Patent Publication Number: US-8967961-B2

Title: Ceramic matrix composite airfoil structure with trailing edge support for a gas turbine engine

Description:
BACKGROUND 
     The present disclosure relates to a gas turbine engine, and more particularly to Ceramic Matrix Composite (CMC) components therefor. 
     The turbine section of a gas turbine engine includes a multiple of airfoils which operate at elevated temperatures in a strenuous, oxidizing type of gas flow environment and are typically manufactured of high temperature superalloys. CMC materials provide higher temperature capability than metal alloys and a high strength to weight ratio. CMC materials, however, may require particular manufacturing approaches as the fiber orientation primarily determines the strength capability. 
     CMC airfoil designs have struggled to create a thin trailing edge which is strong enough to avoid splitting due to thermal-mechanical loads. A natural geometric stress concentration occurs where the pressure and suction side airfoil walls come together into a sharp trailing edge feature. The stress concentration may be difficult to overcome with 2D, 2.5D and 3D fiber architectures. 
     SUMMARY 
     An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes a pressure side formed of at least one Ceramic Matrix Composite ply, a suction side formed of at least one Ceramic Matrix Composite ply and an aft trailing edge support between the pressure side and the suction side. 
     An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes a pressure side formed of at least one Ceramic Matrix Composite ply, a suction side formed of at least one Ceramic Matrix Composite ply and an aft trailing edge support between the pressure side and the suction side and a forward trailing edge support between said pressure side and said suction side. 
     A method of assembling a Ceramic Matrix Composite airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure including venting an airfoil aft of an aft trailing edge support between a pressure side and a suction side. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of a gas turbine engine; 
         FIG. 2  is an enlarged sectional view of a Low Pressure Turbine section of the gas turbine engine; 
         FIG. 3  is an enlarged perspective view of an example rotor disk of the Low Pressure Turbine section; 
         FIG. 4  is an enlarged perspective view of an example stator vane structure of the Low Pressure Turbine section; 
         FIG. 5  is a perspective view of a CMC vane structure; 
         FIG. 6  is a sectional view of the stator vane structure of  FIG. 5 ; 
         FIG. 7  is a sectional view of a trailing edge of the stator vane structure; 
         FIG. 8  is a sectional view of a trailing edge of another disclosed non-limiting embodiment of the stator vane structure; 
         FIG. 9  is a sectional view of the trailing edge of another disclosed non-limiting embodiment of the stator vane structure illustrating a split trailing edge; and 
         FIG. 10  is a sectional view of a trailing edge of another disclosed non-limiting embodiment of the stator vane structure illustrating a vent. 
         FIG. 11  is a sectional view of a trailing edge of another disclosed non-limiting embodiment of the stator vane structure; and 
         FIG. 12  is a sectional view of a trailing edge of another disclosed non-limiting embodiment of the stator vane structure. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path. The turbines  54 ,  56  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     With reference to  FIG. 2 , the low pressure turbine  46  generally includes a low pressure turbine case  60  with a multiple of low pressure turbine stages. The stages include a multiple of rotor structures  62 A,  62 B,  62 C interspersed with vane structures  64 A,  64 B. Each of the rotor structures  62 A,  62 B,  62 C and each of the vane structure  64 A,  64 B may include airfoils  66  manufactured of a ceramic matrix composite (CMC) material ( FIGS. 3 and 4 ). It should be understood that examples of CMC material for componentry discussed herein may include, but are not limited to, for example, the CMC material S200 manufactured by COI Ceramic and a Silicon Carbide Fiber in a Silicon Carbide matrix (SiC/SiC). Although depicted as a low pressure turbine in the disclosed embodiment, it should also be understood that the concepts described herein are not limited to use with low pressure turbines as the teachings may be applied to other sections such as high pressure turbines, high pressure compressors, low pressure compressors, the mid turbine frame  57 , as well as intermediate pressure turbines and intermediate pressure compressors of a three-spool architecture gas turbine engine. 
     With reference to  FIG. 5 , one CMC airfoil  66  “singlet” is illustrated, however, it should be understood that other vane structures with, for example a ring-strut-ring full hoop structure will also benefit herefrom. Although a somewhat generic CMC airfoil  66  will be described in detail hereafter, it should be understood that various rotary airfoils or blades and static airfoils or vanes may be particularly amenable to the fabrication described herein. 
     The CMC airfoil  66  generally includes an airfoil portion  68  defined between a leading edge  70  and a trailing edge  72 . It should be understood that the airfoil portion  68  may include various twist distributions. The airfoil portion  68  includes a generally concave shaped side which forms a pressure side  82  and a generally convex shaped side which forms a suction side  84 . It should be further appreciated that various structures with a trailing edge will also benefit herefrom. 
     Each CMC airfoil  66  may include a fillet section  86  to provide a transition between the airfoil portion  68  and a platform segment  88 . The platform segment  88  may include unidirectional plies which are aligned tows with or without weave, as well as additional or alternative fabric plies to obtain a thicker platform segment if so required. In the disclosed non-limiting embodiment, either or both of the platform segments segment  88  may be of a circumferential complementary geometry such as a chevron-shape to provide a complementary abutting edge engagement for each adjacent platform segment to define the inner and outer core gas path. That is, the CMC airfoils  66  are assembled in an adjacent complementary manner with the respectively adjacent platform segments  88  to form a cascade of airfoils. 
     Pressure distributions to which the CMC airfoil  66  is subjected is generally of a higher pressure and lower velocity along the pressure side  82  and a relatively lower pressure and higher velocity along the suction side  84 . That is, there is a differential pressure across the chord of the CMC airfoil  66 . This differential is also within the significant temperature environment of the turbine section  28  over which the core flow expands downstream of the combustor section  26 . 
     With reference to  FIG. 6 , the pressure side  82  and the suction side  84  may be formed from a respective first and second multiple of CMC plies  90 ,  92  which meet and may be bonded together along at the trailing edge  72  at an essentially line interface  94  (also shown in  FIG. 7 ). Adjacent to the trailing edge  72  and within the CMC plies  90 ,  92  which define the airfoil portion  68  are located a forward trailing edge support  96  and an aft trailing edge support  98 . As defined herein, “fore” to “aft” is in relation to the gas flow direction past the airfoil  66 , such as the hot gas which flows past the turbine blade or vane in operation. 
     The forward trailing edge support  96  and the aft trailing edge support  98  in the disclosed, non-limiting embodiment are generally “C” shaped in which the open portion of the “C” of the forward trailing edge support  96  faces forward, while the open portion of the “C” of the aft trailing edge support  98  face aft to provide a back-to-back relationship. It should be appreciate that the “C” shape is a general description and that other shapes such as an “O”; “0”; “I” or other shape may also be utilized to provide significant surface area to bond with the CMC plies  90 ,  92 . The forward trailing edge support  96  and the aft trailing edge support  98  may alternatively or additionally be formed as a monolithic ceramic material such as a silicon carbide, silicon nitride or alternatively from a multiple of CMC plies. 
     The forward trailing edge support  96  defines an internal pressure vessel  100  within the CMC airfoil  66  between the CMC plies  90 ,  92  to receive, for example a cooling flow therethrough. In another non-limiting alternate embodiment, the forward trailing edge support  96  is not required as the aft trailing edge support  98 ′ provides sufficient support for the expected internal pressure ( FIG. 8 ). 
     The internal pressure vessel  100  strengthens the CMC airfoil  66  to resist the differential pressure generated between the core flow along the airfoil portion  68  and provides a passage for secondary cooling flow which may be communicated through the airfoil portion  68 . It should be appreciated that other passages may be formed to provide a path for wire harnesses, conduits, or other systems. 
     For an uncooled or lightly cooled airfoil  66 , a potential split S in the trailing edge  72  ( FIG. 9 ) has no significant impact to the purpose of turning the flow. However, for hollow airfoils  66  that transport cooling air, the “C” section architecture prevents the loss of cooling air, because even a trailing edge  72  which has split is isolated from the main body cooling flow within the internal pressure vessel  100 . That is, as the forward trailing edge support  96  faces forward and is bonded to the CMC plies  90 ,  92 , the forward trailing edge support  96  facilitates formation of the pressure vessel  100  for the cooling air as the forward trailing edge support  96  may be pressed outward into the CMC plies  90 ,  92 . This is a relatively stronger architecture than the pressure applied to the back side of the aft trailing edge support  98  in which the pressure may tend toward peeling the aft trailing edge support  98  from the CMC plies  90 ,  92 . 
     The aft trailing edge support  98  may be arranged such that the open ends of the “C” touch each other. The aft trailing edge support  98  facilitates usage of a relatively small number of CMC plies  90 ,  92  at the trailing edge  72 , such as 1-4 plies each, to form a sharp trailing edge  72 . 
     The aft trailing edge support  98  provides a desired bending strength through the appropriate consideration of section thickness and permits the trailing edge  72  to actually split, thus relieving stresses which may naturally occur ( FIG. 9 ). The aft trailing edge support  98  prevents the split in the trailing edge  72  from debonding the CMC plies  90 ,  92 . That is, the relatively higher pressure and lower velocity along the pressure side  82  and the relatively lower pressure and higher velocity along the suction side  84  actually forces the split in the trailing edge  72  together as the aft trailing edge support  96  compartmentalizes the external pressure from the internal pressure forward thereof. The trailing edge  72 , once spilt is equalized in pressure and the CMC plies  90  on the pressure side  82 , are pushed onto the aft trailing edge support  98 . Thus, the presence of the aft trailing edge support  98  allows the force on the pressure side  82  to be resisted, and the split sees a compressive load. 
     In another disclosed non-limiting embodiment, a vent  102  is located through the suction side  84  to selectively balance the internal pressure within the aft trailing edge support  98  with the low external core path pressure on the suction side, which further tends to minimize the internal pressurization, and the initial potential for a split in the trailing edge  72  ( FIG. 10 ). 
     In another disclosed non-limiting embodiment, other shapes such as an “O”; “0” ( FIG. 11 ) aft trailing edge support  98 ′; “I” aft trailing edge support  98 ″ ( FIG. 12 ) or other shape may also be utilized to provide significant surface area to bond with the CMC plies  90 ,  92 . 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.