Patent Publication Number: US-2017369179-A1

Title: Gas turbine engine

Description:
FIELD OF THE INVENTION 
     The present invention relates to a gas turbine engine, particularly to a gas turbine engine suitable for use on an aircraft, and an aircraft comprising a gas turbine engine. 
     BACKGROUND TO THE INVENTION 
     With reference to  FIG. 1 , a gas turbine engine is generally indicated at  10 , having a principal and rotational axis  11 . The engine  10  comprises, in axial flow series, an air intake  12 , a propulsive fan  13 , 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  and an exhaust nozzle  24 . A nacelle  21  generally surrounds the engine  10  and defines both the intake  12  and a bypass exhaust nozzle  20 . 
     The gas turbine engine  10  works in the conventional manner so that air entering the intake  12  is accelerated by the fan  13  to produce two air flows: a first air flow A into the intermediate pressure compressor  14  and a second air flow B which passes through a bypass duct defined by an internal space between a radially inner side of the engine nacelle  21  and a radially outer side of a core nacelle  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  24  to provide additional propulsive thrust. The high  17 , intermediate  18  and low  19  pressure turbines drive respectively the high pressure compressor  15 , intermediate pressure compressor  14  and fan  13 , each by suitable interconnecting shafts. The compressors  14 ,  15 , combustor  16  and turbines  17 ,  18 ,  19  define an engine core, and are housed within the core nacelle  22 . The core nacelle  22  defines a core inlet  23  at an axially forward end, and a core exhaust  24  at an axially rearward end. 
     A figure of merit for gas turbine engines is the “bypass ratio”, i.e. the ratio of air mass flow which bypasses the core, relative to the mass flow which flows through the core. In general, at subsonic and transonic speeds, higher bypass ratios result in higher propulsive efficiency, and therefore lower specific fuel consumption. Large bypass ratios imply a large fan diameter for a given overall thrust. Such engines are typically installed on aircraft on pylons located underneath, or slightly forward of the wing. Consequently, in order to provide adequate ground clearance with a high bypass ratio engine installed beneath a wing, long landing gear legs are required, which results in high weight, and may make egress from the aircraft in the event of an emergency difficult, in view of the long distance between the aircraft fuselage and the ground. U.S. Pat. No. 8,402,740 describes one such prior example. 
     It is also desirable to increase the thermal efficiency of the gas turbine engine. It is known to provide one or both of intercooling and recuperation to increase the thermal efficiency. Intercooling arrangements comprise a heat exchanger having a hot side in thermal contact with compressed air, upstream of further compression stages, and a cold side in thermal contact with a cold sink such as bypass flow. For example, the hot side may be located between an outlet of a booster or intermediate pressure compressor, and an inlet of a high pressure compressor. By reducing the temperature of the compressed air prior to further compression, the work required to compress the air further is reduced. Similarly, a recuperator arrangement comprises a heat exchanger having a hot side in thermal contact with an area downstream of the engine combustor (such as downstream of the final turbine stage), and a cold side in thermal contact with a combustor inlet. Consequently, waste heat is recycled into the engine, thereby increasing thermal efficiency. However, such systems add weight and complexity to engines, and are difficult to package in the limited space available. 
     The present invention seeks to provide an aircraft gas turbine engine which overcomes or ameliorates some or all of the above problems. 
     SUMMARY OF THE INVENTION 
     According to a first aspect of the present invention, there is provided an aircraft gas turbine engine comprising: 
     first and second non-coaxial propulsors, each propulsor being driven by a common gas turbine engine core comprising a propulsor drive turbine arranged to drive the first and second propulsors via a propulsor drive coupling; 
     wherein the core comprises a first core module comprising a first compressor and a first turbine interconnected by a first shaft, and a second core module comprising a second compressor and the propulsor drive turbine interconnected by a second shaft, the first and second core modules being axially spaced. 
     Advantageously, such an arrangement provides a gas turbine engine having a high bypass ratio with a relatively small fan diameter, thereby permitting installation underneath an aircraft wing. At the same time, the engine has multiple compressors operating at their respective ideal speeds driven by separate shafts and separate, without requiring a shaft interconnecting the fan drive turbine and the fan which passes through the centre of the other shaft. Consequently, shaft lengths can be reduced, thereby reducing vibration and weight, while disc diameters can also be reduced. 
     Each propulsor may comprise a ducted fan or an un-ducted propeller. 
     The core may comprise a compressor provided axially rearwardly of the propulsor drive turbine. 
     The propulsor drive coupling may be arranged such that the propulsor drive turbine rotates at a higher rotational speed than the propulsor in use. The propulsor drive coupling may comprise a mechanical gearbox. 
     The gas turbine engine gearbox may comprise a common input shaft coupled to the propulsor drive turbine, and first and second output shafts coupled to the first and second propulsors respectively. The gearbox may comprise a bevel gear arrangement. The gearbox may comprise a reduction gearbox, such that the input shaft rotates at a higher speed in use than the first and second output shafts. The gearbox may comprise an input:output ratio of between 1 and 5. It is has been found that the present invention is particularly advantageous where the gearbox comprises a reduction gearbox. Reduction gearboxes permit relatively high speed fan drive turbines to be employed, which increases the efficiency of the turbine, while reducing the number of turbine stages that are required, and reducing the diameter of the turbine, thereby reducing the weight and cost of the fan drive turbine. Consequently, the input shaft which interconnects the fan drive turbine and gearbox rotates at a relatively high speed. As a result, the torque carried by the input shaft is relatively low for a given power. This in turn means that a relatively thin, low diameter input shaft relative to the diameter of the core can be employed. Such shafts reduce weight further, but may result in bending or “whirl” modes of vibration. By employing a turbine engine core in which the fan drive turbine is provided as part of a second core module which is axially spaced from a first core module, the fan drive input shaft length is reduced, thereby ameliorating this issue. 
     Alternatively, the propulsor drive coupling may comprise a propulsor drive turbine driven electrical generator and first and second electrical motors coupled to respective propulsors, the generator being electrically coupled to the first and second motors. Consequently, the propulsor drive turbine can be situated remotely from the propulsors, without requiring a relatively heavy mechanical transmission system. 
     The electrical generator may comprise an AC generator, and each electrical motor may comprise an AC motor, the generator and electrical motors being coupled by an AC electrical interconnector. The electrical motors may comprise a power electronics unit configured to modulate the frequency of electrical power delivered to the respective electric motor. Thereby, the speed of each fan can be controlled independently of the speed of the propulsor drive turbines. 
     The first and second propulsors may be co-planar, being provided at substantially the same axial position, though having non-coincident axes of rotation. Each propulsor may comprise a single stage fan, having an outlet guide vane arrangement downstream. The or each propulsor may comprise fixed or variable pitch rotor blades. The gearbox may comprise a differential drive. Where the gearbox comprises a differential drive, and the first and second fans have variable pitch, the fan speed of the first and second fans can be controlled independently by differentially controlling the pitch of the fans, and thereby differentially controlling the load on the respective output shafts. At the same time, fan operating margin can be controlled using the fan pitch. Furthermore, where the fan pitch mechanism is configured to provide reverse pitch, reverse thrust can be provided, without the requirement for further thrust reverser arrangements. Consequently, reverse thrust, fan operating margin control and differential engine thrust can be achieved using a single actuator mechanism. 
     The engine core may be located between the first and second fans. The engine core may comprise a core inlet configured to receive free stream air from between the first and second fans. 
     The engine core may comprise a low pressure compressor, which may comprise either the first or the second compressor. The low pressure compressor may be coupled to the fan drive turbine by a low pressure shaft. The engine core may further comprise a high pressure turbine which may be coupled to a high pressure compressor by a high pressure shaft, which is independently rotatable relative to the low pressure shaft. The high pressure compressor may comprise one or more axial compressor stages upstream in core flow of one or more centrifugal compressor stages. The high pressure turbine may comprise a two-stage turbine. 
     The low pressure compressor may be provided axially forwardly of the low pressure turbine, and may be provided axially forwardly of the gearbox. 
     Alternatively, the low pressure turbine may be provided axially rearward of the high pressure compressor. The core inlet may be provided downstream of one or both fans, configured to ingest fan air, and may be provided at a downstream end of a fan nacelle. Advantageously, the core inlet provides boundary layer ingestion, thereby reducing nacelle drag. Nacelle drag may be particularly large in the case of a two-fan arrangement, since the nacelle surface area / fan area ratio is increased relative to single fan or coaxial fan arrangements. 
     The gas turbine engine may comprise a compressor intercooler arrangement configured to cool compressed air from an outlet of the low pressure compressor upstream in core air flow of an inlet of the high pressure compressor. The intercooler arrangement may comprise a compressed air duct extending between the low pressure compressor outlet and the high pressure compressor inlet. The intercooler arrangement may further comprise a cooling air duct configured to exchange heat between cooling air within the cooling duct, and compressor air within the compressed air duct. 
     The cooling air duct may comprise an inlet configured to ingest freestream air. Alternatively, the cooling air duct may comprise an inlet configured to ingest fan air, downstream of either the first or second fan. The cooling air duct may comprise a flow modulation valve configured to modulate air mass flow through the cooling air duct. Advantageously, where the cooling air duct inlet is configured to ingest fan air flow, and a modulating valve is provided, intercooling can be controlled, whilst simultaneously controlling effective fan outlet area using the same valve. Consequently, fan pressure ratio can be controlled, thereby preventing fan flutter, whilst also controlling the temperature of air delivered to the high pressure compressor. It has been found that at high engine thrust settings at low altitude (for example at takeoff), high intercooling (i.e. a large reduction in compressor air temperature) is required, to control high pressure compressor delivery temperatures. Simultaneously, high fan outlet areas are required to control fan flutter. Consequently, the same valve setting can beneficially affect both parameters. On the other hand, at high altitude, lower thrust conditions, intercooling can be reduced, since lower atmospheric temperature allows higher compressor pressure rise without resulting in higher compressor delivery temperatures, whereas a reduced fan outlet area may increase fan efficiency. Consequently, core temperature control and fan efficiency can be advantageously controlled using a single actuator. 
     The cooling air duct may comprise a first inlet configured to ingest freestream air and a second inlet configured to ingest fan air, the cooling air duct comprising a valve configured to modulate airflow from the first inlet and the second inlet. Consequently, intercooler airflow can be maintained for maximum core thermal efficiency, while controlling fan outlet flow for maximum fan efficiency. 
     The gas turbine engine may further comprise a recuperator arrangement configured to exchange heat between air exhausted from the fan drive turbine and air exiting the first or second compressor prior to entering the combustor. The gas turbine engine may comprise an exhaust duct configured to redirect forward flowing exhaust air from the fan drive turbine to a rearward direction. 
     The recuperator arrangement may comprise a recuperator compressor air duct in thermal contact with a fan drive turbine exhaust duct. The fan drive turbine exhaust duct may be configured to duct fan drive turbine exhaust flow in a rearward direction, and the recuperator compressor air duct may be configured to duct compressor air in a forward direction. Consequently, a reverse flow heat exchange arrangement is provided, which maximises heat transfer in a space efficient manner. 
     According to a second aspect of the present invention there is provided an aircraft comprising a gas turbine engine in accordance with the first aspect of the invention. 
     The first and second propulsors of the gas turbine engine may be located underneath a wing of the aircraft. The core inlet may be provided within the aircraft wing. 
     The core inlet may be configured to ingest air adjacent a trailing edge of the wing, adjacent an upper wing surface. 
     Alternatively, the core may be located underneath the wing. 
     The core may be mounted to a fuselage of the aircraft, and may be mounted at an aft portion of the aircraft adjacent an empennage. The core may be located at an upper surface of the fuselage, with the first and second fans being located at port and starboard sides of the fuselage respectively. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows a prior gas turbine engine; 
         FIG. 2  shows a perspective view from above and to the side from a forward end of a first gas turbine engine in accordance with the present disclosure; 
         FIG. 3  shows a cross sectional top view of the gas turbine engine of  FIG. 2 ; 
         FIG. 4  shows a cross sectional side view of the gas turbine engine of  FIG. 2 ; 
         FIG. 5  shows a schematic view of the gas turbine engine of  FIG. 2 ; 
         FIG. 6  shows a schematic view of a second gas turbine engine in accordance with the present disclosure; 
         FIG. 7  shows a schematic view of a third gas turbine engine in accordance with the present disclosure; 
         FIG. 8  shows a schematic view of a fourth gas turbine engine in accordance with the present disclosure 
         FIG. 9  shows a schematic view of a fifth gas turbine engine in accordance with the present disclosure; 
         FIG. 10  shows a schematic view of a sixth gas turbine engine in accordance with the present disclosure; 
         FIG. 11  shows a schematic view of a seventh gas turbine engine in accordance with the present disclosure; 
         FIG. 12  shows a schematic front cross sectional view of part of an aircraft having the gas turbine engine of  FIG. 2 ; 
         FIG. 13  shows a perspective side view from above and from a forward end of part of the aircraft of  FIG. 12 ; 
         FIG. 14  shows a perspective side view from above and from a forward end of part of an aircraft having the engine of  FIG. 11 ; 
         FIG. 15  shows a perspective side view from above and from a forward end of part of an aircraft having an engine installation of an eighth gas turbine engine; 
         FIG. 16  shows a perspective side view from above and from a forward end of part of an aircraft having an engine installation of a ninth gas turbine engine; 
         FIG. 17  shows a perspective side view from above and from a forward end of part of an aircraft having a still further alternative engine installation for an engine in accordance with any of  FIGS. 2 to 11 ; 
         FIG. 18  shows a perspective side view from above and from a forward end of part of an aircraft having a tenth gas turbine engine in accordance with the present disclosure; 
         FIG. 19  shows a schematic side cross section view of part of an aircraft having the gas turbine engine of  FIG. 6 ; 
         FIG. 20  shows an alternative gearbox arrangement for the gas turbine engine of any of  FIGS. 2 to 11 ; 
         FIGS. 21 a  and 21 b    show an intercooler duct outlet of the gas turbine engine of  FIGS. 2 to 5  in a closed position and an open position respectively; 
         FIG. 22  shows a schematic cross sectional side view of a recuperator of the gas turbine engine of  FIG. 5 ; 
         FIG. 23  shows an axial cross section of the recuperator of  FIG. 22 ; 
         FIG. 24  shows a perspective view from the front and the side of an eleventh gas turbine engine; 
         FIG. 25  shows a perspective view from the front and the side of the gas turbine engine of  FIG. 7 ; 
         FIG. 26  shows a schematic view of the eleventh gas turbine engine of  FIG. 24 ; and 
         FIG. 27  shows a schematic view of a twelfth gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
       FIGS. 2, 3, 4 and 5  show a first gas turbine engine  110  in accordance with the present invention. The engine  110  comprises first and second ducted fans  113   a,    113   b  provided within respective fan nacelles  121   a  and  121   b.  The fans  113   a,    113   b,  and nacelles  121  are provided in a common plane, and rotate about parallel rotational axes  111   a,    111   b,  but are non-coaxial, i.e. they do not occupy the same rotational axis. Alternatively, the fans  113   a,    113   b  could be provided at different axial positions, or could be canted relative to one another. 
     Each fan  113   a,    113   b  provides a propulsive air flow B which flows in an axial direction X, which defines a rearward direction. A forward direction is defined by an axial direction counter to this direction. 
     Each fan  113   a,    113   b  comprises a plurality of fan blades  156 . Each fan blade  156  is pivotable about a radially extending axis by a respective pitch change actuator  163   a,    163   b.  The pitch change actuator may be of conventional construction, comprising for instance a hydraulic system similar to that used in variable pitch propellers. The actuator  163   a,    163   b  is coupled to each fan blade, such that the blades  156  are pivoted together, such that the fan pitch can be altered. A separate pitch change actuator is provided for each fan  113   a,    113   b.    
     The engine  110  further comprises an engine core  175 . Referring to  FIG. 5 , the core  175  comprises a first core module  190  comprising a first compressor in the form of a high pressure compressor  129  and a first turbine in the form of a high pressure turbine  131  interconnected by a first shaft in the form of a high pressure shaft  177 . The core  175  further comprises a second core module  191  comprising a second compressor in the form of a low pressure compressor  128  and a fan drive turbine  143  interconnected by a second shaft in the form of a low pressure shaft  127 . The first and second modules  190 ,  191  are separated in an axial direction X. In this embodiment, each component of the first module  190  is provided rearwardly of each component of the second module  191 . Consequently, though the shafts  127 ,  177  rotate about a common engine axis  111   c,  the shafts do not overlap in an axial direction. 
     The core  175  defines a core airflow path A. Each fan  113   a,    113   b  is driven by a fan drive turbine (described in further detail below) via a fan drive coupling. The fan drive coupling comprises respective first and second output shafts  125   a,    125   b.  Each output shaft  125   a,    125   b  is coupled to the respective fan via a respective output bevel drive  139   a,    139   b  (shown schematically in  FIG. 5 ), and at an opposite end is coupled to a gearbox  126 . The gearbox  126  is driven by an input shaft  127 , and is configured to receive input power from the input shaft  127 , and drive the output shafts  125   a,    125   b  at a lower rotational speed than the input shaft  127 . The gearbox comprises a first toothed gear wheel in form of an input bevel gear  188  coupled to the input shaft  127 . The teeth of the input bevel gear  188  are operably mated to teeth of first and second toothed gears in the form of output bevel gears  181   a,    181   b,  which are in turn coupled to respective first and second output shafts  125   a,    125   b.  The input shaft  127  rotates about a third rotational axis  111   c,  corresponding to a core rotational axis. The output shafts  125   a,    125   b  have rotational axes substantially normal to the input shaft rotational axis  111   c,  with the output shafts  125   a,    125   b  being provided in a common radial plane defining an approximately 90° angle relative to one another. It would be understood though that the output shafts  125   a,    125   b  could be provided at a different angle, depending on the diameter of the engines, their distance from the engine core, and the required spacing between the nacelles  121 . 
     The gearbox  126  and bevel drive  139   a,    139   b  are together configured to provide a reduction ratio, such that the ratio between the input shaft  127  rotational speed and the fan  113   a,    113   b  rotational speeds is approximately 4:1. The gearbox  126  may comprise further toothed gear wheels, and may comprise a planetary or star gearbox configuration. Alternatively, the gearbox may comprise a differential drive, or a continuously variable transmission or belt drive. 
     The input shaft  127  is driven by a common gas turbine engine core comprising at least one fan drive turbine, at least one compressor, and at least one combustor. In this first embodiment, the gas turbine engine core comprises the low pressure compressor  128 , high pressure compressor  129 , a combustor  130 , the high pressure turbine  131  and the low pressure fan drive turbine  143 . 
     Referring to  FIG. 3 , the low pressure compressor  128  comprises a multi-stage axial flow compressor comprising three stages, each stage comprising a respective compressor rotor  132  and stator (not shown). The low pressure compressor  128  is located axially forwardly of the gearbox  126 , and defines a core inlet  133  at an axially forward end. The low pressure compressor  128  is driven by the input shaft  127 , or possibly a separate shaft which is driven at the same speed as the input shaft. Air ingested into the core inlet  133  defines a core flow. In operation, the compressor  128  is configured to ingest air from the core inlet  133 , compress the air, and urge the compressed air in a direction parallel to the axial direction X, i.e. rearwardly. 
     At a rearward end of the low pressure compressor  128  is a low pressure compressor outlet  134 . Air from the low pressure compressor  128  is directed in operation to the low pressure compressor outlet  134 , into an inter-compressor core air duct  135 . The inter-compressor core air duct  135  extends rearwardly toward a rear end of the gas turbine engine core. Surrounding at least part of the inter-compressor core air duct  135  is an intercooler duct  136 . The intercooler duct  136  comprises a hollow passage having an inlet  137  at a forward end configured to ingest freestream air from between the first and second fan nacelles  121   a,    121   b  to define an intercooler airflow C. Consequently, air flow within the intercooler duct  136  and air flow within the inter-compressor duct  135  flow in parallel, and are in thermal contact through the walls of the inter-compressor duct  135 . In view of the temperature difference between the high temperature compressed air within the inter-compressor duct  135  and low temperature ambient air within the intercooler duct  137 , heat is exchanged from the compressed air to the ambient air. Consequently, the intercooler duct  136  and inter-compressor core air duct  135  together form a compressor intercooler  157 , thereby reducing the work required by further compressor stages, and increasing thermal efficiency. 
     The intercooler duct  137  further comprises an intercooler cooling flow modulation valve  138  configured to modulate airflow C mass flow rate.  FIGS. 21 a  and 21 b    show a cross section of the region D of  FIG. 4  in a closed and an open position respectively. As can be seen, the flow modulation valve  138  comprises an axially movable exhaust plug  162 , which is moveable between a closed position (shown in  FIG. 21 a   ) and an open position (shown in  FIG. 21 b   ) by a valve actuator  163  in the form of a hydraulic ram. As will be understood, the plug  162  may be moveable to intermediate positions between the open and closed positions. When in the open position, the airflow C mass flow rate is relatively high, resulting in a large amount of compressor air intercooling. On the other hand, when in the closed position, the airflow C mass flow rate is relatively low, or is shut off completely, such that little or no compressor air intercooling is provided. Consequently, the degree of intercooling can be controlled. 
     The exhaust plug  162  is shaped such that, when in the closed position, the intercooler duct  136  and plug  162  form a continuous surface, which tapers in a rearward direction in a “boat tail” configuration. Consequently, the intercooler duct  136  and plug  162  provide minimal drag when in the closed position. Similarly, a front surface  164  is angled downwardly, such that the plug provides minimal drag when in the open position. The shape of the plug  162  may be such that it uses the Coanda effect to redirect airflow C back towards a rearward direction. 
     The inter-compressor duct  135  comprises an elbow  180  at a rearward, downstream in core flow A end, which redirects core flow A at the downstream end by substantially 180° to a forward direction. The core flow A is thereby directed in operation into an inlet  140  of the high pressure compressor  129 . 
     The high pressure compressor  129  comprises a plurality of axial compressor stages  141  at an axially rearward (i.e. upstream in core flow A) end thereof. Forwardly (i.e. downstream in core flow A) of the axial compressor stages  141  is a centrifugal impellor compressor  142 . Together, the axial and centrifugal compressor stages  141 ,  142  further raise the pressure of the core air flow B in operation, and urge the core air flow B forwardly. 
     Downstream in core flow A is a recuperator compressor air passage  144 , which extends generally radially outwardly from an outlet of the high pressure compressor  129 . The recuperator compressor air passage  144  extends axially forwardly, before returning radially inwardly at a downstream end. The purpose of this passage will be described in further detail below. 
     Axially forwardly (i.e. downstream in core flow A) of the high pressure compressor  129  and downstream in core flow A from the recuperator compressor air passage  144  is the combustor  130 , which is of conventional construction. In the combustor, fuel is provided and burnt with the compressed air in operation to increase the temperature of the core air flow A. 
     The high pressure turbine  131  is provided axially forwardly (i.e. downstream in core flow A) of the combustor. The high pressure turbine  131  comprises first and second stages, each comprising a respective rotor and stator. The high pressure turbine  131  directs flow forwardly, while extracting energy from the flow to drive a high pressure shaft  177 , which is coupled to the high pressure compressor  129 , to thereby drive the high pressure compressor  129  in operation. 
     Axially forwardly (i.e. downstream in core flow A) of the high pressure turbine  131  is the low pressure fan drive turbine  143 , which is of similar construction to the high pressure turbine  131 , comprising a plurality of rotors and stators, but generally comprises more stages than the high pressure turbine  131 . The low pressure fan drive turbine is coupled to both the low pressure compressor  128  and the gearbox  126  via the input shaft  127 . Consequently, the low pressure turbine  143  drives both the fans  113   a,    113   b  and the low pressure compressor  128  via the shaft  127  in operation. 
     Axially forwardly (i.e. downstream in core flow A) of the low pressure turbine  143  is a core exhaust passage  145 , which is configured to receive hot combustion products from a downstream end of the low pressure fan drive turbine  143  in the core air flow A. The core exhaust passage  145  comprises an elbow  146 , which turns core air flow A approximately 180°, and so redirects air rearwardly in use. 
     Downstream of the elbow  146  in core air flow A is a core exhaust recuperator passage  147 , shown in more detail in  FIG. 22 . At least part of the core exhaust recuperator passage  147  encloses the recuperator compressor air passage  144 , such that air within the respective passages  144 ,  147  is in thermal contact. The recuperator compressor air passage  144  extends in a radially outward direction from a diffuser  176  provided at an outlet of the high pressure compressor  129 . The passage  144  extends forwardly through the passage, before turning 180°, and extending rearwardly. The passage may comprise further turns, to increase the length of passage within the core exhaust recuperator passage  147 , to thereby increase the surface area in contact with exhaust air, and so increase heat exchange. The recuperator compressor air passage  144  then extends radially inwardly once more into an inlet of the combustor  130 . Consequently, the passages  144 ,  147  together form a recuperator  148 , since heat is exchanged in operation between exhaust flow and compressed air flow upstream of the combustor  130 . Consequently, thermal efficiency is increased relative to prior arrangements. 
     Referring to  FIG. 23 , a plurality of recuperator compressor air passages  144  are provided, which extend from the diffuser  176 , circumferentially around the core axis. As shown in  FIG. 23 , a separate recuperator compressor air passage  144  is provided for each passage defined by adjacent diffuser vanes  178 . Each compressor air passage  144  communicates with a heat exchanger inlet manifold  179 , before emerging again as separate recuperator compressor air passages. 
     Downstream in core flow A and rearwardly of the core exhaust recuperator passage  147  is a core exhaust nozzle  149 , which exhausts core air flow A into the ambient air flow. 
     Consequently, the above arrangement defines a “reverse flow” architecture, in which core flow A flows in a forward direction, i.e. in an opposite direction to the fan efflux, since at least one core turbine is provided forwardly of at least one core compressor. Both intercooling and recuperation can be provided by this arrangement, increasing thermal efficiency. 
       FIGS. 12 and 13  show the engine  110  installed on an aircraft  150 . The aircraft  150  comprises a fuselage  151  and a pair of wings  152 , one of which is shown in  FIG. 12 . The engine core axis  111   c  is below the wing  152 , below a lower  154  surface. The core inlet  133  is provided 
     The core is located beneath the lower wing surface  154  on a pylon  160 . The aircraft  150  is supported by wing mounted main landing gear  155 . The landing gear  155  has a relatively short length, while providing sufficient ground clearance, in view of the relatively small diameter nacelles  121 . Consequently, the weight of the landing gear can be reduced. This has further beneficial effects on aircraft having wing mounted landing gear, since landing gear weight reductions permit wing structural weight reductions, leading to further overall aircraft weight reductions. Furthermore, with the engine  110  located closer to the ground, maintenance of the engine  110  is simplified, since it can be reached more easily. 
       FIG. 6  provides a schematic illustration of a second gas turbine engine  210  in accordance with the present invention. The engine  210  comprises first and second fans (only one of which is shown in the drawing for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan  213  is similarly driven by respective output shafts  225   a,    225   b  which are in turn driven by a gearbox  226 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  227 . The gearbox  226  is again similar to the gearbox  126 . 
     The fan drive turbine shaft  227  is driven by a core engine, having a different architecture from that of the engine  110 . The engine  210  again comprising first and second core modules  290 ,  291 . The first core module comprises a first compressor in the form of a low pressure compressor  228  and a first turbine in the form of a high pressure turbine  231  interconnected by a first shaft in the form of a high pressure shaft  277 . The core  175  further comprises a second core module  291  comprising a second compressor in the form of a high pressure compressor  229  and a fan drive turbine  243  interconnected by a second shaft in the form of a low pressure shaft  227 . The first and second modules  290 ,  291  are separated in an axial direction X. In this embodiment, each component of the first module  290  is provided rearwardly of each component of the second module  291 . Again, though the shafts  227 ,  277  rotate about a common engine axis  111   c,  the shafts do not overlap in an axial direction. 
     The low pressure compressor  228  is configured to ingest a core air flow A. The inlet for the low pressure compressor  228  is located facing a rearward end of the engine, i.e. facing toward the axial direction X, parallel to the fan flow. Consequently, air flows into the compressor  228  in a forward direction, i.e. counter to the fan flow direction X. The low pressure compressor  228  is of conventional construction, and may comprise a centrifugal compressor, axial compressor, or may comprise both axial and centrifugal stages. 
     An outlet of the low pressure compressor  228  leads to a hot side of an intercooler heat exchanger  257 . A cold side of the intercooler heat exchanger is provided with ambient air flow C. Consequently, the intercooler  257  cools the air compressed by the low pressure compressor  228 , by exchanging heat by ambient temperature air. 
     Downstream in core flow (i.e. forwardly) of the intercooler  257  is a high pressure compressor  229 , which may be of similar construction to the high pressure compressor  229  of the first embodiment. The high pressure compressor  229  is configured to compress the intercooled airflow A, and urge the airflow A rearwardly. 
     Downstream in core flow (i.e. rearwardly) of the high pressure compressor  229  outlet is a cold side of a recuperator heat exchanger  248 . The recuperator heat exchanger is configured to raise the temperature of the core airflow A downstream of the high pressure compressor  229 , and upstream of a combustor  230  by exchanging heat with the exhaust flow, similarly to the embodiment of  FIGS. 2 to 5 . The combustor  230  is provided downstream in core flow A (i.e. axially rearwardly) of the recuperator  248 , and is again of conventional construction. The combustor  230  raises the temperature of core air flow A still further. The core flow A is redirected axially forwardly once again prior to flowing through the combustor  230 . Alternatively, the combustor  230  could comprise a radial combustor  230 , with the core airflow A flowing radially inwardly through the combustor  230  in use. 
     High and low pressure turbines  231 ,  243  are provided in series downstream of the combustor. In this region, the core airflow A is forward flowing. Again, the turbines are of similar construction to those of the first embodiment. Air flowing from an exhaust of the low pressure turbine  243  flows through a hot side of the recuperator heat exchanger  248 , and out of an outlet (not shown). 
     Consequently, the gearbox  226  is provided at a forward end of the engine  210 , with the high pressure compressor  229 , low pressure turbine  243 , high pressure turbine  231 , combustor  230  and low pressure compressor  228  being provided in sequence extending rearwardly. However, the core air A is directed through the low pressure compressor  228 , intercooler  257 , high pressure compressor  229 , combustor  230 , high pressure turbine  231  low pressure turbine  243  and recuperator  248  in flow series. 
     The high pressure compressor  229 , low pressure turbine  243  and gearbox  226  are interconnected by a low pressure fan drive shaft  227 . The gearbox drives the output shafts  225   a,    225   b.  A high pressure shaft  277  interconnects the low pressure compressor  228  and high pressure turbine  231 . In view of this arrangement, the high pressure compressor rotates at the same speed as the gearbox input shaft and low pressure turbine. Similarly, the low pressure compressor rotates at the same speed as the high pressure turbine. However, the shafts are relatively short, which reduces weight. 
     The gearbox  226  may be of similar construction to the gearbox  126 . Alternatively, the gearbox  226  could comprise a differential gearbox  226  as shown in  FIG. 20 . The differential gearbox  226  is located within a housing  265  which encloses a first bevel gear  266 , which is coupled to the fan drive turbine shaft  227 . The input bevel gear  266  engages with a sun gear  267 , which is configured to rotate about an axis normal to the axis of rotation of the bevel gear  266 . The sun gear  267  is mounted on a planetary carrier  268 , which mounts first and second planetary gears  269 ,  270  via respective shafts  271 ,  272 . The planetary bevel gears  269 ,  270  are configured to rotate about their respective axes, which are generally parallel to the fan drive turbine shaft  227  axis, whilst also orbiting about the sun gear  267  axis. Each planetary bevel gear,  269 ,  270  engages against a respective output bevel gear  273 ,  274  located at opposite sides. Each output bevel gear  273 ,  274  is coupled to a respective first and second output shaft  225   a,    225   b.    
     Consequently, input torque from the fan drive turbine shaft  227  drives both output shafts  225   a,    225   b  via the input, sun, planetary and output bevel gears  266 ,  267 ,  269 ,  270 ,  273 ,  274 . In operation, the input bevel gear  266  drives the sun gear  267 , which drives the planet carrier  268 , and the planetary gears  269 ,  270  orbit about the sun gear axis. Where the load on each output shaft  225   a,    225   b  is even, each planetary gear  269 ,  270  does not rotate about its respective axis, but only orbits, and so the output bevel gears  273 ,  274  are driven at equal speed, and the output torque is divided evenly between the output shafts  225   a,    225   b.  On the other hand, where the load on the output shafts  225   a,    225   b  are uneven, the output shaft having the lower load is driven at a higher rotational speed than the other output shaft, since the torque is still applied evenly. 
     As will be understood, the output load on each shaft  225   a,    225   b  will be dependent on the aerodynamic loads on the fans  213 . The aerodynamic loads on the fan will be a function of the fan pitch, fan rotational speed, and forward flight speed. The load can therefore be altered by adjusting one or more of these parameters. 
     In this embodiment, each fan  213  comprises a pitch change actuator  263  configured to collectively alter the pitch of the blades of the fan  213 , and so adjust the angle of attack of the blades  213 . In general, increasing the angle of attack of the blades will result in a higher load (i.e. a higher absorbed power), and so a higher torque for a given rotational speed. Since a differential will generally provide an equal torque to each output shaft, the speed of the fan having the higher load will be reduced. Consequently, by adjusting the pitch of a fan, the speed of that fan can be reduced relative to the other fan. 
     This effect can be beneficially employed for various applications. For example, the fan speed relative to the fan load can be adjusted, which may help control surge margin or flutter of a single fan. If the fan rotational speeds were fixed relative to one another, a reduction in speed of one fan would necessitate a reduction in speed of the other, which would be non-optimal, since inlet flows to each fan may be different in view of their different locations on the aircraft, and so the fans may have a different surge margin at given rotational speeds. Furthermore, this effect could be employed to rapidly adjust the thrust provided by each fan relative to the other, since fan thrust is related to fan rotational speed. Since the fans are non-coaxial, each fan may have a different lever arm relative to the aircraft. Consequently, increasing thrust on one engine alone may produce a pitch, yaw or roll force. Since pitch change can be actuated relatively quickly, such a system could be employed to provide additional aircraft control, thereby reducing the use of control surfaces, and minimising drag. 
       FIG. 19  shows a schematic side cross sectional view of an aircraft installation for the engine  210 . The fans  213 ,  213  are provided within nacelles  221  located on pylons  260  provided underneath a lower side  254  of a wing  252 , adjacent a leading edge  256 . Since the low pressure compressor  228  is provided at the rear of the engine  210 , an intake scoop  258  which provides air to the low pressure compressor  228  can be located adjacent a trailing edge  259 , and can thereby be supplied with air from a rearward region of a wing upper surface  253 , thereby energising the boundary layer thereon, and reducing airframe drag. Consequently, the fuel efficiency of the aircraft as a whole may be improved. 
       FIG. 7  shows a schematic illustration of a third gas turbine engine  310  in accordance with the present invention. Again, the engine  310  comprises first and second fans  313   a,    313   b  of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan  313   a,    313   b  is similarly driven by respective output shafts  325   a,    325   b  which are in turn driven by a gearbox  326 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  327 . The gearbox  326  is again similar to the gearbox  126 . 
     Again, the engine  310  comprises a core engine  376  which differs in this embodiment. The core engine  375  comprises first and second core modules  390 ,  391 . The first core module  390  comprises a first compressor in the form of a low pressure compressor  328  and a first turbine in the form of a high pressure turbine  331  interconnected by a first shaft in the form of a high pressure shaft  327 . The core  375  further comprises a second core module  391  comprising a second compressor in the form of a high pressure compressor  329  and a fan drive turbine  343  interconnected by a second shaft in the form of a low pressure shaft  327 . The first and second modules  390 ,  391  are separated in an axial direction X. In this embodiment, each component of the first module  390  is provided forwardly of each component of the second module  391 . Again, though the shafts  327 ,  377  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     The low pressure compressor  328  is provided axially rearwardly of the gearbox  326 , and is configured to ingest fan air, i.e. air directly downstream of the fan  313   a,    313   b  from a forward direction, compress this air, and urge this in a rearward direction. Rearwardly, and downstream in core flow A, is an intercooler  357  of similar construction to the previous embodiments, which is arranged to cool compressed air using freestream airflow C. Downstream of the intercooler  357  is the high pressure compressor  329 , which further compresses the air, and redirects the air forwardly. This compressed air is then passed through a recuperator heat exchanger  348  which further heats the air using exhaust gas heat as in previous embodiments, before being passed to a combustor  330 . Prior to entry to the combustor, the core flow A is again redirected rearwardly. After passing through the combustor in a rearward direction, the core flow A is passed rearwardly through the high and low pressure turbines  331 ,  343  in series, before passing through the hot side of the recuperator  348 . The low pressure compressor  328 , high pressure turbine  331  and gearbox  326  are interconnected by a fan drive shaft  327 . The gearbox drives the output shafts  325   a,    325   b.    
     Consequently, the gearbox  326  is provided at a forward end of the engine  310 , with the low pressure compressor  328 , combustor  330 , high pressure turbine  331 , low pressure turbine  343 , and high pressure compressor  229  being provided in sequence extending rearwardly. However, the core air A is directed through the low pressure compressor  328 , intercooler  357 , high pressure compressor  329 , combustor  330 , high pressure turbine  331  low pressure turbine  343  and recuperator  348  in flow series. 
       FIG. 25  is a perspective illustration of part of an aircraft  350  comprising the engine  310 . As can be seen, a combined core and intercooler inlet  333  is provided within the nacelle  321   b.  A further core inlet (not shown) is provided in the nacelle  321   a  downstream of the fan  313   b.  The core inlets  333  from both nacelles  321   a,    321   b  combine to form a core inlet passage  380 , which communicates with the low pressure compressor  328 . The intercooler  357  also communicates with the core inlet passage  380 . Consequently, high pressure fan air provides both the core and intercooler flows A, C. In general, the inlet is arranged such that air is fed into the core prior to being heated by the intercooler. Advantageously, such an arrangement increases the overall pressure ratio (OPR) of the core, and so increases thermal efficiency. The effect on the intercooler effectiveness may vary depending on the details of the arrangement, since the mass flow will be increased (thereby increasing heat exchange), while cold stream temperature will be increased (thereby reducing heat exchange). 
       FIG. 8  shows a schematic illustration of a fourth gas turbine engine  410  in accordance with the present invention. Again, the engine  410  comprises first and second fans (omitted in the drawing for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan is similarly driven by respective output shafts (not shown) which are in turn driven by a gearbox  426 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  427 . The gearbox  426  is again similar to the gearbox  126 . 
     The engine  410  comprises a core engine  475  having first and second core modules  490 ,  491 . The first core module  490  comprises a first compressor in the form of a high pressure compressor  429  and a first turbine in the form of a high pressure turbine  431  interconnected by a first shaft in the form of a high pressure shaft  477 . The core  475  further comprises a second core module  491  comprising a second compressor in the form of a low pressure compressor  428  and a fan drive turbine  443  interconnected by a second shaft in the form of a low pressure shaft  427 . The first and second modules  490 ,  491  are separated in an axial direction X. In this embodiment, each component of the first module  490  is provided rearwardly of each component of the second module  491 . Again, though the shafts  427 ,  477  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     The low pressure compressor  428  is provided axially rearwardly of the gearbox  426 , and is configured to ingest freestream air from a forward direction, compress this air, and urge this in a rearward direction. Rearwardly, and downstream in core flow A, is an intercooler  457  of similar construction to the previous embodiments, which is arranged to cool compressed air using freestream airflow C. Downstream of the intercooler  457  is the high pressure compressor  429 , which further compresses the air, and redirects the air forwardly. This compressed air is then passed through a recuperator heat exchanger  448  which further heats the air using exhaust gas heat as in previous embodiments, before being passed to a combustor  430 . Air flows forwardly through the combustor  430 , before being passed forwardly through the high and low pressure turbines  431 ,  443  in series, before then passing through the hot side of the recuperator  448 . Again, the low pressure compressor  428 , low pressure turbine  443  and gearbox  426  are interconnected by a low pressure fan drive shaft  427 . The gearbox drives the output shafts. 
     Consequently, the gearbox  426  is provided at a forward end of the engine  410 , with the low pressure compressor  428 , low pressure turbine  443 , high pressure turbine  431 , combustor  430 , and high pressure compressor  429 , being provided in sequence extending rearwardly. However, the core air A is directed through the low pressure compressor  428 , intercooler  457 , high pressure compressor  429 , combustor  430 , high pressure turbine  431  low pressure turbine  443  and recuperator  448  in flow series. This arrangement provides a shorter shaft  427  between the low pressure compressor  428  and low pressure turbine  443 . However, tighter turning of the airflow downstream of the low pressure compressor  428  is required in view of this. 
       FIG. 9  shows a schematic illustration of a fifth gas turbine engine  510  in accordance with the present invention. Again, the engine  510  comprises first and second fans (omitted in the drawing for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan is similarly driven by respective output shafts (not shown) which are in turn driven by a gearbox  526 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  527 . The gearbox  526  is again similar to the gearbox  126 . 
     The engine  510  comprises a core engine  575  having first and second core modules  590 ,  591 . The first core module  590  comprises a first compressor in the form of a high pressure compressor  529  and a first turbine in the form of a high pressure turbine  531  interconnected by a first shaft in the form of a high pressure shaft  577 . The core  575  further comprises a second core module  591  comprising a second compressor in the form of a low pressure compressor  528  and a fan drive turbine  543  interconnected by a second shaft in the form of a low pressure shaft  527 . The first and second modules  590 ,  591  are separated in an axial direction X. In this embodiment, each component of the first module  590  is provided rearwardly of each component of the second module  591 . Again, though the shafts  527 ,  577  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     Again, the core engine arrangement differs in this embodiment. The low pressure compressor  528  is provided axially rearwardly of the gearbox  526 , and is configured to ingest freestream air from a forward direction, compress this air, and urge this in a rearward direction. Rearwardly, and downstream in core flow A, is an intercooler  557  of similar construction to the previous embodiments, which is arranged to cool compressed air using freestream airflow C. Downstream of the intercooler  557  is the high pressure compressor  529 , which further compresses the air, and directs the air rearwardly through a recuperator heat exchanger  548  which further heats the air using exhaust gas heat as in previous embodiments, before being passed to a combustor  530 . Air continues to flow rearwardly through the combustor  530 , before being passed further rearwardly through a high pressure turbine  531 . Downstream of the high pressure turbine  531 , the core airflow A is redirected forwardly to the low pressure fan drive turbine  543  situated axially between the high pressure compressor  529  and low pressure compressor  528 . Air exhausted from the low pressure turbine is passed through a hot side of the recuperator  548 , thereby heating compressor air prior to entry into the combustor  530 . The gearbox  526  drives the output shafts. 
     Consequently, the gearbox  526  is provided at a forward end of the engine  510 , with the low pressure compressor  528 , low pressure turbine  543 , high pressure compressor  529 , combustor  530 , and high pressure turbine  531 , being provided in sequence extending rearwardly. However, the core air A is directed through the low pressure compressor  528 , intercooler  557 , high pressure compressor  529 , combustor  530 , high pressure turbine  531  low pressure turbine  543  and recuperator  548  in flow series. 
     As can be seen from  FIG. 9 , in this arrangement, the high pressure turbine  529  and compressor  531  are situated adjacent one another, with the low pressure turbine  543  and compressor  528  being situated adjacent one another forwardly thereof. Consequently, the shafts  527 ,  577  do not have to pass through one another, and can be relatively short. Furthermore, the ducting between the low pressure compressor  528  and high pressure compressor  529  is shorter, resulting in lower flow losses, and lower weight. Furthermore, the low pressure compressor  528  and turbine  543  are interconnected, as are the high pressure compressor  529  and high pressure turbine  531 . Consequently, the speeds of these components can be matched. On the other hand, hot high pressure turbine exhaust gasses must be redirected forwardly, and run the length of the high pressure engine core (i.e. the compressor  129 , combustor  520  and turbine  531 ) and the length of the low pressure turbine  543 , before running rearwardly again. 
       FIG. 10  shows a schematic illustration of a sixth gas turbine engine  610  in accordance with the present invention. Again, the engine  610  comprises first and second fans (omitted in the drawing for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan is similarly driven by respective output shafts (not shown) which are in turn driven by a gearbox  626 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  627 . The gearbox  626  is again similar to the gearbox  126 . 
     The engine  610  comprises a core engine  675  having first and second core modules  690 ,  691 . The first core module  690  comprises a first compressor in the form of a low pressure compressor  628  and a first turbine in the form of a high pressure turbine  631  interconnected by a first shaft in the form of a high pressure shaft  677 . The core  675  further comprises a second core module  691  comprising a second compressor in the form of a high pressure compressor  629  and a fan drive turbine  643  interconnected by a second shaft in the form of a low pressure shaft  627 . The first and second modules  690 ,  691  are separated in an axial direction X. In this embodiment, each component of the first module  690  is provided rearwardly of each component of the second module  691 . Again, though the shafts  627 ,  677  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     The low pressure compressor  628  is provided at a mid-section of the engine  610 , axially rearwardly of the gearbox  626 , the high pressure compressor  629 , and the low pressure fan drive turbine  643 , and is configured to ingest freestream air from a forward direction, compress this air, and urge this compressed air in a forward direction. Downstream in core flow A, is an intercooler  657  of similar construction to the previous embodiments, which is arranged to cool compressed air using freestream airflow C. Downstream of the intercooler  657  is the high pressure compressor  629 , which further compresses the air, and directs the air rearwardly through a recuperator heat exchanger  648  which further heats the air using exhaust gas heat as in previous embodiments, before being passed to a combustor  630 . Air continues to flow rearwardly through the combustor  630 , before being passed further rearwardly through the high pressure turbine  631 . Downstream of the high pressure turbine  531 , the core airflow A is redirected forwardly to the low pressure fan drive turbine  643  situated axially between the high pressure compressor  629  and low pressure compressor  628 . Air exhausted from the low pressure turbine  643  is turned rearwardly once more, before being passed through a hot side of the recuperator  648 , thereby heating compressor air prior to entry into the combustor  630 . The high pressure compressor  629 , low pressure turbine  643  and gearbox  626  are interconnected by a fan drive shaft  627 . The gearbox  626  drives the output shafts. A high pressure shaft  677  interconnects the low pressure compressor  628  and high pressure turbine  631 . 
     Consequently, the gearbox  626  is provided at a forward end of the engine  610 , with the high pressure compressor  629 , low pressure turbine  643 , low pressure compressor  628 , combustor  630 , and high pressure turbine  631 , being provided in sequence extending rearwardly. However, the core air A is directed through the low pressure compressor  628 , intercooler  657 , high pressure compressor  629 , combustor  630 , high pressure turbine  631  low pressure turbine  643  and recuperator  648  in flow series. 
     As will be understood, the length between the high pressure compressor  629  exit and the combustor  630  are increased, thereby providing more surface area for heat exchange for the recuperator, thereby increasing its effectiveness. On the other hand, the complex airflow may result in flow losses. 
       FIGS. 11 and 14  show a seventh gas turbine engine  710  in accordance with the present invention. The gas turbine engine  710  has a similar core architecture to that of the engine of  FIGS. 2 to 5 , but with a different fan drive coupling and with a different intercooler arrangement. 
     The engine  710  comprises a core engine  775  having first and second core modules  790 ,  791 . The first core module  790  comprises a first compressor in the form of a high pressure compressor  729  and a first turbine in the form of a high pressure turbine  731  interconnected by a first shaft in the form of a high pressure shaft  777 . The core  775  further comprises a second core module  791  comprising a second compressor in the form of a low pressure compressor  728  and a fan drive turbine  743  interconnected by a second shaft in the form of a low pressure shaft  727 . The first and second modules  790 ,  791  are separated in an axial direction X. In this embodiment, each component of the first module  790  is provided rearwardly of each component of the second module  791 . Again, though the shafts  727 ,  777  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     The low pressure compressor  728  is provided at a forward end of the engine  710 , with the low pressure fan drive turbine  743 , high pressure compressor, a combustor  730 , and the high pressure compressor  729 , being provided in sequence extending rearwardly. The engine  710  is configured such that core air A is directed through the low pressure compressor  728 , high pressure compressor  729 , combustor  730 , high pressure turbine  731  low pressure turbine  743  and a recuperator  748  in flow series. A high pressure shaft  777  interconnects the high pressure compressor  729  and high pressure turbine  731 , and a low pressure fan drive shaft  727  interconnects the low pressure compressor  728  and low pressure turbine  743 . 
     The engine  710  is provided with an intercooler duct  780  configured to receive an intercooler cooling airflow C. The airflow C is provided from an intercooler duct inlet  733  located behind the fans  713   a,    713   b  within the fan ducts  721   a,    721   b.  The duct  736  comprises a valve  738  similar to the valve  138 , which modulates flow through the intercooler duct  780 . However, since the intercooler duct inlet  733  receives air from downstream of the fan  713   a,  modulation of cooling air flow through the duct  780  affect the fan pressure ratio (i.e. reducing flow by closing the valve  738  increases back pressure on the fan, thereby reducing fan pressure ratio, and vice versa). Consequently, the valve  738  can be modulated to control fan pressure ratio, thereby controlling fan surge margin. 
     The engine  710  comprises first and second fans  713   a,    713   b  arranged similarly to the fans  113   a,    113   b  of the first embodiment, and driven by the core by a fan drive coupling. However, the fan drive coupling of the engine  710  differs from that of the previously described embodiments. 
     The fan drive coupling comprises the fan drive shaft  727 , which is coupled to, and thereby drives in use, an electrical generator  762 . Optionally, the generator  762  could be driven by a gearbox, and could be located forward of the low pressure compressor  728 . The electrical generator comprises a alternating current (AC) generator, and is electrically coupled to first and second electrical motors  761   a,    761   b  via respective electrical interconnectors  760   a,    760   b.  The electrical motors  761   a,    761   b  are in turn coupled to respective fans  713   a,    713   b,  and thereby drive them in use. Consequently, the fan drive turbine  727  drives the fans  713   a,    713   b  via the generator  762 , electrical interconnectors  760   a,    760   b  and motors  761   a,    761   b.  As will be understood, the pole numbers of the motors  761   a,    761   b  can differ from those of the generator  762 , thereby providing an effective reduction ratio between the fan drive shaft  727  and fans  713   a,    713   b.  It will be understood that DC generators, motors and interconnectors could alternatively be used. Optionally, power electronics units (i.e. AC to AC converters) can be employed to control the electrical frequency of power provided to the electrical motors  761   a,    761   b,  to thereby control speed of the electrical motors  761   a,    761   b,  and therefore fans  713   a,    713   b  independently of the other motor, and of the fan drive turbine  727  speed. 
     Any of the gas turbine engines  2  to  11  could be mounted to an aircraft in accordance with alternative installation configurations, as shown in  FIGS. 15 to 17 . 
       FIG. 15  shows an engine installation for an aircraft  850 . The installation comprises a gas turbine engine  810 , which has a similar core arrangement to any of the previously described engines, having a core  876  and a pair of fans  813   a,    813   b  installed within respective nacelles  821   a,    821   b.  However, in this instance, the core  876  is installed within the wing  852 , such that a core rotational axis  811   c  extends between upper  853  and lower surfaces  854  of the wing  852 . Consequently, core nacelle drag is minimised, as part of the core  876  is contained within the wing  852 . The nacelles  821   a,    821   b  meanwhile project from a pylon  850  when extends beneath the wing  852 . A shaft may be provided within the wing interconnecting the gas turbine engine fan drive turbine (not shown) and a bevel gearbox provided between the engine nacelles  821   a,    821   b.    
     Alternatively, the engine  810  may comprise an electrical generator and electrical motors, similar to the embodiment shown in  FIG. 11 . 
       FIG. 16  shows an engine installation for an aircraft  950  comprising an engine  910 . The engine  910  may be similar to any of the engines shown in  FIGS. 2 to 11 . In this embodiment, the engine  910  is mounted to an aft end of an aircraft fuselage  951 , with an engine core  976  mounted to an upper surface of the fuselage  951  by a core pylon  981 . The core pylon  981  forms a root portion of a vertical tail  982 , which extends upwardly from the core  976 . 
     The engine comprises a pair of fan nacelles  921   a,    921   b,  which are located adjacent a respective side of the fuselage  951 , below the core  976 . Each nacelle  921   a,    921   b  is mounted by a respective nacelle pylon  983 , through which first and second output shafts (not shown) pass. Since only the core  976  is mounted to the tail, the structural penalties for this installation are relatively small. This installation may be provided in addition to an installation such as that shown in any of  FIG. 12 to 15 or 17 , such that an engine is provided underneath each wing, in addition to a fuselage mounted engine. 
       FIG. 17  shows an engine installation for an aircraft  1050  having an engine  1010 , which again may comprise a core  1076  similar to any of the cores illustrating in  FIGS. 2 to 11 . In this instance, the gas turbine engine  1010  comprises a single core inlet/intercooler inlet  1033 . In this case, the inlet coincides with the core engine rotational axis  1011   c,  and is configured to ingest free-stream air, and deliver this air to both a low pressure compressor and an intercooler duct (not shown). Consequently, the low pressure compressor and intercooler are both supplied with low temperature, undisturbed air. 
       FIG. 18  shows a tenth gas turbine engine  1110  installed on an aircraft  1150 . The engine  1110  comprises an engine core  1176  which is arranged in accordance with any of the architectures of the previous embodiments. However, in place of a fan, each fan drive turbine is coupled, via a gearbox, to a respective first and second propulsors in the form of open rotor propeller arrangements  1183 . The propeller arrangements lack nacelles, instead being provided on a respective pod  1186 . Each open rotor propeller arrangement comprises first and second coaxial contra-rotating propeller stages  1184 ,  1185 , driven by a gearbox (not shown). Open rotors generally enable high bypass ratios, without the aerodynamic and weight disadvantages implied by the large diameter nacelle. However, the bypass ratio is generally limited by other considerations, such as ground clearance, noise of the high tip speed rotors, and safety considerations. Each of these problems is ameliorated by having a pair of rotors driven by a common gas turbine engine core. 
       FIG. 24  shows an eleventh gas turbine engine  1210  installed on an aircraft  1250 . The engine  1210  is arranged as shown in  FIG. 26 . The engine  1210  comprises first and second fans  1213  (only one of which is shown in  FIG. 26  for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan  1213  is similarly driven by respective output shafts  1225   a,    1225   b  which are in turn driven by a gearbox  1226 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  1227 . 
     The gearbox  1226  is again similar to the gearbox  126 . As can be seen in  FIG. 24  in particular, the fans  1213 , are located at a rearward end of the engine  1210 , downstream of the gearbox  1226 , and downstream of a core engine inlet  1233 . Advantageously, the relatively heavy core can be mounted forwardly, which may help to damp any wing flutter. 
     The fan drive turbine shaft  1227  is driven by a core engine  1275  provided within a nacelle, the core engine  1276  having a different architecture from that of the previously described engines. Engine  1210  comprises a core engine  1275  having first and second core modules  1290 ,  1291 . The first core module  1290  comprises a first compressor in the form of a low pressure compressor  1228  and a first turbine in the form of a high pressure turbine  1231  interconnected by a first shaft in the form of a high pressure shaft  1277 . The core  1275  further comprises a second core module  1291  comprising a second compressor in the form of a high pressure compressor  1229  and a fan drive turbine  1243  interconnected by a second shaft in the form of a low pressure shaft  1227 . The first and second modules  1290 ,  1291  are separated in an axial direction X. In this embodiment, each component of the first module  1290  is provided forwardly of each component of the second module  1291 . Again, though the shafts  1227 ,  1277  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     The engine  1210  comprises a low pressure compressor  1228 , which ingests a core air flow A. The core engine inlet  1233  for the low pressure compressor  1228  is located facing a forward end of the engine, i.e. facing oppositely to the axial direction X, and is located forwardly of the fans  1213   a,    1213   b.  Consequently, air flows into the compressor  1228  in a rearward direction, and so no turning is required prior to the airflow entering the compressor  1228 , which may improve efficiency and reduce overall engine length. 
     An outlet of the low pressure compressor  1228  leads to a hot side of an intercooler heat exchanger  1257 . A cold side of the intercooler heat exchanger is provided with ambient air flow C from an intercooler inlet  1237 . The intercooler inlet  1237  is configured to ingest freestream air, and is provided forwardly of the fans  1213   a,    1213   b.  Consequently, the intercooler  1257  cools the air compressed by the low pressure compressor  1228 , by exchanging heat by ambient temperature air. Downstream in core flow of the intercooler  1257  is a high pressure compressor  1229 . The high pressure compressor  1229  is configured to compress the intercooled airflow A, and urge the airflow A forwardly. 
     Downstream in core flow of the high pressure compressor  1229  outlet is a cold side of a recuperator heat exchanger  1248 , similar to those of the previous embodiments. The recuperator heat exchanger is configured to raise the temperature of the core airflow A downstream of the high pressure compressor  1229 , and upstream of a combustor  1230  by exchanging heat with the exhaust flow. The combustor  1230  is provided downstream in core flow A of the recuperator  1248 , and is again of conventional construction. The combustor  1230  raises the temperature of core air flow A still further. The core flow A is redirected axially rearwardly once again prior to flowing through the combustor  1230 . 
     High and low pressure turbines  1231 ,  1243  are provided in series downstream of the combustor  1230 . In this region, the core airflow A is generally rearwardly flowing. Air flowing from an exhaust of the low pressure turbine  1243  flows through a hot side of the recuperator heat exchanger  1248 , and out of an outlet (not shown). 
     Consequently, the gearbox  1226  is provided at a rearward end of the engine  1210 , with the high pressure compressor  1229 , low pressure turbine  1243 , high pressure turbine  1231 , combustor  1230  and low pressure compressor  1228  being provided in sequence extending forwardly. However, the core air A is directed through the low pressure compressor  1228 , intercooler  1257 , high pressure compressor  1229 , combustor  1230 , high pressure turbine  1231  low pressure turbine  1243  and recuperator  1248  in flow series. 
     The high pressure compressor  1229 , low pressure turbine  1243  and gearbox  1226  are interconnected by a low pressure fan drive shaft  1227 . The gearbox drives the output shafts  1225   a,    1225   b.  A high pressure shaft  1277  interconnects the low pressure compressor  1228  and high pressure turbine  1231 . Again, in this embodiment, the shafts  1227 ,  1277  are not concentric, and so the shaft arrangement is simplified, and the engine can have a narrower diameter, which decreases core nacelle drag. Referring again to  FIG. 24 , the engine core  1276  is mounted to a lower surface of a wing (not shown) by a pylon  1260 , and each fan  1213   a,    1213   b  is mounted to the core engine  1276 . 
       FIG. 27  illustrates a twelfth gas turbine engine  1310 . The engine  1310  comprises first and second fans (only one of which,  1313   a,  is shown in  FIG. 27  for clarity) of similar construction to the fans  113   a,    113   b  of the engine  110 . Each fan is similarly driven by respective output shafts  1325   a,    1325   b  which are in turn driven by a gearbox  1326 , which is in turn driven by an input shaft in the form of a fan drive turbine shaft  1327 . Again, the fans  1313  are located at a rearward end of the engine  1310 , downstream of the gearbox  1326 . 
     The fan drive turbine shaft  1327  is driven by a core engine  1375 . Engine  1310  comprises a core engine  1375  having first and second core modules  1390 ,  1391 . The first core module  1390  comprises a first compressor in the form of a low pressure compressor  1328  and a first turbine in the form of a high pressure turbine  1331  interconnected by a first shaft in the form of a high pressure shaft  1377 . The core  1375  further comprises a second core module  1391  comprising a second compressor in the form of a high pressure compressor  1343  and a fan drive turbine  1329  interconnected by a second shaft in the form of a low pressure shaft  1327 . The first and second modules  1390 ,  1391  are separated in an axial direction X. In this embodiment, each component of the first module  1390  is provided forwardly of each component of the second module  1391 . Again, though the shafts  1327 ,  1377  rotate about a common engine axis, the shafts do not overlap in an axial direction. 
     While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention. 
     For example, the fans could comprise fixed pitch blades. In such a case, a cold flow thrust reverse mechanism may be provided. Where the fans are powered by electrical motors, the fans could be located remotely from the engine core. For example, the engine core could be located in the tail, with the fans located on the wings. 
     The first and second shafts need not be co-axial, and could be offset relative to one another.