Abstract:
A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a combustor section, a low spool including a fan section and an intercooling turbine section along an engine axis aft of the fan section and forward of a combustor section. The fan section and the intercooling turbine section are coaxial with the low spool. A high spool along the engine axis includes a high pressure compressor section and a high pressure turbine section, the high pressure compressor section is aft of the intercooling turbine section and forward of the combustor section, and the high pressure turbine section is aft of the combustor section. The high pressure compressor section and the high pressure turbine section are coaxial with the high spool. A method of operating a gas turbine engine is also disclosed.

Description:
REFERENCE TO RELATED APPLICATIONS 
     The present disclosure claims priority to U.S. Provisional Patent Application No. 61/551,107, filed Oct. 25, 2011. 
    
    
     BACKGROUND 
     The present disclosure relates to gas turbine engines, and more particularly to a variable cycle gas turbine engine. 
     Variable cycle engines power high performance aircraft over a range of operating conditions yet achieve countervailing objectives such as high specific thrust and low fuel consumption. The variable cycle engine essentially alters a bypass ratio during flight to match requirements. This facilitates efficient performance over a broad range of altitudes and flight conditions to generate high thrust when needed for high energy maneuvers yet also optimize fuel efficiency for cruise and loiter conditions. 
     SUMMARY 
     A gas turbine engine according to an exemplary aspect of the present disclosure includes a low spool along an engine axis with a fan section and an intercooling turbine section forward of a combustor section. A high spool along the engine axis with a high pressure compressor section and a high pressure turbine section, the high pressure compressor section aft of the intercooling turbine section and forward of the combustor section, the high pressure turbine section aft of the combustor section. 
     A gas turbine engine according to an exemplary aspect of the present disclosure includes a low spool along an engine axis with a fan section and an intercooling turbine section forward of said combustor section. A high spool along the engine axis with an intercooling turbine section, a high pressure compressor section and a high pressure turbine section, the intercooling turbine section and the high pressure compressor section forward of the combustor section and the high pressure turbine section aft of the combustor section. 
     A method of operating a gas turbine engine according to an exemplary aspect of the present disclosure includes modulating a guide vane of an intercooling turbine section forward of a high pressure compressor section to reduce the intercooling turbine expansion pressure ratio (ICT PR) during a first flight condition; and modulating the guide vane of the intercooling turbine section to increase the intercooling turbine expansion pressure ratio (ICT PR) during a second flight condition. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a general schematic view of an exemplary variable cycle gas turbine engine according to one non-limiting embodiment; 
         FIG. 2  is a general schematic view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 3  is a temperature-versus-entropy diagram for a high/hot day take off condition with example temperature distributions; and 
         FIG. 4  is a temperature-versus-entropy diagram for a cruise condition with example temperature distributions. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a variable cycle two-spool bypass turbofan that generally includes a fan section  22 , an intercooling turbine section (ICT)  24 , a high pressure compressor section (HPC)  28 , a combustor section  30 , a high pressure turbine section (HPT)  32 , a low pressure turbine section (LPT)  36 , and a nozzle section  38 . Additional sections may include an augmentor section  38 A, various duct sections, and among other systems or features such as a geared architecture  42 G, which may be located in various other engine sections than that shown such as, for example, aft of the LPT. The sections are defined along a central longitudinal engine axis X. 
     The engine  20  generally includes a low spool  42  and a high spool  46  which rotate about the engine central longitudinal axis X relative to an engine case structure  48 . It should be appreciated that other architectures, such as a three-spool architecture, will also benefit herefrom. 
     The engine case structure  48  generally includes an outer case structure  50 , an intermediate case structure  52  and an inner case structure  54 . It should be understood that various structures individually or collectively within the engine may define the case structures  50 ,  52 ,  54  to essentially define an exoskeleton that supports the spools  42 ,  46  for rotation therein. 
     The fan section  22  generally includes a bypass fan  34  and a low pressure compressor (LPC)  44 . The low spool  42  drives the bypass fan  34  directly or through a geared architecture  42 G to drive the bypass fan  34  at a lower speed than the low spool  42 . The bypass fan  34  communicates fan flow into a bypass flow path  56 , a second stream bypass flow path  58 , and a core flow path  60 . 
     The low pressure compressor  44  in the disclosed non-limiting embodiment includes two stages downstream of the bypass fan  34 . It should be appreciated that various fan stages, as well as a low pressure compressor section, may alternatively or additionally be provided. The low pressure compressor  44  is within an intermediate flow path  57  upstream of a split  59 S between the second stream bypass flow path  58  and the core flow path  60  but downstream of a split  57 S between the bypass flow path  56  and the intermediate flow path  57  such that all airflow from the low pressure compressor  44  is expanded through the ICT  24 . The ICT  24  facilitates the expansion of the airflow to a lower temperature than at the exit of the low pressure compressor  44  and therefore the inlet temperature to the HPC  28  is reduced. 
     The ICT  24  communicates fan flow into the second stream bypass flow path  58  and the core flow path  60 . The ICT  24  is downstream of the low pressure compressor  44  such that all flow from the ICT  24  is selectively communicated into the second stream bypass flow path  58  and the core flow path  60 . That is, the ICT  24  is upstream of the split  59 S between the second stream bypass flow path  58  and the core flow path  60  but downstream of the split  57 S between the bypass flow path  56  and the intermediate flow path  57 . 
     The HPC  28 , the combustor section  30 , the HPT  32 , and the LPT  36  are in the core flow path  60 . These sections are referred to herein as the engine core. The core airflow is compressed by the fan section  22 , expanded limitedly by the ICT  24 , compressed monotonically by the HPC  28 , mixed and burned with fuel in the combustor section  30 , then expanded over the HPT  32  and the LPT  36 . The turbines  32 ,  36  rotationally drive the respective high spool  46  and low spool  42  in response to the expansion. The limited expansion of the core flow by the ICT  24  rotationally drives the low spool  42  as a supplement to the LPT  36 . 
     The second stream bypass flowpath  58  permits the match of the ICT  24  exit flow to the flow demand into the HPC  28 . That is, the ICT  24  expands fan section  22  flow to reduce inlet temperatures to the HPC  28 . 
     The bypass flow path  56  is generally defined by the outer case structure  50  and the intermediate case structure  52 . The second stream bypass flowpath  58  is generally defined by the intermediate case structure  52  and the inner case structure  54 . The core flow path  60  is generally defined by the inner case structure  54 . The second stream bypass flow path  58  is defined radially inward of the bypass flow path  56  and the core flow path  60  is radially inward of the bypass flowpath  58 . 
     The nozzle section  38  may include a bypass flow exhaust nozzle  62  (illustrated schematically) which receives flow from bypass flow path  56  and a mixed flow exhaust nozzle  64  which receives a mixed flow from the second stream bypass flowpath  58  and the core flow path  60 . It should be understood that various fixed, variable, convergent/divergent, two-dimensional and three-dimensional nozzle systems may be utilized herewith. 
     The low pressure compressor  44 , the ICT  24 , and the LPT  36  are coupled by a low shaft  66  (illustrated schematically) which is also coupled to the bypass fan  34  directly or through the geared architecture  42 G. In the disclosed non-limiting embodiment, the low pressure compressor  44  includes a first stage guide vane  68 A, a first stage fan rotor  70 A, a second stage guide vane  68 B and a second stage fan rotor  70 B. It should be appreciated that various systems may be utilized to activate the variable inlet guide vanes and variable stators. It should also be understood that other single or multistage architectures may alternatively or additionally be provided such as various combinations of a fixed or variable vanes. 
     The HPC  28  and the HPT  32  are coupled by a high shaft  90  (illustrated schematically) to define the high spool  46 . In the disclosed non-limiting embodiment, the HPC  28  upstream of the combustor section  30  includes a multiple of stages each with a rotor  84  and vane  86 . It should be understood that the HPC  28  may alternatively or additionally include other compressor section architectures which, for example, include additional or fewer stages each with or without various combinations of variable or fixed guide vanes. It should also be understood that each of the turbine sections  32 ,  36  may alternatively or additionally include other turbine architectures which, for example, include additional or fewer stages each with or without various combinations of variable or fixed guide vanes. 
     The HPT  32  in the disclosed non-limiting embodiment, includes a multiple of stages (two shown) with first stage variable high pressure turbine guide vanes  88 A, a first stage high pressure turbine rotor  90 A, second stage variable high pressure turbine guide vanes  88 B and a second stage high pressure turbine rotor  90 B. It is desirable to have variable turbine vanes both in the ICT  24  and the HPT  32  to facilitate an optimum cycle efficiency; however, significant improvement to the thermodynamic cycle is achieved by the presence of the ICT  24  alone even if the HPT  32  is not variable. 
     The LPT  36  in the disclosed non-limiting embodiment, includes a multiple of stages (four shown), each stage with variable low pressure turbine inlet guide vanes  98  upstream of a respective low pressure turbine rotor  100 . It is desirable to have variable turbine vanes both in the ICT  24  and the LPT  36  to facilitate an optimum cycle efficiency; however, significant improvement to the thermodynamic cycle is achieved by the presence of the ICT  24  alone even if the LPT  36  is not variable. The LPT  36  is the last turbine section within the core flow path  60  and thereby communicates with the mixed flow exhaust nozzle  64  which receives a mixed flow from the second stream bypass duct  58  and the core flow path  60 . The augmentor section  38 A among other systems or features may be located downstream of the LPT  36 . 
     The ICT  24  is coupled to the low shaft  66  and is driven with the low spool  42 . In the disclosed non-limiting embodiment, the ICT  24  includes upstream intercooling turbine variable vanes  76 , an intercooling turbine rotor  78  and downstream intercooling turbine variable vanes  80 . The downstream intercooling turbine variable vanes  80  is immediately upstream of the split  59 S between the second stream bypass flow path  58  and the core flow path  60 . It should be appreciated that the intercooling turbine rotor  78  is a cold turbine located forward of the combustor section  30  but includes turbine blades similar in shape to the turbine blades within the HPT  32  and the LPT  36 . 
     In the disclosed non-limiting embodiment, inlet guide vanes (IGVs)  92  immediately downstream of the intercooling turbine variable vanes  80  are within the core flow path  60  downstream of the split  59 S. At cruise, the IGVs  92  are closed, and at takeoff, the IGVs  92  are opened. These IGV settings are similar in function to the downstream intercooling turbine variable vanes  80  such that the downstream intercooling turbine variable vanes  80  may be eliminated in one alternative non-limiting embodiment to simplify the engine architecture. 
     The high spool  46  is independent of the low spool  42 . The HPC  28 , the combustor section  30  and the HPT  32  set the core flow and the HPT  32  is the choke point that the HPC  28  feeds. The speed and power output of the HPT  32  determines the flow pumping rate and the pressure rise of the HPC  28  for a given combustor exit temperature that affects the status of the choke point. 
     The HPC  28  demands a certain inlet flow based on combustor exit temperature. There is a balance between the pressure at the exit of the HPC  28  and the temperature at the choke point. Higher combustor section exit temperature enables higher pressure at the exit of the HPC  28  and a simultaneous solution of continuity (flowrate) and power from the HPT  32 . 
     The low pressure compressor (LPC)  44  communicates the flow required by the HPC  28  but this may result in a flow that stalls the LPC  44 . The flow which exits the ICT  24  is thereby split between the HPC  28  and the second stream bypass flowpath  58  around the HPC  28  to facilitate the ability of the LPC  44  to provide a matched flow to the HPC  28  that is not a stalled condition. That is, the flow which exits the ICT  24  is selectively split into the second stream bypass flowpath  58  to provide a matched flow into the HPC  28 . 
     Opening the upstream intercooling turbine variable vanes  76  and closing the downstream intercooling turbine variable vanes  80  for cruise reduces the intercooling turbine section pressure ratio (ICT PR) and hence reduces the intercooling effect, e.g., the inlet temperature to the HPC  28  will not be significantly decreased. Closing the upstream intercooling turbine variable vanes  76  and opening the downstream intercooling turbine variable vanes  80  for takeoff will increase ICT PR and hence increase the intercooling effect, e.g., the inlet temperature to the HPC  28  will be more significantly decreased. In the disclosed non-limiting embodiment, the intercooling turbine variable vanes  76 ,  80  are modulated between a 5%-25% closed position. That is, the intercooling turbine variable vanes  76 ,  80  are never completely closed. 
     Generally, the intercooling variable vanes  80  are opened for takeoff to increase the pressure ratio and intercooling effect to reduce combustor inlet temperature (T 3 ) on hot day conditions ( FIG. 3 ). The intercooling variable vanes  80  are closed for cruise to reduce the intercooling turbine expansion pressure ratio (ICT PR) and the intercooling effect ( FIG. 4 ). 
     The second stream bypass flowpath  58  permits the match of the ICT  24  exit flow to the flow demand into the HPC  28 . That is, the ICT  24  expands fan section  22  flow to reduce inlet temperatures to the HPC  28 . 
     With reference to  FIG. 2 , in another disclosed, non-limiting embodiment, the ICT  24 H is coupled to the high shaft  90 ′ and is driven with the high spool  46 ′. When the ICT  24 H is on the high spool  46 ′, the ICT  24 H operates relatively faster and the amount of turning in the turbine is relatively smaller as compared to the when the ICT  24  is on the low spool  42  ( FIG. 1 ). The ICT  24 H on the high spool  46 ′ is more efficient thermodynamically but relatively less efficient aerodynamically. There are thereby some tradeoffs involved, but otherwise the architecture and operation is generally equivalent. 
     With reference to  FIGS. 3 and 4 , a conventional engine cycle is defined thermodynamically on an example Temperature-Entropy diagram by the points A, B, C, E, F, G, H. The priority for improvement of the thermodynamic efficiency of the engine is to increase the area enclosed by the points B, C, E, F, G, but especially doing so by “raising the roof” of points (E) and (F) that correspond respectively to an increase in the overall PR of the engine compression system (E) and an increase in the inlet temperature to the HPT  32  (F). It should be appreciated that the temperatures are merely exemplary to one disclosed non-limiting embodiment and should not be considered otherwise limiting. 
     The inventive engine cycle disclosed herein is defined thermodynamically on the Temperature-Entropy diagram by points a, b, c, d, e, f, g, h. The priority is improvement of the cruise condition efficiency where significant fuel is consumed. 
     Both the conventional engine and the inventive engine  20  architectures disclosed herein operate at the hot day takeoff condition ( FIG. 3 ) with the same inlet temperature and pressure to the engine, the same inlet temperature and pressure to the low pressure compressor  44 : TB=Tb; and PB=Pb, as well as the same temperature and pressure at the exit of the low pressure compressor  44 : TC=Tc; and PC=Pc. 
     For the inventive engine disclosed herein the inlet temperature and pressure to the HPC  28  are Td and Pd, respectively. The ICT  24  expands the exit flow of the low pressure compressor  44  so that the inlet temperature and pressure to the HPC  28  of the inventive engine are decreased significantly to achieve an intercooling effect on the temperature of compression, that is, Td&lt;TC and Pd&lt;PC. 
     For both the conventional engine and the inventive engine, the exit condition of the HPC  28  is the inlet condition of the combustor section  30 . Both the conventional engine and the inventive engine operate at the hot day takeoff condition with the same combustor inlet temperature (T 3 ), where TE=Te, and with the same HPT  32  first rotor inlet temperature, (T 4 . 1 ), where TF=Tf. This is consistent with utilization of the same materials and mechanical design technologies for both the conventional and inventive engine. 
     The pressure ratio (PR) of the HPC  28  of the inventive engine is significantly higher than the PR of the conventional engine, that is, Pe:Pd&gt;PE:PC. The temperature ratio (TR) of the HPC  28  of the inventive engine is significantly higher than the TR of the conventional engine, that is, Te:Td&gt;TE:TC. The higher PR of the HPC  28  of the inventive engine  20  is achievable, for example, with additional compressor section stages. 
     Neglecting combustor pressure losses, the pressures, PE and PF for the conventional engine are the same. The pressures, Pe and Pf, for the inventive engine are the same, but PE&gt;Pe and PF&gt;Pf; this is attributable to the pressure expansion in the ICT  24 . 
     Both the conventional engine and the inventive engine operate with the same HPT  32  first rotor inlet temperature (T 4 . 1 ), and TF=Tf at the hot day takeoff condition. At the hot day takeoff condition, both the conventional engine and the inventive engine operate with the same exit pressure from the turbine section so that PG=Pg, but not the same exit temperature from the turbine section, that is, Tg&gt;TG. 
     The thermodynamic cycle efficiency of an engine generally is proportional to the ratio of two areas on the Temperature-Entropy diagram. That is, the numerator area and the denominator area form this ratio of areas. For the conventional engine, the numerator area is enclosed by the points B, C, E, F, and G, while the denominator area is enclosed by the points H, G, B, and A. For the inventive engine, the numerator area is enclosed by the points b, c, d, e, f, and g, while the denominator area is enclosed by the points h, g, b, and a. 
     At the hot day takeoff condition, the numerator area of the conventional engine is greater than or equal to the numerator area of the inventive engine, while the denominator area of the conventional engine is less than the denominator area of the inventive engine; thus, the thermodynamic efficiency of the conventional engine is relatively better than the inventive engine at the hot day takeoff condition ( FIG. 3 ). 
     The priority, however, is to improve the thermodynamic cycle efficiency at the cruise condition where much of the fuel is consumed. Both the conventional engine and the inventive engine operate at the cruise condition with the same inlet temperature and pressure to the engine and the same inlet temperature and pressure to the low pressure compressor  44 ; TB=Tb and PB=Pb. Note that the inlet temperature and pressure at the cruise condition ( FIG. 4 ) are less than the inlet temperature and pressure at the hot day takeoff condition ( FIG. 3 ). 
     At the cruise condition, both the conventional and inventive engine have the same temperature and pressure at the exit of the low pressure compressor  44 ; TC=Tc and PC=Pc. For the inventive engine  20 , the inlet temperature and pressure to the HPC  28  are Td and Pd, respectively. At the cruise condition, the ICT  24  expands the exit flow of the low pressure compressor  44  so that the inlet temperature and pressure to the HPC  28  of the inventive engine are not decreased significantly to obtain a smaller intercooling effect on the temperature within the compressor section; regardless, Td&lt;TC and Pd&lt;PC. 
     The expansion of the ICT  24  is selectively less at the cruise condition and this is obtained by modulation of the variable vanes  76  and  80 . At the cruise condition as well as the hot day takeoff condition, the HPC  28  of the inventive engine  20  has a higher PR than the conventional engine and the higher PR is achieved for example, with additional stages of compression in the HPC  28 . The pressure ratio (PR) of the HPC  28  of the inventive engine is significantly higher than the PR of the conventional engine, that is, Pe:Pd&gt;PE:PC. The temperature ratio (TR) of the HPC of the inventive engine is significantly higher than the TR of the conventional engine, that is, Te:Td&gt;TE:TC. At the cruise condition, the HPC  28  exit pressure and exit temperature of the inventive engine are higher than the conventional engine, that is, Pe&gt;PE and Te&gt;TE. 
     For both the conventional engine and the inventive engine, the exit condition of the HPC  28  is the inlet condition of the combustor section  30 . Neglecting combustor pressure losses, the pressures, PE and PF for the conventional engine are the same. The pressures, Pe and Pf for the inventive engine are the same, but at the cruise condition, Pe&gt;PE and Pf&gt;PF; this is attributed to the deliberately smaller expansion of pressure in the ICT  24  and the higher PR of the HPC  28  of the inventive engine. 
     Application of the same materials and mechanical design technologies to both the conventional and inventive engine is limiting at the hot day takeoff condition but not at the cruise condition provided T 3  and T 4 . 1  at the cruise condition are lower than at the hot day takeoff condition. 
     At the cruise condition, HPT  32  first rotor inlet temperature (T 4 . 1 ) of the inventive engine is greater than T 4 . 1  of the conventional engine; that is, Tf&gt;TF at the cruise condition. At the cruise condition, both the conventional engine and the inventive engine operate with the same exit pressure of the turbine section so that PG=Pg, and the same exit temperature from the turbine section, TG=Tg. 
     With reference to  FIG. 4 , at the cruise condition, the numerator area of the inventive engine is greater than the numerator area of the conventional engine, while the denominator areas of the conventional engine and the inventive engine are the same; thus, the thermodynamic efficiency of the inventive engine is greater than the conventional engine at the cruise condition. The larger numerator area of the inventive engine is evident by comparison between the two sectional areas of the Temperature-Entropy diagram at the cruise condition. 
     The first sectional area is enclosed by the points z, e, f, and F, while the second sectional area is enclosed by the points C, E, z, and d. The first sectional area yields an increase in the numerator area of the inventive engine while the second sectional area yields a reduction in the numerator area of the inventive engine. The first sectional area is greater than the second sectional area to yield a net increase in the numerator area of the inventive engine disclosed herein versus the numerator area of the conventional engine. 
     The ICT  24  effectively “raises the roof” of the thermodynamic cycle of the engine at the cruise condition with the same materials and mechanical design constraints as a conventional engine architecture at the hot day takeoff condition. 
     It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the engine but should not be considered otherwise limiting. 
     Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.