US20160201606A1 - Low weight large fan gas turbine engine - Google Patents

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb.

Description

CROSS-REFERENCE TO RELATED APPLICATION
• This application is a continuation of U.S. patent application Ser. No. 14/428,049, which was filed on Mar. 15, 2015, which is a National Stage Entry of PCT Application No. PCT/US2013/025470, filed on Feb. 10, 2013, which claims priority to U.S. Provisional Application Ser. No. 61/708,288, filed Oct. 1, 2012.
BACKGROUND
• A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
• The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.
• A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.
• Although geared architectures have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.
SUMMARY
• A gas turbine engine according to an example of the present disclosure includes a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan and the low pressure turbine section are configured to rotate at a common speed and in a common direction. The gas turbine engine has a bypass ratio of greater than about 10. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb. The fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.
• In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.
• In a further embodiment of any of the foregoing embodiments, the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.
• In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages.
• In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.
• In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a greater number of stages than the low pressure compressor section.
• In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.45.
• In a further embodiment of any of the foregoing embodiments, the fan includes fewer than 20 fan blades.
• In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a pressure ratio of greater than about 5.
• A gas turbine engine according to an example of the present disclosure includes a fan having fewer than 20 fan blades situated at an inlet of a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The low pressure turbine section has a greater number of stages than the low pressure compressor section. The fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.
• In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.
• In a further embodiment of any of the foregoing embodiments, the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.
• In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages.
• In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.
• In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5.
• In a further embodiment of any of the foregoing embodiments, the gas turbine engine has a bypass ratio of greater than about 10.
• In a further embodiment of any of the foregoing embodiments, the low pressure turbine has a pressure ratio of greater than about 5.
• A method of designing a gas turbine engine according to an example of the present disclosure includes providing a fan configured to deliver airflow to a bypass duct, and providing a core engine configured to rotate the fan. The core engine include a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to rotate the fan at a common speed and in a common direction. The fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65. The fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than 10.
• In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about 5.
• A further embodiment of any of the foregoing embodiments includes a low pressure compressor section driven by the low pressure turbine section. The low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages.
• In a featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter less than 0.50 the size of the fan diameter.
• In another embodiment according to the previous embodiment, the diameter of the fan drive turbine is greater than 0.30 the size of the fan diameter.
• In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is between about 0.35 and about 0.45 the size of the fan diameter.
• In another embodiment according to any of the previous embodiments, the fan drive turbine further comprises a high pressure turbine located upstream of the low pressure turbine.
• In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine.
• In another embodiment according to any of the previous embodiments, a compressor section has a low pressure compressor driven by the low pressure turbine and a combustor in fluid communication with the compressor section.
• In another embodiment according to any of the previous embodiments, a first shaft connects the low pressure turbine, low pressure compressor, and the fan drive geared architecture.
• In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box.
• In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is defined by an outer case surface of the fan drive turbine.
• In another embodiment according to any of the previous embodiments, the fan diameter is defined by an outer peripheral surface of the fan blades.
• In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing.
• In another featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter between about 0.35 and about 0.45 the size of the fan diameter.
• In another embodiment according to the previous embodiment, the fan drive geared architecture has a gear reduction ratio of greater than about 2.3
• In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box.
• In another embodiment according to any of the previous embodiments, a compressor section has at least a first compressor and a second compressor, a combustor in fluid communication with the compressor section, and at least one additional turbine. A first shaft connects the fan drive turbine and the first compressor and a second shaft connects the second compressor and the one additional turbine.
• In another embodiment according to any of the previous embodiments, the second shaft rotates at a faster speed than the first shaft.
• In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine and the one additional turbine comprises a high pressure turbine.
• In another embodiment according to any of the previous embodiments, the fan drive geared architecture couples the first shaft to the fan at a location upstream of the compressor section.
• In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing.
• In another embodiment according to any of the previous embodiments, the pylon mount interface comprises at least a front mount beam and a rear mount beam located aft of the front mount beam.
BRIEF DESCRIPTION OF THE DRAWINGS
• The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
• FIG. 1 schematically illustrates a geared turbofan engine embodiment.
• FIG. 2 schematically illustrates a direct drive turbine engine embodiment.
• FIG. 3 shows a side view of a geared turbofan embodiment in one example mounting configuration.
• FIG. 4 shows an end view ofFIG. 3 in an aft direction looking forward.
DETAILED DESCRIPTION
• FIG. 1 schematically illustrates an examplegas turbine engine20 that includes afan section22, acompressor section24, acombustor section26 and aturbine section28. Alternative engines might include an augmenter section (not shown) among other systems or features. Thefan section22 drives air along a bypass flow path B while thecompressor section24 draws air in along a core flow path C where air is compressed and communicated to acombustor section26. In thecombustor section26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through theturbine section28 where energy is extracted and utilized to drive thefan section22 and thecompressor section24.
• Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including 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 a fan 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.
• Theexample engine20 generally includes alow speed spool30 and ahigh speed spool32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure36 viaseveral bearing systems38. It should be understood that various bearingsystems38 at various locations may alternatively or additionally be provided.
• Thelow speed spool30 generally includes aninner shaft40 that connects afan42 and a low pressure (or first)compressor section44 to a low pressure (or first)turbine section46. Theinner shaft40 drives thefan42 through a speed change device, such as a gearedarchitecture48, to drive thefan42 at a lower speed than thelow speed spool30. The high-speed spool32 includes anouter shaft50 that interconnects a high pressure (or second)compressor section52 and a high pressure (or second)turbine section54. Theinner shaft40 and theouter shaft50 are concentric and rotate via the bearingsystems38 about the engine central longitudinal axis A.
• Acombustor56 is arranged between thehigh pressure compressor52 and thehigh pressure turbine54. In one example, thehigh pressure turbine54 includes at least two stages to provide a double stagehigh pressure turbine54. In another example, thehigh pressure turbine54 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.
• The examplelow pressure turbine46 has a pressure ratio that is greater than about5. The pressure ratio of the examplelow pressure turbine46 is measured prior to an inlet of thelow pressure turbine46 as related to the pressure measured at the outlet of thelow pressure turbine46 prior to an exhaust nozzle.
• Amid-turbine frame58 of the enginestatic structure36 is arranged generally between thehigh pressure turbine54 and thelow pressure turbine46. Themid-turbine frame58 furthersupports bearing systems38 in theturbine section28 as well as setting airflow entering thelow pressure turbine46.
• The core airflow C is compressed by thelow pressure compressor44 then by thehigh pressure compressor52 mixed with fuel and ignited in thecombustor56 to produce high speed exhaust gases that are then expanded through thehigh pressure turbine54 andlow pressure turbine46. Themid-turbine frame58 includesvanes60, which are in the core airflow path and function as an inlet guide vane for thelow pressure turbine46. Utilizing thevane60 of themid-turbine frame58 as the inlet guide vane forlow pressure turbine46 decreases the length of thelow pressure turbine46 without increasing the axial length of themid-turbine frame58. Reducing or eliminating the number of vanes in thelow pressure turbine46 shortens the axial length of theturbine section28. Thus, the compactness of thegas turbine engine20 is increased and a higher power density may be achieved.
• The disclosedgas turbine engine20 in one example is a high-bypass geared aircraft engine. In a further example, thegas turbine engine20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example gearedarchitecture48 is 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 about 2.3.
• In one disclosed embodiment, thegas turbine engine20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of thelow pressure compressor44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
• A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. Thefan section22 of theengine20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—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 about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.
• “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)]^0.5. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
• The example gas turbine engine includes thefan42 that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, thefan section22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment thelow pressure turbine46 includes no more than about6 turbine rotors schematically indicated at34. In another non-limiting example embodiment thelow pressure turbine46 includes about 3 turbine rotors. A ratio between the number offan blades42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The examplelow pressure turbine46 provides the driving power to rotate thefan section22 and therefore the relationship between the number ofturbine rotors34 in thelow pressure turbine46 and the number ofblades42 in thefan section22 disclose an examplegas turbine engine20 with increased power transfer efficiency.
• The configuration shown inFIG. 2 is a directdrive turbine engine25. The directdrive turbine engine25 includes afan section22′, acompressor section24′, acombustor section26′, and aturbine section28′. Thefan section22′ drives air along a bypass flow path B′ while thecompressor section24′ draws air in along a core flow path C′ where air is compressed and communicated to thecombustor section26′. In thecombustor section26′, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through theturbine section28′ where energy is extracted and utilized to drive thefan section22′ and thecompressor section24′.
• The directdrive turbine engine25 generally includes alow speed spool30′ and ahigh speed spool32′ mounted for rotation about an engine central longitudinal axis A′ relative to an engine static structure viaseveral bearing systems38′. Thelow speed spool30′ generally includes an inner shaft that connects afan42′ having a plurality of blades and a low pressure (or first)compressor section44′ to a low pressure (or first)turbine section46′. The inner shaft orlow speed spool30′ directly drives thefan42′, that is, thefan42′ and lowpressure turbine section46′ are driven at the same speed. The high-speed spool32′ includes an outer shaft that interconnects a high pressure (or second)compressor section52′ and a high pressure (or second)turbine section54′. The inner shaft and the outer shaft are concentric and rotate via the bearingsystems38′ about the engine central longitudinal axis A′.
• In the direct drive configuration shown inFIG. 2, a fan drive turbine directly drives thefan section22′, i.e. there is no geared architecture in this configuration. InFIG. 2, the fan drive turbine comprises thelow pressure turbine46′ which is coupled to directly drive thefan42′.
• The geared architecture configuration has increased efficiency that enables the use and fabrication of a smallerlow pressure turbine46 both in diameter and in the number or overall stages as compared to the direct drive turbine engine25 (FIG. 2), which must rotate at a less efficient speed.
• Moreover, the smaller, more efficientlow pressure turbine46 of the gearedturbofan engine20 enables alternate and more efficient mounting configurations. Space limitations for wing mounted engines result from a minimum distance between a bottom of an engine and the runway. Larger landing gear components can be utilized to raise the aircraft and thereby the engine relative to the runway, but larger landing gear components are not a desirable option due to significant weight penalties. Accordingly, as thepropulsor fan section22 grows in size, the mounting options decrease. For engines having the same fan section diameter, the fan drive turbine section of the direct drive engine25 (FIG. 2) is much larger than the fan drive turbine section of a geared turbofan engine20 (FIG. 1).
• This difference becomes significant when defining a mounting configuration for the engine. The core engine section including the fan drive turbine section can be mounted under the wing, with the fan section extending forward of the wing. The larger fan drive turbine section of a direct drive turbine requires that the engine centerline be spaced a further distance from a bottom surface of the wing as compared to a centerline of a geared turbofan engine with the smaller more efficient fan drive turbine. Even modest reductions in this spacing can enable significant weight savings in smaller landing gear lengths and structures.
• The example gearedturbofan engine20 includes a fan diameter62 (FIG. 1) and an exampledirect drive engine25 includes a fan diameter64 (FIG. 2). In one example configuration, both thefan diameter62 of the gearedturbofan engine20 and thefan diameter64 of the directdrive turbine engine25 are of a common size. Further, in this example, the fan pressure ratio and overall pressure ratio through the core are the same. When thesefan diameters62,64 and pressure ratios are the same, the gearedturbofan engine20 includes a fan drive turbine diameter66 (FIG. 1) that is much smaller than a diameter68 (FIG. 2) of the fan drive turbine for thedirect drive engine25. In one example, for a common fan diameter, the fan drive turbine is about 0.35 to about 0.45 thediameter62 of thefan42, wherein a correspondingdirect drive engine25 would include a fan drive turbine between about 0.50 and 0.65 thediameter64 of thefan42′.
• FIGS. 3-4 show thegeared turbofan engine20 in one example mount configuration. Afront mount beam70 and arear mount beam72 are used to connect the engine case74 to apylon76 that is mounted underneath a wing. One relatively important dimension, indicated at80, is the distance between abottom surface82 of the wing and anoutermost surface84 of the fan drive turbine section, that is,low pressure turbine46. For a fan diameter64 (FIG. 2) that is the same as thefan diameter62 for the gearedturbofan engine20 inFIGS. 3-4, the fan drive turbine section, that is,low pressure turbine46′, would have a comparatively greater size as indicated by anoutermost surface78 of thelow pressure turbine46′. The increased turbine size for the direct drive configuration decreases thewing clearance dimension80′ when compared to thedimension80 of the gearedturbo fan engine20.
• Thus, the significance of the difference in size of the two different fan drive turbine sections is illustrated with the required spacing of thecritical dimension80′ for a direct drive turbine indicated between theoutermost surface78, shown by the dashed lines, and thebottom surface82 of the wing. Accordingly, the size of thefan42′ for a directdrive turbine engine25 is limited by the size of the fan drive turbine, i.e. the size of thelow pressure turbine46′. As such, the gearedturbofan engine20 with the smaller more efficient fan drive turbine, i.e.low pressure turbine46, can provide a larger fan in the same space, and/or enable a fan size not possible in a direct drivegas turbine engine25.
• Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

Claims (20)

What is claimed is:

1. A gas turbine engine comprising:

a fan configured to deliver airflow to a bypass passage;

a core engine configured to rotate the fan, the core engine including:

a high pressure turbine section configured to drive a high pressure compressor section, and

a low pressure turbine section configured to drive the fan and a low pressure compressor section, the fan and the low pressure turbine section configured to rotate at a common speed and in a common direction; and

wherein the gas turbine engine has a bypass ratio of greater than about 10, the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.

2. The gas turbine engine as recited inclaim 1, wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.

3. The gas turbine engine as recited inclaim 1, wherein the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.

4. The gas turbine engine as recited inclaim 1, wherein the high pressure turbine section includes two stages.

5. The gas turbine engine as recited inclaim 1, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.

6. The gas turbine engine as recited inclaim 1, wherein the low pressure turbine section has a greater number of stages than the low pressure compressor section.

7. The gas turbine engine as recited inclaim 6, wherein the fan has a pressure ratio of less than about 1.45.

8. The gas turbine engine as recited inclaim 7, wherein the fan includes fewer than 20 fan blades.

9. The gas turbine engine as recited inclaim 1, wherein the low pressure turbine section has a pressure ratio of greater than about 5.

10. A gas turbine engine comprising:

a fan having fewer than 20 fan blades situated at an inlet of a bypass passage;

a core engine configured to rotate the fan, the core engine including:

a high pressure turbine section configured to drive a high pressure compressor section, and

a low pressure turbine section configured to drive the fan and a low pressure compressor section, the low pressure turbine section having a greater number of stages than the low pressure compressor section; and

wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.

11. The gas turbine engine as recited inclaim 10, wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.

12. The gas turbine engine as recited inclaim 10, wherein the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.

13. The gas turbine engine as recited inclaim 10, wherein the high pressure turbine section includes two stages.

14. The gas turbine engine as recited inclaim 10, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.

15. The gas turbine engine as recited inclaim 14, wherein the fan has a pressure ratio of less than about 1.5.

16. The gas turbine engine according toclaim 10, wherein the gas turbine engine has a bypass ratio of greater than about 10.

17. The gas turbine engine according toclaim 16, wherein the low pressure turbine has a pressure ratio of greater than about 5.

18. A method of designing a gas turbine engine comprising:

providing a fan configured to deliver airflow to a bypass duct;

providing a core engine configured to rotate the fan, the core engine including:

a high pressure turbine section configured to drive a high pressure compressor section, and

a low pressure turbine section configured to rotate the fan at a common speed and in a common direction;

wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan, and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65; and

wherein the fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than 10.

19. The method as recited inclaim 18, wherein the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about 5.

20. The method as recited inclaim 18, comprising a low pressure compressor section driven by the low pressure turbine section, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages.

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