Abstract:
A fairing kit for a gas turbine engine, where the engine has a core gas turbine engine, a fan rotor, and a plurality of external fan blades attached to the fan rotor and powered by the core gas turbine engine. The core gas turbine engine has an annular splitter for directing a portion of incoming airflow into the core gas turbine engine. The fairing kit comprises: a) a fairing; b) a plurality of fasteners for securing the fairing to the core gas turbine engine; and c) a conformable seal for sealing mating surfaces of the fairing and the core gas turbine engine.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The technology described herein relates generally to gas turbine engines, and more particularly, to a fairing kit for transporting such engines on ferry flights. 
         [0002]    At least one known gas turbine engine assembly includes a fan assembly that is mounted upstream from a core gas turbine engine. During operation, a portion of the airflow discharged from the fan assembly is channeled downstream to the core gas turbine engine wherein the airflow is further compressed. The compressed airflow is then channeled into a combustor, mixed with fuel, and ignited to generate hot combustion gases. The combustion gases are then channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight. The other portion of the airflow discharged from the fan assembly exits the engine through a fan stream nozzle. 
         [0003]    Gas turbine engines such as described herein are frequently installed on aircraft in pairs or multiples, such that in the course of normal operation the aircraft is propelled in flight by two, three, four, or more gas turbine engines. With such multi-engine installations, the aircraft may in some circumstances be safely operated with fewer than all installed engines operating. 
         [0004]    In service, gas turbine engines are subject to ordinary wear and tear, as well as instances wherein the engine itself may experience unusual wear and tear due to external or internal causes which make continued operation of the engine impossible or inadvisable. Engines which are in need of service or repair to return to satisfactory operating condition frequently must be transported to a suitable service or repair facility, which may be located some distance from where the engine was taken out of service. To transport the engine, therefore, steps must be taken to remove the engine from the aircraft on which it is installed for transportation as cargo or it must be transported by the aircraft while still in its installed location. 
         [0005]    Some aircraft have been configured specifically to carry a non-operating gas turbine engine to transport the engine from one location to another, such as to a location where it is needed for operation or to a service or repair facility to be returned to service. A special fixed or removable mounting pylon may be provided for this purpose. Other aircraft may be configured so as to be able to transport a non-operating gas turbine engine in a conventional mounting location. 
         [0006]    When a non-operating aircraft gas turbine engine is carried aloft in an exposed position on the exterior of an aircraft (as opposed to being carried internally as cargo in a transport container), it is exposed to temperature and humidity changes as well as precipitation, dirt, debris, and other contaminants which may reach the core portion of the engine. Due to the possibility of moisture being present in the core portion of the engine after such a journey, a lengthy heating and drying process is normally required before the engine can be serviced or operated. 
         [0007]    Additionally, the non-operating aircraft gas turbine engine may have internal parts, particularly in the core portion of the engine, which have been subject to wear, damage, or contamination such that free rotation (or windmilling) of the engine due to airflow experienced during a non-operating transport operation may cause further wear and/or damage to such parts. 
         [0008]    Accordingly, there remains a need for a method for preparing and transporting non-operating aircraft gas turbine engines which limits exposure, to moisture and contamination, and free rotation, from transporting externally on an aircraft. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0009]    In one aspect, a fairing kit for a gas turbine engine is described. The engine has a core gas turbine engine, a fan rotor, and a plurality of external fan blades attached to the fan rotor and powered by the core gas turbine engine. The core gas turbine engine has an annular splitter for directing a portion of incoming airflow into the core gas turbine engine. The fairing kit comprises: a) a fairing; b) a plurality of fasteners for securing the fairing to the core gas turbine engine; and c) a conformable seal for sealing mating surfaces of the fairing and the core gas turbine engine. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0010]      FIG. 1  is a cross-sectional illustration of an exemplary gas turbine engine assembly; 
           [0011]      FIG. 2  is a partial cross-sectional elevational view of an exemplary gas turbine engine, illustrating an exemplary embodiment of a fairing installed on the engine; 
           [0012]      FIG. 3  is a cross-sectional view of the fairing of  FIG. 2  in a disassembled condition; 
           [0013]      FIG. 4  is a partial view of the exemplary gas turbine engine of  FIG. 1  illustrating an exemplary borescope motoring pad with a coverplate installed; 
           [0014]      FIG. 5  is a partial view of the borescope motoring pad of  FIG. 4  with the coverplate removed; and 
           [0015]      FIG. 6  is a perspective view of an exemplary locking plate suitable for use with the borescope motoring pad of  FIGS. 4 and 5 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0016]      FIG. 1  is a cross-sectional schematic illustration of an exemplary gas turbine engine assembly  10  having a longitudinal axis  11 . Gas turbine engine assembly  10  includes a fan assembly  12  and a core gas turbine engine  13 . Core gas turbine engine  13  includes a high pressure compressor  14 , a combustor  16 , and a high pressure turbine  18 . In the exemplary embodiment, gas turbine engine assembly  10  also includes a low pressure turbine  20 , and a multi-stage booster compressor  32 , and a splitter  34  that substantially circumscribes booster  32 . 
         [0017]    Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disk  26 , the forward portion of which is enclosed by a streamlined spinner  25 . Gas turbine engine assembly  10  has an intake side  28  and an exhaust side  30 . Fan assembly  12 , booster  22 , and turbine  20  are coupled together by a first rotor shaft  11 , and compressor  14  and turbine  18  are coupled together by a second rotor shaft  22 . 
         [0018]    In operation, air flows through fan assembly  12  and a first portion  50  of the airflow is channeled through booster  32 . The compressed air that is discharged from booster  32  is channeled through compressor  14  wherein the airflow is further compressed and delivered to combustor  16 . Hot products of combustion (not shown in  FIG. 1 ) from combustor  16  are utilized to drive turbines  18  and  20 , and turbine  20  is utilized to drive fan assembly  12  and booster  32  by way of shaft  21 . Gas turbine engine assembly  10  is operable at a range of operating conditions between design operating conditions and off-design operating conditions. 
         [0019]    A second portion  52  of the airflow discharged from fan assembly  12  is channeled through a bypass duct  40  to bypass a portion of the airflow from fan assembly  12  around core gas turbine engine  13 . More specifically, bypass duct  40  extends between a fan casing or shroud  36  and splitter  34 . Accordingly, a first portion  50  of the airflow from fan assembly  12  is channeled through booster  32  and then into compressor  14  as described above, and a second portion  52  of the airflow from fan assembly  12  is channeled through bypass duct  40  to provide thrust for an aircraft, for example. Splitter  34  divides the incoming airflow into first and second portions  50  and  52 , respectively. Gas turbine engine assembly  10  also includes a fan frame assembly  60  to provide structural support for fan assembly  12  and is also utilized to couple fan assembly  12  to core gas turbine engine  13 . 
         [0020]    Fan frame assembly  60  includes a plurality of outlet guide vanes  70  that extend substantially radially between a radially outer mounting flange and a radially inner mounting flange and are circumferentially-spaced within bypass duct  40 . Fan frame assembly  60  may also include a plurality of struts that are coupled between a radially outer mounting flange and a radially inner mounting flange. In one embodiment, fan frame assembly  60  is fabricated in arcuate segments in which flanges are coupled to outlet guide vanes  70  and struts. In one embodiment, outlet guide vanes and struts are coupled coaxially within bypass duct  40 . Optionally, outlet guide vanes  70  may be coupled downstream from struts within bypass duct  40 . 
         [0021]    Fan frame assembly  60  is one of various frame and support assemblies of gas turbine engine assembly  10  that are used to facilitate maintaining an orientation of various components within gas turbine engine assembly  10 . More specifically, such frame and support assemblies interconnect stationary components and provide rotor bearing supports. Fan frame assembly  60  is coupled downstream from fan assembly  12  within bypass duct  40  such that outlet guide vanes  70  and struts are circumferentially-spaced around the outlet of fan assembly  12  and extend across the airflow path discharged from fan assembly  12 . 
         [0022]      FIG. 2  is a partial cross-sectional elevational view of the forward portion of an exemplary aircraft gas turbine engine such as depicted in  FIG. 1 . As shown in  FIG. 2 , a fairing  80  has been installed on the core gas turbine engine  13  in place of the spinner  25  and the fan blades  24  which have been removed and stored separately for transportation. In the embodiment shown in  FIG. 2 , the fairing  80  is of two-piece construction, comprising an inner portion  82  and an outer portion  84 . Inner portion  82  and outer portion  84  are joined to one another as well as to the fan rotor  26  by bolts or other suitable fasteners  88  and suitably located mounting holes provided in the fairing components. Mounting holes  27  in the fan rotor  26 , such as may typically be used to fasten the spinner  25  in place, may be utilized to secure the fairing  80  via fasteners  88 . 
         [0023]    Depending upon the physical configuration of the gas turbine engine upon which the fairing  80  is to be installed, the fairing may be of one-piece (i.e., unitary) construction or multi-piece (i.e., two or more pieces) construction. The fairing  80  has a generally streamlined shape, typically of rounded conical or bullet-shaped design, to minimize drag on the engine during transport. The elements of the fairing are sized, shaped, and adapted to suit the characteristics of the particular engine application desired. 
         [0024]      FIG. 3  is a cross-sectional elevational view of the fairing  80 , similar to the view of  FIG. 2  but showing the fairing  80  in a disassembled condition and depicting the relationship of inner and outer portions  82  and  84 , respectively. In the embodiment shown, the outer portion  84  is of annular configuration and surrounds and is secured to the inner portion  82 . Also shown in  FIG. 3  is an annular seal  86 , which may be formed from rubber, foam, plastic, or other conformable material, which abuts and forms a substantially if not fully airtight seal against splitter  34  when the fairing  80  is installed on the core gas turbine engine  13 . Other configurations may be possible wherein the annular seal  86  is integrally formed with the fairing  80 , such as wherein the fairing  80  is formed from a material which is suitably compliant so as to conform to and sealingly engage the splitter  34 . 
         [0025]    For the embodiment shown in  FIGS. 2 and 3 , the fairing  80  is installed after removing and separately storing the spinner  25  and fan blades  24 . The inner portion  82  of the fairing  80  is installed on the fan rotor  26  via mounting holes  27  and a plurality of fasteners  88 . The outer portion of the fairing  84  is then installed over and around the inner portion  82  and is fastened to the inner portion  82  via fasteners  88 . As the outer portion  84  is secured in position, outer portion  84  moves axially toward the core gas turbine engine and the annular seal  86  comes into contact with and sealingly engages the forward portion of the splitter  34 , so as to seal the opening of the booster  32  and prevent introduction of moisture, contaminants, and airflow into the booster  32  and thus into the core gas turbine engine  13 . As such, for the embodiment shown the process of securing the fairing  80  to the core gas turbine engine is a two phase operation wherein the first portion is secured first, followed by the second portion. 
         [0026]    By effectively sealing the front of the booster  32 , the fairing  80  prevents air from flowing through the booster  32  during flight and imparting any rotational forces to the booster  32  (and any other rotating turbomachinery or accessories which share shaft  21  with booster  32 ). Air is also prevented from flowing through the compressor  14 , as well as through both high pressure turbine  18  and low pressure turbine  20 , and thereby avoiding imparting any rotational forces to those components (and any other rotating turbomachinery or accessories connected to shafts  21  or  22 ). 
         [0027]    Additionally, the sealing engagement of the fairing  80  with the splitter  34  coupled with the mechanical engagement of the fairing  80  with the fan rotor  26  (via fasteners  88 ) provides mechanical resistance to rotation for fan rotor  26  and any other components within the core gas turbine engine  13  which are associated with the shaft  21  to which the fan rotor  26  is attached. 
         [0028]    Many aircraft gas turbine engines have a location where a device can be attached to rotate the core gas turbine engine  13  for inspection and/or service operations, such as borescope inspection. This location (shown in  FIGS. 4 and 5 ) is commonly referred to as a borescope motoring or turning pad  90 . Borescope motoring pad  90  is a location where a coverplate  92  can be removed via fasteners  94  (such as bolts, pins, screws, or other fasteners) to expose a socket  96  which is adapted for engagement by a tool (such as a ½ inch square drive ratchet, wrench, or bar) or a motorized drive assembly to rotate the core gas turbine engine  13  through a suitable gearbox assembly. 
         [0029]    To provide additional optional mechanical security against rotation of the core gas turbine engine  13  during a transport operation, a rotor lock  91  of suitable configuration for the particular engine application may be utilized. Such a rotor lock has suitable mounting holes  93 , a seal  95 , and a knob  97  of suitable configuration to engage the socket  96 . The rotor lock  91  may be installed on the borescope motoring pad  90 , engaging the socket  96  and sealing the opening via seal  95 , and secured in place via fasteners  94 , thereby mechanically preventing rotation of the socket  96  and all other rotating parts mechanically coupled to the socket  96 . Rotor locks such as rotor lock  90  may be fabricated of any suitable material, such as metallic or composite materials, having suitable strength to resist turning motion of the socket  96 . 
         [0030]    In the exemplary embodiment, the fairing  80  may be a fiberglass material, a graphite material, a carbon material, a ceramic material, an aromatic polyamid material such as KEVLAR, a thin metallic material such as, but not limited to, titanium, aluminum, and/or a Metal Matrix Composite (MMC) material, and/or mixtures thereof. Any suitable thermosetting polymeric resin can be used in forming fairing  80 , for example, vinyl ester resin, polyester resins, acrylic resins, epoxy resins, polyurethane resins, bismalimide resin, and mixtures thereof. Overall, the material is selected such that an exterior surface of fairing  80  is resistant to wear and or damage that may be caused by foreign objects ingested into gas turbine engine assembly  10 . Alternate fairing configurations may use a thin metal wrap over a composite fairing to protect against such wear or damage. 
         [0031]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.