Abstract:
A case is provided for a gas turbine engine. The case includes a wall defining a through-hole. The case also includes first and second pockets adjacent to, and on opposite sides, of the through hole. A method of reducing stress in a case of a gas turbine engine is also provided that includes reducing stress about a through-hole by providing a concavity on each side of the through-hole.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/877,696 filed Sep. 13, 2013, which is hereby incorporated herein by reference in its entirety. 
     
    
     BACKGROUND 
       [0002]    The present disclosure relates to a gas turbine engine and, more particularly, to a case therefor. 
         [0003]    Gas turbine engines, such as those that power modem commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to bum a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. 
         [0004]    The combustor section typically includes an outer shell lined with heat shields to form a combustion chamber. The combustion chamber is surrounded by a diffuser case formed of an inner and outer case, where the inner case defines a pre-diffuser and the outer case serves as structural containment. Together the cases form the flowpath and necessary volume to mitigate unrecoverable compressor surge. Although effective, the diffuser case includes multiple through-holes which may form undesirable stress concentrations. 
       SUMMARY 
       [0005]    A case for a gas turbine engine, according to one disclosed non-limiting embodiment of the present disclosure, includes a wall defining a through-hole. The case also includes first and second pockets adjacent to, and on opposite sides, of the through hole. 
         [0006]    In a further embodiment of the present disclosure, at least one of the first or the second pockets each includes a circular periphery. 
         [0007]    In a further embodiment of the present disclosure, at least one of the first or the second pockets each includes a race track shaped periphery. 
         [0008]    In a further embodiment of any of the foregoing embodiments of the present disclosure, at least one of the first or the second pockets each includes a rectilinear shaped periphery. 
         [0009]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the first or the second pockets circumferentially flank the through-hole. 
         [0010]    In a further embodiment of any of the foregoing embodiments of the present disclosure, at least one of the first or second pockets each extends a depth from the inner surface of between 10%-50% of a thickness of the wall. 
         [0011]    In a further embodiment of any of the foregoing embodiments of the present disclosure, at least one of the first or the second pocket is 100%-500% a diameter of the through-hole. 
         [0012]    In a further embodiment of any of the foregoing embodiments of the present disclosure, an outer periphery of at least one of the first or second pockets is each circumferentially spaced a distance from an outer diameter of the through-hole along a hoop line. The distance is between 10%-100% the diameter of the through-hole. 
         [0013]    In a further embodiment of any of the foregoing embodiments of the present disclosure, at least one pocket extends for a depth from the inner surface between 10%-50% a thickness of the wall and has a diameter of 100%-500% of a diameter of the through-hole. The pocket has an outer edge circumferentially spaced a distance from a surface defining the through-hole along a hoop line. The distance is between 10%-400% of the diameter of the through-hole. 
         [0014]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the wall is an outer wall of a diffuser case. 
         [0015]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the through-hole is located through a boss. 
         [0016]    A method of reducing stress in a case of a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes reducing stress about a through-hole by providing a concavity on each side of the through-hole. 
         [0017]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes defining each pocket in an inner surface of a wall. 
         [0018]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes defining each pocket for a depth from an inner surface of the case between 10%-50% of a thickness of the wall. 
         [0019]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes defining each pocket with an outer edge having a diameter of between 100%-500% of a diameter of the through-hole. 
         [0020]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes circumferentially spacing each pocket a distance along a hoop line from between 10%-400% of a diameter of the through-hole. 
         [0021]    The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]    Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
           [0023]      FIG. 1  is a schematic cross-section of an example gas turbine engine architecture; 
           [0024]      FIG. 2  is a schematic cross-section of another example gas turbine engine architecture; 
           [0025]      FIG. 3  is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment that may be used with the example gas turbine engine architectures shown in  FIGS. 1 and 2 ; 
           [0026]      FIG. 4  is a schematic view of a gas turbine engine case assembly; 
           [0027]      FIG. 5  is an expanded schematic view of a case; 
           [0028]      FIG. 6  is an expanded outer perspective view of a through-hole in the case; 
           [0029]      FIG. 7  is an expanded inner perspective view of a through-hole in the engine case and pockets formed therein; 
           [0030]      FIG. 8  is an expanded inner perspective view of the through-holes with corresponding pockets according to one disclosed non-limiting embodiment; 
           [0031]      FIG. 9  is an expanded inner perspective view of the through-holes and pockets according to another disclosed non-limiting embodiment; 
           [0032]      FIG. 10  is an expanded inner perspective view of the through-holes and pockets according to another disclosed non-limiting embodiment; 
           [0033]      FIG. 11  is an inner perspective view of a through hole with corresponding pockets defined in the case assembly according to another disclosed non-limiting embodiment; 
           [0034]      FIG. 12  is a lateral sectional view showing the pockets of  FIG. 11 ; and 
           [0035]      FIG. 13  is a perspective sectional view showing example stress concentrations adjacent to the pockets of  FIG. 11 . 
       
    
    
     DETAILED DESCRIPTION 
       [0036]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbo fan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Another alternative engine architecture  200  might include an augmentor section  12 , an exhaust duct section  14  and a nozzle section  16  in addition to the fan section  22 ′, compressor section  24 ′, combustor section  26 ′ and turbine section  28 ′ (see  FIG. 2 ). Although depicted as an aero engine in the disclosed non-limiting embodiments, it should be understood that the concepts described herein are not so limited and the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans with an intermediate spool. 
         [0037]    Referring to  FIG. 1 , the fan section  22  drives air along a bypass flowpath and into the compressor section  24 . The compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26 , which then expands and directs the air through the turbine section  28 . The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine case assembly  36  via several bearing structures  38 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor (“LPC”)  44  and a low pressure turbine (“LPT”)  46 . The inner shaft  40  may drive the fan  42  directly (see  FIG. 2 ) or through a geared architecture  48  (see  FIG. 1 ) to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
         [0038]    The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (“HPC”)  52  and a high pressure turbine (“HPT”)  54 . A combustor  56  is arranged between the HPC  52  and the HPT  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0039]    Core airflow is compressed by the LPC  44  then the HPC  52 , mixed with the fuel and burned in the combustor  56 , then expanded over the HPT  54  and the LPT  46 . The LPT  46  and the HPT  54  rotationally drive the respective low spool  30  and the high spool  32  in response to the expansion. The main engine shafts  40 ,  50  are supported at a plurality of points by the bearing structures  38  within the static structure  36 . 
         [0040]    In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine with a bypass ratio greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3:1, and in another example, is greater than about 2.5:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the LPC  44  and the LPT  46  to render increased pressure in a fewer number of stages. 
         [0041]    A pressure ratio associated with the LPT  46  is pressure measured prior to the inlet of the LPT  46  as related to the pressure at the outlet of the LPT  46  prior to an exhaust nozzle of the gas turbine engine  20 . In another non-limiting example, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC  44 , and the LPT  46  has a pressure ratio greater than about five (5:1). It should be appreciated, however, that the above parameters are only exemplary of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
         [0042]    In an example high-bypass turbofan embodiment, significant thrust is provided by the bypass flow path due to the high bypass ratio as the fan section  22  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
         [0043]    Fan Pressure Ratio is the pressure ratio across a fan blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“Tram”/518.7) 0.5 . The Low Corrected Fan Tip Speed according to the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
         [0044]    With reference to  FIG. 3 , the combustor section  26  generally includes a combustor  56  with an outer combustor wall assembly  60 , an inner combustor wall assembly  62  and a diffuser case  64 . The outer combustor wall assembly  60  and the inner combustor wall assembly  62  are spaced apart such that an annular combustion chamber  66  is defined therebetween. 
         [0045]    The outer combustor wall assembly  60  is spaced radially inward from an outer diffuser case  64 A of the diffuser case  64  to define an outer annular plenum  76 . The inner combustor wall assembly  62  is spaced radially outward from an inner diffuser case  64 B of the diffuser case  64  to define an inner annular plenum  78 . It should be understood that although a particular combustor is illustrated, other combustor types with various combustor wall and diffuser case arrangements will also benefit herefrom. 
         [0046]    The combustor wall assemblies  60 ,  62  contain the combustion products for direction toward the turbine section  28 . Each combustor wall assembly  60 ,  62  generally includes a respective support shell  68 ,  70  which supports one or more liner panels  72 ,  74  mounted within the respective support shell  68 ,  70 . Each of the liner panels  72 ,  74  may be generally rectilinear with a circumferential arc and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array. In one disclosed non-limiting embodiment, the liner array includes a multiple of forward liner panels  72 A and a multiple of aft liner panels  72 B that are circumferentially staggered to line the outer shell  68 . A multiple of forward liner panels  74 A and a multiple of aft liner panels  74 B are circumferentially staggered to line the inner shell  70 . 
         [0047]    The combustor  56  further includes a forward assembly  80  immediately downstream of the compressor section  24  (see  FIG. 1 ) to receive compressed airflow therefrom. The forward assembly  80  generally includes an annular hood  82  and a bulkhead assembly  84  that support a multiple of fuel nozzles  86  (one shown) and a multiple of swirlers  90  (one shown). The annular hood  82  extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies  60 ,  62 . The annular hood  82  includes a multiple of circumferentially distributed hood ports  94  that accommodate the respective fuel nozzle  86  and introduce air into the forward end of the combustion chamber  66  through a respective swirler  90 . The bulkhead assembly  84  includes a bulkhead support shell  96  secured to the combustor wall assemblies  60 ,  62 , and a multiple of circumferentially distributed bulkhead liner panels  98  secured to the bulkhead support shell  96 . Each fuel nozzle  86  may be secured to the diffuser case  64  and project through one of the hood ports  94  and respective swirlers  90 . 
         [0048]    The forward assembly  80  introduces core combustion air into the forward section of the combustion chamber  66  while the remainder enters the outer annular plenum  76  and the inner annular plenum  78 . The multiple of fuel nozzles  86  and adjacent structure generate a fuel-air mixture that supports stable combustion in the combustion chamber  66 . 
         [0049]    Opposite the forward assembly  80 , the outer and inner support shells  68 ,  70  are mounted to a first row of Nozzle Guide Vanes (NGVs)  54 A in the HPT  54  (see  FIG. 1 ). The NGVs  54 A are static engine components which direct core airflow combustion gases onto turbine blades in the turbine section  28  to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs  54 A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. 
         [0050]    With reference to  FIG. 4 , the engine case assembly  36  generally includes a multiple of cases or modules in addition to the outer diffuser case  64 A to include, for example, a fan case  100 , an intermediate case  102 , a HPC case  104 , the outer diffuser case  64 A, a HPT case  106 , a mid turbine frame (MTF) case  108 , a LPT case  110 , and a Turbine Exhaust Case (TEC)  112 . The fan case  100  is bolted to the intermediate case  102 , which is bolted to the HPC case  104 , which is bolted to the outer diffuser case  64 A, which is bolted to the HPT case  106 , which is bolted to the MTF case  108 , which is bolted to the LPT case  110 , which is bolted to the TEC  112  each at a respective flange. It should be understood that the order of assembly may not necessarily follow the disclosed description and that various additional or alternative cases may be provided. 
         [0051]    With reference to  FIG. 5 , the outer diffuser case  64 A generally includes a multiple of through-holes  120  which penetrate through a wall  122  typical of holes configured to receive instrumentation such as a borescope, threaded holes for bolts to mount various components such as the fuel injectors and other types of apertures. The through-holes  120  may be defined through a boss  124  or other feature which extends from an outer surface  126  of the wall  122  (see  FIG. 6 ). It should be appreciated that various through-holes inclusive of those through a boss, not through a boss, or other through-holes will benefit herefrom. 
         [0052]    The outer diffuser case  64 A is pressurized, which produces hoop stresses in the wall  122 . At the holes  120 , stresses are relatively high. The through-holes  120  create high stress concentrations in the wall material that may otherwise reduce the strength and life of the component. To reduce these stresses, an inner surface  128  of the wall  122  includes pockets  130  which are operatively disposed adjacent to and circumferentially flank each through-hole  120  (see  FIG. 7 ). A single through-hole  120  requires two pockets  130 . That is, the pockets  130  are located on either side of the through-hole  120  in the hoop direction along a hoop line H such that the pockets  130  are aligned with the stress state such as the hoop stresses to break or otherwise shield the through-hole  120  from the nominal local stresses. It will be appreciated by those skilled in the art that such pockets  130  provide space for material of wall  122  to deform, expand, and/or contract, and thus reduce stresses in the wall material defining the through-holes  120 . 
         [0053]    With reference to  FIG. 7 , the pockets  130  have a circular outer edge. However, race track pockets  130 A ( FIG. 8 ), rectilinear pockets  130 B ( FIG. 9 ) or other shapes can alternatively or additionally be utilized. That is, an arrangement of multiple through holes  120 C in proximity may blend multiple pockets  130 C ( FIG. 10 ) which form other shield pocket geometries, e.g., race track, rectilinear etc. 
         [0054]    With reference to  FIG. 11 , in one disclosed non-limiting embodiment, the pockets  130  are generally larger than the through-hole  120  and, preferably, are of a diameter D 100%-500% of a diameter D of the through-hole  120 . The pockets  130  may be dimple shaped and extend from the inner surface  128  for a depth T of between 10%-50% a thickness ‘t’ of the wall  122  ( FIG. 12 ). It should be appreciated that the pockets  130  may alternatively include a flat bottom, a curved bottom, or be spherical in shape. An outer diameter  132  of each of the pockets  130  are also circumferentially spaced a distance Z from an outer diameter  134  of the through-hole  120  along the hoop line H from between 10%-100% the diameter of the through-hole  120 . It should be appreciated that various combinations of the above non-limiting embodiment parameters as well as others will also benefit herefrom. 
         [0055]    With reference to  FIG. 13 , the pockets  130  reduce the stress in and around material disposed about an inner periphery of the through-hole  120 , which increases the strength and reduces crack initiation at that location. The pockets  130 , in one tested example, reduce stress by approximately 10% at the critical stress location  140 , but increase stress in non-critical areas, locations  142  and  144 . Relatively deeper pockets  130  drive higher stress into locations  142  and  144  to reduce stress at location  140  such that optimal pocket  130  depth T results in a desired balance of stress. The pockets  130  also advantageously reduce weight. 
         [0056]    The use of the terms “a” and “an” and “the” and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
         [0057]    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. 
         [0058]    It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
         [0059]    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. 
         [0060]    The foregoing description is exemplary rather than defined by the features 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 appreciated 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.