CN104011337B - There is the gas-turbine unit of fan variable area nozzle - Google Patents
There is the gas-turbine unit of fan variable area nozzle Download PDFInfo
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- CN104011337B CN104011337B CN201280065397.1A CN201280065397A CN104011337B CN 104011337 B CN104011337 B CN 104011337B CN 201280065397 A CN201280065397 A CN 201280065397A CN 104011337 B CN104011337 B CN 104011337B
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- fan
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- gear train
- nozzle
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/06—Varying effective area of jet pipe or nozzle
- F02K1/09—Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/075—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type controlling flow ratio between flows
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Control Of Turbines (AREA)
Abstract
A kind of cabin components for high bypass gas-turbine unit of the illustrative aspects according to the disclosure, including: core cabin, described core cabin is defined around engine center bobbin thread；Fan cabin, described fan cabin at least partly surrounds described core cabin and is mounted, to limit the fan by-pass flow path for fan by-pass air stream；And fan variable area nozzle, described fan variable area nozzle axially may move relative to described fan cabin, to change fan nozzle discharge area during power operation and to regulate the fan pressure ratio of described fan by-pass air stream, described fan variable area nozzle is operable as changing described fan nozzle discharge area with about 20%.
Description
Cross-Reference to Related Applications
This application claims the priority of the U. S. application No. 13/340,798 submitted on December 30th, 2011, this U. S. application is the part continuation application of the U. S. application No. 13/314,365 that on December 8th, 2011 submits to.
Background of invention.
Technical field
The present invention relates to gas-turbine unit, and relate more specifically to the turbofan with fan variable area nozzle (VAFN), this nozzle axially moves to change its bypass flow path area.
Background technology
Conventional gas-turbine unit generally includes fan section and core-engine, and the diameter of fan section is more than the diameter of core-engine.Fan section and core-engine are within longitudinal axis arranges and is encapsulated in engine compartment components.
Burning gases are discharged through core exhaust nozzle from core-engine, are simultaneously located at the radially outside ring-type fan in main air flow path and flow through the discharge of ring-type fan exhaust nozzle, and this ring-type fan exhaust nozzle is limited between fan cabin and core cabin.Through the major part of the pressurization fan air generation thrust that fan exhaust nozzle is discharged, and the burning gases discharged through core exhaust nozzle provide thrust-drag margin.
The fan nozzle of conventional gas-turbine unit has fixing geometry.The fan nozzle of fixed geometirc structure is suitable for the half-way house of take-off and landing condition and Cruise Conditions.Some gas-turbine units have been carried out fan variable area nozzle.Fan variable area nozzle provides less fan outlet nozzle diameter during Cruise Conditions, and provides bigger fan outlet nozzle diameter during take-off and landing condition.Existing fan variable area nozzle is typically with relative complex mechanism, its degree increasing to engine total weight to offset the fuel efficiency thus increased.
Summary of the invention
A kind of cabin components for high bypass gas-turbine unit of an illustrative aspects according to the disclosure, including: core cabin, described core cabin is defined around engine center bobbin thread；Fan cabin, described fan cabin at least partly surrounds described core cabin and is mounted, to limit the fan by-pass flow path for fan by-pass air stream；And fan variable area nozzle, described fan variable area nozzle axially may move relative to described fan cabin, to change fan nozzle discharge area during power operation and to regulate the fan pressure ratio of described fan by-pass air stream, described fan variable area nozzle is operable as changing described fan nozzle discharge area with about 20%.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include controller, described controller is operable to control described fan variable area nozzle, in order to changes described fan nozzle discharge area and regulates the pressure ratio of described fan by-pass air stream.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described controller can be operable to reduce described fan nozzle discharge area under cruise flight condition.Additionally or alternatively, described controller can be operable to control described fan nozzle discharge area, in order to reduces the unstability of fan.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described fan pressure ratio can be less than about 1.45.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described fan variable area nozzle can limit the trailing edge in described fan cabin.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include controller, described controller is operable axially to move described fan variable area nozzle, in order in response to fan nozzle discharge area described in mission requirements change.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described cabin components may further include gear train, described gear train is driven by the core-engine within described core cabin, in order to drive the fan within described fan cabin.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described fan high-ranking officers positive fan leaf point speed is limited to less than 1150 feet per seconds.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described core-engine includes that low-pressure turbine, described low-pressure turbine define the pressure ratio more than about five (5).
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described core-engine includes low-pressure turbine, and described low-pressure turbine limits the pressure ratio more than five (5).
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described gear train defines the gear reduction ratio more than or equal to about 2.3.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, described assembly may further include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described gear train defines the gear reduction ratio more than or equal to about 2.5.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, by-pass ratio can be defined to more than about six (6) by described bypass stream.
In the further non-limiting embodiments of the embodiment of arbitrary aforementioned components, by-pass ratio can be defined to more than about ten (10) by described bypass stream.Additionally or alternatively, by-pass ratio can be defined to more than about ten (10) by described bypass stream.
Accompanying drawing explanation
For those skilled in the art, the various feature and advantage of the present invention will be apparent from from the detailed description of the invention subsequently of current preferred mode.Can be briefly described below with the accompanying drawing of this detailed description of the invention:
Figure 1A is the general schematically imperfect view in local of the exemplary gas turbogenerator embodiment for using with the present invention；
Figure 1B is the rearview of this engine；
Fig. 1 C is the side view of this engine integrated with suspension bracket；
Fig. 1 D is the stereogram of this engine integrated with suspension bracket；
Fig. 2 A is in the side cross-sectional view of the VAFN of closing position；
Fig. 2 B is in the side cross-sectional view of the VAFN of open position；And
Fig. 3 is the curve map of by-pass conduit normalization (normalized) cross-sectional area distribution；
Fig. 4 is the curve map that effective area increases to nozzle translation；
Fig. 5 is the curve map of conduit area distributions；
Fig. 6 A is the schematic geometric view of auxiliary port position；
Fig. 6 B is the schematic geometric view of auxiliary port entering angle；And
Fig. 6 C is the schematic geometric view of VAFN outer surface curvature.
Detailed description of the invention
Figure 1A shows the imperfect schematic diagram in general local of fan gas turbine engine 10, within fan gas turbine engine 10 is suspended on engine compartment components N from engine lifting bracket P, this engine compartment components N is typical for being designed for the airborne vehicle of subsonic speed operation.
Turbofan 10 includes core-engine within accommodating the core cabin 12 of low rotor 14 and high rotor 24.Low rotor 14 includes low pressure compressor 16 and low-pressure turbine 18.Low rotor 14 drives fan section 20 by gear train 22.High rotor 24 includes high pressure compressor 26 and pressure turbine 28.Burner 30 is arranged between high pressure compressor 26 and pressure turbine 28.Low rotor and high rotor 14,24 rotate around engine rotation axis A.
Engine 10 is the gear-driven aircraft engine of high bypass.In a disclosed non-limiting embodiments, the by-pass ratio of engine 10 is more than about six (6), one of them Example embodiments is more than ten (10), gear train 22 is epicyclic train of gears (such as planetary gear system) or gear reduction ratio other gear trains more than about 2.3, and low-pressure turbine 18 has the pressure ratio more than about 5.In a disclosed embodiment, the by-pass ratio of engine 10 is more than about ten (10:1), and turbofan diameter is noticeably greater than the diameter of low pressure compressor 16, and low-pressure turbine 18 has the pressure ratio more than about 5:1.Gear train 22 can be epicyclic train of gears (such as planetary gear system) or gear reduction ratio other gear trains more than about 2.5:1.It should be appreciated, however, that parameter above is only an illustrative embodiments of gear drive framework engine, and present invention can be suitably applied to include directly driving other gas-turbine units of turbofan.
Air stream enters fan cabin 34, at least partly about core cabin, fan cabin 34 12.Airflow is passed in core cabin 12 by fan section 20, in order to provide power for low pressure compressor 16 and high pressure compressor 26.The core air stream compressed by low pressure compressor 16 and high pressure compressor 26 mixes with the fuel in burner 30, and expands on pressure turbine 28 and low-pressure turbine 18.Turbine 28,18 is connected into and rotates with corresponding rotor 24,14, in order to is rotatably driven compressor 26,16 in response to expansion and is rotatably driven fan section 20 by gear train 22.The core nozzle 43 that core-engine exhaust E is passed through between core cabin 12 and tail cone 32 limiting leaves core cabin 12.
Within core cabin 12 is supported on fan cabin 34 by structure 36, structure 36 is typically commonly referred to as fan outlet stator (FEGV).Bypass flow path 40 is limited between core cabin 12 and fan cabin 34.Engine 10 generates the high bypass flow arrangement with by-pass ratio, and wherein, about the 80% of the air stream in entrance fan cabin 34 becomes bypass stream B.Bypass stream B is transmitted through the fan by-pass flow path 40 of general toroidal, and discharging through fan variable area nozzle (VAFN) 42 from engine 10, fan variable area nozzle 42 defines the fan nozzle discharge area 44 between fan cabin 34 and core cabin 12 at fan cabin end section 34S in the fan cabin 34 in fan section 20 downstream.
Thrust is the function of density, speed and area.One or more amount and direction that can be manipulated to change the thrust provided by bypass stream B in these parameters.Variable area fan nozzle (" VAFN ") 42 operates effectively to change the area of fan nozzle discharge area 44 in response to controller C, in order to be selectively adjusted the pressure ratio of bypass stream B.Low-pressure is than turbofan for its high propulsive efficiency but desirably.But, low-pressure may be inherently prone to by fan stability/Flutter Problem at low-power and low flying speed than fan.VAFN42 allows engine to change at low-power to fan operation line advantageously, it is to avoid unstability region, and still provides relatively small nozzle area necessary for acquisition high efficiency fan operation line when in cruise.
Due to high by-pass ratio, bypass stream B provides the thrust of significant quantity.The fan section 20 of engine 10 is designed to special flying condition typically with about 0.8 Mach and about 35,000 feet cruise.0.8 Mach and 35, the flying condition of 000 foot, wherein engine is at its optimum fuel and also referred to as " stablizes cruise thrust specific fuel consumption (bucket
Cruise Thrust Specific Fuel Consumption, ' TSFC ') " it is the industry standard parameters of the lbt (lbf) that fuel pound quality (lbm) burnt produces at this minimum point divided by engine." low fan pressure ratio " is the pressure ratio individually crossing fan blade when not having fan outlet stator (" FEGV ") system 36.As according to a non-limiting embodiments disclosed herein, low fan pressure is than less than about 1.45." low correction fan tip speed " is that the actual fan tip speed in terms of feet per second is divided by [(TEnvironment deg R) / 518.7)Λ0.5] industry standard temperature correction.As according to a non-limiting embodiments disclosed herein, " low correction fan tip speed " be less than about 1150 feet per seconds.
Owing to the fan blade within fan section 20 is designed to the most fixing negative sweep (stagger for efficient Cruise Conditions efficiently
Angle) place, VAFN42 is operable to effectively change fan nozzle discharge area 44, so as regulation fan by-pass air stream, the angle of attack in fan blade or incidence angle are retained close to what the high efficience motor under other flying conditions (such as land and take off) operated and are designed into firing angle, in order to thus provide the optimized power operation in flying condition scope about performance and other operating parameters (such as noise level).
VAFN42 is separated at least two sector 42A-42B(Figure 1B being limited between suspension bracket P and lower Bi-Fi current divider L), this lower Bi-Fi current divider L typically makes the core cowl of the reverse radome fairing of larger-diameter fan conduit and small diameter interconnect (Fig. 1 C and 1D).Each of at least two sector 42A-42B is independently adjusted asymmetricly to change fan nozzle discharge area 44, in order to generate vectored thrust.Although should be appreciated that and showing two sections, but can alternatively or additionally provide any amount of sections.
In operation, VAFN42 communicates with controller C or the like, in order to regulate fan nozzle discharge area 44 with symmetrical and asymmetric manner.Other control system including engine controller or aircraft control system can also be used in conjunction with.By regulating the whole circumference of VAFN42 symmetrically, wherein, all sectors are moved equably, and thrust efficiency and fuel economy are maximized during each flying condition.By regulating the circumferential sectors 42A-42B of VAFN42 individually to provide asymmetrical fan nozzle discharge area 44, engine bypass stream is by optionally vector quantization, in order to be such as provided solely for trimmed equilibrium or the controlled manipulation of thrust strengthens terrestrial operation or short field performance.
VAFN42 generally includes auxiliary port assembly 50, and it has the first fan cabin section 52 and the second fan cabin section 54 being moveably mounted relative to the first fan cabin section 52.Second fan cabin section 54 is axially slided along engine axis A relative to the first fixing fan cabin section 52, in order to change the effective area of fan nozzle discharge area 44.Second fan cabin section 54 schematically shows in response to actuator 58() schematically show in Fig. 1 C and 1D at track commutator segment cover 56A, 56B() on rearwardly slide.Track commutator segment cover 56A, 56B are adjacent to corresponding suspension bracket P and lower Bi-Fi current divider L and extend (Fig. 1 D) from the first fan cabin section 52.
VAFN42 changes physical area and the geometry of bypass flow path 40 during special flying condition.By making the second fan cabin section 54 slide between closing position (Fig. 2 A) and open position (Fig. 2 B) relative to the first fan cabin section 52, bypass stream B is changed effectively.By the second fan cabin section 54 being orientated as with the first fan cabin section 52 in line so that fan nozzle discharge area 44 to be defined to discharge area F0, auxiliary port assembly 50 is closed (Fig. 2 A).
By making the second fan cabin section 54 rearwardly move to open auxiliary port 60 along track commutator segment cover 56A, 56B away from the first fan cabin section 52, VAFN42 is opened, and auxiliary port 60 extends substantially to provide the discharge area F1 of the fan nozzle discharge area 44 of increase between the second fan cabin section 54 opened is relative to the first fan cabin section 52.It is to say, utilize the discharge area F1 of port 60 more than discharge area F0(Fig. 2 B).
In a disclosed embodiment, auxiliary port 60 is incorporated in the gas extraction system of high by-pass ratio business turbofan, within the by-pass conduit of fan outlet stator (FEGV) afterbody (Fig. 2 A, 2B).Auxiliary port 60 is positioned in the tail-section of by-pass conduit outer wall.
With reference to Fig. 3, the increase of by-pass conduit area distributions, effective area is adapted to provide suitable flow field to translation (Fig. 4), area distributions (Fig. 5) and the position (Fig. 6 A) of auxiliary port 60 and wall curvature (Fig. 6 B-6C), and it allows auxiliary port 60 to obtain required extra effectively discharge area.Due to translation, auxiliary port 60 will substantially make effective area gain double.Auxiliary port 60 provides the method for relatively low weight, and the method provides the discharge area increased not cause high system loss or unacceptable airborne vehicle installation question to control fan operation line.By adjusting by-pass conduit area distributions and outer wall curvature, the stroke at auxiliary port 60 reaches to achieve desired maximum effective area before its effective area increases the limit to be increased.
Auxiliary port pelvic outlet plane 44B(is defined as the plane between the trailing edge of static section and the leading edge of mobile section) initially there is opening, wherein, pelvic outlet plane normal vector is the most axial, but when stroke increases, normal vector becomes more to tilt and the vector of close almost radial direction.Pelvic outlet plane normal, once becoming almost radially, has reached maximum auxiliary port validity.Once reached this point, then the effective area ratio to translating becomes the mild ratio of " only main burner " from the abrupt slope of " port of good design ", because additional areas will be provided due to the inside slope in core cabin 12 by main burner 44A.The auxiliary port nozzle of good design will realize the effective area of about+25% before reaching the port validity limit.It is to say, there is the limited range of stroke, wherein, auxiliary port makes the ratio of extra validity double.Outside this range, the ratio of extra validity can be equivalent to the translation nozzle without auxiliary port.Or in other words, auxiliary port shortens and realizes stroke necessary to expectation effective area for pure flat shifting nozzle.
It is adjusted to ensure the conduit cross section area in auxiliary port 60 front more than port openings cross-sectional area with reference to Fig. 5, the maximum demand effective area more than VAFN42 of the cross-sectional area at auxiliary port 60, and by-pass conduit area distributions.This avoids upstream internal cross section and becomes the situation controlling flow area (that is, less than discharge area), and this situation can result in operating limit and structure problem.
With reference to Fig. 6 A, the auxiliary port 60 in disclosed embodiment is orientated as unlike 0.1 DEL_X/L_DUCT more forward, and 0.1
Point D at maximum radius Rmax of the ring-type fan bypass flow path 40 that DEL_X/L_DUCT limits from the second fan cabin section 54 limits.Rmax is defined through a D and is perpendicular to engine axis A.When the second fan cabin section 54 is in the close position, the some D in disclosed non-limiting embodiments is positioned in inner wall surface 54I of the second fan cabin section 54.DEL_X is the axial distance of the First Point of 60 from Rmax to auxiliary port.L_DUCT is total axial length of ring-type fan bypass flow path 40.Average angle between port lines and fan conduit outer wall is relatively low, to provide the low-loss outlet stream run well.In disclosed embodiment, the entering angle (Theta_in) of the auxiliary port 60 of the wall relative to fan by-pass conduit OD is less than 20 degree (Fig. 6 B), and outer VAFN surface has R_ARC/CHORD > 0.7, wherein, R_ARC is the radial distance of the radial outer wall surface 54O from engine axis A to second fan cabin section 54, and CHORD is the chord length of the second fan cabin section 54.(Fig. 6 C).The curvature of outer wall surface 54O near auxiliary port 60 promotes to flow through auxiliary port 60.In a disclosed embodiment, the stroke of the second fan cabin section 54 necessary for obtaining extra 20% effective discharge area is about 8.4 inches.
In operation, VAFN42 and controller C communicates and moves the second fan cabin section 54 with the first fan cabin section 52 relative to auxiliary port assembly 50, in order to effectively change the area limited by fan nozzle discharge area 44.Various control systems including engine controller or aircraft control system can also be used in conjunction with.By regulating the axial location of the whole circumference of the second fan cabin section 54, wherein, all sectors are moved simultaneously, and by changing fan nozzle discharge area, motor power and fuel economy are maximized during each state of flight.By regulating each sector of the second fan cabin section 54 individually to provide asymmetrical fan nozzle discharge area 44, engine bypass stream is by optionally vector quantization, in order to be such as provided solely for manipulation, the terrestrial operation of enhancing and short field performance that trimmed equilibrium, thrust are controlled.
Description above is exemplary rather than being limited by its interior limiting factor.In view of teaching above, many amendments of the present invention and modification are possible.The preferred embodiment of the present invention has been disclosed, but, it will be appreciated by those of ordinary skill in the art that some amendment will fall within the scope of the present invention.It is understood, therefore, that within the scope of the appended claims, the present invention can by from be particularly described different in the way of put into practice.For this reason, claims should be studied to determine true scope and the content of the present invention.
Claims (14)
1. for a cabin components for high bypass gas-turbine unit, including:
Core cabin, described core cabin is defined around engine center bobbin thread；
Fan cabin, described fan cabin at least partly surrounds described core cabin and is mounted, to limit the fan by-pass flow path for fan by-pass air stream；
Gear train, described gear train is driven by the core-engine within described core cabin, in order to drive the fan within described fan cabin, described fan high-ranking officers positive fan leaf point speed to be limited to less than 1150 feet per seconds；And
Fan variable area nozzle, described fan variable area nozzle axially can move relative to described fan cabin, to change fan nozzle discharge area during power operation and to regulate the fan pressure ratio of described fan by-pass air stream less than in the scope of 1.45, described fan variable area nozzle is operable as changing described fan nozzle discharge area with 20%.
Assembly the most according to claim 1, farther includes controller, and described controller is operable to control described fan variable area nozzle, in order to changes described fan nozzle discharge area and regulates the described pressure ratio of described fan by-pass air stream.
Assembly the most according to claim 2, wherein, described controller is operable to reduce described fan nozzle discharge area under cruise flight condition.
Assembly the most according to claim 2, wherein, described controller is operable to control described fan nozzle discharge area, in order to reduce fan unstability.
Assembly the most according to claim 1, wherein, described fan variable area nozzle defines the trailing edge in described fan cabin.
Assembly the most according to claim 1, farther includes controller, and described controller is operable to axially move described fan variable area nozzle, in order in response to fan nozzle discharge area described in mission requirements change.
Assembly the most according to claim 1, farther includes gear train, and described gear train is driven by the core-engine within described core cabin, in order to drive the fan within described fan cabin.
Assembly the most according to claim 1, farther include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described core-engine includes that low-pressure turbine, described low-pressure turbine define the low-pressure turbine pressure ratio more than five (5).
Assembly the most according to claim 1, farther include gear train, described gear train is driven by the core-engine within described core cabin, to drive the fan within described fan cabin, described core-engine includes that low-pressure turbine, described low-pressure turbine define the low-pressure turbine pressure ratio more than five (5).
Assembly the most according to claim 1, farther include gear train, described gear train is driven by the core-engine within described core cabin, in order to driving the fan within described fan cabin, described gear train defines the gear reduction ratio more than or equal to 2.3.
11. assemblies according to claim 1, farther include gear train, described gear train is driven by the core-engine within described core cabin, in order to driving the fan within described fan cabin, described gear train defines the gear reduction ratio more than or equal to 2.5.
12. assemblies according to claim 1, wherein, by-pass ratio is defined to more than six (6) by described fan by-pass air stream.
13. assemblies according to claim 1, wherein, by-pass ratio is defined to more than ten (10) by described fan by-pass air stream.
14. assemblies according to claim 1, wherein, by-pass ratio is defined to more than ten (10) by described fan by-pass air stream.
Applications Claiming Priority (4)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US13/340798 | 2011-12-30 | ||
| US13/340,798 US20130149112A1 (en) | 2011-12-08 | 2011-12-30 | Gas turbine engine with fan variable area nozzle |
| US13/340,798 | 2011-12-30 | ||
| PCT/US2012/071946 WO2013141934A2 (en) | 2011-12-30 | 2012-12-28 | Gas turbine engine with fan variable area nozzle |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN104011337A CN104011337A (en) | 2014-08-27 |
| CN104011337B true CN104011337B (en) | 2016-08-24 |
Family
ID=49223433
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201280065397.1A Active CN104011337B (en) | 2011-12-30 | 2012-12-28 | There is the gas-turbine unit of fan variable area nozzle |
Country Status (4)
| Country | Link |
|---|---|
| EP (1) | EP2798162A4 (en) |
| CN (1) | CN104011337B (en) |
| SG (1) | SG11201403545SA (en) |
| WO (1) | WO2013141934A2 (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108626023B (en) * | 2017-03-23 | 2019-12-10 | 中国航发商用航空发动机有限责任公司 | turbofan engine with large bypass ratio and exhaust device with variable bypass exhaust area |
| CN114320612B (en) * | 2021-12-08 | 2022-08-23 | 北京航空航天大学 | Engine outer duct backflow control method and device, engine and aircraft |
| CN114992670B (en) * | 2022-05-12 | 2023-06-13 | 中国航发四川燃气涡轮研究院 | Stress application integrated support plate with variable blending function |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US3747343A (en) * | 1972-02-10 | 1973-07-24 | United Aircraft Corp | Low noise prop-fan |
| CN101523041A (en) * | 2006-09-29 | 2009-09-02 | 空中客车法国公司 | Jet engine nacelle for an aircraft and aircraft comprising such a nacelle |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6901739B2 (en) * | 2003-10-07 | 2005-06-07 | General Electric Company | Gas turbine engine with variable pressure ratio fan system |
| EP2064434B1 (en) * | 2006-10-12 | 2012-06-27 | United Technologies Corporation | Operational line management of low pressure compressor in a turbofan engine |
| WO2008045071A1 (en) * | 2006-10-12 | 2008-04-17 | United Technologies Corporation | Gas turbine engine with a variable exit area fan nozzle, nacelle assembly of such a engine, and corresponding operating method |
| EP2069630B1 (en) * | 2006-10-12 | 2013-11-20 | United Technologies Corporation | Nacelle assembly and corresponding method |
| FR2910937B1 (en) * | 2007-01-02 | 2009-04-03 | Airbus France Sas | AIRCRAFT REACTOR NACELLE AND AIRCRAFT COMPRISING SUCH A NACELLE |
| US8074440B2 (en) * | 2007-08-23 | 2011-12-13 | United Technologies Corporation | Gas turbine engine with axial movable fan variable area nozzle |
| US9010126B2 (en) * | 2008-02-20 | 2015-04-21 | United Technologies Corporation | Gas turbine engine with variable area fan nozzle bladder system |
| US8141366B2 (en) * | 2008-08-19 | 2012-03-27 | United Technologies Corporation | Gas turbine engine with variable area fan nozzle |
| US8973364B2 (en) * | 2008-06-26 | 2015-03-10 | United Technologies Corporation | Gas turbine engine with noise attenuating variable area fan nozzle |
| US8739515B2 (en) * | 2009-11-24 | 2014-06-03 | United Technologies Corporation | Variable area fan nozzle cowl airfoil |
| GB201117824D0 (en) * | 2011-10-17 | 2011-11-30 | Rolls Royce Plc | Variable area nozzle for gas turbine engine |
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2012
- 2012-12-28 SG SG11201403545SA patent/SG11201403545SA/en unknown
- 2012-12-28 CN CN201280065397.1A patent/CN104011337B/en active Active
- 2012-12-28 WO PCT/US2012/071946 patent/WO2013141934A2/en active Application Filing
- 2012-12-28 EP EP12871636.2A patent/EP2798162A4/en not_active Withdrawn
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3747343A (en) * | 1972-02-10 | 1973-07-24 | United Aircraft Corp | Low noise prop-fan |
| CN101523041A (en) * | 2006-09-29 | 2009-09-02 | 空中客车法国公司 | Jet engine nacelle for an aircraft and aircraft comprising such a nacelle |
Also Published As
| Publication number | Publication date |
|---|---|
| SG11201403545SA (en) | 2014-07-30 |
| WO2013141934A3 (en) | 2013-11-07 |
| CN104011337A (en) | 2014-08-27 |
| EP2798162A2 (en) | 2014-11-05 |
| WO2013141934A2 (en) | 2013-09-26 |
| EP2798162A4 (en) | 2015-08-26 |
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