Abstract:
A connector between a first object having a first coefficient of thermal expansion and a second object having a second coefficient of thermal expansion has a slot attaching to the second object. The slot formed by an inner shell having a slot shape and a first portion that blends into an outer surface of the second object, the slot having an enclosed portion facing forward and an aft portion that is open.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 61/707,579, which was filed 28 Sep. 2012 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     Turbofans are a type of gas turbine engine commonly used in aircraft, such as jets. The turbofan generally includes high and low pressure compressors, high and low pressure turbines, high and low speed spool shafts, a fan, and a combustor. The high-pressure compressor (HPC) is connected to the high-pressure turbine (HPT) by the high speed spool rotatable shaft, and together act as a high-pressure system. Likewise, the low-pressure compressor (LPC) is connected to the low-pressure turbine (LPT) by the low speed spool rotatable shaft, and together act as a low-pressure system. The low speed spool shaft is housed within the high speed spool shaft and is connected to the fan such that the HPC, HPT, LPC, LPT, and high and low spool shafts are coaxially aligned. 
     Air is drawn into the gas turbine engine by the fan and/or the LPC. The HPC further increases the pressure of the air drawn into the system. The high-pressure air then enters the combustor, which burns fuel and emits exhaust gas. The exhaust gas flows from the combustor into the HPT where it rotates the high spool shaft to drive the HPC. After the HPT, the exhaust gas is exhausted to the LPT. The LPT uses the exhaust gas to turn the low spool shaft, which powers the LPC and the fan to continually bring air into the system. Air brought in by the fan bypasses the LPC and HPC, and acts to increase the engine&#39;s thrust, driving the jet forward. 
     In order to support the high and low pressure systems, bearings are located within the gas turbine engine to help distribute the load created by the high and low pressure systems. The bearings are connected to an engine casing that houses a mid-turbine frame located between the HPT and the LPT by bearing support structures. The bearing support structures can be, for example, bearing cones or struts. The load from the bearing support structures are transferred to the engine casing through the mid-turbine frame. 
     SUMMARY 
     According to a non-limiting embodiment shown herein, a connector between a first object having a first coefficient of thermal expansion and a second object having a second coefficient of thermal expansion has a slot attaching to the second object, the slot formed by an inner shell having a slot shape and a first portion that blends into an outer surface of the second object, the slot having an enclosed portion facing forward and an aft portion that is open. 
     According to any previous claim provided herein, the first portion blends into the outer surface within the slot. 
     According to any previous claim provided herein, a pin extends from the first object, the pin is inserted in the slot wherein the pin minimizes axial and circumferential movement between the first object and the second object while permitting radial movement therebetween during changes in temperature. 
     According to any previous claim provided herein, an outer shell encloses the inner shell. 
     According to any previous claim provided herein, the outer shell has a second portion blending into the second object. 
     According to any previous claim provided herein, a second portion blends away from the slot, the second portion diverging from the first portion. 
     According to any previous claim provided herein, a filler is disposed between the inner shell and the outer shell. 
     According to any previous claim provided herein, the inner shell and the outer shell have a u-shape and a nadir of the u-shape is vertically oriented. 
     According to any previous claim provided herein, the first shell and the second shell are formed of a CMC material. 
     According to any previous claim provided herein, at least one of the first shell or the second shall have a warp that is oriented in parallel to an axis passing through the second object. 
     According to any previous claim provided herein, the first shell is formed of a CMC material. 
     According to any previous claim provided herein, the first shell has a warp that is oriented in parallel to an axis passing through the second object. 
     According to further non-limiting embodiment disclosed herein a connector between a turbine frame and an engine casing has a slot attaching to the turbine frame, the slot formed by an inner shell having a slot shape and a first portion that blends into an outer surface of the turbine frame, the slot having an enclosed portion facing forward and an aft portion that is open. 
     According to any previous claim provided herein, the first portion blends into the outer surface within the slot. 
     According to any previous claim provided herein, a pin extends from the engine casing, the pin is inserted in the slot. 
     According to any previous claim provided herein, an outer shell encloses the inner shell. 
     According to any previous claim provided herein, the outer shell has a second portion blending into the turbine frame. 
     According to any previous claim provided herein, the second portion blends away from the slot, the second portion diverging from the first portion. 
     According to any previous claim provided herein, a filler is disposed between the inner shell and the outer shell. 
     According to any previous claim provided herein, the inner shell and the outer shell have a u-shape and a nadir of the u-shape is vertically oriented. 
     According to any previous claim provided herein, the first shell and the second shell are formed of a CMC. 
     According to any previous claim provided herein, the first shell is formed of a CMC material. 
     According to any previous claim provided herein, the turbine frame is a mid-turbine frame. 
     According to any previous claim provided herein, a pin is trapped between a forward end of the slot and the turbine frame. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic, partially cut-away view of a gas turbine engine incorporating a non-limiting embodiment described herein. 
         FIG. 2  is a schematic view of a mid-turbine frame and connector of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a perspective view of the connector of  FIG. 2 . 
         FIG. 4  is a cut-away view of the connector of  FIG. 3 . 
     
    
    
     DESCRIPTION 
       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. 
     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 or geared turbofan architectures. 
     The fan section  22  drives air along a bypass flowpath B 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 . 
     The gas turbine 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. 
     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 a 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  58  (“MTF”) of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The MTF  58  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. 
     The core airflow C 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 MTF  58  includes airfoils  60  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. 
     The gas turbine engine  20  is in one example a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is greater than about six (6:1) with an example embodiment being greater than ten (10:1). 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 (2.3:1). The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). The 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. 
     In one disclosed embodiment, the gas turbine engine  20  bypass ratio is greater than about ten (10:1), and the fan diameter is significantly larger than that of the low pressure compressor  44 . The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). 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 (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. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the gas turbine 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 feet, with the engine at its best fuel consumption, also known as bucket cruise Thrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standard parameter of lbm of fuel being burned divided by lbf 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 feet per second 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 feet per second (351 meters per second). 
     Referring now to  FIG. 2 , the MTF  58  includes a first duct  100 , a second duct  105 , and the airfoil  60  (used herein as a vane) connecting the first duct  100  to the second duct  105 . The first and second ducts may be made of a ceramic material such as a ceramic matrix composite (“CMC”) and the like, and the second duct  105  is radially inboard and passes through the first duct  100 . CMC material is used in the MTF because of the intense heat of the gasses passing from the combustor  56  and the high pressure turbine  54  through the MTF  58 . 
     The MTF  58  is supported and surrounded by an outer case  115 , which may be metallic, a pin  120  extending from the outer case  115  and a connector  125  attaching to the first duct  100 , as will be discussed infra. The connector  125  is also made of a ceramic material such as CMC and the like. The pin  120 , which must withstand high axial and circumferential loads caused by the force of the gas turbine engine  20  is not a CMC and may be metallic, such as a nickel alloy. The pin  120  is conventionally attached to the outer case  115  such as by gluing, mechanical means, welding or brazing. 
     Referring now to  FIGS. 2-4 , details of the connector  125  are shown. The connector  125  is made of the same material as the MTF in the non-limiting example shown, though other compatible, high-temperature resistant materials may be used. The material may be CMC fabric or unitape that is located upon an outer surface  130  the first duct  100 . The fabric or unitape has strands forming warp  131  and weft  132  that are organized at 90° angles relative to each other. Alternate example CMC fabrics can be organized at alternate angles and achieve the same affect. CMC performs well in tension and the connector  125  is organized to engage the pin  120  such that the connector  125  is maximally placed in tension against the pin  120  such that the strands of the warp  131  are in tension. 
     The connector  125  has an inner shell  135  that has a u-shape and an outer shell  140  that also has a u-shape. The inner shell  135  and the outer shell  140  are in register with each other so that a radial outer surface  145  of the inner shell  135  is in close proximity to a radial outer surface  150  of the outer shell  140 . The outer shell  140  encloses the inner shell circumferentially about the first duct  100 . 
     The first duct  100  angles radially outwardly moving axially aft along the outer surface  130 . As seen in  FIG. 2 , it appears that the top  155  of the inner shell  135  and the top  160  of the outer shell  140  extend horizontally until they reach the outer surface  130  of the first duct  100 . At a nadir  165  of the u-shape of the inner shell  135  and a nadir  170  of the u-shape of the outer shell  140 , the inner and outer shells  135 ,  140  are disposed vertically relative to the first duct  100 . Near the bottom of the inner shell  135  and the outer shell  140 , the inner shell  135  and the outer shell  140  flare outwardly to blend with the outer surface  130 . An inner blending surface  175  flares from the inner shell  135  away from the outer shell  140  to meld into the inner surface within a slot  180  within the u-shaped inner shell  135 . The slot has an open end  185  that opens axially aft. An outer blending surface  190  flares from the outer shell  140  away from the inner shell  135  to meld into the outer surface outside of the slot  180 . The inner blending surface  175  tends to appear to close the u-shaped inner shell  135  as viewed from a top  195  (see  FIGS. 2 and 3 ) and the outer blending surface  190  tends to make the outer shell appear to have a bell-shaped surface viewed from the top  195 . The blending surfaces  190  start at a point  197  below a midpoint  200  of the inner and outer shells  135 ,  140  between the nadirs  165 ,  170  thereof and the first duct  100 . The points  197  form a line that is in parallel to the outer surface  130  of the first duct  100 . The concept provided for herein will work whether the first duct  100  is angled or not provided the pin  120  is oriented radially from the first duct  100 . 
     The pin  120  is parallel and in contact with the inner shell  135  and the outer shell  140  at the nadirs  165 ,  170  thereof. The pin  120  has no sharp edges that might tend to shear the connector  125 . The pin  120  and the connector  125  move radially relatively to each other to allow for different coefficients of expansion due to heat, but the pin  120  minimizes axial and circumferential movement of the connector  125  and the MTF  58  due to the connection and the CMC material of the connector  125 . The pin  120  is trapped in the slot  180  by the nadir  165  and the inner shell  135  (see  FIG. 2 ). 
     The area  202  between the shells may be filled with non-compressible filler, such as resin, fiber, cloth, tape, or the like, as the inner and outer shells bear the axial and circumferential load of the gas turbine engine  20 . The inner and outer blending surfaces  175  and  190  create a large footprint on the outer surface  130  as the inner and outer blending surfaces separate from each other as they approach the outer surface  130 . The large footprint minimizes bending of the inner and outer shells  135 ,  140  as axial loads of the gas turbine engine are transferred therefrom to the first duct  100  thereby minimizing a probability that the connector  125  will fail by crushing, bending, shearing or the like. 
     The connector warps  131  are in tension axially, a favorable direction in which the CMC is strong, as opposed to putting the warp  131  in bending or shear. The pin  120  is trapped in the slot  180  of the connector  125 , which is in tension along the warp  131 , and there is little shearing and fewer stress concentrations. 
     The connector  125  may be pre-formed and then glued to the outer surface  130  or may be formed in one-piece with the first duct  100 . The MTF  58  is annular and there as many connectors  125  and pins  120  as necessary to manage the axial load of the gas turbine engine  20 . 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.