Abstract:
A gas turbine engine includes a turbomachinery core operable to generating a first flow of pressurized combustion gases, the core having an exit plane; a fan disposed upstream of the core adapted to extract energy from the core and generate a first flow of pressurized air; a bypass duct surrounding the core which receives a portion of the flow of pressurized air from the fan; a duct burner disposed in the bypass duct, upstream of the exit plane, for receiving the first flow of pressurized air and generating a second flow of pressurized combustion gases; and an exhaust duct disposed downstream of the core and operable to receive and the first and second flows of pressurized combustion gases and to discharge the combined flows downstream.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines and more particularly to a turbofan engine having a duct burner. 
   Some new aircraft designs under study have diverse propulsion needs in terms of sea-level static (SLS) take-off thrust levels, cruise thrust and specific fuel consumption (SFC) levels, engine diameter and length restrictions, and exhaust system shaping (i.e. for “low observables” purposes) that make it difficult to utilize an existing engine or even to define a new non-augmented engine that meets all of these needs. In particular, demanding “hot day”, high altitude, short runway requirements can be difficult to meet with unaugmented or “dry” versions of existing engines that also meet applicable fan diameter limitations. Augmented (afterburning) versions of these existing engines are typically too long and are not as amenable to special exhaust system shaping as non-afterburning versions. Non-augmented, higher fan pressure ratio engines can be defined that supply the needed thrust within the diameter and length constraints but they will have higher than desired cruise segment SFC levels. 
   What is needed in such situations is an augmentation system that can supply a modest (e.g. about 15-25%) increase in take-off thrust for either existing engines, such as low-bypass military turbofan engines, or can be incorporated into a new engine design so a more optimum cruise cycle can be utilized. This augmentation concept must also be compatible with the length and shaping needs of the exhaust system. 
   Current engine augmentation systems are located in the engine tailpipe downstream of the rear frame. They can easily be sized to provide well in excess of the 15-25% thrust increase mentioned above, but will add appreciable length as well as not being highly adaptable to exhaust system shaping 
   “Duct burners”, i.e. augmentation systems placed in a bypass duct of an engine, have been demonstrated in the prior art. However, these duct burners required two separate exhausts, one for the primary stream and one for the augmented bypass stream, increasing the weight, complexity, and cost of the engine. 
   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 generating a first flow of pressurized combustion gases, the core having an exit plane; a fan disposed upstream of the core adapted to extract energy from the core and provide a first flow of pressurized air; a bypass duct surrounding the core which receives a portion of the flow of pressurized air from the fan; a duct burner disposed in the bypass duct, upstream of the exit plane, for generating a second flow of pressurized combustion gases; and an exhaust duct disposed downstream of the core and operable to receive and the first and second flows of pressurized combustion gases and to discharge the combined flows downstream. 
   According to another aspect of the invention, a gas turbine engine includes: a turbomachinery core operable to generating a first flow of pressurized combustion gases, the core having an exit plane; a fan disposed upstream of the core adapted to extract energy from the core and provide a first flow of pressurized air; a bypass duct surrounding the core which receives a portion of the flow of pressurized air from the fan a duct burner disposed in the bypass duct, upstream of the exit plane, for generating a second flow of pressurized combustion gases; an exhaust duct disposed downstream of the core and operable to receive and the first and second flows of pressurized combustion gases and to discharge the combined flows downstream; a first exhaust nozzle disposed downstream of the exhaust duct; a fan outer duct surrounding the bypass duct; a flade stage comprising a supplementary fan disposed in the fan outer duct and driven by the fan for generating a second flow of pressurized air; and a second exhaust nozzle disposed in the exhaust nozzle and adapted to discharge the second flow of pressurized air in a downstream direction. 
   According to another aspect of the invention, a method of operating a gas turbine engine includes: burning a fuel in a turbomachinery core having an exit plane, to produce a first flow of pressurized combustion gases; generating a first flow of pressurized air using a fan disposed upstream of the core; channeling a portion of the first flow of pressurized air to a duct burner disposed upstream of the exit plane; burning a fuel in the duct burner to produce a second flow of pressurized combustion gases; and combining the first and second flows of pressurized combustion gases in to a mixed exhaust flow downstream of the exit plane. 

   
     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 side cross-sectional view of a gas turbine engine constructed according to an aspect of the present invention; 
       FIG. 2  is a schematic side cross-sectional view of a gas turbine engine constructed according to another aspect of the present invention; and 
       FIG. 3  is a schematic side 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 representative gas turbine engine, generally designated  10 . The engine  10  has a longitudinal center line or axis A and an outer stationary annular casing  12  disposed concentrically about and coaxially along the axis A. The engine  10  has a fan  14 , compressor  16 , combustor  18 , high pressure turbine  20 , and low pressure turbine  22  arranged in serial flow relationship. In operation, pressurized air from the compressor  16  is mixed with fuel in the combustor  18  and ignited, thereby generating pressurized combustion gases. Some work is extracted from these gases by the high pressure turbine  20  which drives the compressor  16  via an outer shaft  24 . The combustion gases then flow into the low pressure turbine  22 , which drives the fan  14  via an inner shaft  26 . 
   A portion of the fan discharge flows through the compressor  16 , combustor  18 , and high-pressure turbine  20 , which are collectively referred to as the “core”  28  of the engine  10 . Another portion of the fan discharge flows through an annular bypass duct  30  which surrounds the core  28 . While the illustrated engine  10  has a conventional three-stage fan as is often found in military low-bypass turbofan engines, the principles of the present invention are equally applicable to other engine configurations so long as a bypass flow is present. 
   A duct burner  32  comprising one or more fuel injectors  34  and flameholders  36  of a known configuration is disposed within the bypass duct  30 , at a point upstream of an exit plane “E” of the flow from the core  28 . Fuel feed and ignition fore the duct burner  32  are provided in a known manner, for example using the controls of the engine  10 , similar to the manner in which a prior art afterburner would be controlled. If desired, the duct burner  32 , or components thereof, may be configured to selectively fold or retract out of the bypass duct  30 , so as to minimize flow losses therein when the duct burner  32  is not being used. Cooling air for the duct burner liners and downstream exhaust system needs can be extracted in front of the duct burner  32  to obtain a cooling source at the appropriate temperature and pressure levels. 
   An exhaust duct  38  is disposed downstream of the core  28 , and receives the mixed air flow from both the core  28  and the bypass duct  30 . A mixer  40  (for example a lobed or chute-type mixer) is disposed at the juncture of the core  28  and bypass duct  30  flow streams to promote efficient mixing of the two streams. If needed, the mixer  40  may be of the type which can selectively vary its open area, so as to control the back pressure on the fan  14 . This type of mixer is sometimes referred to as a variable area bypass injector (“VABI”). 
   A nozzle  42  having an inlet  44 , a throat  46 , and an exit  48  is disposed downstream of the exhaust duct  38 . The throat area, denoted “A 8 ” in accordance with conventional practice, may be variable through the use of moveable components in the nozzle  42 , in order to accommodate changes in the operating cycle flow when the duct burner  32  is cycled on and off. It may also be possible, depending on the selected duct temperature, to define a fixed A 8  that would provide a useful level of maximum power augmented performance and an acceptable level of dry thrust and cruise SFC. In the illustrated example, the nozzle  42  is a so-called “2-D” design incorporating a serpentine flow path as “low observable” feature to reduce or prevent detection of the hot engine exhaust, and may include a thrust reverser or vectoring feature. However, the present invention may also be used with a conventional axisymmetric nozzle design (not shown). 
     FIG. 2  illustrates an alternative gas turbine engine which includes a duct burner, generally designated  110 . The engine  110  is generally similar in construction to the engine  10  and includes outer stationary annular casing  112 , a fan  114 , compressor  116 , combustor  118 , high pressure turbine  120 , and low pressure turbine  122  arranged in serial flow relationship. The engine  110  also includes a duct burner  132  as described above, disposed within a bypass duct  130 , an exhaust duct  138 , a mixer  140 , and a nozzle  142  having an inlet  144 , a throat  146 , and an exit  148 . The throat area A 8  may be variable as described above. In the illustrated example, the nozzle  142  is a “2-D” design incorporating a serpentine flow path. 
   The engine  110  also includes a supplementary fan, referred to as a “FLADE” stage  150  in the form of a ring of airfoils extending radially outwardly from an annular shroud  152  and driven by the fan  114 . The FLADE stage  150  is positioned in a fan outer duct  154  which surrounds the bypass duct  130 . The FLADE stage  150  provides an additional flow stream at a different flow and pressure ratio that than of the fan  114 . The FLADE stage  150  can be used for optimizing installation losses (i.e. to allow the engine  110  to “swallow” excess airflow from a fixed geometry inlet at high speeds), and to provide additional nozzle cooling. For example, airflow from the FLADE stage  150  may be discharged into an interior space  156  of the nozzle  142  to cool the nozzle surfaces, and then ejected through a slot or FLADE nozzle  158  in the nozzle  142  to provide some supplemental thrust. 
   Optionally, a heat exchanger  160  may be placed inside the interior space  156  of the nozzle  142 . The heat exchanger  160 , for example a liquid-to-air type, may be connected to a heat transfer fluid (e.g. fuel or oil) through lines  162  and  164 . This allows waste heat to be rejected from the airframe (not shown) and also provides some thrust increase by increasing the temperature of the FLADE discharge flow. 
     FIG. 3  illustrates another alternative gas turbine engine  210  which includes a duct burner. The engine  210  is generally similar in construction to the engine  10  and includes outer stationary annular casing  212 , a fan  214 , compressor  216 , combustor  218 , high pressure turbine  220 , and low pressure turbine  222  arranged in serial flow relationship. The engine  210  also includes a duct burner  232  as described above, disposed within a bypass duct  230 , and an exhaust duct  238  which receives the mixed flow from the bypass duct  230  and the engine core  228 . 
   The mixed flow exhaust is discharged through an outer nozzle  242 . While various types of nozzles may be used, in this example the outer nozzle  242  is a plug type and includes a centerbody  243 , an inner shroud  245 , and an outer shroud  247 . The centerbody  243  is centered along the longitudinal axis A of the engine  210  and extends in an aft direction. The centerbody  243  includes, sequentially, a small-diameter tapered forward section, a throat section of increased diameter, and an aft section which tapers in diameter to form an aft-facing conical shape. The inner and outer shrouds  245  and  247  may be independently translated to achieve independently variable throat (A 8 ) and exit plane (conventionally denoted “A 9 ”) areas during different operating conditions. The construction and operation of such a nozzle is described in published US Patent Application No. US2006/0016171A1, which is assigned to the assignee of the present invention. 
   The engine  210  also includes a supplementary fan, referred to as a “FLADE” stage  250  in the form of a ring of airfoils extending radially outwardly from an annular shroud  251  and driven by the fan  214 . The FLADE stage  250  is positioned in a fan outer duct  254  which surrounds the bypass duct  230 . The FLADE stage  250  provides an additional flow stream at a different flow and pressure ratio that than of the fan  214 . The FLADE stage  250  can be used for optimizing installation losses (i.e. to allow the engine  210  to “swallow” excess airflow from a fixed geometry inlet at high speeds), to provide additional nozzle cooling, and to provide additional thrust. 
   In the illustrated example, flow from the fan outer duct  254  is discharged thorough one or more radially-extending hollow struts  266  to an inner nozzle  268 . The inner nozzle  266  is a plug type having a shroud  270  (which in this case is formed by a portion of the same structure that defines the centerbody  243  of the outer nozzle  242 ), and a plug  272 . A nozzle throat  274  is defined between the narrowest portion of the shroud  270  and the widest portion of the plug  272 . The plug  272  may be translated fore or aft using an actuator  276  to vary the throat area A 8  in a known manner. The discharge of the FLADE stage flow through the fan outer duct  254 , struts  266 , and inner nozzle  268  in this manner provides cooling to structure of the entire exhaust system. This in turn allows higher exhaust temperatures to be used in the outer nozzle. Analytical studies and scale model tests have indicated that at a given level of jet velocity, jet noise can be reduced if the temperature of the jet exhaust is increased while pressure levels are decreased. For the engine  210  shown, fan pressure levels and duct burner temperature levels would be adjusted to arrive at the optimum combination of parameters to minimize jet noise at the prescribed level of mixed stream jet velocity needed to produce adequate take-off thrust. The duct burner  232  could also be used to minimize the engine size and time needed for transonic acceleration and climb to an optimum supersonic cruise flight condition. 
   The duct burner systems described above can supply modest thrust increases for either existing engines, like military low-bypass military turbofan engines, or can be incorporated into a new engine design so a more optimum cruise cycle can be utilized. For example, SLS thrust increases of about 15%-25% can be achieved with duct temperatures in the 1000° to 1300° C. to (2000° to 2400° F.) range. 
   Engine designs that are fan flow and/or length limited due to airframe restrictions could benefit from this duct burning thrust augmentation concept by allowing a more cruise efficient, lower fan pressure ratio, higher bypass ratio engine to be used in place of the same thrust and fan flow size higher fan pressure ratio dry engine. 
   For example, a 3.5 fan pressure ratio, moderate bypass ratio engine with a duct burner that provides about a 20% SLS thrust increase would have the same fan flow and take-off thrust as a 4.5 fan pressure ratio, lower bypass dry engine, while also providing about a 10% lower subsonic cruise SFC. 
   The foregoing has described a duct burner and a gas turbine engine incorporating a duct burner. 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.