A gas turbine engine includes a compressor, a combustor and a turbine. The compressor includes a plurality of disks, each with arcuate blades extending from its periphery. The disks rotate rapidly on a shaft, drawing in air, compressing it and moving the highly compressed air downstream toward the combustor. In the combustor, the compressed air is mixed with metered fuel which is burned, generating hot gases. The hot gases flow to the turbine which comprises at least one disk having arcuate blades extending from its periphery. Energy is extracted from the hot gases by the blades, the hot gases striking the blades causing the disks to turn, which in turn rotates the shaft, powering the engine. The remaining gases passing through the turbine generate thrust to propel an aircraft.
The materials used in the turbine section, because of their exposure to high temperatures and because of the rapid rotation, have typically been comprised of high temperature superalloys. However, because of their light weight and high temperature capabilities, ceramic composite materials, such as SiC/SiC ceramic matrix composites, which exhibit favorable characteristics, have been considered for use in the turbine portion of the engine, such as in turbine engine blade applications. One of the drawbacks of this material has been its poor interlaminar properties. The primary cause for low interlaminar strength is the presence of a boron nitride coating, which is typically applied over the fibers to form an interface between the fibers and matrix, thereby increasing fracture toughness by allowing the load to be transferred to the fibers and absorb energy by promoting crack propagation along the fibers, or within the weaker fiber coating. This low interlaminar strength, by improving fracture toughness, reduces the tendency of the material to suddenly fail in a brittle mode. In many of the hot section applications such as combustion liners, HPT vanes, LP blades and shrouds, the thermal gradients and mechanical loads can result in significant local interlaminar stresses. Therefore, it is desirable to enhance the interlaminar strength of ceramic composites in local areas for many of these applications.
A number of techniques have been used in the past to manufacture turbine engine components, such as turbine blades using ceramic matrix composites. However, such turbine components, under normal operating conditions, are not subjected to uniform stress patterns, instead experiencing varying degrees of local stresses at different times and at different locations within the part during normal turbine operation. A turbine blade generally has a dovetail portion, an airfoil portion opposite the dovetail portion and an optional platform located between the dovetail portion and the airfoil portion. In the dovetail portion of turbine blades, relatively higher tensile stress regions are located in the outermost portion of the dovetail section. Ideally, the CMC component should be designed such that the component has a higher tensile strength in the region experiencing the higher tensile stresses. One method of manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated herein by reference and assigned to the assignee of the present invention, relates to the production of silicon carbide matrix composites containing fibrous material that is infiltrated with molten silicon, the process herein referred to as the Silcomp process. The fibers generally have diameters of about 140 micrometers or greater, which prevents the manufacture of intricate, complex shapes, such as turbine blade components, by the Silcomp process.
Another technique of manufacturing CMC turbine blades is the method known as the slurry cast melt infiltration (MI) process. A technical description of such a MI method is described in detail in U.S. Pat. No. 6,280,550 B1, which is assigned to the Assignee of the present invention and which is incorporated herein by reference. In one method of manufacturing using the MI method, CMCs are produced by initially providing plies of balanced two-dimensional (2D) woven cloth comprising silicon carbide (SiC)-containing fibers, having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers running in both directions of the weave. By “silicon carbide-containing fiber” is meant a fiber having a composition that includes silicon carbide, and preferably is substantially only silicon carbide. The fiber may have a silicon carbide core surrounded with carbon, or in the reverse, the fiber may have a carbon core surrounded by, or encapsulated with, silicon carbide. These examples are exemplary of the term “silicon carbide-containing fiber” and are not limited to this specific combination. Other fiber compositions are contemplated, so long as they include silicon carbide.
Prior ceramic matrix composites, such as U.S. Pat. No. 4,642,271, to Rice may be suitable for producing a homogenous composite with favorable toughness characteristics and other inplane properties, but lack the interlaminar properties required for many turbine engine applications. Typical methods of improving interlaminar strength of SiC/SiC composites have utilized through thickness fiber reinforcement. T-forming and Z-pinning are examples of techniques used to introduce load carrying fibers in the through-thickness directions of composites and, thus, enhance interlaminar strength within desired regions. T-forming, as set forth in U.S. Pat. No. 6,103,337 to Moody, is a method by which fibers are inserted directly into a preform so that spacing, depth of penetration, and orientation can be controlled to produce 3-D fiber architectures with improved interlaminar strength. Z-pinning is a technique used to reinforce a composite structure to prevent various layers rigidly connected to one another and from delaminating. These methods, however, require trade-offs in in-plane mechanical properties and result in significant increases in fiber and/or manufacturing costs.
Accordingly, there is a need for a method of producing a composite that possesses regions of favorable in-plane properties and regions of favorable interlaminar properties, thereby overcoming the inadequacies of the prior art.