Abstract:
A gas turbine engine includes a turbomachinery core operable to generate a flow of pressurized combustion gases at a variable flow rate, while maintaining a substantially constant core pressure ratio; a rotating fan disposed upstream of the core, the fan adapted to extract energy from the core and generate a first flow of air which is compressed at a first pressure ratio; and at least a first bypass duct surrounding the core downstream of the fan adapted to selectively receive at least a first selected portion of the first flow which is compressed at a second pressure ratio lower than the first pressure ratio, and to bypass the first selected portion around the core, thereby varying a bypass ratio of the engine. The fan is adapted to maintain a flow rate of the first flow substantially constant, independent of the bypass ratio.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engines and more particularly to a turbofan engine in which the fan flow can be modulated independent of the fan speed. 
     Future mixed mission morphing aircraft as well as more conventional mixed mission capable military systems that have a high value of take-off thrust/take-off gross weight, i.e., a thrust loading in the 0.8-1.2 category, present many challenges to the propulsion system. They need efficient propulsion operation at diverse flight speeds, altitudes, and particularly at low power settings where conventional engines operate at inefficient off-design conditions both in terms of uninstalled performance and, to an even greater degree, fully installed performance that includes the impact of spillage drag losses associated with supersonic inlets. 
     When defining a conventional engine cycle and configuration for a mixed mission application, compromises have to be made in the selection of fan pressure ratio, bypass ratio, and overall pressure ratio to allow a reasonably sized engine to operate effectively at both subsonic and supersonic flight conditions. In particular, the fan pressure ratio and related bypass ratio selection needed to obtain a reasonably sized engine capable of developing the thrusts needed for combat maneuvers and supersonic operation are non-optimum for efficient low power subsonic flight. Basic uninstalled subsonic engine performance is compromised and fully installed performance suffers even more due to the inlet/engine flow mismatch that occurs at reduced power settings. 
     BRIEF SUMMARY OF THE INVENTION 
     The above-mentioned shortcomings in the prior art among others are addressed by the present invention, which according to one aspect provides a gas turbine engine, including: a turbomachinery core operable to generate a flow of pressurized combustion gases at a variable flow rate, while maintaining a substantially constant core pressure ratio; a rotating fan disposed upstream of the core, the fan adapted to extract energy from the core and generate a first flow of air which is compressed at a first pressure ratio; and at least a first bypass duct surrounding the core downstream of the fan adapted to selectively receive at least a first selected portion of the first flow which is compressed at a second pressure ratio lower than the first pressure ratio, and to bypass the first selected portion around the core, thereby varying a bypass ratio of the engine. The fan is adapted to maintain a flow rate of the first flow substantially constant, independent of the bypass ratio. 
     According to another aspect of the invention, a gas turbine engine includes: a first turbomachinery core operable to generate a flow of pressurized combustion gases at a first design flow rate; a second turbomachinery core operable to generate a flow of pressurized combustion gases at a second design flow rate substantially greater than said first flow rate; a low pressure turbine disposed downstream of the first and second cores and adapted to extract energy from at least one of the cores; a rotating fan disposed upstream of the first and second cores and adapted to be mechanically driven by the low pressure turbine to generate a first flow of air which is compressed at a first pressure ratio; and means for selectively ducting a portion of the first flow of air through one of the cores to the low pressure turbine. 
     According to another aspect of the invention, a method of operating a gas turbine engine includes: burning a fuel in at least one turbomachinery core to produce a first flow of pressurized combustion gases; extracting energy from the first flow of pressurized combustion gases and using the energy to generate a first flow of pressurized air with a rotating fan, wherein the first flow is compressed at a first pressure ratio; passing a first portion of the first flow through the at least one core; and bypassing at least a selected, variable second portion of the first flow around the at least one core through a first bypass duct, resulting in the engine operating at a selected bypass ratio; wherein a total flow rate of the first flow is maintained substantially constant regardless of the bypass ratio, and wherein an operating pressure ratio of the engine remains substantially constant during engine operation. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
         FIG. 1  is a schematic cross-sectional view of a convertible gas turbine engine constructed according to an aspect of the present invention; 
         FIG. 2A  is an enlarged portion of  FIG. 1 , showing a fan of the engine of  FIG. 1  in a partial power operating mode; 
         FIG. 2B  is an enlarged portion of  FIG. 1 , showing a fan of the engine of  FIG. 1  in a maximum power operating mode; 
         FIG. 3  is a graph depicting a specific fuel consumption versus thrust characteristic of a gas turbine engine of the present invention compared to a prior art gas turbine engine; 
         FIG. 4  is a schematic cross-sectional view of a gas turbine engine constructed according to another aspect of the present invention; 
         FIG. 5  is a schematic cross-sectional view of a gas turbine engine constructed according to another aspect of the present invention; and 
         FIG. 6  is a schematic cross-sectional view of a gas turbine engine constructed according to yet another aspect of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates a portion of an exemplary convertible gas turbine engine, generally designated  10 . The engine  10  has a fan  12 , a first core  14 A including a first compressor  16 A, first combustor  18 A, and first high pressure turbine  20 A, and a second core  14 B which includes a second compressor  16 B, second combustor  18 B, and second high pressure turbine  20 B. The fan  12  is driven by a low pressure turbine  22  disposed downstream of the cores  14 A and  14 B. The cores  14 A and  14 B can optionally have the same design pressure ratios, but have different design flow rates, with the second core  14 B having a higher design and maximum flow. An exhaust duct  24  is disposed downstream of the low pressure turbine  22 , and a convergent-divergent exhaust nozzle  26  is disposed downstream of the exhaust duct  24 . The throat  28  of the exhaust nozzle  26  may have a variable area “A 8 ”, and if desired an afterburner (not shown) may be incorporated upstream of the exhaust nozzle  26 . 
     Core inlet ducting  30  provides a flow path between the fan  12  and the first and second cores  14 A and  14 B. The core inlet ducting  30  is provided with inlet flow control flaps  32  that permit the fan flow to be selectively ducted to either the first core  14 A or the second core  14 B. Similarly, core outlet ducting  34  provides a flow path between the first and second cores  14 A and  14 B and the low pressure turbine  22 . The core outlet ducting  34  is provided with outlet flow control flaps  36  that permit the low pressure turbine  22  to be selectively connected to either the first core  14 A or the second core  14 B. 
     An inner bypass duct  38  surrounds the cores  14 A and  14 B and forms a flowpath from the fan  12  to the exhaust duct  24 . An outer bypass duct  40  surrounds the inner bypass duct  38  and forms a second, separate flowpath between the fan  12  and the exhaust duct  24 . The inner and outer bypass ducts  38  and  40  merge downstream of the fan  12  to form a single main bypass duct  41 . A front mixer  43  is disposed at the juncture of the core and bypass duct flow streams to promote efficient mixing of the two streams. If needed, the front mixer  43  may be of the type which can selectively vary its open area, This type of mixer is sometimes referred to as a variable area bypass injector (“VABI”). A mixer  42  (for example a lobed or chute-type mixer) is disposed downstream of the low pressure turbine  22 , at the juncture of the core and bypass duct flow streams, to promote efficient mixing of the two streams. The mixer  42  may also be a VABI, which can be used to control the back pressure on the fan  112 . 
     The fan  12 , shown in more detail in  FIGS. 2A and 2B , is of a “convertible” type. The pressure ratio of the flow discharged into the bypass duct (i.e. the fan tip overall PR) can be varied while core supercharge pressure ratio is maintained and the total mass flow rate of the fan  12  is held constant. For purposes of explanation the fan  12  is considered to include a “primary section”  44  and “secondary section”  46 , both contained within a fan duct  48 . The exact number of fan stages in each section, the design pressure ratio of the individual sections, and other design characteristics may be varied to suit a particular application. In the illustrated example, The primary section  44  includes, in flow sequence, a row of non-rotating fan inlet guide vanes or “IGVs”  50 , a first stage of rotating fan blades  52 , a row of non-rotating interstage vanes  54 , and a second stage of rotating fan blades  56 . The inlet guide vanes  50  may have their angle of attack with respect to the airflow and their open flow area selectively changed by using an actuator  58  of a known type. Optionally, the interstage vanes  54  may have their angle of attack with respect to the airflow and their open flow area selectively changed by using an actuator of a known type (not shown). 
     The outer bypass duct  40  is connected to the fan duct  48  between the primary section  44  and the secondary section  46 . A mode control valve  60  may be selectively moved between an open position, in which a portion of the discharge from the primary section  44  can flow into the outer bypass duct  40 , and a closed position, in which all of the discharge from the forward portion flows into the secondary section  46 . 
     The secondary section  46  includes, in flow sequence, a row of non-rotating secondary guide vanes  62  with radially inner and outer sections  62 A and  62 B, and a stage of rotating fan blades  64 . Independent of each other, the inner and outer sections  62 A and  62 B of the secondary guide vanes  62  may have their angle of attack with respect to the airflow and their open flow area selectively changed by using an actuator of a known type (not shown). 
     A fan flow control flap  66  is positioned downstream of the secondary section  46  and serves to selectively alter the proportion of fan discharge flow between the core inlet ducting  30  and the inner bypass duct  38 . In conjunction with a stationary splitter structure  68 , the fan flow control flap  66  effectively divides the discharge flow of the secondary section  46  into radially inner, center, and outer streams. In a “low flow” position, i.e. a high bypass mode, shown in  FIG. 2A , the inner stream is passed to the core inlet ducting  30 , while the center and outer streams pass to the inner bypass duct  38 . In a “high flow” position, i.e. a low bypass mode, shown in  FIG. 2B , the inner and center streams pass to the core inlet ducting  30 , while the outer stream passes to the inner bypass duct  38 . The portion of the total fan flow entering one or both of the bypass ducts  38  or  40  is referred to as the “bypass flow”, while the remainder is referred to as the “core flow”. 
     The convertible engine  10  operates in one of two modes. In a first mode, intended for partial-power operation, the second core  14 B is shut off. The core flow from the inner stream is passed to the core inlet ducting  30 , while the center and outer streams pass to the inner bypass duct  38 . The mode control valve  60  is open and a portion of the flow from the primary section  44  of the fan  12  flows through the outer bypass duct  40 . The flow split between the inner and outer bypass ducts  38  and  40  can be modulated by varying the position of the secondary guide vanes  62 . The inlet flow control flaps  32  duct the core flow to the first core  14 A. The core flow is pressurized by the first compressor  16 A, mixed with fuel in the first combustor  18 A, and ignited, thereby generating pressurized combustion gases. Some work is extracted from these gases by the first high pressure turbine  20 A which drives the second compressor  16 B via a shaft (not shown). The outlet flow control flaps  36  connect the first core  14 A to the low pressure turbine  22 , which in turn drives the fan  12  via an inner shaft  70 . This mode of operation is termed “double bypass” and has a relatively low fan pressure ratio and high bypass ratio. 
     In a second mode, intended for maximum-power operation, the first core  14 A is shut off. The flow from the inner and center streams are passed to the core inlet ducting  30 , while the outer stream passes to the inner bypass duct  38 . The mode control valve  60  is closed. The flow split between the inner bypass duct  38  and the core flow can be modulated by varying the position of the guide vanes  62 . The inlet flow control flaps  32  duct the fan flow to the second core  14 B. The core flow is pressurized by the second compressor  16 B, mixed with fuel in the second combustor  18 B, and ignited, thereby generating pressurized combustion gases. Some work is extracted from these gases by the second high pressure turbine  20 B which drives the second compressor  16 B via a shaft (not shown). The outlet flow control flaps  36  connect the second core  14 B to the low pressure turbine  22 , which in turn drives the fan  12  via an inner shaft  70 . This mode of operation is termed “single bypass” and has a relatively high fan pressure ratio and low bypass ratio. 
     Means are provided for effecting a changeover between the first and second cores  14 A and  14 B while the engine  10  operates continuously. As illustrated in  FIG. 1 , the first and second cores  14 A and  14 B are provided with first and second core exhaust ducts  72 A and  72 B, respectively, operated by shutoff valves  74 A and  74 B. Means are also provided for supplying intake air to the “inactive” core, for example through the core inlet ducting  30 . In practice, when one core is operating and it is desired to change modes, the other core would be started and brought up to operating speed, then the core flow from the fan would be switched over to that core, and finally the previously “active” core would be shut down until required again. 
       FIG. 3  illustrates schematically the differences in the operation of the engine  10  as compared to a prior art gas turbine engine. The prior art Brayton-cycle engine shows a trough-shaped plot (indicated at “P”) of specific fuel consumption (SFC) versus thrust, with higher SFCs occurring at thrust levels higher or lower than a design point minimum (labeled “M”). For an aircraft which needs to operate over a wide speed range, it is possible that the required cruise thrust could be well below the “design point” of the engine, causing the resultant SFC (intersecting the line labeled “X”) to be much higher than desired. This characteristic makes it difficult to design a single engine having high fuel efficiency at both high and low speeds. 
     Plot B of  FIG. 3  illustrates a theoretical model of the operation of a convertible engine  10  as described above. Maximum power would be the same as for the prior art engine, with the same maximum fan flow and fan overall pressure ratio (i.e. the pressure at the exit of the fan secondary section  46  divided by the pressure upstream of the fan  12 ). At high thrust levels, the engine  10  operates in “single bypass” mode as described above, and the right-hand portion of the SFC curve is essentially the same as the prior art engine. When reduced power is required, for example for long-range cruising flight, the convertible engine  10  may be operated in “double bypass” mode, maintaining a constant total fan flow rate, reducing the fan overall pressure ratio in the bypass duct, maintaining a constant core pressure ratio (i.e. the pressure at the combustor inlet divided by the pressure upstream of the compressor  16 ) and a constant overall pressure ratio (i.e. the pressure at the combustor inlet divided by the pressure upstream of the fan  12 ), and increasing the bypass ratio. In this condition, the pressure ratio across the fan  12  entering the core  14 , referred to as the “hub pressure ratio” or “core supercharge”, remains as the same value as for maximum power. In this operating regime, the convertible engine  10  follows a much lower SFC curve, with a substantial reduction in fuel consumption over the prior art engine at the same reduced thrust level, as shown at the point where curve B intersects line X. 
     It is noted that the pressure ratio “split” of the primary and secondary sections  44  and  46  of the fan  12  may be varied to suit a particular design requirement for reduced power operation while maintaining the same design overall fan pressure ratio, and thus the same thrust and SFC at maximum power. For example, the convertible engine shown at plot “B” has a primary section pressure ratio slightly greater than the secondary section pressure ratio. An alternative engine shown at plot “A” may have a higher primary section pressure ratio, with minimum SFC achieved at a higher thrust level, while another alternative engine “C” may have a lower primary section pressure ratio, achieving minimum SFC at a lower thrust level. 
     The operating principles of the convertible engine  10  may be embodied in a number of different physical configurations. For example,  FIG. 4  illustrates an alternative convertible engine  110  having a fan  112 , a core  114  including a compressor  116 , combustor  118 , and high pressure turbine  120 . The fan  112  is driven by a low pressure turbine  122  disposed downstream of the core  114 . An exhaust duct  124  is disposed downstream of the low pressure turbine  122 , and a convergent-divergent exhaust nozzle  126  is disposed downstream of the exhaust duct  124 . The throat  128  of the exhaust nozzle  126  may have a variable area “A 8 ”, and if desired an afterburner (not shown) may be incorporated upstream of the exhaust nozzle  126 . 
     Core inlet ducting  130  provides a flow path between the fan  112  and the core  114 . An inner bypass duct  138  surrounds the core  114 A and forms a flowpath from the fan  112  to the exhaust duct  124 . An outer bypass duct  140  surrounds the inner bypass duct  138  and forms a second, separate flowpath between the fan  112  and the exhaust duct  124 . The inner and outer bypass ducts  138  and  140  merge downstream of the fan  112  to form a single main bypass duct  141 . A front mixer  143  is disposed at the juncture of the core and bypass duct flow streams to promote efficient mixing of the two streams. If needed, the front mixer  143  may be of the type which can selectively vary its open area, This type of mixer is sometimes referred to as a variable area bypass injector (“VABI”). A mixer  412  (for example a lobed or chute-type mixer) is disposed downstream of the low pressure turbine  22 , at the juncture of the core and bypass duct flow streams, to promote efficient mixing of the two streams. The mixer  42  may also be a VABI, which can be used to control the back pressure on the fan  112 . 
     The fan  112  is of a “convertible” type as described above and includes a “primary section”  144  and “secondary section”  146 , both contained within a fan duct  148 . The exact number of fan stages in each section, the design pressure ratio of the individual sections, and other design characteristics may be varied to suit a particular application. 
     The outer bypass duct  140  is connected to the fan duct  148  between the primary section  144  and the secondary section  146 . A mode control valve  160  may be selectively moved between an open position, in which a portion of the discharge from the primary section  144  can flow into the outer bypass duct  140 , and a closed position, in which all of the discharge from the primary section  144  flows into the secondary section  146 . 
     A fan flow control flap  166  is positioned downstream of the secondary section  146  and serves to selectively alter the proportion of fan discharge flow between the core  114  and the inner bypass duct  138 . In conjunction with a stationary splitter structure  168 , the fan flow control flap  166  effectively divides the discharge flow of the secondary section  146  into radially inner, center, and outer streams. in a “high flow” position, the inner and center streams are passed to the core  114 , while the outer stream passes to the inner bypass duct  138 . In a “low flow” position, the inner stream passes to the core  114  while the center and outer streams pass to the inner bypass duct  138 . The portion of the total fan flow entering one or both of the bypass ducts  138  or  140  is referred to as the “bypass flow”, while the remainder is referred to as the “core flow”. 
     The compressor  116  is a positive-displacement pump of a known type, such as a so-called “worm” compressor that includes an inner and outer body compression system. A key aspect of this type of compressor  116 , as compared to a prior-art axial-flow turbomachinery compressor, is that it is able to maintain a substantially constant pressure ratio while accommodating varying flow rates. The two-piece compressor  116  is driven by the high pressure turbine  120  and a gearset  172  which maintains the required inner and outer body speed relationship. The high pressure turbine  120  is provided with variable area HPT nozzles (VATN)  174  which may have their open flow area selectively changed by using an actuator  176  of a known type, in order to effectuate a change in the operating speed of the high pressure turbine  120  and compressor  116 . The low pressure turbine  122  may also incorporate variable area LPT vanes  178 . 
     The convertible engine  110  operates in one of two modes, in a manner somewhat similar to that of the convertible engine  10  described above. In particular, the operation of the convertible fan  112  is identical to that of the fan  12 . In “double bypass” mode, intended for partial-power operation, the core flow from the inner stream is passed to the core  114 , while the center and outer streams pass to the inner bypass duct  138 . The mode control valve  160  is open and a portion of the flow from the primary section  144  of the fan  112  flows through the outer bypass duct  140 . The flow split between the inner and outer bypass ducts  138  and  140  can be modulated by varying the position of the secondary guide vanes  162 . The inlet flow control flaps  132  duct the core flow to the core  114 . The variable area HPT nozzles  174  are closed as needed as the speed and flow of the compressor  116  are reduced, and the mixer  142  and/or variable area LPT vanes  178  are adjusted as necessary. The unique worm compression system maintains constant core pressure ratio as flow is reduced, effectively operating as a continuous family of individual compressor sizes. 
     In “single bypass” mode, intended for maximum-power operation, the flow from the inner and center streams are passed to the core  114 , while the outer stream passes to the inner bypass duct  138 . The mode control valve  160  is closed. The flow split between the inner bypass duct  138  and the core flow can be modulated by varying the position of the guide vanes  162 . The variable area HPT nozzles  174  are fully open so that the compressor  116  operates at full speed. The mixer  142  and variable are LPT vanes  178  are fully open. 
       FIG. 5  illustrates another alternative convertible engine  210  similar in construction to the convertible engine  110  and having a fan  212 , a core  214  including a positive-displacement compressor  216  (such as a worm compressor), combustor  218 , and high pressure turbine  220 . The fan  212  is driven by a low pressure turbine  222  disposed downstream of the core  214 . An exhaust duct  224  is disposed downstream of the low pressure turbine  222 , and a convergent-divergent exhaust nozzle  226  is disposed downstream of the exhaust duct  224 . The throat  228  of the exhaust nozzle  126  may have a variable area “A 8 ”, and if desired an afterburner (not shown) may be incorporated upstream of the exhaust nozzle  226 . 
     The primary difference between the convertible engine  210  and the convertible engine  110  is that the high-pressure turbine  220  is a positive-displacement device of a known type, such as a so-called “worm” turbine. A key aspect of this type of turbine  220 , as compared to a prior-art axial-flow turbine, is that it is able to maintain a substantially constant pressure drop while accommodating varying flow rates. the high pressure turbine  220  is interconnected to the low pressure turbine  222  by a gearset  272  which provides control of the speed ratio between the two. The low pressure turbine  222  is provided with variable area LPT nozzles (VATN)  278  which may have their open flow area selectively changed by using an actuator of a known type (not shown), in order to effectuate a change in the operating speed of the engine compressor  216 . An advantage of this configuration is that no variable-area hardware is required in the “hot section” of the engine  210 , and the gearset  272  will transmit lower power levels than the gearset  172 . 
       FIG. 6  illustrates another alternative convertible engine  310  similar in construction to the convertible engine  210  and having a fan  312 , a core  314 , a low pressure turbine  322 , exhaust duct  324 , exhaust nozzle  326 , inner bypass duct  338 , and outer bypass duct  340 . The engine  310  includes a supplementary fan, referred to as a “FLADE” stage  380  in the form of a ring of airfoils extending radially outwardly from an annular shroud  382  and driven by the fan  312 . The FLADE stage  380  is positioned in a fan outer duct  384  which surrounds the outer bypass duct  330 . The FLADE stage  380  provides an additional flow stream at a different flow and pressure ratio that than of the fan  312 . Other fan stage counts with possibly FLADE stages on more than one fan blade could also be used, depending on the final selection of fan and FLADE pressure ratios. The FLADE stage flow is sized to provide sufficient bleed air pressure and flow for a selected aircraft bleed-air powered system of a known type (not shown). A row of variable-angle FLADE inlet guide vanes  386 , operated by an actuator  388 , are moveable between open and closed positions to vary the flow through the FLADE stage  380 . A FLADE nozzle  390  is disposed around the exhaust nozzle  326  and is connected to the fan outer duct  384 . the FLADE stage discharge exits the FLADE nozzle  390 . This flow could be used for cooling the exhaust nozzle  326 , for in-flight performance enhancements, and/or to tailor engine inlet flow to minimize inlet spillage and throat bleed drag losses. 
     The foregoing has described a convertible gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.