CA2060884A1 - Dual alloy turbine blade - Google Patents
Dual alloy turbine bladeInfo
- Publication number
- CA2060884A1 CA2060884A1 CA002060884A CA2060884A CA2060884A1 CA 2060884 A1 CA2060884 A1 CA 2060884A1 CA 002060884 A CA002060884 A CA 002060884A CA 2060884 A CA2060884 A CA 2060884A CA 2060884 A1 CA2060884 A1 CA 2060884A1
- Authority
- CA
- Canada
- Prior art keywords
- single crystal
- core
- turbine blade
- attachment section
- polycrystalline
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A composite turbine blade (20) having a single crystal airfoil section (22), a single crystal platform (26), and a composite attachment section (24). The attachment section (24) is comprised of a thin layer (34) of single crystal material overlying and metallurgically bonded, along interfacial bond lines (36), to a core (30) made of a fined grained, polycrystalline superalloy. The layer (34) has an external configuration having ridges (27) and grooves (28) for removably attaching to a complementary groove in a turbine disk. The blade is prepared by casting a single crystal body with a cavity within the attachment section (24), and then filling the attachment section with the polycrystalline superalloy to form a composite structure. Filling is preferably accomplished by plasma spraying the cavity with the superalloy, and hot isostatically compacting the sprayed superalloy to minimize porosity. The composite structure is then heat treated to develop an optimized microstructure in the dual alloy attachment section (24). The resulting turbine blade (20) has improved life resulting from reduced low cycle fatigue susceptibility of the composite attachment section (24).
Description
DUAL ALLOY _TURB I NE BLADE
TECHNICAL FIELD
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This inventiQn relates generally to gas turbine power plants, and, more particularly, to turblne blades used in high performance gas turbine engines.
BACKGROUN~ OF THE INVENTION
Gas turbine power plants are used as the primary propulsive power source for aircraft, in the forms of jet engines and turboprop engines, as auxiliary power sources for driving air compressors, hydraulic pumps, etc. on aircraft, and as stationary power supplies such as backup electrical generators for hospitals and the like. The same basic power generation principles apply for all of these types of gas turbine power plants.
lS Compressed air i~ mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turhine vanes in the engine. The vanes turn the ~high velocity gas flow partially sideways to impinge upon turbine blades mounted on a turbine disk or wheel that is free to rotate.
The force of the impinging gas causes the turbine disk to sp~n at high speed. Jet propulsion engines use this power to draw more air into the engine and then high velocity combustion ga~ is passed out the aft end of the gas turbine, creating forward thrust. Other engines use this power to turn a propeller or an electric generator.
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~; Thè turbine blades and vanes lie at the héart of ~ . , the power plant, and ~t is well established that in most , ' , .
~ 2-ca~ses, they are one of the limiting factors in achieving improved power plant efficiency. In particular, because they are subjected to high heat and stress loadings as they are rotated and impacted by the hot gas, there is a continuing effort to identify improvements to the construction and/or design of turbine blades to achieve ever higher performance.
Much research and engineering has been directed to the problem of improved turbine blade materials. The earliest turbine blades were made of simple cast alloys having relatively low maximum operating temperatures.
The alloy materials have been significantly improved over a period of years, resulting in various types of nickel-based and cobalt-based superalloys that are in use today.
As the alloy materials were improved, the metallurgical microstructure of the turbine blades was also improved. First, the polycrystalline grain structures were modified by a wide variety of treatments to optimize their performance. Directionally solidified ôr orlented polycrystalline blades were then developed, having elongated grains with deformation-resistant orientations parallel to the radial axis of the blade in order to best resist the centrifugal stresses. Each of the~e advancements led to improved performance of the blades. Polycrystalline and oriented polycrystalline '-~ blade are widely used in most commercial and many ~ m,ilitary alrcraft engine~ today.
, -~ - It 'has been proposed to improve polycrystalline ,~
30 blades by including reinforcing ceramic fibers or the ~, like in the structure but such approaches ;have not met with success primarily because of the ~problems in .
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adeguately bonding such differing materials so that operating stresses are evenly distributed.
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More recently, single crystal turbine blades have been introduced as a result of the development of practical techniques to cast them in large qUantities.
These turbine blades have the advantage of eliminating grain boundaries entirely, which are one of the important causes of creep deformation and failure of the airfoil. The elimination of grain boundaries allows the chemical composition of the single crystal blade to be ad~usted to achieve improved creep and high-cycle fatigue performance at the highest engine operating temperatures. Single crystal turbine blades are now used in high performance military aireraft and may eventually be intFoduced into commercial applications.
Whlle the single crystal turbine blades have provided improved overall airfoil performance as compared with polycrystalline blades, they still exhibit problem areas. In many applications, the highly loaded attachment area is subject to low cycle fatigue failures. As a result, there is a continuing need to provide yet further improvements to achieve higher operati:ng temperatures and lengthened operating lives in the blades used in high performance gas turbine engine~.
It i~ therefore an object of the present invention to provide a novel turbine blade, and method of making samé, whlch has an increased operating life.
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Another obiect of the invention is to provide a single~ crystal turbine blade having a reduced susceptibility to fallure in its attachment area.
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A further ob~ect of the invention is to provide a composite structure in at least a portion of the attachment section of a single crystal turbine blade to retard creep and/or crack growth in sald portion.
SUMMARY OF THE INVENTION
The present invention resides in an improved gas turbine blade that utilizes a sinqle crystal alloy body optimized for high temperature performance of the airfoil section, with a reinforcing polycrystalline alloy core within the interior of at least a portion of the attachment or root section in order to form a composite structure. The resulting turbine blade is physically interchangeable with prior blades, but has improved strength, stiffness and low cycle fatigue resistance in the attachment section.
While turbine blade is a unitary structure, it may be conveniently de~cribed ac having two sections: an airfoil section and an attachment or root section. The airfoil sectlon is elongated and curved slightly into a shape suitable for reacting against the flow of the hot combustion gas. The root section attaches the airfoil section to the rotatable turbine disk or hubo The most widely used attachment is a "firtree" shape, wherein the attachment section of the blade has a serie~ of enlarged ridge~ that fit into a conforming receptacle in the rim of the turbine disk. The blade is held in place by the ` ~physical interlocking of the ridges and the receptacle, yet is reiatively easy to insert and remove when necessary.
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The airfoil section of the turbine blade is subjected to a combination of stresses induced by centrifugal forces and hot gas impingement. Centrifugal forces induce slow creep deformation and, if rotational speeds are high enough, failure by stress rupture. Hot gas impingement combined with centrifugal loading can lead to high-cycle (low-amplitude strain) fatigue. The single crystal alloys have been .optimized to resist these mechanisms of failure. However, it has been observed that the attachment section is susceptible to another, completely different failure mechanism: low cycle (high amplitude strain) fatigue. Existing single crystal turbine blades have their lives limited in many cases, by this low cycle fatigue mode. Because the tur~ine. blade single crystal alloy is optimized to resist other failure mechanisms, low cycle fatique failure of the attachment section becomes a more prominent concern in high performance gas turbine engines.
While the inventor does not wish to be held to any particular theory, it ls believed that the source of . the low cycle fatigue performance improvement arises .- from the inherent differences between the lower modulus - : single crystal and higher modulus polycrystalline micro tructures. Low ..cycle fatigue occurs under .. condltions .of high cyclic load and the .related large plastic strai~ns.~. The absence of.grain boundaries in the : slngle crystal~ materiaI has the effect of increasing the .straln at any given stre~ and eliminating a major ;. .30 .. ..mlcrostructural restraint to the growth of micro cracks ~whlch are formed.during high plastic.strain. .The fine . grained polycrystal}ine core material.is much stiffer . and~stherefore ::attracts a larger :share of .the radial . brood belng transferred through the blade. This reduces ~ 35 the ~critical stresses in the softer single crystal :
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material of the attachment areas which increases the low cycle fatigue life of the composite blade.
In accordance with the present invention, a turbine blade comprises a low modulus, cast single S crystal body having an airfoil section and an attachment section, and a higher modulus structural core of a polycrystalline alloy bonded within said attachment section.
The turbine blade of the present invention has a single crystal body having a composition, orientation, and structure optimized to provide excellent creep and high-cycle fatigue resistance in the airfoil section.
This blade is grown by existing single crystal growth techniques, such as those reported in U.S. Patents Nos.
4,412,577 and 3,494,709, whose disclosures are incorporated herein by reference. However, the blade is grown with the attachment section containing a hollow cavity. Alternately, a cavity may be later machined into the blade.
, A core of a polycxystalline superalloy is applied within the center of the attachment section. The thickness, composition and microstructure of the core are optimized to be resistant to low cycle, moderate temperature fatigue damage and other failure mechanisms that are predominant in the attachment section. The entire attachment section is preferably not made of the polycrystalline material. The lower-modulus sin~le-cry~tal material receives the airfoil attachmen~
load from the stiffer, higher-modulus, polycrystalline core. Notch-root stres~es are minimized in the single crystal material by the support provided by the high-modulu~ core. Reduced notch-sensitivity ~is also .
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achieved by the use of the low-modulus single-crystal material.
The polycrystalline core can be applied by any number of techniques, but prefe~ably by plasma spraying. The core material can then be metallurgically refined .to improve the microstructure to be more resistant to failure, for example by hot isostatic pressing or heat treating.
In accordance with the processing aspect of the present invention, a .process for preparing a turbine blade generally comprises the steps of casting a single crystal body having an airfoil section and an attachment section, forming a cavity within the core of the attachment section, reinforcing the core of the attachment section by filIing the cavity with a polycrystalline alloy, metallurgically refining the polycrystalline core and, finally, machining the attachment section into a desired final configuration for attachment to a turbine disk. In a preferred approach, a process for preparing a turbine blade comprises the steps of casting a single crystal body .having an.airfoil section and an attachment section, plasma spraying a hlgh strength..polycrystalline alloy : : .into ;a.core cavity formed within the control portion of 25~ the~ attachment section, and hot isostatic pressing the body to~con-olldat- the polycrystalline alloy core.
: : In the most preferred approach, the single c nstal portion of the ..blade is of SC180 composition super~alloy (described in EPO Patent Appln. No. 246~082) ;30~ having a [OOl] crystallographic orientation parallel to the blade's longitudinal.axis. The polycrystalline core ls~preferably of U=270 superalloy since its composition P~ is compatible to SCl80. The polycrystalline core is ~, ' ' .; ' , applied by vacuum plasma spray deposition and then consolidated by hot isostatic pressing, so that the core is dense and welI bonded to the sin~le crystal portion of the attachment section.
It should be appreciated that the turbine blade of the invention achieves improved performance and life by incorporating the best features of two different approaches, while minimizing the detractions of each.
Optimized airfoil section performance is attained by using an optimized single crystal alloy, and optimized attachment section performance is attained by using an optimized polycrystalline alloy in the core to provide additional strength. This composite structure behaves in a complex fa~hion which is not entirely predictable by only considering the individual properties of the single crystal material or the polycrystalline material. Initially the single crystal layer resists the centrifugal stresses but after some small amount of creep, the stresses are transferred into the stronger polycrystalline core. Other features and advantages of the present invention will be apparent from the following more detailed description of a presently preferred embodiment, taken in conjunction with the accompanying drawlngs, which illustrate, by way of example and not limitation, the principles of the invention.
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BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a perspective view of a conventlonal single crystal turbine blade;
-FIG. 2 is a partial perspective view of a single crystal turbine blade of the present invention; and '~, .
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_g_ FIG. 3 is an enlarged sectional view of the attachment region o the blade shown ln FIG. 2.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
By way of background, FIG. 1 illustrates a prior S single crystal turbine blade (10). The blade (10) has an airfoil section (12), an attachment or root section (14), and, usually, a platform or stabilizer (16) between the two sections. The attachment section (14) has the pattern of alternating ridges (17) and depressions (18) ~hat form a "firtree" shape for removable attachment to complementary grooves in a turbine disk (not shown). The blade (10) is ~abricated entirely of a piece of single crystal superalloy, preferably with a [001] crystallographic direction parallel to the blade's longitudinal axis.
As used herein, a single crystal article is one in which substantially all of the article ha~ a single crystallographic orientation through the load bearing portions, without the presence of high angle grain boundaries. A small amount of low angle grain boundaries, such aq tilt or twist boundaries, are permitted within such a single crystal article, but are preferably not present. However, such low angle boundaries are often present after solidification and 25 ~formation ~of the single crystal article, or after some deformation of the article durlng creep or other light -- deformation process. Other minor irregularities are also permitted within the scope of the term "single crystal". For example, small areas of high angle grain boundaries may be formed in various portion~ of the article, due to the inability of the single crystal to grow perf~ctly near corners and the like. Such deviations -Srom a perfect single crystal, which are ., -found in normal commercial production operations are within the scope of the term "single crystal" as used herein.
FIG. 2 illustrates a dual alloy, dual structure S turbine blade (20), which also has an airfoil section (22), an attachment section (24), and a platform or stabilizer (26). The attachment section (24) has a firtree of the same outward configuration and dimensions as the firtree of the prior blade (10). The physical appearance and configuration of the blade (20) may be identical with that of a prior blade (10), 90 that the improved blade can directly replace the prior blade in existing turbine wheels.
From the enlarged cross-sectional illustration of lS FIG. 3, however, it is apparent that the structure of the blade (20) differs from that of the blade (10). The airfoil sections (12) and (22) are identical, but the attachment sections (14) and (24) are not metallurgically identical. ~he attachment section (24) is formed with an~polycrystalline core (30) that extends from the base of the blade up towards the platform (26) beyond the firtree. The core (30) is preferably formed of a size ~ust smaller than the entire attachment sect$on (14) but large enough to provide re$nforcement thereto. The core (30) preferably tapers sufficiently ~-~ to form a mechanical interlocking structure with the outer layer;of single crystal material. Overlying the ~~core (30) $s at least a th$n layer of the single crystal mater$al -(34). The layer (34) has its external conf$gurat$on mach$ned with the same ridges (27) and grooves (28) as the prlor art blade (10).
The polycrystail$ne metallic alloy core (30) must be metallurg$cally bonded to the single crystal along . : : ' , - . ' . . -., ~
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The single crystal material may be any acceptable superalloy that can be prepared as a single crystal.
The preferred single crystal materials are those that have compositions tailored to yield optimal high temperature properties in the single crystal airfoil section (22) but have a relatively low modulus in the transverse [100] grain direction. The most preferred single crystal material is an alloy known as SC180, disclo~ed in European Patent Application No. 246,082.
In its most preferred form SC180 has a nominal composition of about 10~ Co, 5% Cr, 1.7~ Mo, 53 W, 8.5%
Ta, 5.2% Al, 3% Re, 1.0% Ti, 0.1% Hf and the balance nickel. Its modulus is relatively low at about 14.8 x 106 cm/cm. The crystalllne orientation of the single crystal is preferably with the [001] direction parallel to the blade's lonqitudinal axis. Other acceptable single crystal materials are well known in the art.
See, for example, U.S. Patents Nos. 4,582,S48;
4,643,782; and 4,719,080.
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The polycrystalline material for u~e in the core (30) may be any acceptable superalloy that can be prepared with a fine grain. The preferred polycrystalline- materials are tho~e that have composition~, grain sizes, and processing optimized to yield ~maxlmum performance as -an ~attachment- section alloy.~ ~This criterion implieslan alloy having high strength and excell~nt low cycle fatigue performance.
The most preferred polycrystalline material is U-720 which:haQ a nominal composition of about 14.5% Co, 18.0~
Cr,;3.0% Mo,-1.2% W, 2.5% Al, 5.0% Ti and minor amounts of B, C, and Zr in a nickel matrix. This alloy has a . ~ ~ relatively high modulus of about 28.2 x 106 cm/cm. In ~: : :
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?~a addition, the chemical composition is similar enough to SCl80 to minimize phase instability near the interfacial bond line (36). Other acceptable polycrystalline superalloys include, but are not limited to well-known wrought disk alloys such as those sold under the trademarks or tradenames MAR M-247, Waspoloy, IN-100, and Astroloy.
The turbine blade of the invention is fabricated by first casting a single crystal piece having the shape of the airfoil section (22~, platform (2~), and preferably a channel or cavity for the tapered core (30~
in the attachment section (24). If the cavity is not formed during the casting process, it may later be electrochemically machined into the solid attachment lS section (14). A more preferred process is to initially cast a small undersized cavity in the blade and then later machine the cavity to a desired final size and shape to ensure greater uniformity in production blades.
Any fabrication technique which produces a substantially slngle crystal article is operable in con~unction with the present invention. The preferred ~technique, u~ed to prepare the single crystal articles descri~ed herein, is the high thermal gradLent solldlflcatlon method. : Molten metal of the desired -25 composltion is placed into a heat resistant ceramic mold : having essentially the desired shape of the final : fabricated component. -~he mold and metal contained thereln ~are~placed withln a furnace, induction heating ~coil,~ or other heating device to melt the metal, and the : 30 mold -and molten metal are gradually cooled in a - controiled temperature gradient. In this process, metal ad~acent the-cooler end of the mold solldifies first, and the lnterface -between the solidified and liquid metal gradually moves through the metal as cooling : ~
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' It is known that certain preferred crystallographic orientations such as 1001~ can be grown to the exclusion of others during such a gradient soIidification process, so that a single grain becomes dominant throughout the article. Techniques have been developed to promote the formation of the single crystal orientation rapidly, so that substantially all of the article has the same single crystal orientation. Such techniques include seeding, described in U.S. Patent No.
4,412,577, whereby an oriented ~ingle crystal starting material is positioned ad~acent the metal first solidlfied, so that the metal initially develops that orientation. Another' approach is a geometrical selection proces~ such as described in U.S.'I''Patent No.
3,494,709.
`'As'lndicated, all' other techniques for forming a single crystal are acceptable for-use in con~unction ~With the ~present invention. The floating'zone technique mayS~bë~used~wherein a molten zone is passed through a polycr'ystalline piece' of 'metal-'"to produce' a- movinq solidificatlon 'front. Solid-state techniques are also permltted wherein a solid piece of polycrystalline 30~ ~material i~ transformed to a single crystal in the solid s~te.~ The-~-solid 'state' approach is not preferred becau~ë it 'ls'typically slow andSproduces a relatively 'impêrfect`single crystal. ~' ~ '--' - ' , ~
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_ The polycrystalline core (30) is applied by any technique that produces a sound microstructure that is well bonded to the underlying single crystal substrate.
The preferred approach is vacuum plasma spray deposition. The target to be coated, here the tapered cavity of the blade (20), is placed into a vacuum chamber which is evacuated to a relatively low pressure. A plasma gun that melts metal fed thereto is aimed at the target substrate, typically positioned several centimeters from the plasma gun. Particles of metal of the desired final composition are fed to the plasma gun, which melts, ox at least softens, the particles and propels them toward the target to impact thexeupon. Different blends of particles can also be used, but a single particulate feed material is preferred for uniformity.
The plasma deposition process is continued for as long as necessary to fill up the core caYity. By way of example and not of limitation, a typical blade (20) may be 5 to 10 centimeters long, and the depth of the core (30) may be about 1.3 to 3.8 centimeters.
Such a blade was analyzed and calculated to have about 10~ less stress in the attachment grooves (28) which would increase~the low cycle fatigue life of ~he attachment section by a factor of about 2. Of course other blade designs will have to be analyzed to determine the optimum proportions for the core and the amount of increasod li~e provided thereby.
The as-deposited core may have a slight degree of 3~ porosity and possibly unmelted particles. To remove the porosity and-irregularities, the blade (20) is placed into a pressure chambex and hot i~ostatically pressed.
The hot isostatic pressing is conducted at an elevated .-. . , . . ........ - . ~ . . .
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pressure, typically 1034 to 1724 bars, and at anelevated temperature, typically 1080C to 1221C, for a sufficient time, such as 4 hours. The exact temperature and time may vary depending upon heat treatment requirements for the single crystal and the core materials. An acceptable and preferred hot isostatic pressing treatment is 1221C and 1034 bars for 4 hou~s.
Upon completion of this treatment the porosity in the core should be completely closed, with good bonding at the bond line (36). After pressing, the composite blade is preferably solution heat-treated and aged at about 649C to 1260C (more preferably 760C to 871C) to optimize the polycrystalline microstructure. Care must be taken to avoid incipient melting of the single crystal material, and the appropriate combination of pressing and heat treatment parameters will depend upon the materials selected for the single crystal and polycrystalline core in any particular case.
Any other acceptable procedure may also be used to fill the single crystal cavity with the polycrystalline material. Such other techniques include, but are not limited to, vapor deposition, plasma transfer arc, electrodeposition, deposition from solution, and powder spraying.
As should now be appreciated, the turbine blades of the invention provide an improved dual alloy composlte structure and therefore improved performance ` compared to prior blades. Although a particular embodiment of the invention has been described in detail for purposes of illustration, various modifications may be made without departing from the spirit and scope of the invention. For example, some stationary vanes or other components in a gas turbine engine may experience attachm nt problems which could be solved by adding a : ~
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TECHNICAL FIELD
.
This inventiQn relates generally to gas turbine power plants, and, more particularly, to turblne blades used in high performance gas turbine engines.
BACKGROUN~ OF THE INVENTION
Gas turbine power plants are used as the primary propulsive power source for aircraft, in the forms of jet engines and turboprop engines, as auxiliary power sources for driving air compressors, hydraulic pumps, etc. on aircraft, and as stationary power supplies such as backup electrical generators for hospitals and the like. The same basic power generation principles apply for all of these types of gas turbine power plants.
lS Compressed air i~ mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turhine vanes in the engine. The vanes turn the ~high velocity gas flow partially sideways to impinge upon turbine blades mounted on a turbine disk or wheel that is free to rotate.
The force of the impinging gas causes the turbine disk to sp~n at high speed. Jet propulsion engines use this power to draw more air into the engine and then high velocity combustion ga~ is passed out the aft end of the gas turbine, creating forward thrust. Other engines use this power to turn a propeller or an electric generator.
~ . .. . . ....... .. . .
~; Thè turbine blades and vanes lie at the héart of ~ . , the power plant, and ~t is well established that in most , ' , .
~ 2-ca~ses, they are one of the limiting factors in achieving improved power plant efficiency. In particular, because they are subjected to high heat and stress loadings as they are rotated and impacted by the hot gas, there is a continuing effort to identify improvements to the construction and/or design of turbine blades to achieve ever higher performance.
Much research and engineering has been directed to the problem of improved turbine blade materials. The earliest turbine blades were made of simple cast alloys having relatively low maximum operating temperatures.
The alloy materials have been significantly improved over a period of years, resulting in various types of nickel-based and cobalt-based superalloys that are in use today.
As the alloy materials were improved, the metallurgical microstructure of the turbine blades was also improved. First, the polycrystalline grain structures were modified by a wide variety of treatments to optimize their performance. Directionally solidified ôr orlented polycrystalline blades were then developed, having elongated grains with deformation-resistant orientations parallel to the radial axis of the blade in order to best resist the centrifugal stresses. Each of the~e advancements led to improved performance of the blades. Polycrystalline and oriented polycrystalline '-~ blade are widely used in most commercial and many ~ m,ilitary alrcraft engine~ today.
, -~ - It 'has been proposed to improve polycrystalline ,~
30 blades by including reinforcing ceramic fibers or the ~, like in the structure but such approaches ;have not met with success primarily because of the ~problems in .
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- . - . . .
- - . ~ .
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adeguately bonding such differing materials so that operating stresses are evenly distributed.
. .
More recently, single crystal turbine blades have been introduced as a result of the development of practical techniques to cast them in large qUantities.
These turbine blades have the advantage of eliminating grain boundaries entirely, which are one of the important causes of creep deformation and failure of the airfoil. The elimination of grain boundaries allows the chemical composition of the single crystal blade to be ad~usted to achieve improved creep and high-cycle fatigue performance at the highest engine operating temperatures. Single crystal turbine blades are now used in high performance military aireraft and may eventually be intFoduced into commercial applications.
Whlle the single crystal turbine blades have provided improved overall airfoil performance as compared with polycrystalline blades, they still exhibit problem areas. In many applications, the highly loaded attachment area is subject to low cycle fatigue failures. As a result, there is a continuing need to provide yet further improvements to achieve higher operati:ng temperatures and lengthened operating lives in the blades used in high performance gas turbine engine~.
It i~ therefore an object of the present invention to provide a novel turbine blade, and method of making samé, whlch has an increased operating life.
: : :
Another obiect of the invention is to provide a single~ crystal turbine blade having a reduced susceptibility to fallure in its attachment area.
::
. , .
-: . , . : . .
A further ob~ect of the invention is to provide a composite structure in at least a portion of the attachment section of a single crystal turbine blade to retard creep and/or crack growth in sald portion.
SUMMARY OF THE INVENTION
The present invention resides in an improved gas turbine blade that utilizes a sinqle crystal alloy body optimized for high temperature performance of the airfoil section, with a reinforcing polycrystalline alloy core within the interior of at least a portion of the attachment or root section in order to form a composite structure. The resulting turbine blade is physically interchangeable with prior blades, but has improved strength, stiffness and low cycle fatigue resistance in the attachment section.
While turbine blade is a unitary structure, it may be conveniently de~cribed ac having two sections: an airfoil section and an attachment or root section. The airfoil sectlon is elongated and curved slightly into a shape suitable for reacting against the flow of the hot combustion gas. The root section attaches the airfoil section to the rotatable turbine disk or hubo The most widely used attachment is a "firtree" shape, wherein the attachment section of the blade has a serie~ of enlarged ridge~ that fit into a conforming receptacle in the rim of the turbine disk. The blade is held in place by the ` ~physical interlocking of the ridges and the receptacle, yet is reiatively easy to insert and remove when necessary.
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The airfoil section of the turbine blade is subjected to a combination of stresses induced by centrifugal forces and hot gas impingement. Centrifugal forces induce slow creep deformation and, if rotational speeds are high enough, failure by stress rupture. Hot gas impingement combined with centrifugal loading can lead to high-cycle (low-amplitude strain) fatigue. The single crystal alloys have been .optimized to resist these mechanisms of failure. However, it has been observed that the attachment section is susceptible to another, completely different failure mechanism: low cycle (high amplitude strain) fatigue. Existing single crystal turbine blades have their lives limited in many cases, by this low cycle fatigue mode. Because the tur~ine. blade single crystal alloy is optimized to resist other failure mechanisms, low cycle fatique failure of the attachment section becomes a more prominent concern in high performance gas turbine engines.
While the inventor does not wish to be held to any particular theory, it ls believed that the source of . the low cycle fatigue performance improvement arises .- from the inherent differences between the lower modulus - : single crystal and higher modulus polycrystalline micro tructures. Low ..cycle fatigue occurs under .. condltions .of high cyclic load and the .related large plastic strai~ns.~. The absence of.grain boundaries in the : slngle crystal~ materiaI has the effect of increasing the .straln at any given stre~ and eliminating a major ;. .30 .. ..mlcrostructural restraint to the growth of micro cracks ~whlch are formed.during high plastic.strain. .The fine . grained polycrystal}ine core material.is much stiffer . and~stherefore ::attracts a larger :share of .the radial . brood belng transferred through the blade. This reduces ~ 35 the ~critical stresses in the softer single crystal :
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material of the attachment areas which increases the low cycle fatigue life of the composite blade.
In accordance with the present invention, a turbine blade comprises a low modulus, cast single S crystal body having an airfoil section and an attachment section, and a higher modulus structural core of a polycrystalline alloy bonded within said attachment section.
The turbine blade of the present invention has a single crystal body having a composition, orientation, and structure optimized to provide excellent creep and high-cycle fatigue resistance in the airfoil section.
This blade is grown by existing single crystal growth techniques, such as those reported in U.S. Patents Nos.
4,412,577 and 3,494,709, whose disclosures are incorporated herein by reference. However, the blade is grown with the attachment section containing a hollow cavity. Alternately, a cavity may be later machined into the blade.
, A core of a polycxystalline superalloy is applied within the center of the attachment section. The thickness, composition and microstructure of the core are optimized to be resistant to low cycle, moderate temperature fatigue damage and other failure mechanisms that are predominant in the attachment section. The entire attachment section is preferably not made of the polycrystalline material. The lower-modulus sin~le-cry~tal material receives the airfoil attachmen~
load from the stiffer, higher-modulus, polycrystalline core. Notch-root stres~es are minimized in the single crystal material by the support provided by the high-modulu~ core. Reduced notch-sensitivity ~is also .
..
.
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achieved by the use of the low-modulus single-crystal material.
The polycrystalline core can be applied by any number of techniques, but prefe~ably by plasma spraying. The core material can then be metallurgically refined .to improve the microstructure to be more resistant to failure, for example by hot isostatic pressing or heat treating.
In accordance with the processing aspect of the present invention, a .process for preparing a turbine blade generally comprises the steps of casting a single crystal body having an airfoil section and an attachment section, forming a cavity within the core of the attachment section, reinforcing the core of the attachment section by filIing the cavity with a polycrystalline alloy, metallurgically refining the polycrystalline core and, finally, machining the attachment section into a desired final configuration for attachment to a turbine disk. In a preferred approach, a process for preparing a turbine blade comprises the steps of casting a single crystal body .having an.airfoil section and an attachment section, plasma spraying a hlgh strength..polycrystalline alloy : : .into ;a.core cavity formed within the control portion of 25~ the~ attachment section, and hot isostatic pressing the body to~con-olldat- the polycrystalline alloy core.
: : In the most preferred approach, the single c nstal portion of the ..blade is of SC180 composition super~alloy (described in EPO Patent Appln. No. 246~082) ;30~ having a [OOl] crystallographic orientation parallel to the blade's longitudinal.axis. The polycrystalline core ls~preferably of U=270 superalloy since its composition P~ is compatible to SCl80. The polycrystalline core is ~, ' ' .; ' , applied by vacuum plasma spray deposition and then consolidated by hot isostatic pressing, so that the core is dense and welI bonded to the sin~le crystal portion of the attachment section.
It should be appreciated that the turbine blade of the invention achieves improved performance and life by incorporating the best features of two different approaches, while minimizing the detractions of each.
Optimized airfoil section performance is attained by using an optimized single crystal alloy, and optimized attachment section performance is attained by using an optimized polycrystalline alloy in the core to provide additional strength. This composite structure behaves in a complex fa~hion which is not entirely predictable by only considering the individual properties of the single crystal material or the polycrystalline material. Initially the single crystal layer resists the centrifugal stresses but after some small amount of creep, the stresses are transferred into the stronger polycrystalline core. Other features and advantages of the present invention will be apparent from the following more detailed description of a presently preferred embodiment, taken in conjunction with the accompanying drawlngs, which illustrate, by way of example and not limitation, the principles of the invention.
.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a perspective view of a conventlonal single crystal turbine blade;
-FIG. 2 is a partial perspective view of a single crystal turbine blade of the present invention; and '~, .
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_g_ FIG. 3 is an enlarged sectional view of the attachment region o the blade shown ln FIG. 2.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
By way of background, FIG. 1 illustrates a prior S single crystal turbine blade (10). The blade (10) has an airfoil section (12), an attachment or root section (14), and, usually, a platform or stabilizer (16) between the two sections. The attachment section (14) has the pattern of alternating ridges (17) and depressions (18) ~hat form a "firtree" shape for removable attachment to complementary grooves in a turbine disk (not shown). The blade (10) is ~abricated entirely of a piece of single crystal superalloy, preferably with a [001] crystallographic direction parallel to the blade's longitudinal axis.
As used herein, a single crystal article is one in which substantially all of the article ha~ a single crystallographic orientation through the load bearing portions, without the presence of high angle grain boundaries. A small amount of low angle grain boundaries, such aq tilt or twist boundaries, are permitted within such a single crystal article, but are preferably not present. However, such low angle boundaries are often present after solidification and 25 ~formation ~of the single crystal article, or after some deformation of the article durlng creep or other light -- deformation process. Other minor irregularities are also permitted within the scope of the term "single crystal". For example, small areas of high angle grain boundaries may be formed in various portion~ of the article, due to the inability of the single crystal to grow perf~ctly near corners and the like. Such deviations -Srom a perfect single crystal, which are ., -found in normal commercial production operations are within the scope of the term "single crystal" as used herein.
FIG. 2 illustrates a dual alloy, dual structure S turbine blade (20), which also has an airfoil section (22), an attachment section (24), and a platform or stabilizer (26). The attachment section (24) has a firtree of the same outward configuration and dimensions as the firtree of the prior blade (10). The physical appearance and configuration of the blade (20) may be identical with that of a prior blade (10), 90 that the improved blade can directly replace the prior blade in existing turbine wheels.
From the enlarged cross-sectional illustration of lS FIG. 3, however, it is apparent that the structure of the blade (20) differs from that of the blade (10). The airfoil sections (12) and (22) are identical, but the attachment sections (14) and (24) are not metallurgically identical. ~he attachment section (24) is formed with an~polycrystalline core (30) that extends from the base of the blade up towards the platform (26) beyond the firtree. The core (30) is preferably formed of a size ~ust smaller than the entire attachment sect$on (14) but large enough to provide re$nforcement thereto. The core (30) preferably tapers sufficiently ~-~ to form a mechanical interlocking structure with the outer layer;of single crystal material. Overlying the ~~core (30) $s at least a th$n layer of the single crystal mater$al -(34). The layer (34) has its external conf$gurat$on mach$ned with the same ridges (27) and grooves (28) as the prlor art blade (10).
The polycrystail$ne metallic alloy core (30) must be metallurg$cally bonded to the single crystal along . : : ' , - . ' . . -., ~
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3 ,,, ;~; _ the interfacial bond lines (36) without substantial porosity and defects.
The single crystal material may be any acceptable superalloy that can be prepared as a single crystal.
The preferred single crystal materials are those that have compositions tailored to yield optimal high temperature properties in the single crystal airfoil section (22) but have a relatively low modulus in the transverse [100] grain direction. The most preferred single crystal material is an alloy known as SC180, disclo~ed in European Patent Application No. 246,082.
In its most preferred form SC180 has a nominal composition of about 10~ Co, 5% Cr, 1.7~ Mo, 53 W, 8.5%
Ta, 5.2% Al, 3% Re, 1.0% Ti, 0.1% Hf and the balance nickel. Its modulus is relatively low at about 14.8 x 106 cm/cm. The crystalllne orientation of the single crystal is preferably with the [001] direction parallel to the blade's lonqitudinal axis. Other acceptable single crystal materials are well known in the art.
See, for example, U.S. Patents Nos. 4,582,S48;
4,643,782; and 4,719,080.
.
The polycrystalline material for u~e in the core (30) may be any acceptable superalloy that can be prepared with a fine grain. The preferred polycrystalline- materials are tho~e that have composition~, grain sizes, and processing optimized to yield ~maxlmum performance as -an ~attachment- section alloy.~ ~This criterion implieslan alloy having high strength and excell~nt low cycle fatigue performance.
The most preferred polycrystalline material is U-720 which:haQ a nominal composition of about 14.5% Co, 18.0~
Cr,;3.0% Mo,-1.2% W, 2.5% Al, 5.0% Ti and minor amounts of B, C, and Zr in a nickel matrix. This alloy has a . ~ ~ relatively high modulus of about 28.2 x 106 cm/cm. In ~: : :
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?~a addition, the chemical composition is similar enough to SCl80 to minimize phase instability near the interfacial bond line (36). Other acceptable polycrystalline superalloys include, but are not limited to well-known wrought disk alloys such as those sold under the trademarks or tradenames MAR M-247, Waspoloy, IN-100, and Astroloy.
The turbine blade of the invention is fabricated by first casting a single crystal piece having the shape of the airfoil section (22~, platform (2~), and preferably a channel or cavity for the tapered core (30~
in the attachment section (24). If the cavity is not formed during the casting process, it may later be electrochemically machined into the solid attachment lS section (14). A more preferred process is to initially cast a small undersized cavity in the blade and then later machine the cavity to a desired final size and shape to ensure greater uniformity in production blades.
Any fabrication technique which produces a substantially slngle crystal article is operable in con~unction with the present invention. The preferred ~technique, u~ed to prepare the single crystal articles descri~ed herein, is the high thermal gradLent solldlflcatlon method. : Molten metal of the desired -25 composltion is placed into a heat resistant ceramic mold : having essentially the desired shape of the final : fabricated component. -~he mold and metal contained thereln ~are~placed withln a furnace, induction heating ~coil,~ or other heating device to melt the metal, and the : 30 mold -and molten metal are gradually cooled in a - controiled temperature gradient. In this process, metal ad~acent the-cooler end of the mold solldifies first, and the lnterface -between the solidified and liquid metal gradually moves through the metal as cooling : ~
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~ _ }; 'v~`;l;; V~ J._ . . . .. . . , .,, , ~, a continues. Such gradient solidification can be accomplished by placing a chill block adjacent one end of the mold and then turning off the heat source, allowing the mold and molten metal to cool and solidify in a temperature grad~ent. Alternatively, the mold and molten metal can be gradually withdrawn from the heat source.
~, .
' It is known that certain preferred crystallographic orientations such as 1001~ can be grown to the exclusion of others during such a gradient soIidification process, so that a single grain becomes dominant throughout the article. Techniques have been developed to promote the formation of the single crystal orientation rapidly, so that substantially all of the article has the same single crystal orientation. Such techniques include seeding, described in U.S. Patent No.
4,412,577, whereby an oriented ~ingle crystal starting material is positioned ad~acent the metal first solidlfied, so that the metal initially develops that orientation. Another' approach is a geometrical selection proces~ such as described in U.S.'I''Patent No.
3,494,709.
`'As'lndicated, all' other techniques for forming a single crystal are acceptable for-use in con~unction ~With the ~present invention. The floating'zone technique mayS~bë~used~wherein a molten zone is passed through a polycr'ystalline piece' of 'metal-'"to produce' a- movinq solidificatlon 'front. Solid-state techniques are also permltted wherein a solid piece of polycrystalline 30~ ~material i~ transformed to a single crystal in the solid s~te.~ The-~-solid 'state' approach is not preferred becau~ë it 'ls'typically slow andSproduces a relatively 'impêrfect`single crystal. ~' ~ '--' - ' , ~
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_ The polycrystalline core (30) is applied by any technique that produces a sound microstructure that is well bonded to the underlying single crystal substrate.
The preferred approach is vacuum plasma spray deposition. The target to be coated, here the tapered cavity of the blade (20), is placed into a vacuum chamber which is evacuated to a relatively low pressure. A plasma gun that melts metal fed thereto is aimed at the target substrate, typically positioned several centimeters from the plasma gun. Particles of metal of the desired final composition are fed to the plasma gun, which melts, ox at least softens, the particles and propels them toward the target to impact thexeupon. Different blends of particles can also be used, but a single particulate feed material is preferred for uniformity.
The plasma deposition process is continued for as long as necessary to fill up the core caYity. By way of example and not of limitation, a typical blade (20) may be 5 to 10 centimeters long, and the depth of the core (30) may be about 1.3 to 3.8 centimeters.
Such a blade was analyzed and calculated to have about 10~ less stress in the attachment grooves (28) which would increase~the low cycle fatigue life of ~he attachment section by a factor of about 2. Of course other blade designs will have to be analyzed to determine the optimum proportions for the core and the amount of increasod li~e provided thereby.
The as-deposited core may have a slight degree of 3~ porosity and possibly unmelted particles. To remove the porosity and-irregularities, the blade (20) is placed into a pressure chambex and hot i~ostatically pressed.
The hot isostatic pressing is conducted at an elevated .-. . , . . ........ - . ~ . . .
; ~. "-' .... '. ' ,~
,, 3iv~v__ .. . . . .. , . . ... ~ ~ ~ _ _ . .
~ ~J V ~i C~ () 1'~
pressure, typically 1034 to 1724 bars, and at anelevated temperature, typically 1080C to 1221C, for a sufficient time, such as 4 hours. The exact temperature and time may vary depending upon heat treatment requirements for the single crystal and the core materials. An acceptable and preferred hot isostatic pressing treatment is 1221C and 1034 bars for 4 hou~s.
Upon completion of this treatment the porosity in the core should be completely closed, with good bonding at the bond line (36). After pressing, the composite blade is preferably solution heat-treated and aged at about 649C to 1260C (more preferably 760C to 871C) to optimize the polycrystalline microstructure. Care must be taken to avoid incipient melting of the single crystal material, and the appropriate combination of pressing and heat treatment parameters will depend upon the materials selected for the single crystal and polycrystalline core in any particular case.
Any other acceptable procedure may also be used to fill the single crystal cavity with the polycrystalline material. Such other techniques include, but are not limited to, vapor deposition, plasma transfer arc, electrodeposition, deposition from solution, and powder spraying.
As should now be appreciated, the turbine blades of the invention provide an improved dual alloy composlte structure and therefore improved performance ` compared to prior blades. Although a particular embodiment of the invention has been described in detail for purposes of illustration, various modifications may be made without departing from the spirit and scope of the invention. For example, some stationary vanes or other components in a gas turbine engine may experience attachm nt problems which could be solved by adding a : ~
~ - ,,- .
.
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.
.
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`- !` ` ' "` '" - - : -" ' reinforcing core of polycrystalline alloy. -Accordingly, the invention is not to be limited except as by the appended claims. t .
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Claims (10)
1. A composite turbine blade (20) having a single crystal airfoil section (22) and a single crystal platform (26), further comprising a composite attachment section (24) having a layer (34) of single crystal material overlying and metallurgically bonded, along interfacial bond lines (36), to a core (30) made from a polycrystalline alloy, the layer (34) having an exterior surface configured for removably attaching to a complementary groove in a turbine disk.
2. The turbine blade (20) of Claim 1 wherein said polycrystalline core (30) has a greater modulus than said single crystal material.
3. The turbine blade (20) of Claim 1 wherein said core (30) of polycrystalline alloy at least doubles the low cycle fatigue life of the attachment section (24) as compared to a blade (10) of the same size and shape but without such a core.
4. The turbine blade (20) of Claim 1 wherein said polycrystalline alloy is selected from the group consisting of MAR M-247, U-720, IN-100, Astroloy and Waspoloy.
5. The turbine blade (20) of Claim 1 wherein said polycrystalline alloy has been consolidated by hot isostatic pressing.
6. The turbine blade (20) of Claim 1 wherein said core (30) is tapered to mechanically interlock with said single crystal layer (34).
7. The turbine blade (20) of Claim 1 wherein the polycrystalline alloy is U-720.
8. A process for manufacturing a composite turbine blade (20), comprising the steps of:
casting a single crystal body having an airfoil section (22), a platform (26) and an attachment section (24) having an exterior surface configured for removably attaching to a complementary groove in a turbine disk;
forming a cavity within the attachment section (24);
plasma spray-filling the cavity within the attachment section (24) with a polycrystalline alloy to form a core (30); and metallurgically refining the polycrystalline core (30).
casting a single crystal body having an airfoil section (22), a platform (26) and an attachment section (24) having an exterior surface configured for removably attaching to a complementary groove in a turbine disk;
forming a cavity within the attachment section (24);
plasma spray-filling the cavity within the attachment section (24) with a polycrystalline alloy to form a core (30); and metallurgically refining the polycrystalline core (30).
9. The process of Claim 8 wherein said refining step includes hot isostatic pressing followed by heat treating so that the microstructure of the polycrystalline core (30) is consolidated and fine grained.
10. A composite turbine blade (20) made by the process of Claim 8.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US38559889A | 1989-07-25 | 1989-07-25 | |
| US385,598 | 1989-07-25 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2060884A1 true CA2060884A1 (en) | 1991-01-26 |
Family
ID=23522087
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002060884A Abandoned CA2060884A1 (en) | 1989-07-25 | 1990-07-18 | Dual alloy turbine blade |
Country Status (3)
| Country | Link |
|---|---|
| EP (1) | EP0511958A1 (en) |
| CA (1) | CA2060884A1 (en) |
| WO (1) | WO1991001433A1 (en) |
Families Citing this family (25)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2237846B (en) * | 1989-11-09 | 1993-12-15 | Rolls Royce Plc | Rim parasitic weight reduction |
| JP2000512707A (en) * | 1996-06-21 | 2000-09-26 | シーメンス アクチエンゲゼルシヤフト | Rotor of turbine machine having blades mountable in groove and rotor blades |
| US6924046B2 (en) | 2001-10-24 | 2005-08-02 | Siemens Aktiengesellschaft | Rhenium-containing protective layer for protecting a component against corrosion and oxidation at high temperatures |
| EP1524334A1 (en) | 2003-10-17 | 2005-04-20 | Siemens Aktiengesellschaft | Protective coating for protecting a structural member against corrosion and oxidation at high temperatures and structural member |
| EP1783236A1 (en) | 2005-11-04 | 2007-05-09 | Siemens Aktiengesellschaft | Alloy, protecting coating for a component protection against corrosion and oxidation at high temperature and component |
| EP1790743A1 (en) | 2005-11-24 | 2007-05-30 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| EP1790746B1 (en) | 2005-11-24 | 2010-11-10 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| DE502005005646D1 (en) | 2005-12-14 | 2008-11-20 | Siemens Ag | Alloy, protective layer and component |
| EP1806418A1 (en) | 2006-01-10 | 2007-07-11 | Siemens Aktiengesellschaft | Alloy, protective coating for protecting a structural member against corrosion and oxidation at high temperatures and structural member |
| EP1806419B1 (en) | 2006-01-10 | 2009-08-26 | Siemens Aktiengesellschaft | Alloy, protective coating for protecting a structural member against corrosion and oxidation at high temperatures and structural member |
| JP4918806B2 (en) * | 2006-04-06 | 2012-04-18 | 株式会社日立製作所 | Turbine rotor and turbine blade |
| GB0614186D0 (en) * | 2006-07-18 | 2006-08-23 | Rolls Royce Plc | Blades |
| EP2216421A1 (en) | 2009-01-29 | 2010-08-11 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| EP2354260A1 (en) | 2010-01-12 | 2011-08-10 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| EP2345748A1 (en) | 2010-01-14 | 2011-07-20 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| EP2392684A1 (en) | 2010-06-02 | 2011-12-07 | Siemens Aktiengesellschaft | Alloy, protective layer and component |
| US20130337286A1 (en) | 2010-11-02 | 2013-12-19 | Friedhelm Schmitz | Alloy, protective coating, and component |
| EP2474413A1 (en) | 2011-01-06 | 2012-07-11 | Siemens Aktiengesellschaft | Alloy, protective coating and component |
| EP2474414A1 (en) | 2011-01-06 | 2012-07-11 | Siemens Aktiengesellschaft | Alloy, protective coating and component |
| CN103649372B (en) | 2011-07-08 | 2015-10-21 | 西门子公司 | There is the layer system of double-deck metal level |
| EP2557201A1 (en) | 2011-08-09 | 2013-02-13 | Siemens Aktiengesellschaft | Alloy, protective coating and component |
| EP2568054A1 (en) | 2011-09-12 | 2013-03-13 | Siemens Aktiengesellschaft | Alloy, protective coating and component |
| RU2597459C2 (en) | 2011-09-12 | 2016-09-10 | Сименс Акциенгезелльшафт | SYSTEM OF LAYERS WITH DOUBLE METAL MCrAlY-COATING |
| CN104858399B (en) * | 2015-04-17 | 2016-04-20 | 北京航空航天大学 | A kind of functionally gradient single crystal blade material be combined into by two kinds of alloys and preparation method thereof with realize equipment |
| EP3575556A1 (en) * | 2018-06-01 | 2019-12-04 | Siemens Aktiengesellschaft | Turbine blade assembly and method for manufacturing such blades |
Family Cites Families (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB945561A (en) * | 1961-04-27 | 1964-01-02 | Rolls Royce | Improvements in or relating to axial flow compressors |
| CH544217A (en) * | 1971-04-08 | 1973-11-15 | Bbc Sulzer Turbomaschinen | Gas turbine blade |
| US3713752A (en) * | 1971-10-28 | 1973-01-30 | United Aircraft Corp | Composite blade for a gas turbine engine |
| DE2834843A1 (en) * | 1978-08-09 | 1980-06-26 | Motoren Turbinen Union | COMPOSED CERAMIC GAS TURBINE BLADE |
| US4343593A (en) * | 1980-01-25 | 1982-08-10 | The United States Of America As Represented By The Secretary Of The Air Force | Composite blade for turbofan engine fan |
| US4787821A (en) * | 1987-04-10 | 1988-11-29 | Allied Signal Inc. | Dual alloy rotor |
| US4921405A (en) * | 1988-11-10 | 1990-05-01 | Allied-Signal Inc. | Dual structure turbine blade |
-
1990
- 1990-07-18 CA CA002060884A patent/CA2060884A1/en not_active Abandoned
- 1990-07-18 WO PCT/US1990/004032 patent/WO1991001433A1/en not_active Application Discontinuation
- 1990-07-18 EP EP90912062A patent/EP0511958A1/en not_active Withdrawn
Also Published As
| Publication number | Publication date |
|---|---|
| WO1991001433A1 (en) | 1991-02-07 |
| EP0511958A1 (en) | 1992-11-11 |
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