Abstract:
A gas turbine engine has a first component and a second component. The first and second components have a high-pressure chamber on one side and a low pressure chamber on an opposed side. A three sided seal has one side facing each of the first and second components, and a third side facing a third component. At least one non-metallic wear surface is between one of the three sides of the seal and the facing component.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a continuation of U.S. patent application Ser. No. 13/688,340, filed Nov. 29, 2012. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0002]    This invention was made with government support under Contract No. N00019-02-C-3003 awarded by the United States Navy. The Government has certain rights in this invention. 
     
    
     BACKGROUND 
       [0003]    This application relates to a pressure seal having three surfaces in sliding contact, with at least one of the surfaces being provided with a non-metallic wear surface. 
         [0004]    Gas turbine engines are known, and typically include a fan delivering air into a compressor. The air is compressed in the compressor, and delivered into a combustion section. The air is mixed with fuel in the combustion section and ignited. Products of the combustion pass downstream over turbine rotors, driving them to rotate. 
         [0005]    There are a number of areas within a gas turbine engine where a high pressure chamber must be maintained separate from a low pressure chamber. Thus, various locations require a reliable seal. 
         [0006]    One type of seal is a U or J-shaped seal having three sides. Two components contact two of the sides of the seal, and some form of retention member may contact a third side. The seal is generally moveable along the three surfaces. 
         [0007]    In the past, seals in this particular application have always been formed of metallic materials. This has led to undue wear. 
         [0008]    While non-metallic seals have been proposed for many applications, they have not been proposed in gas turbine engines between high and low pressure chambers, where a three sided sealing application is used. 
       SUMMARY 
       [0009]    In a featured embodiment, a gas turbine engine has a first component and a second component. The first and second components have a high-pressure chamber on one side and a low pressure chamber on an opposed side. A three sided seal has one side facing each of the first and second components, and a third side facing a third component. At least one non-metallic wear surface is between one of the three sides of the seal and the facing component. 
         [0010]    In another embodiment according to the previous embodiment, there are non-metallic wear surfaces between each of the three sides of the seal and the facing component. 
         [0011]    In another embodiment according to any of the previous embodiments, one of the components is a liner segment of an exhaust system in the gas turbine engine. 
         [0012]    In another embodiment according to any of the previous embodiments, another the component is a moving liner segment. 
         [0013]    In another embodiment according to any of the previous embodiments, the third component is a retention element for positioning the seal. 
         [0014]    In another embodiment according to any of the previous embodiments, one of the components is a case segment for the gas turbine engine. 
         [0015]    In another embodiment according to any of the previous embodiments, one of the components is a duct segment for the gas turbine engine. 
         [0016]    In another embodiment according to any of the previous embodiments, the wear surface is on the seal. 
         [0017]    In another embodiment according to any of the previous embodiments, the wear surface is on at least one of the components. 
         [0018]    In another embodiment according to any of the previous embodiments, the third side may sometimes be in contact with the third component, and sometimes be spaced from the third component during different operational points in the operation of the gas turbine engine. 
         [0019]    In another embodiment according to any of the previous embodiments, the retention segment has a pair of retention elements on opposed position faces of one of the three sides of the seal. 
         [0020]    In another embodiment according to any of the previous embodiments, the three sided seal floats relative to each of the first, second and third components such that it is movable relative to each of the components. 
         [0021]    These and other features of this application will be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]      FIG. 1  schematically shows a gas turbine engine. 
           [0023]      FIG. 2  shows a location in the  FIG. 1  gas turbine engine receiving a seal. 
           [0024]      FIG. 3A  shows a first seal embodiment. 
           [0025]      FIG. 3B  shows another seal embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0026]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flowpath C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine 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 engines including three-spool architectures. 
         [0027]    The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0028]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0029]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0030]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0031]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm (pounds mass) of fuel being burned divided by lbf (pounds force) of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
         [0032]      FIG. 2  shows a location within a gas turbine engine, which may be in a nozzle portion of the gas turbine engine, downstream of the turbine section  28 . A high pressure area  70  is separated from a lower pressure area  72  by a seal  74 . Seal  74  has three sides  77 ,  79  and  81 , all of which may be in contact with a seal surface on an associated component  76 ,  78  and  80 . As shown, the sides  77 ,  79  and  81  may be also out of contact with a component  77 ,  78  and  80  but each face a component. 
         [0033]    Component  78  may be a static liner segment of an exhaust system, and has a sealing surface  82  positioned to be in contact with the surface  79  on the seal  74 . A pair of retention elements  76  face, and may contact, the surface  77  on the seal  74 . As shown in  FIG. 2 , the retention elements  76  may also be spaced away from the surface  77  on the seal  74  under much of its operational life. Another component  80  may be a moving liner segment, and have an end surface in contact with a surface  81  on the seal  74 . 
         [0034]    In other embodiments, one of the components may be a case segment for a gas turbine engine, or a duct segment for a gas turbine. 
         [0035]      FIG. 3A  is an enlarged view of the seal embodiment. 
         [0036]    As shown in  FIG. 3A , the seal floats between the three surfaces, and as it moves there is wear. Thus, the provision of a non-metallic wear surface  82  between the surfaces  78  and  79  provides longer life for the seal and the components. 
         [0037]    In embodiments, wear surfaces may be applied on each of the components  76 ,  80 , and  78 . The wear surfaces may be polytetraflouroethylene (PTFE) or thermoplastic polymer or other appropriate non-metallic materials. The non-metallic surfaces may be formed of carbon, silicon, ceramic or other composite-based materials and may be selected due to wear and lubrication characteristics. The materials may be applied as pads, segmented strips, or continuous strips. 
         [0038]      FIG. 3B  shows another embodiment  174 . In embodiment  174 , the three sides  176 ,  178  and  180  of the seal receive non-metallic pads  177 ,  179  and  181 , respectively. 
         [0039]    The pads or other non-metallic materials may be mechanically secured to the seal or other components via rivets, fasteners, and etcetera. 
         [0040]    Further, the entire seal may be formed of non-metallic materials, or alternatively, the members  76 ,  78  and  80  may be formed of non-metallic materials. That is, there need not be a non-metallic material secured to the underlying substrate, but rather the substrate itself may be formed of the non-metallic materials. 
         [0041]    Although embodiments of this invention have 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.