Abstract:
In order to avoid external bulging ( 5 ) to accommodate accessory mechanisms ( 27 ) and gearboxes ( 30 ) to drive these mechanisms ( 27 ) in accordance with the present invention splitter fairings ( 26 ) are located within a bypass duct ( 23 ) of an engine ( 20 ). The bypass duct ( 23 ) is defined between a casing ( 21 ) and compressor/turbine propulsion core ( 22 ). The fairings ( 26 ) are of sufficient dimensions to accommodate the accessory mechanisms ( 27 ) whilst the bypass duct ( 23 ) is appropriately shaped axi-symmetrically to eliminate and balance any blocking effect of these fairings ( 26 ) within the duct ( 23 ) upon air flow ( 24 ). Further fairings ( 26′ ) may be provided to accommodate oil tank reservoirs ( 34 ) as well as filter/heat exchanger mechanisms ( 35 ) for the engine ( 20 ). In such circumstances, a notional elongate cylindrical profile for the engine ( 20 ) is maintained such that a reduced cross section is require for that engine ( 20 ) and so allowing a smaller airframe with better sonic boom signature.

Description:
This application is a continuation of PCT/GB04/02801, filed Jun. 30, 2004. 

   FIELD OF THE INVENTION 
   The present invention relates to aircraft engine arrangements and more particularly to engine arrangements used in relatively high speed or super-sonic aircraft applications. 
   BACKGROUND OF THE INVENTION 
   In a relatively modern aero-gas turbine engine certain accessories, such as the gearbox and electrical starter/generator, are mounted outwardly of the fan casing within the nacelle or airframe within which the engine is embedded. Accessory services, such as oil feed pipes and electrical cables, are routed through fairings spanning across the bypass duct. These fairings do not carry structural loads, but provide an aerodynamic shape around the accessory services. 
   To minimise aerodynamic drag the nacelle or airframe is tightly wrapped around the engine, minimising frontal area. However, one disadvantage is that aerodynamic shape of the airframe or nacelle is compromised with bulging to accommodate the accessories. Clearly, any bulging can be streamlined but by implication will be detrimental to the aircraft drag coefficient due to steeper cowl angles required to clear the engine accessories. For super-sonic aircraft, such bulging is also known to increase the sonic boom of the nacelle. 
   GB744,695 discloses a compact two-circuit gas turbine engine comprising a core engine having, in downstream flow sequence, a compressor, a combustor and a turbine. The core gas flow is turned and directed forwardly to flow through the combustor which is housed in an array of discrete tubes. The engine further comprises discrete bypass flow tubes which are, in a circumferential direction, alternately spaced between the combustor tubes. As the combustor tubes extend only an axial portion of the bypass tubes, engine accessories are housed between the bypass tubes and axially forward of the combustor tubes. Although this engine configuration is shorter by virtue of the reverse flow combustor, it is seriously disadvantaged as reversing the gas flow induces substantial flow energy losses and causes gas flow disruption into the combustor. Further, the circumferentially alternating bypass and combustor tubes mean that for any given air flow through the core engine, not only is there an annular inlet, but also the reversed combustor core flow occupies a substantial portion of what would be, in a modern conventional gas turbine engine, a substantially annular bypass duct. Thus the bypass gas flow is subject to substantial energy losses ingressing, flowing through and egressing the discrete bypass tubes. Thus the frontal area of this engine would be significantly greater than a conventional gas turbine engine having an annular bypass duct and no reversed combustor flow. Furthermore, GB744,695 does not disclose either a fairing spanning across a bypass duct or of mounting accessories within such a fairing. The engine of GB744,695 is not suitable for high-speed or super-sonic flight. 
   SUMMARY OF THE INVENTION 
   In accordance with the present invention there is provided a gas turbine engine comprising a rotational axis, a fan, a core engine surrounded by an outer casing thereby defining a bypass duct, engine accessories and a fairing, the fairing extends generally radially between the core engine and the outer casing characterised in that the engine accessories are housed within the fairing. The accessories are drivingly connected to the core engine via a drive shaft. 
   The accessories comprise a gearbox and drivingly mounted thereon other accessories. Preferably, the accessories are arranged substantially axially with respect to the engine rotational axis and the other accessories are arranged substantially axially along the gearbox to minimise cross sectional area of the fairing. 
   Alternatively, the accessories are arranged substantially perpendicular with respect to the engine rotational axis. 
   Alternatively, the accessories are angled between perpendicular and parallel to the engine rotational axis. 
   Preferably, the other accessories are arranged with respect to their size to define an aerodynamic shape of the fairing. 
   Preferably, at least two fairings are provided and where and where at least two fairings are provided a conventional annular array of guide vanes is advantageously not required. 
   Preferably, the fairing is capable of transferring engine loads between the core engine and the outer casing, the structural loads comprise any one or more from the group thrust, lateral, vertical or torsional loads. Herein, the fairings are curved and arranged to straighten the bypass air flow from the fan. 
   Preferably, the engine is surrounded by a nacelle to minimise aerodynamic drag. 
   Preferably, the fairings are in aerodynamic balance across the engine. 
   Preferably, at least one casing is adapted for airflow normalisation across the bypass duct and such adaptation is by barrelling of the at least one casing. 
   Preferably, the fairings and/or a gearbox casing provide heat shielding for the accessory mechanisms. 
   Alternatively, the fairings accommodate an oil tank and/or fuel oil heat exchangers. 
   Preferably, a portion of the bypass duct is movable to allow access to the fairing. 
   Alternatively, an access door is provided in the casing and an access door is provided in the fairing ( 26 ). 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     Embodiments of the present invention will now be described by way of example with reference to the accompanying drawings in which: 
       FIG. 1  is a schematic side section of a prior art gas turbine engine mounted with in a nacelle; 
       FIG. 2  is a schematic side view of a possible high speed aircraft turbine engine configuration upon a wing; 
       FIG. 3  is a schematic longitudinal cross-section taken through the Horizontal centre line of an aircraft turbine engine arrangement in accordance with the present invention; 
       FIG. 3   a , is a section through a fairing along A-A of  FIG. 3 ; 
       FIG. 4  is a schematic front section of an aircraft turbine engine arrangement in accordance with the present invention; 
       FIG. 5  is a schematic section along B-B in  FIG. 4  of a fairing in accordance with a further embodiment of the present invention; 
       FIG. 6  is a schematic front section taken through the Horizontal centre line of an aircraft turbine engine arrangement in accordance with the present invention. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   With reference to  FIG. 1 , a conventional prior art ducted fan gas turbine engine generally indicated at  10  has a principal and rotational axis  11 . The engine  10  comprises, in axial flow series, an air intake  12 , a propulsive fan  13 , and a core engine  8  itself comprising an intermediate pressure compressor  14 , a high-pressure compressor  15 , combustion equipment  16 , a high-pressure turbine  17 , and intermediate pressure turbine  18 , a low-pressure turbine  19 ; the engine  10  further comprises an exhaust nozzle  20 . A nacelle  21  generally surrounds the engine  10  and defines both the intake  12  and the exhaust nozzle  20 . 
   The gas turbine engine  10  works in a conventional manner so that air entering the intake  12  is accelerated by the fan  13  to produce two air flows: a first air flow into the core engine  8  and on through the intermediate pressure compressor  14  and a second air flow which passes through a bypass duct  22  to provide propulsive thrust. The intermediate pressure compressor  14  compresses the air flow directed into it before delivering that air to the high pressure compressor  15  where further compression takes place. 
   The compressed air exhausted from the high-pressure compressor  15  is directed into the combustion equipment  16  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines  17 ,  18 ,  19  before being exhausted through the nozzle  20  to provide additional propulsive thrust. The high, intermediate and low-pressure turbines  17 ,  18 ,  19  respectively drive the high and intermediate pressure compressors  15 ,  14  and the fan  13  by associated interconnecting shafts  23 ,  24 ,  25 . 
   The fan  13  is circumferentially surrounded by a structural member in the form of a fan casing  41 , which is supported by an annular array of outlet guide vanes  9  spanning between a casing  39  that surrounds the core engine  8 . 
   The engine  10  further comprises a gearbox/generator assembly  28  used for engine start up and for generating electricity once the engine has been started and working in convention fashion. The generated electricity is used for engine and associated aircraft electrical accessories as well known in the art. The gearbox/generator assembly  28  is drivingly connected to the high-pressure shaft  24  via drive means  35 , however, in other embodiments the gearbox/generator assembly  28  may be driven by any one or more of the shafts  24 ,  25 . In this embodiment, the gearbox/generator assembly  28  comprises an internal gearbox  29  connecting a first drive shaft  30  to the high-pressure shaft  23 , an intermediate gearbox  31  connecting the first drive shaft  30  to a second drive shaft  32  and an external gearbox  33  drivingly connected to the second drive shaft  32 . The external gearbox  33  is drivingly connected to a generator  34  that is capable of the aforesaid operation. The generator  34  and external gearbox  33  are mounted on the fan casing  41  and housed within the nacelle  21 . The first drive shaft  30 , intermediate gearbox  31  and the second drive shaft  32  are housed within a bypass duct splitter fairing  40 . 
   With reference to “The Jet Engine” book, Rolls-Royce plc, 1986 5 th  edition, ISBN 0902121235, pages 66-71, not only does the gearbox  33  drive the starter and generator  36 , but also drives other accessories such as numerous pumps. Conventionally, the gearbox  33  and driven accessories ( 36 ) are arranged circumferentially about the fan casing  41  and generally at the bottom of the engine  10 . 
   Other engine accessories  36 , as known in the art, are also mounted on the fan casing  41 . 
   Generally, a turbine engine includes a number of rotating compressor blades  13 ,  14 ,  15  and turbine blades  17 ,  18 ,  19  arranged about a common axis  11 . In such circumstances notionally a turbine engine is cylindrical. Thus, the base shape for a turbine engine is a longitudinal cylinder and any accessory mechanisms  28 ,  36  to that base cylindrical shape will protrude externally. With regard to high speed aircraft, its aerodynamic profile and envelope is highly important with regard to drag coefficient as well as sonic boom/noise. In such circumstances, previous protrusions and bumps caused by gearbox and accessory mechanisms to the base engine cylindrical profile cause problems when attempting to minimise aerodynamic profile. 
     FIG. 2  illustrates a typical high speed engine arrangement on a wing  2  of an aircraft  3 . As can be seen, the wing  2  is associated with a turbine engine  10 . With high speed and potentially supersonic speed a conventional pitot type nacelle intake is unsuitable due to the severity of shock wave formation and therefore progressively reduced intake efficiency is experienced as inlet air speed increases. Thus, at high speeds so-called external/internal compression intake configurations, where supersonic airflow in to the intake is substantially reduced to subsonic to match the engine compressor requirements, are preferred. This type of intake arrangement, as shown in  FIG. 2 , produces a series of mild shockwaves without excessively reducing the compressor intake efficiency. 
   In order to reduce aerodynamic drag the fan diameter is kept to a minimum, often resulting in a relatively long engine length. The relatively long and thin aspect of the engine  10  is compromised by the requirement to provide accessory mechanisms within the nacelle  21  which results in at least one protrusion bulge  5 , in this example, below the engine  10 . This bulge  5 , although aerodynamically smoothed, still intensifies the aerodynamic drag co-efficient as well as causing increased sonic boom intensity. 
   Ideally, the engine profile within a nacelle should be minimised in order to achieve as low an aerodynamic drag coefficient as possible as well as at high speeds reduce environmental noise problems with respect to sonic boom. The present invention relates to an arrangement of the engine wherein the accessory mechanisms are incorporated within the base cylindrical profile of the engine, thus significantly reducing aerodynamic drag helping to minimise sonic boom. 
   Referring now to  FIGS. 3 and 4  it can be seen that the generally cylindrical profile of the nacelle or casing  21  of the engine  10  is maintained whilst accessory mechanisms are arranged within that profile. The engine  10  is generally configured as described with reference to  FIG. 1 , however those differences attributed to the present invention will now be discussed. 
   In accordance with the present invention, a fairing  26  is provided and located within the bypass duct  22  which house accessory mechanisms  27 . These accessory mechanisms  27  include the gearbox/generator  28  as well as other accessories  36  such as pumps for oil, fuel, dedicated airframe electrical generators and hydraulic actions. The gearbox  28  is now substantially axially aligned (to axis  11 ) and each of the driven accessories  36  is also substantially axially aligned within the fairing  26 . Thus the axes of rotation of the accessories  36  driven from the gearbox  28  are substantially normal to the engine axis  11 . 
   Although it is preferable for the gearbox and accessories to be aligned generally parallel to the axis  11 , it is also possible to align them generally perpendicular or even at an angle between parallel and perpendicular. An advantage of this is that the drive arm  54  engages the gearbox  28  at an advantageous and desired angle ( FIG. 3 ) depending on where the drive arm  54  engages the core engine  8  and where the gearbox  28  is mounted within the fairing  26 . 
   The fairings  26  are located within the general cylindrical profile of the engine  10  and do not create protrusion bumps  5  as described with reference to  FIG. 2 . In contrast with the prior art arrangement, the present invention allows a more cylindrical nacelle profile, which significantly reduces aerodynamic drag and/or reduce the aircraft&#39;s sonic boom signature. 
   The accessory mechanisms  27  are coupled to provide their necessary function in accordance with known procedures. 
     FIG. 3   a , shows a preferred arrangement of accessories  28 ,  36  within the fairing and the profile of the fairing  26  itself. The gearbox  28  is radially inwardly positioned to the accessories  36 . The gearbox  28  is drivingly connected to the core engine  8  via drive shaft  54  and is generally axially aligned and arranged, thereby presenting the least area to the bypass flow stream. Each accessory  36  driven through the gearbox  28  is positioned such that the size of each accessory  36  conveniently defines an aerodynamic profile for the fairing  26 . Such an arrangement of the accessories  28 ,  36  is particularly advantageous in minimising the amount of blockage in the bypass duct  22 . 
   It should be appreciated that at least one other fairing  26 ′ may be incorporated into the engine, the fairing including other accessories  27 ′. 
   Conventionally, the annular array of guide vanes  9  ( FIG. 1 ) is capable of transmitting structural, aero and dynamic loads between the core engine  8  and the outer fan casing  41  and there to aircraft mounting architecture  58  ( FIG. 4 ). A further advantage of the present invention is that the fairings  26 ,  26 ′ are designed to carry structural, aero and dynamic engine loads. For the present invention at least some of the guide vanes  9  may be replaced by the fairings  26 ,  26 ′ although it is possible that the entire array of guide vanes is replacable particularly where more than one fairing  26 ,  26 ′ is provided. 
   In this case ( FIGS. 4 and 6 ) the fairings  26 ,  26 ′ are rigidly connected between the core casing  39  and fan casing  41  or casing  21 . The fairings  26 ,  26 ′ comprise a rigid box-like structure  60  capable of carrying thrust, vertical and horizontal loads as well as torsional engine loads. It should be appreciated by one skilled in the art that many different structural forms are possible, but such alternative forms are to be readily understood to be a means to transfer engine loads between the core engine  8  and the fan casing  41 . The fairings  26 ,  26 ′ are therefore rigidly connected to the outer casing  41  or  21  and the core casing  39 , each casing being substantially annular and inherently very stiff. As the fairings  26 ,  26 ′ extend axially a relatively long way compared to the prior art fairing  40  ( FIG. 2 ) additional benefits of enhanced core engine stiffness are realised. Such benefits include improved control of blade tip clearances and therefore improved economy. 
   Referring to  FIG. 5 , it is well known that the guide vanes  9  are also provided to straighten the flow of bypass air, issuing from the fan  13 . In a further embodiment of the present invention and as an additional advantage, the fairings  26 ,  26 ′ are also curved to accomplish similar bypass air flow straightening. 
   Referring now to  FIG. 6 , the present invention allows the accommodation of the accessory mechanisms  27  within the fairing  26  but it would be understood that the fairings  26 , being arranged within the bypass duct  22 , would cause turbulence, blocking and non uniformity in the airflow  90 . In such circumstances, internal shaping within the generally concentric relationship between the nacelle casing  21  and the core casing  39  to form the bypass duct  22  is configured to act to control airflow  90  for efficient engine  10  operation. The internal shaping involves barrelling of the concentric relationship between the casing  21  and a core casing  39  to effectively vary the bypass duct  22  radial extents at different circumferential positions to limit the effect of introduction of fairings  26  within that duct  22 . This barrelling comprises the radial extent of the bypass duct  22  varying between  44  and  43  positioned generally away from the fairings  26 ,  26 ′ and immediately adjacent the fairings  26 ,  26 ′ respectively. Radial extent  43  is greater than extent  44 . 
   This barrelling is produced either by shaping of the casing  21 , the core casing  39  or both the casings  21 ,  39  being barrelled. Where the fairings  26 ,  26 ′ vary in circumferential width in the downstream direction, due to the varying size of accessories housed therein, the degree of barrelling is also varied to maintain a constant or otherwise desired air flow  90  cross sectional profile. It should be appreciated that the amount of barrelling is relatively small and that the external profile of the nacelle is maintained cylindrical as hereinbefore advantageously described. 
   It will also be understood that although there would be greater detrimental effects with regard to airflow  90  and conventional (centre-line horizontal) orientation of the accessories it may be possible to provide three fairings in a 120° relationship or even four fairings with a spacing 90° from each other. Alternatively, the fairings in accordance with the present invention may be unbalanced in terms of the blocking cross-section with such asymmetric variations accommodated by varying in the bypass duct cross-section or otherwise. 
   In addition to fairings  26  which accommodate accessory mechanisms  27  it will also be understood that fairings  26  ( FIG. 4 ) could be included which simply act as lubricant oil reservoir tanks  94  or accommodate oil filters  35  or provide appropriate positioning for heat exchangers  45  for oil or fuel cooling. With any high maintenance features it is desirable to locate them close to dedicated cut-out panels or access panels  50  in the engine casings. 
   It will be appreciated that the core engine  8 , incorporating the combustor  16 , turbines  17 ,  18 ,  19  and other devices, will become relatively hot. In such circumstances, the fairings  26  incorporate appropriate shielding means  91 ,  52  of the accessory mechanisms  27  from core  8  temperatures. In one embodiment this is achieved by using the gearbox casing  91  and seals with the core casing  39  and the fairings  26  to shield the engine accessories in a separate zone. Nevertheless, it will be appreciated that the airflow  90  through the ducts  22  will itself provide cooling of the fairing  26  and this in turn should limit heating problems with regard to the accessory mechanisms held within the fairings  26 . 
   Generally, the mechanisms  27  held within the fairings  26  will be coupled to the propulsive power of the core engine  8  through adjacent gearbox  28 . Thus, respective radial drive arms  54  ( FIG. 4 ) from the core  8  drive these gearboxes  28  and so the accessory mechanisms  27  within the fairings  26 . Alternatively, the accessories  27  could each be driven individually by an electric motor  56 , rather than an engine powered radial drive. 
   Turbine engine  10  operations would be in accordance with normal procedure except that the fairings  26  allow location of accessory mechanisms  27  within the normal engine  10  cowling profile. In short, the accessory mechanisms  27  are located within the fairings  26  which span the bypass duct  22 . Air flow  90  is maintained by appropriate asymmetric shaping and barrelling of the duct  22  to alleviate the blockage caused by the fairings  26 . In such circumstances, even with such barrelling of the casing  21  the engine  10  has a smaller diameter than previously set by other considerations, (e.g. fan blade-off deflection profiles or engine pipework or ducting routed between the nacelle cowl  37  and engine casings  21 ). This diameter dictates the minimum nacelle size. 
   Clearly, it will be necessary to maintain the gearbox  28  as well as the accessory mechanisms  27  held within the fairings  26 . In such circumstances, access to these fairings  26  and mechanisms  27  is made through dedicated access doors  50 . These access doors  50  are disposed within and form part of the casing  21  structure which defines the bypass duct  22 . The doors  50  are in the form of hinged ducts portions rotatably mounted about the nacelle casing  21 . Alternatively the access doors  50  are removable. The access doors  50  provide improved duct  22  stiffness during flight whilst fixed sections  41   a  of the casing  41  provide structural strength for support of the engine components (e.g. thrust reverser/variable nozzle). The doors  50  provide access during maintenance activities to the fairing  26  housed accessories as well as the core engine  8  components. 
   With reference to  FIG. 6 , an alternative access configuration comprises a movable portion  21   a  of the nacelle  21 , a movable panel  62  of the casing  41 , a movable panel  64  of the fairing  26  for access to the accessories  27  and a movable panel  66  of the core casing  39 . Although described as movable all these access panels  21   a ,  41 ,  62 ,  64  may be rotatably mounted or removable and securable by mechanisms as known in the art. 
   With a more desirable engine  10  profile, and therefore airframe or nacelle cross-section in which that engine  10  is located, it will be appreciated that there may be a reduced sonic boom signature compared with conventional previous high speed aircraft turbine engine arrangements. In addition, avoiding the detrimental aerodynamic effects of an exterior bulge causing increased cowling or casing drag should improve aircraft performance. Furthermore, if there is any bulging for airflow uniformity it will be spread laterally rather than vertically, that is to say across the air frame, fuselage or wing. A more regular engine  10  profile consistent with a base cylindrical shape allows necessary cross sectional area of the nacelle  21  formed around the engine  10  to be reduced which in turn allows the airframe fuselage profile to be defined within accepted aircraft design rules but with consequent reduction in sonic boom intensity with particular advantages for supersonic flight.