Patent Document

CROSS REFERENCE TO RELATED APPLICATION 
     This application claims the benefit of and incorporates by reference herein the disclosure of U.S. Ser. No. 61/915,885, filed Dec. 13, 2013. 
    
    
     GOVERNMENT LICENSE RIGHTS 
     This invention was made with government support under Contract No. FA8650-09-D-2923 Order 0013 awarded by the United States Air Force. The government has certain rights in the invention. 
    
    
     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 
     In one embodiment, a gas turbine engine is disclosed comprising: a combustor section; a first spool along an engine axis with a first turbine section, said first turbine section forward of said combustor section to receive a core flow along a core flow path; a second spool along said engine axis with a low pressure compressor section and a second turbine section, said low pressure compressor section aft of said combustor section to receive said core flow along said core flow path; a first intercooling turbine section forward of said combustor section to receive said core flow along said core flow path; a first intercooling turbine section bypass to selectively bypass at least a portion of said core flow through a first intercooling turbine section bypass path around said first intercooling turbine section; a second intercooling turbine section aft of said combustor section to receive said core flow along said core flow path; and a second intercooling turbine section bypass to selectively bypass at least a portion of said core flow through a second intercooling turbine section bypass path around said second intercooling turbine section. 
     In another embodiment, a gas turbine engine is disclosed comprising: a combustor section; a first spool along an engine axis with a first turbine section, said first turbine section downstream of said combustor section to receive a core flow along a core flow path; a second spool along said engine axis with a low pressure compressor section and a second turbine section, said low pressure compressor section upstream of said combustor section to receive said core flow along said core flow path; a first intercooling turbine section upstream of said combustor section to receive said core flow along said core flow path; a first intercooling turbine section bypass to selectively bypass at least a portion of said core flow through a first intercooling turbine section bypass path around said first intercooling turbine section; a second intercooling turbine section upstream of said combustor section to receive said core flow along said core flow path; and a second intercooling turbine section bypass to selectively bypass at least a portion of said core flow through a second intercooling turbine section bypass path around said second intercooling turbine section. 
     Other embodiments are also disclosed. 
    
    
     
       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 temperature-versus-entropy diagram for a high/hot day take off condition; 
         FIG. 3  is a temperature-versus-entropy diagram for a cruise condition; 
         FIG. 4  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 5  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 6  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 7  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 8  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 9  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; 
         FIG. 10  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment; and 
         FIG. 11  is a schematic close-up view of an exemplary variable cycle gas turbine engine according to another non-limiting embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein in some embodiments as a variable cycle reverse core four-spool high bypass turbofan that generally includes a fan section  22 , an intercooling turbine section (ICT)  24 , a low pressure compressor section (LPC)  26 , a high pressure compressor section (HPC)  28 , a combustor section  30 , a high pressure turbine section (HPT)  32 , and a low pressure turbine section (LPT)  36 . Additional sections may include a nozzle section (not shown), an augmentor section (not shown), various duct sections (not shown), and a geared architecture  42 G among other systems or features. The sections are defined along a central longitudinal engine axis X. 
     The engine  20  generally includes a first spool coupled by shaft  38  (illustrated schematically), a second spool coupled by shaft  40  (illustrated schematically), a third spool coupled by shaft  42  (illustrated schematically) and a fourth spool coupled by shaft  46  (illustrated schematically), all of which rotate about the engine central longitudinal axis X relative to an engine case structure  48 . 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 first through fourth spools for rotation therein. 
     The fan section  22  generally includes a bypass fan  34  and a multi-stage fan  43 . The second spool shaft  40  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 second spool shaft  40 . The geared architecture  42 G may comprise a planetary gear or a star gear. The bypass fan  34  communicates fan flow through an exit guide vane  35  and into a bypass flow path  56 , a second stream bypass flow path  58 , and a core flow path  60 . 
     The ICT  24  communicates fan flow into the core flow path  60 . The ICT  24  facilitates the selective expansion of the airflow to a lower temperature than at the exit of the multi-stage fan  43  and therefore the inlet temperature to the LPC  26  is reduced. The ICT  24  comprises a first ICT  24 A and a second ICT  24 B in serial flow communication within the core flow path  60 . The ICT section  24  is upstream of the LPC  26  such that all flow from the ICT section  24  is communicated to the input of the LPC  26 . 
     The HPC  28 , the combustor section  30 , the HPT  32 , the first LPT  36 A, the second LPT  36 B and the third LPT  36 C are also in the core flow path  60 . The core airflow is compressed by the fan section  22 , expanded limitedly by the ICT  24 , compressed monotonically by the LPC  26  and 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 A,  36 B and  36 C rotationally drive respectively the shafts  46 ,  42 ,  40  and  38  in response to the expansion. The limited expansion of the core flow by the ICT  24 A rotationally drives the first shaft  38  as a supplement to the LPT  36 C, while the limited expansion of the core flow by the ICT  24 B rotationally drives the third shaft  42  as a supplement to the LPT  36 A. 
     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 . 
     Hot core gases exiting the LPT  36  may be mixed with the second stream bypass flow path  58  at mixed flow exhaust nozzle  61 . It should be understood that various fixed, variable, convergent/divergent, two-dimensional and three-dimensional nozzle systems may be utilized herewith. 
     The LPT  36 C, the ICT  24 A and the multi-stage fan  43  are coupled by shaft  38  to define the first spool. The LPT  36 B is coupled by shaft  40  to the bypass fan  34  directly or through the geared architecture  42 G to define the second spool. The ICT  24 B, the LPC  26 , and the LPT  36 A are coupled by shaft  42  to define the third spool. The HPC  28  and the HPT  32  are coupled by shaft  46  to define the fourth spool. The LPT  36 C is the last turbine section within the core flow path  60  and thereby communicates with the mixed flow exhaust nozzle  61  which receives a mixed flow from the second stream bypass flow path  58  and the core flow path  60 . 
     In an embodiment, the ICT  24 A includes a bypass flow path  62  formed therein and the ICT  24 B includes a bypass flow path  64  formed therein. It should be appreciated that the ICT  24 A and  24 B each comprise a cold turbine located upstream of the combustor section  30  in the core flow path  60  but each includes turbine blades similar in shape to the turbine blades within the HPT  32  and the LPT  36 . 
     Allowing bypass flow through the bypass flow paths  62  and  64  within respective ICT  24 A and  24 B for cruise (at altitudes where the ambient air temperature is significantly lower than the ambient air temperature at ground level during takeoff) reduces the intercooling turbine section pressure ratio (ICT PR) and hence reduces the intercooling effect, e.g., the inlet temperature to the LPC  26  will not be significantly decreased. Reducing or eliminating bypass flow through the bypass flow paths  62  and  64  within respective ICT  24 A and  24 B for takeoff will increase ICT PR and hence increase the intercooling effect, e.g., the inlet temperature to the LPC  26  will be more significantly decreased. 
     Generally, bypass flow through the bypass flow paths  62  and  64  is reduced or eliminated for takeoff to increase the pressure ratio and intercooling effect to reduce combustor inlet temperature (T 3 ) on hot day takeoff conditions ( FIG. 2 ). Bypass flow through the bypass flow paths  62  and  64  is permitted for cruise to reduce the intercooling turbine expansion pressure ratio (ICT PR) and the intercooling effect ( FIG. 3 ). 
     With reference to  FIG. 4 , an enlarged view of the region of the engine  20  surrounding the first ICT  24 A is illustrated. In an embodiment, a first bypass door  66  is provided which when closed reduces or eliminates flow through the bypass flow path  62 . As shown in  FIG. 4 , the first bypass door  66  may be closed during hot day takeoff conditions, forcing all or most of the core gas flow to pass through the turbine section  68  of ICT  24 A, causing the core flow gas to be expanded and producing an intercooling effect to the hot ambient air entering the core  60 . Similarly, with reference to  FIG. 5 , an enlarged view of the region of the engine  20  surrounding the second ICT  24 B is illustrated. In an embodiment, a second bypass door  70  is provided which when closed reduces or eliminates flow through the bypass flow path  64 . As shown in  FIG. 5 , the second bypass door  70  may be closed during hot day takeoff conditions, forcing all or most of the core gas flow to pass through the turbine section  72  of ICT  24 B, causing the core flow gas to be expanded and producing an additional intercooling effect to the hot ambient air flowing in the core  60 . 
     With reference to  FIG. 6 , an enlarged view of the region of the engine  20  surrounding the first ICT  24 A is illustrated. As shown in  FIG. 6 , the first bypass door  66  may be opened during cruise conditions, allowing some of the core gas flow to pass through the bypass flow path  62 . The core flow passing through the bypass flow path  62  does not pass through the turbine section  68  of ICT  24 A, causing less of the core flow gas to be expanded and producing a reduced intercooling effect to the relatively cold ambient air entering the core  60 . Similarly, with reference to  FIG. 7 , an enlarged view of the region of the engine  20  surrounding the second ICT  24 B is illustrated. As shown in  FIG. 7 , the second bypass door  70  may be opened during cruise conditions, allowing some of the core gas flow to pass through the bypass flow path  64 . The core flow passing through the bypass flow path  64  does not pass through the turbine section  72  of ICT  24 B, causing less of the core flow gas to be expanded and producing a reduced additional intercooling effect to the relatively cold ambient air flowing in the core  60 . 
     With reference to  FIG. 8 , an enlarged view of the region of the engine  20  surrounding the first ICT  24 A is illustrated. In an embodiment, a third bypass door  74  is provided which when closed reduces or eliminates flow through the turbine section  68 . As shown in  FIG. 8 , the first bypass door  66  may be closed and the third bypass door  74  may be opened during hot day takeoff conditions, forcing all or most of the core gas flow to pass through the turbine section  68  of ICT  24 A, causing the core flow gas to be expanded and producing an intercooling effect to the hot ambient air entering the core  60 . Similarly, with reference to  FIG. 9 , an enlarged view of the region of the engine  20  surrounding the second ICT  24 B is illustrated. 
     In an embodiment, a fourth bypass door  76  is provided which when closed reduces or eliminates flow through the turbine section  72 . As shown in  FIG. 9 , the second bypass door  70  may be closed and the fourth bypass door  76  may be opened during hot day takeoff conditions, forcing all or most of the core gas flow to pass through the turbine section  72  of ICT  24 B, causing the core flow gas to be expanded and producing an additional intercooling effect to the hot ambient air flowing in the core  60 . 
     With reference to  FIG. 10 , an enlarged view of the region of the engine  20  surrounding the first ICT  24 A is illustrated. As shown in  FIG. 10 , the first bypass door  66  may be opened and the third bypass door  74  may be closed during cruise conditions, forcing all or most of the core gas flow to pass through the bypass flow path  62 . The core flow passing through the bypass flow path  62  does not pass through the turbine section  68  of ICT  24 A, therefore little or none of the core flow gas is expanded, thereby producing little or no intercooling effect to the relatively cold ambient air entering the core  60 . Similarly, with reference to  FIG. 11 , an enlarged view of the region of the engine  20  surrounding the second ICT  24 B is illustrated. As shown in  FIG. 11 , the second bypass door  70  may be opened and the fourth bypass door  76  may be closed during cruise conditions, forcing all or most of the core gas flow to pass through the bypass flow path  64 . The core flow passing through the bypass flow path  64  does not pass through the turbine section  72  of ICT  24 B, therefore little or none of the core flow gas to be expanded, thereby producing little or no additional intercooling effect to the relatively cold ambient air flowing in the core  60 . 
     In an embodiment, either ICT  24 A or ICT  24 B may be removed from the engine  20 , with the remaining ICT  24  providing the intercooling effect. Additionally, LPT  36 B and LPT  36 C may be replaced with a single turbine to form a three spool engine in an embodiment, with the single turbine driving all of the forward components. The bypass doors  66 ,  70 ,  74  and  76  are disposed circumferentially around the engine axis X, and each may comprise multiple door sections disposed between intervening support struts (not shown) in an embodiment. In another embodiment, the bypass doors  66 ,  70 ,  74  and  76  may comprise variable vanes. 
     With reference to  FIGS. 2 and 3 , a conventional engine cycle is defined thermodynamically on a 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). 
     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. 2 ) with the same inlet temperature and pressure to the engine: TB=Tb; and PB=Pb, as well as the same temperature and pressure at the exit of the multi-stage fan (MSF)  43 : TC=Tc; and PC=Pc. 
     For the inventive engine disclosed herein the inlet temperature and pressure to the LPC  26  are Td and Pd, respectively. The ICT  24 A/ 24 B expands the core flow so that the inlet temperature and pressure to the LPC  26  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 LPC  26  and 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 LPC  26  and 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 LPC  26  and 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. 2 ). 
     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: TB=Tb and PB=Pb. Note that the inlet temperature and pressure at the cruise condition ( FIG. 3 ) are less than the inlet temperature and pressure at the hot day takeoff condition ( FIG. 2 ). 
     At the cruise condition, both the conventional and inventive engine have the same temperature and pressure at the exit of the MSF  43 ; TC=Tc and PC=Pc. For the inventive engine  20 , the inlet temperature and pressure to the LPC  26  are Td and Pd, respectively. At the cruise condition, the ICT  24  expands the exit flow of the MSF  43  so that the inlet temperature and pressure to the LPC  26  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 use of the bypass doors  66 ,  70 ,  74  and  76 . At the cruise condition as well as the hot day takeoff condition, the LPC  26  and 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 LPC  26  and HPC  28 . The pressure ratio (PR) of the LPC  26  and 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 LPC  26  and HPC  28  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 LPC  26  and 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. 3 , 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. 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. 
     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 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. 
     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.

Technology Category: 4