Patent Publication Number: US-2017370284-A1

Title: Gas turbine engine

Description:
The present disclosure concerns a gas turbine engine. 
     There is a continual need to both increase the fuel efficiency of gas turbine engines, and reduce their cost. One known technology for at least increasing fuel efficiency is to employ a reduction gearbox between a fan and a fan drive turbine, such that the fan can be operated at a lower rotational speed than the fan drive turbine. One such engine is described in U.S. Pat. No. 8,176,725, which describes a gas turbine engine having a fan drive gear system, a low spool connected to the fan drive gear system, and a high spool disposed aft of the low spool. The low spool comprises a rearward-flow low pressure compressor disposed aft of the fan drive gear system, and a forward-flow low pressure turbine disposed aft of the low pressure compressor. The high spool comprises a forward-flow high pressure turbine disposed aft of the low pressure turbine, a combustor disposed aft of the high pressure turbine, and a forward-flow high pressure compressor disposed aft of the combustor. The engine comprises a heat exchanger configured to exchange heat between the compressed air in the core, and the fan flow, to thereby provided intercooling. 
     The present invention seeks to provide an improved gas turbine engine, which has high fuel efficiency, and low cost. 
     According to a first aspect of the invention there is provided a gas turbine engine comprising: 
     a low pressure module comprising a low pressure compressor comprising a centrifugal compressor stage, and a low pressure turbine configured to drive a load
 
a high pressure module comprising a high pressure turbine, a combustor and a high pressure compressor, the high pressure module being provided axially spaced from the low pressure module; and
 
an intercooler heat exchanger configured to exchange heat between core airflow exiting the low pressure compressor and ambient airflow, wherein the high pressure turbine and high pressure compressor are coupled by a high pressure shaft, and the low pressure turbine, low pressure compressor and load are coupled by a low pressure coupling.
 
     Advantageously, the above arrangement provides a gas turbine engine which is thermally efficient, compact and low cost. 
     The load driven by the low pressure turbine may comprise a fan. The fan may define a forward end of the engine, and may be provided forwardly of the low pressure turbine. The fan may be configured to provide a fan outlet flow in a rearward direction. The high pressure compressor may be configured to transfer flow in a forward direction generally opposite to the rearward direction. Alternatively, the high pressure compressor may be configured to transfer flow in a radially outward or inward direction. 
     The combustor may be configured to receive core flow from the high pressure compressor, and deliver flow from a compressor outlet in the forward direction. 
     The high pressure turbine may be configured to receive flow from the combustor and deliver flow to the low pressure turbine. 
     The low pressure turbine load may be coupled to the low pressure turbine by a reduction gearbox. The reduction gearbox may be provided between the low pressure compressor and the load. The low pressure turbine may be provided between the combustor and the low pressure compressor. 
     The high pressure compressor may comprise one or both of at least one axial flow compressor stage and at least one centrifugal flow compressor stage. 
     The low pressure compressor may comprise a single stage centrifugal impellor. The low pressure compressor may further comprise an axial flow compressor upstream of the centrifugal flow compressor, and coupled to the low pressure shaft. 
     The engine may comprise an interstage duct extending between a low pressure compressor outlet and a high pressure compressor inlet. The intercooler heat exchanger may comprise the interstage duct. The interstage duct may be provided radially outwardly of the high pressure compressor and high pressure turbine. 
     The engine may comprise a core exhaust duct configured to redirect core exhaust from a low pressure turbine outlet in the rearward direction. The core exhaust duct may terminate downstream of the interstage duct. The interstage duct and core exhaust duct may extend generally parallel to one another. A plurality of core exhaust ducts and interstage ducts may be provided, and may be arranged alternately with one another, and may be circumferentially spaced around the engine. 
     The engine may comprise a bypass ratio of 10 or greater. 
     The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects of the invention may be applied mutatis mutandis to any other aspect of the invention. 
    
    
     
       Embodiments of the invention will now be described by way of example only, with reference to the Figures, in which: 
         FIG. 1  is a sectional side view of a first gas turbine engine in accordance with the present invention; 
         FIG. 2  is a sectional frontal view of the gas turbine of  FIG. 1  along the line  8 ; and 
         FIG. 3  is a sectional side view of a second gas turbine engine in accordance with the present invention; and 
         FIG. 4  is a sectional side view of a third gas turbine engine in accordance with the present invention. 
     
    
    
     With reference to  FIG. 1 , a gas turbine engine  10  comprises a low pressure module  60  comprising a low pressure turbine  44 , low pressure compressor  24  and load in the form of a fan  12  interconnected by a low pressure coupling comprising a low pressure shaft  22 , reduction gearbox  56  and fan shaft  58 . The fan  12  is configured to accelerate air entering an engine inlet  14  and provide a fan bypass flow A and core flow B, both of which flow initially in a first axial direction X. The fan  12  defines a forward end of the engine, and so the first axial direction defines a rearward direction, with a forward direction being defined by a direction opposite (i.e. 180°) to direction X. Rearwardly, and downstream in bypass flow A of the fan  12  is an outlet guide vane  17 , which straightens the flow from the fan  12 , and supports an engine core. Also downstream of the fan  12  is a core engine inlet  16  which diverts part of the fan flow into the core. The core (comprising compressors  18 ,  32 , combustor  40  and turbines  42 ,  44 , described in further detail below) is housed within a generally annular core nacelle  13 . The fan  12  is housed within a generally annular fan nacelle  15 , which extends axially downstream of the fan  12 , and is driven by a low pressure fan drive turbine  44  via a gearbox  56 . The fan flow A is defined by the region bounded by the nacelles  13 ,  15 . A bypass ratio is defined by the ratio between the mass flow rate of air drawn through the fan disk that bypasses the engine core (bypass flow A) to the mass flow rate passing through the engine core (core flow B). In the described embodiments, the engine has a bypass ratio of approximately 10. 
     The engine  10  further comprises a high pressure module  62  comprising, in flow series, a high pressure compressor  32 , combustor  40 , and high pressure turbine  42 . The high pressure compressor  32  and high pressure turbine  42  are coupled by a high pressure shaft  36 . The high pressure module  62  and low pressure module  60  are axially spaced, i.e. their respective axes do not overlap. 
     Downstream in core flow B of the core inlet  16  is the low pressure compressor  18 . The low pressure compressor comprises a single stage axial flow compressor comprising a rotor  20  and stator  22 . Downstream of the stator  22  is a centrifugal compressor stage comprising an impellor  24  and diffuser  25 , which further compresses the core airflow B. The centrifugal compressor impellor  24  is of conventional construction, and is arranged to ingest air provided to the centrifugal compressor from the axial direction X, and expel air at an outlet  26  in a generally radial direction Y, through the diffuser  25 . Centrifugal compressors generally have a higher stage ratio than axial compressors, and so a single rotor component can raise the pressure to a greater degree than an axial compressor in a given length. However, centrifugal compressors must either operate at high speeds, or have a large tip diameter in order to operate efficiently. In addition, the diffuser  25  generally increases the diameter still further. Consequently, the centrifugal compressor has a relatively large diameter compared to other rotor components of the gas turbine engine. In particular, the centrifugal compressor outlet has a larger radial extent than the final stage of the low pressure turbine  44 . In particular, the tip of the centrifugal impellor  24  has a greater radius than the tips of the rotor blades of the low pressure turbine  44 . 
     The outlet  26  of the centrifugal compressor impellor stage (i.e. the diffuser  25 ) provides core airflow B to an intercooler in the form of an interstage duct  28 . The interstage duct  28  carries compressed core airflow B from the outlet  26  of the centrifugal compressor  24  to an inlet  30  of a high pressure centrifugal compressor  32 . An intermediate portion  34  of the duct  28  extends in a generally axial direction, and carries core airflow B in the axial direction X. The intermediate portion  34  is located in thermal contact with the core nacelle  13 , and so the intermediate portion acts as a parallel flow heat exchanger, exchanging heat between the relatively hot compressed core flow B with the relatively cool fan flow A. Fins, turbulators or other flow control devices may be provided within the intermediate portion  34  and/or on an outer surface of the core nacelle  13  to increase the surface area of the intercooler heat exchanger hot and/or cold sides, and thereby facilitate heat exchange. 
     A downstream end of the interstage duct  28  comprises an elbow connector  36  which is configured to turn the core airflow B approximately 180° from being generally parallel to the axial flow direction X, to counter to the axial flow direction X for ingestion into the inlet  30  of the high pressure centrifugal compressor  32 . The high pressure centrifugal compressor  32  again comprises a centrifugal impellor and diffuser configured to raise the pressure of the core airflow B by redirecting the air from generally counter to the axial flow direction X, to the generally radial direction Y. The high pressure centrifugal compressor  32  has a smaller diameter than the low pressure impellor  24 , as the high pressure impellor rotates at a higher speed in use. 
     An outlet  36  of the high pressure compressor  32  delivers air to an inlet of the combustor  40 , which adds fuel to the core airflow B, to thereby burn the fuel and increase the temperature of the core airflow B. A high pressure turbine  42  is provided downstream of the combustor  40 , which is acted upon by the core airflow B to thereby drive the turbine  42 . The turbine in the described embodiment comprises a single stage axial flow rotor, though multi stage or centrifugal flow rotors could be employed. 
     The low pressure turbine  44  is provided downstream in core flow B of the high pressure turbine  42 . The low pressure turbine  44  generally comprises a plurality of axial flow turbine rotors and interposed stators. 
     A core exhaust duct  46  is provided downstream of an outlet of the low pressure turbine  44 . The core exhaust comprises an elbow  48  at an upstream portion, which is configured to redirect the forward flow from the low pressure turbine  44  outlet 180° and radially outwardly, to the axial, rearward direction X. An intermediate portion  48  extends from the elbow  48  in the axial direction X, and exhausts the core airflow at an aft end of the engine. As can be seen from  FIG. 1 , the intermediate portions  34 ,  48  of the interstage duct  28  and core exhaust duct  46  extend in the axial direction at substantially the same radial position. Consequently, a plurality of interstage and core exhaust ducts  28 ,  46  are arranged alternately around the circumference of the engine core casing  13 , as shown in  FIG. 2 . The engine exhaust duct  46  is generally spaced from both the interstage duct  28  and engine core casing  13  and may also be insulated to avoid heating either the fan flow A or the core flow B within the interstage duct  28 . 
     The high pressure compressor  32  and turbine  42  are coupled by a high pressure fan drive shaft  52 . Consequently, the high pressure compressor  32  is driven by the turbine  42  via the shaft  52 . The low pressure turbine  44  and low pressure compressor  24  are coupled by a low pressure shaft  54 , and so the low pressure compressor  24  is driven by the low pressure turbine  44  via the low pressure shaft  54 . The low pressure shaft  54  is also coupled to a reduction gearbox  56 , which is in turn coupled to the fan  12  via a fan shaft  58 . The reduction gearbox is configured to provide a reduction ratio of at least 2.5:1. Consequently, the fan  12  is driven by the low pressure turbine  44 , but the fan  12  rotational speed is different from the turbine  44  rotational speed. Consequently, the fan drive low pressure turbine  44  can rotate at a relatively high speed compared to conventional high pressure ratio fan gas turbine engines, and so the low pressure turbine can have a smaller diameter. Consequently, the radius of curvature of the core flow exhaust elbow  48  can be relatively large, without resulting in an excessively large diameter engine core. 
     The engine  10  is consequently arranged as follow. The fan  12  is located at an axially forward end of the engine  10 , with the reduction gearbox  56  provided at the same axial position, or rearwardly thereof. The low pressure compressor  18  is provided rearwardly of the fan  12  and gearbox  56 . The core exhaust elbow  48  is provided rearwardly of the low pressure compressor  24 , and the low pressure turbine  44  is provided axially rearwardly of both the elbow  48  and the low pressure compressor  24 . The low pressure shaft  54  extends between a forward end of the low pressure turbine  44  and the gearbox  56 . Consequently, the low pressure shaft  54  is relatively short, which reduces engine weight, and reduces shaft whirling. 
     The high pressure turbine  42  is provided rearwardly of the low pressure turbine  44 , with the combustor  40  being provided further rearwardly, followed by the high pressure compressor  32  still further rearwardly, and the interstage duct elbow  36  is provided at the rear of the engine  10 . The high pressure shaft  52  extends between the high pressure turbine  32  and high pressure compressor  32 , through the combustor  40 , and so is again relatively short. Since the modules  60 ,  62 , and so the shafts  52 ,  54  are axially spaced, they do not have to be arranged concentrically, which simplifies bearing and oil design. Furthermore, the engine  10  can be modular, with the high pressure compressor  32  and turbine  42  of the engine core being removeable from the remainder of the engine, without having to remove the low pressure section (i.e. the low pressure turbine  44 , compressor  18 , gearbox  56  and fan  12 ). 
     In general, reverse flow architecture cores require relatively large diameter cores, particularly at their mid-sections, since the low pressure turbine generally requires a large number of large diameter stages, and this is located relatively forward within the engine core nacelle  13  in a reverse flow architecture. This large diameter is exacerbated by the requirement for an exhaust duct  46  which turns the flow 180° to be exhausted in the axial direction X. The turning of the flow also results in aerodynamic inefficiency within the exhaust duct, which in turn results in higher backpressure at the low pressure turbine exhaust, resulting in lower turbine efficiency, and so lower overall cycle efficiency or increased turbine stage count. This can be ameliorated by increasing the radius of curvature of the exhaust duct elbow  48 , but results in a still larger diameter engine core. 
     These advantages and disadvantages are thought to largely cancel one another in conventional reverse flow architectures, such that significant benefits cannot be achieved, particularly in large, high bypass ratio engines. 
     The current invention overcomes these limitations by utilising the large diameter low pressure centrifugal compressor impellor  24 , in place or in addition to a conventional axial flow low pressure compressor  20 . Centrifugal compressors generally have higher efficiencies (both in terms of stage pressure rise and thermodynamic efficiency) where there tip speed is high. This can be achieved by providing a large diameter rotor, or be rotating at high speeds. In the present invention, the large diameter core provides an opportunity to provide a high efficiency centrifugal compressor, which can efficiently develop a high pressure ratio. The resultant engine has further weight and cost savings, which would be expected to outweigh the above disadvantages. Furthermore, in view of the desirability of a high rotational speed for the centrifugal compressor  24 , the rotational speed of the low pressure turbine  44  can also be increased, which may result in fewer turbine stages being required for a given power, or a reduced diameter. This could in turn be used to reduce the diameter of the core engine, or increase the radius of curvature of the exhaust duct elbow  48 , thereby further resolving the above disadvantages. 
     The reduction gearbox  56  also provides distinct advantages in combination with other features of the invention, in that the turbine speed can be increased, without requiring an increase in fan rotational speed. Such an advantage is particularly desirable in high bypass turbofans, since the resultant high tip speeds in a directly driven high speed fan would otherwise result in high noise levels and reduced efficiency. 
       FIG. 3  shows a second gas turbine engine  110  in accordance with the present disclosure. The second gas turbine engine  110  is similar in many respects to the gas turbine engine  10  of  FIGS. 1 and 2 , and comprises a fan  112  configured to provide a fan bypass flow A and core flow B, both of which flow initially in a first axial direction X. Downstream of the fan  112  is an outlet guide vane  117 , which straightens the flow from the fan  112 . Also downstream of the fan  112  is a core engine inlet  116  which diverts part of the fan flow into the core. The core (comprising compressors  118 ,  132 , combustor  140  and turbines  142 ,  144 ) is housed within a generally annular core nacelle  113 . Low pressure turbine  144  is coupled to the low pressure compressor  124  by a low pressure shaft  154  which is in turn coupled to the fan  112  by a reduction gearbox  156 , and the high pressure turbine  142  is coupled to the high pressure compressor  132  by a high pressure shaft  152 . The fan  112  is housed within a generally annular fan nacelle  115 , which extends axially downstream of the fan  112 . The fan flow A is defined by the region bounded by the nacelles  113 ,  115 . Again, the engine has a bypass ratio of approximately 10. Again, the high pressure compressor  132 , combustor  140 , and high and low pressure turbines  142 ,  144  are reverse flow, while the fan  112  and low pressure compressor  124  are of conventional flow, providing flow in the aft direction X. Ducts  128 ,  146  similar to the ducts  28 ,  46  are provided. 
     The core and fan  112  are similar to that of the engine  10 , but the high pressure compressor  132  and turbine  142  differ to those of the engine  10 . The high pressure compressor  132  is of axial type, having a plurality of axial compressor stages, each stage comprising a rotating compressor rotor comprising a plurality of compressor blades, and a compressor stator comprising a plurality of stationary stator blades. The rotors are driven by the high pressure turbine  142 . The high pressure turbine  142  is of axial type, and has first  142   a  and second  142   b  stages, such that increased power can be provided to the high pressure compressor. Consequently, the high pressure compressor  132  may be capable of generating a higher compression ratio than the high pressure compressor  32  of the engine  10  at the cost of increased engine length. However, the length of the high pressure shaft  152  is substantially unaffected, since the high pressure shaft  152  extends between a forward end of the high pressure compressor  132  and an aft end of the high pressure turbine  142 . 
       FIG. 4  shows a third gas turbine engine  210 . The third gas turbine engine  210  is similar in many respects to the gas turbine engine  10  of  FIGS. 1 to 3 , and again comprises a low pressure module  260  comprising a low pressure turbine  244 , low pressure compressor  224  and load in the form of an electrical generator  264  interconnected by a low pressure coupling comprising a low pressure shaft  222 . The engine further comprises a high pressure module  262  comprising, in flow series, a high pressure compressor  232 , combustor  240 , and high pressure turbine  242 . The high pressure compressor  232  and high pressure turbine  242  are coupled by a high pressure shaft  236 . The high pressure module  262  and low pressure module  260  are axially spaced, i.e. their respective axes do not overlap 
     A core engine inlet is provided  216  which ingests airflow into the core. An intercooler duct  228  is provided, which again cools air between the low and high pressure compressors  224 ,  232 , as ambient air flows thereover. A fan may be provided to blow cold ambient air over the intercooler ducting  228 . 
     The high pressure shaft  236  is arranged to rotate about an axis generally perpendicular to a rotational axis of the low pressure shaft  222  and fan shaft  258 . An inlet  230  to the high pressure compressor  232  is provided at a radially outer end of the compressor  232 , with the compressor  232 , combustor  240  and high pressure turbine  242  being configured to direct air radially inwardly. Downstream of the high pressure turbine  242 , a duct  266  is provided to direct air forwardly toward an inlet of the low pressure turbine  244 , before it is exhausted through an exhaust duct  246 . Consequently, less bending is required relative to the embodiments shown in  FIGS. 1 to 3 . Such an arrangement is particularly suitable for a land or ship based application, in which engine diameter is of less importance. Alternatively, the generator  264  could be replaced by a gearbox driving a helicopter rotor blade in a helicopter application. 
     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 bypass ratio could be altered. The turbines and compressors could have different numbers of stages. The gearbox may have a different reduction ratio, or may be of a different arrangement. The gearbox could be omitted in some cases. 
     Though the load driven by the low pressure turbine is in the form of a fan, it will be understood that the load could alternatively comprise one or more of an electrical generator, a marine propeller, or any other suitable load. 
     It will be understood that the high pressure shaft could be provided having substantially any rotational axis relative to the low pressure shaft.