Abstract:
One embodiment of the present invention is a unique gas turbine engine with a bleed air powered auxiliary engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for bleed air powered auxiliary engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims the benefit of U.S. Provisional Patent Application  61 / 291 , 200 , filed Dec. 30, 2009, and is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to an auxiliary engine powered by bleed air, e.g., from a gas turbine engine. 
       BACKGROUND 
       [0003]    Gas turbine engine bleed air powered systems remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
       SUMMARY 
       [0004]    One embodiment of the present invention is a unique gas turbine engine with a bleed air powered auxiliary engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for bleed air powered auxiliary engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically depicts a non-limiting example of a gas turbine engine having a bleed air power powered auxiliary engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  schematically depicts a non-limiting example of the auxiliary engine of  FIG. 1 . 
       
    
    
     DETAILED DESCRIPTION 
       [0008]    For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
         [0009]    Referring now to the drawings, and in particular  FIG. 1 , a non-limiting example of a gas turbine engine system  10  in accordance with an embodiment of the present invention is schematically depicted. Gas turbine engine system  10  includes a gas turbine engine  12 , which is an aircraft propulsion power plant. In one form, engine  12  is an axial flow turbofan engine. In other embodiments, engine  12  may be, for example, a turbojet engine, a turboprop engine, and/or a turboshaft engine having axial, centrifugal and/or axi-centrifugal flow compressors and/or turbines. In addition to aero gas turbine engines, embodiments of the present invention are applicable to marine gas turbine engines and land-based gas turbine engines. 
         [0010]    In the illustrated embodiment, gas turbine engine system  10  also includes an auxiliary engine  14  that operates using bleed air extracted from gas turbine engine  12 . In one form, gas turbine engine  12  is a two-spool engine. In other embodiments, engine  12  may have a greater or lesser number of spools, e.g., such as a single-spool engine or a three-spool engine. In one form, gas turbine engine  12  includes a fan  16 , a compressor  18  with outlet guide vane (OGV)  20 , a pressurized bleed air source  22 , a diffuser  24 , a combustor  26 , a high pressure (HP) turbine  28 , a low pressure (LP) turbine  30 , an exhaust nozzle  32  and a bypass duct  34 . Diffuser  24  and combustor  26  are fluidly disposed between OGV  20  of compressor  18  and HP turbine  28 . LP turbine  30  is drivingly coupled to fan  16  via an LP shaft  36 . HP turbine  28  is drivingly coupled to compressor  18  via an HP shaft  38 . Compressor  18 , HP shaft  38  and HP turbine  28  form, in part, an HP spool. Fan  16 , LP shaft  36  and LP turbine  30  form, in part, an LP spool. 
         [0011]    Compressor  18  includes a plurality of blades and vanes  40  for compressing air. During the operation of gas turbine engine  12 , air is drawn into the inlet of fan  16  and pressurized by fan  16 . Some of the air pressurized by fan  16  is directed into compressor  18  and the balance is directed into bypass duct  34 . Bypass duct  34  directs the pressurized air to exhaust nozzle  32 , which provides a component of the thrust output by gas turbine engine  12 . Compressor  18  receives some of the pressurized air from fan  16 , which is compressed by blades and vanes  40 . 
         [0012]    The pressurized air discharged from compressor  18  is directed downstream by OGV  20  to diffuser  24 , which diffuses the airflow, reducing its velocity and increasing its static pressure. The diffused airflow is directed into combustor  26 . Fuel is mixed with the pressurized air in combustor  26 , which is then combusted in a combustion liner (not shown). The hot gases exiting combustor  26  are directed into HP turbine  28 , which extracts power from the hot gases in the form of mechanical shaft power to drive compressor  18  via HP shaft  38 . The hot gases exiting HP turbine  28  are directed into LP turbine  30 , which extracts power from the hot gases in the form of mechanical shaft power to drive fan  16  via LP shaft  36 . The hot gases exiting LP turbine  30  are directed into nozzle  32 , and provide a component of the thrust output by gas turbine engine  12 . 
         [0013]    In one form, pressurized bleed air source  22  is a compressor bleed. Compressor bleed  22  is in fluid communication with compressor  18 , and is operative to bleed pressurized air from compressor  18 . In one form, compressor bleed  22  bleeds interstage air from compressor  18 , e.g., from one or more stages of blades and vanes  40 . In another form, compressor bleed  22  bleeds air discharged from compressor  18 , e.g., in addition to or in place of interstage air. In other embodiments, pressurized bleed air source  22  may be any source of pressurized air, for example and without limitation, motor and/or engine driven pumps and/or compressors, and/or other pressurized air sources, such as compressed air storage tanks and/or other compressed air systems/facilities. Auxiliary engine  14  is a turbine engine system in fluid communication with bleed air source  22   22  via ducting  42 . Ducting  42  supplies the pressurized air bled from compressor  18  to auxiliary engine  14 . In other embodiments, it is contemplated that auxiliary engine  14  may be coupled directly to bleed air source  22  without intervening ducting  42 . 
         [0014]    Referring now to  FIG. 2 , auxiliary engine  14  in a non-limiting exemplary elemental form includes a combustor  44  and a turbine  46 . Other embodiments may include additional components. Combustor  44  is in fluid communication with ducting  42 . Turbine  46  is in fluid communication with combustor  44 . In one form, turbine  46  is a two-stage turbine, although turbines having a greater or lesser number of stages may alternatively be employed. A valve  48 , such as a fast-acting air valve, controls the flow of bleed air into combustor  44 . In one form, valve  48  is configured to operate between a fully closed position and a fully open position in order to modulate the flow of bleed air into combustor  44  to a desired level, e.g., in response to a control input based on a desired output of auxiliary engine  14 . In other embodiments, valve  48  may be an on/off valve, or any valve operable between a maximum flow condition and a minimum flow condition. In one form, combustor  44  includes a plurality of fuel injectors (not shown), which add fuel to the pressurized air received from ducting  42 , which is ignited in combustor  44 , e.g., in a combustion liner (not shown). In other embodiments, only a single fuel injector may be employed. The resultant hot gas stream is expanded in turbine  46 , which extracts power from the hot gases in the form of mechanical shaft power. 
         [0015]    In one form, auxiliary engine  14  also includes a turbine  50 , e.g., downstream of turbine  46 , a gearbox  52  and a reduction gearbox  54 . In one form, turbine  50  is on a different spool than turbine  46 , i.e., supported by bearings that allow rotation independent of turbine  46 . In other embodiments, auxiliary engine  14  may not include another turbine, such as turbine  50 , and/or may not include one or both of gearbox  52  and a reduction gearbox  54 . Turbine  50  extracts additional power from the hot gas stream. Turbine  46  is coupled to gearbox  52  via a shaft  56 . In one form, turbine  50  is a two-stage turbine, although turbines having a greater or lesser number of stages may be employed in other embodiments. Turbine  50  is coupled to gearbox  52  via a shaft  58 . In one form, gearbox  52  maintains a constant speed ratio between turbine  46  and turbine  50 . In other embodiments, gearbox  52  may control a speed ratio between turbine  46  and turbine  50  that may be constant or may be variable. Gearbox  52  is coupled to reduction gearbox  54  via a shaft  60 , and provides the combined power output from turbine  46  and turbine  50  to reduction gearbox  54 . In one form, a generator  62  is coupled to the output of reduction gearbox  54 . 
         [0016]    Auxiliary engine  14  produces power from bleed air received from aircraft or ground based air producing machinery, such as gas turbine engines, motor and/or engine driven pumps and/or compressors, and/or other pressurized air sources, such as compressed air storage tanks and/or other compressed air systems/facilities. Compressed air introduced into and mixed with fuel combusted in combustor  44  provides energy to power one or more turbines, e.g., turbine  46  and turbine  50 , which may be used to operate machinery, such as generator  62 . Bleed air is traditionally available in aircraft and some ground and seaborne applications. In other embodiments, other machines and/or devices may be powered by auxiliary engine  14  in addition to or in place of generator  62 . 
         [0017]    In various forms, auxiliary engine  14  may include an air duct inlet that is connected to the pressurized air source and sized appropriately for the flow and temperature and a fast-acting air valve to admit and modulate bleed air flow. The duct may be connected to a combustion casing that contains a combustor similar to normal gas turbine engine combustors, one or more fuel nozzles to provide fuel to the combustor to burn the air/fuel mixture, a single or multi-stage, single or multi-spool turbine section, an exhaust, and where desired, may also include a reduction gearbox that provides power output. When a multi-spool turbine is employed, the auxiliary engine may employ a gearbox that maintains a constant speed ratio between the various spools while providing a single or multiple output speed. Alternatively, when multiple spools are employed, power may be absorbed directly, e.g., from each spool, with turbine speed control from the power input device. 
         [0018]    In the case of aircraft, substantial high pressure bleed air is available in cruise flight, since aircraft are designed to fly with one engine out. At least two bleed flow levels are envisioned; e.g., one for idle/warmup, and the other for full power operation. Part power operation could also be achieved through duct air valve modulation. A control system may be employed to control operation from stop to idle, accelerating from idle to part or full power, and from full or part power to idle, e.g., to prevent flame stability or turbine integrity issues. It is also envisioned that an auxiliary engine such as auxiliary engine  14  may operate at constant speed from idle to full power to improve system response time, or variable speed if the application requirements demand variable speed input. 
         [0019]    In one aspect the present application provides a novel way to extract large amounts of power from a small package. For example, some models of the Rolls-Royce Model 250 C30 gas turbine engine produce approximately 650 HP, which is produced from the energy remaining from the high pressure core compressor and turbine, by the engine&#39;s low pressure turbine. The compressor of the C30 engine requires in the range of 1200 HP to run, that is, the high pressure turbine produces 1200 HP to power the compressor. Embodiments of the present invention may make it possible to use the power from the high pressure turbine for power output from the auxiliary engine, so that the total power output of the machine is approximately 1850 HP. Embodiments of the present invention may provide a novel way to provide high power density, modular power for applications such as megawatt power generators or pumping where space limitations exist. Power outputs on the order of 10 HP per pound are achievable with this approach, providing instant power on demand. This present application provides an additional power source for high altitude operation of a gas turbine powered application where very low Reynolds Number values make a traditional auxiliary power unit (APU) impractical. The large gas turbine compressor, which is the bleed source for the pressurized air employed by the auxiliary engine, is less subject to the impact of low Reynolds Number operation, and operates with improved stability due to the bleed offtake. 
         [0020]    As one example, a system can be obtained by using a stock Rolls-Royce Model 250 C30 turbine, combustor and combustor casing, fuel nozzle and exhaust assembly mated to a reduction gearbox that mechanically maintains a constant speed ratio between the high pressure and low pressure turbines of the C30 while providing a single output speed, all fed by an external bleed air source. This design features a high pressure turbine and a low pressure turbine. In one form, a bleed source of 5 lbs/sec at 8 atmospheres may be employed, although higher or lower pressures and/or flows may be employed in other embodiments. The air duct/duct air valve may interface with the existing C30 combustor casing. 
         [0021]    As another example, an air duct leading to a combustor that feeds into a single spool turbine and exhaust may be employed, all fed by an external bleed air source. The single spool turbine can either directly drive a generator or other powered device such as a pump, or can drive a gearbox that in turn drives either a generator or other powered device such as a pump. 
         [0022]    Embodiments of the present invention include a gas turbine engine system, including a compressor; a first combustor in fluid communication with the compressor; a second combustor in fluid communication with the compressor in parallel with the first combustor; a first turbine in fluid communication with the first combustor; and a second turbine in fluid communication with the second combustor. 
         [0023]    In a refinement, the gas turbine engine system further includes a bleed system in fluid communication with the compressor, wherein the second combustor is in fluid communication with the compressor via the bleed system. In another refinement, the second combustor is in fluid communication with the compressor in parallel with the first combustor. 
         [0024]    In another refinement, the gas turbine engine system further includes a duct fluidly coupling the second combustor with the compressor. 
         [0025]    In yet another refinement, the gas turbine engine system further includes a reduction gearbox coupled to the second turbine. 
         [0026]    In still another refinement, the gas turbine engine system further includes a third turbine in fluid communication with the second combustor, wherein the third turbine is on a different spool than the second turbine. In a further refinement, the third turbine is downstream of the second turbine. In another refinement, a gearbox is coupled to both the second turbine and the third turbine, wherein the gearbox is structured to maintain a constant speed ratio between the second turbine and the third turbine. In yet another refinement, a reduction gearbox is coupled to at least one of the second turbine and the third turbine. In still another refinement, the reduction gearbox is coupled to both the second turbine and the third turbine. 
         [0027]    In yet still another refinement, the gas turbine engine system further includes a valve structured to control the flow of bleed air into the second combustor. In a further refinement, the valve is a fast-acting air valve. 
         [0028]    Embodiments of the present invention include a turbine engine system, comprising: a pressurized bleed air source; a combustor in fluid communication with the pressurized bleed air source; a valve disposed between the pressurized bleed air source and the combustor, wherein the valve is configured to control a flow of pressurized air from the pressurized bleed air source into the combustor; and a first turbine in fluid communication with the combustor. 
         [0029]    In a refinement, the pressurized bleed air source is a compressor bleed from a compressor of a gas turbine engine. 
         [0030]    In another refinement, the turbine engine system further comprises a second turbine in fluid communication with the first turbine. 
         [0031]    In yet another refinement, the turbine engine system further comprises a gearbox coupled to both the first turbine and the second turbine, wherein the gearbox is structured to control a speed ratio between the first turbine and the second turbine. 
         [0032]    In still another refinement, the first turbine and the second turbine are configured to supply power to a machine. 
         [0033]    In yet still another refinement, the turbine engine system further comprises a reduction gearbox is coupled to both the first turbine and the second turbine. 
         [0034]    In a further refinement, the turbine engine system is configured to supply power to a machine. 
         [0035]    Embodiments of the present invention include a turbine engine system, comprising: means for providing pressurized air; a valve in fluid communication with the means for providing pressurized air, wherein the valve is configured to control a flow of pressurized air from the means for providing pressurized air; a combustor in fluid communication with the valve; and a turbine in fluid communication with the combustor. 
         [0036]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.