Patent Publication Number: US-2016245184-A1

Title: Geared turbine engine

Description:
BACKGROUND OF THE INVENTION 
     1. Technical Field 
     This disclosure relates generally to a geared turbine engine. 
     2. Background Information 
     Various types of turbine engines for propelling an aircraft are known in the art. One exemplary turbine engine type is a geared turbofan turbine engine. A typical geared turbofan turbine engine includes a gear train, a fan rotor and a core. Typically, the core consists essentially of a low speed spool and a high speed spool. The gear train connects the fan rotor to the low speed spool and enables the low speed spool to drive the fan rotor at a slower rotational velocity than that of the low speed spool. Another example of a geared turbofan turbine engine is disclosed in U.S. Pat. No. 8,869,504 to Schwarz et al., which is hereby incorporated herein by reference in its entirety. While such turbine engines have various advantages, there is still a need in the art for improvement. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the disclosed invention, a turbine engine is provided that includes a fan rotor, a first compressor rotor, a second compressor rotor, a third compressor rotor, a first turbine rotor, a second turbine rotor, a third turbine rotor and a gear train. The fan rotor and the first compressor rotor are connected to the first turbine rotor through the gear train. The second compressor rotor is connected to the second turbine rotor. The third compressor rotor is connected to the third turbine rotor. 
     According to another aspect of the disclosed invention, another turbine engine is provided that includes a first rotating assembly, a second rotating assembly and a third rotating assembly. The first rotating assembly includes a fan rotor, a first compressor rotor, a first turbine rotor and a gear train. The second rotating assembly includes a second compressor rotor and a second turbine rotor. The third rotating assembly includes a third compressor rotor and a third turbine rotor. 
     According to still another aspect of the disclosed invention, a method for manufacturing is provided. This method includes steps of manufacturing a first turbine engine configured for a first thrust rating and manufacturing a second turbine engine configured for a second thrust rating which is different than the first thrust rating. The first turbine engine includes a rotating assembly and a first multi-spool core, where the rotating assembly includes a fan rotor, a compressor rotor, a turbine rotor and a gear train. The second turbine engine includes a second multi-spool core. An upstream-most set of compressor blades of the first multi-spool core defines a first area. An upstream-most set of compressor blades of the second multi-spool core defines a second area that is within plus/minus twenty percent of the first area. The first and the second turbine engines are manufactured by and/or for a common entity. 
     A first shaft may be included and connect the gear train to the first turbine rotor. A second shaft may be included and connect the second compressor rotor to the second turbine rotor. A third shaft may be included and connect the third compressor rotor to the third turbine rotor. 
     The second shaft may extend through the third shaft. The first shaft may extend through the second shaft. 
     The first compressor rotor may be connected to the gear train through the fan rotor. 
     The first compressor rotor may be connected to the gear train independent of the fan rotor. 
     The first compressor rotor may include a first set of compressor blades and a second set of compressor blades downstream of the first set of compressor blades. 
     The fan rotor may include one or more variable pitch fan blades. 
     The variable pitch fan blades may be configured to move between a first position and a second position. The fan rotor may be operable to provide forward thrust where the variable pitch fan blades are in the first position. The fan rotor may be operable to provide reverse thrust where the variable pitch fan blades are in the second position. 
     A nacelle may be included housing the fan rotor. The nacelle may include a translating sleeve configured to open a passageway through the nacelle where the variable pitch fan blades are in the second position. The translating sleeve may be configured to close the passageway where the variable pitch fan blades are in the first position. 
     A leading edge of a first of the variable pitch fan blades may move in a forward direction as that blade moves from the first position to the second position. 
     The gear train may connect the fan rotor and the first compressor rotor to the first turbine rotor. 
     The first compressor rotor may be connected to the gear train through the fan rotor. 
     The fan rotor and the first compressor rotor may be connected to the gear train in parallel. 
     The gear train may connect the fan rotor to a shaft. The shaft may connect the gear train and the first compressor rotor to the first turbine rotor. 
     The first compressor rotor may consist essentially of (only include) a rotor disk and a set of compressor blades arranged around and connected to the rotor disk. 
     The first compressor rotor may include a rotor disk, a first set of compressor blades and a second set of compressor blades. The first set of compressor blades may be arranged around and connected to the rotor disk. The second set of compressor blades may be arranged around and connected to the rotor disk downstream of the first set of compressor blades. 
     More than fifty percent of components included in the first turbine engine may be configured substantially similar to corresponding components in the second turbine engine. 
     The foregoing features and the operation of the invention will become more apparent in light of the following description and the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a partial sectional illustration of an example geared turbofan turbine engine. 
         FIG. 2  is a partial schematic illustration of the turbine engine of  FIG. 1 . 
         FIG. 3  is a partial schematic illustration of an example of a low pressure compressor section. 
         FIG. 4  is a partial schematic illustration of another example geared turbofan turbine engine. 
         FIG. 5  is a partial schematic illustration of another example geared turbofan turbine engine. 
         FIG. 6  is a partial schematic illustration of another example geared turbofan turbine engine providing forward thrust. 
         FIG. 7  is a partial schematic illustration of the turbine engine of  FIG. 6  providing reverse thrust. 
         FIG. 8  is a partial schematic illustration of another example geared turbofan turbine engine providing forward thrust. 
         FIG. 9  is a partial schematic illustration of the turbine engine of  FIG. 8  providing reverse thrust. 
         FIG. 10  is a flow diagram of a method for manufacturing a plurality of turbine engines. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a partial sectional illustration of a geared turbofan turbine engine  20 . The turbine engine  20  extends along an axial centerline  22  between an upstream airflow inlet  24  and a downstream airflow exhaust  26 . The turbine engine  20  includes a fan section  28 , a compressor section, a combustor section  32  and a turbine section. The compressor section includes a low pressure compressor (LPC) section  29 , an intermediate pressure compressor (IPC) section  30  and a high pressure compressor (HPC) section  31 . The turbine section includes a high pressure turbine (HPT) section  33 , an intermediate pressure turbine (IPT) section  34  and a low pressure turbine (LPT) section  35 . 
     The engine sections  28 - 35  are arranged sequentially along the centerline  22  within an engine housing  36 . This housing  36  includes an inner (e.g., core) casing  38  and an outer (e.g., fan) casing  40 . The inner casing  38  houses the LPC section  29  and the engine sections  30 - 35 , which form a multi-spool core of the turbine engine  20 . The outer casing  40  houses at least the fan section  28 . The engine housing  36  also includes an inner (e.g., core) nacelle  42  and an outer (e.g., fan) nacelle  44 . The inner nacelle  42  houses and provides an aerodynamic cover for the inner casing  38 . The outer nacelle  44  houses and provides an aerodynamic cover the outer casing  40 . The outer nacelle  44  also overlaps a portion of the inner nacelle  42  thereby defining a bypass gas path  46  radially between the nacelles  42  and  44 . The bypass gas path  46 , of course, may also be partially defined by the outer casing  40  and/or other components of the turbine engine  20 . 
     Each of the engine sections  28 - 31  and  33 - 35  includes a respective rotor  48 - 54 . Each of these rotors  48 - 54  includes a plurality of rotor blades (e.g., fan blades, compressor blades or turbine blades) arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s). 
     Referring to  FIG. 2 , the rotors  48 - 54  are respectively configured into a plurality of rotating assemblies  56 - 58 . The first rotating assembly  56  includes the fan rotor  48 , the LPC rotor  49  and the LPT rotor  54 . The first rotating assembly  56  also includes a gear train  60  and one or more shafts  62  and  63 , which gear train  60  may be configured as an epicyclic gear train with a planetary or star gear system. The LPC rotor  49  is connected to the fan rotor  48 . The fan rotor  48  is connected to the gear train  60  through the fan shaft  62 . The LPC rotor  49  is therefore connected to the gear train  60  through the fan rotor  48  and the fan shaft  62 . The gear train  60  is connected to and driven by the LPT rotor  54  through the low speed shaft  63 . 
     The second rotating assembly  57  includes the IPC rotor  50  and the IPT rotor  53 . The second rotating assembly  57  also includes an intermediate speed shaft  64 . The IPC rotor  50  is connected to and driven by the IPT rotor  53  through the intermediate speed shaft  64 . 
     The third rotating assembly  58  includes the HPC rotor  51  and the HPT rotor  52 . The third rotating assembly  58  also includes a high speed shaft  65 . The HPC rotor  51  is connected to and driven by the HPT rotor  52  through the high speed shaft  65 . 
     Referring to  FIG. 1 , one or more of the shafts  62 - 65  may be coaxial about the centerline  22 . One or more of the shafts  63 - 65  may also be concentrically arranged. The low speed shaft  63  is disposed radially within and extends axially through the intermediate speed shaft  64 . The intermediate speed shaft  64  is disposed radially within and extends axially through the high speed shaft  65 . The shafts  62 - 65  are rotatably supported by a plurality of bearings; e.g., rolling element and/or thrust bearings. Each of these bearings is connected to the engine housing  36  (e.g., the inner casing  38 ) by at least one stationary structure such as, for example, an annular support strut. 
     During operation, air enters the turbine engine  20  through the airflow inlet  24 . This air is directed through the fan section  28  and into a core gas path  66  and the bypass gas path  46 . The core gas path  66  flows sequentially through the engine sections  29 - 35 . The air within the core gas path  66  may be referred to as “core air”. The air within the bypass gas path  46  may be referred to as “bypass air”. 
     The core air is compressed by the compressor rotors  49 - 51  and directed into a combustion chamber  68  of a combustor  70  in the combustor section  32 . Fuel is injected into the combustion chamber  68  and mixed with the compressed core air to provide a fuel-air mixture. This fuel air mixture is ignited and combustion products thereof flow through and sequentially cause the turbine rotors  52 - 54  to rotate. The rotation of the turbine rotors  52 - 54  respectively drive rotation of the compressor rotors  51 - 49  and, thus, compression of the air received from a core airflow inlet  72 . The rotation of the turbine rotor  54  also drives rotation of the fan rotor  48 , which propels bypass air through and out of the bypass gas path  46 . The propulsion of the bypass air may account for a majority of thrust generated by the turbine engine  20 , e.g., more than seventy-five percent (75%) of engine thrust. The turbine engine  20  of the present disclosure, however, is not limited to the foregoing exemplary thrust ratio. 
     The exemplary LPC rotor  49  of  FIG. 1  includes a rotor disk  74  and one set of compressor blades  76 . These compressor blades  76  are arranged around and connected to the rotor disk  74  as described above. The compressor blades  76  are adjacent to and downstream of a set of stator vanes  78 . These stator vanes  78  may be positioned generally at the inlet  72  to the core gas path  66 . In other embodiments, however, the stator vanes  78  may be positioned downstream of the compressor blades  76 . In still other embodiments, an additional set of stator vanes may be positioned downstream and adjacent the compressor blades  76 . 
     While the LPC rotor  49  is described above as including a single set of compressor blades  76 , the turbine engine  20  of the present disclosure is not limited to such a configuration. For example, referring to  FIG. 3 , the LPC rotor  49  may alternatively include two sets of compressor blades  76  and  80  disposed at different axial locations along the rotor disk  74 . The first set of compressor blades  76 , for example, is positioned adjacent and downstream of the stator vanes  78 . The second set of compressor blades  80  is positioned downstream of the first set of compressor blades  76 . The two sets of compressors blades  76  and  80  may be separated by another set of stator vanes  82  so as to provide the LPC section  29  with two stages. Of course, in other embodiments, the LPC rotor  49  may include more than two sets of compressor blades and provide the LPC section  29  with more than two stages. 
     Referring to  FIG. 4 , in some embodiments, the LPC rotor  49  may be connected to the fan shaft  62  and the gear train  60  independent of the fan rotor  48 . The LPC rotor  49  and the fan rotor  48 , for example, may be connected to the fan shaft  62  and, thus, the gear train  60  in parallel. 
     Referring to  FIG. 5 , in some embodiments, the LPC rotor  49  may be connected directly to the low speed shaft  63  and, thus, independent of the gear train  60 . With this configuration, the LPC rotor  49  and the LPT rotor  54  rotate at the same rotational velocity. In contrast, the LPC rotor  49  of  FIGS. 1, 2 and 4  rotates at a slower rotational velocity than the LPT rotor  54  due to reduction gearing of the gear train  60 . 
     In some embodiments, the fan blades  84  may be configured as fixed blades and fixedly connected to the fan rotor  48  as illustrated in  FIG. 5 . In other embodiments, referring to  FIG. 6 , one or more of the fan blades  84  may be configured as variable pitch fan blades and pivotally connected to a hub of the fan rotor  48 . With this configuration, a pitch of each respective fan blade  84  may be changed using an actuation system  86  within the hub of the fan rotor  48 . This actuation system  86  may be configured for limited variable pitch. Alternatively, the actuation system  86  may be configured for full variable pitch where, for example, fan blade pitch may be partially or completely reversed. 
     By reversing fan blade pitch, the fan blades  84  may be moved between a first (e.g., forward thrust) position as shown in  FIG. 6  and a second (e.g., reverse thrust) position as shown in  FIG. 7 . In the first position of  FIG. 6 , the fan blades  84  and the fan rotor  48  may be operable to provide forward thrust; e.g., push air through an exhaust  88  of the bypass gas path  46  as described above. Leading edges  90  of the fan blades  84 , for example, may be axially forward of trailing edges  92  of the fan blades  84 . In the second position of  FIG. 7 , the fan blades  84  and the fan rotor  48  may be operable to provide reverse thrust; e.g., push air through the airflow inlet  24 . The leading edges  90  of the fan blades  84 , for example, may be axially aft of the trailing edges  92  of the fan blades  84 . With such a configuration, the turbine engine  20  may be configured without a traditional thrust reverser in the outer nacelle  44 . 
     When providing reverse thrust, air may flow into the bypass gas path  46  through the exhaust  88  as shown in  FIG. 7 . Alternatively, the outer nacelle  44  may include an aft translating sleeve  94  as shown in  FIGS. 8 and 9 . When the fan blades  84  are in the second position (see  FIG. 9 ) for providing reverse thrust, the sleeve  94  may be translated aft so as to open a passageway  96  through the outer nacelle  44 . This passageway  96  may include one or more turning scoops  98  so as to assist in redirecting air into the bypass gas path  46 . These turning scoops  98  may be in the form of stationary turning vanes and/or radially deployable turning vanes. However, when the fan blades  84  are in the first position (see  FIG. 8 ) for providing forward thrust, the sleeve  94  may be translated forwards so as to close the passageway  96  and stow the turning scoops  98 . 
     As the fan blades  84  move from the first position to the second position, the leading edges  90  may turn in a forward direction. Here the forward direction is “forward” relative to rotation of the fan rotor  48 . For example, if the fan rotor  48  is turning in a clockwise direction, the leading edge  90  of each respective fan blade  84  may start moving in a clockwise direction before it reverses pitch and moves in a counter-clockwise direction. However, in other embodiments, as the fan blades  84  move from the first position to the second position, the leading edges  90  may turn in a reverse direction. 
       FIG. 10  is a flow diagram of a method  1000  for manufacturing a plurality of turbine engines. These turbine engines may be manufactured by a common entity; e.g., a manufacturer. The turbine engines may also or alternatively be manufactured for a common entity; e.g., a customer or end user. The turbine engines may still also or alternatively be manufactured generally contemporaneously, in common production run/cycle and/or during back-to-back production runs/cycles. 
     In step  1002 , a first turbine engine is manufactured. In step  1004 , a second turbine engine is manufactured. The first turbine engine and/or the second turbine engine may each have a configuration generally similar to the turbine engine  20  embodiments described above. However, the first turbine engine may be configured for a first thrust rating whereas the second turbine engine may be configured for a second thrust rating that is different (e.g., lower) than the first thrust rating. For example, the first thrust rating may be 10× whereas the second thrust rating may be 7×. The method  1000  of the present disclosure, however, is not limited to the foregoing exemplary thrust rating ratio. 
     The thrust ratings of the first and the second turbine engines may be dependent upon various parameters. These parameters may include, but are not limited to, the following:
         Geometry (e.g., shape and size) of the fan blades (e.g.,  84 );   Number of the fan blades (e.g.,  84 );   Geometry of the compressor blades (e.g.,  76 ,  80 );   Number of the compressor blades (e.g.,  76 ,  80 );   Number of stages in the LPC section (e.g.,  29 );   Configuration of the gear train (e.g.,  60 ); and   Configuration components in and operation of the core.       

     The first and the second turbine engines may each be configured for its specific thrust rating by changing one or more of the foregoing parameters. However, if the first and the second turbine engines are each configured with substantially similar cores and one or more other parameters (e.g., geometry of compressor blades and/or fan blades, number of LPC stages, gear train gearing, etc.) are changed to achieve the desired thrust ratings, then time and costs associated with engineering and/or manufacturing the first and the second turbine engines may be reduced. For example, if the first and the second turbine engines are configured with multi-spool cores having substantially similar configurations, then more than about fifty percent (50%) of the components included in the first turbine engine may be substantially similar to corresponding components in the second turbine engine. Thus, a single set of core components and/or other components may be engineered and manufactured for use in both the first and the second turbine engines. This commonality in turn may reduce research and development time and costs as well as manufacturing time and costs. 
     The phrase “substantially similar” is used herein to describe a set of components with generally identical configurations; e.g., sizes, geometries, number of rotor stages, etc. However, the components need not be completely identical. For example, in some embodiments, substantially similar components may be made of different materials and/or have different coatings. In some embodiments, substantially similar components may include different accessory mounts and/or locate accessories at different positions. In some embodiments, substantially similar components may include different cooling passages, different seals, different cooling features (e.g., turbulators or fins), etc. 
     In some embodiments, the first and the second turbine engines may include substantially similar multi-spool cores as described above. For example, the inner case of the first turbine engine and the inner case of the second turbine engine may have substantially similar configurations. The combustor  70  of the first turbine engine and the combustor  70  of the second turbine engine may also or alternatively have substantially similar configurations. 
     However, corresponding rotor blades in one or more of the engine sections may be slightly different. For example, an upstream-most set of compressor blades  100  (see  FIG. 1 ) in the core of the first turbine engine may define a cross-sectional annular first area. An upstream-most set of compressor blades  100  in the core of the second turbine engine may define a cross-sectional annular second area which is slightly different than the first area. The second area, for example, may be within plus/minus twenty percent (+/−20%) of the first area. 
     The first and the second turbine engines are described above with certain commonalities and certain differences. These commonalities and differences, however, may change depending upon the specific thrust rating requirements, customer requirements, government agency requirements, etc. The present disclosure therefore is not limited to the exemplary embodiments described above. 
     The present disclosure is not limited to the exemplary turbine engine  20  configurations described above. In some embodiments, for example, the core may include more than two rotating assemblies; e.g., three spools, four spools, etc. The core, for example, may include an additional intermediate compressor rotor and an additional intermediate turbine rotor connected together by an additional intermediate speed shaft. In some embodiments, the rotating assembly may include at least one additional compressor rotor where, for example, the LPC rotor  49  and the additional compressor rotor are arranged on opposite sides of the gear train  60 . Furthermore, the present disclosure is not limited to a typical turbine engine configuration with the fan section  28  forward of the core (e.g., engine sections  30 - 35 ). In other embodiments, for example, the turbine engine  20  may be configured as a geared pusher fan engine or another type of gear turbine engine. The present invention therefore is not limited to any particular types or configurations of turbine engines. 
     While various embodiments of the present invention have been disclosed, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the invention. For example, the present invention as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present invention that some or all of these features may be combined with any one of the aspects and remain within the scope of the invention. Accordingly, the present invention is not to be restricted except in light of the attached claims and their equivalents.