US20130192191A1 - Gas turbine engine with high speed low pressure turbine section and bearing support features - Google Patents

Abstract

A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section.

Description

CROSS-REFERENCE TO RELATED APPLICATION
• This application is a continuation-in-part of U.S. patent application Ser. No. 13/363,154, filed on Jan. 31, 2012 and entitled “Gas Turbine Engine With High Speed Low Pressure Turbine Section.”
BACKGROUND OF THE INVENTION
• This application relates to a gas turbine engine wherein the low pressure turbine section is rotating at a higher speed and centrifugal pull stress relative to the high pressure turbine section speed and centrifugal pull stress than prior art engines.
• Gas turbine engines are known, and typically include a fan delivering air into a low pressure compressor section. The air is compressed in the low pressure compressor section, and passed into a high pressure compressor section. From the high pressure compressor section the air is introduced into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over a high pressure turbine section, and then a low pressure turbine section.
• Traditionally, on many prior art engines the low pressure turbine section has driven both the low pressure compressor section and a fan directly. As fuel consumption improves with larger fan diameters relative to core diameters it has been the trend in the industry to increase fan diameters. However, as the fan diameter is increased, high fan blade tip speeds may result in a decrease in efficiency due to compressibility effects. Accordingly, the fan speed, and thus the speed of the low pressure compressor section and low pressure turbine section (both of which historically have been coupled to the fan via the low pressure spool), have been a design constraint. More recently, gear reductions have been proposed between the low pressure spool (low pressure compressor section and low pressure turbine section) and the fan.
SUMMARY
• In a featured embodiment, a turbine section of a gas turbine engine has a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed. A first performance quantity is defined as the product of the fan drive turbine's speed squared and the fan drive turbine's exit area. A second performance quantity is defined as the product of the second speed squared and the second area. A ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5. The second turbine section drives a shaft which is mounted on a bearing on an outer periphery of the first shaft at a location upstream of a point where the first shaft connects to a hub carrying turbine rotors associated with said second turbine section.
• In another embodiment according to the previous embodiment, the ratio is above or equal to about 0.8.
• In another embodiment according to any of the previous embodiments, the fan drive turbine section has at least 3 stages.
• In another embodiment according to any of the previous embodiments, the fan drive turbine section has up to 6 stages.
• In another embodiment according to any of the previous embodiments, the second turbine section has 2 or fewer stages.
• In another embodiment according to any of the previous embodiments, a pressure ratio across the first fan drive turbine section is greater than about 5:1.
• In another embodiment according to any of the previous embodiments, a second shaft associated with the fan drive turbine is supported by a second bearing at an end of the second shaft, and downstream of the fan drive turbine.
• In another embodiment according to any of the previous embodiments, the fan drive turbine and second turbine sections are configured to rotate in opposed directions.
• In another embodiment according to any of the previous embodiments, there is no mid-turbine frame positioned intermediate the fan drive turbine and second turbine sections.
• In another featured embodiment, a gas turbine engine has a fan, a compressor section in fluid communication with the fan, a combustion section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustion section. The turbine section includes a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is higher than the first speed. A first performance quantity is defined as the product of the fan drive turbine's speed squared and the fan drive turbine's area. A second performance quantity is defined as the product of the second turbine's speed squared and the second turbine's area. A ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5. The second turbine section drives a shaft which is mounted on a bearing on an outer periphery of the first shaft at a location upstream of a point where the first shaft connects to a hub carrying turbine rotors associated with said second turbine section.
• In another embodiment according to the previous embodiment, the ratio is above or equal to about 0.8.
• In another embodiment according to any of the previous embodiments, the compressor section includes a first and second compressor sections. The fan drive turbine section and the first compressor section are configured to rotate in a first direction. The second turbine section and the second compressor section and are configured to rotate in a second opposed direction.
• In another embodiment according to any of the previous embodiments, a gear reduction is included between the fan and a low spool driven by the fan drive turbine section such that the fan is configured to rotate at a lower speed than the fan drive turbine section.
• In another embodiment according to any of the previous embodiments, the fan rotates in the second opposed direction.
• In another embodiment according to any of the previous embodiments, a second shaft associated with the fan drive turbine is supported by a second bearing at an end of the second shaft, and downstream of the fan drive turbine.
• In another embodiment according to any of the previous embodiments, a third bearing supports the second compressor section on an outer periphery of the first shaft driven by the second turbine section.
• In another embodiment according to any of the previous embodiments, a fourth bearing is positioned adjacent the first compressor section, and supports an outer periphery of the second shaft which is configured to rotate with the fan drive turbine section.
• In another embodiment according to any of the previous embodiments, there is no mid-turbine frame positioned intermediate the first and second turbine sections.
• In another featured embodiment, a gas turbine engine has a fan, a compressor section in fluid communication with the fan, a combustion section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustion section. The turbine section includes a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. A second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is higher than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area. A ratio of the first performance quantity to the second performance quantity is between about 0.8 and about 1.5. The compressor section includes first and second compressor sections. The fan drive turbine section and the first compressor section will rotate in a first direction and the second turbine section and the second compressor section will rotate in a second opposed direction. A gear reduction is included between the fan and first compressor section, such that the fan will rotate at a lower speed than the fan drive turbine section, and rotate in the second opposed direction.
• In another embodiment according to the previous embodiment, a gear ratio of the gear reduction is greater than about 2.3.
• These and other features of this disclosure will be better understood upon reading the following specification and drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
• FIG. 1 shows a gas turbine engine.
• FIG. 2 schematically shows the arrangement of the low and high spool, along with the fan drive.
• FIG. 3 shows a schematic view of a mount arrangement for an engine such as shown inFIGS. 1 and 2.
DETAILED DESCRIPTION
• FIG. 1 schematically illustrates agas turbine engine20. Thegas turbine engine20 is disclosed herein as a two-turbine turbofan that generally incorporates afan section22, acompressor section24, acombustor section26 and aturbine section28. Alternative engines might include an augmentor section (not shown) among other systems or features. Thefan section22 drives air along a bypass flow path B while thecompressor section24 drives air along a core flow path C for compression and communication into thecombustor section26 then expansion through theturbine section28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, 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 including three-turbine architectures.
• Theengine20 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 aninnermost shaft40 that interconnects afan42, a low pressure (or first)compressor section44 and a low pressure (or first)turbine section46. Noteturbine section46 will also be known as a fan drive turbine section. Theinner shaft40 is connected to thefan42 through a gearedarchitecture48 to drive thefan42 at a lower speed than the low speedfan drive turbine46. Thehigh speed spool32 includes a moreouter shaft50 that interconnects a high pressure (or second)compressor section52 and high pressure (or second)turbine section54. Acombustor56 is arranged between the highpressure compressor section52 and the highpressure turbine section54. As used herein, the high pressure turbine section experiences higher pressures than the low pressure turbine section. A low pressure turbine section is a section that powers afan42. Theinner shaft40 and theouter shaft50 are concentric and rotate via bearingsystems38 about the engine central longitudinal axis A which is collinear with their longitudinal axis.
• The core airflow C is compressed by the lowpressure compressor section44 then the highpressure compressor section52, mixed and burned with fuel in thecombustor56, then expanded over the highpressure turbine section54 and lowpressure turbine section46.
• Theengine20 in one example is a high-bypass geared aircraft engine. The bypass ratio is the amount of air delivered into bypass path B divided by the amount of air into core path C. In a further example, theengine20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the gearedarchitecture48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the lowpressure turbine section46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, theengine20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the lowpressure compressor section44, and the lowpressure turbine section46 has a pressure ratio that is greater than about 5:1. In some embodiments, the high pressure turbine section may have two or fewer stages. In contrast, the lowpressure turbine section46, in some embodiments, has between 3 and 6 stages. Further the lowpressure turbine section46 pressure ratio is total pressure measured prior to inlet of lowpressure turbine section46 as related to the total pressure at the outlet of the lowpressure turbine section46 prior to an exhaust nozzle. The gearedarchitecture48 may be an epicycle gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine
• 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 (“TSFC”). TSFC is the industry standard parameter of the rate of 1 bm of fuel being burned per hour divided by 1 bf of thrust the engine produces at that flight condition. “Low fan pressure ratio” is the ratio of total pressure across the fan blade alone, before the fan exit guide vanes. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment 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 [(Ram Air Temperature deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. Further, thefan42 may have26 or fewer blades.
• Anexit area400 is shown, inFIG. 1 andFIG. 2, at the exit location for the highpressure turbine section54 is the annular area of the last blade ofturbine section54. An exit area for the low pressure turbine section is defined atexit401 for the low pressure turbine section is the annular area defined by the last blade of thatturbine section46. As shown inFIG. 2, theturbine engine20 may be counter-rotating. This means that the lowpressure turbine section46 and lowpressure compressor section44 rotate in one direction (“−”), while thehigh pressure spool32, including highpressure turbine section54 and highpressure compressor section52 rotate in an opposed direction (“+”). Thegear reduction48, which may be, for example, an epicyclic transmission (e.g., with a sun, ring, and star gears), is selected such that thefan42 rotates in the same direction (“+”) as thehigh spool32. With this arrangement, and with the other structure as set forth above, including the various quantities and operational ranges, a very high speed can be provided to the low pressure spool. Low pressure turbine section and high pressure turbine section operation are often evaluated looking at a performance quantity which is the exit area for the turbine section multiplied by its respective speed squared. This performance quantity (“PQ”) is defined as:
• 
  PQ_ltp=(A_lpt×V_lpt²)  Equation 1
• 
  PQ_hpt=(A_hpt×V_hpt²)  Equation 2
• where A_lptis the area of the low pressure turbine section at the exit thereof (e.g., at401), where V_lptis the speed of the low pressure turbine section, where A_hptis the area of the high pressure turbine section at the exit thereof (e.g., at400), and where V_hptis the speed of the low pressure turbine section.
• Thus, a ratio of the performance quantity for the low pressure turbine section compared to the performance quantify for the high pressure turbine section is:
• 
  (A_lpt×V_lpt²)/(A_hpt×V_hpt²)=PQ_ltp/PQ_hpt  Equation 3
• In one turbine embodiment made according to the above design, the areas of the low and high pressure turbine sections are 557.9 in²and 90.67 in², respectively. Further, the speeds of the low and high pressure turbine sections are 10179 rpm and 24346 rpm, respectively. Thus, using Equations 1 and 2 above, the performance quantities for the low and high pressure turbine sections are:
• 
  PQ_ltp=(A_lpt×V_lpt²)=(557.9 in²)(10179 rpm)²=57805157673.9 in²rpm²  Equation 1
• 
  PQ_hpt=(A_hpt×V_hpt²)=(90.67 in²)(24346 rpm)²=53742622009.72 in²rpm²  Equation 2
• and using Equation3 above, the ratio for the low pressure turbine section to the high pressure turbine section is:
• 
  Ratio=PQ_ltp/PQ_hpt=57805157673.9 in²rpm²/53742622009.72 in²rpm²=1.075
• In another embodiment, the ratio was about 0.5 and in another embodiment the ratio was about 1.5. With PQ_ltp/PQ_hptratios in the 0.5 to 1.5 range, a very efficient overall gas turbine engine is achieved. More narrowly, PQ_ltp/PQ_hptratios of above or equal to about 0.8 are more efficient. Even more narrowly, PQ_ltp/PQ_hptratios above or equal to 1.0 are even more efficient. As a result of these PQ_ltp/PQ_hptratios, in particular, the turbine section can be made much smaller than in the prior art, both in diameter and axial length. In addition, the efficiency of the overall engine is greatly increased.
• The low pressure compressor section is also improved with this arrangement, and behaves more like a high pressure compressor section than a traditional low pressure compressor section. It is more efficient than the prior art, and can provide more compression in fewer stages. The low pressure compressor section may be made smaller in radius and shorter in length while contributing more toward achieving the overall pressure ratio design target of the engine.
• As shown inFIG. 3, the engine as shown inFIG. 2 may be mounted such that thehigh pressure turbine54 is “overhung” bearing mounted. As shown, the high spool andshaft32 includes abearing142 which supports thehigh pressure turbine54 and thehigh spool32 on an outer periphery of a shaft that rotates with thehigh pressure turbine54. As can be appreciated, the “overhung” mount means that thebearing142 is at an intermediate location on the spool including the shaft, thehigh pressure turbine54, and thehigh pressure compressor52. Stated another way, thebearing142 is supported upstream of apoint501 where theshaft32 connects to ahub500 carrying turbine rotors associated with the high pressure turbine (second)turbine section54. Notably, it would also be downstream of thecombustor56. Note that the bearing142 can be positioned inside anannulus503 formed by theshaft32 and thehub assembly500 so as to be between the shaft and the feature numbered106 and it still would be an “overhung” configuration.
• The forward end of thehigh spool32 is supported by abearing110 at an outer periphery of theshaft32. Thebearings110 and142 are supported onstatic structure108 associated with the overall engine casings arranged to form the core of the engine as is shown inFIG. 1. In addition, theshaft30 is supported on abearing100 at a forward end. Thebearing100 is supported onstatic structure102. A rear end of theshaft30 is supported on abearing106 which is attached tostatic structure104.
• With this arrangement, there is no bearing support struts or other structure in the path of hot products of combustion passing downstream of thehigh pressure turbine54, and no bearing compartment support struts in the path of the products of combustion as they flow across to thelow pressure turbine46.
• As shown, there is no mid-turbine frame or bearings mounted in thearea402 between theturbine sections54 and46.
• While this invention has been disclosed with reference to one embodiment, it should be understood that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims (20)

1. A turbine section of a gas turbine engine comprising:

a fan drive turbine section; and

a second turbine section,

wherein said fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed,

wherein said second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed,

wherein a first performance quantity is defined as the product of the first speed squared and the first area,

wherein a second performance quantity is defined as the product of the second speed squared and the second area;

wherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5; and

said second turbine section driving a first shaft , and said first shaft being supported on a bearing , said bearing being mounted on an outer periphery of said first shaft at a location that is upstream of point where the first shaft connects to a hub carrying turbine rotors associated with said second turbine section.

2. The turbine section as set forth inclaim 1, wherein said ratio is above or equal to about 0.8.

3. The turbine section as set forth inclaim 1, wherein said fan drive turbine section has at least three stages.

4. The turbine section as set forth inclaim 1, wherein said fan drive turbine section has up to six stages.

5. The turbine section as set forth inclaim 1, wherein said second turbine section has two or fewer stages.

6. The turbine section as set forth inclaim 1, wherein a pressure ratio across the fan drive turbine section is greater than about 5:1.

7. The turbine section as set forth inclaim 1, wherein a second shaft associated with said fan drive turbine is supported by a second bearing at an end of said second shaft, and downstream of said fan drive turbine.

8. The turbine section as set forth inclaim 1, wherein said fan drive and second turbine sections are configured to rotate in opposed directions.

9. The turbine section as set forth inclaim 1, wherein there is no bearing support structure positioned intermediate said fan drive and second turbine sections.

10. A gas turbine engine comprising:

a fan;

a compressor section in fluid communication with the fan;

a combustion section in fluid communication with the compressor section;

a turbine section in fluid communication with the combustion section,

wherein the turbine section includes a fan drive turbine section and a second turbine section,

wherein said fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed,

wherein said second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is higher than the first speed,

wherein a first performance quantity is defined as the product of the first speed squared and the first area,

wherein a second performance quantity is defined as the product of the second speed squared and the second area;

wherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5;

said second turbine section driving a first shaft, and said first shaft being mounted on a bearing, said bearing being mounted on an outer periphery of said first shaft at a location upstream of a point where the first shaft connects to a hub carrying turbine rotors associated with said second turbine section.

11. The engine as set forth inclaim 10, wherein said ratio is above or equal to about 0.8.

12. The engine as set forth inclaim 10, wherein the compressor section includes a first compressor section and a second compressor section, wherein the fan drive turbine section and the first compressor section are configured to rotate in a first direction, and wherein the second turbine section and the second compressor section are configured to rotate in a second opposed direction.

13. The engine as set forth inclaim 12, wherein a gear reduction is included between said fan and said second shaft driven by the fan drive turbine section such that the fan rotates at a lower speed than the fan drive turbine section.

14. The engine as set forth inclaim 13, wherein said fan rotates in the second opposed direction.

15. The engine as set forth inclaim 10, wherein a second shaft associated with said fan drive turbine is supported by a second bearing at an end of said second shaft, and downstream of said fan drive turbine.

16. The engine as set forth inclaim 15, wherein a third bearing supports said second compressor section on an outer periphery of said first shaft driven by said second turbine section.

17. The engine as set forth inclaim 16, wherein a fourth bearing is positioned adjacent said first compressor section, and supports an outer periphery of said second shaft which is configured to rotate with said fan drive turbine section.

18. The engine as set forth inclaim 12, wherein there is no bearing support structure positioned intermediate said first and second turbine sections.

19. A gas turbine engine comprising:

a fan;

a compressor section in fluid communication with the fan;

a combustion section in fluid communication with the compressor section;

a turbine section in fluid communication with the combustion section,

wherein the turbine section includes a fan drive turbine section and a second turbine section,

wherein said fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed,

wherein said second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is higher than the first speed,

wherein a first performance quantity is defined as the product of the first speed squared and the first area,

wherein a second performance quantity is defined as the product of the second speed squared and the second area;

wherein a ratio of the first performance quantity to the second performance quantity is between about 0.8 and about 1.5; and

the compressor section including a first compressor section and a second compressor section, wherein the fan drive turbine section and the first compressor section will rotate in a first direction and the second turbine section and the second compressor section will rotate in a second opposed direction, a gear reduction included between said fan and first compressor section, such that the fan will rotate at a lower speed than the fan drive turbine section, and said fan will rotate in the second opposed direction.

20. The engine as set forth inclaim 19, wherein a gear ratio of said gear reduction is greater than about 2.3.

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