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Patent
United Technologies | Date: 2017-03-08

A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

Claims which contain your search:

2. The engine as set forth in claim 1, wherein said power density is a ratio of a thrust provided by said engine (20) to a volume of a turbine section (28) including both said high pressure turbine (54) and said low pressure turbine (46), said thrust is sea level take-off, flat-rated static thrust.

3. The engine (20) as set forth in claim 1 or 2, wherein said fan section delivers a portion of air into a bypass duct and a portion of the air into said low pressure compressor (44) as core flow, and has a bypass ratio greater than about 6.0.

...

Patent
United Technologies | Date: 2017-01-11

A turbofan engine (20) includes an engine case (22), a gaspath through the engine case (22), a fan (42) having a circumferential array of fan blades, a compressor in fluid communication with the fan (42), a combustor (48) in fluid communication with the compressor, and a turbine in fluid communication with the combustor (48). The turbine has a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204). A speed reduction mechanism (46) couples the fan drive turbine section (27) to the fan (42). A bypass area ratio is between 8.0 and 20.0. A ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan (42) is less than 0.50. A ratio of a turbine section airfoil count to the bypass area ratio is between 10 and 170.

Claims which contain your search:

1. A turbofan engine (20) comprising:an engine case (22);a gaspath through the engine case (22);a fan (42) having a circumferential array of fan blades;a compressor in fluid communication with the fan (42);a combustor (48) in fluid communication with the compressor;a turbine in fluid communication with the combustor (48), the turbine having a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204); anda speed reduction (46) mechanism coupling the fan drive turbine section (27) to the fan (42), wherein a bypass area ratio is between 8.0 and 20.0, a ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan (42) is less than 0.50, and a ratio of a turbine section airfoil count to the bypass area ratio is between 10 and 170, said fan drive turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine section (27).

3. The engine (20) of claim 1 or 2, wherein a hub-to-tip ratio (RI:RO) of the fan drive turbine section (27) is between 0.4 and 0.5 measured at the maximum RO axial location in the fan drive turbine section (27).

13. A turbofan engine (20) comprising:a fan case (40); anda gas generator including a core cowl, wherein the fan case (40) and core cowl are configured so that a flow-path bypass ratio therebetween is between 8.0 and 20.0; the gas generator includes a fan drive turbine (27) having at least three blade stages (200, 202, 204) and configured so that a ratio of a turbine airfoil count to the bypass ratio between 10 and 170, said turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine (27) and a ratio of maximum gaspath radius along the fan drive turbine (27) to maximum radius of the fan is less than 0.50, and there being a second turbine section.

...
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Name Score Publications Conferences Grants Patents Trademarks News Webs
2320.8 10 10 10 10 10 10 10
192.4 10 10 10 10 10 10 10
113.0 10 10 10 10 10 10 10
88.8 10 10 10 10 10 10 10
63.3 10 10 10 10 10 10 10
55.1 10 10 10 10 10 10 10
54.1 10 10 10 10 10 10 10
52.7 10 10 10 10 10 10 10
45.1 10 10 10 10 10 10 10
43.7 10 10 10 10 10 10 10
40.4 10 10 10 10 10 10 10
40.1 10 10 10 10 10 10 10
39.7 10 10 10 10 10 10 10
39.5 10 10 10 10 10 10 10
38.5 10 10 10 10 10 10 10
38.3 10 10 10 10 10 10 10
36.0 10 10 10 10 10 10 10
31.9 10 10 10 10 10 10 10
28.6 10 10 10 10 10 10 10
27.1 10 10 10 10 10 10 10
24.6 10 10 10 10 10 10 10
23.5 10 10 10 10 10 10 10
22.6 10 10 10 10 10 10 10
20.5 10 10 10 10 10 10 10
20.4 10 10 10 10 10 10 10
20.1 10 10 10 10 10 10 10
19.6 10 10 10 10 10 10 10
18.5 10 10 10 10 10 10 10
18.4 10 10 10 10 10 10 10
17.5 10 10 10 10 10 10 10
17.1 10 10 10 10 10 10 10
16.2 10 10 10 10 10 10 10
15.9 10 10 10 10 10 10 10
14.5 10 10 10 10 10 10 10
13.9 10 10 10 10 10 10 10
13.8 10 10 10 10 10 10 10
13.8 10 10 10 10 10 10 10
13.5 10 10 10 10 10 10 10
13.3 10 10 10 10 10 10 10
12.9 10 10 10 10 10 10 10
12.8 10 10 10 10 10 10 10
12.5 10 10 10 10 10 10 10
12.5 10 10 10 10 10 10 10
12.0 10 10 10 10 10 10 10
11.9 10 10 10 10 10 10 10
11.6 10 10 10 10 10 10 10
11.6 10 10 10 10 10 10 10
11.4 10 10 10 10 10 10 10
11.3 10 10 10 10 10 10 10
11.3 10 10 10 10 10 10 10
11.3 10 10 10 10 10 10 10
11.2 10 10 10 10 10 10 10
11.1 10 10 10 10 10 10 10
11.0 10 10 10 10 10 10 10
10.9 10 10 10 10 10 10 10
10.7 10 10 10 10 10 10 10
10.6 10 10 10 10 10 10 10
10.6 10 10 10 10 10 10 10
10.6 10 10 10 10 10 10 10
10.5 10 10 10 10 10 10 10
Scientific Production Entreprise Aerosila Oao
10.4 - - 1 10 10 10 10
Imperial College London
10.4 29 - - 10 10 10 10
Institutul National Of Cercetare Dezvoltare Turbomotoare Comoti
10.3 - - 1 10 10 10 10
University of Ulsan
10.3 18 - - 10 10 10 10
Yanmar Co.
10.2 - - - 10 10 10 10
Cernium
10.1 - 1 2 10 10 10 10
TU Berlin
10.1 6 8 1 10 10 10 10
Winterthur Gas & Diesel Ltd.
10.1 - - - 10 10 10 10
CAS Institute of Engineering Thermophysics
10.0 5 2 - 10 10 10 10
University of Oxford
9.8 10 1 1 10 10 10 10
Zagazig University
9.7 4 3 - 10 10 10 10
Tsinghua University
9.7 6 4 - 10 10 10 10
Archimhdhs Kentron Kainotomias Kaidimiourgias Archimedes Center For Innovation And Creation
9.6 - - 1 10 10 10 10
THERMOCOAX SAS
9.5 - - 1 10 10 10 10
Bombardier
9.4 1 - 1 10 10 10 10
BAE Systems
9.4 - - 1 10 10 10 10
Compania Espanola De Sistemas Aeronauticos
9.3 - - 1 10 10 10 10
Meggitt Aerospace Ltd
9.3 - - 1 10 10 10 10
ITP Group
9.3 - - 2 10 10 10 10
Mitsubishi Group
9.2 - 1 - 10 10 10 10
N&R Engineering
9.1 - - 1 10 10 10 10
TU Darmstadt
9.1 - 1 2 10 10 10 10
Kobe University
9.0 7 - - 10 10 10 10
Arizona State University
8.9 - 5 - 10 10 10 10
Techspace Aerospace SA
8.8 - - 2 10 10 10 10
Avio S. P. A.
8.7 - 1 2 10 10 10 10
TU Dresden
8.6 - - 2 10 10 10 10
RWTH Aachen
8.6 2 2 1 10 10 10 10
AVIC Commercial Aircraft Engine Co.
8.4 2 4 - 10 10 10 10
Chienkuo Technical University
8.4 3 - - 10 10 10 10
Rohr Inc.
8.2 - - - 10 10 10 10
University of Genoa
8.1 5 2 1 10 10 10 10
Beth Israel Deaconess Medical Center
8.1 23 - - 10 10 10 10
Toshiba Corporation
8.0 1 1 - 10 10 10 10
University of Bologna
7.9 13 - - 10 10 10 10
Shanghai JiaoTong University
7.9 16 - - 10 10 10 10
Mazda Motor Corporation
7.9 - - - 10 10 10 10
Panasonic
7.8 - - - 10 10 10 10
Purdue University
7.8 1 3 - 10 10 10 10
Emory University
7.7 20 - - 10 10 10 10

Patent
United Technologies | Date: 2017-03-08

A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

Claims which contain your search:

2. The engine as set forth in claim 1, wherein said power density is a ratio of a thrust provided by said engine (20) to a volume of a turbine section (28) including both said high pressure turbine (54) and said low pressure turbine (46), said thrust is sea level take-off, flat-rated static thrust.

3. The engine (20) as set forth in claim 1 or 2, wherein said fan section delivers a portion of air into a bypass duct and a portion of the air into said low pressure compressor (44) as core flow, and has a bypass ratio greater than about 6.0.


Patent
United Technologies | Date: 2017-01-11

A turbofan engine (20) includes an engine case (22), a gaspath through the engine case (22), a fan (42) having a circumferential array of fan blades, a compressor in fluid communication with the fan (42), a combustor (48) in fluid communication with the compressor, and a turbine in fluid communication with the combustor (48). The turbine has a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204). A speed reduction mechanism (46) couples the fan drive turbine section (27) to the fan (42). A bypass area ratio is between 8.0 and 20.0. A ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan (42) is less than 0.50. A ratio of a turbine section airfoil count to the bypass area ratio is between 10 and 170.

Claims which contain your search:

1. A turbofan engine (20) comprising:an engine case (22);a gaspath through the engine case (22);a fan (42) having a circumferential array of fan blades;a compressor in fluid communication with the fan (42);a combustor (48) in fluid communication with the compressor;a turbine in fluid communication with the combustor (48), the turbine having a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204); anda speed reduction (46) mechanism coupling the fan drive turbine section (27) to the fan (42), wherein a bypass area ratio is between 8.0 and 20.0, a ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan (42) is less than 0.50, and a ratio of a turbine section airfoil count to the bypass area ratio is between 10 and 170, said fan drive turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine section (27).

3. The engine (20) of claim 1 or 2, wherein a hub-to-tip ratio (RI:RO) of the fan drive turbine section (27) is between 0.4 and 0.5 measured at the maximum RO axial location in the fan drive turbine section (27).

13. A turbofan engine (20) comprising:a fan case (40); anda gas generator including a core cowl, wherein the fan case (40) and core cowl are configured so that a flow-path bypass ratio therebetween is between 8.0 and 20.0; the gas generator includes a fan drive turbine (27) having at least three blade stages (200, 202, 204) and configured so that a ratio of a turbine airfoil count to the bypass ratio between 10 and 170, said turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine (27) and a ratio of maximum gaspath radius along the fan drive turbine (27) to maximum radius of the fan is less than 0.50, and there being a second turbine section.


Patent
United Technologies | Date: 2017-01-11

A turbofan engine (20) includes an engine case (22), a gaspath through the engine case (22), a fan (42) having an array of fan blades, a compressor in fluid communication with the fan (42), a combustor (48) in fluid communication with the compressor, and a turbine in fluid communication with the combustor (48). The turbine has a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204) and a second turbine section (26). A speed reduction mechanism (46) couples the fan drive turbine (27) section to the fan (42). A ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan blades is less than 0.55. A bypass area ratio is greater than 6.0. A ratio of a fan drive turbine section airfoil count to the bypass area ratio is less than 170.

Claims which contain your search:

1. A turbofan engine (20) comprising:an engine case (22);a gaspath through the engine case (22);a fan (42) having an array of fan blades;a compressor in fluid communication with the fan (42);a combustor (48) in fluid communication with the compressor;a turbine in fluid communication with the combustor (48), the turbine having a fan drive turbine section (27) having 3 to 6 blade stages (200, 202, 204);a speed reduction mechanism (46) coupling the fan drive turbine section (27) to the fan, wherein a ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan blades is less than 0.55, a bypass area ratio is greater than 6.0, and a ratio of a fan drive turbine section airfoil count to the bypass area ratio is less than 170; anda second turbine section.

2. The engine (20) of claim 1, wherein the bypass area ratio is:greater than about 8.0; orbetween 8.0 and 20.0.

4. The engine (20) of claim 1, 2 or 3, wherein a hub-to-tip ratio (RI:RO) of the fan drive turbine section (27) is between 0.4 and 0.5 measured at the maximum RO axial location in the fan drive turbine section (27).

7. The engine (20) of any preceding claim, wherein said ratio of maximum gaspath radius along the fan drive turbine section (27) to maximum radius of the fan (42) is less than 0.50, optionally between 0.35 and 0.50.

8. The engine (20) of any preceding claim, wherein:said ratio of fan drive turbine section airfoil count to bypass area ratio is between 10 and 150; and/oran airfoil count of the fan drive turbine section (27) is below 1600.

15. A turbofan engine (20) comprising:a fan (42) having fan blades;a fan case (40); anda gas generator including a core cowl, wherein the fan case (40) and core cowl are configured so that a flow-path bypass ratio therebetween is greater than 6.0, a ratio of maximum gaspath radius along a fan drive turbine section (27) to maximum radius of the fan blades is less than 0.55, and the fan drive turbine section (27) is configured so that a ratio of a turbine airfoil count to the bypass ratio is less than 170, optionally wherein the engine (20) has only a single fan stage, and the fan drive turbine section (27) has blade stages (200, 202, 204) interspersed with vane stages (206, 208).


Patent
United Technologies | Date: 2017-03-22

A gas turbine engine (20) comprises a fan (42) and a turbine section (28) having a first turbine rotor (46). The first turbine rotor (46) drives a compressor rotor (44). A gear reduction (48) effects a reduction in the speed of the fan (42) relative to an input speed from a fan drive turbine rotor. The compressor rotor (44) has a number of compressor blades in at least one of a plurality of rows of the compressor rotor (44). The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor: A compressor module and a method of designing a gas turbine engine (20) are also disclosed.

Claims which contain your search:

3. The gas turbine engine as set forth in claim 1 or 2, wherein the gear reduction (48) has a gear ratio of greater than about 2.3, or greater than about 2.5.

4. The gas turbine engine as set forth in any preceding claim, wherein the fan (42) delivers air into a bypass duct, and a portion of air into the compressor rotor (44), with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor rotor (44), and the bypass ratio being greater than about 6, or greater than about 10.


Patent
United Technologies | Date: 2017-02-15

A turbine engine (10) has a fan shaft (20) and at least one tapered bearing mounted on the fan shaft (20). The fan shaft (20) includes at least one passage extending in a direction having at least a radial component, and adjacent the at least one tapered bearing. A fan (18) is mounted for rotation on the tapered bearing. An epicyclic gear train (22) is coupled to drive the fan (18). The epicyclic gear train (22) includes a carrier (26) supporting intermediate gears (32) that mesh with a sun gear (30) and a ring gear (38) that surrounds and meshes with the intermediate gears (32). Each of the intermediate gears (32) are supported on a respective journal bearing (34). The epicyclic gear train (22) defines a gear reduction ratio of greater than or equal to about 2.3. A turbine section is coupled to drive the fan (18) through the epicyclic gear train (22) and has a fan drive turbine that includes a pressure ratio that is greater than about 5. The fan (18) includes a pressure ratio that is less than about 1.45, and a bypass ratio of greater than about ten.

Claims which contain your search:

8. The turbine engine (10) as recited in any preceding claim, wherein the epicyclic gear train (22) has a gear reduction ratio of greater than or equal to 2.3.

1. A turbine engine (10) comprising:a fan shaft (20);at least one tapered bearing mounted on the fan shaft (20), the fan shaft (20) including at least one passage extending in a direction having at least a radial component, and adjacent the at least one tapered bearing;a fan (18) mounted for rotation on the tapered bearing;an epicyclic gear train (22) coupled to drive the fan (18), the epicyclic gear train (22) including a carrier (26) supporting intermediate gears (32) that mesh with a sun gear (30), and a ring gear (38) surrounding and meshing with the intermediate gears (32), each of the intermediate gears (32) being supported on a respective journal bearing (34), wherein the epicyclic gear train (22) defines a gear reduction ratio of greater than or equal to about 2.3; anda turbine section coupled to drive the fan (18) through the epicyclic gear train (22), the turbine section having a fan drive turbine (208) that includes a pressure ratio that is greater than 5, the fan (18) includes a pressure ratio that is less than 1.45, and the fan (18) has a bypass ratio of greater than ten.

10. The turbine engine (10) as recited in any of claims 1 to 7, wherein the epicyclic gear train (22) has a gear reduction ratio of greater than or equal to 2.5.

11. The turbine engine (10) as recited in any preceding claim, wherein the fan (18) defines a bypass ratio of greater than about 10.5:1 with regard to a bypass airflow and a core airflow.

9. The turbine engine (10) as recited in any of claims 1 to 7, wherein the epicyclic gear train (22) has a gear reduction ratio of greater than or equal to about 2.5.


Patent
United Technologies | Date: 2017-01-11

A turbofan engine (20) comprises a fan (42) having fan blades (70). A compressor is in communication with the fan section. The fan (42) is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is greater than 6.0. A combustor (48) is in fluid communication with the compressor. A turbine is in communication with the combustor (48). The turbine has a first turbine section (27) that includes two or more stages (200, 202, 204) and a second turbine section (26) that includes at least two stages. A ratio of airfoils in the first turbine section (27) to the bypass ratio is less than 170. The first turbine section (27) includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades (70) is less than 0.50. A speed reduction mechanism (46) is coupled to the fan (42) and rotatable by the turbine.

Claims which contain your search:

1. A turbofan engine (20) comprising:a fan (42) having fan blades (70);a compressor in communication with the fan section, wherein the fan (42) is configured to communicate a portion (526) of air into a bypass path (528) defining a bypass area outwardly of the compressor and a portion (522) into the compressor and a bypass ratio defined as air communicated through the bypass path (528) relative to air communicated to the compressor is greater than 6.0;a combustor (48) in fluid communication with the compressor;a turbine in communication with the combustor (48), the turbine having a first turbine section (27) that includes two or more stages (200, 202, 204) and a second turbine section (26) that includes at least two stages, wherein a ratio of airfoils in the first turbine section (27) to the bypass ratio is less than 170 and the first turbine section (27) includes a maximum gas path radius and a ratio of the maximum gas path radius to a maximum radius of the fan blades (70) is less than 0.50; anda speed reduction (46) mechanism coupled to the fan (42) and rotatable by the turbine.

5. A turbofan engine (20) comprising:a fan (42) having fan blades (70);a compressor in communication with the fan section, wherein the compressor includes a first compressor section (30) including at least two stages and a second compressor section (28) including at least five stages, the second compressor section (28) is configured to operate at a higher pressure than the first compressor section (30), the fan (42) is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor, and a bypass ratio defined as air communicated through the bypass path relative to air communicated to the compressor is greater than 6.0;a combustor (48) in fluid communication with the compressor;a turbine in communication with the combustor (48), the turbine having a first turbine section (27) and a second turbine section (26), wherein a ratio of airfoils in the first turbine section (27) to the bypass ratio is less than 170 and the first turbine section (27) includes a maximum gas path radius and a ratio of the maximum gas path radius to a maximum radius of the fan blades (70) is less than 0.50; anda speed reduction mechanism (46) coupled to the fan (42) and rotatable by the turbine.

9. The turbofan engine (20) as recited in claim 8, wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section (26) is between 0.4 and 0.5, measured at a maximum Ro axial location within the low pressure turbine.

12. The turbofan engine (20) as recited in any preceding claim, wherein the bypass ratio is greater than 8.0.

13. The turbofan engine (20) as recited in any preceding claim, wherein a fan pressure ratio across the fan (42) is less than 1.45.

15. The turbofan engine (20) as recited in claim 14, wherein the epicyclic gearbox provides a speed reduction ratio between 2:1 and 5:1.


Patent
United Technologies | Date: 2017-01-04

A geared architecture (22) for a gas turbine engine (10) comprises a fan shaft (20) and a fan (18) supported on the fan shaft (20) and defining a bypass flow ratio greater than six. A frame (108) supports the fan shaft (20). A gear system (100) drives the fan shaft (20). The gear system (100) has a gear reduction ratio of greater than or equal to about 2.3. A torque frame (28) at least partially supports the gear system (100). An input (102) is coupled to the gear system (100). A downstream turbine (27a) is coupled to rotatably drive the input coupling (102). The downstream turbine (27a) defines a pressure ratio that is greater than five.

Claims which contain your search:

1. A geared architecture (22) for a gas turbine engine (10) comprising:a fan shaft (20) and a fan (18) supported on said fan shaft (20) and defining a bypass flow ratio greater than six;a frame (108) which supports said fan shaft (20);a gear system (100) which drives said fan shaft (20), said gear system (100) having a gear reduction ratio of greater than or equal to about 2.3;a torque frame (28) which at least partially supports said gear system (100);an input coupling (102) to said gear system (100); anda downstream turbine (27a) coupled to rotatably drive said input coupling (102), said downstream turbine (27a) defining a pressure ratio that is greater than five.

11. The geared architecture (22) as recited in any preceding claim, wherein said bypass flow ratio is greater than ten.

12. The geared architecture (22) as recited in any preceding claim, wherein said gear reduction ratio is greater than or equal to about 2.5.


Patent
United Technologies | Date: 2017-01-10

A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section. The engine also includes a rotating element and at least one bearing compartment including a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. A method and section for a gas turbine engine are also disclosed.

Claims which contain your search:

1. A gas turbine engine comprising: a fan section, a bypass passage, a compressor section, and a turbine section arranged along an engine longitudinal axis; a rotating element and at least one bearing compartment including a bearing for supporting said rotating element; wherein said at least one bearing compartment has a first seal and a second seal each associated with a corresponding one of two opposed axial ends, on either axial side of said bearing relative to said engine longitudinal axis, at least one of said first seal and said second seal being a non-contacting seal having a seal face facing a rotating face of said rotating element; and wherein a bypass ratio is defined as the volume of air passing into said bypass passage compared to the volume of air passing into said compressor section, wherein said bypass ratio is greater than 10 at a cruise condition.

14. The gas turbine engine as set forth in claim 8, wherein said fan drive turbine is configured to drive said gear arrangement, said fan drive turbine defining a turbine pressure ratio greater than 5:1, measured prior to an inlet of said fan drive turbine as related to a pressure at an outlet of said fan drive turbine prior to an exhaust nozzle.

16. The gas turbine engine as set forth in claim 1, wherein said fan section comprises at least one fan blade, with a low fan pressure ratio of less than 1.45, the low fan pressure ratio measured across the at least one fan blade alone.

17. The gas turbine engine as set forth in claim 16, wherein said rotating element is configured to rotate at a velocity greater than or equal to about 450 feet per second, and said gear arrangement defines a gear reduction ratio of greater than 2.3:1.

18. A method of operating a gas turbine engine, the method comprising the steps of: arranging a bearing within a bearing compartment to support a rotating element, said rotating element defining a rotating face, said bearing compartment having a first seal and a second seal each associated with a corresponding one of two opposed axial ends, on either axial side of said bearing; rotating said rotating face relative to at least one of said first seal and said second seal; sealing said bearing compartment with said first seal and said second seal, at least one of said first seal and said second seal being a non-contacting seal configured to resist leakage of lubricant outwardly of said bearing compartment and to allow air to flow from a chamber adjacent said seal and into said bearing compartment, said non-contacting seal defining a seal face facing said rotating face; and communicating air from a fan to a bypass passage and to compressor section, wherein a bypass ratio is defined as the volume of air passing into said bypass passage compared to the volume of air passing into said compressor section, said bypass ratio greater than 10 at a cruise condition.

20. The method as set forth in claim 18, wherein said step of rotating comprises rotating said rotating element at a velocity greater than or equal to 450 feet per second, and said fan comprises at least one fan blade, with a low fan pressure ratio of less than 1.45, the low fan pressure ratio measured across the at least one fan blade alone.


Patent
United Technologies | Date: 2017-06-21

A gas turbine engine (20,100,105) comprises a nacelle (94,99) and a fan rotor (42) carrying a plurality of fan blades (98,114). The nacelle (94,99) is formed such that one portion (96,106) extends axially further from the fan blades (98,114) than does another portion (97,108). The nacelle (94,99) has inner periphery (107) that is substantially axially symmetric about a center axis (C) of the rotor (42) from either a throat (110) of the nacelle (94,99) at a substantially bottom dead center location, or a point of inflection (112) at which the inner periphery (107) of the nacelle (94,99) at substantially bottom dead center merges a convex portion into a concave portion.

Claims which contain your search:

6. The gas turbine engine as set forth in any preceding claim, wherein a second distance (L) is defined from a plane (X) defined by leading edges (102,116) of said fan blades (98,114) to an axial location of a forwardmost part (96,97,106,108) of said nacelle (94,99), and an outer diameter (D) of said fan blades (98,114) is defined, and a ratio of said distance (L) to said outer diameter (D) is between 0.2 and 0.5.

7. The gas turbine engine as set forth in claim 6, wherein said ratio is greater than 0.25.

8. The gas turbine engine as set forth in claim 7, wherein said ratio is greater than 0.30.

9. The gas turbine engine as set forth in claim 6, 7 or 8, wherein said ratio is less than 0.40.

10. The gas turbine engine as set forth in any of claims 6 to 9, wherein the second distance (L) measured to the one portion (96,106) of said nacelle still results in said ratio being less than 0.45, and the second distance (L) measured to the other portion (97,108) of said nacelle (94,99) still results in said ratio being greater than 0.2.

12. The gas turbine engine as set forth in claim 11, wherein a gear ratio of said gear reduction (48) is greater than 2.3.

13. The gas turbine engine as set forth in claim 11 or 12, wherein a pressure ratio across said fan drive turbine (46) is greater than 5.

14. The gas turbine engine as set forth in any preceding claim, wherein said fan rotor (42) delivers air into a bypass duct as bypass air (B), and into a core engine including a compressor (24), and a bypass ratio being defined as the volume of air (B) being delivered into said bypass duct to the volume of air (C) delivered into said core engine, with said bypass ratio being greater than 6.


Patent
United Technologies | Date: 2017-06-14

A gas turbine engine (80) comprises a compressor section (82) and a turbine section (93), with the turbine section having a first stage blade row (91) and a downstream blade row (110). A higher pressure tap (86) is tapped from a higher pressure first location in the compressor (82). A lower pressure tap (88) is tapped from a lower pressure location in the compressor (82) which is at a lower pressure than the first location. The higher pressure tap (86) passes through a heat exchanger (84), and then is delivered to cool the first stage blade row (91) in the turbine section (93). The lower pressure tap (84) is delivered to at least partially cool the downstream blade row (110).

Claims which contain your search:

8. The gas turbine engine as set forth in claim 7, wherein said bypass ratio is greater than or equal to about 10.0.

10. The gas turbine engine as set forth in claim 9, wherein a gear ratio of said gear reduction (48) is greater than or equal to about 2.3:1.

7. The gas turbine engine as set forth in any preceding claim, wherein a fan is positioned upstream of said compressor section (82) and said fan delivering air into a bypass duct as propulsion air, and into said compressor section (82) with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0.