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Publication numberUS3394918 A
Publication typeGrant
Publication dateJul 30, 1968
Filing dateApr 13, 1966
Priority dateApr 13, 1966
Publication numberUS 3394918 A, US 3394918A, US-A-3394918, US3394918 A, US3394918A
InventorsWiseman Robert L
Original AssigneeHowmet Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Bimetallic airfoils
US 3394918 A
Abstract  available in
Images(1)
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Claims  available in
Description  (OCR text may contain errors)

July 3o, 196s FIG. 1

R. L. WISEMAN BIMETALLIC AIRFOILS Filed April 13. 1966 iilrl.

o 0 o ooo l ooo o o o o o n o o o INVENTOR ROBT L. WISEMAN l wird l UA, BY l 1&1 w ALM ATTO RN EYS United States Patent O 3,394,918 BIMETALLIC AIRFOILS Robert L. Wiseman, Westport, Conn., assigner to Howmet Corporation, a corporation of New Jersey Continuation-impart of application Ser. No. 311,532,

Sept. 25, 1963. This application Apr. 13, 1966, Ser. No. 542,377

13 Claims. (Cl. 253-77) This is a continuation-in-part of my earlier filed copending application, Ser. No. 311,532, filed Sept. 25, 1963.

This invention relates to bimetallic airfoils, and more particularly it relates to airfoils which are subject to impingement of hot gas on their surface areas which define hot zones where temperatures of the impinging gases are above about l100 F. and cool zones where the temperatures of the impinging gases are below this temperature and are formed of a first metal composition in the cool zone, and a second metal composition in the cool zone with the rst metal composition having ductile and impact resistant properties in excess of said second metal composition and the -second metal composition having thermal fatigue properties above 1100" F.

Airfoils, such as those used in gas turbines, are subjected to very high temperatures of from about 1100 ll?. to about 2000 F. and have relatively heavy stresses applied for extended periods of time during its operation at very high speeds. After extended operation of a turbine engine, these conditions often have caused bowing of the airfoil. To prevent this distortion in the airfoil, efforts have heretofore been made to find a metal composition for the airfoil which has the properties of high heat resistance and excellent material strength. The alloy compositions which are now being used in the construction of airfoils possess either the high heat resistant properties or the required material strength but no satisfactory alloy has been found which satisfactorily combines both these properties to the extent necessary.

I have found that certain portions of the airfoil are subjected to very high stresses and require an alloy composition with particularly good ductile and impact resistance in that area whereas other portions of airfoil section are subjected to very high temperatures which require an alloy composition along that area which possesses thermal fatigue properties at temperatures in excess of about 1100 F. Heretofore, airfoils have been constructed with roots of different compositions than the airfoil section but this did not prevent distortion of the airfoil section per se to which the variable extremes of heat and stress were applied.

It is accordingly Van object of this invention to provide an integral unitary airfoil structure of two distinct metal compositions, a first possessing ductile and impact resistant properties in excess of the other, and a second possessing thermal fatigue properties above about 1100 F.

Broadly stated, the invention is in a bimetallic airfoil which is subject to impingement of hot gas on its surface areas defining a first zone where the temperature of the impinging gas is above about 1100 F. and a cool zone where the temperature of the impinging gas is below this temperature. The airfoil is comprised of an integral unitary airfoil structure of two metal compositions with a first integral portion composed of a first metal composition and a second integral portion composed of a second metal composition. The second metal composition rice defines the surface area along the hot zone of the airfoil and is fused to the first metal composition by a metallurgical bond along a fuse zone located in the cool zone. The first metal composition has ductile and impact resistant properties in excess of the second metal cornposition and the second metal composition has thermal fatigue properties above ll00 F.

The hot zone varies with the airfoil. With some airfoils it is located at the leading and trailing edges and with others the entire intermediate portion of the airfoil is subject to the impingement of hot gases in excess of about ll00 F. In each of the above cases the metal possessing the high thermal fatigue properties is provided in the hot zone, but the fuse Zone is located in the cool zone.

A particular type of airfoil to which the invention has application is in the fourth stage of a turbo fan engine. An airfoil called a blucket is used which is a combination of a blade and a turbine bucket. This part has an integral fan section at one end of the airfoil formed of the `first metal composition and a blade section at the opposite end of the airfoil formed of the second metal composition with a block formed between the fan and blade sections along which the metallurgical bond is effected.

By virtue of the invention, airfoils can be designed and fabricated into a specific shape incorporating different alloys in different areas as dictated by the environment of each separate area with no sacrifice in design parameters, i.e. the smoothness of the general contour of the airfoil, particularly at the point of joinder of the two alloys.

By metallurgical bond as referred to herein along the fusion Zone between the two metal compositions it is intended to include electron beam welding, diffusion bonding, fusion by pouring molten metal directly against a preformed member or any other method by which the two metal compositions are bonded together metallurgically along a fuse zone.

A preferred embodiment of the invention is described hereinbelow with reference to the drawing wherein:

lFIG. l is a perspective view of a typical turbine vane with an airfoil section constructed from two distinct metal compositions which are welded together;

FIG. 2 is a plan view of the turbine vane shown in FIG. 1 looking into the end of the airfoil section and showing the channels formed therein;

IFIG. 3 is a perspective view of a second embodiment of a turbine vane in which a distinct metal composition is welded along the leading and trailing edges respectively of the airfoil section; and

FIG. 4 is a perspective view of a turbine vane which has an airfoil section and a turbine section and a fan section composed of distinct metals.

The turbine vane shown in FIG. l consists of an airfoil section 10 which is mounted in blocks 11 and 12 which are located at the respective ends of the airfoil section. The blocks 11 and 12 may be separate members which are affixed to the ends of the airfoil section or they may be cast as part of the airfoil section forming a unitary fused construction.

The airfoil section 10 is constructed with end portions 13 and 14 which are the portions extending from the respective blocks 11 and 12, and an integral intermediate portion 15 which is metallurgically bonded to the end portions and is constructed from a different metal composition than the metal composition used for the construction of the end portions. It has been found that the usual electron beam welding technique is satisfactory to join the end portions and the intermediate portion together permanently. The metal composition which is used for the i11- termediate section 15 is characterized by having high thermal fatigue properties at temperatures above 1100" F. to 2000 F. thereby permitting it to withstand impingement of hot gases upon the'surface of the intermediate section at temperatures above 1100 F. which are characteristic of the hot Zone. The temperature to which the end portions 13 and 14 are subjected ordinarily lies in the range 1000 F., therefore in the cool zone, while the blocks 11 and 12 which are somewhat embedded in the engine will generally experience temperatures of about 500 F. The blocks 11 and 12 and the flange portions 13 and 14, however, are constructed of a metal composition which has very good ductility and impact resistance so as to permit the airfoil section to be held within the engine and substantially to preclude distortion of the airfoil section due to the heavy stresses to which it will be subjected during operation. It is to be noted that the end portions constitute that part of the airfoil section which is ordinarily vulnerable to the mechanical forces during operation and thus, the metal composition has superior ductility and impact resistance. Extending longitudinally through the air foil are a plurality of channels 16. These channels permit a coolant to be introduced into and circulated through the channels to aid in keeping the airfoil cool and resistant to distortion by the high temperatures. Fuse zones 18 and 19 are formed along the area where the end portion is joined to the intermediate portion to form a metallurgical bond; the fuse zones are located in a cool zone.

Of the metal compositions which can be used in the construction of these airfoil sections, the compositions set forth below by way of example have proven particularly satisfactory in the construction shown in FIG. 1 in which composition A comprised the end portions 13 and 14 and composition B formed the intermediate section 15.

Composition A: Percent Cr 11 to 14 A1 5.5 to 6.5 Mo 3.5 to 5.5 Cb|Ta 1.0 to 8 Ti 0.25 to 1.25

C 0.12 to 0.17 Zr 0.03 to 0.08

B 0.01 to 0.02

Ni Balance.

Composition B:

W 11.5 to 13.50

Co 9 to 11 Cr 8 to 11 Al 4.75 to 5.25 Ti 1.75 to 2.25 Cb 0.75 to 1.25 C 0.12 to 0.17 Zr 0.03 to 0.08 B 0.01 to 0.02

Ni Balance.

In FIG. 3 the turbine vane shown consists of an airfoil section 20 with blocks 21 and 22 mounted on the respective ends of the airfoil section. The airfoil section shown has an airfoil cross section which defines a leading edge 23 and a trailing edge 24. 'Ihe blocks 21 and 22 and Sulbstantially all of the airfoil section 20 are constructed of a rst metal composition which lis characterized by having good mechanical strength, particularly high ductility and impact resistance, Substantially defining the leading edge 23 and the trailing edge 24 are an integral leading edge portion 25 and an integral trailing edge portion 26 which are welded to the airfoil section by electron beam welding such that they form an integral unitary construction with the airfoil section 20 and fuse zones 27 and 28 at which the two metals are joined in a metallurgical bond. Although greater advantages are realized by forming the integral leading and trailing edge sections 25 and 26 on the respective leading and trailing edges 23 and 24 of the airfoil section 20, similar good advantages can be realized by forming the integral leading edge fan section 25 alone. It has been found that the temperature of gas impinging upon the leading edge 23 of the blade is around l600 to 1700'F. with a lower temperature of around 12.00 to 1300 F. on the trailing edge 24, therefore delining hot zones along these areas. Between the leading edge 23 and the trailing edge 24 and along the end portions, the temperature on the surface of the airfoil section is usually below 1000 F. or a cool zone because the hot gases do not impinge directly on the surfaces. The fuse zone 27 is located in the cool zone. Thus it is seen that by forming the integral leading and trailing edge fan portions 25 and 26 of a metal composition which has high heat resistant properties, those areas which will be subjected to the highest temperatures will be able to withstand these temperatures. It is also to be noted that according to this construction the airfoil section 20 is generally constructed of a metal composition which will withstand the mechanical stresses to which the blade will lbe subjected.

By way of example, turbine blades have been constructed from the following metal compositions in which composition C was the metal used for the airfoil section 20 and composition D was the metal used for the integral leading and trailing edge portions.

Composition C: Percent Cr 20 to 22 VJ l0 to 12 Cb-l-Ta 1.5 to 2.5 Fe 1 to 2.5 C 0.4 to 0.5

Co Balance.

Composition D:

Cr 20 to 23 W 9 to 11 Ta 8 to l0 Fe 0.75 to 1.5 C 0.75 to 0.95 Si 0.1 to 0.4 Zr 0.1 to 0.3 Co Balance.

As shown in FIG. 4 some turbine blades are constructed with an airfoil section 30 which has a turbine section 31 and a fan section 32. A base 34 is formed on the root end of the airfoil section 30 and separating the turbine section from the fan section is block 35. In the operation of the turbine blade shown in FIG. 4. The turbine section is subjected to the high temperatures of the gas in a hot zone at temperatures in the range of l600 to 1700 F. and the fan section 32 is exposed to ambient temperatures ranging from F. to 90 F. in a cool zone. Thus it is desirable to form the turbine section from a metal composition which has high thermal fatigue properties whereas the fan section 32 being unsupported at its extreme end during operation must have good material strength to withstand the mechanical stresses during its operation. Here the turbine section is constructed from a rst metal composition and the fan section is constructed from a second metal composition and they are welded by electron beam welding along the block 35 to form a metallurgical bond between the two metal compositions and define a fuse zone 36 therebetween which is located out of the hot zone in the cool zone.

By way of example, listed below is a nickel base alloy composition E from which the turbine Section was constructed and an iron base alloy composition F from which the fan section was constructed.

Another alloy composition which has proven to be particularly Well-suited for construction of the roots of the flange portions of the airfoil sections shown is alloy composition G in the approximate percentages listed below.

Composition G: Percent lt is also intended that titanium and titanium base alloy can be used for the iiange or roots of the airfoil section with particular advantages in aircraft applications because of the importance of Weight and corrosion resistance in this use.

Specific existing alloy compositions have been given above and set out below is a percentage range of metals. Composition H would be used in the hot zone and Composition I would be used in the cold zone.

Composition I,

Composition H,

percent percent Metals Room vtemperature properties of these compositions giving the desired thermal fatigue properties and mechanical strength are as follows:

Composition H:

UTS 120,000 p.s.i.

.2% Y.S 105,000 p.s.i.

Percent E 5 min.

SR hrs. at

1800o F. and

29,000 p.s.i.

6 Composition I:

UTS 230,000 p.s.i. .2% Y.S 200,000 p.s.i. Percent E 18 min.

SR NIL.

Other important considerations in connection with the blades and vanes having airfoil sections as claimed lies in methods of making these yblades and vanes. For example when blades and vanes having airfoil sections of a single metallic structure wear out along the leading and trailing edges, they can be repaired by electron beam welding an alloy composition having high heat resistant properties to give the structure shown in FIG. 3.

It has further been proposed that the over-all characteristics of the blade or vane can be improved by chromizing the blades and vanes having the bimetallic airfoil section to form a high chomium surface alloy. It has been found that by such a chromizing operation a much more durable product is realized.

It is to -be noted further that in joining the four stage bluckets, a whole wheel or hub can be cast of the first metal composition having the high material strength and then the second lmetal composition can be separately welded to each of the airfoil sections at the desired location.

I cla-im:

1. A bimetallic airfoil which is subject to the impingement of hot `gas on its surface areas defining a -hot zone where the temperatures of the impinging gases are above about 1l00 F. and a cool zone where the temperature of the -impinging gases are below this temperature compais-ing an integral unitary airfoil structure of two metal alloy compositions, a first integral portion composed of a first metal alloy composition, and a second integral portion composed of a `second metal alloy composition, sa-id second metal composition defining the surface area along the hot zone of the airfoil and is fused to said first metal composition by a metallurgical bond along a fuse zone located in the cool zone, said first metal composition having ductile and impact resistant properties in excess of said second metal composition and saiid second metal composition having thermal fatigue properties above 1100o F.

2. A bimetallic airfoil according to claim 1 wherein said first metal composition defines end portions of the airfoil and said second metal composition defines an intermediate portion of the airfoil.

3. A bimetallic airfoil according to claim 1 wherein said second metal composition is provided along a leading edge of the airfoil.

4. A bimetallic airfoil according to claim 3 wherein said second metal composition is provided along a trailing edge of the airfoil.

5. A bimetall-ic airfoil according to claim 1 wherein said airfoil has an integral fan section at one end of said airfoil composed of said first metal composition and a blade section at the opposite end of said airfoil formed of said second metal composition, and a block formed between said fan and blade sections along which the fuse zone extends.

6. A bimetallic airfoil which is subject to the impingement of hot gas on its surface areas defining a hot zone where the temperatures of the impinging gases are above about 1100 F. and a cool zone where the temperature of the impinging gases are below this temperature comprising an integral unitary airfoil structure of two metal alloy compositions, a first Iintegral portion composed of a first metal alloy composition, and a second integral portion composed of a second metal alloy composition, said second metal composition defining the surface area along the hot zone of the airfoil and is fused to said first metal composition by a metallurgical bond along a fuse zone located in the cool zone, said first metal composition having ductile yand impact resistant properties in excess of said second metal composition and said second metal composition having thermal fatigue properties above 1100 F., said rst and second metal compositions having the following percentages by weight:

First composition: Percent C 04 to .14

P, max .02 S, max .02 Si 25 to 10 Mn 50 to 1.50

N 04 to .10 Mo 1 0 to 6 0 Ni 2 0 to 6 0 Cr 10.0 to 20.0 Fe 50.0 to 80.0

Al 1.0 to 7.0

Zr to 8.0 V 0 to 16.0

Second composition:

C 05 to .60

P, max .04 S, max .04 Si, max 1.0 Mn, max 1.0 Mo 1.0 to 10.0 Ni 5.0 to 15.0 Cr 5.0 to 30.0 Fe 1.0` to 5.0 B 0.008 to .020 Co 5.0 to 60.0 Al .10 to 6.0 Ti .10 to 6.0 W .10 to 10.0 Cb .05 to 1.0 Ta 1.0 to 6.0 Zr .05 to 1.0

7. A bimetallic airfoil according to claim 6 wherein said first yrnetal composition defines end portions of the arfoil and said second metal composition defines an intermediate portion of the airfoil.

8.. A bimetallic vairfoil according to claim 6 wherein said second metal composition is provided along a leading edge of the airfoil.

9. A bimetallic airfoil according to claim 8 wherein said second metal composition is provided along a trailing edge of the a'irfoil.

10. A bimetallic airfoil according to claim 6 wherein said airfoil has an integral fan section at one end of said airfoil composed of said rst metal composition and a blade section at the opposite end of said airfoil formed of said second metal composition, and a block formed between said fan and blade sections along which the fuse zone extends.

11. A -bimetallic a-irfoil according to claim 6 wherein said rst metal and said second metal compositions have the following nominal percentages by weight:

First composition: Percent Cr 11 to 14 Al 5.5 to 6.5 Mo 3.5 to 5.5 Cb-t-Ta 1.0 to 8 Ti 0.25 to 1.25 C 0.12 to 0.17 Zr 0.03 to 0.08 B 0.01 to 0.02 Ni Balance 8 Second composition:

W 11.5 to 13.50 Co 9 to 11 Cr S to 11 Al 4.75 to 5.25 Ti 1.75 to 2.25 Cb 0.75 to 1.25 C 0.12 to 0.17

Zr 0.03 to 0.08

B 0.01 to 0.02

Ni Balance 12. A bimetallic airfoil according to claim 6 wherein said first and second metal compositions have the following nominal percentages by weight:

First composition: Percent Cr 20 to 22 W 10 to 12 Cb-f-Ta 1.5 to 2.5 Fe 1 to 2.5

C 0.4 to 0.5

Co Balance Second composition:

Cr 20` to 23 W 9 to 11 Ta 8 to 10 Fe 0.75 to 1.5

C 0.75 to 0.95 Si 0.1 to 0.4 Zr 0.1 to 0.3

Co Balance 13. A bimetallic airfoil section according to claim 6 wherein said lirst and second metal compositions have the following nominal percentages by weight:

First composition: Percent Cr 15.5 to 16.7 Ni 3.6 to 4.6 Cu 2.8 to 3.5 Si 0.5 to 1 Cb-i-Ta 0.15 to 0.40 Fe Balance Second composition:

Co 24 to 28 Cr 13.5 to 16.5

Mo 4 to 5 Al 4 to 4.75

Ti 2 to 2.75 C 0.5 to 0.11

Ni Balance References Cited UNITED STATES PATENTS 2,853,160 9/1958 Matters 253-77 X 2,971,745 2/1961 Warren et al 253-77 3,002,675 10/1961 Howell et al. 3,112,865 12/1963 Gisslen 253-77 3,215,511 11/1965 Chisholm et al 253-77 X EVERETTE A. POWELL, JR., Primary Examiner.

M. P. SCHWADRON, Assistant Examiner.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2853160 *Feb 16, 1956Sep 23, 1958Allis Chalmers Mfg CoMeans for and method of fabrication
US2971745 *Jan 12, 1959Feb 14, 1961Gen ElectricFabricated blade and bucket rotor assembly
US3002675 *Nov 3, 1958Oct 3, 1961Power Jets Res & Dev LtdBlade elements for turbo machines
US3112865 *Oct 3, 1961Dec 3, 1963Gen ElectricBlade platform structure
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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3524712 *May 8, 1967Aug 18, 1970Rolls RoyceCompressor blade for a gas turbine engine
US3768147 *Dec 20, 1971Oct 30, 1973Gen ElectricMethod of friction welding
US3847203 *Jun 22, 1972Nov 12, 1974Secr DefenceMethod of casting a directionally solidified article having a varied composition
US3982854 *Jun 18, 1973Sep 28, 1976General Electric CompanyFriction welded metallic turbomachinery blade element
US4305697 *Mar 19, 1980Dec 15, 1981General Electric CompanyMethod and replacement member for repairing a gas turbine engine vane assembly
US4326833 *Mar 19, 1980Apr 27, 1982General Electric CompanyMethod and replacement member for repairing a gas turbine engine blade member
US4850802 *Apr 21, 1983Jul 25, 1989Allied-Signal Inc.Composite compressor wheel for turbochargers
US5209645 *Jun 15, 1990May 11, 1993Hitachi, Ltd.Ceramics-coated heat resisting alloy member
US5221188 *Dec 23, 1991Jun 22, 1993Mtu Motoren-Und Turbinen-Union Munchen GmbhBlade device for turbo-engines
US7959409 *Mar 1, 2007Jun 14, 2011Honeywell International Inc.Repaired vane assemblies and methods of repairing vane assemblies
US8267663 *Apr 28, 2008Sep 18, 2012Pratt & Whitney Canada Corp.Multi-cast turbine airfoils and method for making same
US20090269193 *Apr 28, 2008Oct 29, 2009Larose JoelMulti-cast turbine airfoils and method for making same
Classifications
U.S. Classification416/241.00R, 416/231.00R
International ClassificationF16H41/00, F16H41/28, B32B15/01, F01D5/28, C22C19/00
Cooperative ClassificationB32B15/013, F16H41/28, C22C19/00, F01D5/28
European ClassificationF16H41/28, B32B15/01D, F01D5/28, C22C19/00
Legal Events
DateCodeEventDescription
Jul 28, 1983ASAssignment
Owner name: HOWMET TURBINE COMPONENTS CORPORATION 825 THIRD AV
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST. SUBJECT TO AGREEMENT DATED DECEMBER 31, 1975.;ASSIGNOR:HOWMET CORPORATON A CORP. OF DE;REEL/FRAME:004164/0321
Effective date: 19830705