Abstract:
A method facilitates generating thrust from a gas turbine engine using a pulse detonation system. The method includes introducing fuel and air to the engine, mixing fuel and air in a pulse detonation system deflagration chamber positioned radially outward from an engine exhaust centerbody, and detonating the fuel and air mixture within the pulse detonation system to facilitate increasing the temperature and pressure within the engine and to generate engine thrust.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines and more particularly, to a pulse detonation system for a turbofan engine. 
   Variable cycle turbofan ramjet engines may be used to provide aircraft flight speeds between low subsonic Mach numbers to high supersonic Mach numbers of about Mach 6. Known engines, as described in U.S. Pat. No. 5,694,768, include a core engine system and a dual mode augmentor. The dual mode augmentor provides additional heat to exhaust airflow exiting the core engine system to increase engine thrust. The core engine system provides power to drive a fan assembly and typically includes in serial, axial flow relationship, a compressor, a combustor, a high pressure turbine, and a low pressure turbine. The dual mode augmentor is positioned downstream from the core engine and receives air from the core engine and a bypass duct surrounding the core engine. 
   Known engines can operate over a wide range of flight speed operations if several different combustion systems are utilized. During flight speed operations from take-off to approximately Mach 3, the core engine and an engine fan system provide airflow at a pressure and quantity used by the augmentor to produce thrust for the engine. However, augmentor performance may be limited by the constraints of existing engine components. More specifically, at least some known engines include a conventional bluff centerbody that extends aftward from the core engine and enables the engine to achieve pressure ratios necessary for engine operations. 
   BRIEF SUMMARY OF THE INVENTION 
   In one aspect, a method for generating thrust from a gas turbine engine using a pulse detonation system is provided. The method comprises introducing fuel and air to the engine, mixing fuel and air in a pulse detonation system deflagration chamber positioned radially outward from an engine exhaust centerbody, and detonating the fuel and air mixture within the pulse detonation system to facilitate increasing the temperature and pressure within the engine and to generate engine thrust. 
   In another aspect of the invention, a pulse detonation system for a gas turbine engine is provided. The pulse detonation system is configured to create a temperature rise and a pressure rise within the gas turbine engine and to increase gas turbine engine thrust. The pulse detonation system includes at least one deflagration chamber radially outward from an engine exhaust centerbody. 
   In a further aspect, a gas turbine engine is provided. The gas turbine engine includes an inlet portion, an exhaust portion, a centerline axis of symmetry, an exhaust centerbody, and a pulse detonation system. The exhaust portion is positioned co-axially with the inlet portion. The exhaust centerbody is concentrically aligned with the exhaust portion and extends axially along the centerline axis of symmetry into the exhaust portion. The pulse detonation system is positioned between the engine inlet portion and the engine exhaust portion, and is configured to create a temperature rise and a pressure rise within said engine and to increase engine thrust. The pulse detonation system includes at least one deflagration chamber that is radially outward from the engine exhaust centerbody. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a cross-sectional side view of a gas turbine engine including a pulse detonation system in a first mode of engine operation; 
       FIG. 2  is an enlarged partial cross-sectional side view the engine shown in  FIG. 1  in a second mode of engine operation; 
       FIG. 3  is a cross sectional view of an exemplary embodiment of the deflagration chamber taken along lines  3 — 3  shown in  FIG. 1 ; 
       FIG. 4  is a cross sectional view of an alternative embodiment of the deflagration chambers taken along lines  3 — 3  shown in  FIG. 1 ; and 
       FIG. 5  is a detailed view of the exemplary two-stage pulse detonation system shown in area A of FIG.  1 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a cross-sectional side view of a gas turbine turbofan engine  10  including a pulse detonation system  12  in a first mode of engine operation.  FIG. 2  is an enlarged partial cross-sectional side view of engine  10  in a second mode of engine operation. In one embodiment, engine  10  is an F110 engine and is available from General Electric Aircraft Engines, Cincinnati, Ohio. Engine  10  has a generally longitudinally extending axis or centerline  14  extending from an inlet end  16  of engine  10  aftward to an exhaust end  18  of engine  10 . Engine  10  includes a core engine  30  which includes a high pressure compressor  34 , a combustor  36 , a high pressure turbine  38 , and a power turbine or a low pressure turbine  40 , all arranged in a serial, axial flow relationship. Engine  10  also includes a bypass duct  42  surrounding the core engine  30 . In alternative embodiments, engine  10  also includes a core fan assembly. 
   An exhaust nozzle  50  extends aftward from core engine  30  and includes a nozzle portion  52 . Nozzle portion  52  extends between nozzle  50  and core engine  30  and defines a portion of an outer boundary of an engine exhaust flowpath  54 . More specifically, nozzle portion  52  directs combustion gases discharged from core engine  30  and airflow exiting bypass duct  42  downstream through exhaust nozzle  50 . 
   A bluff body or centerbody  56  extends aftward from core engine  30  to an apex  58  formed at an aft end  60  of centerbody  56 . More specifically, centerbody  56  is concentrically aligned with respect to nozzle  50  and extends aftward along engine centerline  14 . Centerbody  56  is contoured and has a variable width  66  measured axially along centerbody  56  such that centerbody  56  defines a convergent-divergent path through nozzle  50 . Accordingly, an outer surface  68  of centerbody  56  defines an inner boundary of engine exhaust flowpath  54 . More specifically, a nozzle throat area  70  is defined between centerbody outer surface  68  and nozzle  50 . 
   Centerbody  56  is axially translatable between a first position  80  and a second position  82  depending upon a mode of engine operation of engine  10 . More specifically, when engine  10  is in a first mode of operation, also known as a dry mode of operation, centerbody  56  is positioned at an aft first position  80 , as illustrated in FIG.  1 . When engine  10  is in a second mode of operation, known as a reheat or augmented mode of operation, centerbody  56  is positioned at a forward second position  82 , as illustrated in FIG.  2 . More specifically, centerbody  56  is moveable a distance  84  between aft first position  80  and upstream second position  82 . 
   Pulse detonation system  12  is disposed downstream from core engine  30  and receives core engine combustion gases discharged from core engine  30  and airflow exiting bypass duct  42 . System  12  is known as a two-stage detonation augmentor, and creates a temperature rise and a pressure rise within engine  10  without the use of turbomachinery included within core engine  30  to generate thrust from engine  10 . Specifically, pulse detonation system  12  includes a hollow deflagration chamber  100  and a hollow detonation chamber  102  that facilitate increasing the performance and operating range of engine  10 . 
   Deflagration chamber  100  is contoured and is positioned radially outwardly from centerbody  56  in flow communication with core engine  30 . Thus, because chamber  100  extends into flowpath  54 , the contour of chamber  100  directs the flow of core engine combustion gases discharged from core engine  30  and airflow exiting bypass duct  42 . Furthermore, because of the contour of chamber  100 , an upstream end  104  of deflagration chamber  100  is positioned a farther distance from centerbody  56  than a downstream end  106  of deflagration chamber  100 . In the exemplary embodiment, deflagration chamber  100  is annular and extends circumferentially around centerbody  56  within engine nozzle  50 . In an alternative embodiment, deflagration chamber  100  is non-annular and engine  10  includes a plurality of deflagration chambers  100  extending axi-symmetrically and circumferentially around centerbody  56  within engine nozzle  50 . Deflagration chamber  100  is coupled in flow communication with a fuel source (not shown) and an air source (not shown) used for atomization. 
   Detonation chamber  102  is positioned at deflagration chamber downstream end  106  in flow communication with deflagration chamber  100 , such that flow exiting deflagration chamber  100  is discharged through detonation chamber  102 . More specifically, deflagration chamber  100  includes a vaneless radial nozzle (not shown) that accelerates and directs flow from chamber  100  into detonation chamber  102 . Detonation chamber  102  is in serial, axial flow relationship with deflagration chamber  100 . Detonation chamber  102  is also in flow communication with a reversed flap  110  positioned downstream from chambers  100  and  102 . 
   Flap  110  is translatable between a first position  112  and a second position  114  depending upon a mode of engine operation. More specifically, flap  110  translates to first position  112  during dry mode of engine operation, and second position  114  during reheat mode of engine operation. Flap  110  is contoured and when in a first position  112 , flap  110  extends radially inwardly from an inner surface  116  of nozzle  50  towards an inner surface  120  of deflagration chamber  100 . More specifically, when flap  110  is in first position  112 , the contour of flap  110  substantially mirrors the contour of centerbody  56 . Accordingly, engine combustion gases discharged from core engine  30  and flowing past deflagration chamber  100  along flowpath  54  are channeled downstream between flap  110  and centerbody  56 . 
   When flap  110  is in first position  112 , flap  110  facilitates preventing airflow from backflowing to contact detonation chamber  102 , and thus, essentially prevents flow communication between detonation chamber  102  and engine flowpath  54 . Alternatively, when flap  110  is in second position  114 , flap  110  is considered “stowed” in close proximity to nozzle inner surface  116 , and thus, detonation chamber  100  is returned to flow communication with flowpath  54  and pulse detonation system  12  receives combustion gases discharged from core engine  30  and airflow exiting bypass duct  42 . 
   During operation, engine  10  is initially operated in a dry mode of operation, and no fuel is supplied to pulse detonation system  12 , or more specifically, no fuel is supplied to deflagration chamber  100 . During the dry mode of engine operation, reversed flap  110  is positioned at first position  112 , and facilitates directing flow passing deflagration chamber  100  downstream along flowpath  54 . Furthermore, during the dry mode of engine operation, centerbody  56  is positioned in aft first position  80 , and combustion gases discharged from core engine  30  and airflow exiting fan bypass duct  42  flow through the convergent-divergent path defined between centerbody  56  and nozzle  50 . During the dry mode of engine operation, axial movements of centerbody  56  provide throat area modulation. 
   In the augmented or reheat mode of engine operation, flap  110  is translated to second position  114 , or the stowed position, and detonation chamber  102  is returned to flow communication with flowpath  54 . Fuel is supplied to deflagration chamber  100  such that chamber  100  is operated in a fuel-rich mode of operation. Flow exiting deflagration chamber  100  enters detonation chamber  102  through the vaneless radial nozzle which operates above a critical pressure ratio, and combustion is initiated within detonation chamber  102 . Because centerbody  56  is translated to second position  82  during the reheat mode of engine operation, the pressure ratio across the vaneless radial nozzle is increased. When this pressure ratio reaches the critical value, detonation occurs within detonation chamber  102 . The resulting detonation shock pattern results in the temporary interruption of flow into chamber  102 , the discharge of detonation products aftwards, and the initiation of a fresh charge of deflagration products through the radial nozzle. The cycle is repeated at a high frequency such that an amount of thrust from engine  10  is increased without impacting operation of core engine  30 . As a result, operation of pulse detonation system  12  creates a pressure and temperature rise within engine  10 , which facilitates increasing an amount of thrust from engine  10 . 
     FIG. 3  is a cross sectional view of an exemplary embodiment of deflagration chamber  100  taken along lines  3 — 3  shown in FIG.  1 . The cross-sectional view represents a view taken from core engine  30  (shown in  FIG. 1 ) towards exhaust nozzle  50 . Centerbody  56  extends aftward from core engine  30  and is substantially concentrically aligned with respect to nozzle  50  along engine centerline  14 . Centerbody  56  is contoured and has a variable width  66  measured axially along centerbody  56 . Deflagration chamber  100  is contoured and is positioned radially outwardly from centerbody  56  in flow communication with core engine  30 . In the exemplary embodiment, deflagration chamber  100  is annular and extends circumferentially around centerbody  56  within engine nozzle  50 . 
     FIG. 4  is a cross sectional view of an alternative embodiment of the deflagration chambers taken along lines  3 — 3  shown in FIG.  1 . The cross-sectional view represents a view taken from core engine  30  (shown in  FIG. 1 ) towards exhaust nozzle  50 . Centerbody  56  extends aftward from core engine  30  and is concentrically aligned with respect to nozzle  50  and extends aftward along engine centerline  14 . Centerbody  56  is contoured and has a variable width  66  measured axially along centerbody  56 . In this embodiment, deflagration chamber  100  is non-annular and engine  10  includes a plurality of deflagration chambers  100  extending axi-symmetrically and circumferentially around centerbody  56  within engine nozzle  50 . Deflagration chamber  100  is coupled in flow communication with the fuel source (not shown) and the air source (not shown) used for atomization. 
     FIG. 5  is a detailed view of the exemplary two-stage pulse detonation system shown in area A of FIG.  1 . Detonation chamber  102  is positioned at deflagration chamber downstream end  106  in flow communication with deflagration chamber  100 , such that flow exiting deflagration chamber  100  is discharged through detonation chamber  102 . More specifically, deflagration chamber  100  includes a vaneless radial nozzle  500  that accelerates and directs flow from chamber  100  into detonation chamber  102 . In the augmented or reheat mode of engine operation, flap  110  is translated to second position  114 , or the stowed position, and detonation chamber  102  is returned to flow communication with flowpath  54 , which includes combustion gases discharged from core engine  30  and airflow exiting bypass duct  42 . Fuel is supplied to deflagration chamber  100  such that chamber  100  is operated in a fuel-rich mode of operation. Flow exiting deflagration chamber  100  enters detonation chamber  102  through vaneless radial nozzle  500 , which based on inlet, outlet, and throat dimensions, and upstream and downstream pressures, operates above the critical pressure ratio, and combustion is initiated within detonation chamber  102 . Because centerbody  56  is translated to second position  82  during the reheat mode of engine operation, the pressure ratio across the vaneless radial nozzle is increased. When this pressure ratio reaches the critical value, detonation occurs within detonation chamber  102 . The resulting detonation shock pattern results in the temporary interruption of flow into chamber  102 , the discharge of detonation products aftwards, and the initiation of a fresh charge of deflagration products through the radial nozzle. The cycle is repeated at a high frequency during operation in the augmented mode. 
   The above-described pulse detonation system includes at least one deflagration chamber in serial, axial-flow with a detonation augmentor which produces engine thrust without the use of turbomachinery. As a result, engines using the pulse detonation system obtain increased thrust over baseline engines operating without the pulse detonation system. As a result, a pulse detonation system is provided which permits an engine to operate with a high efficiency and thus facilitates increasing performance over a wide range of operating flight speeds. 
   While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.