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

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims benefit of U.S. Provisional Patent Application No. 61/427,724, filed Dec. 28, 2010, entitled AIRCRAFT AND GAS TURBINE ENGINE, which is incorporated herein by reference. 
    
    
     GOVERNMENT RIGHTS 
     The present application was made with the United States government support under Contract No. F33615-03-D-2357 0004, awarded by the U.S. Air Force. The United States government may have certain rights in the present application. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to aircraft and gas turbine engines, and more particularly, to electrical power generation and windmill starting in gas turbine engine powered aircraft. 
     BACKGROUND 
     Gas turbine engines and aircraft powered by gas turbine engines that generate power for the aircraft and are required to perform windmill starts 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 
     One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique aircraft. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine powered aircraft. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
         FIG. 1  illustrates some aspects of a non-limiting example of an aircraft in accordance with an embodiment of the present invention. 
         FIG. 2  schematically illustrates some aspects of non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 3  schematically illustrates some aspects of a non-limiting example of an electrical rotor machine coupled to an aircraft electrical system in accordance with an embodiment of the present invention. 
         FIG. 4  schematically illustrates some aspects of a non-limiting example of an accessory gearbox of the gas turbine engine embodiment of  FIG. 2  having a starter motor coupled to the electrical rotor machine of  FIG. 3  for starting the gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
     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. 
     Referring to  FIG. 1 , there are illustrated some aspects of a non-limiting example of an aircraft  10  in accordance with an embodiment of the present invention. In one form, aircraft  10  includes flight structures in the form of a fuselage  12 , wings  14  and an empennage  16 . Aircraft  10  also includes a gas turbine propulsion system  18 . In one form, aircraft  10  is an unmanned single engine air vehicle. In other embodiments, aircraft  10  may be any fixed-wing aircraft, including turbofan aircraft, turbojet aircraft and turboprop aircraft. In still other embodiments, aircraft  10  may be a rotary-wing aircraft, a combination rotary-wing/fixed-wing aircraft, a missile, or any air vehicle. In various embodiments, aircraft  10  may have a single propulsion engine or a plurality of propulsion engines. In addition, in various embodiments, aircraft  10  may employ any number of wings  14 . Empennage  16  may employ a single or multiple flight control surfaces. 
     Referring to  FIG. 2 , there are illustrated some aspects of a non-limiting example of a propulsion system  18  in accordance with an embodiment of the present invention. Propulsion system  18  includes a gas turbine engine  20 . Engine  20  is a primary propulsion engine that provides thrust for flight operations of aircraft  10 . In one form, engine  20  is a two spool engine having a high pressure (HP) spool  24  and a low pressure (LP) spool  26 . In other embodiments, engine  20  may include three or more spools, e.g., may include an intermediate pressure (IP) spool and/or other spools. In one form, engine  20  is a turbofan engine, wherein LP spool  26  is operative to drive a propulsor  28  in the form of a turbofan (fan) system, which may be referred to as a turbofan, a fan or a fan system. In other embodiments, engine  20  may be a turboprop engine, wherein LP spool  26  powers a propulsor  28  in the form of a propeller system (not shown), e.g., via a reduction gearbox (not shown). In yet other embodiments, LP spool  26  powers a propulsor  28  in the form of a propfan. In still other embodiments, propulsor  28  may take other forms, such as one or more helicopter rotors or tilt-wing aircraft rotors. In one form, a single propulsion system  18  is coupled to fuselage  12  of aircraft  10 . In other embodiments, one or more propulsion system  18  may be coupled to each wing  14 . In still other embodiments, one or more propulsion systems  18  may be coupled to the fuselage and/or the empennage in addition to or in place of wing-mounted propulsion systems  18 . 
     In one form, engine  20  includes, in addition to fan system  28 , a bypass duct  30 , a compressor system  32 , a diffuser  34 , a combustion system  36 , a high pressure (HP) turbine system  38 , a low pressure (LP) turbine system  40 , a nozzle  42 A, and a nozzle  42 B. In other embodiments, there may be, for example, an intermediate pressure spool having an intermediate pressure turbine system. Engine  20  also includes an electrical rotor machine  44  and a tail cone  46 . Electrical rotor machine  44  is coupled to LP spool  26 . In one form, electrical rotor machine  44  is integrated within engine tail cone  46 . In other embodiments, electrical rotor machine  44  may be disposed in other locations, for example and without limitation, upstream or downstream of propulsor  28 , or otherwise upstream of combustion system  36 , e.g., in order to provide a cooler environment for electrical rotor machine  44 . In one form, electrical rotor machine  44  is configured to convert mechanical power to electrical power. In other embodiments, electrical rotor machine  44  may also be configured to convert electrical power to mechanical power, e.g., as in a motor/generator. 
     In the depicted embodiment, engine  20  core flow is discharged through nozzle  42 A, and the bypass flow is discharged through nozzle  42 B. In other embodiments, other nozzle arrangements may be employed, e.g., a common nozzle for core and bypass flow; a nozzle for core flow, but no nozzle for bypass flow; or another nozzle arrangement. Bypass duct  30  and compressor system  32  are in fluid communication with fan system  28 . Nozzle  42 B is in fluid communication with bypass duct  30 . Diffuser  34  is in fluid communication with compressor system  32 . Combustion system  36  is fluidly disposed between compressor system  32  and turbine system  38 . Turbine system  40  is fluidly disposed between turbine system  38  and nozzle  42 A. In one form, combustion system  36  includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustion system  36  may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, a pulse detonation combustion system or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
     Fan system  28  includes a fan rotor system  48  driven by LP spool  26 . In various embodiments, fan rotor system  48  includes one or more rotors (not shown) that are powered by turbine system  40 . Fan system  28  may include one or more vanes (not shown). Bypass duct  30  is operative to transmit a bypass flow generated by fan system  28  around the core of engine  20 . Compressor system  32  includes a compressor rotor system  50 . In various embodiments, compressor rotor system  50  includes one or more rotors (not shown) that are powered by turbine system  38 . Turbine system  38  includes a turbine rotor system  52 . In various embodiments, turbine rotor system  52  includes one or more rotors (not shown) operative to drive compressor rotor system  50 . Turbine rotor system  52  is drivingly coupled to compressor rotor system  50  via a shafting system  54 . Turbine system  40  includes a turbine rotor system  56 . In various embodiments, turbine rotor system  56  includes one or more rotors (not shown) operative to drive fan rotor system  48 . Turbine rotor system  56  is drivingly coupled to fan rotor system  48  via a shafting system  58 . In various embodiments, shafting systems  54  and  58  include a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed in one or both of shafting systems  54  and  58 . Turbine system  40  is operative to discharge the engine  20  core flow to nozzle  42 A. 
     During normal operation of gas turbine engine  20 , air is drawn into the inlet of fan system  28  and pressurized by fan rotor system  48 . Some of the air pressurized by fan rotor system  48  is directed into compressor system  32  as core flow, and some of the pressurized air is directed into bypass duct  30  as bypass flow. Compressor system  32  further pressurizes the portion of the air received therein from fan system  28 , which is then discharged into diffuser  34 . Diffuser  34  reduces the velocity of the pressurized air, and directs the diffused core airflow into combustion system  36 . Fuel is mixed with the pressurized air in combustion system  36 , which is then combusted. The hot gases exiting combustion system  36  are directed into turbine systems  38  and  40 , which extract energy in the form of mechanical shaft power to drive compressor system  32  and fan system  28  via respective shafting systems  54  and  58 . The hot gases exiting turbine system  40  are discharged through nozzle system  42 A, and provide a component of the thrust output by engine  20 . 
     Referring to  FIG. 3 , aircraft  10  includes an electrical system  60  having an electrical bus  62 . Electrical system  60  may include, for example and without limitation, avionics and other aircraft control systems; intelligence data collection systems such as various sensor payloads, synthetic aperture radar (SAR) systems, electro-optical/infrared (EO/IR) imagery; command, communication and control uplink and downlink systems; and weapon systems. Electrical rotor machine  44  is coupled to electrical system  60  via electrical bus  62 . One or more electrical power conditioning devices or other electrical or electronic devices (not shown) may be disposed between electrical rotor machine  44  and electrical bus  62 , e.g., to convert the electrical power output by electrical rotor machine  44  into a form suitable for use on electrical bus  62  and by electrical system  60 . 
     The inventors have determined that during normal flight operations at high altitude, e.g., 60,000 feet and 0.6 Mach number, extracting power from LP spool  26  to supply aircraft  10  with electrical power results in a lower thrust reduction than if the same amount of power was extracted from HP spool  24 . In a particular example, the thrust reduction was reduced by a factor of four (4). In addition, the inventors have determined that the adverse effect on thrust specific fuel consumption (TSFC) is lower. That is, extracting power from LP spool  26  to supply aircraft  10  with electrical power results in a smaller adverse impact on TSFC than if the same amount of power was extracted from HP spool  24 . In a particular example, the adverse impact was reduced by a factor of two (2). Accordingly, in one form, electrical rotor machine  44  is configured to provide electrical power to aircraft  10  during flight operations, including providing power to electrical system  60 . In other embodiments, other power sources may be used in addition to or in place of electrical rotor machine  44  to power aircraft  10  during flight operations. 
     Referring to  FIG. 4 , engine  20  also includes an accessory gearbox  70 . Accessory gearbox  70  is mechanically coupled to HP spool  24  via a shafting system  72 . Accessory gearbox  70  is configured to drive a plurality accessories mounted thereon, including, for example and without limitation, a lube pump  74 , a hydraulic pump  76 , and a fuel pump and metering unit  78 . Also mounted on accessory gearbox  70  is an electrical rotor machine  80 . In one form, electrical rotor machine  80  is configured to convert electrical power into mechanical power. In a particular form, electrical rotor machine  80  is a starter motor. Electrical rotor machine  80  is configured to supply sufficient mechanical power to HP spool  24  via accessory gearbox  70  and shafting system  72  to impart sufficient rotation to HP spool  24  to start engine  20 . In other embodiments, electrical rotor machine  80  may also or alternatively be configured to convert mechanical power into electrical power, and may be, for example and without limitation, a starter/generator, a generator or an alternator. 
     During flight operations of aircraft  10 , events may occur that result in engine  20  shutting down. For example, engine  20  may be commanded to shut down under certain circumstances, or an uncommanded shutdown of engine  20  may occur, e.g., a flameout resulting from adverse ambient and/or engine  20  inlet conditions. In such cases, it is desirable to restart engine  20 . Many aircraft, such as aircraft  10  employ engines that are started via a ground cart that supplies electrical energy or pressurized air for a pneumatic starter. However, once airborne, such facilities may not be available. In order to reduce aircraft weight, many aircraft, such as aircraft  10 , do not retain onboard batteries for effecting an in-flight engine start. Rather, many aircraft rely on conventional windmill starting techniques, wherein during the windmilling event, the aircraft is guided to increase its velocity, resulting in increased air velocity through the engine. The increased air velocity is employed to impart sufficient rotational velocity to the high pressure spool to allow fuel introduction and ignition, thereby performing an in-flight engine start. 
     However, some aircraft, such as aircraft  10 , have configurations that prevent a conventional windmill start, e.g., due to adverse engine inlet conditions, and/or flight at high altitudes, which may be less conducive to rotating the high pressure spool sufficiently for a windmill start. In order to overcome such deficiencies, some embodiments of the present invention provide a nonconventional windmill start for the propulsion engines. For example, electrical rotor machine  44  is configured to extract mechanical power from LP spool  26  during a windmilling event, convert the mechanical power to electrical power, and supply the electrical power to HP spool  24  for use in effecting a windmill start of engine  20 . Because propulsor  28  rotates with LP spool  26 , and because propulsor  28  has a substantially larger diameter than HP spool  24 , LP spool  26  is able to extract a substantial amount of power from the air rushing through propulsor  28 . This power is captured by electrical rotor machine  44  and converted to electrical power. As illustrated in  FIG. 4 , electrical rotor machine  44  is electrically coupled to electrical rotor machine  80 , as indicated by line  82 . Switches, conditioning units and/or control systems (not shown) may be electrically disposed between electrical rotor machine  44  and electrical rotor machine  80 , depending upon the application. The electrical power generated by electrical rotor machine  44  is supplied to electrical rotor machine  80 . Electrical rotor machine  80  is configured and operative to supply mechanical power to rotate HP spool  24 , using the electrical power generated by electrical rotor machine  44 , during a windmill event to achieve sufficient rotational velocity to inject and ignite fuel in combustion system  36  and to start engine  20 . 
     Embodiments of the present invention include a gas turbine engine, comprising: a first spool configured as a high pressure spool; a second spool configured to operate at lower pressures than the high pressure spool; an electrical rotor machine coupled to the second spool, wherein the electrical rotor machine is configured to extract mechanical power from the second spool during a windmilling event, convert the mechanical power to electrical power, and supply the electrical power to the first spool for use in effecting a windmill start of the gas turbine engine. 
     In a refinement, the electrical rotor machine is configured to supply electrical power to an aircraft during normal gas turbine engine operation. 
     In another refinement, the gas turbine engine further comprises a starter motor configured to rotate the first spool for starting the gas turbine engine, wherein the electrical power is supplied from the electrical rotor machine to the starter motor. 
     In yet another refinement, the gas turbine engine further comprises an accessory gearbox, wherein the starter motor is mounted on the accessory gearbox and operative to supply mechanical power to the first spool via the accessory gearbox for effecting the windmill start of the gas turbine engine. 
     In still another refinement, the starter motor is configured as a starter/generator. 
     In yet still another refinement, the gas turbine engine further comprises an engine tail cone, wherein the electrical rotor machine is positioned within the engine tail cone. 
     In a further refinement, the gas turbine engine is configured as a two-spool engine, wherein the second spool is configured as a low pressure spool. 
     In a yet further refinement, the gas turbine engine is a turbofan engine, and the second spool includes a turbofan. 
     In a still further refinement, the gas turbine engine is configured as a three-spool engine having the high pressure spool, an intermediate pressure spool and a low pressure spool, wherein the second spool is configured as the low pressure spool. 
     Embodiments of the present invention include a gas turbine engine, comprising: a fan rotor; a high pressure spool; an electric starter coupled to the high pressure spool and configured to rotate the high pressure spool to start the gas turbine engine; and an electrical rotor machine coupled to the fan rotor and configured to extract mechanical power from the fan rotor during a windmilling event, convert the mechanical power to electrical power, and supply the electrical power to the electric starter for use in effecting a windmill start of the gas turbine engine. 
     In a refinement, the gas turbine engine further comprises an accessory gearbox. 
     In another refinement, the electric starter is mounted on the accessory gearbox. 
     In yet another refinement, the gas turbine engine further comprises an engine tail cone, wherein the electrical rotor machine is integrated into the engine tail cone. 
     In still another refinement, the electrical rotor machine is configured to extract mechanical power from the fan rotor and supply electrical power to an aircraft during normal flight operations. 
     In yet still another refinement, the gas turbine engine is configured as a two spool engine. 
     Embodiments of the present invention include an aircraft, comprising: a flight structure; and a gas turbine engine having a first spool and a propulsor coupled to the first spool; an electrical rotor machine coupled to the first spool and configured to extract mechanical power from the first spool, convert the mechanical power to electrical power, and to supply electrical power to the aircraft during flight operations as the sole source of electrical power for the aircraft. 
     In a refinement, the gas turbine engine has a high pressure spool; and wherein the electrical rotor machine is configured to supply the electrical power to the high pressure spool for use in effecting a windmill start of the gas turbine engine subsequent to an in-flight shutdown. 
     In another refinement, the aircraft further comprises a starter motor configured to rotate the high pressure spool for starting the gas turbine engine, wherein the electrical power is supplied from the electrical rotor machine to the high pressure spool to effect the windmill start of the gas turbine engine. 
     In yet another refinement, the first spool is a low pressure spool. 
     In still another refinement, the gas turbine engine includes an engine tail cone; and wherein the electrical rotor machine is integrated into the engine tail cone. 
     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.