Abstract:
A method of assembly for a gas turbine engine and nacelle includes the steps of aligning the nacelle assembly and the engine substantially parallel to the rotational axis of the engine and translating the nacelle assembly along the axis to engage a first and a second attachment and securing the first attachment.

Description:
This is a division of application Ser. No. 09/657,266, filed Sep. 7, 2000. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to a ducted gas turbine engine and includes a nacelle assembly which is detachably connected to a ducted gas turbine engine. 
     BACKGROUND OF THE INVENTION 
     Ducted gas turbine engines usually comprise a core engine which drives a propulsive fan assembly. The fan assembly comprises a number of radially extending aerofoil blades mounted on a common hub and enclosed within a generally cylindrical casing assembly. The fan assembly and casing assembly are encircled by a generally annular nacelle assembly which forms the air intake of the engine and is aerodynamically shaped. The nacelle may extend both forward and rearward, relative to the direction of airflow, of the fan assembly. 
     There is a remote possibility with such engines that part or all of one or more of the fan blades could become detached from the remainder of the fan assembly. The occurrence of a part or all of one or more of the fan blades becoming detached from the fan assembly and impacting the casing assembly is hereinafter termed a FBO (fan blade off) event. The casing assembly surrounding the fan assembly is specifically designed to contain the detached blade or blade portion. However, it is important that the nacelle is not damaged during the FBO event as the casing assembly is subject to distortion. It is also important to remove the possibility of further damage to the nacelle, after the FBO event, resulting from vibrations during run down and subsequent windmilling due to the fan assembly being out of balance. Run down being hereinafter defined as the deceleration of axial rotational speed of the engine from the rotational speed at which a fan blade or part of a fan blade has been released and caused safety systems to shut down the engine. Windmilling being hereinafter defined as the axial rotation of the fan assembly arising from air ingressing the engine due to the forward speed of the aircraft after engine shut down. 
     Typically the nacelle assembly may be attached to a component of the engine and/or an engine support pylon assembly with the necessary access to the engine and engine core mounted accessories usually made by either opening fan cowl doors located in the body of the nacelle as described in WO93/02920)or by the nacelle assembly comprising two part-circular portions acting in a clam-shell like manner as described in U.S. Pat No. 5,205,513. Furthermore, the nacelle assembly is commonly attached to the fan casing as described in U.S. Pat No. 4,044,973, with such attachments being required to be particularly robust to maintain attachment after a FBO event. The nacelle designs, in particular the attachment means to the engine and/or pylon, of the prior art herein cited lend themselves to complex and heavy, thus expensive, assemblies. The nacelle assemblies also appear to be prone to damage during a FBO event and subsequent vibrational damage caused by windmilling of the out of balance fan assembly during fly home. 
     SUMMARY OF THE INVENTION 
     It is an object of the present invention to provide a method of assembly for a gas turbine engine and nacelle that is quickly accomplished by aligning and translating the nacelle relative to the engine until one of two attachments are enagaged. 
     According to the present invention there is provided a nacelle assembly adapted for mounting on a ducted fan gas turbine engine comprising a generally annular body having an air inlet and an air outlet, the generally annular body encircling a region of the engine when working in operative association with the engine and has a first attachment means to a rigid member and a second attachment means to a casing assembly on the engine wherein the second attachment means is frangible. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the casing assembly comprises a containment casing and surrounds a fan assembly. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the second attachment means provides support in the radial direction. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine wherein the second attachment means detaches the nacelle assembly from the casing assembly during a FBO event. 
     Preferably a nacelle assembly adapted for mounting on a gas turbine wherein the rigid member is a component of the engine. Alternatively the nacelle assembly is adapted for mounting on a gas turbine engine wherein the rigid member is a component of a pylon structure or an aircraft structure. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the first attachment means provides support for the nacelle in the radial, axial and circumferential directions. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the first attachment means is a releasable attachment. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the annular body comprises a radially outer facing and a radially inner facing defining a space therebetween. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the annular body comprises the outer facing and inner facing joining and extending rearward of the space to form a single skin. Alternatively the nacelle assembly is adapted for mounting on a gas turbine engine wherein the outer facing and inner facing are constructed from sandwich constructions. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the space contains a lightweight core, the lightweight core attached to both the outer facing and the inner facing. Alternatively the nacelle assembly is adapted for mounting on a gas turbine engine wherein the space contains a connector, the connector attached to both the outer facing and the inner facing. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine engine wherein the connector extends substantially in the axial direction. Alternatively the nacelle assembly is adapted for mounting on a gas turbine engine wherein the connector extends substantially in the circumferential direction. 
     Preferably the nacelle assembly is adapted for mounting on a gas turbine wherein the annular body includes an access panel. Alternatively the nacelle assembly is adapted for mounting on a gas turbine engine wherein an engine accessory is operationally located within the space in the annular body. 
     Preferably a method for assembling a nacelle assembly with an engine comprises the steps aligning the nacelle assembly and the engine substantially parallel to the engine rotation axis, translating the nacelle assembly along the axis to engage the first and second attachments, and securing the first attachment. 
     Preferably a method for removing a nacelle assembly from an engine comprises the steps releasing the first attachment, translating the nacelle assembly substantially parallel to the axis of the engine. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     A specific embodiment of the invention will now be described by way of example with reference to the accompanying drawing in which: 
     FIG. 1 is a schematic axial cross section side view of a ducted gas turbine engine in accordance with the present invention. 
     FIG. 2 is a schematic axial cross section side view of the front portion of a ducted gas turbine engine in accordance with the present invention. 
     FIG. 3 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 4 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 5 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 6 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 6A is a schematic axial cross section side view enlargement of a portion of the embodiment, shown in FIG. 6, of a nacelle body construction in accordance with the present invention. 
     FIG. 7 is a schematic cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 7A is a schematic axial cross section side view as shown on FIG. 7 of the embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 8 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 9 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
     FIG. 10 is a schematic axial cross section side view of another embodiment of a nacelle body construction in accordance with the present invention. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1, a ducted gas turbine engine  20  of known general configuration and method of operation, comprises a rotational axis  21  of the engine  20 , an engine core  22  surrounded by a core casing  24  and which drives a propulsive fan assembly  26 . The fan assembly  26  comprises a retention disc  28  with an array of radially extending aerofoil blades  30 . The engine  20  is secured to an aircraft wing (not shown) from the engine core casing  24  by an engine support pylon assembly  32  in known manner. Alternatively, the engine  20  may be mounted to the aircraft structure (not shown) A nacelle assembly  40  encircles a region of the engine  20 . 
     Referring to FIG. 2, a casing assembly  34  surrounding the fan assembly  26  and secured to an annular array of radially extending vanes  36  comprises a containment casing  38  for retention of a blade  30  or a portion of a blade  30  during a FBO event. 
     The nacelle assembly  40  comprises an air inlet  42  and an air outlet  44  and a generally annular body  46  which encircles a region of the engine  20 . In particular the nacelle  40  encircles a region of both the casing assembly  34  and the fan assembly  26  and is extended rearwards for attachment by a first attachment means  48  to a strengthening ring  50 . The first attachment  48  means is made by conventional means as known in the art so as to provide axial, radial and circumferential support to the nacelle assembly  40 . The nacelle assembly  40  is also attached at the front of the casing assembly  34  by a second attachment means  52 , the second attachment means  52  is a frangible attachment  52  of construction as known in the art. The second attachment means  52 , located upstream of the first attachment means  48 , provides support in the radial direction assisting alignment of a gas washed inner nacelle surface  54  and a casing assembly inner surface  56 . The frangible attachment  52  is designed to detach the nacelle  40  and the engine casing assembly  34  during a FBO event. 
     During a FBO event a blade  30  or blade portion  30  is released from the fan assembly  26  and strikes the containment casing  38 , part of the casing assembly  34 , causing the containment casing  38  and the casing assembly  34  to distort their original shape. It is an advantage of the present invention that the nacelle  40  detaches from the engine casing assembly  34  during the FBO event so that the nacelle assembly  40  is not damaged. It is another advantage that the nacelle assembly  40  is no longer attached to the casing assembly  34  after the FBO event as it is not subject to the consequential vibrations arising from the out of balance of the fan assembly during run down and windmilling. 
     The strengthening ring  50  is attached to the engine core casing  24  by a rigid member  58  as known in the art. The nacelle assembly  40  also comprises an acoustic lining  60  configured and implemented as known in the art. The nacelle assembly  40  also comprises an anti-icing means  62  as known in the art. 
     The nacelle assembly  40  is configured to form an annular space  64  radially outward of the casing assembly  34  to accommodate an engine accessory  66 . The annular space  64  also provides a space for the casing assembly  34  to deflect without contacting the nacelle assembly  40  during a FBO event. 
     The nacelle assembly  40  provides an aerodynamic external profile for the engine  20  and an aerodynamic air inlet  42  and air outlet  44  for the propulsive fan assembly  26 . 
     The construction of the nacelle assembly  40  and in particular the annular body  46  is intrinsic to the implementation of the invention. It is intended that the nacelle body  46  is both lightweight and strong. The following descriptions with reference to FIGS. 3 to  10  give details of further embodiments of the annular body  46  in accordance with the present invention. 
     During normal operation of an engine  20  the nacelle assembly  40  carries aerodynamic loads and loads generated from flexural displacements of the engine  20  and/or the pylon assembly  32 . 
     The method for removal of the nacelle assembly  40 , particularly for access to the engine  20 , is by way of releasing the first attachment means  48  and translating the nacelle assembly  40  in a generally forward axial direction relative to the engine  20 . The second attachment means  52  being so arranged as to disengage the nacelle assembly from the casing assembly  34  when the nacelle assembly  40  is translated forward with respect to the engine  20 . Similarly, the method for attachment of the nacelle assembly  40  to the engine  20  is by way of translating the nacelle assembly  40  in a generally rearward axial direction relative to the engine  20  thereby engaging the second attachment means  52  and first attachment means  48 . It is preferable for the first attachment means  48  to be relatively easy and quick to release, such attachment means may be conventional clamps, “V”-blades or latches. 
     Other embodiments of the present invention described hereinafter describe configurations of the nacelle assembly&#39;s  40  annular body  46  which perform the aforementioned load carrying. It is important for the annular body  46  to be lightweight and relatively strong particularly after a FBO event. The annular body  46  is required to remain intact and operational throughout the remainder of the flight of the aircraft (not shown). After a FBO event the annular body  46  is attached only by the first attachment means  48  and is subject to aerodynamic loads and loads generated from flexural displacements of the engine  20  and/or the pylon assembly  32 . The construction of the annular body  46  is therefore required to be lightweight and strong and the following embodiments hereafter of the present invention describe such constructions. 
     In a another embodiment of the present invention referring to FIG. 3, the construction of the nacelle body  46  is generally annular with respect to the rotational axis  21  of the engine  20  and comprises a radially outer facing  68  and radially inner facing  70 . Both the outer facing  68  and the inner facing  70  are relatively thin, strong and stiff and define an internal space  86  therebetween. 
     The configuration of the nacelle assembly  46  is designed to form a annular space  64  (FIG. 2) radially outward of the casing assembly  34  to accommodate engine accessories  66 . This is achieved by discontinuing the internal space  86  in the region of the second attachment means  52  and joining the outer facing  68  with the inner facing  70  to form a single skin  74 . The single skin  74  extends rearward to the first attachment means  48  at the strengthening ring  50 . 
     In a another embodiment of the present invention referring to FIG. 4, the construction of the nacelle body  46  comprises a relatively thin, strong and stiff radially outer facing  68  and radially inner facing  70  generally surrounding a lightweight main core  72  as known in the art as a sandwich construction. The purpose of the main core  66  being to transfer bending shear, torque, compressive and tensile stresses and loads between the outer facing  68  and the inner facing  70 . 
     In another embodiment of the present invention, referring to FIG. 5, an annular body  46  as described with reference to the embodiment shown in FIG. 3 having an access panel  76  located in the single skin  74 . The access panel  76  allowing access to the engine accessory  66  without removal of the nacelle assembly  40 . 
     In another embodiment of the present invention, referring to FIG.  6  and FIG. 6A, an annular body  46  as described with reference to the embodiment shown in FIGS. 3,  4  and  5  comprising items that are common to both, the outer facing  68  and the inner facing  70  are formed from sandwich constructions themselves with an outer sub-facing  78  and an inner sub-facing  80  generally surrounding a sub-core  82 . The outer sub-facing  78  relating to an exterior surface  84  of the annular body  46 . The embodiment described with reference to FIG. 5 may also comprise an internal space  86  rather than a main core  72 . The embodiment described with reference to FIG. 5 may also comprise an access panel  76  as described with reference to FIG.  5 . 
     In another embodiment of the present invention, referring to FIG. 7, an annular body  46  as described with reference to the embodiments shown in FIGS. 3,  5 ,  6 ,  6 A comprising an annular array of webs  88  connecting the outer facing  68  and the inner facing  70 . Each web  88  extends axially to thereby define an array of voids  90  (FIG.  7 A). The webs  88  may extend for the entire axial distance of the void  90  (FIG. 7A) or may extend for a portion of the axial distance of the void  90 , so that the voids  90  are interconnected with each other. 
     In another embodiment of the present invention, referring to FIG. 8, an annular body  46  as described with reference to the embodiments shown in FIGS. 3,  5 ,  6 ,  6 A comprising a substantially annular connector  92  connecting the outer facing  68  and the inner facing  70 . The connector  92  extends substantially radially between the outer facing  68  and the inner facing  70 . 
     In another embodiment of the present invention, referring to FIG. 9, an annular body  46  as described with reference to the embodiments shown in FIGS. 3,  4 ,  5 ,  6 ,  6 A,  7 ,  7 A,  8  comprising extending the internal space  86  rearward in the annular body  46 . For this embodiment it is intended for the internal space to be extended to the region of the first attachment means  48 . The internal space  86  may also comprise a main core  72  or any of the features such as the web  90  or the connector  92 . 
     In another embodiment of the present invention, referring to FIG. 10, an annular body  46  as described with reference to the embodiments shown in FIGS. 3,  4 ,  5 ,  6 ,  6 A,  7 ,  7 A,  8 ,  9  comprising arranging the engine accessory  66  between the outer facing  68  and inner facing  70  of the annular body  46 . With reference to the aforesaid embodiments the engine accessory  66  may be positioned substantially within the internal space  86 , the void  92  or the annular void  94 . 
     Although the present invention has been described with reference to the first attachment means  48  being releasably attached to the strengthening ring  50  the first attachment means  48  may also be attached in a similar manner to any relatively rigid engine  20  component, such as the casing assembly  34 , the annular array of vanes  36  or the rigid member  58 . 
     Suitable materials for the facing  68 ,  70 ,  78 ,  80  and single skin  74 , access panel  76  and the web  88  and the connector  92  include thermoplastics and thermosets (eg. polythene, polycarbonate, polyethersulphone, polyetheretherketone (PEEK), polyvinylchloride (PVC), epoxy resin cured by amines, nylon, polytetraflouroethelene (PTFE)), resins (e.g. Epoxy, polyamides, phenolic, silicone, cyanoacrylates, anaerobics and acrylics), ceramics (e.g. silicon nitride, silicon carbide, glass-ceramics), aluminium alloys (e.g. Al—Cu, Al—Mg, AL—Mg—Si, Al—Zn—Mg, Al—Li), magnesium alloys, titanium alloys and nickel, which may be reinforced with the following materials: glass, aramid, carbon, alumina, silicon carbide. Suitable materials for the main core  72  and the sub-core  82  include expanded plastics (e.g. polyurethane), low density woods, honeycomb structures (e.g. aluminium, paper).