Abstract:
A fan case for a gas turbine engine includes a fan case surrounding a fan with fan blades. A liner is disposed between the fan case and the fan and is spaced a radial distance from the fan case. A torque stop is arranged between the fan case and the liner. A method for reducing fan case liner loads is also disclosed.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a gas turbine engine component, such as a liner for a fan case. 
         [0002]    Gas turbine engines typically include a fan section, a compressor section, a combustor section and a turbine section. The fan section may be housed in a fan case. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
         [0003]    A thermally conformable liner (TCL) may be present on the interior of the fan case. The TCL facilitates consistent thermal growth between the fan blades and rubstrips located on the TCL, allowing the tip clearance gap to remain constant with temperature. Currently, the TCL may be arranged directly adjacent to the fan case. During fan blade out (FBO) events, which may occur as a result of a bird strike, for example, a fan blade may become disconnected from the fan. A fan blade tip can dig into the rubstrips. This interaction generates a deceleration torque as well as axial loads. These FBO-induced loads tend to be very high, for example, several times the operating loads of the TCL. 
       SUMMARY 
       [0004]    In one exemplary embodiment, a fan section for a gas turbine engine includes a fan case surrounding a fan with fan blades. A liner is disposed between the fan case and the fan and is spaced a radial distance from the fan case. At least one torque stop is arranged between the fan case and the liner. 
         [0005]    In a further embodiment of the above, the ratio of the radial distance to a radius of the fan blades is 0.25:40. 
         [0006]    In a further embodiment of any of the above, at least one of a radially inner surface of the fan case and a radially outer surface of the liner includes a low-friction coating. 
         [0007]    In a further embodiment of any of the above, the low-friction coating is a Teflon® spray. 
         [0008]    In a further embodiment of any of the above, at least one torque stop is frangible. 
         [0009]    In a further embodiment of any of the above, at least one torque stop prevents rotation of the liner relative to the fan case. 
         [0010]    In a further embodiment of any of the above, at least a portion of the liner is bonded to the fan case with an adhesive. 
         [0011]    In a further embodiment of any of the above, at least one torque stop and the adhesive can withstand a tangential load given by the equation (1+S)*T Rub /R FanCase /N TorqueStops , where S is a safety factor. T Rub  is a rub torque generated by a 2.5 lb (1.3 kg) bird strike. R FanCase  is a radius of the fan case. N TorqueStops  is the number of torque stops. 
         [0012]    In a further embodiment of any of the above, S is 0.35. 
         [0013]    In a further embodiment of any of the above, the liner includes one or more rails which contact the fan case. 
         [0014]    In a further embodiment of any of the above, the liner includes an abradable rubstrip adjacent to tips of the fan blades. 
         [0015]    In a further embodiment of any of the above, the abradable rubstrip is arranged radially inward from at least one torque stop. 
         [0016]    In a further embodiment of any of the above, a honeycomb structure is arranged between the t least one torque stop and the abradable rubstrip. 
         [0017]    In a further embodiment of any of the above, an aluminum septum is arranged between the one torque stop and the abradable rubstrip. 
         [0018]    In a further embodiment of any of the above, a Kevlar® layer is arranged between the one torque stop and the fan case. 
         [0019]    In one exemplary embodiment, a method for reducing fan case liner loads includes the steps of providing a fan case surrounding a fan with fan blades, providing a liner disposed between the fan case and the fan and spaced radially apart from the fan case, providing at least one frangible stop arranged between the fan case and the liner and providing an anti-friction coating on one of the radially inner surface of the fan case and the radially outward surface of the liner. 
         [0020]    In one exemplary embodiment, the anti-friction coating is a Teflon® spray. 
         [0021]    In one exemplary embodiment, at least one stop prevents rotation of the liner relative to the fan case. 
         [0022]    In one exemplary embodiment, a portion of the liner is bonded to the fan case with an adhesive. 
         [0023]    In one exemplary embodiment, the liner includes one or more rails which contact the fan case. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0025]      FIG. 1  schematically illustrates an example gas turbine engine embodiment. 
           [0026]      FIG. 2  illustrates a schematic detail view of the prior art gas turbine engine fan case. 
           [0027]      FIG. 3  illustrates a schematic detail view of the inventive gas turbine engine fan case. 
           [0028]      FIG. 4  illustrates a schematic view of a thermally conformable liner. 
           [0029]      FIG. 5  illustrates an alternate schematic view of the thermally conformable liner of  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  is arranged in a fan case  23 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0031]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0032]    The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis X 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. 
         [0033]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The fan  42  includes fan blades with tips  43 . The inner shaft  40  drives the fan  42  through a speed change device, such as 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 (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis X. 
         [0034]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0035]    The example low pressure turbine  46  has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0036]    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  as well as setting airflow entering the low pressure turbine  46 . 
         [0037]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  59  of the mid-turbine frame  57  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  57 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0038]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0039]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0040]    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 (&#39;TSFC&#39;)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0041]    “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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
         [0042]    “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. 
         [0043]      FIG. 2  shows a schematic fan case  23  as is known in the prior art. A liner assembly  96  is directly adjacent to the fan case  23  and includes a rubstrip. Velocity ω is provided by the fan (not shown) in the forward direction. During a fan blade out (FBO) event, fan blade tips (not shown) dig into the rubstrip on the liner assembly  96 , creating reactive rub torque forces T 1  and T 2 . These torque forces T 1  and T 2  can be very high. 
         [0044]    Referring to  FIGS. 3 and 4  schematically depicting the disclosed embodiment, the fan case  23  includes a liner assembly  96  with a thermally conformable liner (TCL)  100 . The TCL  100  may reduce rub torque forces translated to the fan case  23  in an FBO event. The TCL  100  may be cantilevered and “float” such that there is a gap  103  between the TCL  100  and the fan case  23 . The TCL may include a rub strip  98 , one or more torque stops  104 , an aluminum septum  106 , a honeycomb  108 , and an aluminum liner  110 . The rub strip  98  may include an abradable material. The aluminum liner  110  and aluminum septum  106  may be 0.04 inches (1.02 mm) thick. A high strength material  112  such as Kevlar® may also be present adjacent to the fan case  23  and may be bonded to the fan case  23  with an adhesive. An acoustic liner  114   a  may be radially inward of the aluminum liner  112 . An acoustic liner  114   b  may also be bonded to the fan case  23  with an adhesive. An aluminum support shell  116  may be radially inward of the acoustic liner  114   a.    
         [0045]    Referring to  FIG. 5 , rails  118  may be glued to the aluminum liner  110 . The rails  118  may include blocks  120 , which may be glued to the fan case  23  with an adhesive. 
         [0046]    Certain features including the size of the gap  103 , the application of a coating or surface treatment to reduce friction in the gap  103 , the material properties of the stop  104 , and the flexibility of the TCL  100  can be analytically determined for each engine application to significantly reduce the FBO loads transmitted to the fan case  23 . In one example, the ratio of the gap  103  width to the fan  42  blade radius is 0.25:40. In another example, the fan case  23  or the TCL  100  may have a surface coating or treatment to reduce surface-to-surface friction in the gap  103 , for example, a Teflon® coating on one or both of the fan case  23  radially inner surface and the TCL  100  radially outer surface. 
         [0047]    In a further example, the one or more torque stops  104  may be frangible, anti-rotation stops, which may reduce loads transmitted between the fan blade tip  43  and the rubstrip  98 . The TCL  100  may be connected to the fan case  23  by an adhesive as was described previously. In one example, the stop  104  and the adhesive may be designed to withstand a tangential load given by the equation (1+S)*T Rub /R FanCase /N TorqueStops  where S is a safety factor to ensure that a small bird strike does not damage the TCL  100 , T Rub  is the rub torque generated by a 2.5 lb (1.3 kg) bird strike, R FanCase  is the radius of the fan case, and N TorqueStops  is the number of torque stops  104 . T Rub  is variable and may depend on blade flexibility, material composition, and other factors. In one example, S is 0.35. 
         [0048]    As a result of lower fan case  23  rub torque peaks, other hardware in the load path, including mounts, can also be decreased by about 5-20%. 
         [0049]    Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that and other reasons, the following claims should be studied to determine their true scope and content.