CA2042218A1

CA2042218A1 - Composite airfoil with increased shear capability

Info

Publication number: CA2042218A1
Application number: CA002042218A
Authority: CA
Inventors: Jan C. Schilling; Charles E. Steckle; Paul S. Stephens; Walter D. Howard
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1990-07-20
Filing date: 1991-05-09
Publication date: 1992-01-21
Also published as: DE4122652A1; ITMI911655A0; SE9102142D0; JPH04232400A; GB2249592A; FR2664941A1; BE1004771A3; ITMI911655A1; SE9102142L; GB9115258D0; IT1248496B

Abstract

ABSTRACT OF THE DISCLOSURE

An airfoil blade and method of manufacturing same having greater tolerance to impact damage. The blade comprises a plurality of layers of composite laminates sandwiching at least one layer of a less than rigid bonding agent. The resilient bonding agent allows for relative movement between adjacent layers allowing loading of the laminate to its full strain potential and thereby reducing the shock loading normally transferred to the adjacent laminates by a hard-setting epoxy bonding material. In one form, the resilient material is selectively positioned between layers of laminates at different predetermined locations of the blade. The blade may be stitched to provide additional layer-to-layer support using synthetic cording such as an aeromatic polyamide.

Description

2 ~ 8 - 1 - 13D~l-9676 Co~Po~ITE AIRFOIII WIq!~ INCR~A8E:D ~XE:~ C~PABILITY

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 U58 high strain ~ibers and toughened resins. While these methods increase the blade resistance to ~oreign object damage, they do not increase the blades' ability to inhibit delamination.
Each laminate consists of a single layer of generally ~2~
- 2 - 13DV~9676 longi~udinal fiber elements. The l~minates 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 ~he thickness of the structure by shear forces throuyh the resin system. The resin is weaker in shear than the ~iber and is therefore a weak link in the struckure if transverse loads are applied. Also, the resin is brittle in nature and does not elongate (yield) but breaks. Impac loading caused by birds, ice, or other foreign objects results in very high transverse loading to the blade f iber layers resulting in fracture of the blade. Blade construction and processing also introduces areas which are sub3ect to delamination with impact loading. These areas usually are found between laminates or between laminates and ~ibers, formers, and spars where there i5 a substantial change in strength or shear capability.
FIG. 1 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 propagat~ng. The materials themselves are harder to process due to the toughening additives.

%~J ~l ~
~ 3 - 13DV~9676 SUMM~RY OF ~ V~TIO~

It is an object of the present invention to provide a method and apparatus which inhibit the delamination propagation disa~vantages 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 pres~lected areas alternating layers of a resilient bonding materîal. 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 2S substantially different properties as compaxed to their adjacent laminates, for these sections a more compliant layer of adhesive is placed between laminates to transition/distribute the load.

2 ~
- 4 ~ 13DV-9676 BRI~F ~ESCRIPTION OF ~E D~AWI~G~

For a better understanding o~ the present invention, reference may be had to the following de~ailed descrip~ion 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 interm~diate cloth fiber layer for transitioning from the composite layers to a blade insert;
FIG. 2 is a plan view of one form of air~oil blade with which the present invention may be used;
FIG. 3 is a cross-sectional view along the line 3-3 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 3 6 of FIG r 7 DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

One form of fan blade lO 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 intershangeable with the terms "fan" or 30 ~propeller". Furthermore, while described in application to a propeller or fan blade, the invention 2~2~

is also applicable to other types of air~oil blades such as both ducted and unducted f an blades and compressor blades in gas tuxhine engines. Blade lO
has an airfoil section 12, including a tip ~nd 14, and a root section 16~ The airfoil section 12 ~as 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 arc 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 ko the trailing surface 20, extending radially from root section 16 to tip section 14~ forming a swept back propeller f an blade configuration. The surfaces de~ined by edges 22, 205 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 la~inates 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 iber failure hy 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.
~ypical 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 hiyh sweep or sma~l 2~22~
- 6 - 13D~-9676 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 -45, 0, +45~, 0 with respect to a reference axis.
Two consecutive layers may be layered at the same angle. This form of layering produces an aeroelas$ically 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 ~or 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 for~ of construGtion is illustrat~d in U.S. Patent Serial No. (13DV-9601) assigned to the as~ignee of the present invention.
The cloth 7 in FIG~ 1 may be a tran~ition layer to a foam insert or to a spar or to the outside blade surface.
Turning now to FIG. 3, there is shown a cross-sectional 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 i5 somewhat resilient. Suitable bonding mat~rials may be thermoplastic or thermosetting bonding agents such as polyuret~ane or similar materials having a rubber-like ?,~22~

charackeristic, 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 interm~diate mastic layers 32 of a resilient matPrial betwe~n 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 det~rmined by destructive testing of the airfoil, for example, by impacting th~ blade with an object or by analytical technique~. The u~e 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 successPul in inhibiting delamination propagation is to stitch the laminate layers together. FIG. 5 illustrates th~
blade of FIG. 2 in which stitahing i5 indicated by the dashed lines 34. A high strength synthetic thread may b~ used to sew the laminate layers together in selected locations prior to final bonding or molding of the blade. Stitching alon~ may not increase the blade's capability to deter the onset of d~lamination (local) hut it will increase the bladels 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 impact~ [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, larye 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 o~ the structure versus potential tearing of the blade resulting in liberation of a large klade segment. Stitching is useful in inhibiti~g delamination from everyday type occurrences which only have the potential for damags to the blade leading edge 22, trailing edge 24, and tip region 14O
Stitching i5 not desirable across the main air~oil body of the blade so that the aforementioned delamination can occur in situations where the impact load must be absorbed in order to prevent liberation o~ the entire blade or blade assembly.
FIG. 6 is a cross-~ectional view taken along the lines 6-6 in FIG. 5 and illustrates the through blade arrangement of stitching or threads 34O 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 thP area indicated at 36 in FIG. 7 showing the laminate layers 30, semi-ductile matrix layers 32, and stitching 34.
For blades lO in which stitching is used to inhibit delamination propagation, the blade~ 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 executP the described methods although resin transfer molding i~ the preferred process for blade bonding utilizing the stitching method.

2 ~ ~ C2 ~ ~ 8 While the principles of the invention have now been made clear in an illus~rative embodiment, it will become apparent to those skilled in the art that many modifications of the structures, arrangements, and S components present~d in ~he 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 as set forth in the claims which follow.

Claims

1. An airfoil blade comprising:
a plurality of overlying layers of a non-metallic 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.

2. The airfoil 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 1 wherein said composite material is selected from the group comprising graphite composite material, glass composite material, and boron composite material.

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 1 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. The invention as defined in any of the preceding claims including any further features of novelty disclosed.