Ni-based superalloy for an aircraft engine casing according to any one of the above (1) to (3), which has a tensile strength of 1000MPa or more at 750 ℃;

(5) an aircraft engine casing made of the Ni-based super heat resistant alloy for an aircraft engine casing according to any one of the above (1) to (4).

The term "· · or less" in the above (1) means "0% or more and · · · · · · · · · · · · · · · · · · · · · · · · · · · · ·.

ADVANTAGEOUS EFFECTS OF INVENTION

The Ni-based superalloy for an aircraft engine casing according to the present invention has excellent tensile properties (high-temperature strength), low cycle fatigue properties, and excellent workability in a high-temperature region, and therefore, a large-sized engine casing can be easily produced, and moreover, since the engine can be increased in size, it contributes greatly to an improvement in the combustion consumption rate of an aircraft due to an increase in the combustion temperature.

In addition, since the material has fracture characteristics and can obtain excellent high-temperature characteristics such as less characteristic change in a high-temperature region, the material is suitable for use as a material for forming an aircraft engine casing, which requires a high combustion temperature for improving the fuel economy.

Drawings

FIG. 1 is a photomicrograph showing the metallographic structure of the Alloy of the invention (Alloy A).

Fig. 2 is a micrograph showing a metallographic structure of a conventional alloy (Waspaloy equivalent alloy).

FIG. 3 is a graph showing experimental results of static grain growth of the alloy of the present invention and a conventional alloy.

Detailed Description

The reasons for limiting the composition of the Ni-based superalloy for an aircraft engine casing according to the present invention will be described in detail below.

Co：

The Co component is mainly dissolved in the matrix (γ phase) to improve the fracture characteristics, but when the content is less than 4.0% by mass (hereinafter simply referred to as "%" as "mass%"), the sufficient fracture characteristics cannot be provided, which is not preferable, while when the content exceeds 11.0%, the hot workability is lowered, which is not preferable. Therefore, the content of Co is set to 4.0 to 11.0%. In order to more reliably obtain the effect of Co, the lower limit is preferably 8.0%, and more preferably 8.2%. The preferred upper limit of Co is 10.0%.

Cr：

Cr component forms a good protective film to improve high-temperature corrosion resistance such as high-temperature oxidation resistance and high-temperature vulcanization resistance of the alloy, and forms M with C₆Carbide of type C and increase M₆Solid solution temperature of C-type carbide in matrix by M having pinning effect₆The dispersion of the C-type carbide suppresses grain growth, contributes to grain size regulation of grains by the occurrence and progress of recrystallization, and improves grain boundary strength, but when the content is less than 12.0%, the desired high temperature corrosion resistance cannot be ensured, whereas when the content exceeds 17.0%, harmful phases such as σ phase and μ phase precipitate, and conversely, the high temperature corrosion resistance is lowered, so the content is set to 12.0 to 17.0%. In order to more reliably obtain the effect of Cr, the lower limit is preferably 13.0%, and more preferably13.5 percent. The upper limit of Cr is preferably 16.0%, more preferably less than 15.0%, and still more preferably 14.8%.

Al：

The Al component has a gamma' -phase (Ni) as a main precipitation hardening phase by aging treatment₃Al) to improve high-temperature tensile properties, low cycle fatigue properties, and fracture properties, and to bring about high-temperature strength, but when the content is less than 2.0%, the precipitation ratio of the γ 'phase is insufficient, so that a desired high-temperature strength cannot be secured, while when the content exceeds 4.0%, hot workability is lowered, and the amount of the γ' phase produced is too large, so that ductility is lowered, which is not preferable. Therefore, the Al content is set to 2.0 to 4.0%. In order to obtain the effect of Al more reliably, the lower limit is preferably 2.2%, and more preferably 2.6%. The preferable upper limit of Al is 3.3%, and more preferably 3.0%.

Ti：

The Ti component mainly has an effect of being dissolved in the γ 'phase to improve the high-temperature tensile property, the low cycle fatigue property and the fracture property, but when the content is less than 2.0%, the precipitation ratio of the γ' phase is insufficient, and thus the desired high-temperature strength cannot be secured, while when the content exceeds 4.0%, the hot workability is lowered, which is not preferable. Therefore, the content of Ti is set to 2.0 to 4.0%. In order to more reliably obtain the effect of Ti, the lower limit is preferably 2.6%, and more preferably 2.9%. The upper limit of Ti is preferably 3.6%, more preferably 3.4%.

Al+Ti：

The Al content and the Ti content are respectively 2.0 to 4.0% as described above, but in the present invention, the total content of Al and Ti is further set to 4.6 to 6.7%.

Gamma prime phase with Ni₃Since the Al site is replaced with Ti based on Al to increase the strengthening ability as a precipitated phase, the high-temperature strength can be further improved by the composite addition of Al and Ti. In addition, the total amount of the γ ' phase is also an important factor in imparting strength, and the strength is improved because the amount of the γ ' phase increases as the total content of Al and Ti increases, but if the total content is less than 4.6%, the amount of the γ ' phase precipitated is insufficient,therefore, the desired high-temperature strength cannot be secured, but if the content exceeds 6.7%, the hot workability is lowered, which is not preferable. Since Al and Ti have a composite addition effect, it is necessary to set the total content range in addition to the respective contents of Al and Ti.

Further, by setting the total content of Al and Ti to 4.6% to 6.7%, the Ni-based superalloy of the present invention can exhibit the characteristic functional effects of the Ni-based superalloy, that is, excellent high-temperature strength characteristics such as high-temperature tensile characteristics and high-temperature low-cycle fatigue characteristics, and also excellent hot-workability. The lower limit of the total amount of Al and Ti is preferably 5.2%, and more preferably 5.5%. The preferable upper limit of the total amount of Al and Ti is 6.3%.

If the Al equivalent is Al +0.56 × Ti, the "total content of Al and Ti: 4.6-6.7% ", Al equivalent is 3.8-5.5%. The lower limit of the Al equivalent is preferably 3.9, and more preferably 4.2. The preferred upper limit of the Al equivalent is 4.8.

Further, the ratio of the amount of Al to the amount of Ti is preferably 0.7 to 1.1.

This is because Al/Ti is roughly expressed in terms of Ni₃The ratio of Al to Ti in the γ' phase Al site based on Al is further improved. The lower limit of the Al/Ti value is preferably 0.8 and the upper limit is preferably 1.0.

Mo：

The Mo component has a function of improving high-temperature tensile properties, low-cycle fatigue properties, and fracture properties by being dissolved in a matrix (γ phase), and the function can exert a combined effect particularly in the coexistence with W. And has the formation of M with C₆The C-type carbide enhances grain boundaries, suppresses grain growth, and contributes to grain formation by the occurrence and progress of recrystallization, and when the content is 5.5% or less, sufficient high-temperature ductility and low cycle fatigue characteristics cannot be imparted, while when the content exceeds 10.0%, hot workability is lowered, and harmful phases such as μ phase precipitate to cause embrittlement, which is not preferable. Therefore, the content of Mo is set to be more than 5.5% and 10.0% or less. In order to obtain the effect of Mo more reliably, the lower limit is preferably 6.0%,further preferably 6.3%, and more preferably 6.9%. The preferable upper limit of Mo is 8.0%, and more preferably 7.4%.

W：

The W component has the effect of improving high-temperature tensile properties, low-cycle fatigue properties and fracture properties by being dissolved in a matrix (gamma phase) and a gamma' phase in a solid solution, and also has the effect of exhibiting composite strengthening by solid solution strengthening in a matrix in the presence of Mo, and forming M with C₆The C-type carbide cannot provide sufficient low cycle fatigue characteristics when it contains no W component because it strengthens the grain boundary, suppresses grain growth, and contributes to grain formation by the occurrence and progress of recrystallization, while it is not preferable because when it exceeds 4.0%, hot workability is lowered and ductility is also lowered. Therefore, W is contained in a range of 4.0% or less (i.e., more than 0% and 4.0% or less). In order to more reliably obtain the effect of W, the lower limit is preferably 1.1%, and more preferably 1.7%. The preferable upper limit of W is 2.7%, and more preferably 2.3%.

The ratio of Mo to W is preferably 4.6 or less in Mo/W ratio.

The reason for this is that if it exceeds 4.6, the effect of the composite addition of Mo and W contributing to the strength characteristics is reduced. The lower limit of the Mo/W value is more preferably 3.5, and the upper limit is preferably 4.3.

B：

The component B has M formed with Cr, Mo, etc₃B₂The boride has an effect of improving grain boundary strength and an effect of suppressing grain growth, but when the content is less than 0.001%, the amount of boride is not sufficient, and a sufficient grain boundary strengthening function and a sufficient pinning effect of grain boundaries cannot be obtained, while when the content exceeds 0.04%, the amount of boride is too large, and hot workability, weldability, ductility and the like are deteriorated, which is not preferable. Therefore, the content of B is 0.001 to 0.040%. In order to more reliably obtain the effect of B, the lower limit is preferably 0.003%, and more preferably 0.004%. The upper limit of B is preferably 0.020%, and more preferably 0.015%.

C：

C component toolWith M having a pinning effect with Ti, Mo, or the like₆C. MC type carbide, which suppresses grain growth, contributes to grain formation by the occurrence and progress of recrystallization, and has the effect of improving grain boundary strength, and also has the effect of newly forming M by aging treatment₂₃C₆The carbide thus strengthens the grain boundary, but when the content is less than 0.02%, M is included₆C. The precipitation ratio of the MC type carbide is not sufficient, and a sufficient grain boundary strengthening function and a sufficient pinning effect of grain boundaries cannot be obtained, while if it exceeds 0.06%, the amount of carbide produced becomes too large, and hot workability, weldability, ductility and the like are lowered, which is not preferable. Therefore, the content of C is 0.02 to 0.06%. In order to more reliably obtain the effect of C, the lower limit is preferably 0.025%, and more preferably 0.035%. The upper limit of C is preferably 0.055%, and more preferably 0.050%.

Zr：

The Zr component is contained as necessary because it has an effect of improving grain boundary strength as in B, but if the content of Zr exceeds 0.05%, the melting point is lowered to inhibit high temperature strength and hot workability, and therefore the content of Zr is 0 to 0.05%. In order to reliably obtain the effect of adding Zr, the lower limit may be set to 0.005%. The preferable upper limit when Zr is contained is 0.03%.

Mg：

The Mg component has an effect of improving hot ductility by fixing S, which is an unavoidable impurity that segregates at grain boundaries and hinders hot ductility, as a sulfide, and therefore can be contained if necessary, but if the content of Mg exceeds 0.005%, the remaining Mg becomes a factor that hinders hot ductility, and therefore the content of Mg is set to 0 to 0.005%. In order to reliably obtain the effect of Mg addition, the lower limit may be set to 0.0002%. The preferable upper limit in the case of containing Mg is 0.003%.

P：

P is segregated in grain boundaries to increase the grain boundary strength and improve fracture characteristics, and therefore, it may be contained if necessary, but if it exceeds 0.01%, a harmful phase is formed to inhibit hot workability and high-temperature corrosion resistance, and therefore the content of P is set to 0 to 0.01%. In order to reliably obtain the effect of adding P, the lower limit may be set to 0.0002%. The preferable upper limit when P is contained is 0.005%.

Nb：

The Nb component is added as needed because it has the effect of improving the high-temperature tensile properties, low cycle fatigue properties, and fracture properties by being dissolved in the matrix (γ phase) and γ' phase in a solid state, and imparting high-temperature strength, but if the content thereof exceeds 1.0%, the hot workability is lowered, which is not preferable. Therefore, the content of Nb is set to 0 to 1.0%. In order to reliably obtain the effect of adding Nb, the lower limit may be set to 0.005%. The preferable upper limit of Nb content is 0.2%.

Ta：

The Ta component is added as needed because it has the effect of improving the high-temperature tensile properties, low cycle fatigue properties, and fracture properties by being dissolved in a matrix (γ phase) and γ' phase in a solid state as in Nb, and brings about high-temperature strength, but if the content thereof exceeds 1.0%, the hot workability is lowered, which is not preferable. Therefore, the content of Ta is set to 0 to 1.0%. In order to reliably obtain the effect of adding Ta, the lower limit may be set to 0.002%. The preferable upper limit when Ta is contained is 0.2%.

Fe：

The Fe component is inexpensive and economical, and has an effect of improving hot workability, and therefore, it may be added as needed, but if the content exceeds 2.0%, the high-temperature strength is deteriorated, which is not preferable. Therefore, the content of Fe is set to 0 to 2.0% or less. In order to reliably obtain the effect of adding Fe, the lower limit may be set to 0.01%, and more preferably 0.02%. The preferable upper limit when Fe is contained is 0.6%.

M₆Solid solution temperature T1 of C-type carbide and solid solution temperature T2 of γ' phase:

in the present invention, an aircraft engine casing is produced by hot forging an ingot made of a Ni-based super heat resistant alloy to obtain a forged strip blank, and then further repeatedly hot working such as hot forging and hot rolling the forged strip blank. The high-strength γ - γ 'Ni-based superalloy has high deformation resistance when the hot working temperature is equal to or less than the solution temperature T2 of the γ' phase, and therefore, hot working may be performed at a temperature exceeding the solution temperature T2 of the γ 'phase, but when heated at a temperature slightly lower than or equal to the solution temperature T2 of the γ' phase, partial or complete grain growth occurs, which may cause unevenness in the metallographic structure or strength characteristics. In particular, in terms of improving the low cycle fatigue characteristics, it is preferable that the crystal grains are fine, but even if the grains are distributed evenly as fine grains, if coarse grains are present locally, a phenomenon occurs in which the coarse grains greatly affect the number of fracture cycles.

Therefore, in order to obtain a uniform metallographic structure, it is preferable to use M₆The solid solution temperature T1 of the C-type carbide is set in the vicinity of the solid solution temperature T2 of the γ' phase to pass through M₆The dispersion of the C-type carbide suppresses the tendency of the grain growth to increase with the disappearance of the γ' phase.

Further, we have repeatedly conducted various studies and found that: m calculated from thermodynamic equilibrium calculation based on the composition of Ni-based superalloy by, for example, CALPHAD method₆When the difference (T1-T2) between the solid solubility temperature T1 of the C-type carbide and the solid solubility temperature T2 of the gamma' phase is equal to or lower than-30 ℃ (T1-T2) or lower than +40 ℃, preferably equal to or lower than-20 ℃ (T1-T2) or lower than +20 ℃), cracks in plastic working are not easily generated, a metallurgical structure with uniform grain size and precipitation can be obtained even at a low working ratio, and M is a rare earth metal₆The dispersion of the type C carbide not only has an effect of suppressing the grain growth, but also has an effect of serving as a site for recrystallization and facilitating recrystallization even under conditions where the apparent work ratio is small.

However, in the hot working temperature region, M is in the range of (T1-T2) < -30 ℃₆The C-type carbide does not bring about the effect of compensating for the disappearance of the gamma' phase, and in the case of +40 ℃ < (T1-T2), M is₆Since the C-type carbide is coarse and has a strong tendency to be unevenly dispersed and the effect is reduced, it is preferable to satisfy-30 ℃ or less (T1-T2) or less and +40 ℃.

Further preferably satisfies the condition of-20 ℃ or lower (T1-T2) or lower +20 ℃. A more preferable lower limit is-15 ℃ and an upper limit is +10 ℃.

Accordingly, in the Ni-based superalloy of the present invention, the grain size number in ASTM may be 6 or more, preferably 7 or more, more preferably 8 or more, and still more preferably 9 or more. In addition, in the processed product such as an aircraft engine casing of the Ni-based superalloy of the present invention, the crystal grain size is preferably 30 μm or less, more preferably 20 μm or less, and still more preferably 10 μm or less.