GB2249592A

GB2249592A - Composite airfoil blade.

Info

Publication number: GB2249592A
Application number: GB9115258A
Authority: GB
Inventors: Jan Christopher Schilling; Charles Evan Steckle; Paul Stanley Stephens; Walter Douglas Howard
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1990-07-20
Filing date: 1991-07-15
Publication date: 1992-05-13
Also published as: CA2042218A1; DE4122652A1; BE1004771A3; SE9102142L; IT1248496B; GB9115258D0; SE9102142D0; FR2664941A1; JPH04232400A; ITMI911655A0; ITMI911655A1

Description

224 95 92 - 1 COXPOSITE AIRFOIL WITH INCREASED SHEAR CAPABILITY The

present invention relates to blades for fluid flow 'machines and, more particularly, to a blade of non-metallic composite material laminates embedded in a resin matrix with increased capability for tolerating foreign object impact.

Propeller and gas turbine engine fan blades of composite materials such as.graphite or fiberglass are desirable replacements for metal blades. These composite materials have high strength characteristics and are significantly lighter in weight than their metal counterparts. However, one area in which composite blades are not as satisfactory as metal blades is their resistance to foreign object impact. The composite blade when impacted tends to delaminate primarily at the laminate to resin interface.

Typically, a composite blade is fabricated by bonding together a plurality of substantially parallel filament laminates. Standard methods applicable to blades are to use high strain fibers and toughened resins. While these methods increase the blade resistance to foreign object damage, they do not increase the blades' ability to inhibit delamination. Each laminate consists of a single layer of generally 1 longitudinal fiber elements. The laminates are joined together by means of a resin matrix. When the structure is loaded normal to the laminate direction, the load must be transferred through the thickness of the structure by shear forces through the resin system. The resin is weaker in shear than the fiber and is therefore a weak link in the structure if transverse loads are applied. Also, the resin is brittle in nature and does not elongate (yield) but breaks. Impact loading caused by birds, ice, or other foreign objects results in very high transverse loading to the blade fiber layer-s resulting in fracture of the blade. Blade construction a-nd processing also introduces areas which are subject to delamination with impact loading. These areas usually are found between laminates or between laminates and fibers, formers, and spars where there is a substantial change in strength or shear capability. FIG. I illustrates a typical transition area in which two overlaying unidirectional high-strength laminate layers 8, 9 transition to a cloth 7 and create an interface of high load transition.

one proposed solution to impact loading is to construct the blade of higher impact strength (tougher) matrix materials. These materials greatly increase the threshold for delamination initiation but do not inhibit the delamination from propagating. The materials themselves bLre harder to process due to the toughening additives.

i h9UMMARY OF TM INVENTION is It is an object of the present invention to provide a method and apparatus which inhibit the delamination propagation disadvantages of composite blades upon impact.

It is another object of the present invention to provide increased capability for a fan (prop) blade to withstand the high impact energies of foreign objects (delamination threshold, i.e., increased toughness) and further to provide blade integrity after onset of delamination (inhibit delamination propagation).

In general, the above and other objects are attained in a composite airfoil blade comprising a plurality of layers of a composite laminate which incorporates in preselected areas alternating layers of a resilient bonding material. In one form, the blade may be stitched through the alternating layer areas with a high strength, resilient thread to link the layers prior to molding of the blade. In another form, the laminates may be braided in a three dimensional matrix and a resin injected into the matrix using resin transfer molding. Braiding of the laminates adds through thick load capability using the base fiber elements. Some fiber laminates have substantially different properties as compared to their adjacent laminates, for these sections a more compliant layer of adhesive is placed between laminates to transition/distribute the load.

BRIEF DESCRIPTION QE M DRAWINGS

For a better understanding of the present invention, reference may be had to the following detailed description taken in conjunction with the accompanying drawings in which:

FIG. 1 illustrates a pair of overlying layers of composite material in an alternating pattern and an intermediate cloth f iber layer for transitioning from the composite layers to a blade insert; FIG. 2 is a plan view of one form of airfoil blade with which the present invention may be used; FIG. 3 is a cross-sectional view along the line 33 of FIG. 2 showing blade construction in accordance with the present invention; FIG. 4 is an expanded view of the area 28 of FIG.

3; FIG. 5 is an alternate embodiment of the blade of FIG. 2 illustrating the use of stitching to assist in inhibiting delamination of blade layers; FIG. 6 is a cross-sectional View of FIG. 5 taken along the line 6-6; FIG. 7 is a cross-sectional view of FIG. 5 taken along line 7-7; and FIG. 8 is an expanded view of area 36 of FIG. 7.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

One form of fan blade 10 with which the present invention may be used is illustrated in FIG. 2. Reference made herein to "fan blade" is intended to be synonymous and interchangeable with the terms "fan" or Opropellern. Furthermore, while described in application to a propeller or fan blade, the invention is also applicable to other types of airfoil blades such as both ducted and unducted fan blades and compressor blades in gas turbine engines. Blade 10 has an airfoil section 12, including a tip end 14, and a root section 16. The airfoil section 12 has a leading surface 22 and a trailing surface 20. The blade 10 is comprised of a plurality of angle plied composite laminates of continuous fibers which are embedded in a matrix material. In the illustrative embodiment, the continuous fibers of the composite laminates extend across the entire airfoil from the leading surface 22 to the trailing surface 20, extending radially,from root section 16 to tip section 14, forming a swept back propeller fan blade configuration. The surfaces defined by edges 22, 20, 14, and 16 are the airfoil pressure side surface 18 and the airfoil suction side surface 24. In some embodiments, the surfaces 18 and 24 form a composite shell in which a foam/metal blade spar 26 may be inserted and bonded to the inner laminates for shaping the blade or for providing a structural attachment for a blade hub.

Typically, fibers of a composite laminate are unidirectional, side-by-side, parallel, and encased in a semi-ductile low strength, low modulus matrix material which transfers and localizes the effect of a single fiber failure by redistributing the load near the failed fiber to adjacent-fibers. The fibers have a modulus of elasticity from about 10 x 106 p.s.i. for glass to about 44 x 106 P.s.i. for modern graphite.

Typical fibers are comprised of graphite, boron, or S glass. Graphite fibers with a modulus of elasticity 44 x 106 p.s.i. is preferable. A higher modulus and therefore a higher strength fiber allows for greater geometry accommodations such as high sweep or small edge thickness. The matrix material is typically a thermo-set resin, but could be thermo-plastic.

The laminates may be layered with fibers of each layer being aligned in an alternating pattern such as -450, 00, +450, 00 with respect to a reference axis. Two consecutive layers may be layered at the same angle. This form of layering produces an aeroelastically stable blade with well tuned vibratory modes. FIG. 1 illustrates two adjacent layers 8, 9 forming an alternating pattern. The fibers may be braided in three dimensions such that some fibers penetrate intermediate layers of the composites for added resistance to delamination.

The composite blades may be formed as solid composite blades or may include foam, hollow, or other inserts to reduce weight and/or metal inserts to increase strength or provide a medium for connection of a blade hub. One form of construction is illustrated in U.S. Patent Serial No. (13DV-9601) assigned to the assignee of the present invention. The cloth 7 in FIG. 1 may be a transition layer to a foam insert or to a spar or to the outside blade surface.

Turning now to FIG. 3, there is shown a crosssectional view of the blade of FIG. 2 taken along the lines 3-3. The blade 10 can be seen to comprise a plurality of laminates 30, which may be laminate layers such as those at 8, 9 in FIG. 1. The laminates 30 are bonded together in an epoxy matrix 32. In Applicants' inventive blade, the laminates 30 are bonded by an adhesive material, in selected areas, which is less than rigid. Preferably, the material is somewhat resilient. Suitable bonding 'materials may be thermoplastic or thermosetting bonding agents such as polyurethane or similar materials having a rubber-like characteristic, i.e., a material which can be strained without breaking. FIG. 4 is an enlarged view of the area 28 in FIG. 3 and better illustrates the intermediate mastic layers 32 of a resilient material between laminates 30. The layers 32 are no thicker than the laminates 30. Typically, the laminates are about ten mils thick although they can vary between five mils and twenty mils. The resilient material of layers 32 is selectively positioned between the laminates 30 during the laying up process of the blade. These selected locations can be determined by destructive testing of the airfoil, for example, by impacting the blade with an object or by analytical techniques. The use of the resilient layers 32 increases the capability of the fan blade to withstand higher impact loads without delamination but does not inhibit delamination propagation once initiated.

One method which has been found successful in inhibiting delamination propagation is to stitch the laminate layers together. FIG. 5 illustrates the blade of FIG. 2 in which stitching is indicated by the dashed lines 34. A high strength synthetic thread may be used to sew the laminate layers together in selected locations prior to final bonding or molding of the blade. Stitching alone may not increase the blade's capability to deter the onset of delamination (local) but it will increase the blade's capability to inhibit delamination propagation caused by outside stimulus such as blade vibration. Stitching also greatly increases the capability to absorb larger impacts than those causing the onset of delamination without any further delamination.

Stitching must be used very judiciously in fan or prop blades. From a maintenance viewpoint, average impacts (small to medium birds (4 oz. to 2. 5 lbs)] must be taken with either no damage or damage which does not cause secondary failure with time. For large impacts, 4 to 8 lbs, safety is extremely important. For open rotor systems, large impacts must not cause liberation of the whole part which could cause other damage. Therefore, blade delamination under these circumstances is desirable. During the event, energy is absorbed by delamination of the structure versus potential tearing of the blade resulting in liberation of a large blade segment. Stitching is useful in inhibiting delamination from everyday type occurrences which only have the potential for damage to the blade leading edge 22, trailing edge 24, and tip region 14. Stitching is not desirable across the main airfoil body of the blade so that the aforementionid delamination can occur in situations where the impact load must be absorbed in order to prevent liberation of the entire blade or blade assembly.

FIG. 6 is a cross-sectional view taken along the lines 6-6 in FIG. 5 and illustrates the through blade arrangement of stitching or threads 34. FIG. 7 is a cross-sectional view taken along the lines 7-7 in FIG. 5 and also illustrates the stitching formation in the blade 10. FIG. 8 is an enlarged view of the area indicated at 36 in FIG. 7 showing the laminate layers 30, semi-ductile matrix layers 32, and stitching 34. For blades 10 in which stitching is used to inhibit delamination propagation, the blades may be bonded together after stitching using injection or compression molding.

compression molding or autoclaving and injection molding or resin transfer molding can be used to execute the described methods although resin transfer molding is the preferred process for blade bonding utilizing the stitching method.

While the principles of the invention have now been made clear in an illustrative embodiment, it will become apparent to those skilled in the art that many modifications of the structures, arrangements, and components presented in the above illustrations may be made in the practice of the invention in order to develop alternate embodiments suitable to specific operating requirements without departing from the spirit and scope of the invention, -

Claims

1. An airfoil blade comprising:

a plurality of overlying layers of a non-metal composite material; and at least one layer of less than rigid material positioned between adjacent layers of composite laminates at preselected areas of the blade such that said blade comprises alternating. layers of composite laminates and less than rigid material in said preselected areas.

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2. The ait " foil blade of claim 1 and including a plurality of layers of less than rigid bonding material selectively positioned between layers of laminates.

3. The airfoil blade of claim I wherein said composite material is selected from the group comprising graphite composite material, glass composite material, and boron composite material.

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4. The airfoil blade of claim 1 wherein said resilient material is a thermoplastic or thermosetting binding unit.

5. The airfoil blade of claim 1 and including a plurality of rows of a flexible thread stitched through said blade and interconnecting said plurality of layers.

6. The airfoil blade of claim 4 wherein said thread comprises an aeromatic polyamide.

7. The airfoil blade of claim 1 wherein said layers of composite material comprise fibers braided in three-dimensions with fibers penetrating intermediate layers.

8. The airfoil blade of claim I wherein said resilient layers comprise a resin material deposited on said layers of composite material prior to positioning said layers on said former.

9. The airfoil blade of claim 1 wherein each of said layers of composite material has a thickness between about five and twenty mils.

10. The airfoil blade of claim 8 wherein said resilient layers are no greater than the thickness of said layers of composite material.

11. A method of manufacturing an airfoil blade comprising the steps of: alternately overlaying a plurality of layers of a non-metallic composite material and at least one layer of a resilient bonding material in preselected areas; and stitching through the layers of composite and resilient material to create a preformed blade.

12. An airfoil blade substantially as hereinbefore described with reference to the accompanying drawings.

13. A method of manufacturing an airfoil blade substantially as hereinbefore described with reference to the accompanying drawings.

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