Abstract:
A hollow core fan blade for a gas turbine engine, having stitched composite skins and substructure, is fabricated using a resin transfer molding approach which results in a damage tolerant and cost-efficient structure. The fan blade is comprised of a stitched composite cover that is in-turn stitched to a spar and rib-like substructure made of similar carbon fiber materials. A titanium leading edge, root section, and tip closeout member are added to the stitched carbon fibers and placed inside a forming die. Resin is infused, after which the blade assembly, having predetermined design characteristics, is cured by heating the forming die and pressurizing the part internally with a set of inflatable bladders which ultimately yields a rigid fan blade component.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This patent application claims the benefits of Provisional Patent Application No. 60/136,825, filed Jun. 1, 1999. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was not made by an agency of the United States Government or under a contract with an agency of the United States Government. 
    
    
     REFERENCE TO MICROFICHE APPENDIX 
     Not Applicable 
     BACKGROUND OF THE INVENTION 
     This invention relates to the design features of a hollow-core structure, and more particularly to the design features for a stitched carbon fiber reinforced hollow core fan blade for a turbofan engine. 
     In a continuous development cycle to improve turbofan engine operating efficiencies for jet aircraft, engine manufacturers have been designing increasingly higher thrust engines. This new generation of engines, known as “high bypass engines” or “very high bypass engines”, typically operate at significantly higher bypass ratios than their predecessors. To achieve these higher bypass ratios, airflow into the engine must be substantially increased. Generally, this is accomplished by increasing the inlet diameter of the engine. While this effectively increases the operating performance of the engine, it also requires that the first stage engine fan blades operate in a more demanding structural environment, as they increase in length and size. As a result, blade designs that worked in past applications are no longer able to meet these more difficult design parameters. Increases in blade size and mass result in dynamic-induced load conditions that can exceed the strength capability of current high-performance materials, such as titanium alloys or carbon fiber composites. Therefore, design changes in the structural geometry or shape of the blade are necessary to meet these new design requirements. While much work has been done on changing the exterior shape of fan blades by employing a wide-chord design philosophy, very little has been done to change the interior arrangement of the blade. One of the most effective ways to improve the structural performance of a rotating component is to reduce the structural mass of the part. This results in lower rotational inertia forces which reduces the internal loads experienced along the length of the blade. For rotating equipment where the rotational velocity is held constant and material properties are generally fixed, a reduction in mass is the single most effective way of accomplishing significant improvements in structural performance. 
     Reductions in part mass are especially beneficial because they produce a compounding-effect. As the mass is reduced, the magnitude of the internal loads drops, which in turn leads to further reductions in structural mass. A practical limit for this iteration is reached when the part reaches a fully stressed design point for a given set of material properties and load conditions. Once this design point is reached, the structural geometry or shape of the part is optimum. Any further improvements in the overall structural performance can only be achieved by changing or improving the material properties of the as-fabricated part. Thus it is important to note, that its geometric shape and specific material properties primarily influence the structural efficiency of a given component. Once the part geometry is optimized, only improvements in the strength and stiffness properties of the material will result in improvements in the structural performance. The premise of optimizing both the structural geometry and material properties is the overall basis for the stitched-composite hollow-core fan blade design described herein. It combines the superior mechanical properties found in carbon fiber based stitched composite materials with the optimum structural geometry inherent in a hollow-core structural arrangement. Through the combination of these two important design features, the optimum structural design point for a fan blade component is achieved. 
     Current state-of-the-art construction techniques for fan blade fabrication are comprised of hollow-core titanium designs and solid laminate carbon fiber designs. While both methodologies employ a wide-chord design geometry to improve the structural efficiency of the blade element, their respective design philosophies diverge based on their choice of materials. In each case, the design is driven by the limitations of their respective material systems rather than by the objectives of the design problem. For instance, the hollow-core titanium design approach uses the best structural geometry to optimize the internal loading through the structure, but the specific material properties (divided by density) of the titanium alloys are inferior to those of typical high-performance carbon fiber materials. In contrast, the carbon fiber design approach makes use of the high specific material properties of the carbon fibers, but does not take advantage of the hollow-core design approach to optimize the structural loading in the blade body. The optimum design approach would be to combine the hollow-core geometry with the high specific mechanical properties of the carbon fiber materials. The resulting blade design would provide the most structurally efficient fan blade possible for a given structural volume and set of design parameters. 
     While a hollow core carbon fiber fan blade is highly desirable, there are several limitations that have previously precluded its development. The primary concern for such a component is the limited damage tolerance behavior of laminated carbon fiber materials. As carbon fibers are known to be more brittle than metallic materials, meeting the impact damage design requirements for the bird-strike load condition has proved difficult. In order to meet these requirements, existing carbon fiber fan blade designs make two compromises: 1) blade sections are kept solid to maximize the amount of material at a given cross-section, and 2) the rotational velocity of the fan blade assembly is lowered to decrease the impact energy of the bird-strike. Each of these compromises is made because of the limited damage tolerance capability of the carbon fiber material systems and ultimately results in degraded engine performance. The key to removing these impediments is to improve the through-the-thickness mechanical strength of carbon fiber laminates. Prior efforts to accomplish this were focused on using toughened resin systems. While this approach has proven somewhat effective, the attained improvements are still significantly lower than the mechanical strengths realized in the primary fiber directions of the laminate. Since fiber properties are superior to than those of the resin, placing fibers in the z-direction (perpendicular to the lay-up plane) of the laminate offers the highest potential for improving the through-the-thickness mechanical properties of the carbon fiber material system. Indeed, the use of fiber reinforcement in the z-direction is the enabling advancement of the art that permits a hollow-core design approach to be successfully implemented using carbon fiber materials. 
     Another concern is the difficulty of fabricating a hollow core structure that is both damage tolerant and cost-effective using carbon fiber material systems. While several hollow core design approaches have been proposed in prior art, none of them is capable of satisfying the damage tolerance requirements necessary to meet the bird-strike load case because they rely solely on the resin interface to provide the through-the-thickness mechanical strength for the part. Since resin properties are not capable of providing an adequate level of damage resistance, further development of those particular fan blade design approaches has not resulted. Without significantly advancing the state-of-the-art in damage arrestment and residual strength, further development of carbon fiber hollow core fan blade concepts is unlikely. 
     To advance the state-of-the-art regarding damage tolerance, the stitched composite fan blade design described herein employs three key design features: 1) it uses through-the-thickness stitching to improve the z-direction mechanical strength, 2) it has a multi-element substructure design to provide structural redundancy, and 3) it has a continuous cover skin load path around the root section at each spar location. All of these advancements in the art were pursued to achieve the requisite level of damage tolerance necessary to make a composite hollow core fan blade operationally feasible. The out-of-plane mechanical properties are enhanced by the stitching and are no longer dictated by the inferior resin properties at the interface. The multi-element substructure arrangement not only reduces blade mass and internal loads it provides structural redundancy, whereby a single spar element failure does not preclude failure of the entire blade body. The continuous cover skin load path around the root section at each spar location maintains the load continuity in the highly loaded covers and reduces the structural fragility normally encountered at the root joint. All of these design features are extremely advantageous for the bird-strike load case because they work together to improve the damage characteristics of the materials as well as the overall load distribution within the structure. This will result in higher residual strengths and more resilient failure modes, as the redundant substructure elements redistribute internal loads to prevent the blade body from breaking-away during an impact. The synergy provided by these design features advance the art to the point where a hollow-core blade geometry made using carbon fiber composites is now feasible. 
     It is known that in prior art to design and manufacture hollow core titanium fan blades for large gas turbine engines by machining matching cavities in titanium plates, then diffusion bonding the halves together inside a die cavity. It is also known that laminating pre-plied carbon fibers together can produce a solid carbon fiber fan blade design. It is also know that hollow-core blade designs of pre-plied materials have also been proposed. However, the titanium-based designs and manufacturing methods do not utilize the high specific properties of composite materials, while the composite material-based approaches do not meet the rigorous damage tolerant requirements for the bird-strike load case. In either case, the resulting blades tend to be heavier than a fully optimized design would be. What is needed, therefore, is a reinforced composite hollow core fan blade design that can be cost-effectively manufactured and which also meets the more stringent damage tolerance requirements imposed by the design conditions of the high bypass ratio engine. 
     BRIEF SUMMARY OF THE INVENTION 
     This invention solves the problem outlined above by combining the efficient load-carrying arrangement of a hollow-core multi-element substructure with the superior mechanical properties of stitched carbon fiber materials to produce a fan blade that is both highly damage tolerant and relatively inexpensive to manufacture. The blade is comprised of a multi-element substructure consisting of vertical spars and horizontal ribs that are enclosed within a cover skin, all of which is stitched together to form a dry preform assembly. Three titanium detail parts are added to the preform assembly to closeout the edges. The entire part assembly is infused with resin and cured to produce a rigid part. There are three important features that enable this design to be extremely durable and damage tolerant: 1the carbon fiber stacks that comprise the cover skin and substructure are stitched together to enhance the through-the-thickness mechanical properties, 2the substructure is a multi-element design to provide structural redundancy, and 3the cover skins are continuous at each spar location to provide improved load transfer around the root fitting joint. 
     Another important advantage of the invention is the hollow-core substructure design. It not only enhances the loading aspects of the part, it also permits the part to be manufactured using an internal pressure apparatus with a removable bladder system. The advantages of this are, smooth and uniform inner surfaces, as well as a significant reduction in part cost because it makes out-of-autoclave processing possible. The detail design nature of the substructure passage network is also a very important design features because it ultimately enables the fabrication of this invention. 
    
    
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWING 
     FIG. 1 is a perspective view showing the overall blade assembly with the lower portion of the suction side cover removed; 
     FIG. 2 is a cross-sectional view along lines  2 — 2  of FIG. 1, showing stitching through the pressure side cover; 
     FIG. 3 is a cross-sectional view along lines  3 — 3  of FIG. 1, showing stitching through the cover and substructure layers; 
     FIG. 4 is an exploded perspective view showing the individual details and regions that comprise the blade assembly; 
     FIG. 5 is a perspective view of the overall blade assembly showing the location and orientation of the cross-sectional views; 
     FIG. 6 is a cross-sectional view along lines  6 — 6  of FIG. 5, showing a typical airfoil section through the hollow core regions; 
     FIG. 7 is a cross-sectional view along lines  7 — 7  of FIG. 5, showing a typical cover and spar interface; 
     FIG. 8 is a cross-sectional view along lines  8 — 8  of FIG. 5, showing the leading edge detail and forward spar interface; 
     FIG. 9 is a cross-sectional view along lines  9 — 9  of FIG. 5, showing the tip closeout detail and tip rib interface; 
     FIG. 10 is a cross-sectional view along lines  10 — 10  of FIG. 5, showing the cover and spar runout at the root fitting; 
     FIG. 11 is a cross-sectional view along lines  11 — 11  of FIG. 5, showing the access hole through the root fitting and root rib; 
     FIG. 12 is a cross-sectional view along lines  12 — 12  of FIG. 5, showing the access hole through the mid rib; 
     FIG. 13 is a diagrammatic view showing the internal cavities of the blade assembly inside the forming die; 
     FIG. 14 is a cross-sectional view showing the corner of the root fitting and forming die; 
     FIG. 15 is a cross-sectional view similar to FIG. 3, showing an alternative embodiment; 
     FIG. 16 is a cross-sectional view similar to FIG. 8, showing an alternative embodiment; 
     FIG. 17 is a cross-sectional view similar to FIG. 10, showing an alternative embodiment; 
     FIG. 18 is a cross-sectional view similar to FIG. 10, showing an alternative embodiment; 
     FIG. 19 is a cross-sectional view similar to FIG. 6, showing an alternative embodiment. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention provides a design for a stitched carbon fiber fan blade component for a high bypass turbofan engine that is both damage tolerant and cost-effective to manufacture. It relies on a hollow-core geometry to efficiently distribute internal loads, through-the-thickness stitching to provide superior out-of-plane mechanical strength, and a continuous cover skin at each spar location to maintain the load path around the root fitting. 
     A perspective view of the overall blade assembly  10  is shown in FIG. 1 with the lower portion of the suction side cover removed to expose the hollow areas of the substructure  12 . The blade assembly  10  is comprised of a carbon fiber cover  11 , a carbon fiber substructure  12 , a titanium leading edge detail  13 , a titanium tip closeout detail  14 , and a titanium root fitting  15  at the base. The carbon fiber stack materials are reinforced with through-the-thickness penetration thread  23  everywhere except along the root fitting  15 . Each side of the cover  11  is individually stitched together in the regions encompassed by the hollow areas. FIG. 2 shows how the individual skin layers are stitched with a vertical penetration thread  23  of either carbon or aramid fibers and then restrained from pulling through the backside with a locking thread  24  of similar material. Although the figure only shows two discrete layers of material, the cover  11  could be comprised of many layers stitched together or just one layer which would not be stitched; or any combination of layers thereof within a single part. In regions where the cover  11  and substructure  12  overlap to create a solid thickness, through-the-thickness stitching with a penetration thread  23  and locking thread  24  would be used to reinforce the attachment of the cover  11  to the substructure  12  as shown in FIG.  3 . 
     Once the carbon fiber cover  11  and substructure  12  is stitched together to form a dry fiber preform, the three titanium details are inserted into the preform in the relative positions shown in FIG. 4 to create the blade assembly  10 . In this figure, the cover  11  is shown in a position split about the root fitting  15  mid-plane as well as not stitched to the substructure  12 . This was done to better show the alignment of the internal parts. In reality once stitched, the cover  11  forms a U-shaped geometry that encloses all of the other detail parts. The titanium leading edge detail  13  butts up against the edge of the substructure  12  and is partially overlapped on both sides by the cover  11 . In a similar manner, the tip closeout detail  14  is inserted at the top of the preform assembly. At the base, the root fitting  15  is slipped into the gap between the substructure  12  runout and the loop of the cover  11  as it transitions from the suction to pressure side of the blade. The combination of these parts creates a structural configuration with a multi-spar and rib arrangement for structural redundancy. The substructure is comprised of multiple regions identified as a forward spar  16 , mid spar  17 , aft spar  18 , tip rib  19 , mid rib  20 , and a root rib  21 . 
     Once the titanium parts are loosely inserted into the fiber preform, the entire assembly is placed inside a die and infused with resin. As the part is heated and infused, a set of internal bladders is expanded to push out excess resin and provide the uniform internal pressure needed to cure the part. After adequate temperature and pressure are applied, a cured part of the configuration shown in FIG. 5 is achieved. In this figure, the lower portion of the suction side cover  11  is cut away to show the internal features of the substructure. The detail feature cross-sectional drawings for the assembly, which are referenced in FIG. 5, are shown in FIGS. 6 through 11 and  14 , and described in the following text. 
     A typical airfoil cross-sectional view taken through the hollow regions  22  is shown in FIG.  6 . Here, individual layers of material fabric are laid upon one another to create three solid spar regions, forward spar  16 , mid spar  17 , and aft spar  18 . The individual layers that are used to create the substructure  12  come from single plies of warp-knit material with cutouts in them corresponding to the location of the hollow regions  22  in the substructure. Each layer of the substructure  12  is dropped off to create a variation in height across the airfoil section as it approaches the leading or trailing edge. A nominal step is achieved at each stack drop-off. The trailing edge runout at the aft spar  18  is a solid stitched stack thickness created as the individual layers of the substructure  12  drop off and until only the layers of the cover  11  are remaining. The cover  11  is terminated at three edges and wraps around the root section in a continuous fashion at the fourth edge to render the outer surface of the blade assembly  10 . The hollow regions  22  between the spar and rib elements are maintained as the cutout regions in the individual layers coleus upon one another. The interface between the cover  11  and substructure  12  is stitched with penetration thread  23 , as are the cover  11  areas that encompass the hollow regions  22 . 
     A cross-sectional view of the mid spar  17  is shown in FIG.  7 . The collection of fabric layers that comprise the substructure  12  are built-up to create a solid thickness which is bounded on both surfaces by the cover  11 . The individual layers are held together by penetration thread  23  and resin. The spar solid elements run from the root fitting  15  to the tip closeout detail  14 . Along the sides, as the spar transitions from a solid to a hollow section, the individual material layers are dropped off side-to-side in a staggered fashion. This is done to create a substructure fillet  26  which is typical around the periphery of the hollow regions. The frayed edges of the trimmed fabric layers will mix with resin to form a gradual transition at this location. This results in a natural radius for the inflatable bladder to form up against. This feature is important because it prevents the bladder from bridging across the drop-off steps and failing during the cure cycle. 
     The forward edge of the blade assembly  10  is closed out with a titanium leading edge detail  13  that is cocured between the suction and pressure sides of the cover  11  as illustrated in the cross-sectional view of FIG.  8 . Here, the edges of the forward spar  16  terminate at the back edge of the leading edge detail  16 . Penetration thread  23  stitching for the cover  11  continues to within 0.25 inches of the stack edge. Penetration thread  23  stitching through the cover  11  and substructure  12  continues to within .25 inches of the substructure  12  edges. The stitching in this region is important for improving the out-of-plane strength at the forward spar  16  region for the bird-strike load case. The aft edges of the forward spar  16  stacks are cut in the side-to-side staggered fashion to form the substructure fillet  26  typically found around the hollow regions  22 . 
     The tip closeout design of the blade assembly  10  is developed in a manner similar to the leading edge closeout. As illustrated in FIG. 9, the tip rib  19  stacks are terminated at the edge of the tip closeout detail  14 , while the cover  11  continues slightly farther up until it sandwiches a portion of the titanium tip closeout detail  14 . The joint between the carbon fiber materials and the titanium details is a cocured resin interface. The portion of the tip closeout detail  14  that extends beyond the cover  11  is bare titanium. This provides a concentrated mass that can be machined as necessary to balance the rotational inertia of the blade assembly  10 . It also provides a rub surface for final sizing of the blade length during operational use. 
     The transfer of load from the carbon fiber material to the titanium root fitting  15  is a critical element of the blade design. Eventually all of the load carried by the carbon fibers must be transmitted through the root fitting  15 . To maximize load transfer, the substructure  12  is trimmed in a staggered fashion and mated to machined steps in the root fitting  15  to create the stepped-lap joint shown in the FIG. 10 cross-sectional view. The cover  11  is not spliced, but continues around the root fitting  15  at each spar location to provide a continuous load path from the pressure to the suction sides of the blade assembly  10 . The uninterrupted cover  11  load path significantly improves the joint strength at the root, which is typically a weakness in composite fan blade designs. The interfaces between the carbon fiber materials and titanium details are cocured. 
     The relationship between the blade assembly  10  and the fabrication apparatus is important because it dictates many of the design features that are incorporated into the blade design. To accommodate the internal bladder apparatus, a passage way is needed into the hollow regions  22  of the blade assembly  10 . Once the cover  11  is stitched to the substructure  12  access can only be gained through an access hole  37  through the root fitting  15  as illustrated in the FIG. 11 cross-sectional view. Each access hole  37  is drilled in the root fitting  15  between the attachment lugs. Access through the root rib  21  is provided by locally trimming slots in the inner most substructure  12  layers; in line with the root fitting  15  access hole  37 . Small access holes are also required through the substructure  12  at the mid rib  20  location. A typical rib access hole  27  is illustrated in the FIG. 12 cross-sectional view. The hole is formed by locally trimming slots in the inner most layers of the substructure  12  in a perpendicular direction to the mid rib  20 . It is through these access holes that the forming bladder apparatus must be inserted and later removed. 
     The positional relationship between the blade assembly  10  and the forming apparatus is shown diagrammatically in FIG.  13 . The outline and internal cavities of the blade assembly  10  are shown inside an approximate cross-section of the forming die  33 . Representative locations for the resin injection ports  28  and the resin exit port  29  are shown. The part is surrounded by a resin seal  32  to prevent leakage during resin infusion. The internal forming apparatus consists of the inflatable bladder  31  and the bladder support tube  30 . These elements are shown positioned inside the hollow regions  22  prior to inflation. As the forming die  33  is heated and resin is infused, the inflatable bladder  31  is expanded to fill the hollow regions  22  and to provide the required internal pressure necessary to cure the part. After the blade assembly  10  is cured, the bladder apparatus is removed through the rib and root fitting access holes  27  and  37 . 
     Another important relationship between the forming die  33  and the root fitting  15  is the locating feature used to properly position the root fitting  15  inside the forming die  33  and to prevent resin from entering the root fitting attachment hole  36  during the resin infusion process. While the primary purpose of the root fitting attachment hole  36  is to provide a means of fastening the blade assembly  10  onto the fan rotor hub, it also can be used in conjunction with the locator hole  38  to index the root fitting  15  inside the forming die  33 . The FIG. 14 cross-sectional view depicts one comer of the root fitting  15  and forming die  33  interface. For this design feature, the left and right sides of the forming die  33  at the root section would have matching holes drilled to accommodate two large locating pins  34 . Then, one side of the forming die  33  would also be drilled to accommodate the small locating pin  39  shown in the figure. 
     Now referring to FIG. 15, which shows an alternative stitching method to the preferred embodiment shown in FIG.  3 . Here, the outer most material layer of the cover  11  is not stitched with the penetration thread  23 . Although this would reduce the damage resistance capability of the blade assembly  10 , it could improve the surface roughness and enhance the aerodynamic performance. All other aspects of the design would be identical to FIGS. 1 through 14. 
     Now referring to FIG. 16, which shows an alternative leading edge joint to the preferred embodiment shown in FIG.  8 . Here, a stepped-lap joint is used instead of a simple lap joint. This approach could be used to improve the shear load transfer between the leading edge detail  25  and the cover  11 . All other aspects of the design would be identical to FIGS. 1 through 14. 
     Now referring to FIG. 17, which shows an alternative root fitting detail  40  runout design to the preferred embodiment shown in FIG.  10 . Here, a simple butt splice is used instead of a stepped-lap joint at the root rib runout. This approach could be used to reduce the fabrication cost of the root fitting. All other aspects of the design would be identical to FIGS. 1 through 14. 
     Now referring to FIG. 18, which shows an alternative for the cover  11  design to the preferred embodiment shown in FIG.  10 . Here, the cover  11  is not continuous at the root fitting  15 , but rather split  41  along the centerline of the parting-plane between the pressure and suction sides of the blade. Although this approach would significantly reduce the load capability around the root fitting  15 , it could simplify the manufacturing complexity of the cover  11 . All other aspects of the design would be identical to FIGS. 1 through 14. 
     Now referring to FIG. 19, which shows an alternative substructure arrangement to the preferred embodiment shown in FIG.  6 . Here, the hollow regions  22  are not created in the substructure  12  and the substructure element of the blade assembly is kept solid. Although this would reduce the structural efficiency of the blade assembly  10 , it could result in lower manufacturing costs. All other aspects of the design would be identical to FIGS. 1 through 14 except that the hollow regions  22  would not be present. The forming pressure required to cure the part would be provided by applying mechanical pressure to the forming die  33  to squeeze the stack layers together. 
     Although exemplary embodiments of the invention have been shown and described, many changes, modifications, and substitutions may be made by one having ordinary skill in the art without departing from the spirit and scope of the invention. Therefore, the scope of the invention is to be limited only in accordance with the following claims. 
     Part Number Listing: 
       10 . Blade Assembly 
       11 . Cover 
       12 . Substructure 
       13 . Leading Edge Detail 
       14 . Tip Closeout Detail 
       15 . Root Fitting 
       16 . Forward Spar 
       17 . Mid Spar 
       18 . Aft Spar 
       19 . Tip Rib 
       20 . Mid Rib 
       21 . Root Rib 
       22 . Hollow Region 
       23 . Stitching 
       24 . Lock Stitch 
       25 . Alternate Leading Edge Design 
       26 . Substructure Fillet 
       27 . Rib Access Hole 
       28 . Resin Port In 
       29 . Resin Port Out 
       30 . Bladder Support Tubes 
       31 . Inflatable Bladder 
       32 . Resin Seal 
       33 . Forming Tool 
       34 . Large Locating Pin 
       35 . Resin Voids 
       36 . Root Fitting Attachment Hole 
       37 . Main Hole Through Root Fitting 
       38 . Locator Hole through Fitting 
       39 . Small Locator Pin 
       40 . Alternate root fitting 
       41 . Split of cover stacks 
       42 . Solid blade configuration