Patent Publication Number: US-2017370290-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. However, as the bypass ratio increases, the fan pressure ratio necessarily decreases. Low pressure ratio fans are particularly susceptible to operability issues such as stall and/or flutter during some operational conditions. Consequently, it has been proposed to provide such fan with one or both of variable pitch, and a variable area nozzle. In both cases, these devices can be used to adjust the pressure ratio on the fan, and so prevent stall and/or flutter. However, these devices are relatively heavy and expensive, and do not generally contribute to the performance of the engine in themselves, and so represent dead weight. 
     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: 
     a fan arranged to be driven by a gas turbine engine core, the core comprising a first core module comprising a first compressor and a fan drive turbine interconnected by a first shaft, and a second core module comprising a second compressor and a second turbine interconnected by a second shaft, the first and second core modules being axially spaced; 
     an intercooler arrangement configured to cool core airflow between the first and second compressors, the intercooler arrangement comprising a cooling air duct provided in heat exchange relationship with a compressor duct provided between the first and second compressors, the cooling air duct comprising a fan air inlet configured to ingest fan air, downstream of the fan, wherein the cooling air duct comprises a flow modulation valve configured to modulate air mass flow through the fan air inlet. 
     Advantageously, such an arrangement provides a compact gas turbine engine, in which intercooling is provided to improve thermodynamic by enabling a higher overall pressure ratio while maintaining an acceptable compressor delivery air temperature. Meanwhile, the intercooler flow modulation valve provides control over intercooling, whilst simultaneously controlling effective fan outlet area using the same valve, since the intercooler cooling duct inlet is provided downstream of the fan. 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 operability (stall and 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 core may comprise a compressor provided axially rearwardly of the fan drive turbine. 
     The fan drive coupling may comprise a gearbox arranged such that the fan drive turbine rotates at a higher rotational speed than the fan in use. The gearbox may have an input:output ratio of between 1 and 5, and preferably has a ratio greater than 2. 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. 
     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. 
     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. 
     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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows a prior gas turbine engine; 
         FIG. 2  shows a schematic cross sectional side view of a first gas turbine engine in accordance with the present disclosure; and 
         FIG. 3 a    and  FIG. 3 b    show a close-up of the part of the component of  FIG. 2  shown in box D; 
     
    
    
     DETAILED DESCRIPTION 
       FIGS. 2 and 3  show a first gas turbine engine  110  in accordance with the present invention. The engine  110  comprises a ducted fan  113  provided within a fan nacelle  121  which defines a bypass passage  148 . The fan  113  provides a propulsive air flow B which flows parallel to an axial direction X. A forward direction is defined by an axial direction anti-parallel to this direction. 
     The engine  110  further comprises an engine core  175 . The core  175  comprises a first core module  190  comprising a first compressor in the form of a low pressure compressor  114  configured to draw core flow air A into the core  175  from an inlet  149  positioned downstream of the fan  113 . The first core module  190  further comprises a first turbine in the form of a low pressure fan drive turbine  119  interconnected by a first shaft in the form of a low pressure shaft  177 . The core  175  further comprises a second core module  191  comprising a second compressor in the form of a high pressure compressor  115  and a high pressure turbine  117  interconnected by a second shaft in the form of a high pressure shaft  127 . The first and second modules  190 ,  191  are separated in an axial direction X, i.e. they do not overlap in the axial direction. In this embodiment, each component of the first module  190  is provided forwardly (i.e. in a direction opposite to the axial direction X) of each component of the second module  191 . Consequently, though the shafts  127 ,  177  rotate about a common engine axis  111 , the shafts do not overlap in an axial direction. 
     The core  175  defines a core airflow path A. The fan  113  is driven by the fan drive turbine  119  via a fan drive coupling. The fan drive coupling comprises an output shaft  125  which is coupled to the fan  113  via a reduction gearbox  126 . The gearbox  126  is driven by the fan drive shaft  177 , and is configured to drive the output shaft, and so the fan, at a lower rotational speed than the input shaft  177 . The gearbox  126  provides a reduction ratio, such that the ratio between the input shaft  177  rotational speed and the fan  113  rotational speed 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. 
     Both the compressors  114 ,  115  generally comprise multi-stage axial flow compressors. At a rearward end of the low pressure compressor  114  is a low pressure compressor outlet  134 . Air from the low pressure compressor  114  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  175 . Surrounding at least part of the inter-compressor core air duct  135  is an intercooler fan air duct  136 . The intercooler fan air duct  136  comprises a hollow passage having an inlet  137  at a forward end configured to ingest fan air from within the fan nacelle  121 , downstream of the fan  113 , to define an intercooler airflow C. 
     A heat exchanger matrix  150  is provided at an aft end of the intercooler fan air duct  136  and inter-compressor core air duct  135 . Air from the ducts  135 ,  136  is in thermal contact within the heat exchanger matrix. In view of the temperature difference between the high temperature compressed airflow A within the inter-compressor duct  135  and low temperature intercooler airflow C within the intercooler duct  137 , heat is exchanged from the compressed airflow A to the intercooler airflow C. Consequently, the intercooler duct  136  and inter-compressor core air duct  135  together form a compressor intercooler  150 , thereby reducing the work required by further compressor stages, and increasing thermal efficiency. 
     The intercooler duct  136  further comprises an intercooler cooling flow modulation valve  138  configured to modulate intercooler airflow C mass flow rate.  FIGS. 3 a  and 3 b    show a cross section of the region D(i) and D(ii) of  FIG. 2  respectively. It will be understood that the positions shown in the top and bottom half of  FIG. 2  are for illustrative purposes, and in practice, the flow modulate valve is likely to be in the same position at all engine circumferential positions. 
     In  FIGS. 3 a  and 3 b   , the valve  138  is shown 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. 3 a   ) and an open position (shown in  FIG. 3 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 of the high pressure compressor  115 . In operation, the high pressure compressor  115  further raises the pressure of the core air flow A in operation, and urge the core air flow A forwardly. 
     Axially forwardly (i.e. downstream in core flow A) of the high pressure compressor  115  is the combustor  130 , which is of conventional construction. In the combustor  130 , fuel is provided and burnt with the compressed air from the high pressure compressor  115  in operation to increase the temperature of the core air flow A. 
     The high pressure turbine  117  is provided axially forwardly (i.e. downstream in core flow A) of the combustor  130 . In use, the high pressure turbine  117  directs flow forwardly, while extracting energy from the flow to drive the high pressure shaft  127 , which is coupled to the high pressure compressor  115 , to thereby drive the high pressure compressor  115  in operation. 
     Axially forwardly (i.e. downstream in core flow A) of the high pressure turbine  117  is the low pressure fan drive turbine  119 , which is of similar construction to the high pressure turbine  117 , comprising a plurality of rotors and stators. The low pressure fan drive turbine  119  is coupled to both the low pressure compressor  114  and the gearbox  126 . Consequently, the low pressure turbine  119  drives the fans  113  and the low pressure compressor  114  via the shaft  177  in operation. 
     Between the low pressure turbine  119  and the low pressure compressor  114  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  119  in the core air flow A. The core exhaust passage  145  turns core air flow A approximately 180°, and so redirects air rearwardly in use. Core air A from the core exhaust passage  145  mixes with fan air B downstream within the nacelle  121 . 
     Consequently, the above arrangement defines a “reverse flow” architecture, in which core flow A flows in a forward direction during at least part of the compression and expansion processes, i.e. in an opposite direction to the fan efflux, since at least one core turbine  117 ,  119  is provided forwardly of at least one core compressor. 
     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 variable pitch blades. In such a case, a cold flow thrust reverse mechanism may be provided. 
     The first and second shafts need not be co-axial, and could be offset relative to one another. 
     A plurality of fans could be provided, each being driven by the fan drive turbine. A recuperator could be provided, configured to exchange heat from relatively high temperature exhaust air downstream of the fan drive turbine, and relatively low pressure compressed core air downstream of the high pressure compressor.