Patent Publication Number: US-2017363043-A1

Title: Gas turbine engine

Description:
The present disclosure concerns a gas turbine engine having multiple combustion chambers 
     There is a continuing desire for gas turbine engines having improved specific fuel consumption (i.e. reduced fuel consumption for a given amount of thrust), to provide improved range, and reduced operating costs. This is particularly the case for aircraft which must operate over a wide range of speeds, such as supersonic aircraft. It is also desirable for gas turbine engines which operate at high speeds to have a high specific thrust (i.e. a high thrust for a given engine total intake air mass flow). 
       FIG. 1  shows a prior gas turbine engine  1  for a proposed supersonic business jet. The engine  1  comprises an engine core comprising a high pressure compressor  10 , which supplies compressed air to a main combustor  12 . Combustion products from the combustor  12  flow downstream in use to a high pressure turbine  14 , which drives the high pressure compressor  10  via a shaft  16 . Downstream of the high pressure turbine  14  is a low pressure turbine  18 , which drives a low pressure compressor in the form of a fan  20  via a low pressure shaft  22 . The fan  20  supercharges the core, and provides airflow through a bypass duct  24  which extends around the core. The bypass flow and core flow are combined in a mixer duct  26  located downstream of the low pressure turbine  18 . A further combustor  28  (known as an afterburner or reheat combustor) is provided downstream of the mixer duct  26 , which further raises the temperature and velocity of the exhaust. Air exits the engine through a nozzle  30 , which may of convergent-divergent type, and typically has a variable area. The variable area is generally necessary to control backpressure on the turbines  14 ,  18  when the afterburner  28  is in operation. 
     Such an arrangement may have a relatively large length in view of the high velocity airflow within the afterburner, and the relatively long burn time of typical aviation fuels. Furthermore, such systems are relatively inefficient when the afterburner is in operation. Such engines typically have relatively low bypass ratios (i.e. the ratio between the core mass flow and bypass mass flow in use) of around 1:1 or less, so that they can operate efficiently at high speed (generally greater than Mach 1). However, such engines are consequently relatively inefficient at low speed (i.e. less than Mach 1) compared to engines having higher bypass ratios. 
     Consequently, there is a need to provide an aircraft gas turbine engine having high efficiency across a broad range of flight speeds. 
     In accordance with the present invention there is provided an aircraft gas turbine engine comprising a high pressure compressor driven by a high pressure turbine via a high pressure shaft, a first combustor provided downstream of the high pressure compressor and upstream of the high pressure turbine, a low pressure compressor driven by a low pressure turbine via a low pressure shaft, the low pressure compressor being configured to provide air to the high pressure compressor and to a bypass flow, the low pressure turbine comprising at least first and second turbine stages, and a second combustor provided downstream of the first stage of the low pressure turbine and upstream of the second stage of the low pressure turbine, wherein the gas turbine engine comprises a shaft coupling arrangement configured to transfer power between the high and low pressure shafts. 
     Advantageously, the arrangement provides a gas turbine engine having a high specific thrust and a low fuel consumption in view of the provision on a second combustor located between turbine stages of the low pressure turbine. Consequently, the specific thrust of the engine can be increased using the second combustor, while allowing for a relatively high bypass ratio, thereby providing high efficiency at both supersonic and subsonic speeds. 
     A bypass ratio is defined by a ratio of mass flow through the bypass to mass flow through high pressure compressor. The engine may comprise a bypass ratio greater than 1, and may comprise a bypass ratio of approximately 3. Due to the relatively high bypass ratio (compared to less than 1 for conventional high specific thrust aircraft gas turbine engines), the engine has a relatively low specific fuel consumption at subsonic speeds due to the improved propulsive efficiency achievable with a higher bypass ratio, and the ability to generate high thrust at high speeds without recourse to using an afterburner. 
     The low pressure compressor may comprise at least one fan stage configured to provide air to the high pressure compressor and the bypass, and at least one core stage configured to provide air to the high pressure compressor only. 
     The engine may comprise an afterburner provided downstream of the low pressure turbine in the core flow. Alternatively or in addition, the engine may comprise a duct burner located in the bypass, downstream of the low pressure compressor. 
     The engine may comprise a mixer configured to mix bypass and core flows downstream of the low pressure compressor and the low pressure turbine. The mixer may be provided upstream of the afterburner. 
     The engine may comprise a variable area exhaust nozzle downstream of the low pressure turbine. 
     The coupling arrangement may comprise a fluid coupling such as a torque converter comprising an input shaft coupled to the one of the high pressure shaft and the low pressure shaft and an output shaft coupled to the other shaft. Alternatively or in addition, the shaft coupling arrangement may comprise a mechanical clutch and/or a continuously variable transmission or a gearbox having a plurality of discrete ratios. The coupling arrangement may comprise an electric generator coupled to one of the high and low pressure shafts, and an electric motor coupled to the other shaft, the electric generator being electrically coupled to the electric motor to thereby drive the electric motor. 
     The fluid coupling may comprise a first rotor coupled to the input shaft and a second rotor coupled to the output shaft, the first and second rotors being immersed in a transmission fluid within a fluid coupling housing. The transmission fluid may comprise aviation fuel. The gas turbine engine may comprise a fuel system configured to provide fuel from a fuel tank to an engine injector via the fluid coupling housing. Advantageously, any heat generated due to transmission inefficiencies of the fluid coupling is transferred to the combustor and thereafter to the turbines, thereby increasing the thermodynamic efficiency of the arrangement. 
     The fluid coupling may comprise a stator immersed within the transmission fluid. One or more of the first and second rotor and the stator may comprise a bladed disc. One or more of the first and second rotor and the stator may comprise a variable pitch mechanism configured to vary the pitch of blades of the first or second rotor or stator. Consequently, power transmitted from the high pressure shaft to the low pressure shaft can be selectively varied. 
     The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein. 
    
    
     
       Embodiments will now be described by way of example only, with reference to the Figures, in which: 
         FIG. 1  is a schematic side view of a prior gas turbine engine; 
         FIG. 2  is a schematic side view of a first gas turbine engine in accordance with the present disclosure; 
         FIG. 3  is a schematic side view of a second gas turbine engine in accordance with the present disclosure; 
         FIG. 4  is a schematic side view of a first shaft coupling arrangement suitable for use with the gas turbine engine of  FIG. 2  or  FIG. 3 ; 
         FIG. 5  is a schematic side view of a second shaft coupling arrangement suitable for use with the gas turbine engine of  FIG. 2  or  FIG. 3 ; and 
         FIG. 6  is a schematic side view of a third shaft coupling arrangement suitable for use with the gas turbine engine of  FIG. 2  or  FIG. 3 . 
     
    
    
     With reference to  FIG. 2 , a gas turbine engine is generally indicated at  100 . The engine  100  comprises, in axial flow series, an air intake  112 , a propulsive fan  11 , a high-pressure compressor  116 , a first combustor  118 , a high-pressure turbine  120 , a first stage  122  of a low-pressure turbine, a second combustor  124 , a second stage  126  of the low pressure turbine, a mixer  128 , an afterburner  130 , and an exhaust nozzle  132 . A nacelle  134  generally surrounds the engine  100  and defines the intake  112 , mixer,  128 , afterburner  130 , nozzle  132  and a bypass duct  136 . 
     The fan  114  comprises a multi stage axial compressor comprising a plurality of rotors and stators of conventional construction, which is arranged to provide airflow to both the high pressure compressor  116  (i.e. the “core”) and the bypass passage  136 . The fan  114  is coupled to both the first and second stages  122 ,  126  of the low pressure turbine, and is thereby driven, by a low pressure shaft  138 . In some embodiments, the low pressure compressor  114  may further comprise a “booster compressor” driven by the low pressure shaft  138 , which is configured to compress engine core air, and deliver compressed air to the high pressure stage  116  only, and not to the bypass duct  136 . The booster compressor may comprise one or more stages, and may be of axial or centrifugal type. The fan  114  and compressor  116  are sized in the described embodiment such that the ratio of mass of flow air in the bypass flow relative to the air in the core flow (i.e. the bypass ratio) in use is approximately 3. More generally, the bypass ratio is likely to be greater than 1. In view of the relatively high bypass ratio, the engine is thought to provide efficiency savings during subsonic flight of up to 30%. 
     Similarly, the high pressure compressor  116  typically comprises a multi stage axial flow compressor comprising a plurality of rotor and stator stages (not shown). The high pressure compressor rotor  116  is coupled to the high pressure turbine  120  by a high pressure shaft  140 . 
     The first combustor  118  comprises a substantially constant pressure combustor, which may be of conventional construction, such as a rich burn, lean burn or rich-quick-quench-lean burn (RQL) combustor. The first combustor  118  is configured to burn fuel with air delivered by the high pressure compressor  116 , and deliver hot combustion products to the high pressure turbine  120 . 
     The high pressure turbine  120  is provided downstream of the first combustor  118 , and is configured to receive hot combustion gasses from the combustor  118 , and expand those gasses to drive the high pressure compressor  116 . The turbine may be of conventional construction, and typically consists of at least one turbine rotor (which may be of axial or radial type) and a turbine stator or diffusor. The turbine  122  may comprise one or more turbine stage. 
     The first stage  122  of the low pressure turbine  122  is provided downstream of the high pressure turbine  122 , and again comprises at least one radial or axial flow rotor, and a stator or diffuser. Further low pressure turbine stages may optionally be provided upstream of the second combustor  124 . 
     The second combustor  124  is provided downstream of the first stage of the low pressure turbine  122 . The second combustor  124  is similar to the first combustor  118 , being of a substantially constant pressure type. 
     The second stage of the low pressure turbine  126  is provided downstream of the second combustor  124 , and again, is of conventional construction. Again, further turbine stages may be provided downstream. 
     The mixer  128  is provided downstream of the second stage of the low pressure turbine  126 , and downstream of the bypass duct  136 , such that the mixer  128  is configured to receive engine core exhaust from the turbine  126 , and bypass exhaust from the fan  114 . The flows are mixed together in this region prior to delivery to the afterburner  130 . 
     The afterburner  130  is generally of a different construction to the first and second combustors  118 ,  124 , in view of the relatively high velocity airflow through the afterburner, but again is of a known type suitable for use as an afterburner. 
     The exhaust nozzle  132  is of conventional construction, being of a convergent-divergent type, and having a variable area, such that the outlet area can be varied in dependence on operating conditions. 
     The engine  100  further comprises a shaft coupling arrangement  142  configured to couple the low and high pressure shafts  138 ,  140  together, to thereby selectively transfer power between the low and high pressure shafts, while allowing each shaft  138 ,  140  to rotate at a different speed, independently of the other shaft. It will be understood that any suitable mechanical, electrical, magnetic, hydraulic or pneumatic transmission means may be employed, providing the transmission is capable of selectively transferring power while permitting rotation of the low and high pressure shafts  138 ,  140  at different speeds. 
       FIG. 4  shows a first proposed coupling arrangement  142   a . The coupling arrangement  142   a  comprises a radial drive shaft  152  driven by the low pressure shaft  138  via a first bevel gear arrangement  154 . The radial drive shaft  154  is in turn coupled to an input shaft  156  of a torque converter  158  via a second bevel arrangement  160 . 
     The torque converter  158  comprises a fluid filled housing  161  housing an input impeller  162  comprising a bladed rotor configured to impart swirl into the transmission fluid, a bladed stator  164  configured to control swirl of the transmission fluid, and an output turbine  168 , configured to convert swirl of the transmission fluid to torque. The output turbine  168  is coupled to the high pressure shaft  140  via a third bevel arrangement  170 , output radial drive  172  and fourth bevel arrangement  174 . 
     In this embodiment, the transmission fluid comprises gas turbine engine fuel such as aviation fuel, and is transmitted to the housing  161  via a fuel supply conduit  176 . A further fluid supply conduit  178  is provided, which extends between the housing  161  and the first combustor  118 . Consequently, a continuous flow of aviation fuel is provided through the housing  161 , which both provides a medium within which the rotors  162 ,  168  and stator  164  may operate, and cooling to dissipate heat generated by the swirling of the fluid. Since this fluid is then transferred to the combustor  118 , the heat produced by the torque converter  158  is conserved within the thermodynamic cycle of the engine  100 , and thus any inefficiency in the torque converter arrangement is at least partially recovered by the turbines  120 ,  122 ,  128 . Consequently, the above arrangement provides a highly efficient system. The torque converter could also include a “lock-up” arrangement for mechanically locking the first and second shaft via a clutch, to thereby directly link the first and second shafts. One of the radial drives  152 ,  172  could be coupled to the respective shaft  138 ,  140  via gearing, which may increase or reduce the speed of the respective radial drive relative to the respective shaft. 
       FIG. 5  shows a first alternative coupling arrangement  142   b . The arrangement  142   b  comprises a radial drive shaft  252  driven by the low pressure shaft  138  via a first bevel gear arrangement  254 . The radial drive  252  is in turn coupled to an input shaft  256  of a continuously variable transmission arrangement (CVT) via a second bevel gear arrangement  260 . 
     The CVT comprises first and second conical input rotors  280 ,  282  which are driven by the input shaft  256  and coupled to one another by a drive chain  284 . A first actuator  286  varies the axial distance between the first and second rotors  280 ,  282  to thereby vary the effective outer diameter of the rotors  280 ,  282  where they engage with the chain  284 , and so the input gearing of the CVT. The chain  284  also extends around third and fourth conical output rotors  286 ,  288 , which are thereby rotated by the chain  284 . The fourth rotor  288  is coupled to the high pressure shaft  140  via an output shaft, a third bevel arrangement  270 , a second radial drive shaft  272  and a fourth bevel arrangement  274 . The rotors  286 ,  288  are coupled to a second actuator  290  which again varies the axial distance between the conical rotors  286 ,  288  to thereby vary the output gearing of the CVT. In use, the actuators  286 ,  290  are operated synchronously, such that the gearing of the CVT can be continuously varied without introducing slack or excessive tension into the chain  284 . Torque from the low pressure shaft  138  is thus transferable to the high pressure shaft  140  via the coupling arrangement  242   b.    
       FIG. 6  shows a second alternative coupling arrangement  142   c . The arrangement  142   c  comprises a radial drive shaft  352  driven by the low pressure shaft  138  via a first bevel gear arrangement  354 . An electrical generator  392  is driven by the drive shaft  352 , and provides electrical power to a variable speed electric motor  394  via an electrical interconnector  396 . The variable speed motor  394  is in turn mechanically coupled to the high pressure shaft  140  via a shaft  372  and bevel gear arrangement  374 . Consequently, power from the low pressure shaft  138  can be transferred to the high pressure shaft  140  via the coupling arrangement  142   c.    
       FIG. 3  shows an alternative gas turbine engine in accordance with the present invention. The engine  200  again comprises a low pressure fan  214  and a high pressure compressor  216 . The high pressure compressor  216  is driven by a high pressure turbine comprising first  220  and second  221  stages. The low pressure fan  214  is driven by a low pressure turbine comprising first  222  and second  226  stages. The airflow through the high pressure compressor  216  and turbine stages  220 ,  221 ,  222 ,  226  defines a core flow path. A first combustor  218  is provided in the core flow path between the high pressure compressor  216  and the first stage  220  of the high pressure turbine. A re-heat combustor  224  is provided in the core flow path between the first and second stages  220 ,  221  of the high pressure turbine. A further reheat combustor  230  is provided in the core flow path between the first  222  and second  226  stages of the low pressure turbine. Further low pressure turbine stages (not shown) may optionally be provided downstream of the second stage of the low pressure turbine  226 . Again, the shafts  238 ,  240  are interconnected by a shaft coupling arrangement  140 , such as the arrangements  140   a ,  140   b  or  140   c  described above. 
     The airflow that flows through the fan  214  but not the compressor  216  or turbines  220 ,  221 ,  222 ,  226  defines a bypass flow path, which is defined by a nacelle  234 . A duct burner combustor  231  is located in the bypass flow path, downstream of the fan  214 . Each of the first, second and afterburner combustors  218 ,  224 ,  230  may be of a conventional type suitable for use as a core combustor, while the duct burner combustor may be of a type suitable for use as an afterburner combustor. Downstream of both the duct burner combustor  232  and the second stage  226  of the low pressure turbine is a mixing duct  228 , configured to mix the core and bypass streams. The engine terminates in a nozzle  232 , which is configured to expel core and bypass flows to generate thrust. 
     In view of the provision of a fourth combustor in the bypass duct, a high specific thrust engine can be provided, while having a relatively short engine length. 
     It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein. 
     For example, the afterburner and/or duct burner combustors may be eliminated. The fan could be a single stage fan, comprising a single rotor and stator.