Patent Publication Number: US-2017350265-A1

Title: Flow metering and directing ring seal

Description:
BACKGROUND 
     The present disclosure relates to a gas turbine engine and more particularly, to the metering and directing of cooling flow within the gas turbine engine through the use of a ring seal. 
     A gas turbine engine may include a secondary flow system that is configured to provide the internal components of the engine with cooling air in order to prevent or limit deterioration and damage to those components. However, the cooling air is sometimes taken from the compressor, so the cooling air can be at a higher temperature than the engine components in contact with ambient air, such as the turbine case. The secondary flow system allows for cooling air to flow through various passages within the engine to cool the engine components as needed. In the turbine section of the gas turbine engine, axial ring seals, also known as dog bone seals, are commonly used to meter flow through the use of dimensionally controlled holes extending through the ring seals. However, the cooling flow through the axial ring seals can unintentionally contact the turbine case (as shown in  FIG. 3 ), causing impingement and high convective heat transfer that results in the turbine case expanding due to the high temperature of the cooling air and the relatively low temperature of the turbine case. Due to thermal expansion in a radially outward direction, the turbine case can pull the turbine blade outer air seal radially outward, increasing the space between the turbine blade and the blade outer air seal. The increase in space between the turbine blade and the blade outer air seal allows air to seep passed the turbine blades without exerting force on the turbine blades, causing a decrease in efficiency of the engine. Therefore, it is advantageous to prevent the cooling air from contacting the turbine case and causing the turbine case to thermally expand. 
     SUMMARY 
     A seal ring includes a disk having a radially outer end and an orifice at a center forming a radially inner end, a plurality of holes extending through the disk, and a rail having an annular shape extending axially away from the disk at a position that is between the radially outer end and the plurality of holes extending through the disk with the rail being configured to direct the flow of fluid. 
     A metering seal in a turbine section of a gas turbine engine includes a sealing ring extending between a vane support on a radially outer end and a vane outer shroud on a radially inner end, a series of holes extending through the sealing ring, and a wall connected to the sealing ring. The wall has an annular shape and extending in an axial direction and is configured to direct the flow of a cooling fluid to at least partially prevent the cooling fluid from contacting a turbine case radially outward from the metering seal. 
     A seal includes a ring with an enlarged radially outer end, an enlarged radially inner end, and a middle section between the radially outer end and the radially inner end and having a thickness that is less than a thickness of the radially outer end and a thickness of the radially inner end to form a dog-bone shape, the middle section having a plurality of holes. The seal also includes a rail having an annular shape that extends away from the middle section of the ring. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a side elevation cutaway view of a gas turbine engine. 
         FIG. 2A  is a cross-sectional view of a first stage of a high pressure turbine section with a prior art ring seal. 
         FIG. 2B  is a partial perspective view of the prior art ring seal interacting with a turbine case. 
         FIG. 3  is a cross-sectional view of a first stage of a high pressure turbine section with a first embodiment of a flow metering and directing axial ring seal. 
         FIG. 4  is a partial perspective view of the flow metering and directing axial ring seal. 
         FIG. 5  is a partial perspective view of the flow metering and directing axial ring seal interacting with a turbine case. 
         FIG. 6  is a cross-sectional view of the flow metering and directing axial ring seal. 
         FIG. 7  is a cross-sectional view of another embodiment of the flow metering and directing axial ring seal. 
     
    
    
     DETAILED DESCRIPTION 
     A flow metering and directing ring seal for use in a gas turbine engine or another type of engine is disclosed herein that includes a disk with a plurality of flow metering holes extending through the disk and a flow directing annular rail extending away from the disk. The flow metering holes allow cooling air to flow through the gas turbine engine, including the internal components of the turbine section. The rail directs the flow of cooling air away from a turbine case radially outward from the ring seal and towards internal components of a turbine section axially downstream from the ring seal, preventing the cooling air from contacting the turbine case and ensuring the cooling air reaches the internal components of the turbine section to reduce deterioration and damage to the engine. 
     Without the rail, the cooling air can unintentionally interact with the turbine case, causing impingement and high convective heat transfer that results in the turbine case expanding due to the high temperature of the cooling air (the cooling air is usually taken from a compressor section of the engine, which results in the cooling air having a higher temperature than the ambient air surrounding the turbine case). Due to thermal expansion, the turbine case can expand radially outward and pull on a blade outer air seal, which is intended to limit a gap between the rotating turbine rotor blade and the outer shroud of a core gas path. With the blade outer air seal pulled radially outward, the gap between the turbine rotor blade and the outer shroud increases, allowing the air flowing through the core gas path to seep passed the turbine blade resulting in a decrease in efficiency of the engine. Therefore, the flow metering and directing ring seal with the rail is important because the rail directs cooling air towards the internal components of the turbine section and away from the turbine case, reducing deterioration and damage to the engine and increasing efficiency of the engine by preventing the increase in the gap between the turbine rotor blade and the outer shroud of the core gas path through the turbine section. 
       FIG. 1  is a side elevation cutaway view of gas turbine engine  10 , which is shown as a turbofan engine for an aircraft propulsion system. Gas turbine engine  10  extends along axial centerline  12  between upstream airflow inlet  14  and downstream airflow exhaust  16 . 
     Gas turbine engine  10  includes fan section  18 , compressor section  19 , combustor section  20 , and turbine section  21 . Compressor section  19  includes low pressure compressor (“LPC”) section  19 A and high pressure compressor (“HPC”) section  19 B. Turbine section  21  includes high pressure turbine (“HPT”) section  21 A and low pressure turbine (“LPT”) section  21 B. 
     Fan section  18 , compressor section  19 , combustor section  20 , and turbine section  21  are arranged sequentially along centerline  12  within engine housing  22 . Engine housing  22  includes inner case  24  (e.g., a core case) and outer case  26  (e.g., a fan case). Inner case  24  may house one or more of fan section  18 , compressor  19 , combustor section  20 , and turbine section  21  (e.g., an engine core). Outer case  26  may house at least fan section  18 . 
     Each of gas turbine engine sections  18 ,  19 A,  19 B,  21 A and  21 B includes fan rotor  28 , LPC rotor  29 , HPC rotor  30 , HPT rotor  31 , and LPT rotor  32 , respectively. Each of these rotors  28 ,  29 ,  30 ,  21 , and  32  includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered, and/or otherwise attached to the respective rotor disk(s). 
     Fan rotor  28  is connected to gear train  34 , for example, through fan shaft  36 . Gear train  34  and LPC rotor  29  are connected to and driven by LPT rotor  32  through low speed shaft  37 . The combination of at least LPC rotor  29 , LPT rotor  32 , and low speed shaft  37  may be referred to as “a low speed spool.” HPC rotor  30  is connected to and driven by HPT rotor  31  through high speed shaft  38 . The combination of at least HPC rotor  30 , HPT rotor  31 , and high speed shaft  38  may be referred to as “a high speed spool.” Shafts  36 ,  37 , and  38  are rotatably supported by a plurality of bearings  40 , which can be rolling element bearings, thrust bearings, or other types of bearings. Each of these bearings  40  is connected to engine housing  22  by at least one stationary structure such as, for example, an annular support strut. 
     During operation, air enters gas turbine engine  10  through airflow inlet  14 . Air is directed through fan section  18  and is then split into either core gas path  42  or bypass gas path  44 . Core gas path  42  flows sequentially through fan section  18 , compressor section  19 , combustor section  20 , and turbine section  21 . The air within core gas path  42  may be referred to as “core air.” Bypass gas path  44  flows through a duct between inner case  24  and outer case  26 . The air within bypass gas path  44  may be referred to as “bypass air.” 
     The core air is compressed by LPC rotor  29  and HPC rotor  30  and directed into combustion chamber  46  of combustor  48  in combustor section  20 . Fuel is injected into combustion chamber  46  and mixed with the core air that has been compressed by compressor section  19  to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof expand and flow through and sequentially cause HPT rotor  31  and LPT rotor  32  to rotate. The rotation of HPT rotor  31  and LPT rotor  32  respectively drive rotation of LPC rotor  29  and HPC rotor  30  and compression of the air received from core gas path  42 . The rotation of LPT rotor  32  also drives rotation of fan rotor  28 , which propels bypass air through and out of bypass gas path  44 . The propulsion of the bypass air may account for a majority of thrust generated by gas turbine engine  10 , which can be more than 75% of engine thrust. Gas turbine engine  10  of the present disclosure, however, is not limited to the foregoing exemplary thrust ratio. 
     During operation, turbine section  21  can experience elevated temperatures due to the elevated temperature of the core air flowing through core gas path  42  after the core air has flowed through combustor section  20 . To cool turbine section  21  to reduce deterioration and damage, cooling air is directed through the internal components of the turbine section. One strategy in the prior art for channeling cooling air was to use prior art ring seal  50 , as shown in  FIGS. 2A and 2B .  FIG. 2A  is a cross-sectional view of high pressure turbine section  21 A with prior art ring seal  50 , while  FIG. 2B  is a partial perspective view of prior art ring seal  50  interacting with inner case  24  (with the portion surrounding turbine section  21  referred to as turbine case  52 ). High pressure turbine section  21 A includes prior art ring seal  50 , turbine case  52  (with tab  54 ), vane  56 , vane outer shroud  58 , vane support  60 , rotor blade  62 , blade outer air seal  64 , coating  66 , and internal components  68 . Cooling air flow  70  flows through prior art ring seal  50 , and gap  72  is present between rotor blade  62  and coating  66  on blade outer air seal  64 . 
     The prior art configuration meters cooling air flow  70  (taken from compressor section  19 ) flowing through holes in prior art ring seal  50 . However, after flowing through holes in prior art ring seal  50 , cooling air flow  70  can broaden and come into contact with an inner side of turbine case  52  and tab  54  instead of flowing axially rearward towards internal components  68  and blade outer air seal  64  to cool those components. Because a radially outer side of turbine case  52  of inner case  24  is in contact with the near ambient temperature bypass air flowing through bypass gas path  44 , turbine case  52  is at a relatively low temperature as compared to the temperature of cooling air flow  70 . Therefore, when the elevated temperature cooling air flow  70  contacts turbine case  52  and results in impingement and high convective heat transfer, turbine case  52  undergoes thermal expansion in a radially outward direction. With turbine case  52  being the structural component to which blade outer air seal  64  is attached, blade outer air seal  64  is pulled radially outward by turbine case  52 . Because blade outer air seal  64  (with coating  66 ) is radially outward than when in ideal operation, gap  72  gets larger, allowing core air to seep passed rotor blade  72 . With core air seeping passed rotor blade  72 , less energy is transferred to rotor blade  72  by the core air, reducing the efficiency of gas turbine engine  10 . Further, with cooling air flow  70  interacting with turbine case  52  instead of flowing to and cooling internal components  68 , internal components  68  have an increased risk of deterioration and becoming damaged. Thus, a ring seal; such as ring seal  150  in  FIGS. 3, 4, 5, and 6 ; that meters and directs flow is advantageous. 
       FIG. 3  is a cross-sectional view of high pressure turbine section  21 A with a first embodiment of a flow metering and directing ring seal  150 ,  FIG. 4  is a partial perspective view of ring seal  150 ,  FIG. 5  is a partial perspective view of ring seal  150  interacting with turbine case  52 , and  FIG. 6  is a cross-sectional view of ring seal  150 . High pressure turbine section  21 A is shown in  FIG. 3  with ring seal  150 , turbine case  52  (with tab  54 ), vane  56 , vane outer shroud  58 , vane support  60 , rotor blade  62 , blade outer air seal  64 , coating  66 , and internal components  68 . Cooling air flow  70  flows through ring seal  150 , and gap  72  is present between rotor blade  62  and coating  66  on blade outer air seal  64 . Ring seal  150  includes disk  174 , plurality of holes  176 , rail  178 , radially outer end  180 A, radially inner end  180 B, disk hole fillets  182 , rail taper  184 , and hard-faced coating  186 . The functionality and configuration of high pressure turbine section  21 A, turbine case  52 , vane  56 , vane outer shroud  58 , vane support  60 , rotor blade  62 , blade outer air seal  64  (with coating  66 ), and internal components  68  in  FIGS. 3 and 5  are the same as those components in  FIGS. 1, 2A, and 2B . 
     Ring seal  150  has an annular shape and is located in a similar position to prior art ring seal  50  of  FIGS. 2A and 2B , with ring seal  150  spanning and sealing a substantially radial distance between vane outer shroud  58  and vane support  60 . Ring seal  150  can be constructed from a variety of materials, including a nickel based alloy or another type of alloy, but should have sufficient strength and resilience to bridge this distance between vane outer shroud  58  and vane support  60  while also maintaining contact with both elements to provide a complete and satisfactory seal. Ring seal  150  should be able to handle the high temperatures of cooling air flow  70  without becoming damaged while still providing a satisfactory seal. Further, ring seal  150  should be constructed from a material that can be in contact with and does not damage vane outer shroud  58  and vane support  60 . Ring seal  150  can include hard-faced coating  186  at contact points to ensure ring seal  150  does not damage vane outer shroud  58  and vane support  60  (or vice-versa) and a satisfactory seal is formed and maintained. The components of ring seal  150  discussed below can be individual pieces fastened together to form ring seal  150 , or ring seal  150  can be one continuous and monolithic piece formed during the manufacturing process. Ring seal  150  can be constructed using a variety of manufacturing processes, including additive manufacturing. While ring seal  150  is disclosed as sealing the distance between vane outer shroud  58  and vane support  60 , ring seal  150  can be configured to seal any distance in other areas of gas turbine engine  10  while also metering and directing a cooling air flow to necessary components. 
     Disk  174  (also referred to as a sealing ring or a ring) of ring seal  150  is an annular-shaped component that spans the distance between vane outer shroud  58  and vane support  60 . As shown most easily in  FIG. 4 , disk  174  has an orifice (thus making disk  174  an annular shape) and is centered about centerline  12  of gas turbine engine  10 . Disk  174  can be entirely in a radial direction or can be angled such that radially outer end  180 A is axially forward from radially inner end  180 B to form a frustoconical shape. Further, if the design of gas turbine engine requires, disk  174  can have another shape and configuration to span and seal a distance other than that shown in the disclosed drawings. A thickness of disk  174  can vary depending on the material used to construct disk  174 , the distance that disk  174  must span, the number and surface area of the plurality of holes  176  in disk  174 , the temperature of cooling air flow  70 , and other factors. In the disclosed embodiment, the thickness of disk  174  at a middle section is approximately 0.152 centimeters (0.060 inches) with ring seal  150  being constructed from a nickel based alloy. 
     Radially outer end  180 A of disk  174  is in contact with and adjacent to vane support  60 , and radially inner end  180 B of disk  174  is in contact with and adjacent to vane outer shroud  58 . Radially outer end  180 A can have a thickness that is the same as the thickness of disk  174  or can have an increase in thickness as compared to the middle section of disk  174  to aid in sealing ring seal  150  to vane support  60  and to provide additional structural rigidity and strength to ring seal  150 . The increase in thickness can be a bulge that extends axially forward (as shown in  FIG. 4 ), a bulge that extends axially rearward, or both. Similarly, radially inner end  180 B can have a thickness that is the same as the thickness of disk  174  or can have an increase in thickness as compared to the middle section of disk  174  to aid in sealing ring seal  150  to vane outer shroud  58  and to provide additional structural rigidity and strength to ring seal  150 . As with radially outer end  180 A, radially inner end  180 B can have the increase in thickness as a bulge that extends axially forward, a bulge that extends axially rearward, or both (as shown in  FIG. 4 ). As shown in  FIG. 4 , disk  174  has a dog-bone shape with the radially outer end  180 A and the radially inner end  180 B having a bulb-shaped configuration. Radially outer end  180 A and radially inner end  180 B can have hard-faced coating  186 . Hard-faced coating  186  is on a contact surface between ring seal  150  and vane outer shroud  58  and vane support  60 ; specifically, hard-faced coating  186  is on a forward facing surface of radially outer end  180 A and on a rearward facing surface of radially inner end  180 B. Hard-faced coating  186  aids in establishing and maintaining a complete seal while also protecting ring seal  150 , vane outer shroud  58 , and vane support  60  from damage due to contact and friction between components. Hard-faced coating  186  can be any type of coating that is configured to seal and protect ring seal  150 , vane outer shroud  58 , and vane support  60 , including chromium carbide. Hard-faced coating  186  can be applied through a variety of means, including plasma spraying. While the disclosed embodiment shows hard-faced coating  186  only on ring seal  150  at radially outer end  180 A and radially inner end  180 B, hard-faced coating  186  and/or other coatings can be applied to ring seal  150  at other locations. 
     Disk  174  includes a plurality of holes  176  that extend through the middle section of disk  174 . The plurality of holes  176  meter cooling air flow  70  through high pressure turbine section  21 A. The plurality of holes  176  are spaced around disk  174  as most easily seen in  FIG. 4 , but the plurality of holes  176  do not need to be spaced equally from adjacent holes and the holes do not need to be the same radial distance from radially inner end  180 B. Disk  174  can have any number of holes of the plurality of holes  176  depending on various design considerations, including the necessary amount of cooling air flow  70 , the strength and rigidity requirements of ring seal  150 , and others. In one embodiment, the number of the holes of the plurality of holes  176  is between 40 and 100. The plurality of holes  176  can have a configuration in which one hole is present per circumferential location (as shown in  FIG. 4 ) or multiple holes are at the same circumferential location with the holes having a different radial distance from radially inner end  180 B. Further, while the disclosed embodiment shows the plurality of holes  176  all having a similarly-sized circular shape, each of the plurality of holes  176  can have a different size (i.e., surface area) and/or shape other than a circular shape, depending on design considerations and the amount of cooling air flow  70  that is needed by internal components  68  and other components of gas turbine engine  10 . Also, the size and shape of each hole of the plurality of holes  176  can be different than adjacent holes. Each of the plurality of holes  176  can have a consistent cross-section as the holes extend through disk  174 , or the cross-section can vary to control/meter cooling air flow  70  as desired through the plurality of holes  176 . The plurality of holes can extend through disk  174  at a variety of angles when measured from a surface of the middle section of disk  174 , but the plurality of holes  176  as shown in  FIGS. 3, 4, 5, and 6  extend through disk  174  at an angle that is approximately perpendicular to the surface of the middle section. The plurality of holes  176  can extend through disk  174  at a location that is closer to radially inner end  180 B, closer to radially outer end  180 A, or equidistant between the ends of disk  174 . However, ring seal  150  with the plurality of holes  176  at a location that is approximately equidistant between radially outer end  180 A and radially inner end  180 B likely has a greater rigidity and strength than if the plurality of holes  176  were at another radial location. 
     Disk hole fillets  182  are located at a transition area between disk  174  and each of the plurality of holes  176 . Disk hole fillets  182  are an annular, curved fillet on an axially forward transition area between an axially forward surface of the middle section of disk  174  and each of the plurality of holes  176  and on an axially rearward transition area between an axially rearward surface of the middles section of disk  174  and each of the plurality of holes  176 . A radius of curvature of each disk hole fillet  182  can be constant around the annular fillet and the same on the axially forward surface as the fillet is on the axially rearward surface, or the radius of curvature can vary as each disk hole fillet  182  extends around each of the plurality of holes  182 . Disk hole fillets  182  aid in metering cooling air flow  70  and in providing a smooth, laminar flow path through the plurality of holes  176 . While the disclosed embodiment has disk hole fillets  182  on both sides of disk  174  and on each of the plurality of holes  176 , seal ring  150  can include disk hole fillets  182  on only one or on neither of the sides of disk  174  and/or on only a portion of the plurality of holes  176 . Disk hole fillets  182  can be formed through a variety of manufacturing processes, including additive manufacturing, molding or another type of forming, machining the plurality of holes  176  and disk hole fillets  182  after the rest of ring seal  150  has been constructed, or another method. 
     Rail  178  (also referred to as a wall) has a hollow cylindrical shape and is connected to a rearward side of the middle section of disk  174 . Rail  178  direct cooling air flow  70  flowing through the plurality of holes  176  away from turbine case  52  and towards internal components  68 . Cooling air flow  70  flowing through the plurality of holes  176  contacts rail  178  and is redirected to flow substantially axially rearward. Rail  178  extends away from disk  174  at a location between the plurality of holes  176  and radially outer end  180 A, and rail  178  can extend away from disk  174  at any angle, including an angle that is perpendicular to disk  174 , an angle that is an acute angle when measured from a radially inward portion of the middle section of disk  174 , or another angle. As shown in the disclosure, disk  174  can be slightly angled with radially inner end  180 B being rearward from radially outer end  180 A, and rail  178  can extend rearward in a substantially axial direction (forming an acute angle between rail  178  and the radially inward portion of disk  174 ). 
     Rail  178  can have any length, such as a length that is 10% of the length of disk  174 , making a ratio of the length of rail  178  to a distance between radially outer end  180 A and radially inner end  180 B (i.e., a length of disk  174 ) approximately 0.1. Further, rail  178  can have a length that is equal to the length of disk  174  for a ratio of 1.0. Rail  178  can have another length, such as a length that is between 10% and 100% of the length of disk  174  or a length that is larger. No matter what the length of rail  178  is, the length should be sufficiently long to direct cooling air flow  70  away from turbine case  52  and towards internal components  68  and other components in need of cooling air flow  70 . The length of rail  178  can depend on a number of factors, including the diameter/surface area of each of the plurality of holes  176 , the flow velocity of cooling air flow  70 , the angle at which each of the plurality of holes  176  extends through disk  174 , a distance from the plurality of holes  176  to rail  178 , whether disk  174  is angled, and other factors. 
     A thickness of rail  178  (i.e., radial thickness) should be sufficient to withstand the high temperatures of cooling air flow  70  while also maintaining structural rigidity and strength. The thickness of rail  178  can vary depending on the material used to construct rail  178 , the length of rail  178 , the temperature of cooling air flow  70 , and other factors. The thickness of rail  178  can be constant along the axial length of rail  178  or can vary such that the thickness of rail  178  is larger at a point closer to disk  174  and smaller at a point farther from disk  174 . Further, the thickness of rail  178  can have other configurations, including a thickness that is greatest near a tip of rail  178 , a thickness that is equal to the thickness of the middle section of disk  174 , or other configurations. The thickness of rail  178  can be the same as the thickness of the middles section of disk  174  such that the ratio of the thickness of rail  178  to the thickness of disk  174  is approximately 1. In the disclosed embodiment, the thickness of rail  178  is approximately 0.152 centimeters (0.060 inches). Rail  178  can also include rail taper  184 , which is an angled or curved fillet at the tip of rail  178 . Rail taper  184  can be present on a radially inner side of rail  178 , a radially outer side of rail  178 , on both sides, or on neither side. Rail taper  184  aids in transitioning cooling air flow  70  to internal components  68 . 
     Rail  178  can be fastened to disk  174  through a variety of means, including welding, braising, glue, bolts, or another fastener. Additionally, disk  174  and rail  178  can be one continuous and monolithic piece formed during the manufacturing process. A transition from disk  174  to rail  178  can include a fillet on one or both of the radially outer side and the radially inner side to increase rigidity and strength in that area. Rail  178  can be constructed from a variety of materials, including a nickel based alloy or another type of alloy, but should have sufficient strength and resilience to extend rearward from disk  174  while also maintaining structural rigidity and strength to direct cooling air flow  70 . 
     As mentioned above, rail  178  directs cooling air flow  70  after the air has passed through the plurality of holes  176 . Rail  178  directs cooling air flow  70  away from turbine case  52 , which is radially outward from rail  178 , and towards internal components  68  in need of cooling air. Rail  178  prevents cooling air flow  70  from contacting turbine case  52  and ensures cooling air flow  70  reaches internal components  68  to reduce deterioration and damage to gas turbine engine  10 . Depending on design considerations and the metering and directing needs of the seal, ring seal  150  can have a variety of shapes and configurations, such as the embodiment of  FIG. 7 . 
       FIG. 7  is a cross-sectional view of another embodiment of a ring seal. Ring seal  250  includes disk  274 , plurality of holes  276 , rail  278 , radially outer end  280 A, radially inner end  280 B, disk hole fillets  282 , rail taper  284 , and hard-faced coating  286 . The functionality and configuration of ring seal  250  is similar to that of ring seal  150  of  FIGS. 3-6 . However, ring seal  250  has rail  278  with a length that is longer than rail  178  and the configuration of the plurality of holes  276  are different than the plurality of holes  176  of ring seal  150 . 
     Disk  274  of ring seal  250  includes the plurality of holes  276  that have an increased diameter over the plurality of holes  176  of ring seal  150 . Also, the plurality of holes  276  extend through disk  274  at a nonperpendicular angle when measured in relation to a surface of disk  274  (and approximately horizontal when viewed in relation to gas turbine engine  10  in which ring seal  250  is implemented). The plurality of holes  276  can have any size, shape, and configuration that meters cooling air flow  70 , and the size, shape, and configuration is dependent on the amount of cooling air flow  70  needed for internal components  68 , the temperature of cooling air flow  70 , the velocity of cooling air flow  70 , and other factors. 
     Rail  278  of ring seal  250  has a length that is longer than rail  178  of ring seal  150 , with rail  278  having a length that is approximately equal to a distance between radially outer end  280 A and radially inner end  280 B (i.e., a length of disk  274 ). Ring seal  250  can have any length of rail  278 , but rail  278  should have a length that is sufficient to direct cooling air flow  70  away from turbine case  52  and towards internal components  68 . The length of rail  278  can depend on a number of factors, including the diameter/surface area of each of the plurality of holes  276 , the flow velocity of cooling air flow  70 , the angle at which each of the plurality of holes  276  extends through disk  274 , a distance from the plurality of holes  276  to rail  278 , whether disk  274  is angled, and other factors. 
     The flow metering and directing ring seal  150 / 250  for use in gas turbine engine  10  or another type of engine is disclosed herein that includes disk  174 / 274  with the plurality of holes  176 / 276  extending through disk  174 / 274  to meter cooling air flow  70  and the flow directing annular rail  178 / 278  extending away from disk  174 / 274 . The flow metering plurality of holes  176 / 276  allow cooling air flow  70  to flow through gas turbine engine  10  to internal components  68  of turbine section  21 . Rail  178 / 278  directs cooling air flow  70  away from turbine case  52  radially outward from ring seal  150 / 250  and towards internal components  68  of turbine section  21  axially downstream from ring seal  150 / 250 , preventing cooling air flow  70  from contacting turbine case  52  and ensuring cooling air flow  70  reaches internal components  68  of turbine section  21  to reduce deterioration and damage to gas turbine engine  10 . 
     Without rail  178 / 278 , cooling air flow  70  can unintentionally interact with turbine case  52 , causing impingement and high convective heat transfer that results in turbine case  52  expanding due to the high temperature of cooling air flow  70 . Due to thermal expansion, turbine case  52  can expand radially outward and pull on blade outer air seal  64 , which is intended to limit gap  72  between the rotating turbine rotor blade  62  and blade outer air seal  64  of core gas path  42 . With blade outer air seal  64  pulled radially outward, gap  72  between rotor blade  62  and blade outer air seal  64  increases, allowing the air flowing through core gas path  42  to seep passed rotor blade  62  resulting in a decrease in efficiency of gas turbine engine  10 . Therefore, the flow metering and directing ring seal  150 / 250  with rail  178 / 278  is important because rail  178 / 278  directs cooling air flow  70  towards internal components  68  of turbine section  21  and away from turbine case  52 , reducing deterioration and damage to gas turbine engine  10  and increasing efficiency of gas turbine engine  10  by preventing the increase in gap  72  between rotor blade  62  blade outer air seal  64  of core gas path  42  through turbine section  21 . 
     Discussion of Possible Embodiments 
     The following are non-exclusive descriptions of possible embodiments of the present invention. 
     A seal ring includes a disk having a radially outer end and an orifice at a center forming a radially inner end, a plurality of holes extending through the disk, and a rail having an annular shape extending axially away from the disk at a position that is between the radially outer end and the plurality of holes extending through the disk with the rail being configured to direct the flow of fluid. 
     The seal ring of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     The rail is connected to the disk at a position that is closer to the radially outer end than to the radially inner end of the disk. 
     The radially outer end of the disk is axially forward from the radially inner end so the disk is a frustoconical shape. 
     Each of the plurality of holes includes a fillet configured to direct the flow of fluid through the plurality of holes. 
     The plurality of holes is arranged circumferentially around the disk and extends at an angle that is perpendicular to the disk. 
     Each of the plurality of holes are equally spaced from adjacent holes. 
     The number of the plurality of holes is between 40 and 100. 
     Each of the plurality of holes have an equal cross-sectional area to one another. 
     A ratio of a length of the rail to a length of the disk is approximately 1. 
     A tip of the rail is tapered. 
     A coating on an axially forward facing surface of the radially outer end and on an axially rearward facing surface of the radially inner end. 
     A metering seal in a turbine section of a gas turbine engine includes a sealing ring extending between a vane support on a radially outer end and a vane outer shroud on a radially inner end, a series of holes extending through the sealing ring, and a wall connected to the sealing ring. The wall has an annular shape and extending in an axial direction and is configured to direct the flow of a cooling fluid to at least partially prevent the cooling fluid from contacting a turbine case radially outward from the metering seal. 
     The metering seal of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     A size of each hole of the series of holes is dependent on an amount of cooling fluid needed to be supplied to the turbine section. 
     A ratio of a radial thickness of the wall to an axial thickness of the sealing ring is approximately 1. 
     The sealing ring and wall are one continuous and monolithic component. 
     A seal includes a ring with an enlarged radially outer end, an enlarged radially inner end, and a middle section between the radially outer end and the radially inner end and having a thickness that is less than a thickness of the radially outer end and a thickness of the radially inner end to form a dog-bone shape. The seal also includes a rail having an annular shape that extends away from the middle section of the ring. 
     The seal of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     A hard-faced coating on an axially forward facing surface of the radially outer end and on an axially rearward facing surface of the radially inner end. 
     The radially outer end is axially forward from the radially inner end and the plurality of holes extend through the middle section at an angle that is perpendicular to the middle section. 
     A tip of the rail is axially rearward from the radially inner end of the ring. 
     A ratio of a length of the rail to a length of the ring is between approximately 0.1 and 1.0. 
     Any relative terms or terms of degree used herein, such as “substantially,” “essentially,” “generally,” “approximately,” and the like should be interpreted in accordance with and subject to any applicable definitions or limits expressly stated herein. In all instances, any relative terms or terms of degree used herein should be interpreted to broadly encompass any relevant disclosed embodiments as well as such ranges or variations as would be understood by a person of ordinary skill in the art in view of the entirety of the present disclosure, such as to encompass ordinary manufacturing tolerance variations; incidental alignment variations; alignment or shape variations induced by thermal, rotational, or vibrational operational conditions; and the like. 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.