• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Turbine section of a gas turbine engine and gas turbine engine. A gas turbine engine includes a very high speed low pressure turbine such that an amount defined by the outlet area of the low pressure turbine multiplied by the square of the rotational speed of the low pressure turbine compared to the same parameters for a gas turbine highest pressure is in a ratio between about 0.5 and about 1.5. further, the lower pressure turbine is mounted with a first bearing mounted in a medium turbine frame, and a second bearing mounted within a turbine exhaust box.
    公开号:BR112015010811B1
    申请号:R112015010811-3
    申请日:2013-11-07
    公开日:2021-08-31
    发明作者:Frederick M. Schwarz；Jorn A. Glahn
    申请人:United Technologies Corporation；
    IPC主号:
    专利说明:

    FUNDAMENTALS OF THE INVENTION
    [001] This application relates to a gas turbine engine in which the low pressure turbine section is rotating at a high speed and centrifugal traction voltage with respect to the speed of the high pressure turbine section and centrifugal traction voltage of the than prior art engines.
    [002] Gas turbine engines are known, and typically include a fan that distributes air to a low pressure compressor section. Air is compressed in the low pressure compressor section, and passed to a high pressure compressor section. From the high pressure compressor section air is introduced to a combustion section where it is mixed with fuel and combusted. The products of this combustion pass downstream over a high-pressure turbine section, and then a low-pressure turbine section.
    [003] Traditionally, in many prior art engines the low pressure turbine section drives 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 a trend in the industry to increase fan diameters. However, when the fan diameter is increased, high fan blade tip speeds can result in a decrease in efficiency due to compressibility effects. Appropriately, the fan speed, and thus the speed of the low pressure compressor section and low pressure turbine section (both of which historically were coupled with the fan via the low pressure coil), were a design constraint. . More recently, gear reductions have been proposed between the low pressure coil (low pressure compressor section and low pressure turbine section) and the fan. SUMMARY
    [004] In a presented embodiment, a turbine section of a gas turbine engine has a first and a second turbine section. The first turbine section has a first exit area and rotates at a first speed. The second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed. A first performance quantity is defined as the product of the first velocity squared and the first area. A second performance quantity is defined as the product of the second velocity squared and the second area. A ratio of the first amount of performance to the second amount of performance is between about 0.5 and about 1.5. The first turbine section is supported on two bearings, with a first bearing mounted on a medium turbine frame that is positioned intermediate the first turbine section and the second turbine section, and a second bearing mounting the first turbine section, with the second bearing having a support extending downstream of the first turbine section.
    [005] In another modality according to the previous modality, the ratio is above or equal to about 0.8.
    [006] In another modality according to any of the previous modality, the first turbine section has at least three stages.
    [007] In another modality according to any of the previous modality, the first turbine section has up to six stages.
    [008] In another modality according to any of the previous modality, the second turbine section has two or less stages.
    [009] In another modality according to any of the previous modality, a pressure ratio across the first turbine section is greater than about 5:1.
    [0010] In another presented embodiment, a gas turbine engine has a fan, and a compressor section in fluid communication with the fan. A combustion section is in fluid communication with the compressor section. A turbine section is 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 greater than the first speed. A first performance quantity is defined as the product of the first velocity squared and the first area. A second performance quantity is defined as the product of the second velocity squared and the second area. A ratio of the first amount of performance to the second amount of performance is between about 0.5 and about 1.5. The first turbine section is supported on two bearings, with a first bearing mounted on a medium turbine frame that is positioned intermediate the first turbine section and the second turbine section, and a second bearing mounting the first turbine section, with the second bearing supported in an exhaust box downstream of the first turbine section.
    [0011] In another modality according to the previous modality, the ratio is above or equal to about 0.8.
    [0012] In another mode according to any of the previous embodiments, the compressor section includes a first compressor section and a second compressor section. The first turbine section and the first compressor section rotate in a first direction. The second turbine section and the second compressor section rotate in an opposite second direction.
    [0013] In another embodiment according to any of the previous embodiments, a gear reduction is included between the fan and a low coil driven by the first turbine section such that the fan rotates at a slower speed than the first turbine section .
    [0014] In another mode according to any of the previous modes, the fan rotates in the second opposite direction.
    [0015] In another embodiment according to any of the above embodiments, a gear ratio of the gear reduction is greater than about 2.3.
    [0016] In another embodiment according to any of the above embodiments, the gear ratio is greater than about 2.5.
    [0017] In another modality according to any of the previous modality, the ratio is above or equal to about 1.0.
    [0018] In another modality according to any of the previous modality, the fan distributes a portion of air to a diversion duct. A bypass ratio is defined as the portion of air delivered to the bypass duct divided by the amount of air delivered to the compressor section, with the bypass ratio being greater than about 6.0.
    [0019] In another modality according to any of the previous modalities, the deviation ratio is greater than about 10.0.
    [0020] In another modality, according to any of the previous modality, the fan has 26 or less blades.
    [0021] In another modality according to any of the previous modality, the first turbine section has at least three stages.
    [0022] In another modality according to any of the previous modality, the first turbine section has up to six stages.
    [0023] In another embodiment according to any of the previous embodiments, a pressure ratio across the first turbine section is greater than about 5:1.
    [0024] These and other features can be better understood from the following specification and drawings, the sequence being a brief description. BRIEF DESCRIPTION OF THE DRAWINGS
    [0025] Figure 1 shows a gas turbine engine.
    [0026] Figure 2 schematically shows the arrangement of the low and high coil, along with the fan drive.
    [0027] Figure 3 shows an assembly functionality. DETAILED DESCRIPTION
    [0028] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed here as a double coil turbofan which generally incorporates a fan section 22, a compressor section 24, a section of combustor 26 and a turbine section 28. Alternative engines shall include an incrementer section (not shown) among other systems or functionality. The fan section 22 drives air along a bypass flow path B while the compressor section 24 drives air along the core flow path C for compression and communication with the combustor section 26 then expansion through the turbine 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting modality, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings can be applied to other types of turbine engines including three-coil architectures.
    [0029] Motor 20 generally includes a low-speed coil 30 and a high-speed coil 32 mounted for rotation about a central longitudinal axis of motor A with respect to a static structure of motor 36 through various systems of bearing 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
    [0030] The low speed coil 30 generally includes an inner shaft 40 which interconnects a fan 42, a low pressure (or first) compressor section 44 and a low pressure (or first) turbine section 46. The shaft internal 40 is connected with fan 42 through a geared architecture 48 to drive fan 42 at a slower speed than low speed coil 30. High speed coil 32 includes an external shaft 50 which interconnects a compressor section of high pressure (or second) 52 and a high pressure (or second) turbine section 54. A combustor 56 is arranged between the high pressure compressor section 52 and the high pressure turbine section 54. A medium turbine frame 57 of the static structure of the engine 36 is generally arranged between the high pressure turbine section 54 and the low pressure turbine section 46. The medium turbine frame 57 additionally supports bearing systems 38 in the turbine section 28. As used on here, the high pressure turbine section experiences higher pressures than the low pressure turbine section. A low pressure turbine section is a section that powers a fan 42. The inner shaft 40 and outer shaft 50 are concentric and rotate through bearing systems 38 around the central longitudinal axis of motor A which is collinear with theirs. longitudinal axes. High and low coils can be either co-rotating or counter-rotating.
    [0031] The core air flow C is comprised of the low pressure compressor section 44 then the high pressure compressor section 52, mixed and combusted with fuel in the combustor 56, then expanded over the high pressure turbine section 54 and the low pressure turbine section 46. The middle turbine frame 57 includes airfoils 59 that are in the core airflow path. Turbine sections 46, 54 rotationally drive respective low speed coil 30 and high speed coil 32 in response to expansion.
    [0032] Engine 20 in one example is a high-shift geared aircraft engine. The bypass ratio is the amount of air distributed to bypass path B divided by the amount of air to core path C. In a further example, the bypass ratio of motor 20 is greater than about six (6) , with an example modality being greater than ten (10), the 48 geared architecture is an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of more than that about 2.3 and the low pressure turbine section 46 has a pressure ratio that is greater than about 5. In a disclosed embodiment, the drift ratio of the engine 20 is greater than about ten (10 :1), the fan diameter is significantly larger than that of the low pressure compressor section 44, and the low pressure 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 low pressure turbine section 46, in some embodiments, has between 3 and 6 stages. Additionally the low pressure turbine section 46 pressure ratio is the total pressure measured before the inlet of the low pressure turbine section 46 as it is related to the total pressure at the outlet of the low pressure turbine section 46 before a nozzle. exhaustion. The 48 geared architecture can be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of more 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 can be applied to other gas turbine engines including direct drive turbofans.
    [0033] A significant amount of impulse is provided by the deflection flow B due to the high deflection ratio. The fan section 22 of the engine 20 is designed for a particular flying condition - typically cruising at about 0.8 Mach and about 35,000 feet. The 0.8 Mach, 35,000 ft flight condition, with the engine at its best fuel consumption - also known as “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 the lbf of thrust the engine produces in that flight condition. “Low fan pressure ratio” is the ratio of the total pressure by the fan blade alone, before the fan leaves the guide fins. The low fan pressure ratio as disclosed herein in accordance with a non-limiting embodiment is less than about 1.45. “Low Corrected Fan Tip Speed” is the actual fan tip speed in feet/second divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)]0.5. The "low corrected fan tip speed" as disclosed herein in accordance with a non-limiting embodiment is less than about 1150 ft/second. Additionally, fan 42 can have 26 or fewer blades.
    [0034] An exit area 400 is shown, in Figure 1 and Figure 2, at the exit location for the high pressure turbine section 54. An exit area for the low pressure turbine section is defined at the exit 401 for low pressure turbine section. As shown in Figure 2, turbine engine 20 can be counter-rotating. This means that the low pressure turbine section 46 and the low pressure compressor section 44 rotate in one direction, while the high pressure coil 32 including the high pressure turbine section 54 and the high pressure compressor section pressure 52 rotate in an opposite direction. The reduction gear 48, which can be, for example, an epicycle transmission (eg, with a sun, ring, and star gears), is selected such that fan 42 rotates in the same direction as high coil 32 With this arrangement, and with the other structure as defined above, including the various quantities and operating ranges, a very high speed can be provided for the low pressure coil. The low pressure turbine section and high pressure turbine section operation are generally evaluated by looking at the amount of performance that is the output area for the turbine section multiplied by its respective speed squared. This amount of performance (“PQ”) is defined as: Equation 1: PQltp = (Alpt x Vlpt) Equation 2: PQhpt = (Ahpt x Vhpt2) where Alpt is the area of the low pressure turbine section at its outlet ( for example, 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 its output (eg at 400), and where Vhpt is the speed of the low pressure turbine section.
    [0035] Thus, a ratio of the amount of performance for the low pressure turbine section compared to the quantified performance for the high pressure turbine section is: Equation 3: (Alpt x Vlpt2)/(Ahpt x Vhpt2) = PQltp / PQhpt
    [0036] In a turbine embodiment made in accordance with the above design, the areas of the low pressure and high pressure turbine sections are 557.9 square inches and 90.67 square inches, respectively. Additionally, the speeds of the low pressure and high pressure turbine sections are 10179 rpm and 24346 rpm, respectively. Thus, using Equations 1 and 2 above, the performance that quantifies the low-pressure and high-pressure turbine sections is: Equation 1: PQltp = (Alpt x Vlpt2) = (557.9 in2)(10179 rpm)2= 57805157673.9 in2rpm2Equation 2: PQhpt = (Ahpt x Vhpt2) = (90.67 in2)(24346rpm)2 = 53742622009.72 in2 rpm2e using Equation 3 above, the ratio for the low pressure turbine section to the turbine section of high pressure is: Ratio = PQltp/ PQhpt = 57805157673.9 in2 rpm2 / 53742622009.72 in2rpm2 = 1.075
    [0037] In another modality the ratio was about 0.5 and in another modality the ratio was about 1.5. With PQltp/PQhpt ratios in the range of 0.5 to 1.5, a very efficient global gas turbine engine is achieved. More narrowly, PQltp/PQhpt ratios of more than 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 overall engine efficiency is greatly increased.
    [0038] 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 technique, and can provide more work in fewer stages. The low pressure compressor section can be made smaller in radius and shorter in size while contributing more to achieving the engine's overall pressure ratio design target.
    [0039] Figure 3 shows a mounting arrangement for an engine having the functionalities as defined here above. With developments for turbine sections, the shaft to drive the fan from the lowest pressure turbine has become increasingly thinner and longer.
    [0040] In Figure 3, a higher pressure shaft 102 is shown supported by a front bearing 116 mounted somehow at 114. The shaft 102 is also mounted by a bearing 104 at a downstream location, and preferably through a frame of medium turbine 100. The medium turbine frame 100 is shown to be intermediate to the downstream end of the high pressure turbine 54, and an upstream end of the low pressure turbine 46. The medium turbine frame 100 also mounts a bearing 108 which supports the lower pressure or fan drive shaft 106. A downstream end of the fan drive shaft 106 is supported in a bearing 112 which is mounted 113 within a turbine exhaust housing 110. bearing 113 extends downstream of the low pressure turbine 46.
    [0041] In known engines, the fan drive shaft 106 is supported by two bearings mounted inside the turbine exhaust box. With such an arrangement, a hub 115 limits the distance along the axis of the shaft 106 that the two bearings can be spaced apart. In the prior art, these bearings must withstand a critical drive speed of the low pressure fan or turbine section 46. The two bearings are not always spaced sufficiently far apart with such an assembly.
    [0042] By mounting the bearing 108 in the middle turbine frame, and the bearing 112 downstream in the turbine exhaust box, a much wider "wheel base" is provided between the two bearings 108 and 112, and there is much support. better to withstand critical speed issues.
    [0043] While Figure 3 shows this arrangement in an engine having two turbine sections, the functionalities can apply equally to an engine having three turbine sections. In a three-section turbine engine, the middle turbine frame can be between an intermediate turbine and the lower pressure turbine. Additionally, in such an engine, the area and speed ratios as described above can also be true with respect to the intermediate turbine section and the lower pressure turbine section.
    [0044] Although an embodiment has been disclosed, one skilled in the art may recognize that certain modifications may be within the scope of this invention. For this reason, the following claims should be studied to determine the actual scope and content of this invention.
    权利要求:
    Claims (15)
    [0001]
    1. Turbine section (28) of a gas turbine engine (20), comprising: a first turbine section (46); and a second turbine section (54), wherein the first turbine section (46) has a first output area (401) and is configured to rotate at a first speed, the second turbine section (54) has a second area. output (401) and is configured to spin at a second speed, which is faster than the first speed, a first performance amount is defined as the product of the first speed squared and the first area (401), a second amount of performance is defined as the product of the second speed squared and the second area (400), and the first turbine section (46) is supported on two bearings (108, 112), with a first bearing (108) mounted on a middle turbine frame (57) which is positioned intermediate the first turbine section (46) and the second turbine section (54), and a second bearing (112) mounting the first turbine section (46) with the second bearing (112) having a support extending downstream of the first turbin section a (46), characterized by the fact that a ratio of the first amount of performance to the second amount of performance is equal to 0.8 or between 0.8 and 1.5.
    [0002]
    2. Turbine section (28), according to claim 1, characterized in that the first turbine section (46) has at least 3 stages.
    [0003]
    3. Turbine section (28), according to any one of claims 1 or 2, characterized in that the first turbine section (46) has up to 6 stages.
    [0004]
    4. Turbine section (28), according to any one of claims 1 to 3, characterized in that the second turbine section (54) has 2 or fewer stages.
    [0005]
    5. Turbine section (28) according to any one of claims 1 to 4, characterized in that a pressure ratio across the first turbine section (46) is greater than 5:1.
    [0006]
    6. Gas turbine engine (20), characterized in that it comprises: a fan (42); a compressor section (24) in fluid communication with the fan (42); a combustion section (26) in communication fluidics with the compressor section (24); and a turbine section (28) as defined in any one of claims 1 to 5, in fluid communication with the combustion section (26), wherein the first exit area (401) is at a first exit point and the second exit area (401) is at a second exit point.
    [0007]
    7. Engine (20) according to claim 6, characterized in that the compressor section (24) includes a first compressor section (24) and a second compressor section (24), wherein the first section the turbine (46) and the first compressor section (24) rotate in a first direction, and the second turbine section (54) and the second compressor section (24) rotate in an opposite second direction.
    [0008]
    8. Motor (20) according to claim 7, characterized in that a gear reduction (48) is included between the fan (42) and a low coil driven by the first turbine section (46) such that the fan (42) rotates at a slower speed than the first turbine section (46).
    [0009]
    9. Motor according to any one of claims 7 or 8, characterized in that the fan (42) rotates in the second opposite direction.
    [0010]
    10. Motor (20) according to any one of claims 8 or 9, characterized in that a gear ratio of the gear reduction (48) is greater than 2.3.
    [0011]
    11. Motor (20) according to any one of claims 8 or 9, characterized in that the gear ratio of the gear reduction (48) is greater than 2.5.
    [0012]
    12. Engine (20), according to any one of claims 6 to 11, characterized in that the ratio of the first amount of performance to the second amount of performance is above or equal to 1.0.
    [0013]
    13. Motor (20) according to any one of claims 6 to 12, characterized in that the fan (42) distributes a portion of air to a diversion duct, and a diversion ratio, being defined as the portion of air distributed to the bypass duct divided by the amount of air distributed to the compressor section (24), is greater than 6.0.
    [0014]
    14. Motor (20), according to claim 13, characterized by the fact that the deviation ratio is greater than 10.0.
    [0015]
    15. Motor (20), according to any one of claims 6 to 14, characterized in that the fan (42) has 26 or fewer blades.
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    CA2945264A1|2015-11-05|2017-05-05|United Technologies Corporation|Gas turbine engine with mount for low pressure turbine section|
    US9897001B2|2014-03-04|2018-02-20|United Technologies Corporation|Compressor areas for high overall pressure ratio gas turbine engine|
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    法律状态:
    2018-11-21| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2020-03-24| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|
    2021-06-22| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2021-08-31| B16A| Patent or certificate of addition of invention granted [chapter 16.1 patent gazette]|Free format text: PRAZO DE VALIDADE: 20 (VINTE) ANOS CONTADOS A PARTIR DE 07/11/2013, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    US201261726211P| true| 2012-11-14|2012-11-14|
    US61/726,211|2012-11-14|
    US13/719,620|2012-12-19|
    US13/719,620|US20140130479A1|2012-11-14|2012-12-19|Gas Turbine Engine With Mount for Low Pressure Turbine Section|
    PCT/US2013/068838|WO2014078157A1|2012-11-14|2013-11-07|Gas turbine engine with mount for low pressure turbine section|
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