US20120233981A1 - Gas turbine engine with low fan pressure ratio - Google Patents
Gas turbine engine with low fan pressure ratio Download PDFInfo
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- US20120233981A1 US20120233981A1 US13/484,308 US201213484308A US2012233981A1 US 20120233981 A1 US20120233981 A1 US 20120233981A1 US 201213484308 A US201213484308 A US 201213484308A US 2012233981 A1 US2012233981 A1 US 2012233981A1
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- Prior art keywords
- fan
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- nacelle
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D17/00—Regulating or controlling by varying flow
- F01D17/10—Final actuators
- F01D17/12—Final actuators arranged in stator parts
- F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
- F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
- F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
- F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
Definitions
- FIG. 4A is a sectional view of another embodiment of a single FEGV slatted airfoil with a;
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
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- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
A turbofan engine includes a fan variable area nozzle axially movable relative to the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.
Description
- The present disclosure is a continuation of U.S. patent application Ser. No. 13/340,761, filed Dec. 30, 2011, which is a continuation in part of U.S. patent application Ser. No. 11/829213, filed Jul. 27, 2007.
- The present invention relates to a gas turbine engine, and more particularly to a turbofan engine having a variable geometry fan exit guide vane (FEGV) system to change a fan bypass flow path area thereof.
- Conventional gas turbine engines generally include a fan section and a core section with the fan section having a larger diameter than that of the core section. The fan section and the core section are disposed about a longitudinal axis and are enclosed within an engine nacelle assembly. Combustion gases are discharged from the core section through a core exhaust nozzle while an annular fan bypass flow, disposed radially outward of the primary core exhaust path, is discharged along a fan bypass flow path and through an annular fan exhaust nozzle. A majority of thrust is produced by the bypass flow while the remainder is provided from the combustion gases.
- The fan bypass flow path is a compromise suitable for take-off and landing conditions as well as for cruise conditions. A minimum area along the fan bypass flow path determines the maximum mass flow of air. During engine-out conditions, insufficient flow area along the bypass flow path may result in significant flow spillage and associated drag. The fan nacelle diameter is typically sized to minimize drag during these engine-out conditions which results in a fan nacelle diameter that is larger than necessary at normal cruise conditions with less than optimal drag during portions of an aircraft mission.
- A gas turbine engine according to an exemplary aspect of the present disclosure includes a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow, and a fan variable area nozzle axially movable relative the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation, the fan pressure ratio less than about 1.45, the fan bypass airflow defines a bypass ratio greater than about six (6).
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a multiple of fan exit guide vanes in communication with the fan bypass flow path, the multiple of fan exit guide vane rotatable about an axis of rotation to vary the fan bypass flow path.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the multiple of fan exit guide vanes may be simultaneously rotatable. Additionally or alternatively, the multiple of fan exit guide vanes may be mounted within an intermediate engine case structure. Additionally or alternatively, each of the multiple of fan exit guide vanes may include a pivotable portion rotatable about the axis of rotation relative a fixed portion.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the pivotable portion may include a leading edge flap.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a controller operable to control a fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the controller may be operable to reduce the fan nozzle exit area at a cruise flight condition. Additionally or alternatively, the controller may be operable to control the fan nozzle exit area to reduce a fan instability.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, the fan defines a corrected fan tip speed less than about 1150 ft/second.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.3.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.5.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a gear system driven by the core engine to drive the fan, the gear system defines a gear reduction ratio of greater than or equal to 2.5.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the core engine may include a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than about five (5). Additionally or alternatively, the core engine may include a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than five (5).
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan bypass airflow may define a fan pressure ratio less than about 1.45.
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine bypass flow may define a bypass ratio greater than about ten (10). Additionally or alternatively, the bypass flow may define a bypass ratio greater than ten (10).
- In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a multiple of fan exit guide vanes in communication with the fan bypass flow path, the multiple of fan exit guide vanes rotatable about an axis of rotation to vary said fan bypass flow path.
- The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:
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FIG. 1A is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention; -
FIG. 1B is a perspective side partial fragmentary view of a FEGV system which provides a fan variable area nozzle; -
FIG. 2A is a sectional view of a single FEGV airfoil; -
FIG. 2B is a sectional view of the FEGV illustrated inFIG. 2A shown in a first position; -
FIG. 2C is a sectional view of the FEGV illustrated inFIG. 2A shown in a rotated position; -
FIG. 3A is a sectional view of another embodiment of a single FEGV airfoil; -
FIGS. 3B is a sectional view of the FEGV illustrated inFIG. 3A shown in a first position; -
FIG. 3C is a sectional view of the FEGV illustrated inFIG. 3A shown in a rotated position; -
FIG. 4A is a sectional view of another embodiment of a single FEGV slatted airfoil with a; -
FIGS. 4B is a sectional view of the FEGV illustrated inFIG. 4A shown in a first position; and -
FIG. 4C is a sectional view of the FEGV illustrated inFIG. 4A shown in a rotated position. -
FIG. 1 illustrates a general partial fragmentary schematic view of agas turbofan engine 10 suspended from an engine pylon P within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation. - The
turbofan engine 10 includes a core section within acore nacelle 12 that houses alow spool 14 andhigh spool 24. Thelow spool 14 includes alow pressure compressor 16 andlow pressure turbine 18. Thelow spool 14 drives afan section 20 directly or through agear train 22. Thehigh spool 24 includes ahigh pressure compressor 26 andhigh pressure turbine 28. Acombustor 30 is arranged between thehigh pressure compressor 26 andhigh pressure turbine 28. The low andhigh spools - The
engine 10 is a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, theengine 10 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), thegear train 22 is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine 18 has a pressure ratio that is greater than about five (5). Theengine 10 in the disclosed embodiment is a high-bypass geared turbofan aircraft engine in which theengine 10 bypass ratio is greater than ten (10), the turbofan diameter is significantly larger than that of thelow pressure compressor 16, and thelow pressure turbine 18 has a pressure ratio greater than five (5).Low pressure turbine 18 pressure ratio is pressure measured prior to inlet oflow pressure turbine 18 as related to the pressure at the outlet of thelow pressure turbine 18 prior to exhaust nozzle. Thegear train 22 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.5. It should be understood, however, that the above parameters are exemplary of only one geared turbofan engine and that the present invention is likewise applicable to other gas turbine engines including direct drive turbofans. - Airflow enters a
fan nacelle 34, which may at least partially surround thecore nacelle 12. Thefan section 20 communicates airflow into thecore nacelle 12 for compression by thelow pressure compressor 16 and thehigh pressure compressor 26. Core airflow compressed by thelow pressure compressor 16 and thehigh pressure compressor 26 is mixed with the fuel in thecombustor 30 then expanded over thehigh pressure turbine 28 andlow pressure turbine 18. Theturbines respective spools compressors gear train 22, thefan section 20 in response to the expansion. A core engine exhaust E exits thecore nacelle 12 through acore nozzle 43 defined between thecore nacelle 12 and atail cone 32. - A
bypass flow path 40 is defined between thecore nacelle 12 and thefan nacelle 34. Theengine 10 generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering thefan nacelle 34 becomes bypass flow B. The bypass flow B communicates through the generally annularbypass flow path 40 and may be discharged from theengine 10 through a fan variable area nozzle (FVAN) 42 which defines a variable fannozzle exit area 44 between thefan nacelle 34 and thecore nacelle 12 at anaft segment 34S of thefan nacelle 34 downstream of thefan section 20. - Referring to
FIG. 1B , thecore nacelle 12 is generally supported upon a coreengine case structure 46. Afan case structure 48 is defined about the coreengine case structure 46 to support thefan nacelle 34. The coreengine case structure 46 is secured to thefan case 48 through a multiple of circumferentially spaced radially extending fan exit guide vanes (FEGV) 50. Thefan case structure 48, the coreengine case structure 46, and the multiple of circumferentially spaced radially extending fanexit guide vanes 50 which extend therebetween is typically a complete unit often referred to as an intermediate case. It should be understood that the fanexit guide vanes 50 may be of various forms. The intermediate case structure in the disclosed embodiment includes a variable geometry fan exit guide vane (FEGV)system 36. - Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 20 of theengine 10 is nominally designed for a particular flight condition——typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without the fan exit guide vane (FEGV)system 36. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. - As the
fan section 20 is efficiently designed at a particular fixed stagger angle for an efficient cruise condition, theFEGV system 36 and/or theFVAN 42 is operated to adjust fan bypass air flow such that the angle of attack or incidence of the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff. TheFEGV system 36 and/or theFVAN 42 may be adjusted to selectively adjust the pressure ratio of the bypass flow B in response to a controller C. For example, increased mass flow during windmill or engine-out, and spoiling thrust at landing. Furthermore, theFEGV system 36 will facilitate and in some instances replace theFVAN 42, such as, for example, variable flow area is utilized to manage and optimize the fan operating lines which provides operability margin and allows the fan to be operated near peak efficiency which enables a low fan pressure-ratio and low fan tip speed design; and the variable area reduces noise by improving fan blade aerodynamics by varying blade incidence. TheFEGV system 36 thereby provides optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels. - Referring to
FIG. 2A , each fanexit guide vane 50 includes arespective airfoil portion 52 defined by an outerairfoil wall surface 54 between theleading edge 56 and a trailingedge 58. Theouter airfoil wall 54 typically has a generally concave shaped portion forming a pressure side and a generally convex shaped portion forming a suction side. It should be understood thatrespective airfoil portion 52 defined by the outerairfoil wall surface 54 may be generally equivalent or separately tailored to optimize flow characteristics. - Each fan
exit guide vane 50 is mounted about a vane longitudinal axis ofrotation 60. The vane axis ofrotation 60 is typically transverse to the engine axis A, or at an angle to engine axis A. It should be understood that various support struts 61 or other such members may be located through theairfoil portion 52 to provide fixed support structure between the coreengine case structure 46 and thefan case structure 48. The axis ofrotation 60 may be located about the geometric center of gravity (CG) of the airfoil cross section. An actuator system 62 (illustrated schematically;FIG. 1A ), for example only, a unison ring operates to rotate each fanexit guide vane 50 to selectively vary the fan nozzle throat area (FIG. 2B ). The unison ring may be located, for example, in the intermediate case structure such as within either or both of the coreengine case structure 46 or the fan case 48 (FIG. 1A ). - In operation, the
FEGV system 36 communicates with the controller C to rotate the fanexit guide vanes 50 and effectively vary the fannozzle exit area 44. Other control systems including an engine controller or an aircraft flight control system may also be usable with the present invention. Rotation of the fanexit guide vanes 50 between a nominal position and a rotated position selectively changes the fanbypass flow path 40. That is, both the throat area (FIG. 2B ) and the projected area (FIG. 2C ) are varied through adjustment of the fan exit guide vanes 50. By adjusting the fan exit guide vanes 50 (FIG. 2C ), bypass flow B is increased for particular flight conditions such as during an engine-out condition. Since less bypass flow will spill around the outside of thefan nacelle 34, the maximum diameter of the fan nacelle required to avoid flow separation may be decreased. This will thereby decrease fan nacelle drag during normal cruise conditions and reduce weight of the nacelle assembly. Conversely, by closing theFEGV system 36 to decrease flow area relative to a given bypass flow, engine thrust is significantly spoiled to thereby minimize or eliminate thrust reverser requirements and further decrease weight and packaging requirements. It should be understood that other arrangements as well as essentially infinite intermediate positions are likewise usable with the present invention. - By adjusting the
FEGV system 36 in which all the fanexit guide vanes 50 are moved simultaneously, engine thrust and fuel economy are maximized during each flight regime. By separately adjusting only particular fanexit guide vanes 50 to provide an asymmetrical fanbypass flow path 40, engine bypass flow may be selectively vectored to provide, for example only, trim balance, thrust controlled maneuvering, enhanced ground operations and short field performance. - Referring to
FIG. 3A , another embodiment of theFEGV system 36′ includes a multiple of fanexit guide vane 50′ which each includes a fixedairfoil portion 66F and pivotingairfoil portion 66P which pivots relative to the fixedairfoil portion 66F. The pivotingairfoil portion 66P may include a leading edge flap which is actuatable by anactuator system 62′ as described above to vary both the throat area (FIG. 3B ) and the projected area (FIG. 3C ). - Referring to
FIG. 4A , another embodiment of theFEGV system 36″ includes a multiple of slotted fanexit guide vane 50″ which each includes a fixed airfoil portion 68F and pivoting and slidingairfoil portion 68P which pivots and slides relative to the fixed airfoil portion 68F to create aslot 70 vary both the throat area (FIG. 4B ) and the projected area (FIG. 4C ) as generally described above. This slatted vane method not only increases the flow area but also provides the additional benefit that when there is a negative incidence on the fanexit guide vane 50″ allows air flow from the high-pressure, convex side of the fanexit guide vane 50″ to the lower-pressure, concave side of the fanexit guide vane 50″ which delays flow separation. - The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.
Claims (18)
1. A gas turbine engine comprising:
a core nacelle defined about an engine centerline axis;
a fan nacelle mounted at least partially around said core nacelle to define a fan bypass airflow path for a fan bypass airflow having a bypass ratio greater than about six (6); and
a fan variable area nozzle axially movable relative said fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation, the fan pressure ratio less than about 1.45.
2. The engine as recited in claim 1 , further comprising a multiple of fan exit guide vanes in communication with said fan bypass flow path, said multiple of fan exit guide vanes rotatable about an axis of rotation.
3. The engine as recited in claim 2 , wherein said multiple of fan exit guide vanes are simultaneously rotatable.
4. The engine as recited in claim 2 , wherein said multiple of fan exit guide vanes are mounted within an intermediate engine case structure.
5. The engine as recited in claim 2 , wherein each of said multiple of fan exit guide vanes include a pivotable portion rotatable about said axis of rotation relative to a fixed portion.
6. The engine as recited in claim 5 , wherein said pivotable portion includes a leading edge flap.
7. The engine as recited in claim 1 , further comprising a controller operable to control a fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.
8. The engine as recited in claim 7 , wherein said controller is operable to reduce said fan nozzle exit area at a cruise flight condition.
9. The engine as recited in claim 7 , wherein said controller is operable to control said fan nozzle exit area to reduce a fan instability.
10. The assembly as recited in claim 1 , further comprising a gear system driven by a core engine within the core nacelle to drive a fan within said fan nacelle, said fan defines a corrected fan tip speed less than about 1150 ft/second.
11. The engine as recited in claim 1 , further comprising a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, said gear system defines a gear reduction ratio of greater than or equal to about 2.3.
12. The engine as recited in claim 1 , further comprising a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, said gear system defines a gear reduction ratio of greater than or equal to about 2.5.
13. The engine as recited in claim 1 , further comprising a gear system driven by said core engine to drive said fan, said gear system defines a gear reduction ratio of greater than or equal to 2.5.
14. The engine as recited in claim 1 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than about five (5).
15. The engine as recited in claim 1 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than five (5).
16. The engine as recited in claim 1 , wherein said fan bypass airflow defines a bypass ratio greater than about ten (10).
17. The engine as recited in claim 1 , wherein said fan bypass airflow defines a bypass ratio greater than ten (10).
18. The engine as recited in claim 17 , further comprising a multiple of fan exit guide vanes in communication with said fan bypass flow path, said multiple of fan exit guide vanes rotatable about an axis of rotation.
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US13/484,308 US20120233981A1 (en) | 2007-07-27 | 2012-05-31 | Gas turbine engine with low fan pressure ratio |
US14/602,625 US20150132106A1 (en) | 2007-07-27 | 2015-01-22 | Gas turbine engine with low fan pressure ratio |
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US11/829,213 US8347633B2 (en) | 2007-07-27 | 2007-07-27 | Gas turbine engine with variable geometry fan exit guide vane system |
US13/340,761 US8459035B2 (en) | 2007-07-27 | 2011-12-30 | Gas turbine engine with low fan pressure ratio |
US13/484,308 US20120233981A1 (en) | 2007-07-27 | 2012-05-31 | Gas turbine engine with low fan pressure ratio |
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US13/340,761 Continuation US8459035B2 (en) | 2007-07-27 | 2011-12-30 | Gas turbine engine with low fan pressure ratio |
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US14/602,625 Continuation-In-Part US20150132106A1 (en) | 2007-07-27 | 2015-01-22 | Gas turbine engine with low fan pressure ratio |
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US13/484,308 Abandoned US20120233981A1 (en) | 2007-07-27 | 2012-05-31 | Gas turbine engine with low fan pressure ratio |
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US20120222397A1 (en) | 2012-09-06 |
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