Abstract:
One exemplary aspect of this disclosure relates to an assembly for a gas turbine engine having an engine axis. The assembly includes a case including an integrally formed projection configured to extend transverse to the engine axis. The assembly further includes an engine component including a flange configured for contact with the projection to limit motion of the component along the engine axis.

Description:
STATEMENT REGARDING GOVERNMENT SUPPORT 
     This invention was made with government support under Contract No. FA8650-09-D-DO0021 awarded by the United States Air Force. The government has certain rights in this invention. 
    
    
     BACKGROUND 
     Gas turbine engines typically include a compressor section, a combustor section, and a turbine section. During operation, air is pressurized in the compressor section, and mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Both the compressor and turbine sections include alternating arrays of rotating blades and stationary vanes that extend into a core airflow path of the gas turbine engine. During a surge condition, wherein fluid in the core airflow path flows opposite the intended direction, it is possible for the stationary vanes to move axially and cause damage to adjacent components. One example system for limiting vane movement is in U.S. Pat. No. 7,854,586, assigned to United Technologies Corporation and hereby incorporated by reference in its entirety. In the &#39;586 patent, an arm 94, which is formed separately from an engine case 28, is attached to the case 28 to prevent undesired vane movement. 
     SUMMARY 
     One exemplary aspect of this disclosure relates to an assembly for a gas turbine engine having an engine axis. The assembly includes a case including an integrally formed projection configured to extend transverse to the engine axis. The assembly further includes an engine component including a flange configured for contact with the projection to limit motion of the component along the engine axis. 
     In a further non-limiting embodiment of the foregoing assembly, the case and the projection are monolithically formed. 
     In a further non-limiting embodiment of the foregoing assembly, the engine component includes at least one vane. 
     In a further non-limiting embodiment of the foregoing assembly, the at least one vane includes a plurality of variable area vanes. 
     In a further non-limiting embodiment of the foregoing assembly, each vane includes a respective platform having a respective flange, and each of the flanges contacts the projection. 
     In a further non-limiting embodiment of the foregoing assembly, the assembly further includes a plurality of rotor blades downstream of the vanes. 
     In a further non-limiting embodiment of the foregoing assembly, an aft surface of the flange is configured to contact a fore surface of the projection. 
     In a further non-limiting embodiment of the foregoing assembly, the aft surface of the flange and the fore surface of the projection are configured for contacting one another along an inclined interface relative to the engine axis. 
     In a further non-limiting embodiment of the foregoing assembly, the inclined interface is inclined at an acute angle relative to the engine axis. 
     In a further non-limiting embodiment of the foregoing assembly, the case is configured to be mounted adjacent a low pressure compressor and a high pressure compressor of the engine. 
     Another exemplary aspect of this disclosure relates to a case for a gas turbine engine having an engine axis. The case includes a stop monolithically formed with the case and configured to circumferentially extend about the engine axis with a fore surface of the stop oriented at an angle relative to the engine axis. 
     In a further non-limiting embodiment of the foregoing case, wherein the stop extends from a main body of the case to a free end. 
     In a further non-limiting embodiment of the foregoing case, the stop is configured to contact a flange of an engine component to limit axial movement thereof. 
     In a further non-limiting embodiment of the foregoing case, the stop is oriented at an acute angle relative to the engine axis. 
     Yet another exemplary aspect of this disclosure relates to a component for a gas turbine engine having an engine axis. The component includes an inner platform including a flange. The flange projects from a radially inner surface of the inner platform, and the flange is configured to be mounted in the engine with an aft surface thereof oriented at an angle relative to the engine axis. 
     In a further non-limiting embodiment of the foregoing component, the component is a variable area vane. 
     In a further non-limiting embodiment of the foregoing component, the vane includes an airfoil section, a root section, and a bushing adjacent the root section. The bushing is configured to radially retain the root section while allowing rotation of the root section. 
     In a further non-limiting embodiment of the foregoing component, the root section is configured for rotation about an axis normal to the engine axis. 
     In a further non-limiting embodiment of the foregoing component, the flange is configured to contact a retainer of a case. 
     The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The drawings can be briefly described as follows: 
         FIG. 1  schematically illustrates an example gas turbine engine. 
         FIG. 2  is a partial, cross-sectional view of a section of an example gas turbine engine. 
         FIG. 3  is a close-up view of the encircled area in  FIG. 2 . 
     
    
    
     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 flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core airflow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  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 C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
       FIG. 2  is a partial, schematic view of a section  60  of the engine  20 . In this example, the section  60  includes an array of vanes  62  and an array of blades  64  downstream of the vanes  62 . The vanes  62  each include an airfoil section  66  projecting into the core airflow path C, and a root section  68 . The vanes  62  in this example are variable area vanes. That is, the vanes  62  are rotatable about an axis X, which extends parallel to a radial direction R (which is normal to the engine central longitudinal axis A) to vary the effective area of the core airflow path C. The root section  68  is rotatable relative to a bushing  70 , which also radially retains the root section  68 . The bushing  70  is provided within an inner platform  72 . While variable area vanes are shown, this disclosure extends to other types of vanes. 
     The blades  64  each include an airfoil section  74  projecting into the core airflow path C from an inner platform  76 . The inner platform  76  is connected to a disk  78 , which is configured to rotate about the engine central longitudinal axis A. 
     In order to prevent unwanted axial movement of the vanes  62 , a case  80  of the engine  20  includes a retainer  82 . In this example, the retainer  82  includes a projection  84  which functions as a retainer or stop as further discussed below. The case  80  is an intermediate case in this example, and is located between the low pressure compressor  44  and the high pressure compressor  52 . This disclosure is not limited to intermediate cases, however. 
     In accordance with various embodiments, the case  80  is integrally or monolithically formed with the retainer  82 . That is, the case  80  and the projection  84  can be formed as a single, monolithic structure, without mechanical joints or seams. In one example, the case  80  and the projection  84  are formed together as part of the same casting process. While  FIG. 2  only illustrates the case  80  in cross section, the case  80  and the projection  84  circumferentially extend about the engine central longitudinal axis A. In one example, the case  80  and the projection  84  are annular, and extend around the entirety of the engine central longitudinal axis A. It is also possible to scallop adjacent sections of the case  80 , which may provide a weight reduction. That is, a single projection  84  may contact more than one vane  62 . 
     The projection  84  contacts a flange  86  projecting from a radially inner surface  88  of the inner platforms  72  to limit vane movement. The projection  84 , in this example, extends generally in an aft direction (to the right in  FIG. 2 ) from a main body  83  of the case  80  to a free end  85 . Further, as illustrated in  FIG. 3 , the projection  84  includes a fore surface  90  and an aft surface  92 . The fore surface  90  is inclined at an angle A 1  relative to the engine central longitudinal axis A. 
     The flange  86  includes a fore surface  94  and an aft surface  96 , which is also inclined at the angle A 1 . The angle A 1 , in one example, is non-parallel with the engine central longitudinal axis A. That is, A 1  is greater than 0° and less than 90°. In one example, the angle A 1  is approximately 60°. 
     As illustrated, the fore surface  90  of the projection  84  is in direct contact with the aft surface  96  of the flange  86 . The surfaces  90  and  96  provide a bearing surface between the flange  86  and the projection  84 , which prevents the vanes  62  from moving in an aft direction (to the right in  FIG. 2 ) during a surge condition, for example, and thus prevents damage to adjacent engine components, such as the blades  64 . 
     In a surge condition, the flow of fluid within the core airflow path C reverses. A reversal of flow is illustrated in phantom at S in  FIG. 2 . In a surge condition, the vanes  62  may deflect in a fore direction (to the left in  FIG. 2 ), such that a fore face  98  of the inner platform  72  contacts an aft face  100  of the case  80 . Absent retaining projection  84 , such contact between faces  98 ,  100  may cause the vanes  62  to subsequently move in an axially aft direction toward the rotor blades  64 , which may cause damage to the rotor blades  64  and other engine components. 
     The retainer  82  prevents unwanted axial movement of the vane  62 , and thus prevents the vane  62  from damaging the engine during a surge condition. Further, because the retaining projection  84  is monolithically formed with the case  80 , ease of assembly is increased, and weight is reduced (due to the elimination of additional parts, scalloping adjacent case sections, etc.). The inclined surfaces  90 ,  96  also increase the ease of aligning the flange  86  and projection  84  during assembly, during which the vane  62  may be loaded from a radial outer location. 
     It should be understood that terms such as “fore,” “aft,” “axial,” and “radial,” are used above with reference to the normal operational attitude of the engine  20 . Further, these terms have been used herein for purposes of explanation, and should not be considered otherwise limiting. Terms such as “approximately” are not intended to be boundaryless terms, and should be interpreted consistent with the way one skilled in the art would interpret the term. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     One of ordinary skill in this art would understand that the above-described embodiments are exemplary and non-limiting. That is, modifications of this disclosure would come within the scope of the claims. Accordingly, the following claims should be studied to determine their true scope and content.