Abstract:
A turbine comprising a first turbine component being of a first material having a first coefficient of thermal expansion. A second turbine component being of a second material having a second coefficient of thermal expansion, said second turbine component adjacent said first turbine component. A space between said first and second turbine components. A seal assembly sealing said space, wherein at least a portion of said seal assembly has a coefficient of thermal expansion substantially similar to at least one of said first or second turbine components to thereby maintain a seal in said space during thermal expansion or contraction of said first and second turbine components.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0001]    This invention was made with Government support under Contract No. W911W6-11-2-0009 awarded by the United States Army. The Government has certain rights in this invention. 
     
    
     TECHNICAL FIELD 
       [0002]    The application relates to turbines, and more specifically, preventing leakage of air in a turbine between multiple turbine components. 
       BACKGROUND 
       [0003]    The aircraft and aircraft engine industry consistently seeks to make improvements to increase fuel efficiency, or reduce specific fuel consumption (SFC) of its technology. Like the automobile industry, much of the efforts to reduce SFC in the aircraft and aircraft engine industry have focused on increasing the overall efficiency of the engine itself. In striving towards that goal, much of the attention is also directed towards reducing the overall weight of the engine. Due to advances in materials design, much attention has been focused on replacing heavier, metal parts with lighter materials, such as ceramic and composite materials, that can withstand the heat and forces that occur in an aircraft turbine engine. While replacing certain parts of the engine with a lighter material reduces the weight of the engine, certain issues arise when the materials are exposed to the high temperature environment within the aircraft engine. One of the issues is the disparity of relative thermal expansions between metal parts and parts made of, for example, ceramic or composite materials. When a metal part is adjacent to a part comprised of a ceramic or composite material, the metal and ceramic or composite parts will expand a different amount and at a different rate, thereby potentially creating unwanted space therebetween. Depending on the location within the engine of the unwanted space or opening, the space or opening may lead to air leaks or other airflow issues therein. Leaks and other airflow issues may reduce the efficiency and therefore increase the SFC of the engine. Due to the disadvantageous nature of leakage flows and the relative motion a thin, compliant seal such as a convoluted, or “W” seal is generally utilized. However in a system where the “W” seal is centered between a metal and ceramic or composite components the relative thermal expansions will cause the thin seal to roll, thereby causing high stresses on the seal and reducing the seal effectiveness and life. Therefore, there is a need in the art for a device and method to counteract the disadvantageous behavior of materials used in turbines which have disparate coefficients of thermal expansion. 
       SUMMARY 
       [0004]    It is therefore desirable to provide a device and method to prevent leakage of air within a turbine. A turbine is provided and comprises a first turbine component being of a first material having a first coefficient of thermal expansion and a second turbine component being of a second material having a second coefficient of thermal expansion, the second turbine component adjacent the first turbine component. A space is between the first and second turbine components. The turbine further comprises a seal assembly sealing the space. At least a portion of the seal assembly has a coefficient of thermal expansion substantially similar to at least one of the first or second turbine components to thereby maintain a seal in the space during thermal expansion or contraction of the first and second turbine components. 
         [0005]    A method of preventing leakage in a turbine is also provided and comprises situating a seal assembly in a space between first and second turbine components, thereby sealing the space. The first and second turbine components are of first and second materials having first and second coefficients of thermal expansion, respectively. A portion of said seal assembly has a coefficient of thermal expansion substantially similar to at least one of said first or second turbine components. The method further comprises maintaining a seal in the space during thermal expansion or contraction of said first and second turbine components. 
         [0006]    A seal assembly configured to seal a space between first and second objects is provided and comprises a seal member having first and second ends, the second end spaced from the first end along an axis. The assembly further comprises a seal carrier enveloping at least a portion of the seal member. The seal carrier is adapted to prevent relative movement between the first and second ends in a direction transverse to the axis. 
     
    
     
       BRIEF DESCRIPTION 
         [0007]      FIG. 1  shows a perspective view of one embodiment of a turbine. 
           [0008]      FIG. 2  shows a side cross-sectional view of the turbine of  FIG. 1  along lines  2 - 2 . 
           [0009]      FIG. 3  shows a side cross-sectional view of a turbine of the prior art as it experiences a change in ambient temperature. 
           [0010]      FIG. 4  shows a side cross-sectional view the turbine of  FIG. 2  as it experiences a change in ambient temperature. 
           [0011]      FIGS. 5 and 8  show perspective views of alternative embodiments of a turbine. 
           [0012]      FIGS. 6 and 9  show side cross-sectional views of alternative embodiments of the turbine of  FIGS. 5 and 8 , along lines  6 - 6  and  9 - 9 , respectively. 
           [0013]      FIGS. 7 and 10  show side cross-sectional views of alternative embodiments of the turbine of  FIGS. 5 and 8 , along lines  7 - 7  and  10 - 10 , respectively. 
           [0014]      FIGS. 11 and 12  show side cross-sectional views of steps of the assembly method of the embodiment of  FIG. 1 . 
           [0015]      FIGS. 13 and 14  show side cross-sectional views of steps of the assembly method of the embodiments shown in  FIGS. 5 and 8 , respectively. 
       
    
    
     DETAILED DESCRIPTION 
       [0016]      FIGS. 1 ,  2  and  4  show one embodiment of a portion of a turbine  10 . The turbine includes a combustor (not shown) having a liner. The liner  12  is situated at an angle relative to a center axis (not shown) of the turbine  10 . The liner  12  includes an outer radial side  14  and an inner radial side  16 . The inner radial side  16  communicates with the channel  23  through which the combusted gases may flow to the nozzle  18 . The turbine  10  includes a first stage nozzle  18  aft of the liner  12 . The nozzle  18  includes an outer radial side  20  and an inner radial side  22 . The nozzle  18  comprises a first portion  24  situated at an angle relative to the center axis of the channel  23  and a second portion  26  extending radially outward from the first portion  24  of the nozzle  18 . There is a space  27  ( FIG. 4 ) between the liner  12  and the nozzle  18 . 
         [0017]    The turbine  10  further includes a seal assembly  28 . The seal assembly  28  includes a carrier  30  and a seal member  32 . The carrier  30  includes a generally flat, axial flange  33  and first and second flanges  44 ,  50  extending radially inward from the radially outward portion  33 . The seal member  32  as shown in  FIG. 2  includes a convoluted portion  34  including multiple folds or convolutions  35  such that it is a generally “W” shaped member. However, there may be more convolutions  35  than shown in  FIG. 2 . The seal member  32  includes first and second ends  36 ,  38 . The first end  36  is forward of the convoluted portion  34  and the second end  38  is aft of the convoluted portion  34 . A generally open portion  40  is disposed generally radially away from the convoluted portion  34  and between the first and second ends  36 ,  38 . 
         [0018]    The seal carrier  30  is preferably situated at the space  27  such that a forward face  42  of the first flange  44  of the carrier  30  is engaged with a contact portion  46  of the liner  12 . The forward face  48  of the second flange  50  may be engaged with the aft face  51  of the second portion  26  of the nozzle  18 . However, the aft face  51  of the second portion  26  and forward face  48  of second flange  50  need not be engaged, as the forward face  48  of the second flange  50  may create an axial restraint with aft face  51  of the second portion  26 . The axial restraint created therebetween may ensure that the carrier  30  does become disassembled during the installation or assembly process. At least part of the seal member  32  is enveloped by the carrier  30 . More specifically, the axial flange  33  is radially outward of the open portion  40  and the aft face  54  of first flange  44  of the carrier  30  is engaged with the first end  36  of the seal member  32 . The second end  38  of the seal member  32  is engaged with a forward face  56  of the second portion  26  of the nozzle  18 . The configuration of the seal assembly  28  provides for multiple points where a seal is provided, thereby preventing leakage of air in at least the radial direction. More specifically, in the embodiment shown in  FIG. 2 , a seal may be provided between the contact portion  46  of the liner  12  and the forward face  42  of the first flange  44 , between the first end  36  of the seal member  32  and the aft face  54  of the first flange  44 , and between the second end  38  of the seal member  32  and the forward face  56  of the second portion  26  of the nozzle  18 . These seals thereby prevent the disadvantageous leakage of air in the radial direction. Furthermore, in the embodiment shown in  FIGS. 2 and 4 , the carrier  30  further includes a third flange  58  configured to act as a thermal barrier between the space  27  and at least a portion of the seal member  32 . 
         [0019]    The invention is particularly advantageous for preventing leakage in the radial direction between a liner  12  of a combustor and a first stage nozzle  18 . For this reason, the system and method is described herein with a frame of reference to such components of a turbine. Moreover, terms such as radial, circumferential and axial are used to describe the system in the chosen frame of reference. The invention, however, is not limited to the chosen frame of reference and descriptive terms, and may be used on turbine components other than the liner  12  of a combustor and a first stage nozzle  18 , and in other orientations in a turbine. Those of ordinary skill in the art will recognize that descriptive terms used herein may not directly apply when there is a change in the frame of reference. Nevertheless, the disclosure is intended to be independent of location and orientation within a turbine and the relative terms used to describe the system and method are to merely provide an adequate description of the disclosure. 
         [0020]    With reference to  FIG. 2 , the liner  12  comprises a first material and the nozzle  18  comprises a second material. Preferably, the liner  12  comprises a ceramic matrix composite (CMC) material and the nozzle  18  comprises a metal. Because CMC material may be prone to wear, a lubricious coating may be provided on the contact portion  46  of the liner  12  or the forward face  42  of first flange  44 , or between the two components. Due to the differing coefficients of thermal expansion between these two materials, the liner  12  and the nozzle  18  may expand at different rates and different amounts when subjected to an ambient temperature change, such as when hot, combusted gas travels from the combustor to the first stage nozzle  18 . For example, as shown in  FIG. 3 , the liner  12 ′ may expand a distance ΔR LINER  and the nozzle  18 ′ may expand a distance ΔR NOZZLE . Issues may arise due to disparate thermal expansion between the liner  12 ′ and the nozzle  18 ′, and more specifically, issues relative to sealing the space therebetween, as discussed in further detail below. 
         [0021]    A prior art turbine is shown in  FIG. 3  with a seal member  32 ′ provided between the liner  12 ′ and the nozzle  18 ′. As the nozzle  18 ′ expands a greater amount than the liner  12 ′, the second end  38 ′ of the seal member  32 ′ moves radially relative to the first end  36 ′. Seal members  32 ′ with convolutions  35 ′ oriented in the axial direction generally cannot withstand large relative radial movement between axially spaced portions of the seal member  32 ′. This relative radial movement between the first and second ends  36 ′,  38 ′ of the seal member  32 ′ may be referred to as seal roll. Seal roll is disadvantageous because it may cause a moment in the circumferential direction, thereby essentially twisting the seal member  32 ′ in the circumferential direction. This moment M SEAL  may result in the seal member  32 ′ failing by becoming displaced and perhaps overturning within the space  27 ′, or by tearing or otherwise breaking 
         [0022]    The seal assembly as described herein substantially prevents seal roll, thereby potentially preventing failure of the seal member, and maintaining a seal at the space between the liner  12  and the nozzle  18 . The carrier  30  in each embodiment may comprise a material having the same or substantially similar coefficient of thermal expansion as the nozzle  18 . For example, the carrier  30  may be the same material as the nozzle  18 , such as metal, while the liner  12  comprises a different material, such as a ceramic, composite, or CMC. In an alternative embodiment, however, the carrier  30  may comprise a material having a same or substantially similar coefficient of thermal expansion as the liner  12 . The amount of thermal expansion can be calculated by ΔL=L*α*ΔT, where L is the length of the object in question, α is the coefficient of thermal expansion, and ΔT is the change in temperature. In one embodiment, the carrier  30  may comprise a material having a same or substantially similar coefficient of thermal expansion as the nozzle  18  such that the difference between ΔR LINER  and ΔR NOZZLE  is less than or equal to 0.030″ (0.762 mm). Therefore, because thermal expansion is dependent upon at least three variables, including the coefficient of thermal expansion, the difference between ΔR LINER  and ΔR NOZZLE  depends on more than just the coefficient of thermal expansion. Therefore, persons skilled in the art will recognize that providing such a difference between ΔR LINER  and ΔR NOZZLE  may be accomplished by altering the other variables on which the thermal expansion is dependent. However, it also may be appreciated by persons skilled in the art that a difference less than or equal to 0.030″ (0.762 mm) between ΔR LINER  and ΔR NOZZLE  is limited to one embodiment described herein and is not meant to limit other embodiments where the relative movement may be more or less. Moreover, a different amount of relative movement may be prescribed in an embodiment where the first and second turbine components are something other than a nozzle and a liner of a combustor. Moreover, the turbine components referred to herein are not limited to solely gas turbines engines used in aircrafts, but may also refer to turbine components in gas turbine engines for other applications, such as other types of machinery that utilize gas turbine engines. 
         [0023]    As the temperature of the ambient environment of the turbine increases, the nozzle  18  expands a distance ΔR NOZZLE  and the carrier expands a substantially similar or same amount ΔR CARRIER , while the liner  12  expands a different, and preferably lesser. amount ΔR LINER.  Because the first end  36  of the seal member  32  is coupled with the aft face  54  of the first flange  44  of the carrier  30 , and the second end  38  of the seal member  32  is coupled with the forward face  56  of the second portion  26  of the nozzle  18 , and the first and second ends  36 ,  38  may move a substantially same or similar amount such that relative movement between the first and second ends  36 ,  38  is minimal. This prevention of relative movement between the first and second ends  36 ,  38  of the seal member  32  thereby prevents the potential problems of seal roll described above. 
         [0024]    The embodiment shown in  FIG. 1  is one embodiment configured to maintain the seal in the space between the liner  12  and the nozzle  18  as described herein. Alternatively, the seal assembly  32  may be configured such that the carrier  30  substantially follows or mimics the expansion of the liner  12  and prevents relative radial movement between the first and the second ends  36 ,  38  of the seal member  32 . 
         [0025]    Alternative embodiments of a turbine are shown in  FIGS. 4 ,  6  &amp;  7  and  FIGS. 6 ,  9  &amp;  10 . In each embodiment, the turbine  110  includes a combustor (not shown) having a liner  112 . The liner  112  is situated at an angle relative to a center axis of the turbine  110 . The liner  112  includes an outer radial side  114  and an inner radial side  116 . The inner radial side communicates with the channel  123  through which the combusted gases may flow into the first stage. The turbine  110  includes a nozzle  118  aft of the liner  112 . The nozzle  118  includes an outer radial side  120  and an inner radial side  122 . The nozzle  118  comprises a first portion  124  situated at an angle relative to the center axis of the turbine  110  and a second portion  126  extending radially outward from the first portion  124  of the nozzle  118 . There is a space  127  between the liner  112  and the nozzle  118 . 
         [0026]    The turbine  110  further includes a seal assembly  128 . The seal assembly  128  includes a carrier  130  and a seal member  132 . The carrier  130  comprises a first member  134  and a second member  136 . The first member  134  includes an axial flange  138  and a radial flange  140  extending in the radially inward direction from the axial flange  138 . The second member  136  of the seal carrier  130  includes forward and aft radial flanges  142 ,  144  and an axial flange  146  between the forward and aft radial flanges  142 ,  144 . The forward and aft radial flanges  142 ,  144  and the axial flange  146  of the second member  136  essentially envelop the second portion  126  of the nozzle  118 . The forward radial flange  142  includes a second axial flange  147  extending in the axially forward position. 
         [0027]    There is a seal member  132  in the space  127 , a portion of which is engaged with the seal carrier  130 . The seal member  132  as shown in  FIG. 8  includes a convoluted portion  148  including multiple folds or convolutions  150  such that it is a generally “W” shaped member. However, there may be more convolutions  150  than shown in  FIGS. 6 and 9 . The seal member  132  includes first and second ends  152 ,  154 . The first end  152  is forward of the convoluted portion  148  and the second end  154  is aft of the convoluted portion  148 . A generally open portion  156  is disposed generally radially away from the convoluted portion  148  and between the first and second ends  152 ,  154 . 
         [0028]    The seal carrier  130  essentially envelops at least a portion of the seal member  132 . More specifically, the seal carrier  130  is preferably situated at the space  127  such that a forward face  158  of the radial flange  140  of the first member  134  is engaged with a contact portion  160  of the liner  112 . The seal member  132  and carrier  130  are positioned such that the axial flange  138  is radially outward of the open portion  156  and the aft face  162  of radial flange  140  of first member  134  is engaged with the first end  152  of the seal member  132 . The second end  154  of the seal member  132  is engaged with a forward face  164  of the forward radial flange  142  of the second member  136 . The configuration of the seal assembly  128  provides for multiple points where a seal is provided, thereby preventing leakage of air in at least the radial direction. More specifically, a seal is provided between the contact portion  160  of the liner  112  and the forward face  158  of the radial flange  140  of the first member  134 , between the first end  152  of the seal member  132  and the aft face  162  of the radial flange  140  of the first member  134 , and between the second end  154  of the seal member  132  and the forward face  164  of the forward radial flange  142  of the second member  136 , as well as between the aft face  165  of the forward radial flange  142  and the second portion  126  of the nozzle  118 . These seals thereby prevent the disadvantageous leakage of air in the radial direction. The second axial flange  147  of the second member  136  is adapted to provide a thermal barrier for at least a portion of the seal member  132 . 
         [0029]    The carrier  130  in may comprise a material having the same or substantially similar coefficient of thermal expansion as the nozzle  118 . For example, the carrier  130  may be the same material as the nozzle  118 , such as metal, while the liner  112  comprises a different material, such as a ceramic, composite, or CMC. In an alternative embodiment, however, the carrier  130  may comprise a material having a same or substantially similar coefficient of thermal expansion as the liner  112 . 
         [0030]    As shown in  FIGS. 5 &amp; 7  and  6  &amp;  10 , the axial flange  138  of the first member  134  is provided with circumferentially spaced resilient portions  166  ( FIG. 5 ),  166 ′ ( FIG. 8 ). In the embodiment as shown in  FIGS. 4 and 7 , each resilient portion  166  includes an axial elongate member  168  between a pair of axially oriented slots  170 . The slots  170  may be provided in order to increase the amount of radial deflection of the elongate members  168 . In the embodiment shown in  FIGS. 8 &amp; 10 , the resilient portions  166 ′ may include essentially circumferential elongate member  168 ′ defined in part by slots  170 ′. More specifically, there are a plurality of slots  170 ′, each slot  170 ′ including an axial portion  172   a,  an elongate circumferential portion  172   b,  and a curved transition portion  172   c  therebetween. 
         [0031]    The methods of assembling each embodiment are shown in  FIGS. 11 ,  12 ,  13  and  14 .  FIGS. 11 and 12  show the assembly of the turbine  10  shown in  FIGS. 1 ,  2  and  4 . The seal carrier  30  is directed into engagement with the liner  12  such that the forward face  42  of the first flange  44  may be in contact with the contact portion  46  of the liner  12 . Concurrently, or thereafter, the first end  36  of the seal member  32  is brought into engagement with the aft face  54  of the first flange  44  of the carrier  30 . The nozzle  18  is then brought into engagement with the second end  38  of the seal member  32 . More specifically, the forward face  56  of the second portion  26  of the nozzle  18  may be brought into engagement with the second end  38  of the seal member  32 . To secure the nozzle  18  relative to the combustor and the liner  12 , as well as to secure the seal assembly  28 , the axial flange  33  is bent, thereby forming the second flange  50 . Once assembled, the turbine  10  is configured as shown in  FIGS. 2 and 4 . Assembly may be accomplished in several manners other than that disclosed above. There are two additional ways this device may be assembled. First, the seal member  32  may be engaged or assembled into the carrier  30 . Next, the nozzle  18  may be assembled such that it is in engagement with the seal member  32 . The axial flange  33  may then be bent down, thereby forming the second flange. The second flange is then used to fix the carrier  30  and seal member  32  relative to the nozzle  18 . The nozzle  18 , carrier  30  and seal member may then be engaged with the liner  16 . In yet another alternative method of assembly, the second flange  50  is pre-bent in the configuration shown in  FIGS. 2 and 4 . The seal member  32  may then be assembled or engaged with the carrier  30 . Where the nozzle portion is segmented (not shown), each nozzle segment may be installed individually by aligning the nozzle to an area without second flange  50 , then clocking (rotating the nozzle  18  relative to the center axis) the nozzle  18  such that aft face  51  of the second portion  26  aligns with forward face  48  of second flange  50 . Preferably, the clocking may takes place when installing the last segment (not shown) of the nozzle  18 . 
         [0032]      FIGS. 13 and 14  show the assembly of the embodiments as shown in  FIGS. 6 and 9 , respectively. The first member  134  of seal carrier  130  is directed into engagement with the liner  112  such that the forward face  158  of the radial flange  140  may be in contact with the contact portion  160  of the liner  112 . Concurrently, or thereafter, the first end  152  of the seal member  132  is brought into engagement with the aft face  162  of the radial flange  140  of the carrier  130 . The second member  136  of the carrier  130  is directed into engagement with the nozzle  118  such that the forward and aft radial flanges  142 ,  144  and the axial flange  146  of the second member  136  essentially envelop the second portion  126  of the nozzle  118 . Thereafter, the second member  136  and nozzle  118 , as assembled together, are directed in the axial direction, as shown by arrows  176 , as the resilient portions  166 ,  166 ′, and more specifically the elongate members  168 ,  168 ′ are flexed in the radial direction, as indicated by arrows  178 . The radial deflection of the elongate members  168 ,  168 ′ allows for the second member  136  to be directed into engagement with the second end  154  of the seal member  132 . More specifically, the second end  154  of the seal member  132  is engaged with the forward face  164  of the forward radial flange  142  of the second member  136 . As shown in  FIGS. 7 and 10 , the elongate members  168 ,  168 ′ each include hooked portions  174 ,  174 ′, respectively. The hooked portions  174 ,  174 ′ may be configured to resist the axial movement of the nozzle  118  by engaging with the second member  136  as the nozzle  118  and the second member  136  may move axially in the aft direction. An alternate method of assembly includes engaging the seal member  132  to the carrier  134 , so that forward face  152  of seal member  132  is in contact with aft face  162  of carrier  134 . Then, second member  136  may be engaged with the assembly such that forward face  164  of second member  136  is in contact with aft seal face  154 . Nozzle  118  can be engaged, prior to, during or after engagement of second member  136 . Nozzle  118  shall be engaged such that radial portion  126  is enveloped by second member  136  between radial flanges  142  and  144 . 
         [0033]    The seal assembly  128  as described herein is not limited to sealing a space  127  between a combustion liner  112  and a first stage nozzle  118 . Rather, the seal assembly  128  as described herein may be configured to seal between first and second turbine components. The seal assembly may be configured to provide a seal where providing a seal is desired, preferably in a radial direction. Moreover, the turbine components referred to herein are not limited to solely gas turbines engines used in aircrafts, but may also refer to turbine components in gas turbine engines for other applications, such as other types of machinery that utilize gas turbine engines. 
         [0034]    While the present invention has been illustrated by a description of various preferred embodiments and while these embodiments have been described in some detail, it is not the intention of the Applicant to restrict or in any way limit the scope of the appended claims to such detail. Additional advantages and modifications will readily appear to those skilled in the art. The various features of the invention may be used alone or in any combination depending on the needs and preferences of the user. This has been a description of the present invention, along with the preferred methods of practicing the present invention as currently known. However, the invention itself should only be defined by the appended claims.