Abstract:
A method facilitates assembling a gas turbine engine. The method comprises mounting a core engine to a vehicle, coupling a fuselage radially outward and around the core engine, and coupling an exhaust nozzle to the core engine to channel exhaust gases discharged from the core engine. In addition, the method also comprises coupling an infrared suppression system in flow communication with the engine exhaust nozzle for channeling exhaust gases discharged from said exhaust nozzle to facilitate suppressing an exhaust infrared signature of the core engine during operation, wherein the infrared suppression system includes an access door and a flow channel that is coupled to the access door such that the flow channel is movable with the access door from a closed position to an open position.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines, and more specifically to methods and apparatus for exhausting gases from gas turbine engines.  
         [0002]     The exhaust nozzle and plume from gas turbine engines is a potential source of high infrared energy which may be used for targeting and/or tracking purposes. More specifically, the infrared energy may be used for targeting and/or tracking by heat seeking missiles and/or various forms of infrared imaging systems. Because the military mission of helicopters may involve flying at low altitudes and at reduced speed in comparison to other military aircraft, helicopters are susceptible to ground-to-air, infrared-guided missiles. For example, within at least some known helicopters, the exposed metal surfaces of the gas turbine engine exhaust may operate in excess of 800° F., and thus emit infrared electromagnetic radiation at virtually all wavelengths as hot exhaust gases flow past the exposed surfaces. Moreover, continued heating of aircraft surfaces, including the fuselage, during hover or flight may also create structural issues.  
         [0003]     Accordingly, within at least some known gas turbine engines, infrared signature reduction methods have been employed to facilitate reducing the infrared signature of a gas turbine engine. More specifically, at least some known gas turbine engines use complicated cooling schemes to supply cooling air to facilitate cooling directly visible surfaces and to dilute the high temperature exhaust gases. Other known gas turbine engines use infrared suppressors which change the direction of the exhaust flow discharged from the engine to facilitate hiding the hottest exposed surfaces with cooler surfaces.  
         [0004]     However, generally, any benefits gained by such systems may be offset by losses created in acquiring the reduced infrared signature. More specifically, when the exhaust gases are cooled by cooling air, the air may be provided at a substantial engine power loss or weight penalty. Furthermore, in other known systems, the benefits gained by such systems may be offset by comparatively large installation space requirements, complex ducting, and/or substantial weight penalties. Moreover, the weight and physical size of such suppression systems may limit access to the gas turbine engine for routine maintenance and inspections.  
       BRIEF SUMMARY OF THE INVENTION  
       [0005]     In one aspect, a method for assembling a gas turbine engine is provided. The method comprises mounting a core engine to a vehicle, coupling a fuselage radially outward and around the core engine, and coupling an exhaust nozzle to the core engine to channel exhaust gases discharged from the core engine. In addition, the method also comprises coupling an infrared suppression system in flow communication with the engine exhaust nozzle for channeling exhaust gases discharged from said exhaust nozzle to facilitate suppressing an exhaust infrared signature of the core engine during operation, wherein the infrared suppression system includes an access door and a flow channel that is coupled to the access door such that the flow channel is movable with the access door from a closed position to an open position.  
         [0006]     In another aspect, an exhaust assembly for a gas turbine engine including a turbine rear frame is provided. The exhaust assembly includes an engine exhaust nozzle extending downstream from the turbine rear frame, and an infrared suppression system coupled in flow communication with the engine exhaust nozzle for channeling exhaust gases discharged from the exhaust nozzle. The suppression system includes a flow channel coupled to an access door, such that the flow channel is movable with the access door from a closed position to an open position wherein the access door forms a work platform configured to support a user thereon. The suppression system facilitates suppressing an exhaust infrared signature of the gas turbine engine.  
         [0007]     In a further aspect, a gas turbine engine configured to couple to a fuselage is provided. The gas turbine engine includes a core engine and an exhaust assembly that extends downstream from the core engine for discharging exhaust gases from the core engine. The exhaust assembly includes an exhaust nozzle that is coupled to the core engine and an infrared suppression system that is coupled in flow communication downstream from the engine exhaust nozzle for channeling exhaust gases discharged from the exhaust nozzle. The infrared suppression system includes a flow channel and an access door. The flow channel is coupled to the access door, such that the flow channel is movable with the access door from a closed position to an open position. Both the flow channel and access door are coupled to the fuselage. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0008]      FIG. 1  is a plan view of a gas turbine engine assembly including an access door that may be used with a helicopter;  
         [0009]      FIG. 2  is a perspective view of a core engine and an exemplary exhaust nozzle that may be used with the gas turbine engine assembly shown in  FIG. 1 ;  
         [0010]      FIG. 3  is a partial front view of an exemplary helicopter including the access door shown in  FIG. 1  in a closed position;  
         [0011]      FIG. 4  is a partial front view of the helicopter shown in  FIG. 3  including the access door shown in  FIG. 1  in an open position;  
         [0012]      FIG. 5  is an alternative embodiment of a turbine exhaust nozzle that may be used with the gas turbine engine assembly shown in  FIG. 1 ;  
         [0013]      FIG. 6  is another alternative embodiment of an exemplary turbine exhaust nozzle that may be used with the gas turbine engine assembly shown in  FIG. 1 ; and  
         [0014]      FIG. 7  is a further alternative embodiment of an exemplary turbine exhaust nozzle that may be used with the gas turbine engine assembly shown in  FIG. 1 .  
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0015]      FIG. 1  is a plan view of a helicopter  10  that includes two gas turbine engine assemblies  42  and access doors  12 .  FIG. 2  is a perspective view of a core engine  40  and exhaust nozzle  76  that may be used with gas turbine engine assembly  42 .  FIG. 3  is a partial front view of an exemplary helicopter  14  including access door  12  shown in a closed position  16 .  FIG. 4  is a partial front view of helicopter  14  including access door  12  shown in an open position  22 .  FIGS. 5, 6 , and  7  are alternative embodiments of turbine exhaust nozzles  76  that may be used with gas turbine engine assembly  42 .  
         [0016]     In the exemplary embodiment, helicopter  14  includes a pair of gas turbine engines  40  that each include an inlet end  44  and an exhaust end  46 . Engines  40  are symmetrical with respect to an axis of symmetry  47  extending between engines  40 . Core engines  40  are mounted within an engine compartment  48  defined by a helicopter fuselage  50 . Specifically, in the exemplary embodiment, gas turbine engine assembly  42  includes core engines  40  and an exhaust assembly  52  that extends downstream from engines  40  for discharging exhaust gases from engines  40 . In one embodiment, each core engine  40  is a T58 engine commercially available from General Electric Aircraft Engines, Lynn, Mass. A rear drive shaft  60  extends from engine  40  to a main transmission  62 .  
         [0017]     Exhaust assembly  52  includes a pair of exhaust nozzle assemblies  70  and suppression system  72 . Suppression system  72 , as described in more detail below, facilitates suppressing an exhaust infrared signature of gas turbine engine assembly  42  during engine operation. As used herein, the term suppression mean that the infrared signature emanating from gas turbine engine assembly  10  is facilitated to be reduced below a pre-determined threshold value which is indicative of the acquisition, tracking, and/or targeting capability of a particular infrared threat.  
         [0018]     Each exhaust nozzle assembly  70  includes a turbine rear frame housing  75  that includes a drive shaft tunnel  74 , and a primary nozzle  76 . Specifically, exhaust discharged from each engine  40  is initially channeled through rear frame housing  75  and around drive shaft tunnel  74  prior to entering primary nozzle  76 . In the exemplary embodiment, each drive shaft tunnel  74  is integrally formed with an elbow  78  such that exhaust entering each rear frame  75  is discharged outwardly at an oblique angle θ measured with respect to axis of symmetry  47 . More specifically, in the exemplary embodiment, angle θ is approximately sixty degrees.  
         [0019]     Exhaust discharged from engines  40  is channeled into a pair of primary nozzles  76  that are each coupled to turbine rear frame housing  75 . In the exemplary embodiment, each primary nozzle is a single-wall nozzle that includes an elbow  82 . Elbow  82  causes the direction of exhaust flowing through nozzle  76  to be discharged in a direction that is substantially parallel to centerline axis of symmetry  47 . Accordingly, a length L N  of primary nozzle  76 , measured between an inlet end  84  and a discharge end  86  that is downstream from inlet end  84 , is variably selected to enable flow to be discharged substantially axially therefrom. Moreover, nozzle length L N  ensures that an exit aperture defined at nozzle discharge end  86  is oriented substantially perpendicularly to a direction of exhaust flow discharged therethrough. In addition, the combination of elbow  82  and nozzle length L N  causes nozzle  76  to extend through fuselage  50  such that exhaust discharged from nozzle  76  is accelerated and then discharged adjacent an external surface  98  of fuselage  50 . In an alternative embodiment, depending on the application of gas turbine engine assembly  10 , flow through rear frame  75  and primary nozzles  76  remains substantially axial, as neither rear frame  75  and/or primary nozzles  76  include respective elbows  78  and  82 .  
         [0020]     A cross-sectional area defined of the nozzle exit aperture defined at discharge end  86  may be any cross-sectional shape that enables nozzle  76  to function as describe herein. More specifically, the nozzle exit aperture facilitates inducing mixing of exhaust flow discharged therefrom, without promoting an outward propagation of exhaust gases discharged therefrom. For example, as shown in  FIGS. 5 and 6 , the nozzle exit aperture cross-sectional area may be, but is not limited to being, circular, elliptical, rectangular, or daisy-shaped. Additionally, discharge end  86  may also include other mixing enhancement features such as, but not limited to, lobes, scalloped edges, turbulators, and/or chevrons. Moreover, in another alternative embodiment, discharge end  86  includes a convergent lobe design which facilitates mixing exhaust gases discharged therethrough with ambient cooling air introduced to gas turbine engine assembly  10 , as described in more detail below.  
         [0021]     Exhaust exiting primary nozzles  76  is channeled into suppression system  72 . Suppression system  72  includes a pair of flow channels  90  that are each coupled to an access door  12 . More specifically, in the exemplary embodiment, flow channel  90  is formed integrally with door  12 . Each flow channel  90  is coupled in flow communication with primary nozzles  76  such that flow exiting nozzles  76  is routed through flow channels  90  before being discharged to the atmosphere. More specifically, a cross-sectional area of each flow channel is selected to form an annulus with each respective primary nozzle  76 , such that flow exiting nozzles  76  forms a venturi effect which creates a local low pressure immediately downstream from each nozzle discharge end  86 . Accordingly, in one embodiment, each flow channel  90  is tapered from an inlet end  12  coupled to primary nozzle  76 , through an exit aperture or discharge end  94 . More specifically, in the exemplary embodiment, flow channel  90  is tapered such that a cross-sectional area defined within flow channel  90  by an inner surface  96  of flow channel  90  is progressively decreased from inlet end  12  to discharge end  94 . Accordingly, the tapering facilitates ensuring a constant exhaust flow path velocity is maintained within flow channel  90 .  
         [0022]     The cross-sectional area defined within flow channel  90  may be any cross-sectional shape that enables flow channel  90  to function as described herein, such as, but not limited to substantially circular, elliptical, or square. In addition, in the exemplary embodiment, discharge end  94  is formed with a substantially rectangular cross-sectional profile, and as such, in the exemplary embodiment, the cross-sectional shape of flow channel  90  varies along a length L C  of flow channel  90  to facilitate providing a smooth transition from inlet end  12  to discharge end  94 . Moreover, the variable cross-sectional area of flow channel  90  also facilitates optimizing engine backpressure within gas turbine engine assembly  42 , while providing a reduced cooling slot exit static pressure to facilitate achieving a desired cooling flow, as described in more detail below. Accordingly, by optimizing system backpressure, flow channel  90  also facilitates maintaining a desired engine operating efficiency.  
         [0023]     Each flow channel L C  is measured between inlet and exit ends  12  and  94 , respectively. Channel length L C  ensures that exhaust discharged from core engines  40  is discharged downstream from, and does not impinge upon, transmission  62 . Channel length L C  also facilitates mixing between exhaust discharged from core engine  40  and ambient cooling air introduced to each flow channel  90 , as described in more detail below, to facilitate reducing an operating temperature of exhaust flowing therethrough. The exact channel length L C  is a function of a plurality of parameters, including, but not limited to, the particular installation, available power penalty, and desired infrared and radar cross-sectional reduction goals.  
         [0024]     In the exemplary embodiment, each flow channel  90  also includes a plurality of cooling baffles  100 , cooling slots  104 , and an aft elbow  102 . Elbow  102  changes a direction of exhaust flowing through flow channel  90 , such that exhaust entering each flow channel  90  is discharged outwardly with respect to axis of symmetry  47  to facilitate preventing the exhaust gases from impinging against, or contacting, fuselage  50 . Cooling slots  104  extend between flow channel inner surface  96  and an outer surface  106  of flow channel  90  to facilitate admitting cooling air into flow channel  90 .  
         [0025]     Slots  104  are aft facing such that exhaust gases entering flow channel  90  are prevented from exiting flow channel  90  through slots  100 . More specifically, air entering slots  104  forms a cooling boundary layer to facilitate cooling those portions of flow channel inner surface  96  that are directly visible through flow channel exit aperture  94 . Accordingly, the combination of exit aperture  94 , flow channel length L C , and elbow  102  facilitate obstructing or preventing direct line-of-sight viewing of uncooled portions of flow channel inner surface  96  through exit aperture  94 . In addition, in the exemplary embodiment, at least a portion of flow channel inner surface  96  is coated with a high emissivity coating to substantially prevent infrared reflections through exit aperture  94  that may be emitted or originate from hotter “hidden” surfaces. In an alternative embodiment, channel inner surface  96  includes a surface characteristic that substantially prevents infrared reflections through exit aperture  94  that may be emitted or originate from hotter “hidden” surfaces.  
         [0026]     Primary nozzles  76  and flow channels  90  are surrounded by an insulated cowl  120  such that nozzles  76  and flow channels  90  are externally obstructed from direct view. More specifically, cowl  120  is coupled around primary nozzles  76  and flow channels  90  such that at least one cooling passage  126  is defined between an inner surface  128  of cowl  120  and nozzles and flow channels  76  and  90 , respectively. More specifically, cooling passage  126  is coupled in flow communication with flow channel slots  100 . Moreover, cowl  120  facilitates preventing hot surfaces extending over nozzles  76  and flow channels  90  from emitting infrared signals radially outwardly. Cowl  120  includes a fairing or boat tail portion  122  and an inlet mixing portion  124 . Boat tail portion  122  extends between fuselage  50  and flow channel elbow  102  to provide structural support to flow channel  90 . In the exemplary embodiment, boat tail portion  122  is tapered to a thin trailing edge  126  to facilitate reducing drag during flight operations.  
         [0027]     Cowl mixing portion  124  includes a plurality openings  130  that are defined along an upstream side  132  of cowl  120 . In an alternative embodiment, cowl mixing portion  124  includes only one opening  130 . Specifically, openings  130  are generally forward facing to prevent exhaust gases from being discharged therethrough, and such that openings  130  function as a ram air scoop to enable ambient air to be admitted via a ram effect, or through natural flow, into primary nozzles  76  and flow channels  90 .  
         [0028]     Ambient air channeled through openings  130  facilitates annulus mixing and flow channel cooling. More specifically, a portion of ambient air entering openings  130 , represented by arrow  140 , is channeled into an annulus surrounding primary nozzles  76 , and the remaining ambient air entering openings  130 , represented by arrow  142 , is channeled into cooling passage  126  and channeled to cooling slots  100 . Air  140  is directed into the annulus surrounding primary nozzles  76  to facilitate mixing with exhaust gases discharged from primary nozzles  76 .  
         [0029]     An insulated blocking panel  150  extends from fuselage  50  towards cowl inner surface  128  adjacent openings  30  to facilitate preventing a direct line-of-sight viewing of primary nozzles  76  or flow channels  90  through openings  130 . Moreover, in the exemplary embodiment, cowl inner surface  128  is coated with a high emissivity coating to substantially prevent infrared reflections through exit aperture that may be emitted or originate from higher temperature surfaces. In an alternative embodiment, cowl inner surface  128  includes a surface characteristic that substantially prevents infrared reflections through openings  130  that may be emitted or originate from higher temperature surfaces.  
         [0030]     Flow channels  90  are each coupled to access door  12 , and as such, are moveable with access door  12  between open position  22  and closed position  16 . More specifically, when access door  12  is in closed position  16 , flow channel  90  is coupled in position to capture exhaust flow discharged from primary nozzles  76 , as described above. However, during helicopter non-flight operations, because access door  12  is hingedly coupled to fuselage  50 , access door  12  may be rotated from closed position  16  to open position  22  to provide access to components within gas turbine engine assembly  10 . Moreover, as door  12  is rotated to open position  22  from closed position  16 , flow channel  90  and cowl  120  are each moved with door  12 , while primary nozzle  76  remains coupled in position to engine rear housing  75 . Accordingly, primary nozzle  76 , flow channel  90 , blocking panel  150 , and cowl  120  are designed for clearance to enable door  12  to be opened, yet retain suppressor flow functionality when door  12  is closed.  
         [0031]     In an alternative embodiment, flow channel  90  includes a plurality of hollow baffles which are internally cooled. The baffles are positioned across the flowpath defined within flow channel  90  such that the baffles actually define a plurality of flowpath passages through flow channel  90 . During operation, because the baffles are internally cooled, exhaust flowing past the baffles is convectively cooled.  
         [0032]     In the exemplary embodiment, access door  12  is substantially rectangular, and includes a substantially planar inner surface  160 . Accordingly, when rotated to open position  22 , planar surface  160  extends substantially perpendicularly from fuselage  50  and is substantially parallel to the ground beneath helicopter  14 . Moreover, when access door  12  is in open position  22 , access door  12  is fabricated with enough strength to support a user on inner surface  160 , and as such, may be used as a work platform.  
         [0033]     During operation, cooling air is supplied to gas turbine engine assembly  10  through cowl openings  130 . A portion  140  of such ambient air is channeled into the annulus surrounding primary nozzles  76  to facilitate reducing an operating temperature of external surfaces of primary nozzles  76 . More specifically, the low pressure area created by the venturi effect created as exhaust flow exits primary nozzles  76  facilitates drawing additional ambient air  140  into the channel extending downstream from primary nozzles  76 . The nozzle exit aperture defined at discharge end  86  facilitates inducing mixing of ambient cooling air  140  and exhaust gases discharged from core engine  40  such that hot exhaust gases at primary nozzle discharge end  86  are facilitated to be suppressed. In addition, the mixing enhancement features included at nozzle discharge end  86  facilitate enhancing shearing and mixing between exhaust and ambient air flows.  
         [0034]     In addition, a portion  142  of such ambient air is channeled through passage  126  and to slots  100 , during operation, wherein remaining air  142  entering flow channel  90  provides a layer of cooling air to facilitate cooling aft portions of flow channel inner surface  96  that are visible through exit aperture  94 . Accordingly, slots  100  facilitate reducing an operating temperature of exhaust flow path surfaces. Additional suppression is achieved through the combination of exit aperture  94 , flow channel length L C , and elbow  102 , which facilitate obstructing or preventing direct line-of-sight viewing of uncooled portions of flow channel inner surface  96  through exit aperture  94 . Accordingly, suppression system  72  facilitates the operating temperature of engine exhaust through gas turbine engine assembly  10 , thus suppressing the infrared signature generated by core engines  40 .  
         [0035]     In the exemplary embodiment, flow exit aperture  86  of primary nozzle  76  has either a substantially circular cross-sectional profile or a substantially elliptical cross-sectional profile. Alternatively, exit aperture  86  may have any cross-sectional profile that enables primary nozzle  76  to function as described herein.  
         [0036]     Moreover, there are several mixing enhancement features included in this invention to facilitate enhancing shearing and mixing between primary nozzle exhaust and ambient air flows  140 . For example and referring to  FIG. 5 , nozzle exit aperture  86  facilitates enhances mixing of ambient cooling air  140  and exhaust gases discharged from core engine  40  using chevron-shaped extensions of primary nozzle  76 . In this embodiment, each chevron-shaped extension is cup- or spoon-shaped and includes a concave surface that faces inwardly towards the hot primary nozzle exhaust flow.  
         [0037]     Referring to  FIG. 6 , nozzle exit aperture  86  is substantially rectangular in cross-section and facilitates enhances mixing of ambient cooling air  140  and exhaust gases discharged from core engine  40  via corrugated surfaces of primary nozzle  76 . In this embodiment, each corrugation is aligned such that the axis of corrugation extends substantially in the same direction as that of the hot primary nozzle exhaust flow. However, the enhanced mixing may be accomplished with or with out the use of corrugations and regardless of the cross-sectional shape of nozzle  76  adjacent aperture  86 . For example, in the exemplary embodiment illustrated in  FIG. 7 , nozzle  76  has a substantially circular cross sectional profile adjacent exit aperture  86  and includes a plurality of corrugations.  
         [0038]     The above-described gas turbine engine assemblies are cost-effective and highly reliable. Each assembly includes a exhaust assembly that facilitates suppressing an infrared signature generated by the core engines. Moreover, in the exemplary embodiment, the exhaust assembly initially turns and accelerates the exhaust prior to mixing the exhaust with an ambient airflow. Additional cooling air facilitates cooling flowpath surfaces that are visible through the exhaust assembly discharge. As a result, the exhaust assembly system facilitates suppressing an infrared signature of the engine in a cost-effective and reliable manner.  
         [0039]     Exemplary embodiments of gas turbine assemblies are described above in detail. The assemblies are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each gas turbine engine assembly component can also be used in combination with other gas turbine engine assembly components.  
         [0040]     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.