WO2008045091A1 - Gas turbine engine fan variable area nozzle with swivalable insert system - Google Patents
Gas turbine engine fan variable area nozzle with swivalable insert system Download PDFInfo
- Publication number
- WO2008045091A1 WO2008045091A1 PCT/US2006/040249 US2006040249W WO2008045091A1 WO 2008045091 A1 WO2008045091 A1 WO 2008045091A1 US 2006040249 W US2006040249 W US 2006040249W WO 2008045091 A1 WO2008045091 A1 WO 2008045091A1
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- Prior art keywords
- fan
- inserts
- engine
- recited
- bypass flow
- Prior art date
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Classifications
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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
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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
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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
- F02K1/52—Nozzles specially constructed for positioning adjacent to another nozzle or to a fixed member, e.g. fairing
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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
- the present invention relates to a gas turbine engine, and more particularly to a turbofan engine having an insert system which swivels relative a fan bypass flow to vary the effective area of a fan nozzle exit area.
- Conventional gas turbine engines generally include a fan section and a core engine with the fan section having a larger diameter than that of the core engine.
- the fan section and the core engine are disposed about a longitudinal axis and are enclosed within an engine nacelle assembly.
- Combustion gases are discharged from the core engine through a primary airflow path and are exhausted through a core exhaust nozzle.
- An annular fan flow path disposed radially outwardly of the primary airflow path, passes through a radial outer portion between a fan nacelle and a core nacelle and is discharged through an annular fan exhaust nozzle defined at least partially by the fan nacelle and the core nacelle to generate fan thrust.
- a majority of propulsion thrust is provided by the pressurized fan air discharged through the fan exhaust nozzle, the remaining thrust provided from the combustion gases discharged through the core exhaust nozzle.
- the fan nozzles of conventional gas turbine engines have fixed geometry.
- the fixed geometry fan nozzles must be suitable for take-off and landing conditions as well as for cruise conditions.
- the requirements for take-off and landing conditions are different from requirements for the cruise condition.
- Optimum performance of the engine may be achieved during different flight conditions of an aircraft by tailoring the fan exhaust nozzle for the specific flight regimes.
- a turbofan engine includes a fan variable area nozzle (FVAN) which effectively changes the physical area and geometry within a fan bypass flow path to manipulate the pressure ratio of a fan bypass flow therethrough.
- the FVAN generally includes a multitude of aerodynamically shaped inserts circumferentially located about the core nacelle within the fan bypass flow path.
- the inserts at a fully stowed position take up a minimum of area within the fan bypass flow path to effectively maximize the fan nozzle exit area while in the fully deployed position (cruise) take up a maximum of area within the fan bypass flow path to effectively minimize the fan nozzle exit area of the FVAN.
- the FVAN provides an asymmetrical fan nozzle exit area such that engine fan bypass flow is selectively vectored.
- the present invention therefore provides an effective, lightweight fan variable area nozzle for a gas turbine engine.
- Figure IA is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention.
- Figure IB is an expanded sectional view of the engine
- Figure 2A is a sectional view of the engine of Figure IA through line 2-2 of the FVAN in a closed position
- Figure 2B is a sectional view of the engine of Figure IA through line 2-2 of the FVAN in a thrust vectored position.
- Figure IA illustrates a general partial fragmentary schematic view of a gas 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 engine within a core nacelle 12 that houses a low spool 14 and high spool 24.
- the low spool 14 includes a low pressure compressor 16 and low pressure turbine 18.
- the low spool 14 drives a fan section 20 through a gear train 22.
- the high spool 24 includes a high pressure compressor 26 and high pressure turbine 28.
- a combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28.
- the low and high spools 14, 24 rotate about an engine axis of rotation A.
- the engine 10 is preferably a high-bypass geared turbofan aircraft engine.
- the engine 10 bypass ratio is greater than ten (10)
- the turbofan diameter is significantly larger than that of the low pressure compressor 16
- the low pressure turbine 18 has a pressure ratio that is greater than 5.
- the gear train 22 is preferably an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5. It should be understood, however, that the above parameters are only exemplary of a preferred geared turbofan engine and that the present invention is likewise applicable to other gas turbine engines.
- Airflow enters a fan nacelle 34, which at least partially surrounds the core nacelle 12.
- the fan section 20 communicates airflow into the core nacelle 12 to power the low pressure compressor 16 and the high pressure compressor 26.
- Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 and expanded over the high pressure turbine 28 and low pressure turbine 18.
- the turbines 28, 18 are coupled for rotation with, respective, spools 24, 14 to rotationally drive the compressors 26, 16 and through the gear train 22, the fan section 20 in response to the expansion.
- a core engine exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and a tail cone 32.
- the core nacelle 12 is supported within the fan nacelle 34 by a pylon structure 36 often generically referred to as an upper bifurcation 36U and lower bifurcation 36L, however, other types of pylons at various radial locations may likewise be usable with the present invention.
- a bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle 34.
- the engine 10 generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering the fan nacelle 34 becomes bypass flow B.
- the bypass flow B communicates through the generally annular fan bypass flow path 40 and is discharged from the engine 10 through an annular fan nozzle 42 which defines a fan nozzle exit area 44 between the fan nacelle 34 and the core nacelle 12.
- 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.
- the fan section 20 of the engine 10 is preferably designed for a particular flight condition — typically cruise at 0.8M and 35,000 feet.
- the fan section 20 includes fan blades which are designed at a particular fixed stagger angle for an efficient cruise condition.
- a fan variable area nozzle (FVAN) 50 within the bypass flow B operates to change the effective area thereof.
- the FVAN 50 is operated to effectively vary the fan nozzle exit area 44 to adjust fan bypass air flow such that the angle of attack or incidence on the fan blade is maintained close to design incidence at other flight conditions such as landing and takeoff, thus enabling optimized engine operation over a range of flight condition with respect to performance and other operational parameters such as noise levels.
- the bypass flow B through the nozzle exit area 44 is effectively altered through articulation of the FVAN 50. That is, articulation of the FVAN 50 effectively changes the physical area and geometry of the fan bypass flow path 40 to manipulate the area through which the bypass flow B must pass.
- the FVAN 50 generally includes a multitude of inserts 52, with an associated linkage 54 and an actuator system 56 (also shown in Figure IB). Each inserts 52 is movably mounted to the core nacelle 12 through the linkage 54 in response to the actuator system 56.
- the inserts 52 are circumferentially located about the core nacelle 12 ( Figure 2A).
- the inserts 52 may be mounted to the core engine structure with cut-outs in the core nacelle 12 such that the inserts 52 extend therefrom. In other words, the inserts 52 may be located at various locations adjacent the bypass flow path so long as the inserts 52 may selectively be extended to effect the bypass flow.
- the inserts 52 are preferably aerodynamically shaped members located upstream of the annular fan nozzle 42 between the fan nacelle 34 and the core nacelle 12.
- the linkage 54 of each insert 52 is preferably driven by the actuator system 56 to move the insert 52 between a fully extended and a fully stowed position.
- the inserts 52 are also shaped to interface with the core nacelle 12 to provide a flush aerodynamic surface when in the stowed position.
- the linkage 52 also preferably orients each insert 52 in response to the bypass flow B at each of a multiple of positions such as take-off/landing, climb and cruise.
- the FVAN 50 preferably provides an approximately 20% (twenty percent) effective change in the fan nozzle exit area 44. That is, the FVAN 50 at the fully deployed position (cruise) takes up a maximum area within the fan bypass flow path 40 to effectively minimize the fan nozzle exit area 44. The FVAN 50 at the fully stowed position (takeoff/landing) takes up a minimum of area within the bypass flow path 40 to effectively maximize the fan nozzle exit area 44. It should be understood that an essentially infinite number of positions intermediate the fully deployed and fully stowed position may be utilized for other flight conditions such as climb or maneuvering. It should also be understood that other insert orientations as well as thrust vectored positions in which some circumferential insets 52 of the FVAN 50 are differentially positioned relative other circumferential inserts 52 ( Figure 2B) are likewise usable with the present invention.
- the FVAN 50 communicates with the controller C to effectively vary the fan nozzle exit area 44 in response to movement of the inserts 52 within the fan bypass flow path 40.
- Other control systems including an engine controller flight control system may likewise be usable with the present invention.
- a nacelle assembly for a gas turbine engine comprising: a core nacelle defined about an axis; a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and an fan variable area nozzle for movement between a stowed position and a deployed position within said fan bypass flow path to selectively vary a fan nozzle exit area.
- said fan variable area nozzle includes a multiple of inserts movably mounted to said core nacelle, each of said multiple of inserts forming a generally flush aerodynamic surface with said core nacelle when in said stowed position.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Control Of Turbines (AREA)
Abstract
A turbofan engine (10) includes a fan variable area nozzle (FVAN) (50) which effectively changes the physical area and geometry within a fan bypass flow path (40) to manipulate the pressure ratio of the bypass flow. The FVAN generally includes a multitude of aerodynamical Iy shaped inserts (52) circumf erentially located about the core nacelle (12). The FVAN at a fully stowed position (takeoff /landing) takes up a minimum of area within the fan bypass flow path to effectively maximize the fan nozzle exit area (44) while in the fully deployed position (cruise) takes up a maximum of area within the bypass flow path to effectively minimize the fan nozzle exit area. By separately adjusting each of the multiple of inserts relative the other inserts the FVAN provides an asymmetrical fan nozzle exit area to selectively vector the fan bypass flow.
Description
GAS TURBINE ENGINE FAN VARIABLE AREA NOZZLE WITH SWIVALABLE INSERT SYSTEM
BACKGROUND OF THE INVENTION
The present invention relates to a gas turbine engine, and more particularly to a turbofan engine having an insert system which swivels relative a fan bypass flow to vary the effective area of a fan nozzle exit area.
Conventional gas turbine engines generally include a fan section and a core engine with the fan section having a larger diameter than that of the core engine. The fan section and the core engine are disposed about a longitudinal axis and are enclosed within an engine nacelle assembly.
Combustion gases are discharged from the core engine through a primary airflow path and are exhausted through a core exhaust nozzle. An annular fan flow path, disposed radially outwardly of the primary airflow path, passes through a radial outer portion between a fan nacelle and a core nacelle and is discharged through an annular fan exhaust nozzle defined at least partially by the fan nacelle and the core nacelle to generate fan thrust. A majority of propulsion thrust is provided by the pressurized fan air discharged through the fan exhaust nozzle, the remaining thrust provided from the combustion gases discharged through the core exhaust nozzle.
The fan nozzles of conventional gas turbine engines have fixed geometry. The fixed geometry fan nozzles must be suitable for take-off and landing conditions as well as for cruise conditions. However, the requirements for take-off and landing conditions are different from requirements for the cruise condition. Optimum performance of the engine may be achieved during different flight conditions of an aircraft by tailoring the fan exhaust nozzle for the specific flight regimes.
Accordingly, it is desirable to provide an effective, lightweight fan variable area nozzle for a gas turbine engine.
SUMMARY OF THE INVENTION
A turbofan engine according to the present invention includes a fan variable area nozzle (FVAN) which effectively changes the physical area and geometry within a fan bypass flow path to manipulate the pressure ratio of a fan bypass flow therethrough. The FVAN generally includes a multitude of aerodynamically shaped inserts circumferentially located about the core nacelle within the fan bypass flow path.
The inserts at a fully stowed position (takeoff/landing) take up a minimum of area within the fan bypass flow path to effectively maximize the fan nozzle exit area while in the fully deployed position (cruise) take up a maximum of area within the fan bypass flow path to effectively minimize the fan nozzle exit area of the FVAN.
By adjusting the inserts symmetrically in which all inserts are moved uniformly, thrust efficiency and fuel economy are maximized during each flight condition. By separately adjusting each of the multiple of inserts relative to the other inserts the FVAN provides an asymmetrical fan nozzle exit area such that engine fan bypass flow is selectively vectored.
The present invention therefore provides an effective, lightweight fan variable area nozzle for a gas turbine engine.
BRIEF DESCRIPTION OF THE DRAWINGS
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:
Figure IA is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention;
Figure IB is an expanded sectional view of the engine; Figure 2A is a sectional view of the engine of Figure IA through line 2-2 of the FVAN in a closed position; and
Figure 2B is a sectional view of the engine of Figure IA through line 2-2 of the FVAN in a thrust vectored position.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
Figure IA illustrates a general partial fragmentary schematic view of a gas 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 engine within a core nacelle 12 that houses a low spool 14 and high spool 24. The low spool 14 includes a low pressure compressor 16 and low pressure turbine 18. The low spool 14 drives a fan section 20 through a gear train 22. The high spool 24 includes a high pressure compressor 26 and high pressure turbine 28. A combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28. The low and high spools 14, 24 rotate about an engine axis of rotation A.
The engine 10 is preferably a high-bypass geared turbofan aircraft engine. Preferably, the engine 10 bypass ratio is greater than ten (10), the turbofan diameter is significantly larger than that of the low pressure compressor 16, and the low pressure turbine 18 has a pressure ratio that is greater than 5. The gear train 22 is preferably an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5. It should be understood, however, that the above parameters are only exemplary of a preferred geared turbofan engine and that the present invention is likewise applicable to other gas turbine engines. Airflow enters a fan nacelle 34, which at least partially surrounds the core nacelle 12. The fan section 20 communicates airflow into the core nacelle 12 to power the low pressure compressor 16 and the high pressure compressor 26. Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 and expanded over the high pressure turbine 28 and low pressure turbine 18. The turbines 28, 18 are coupled for rotation with, respective, spools 24, 14 to rotationally drive the compressors 26, 16 and through the gear train 22, the fan section 20 in response to
the expansion. A core engine exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and a tail cone 32.
The core nacelle 12 is supported within the fan nacelle 34 by a pylon structure 36 often generically referred to as an upper bifurcation 36U and lower bifurcation 36L, however, other types of pylons at various radial locations may likewise be usable with the present invention.
A bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle 34. The engine 10 generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering the fan nacelle 34 becomes bypass flow B. The bypass flow B communicates through the generally annular fan bypass flow path 40 and is discharged from the engine 10 through an annular fan nozzle 42 which defines a fan nozzle exit area 44 between the fan nacelle 34 and the core nacelle 12.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. 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. The fan section 20 of the engine 10 is preferably designed for a particular flight condition — typically cruise at 0.8M and 35,000 feet. The fan section 20 includes fan blades which are designed at a particular fixed stagger angle for an efficient cruise condition. A fan variable area nozzle (FVAN) 50 within the bypass flow B operates to change the effective area thereof.
The FVAN 50 is operated to effectively vary the fan nozzle exit area 44 to adjust fan bypass air flow such that the angle of attack or incidence on the fan blade is maintained close to design incidence at other flight conditions such as landing and takeoff, thus enabling optimized engine operation over a range of flight condition with respect to performance and other operational parameters such as noise levels.
The bypass flow B through the nozzle exit area 44 is effectively altered through articulation of the FVAN 50. That is, articulation of the FVAN 50 effectively changes the physical area and geometry of the fan bypass flow path 40 to manipulate the area through which the bypass flow B must pass. The FVAN 50 generally includes a multitude of inserts 52, with an associated linkage 54 and an actuator system 56 (also shown in Figure IB). Each inserts 52 is movably mounted
to the core nacelle 12 through the linkage 54 in response to the actuator system 56. Preferably, the inserts 52 are circumferentially located about the core nacelle 12 (Figure 2A). Alternatively, the inserts 52 may be mounted to the core engine structure with cut-outs in the core nacelle 12 such that the inserts 52 extend therefrom. In other words, the inserts 52 may be located at various locations adjacent the bypass flow path so long as the inserts 52 may selectively be extended to effect the bypass flow.
The inserts 52 are preferably aerodynamically shaped members located upstream of the annular fan nozzle 42 between the fan nacelle 34 and the core nacelle 12. The linkage 54 of each insert 52 is preferably driven by the actuator system 56 to move the insert 52 between a fully extended and a fully stowed position. The inserts 52 are also shaped to interface with the core nacelle 12 to provide a flush aerodynamic surface when in the stowed position. The linkage 52 also preferably orients each insert 52 in response to the bypass flow B at each of a multiple of positions such as take-off/landing, climb and cruise.
The FVAN 50 preferably provides an approximately 20% (twenty percent) effective change in the fan nozzle exit area 44. That is, the FVAN 50 at the fully deployed position (cruise) takes up a maximum area within the fan bypass flow path 40 to effectively minimize the fan nozzle exit area 44. The FVAN 50 at the fully stowed position (takeoff/landing) takes up a minimum of area within the bypass flow path 40 to effectively maximize the fan nozzle exit area 44. It should be understood that an essentially infinite number of positions intermediate the fully deployed and fully stowed position may be utilized for other flight conditions such as climb or maneuvering. It should also be understood that other insert orientations as well as thrust vectored positions in which some circumferential insets 52 of the FVAN 50 are differentially positioned relative other circumferential inserts 52 (Figure 2B) are likewise usable with the present invention.
In operation, the FVAN 50 communicates with the controller C to effectively vary the fan nozzle exit area 44 in response to movement of the inserts 52 within the fan bypass flow path 40. Other control systems including an engine controller flight control system may likewise be usable with the present invention.
By adjusting the inserts 52 axi-symmetrically such that all inserts are moved uniformly, thrust efficiency and fuel economy are maximized during each flight condition. By separately adjusting the circumferential inserts 52 within, for example, each sectors 42A-42D (Figure 2B) to provide an asymmetrical fan nozzle exit area 44, engine fan bypass flow is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering, enhanced ground operations or short field performance.
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.
What is claimed is:
1. A nacelle assembly for a gas turbine engine comprising: a core nacelle defined about an axis; a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and an fan variable area nozzle for movement between a stowed position and a deployed position within said fan bypass flow path to selectively vary a fan nozzle exit area.
2. The assembly as recited in claim 1, wherein said fan variable area nozzle includes a multiple of inserts movably mounted to said core nacelle.
3. The assembly as recited in claim 1, wherein said fan variable area nozzle includes an aerodynamically shaped insert.
4. The assembly as recited in claim 3, further comprising a linkage between said aerodynamically shaped insert and said core nacelle, said linkage defining a multiple of orientations of said aerodynamically shaped insert.
5. The assembly as recited in claim 4, further comprising an actuator system mounted to said core nacelle to drive said linkage.
6. The assembly as recited in claim 1, wherein said fan variable area nozzle includes a multiple of inserts movably mounted to said core nacelle, each of said multiple of inserts forming a generally flush aerodynamic surface with said core nacelle when in said stowed position.
Claims
7. The assembly as recited in claim 1, wherein said fan variable area nozzle includes a multiple of inserts movably mounted to said core nacelle, each of said multiple of inserts located at a circumferential position about said core nacelle.
8. A gas turbine engine comprising: a core engine defined about an axis; a turbofan driven about said axis; a core nacelle defined at least partially about said core engine; a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and a multiple of aerodynamically shaped inserts for movement between a stowed position and a deployed position within said fan bypass flow path to selectively vary a fan nozzle exit area.
9. The engine as recited in claim 8, wherein said multiple of inserts are each independently moveably relative said core nacelle.
10. The engine as recited in claim 9, further comprising a controller in communication with an actuator system to independently position each of said multiple of inserts in response to a flight condition.
11. The engine as recited in claim 8, wherein each of said multiple of inserts form a flush aerodynamic surface with said core nacelle when in said stowed position.
12. The engine as recited in claim 8, further comprising a gear system driven by said core engine, said turbofan driven by said gear system about said axis.
13. A method of varying a fan nozzle exit area of a gas turbine engine comprising the steps of:
(A) deploying a multiple of aerodynamically shaped inserts to a deployed position within a fan bypass flow path to selectively vary a fan nozzle exit area in response to a flight condition.
14. A method as recited in claim 13, wherein said step (A) further comprises:
(a) deploying the a multiple of aerodynamically shaped inserts from a stowed position aerodynamically flush with a core nacelle.
15. A method as recited in claim 13, wherein said step (A) further comprises:
(a) deploying the multiple of aerodynamically shaped inserts to a maximum deployed position in response to a cruise flight condition.
15. A method as recited in claim 12, wherein said step (A) further comprises:
(a) stowing the multiple of aerodynamically shaped inserts to a stowed position aerodynamically flush with a core nacelle in response to a landing and takeoff flight condition.
16. A method as recited in claim 12, wherein said step (A) further comprises:
(a) deploying the multiple of aerodynamically shaped inserts to an intermediate position in response to a non-cruise flight condition.
17. A method as recited in claim 12, wherein said step (A) further comprises: (a) asymmetrically deploying one of the multiple of aerodynamically shaped inserts relative another of the multiple of aerodynamically shaped inserts to vector a fan thrust.
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/441,607 US8272202B2 (en) | 2006-10-12 | 2006-10-12 | Gas turbine engine fan variable area nozzle with swivalable insert system |
EP06850521.3A EP2069629B1 (en) | 2006-10-12 | 2006-10-12 | Gas turbine engine fan variable area nozzle with swivalable insert system |
PCT/US2006/040249 WO2008045091A1 (en) | 2006-10-12 | 2006-10-12 | Gas turbine engine fan variable area nozzle with swivalable insert system |
US13/625,343 US8662417B2 (en) | 2006-10-12 | 2012-09-24 | Gas turbine engine fan variable area nozzle with swivable insert system |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
PCT/US2006/040249 WO2008045091A1 (en) | 2006-10-12 | 2006-10-12 | Gas turbine engine fan variable area nozzle with swivalable insert system |
Related Child Applications (2)
Application Number | Title | Priority Date | Filing Date |
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US11/441,607 A-371-Of-International US7673772B2 (en) | 2006-02-10 | 2006-05-26 | Method for dispensing medications |
US13/625,343 Continuation US8662417B2 (en) | 2006-10-12 | 2012-09-24 | Gas turbine engine fan variable area nozzle with swivable insert system |
Publications (1)
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WO2008045091A1 true WO2008045091A1 (en) | 2008-04-17 |
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PCT/US2006/040249 WO2008045091A1 (en) | 2006-10-12 | 2006-10-12 | Gas turbine engine fan variable area nozzle with swivalable insert system |
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US (2) | US8272202B2 (en) |
EP (1) | EP2069629B1 (en) |
WO (1) | WO2008045091A1 (en) |
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EP2383442A2 (en) | 2010-04-30 | 2011-11-02 | Rolls-Royce plc | Gas turbine engine |
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CN114718757A (en) * | 2022-03-23 | 2022-07-08 | 西北工业大学 | Design scheme of tail spray pipe of micro turbojet engine |
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Also Published As
Publication number | Publication date |
---|---|
EP2069629A1 (en) | 2009-06-17 |
US8662417B2 (en) | 2014-03-04 |
EP2069629B1 (en) | 2014-09-24 |
US20100043394A1 (en) | 2010-02-25 |
US8272202B2 (en) | 2012-09-25 |
US20130047579A1 (en) | 2013-02-28 |
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