US20180045120A1 - Gas turbine engine with high speed low pressure turbine section - Google Patents
Gas turbine engine with high speed low pressure turbine section Download PDFInfo
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- US20180045120A1 US20180045120A1 US15/791,825 US201715791825A US2018045120A1 US 20180045120 A1 US20180045120 A1 US 20180045120A1 US 201715791825 A US201715791825 A US 201715791825A US 2018045120 A1 US2018045120 A1 US 2018045120A1
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- section
- turbine section
- low pressure
- pressure turbine
- fan
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- 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
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
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- 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
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
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- 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
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- 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
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- 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
Definitions
- a gas turbine engine comprises 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 first turbine section and a second turbine section.
- the first turbine section has a first exit area at a first exit point and rotates at a first speed.
- the second turbine section has a second exit area at a second exit point and rotates 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 0.5 and about 1.5.
- a gear reduction is included between the fan and a low spool is driven by the first turbine section such that the fan rotates at a lower speed than the first turbine section.
- FIG. 1 shows a gas turbine engine
- FIG. 4 shows yet another embodiment.
- the core airflow C is compressed by the low pressure compressor section 44 then the high pressure compressor section 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine section 54 and low pressure turbine section 46 .
- the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
- the turbine sections 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
- the engine 20 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.
- the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
- the geared architecture 48 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 low pressure turbine section 46 has a pressure ratio that is greater than about 5.
- the engine 20 bypass ratio is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor section 44
- the low pressure turbine section 46 has a pressure ratio that is greater than about 5:1.
- 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, the fan 42 may have 26 or fewer blades.
- An exit area 400 is shown, in FIG. 1 and FIG. 2 , at the exit location for the high pressure turbine section 54 .
- An exit area for the low pressure turbine section is defined at exit 401 for the low pressure turbine section.
- the turbine engine 20 may be counter-rotating. This means that the low pressure turbine section 46 and low pressure compressor section 44 rotate in one direction, while the high pressure spool 32 , including high pressure turbine section 54 and high pressure compressor section 52 rotate in an opposed direction.
- the gear reduction 48 which may be, for example, an epicyclic transmission (e.g., with a sun, ring, and star gears), is selected such that the fan 42 rotates in the same direction as the high spool 32 .
- a lpt is the area of the low pressure turbine section at the exit thereof (e.g., at 401 ), where V lpt is the speed of the low pressure turbine section, where A hpt is the area of the high pressure turbine section at the exit thereof (e.g., at 400 ), and where V hpt is the speed of the low pressure turbine section.
- the areas of the low and high pressure turbine sections are 557.9 in 2 and 90.67 in 2 , respectively. Further, the speeds of the low and high pressure turbine sections are 10179 rpm and 24346 rpm, respectively.
- the performance quantities for the low and high pressure turbine sections are:
- FIG. 3 shows an embodiment 200 , wherein there is a fan drive turbine 208 driving a shaft 206 to in turn drive a fan rotor 202 .
- a gear reduction 204 may be positioned between the fan drive turbine 208 and the fan rotor 202 .
- This gear reduction 204 may be structured and operate like the gear reduction disclosed above.
- a compressor rotor 210 is driven by an intermediate pressure turbine 212
- a second stage compressor rotor 214 is driven by a turbine rotor 216 .
- a combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216 .
- FIG. 3 or 4 engines may be utilized with the features disclosed above.
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- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
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- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
- This application is a continuation of U.S. application Ser. No. 14/573,003, filed Dec. 17, 2014, which is a continuation-in-part of U.S. application Ser. No. 13/363,154, filed Jan. 31, 2012.
- 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.
- In a featured embodiment, a gas turbine engine comprises 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 first turbine section and a second turbine section. The first turbine section has a first exit area at a first exit point and rotates at a first speed. The second turbine section has a second exit area at a second exit point and rotates 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 0.5 and about 1.5. A gear reduction is included between the fan and a low spool is driven by the first turbine section such that the fan rotates at a lower speed than the first turbine section.
- In another embodiment according to the previous embodiment, a gear ratio of the gear reduction is greater than about 2.3.
- In another embodiment according to any of the previous embodiments, the gear ratio is greater than about 2.5.
- In another embodiment according to any of the previous embodiments, there is a third turbine section. The first turbine section drives the fan, and the second and third turbine sections each drive a compressor rotor of the compressor section.
- In another embodiment according to any of the previous embodiments, the gear reduction is positioned intermediate the fan and a compressor rotor driven by the first turbine section.
- In another embodiment according to any of the previous embodiments, the gear reduction is positioned intermediate the first turbine section and a compressor rotor driven by the first turbine section.
- These and other features may be best understood from the following drawings and specification.
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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 another embodiment. -
FIG. 4 shows yet another embodiment. -
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B while thecompressor section 24 drives air along a core flow path C for compression and communication into thecombustor section 26 then expansion through theturbine section 28. 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-spool architectures. - The
engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine 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 interconnects afan 42, a low pressure (or first)compressor section 44 and a low pressure (or first)turbine section 46. Theinner shaft 40 is connected to thefan 42 through a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 50 that interconnects a high pressure (or second)compressor section 52 and high pressure (or second)turbine section 54. Acombustor 56 is arranged between the highpressure compressor section 52 and the highpressure turbine section 54. Amid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the highpressure turbine section 54 and the lowpressure turbine section 46. Themid-turbine frame 57 further supports bearingsystems 38 in theturbine section 28. 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 afan 42. Theinner shaft 40 and theouter shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes. The high and low spools can be either co-rotating or counter-rotating. - The core airflow C is compressed by the low
pressure compressor section 44 then the highpressure compressor section 52, mixed and burned with fuel in thecombustor 56, then expanded over the highpressure turbine section 54 and lowpressure turbine section 46. Themid-turbine frame 57 includesairfoils 59 which are in the core airflow path. Theturbine sections low speed spool 30 andhigh speed spool 32 in response to the expansion. - The
engine 20 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, theengine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 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 section 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, theengine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the lowpressure compressor section 44, and the lowpressure turbine section 46 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 section 46, in some embodiments, has between 3 and 6 stages. Further, the lowpressure turbine section 46 pressure ratio is total pressure measured prior to the inlet of lowpressure turbine section 46 as related to the total pressure at the outlet of the lowpressure turbine section 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a planetary 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 and that the present invention is applicable to other gas turbine engines including direct drive turbofans. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 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 lbm of fuel being burned per hour divided by lbf 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, thefan 42 may have 26 or fewer blades. - An
exit area 400 is shown, inFIG. 1 andFIG. 2 , at the exit location for the highpressure turbine section 54. An exit area for the low pressure turbine section is defined atexit 401 for the low pressure turbine section. As shown inFIG. 2 , theturbine engine 20 may be counter-rotating. This means that the lowpressure turbine section 46 and lowpressure compressor section 44 rotate in one direction, while thehigh pressure spool 32, including highpressure turbine section 54 and highpressure compressor section 52 rotate in an opposed direction. Thegear reduction 48, which may be, for example, an epicyclic transmission (e.g., with a sun, ring, and star gears), is selected such that thefan 42 rotates in the same direction as thehigh spool 32. 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 lpt 2) Equation 1: -
PQ hpt=(A hpt ×V hpt 2) Equation 2: - where Alpt is the area of the low pressure turbine section at the exit thereof (e.g., at 401), where Vlpt is the speed of the low pressure turbine section, where Ahpt is the area of the high pressure turbine section at the exit thereof (e.g., at 400), and where Vhpt is 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 lpt 2)/(A hpt ×V hpt 2)=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 in2 and 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 lpt 2)=(557.9 in2)(10179 rpm)2=57805157673.9 in2 rpm2 Equation 1: -
PQ hpt=(A hpt ×V hpt 2)=(90.67 in2)(24346 rpm)2=53742622009.72 in2 rpm2 Equation 2: - and using Equation 3 above, the ratio for the low pressure turbine section to the high pressure turbine section is:
-
Ratio=PQ ltp /PQ hpt=57805157673.9 in2 rpm2/53742622009.72 in2 rpm2=1.075 - In another embodiment, the ratio was about 0.5 and in another embodiment the ratio was about 1.5. With PQltp/PQhpt ratios in the 0.5 to 1.5 range, a very efficient overall gas turbine engine is achieved. More narrowly, PQltp/PQhpt ratios of above or equal to about 0.8 are more efficient. Even more narrowly, PQltp/PQhpt ratios above or equal to 1.0 are even more efficient. As a result of these PQltp/PQhpt ratios, 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 work 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.
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FIG. 3 shows anembodiment 200, wherein there is afan drive turbine 208 driving ashaft 206 to in turn drive afan rotor 202. Agear reduction 204 may be positioned between thefan drive turbine 208 and thefan rotor 202. Thisgear reduction 204 may be structured and operate like the gear reduction disclosed above. Acompressor rotor 210 is driven by anintermediate pressure turbine 212, and a second stage compressor rotor 214 is driven by a turbine rotor 216. A combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216. -
FIG. 4 shows yet anotherembodiment 300 wherein afan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction 306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and ashaft 308 which is driven by a low pressure turbine section. - The
FIG. 3 or 4 engines may be utilized with the features disclosed above. - 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)
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US15/791,825 US20180045120A1 (en) | 2012-01-31 | 2017-10-24 | Gas turbine engine with high speed low pressure turbine section |
Applications Claiming Priority (3)
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US13/363,154 US20130192196A1 (en) | 2012-01-31 | 2012-01-31 | Gas turbine engine with high speed low pressure turbine section |
US14/573,003 US9816442B2 (en) | 2012-01-31 | 2014-12-17 | Gas turbine engine with high speed low pressure turbine section |
US15/791,825 US20180045120A1 (en) | 2012-01-31 | 2017-10-24 | Gas turbine engine with high speed low pressure turbine section |
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US14/573,003 Continuation US9816442B2 (en) | 2012-01-31 | 2014-12-17 | Gas turbine engine with high speed low pressure turbine section |
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US20180045120A1 true US20180045120A1 (en) | 2018-02-15 |
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US15/791,825 Abandoned US20180045120A1 (en) | 2012-01-31 | 2017-10-24 | Gas turbine engine with high speed low pressure turbine section |
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2017
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US20150096303A1 (en) | 2015-04-09 |
US9816442B2 (en) | 2017-11-14 |
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