US20210197976A1 - Pylon shape with geared turbofan for structural stiffness - Google Patents
Pylon shape with geared turbofan for structural stiffness Download PDFInfo
- Publication number
- US20210197976A1 US20210197976A1 US17/198,610 US202117198610A US2021197976A1 US 20210197976 A1 US20210197976 A1 US 20210197976A1 US 202117198610 A US202117198610 A US 202117198610A US 2021197976 A1 US2021197976 A1 US 2021197976A1
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- Prior art keywords
- fan
- gas turbine
- turbine
- pylon
- turbine engine
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- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
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- 239000012530 fluid Substances 0.000 claims description 2
- 239000007789 gas Substances 0.000 description 44
- 230000000712 assembly Effects 0.000 description 6
- 238000000429 assembly Methods 0.000 description 6
- 239000000446 fuel Substances 0.000 description 5
- 230000004323 axial length Effects 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 241000272168 Laridae Species 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
Images
Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/10—Aircraft characterised by the type or position of power plants of gas-turbine type
- B64D27/12—Aircraft characterised by the type or position of power plants of gas-turbine type within, or attached to, wings
-
- 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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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/16—Aircraft characterised by the type or position of power plants of jet type
- B64D27/18—Aircraft characterised by the type or position of power plants of jet type within, or attached to, wings
-
- B64D27/26—
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/40—Arrangements for mounting power plants in aircraft
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep
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- B64D2027/266—
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/40—Arrangements for mounting power plants in aircraft
- B64D27/404—Suspension arrangements specially adapted for supporting vertical loads
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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
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
- F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Definitions
- a minimum distance must be maintained between a bottom of a gas turbine engine and a runway, resulting in space limitations for wing mounted gas turbine engines. Larger landing gear can be employed to raise the aircraft, and therefore the gas turbine engine, relative to the runway. However, this can add weight to the aircraft. As fan section becomes larger, there are fewer options for mounting a gas turbine engine.
- a gas turbine engine includes a mounting structure for mounting the engine to a pylon.
- a propulsor section includes a fan having a fan diameter.
- a geared architecture drives the fan.
- a generator section includes a fan drive turbine that drives the geared architecture.
- a turbine to fan diameter ratio is substantially less than 45%.
- any of the foregoing assemblies includes the gas turbine engine of claim 1 assembled to a pylon mounted to a wing.
- the fan extends forward of the wing.
- a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
- the geared architecture includes an epicyclic gearbox.
- gas turbine engines is mounted to a pylon.
- the pylon includes a forward portion and an aft portion.
- the fan is attached to the forward portion of the pylon.
- the turbine section includes the fan drive turbine which is attached to the aft portion of the pylon.
- the turbine section is a low pressure turbine.
- the turbine to fan diameter ratio is substantially 35% to substantially 40%.
- a gas turbine engine includes a mounting structure for mounting the engine to a pylon.
- the pylon includes a forward portion and an aft portion.
- a propulsor section includes a fan having a fan diameter. The fan is attached to the forward portion of the pylon.
- a geared architecture drives the fan.
- a compressor section is included.
- a combustor is in fluid communication with the compressor section.
- a generator section includes a fan drive turbine that drives the geared architecture.
- a turbine to fan diameter ratio that is substantially less than 45%, and the turbine section is attached to the aft portion of the pylon.
- any of the foregoing assemblies includes the gas turbine engine assembled to a pylon mounted to a wing.
- the fan extends forward of the wing.
- a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
- the geared architecture includes an epicyclic gearbox.
- the turbine section is a low pressure turbine.
- the turbine to fan diameter ratio is substantially 35% to substantially 40%.
- FIG. 1 illustrates a schematic view of an embodiment of a gas turbine engine
- FIG. 2 illustrates a side view of the gas turbine engine mounted to a pylon
- FIG. 3 illustrates a front view of the gas turbine engine mounted to the pylon
- FIG. 1 schematically illustrates an example gas turbine engine 20 , such as a geared turbofan engine, that includes a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engines might include an augmenter section (not shown) among other systems or features.
- the fan section 22 includes a fan 42 and drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to the combustor section 26 .
- the combustor section 26 air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24 .
- the gas turbine engine 20 can have a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive the fan 42 via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
- the example gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about a central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
- the low speed spool 30 generally includes an inner shaft 40 that connects the fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46 .
- the inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48 , to drive the fan 42 at a lower speed than the low speed spool 30 .
- the high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the central longitudinal axis A.
- a propulsor section includes the fan 42 and a portion of the geared architecture 48 .
- a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54 .
- the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54 .
- the high pressure turbine 54 includes only a single stage.
- a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
- a generator section includes the compressors 44 and 52 , the combustor 56 , and the turbines 46 and 54 , as well as a portion of the geared architecture 48 .
- the example low pressure turbine 46 has a pressure ratio that is greater than about 5.
- the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- a mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
- the mid-turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46 .
- the air in the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46 .
- the mid-turbine frame 58 includes vanes 60 , which are in the core flow path C and function as an inlet guide vane for the low pressure turbine 46 . Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58 . Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28 . Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
- the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
- the gas turbine engine 20 includes a bypass ratio greater than substantially six (6), with an example embodiment being greater than substantially ten (10).
- the example geared architecture 48 is an epicyclic gearbox, an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than substantially 2.3.
- the gas turbine engine 20 includes a bypass ratio greater than substantially ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture 48 and that the present disclosure is applicable to other gas turbine engines.
- the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at substantially 0.8 Mach and substantially 35,000 feet.
- TSFC Thrust Specific Fuel Consumption
- Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than substantially 1.50. In another non-limiting embodiment the low fan pressure ratio is less than substantially 1.45.
- the example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than substantially 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than substantially 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than substantially 6 turbine rotors schematically indicated at 34 . In another non-limiting example embodiment the low pressure turbine 46 includes substantially 3 turbine rotors. A ratio between the number of fan blades and the number of low pressure turbine rotors is between substantially 3.3 and substantially 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
- FIGS. 2 and 3 illustrate the gas turbine engine 20 mounted to a pylon 74 , which is mounted to a wing 64 .
- the efficiency of the turbine driven geared architecture 48 disclosed above, enables the use and fabrication of a smaller low pressure turbine 46 , both in diameter and in the number or overall stages, as compared to a direct drive engine that rotates at a less efficient speed. This allows for alternate and more efficient mounting configurations of the gas turbine engine 20 .
- a minimum distance 70 or clearance should be maintained between a bottom 66 of the gas turbine engine 20 and a runway 62 (or ground) and should be taken into consideration when determining a mounting configuration for the gas turbine engine 20 .
- a core engine section including the low pressure turbine 46
- the fan section 22 can be disposed forward of the wing 64 .
- a diameter 68 of the low pressure turbine 46 is much smaller than a diameter 76 of the fan 42 .
- a ratio of the diameter 68 of the low pressure turbine 46 to the diameter 76 of the fan 42 is defined as the turbine to fan diameter ratio.
- the turbine to fan diameter ratio is substantially less than 45%.
- the turbine to fan diameter ratio is substantially 25% to 45%.
- the turbine to fan diameter ratio is substantially 35% to substantially 40%.
- the gas turbine engine 20 can also be located more aft due to the reduced diameter 68 of the low pressure turbine 46 .
- the pylon 74 can also have a height 72 that is greater, or maintained at a given size, than a pylon 74 which mounts a correspondingly direct drive engine, while still maintaining the desired minimum distance 70 relative to the runway 62 .
- the increase in height 72 provides for additional structural support and strength.
- the pylon 74 includes a forward portion 90 and an aft portion 92 .
- the fan 42 is attached to the forward portion 90 of the pylon 74
- the low pressure turbine 46 is attached to the aft portion 92 of the pylon 74 .
- the wing 64 has a sheared or gull wing configuration.
- the wing 64 includes an up angle portion 78 and a flat portion 80 .
- a distance 82 , or gutter height, of substantially 11 inches should be maintained between the wing 64 and an upper portion 88 of the gas turbine engine 20 .
- the up angle portion 78 provides additional space for the gas turbine engine 20 to be mounted under the wing 64 , and even allows the mounting of a larger gas turbine engine 20 , while still maintaining a desired distance 82 and the minimum distance 70 or clearance. That is, a reduced distance 82 between the gas turbine engine 20 and the wing 64 , reducing interference drag.
- the gear reduction, or geared architecture 48 can be considered part of the turbofan architecture without departing from the scope of the disclosed embodiments.
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- Chemical & Material Sciences (AREA)
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- General Engineering & Computer Science (AREA)
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Abstract
A gas turbine includes a mounting structure for mounting the engine to a pylon. A propulsor section includes a fan having a fan diameter. A geared architecture drives the fan. A generator section includes a fan drive turbine that drives the geared architecture. A turbine to fan diameter ratio is substantially less than 45%.
Description
- This application is a continuation of U.S. application Ser. No. 16/246,596 filed on Jan. 14, 2019, which is a continuation of U.S. application Ser. No. 14/432,581 filed on Mar. 31, 2015, now U.S. Pat. No. 10,179,653 granted on Jan. 15, 2019, which is a United States National Phase application of PCT/US2013/025717 filed on Feb. 12, 2013, which claims priority to U.S. Provisional Application No. 61/708,927 filed on Oct. 2, 2012.
- A minimum distance must be maintained between a bottom of a gas turbine engine and a runway, resulting in space limitations for wing mounted gas turbine engines. Larger landing gear can be employed to raise the aircraft, and therefore the gas turbine engine, relative to the runway. However, this can add weight to the aircraft. As fan section becomes larger, there are fewer options for mounting a gas turbine engine.
- A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a mounting structure for mounting the engine to a pylon. A propulsor section includes a fan having a fan diameter. A geared architecture drives the fan. A generator section includes a fan drive turbine that drives the geared architecture. A turbine to fan diameter ratio is substantially less than 45%.
- In a further embodiment of any of the foregoing assemblies includes the gas turbine engine of claim 1 assembled to a pylon mounted to a wing.
- In a further embodiment of any of the foregoing assemblies the fan extends forward of the wing.
- In a further embodiment of any of the foregoing assemblies a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
- In a further embodiment of any of the foregoing gas turbine engines the geared architecture includes an epicyclic gearbox.
- In a further embodiment of any of the foregoing gas turbine engines is mounted to a pylon. The pylon includes a forward portion and an aft portion. The fan is attached to the forward portion of the pylon. The turbine section includes the fan drive turbine which is attached to the aft portion of the pylon.
- In a further embodiment of any of the foregoing gas turbine engines and mounting systems the turbine section is a low pressure turbine.
- In a further embodiment of any of the foregoing gas turbine engines and mounting systems the turbine to fan diameter ratio is substantially 35% to substantially 40%.
- A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a mounting structure for mounting the engine to a pylon. The pylon includes a forward portion and an aft portion. A propulsor section includes a fan having a fan diameter. The fan is attached to the forward portion of the pylon. A geared architecture drives the fan. A compressor section is included. A combustor is in fluid communication with the compressor section. A generator section includes a fan drive turbine that drives the geared architecture. A turbine to fan diameter ratio that is substantially less than 45%, and the turbine section is attached to the aft portion of the pylon.
- In a further embodiment of any of the foregoing assemblies includes the gas turbine engine assembled to a pylon mounted to a wing.
- In a further embodiment of any of the foregoing assemblies the fan extends forward of the wing.
- In a further embodiment of any of the foregoing assemblies a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
- In a further embodiment of any of the foregoing gas turbine engines the geared architecture includes an epicyclic gearbox.
- In a further embodiment of any of the foregoing gas turbine engines the turbine section is a low pressure turbine.
- In a further embodiment of any of the foregoing gas turbine engines the turbine to fan diameter ratio is substantially 35% to substantially 40%.
- These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
-
FIG. 1 illustrates a schematic view of an embodiment of a gas turbine engine; -
FIG. 2 illustrates a side view of the gas turbine engine mounted to a pylon; and -
FIG. 3 illustrates a front view of the gas turbine engine mounted to the pylon; -
FIG. 1 schematically illustrates an examplegas turbine engine 20, such as a geared turbofan engine, that includes afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. Thefan section 22 includes afan 42 and drives air along a bypass flow path B while thecompressor section 24 draws air in along a core flow path C where air is compressed and communicated to thecombustor section 26. In thecombustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through theturbine section 28 where energy is extracted and utilized to drive thefan section 22 and thecompressor section 24. - Although the disclosed non-limiting embodiment depicts a geared turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with geared turbofans as the teachings may be applied to other types of traditional turbine engines. For example, the
gas turbine engine 20 can have a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive thefan 42 via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. - The example
gas turbine engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about a central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided. - The
low speed spool 30 generally includes aninner shaft 40 that connects thefan 42 and a low pressure (or first)compressor section 44 to a low pressure (or first)turbine section 46. Theinner shaft 40 drives thefan 42 through a speed change device, such as a gearedarchitecture 48, to drive thefan 42 at a lower speed than thelow speed spool 30. The high-speed spool 32 includes anouter shaft 50 that interconnects a high pressure (or second)compressor section 52 and a high pressure (or second)turbine section 54. Theinner shaft 40 and theouter shaft 50 are concentric and rotate via thebearing systems 38 about the central longitudinal axis A. A propulsor section includes thefan 42 and a portion of the gearedarchitecture 48. - A
combustor 56 is arranged between thehigh pressure compressor 52 and thehigh pressure turbine 54. In one example, thehigh pressure turbine 54 includes at least two stages to provide a double stagehigh pressure turbine 54. In another example, thehigh pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. A generator section includes thecompressors combustor 56, and theturbines architecture 48. - The example
low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the examplelow pressure turbine 46 is measured prior to an inlet of thelow pressure turbine 46 as related to the pressure measured at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. - A
mid-turbine frame 58 of the enginestatic structure 36 is arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. Themid-turbine frame 58 furthersupports bearing systems 38 in theturbine section 28 as well as setting airflow entering thelow pressure turbine 46. - The air in the core flow path C is compressed by the
low pressure compressor 44 then by thehigh pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expanded through thehigh pressure turbine 54 andlow pressure turbine 46. Themid-turbine frame 58 includes vanes 60, which are in the core flow path C and function as an inlet guide vane for thelow pressure turbine 46. Utilizing the vane 60 of themid-turbine frame 58 as the inlet guide vane forlow pressure turbine 46 decreases the length of thelow pressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of theturbine section 28. Thus, the compactness of thegas turbine engine 20 is increased and a higher power density may be achieved. - The disclosed
gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, thegas turbine engine 20 includes a bypass ratio greater than substantially six (6), with an example embodiment being greater than substantially ten (10). The example gearedarchitecture 48 is an epicyclic gearbox, an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than substantially 2.3. - In one disclosed embodiment, the
gas turbine engine 20 includes a bypass ratio greater than substantially ten (10:1) and the fan diameter is significantly larger than an outer diameter of thelow pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a gearedarchitecture 48 and that the present disclosure is applicable to other gas turbine engines. - A significant amount of thrust is provided by the air in the bypass flow path B due to the high bypass ratio. The
fan section 22 of thegas turbine engine 20 is designed for a particular flight condition—typically cruise at substantially 0.8 Mach and substantially 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. - “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than substantially 1.50. In another non-limiting embodiment the low fan pressure ratio is less than substantially 1.45.
- The example gas turbine engine includes the
fan 42 that comprises in one non-limiting embodiment less than substantially 26 fan blades. In another non-limiting embodiment, thefan section 22 includes less than substantially 20 fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than substantially 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment thelow pressure turbine 46 includes substantially 3 turbine rotors. A ratio between the number of fan blades and the number of low pressure turbine rotors is between substantially 3.3 and substantially 8.6. The examplelow pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number of turbine rotors 34 in thelow pressure turbine 46 and the number of blades in thefan section 22 disclose an examplegas turbine engine 20 with increased power transfer efficiency. -
FIGS. 2 and 3 illustrate thegas turbine engine 20 mounted to apylon 74, which is mounted to awing 64. The efficiency of the turbine driven gearedarchitecture 48, disclosed above, enables the use and fabrication of a smallerlow pressure turbine 46, both in diameter and in the number or overall stages, as compared to a direct drive engine that rotates at a less efficient speed. This allows for alternate and more efficient mounting configurations of thegas turbine engine 20. - A
minimum distance 70 or clearance should be maintained between a bottom 66 of thegas turbine engine 20 and a runway 62 (or ground) and should be taken into consideration when determining a mounting configuration for thegas turbine engine 20. When theengine 20 is mounted to thewing 64, a core engine section, including thelow pressure turbine 46, can be mounted under thewing 64, and thefan section 22 can disposed forward of thewing 64. - A
diameter 68 of thelow pressure turbine 46 is much smaller than adiameter 76 of thefan 42. A ratio of thediameter 68 of thelow pressure turbine 46 to thediameter 76 of thefan 42 is defined as the turbine to fan diameter ratio. In one example, the turbine to fan diameter ratio is substantially less than 45%. In one example, the turbine to fan diameter ratio is substantially 25% to 45%. In one example, the turbine to fan diameter ratio is substantially 35% to substantially 40%. As thelow pressure turbine 46 has asmall diameter 68, this allows the central longitudinal axis A to be located closer (or “close coupled”, which provides for a cost and weight savings) to thewing 64 than a correspondingly capable direct drive engine. Thegas turbine engine 20 can also be located more aft due to the reduceddiameter 68 of thelow pressure turbine 46. - As the
low pressure turbine 46 has a reduceddiameter 68, thepylon 74 can also have aheight 72 that is greater, or maintained at a given size, than apylon 74 which mounts a correspondingly direct drive engine, while still maintaining the desiredminimum distance 70 relative to therunway 62. The increase inheight 72 provides for additional structural support and strength. - The
pylon 74 includes aforward portion 90 and anaft portion 92. Thefan 42 is attached to theforward portion 90 of thepylon 74, and thelow pressure turbine 46 is attached to theaft portion 92 of thepylon 74. - As further shown in
FIG. 3 , thewing 64 has a sheared or gull wing configuration. Thewing 64 includes an upangle portion 78 and aflat portion 80. Adistance 82, or gutter height, of substantially 11 inches should be maintained between thewing 64 and anupper portion 88 of thegas turbine engine 20. The upangle portion 78 provides additional space for thegas turbine engine 20 to be mounted under thewing 64, and even allows the mounting of a largergas turbine engine 20, while still maintaining a desireddistance 82 and theminimum distance 70 or clearance. That is, a reduceddistance 82 between thegas turbine engine 20 and thewing 64, reducing interference drag. - The gear reduction, or geared
architecture 48, can be considered part of the turbofan architecture without departing from the scope of the disclosed embodiments. - The foregoing description is only exemplary of the principles of the invention. Many modifications and variations are possible in light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than using the example embodiments which have been specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.
Claims (15)
1. A gas turbine engine comprising:
a mounting structure for mounting the engine to a pylon; and
a propulsor section including a fan having a fan diameter; and a geared architecture driving the fan; and
a generator section including a fan drive turbine that drives the geared architecture, wherein a turbine to fan diameter ratio is substantially less than 45%.
2. An assembly including the gas turbine engine of claim 1 assembled to a pylon mounted to a wing.
3. The assembly of claim 2 wherein the fan extends forward of the wing.
4. The assembly of claim 2 wherein a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
5. The gas turbine engine as recited in claim 1 wherein the geared architecture includes an epicyclic gearbox.
6. The gas turbine engine as recited in claim 1 , mounted to a pylon, and wherein the pylon includes a forward portion and an aft portion, the fan is attached to the forward portion of the pylon, and the turbine section including the fan drive turbine is attached to the aft portion of the pylon.
7. The gas turbine engine and mounting system as recited in claim 1 wherein the turbine section is a low pressure turbine.
8. The gas turbine engine and mounting system as recited in claim 1 wherein the turbine to fan diameter ratio is substantially 35% to substantially 40%.
9. A gas turbine engine comprising:
a mounting structure for mounting the engine to a pylon, the pylon including a forward portion and an aft portion;
a propulsor section including a fan having a fan diameter, wherein the fan is attached to the forward portion of the pylon;
a geared architecture driving the fan;
a compressor section;
a combustor in fluid communication with the compressor section; and
a generator section including a fan drive turbine that drives the geared architecture, a turbine to fan diameter ratio that is substantially less than 45%, and the turbine section is attached to the aft portion of the pylon.
10. An assembly including the gas turbine engine of claim 1 assembled to a pylon mounted to a wing.
11. The assembly of claim 10 wherein the fan extends forward of the wing.
12. The assembly of claim 10 wherein a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine
13. The gas turbine engine as recited in claim 9 wherein the geared architecture includes an epicyclic gearbox.
14. The gas turbine engine as recited in claim 9 wherein the turbine section is a low pressure turbine.
15. The gas turbine engine as recited in claim 9 wherein the turbine to fan diameter ratio is substantially 35% to substantially 40%.
Priority Applications (1)
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US17/198,610 US20210197976A1 (en) | 2012-10-02 | 2021-03-11 | Pylon shape with geared turbofan for structural stiffness |
Applications Claiming Priority (5)
Application Number | Priority Date | Filing Date | Title |
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US201261708927P | 2012-10-02 | 2012-10-02 | |
PCT/US2013/025717 WO2014055103A1 (en) | 2012-10-02 | 2013-02-12 | Pylon shape with geared turbofan for structural stiffness |
US201514432581A | 2015-03-31 | 2015-03-31 | |
US16/246,596 US11021257B2 (en) | 2012-10-02 | 2019-01-14 | Pylon shape with geared turbofan for structural stiffness |
US17/198,610 US20210197976A1 (en) | 2012-10-02 | 2021-03-11 | Pylon shape with geared turbofan for structural stiffness |
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US16/246,596 Continuation US11021257B2 (en) | 2012-10-02 | 2019-01-14 | Pylon shape with geared turbofan for structural stiffness |
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US20210197976A1 true US20210197976A1 (en) | 2021-07-01 |
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US14/432,581 Active 2033-10-02 US10179653B2 (en) | 2012-10-02 | 2013-02-12 | Pylon shape with geared turbofan for structural stiffness |
US16/246,596 Active 2033-05-24 US11021257B2 (en) | 2012-10-02 | 2019-01-14 | Pylon shape with geared turbofan for structural stiffness |
US17/198,610 Abandoned US20210197976A1 (en) | 2012-10-02 | 2021-03-11 | Pylon shape with geared turbofan for structural stiffness |
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US14/432,581 Active 2033-10-02 US10179653B2 (en) | 2012-10-02 | 2013-02-12 | Pylon shape with geared turbofan for structural stiffness |
US16/246,596 Active 2033-05-24 US11021257B2 (en) | 2012-10-02 | 2019-01-14 | Pylon shape with geared turbofan for structural stiffness |
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EP (2) | EP2904239B1 (en) |
WO (1) | WO2014055103A1 (en) |
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US10001083B2 (en) * | 2014-07-18 | 2018-06-19 | MTU Aero Engines AG | Turbofan aircraft engine |
US20180237120A1 (en) * | 2017-02-22 | 2018-08-23 | General Electric Company | Aircraft with Under Wing Direct Drive Low Pressure Turbine |
GB201805764D0 (en) * | 2018-04-06 | 2018-05-23 | Rolls Royce Plc | A casing |
GB201820924D0 (en) | 2018-12-21 | 2019-02-06 | Rolls Royce Plc | Turbine engine |
US11204037B2 (en) | 2018-12-21 | 2021-12-21 | Rolls-Royce Plc | Turbine engine |
GB201820925D0 (en) * | 2018-12-21 | 2019-02-06 | Rolls Royce Plc | Turbine engine |
US11428160B2 (en) | 2020-12-31 | 2022-08-30 | General Electric Company | Gas turbine engine with interdigitated turbine and gear assembly |
US20240060430A1 (en) * | 2022-08-17 | 2024-02-22 | General Electric Company | Gas turbine engine |
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EP2904239A1 (en) | 2015-08-12 |
US20150239568A1 (en) | 2015-08-27 |
EP3792474B1 (en) | 2023-10-11 |
US11021257B2 (en) | 2021-06-01 |
US20200023981A1 (en) | 2020-01-23 |
EP2904239B1 (en) | 2020-05-06 |
US10179653B2 (en) | 2019-01-15 |
EP2904239A4 (en) | 2015-11-04 |
EP3792474A1 (en) | 2021-03-17 |
WO2014055103A1 (en) | 2014-04-10 |
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