Patent Publication Number: US-10317078-B2

Title: Cooling a multi-walled structure of a turbine engine

Description:
This application claims priority to PCT Patent Application No. PCT/US14/066880 filed Nov. 21, 2014 which claims priority to U.S. Patent Application No. 61/907,228 filed Nov. 21, 2013, which are hereby incorporated herein by reference in their entireties. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Technical Field 
     This disclosure relates generally to a turbine engine and, more particularly, to cooling a multi-walled structure of a turbine engine. 
     2. Background Information 
     A floating wall combustor for a turbine engine typically includes a bulkhead, an inner combustor wall and an outer combustor wall. The bulkhead extends radially between the inner and the outer combustor walls. Each combustor wall includes a shell and a heat shield that defines a respective radial side of a combustion chamber. Cooling cavities extend radially between the heat shield and the shell. These cooling cavities fluidly couple impingement apertures defined in the shell with effusion apertures defined in the heat shield. 
     During turbine engine operation, the impingement apertures direct cooling air from a plenum adjacent the combustor into the cooling cavities to impingement cool the heat shield. The effusion apertures direct the cooling air from the cooling cavities into the combustion chamber to film cool the heat shield. This cooling air subsequently mixes and reacts with a fuel-air mixture within the combustion chamber, thereby leaning out the fuel-air mixture in both an upstream fuel-rich primary zone and a downstream fuel-lean secondary zone. The primary zone of the combustion chamber is located between the bulkhead and the secondary zone, which is generally axially aligned with quench apertures in the combustor walls. 
     In an effort to increase turbine engine efficiency and power, temperature within the combustion chamber may be increased. However, increasing the temperature in the primary zone with a relatively lean fuel-air mixture may also increase NOx, CO and unburned hydrocarbon (UHC) emissions. 
     There is a need in the art for an improved turbine engine combustor. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the invention, an assembly is provided for a turbine engine. This turbine engine assembly includes a body, a shell and a heat shield panel. The panel is attached to the shell with a tapered cooling cavity between the shell and the panel. The panel defines a cooling aperture configured to direct air out of the cooling cavity to impinge against the body. 
     According to another aspect of the invention, another assembly is provided for a turbine engine. This turbine engine assembly includes a body, a shell and a heat shield panel. The panel is attached to the shell with a cooling cavity vertically between the shell and the panel. The panel includes a rail and defines a plurality of cooling apertures, at the rail, through which substantially all air within the cooling cavity is directed out of the cooling cavity to impinge against the body. 
     The cooling aperture may be one of a plurality of cooling apertures defined by the panel and configured to direct air out of the cooling cavity to impinge against the body. 
     Substantially all air entering the cooling cavity may be directed out of the cooling cavity through the cooling apertures. 
     The body may define a plurality of second cooling apertures through which air is directed towards the panel. The cooling apertures may be circumferentially offset from the second cooling apertures. 
     The panel may include a rail that partially defines the cooling cavity. The panel may define the cooling aperture at the rail. The rail may at least partially define the cooling aperture. The panel may also include a base that may partially define the cooling cavity. The base may also or alternatively at least partially define the cooling aperture. 
     A surface of the shell and a surface of the panel may converge towards one another and vertically define at least a portion of the cooling cavity. 
     The body may be configured as or otherwise include a combustor bulkhead. 
     The body may be configured as or otherwise include a second heat shield panel that is attached to the shell. 
     The turbine engine assembly may include a second body. The panel may further define a second cooling aperture configured to direct air from a second cooling cavity between the shell and the panel to impinge against the second body. 
     The second cooling aperture may be one of a plurality of second cooling apertures defined by the panel and configured to direct air out of the second cooling cavity to impinge against the second body. 
     The body may be configured as or otherwise include a combustor bulkhead. In addition or alternatively, the second body may be configured as or otherwise include a second heat shield panel. 
     The body may define a plurality of second cooling apertures through which air is directed towards the panel. The cooling apertures may be circumferentially offset from the second cooling apertures. 
     The rail may at least partially define one or more of the cooling apertures. 
     The panel may include a base that partially defines the cooling cavity. The base may also at least partially define one or more of the cooling apertures. 
     The body may be configured as or otherwise include a combustor bulk head or a second heat shield panel. 
     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 side cutaway illustration of a geared turbine engine; 
         FIG. 2  is a side cutaway illustration of a portion of a combustor section; 
         FIG. 3  is a perspective illustration of a portion of a combustor; 
         FIG. 4  is a side sectional illustration of a portion of the combustor at a first circumferential position; 
         FIG. 5  is a side sectional illustration of the combustor of  FIG. 4  at a second circumferential position; 
         FIG. 6  is an enlarged side sectional illustration of a portion A of the combustor of  FIG. 4 ; 
         FIG. 7  is an enlarged side sectional illustration of a portion B of the combustor of  FIG. 4 ; 
         FIG. 8  is an enlarged side sectional illustration of a portion C of the combustor of  FIG. 5 ; 
         FIG. 9  is a circumferential sectional illustration of a portion of a heat shield panel included in the combustor of  FIG. 4 ; and 
         FIGS. 10-13  are side sectional illustrations of respective portions of alternate embodiment combustors. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a side cutaway illustration of a geared turbine engine  20 . This turbine engine  20  extends along an axial centerline  22  between a forward airflow inlet  24  and an aft airflow exhaust  26 . The turbine engine  20  includes a fan section  28 , a compressor section  29 , a combustor section  30  and a turbine section  31 . The compressor section  29  includes a low pressure compressor (LPC) section  29 A and a high pressure compressor (HPC) section  29 B. The turbine section  31  includes a high pressure turbine (HPT) section  31 A and a low pressure turbine (LPT) section  31 B. The engine sections  28 - 31  are arranged sequentially along the centerline  22  within an engine housing  34 , which includes a first engine case  36  and a second engine case  38 . 
     Each of the engine sections  28 ,  29 A,  29 B,  31 A and  31 B includes a respective rotor  40 - 44 . Each of the rotors  40 - 44  includes a plurality of rotor blades arranged circumferentially around and connected to (e.g., formed integral with or mechanically fastened, welded, brazed, adhered or otherwise attached to) one or more respective rotor disks. The fan rotor  40  is connected to a gear train  46  through a fan shaft  47 . The gear train  46  and the LPC rotor  41  are connected to and driven by the LPT rotor  44  through a low speed shaft  48 . The HPC rotor  42  is connected to and driven by the HPT rotor  43  through a high speed shaft  50 . The shafts  47 ,  48  and  50  are rotatably supported by a plurality of bearings  52 . Each of the bearings  52  is connected to the second engine case  38  by at least one stationary structure such as, for example, an annular support strut. 
     Air enters the turbine engine  20  through the airflow inlet  24 , and is directed through the fan section  28  and into an annular core gas path  54  and an annular bypass gas path  56 . The air within the core gas path  54  may be referred to as “core air”. The air within the bypass gas path  56  may be referred to as “bypass air”. 
     The core air is directed through the engine sections  29 - 31  and exits the turbine engine  20  through the airflow exhaust  26 . Within the combustor section  30 , fuel is injected into a combustion chamber  58  and mixed with the core air. This fuel-core air mixture is ignited to power the turbine engine  20  and provide forward engine thrust. The bypass air is directed through the bypass gas path  56  and out of the turbine engine  20  through a bypass nozzle  60  to provide additional forward engine thrust. Alternatively, the bypass air may be directed out of the turbine engine  20  through a thrust reverser to provide reverse engine thrust. 
       FIG. 2  illustrates an assembly  62  of the turbine engine  20 . This turbine engine assembly  62  includes a combustor  64 . The turbine engine assembly  62  also includes one or more fuel injector assemblies  66 , each of which may include a fuel injector  68  mated with a swirler  70 . 
     The combustor  64  may be configured as an annular floating wall combustor arranged within an annular plenum  72  of the combustor section  30 . The combustor  64  of  FIGS. 2 and 3 , for example, includes an annular combustor bulkhead  74 , a tubular combustor inner wall  76 , and a tubular combustor outer wall  78 . The bulkhead  74  extends radially between and is connected to the inner wall  76  and the outer wall  78 . The inner wall  76  and the outer wall  78  each extends axially along the centerline  22  from the bulkhead  74  towards the turbine section  31 A, thereby defining the combustion chamber  58 . 
       FIG. 4  is a side sectional illustration of a portion of the combustor  64  at a first circumferential position.  FIG. 5  is a side sectional illustration of the combustor  64  portion of  FIG. 4  at a second circumferential position.  FIG. 6  is an enlarged side sectional illustration of a portion A of the combustor  64  of  FIG. 4 .  FIG. 7  is an enlarged side sectional illustration of a portion B of the combustor  64  of  FIG. 4 .  FIG. 8  is an enlarged side sectional illustration of a portion C of the combustor  64  of  FIG. 5 . 
     The inner wall  76  and the outer wall  78  may each be configured as a multi-walled structure; e.g., a hollow dual-walled structure. The inner wall  76  and the outer wall  78  of  FIGS. 2 and 4 , for example, each includes a tubular combustor shell  80 , a tubular combustor heat shield  82 , and one or more cooling cavities  84 - 86  (e.g., impingement cavities). Referring now to FIGS.  2  and  3 , the inner wall  76  and the outer wall  78  may also each include one or more quench apertures  88 , which extend through the wall  76 ,  78  and are disposed circumferentially around the centerline  22 . 
     Referring to  FIG. 2 , the shell  80  extends circumferentially around the centerline  22 . The shell  80  extends axially along the centerline  22  between an axial forward end  90  and an axial aft end  92 . The shell  80  is connected to the bulkhead  74  at the forward end  90 . The shell  80  may be connected to a stator vane assembly  94  or the HPT section  31 A at the aft end  92 . 
     Referring to  FIG. 4 , the shell  80  has a plenum surface  96 , a cavity surface  98  and one or more aperture surfaces  100  and  102  (see also  FIG. 5 ). At least a portion of the shell  80  extends radially between the plenum surface  96  and the cavity surface  98 . The plenum surface  96  defines a portion of the plenum  72 . The cavity surface  98  defines a portion of one or more of the cavities  84 - 86  (see  FIG. 2 ). 
     The aperture surfaces  100  and  102  (see  FIG. 4 ) may be respectively arranged in one or more aperture arrays  104  and  106 . The apertures surfaces  100 ,  102  in each aperture array  104 ,  106  may be disposed circumferentially around the centerline  22 . The aperture surfaces  100  in the first aperture array  104  may be located proximate (or adjacent) to and on a first axial side  108  of a respective heat shield rail  110  (e.g., intermediate rail). The aperture surfaces  102  in the second aperture array  106  may be located proximate (or adjacent) to and on an opposite second axial side  112  of the respective heat shield rail  110 . 
     Each of the aperture surfaces  100 ,  102  defines a respective cooling aperture  114 ,  116 . Each cooling aperture  114 ,  116  extends (e.g., radially) through the shell  80  from the plenum surface  96  to the cavity surface  98 . Each cooling aperture  114 ,  116  may be configured as an impingement aperture. Each aperture surface  100 ,  102  of  FIG. 4 , for example, is configured to direct a jet of cooling air to impinge substantially perpendicularly against the heat shield  82 . 
     Referring to  FIG. 2 , the heat shield  82  extends circumferentially around the centerline  22 . The heat shield  82  extends axially along the centerline  22  between an axial forward end and an axial aft end. The forward end is located at an interface between the wall  76 ,  78  and the bulkhead  74 . The aft end may be located at an interface between the wall  76 ,  78  and the stator vane assembly  94  or the HPT section  31 A. 
     The heat shield  82  may include one or more heat shield panels  118  and  120 , one or more of which may have an arcuate geometry. The panels  118  and  120  are respectively arranged at discrete locations along the centerline  22 . The panels  118  are disposed circumferentially around the centerline  22  and form a forward hoop. The panels  120  are disposed circumferentially around the centerline  22  and form an aft hoop. Alternatively, the heat shield  82  may be configured from one or more tubular bodies. 
     Referring to  FIGS. 4 and 9 , each of the panels  118  has one or more cavity surfaces  122  and  124  and a chamber surface  126 . At least a portion of the panel  118  extends radially between the cavity surfaces  122  and  124  and the chamber surface  126 . Each cavity surface  122  defines at least one side of a respective one of the cooling cavities  84 . Each cavity surface  124  defines at least one side of a portion of a respective one of the cooling cavities  85 . It will be appreciated that the chamber surface  126  similarly defines at least one side of a portion of the combustion chamber  58 . 
     For example, each panel  118  may include a panel base  128  and one or more rails (e.g., rails  110  and  130 - 133 ) with the panel base  128  and the panel rails  110 ,  130 ,  132  and  133  collectively defining cavity surface  122 . Similarly, the panel base  128  and the panel rails  110  and  131 - 133  may collectively define cavity surface  124 , and the panel base  128  may define the chamber surface  126 . 
     The panel base  128  may be configured as a generally curved (e.g., arcuate) plate. The panel base  128  extends axially between an axial forward end  134  and an axial aft end  136 . The panel base  128  extends circumferentially between opposing circumferential ends  138  and  140 . 
     The panel rails may include the axial intermediate rail  110 , one or more axial end rails  130  and  131 , and one more circumferential end rails  132  and  133 . Each of the panel rails  110  and  130 - 133  of the inner wall  76  extends radially in from the respective panel base  128 ; see also  FIG. 2 . Each of the panel rails  110  and  130 - 133  of the outer wall  78  extends radially out from the respective panel base  128 ; see also  FIG. 2 . 
     The axial intermediate and end rails  110 ,  130  and  131  extend circumferentially between and are connected to the circumferential end rails  132  and  133 . The axial intermediate rail  110  is disposed axially (e.g., centrally) between the axial end rails  130  and  131 . The axial end rail  130  is arranged at the forward end  134 . The axial end rail  131  is arranged at the aft end  136 . The circumferential end rail  132  is arranged at the circumferential end  138 . The circumferential rail  133  is arranged at the circumferential end  140 . 
     Still referring to  FIGS. 4 and 9 , each panel  118  may also have one or more aperture surfaces  142  and  144 . These aperture surfaces  142  and  144  may be respectively arranged in one or more aperture arrays  146  and  148 . The aperture surfaces  142 ,  144  in each array  146 ,  148  may be disposed circumferentially around the centerline  22 . Respective aperture surfaces  142  in the forward array  146  may be adjacent (or in or proximate) the respective axial end rail  130  (see also  FIG. 6 ). Respective aperture surfaces  144  in the aft array  148  may be in (or adjacent or proximate) the respective axial end rail  131  (see also  FIG. 7 ). 
     Referring to  FIG. 6 , each of the aperture surfaces  142  defines a cooling aperture  150  in the panel  118  and, thus, the heat shield  82 . Each cooling aperture  150  may extend radially and axially (and/or circumferentially) through the panel base  128 . Alternatively, referring to  FIG. 10 , one or more of the cooling apertures  150  may extend radially and axially (and/or circumferentially) through and be defined in the panel base  128  as well as the axial end rail  130 . Referring to  FIG. 11 , one or more of the cooling apertures  150  may also or alternatively extend axially (and/or circumferentially) through and be defined in the axial end rail  130 . 
     Referring again to  FIG. 6 , one or more of the cooling apertures  150  may each be configured as an impingement aperture. Each aperture surface  142  of  FIG. 6 , for example, is configured to direct a jet of cooling air along a respective trajectory  152  to impinge against a body such as, for example, a heat shield  154  of the bulkhead  74 . 
     Referring to  FIGS. 6 and 8 , the cooling apertures  150  may be laterally (e.g., circumferentially offset) with respect to an array of one or more cooling apertures  156  defined in the bulkhead  74  to reduce or prevent air directed from the apertures  150  and  156  from colliding and directly mixing. Each cooling aperture  150 , for example, may be circumferentially centered between two adjacent cooling apertures  156 , and vice versa. Each cooling aperture  156  may extend radially and axially (and/or circumferentially) through the heat shield  154  and a shell  158  of the bulkhead  74 . Each cooling aperture  156  may be configured as an impingement aperture. Surfaces  160  and  162  defining the cooling aperture  156  of  FIG. 8 , for example, are configured to direct a jet of cooling air along a respective trajectory  164  to impinge against the panel  118 . The trajectory  164  may be substantially parallel and opposite the trajectory  152  in  FIG. 6 , but for example circumferentially offset. 
     Referring to  FIG. 7 , each of the aperture surfaces  144  defines a cooling aperture  166  in the panel  118  and, thus, the heat shield  82 . Each cooling aperture  166  may extend radially and axially (and/or circumferentially) through the panel base  128  and the axial end rail  131 . Alternatively, referring to  FIG. 12 , one or more of the cooling apertures  166  may extend radially and axially (and/or circumferentially) through and be defined in the panel base  128 . One or more of the cooling apertures  166  may also or alternatively extend axially (and/or circumferentially) through and be defined in the axial end rail  131  in a similar manner as illustrated in  FIG. 9 . 
     Referring again to  FIG. 7 , one or more of the cooling apertures  166  may each be configured as an impingement aperture. Each aperture surface  144  of  FIG. 7 , for example, is configured to direct a jet of cooling air along a respective trajectory  168  to impinge against a body such as, for example, a forward portion of a respective one of the panels  120 . Alternatively, one or more of the apertures surfaces  144  may be configured to direct a jet of cooling air into the combustion chamber  58  such that the cooling air forms a film against a downstream portion of the heat shield  82 ; e.g., panels  120 . 
     Referring to  FIG. 2 , the heat shield  82  of the inner wall  76  circumscribes the shell  80  of the inner wall  76 , and defines an inner side of the combustion chamber  58 . The heat shield  82  of the outer wall  78  is arranged radially within the shell  80  of the outer wall  78 , and defines an outer side of the combustion chamber  58  that is opposite the inner side. The heat shield  82  and, more particularly, each of the panels  118  and  120  may be respectively attached to the shell  80  by a plurality of mechanical attachments  170  (e.g., threaded studs respectively mated with washers and nuts); see also  FIG. 4 . The shell  80  and the heat shield  82  thereby respectively form the cooling cavities  84 - 86  in each of the walls  76 ,  78 . 
     Referring to  FIGS. 4, 5 and 9 , each cooling cavity  84  is defined radially by and extends radially between the cavity surface  98  and a respective one of the cavities surfaces  122  as set forth above. Each cooling cavity  84  is defined circumferentially by and extends circumferentially between the end rails  132  and  133  of a respective one of the panels  118 . Each cooling cavity  84  is defined axially by and extends axially between the rails  110  and  130  of a respective one of the panels  118 . In this manner, each cooling cavity  84  may fluidly couple one or more of the cooling apertures  114  with one or more of the cooling apertures  150 . 
     Each cooling cavity  85  is defined radially by and extends radially between the cavity surface  98  and a respective one of the cavities surfaces  124  as set forth above. Each cooling cavity  85  is defined circumferentially by and extends circumferentially between the end rails  132  and  133  of a respective one of the panels  118 . Each cooling cavity  85  is defined axially by and extends axially between the rails  110  and  131  of a respective one of the panels  118 . In this manner, each cooling cavity  85  may fluidly couple one or more of the cooling apertures  116  with one or more of the cooling apertures  166 . 
     Referring to  FIG. 5 , respective portions  172 - 175  of the shell  80  and the heat shield  82  may converge towards one another; e.g., the shell portions  172  and  173  may include concavities. In this manner, a vertical distance between the shell  80  and the heat shield  82  may decrease as each panel  118  extends from the intermediate rail  110  to its axial end rails  130  and  131 . A vertical height of each intermediate rail  110 , for example, may be greater than vertical heights of the respective axial end rails  130  and  131 . The height of each axial end rail  130 ,  131 , for example, is between about twenty percent (20%) and about fifty percent (50%) of the height of the intermediate rail  110 . The shell  80  and the heat shield  82  of  FIG. 5  therefore may define each cooling cavity  84 ,  85  with a tapered geometry. However, in other embodiments, one or more of the cooling cavities  84  and/or  85  may be defined with non-tapered geometries as illustrated, for example, in  FIG. 2 . 
     Referring to  FIG. 4 , core air from the plenum  72  is directed into each cooling cavity  84 ,  85  through respective cooling apertures  114 ,  116  during turbine engine operation. This core air (e.g., cooling air) may impinge against the respective panel base  128 , thereby impingement cooling the panel  118  and the heat shield  82 . 
     The cooling air may flow axially within the respective cooling cavities  84  and  85  from the cooling apertures  114 ,  116  to the cooling apertures  150 ,  166 . The converging surfaces  98  and  122 ,  98  and  124  may accelerate the axially flowing cooling air as it flows towards a respective one of the axial end rails  130 ,  131 . By accelerating the cooling air, thermal energy transfer from the heat shield  82  to the shell  80  through the cooling air may be increased. 
     Referring to  FIG. 6 , respective cooling apertures  150  may direct substantially all of the cooling air within the cooling cavity  84  into the combustion chamber  58  towards the bulkhead  74 . This cooling air may subsequently impinge against the bulkhead  74  (e.g., the heat shield  154 ) and thereby impingement cooling to the bulkhead  74 . The force of the cooling air impinging against the bulkhead  74  may dissipate the kinetic energy of the air, thereby reducing the likelihood that the cool air will mix and react with the relatively hot core air within the combustion chamber  58 . As a result, the temperature within an upstream portion of the combustion chamber  58  may be increased to increase turbine engine efficiency and power without, for example, substantially increasing NOx, CO and unburned hydrocarbon (UHC) emissions of the turbine engine  20 . 
     Referring to  FIG. 7 , respective cooling apertures  166  may direct substantially all of the cooling air within the cooling cavity  85  into the combustion chamber  58  towards the panels  120 . This cooling air may subsequently impinge against the panels  120  and thereby impingement cool a downstream portion of the heat shield  82  and, more particularly, upstream edges of the panels  120 . The force of the cooling air impinging against the panels  120  may dissipate the kinetic energy of the air, thereby reducing the likelihood that the cooling air will mix and react with the relatively hot core air within the combustion chamber  58 . As indicated above, reducing mixing and reactions between the cooling air and the core air may reduce NOx, CO and unburned hydrocarbon (UHC) emissions of the turbine engine  20 . 
     Referring to  FIG. 13 , in some embodiments, one or more of the walls  76  and  78  may each include one or more cooling elements  174 . These cooling elements  174  may be found integral with or attached to the panel base  128 . One or more of the cooling elements  174  may further define the cavity surface  122  of each panel  118 . One or more of the cooling elements  174  may further define the cavity surface  124  of each panel  118 . Each cooling element  174  of  FIG. 13  is configured as a cooling pin. One or more of the cooling elements  174 , however, may alternatively each be configured as a nodule, a rib, a trip strip or any other type of protrusion or device that aids in the cooling of the wall  76 ,  78 . 
     The shell  80  and/or the heat shield  82  may each have a configuration other than that described above. In some embodiments, for example, a respective one of the heat shield portions  174  and  175  may have a concavity that defines the cooling cavity tapered geometry with the concavity of a respective one of the shell portions  172  and  173 . In some embodiments, a respective one of the heat shield portions  174 ,  175  may have a concavity rather than a respective one of the shell portions  172 ,  173 . In some embodiments, one or more of the afore-described concavities may be replaced with a substantially straight radially tapering wall. In some embodiments, each panel  118  may define one or more additional cooling cavities with the shell  80 . In some embodiments, each panel  118  may define a single cooling cavity (e.g.,  84  or  85 ) with the shell  80 , which cavity may taper in a forward or aftward direction. In some embodiments, one or more of the panels  120  may have a similar configuration as that described above with respect to the panels  118 . The present invention therefore is not limited to any particular combustor wall configurations. 
     The terms “forward”, “aft”, “inner”, “outer”, “radial”, circumferential” and “axial” are used to orientate the components of the turbine engine assembly  62  and the combustor  64  described above relative to the turbine engine  20  and its centerline  22 . One or more of these components, however, may be utilized in other orientations than those described above. The present invention therefore is not limited to any particular spatial orientations. 
     The turbine engine assembly  62  may be included in various turbine engines other than the one described above. The turbine engine assembly  62 , for example, may be included in a geared turbine engine where a gear train connects one or more shafts to one or more rotors in a fan section, a compressor section and/or any other engine section. Alternatively, the turbine engine assembly  62  may be included in a turbine engine configured without a gear train. The turbine engine assembly  62  may be included in a geared or non-geared turbine engine configured with a single spool, with two spools (e.g., see  FIG. 1 ), or with more than two spools. The turbine engine may be configured as a turbofan engine, a turbojet engine, a propfan engine, or any other type of 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 within 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.