EP1673486B1 - Procede de realisation et utilisation associes - Google Patents

Procede de realisation et utilisation associes Download PDF

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Publication number
EP1673486B1
EP1673486B1 EP04786887A EP04786887A EP1673486B1 EP 1673486 B1 EP1673486 B1 EP 1673486B1 EP 04786887 A EP04786887 A EP 04786887A EP 04786887 A EP04786887 A EP 04786887A EP 1673486 B1 EP1673486 B1 EP 1673486B1
Authority
EP
European Patent Office
Prior art keywords
disks
substrate
fibre
disk
recess
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 - Fee Related
Application number
EP04786887A
Other languages
German (de)
English (en)
Other versions
EP1673486A1 (fr
Inventor
Joachim Bamberg
Falko Heutling
Josef Mayr
Klaus-Dieter Tartsch
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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 MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Publication of EP1673486A1 publication Critical patent/EP1673486A1/fr
Application granted granted Critical
Publication of EP1673486B1 publication Critical patent/EP1673486B1/fr
Expired - Fee Related legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • C22C47/02—Pretreatment of the fibres or filaments
    • C22C47/025—Aligning or orienting the fibres
    • 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
    • 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
    • C22C47/20—Making alloys containing metallic or non-metallic fibres or filaments by subjecting to pressure and heat an assembly comprising at least one metal layer or sheet and one layer of fibres or 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12444—Embodying fibers interengaged or between layers [e.g., paper, etc.]
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/21—Circular sheet or circular blank
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/21—Circular sheet or circular blank
    • Y10T428/211—Gear
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/21—Circular sheet or circular blank
    • Y10T428/218—Aperture containing

Definitions

  • the invention relates to a method for producing a composite material.
  • titanium alloys The most important materials used today for aircraft engines or other gas turbines are titanium alloys, nickel alloys (also called superalloys) and high-strength steels.
  • the high-strength steels are used in particular for shaft parts and gear parts and for compressor casings and turbine casings.
  • Titanium alloys are typical materials for compressor parts, nickel alloys are suitable for the hot parts of the aircraft engine.
  • Fiber-reinforced composites A very promising group of future generations of aircraft engine material is so-called fiber-reinforced composites.
  • Modern composites have a carrier material, which may be formed as a polymer, a metal or a ceramic matrix, as well as embedded in the carrier material fibers.
  • the present invention relates to the production of a composite material in which the carrier material is formed as a metal matrix.
  • a material is also called a metal matrix composite material - called MMC for short.
  • MMC metal matrix composite material
  • component weight can be reduced by up to 50% over traditional titanium alloys.
  • Reinforcements used are fibers having high strength and high modulus of elasticity.
  • Such fiber-reinforced composite materials are already known from the prior art. So revealed the EP 0 490 629 B1 a preform for a composite with a film, the film having a groove and a thread-like reinforcement disposed in the groove, and wherein the preform has the shape of a ring or a disc. To produce a multilayer composite structure, according to the EP 0 490 629 B1 proceeded so that several such preforms are superimposed, wherein the preforms are solidified under heat and pressure to form a completely dense composite material. Other composites and methods of making the same are known from EP 0 909 826 B1 , of the US 4,697,324 and the US 4,900,599 known.
  • the document US 2002/0031678 A1 also discloses a method for making a composite over film-like preforms having a groove and a reinforcing fiber inserted into the groove.
  • the groove is cut by cutting into the film, wherein the film is clamped on a rotating tool or workbench.
  • several films provided with fibers are stacked on each other and consolidated materially. Safe testing of the multilayer material for fiber breaks or material cracks is practically impossible.
  • the present invention is based on the problem to provide a novel method for producing composite materials, with which an exact position of the fibers and a low scrap of material can be achieved.
  • a recess is introduced into the disc, the depth of which is greater than the diameter of the fiber, such that, when a fiber is inserted into the recess, webs of carrier material protrude beyond the fiber.
  • the disks are stacked in such a way that the fibers of the stacked disks protrude radially differently into the carrier material in a radially outer section for strength-optimizing toothing.
  • the composite material produced according to the invention has a carrier material made of titanium or a titanium alloy as well as a plurality of fibers embedded in the carrier material.
  • the fibers are preferably silicon carbide ceramic fibers.
  • the composite material according to the invention is formed from a plurality of discs of carrier material, wherein a fiber is embedded in each disc. Several such disks with a fiber embedded therein are stacked and bonded together to form the composite. For embedding the fiber in the respective disc of carrier material, a recess is made in the disc. In the recess, the corresponding fiber is inserted and surrounded on all sides by carrier material, so that the fiber is embedded in the disc.
  • Fig. 1 shows a disc 10 of support material, namely titanium, in a highly schematic cross-section. In a central region, the disc 10 has a bore 11th
  • a recess is introduced into a front side 12 of the pane 10 after a first step of the method according to the invention.
  • Fig. 2 shows a greatly enlarged detail of the disc 10 in the region of the end face 12.
  • the recess 13 which is introduced into the end face 12 of the disc 10, is a spiral groove. The spiral groove therefore extends exclusively on an end face 12 of the disc 10 from the inside to the outside.
  • a fiber 14 is inserted into the spiral-shaped recess 13.
  • Fig. 3 can be seen that webs 15 protrude from carrier material with fiber 14 inserted over the fiber 14. The depth of the spiral recess 13 is therefore greater than the diameter of the fiber 14.
  • the arrangement according to Fig. 3 subjected to a superplastic forming process is heated to a forming temperature and uniaxial pressing the webs 15 are so superplastic reshaped that subsequently the fiber 14 in the sense of Fig. 5 Surrounded on all sides by carrier material and thus the fiber 14 is embedded in the carrier material.
  • Fig. 5 can be seen that the position of the fiber 14 is maintained even after the superplastic forming of the webs 15.
  • the carrier material is compacted.
  • Fig. 4 shows a disc 10 made of carrier material with the fiber 10 embedded in the disc 10 in a highly schematic cross-section.
  • the fiber 14 is surrounded on all sides by carrier material and thus embedded in the carrier material.
  • the stacked and stacked disks 10 are then joined by diffusion welding under low axial pressure or connected to each other. This will ultimately provide the composite.
  • the discs 10 Before the stacking of the discs 10 in the sense of Fig. 6 Preferably, an examination of the discs 10 with the fibers 14 embedded in the discs 10 for cracks in the carrier material and for breaks in the fibers 14. This review can done with ultrasound, X-ray or tomography. If such a crack or break found, the disc 10 is discarded. If it is determined during the inspection that there is no crack or break in the fiber 14, then the disc 10 can be used for stacking.
  • Fig. 7 shows a section of the arrangement according to Fig. 6 in the range of three superimposed and interconnected slices 10. So can Fig. 7 it can be seen that the fiber 14 embedded in a disk 10 extends radially offset from the fibers 14 of the two adjacent disks 10. As a result, a hexagonal packing of the fibers 14 can be achieved. As Fig. 7 can be removed, a fiber 14 extends spirally within a disc 10, that in cross-section, the resulting center points of the fiber 14 of a disc 10 between the corresponding centers of the fiber 14 of an adjacent disc 10 are arranged.
  • each fiber 14 within each disc 10 terminates at a distance to an outer, lateral end of the respective disc. According to Fig. 6 this distance is different for each slice. Adjacent to the inner opening 11, however, the lateral distance of the fibers 14 to the opening 11 is the same. Due to the different lateral distances between the fibers 14 and the radially outer, lateral end of the discs 10, gradual changes in the elastic properties of the composite can be achieved. Furthermore, a toothing between the unreinforced and fiber reinforced areas of the composite is achieved, which positively affects the strength properties.
  • Fig. 8 shows a highly schematic cross section through a composite material according to the invention. This was prepared as described above. According to Fig. 8 In an internal section 16 of the composite material, the fibers 14 are embedded in the carrier material. In an outer section 17, however, only the carrier material is present. This means that in the outer portion 17 is present only titanium. This is advantageous if the composite material is subjected to further processing, for example by milling should. When milling, namely the fibers 14 may not be damaged. A subsequent milling of the composite is therefore only in the range of section 17 into consideration, in which only the carrier material is present. Furthermore, can Fig.
  • a fiber made of silicon carbide is inserted into this spiral recess.
  • the disc with the fiber inserted into the disc is consolidated by superplastic forming.
  • the fiber is surrounded on all sides by carrier material or embedded in the carrier material.
  • the discs thus produced with fibers embedded in the discs are checked for cracks in the carrier material and fractures in the fibers. If this test shows that there is neither a crack nor a fiber break, the corresponding discs are stacked in rings.
  • the stacking of a plurality of rings is then subjected to a diffusion welding in a further step of the method according to the invention, so that adjacent disks are joined together.
  • a finishing of the composite material for example by milling.
  • the inventive method is reliable and inexpensive.
  • the method according to the invention is a fully automated process with integrated verification and thus quality assurance. Since each disc can be inspected for quality, defects in the composite can be timely be recognized and thus avoided. Scrap is reduced with it.
  • Another advantage is the fact that an exact position of the fibers in the composite material is predetermined and adhered to.
  • more complex fiber guides for example, star-shaped fiber guides are possible.
  • In the invention can be dispensed with a titanium coating of the fibers, as required in the prior art.
  • Another advantage is that no extremely long fibers need to be used. By guiding the fibers in recesses, fibers of finite length can be used.
  • the composite material produced according to the invention is accordingly distinguished by an exact position of the fibers within the carrier material.
  • the composite material is formed by a plurality of joined discs of carrier material, wherein within each disc a spirally extending fiber is embedded.
  • the fibers terminate at a distance from a lateral, radially outer end of the composite, so that in an outer region of the same only the carrier material is present, in which area a subsequent milling of the composite material can take place.
  • a plurality of fibers can also be embedded in a recess, and that a plurality of interleaved recesses can also be introduced into a disk, wherein each of these recesses can in turn receive one or more fibers.
  • the illustrated embodiment, in which each disc has a recess for receiving a fiber is preferred.
  • the composite material produced according to the invention is particularly suitable for use as a material in the manufacture of rings with integral blading for aircraft engines, which are also referred to as bladed rings (blings).

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • Laminated Bodies (AREA)
  • Manufacture Of Alloys Or Alloy Compounds (AREA)
  • Casting Or Compression Moulding Of Plastics Or The Like (AREA)

Claims (8)

  1. Procédé de fabrication d'un matériau composite composé de plusieurs tranches (10) assemblées en matériau support, au moins un évidement (13) étant réalisé de préférence dans chaque tranche pour recevoir au moins une fibre (14), comportant les étapes suivantes:
    a) mise à disposition de plusieurs tranches (10) en matériau support,
    b) réalisation d'au moins un évidement (13) dans de préférence chaque tranche (10) et introduction consécutive d'au moins une fibre (14) dans le ou chaque évidement (13) de la tranche (10) correspondante,
    c) consolidation de la tranche correspondante (10) de telle sorte que la ou chaque fibre (14) soit entourée de tous côtés par le matériau support ou soit encastrée dans le matériau support de la tranche (10) correspondante,
    d) empilage des tranches (10) consolidées
    e) raccordement des tranches (10) empilées par une étape d'assemblage.
  2. Procédé selon la revendication 1, caractérisé en ce qu'en relation avec l'étape b), un évidement (13) est réalisé dans la tranche (10), dont la profondeur est supérieure au diamètre de la fibre (14) de telle sorte que lorsque la fibre (14) est introduite dans l'évidement (13), des traverses (15) en matériau support dépassent depuis la fibre (14).
  3. Procédé selon la revendication 1 ou 2, caractérisé en ce qu'en relation avec l'étape c), le matériau support est soumis avec la ou chaque fibre (14) introduite dedans à une déformation superplastique de telle sorte que la ou chaque fibre (14) soit entourée de tous côtés par le matériau support.
  4. Procédé selon l'une quelconque des revendications 1 à 3, caractérisé en ce qu'en relation avec l'étape d), les tranches (10) en matériau support avec au moins une fibre (14) encastrée dedans sont disposées l'une sur l'autre, notamment sont empilées pour former un anneau ou un cylindre creux.
  5. Procédé selon l'une quelconque des revendications 1 à 4, caractérisé en ce qu'en relation avec l'étape d), les tranches (10) sont empilées de telle sorte que les fibres (14) des tranches (10) empilées pénètrent radialement à une profondeur différente dans une section extérieure (17) radiale d'une denture optimisant la résistance dans le matériau support.
  6. Procédé selon l'une quelconque des revendications 1 à 5, caractérisé en ce qu'en relation avec l'étape e), les tranches (10) empilées sont assemblées par soudage par diffusion.
  7. Procédé selon l'une quelconque des revendications 1 à 6, caractérisé en ce que des ruptures dans la ou chaque fibre et/ou des fissures dans le matériau support des tranches (10) en matériau support avec au moins une fibre (14) encastrée dedans sont vérifiées avant la liaison de celles-ci avec d'autres tranches, et en ce qu'en cas de constatation d'une fissure ou d'une rupture, la tranche est rebutée.
  8. Procédé selon l'une quelconque des revendications 1 à 7, de fabrication de composants en forme de disque, d'anneau ou à symétrie de révolution avec aubage intégré, c'est-à-dire de dits Bladed Rings (Blings) ou Bladed Disks (Blisks).
EP04786887A 2003-10-18 2004-09-30 Procede de realisation et utilisation associes Expired - Fee Related EP1673486B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10348506A DE10348506A1 (de) 2003-10-18 2003-10-18 Verbundwerkstoff, Verfahren zur Herstellung eines Verbundwerkstoffs und Verwendung desselben
PCT/DE2004/002175 WO2005040444A1 (fr) 2003-10-18 2004-09-30 Materiau composite, procede de realisation et utilisation associes

Publications (2)

Publication Number Publication Date
EP1673486A1 EP1673486A1 (fr) 2006-06-28
EP1673486B1 true EP1673486B1 (fr) 2008-04-09

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EP04786887A Expired - Fee Related EP1673486B1 (fr) 2003-10-18 2004-09-30 Procede de realisation et utilisation associes

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Country Link
US (1) US7524566B2 (fr)
EP (1) EP1673486B1 (fr)
DE (2) DE10348506A1 (fr)
WO (1) WO2005040444A1 (fr)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102008052247A1 (de) * 2008-10-18 2010-04-22 Mtu Aero Engines Gmbh Bauteil für eine Gasturbine und Verfahren zur Herstellung des Bauteils
US9249684B2 (en) 2013-03-13 2016-02-02 Rolls-Royce Corporation Compliant composite component and method of manufacture

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3419952A (en) * 1966-09-12 1969-01-07 Gen Electric Method for making composite material
US4697324A (en) 1984-12-06 1987-10-06 Avco Corporation Filamentary structural module for composites
US4900599A (en) 1986-11-21 1990-02-13 Airfoil Textron Inc. Filament reinforced article
US4919594A (en) 1987-05-15 1990-04-24 Allied-Signal Inc. Composite member, unitary rotor member including same, and method of making
US5337940A (en) * 1990-12-11 1994-08-16 Woods Harlan L Composite preform and method of manufacturing fiber reinforced composite
US5431984A (en) * 1990-12-11 1995-07-11 Avco Corporation Composite preforms with groves for fibers and groves for off-gassing
US6261699B1 (en) * 1999-04-28 2001-07-17 Allison Advanced Development Company Fiber reinforced iron-cobalt composite material system
AU2001296868A1 (en) 2000-09-11 2002-03-26 Allison Advanced Development Company Mechanically grooved sheet and method of manufacture

Also Published As

Publication number Publication date
US20070141298A1 (en) 2007-06-21
DE502004006792D1 (de) 2008-05-21
DE10348506A1 (de) 2005-05-12
US7524566B2 (en) 2009-04-28
EP1673486A1 (fr) 2006-06-28
WO2005040444A1 (fr) 2005-05-06

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