Patent Abstract:
A grommet assembly and method of design to enhance the flow coefficient, thereof, includes a shell having a first side and an opposite second side, and a chamfered grommet projecting through the shell along a centerline and including an annular first end surface spaced outward from the first side and a conical face spanning axially and radially inward from the annular first end surface and axially beyond the second side. The assembly may further include a panel spaced from the shell and defining a cooling cavity therebetween with the conical surface defining at least in-part a hole in fluid communication through the shell and panel and isolated from the cooling cavity. A plurality of cooling channels in the grommet are in fluid communication with the cooling cavity and communicate through the panel. The combination of the conical face and the cooling channels improve the discharge coefficient of the grommet while enhancing grommet cooling.

Full Description:
GROMMET ASSEMBLY AND METHOD OF DESIGN 
       [0001]    This application claims priority to U.S. Patent Appln. No. 61/974,248 filed Apr. 2, 2014. 
     
    
     BACKGROUND 
       [0002]    The present disclosure relates to a grommet assembly and, more particularly, to a dilution air grommet assembly for a combustor and method of design to enhance flow coefficient. 
         [0003]    Gas turbine engines, such as those that power modern commercial and military aircraft, include a fan section to propel the aircraft, a compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and thereby generate thrust. 
         [0004]    The combustor section typically includes a wall assembly having an outer shell lined with heat shields that are often referred to as floatwall panels. Together, the panels define a combustion chamber. A plurality of dilution holes are generally spaced circumferentially about the wall assembly and flow dilution air from a cooling plenum and into the combustion chamber to improve emissions, and reduce and control the temperature profile of combustion gases at the combustor outlet to protect the turbine section from overheating. 
         [0005]    The dilution holes are generally defined by a grommet that extends between the heat shield panel and supporting shell with a cooling cavity defined therebetween. Enhanced cooling of the grommets is desirable for improved engine efficiency, robustness, and durability. 
       SUMMARY  
       [0006]    A grommet according to one, non-limiting, embodiment of the present disclosure includes a core including a chamfered inlet portion having a chamfered ratio equal to or greater than 0.10. 
         [0007]    Additionally to the foregoing embodiment, the chamfered core defines an axial length ratio equal to or greater than 0.25. 
         [0008]    In the alternative or additionally thereto, in the foregoing embodiment, the core includes a cylindrical face and a conical face extending outward from the cylindrical face at a peripheral inner edge, and the cylindrical and conical faces define a hole extending along a centerline through the core. 
         [0009]    In the alternative or additionally thereto, in the foregoing embodiment, the conical face extends transverse to a reference plane normal to the centerline at an angle of about twenty-five to forty-five degrees. 
         [0010]    In the alternative or additionally thereto, in the foregoing embodiment, the conical face is angled from a reference plane disposed normal to the centerline at about thirty degrees. 
         [0011]    In the alternative or additionally thereto, in the foregoing embodiment, a hole communicates through the core along a centerline and is defined at least in-part by a conical face spanning axially and radially outward to an annular end surface carried by the core, and the conical face extends transverse to a reference plane normal to the centerline at an angle of about twenty-five to forty-five degrees. 
         [0012]    In the alternative or additionally thereto, in the foregoing embodiment, a hole communicates through the core along a centerline, and the core carries and extends between opposite annular first and second end surfaces concentrically disposed to the centerline, and wherein the second end surface is located at least in-part radially inward from the first end surface. 
         [0013]    In the alternative or additionally thereto, in the foregoing embodiment, the grommet includes a flange projecting radially outward from the core and spaced axially between the first and second end surfaces. 
         [0014]    In the alternative or additionally thereto, in the foregoing embodiment, the flange includes a peripheral face spanning axially and extending circumferentially around the core, and wherein a plurality of cooling channels are circumferentially spaced from one another and each one of the plurality of cooling channels extends between and communicates through the peripheral face and the second end surface. 
         [0015]    A grommet assembly according to another, non-limiting, embodiment includes a shell having a first side and an opposite second side; a chamfered core projecting through the shell along a centerline and including an annular first end surface spaced outward from the first side and a conical face spanning axially and radially inward from the annular first end surface and axially beyond the second side; and wherein the conical face defines at least in-part a hole in the core and communicating through the shell. 
         [0016]    Additionally the foregoing embodiment, the assembly includes a flange projecting radially outward from the core and spaced axially between the annular first end surface and an opposite, annular, second end surface of the chamfered core. 
         [0017]    In the alternative or additionally thereto, in the foregoing embodiment, the flange carries a peripheral face spanning axially and extending circumferentially around the core, and wherein a plurality of cooling channels are circumferentially spaced from one another and each one of the plurality of cooling channels extend between and communicate through the peripheral face and the second end surface. 
         [0018]    In the alternative or additionally thereto, in the foregoing embodiment, the assembly includes a panel with a cooling cavity defined between the shell and the panel; and wherein the flange is in the cooling cavity. 
         [0019]    In the alternative or additionally thereto, in the foregoing embodiment, the hole is a dilution hole and is in fluid communication between a cooling plenum defined in part by the first side and a combustion chamber defined in-part by the panel. 
         [0020]    In the alternative or additionally thereto, in the foregoing embodiment, the chamfered core has a chamfered ratio equal to or greater than 0.10. 
         [0021]    In the alternative or additionally thereto, in the foregoing embodiment, the conical face spans axially and radially inward to a cylindrical face defining in-part the hole, and the chamfered core has an axial length ratio equal to or greater than 0.25. 
         [0022]    In the alternative or additionally thereto, in the foregoing embodiment, the conical face extends transverse to a reference plane normal to the centerline at an angle of about twenty-five to forty-five degrees. 
         [0023]    In the alternative or additionally thereto, in the foregoing embodiment, the conical face is angled from a reference plane disposed normal to the centerline at about thirty degrees. 
         [0024]    A method of enhancing a discharge coefficient of a grommet assembly design according to another, non-limiting, embodiment includes the steps of choosing an angle between about twenty-five to forty-five degrees wherein an inlet portion of a core of the assembly includes a conical face defining at least in-part a hole extending along a centerline through the core, and wherein the conical face extends transverse to a reference plane normal to the centerline at the angle; choosing a chamfered ratio of a chamfered inlet portion of a core of the grommet assembly; choosing an axial length ratio of the core; choosing a chart based on the chosen angle; and determining the discharge coefficient from the chart displaying axial length ratio verse chamfered ratio. 
         [0025]    Additionally to the foregoing embodiment, the chamfered ratio is equal to or greater than 0.10 and the axial length ratio is equal to or greater than 0.25. 
         [0026]    The foregoing features and elements may be combined in various combination 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 figures are intended to exemplary in nature and non-limiting. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    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: 
           [0028]      FIG. 1  is a schematic cross-section of a gas turbine engine; 
           [0029]      FIG. 2  is a cross-section of a combustor section; 
           [0030]      FIG. 3  is a cross section of a grommet assembly according to one non-limiting example of the present disclosure; 
           [0031]      FIG. 4  is a bottom plan view of a grommet of the grommet assembly; 
           [0032]      FIG. 5  is a partial perspective view of the grommet assembly; and 
           [0033]      FIG. 6  is a graph of an axial length ratio verse a chamfered ratio to determine a discharge coefficient signified by a plurality of charted, curved, lines. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      FIG. 1  schematically illustrates a gas turbine engine  20  disclosed 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 . Alternative engines may 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 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 engine architecture such as turbojets, turboshafts, and three-spool turbofans with an intermediate spool. 
         [0035]    The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine axis A via several bearing structures  38  and relative to a static engine case  36 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42  of the fan section  22 , a low pressure compressor  44  (“LPC”) of the compressor section  24  and a low pressure turbine  46  (“LPT”) of the turbine section  28 . The inner shaft  40  drives the fan  42  directly or through a geared architecture  48  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. 
         [0036]    The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  (“HPC”) of the compressor section  24  and a high pressure turbine  54  (“HPT”) of the turbine section  28 . A combustor  56  of the combustor section  26  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 axis A. 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 HPT  54  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. 
         [0037]    In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is 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 that can increase the operational efficiency of the LPC  44  and LPT  46  and render increased pressure in a fewer number of stages. 
         [0038]    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 one 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 the LPC  44 ; and the LPT  46  has a pressure ratio that is greater than about five (5:1). It should be understood; however, that the above parameters are only exemplary of one example of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
         [0039]    In one non-limiting example, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  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. 
         [0040]    Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a fan exit guide vane system. The low Fan Pressure Ratio according to one non-limiting example of the gas turbine engine  20  is less than 1.45:1. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (T/518.7 0.5 ), where “T” represents the ambient temperature in degrees Rankine The Low Corrected Fan Tip Speed according to one non-limiting example of the gas turbine engine  20  is less than about 1150 fps (351 m/s). 
         [0041]    Referring to  FIG. 2 , the combustor section  26  generally includes an annular combustor  56  with an outer combustor wall assembly  60 , an inner combustor wall assembly  62 , and a diffuser case module  64  that surrounds assemblies  60 ,  62 . The outer and inner combustor wall assemblies  60 ,  62  are generally cylindrical and radially spaced apart such that an annular combustion chamber  66  is defined therebetween. The outer combustor wall assembly  60  is spaced radially inward from an outer diffuser case  68  of the diffuser case module  64  to define an outer annular plenum  70 . The inner wall assembly  62  is spaced radially outward from an inner diffuser case  72  of the diffuser case module  64  to define, in-part, an inner annular plenum  74 . Although a particular combustor is illustrated, it should be understood that other combustor types with various combustor liner arrangements will also benefit. It is further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be so limited. 
         [0042]    The combustion chamber  66  contains the combustion products that flow axially toward the turbine section  28 . Each combustor wall assembly  60 ,  62  generally includes a respective support shell  76 ,  78  that supports one or more heat shields or liners  80 ,  82 . Each of the liners  80 ,  82  may be formed of a plurality of floating panels that are generally rectilinear and manufactured of, for example, a nickel based super alloy that may be coated with a ceramic or other temperature resistant material, and are arranged to form a liner configuration mounted to the respective shells  76 ,  78 . 
         [0043]    The combustor  56  further includes a forward assembly  84  that receives compressed airflow from the compressor section  24  located immediately upstream. The forward assembly  84  generally includes an annular hood  86 , a bulkhead assembly  88 , and a plurality of swirlers  90  (one shown). Each of the swirlers  90  are circumferentially aligned with one of a plurality of fuel nozzles  92  (one shown) and a respective hood port  94  to project through the bulkhead assembly  88 . The bulkhead assembly  88  includes a bulkhead support shell  96  secured to the combustor wall assemblies  60 ,  62  and a plurality of circumferentially distributed bulkhead heat shields or panels  98  secured to the bulkhead support shell  96  around each respective swirler  90  opening. The bulkhead support shell  96  is generally annular and the plurality of circumferentially distributed bulkhead panels  98  are segmented, typically one to each fuel nozzle  92  and swirler  90 . 
         [0044]    The annular hood  86  extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies  60 ,  62 . Each one of the plurality of circumferentially distributed hood ports  94  receives a respective on the plurality of fuel nozzles  92 , and facilitates the direction of compressed air into the forward end of the combustion chamber  66  through a swirler opening  100 . Each fuel nozzle  92  may be secured to the diffuser case module  64  and projects through one of the hood ports  94  into the respective swirler  90 . 
         [0045]    The forward assembly  84  introduces core combustion air into the forward section of the combustion chamber  66  while the remainder of compressor air enters the outer annular plenum  70  and the inner annular plenum  74 . The plurality of fuel nozzles  92  and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber  66 . 
         [0046]    Referring to  FIGS. 3 and 4 , a dilution hole grommet assembly  102  is illustrated and described in relation to the outer wall assembly  60  for simplicity of explanation; however, it is understood that the same grommet assembly may be applied to the inner wall assembly  62  of the combustor  56 . The grommet assembly  102  includes a portion of the support shell  76 , a portion of the heat shield or panel  80 , and a grommet  104 . The grommet assembly  102  generally functions to flow dilution air (see arrow  106 ) from the cooling plenum  70 , through the wall assembly  60 , via the grommet  104 , and into the combustion chamber  66 . This dilution air generally enters the combustion chamber  66  as a jet stream to improve combustion efficiency generally in a core region of the chamber and further serves to cool and/or control the temperature profile of combustion air at the exit of the combustor  56 . 
         [0047]    The heat resistant panel  80  of wall assembly  60  (which may include an array of panels) includes a hot side  108  that generally defines in-part a boundary of the combustion chamber  66  and an opposite cold side  110 . The shell  76  includes an outer side  112  that faces and defines in-part a boundary of the cooling plenum  70  and an opposite inner side  114  that faces and is spaced from the cold side  110  of the heat shield  80 . An annular cooling cavity  116  is located between and defined by the cold side  110  of the heat shield  80  and the inner side  114  of the shell  76 . 
         [0048]    An aperture  118  may communicate through the heat shield  80  and is defined by a circumferentially continuous surface  120  of the heat shield  80  and spanning axially between the hot and cold sides  108 ,  110 . Similarly, an aperture  122  communicates through the shell  76  and is defined by a circumferentially continuous surface  124  of the shell  76  and spanning axially between the outer and inner sides  112 ,  114 . A centerline  126  extends through the apertures  118 ,  122  and may be substantially normal to the wall assembly  60  and may intersect the engine axis A ( FIG. 1 ). 
         [0049]    The grommet  104  of the grommet assembly  102  has a chamfered core  128  that defines a dilution hole  130 , and a flange  132  that projects radially outward from the core  128  and into the cooling cavity  116 . The core  128 , the dilution hole  130  and the flange  132  may all be substantially concentric to the centerline  126 . The cooling cavity  116  does not generally communicate directly with the dilution hole  130 . Thus, the flange  132  may be in circumferentially continuous sealing contact with the inner side  114  of the shell  76  and may be cast as one piece, brazed, or otherwise adhered to the cold side  110  and/or continuous surface  120  of the panel  80 . 
         [0050]    The chamfered core  128  extends into the aperture  118  of the panel  80  and through the aperture  122  of the shell  76  and into the cooling plenum  70 . More specifically, the core  128  carries opposite annular end surfaces  134 ,  136 , both concentric to the centerline  126 , with end surface  134  located in the cooling plenum  70  and spaced outward from the outer side  112  of the shell  76 , and with end surface  136  being substantially flush with the hot side  108  of the panel  80 . The core  128  further includes a substantially conical face  138 , a peripheral inner edge or apex  140 , and a substantially cylindrical face  142  that together generally define the dilution hole  130 . The conical face  138  extends axially and radially inward from the annular end surface  134  and to the inner edge  140 . The cylindrical face  142  extends axially from the inner edge  140  to the end surface  136 . Thus, the chamfered core  128  generally includes a chamfered inlet portion  141  having the conical face  138  and a cylindrical outlet portion  143  having the cylindrical face  142 . 
         [0051]    The grommet  104  further has a plurality of cooling channels  144  spaced circumferentially about the grommet for flowing cooling air from the cooling cavity  116  and into the combustion chamber  66  (see arrow  146 ) for generally cooling the core  128  of the grommet  104  at and/or near the end surface  136 , and which may further enhance penetration of the dilution air jet flow  106  into the combustion chamber  66 . Each cooling channel  144  has an inlet generally defined by an outer circumferential, or peripheral, face  148  of the flange  132  and an outlet defined by the annular end surface  136  of the core  128 . Each channel  144  thus communicates through the flange  132  and the core  128  providing distributed fluid communication between the cooling cavity  116  and the combustion chamber  66 . For ease of manufacturing, each channel  144  may generally be a groove in the grommet  104 , and generally defined between the cold side  110  and continuous surface  120  of the panel  80 , and the flange  132  and core  128  of the grommet  104 . 
         [0052]    Referring to  FIGS. 5 and 6 , more traditional grommet assemblies display low discharge coefficients signifying impaired dilution air jet flow penetration into the core regions of the combustion chamber  66 . Such low discharge coefficients may be attributable to hot combustion air recirculation zones at or near the dilution air grommet that may further cause overheating and degradation of the grommet In accordance with the present disclosure, significant grommet performance and durability improvements (e.g. reduced metal temperatures) can be achieved through use of particular dimensional relationships of the core  128  of the grommet  104  and the cooling channels  144 . 
         [0053]    These dimensional relationships may generally be as follows: 
         [0000]        W/D≧ 1/10;  L/D≧ 1/4;  W≧H; T≧H    
         [0054]    where ‘W/D’ is a chamfered ratio, ‘L/D’ is an axial length ratio, ‘W’ is a distance measured axially (i.e. with respect to centerline  126 ) between the inner edge  140  and the annular end surface  134 , ‘D’ is an outer diameter of the conical face  138  (i.e. where the conical face  138  meets the end surface  134 ), ‘L’ is a distance measured axially between the opposite end surfaces  134 ,  136 , ‘H’ is the distance measured between the outer and inner sides  112 ,  114  of the shell  76  (i.e. shell thickness), and ‘T’ is the distance measured axially between the end surface  134  and the flange  132 . These dimensional relationships may be combined with an angle (see arrow  150  in  FIG. 5 ) of the conical surface  138  (measured from a reference plane that is substantially normal to the centerline  126 ) that falls within a range of twenty-five to forty-five degrees. 
         [0055]    As one non-limiting example, a more traditional discharge coefficient (Cd) can be improved from about  0 . 6  to about  0 . 9 , thereby reducing or eliminating gas recirculation and reducing local metal temperatures from about a melting temperature of the alloy to about a  400  degree Fahrenheit margin below melting temperature when the conical surface angle  150  is about thirty degrees, ‘W’ is equal to or greater than about three times the panel  80  thickness, ‘L’ is equal to or greater than about six times the panel thickness, ‘D’ is equal to or greater than about twenty times the panel thickness, and a hydraulic diameter of the cooling channel  144  (eight illustrated in  FIG. 4 ) is about equal to or greater than  0 . 5  times the panel thickness. As a more specific, non-limiting, example: ‘W’ may be about 0.105 inches (2.667 mm), ‘L’ may be about 0.225 inches (5.715 mm), ‘D’ may be about 0.878 inches (22.301 mm), and the hydraulic diameter of each cooling channel  144  may be about  0 . 020  inches (0.508 mm). 
         [0056]    Referring further to  FIG. 6 , a graph illustrates the chamfered ratio ‘W/D’ verse the axial length ratio with the conical surface angle  150  at about thirty degrees. Empirical data further depicts discharge coefficient values (i.e. 0.7 through 0.95) as a function of the chamfered ratio versus axial length ratio. That is, each discharge coefficient value is represented by a charted, curved, line gathered empirically. Generally, with increasing chamfered and axial length ratios, the discharge coefficient value also increases. Therefore, with predetermined chamfered and axial length ratios, one can determine the discharge coefficient value when the conical surface angle  150  is thirty degrees. 
         [0057]    It is understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude and should not be considered otherwise limiting. It is also understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will also benefit. Although particular step sequences may be shown, described, and claimed, it is understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
         [0058]    The foregoing description is exemplary rather than defined by the limitations described. Various non-limiting embodiments are disclosed; 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 understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For this reason, the appended claims should be studied to determine true scope and content.

Technology Classification (CPC): 8