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Publication numberUS3600103 A
Publication typeGrant
Publication dateAug 17, 1971
Filing dateOct 6, 1969
Priority dateOct 6, 1969
Also published asDE2047209A1
Publication numberUS 3600103 A, US 3600103A, US-A-3600103, US3600103 A, US3600103A
InventorsCharles E Doughty, David F Gray, Frank B Pinney
Original AssigneeUnited Aircraft Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Composite blade
US 3600103 A
Abstract  available in
Images(2)
Previous page
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Claims  available in
Description  (OCR text may contain errors)

llnited States Patent Primary ExaminerEverette A. Powell, J r Attorney-Charles A. Warren ABSTRACT: A compressor or fan blade is made up of layers of high modulus fibers extending in parallel relation, and embedded in a metallic matrix, the successive layers of fibers varying in length and dimension so that, when stacked and compacted together, the desired blade configuration will be obtained. The blade may be made by a process including stacking the layers of fibers which are secured to a metallic supporting sheet, in the proper sequence, and then compacting the assemblage in dies under pressure and high temperature to cause the material of the several backing sheets to fill the voids among the fibers and form the matrix within which the fibersare embedded.

COMPOSITE BLADE BACKGROUND OF-THE INVENTION With the advent of high thrust jet engines for use in aircraft, the need for compressor and fan blades with a high modulus of elasticity has become more apparent. The higher modulus makes a stiffer blade thus reducing vibration problems, and combined with lightness in weight permits a lighter supporting structure thereby materially reducing engine weight.

SUMMARY OF THE INVENTION One feature of the invention is a composite blade made up of several layers of high modulus fibers all embedded in a metallic matrix and with the several layers varying in plan so that when stacked together and embedded in the matrix the finished blade shape is produced. Another feature is a process for producing such composite blades.

According to the invention, the blade is made up of a plurality of layers of high modulus fibers each layer being enclosed in a metallic matrix, with the several layers shaped in plan so that when stacked and compressed in a die the compacted stack will have the shape of the finished blade, the layers will be in substantially parallel relation in the blade, and they will all be bonded together in a metallic matrix.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is an end view of a finished blade in a portion of a supporting disc.

FIG. 2 is a side view of the blade.

FIG. 3 is an assemblage of sectional views along line Zia-3a, 3b3b and Bic-3c of FIG. 1.

H6. 4 is a developed plan view of the blade showing the configuration of several plies in the blade.

FIG. 5 is a sectional view through a die and associated mechanism for producing the blade.

FIG. 6 is an enlarged sectional view through one of the plies of fibers.

FIG. 7 is a diagrammatic view of the assemblage of plies in forming the blade.

FIG. 8 is a fragmentary view of the end of a blade. root showing a modification.

FIG. 9 is a view of the bottom of the blade root at FIG.

DESCRIPTION OF THE PREFERRED EMBODIMENT The invention is shown in a compressor or fan blade 2 for use especially in jet engines, the blade being mounted by a dovetail root portion 4 in a corresponding slot 6 in a supporting disc 8. The blade is a composite, being made up of high modulus fibers in a metallic matrix, with the matrix substantially filling all the voids to form a void-free blade. The fibers extend through the root attachment as will be described. The blade has a platform 10 at its inner end that is inclined from front to rear as shown in FIG. 2 and this platform defines the inner wall of the air path over the airfoil portion of the blade.

As above stated the fibers are high modulus fibers l2, FIG. 6, one type of which is a boron fiber coated with silicon carbide, sometimes with a tungsten core, and these fibers are secured to an aluminum alloy sheet 14 in closely and evenly spaced parallel relation to one another. These fibers are secured to the backing sheet M by a coating 16 applied to the side of the fibers opposite to the backing sheet. This coating may be applied as by plasma spraying an aluminum alloy onto the fibers. The assemblage forms ribbons to be used in making the blade.

The ribbon of fibers combined with the backing sheet and coating is cut into shaped plies 118 which vary in length and width as shown in FIG. 4 and these plies are stacked one above another preferably with the smallest dimension of ply 20 centrally of the blade assembly so that the outermost ply 22, which is the shape of the blade will form a continuous surface over one side of the blade. Although these plies may be stacked directly on one another from the centerline. of the blade, it is desirable to provide an aluminum core 24, which is the shape of the blade, with the individual plies stacked on opposite sides of the core. This core is preferably relatively thin, for example, 5 mils, and may be aluminum or the same aluminum alloy as the backing sheets for the plies.

The shapes of the several plies is so selected that when compacted in the formation of the blade the resulting contour will be the desired contour of the finished blade. Preferably also the same number of plies are located on each side of the core and on both sides the outer ply, when contoured to shape will be that of the finished blade surface.

When the plies have been individually cut to shape they are arranged in stacked relation, as shown in FIG. 9, each ply being desirably tack welded to the surface underneath so that each ply will remain in proper relation to the others during the build up of the plies on both sides of the centerline of the blade. As shown in FIG. 4, each ply projects beyond the airfoil portion 26 of the blade assembly into the root portion 28 in order to form a secure attachment for the blade in the supporting disc. The formation of the dovetail portion is not a part of the present invention; for the purpose of this invention the assemblage of the plies in the root portion may have dovetail and platform forming elements 30 positioned on opposite sides of the plies, as shown in FIG. 1. These elements may be secured either during the forming of the blade in the die or may be diffusion bonded or otherwise attached subsequently.

In either event, the assemblage of the several 1 as in FIG. 7, is placed in a two-part die 32, FIG. 5 in which the assemblage is compacted and unitized by heating. Although only a few plies are shown in this figure, it will be understood that this is shown by way of example, since these plies as formed are only a few mils thick and normally a large number of plies must be used in producing a blade of the desired thickness. The number of plies and the shape of the several plies may well be determined by computer after the compacted thickness of each of the plies is known and the desired configuration of the finished blade is established. The parts of the die 32, receive between them the assembled stack of plies in a cavity 34 that is the configuration of the finished blade, and heat and pressure are applied to compact the plies and to cause the aluminum alloy to flow to fill the voids. The die 30 may be placed between pressure plates 34 and 36 slidable in a housing 38, the latterhaving heating means therein such as an induction coil 40. In the particular embodiment shown a pressure of about 3500 p.s.i. is applied to the dies and the temperature of the dies is raised to a point about 25 below the melting point of the aluminum alloy. At this temperature the alloy will flow around the fibers and fill all the voids among the fibers. The heating and pressing is preferably done in a high vacuum and the pressure is preferably applied only when the desired temperature is reached. With the particular alloy being used the temperature is between l020 and l050.

The temperature and pressure are maintained for about 1% hours, after which the dies and blade are allowed to cool while the pressure is maintained. During the pressing of the blade into finished shape, excess aluminum alloy 12 escapes between the die halves as shown in FIG. 5. The pressure plates 34 and 36 46 of be limited in their movement toward one another by suitable stops 44 one end of the housing 38.

The thickness of the aluminum backing sheet for the fibers is selected so that in the finished blade the volume of fibers will be about 50 percent so that the volumes of fibers and surrounding matrix are about equal.

After the blade is cooled and is removed from the die, the surplusage of aluminum at the edges of the blade are removed, the tip of the blade is cut to shape and the root elements 30, if not added earlier, are secured to opposite surfaces of the root portion. In attachment of these elements it is desirable to form the mating surfaces 44 and 46 of these elements with a compound curvature for more secure retention of the blade within these elements. Thus, as shown in FIG. the root portion is curved longitudinally from end-to-end, and as shown in FIG. 9 the root portion is also cured laterally between opposite ends. This may be accomplished during the pressing of the blade, and shaping the mating surfaces of the elements 30 to fit this compound curvature.

We claim:

1. A composite blade material for a compressor having a plurality of layers of high modulus plies stacked on each side of a thin sheet of aluminum alloy which is located centrally of the blade, the plies comprising high modulus fibers selected from the group consisting of boron and silicon carbide coated boron fibers and a matrix material which is an aluminum alloy, and the plies varying in width and length to provide a varying blade thickness chordwise and lengthwise.

2. A blade according to claim 1 wherein the plies are stacked according to size to form an assemblage of layers graduated in size, with the smaller plies central of the assemblage and the larger plies outer of the smaller plies, the outermost ply being substantially continuous over the blade surface.

3. A blade according to claim 1 having further a root portion at one end of the blade and an airfoil port on, wherein the fibers and matrix form the complete airfoil portion and extend into and form a projecting portion which is a part of the root portion.

4. A blade according to claim 1 wherein the fiber constitutes substantially 50 percent of the volume of the finished blade.

UNITED STATES PATENT OFFICE 5/6 CERTIFICATE OF CORRECTION Patent No. 3,600,103 Dated August 17, 1971 Inventor(s) David F. Gray, et a1.

It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected as shown below:

In column 2, line 14, the numeral "9" should be "7".

In column 2, line 26, the numeral "1" should be "8".

In column 2, line 41, the numeral "34" should be "33".

In column 2, line 73, the numeral "44" should be "45".

In column 3, line 2, the word "cured" should be "curved".

In sheet 2 of the drawings, the numeral "34" immediately to the left of Fig. 5 should be "33".,

In column 2, line 30, the numeral "1" should be "plies".

In column 2, line 62, "46 of" should be "may".

In column 2, line 63, the word "at" should be inserted after the numeral "44" Signed and sealed this 20th day of June 1972.

(SEAL) Attest:

EDWARD M.FLETCHER,JR. ROBERT GOTI'SCHALK Attestlng Officer Commissioner of Patents

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2621140 *Jul 8, 1948Dec 9, 1952Comp Generale ElectriciteMethod for molding propeller blades
US2929755 *Jul 24, 1958Mar 22, 1960Orenda Engines LtdPlastic blades for gas turbine engines
US3098723 *Jan 18, 1960Jul 23, 1963Rand CorpNovel structural composite material
US3248082 *Aug 19, 1965Apr 26, 1966Whitfield Lab IncReinforced gas-turbine blade or vane
GB781333A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3699623 *Oct 20, 1970Oct 24, 1972United Aircraft CorpMethod for fabricating corrosion resistant composites
US3713752 *Oct 28, 1971Jan 30, 1973United Aircraft CorpComposite blade for a gas turbine engine
US3717419 *Jul 9, 1971Feb 20, 1973Susquehanna CorpTurbine blade
US3749518 *Mar 15, 1972Jul 31, 1973United Aircraft CorpComposite blade root configuration
US3752600 *Dec 9, 1971Aug 14, 1973United Aircraft CorpRoot pads for composite blades
US3756745 *Mar 15, 1972Sep 4, 1973United Aircraft CorpComposite blade root configuration
US3799701 *Feb 28, 1972Mar 26, 1974United Aircraft CorpComposite fan blade and method of construction
US3883267 *Aug 6, 1973May 13, 1975SnecmaBlades made of composite fibrous material, for fluid dynamic machines
US3886647 *Apr 25, 1974Jun 3, 1975Trw IncMethod of making erosion resistant articles
US3903578 *Dec 13, 1973Sep 9, 1975United Aircraft CorpComposite fan blade and method of construction
US3942231 *Oct 31, 1973Mar 9, 1976Trw Inc.Contour formed metal matrix blade plies
US3981616 *Oct 22, 1974Sep 21, 1976The United States Of America As Represented By The Secretary Of The Air ForceHollow composite compressor blade
US4006999 *Jul 17, 1975Feb 8, 1977The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationLeading edge protection for composite blades
US4022547 *Oct 2, 1975May 10, 1977General Electric CompanyComposite blade employing biased layup
US4043703 *Dec 22, 1975Aug 23, 1977General Electric CompanyImpact resistant composite article comprising laminated layers of collimated filaments in a matrix wherein layer-layer bond strength is greater than collimated filament-matrix bond strength
US4060413 *Jun 11, 1976Nov 29, 1977Westinghouse Canada LimitedMethod of forming a composite structure
US4108572 *Dec 23, 1976Aug 22, 1978United Technologies CorporationTitanium sheet
US4111606 *Dec 27, 1976Sep 5, 1978United Technologies CorporationComposite rotor blade
US4178667 *Mar 6, 1978Dec 18, 1979General Motors CorporationMethod of controlling turbomachine blade flutter
US4301584 *Jan 31, 1980Nov 24, 1981United Technologies CorporationMethod of forming fiber and metal matrix composite
US4472866 *Mar 1, 1982Sep 25, 1984Trw Inc.Method of making an airfoil
US4583274 *Sep 20, 1984Apr 22, 1986Trw Inc.Method of making an airfoil
US4762268 *May 2, 1986Aug 9, 1988Airfoil Textron Inc.Fabrication method for long-length or large-sized dense filamentary monotapes
US4856162 *Aug 8, 1988Aug 15, 1989United Technologies CorporationFabrication of bonded structures
US4884948 *Mar 29, 1988Dec 5, 1989Mtu Motoren-Und Turbinen Union Munchen GmbhDeflectable blade assembly for a prop-jet engine and associated method
US5141400 *Jan 25, 1991Aug 25, 1992General Electric CompanyWide chord fan blade
US5375978 *Oct 4, 1993Dec 27, 1994General Electric CompanyForeign object damage resistant composite blade and manufacture
US5449273 *Mar 21, 1994Sep 12, 1995United Technologies CorporationComposite airfoil leading edge protection
US5486096 *Jun 30, 1994Jan 23, 1996United Technologies CorporationErosion resistant surface protection
US5840390 *Oct 4, 1995Nov 24, 1998Research Institute Of Advanced Material Gas-Generator Co., Ltd. (Amg)FRM disc preform and manufacturing method thereof
US5876651 *May 29, 1996Mar 2, 1999United Technologies CorporationMethod for forming a composite structure
US7753653 *Jan 12, 2007Jul 13, 2010General Electric CompanyComposite inlet guide vane
US8123463Jul 31, 2008Feb 28, 2012General Electric CompanyMethod and system for manufacturing a blade
US8794925Aug 24, 2010Aug 5, 2014United Technologies CorporationRoot region of a blade for a gas turbine engine
CN101220818BJan 14, 2008Sep 18, 2013通用电气公司Composite inlet guide vane
WO1981002128A1 *Jan 26, 1981Aug 6, 1981United Technologies CorpMethod of forming fiber and metal matrix composite
Classifications
U.S. Classification416/154, 416/220.00R, 416/230, 416/241.00R, 228/190, 416/241.00A, 416/226
International ClassificationC22C47/16, F04D29/38, F01D5/28, F04D29/02
Cooperative ClassificationC22C47/16, F01D5/282, Y02T50/672
European ClassificationF01D5/28B, C22C47/16