US3600103A - Composite blade - Google Patents

Composite blade Download PDF

Info

Publication number
US3600103A
US3600103A US864021A US3600103DA US3600103A US 3600103 A US3600103 A US 3600103A US 864021 A US864021 A US 864021A US 3600103D A US3600103D A US 3600103DA US 3600103 A US3600103 A US 3600103A
Authority
US
United States
Prior art keywords
blade
plies
fibers
layers
assemblage
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.)
Expired - Lifetime
Application number
US864021A
Inventor
David F Gray
Charles E Doughty
Frank B Pinney
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Aircraft Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by United Aircraft Corp filed Critical United Aircraft Corp
Application granted granted Critical
Publication of US3600103A publication Critical patent/US3600103A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C47/00—Making alloys containing metallic or non-metallic fibres or filaments
    • C22C47/16—Making alloys containing metallic or non-metallic fibres or filaments by thermal spraying of the metal, e.g. plasma spraying
    • 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
    • F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
    • 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

Definitions

  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • each ply 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.
  • 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.
  • a two-part die 32 FIG. 5 in which the assemblage is compacted and unitized by heating.
  • 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.
  • 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.
  • 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.
  • 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.
  • the mating surfaces 44 and 46 of these elements are formed with a compound curvature for more secure retention of the blade within these elements.
  • 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.
  • 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.
  • 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.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Metallurgy (AREA)
  • Plasma & Fusion (AREA)
  • Physics & Mathematics (AREA)
  • Organic Chemistry (AREA)
  • Composite Materials (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Pressure Welding/Diffusion-Bonding (AREA)
  • Manufacture Of Alloys Or Alloy Compounds (AREA)

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 fibers are embedded.

Description

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

Claims (4)

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.
US864021A 1969-10-06 1969-10-06 Composite blade Expired - Lifetime US3600103A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US86402169A 1969-10-06 1969-10-06

Publications (1)

Publication Number Publication Date
US3600103A true US3600103A (en) 1971-08-17

Family

ID=25342343

Family Applications (1)

Application Number Title Priority Date Filing Date
US864021A Expired - Lifetime US3600103A (en) 1969-10-06 1969-10-06 Composite blade

Country Status (5)

Country Link
US (1) US3600103A (en)
JP (1) JPS4924923B1 (en)
DE (1) DE2047209A1 (en)
FR (1) FR2065431A5 (en)
GB (1) GB1269172A (en)

Cited By (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3699623A (en) * 1970-10-20 1972-10-24 United Aircraft Corp Method for fabricating corrosion resistant composites
US3713752A (en) * 1971-10-28 1973-01-30 United Aircraft Corp Composite blade for a gas turbine engine
US3717419A (en) * 1971-07-09 1973-02-20 Susquehanna Corp Turbine blade
US3749518A (en) * 1972-03-15 1973-07-31 United Aircraft Corp Composite blade root configuration
US3752600A (en) * 1971-12-09 1973-08-14 United Aircraft Corp Root pads for composite blades
US3756745A (en) * 1972-03-15 1973-09-04 United Aircraft Corp Composite blade root configuration
US3799701A (en) * 1972-02-28 1974-03-26 United Aircraft Corp Composite fan blade and method of construction
US3883267A (en) * 1972-08-04 1975-05-13 Snecma Blades made of composite fibrous material, for fluid dynamic machines
US3886647A (en) * 1971-07-07 1975-06-03 Trw Inc Method of making erosion resistant articles
US3903578A (en) * 1972-02-28 1975-09-09 United Aircraft Corp Composite fan blade and method of construction
US3942231A (en) * 1973-10-31 1976-03-09 Trw Inc. Contour formed metal matrix blade plies
US3981616A (en) * 1974-10-22 1976-09-21 The United States Of America As Represented By The Secretary Of The Air Force Hollow composite compressor blade
US4006999A (en) * 1975-07-17 1977-02-08 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Leading edge protection for composite blades
US4022547A (en) * 1975-10-02 1977-05-10 General Electric Company Composite blade employing biased layup
US4043703A (en) * 1975-12-22 1977-08-23 General Electric Company Impact 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
US4060413A (en) * 1975-12-24 1977-11-29 Westinghouse Canada Limited Method of forming a composite structure
US4108572A (en) * 1976-12-23 1978-08-22 United Technologies Corporation Composite rotor blade
US4111606A (en) * 1976-12-27 1978-09-05 United Technologies Corporation Composite rotor blade
US4178667A (en) * 1978-03-06 1979-12-18 General Motors Corporation Method of controlling turbomachine blade flutter
WO1981002128A1 (en) * 1980-01-31 1981-08-06 United Technologies Corp Method of forming fiber and metal matrix composite
US4472866A (en) * 1982-03-01 1984-09-25 Trw Inc. Method of making an airfoil
US4583274A (en) * 1982-03-01 1986-04-22 Trw Inc. Method of making an airfoil
US4762268A (en) * 1986-05-02 1988-08-09 Airfoil Textron Inc. Fabrication method for long-length or large-sized dense filamentary monotapes
US4856162A (en) * 1985-12-30 1989-08-15 United Technologies Corporation Fabrication of bonded structures
US4884948A (en) * 1987-03-28 1989-12-05 Mtu Motoren-Und Turbinen Union Munchen Gmbh Deflectable blade assembly for a prop-jet engine and associated method
US5141400A (en) * 1991-01-25 1992-08-25 General Electric Company Wide chord fan blade
US5375978A (en) * 1992-05-01 1994-12-27 General Electric Company Foreign object damage resistant composite blade and manufacture
US5449273A (en) * 1994-03-21 1995-09-12 United Technologies Corporation Composite airfoil leading edge protection
US5486096A (en) * 1994-06-30 1996-01-23 United Technologies Corporation Erosion resistant surface protection
US5840390A (en) * 1994-11-02 1998-11-24 Research Institute Of Advanced Material Gas-Generator Co., Ltd. (Amg) FRM disc preform and manufacturing method thereof
US5876651A (en) * 1996-05-29 1999-03-02 United Technologies Corporation Method for forming a composite structure
US20080170943A1 (en) * 2007-01-12 2008-07-17 General Electric Company Composite inlet guide vane
US20100028594A1 (en) * 2008-07-31 2010-02-04 Nicholas Joseph Kray Method and system for manufacturing a blade
US8794925B2 (en) 2010-08-24 2014-08-05 United Technologies Corporation Root region of a blade for a gas turbine engine
CN108087318A (en) * 2017-12-14 2018-05-29 中国航发沈阳发动机研究所 A kind of mixed structure composite material blade

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4000956A (en) * 1975-12-22 1977-01-04 General Electric Company Impact resistant blade
IT1176673B (en) * 1983-09-23 1987-08-18 Gen Electric QUARRY SHOVEL FOR TURBOMACHINE
DE19922012C1 (en) * 1999-05-12 2000-10-19 Mtu Muenchen Gmbh Rotor unit machining takes measurements of the actual surfaces at the fused joint for the flash to be removed and a nominal surface structure is machined from stored data to give the required flow characteristics
FR2867095B1 (en) * 2004-03-03 2007-04-20 Snecma Moteurs METHOD FOR MANUFACTURING A HOLLOW DAWN FOR TURBOMACHINE

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2621140A (en) * 1940-08-28 1952-12-09 Comp Generale Electricite Method for molding propeller blades
GB781333A (en) * 1950-04-06 1957-08-21 Elmer Pershing Warnken Improvements relating to blades of air-foil configuration and the method of making same
US2929755A (en) * 1958-07-24 1960-03-22 Orenda Engines Ltd Plastic blades for gas turbine engines
US3098723A (en) * 1960-01-18 1963-07-23 Rand Corp Novel structural composite material
US3248082A (en) * 1965-08-19 1966-04-26 Whitfield Lab Inc Reinforced gas-turbine blade or vane

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2621140A (en) * 1940-08-28 1952-12-09 Comp Generale Electricite Method for molding propeller blades
GB781333A (en) * 1950-04-06 1957-08-21 Elmer Pershing Warnken Improvements relating to blades of air-foil configuration and the method of making same
US2929755A (en) * 1958-07-24 1960-03-22 Orenda Engines Ltd Plastic blades for gas turbine engines
US3098723A (en) * 1960-01-18 1963-07-23 Rand Corp Novel structural composite material
US3248082A (en) * 1965-08-19 1966-04-26 Whitfield Lab Inc Reinforced gas-turbine blade or vane

Cited By (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3699623A (en) * 1970-10-20 1972-10-24 United Aircraft Corp Method for fabricating corrosion resistant composites
US3886647A (en) * 1971-07-07 1975-06-03 Trw Inc Method of making erosion resistant articles
US3717419A (en) * 1971-07-09 1973-02-20 Susquehanna Corp Turbine blade
US3713752A (en) * 1971-10-28 1973-01-30 United Aircraft Corp Composite blade for a gas turbine engine
US3752600A (en) * 1971-12-09 1973-08-14 United Aircraft Corp Root pads for composite blades
US3799701A (en) * 1972-02-28 1974-03-26 United Aircraft Corp Composite fan blade and method of construction
US3903578A (en) * 1972-02-28 1975-09-09 United Aircraft Corp Composite fan blade and method of construction
US3756745A (en) * 1972-03-15 1973-09-04 United Aircraft Corp Composite blade root configuration
US3749518A (en) * 1972-03-15 1973-07-31 United Aircraft Corp Composite blade root configuration
US3883267A (en) * 1972-08-04 1975-05-13 Snecma Blades made of composite fibrous material, for fluid dynamic machines
US3942231A (en) * 1973-10-31 1976-03-09 Trw Inc. Contour formed metal matrix blade plies
US3981616A (en) * 1974-10-22 1976-09-21 The United States Of America As Represented By The Secretary Of The Air Force Hollow composite compressor blade
US4006999A (en) * 1975-07-17 1977-02-08 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Leading edge protection for composite blades
US4022547A (en) * 1975-10-02 1977-05-10 General Electric Company Composite blade employing biased layup
US4043703A (en) * 1975-12-22 1977-08-23 General Electric Company Impact 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
US4060413A (en) * 1975-12-24 1977-11-29 Westinghouse Canada Limited Method of forming a composite structure
US4108572A (en) * 1976-12-23 1978-08-22 United Technologies Corporation Composite rotor blade
US4111606A (en) * 1976-12-27 1978-09-05 United Technologies Corporation Composite rotor blade
US4178667A (en) * 1978-03-06 1979-12-18 General Motors Corporation Method of controlling turbomachine blade flutter
WO1981002128A1 (en) * 1980-01-31 1981-08-06 United Technologies Corp Method of forming fiber and metal matrix composite
US4301584A (en) * 1980-01-31 1981-11-24 United Technologies Corporation Method of forming fiber and metal matrix composite
US4472866A (en) * 1982-03-01 1984-09-25 Trw Inc. Method of making an airfoil
US4583274A (en) * 1982-03-01 1986-04-22 Trw Inc. Method of making an airfoil
US4856162A (en) * 1985-12-30 1989-08-15 United Technologies Corporation Fabrication of bonded structures
US4762268A (en) * 1986-05-02 1988-08-09 Airfoil Textron Inc. Fabrication method for long-length or large-sized dense filamentary monotapes
US4884948A (en) * 1987-03-28 1989-12-05 Mtu Motoren-Und Turbinen Union Munchen Gmbh Deflectable blade assembly for a prop-jet engine and associated method
US5141400A (en) * 1991-01-25 1992-08-25 General Electric Company Wide chord fan blade
US5375978A (en) * 1992-05-01 1994-12-27 General Electric Company Foreign object damage resistant composite blade and manufacture
US5449273A (en) * 1994-03-21 1995-09-12 United Technologies Corporation Composite airfoil leading edge protection
US5486096A (en) * 1994-06-30 1996-01-23 United Technologies Corporation Erosion resistant surface protection
US5840390A (en) * 1994-11-02 1998-11-24 Research Institute Of Advanced Material Gas-Generator Co., Ltd. (Amg) FRM disc preform and manufacturing method thereof
US5876651A (en) * 1996-05-29 1999-03-02 United Technologies Corporation Method for forming a composite structure
US20080170943A1 (en) * 2007-01-12 2008-07-17 General Electric Company Composite inlet guide vane
US7753653B2 (en) * 2007-01-12 2010-07-13 General Electric Company Composite inlet guide vane
CN101220818B (en) * 2007-01-12 2013-09-18 通用电气公司 Composite inlet guide vane
US20100028594A1 (en) * 2008-07-31 2010-02-04 Nicholas Joseph Kray Method and system for manufacturing a blade
US8123463B2 (en) 2008-07-31 2012-02-28 General Electric Company Method and system for manufacturing a blade
US8794925B2 (en) 2010-08-24 2014-08-05 United Technologies Corporation Root region of a blade for a gas turbine engine
CN108087318A (en) * 2017-12-14 2018-05-29 中国航发沈阳发动机研究所 A kind of mixed structure composite material blade

Also Published As

Publication number Publication date
GB1269172A (en) 1972-04-06
DE2047209A1 (en) 1971-04-15
FR2065431A5 (en) 1971-07-23
JPS4924923B1 (en) 1974-06-26

Similar Documents

Publication Publication Date Title
US3600103A (en) Composite blade
US3942231A (en) Contour formed metal matrix blade plies
US4110505A (en) Quick bond composite and process
US3237697A (en) Helicopter rotor blade
US5375978A (en) Foreign object damage resistant composite blade and manufacture
US3810711A (en) Cooled turbine blade and its manufacture
US5340280A (en) Dovetail attachment for composite blade and method for making
US3799701A (en) Composite fan blade and method of construction
US4301584A (en) Method of forming fiber and metal matrix composite
US3886647A (en) Method of making erosion resistant articles
US2202014A (en) Air propeller blade and material for making same
CN109094058B (en) Light composite material bullet holder and its manufacturing method
US2155375A (en) Pressed article, particularly airscrew blade, of laminated wood, and method of manufacture
US3649425A (en) Arcuate shaped composites of refractory tapes embedded in a metal matrix
US5440806A (en) Method of making blank for the manufacturing of fiber-reinforced coatings or metal components
US3740301A (en) Elongated lightweight structure
GB1561297A (en) Transition reinforcement of composite blade dovetails
JP2003312590A (en) Reinforcement body for aircraft skin panel and manufacturing method for skin panel provided with reinforcement body
GB2039629A (en) Transpiration cooled blade for a gas turbine and method of its fabrication
US3748721A (en) Method of making composites
US6282786B1 (en) Method of making injection formed hybrid airfoil
US2175204A (en) Air propeller blade
WO2014130292A1 (en) Hybrid turbine blade including multiple insert sections
EP0710728B1 (en) Method for manufacturing a fibre reinforced metal disc preform
US4426193A (en) Impact composite blade