Abstract:
A gas turbine engine sealing arrangement which includes a combustor liner establishing a combustion area and a turbine nozzle configured to receive a portion of a combustor liner. A sealing ring arrangement is configured to seal with the turbine nozzle to control fluid flow from the combustion area and reduce leakage.

Description:
[0001]    This invention was made with government support under Contract No. N00019-06-C-0081 awarded by the United States Navy. The Government therefore may have certain rights in this invention. 
     
    
     BACKGROUND 
       [0002]    This application relates generally to sealing an interface in the combustor section of a gas turbine engine. 
         [0003]    Gas turbine engines are known and typically include multiple sections, such as a inlet section, a compression section, a combustor section, a turbine section, and an exhaust nozzle section. The inlet section moves air into the engine. The air is compressed in the compression section. The compressed air is mixed with fuel and is combusted in combustion areas within the combustor section. The products of the combustion are expanded through the turbine section to rotatably drive the engine. 
         [0004]    The combustor section of the gas turbine engine typically includes a combustor liner that establishes combustion areas within the combustor section. The combustion areas extend circumferentially around a centerline of the engine. The combustion areas in a can combustor are separated from each other. The combustion areas in an annular combustor are connected. Turbine nozzles direct the products of combustion from the combustion area to the turbine section in both types of combustors. Substantial leaks at the interfaces between the turbine nozzles and the combustion chambers can cause irregularities in temperature and pressure. The irregularities can reduce the usable life of the turbine nozzles, turbine wheels and other components. Some leakage may be acceptable if the leakage is predictable and relatively uniform. 
       SUMMARY 
       [0005]    An example gas turbine engine sealing arrangement includes a combustor liner mountable adjacent to a turbine nozzle and a flange extending from the combustor liner. The flange establishes a channel that holds a sealing ring assembly moveable to sealed position with the turbine nozzle. 
         [0006]    Another example gas turbine engine sealing arrangement includes a turbine nozzle having a wall that provides spaced apart first and second surfaces. A combustor liner establishes a combustion area and has an annularly extending collar securable adjacent the first surface. A sealing ring arrangement is configured to seal against the second surface to control fluid flow from the combustion area. A flange extending from the combustor liner holds the sealing ring arrangement. 
         [0007]    An example method of sealing an interface within a gas turbine engine includes holding a sealing ring arrangement relative to a turbine nozzle using a flange extending from a combustor liner. The method further includes urging the sealing ring arrangement toward a sealed relationship with the turbine nozzle. 
         [0008]    These and other features of the example disclosure can be best understood from the following specification and drawings, the following of which is a brief description: 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  is a sectional schematic view of an example gas turbine engine. 
           [0010]      FIG. 2  is a sectional side view of a portion of the combustor section in the  FIG. 1  engine. 
           [0011]      FIG. 3  is a close up view of area  3  in the  FIG. 2  combustion section. 
           [0012]      FIG. 4  is a sectional view of the combustor section in the  FIG. 1  engine at line  4 - 4  of  FIG. 2 . 
           [0013]      FIG. 5  is a perspective view of the  FIG. 2  sealing rings. 
       
    
    
     DETAILED DESCRIPTION 
       [0014]      FIG. 1  schematically illustrates an example gas turbine engine  10  including (in serial flow communication) an inlet section  14 , a centrifugal compressor  18 , a combustion section  26 , a turbine wheel  30 , and a turbine exhaust  34 . The gas turbine engine  10  is circumferentially disposed about an engine centerline X. During operation, air is pulled into the gas turbine engine  10  by the inlet section  14 , pressurized by the compressor  18 , mixed with fuel, and burned in the combustion section  26 . The turbine wheel  30  extracts energy from the hot combustion gases flowing from the combustion section  26 . 
         [0015]    In a radial turbine design, the turbine wheel  30  utilizes the extracted energy from the hot combustion gases to power the centrifugal compressor  18 . The examples described in this disclosure are not limited to the radial turbine auxiliary power unit described and may be used in other architectures, such as a single-spool axial design, a two spool axial design, and a three-spool axial design. That is, there are various types of engines that could benefit from the examples disclosed herein, which are not limited to the radial turbine design shown. 
         [0016]    Referring to  FIGS. 2-4  with continuing reference to  FIG. 1 , within the combustion section  26  of the engine  10  an example combustor liner  50  is secured relative to a turbine nozzle  54 . The combustor liner  50  establishes a combustion area  58 . A fuel nozzle  62  is configured to spray fuel into the combustion area  58 . Air is delivered to the combustion area  58  through apertures  66  in the combustion liner  50 . As known, the air pressure within the combustion area  58  is less than the air pressure outside the combustion area  58 . 
         [0017]    An igniter  70  ignites a mixture of fuel and air within the combustion area  58  to generate hot combustion gases G that are forced through the turbine nozzle  54 . The hot combustion gases G drive turbine wheel  30 . 
         [0018]    In this example, eight fuel nozzles  62  are circumferentially arranged about the engine centerline X. The fuel nozzles  62  are arranged such that the spray pattern of fuel from one of the fuel nozzles  62  slightly overlaps the spray pattern of fuel from an adjacent one of the fuel nozzles  62 . Arranging the fuel nozzles  62  in this manner facilitates evenly driving the turbine wheel  30  with the hot combustion gas G moving through the turbine nozzle  54 . 
         [0019]    The combustor liner  50  and the turbine nozzle  54  meet at an interface  74 . In this example, the turbine nozzle  54  provides an annular opening that is defined by spaced apart, concentric outer and inner walls  64  and  65 . The annular opening of the turbine nozzle  54  is configured to receive inner and outer collar portions  80  and  81  of the combustor liner  50 . In this example, the inner and outer collar portions  80  and  81  are placed within the annular turbine nozzle  54  between the outer and inner walls  64  and  65 . The inner collar portion  80  is placed adjacent to a radially outer surface  75  of the inner wall  65  in this example. 
         [0020]    A flange  78  extends from a radially inward face of the combustor liner  50  and is configured to hold a plurality of axially aligned sealing rings  82 , such that the inner wall  65  of the turbine nozzle  54  is positioned radially between the inner collar portion  80  and the sealing rings  82 . 
         [0021]    In this example, a portion of the flange  78  is secured directly to the combustor liner  50 . Welding secures the flange  78  to the combustor liner  50  in this example. Other adhesion techniques are used in other examples. The flange  78  is also formed from a single sheet of material, which, in this example, is the same type of material used to manufacture the combustor liner  50 . 
         [0022]    Another portion of the flange  78  establishes a channel  86  that facilitates holding the sealing rings  50 . In this example, the flange  78  has a J-shaped portion  88  that establishes the channel  86 . The sealing rings  82  are not secured directly to the flange  78  in this example and are thus moveable within the channel  86 . 
         [0023]    In this example the inner collar portion  80 , the outer collar portion  81 , the outer wall  64 , and the inner wall  65  are aligned with the engine centerline X. 
         [0024]    The higher air pressure outside the combustion area  58  exerts forces F on the sealing rings  82 , which urges the sealing rings  82  against the flange  78  and the turbine nozzle  54  to seal the interface  74 . More specifically, the sealing rings  82  are urged against the flange  78  and an inner surface  83  of inner wall  65 . In one example, the inner surface  83  is machined to facilitate maintaining the seal with the sealing rings  82 . 
         [0025]    Referring to  FIG. 5 , the example sealing rings  82  have a break  90 . That is, the example sealing rings  82  are not continuous rings. As known, the interface  74  is exposed to extreme temperature variations, which can cause the sealing rings  82 , and surrounding components, to expand and contract. The break  90  accommodates movements of the sealing rings  82  as the sealing rings  82  expand and contract due to temperature fluctuations within the engine  10 . In another example, the sealing rings  82  are a continuous spiral snap ring. 
         [0026]    In this example, the break  90  of one of the sealing rings  82  is circumferentially offset from the break  90  of another of the sealing rings  82 . Offsetting the breaks in this manner prevents the break  90  from becoming a significant leakage path for air through the interface  74 . That is, area of the break  90  in one of the sealing rings  82  is sealed by another of the sealing rings  82 . 
         [0027]    Two sealing rings  82  are shown in this example. Other examples include using more or fewer sealing rings  82 . Five sealing rings  82  may be arranged together, for example. The radially outer and radially inner faces  94  of the example sealing rings  82  are rounded. In this example, the radially outer face facilitates sealing the sealing rings  82  against the turbine nozzle  54 . Pointed faces or flattened faces are used in other examples. 
         [0028]    In one example, an axially directed face  98  of the sealing rings  82  includes features such as grooves or ribs that limit rotation of the sealing rings  82  relative to each other. 
         [0029]    The example sealing rings  82  are made of a carbon-based material. Other examples include sealing rings  82  made of other materials. The example sealing rings  82  have a radial thickness Tr of about 0.25 inches (0.6 cm) and an axial thickness Ta of about 0.08 inches (0.2 cm). The diameter of the example sealing rings  82  is about 12 inches (30.5 cm). 
         [0030]    Features of the disclosed examples include using a sealing ring to seal an interface between a combustor and a turbine nozzle. Using the sealing ring facilitates assembly of the interface between the combustor and the turbine nozzle because the sealing ring can be moved relative to the combustor liner. If leaks are found when using the sealing ring, the leaks are typically more predictable and uniform than leaks at interfaces in the prior art designs. Controlled leakage amounts can also be created by the sealing rings. Another feature of disclosed examples includes using breaks in the sealing rings to accommodate expansions and contractions. 
         [0031]    Although a preferred embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.