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Publication numberUS20050084379 A1
Publication typeApplication
Application numberUS 10/860,650
Publication dateApr 21, 2005
Filing dateJun 4, 2004
Priority dateJun 6, 2003
Also published asDE10326719A1, DE10339046A1, DE502004010031D1, EP1484475A2, EP1508395A2, EP1508395A3, EP1508395B1, US7211178, US20050016842
Publication number10860650, 860650, US 2005/0084379 A1, US 2005/084379 A1, US 20050084379 A1, US 20050084379A1, US 2005084379 A1, US 2005084379A1, US-A1-20050084379, US-A1-2005084379, US2005/0084379A1, US2005/084379A1, US20050084379 A1, US20050084379A1, US2005084379 A1, US2005084379A1
InventorsKarl Schreiber
Original AssigneeKarl Schreiber
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Compressor blade root for engine blades of aircraft engines
US 20050084379 A1
Abstract
On a compressor blade (1) for gas turbine engines which is made of a fiber-composite core (4) with metallic enclosure (5) and which comprises a blade root (3) connected to an airfoil (2) and held in a recess (9) of a compressor disk (10), the fiber-composite core is made of transversely compressed fibers with subsequent matrix infiltration and is intimately connected via a surface enlargement by a micro-structure (7) to inner surfaces of the enclosure (5). Thus, the forces are introduced into the blade root (3) directly via the fiber material. Opposite side faces of the longitudinally straight blade root (3) each feature circularly arched bearing faces (8) with different centers (M1, M2) which mate with similarly arched bearing areas (11) provided in the recess (9) to provide for a limited, dampened pendular movement for the absorption of the bending forces acting upon the fiber-composite core.
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Claims(27)
1. A compressor blade root connection for a gas turbine engine which comprises: a recess of a compressor disk and a blade having a blade root and comprising a fiber-composite core and a metallic enclosure for engagement with the recess of the compressor disk, wherein fibers of the fiber-composite core are packed sufficiently densely in the enclosure that compressive loads on the blade act directly upon the fibers, the enclosure including a micro-structurally enlarged inner surface to provide an intimate contact with the fiber-composite core to ensure transfer of forces between the enclosure and the fiber-composite core, and the blade root is located pendularly and frictionally dampened in the recess of the compressor disk within a limited angular range to reduce a bending stress on the fiber-composite core of the blade root.
2. A compressor blade root connection in accordance with claim 1, wherein the fiber-composite core has a high fiber share in relation to a plastic matrix to apply the pressure load predominately to the fibers.
3. A compressor blade root connection in accordance with claim 2, wherein the fiber-composite core comprises fibers compressed in transverse direction to the compressor blade, with the plastic matrix having been infiltrated under compression of the fibers.
4. A compressor blade root connection in accordance with claim 3, wherein the fiber-composite core comprises carbon fibers in a plastic matrix.
5. A compressor blade root connection in accordance with claim 1, wherein the inner surface of the enclosure is micro-structurally enlarged through at least one of glass-bead peening, sand blasting or an attached wire mesh.
6. A compressor blade root connection in accordance with claim 1, wherein the blade root is straight in its longitudinal direction and has a plurality of circularly arched bearing faces in the area of the recess which have different centers to limit a pendular movement of the blade root, with each bearing face moveably locating against a conformally rounded bearing area of the recess.
7. A compressor blade root connection in accordance with claim 6, wherein at least one of the rounded bearing faces and the bearing areas are micro-structured for frictional dampening of the pendular movement of the blade root.
8. A compressor blade root connection in accordance with claim 1, and further comprising a reinforcing wedge extending from a bottom of the blade root and inserted into the fiber-composite core to divide the fiber-composite core into two partial strands in a longitudinal direction of the blade root.
9. A compressor blade root connection in accordance with claim 8, wherein the reinforcing wedge is micro-structured at faces mating to the fiber-composite core.
10. A compressor blade root connection in accordance with claim 8, wherein a large end of the reinforcing wedge is positioned below a fit of the blade root to the compressor disk to avoid delamination forces under centrifugal load.
11. A compressor blade root connection in accordance with claim 10, and further comprising a metal plate attached at a bottom of the reinforcing wedge to reduce clamping forces acting upon the fiber-composite core under load.
12. A compressor blade root connection in accordance with claim 3, wherein the blade root is straight in its longitudinal direction and has a plurality of circularly arched bearing faces in the area of the recess which have different centers to limit a pendular movement of the blade root, with each bearing face moveably locating against a conformally rounded bearing area of the recess.
13. A compressor blade root connection in accordance with claim 12, wherein at least one of the rounded bearing faces and the bearing areas are micro-structured for frictional dampening of the pendular movement of the blade root.
14. A compressor blade root connection in accordance with claim 13, and further comprising a reinforcing wedge extending from a bottom of the blade root and inserted into the fiber-composite core to divide the fiber-composite core into two partial strands in a longitudinal direction of the blade root.
15. A compressor blade for a gas turbine engine which comprises:
a metallic enclosure for engagement with a recess of a compressor disk,
a fiber-composite core, wherein fibers of the fiber-composite core are packed sufficiently densely in the enclosure that compressive loads on the blade act directly upon the fibers, the enclosure including a micro-structurally enlarged inner surface to provide an intimate contact with the fiber-composite core to ensure transfer of forces between the enclosure and the fiber-composite core, the blade root being constructed and arranged to be located pendularly and frictionally dampened in the recess of the compressor disk within a limited angular range to reduce a bending stress on the fiber-composite core of the blade root.
16. A compressor blade in accordance with claim 15, wherein the fiber-composite core has a high fiber share in relation to a plastic matrix to apply the pressure load predominately to the fibers.
17. A compressor blade in accordance with claim 16, wherein the fiber-composite core comprises fibers compressed in transverse direction to the compressor blade, with the plastic matrix having been infiltrated under compression of the fibers.
18. A compressor blade in accordance with claim 15, wherein the inner surface of the enclosure is micro-structurally enlarged through at least one of glass-bead peening, sand blasting or an attached wire mesh.
19. A compressor blade in accordance with claim 15, wherein the blade root is straight in its longitudinal direction and has a plurality of circularly arched bearing faces in the area of the recess which have different centers to limit a pendular movement of the blade root, with each bearing face moveably locatable against a conformally rounded bearing area of the recess.
20. A compressor blade in accordance with claim 19, wherein the rounded bearing faces are micro-structured for frictional dampening of the pendular movement of the blade root.
21. A compressor blade in accordance with claim 15, and further comprising a reinforcing wedge extending from a bottom of the blade root and inserted into the fiber-composite core to divide the fiber-composite core into two partial strands in a longitudinal direction of the blade root.
22. A compressor blade in accordance with claim 21, wherein the reinforcing wedge is micro-structured at faces mating to the fiber-composite core.
23. A compressor blade in accordance with claim 22, wherein a large end of the reinforcing wedge is positioned below a fit of the blade root to the compressor disk to avoid delamination forces under centrifugal load.
24. A compressor blade in accordance with claim 23, and further comprising a metal plate attached at a bottom of the reinforcing wedge to reduce clamping forces acting upon the fiber-composite core under load.
25. A compressor blade in accordance with claim 17, wherein the blade root is straight in its longitudinal direction and has a plurality of circularly arched bearing faces in the area of the recess which have different centers to limit a pendular movement of the blade root, with each bearing face moveably locatable against a conformally rounded bearing area of the recess.
26. A compressor blade in accordance with claim 25, wherein the rounded bearing faces are micro-structured for frictional dampening of the pendular movement of the blade root.
27. A compressor blade in accordance with claim 26 and further comprising a reinforcing wedge extending from a bottom of the blade root and inserted into the fiber-composite core to divide the fiber-composite core into two partial strands in a longitudinal direction of the blade root.
Description

This application claims priority to German Patent Application DE10326719.0 filed Jun. 6, 2004, the entirety of which is incorporated by reference herein.

BACKGROUND OF THE INVENTION

This invention relates to a compressor blade root for engine blades of aircraft engines which comprise a fiber-composite core with metallic enclosure and which is held in a recess of a compressor disk.

In order to save weight, the fan blades (compressor blades) of gas turbine engines are, as is generally known, made of fiber-reinforced plastic with a metallic enclosure. The blade is firmly attached via a dovetail-style, longitudinally curved blade root in a conformal recess of the compressor disk (rotor). Centrifugal forces, gas pressure and vibrations of the airfoil excited by the flow medium subject the compressor blades of turbomachines to considerable stresses which, in the case of aircraft engines, can become excessive in the event of a bird strike, i.e. an impingement of a bird onto the rotor wheel, leading to the destruction of the respective blade and to serious consequential damage.

Patent Specification EP 0 353 672 B1, for example, describes a blade made of fiber-reinforced material with a blade core and a blade shaft which extends radially to the rotor and is firmly attached to the rotor. An essentially radial fiber strand extends through the blade shaft and the blade core and is covered by an outer skin which, at the transition to the rotor, is designed as a flange which is fixed to the rotor.

In the case of a fiber-material compressor blade known from Specification U.S. Pat. No. 4,040,770, reinforcing wedges made of a homogenous material, for example titanium, are arranged in the area of the dovetail-style blade root between the fiber layers in order to provide a load-carrying structure in this highly loaded area. On the outside of the blade root, a bearing made of the same homogenous material is provided to protect the fiber layers and to obtain a uniform transfer of forces between the blade root and the compressor disk. The dovetail-style blade root features longitudinal, plane faces at its sides and is, thus, firmly fixed in the conformal recess of the compressor disk, enabling the latter to fully take up the loads acting upon the blade. Between the fiber material and the inserts, or bearings, made of homogenous material, a transition layer is provided whose modulus of elasticity is greater than that of the fiber material and smaller than that of the material of the inserted wedges or bearings. The purpose of this transition layer material is to avoid thermal stresses at the interface between the fiber material and the homogenous reinforcing or bearing material, to improve the load-carrying capacity of the blades and to prolong their service life.

The specific problem with known compressor blades, in particular those featuring a carbon-fiber reinforced plastic core, is the inability of this composite material, or the respective compressor blade, to take up and transfer the high forces occurring under extreme loads. Firstly, fatigue stability and dynamic strength of the fiber-composite core are inadequate, with the high tensile forces encountered by the blade not being properly transferred to the blade root, i.e. from the composite material to the metallic portion of the blade root and from there to the compressor disk. Secondly, the fiber-composite material is not capable of taking up the high bending loads resulting from a bird strike. Finally, high clamping forces and local surface pressures occur at the attachment of the blade root to the compressor disk which result in both fatigue of the fiber-composite material and high fretting wear between the blade root and its location in the compressor disk.

BRIEF SUMMARY OF THE INVENTION

It is a broad aspect of the present invention to provide an engine blade whose compressor blade root features a fiber-composite core and a metallic enclosure such that the blade is capable of taking up and transferring high forces while having a long service life.

It is a particular object of the present invention to provide solution to the above problems by a compressor blade root designed in accordance with the features described herein. Further features and advantageous embodiments of the present invention will become apparent from the description below.

In other words, the concept underlying the present invention with respect to a compressor blade root of the type mentioned above is to have the pressure loads present at the blade root directly act upon the fibers in that the fibers are compressed and the matrix material is infiltrated into the minute cavities subsequently to compression and to introduce the forces transferred by the fiber-composite core to the blade root via an as intimate as possible contact between the fiber-composite core and the enclosure, this intimate contact being produced by a micro-structure applied to the inner surfaces of the enclosure. Since this design of a fiber-composite core is not capable of withstanding high bending loads, it is pendularly located within a limited angular range and frictionally dampened in the recess for the location of the blade root in the compressor disk.

This fiber-composite core enables, on the one hand, high forces to be taken up and transferred to the enclosure in the area of the blade root, i.e. introduced in the metallic part of the blade root. On the other hand, the bending load of the fiber material is reduced by the frictionally dampened pendular movement of the blade root allowed within a limited range in the recess of the compressor disk.

A compressor blade featuring such a compressor blade root is very light, which is favorable for the operation of the gas turbine engine. Such a compressor blade is nevertheless capable of withstanding and transferring, without damage, the high forces to which it is exposed. It has high dynamic strength and fatigue stability, as well as a long service life. Stress and wear in the recess locating the blade root are reduced.

According to the present invention, the bending stresses acting upon the compressor blade are compensated by a frictionally dampened pendular movement of the compressor blade with respect to the compressor disk, which is enabled by the circularly arched side flanks of the longitudinally straight blade root sliding on correspondingly circularly arched bearing areas of the recess in the compressor disk against a frictional resistance.

BRIEF DESCRIPTION OF THE DRAWING

The present invention is more fully described in light of the accompanying drawing showing an embodiment, with the only figure illustrating, in sectional side view, a bottom part of a compressor blade comprising a fiber-composite core and a metallic enclosure, with the root of the compressor blade conformally mating a recess provided on the periphery of a compressor disk.

DETAILED DESCRIPTION OF THE INVENTION

The compressor blade 1 comprises an airfoil 2 and a blade root 3 and includes a fiber-composite core 4 and a metallic enclosure 5 with a thin-walled portion 5 a in the area of the airfoil 2 and with a thick-walled portion 5 b in the area of the blade root 3. In the blade root 3, a reinforcing wedge 6 divides the fiber-composite core 4 into two partial strands 4 a, 4 b to provide for high strength in the root area and to enable the high tensile forces acting upon the airfoil 2 to be effectively introduced into the blade root 3. Upon infiltration, a metal plate 12 is attached (for instance, by welding or other means) at the bottom end of the blade, preferably to the reinforcing wedge, to take up the clamping forces produced by the centrifugal loads and, thus, to reduce the mechanical pressure load on the carbon fibers.

Preferably, a large end of the wedge 6 is positioned below a fit of the blade root 3 to the compressor disk 10 to avoid delamination forces under centrifugal load.

The metallic enclosure 5 and the reinforcing wedge 6 are here made of titanium, while a composite of carbon fibers embedded in plastic is used for the fiber-composite core 4. The fiber-composite core 4 is, however, manufactured such that a fiber strand compressed in transverse direction to the compressor blade 1 is initially produced under high pressure and that a plastic is infiltrated into the remaining cavities of the strand subsequently to the high compression of the fiber material. The high packing density of the fiber structure within the plastic matrix with a relatively low plastic share enables the compressor blade 1, i.e. the fiber material, to take up high tensile forces, with high dynamic strength and fatigue stability being ensured since the compressive stresses have to be transmitted only partially via the plastic matrix.

For the generation of a large specific surface for the purpose of an intimate contact and transfer of forces between the fiber-composite core 4 and the metallic enclosure 5, the inner surfaces of the metallic enclosure 5 are micro-structured at least in the area of the blade root 3 by means of a mechanical treatment such as glass-bead peening or a sand blasting process, or by means of a welded-on fine wire mesh 7, or by other processes/mechanisms, either mechanical, chemical or electrical. Also, the outer surfaces of the reinforcing wedge 6 mating with the fiber structure can be micro-structured by means of such a wire mesh 7, or by means described above or in a different manner.

As becomes apparent from the drawing, the blade root 3 is located via circularly arched bearing faces 8 in a recess 9 of a compressor disk 10 whose two sideward bearing areas 11 are conformally rounded. The radii R1 of the laterally opposite rounded bearing faces 8 are identical, but do not have a common center, i.e. the centers M1, M2 of the circularly arched bearing faces 8 are remote from each other to radially center the blade under centrifugal load. For the functioning of the blade root 3 in the above-described form, it is crucial that the blade root 3 is straight. This shape allows the compressor blade 1 to perform a pendular movement in the recess 9 of the compressor disk 10 which, however, is limited by the different centers M1 and M2, with these remote centers M1, M2 also having the centering effect described above. If the airfoil 2 is subject to a shock, for example in the event of a bird strike, the shock effect is absorbed by the pendular movement of the airfoil 2 within the limited range available and by the friction on a large surface between the bearing faces 8 of the blade root 3 and the bearing areas 11 in the recess 9. Thus, the fiber-composite material, which does not endure bending loads, is relieved to a considerable extent. To increase the friction between the bearing faces of the blade root and the bearing areas 11 of the recess 9, the surfaces of one or both of the bearing faces 8 and the bearing areas 11 can be micro-structured in one of the manners discussed above, or in another manner.

The combination of the above features, i.e. the transversely compressed fiber material with low matrix share, the arrangement of the reinforcing wedge, the surface enlargement at the mating faces between the fiber-composite material, the blade root and the reinforcing wedge and the longitudinally straight blade root with opposite, arched bearing and friction faces, provides a low-weight compressor blade which is capable of effectively transmitting forces and, despite the high loads applied, has a long service life.

List of Reference Numerals

  • 1 Compressor blade (fan blade)
  • 2 Airfoil
  • 3 Blade root
  • 4 Fiber-composite core
  • 4 a, 4 b Partial strands of 4
  • 5 Metallic enclosure
  • 5 a Thin-walled portion of 5
  • 5 b Thick-walled portion of 5
  • 6 Reinforcing wedge
  • 7 Wire mesh, micro-structuring
  • 8 Circularly arched bearing faces (sliding/friction surfaces)
  • 9 Recess of 10
  • 10 Compressor disk (fan disk)
  • 11 Bearing area of 9
  • 12 Metal plate
  • R1 Radii of 8
  • M1, M2 Centers of 8
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US7306434Feb 1, 2005Dec 11, 2007Rolls-Royce PlcReduction of co-efficient of friction to reduce stress ratio in bearings and gas turbine parts
US7828526 *Apr 11, 2007Nov 9, 2010General Electric CompanyMetallic blade having a composite inlay
US8011877 *Nov 24, 2008Sep 6, 2011General Electric CompanyFiber composite reinforced aircraft gas turbine engine drums with radially inwardly extending blades
US8100662Oct 18, 2007Jan 24, 2012Rolls-Royce Deutschland Ltd & Co KgFan blade made of a textile composite material
US8387467 *Feb 25, 2009Mar 5, 2013SnecmaMethod for testing the coating of a vane base
US20100189562 *Jan 11, 2010Jul 29, 2010SnecmaComposite material turbomachine blade with a reinforced root
US20110138926 *Feb 25, 2009Jun 16, 2011SnecmaMethod for testing the coating of a vane base
US20110229337 *Jun 30, 2010Sep 22, 2011General Electric CompanyHybrid ceramic matrix composite turbine blades for improved processibility and performance and process for producing hybrid turbine blades
Classifications
U.S. Classification416/230
International ClassificationB23H3/04, C23C14/00, B23H9/00, B23H3/00, C25B9/00, B64C11/26, F01D5/14, B64C11/04, F01D5/30, F04D29/34, F01D5/28
Cooperative ClassificationY02T50/672, B23H3/00, B23H9/00, F04D29/023, B23H3/04, F01D5/282, B23H9/006, F04D29/34, F01D5/30
European ClassificationF04D29/34, F04D29/02C, F01D5/30, F01D5/28B
Legal Events
DateCodeEventDescription
Nov 2, 2004ASAssignment
Owner name: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG, GERMANY
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:SCHREIBER, KARL;REEL/FRAME:015950/0550
Effective date: 20040811