• 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:
    GAS TURBINE ENGINE A gas turbine engine is described which typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reducing device such as an epicyclic gear assembly can be used to drive the fan section so that the fan section can rotate at a different speed than the turbine section in order 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 epicyclic gear assembly which drives the fan section at a different speed than the turbine section such that both the turbine section and the turbine section fan speeds can rotate at speeds closer to optimal, providing greater performance and performance attributes through desirable combinations of the described features of the various components of the described and described gas turbine engine.
    公开号:BR112014016273B1
    申请号:R112014016273-5
    申请日:2013-01-30
    公开日:2022-02-01
    发明作者:Daniel Bernard Kupratis；Frederick M. Schwarz
    申请人:United Technologies Corporation；
    IPC主号:
    专利说明:

    CROSS REFERENCE TO RELATED ORDER
    [0001] This application is a continuation in part of US patent application no. 13/363,154 filed January 31, 2012 and which claims priority for United States provisional patent application no. 61/653,780 filed on May 31, 2012. FUNDAMENTALS OF THE INVENTION
    [0002] 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 dispensed into the combustion section where it is mixed with fuel and ignited to generate a high velocity exhaust gas stream. The high-velocity exhaust gas flow expands through the turbine section to drive the compressor and fan section. The compressor section typically includes low- and high-pressure compressors, and the turbine section includes both low- and high-pressure turbines.
    [0003] The high pressure turbine drives the high pressure compressor through an outer shaft to form a high coil, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low coil. The inner shaft can also drive the fan section. A direct drive gas turbine engine includes a fan section driven by the inner shaft such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.
    [0004] A speed reducing device such as an epicyclic gear assembly can be used to drive the fan section so that the fan section can rotate at a different speed than the turbine section in order 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 epicyclic gear assembly which drives the fan section at a different speed than the turbine section such that both the turbine section and the turbine section fan can rotate at speeds closer to ideal.
    [0005] Although geared architectures have higher propulsive efficiency, turbine engine manufacturers continue to look for further improvements to engine performance including improvements in heat transfer and propulsive efficiencies. SUMMARY
    [0006] A gas turbine engine according to an exemplary embodiment of this description, among other possible things, includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan driving turbine and a second turbine. The fan driving turbine includes a plurality of turbine rotors. A fan includes a plurality of blades rotating about an axis and a ratio of the number of fan blades to the number of turbine rotors driving the fan is between about 2.5 and about 8.5. A speed change system is driven by the fan drive turbine to rotate the fan about the geometry axis. The fan drive turbine includes a first tail rotor attached to a first shaft. The second turbine includes a second rear rotor attached to a second shaft. A first bearing assembly is arranged axially behind a first connection between the first tail rotor and the first shaft. A second bearing assembly is arranged axially forward of a second connection between the second tail rotor and the second shaft.
    [0007] In an additional embodiment of the exposed motor, the first bearing assembly and the second bearing assembly include roller bearings.
    [0008] In an additional embodiment of any of the engines shown, the compressor section includes a first compressor driven by the fan-driven turbine through the first shaft and a second compressor section driven by the second turbine via the second shaft. The first bearing supports a rear portion of the first axle and the second bearing supports a rear portion of the second axle.
    [0009] In an additional embodiment of any of the exposed engines, a front portion of each of the first and second shafts is supported by a thrust bearing assembly.
    [00010] In an additional embodiment of any of the engines shown, the fan drive turbine has a first outlet area at a first outlet point and rotates at a first speed. The second turbine section has a second exit area at a second exit point and runs at a second speed, which is greater than the first speed. A first performance value is defined as the product of the first velocity squared by the first area. A second performance value is defined as the product of the second velocity squared by the second area. A performance ratio of the first performance value to the second performance value is between about 0.5 and about 1.5.
    [00011] In an additional embodiment of any of the exposed engines, the performance ratio is greater than or equal to about 0.8.
    [00012] In an additional embodiment of any of the exposed engines, the first performance value is greater than or equal to about 4.
    [00013] In an additional embodiment of any of the motors shown, the speed change system includes a gearbox. The fan and fan drive turbine both rotate in a first direction around the axis. The second turbine section rotates in a second direction opposite to the first direction.
    [00014] In an additional embodiment of any of the exposed engines, the speed change system includes a gearbox. The fan, the fan driving turbine section and the second turbine section all rotate in a first direction around the axis.
    [00015] In an additional embodiment of any of the motors shown, the speed change system includes a gearbox. The fan and the second turbine section both rotate in a first direction around the axis. The fan drive turbine rotates in a second direction opposite to the first direction.
    [00016] In an additional embodiment of any of the motors shown, the speed change system includes a gearbox. The fan is rotating in a first direction, and the fan driving turbine and the second turbine section rotate in a second direction opposite to the first direction around the axis.
    [00017] In an additional embodiment of any of the exposed engines, the speed shifting system includes a gear reduction with a gear ratio greater than about 2.3.
    [00018] In an additional embodiment of any of the exposed motors, the fan dispenses a portion of air into a bypass duct. A bypass ratio being defined as the amount of air dispensed into the bypass duct divided by the amount of air dispensed into the compressor section, with the bypass ratio being greater than about 6.0.
    [00019] In an additional embodiment of any of the exposed engines, the deviation ratio is greater than about 10.0.
    [00020] In an additional embodiment of any of the motors shown, a fan pressure ratio across the fan is less than about 1.5.
    [00021] In an additional embodiment of any of the exposed motors, the fan has 26 or fewer blades.
    [00022] In an additional embodiment of any of the exposed engines, the first turbine section has between about 3 and 6 stages.
    [00023] In an additional embodiment of any of the engines shown, a pressure ratio across the first turbine section is greater than about 5:1.
    [00024] In an additional embodiment of any of the exposed engines, a power density greater than about 1.5 IbEin3 (41,520 kg/m3) and less than or equal to about 5.5 lbf/inJ (152,240 kg/m) is included. nf).
    [00025] In an additional embodiment of any of the exposed engines, the second turbine includes at least two stages and works at a first pressure ratio. The fan drive turbine includes more than two stages and works at a second pressure ratio lower than the first pressure ratio.
    [00026] Although the different examples have the specific components shown in the illustrations, embodiments of this description are not limited to those particular combinations. It is possible to use some of the components or features of one of the examples in combination with features or components of another of the examples.
    [00027] These and other features described here can be better understood from the following specification and drawings, after which is a brief description. BRIEF DESCRIPTION OF THE DRAWINGS
    [00028] Figure 1 is a schematic view of an exemplary gas turbine engine.
    [00029] Figure 2 is a schematic view indicating relative rotation between sections of an exemplary gas turbine engine.
    [00030] Figure 3 is another schematic view indicating relative rotation between sections of an exemplary gas turbine engine.
    [00031] Figure 4 is another schematic view indicating relative rotation between sections of an exemplary gas turbine engine.
    [00032] Figure 5 is another schematic view indicating relative rotation between sections of an exemplary gas turbine engine.
    [00033] Figure 6 is a schematic view of a bearing configuration supporting exemplary high and low coil rotation of the exemplary gas turbine engine.
    [00034] Figure 7 is another schematic view of a bearing configuration supporting exemplary high and low coil rotation of the exemplary gas turbine engine.
    [00035] Figure 8A is another schematic view of a bearing configuration supporting exemplary high and low coil rotation of the exemplary gas turbine engine.
    [00036] Figure 8B is an enlarged view of the exemplary bearing configuration shown in Figure 8A.
    [00037] Figure 9 is another schematic view of a bearing configuration supporting exemplary high and low coil rotation of the exemplary gas turbine engine.
    [00038] Figure 10 is a schematic view of a section of the exemplary compact turbine.
    [00039] Figure 11 is a schematic cross section of exemplary stages for the exemplary gas turbine engine described.
    [00040] Figure 12 is a schematic view of an exemplary turbine rotor perpendicular to the geometric axis of rotation. DETAILED DESCRIPTION OF THE INVENTION
    [00041] Figure 1 schematically illustrates an exemplary gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28. Reciprocating engines may include an augmentation section. (not shown) among other systems or features. Fan section 22 drives air along a bypass flow path B while compressor section 24 draws air along a core flow path C where Air is compressed and communicated to a combustor section 26. From the combustor 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and used to drive the fan section 22 and the compressor section 24.
    [00042] Although the described non-limiting embodiment represents a turbofan gas turbine engine, it should be understood that the concepts described here are not limited to use with turbofans, as the precepts can be applied to other types of turbine engines ; for example, a turbine engine including a three-coil architecture in which three coils rotate concentrically about a common axis in such a way that a low coil allows a low pressure turbine to drive a fan via a gearbox, a coil intermediate allows an intermediate pressure turbine to drive a first compressor from the compressor section, and a high coil allows a high pressure turbine to drive a high pressure compressor from the compressor section.
    [00043] Exemplary motor 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about a central longitudinal axis of the motor A with respect to a static frame of the motor 36 via various control systems. bearing 38. It should be understood that various bearing systems 38 at various locations may alternatively, or additionally, be provided.
    [00044] The low speed coil 30 generally includes an internal shaft 40 which connects a fan 42 and a low pressure compressor section (or first section) 44 to a low pressure turbine section (or first section) 46. Inner shaft 40 drives fan 42 through a speed change device, such as geared architecture 48, to drive fan 42 at a slower speed than low speed coil 30. High speed coil 32 includes a shaft outer shaft 50 which interconnects a high pressure compressor section (or second section) 52 and a high pressure turbine section (or second section) 54. The inner shaft 40 and outer shaft 50 are concentric and rotate via the bearing systems. 38 around the central longitudinal axis of engine A.
    [00045] A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a dual stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences greater pressure than a corresponding "low pressure" compressor or turbine.
    [00046] The exemplary low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the exemplary low pressure turbine 46 is measured before a pressure related low pressure turbine 46 inlet. measured at the outlet of the low pressure turbine 46 before an exhaust nozzle.
    [00047] An intermediate turbine frame 58 of the engine static frame 36 is generally arranged between the high pressure turbine 54 and the low pressure turbine 46. The intermediate turbine frame 58 additionally supports bearing systems 38 in the turbine section 28, as well as establishing the flow of air entering the low pressure turbine 46.
    [00048] Airflow from core C is compressed by low pressure compressor 44 and then high pressure compressor 52, mixed with fuel and ignited in combustor 56 to produce high velocity exhaust gases which are then expanded through the high pressure turbine 54 and the low pressure turbine 46. The intermediate turbine frame 58 includes vanes 60, which are in the path of the airflow from the core and serve as an inlet guide vane for the low pressure turbine. 46. Using the blade 60 of the intermediate turbine frame 58 as the inlet guide blade for the low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the intermediate turbine frame 58. or eliminating the number of blades in the low pressure turbine 46 reduces the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density is achieved. ia can be achieved.
    [00049] The gas turbine engine 20 described in an example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an exemplary embodiment being greater than about ten (10). The exemplary geared architecture 48 is an epicyclic gear train, such as a planetary gear system, star gear system, or other known gear system, with a gear reduction ratio greater than about 2.3.
    [00050] In a described embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly greater than an outside diameter of the low pressure compressor 44. It is understood, however, that the cited parameters are only exemplary of an embodiment of a gas turbine engine including a geared architecture and that the present description is applicable to other gas turbine engines.
    [00051] A significant amount of thrust is provided by the bypass flow B because of the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition - typically cruising at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 feet (10,668 meters), with the engine at its best cruising fuel consumption in relation to the thrust it produces - also known as 'specific fuel consumption at cruising speed (' TSFC')” - is the industry standard parameter of pound-mass (Ibm) of fuel per hour that is burned divided by pound-force (Ibf) of thrust that the engine produces at this point of minimum consumption at cruising speed.
    [00052] “Low Fan Pressure Ratio” is the pressure ratio across the fan blade only, without a Fan Output Guide Palette System (“FEGV”). The low fan pressure ratio described herein according to a 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.
    [00053] “Low corrected fan tip speed” is the actual fan tip speed in feet/sec divided by an industry standard temperature correction of [(Tram°R)/518.7)0'51 A “ Low corrected fan tip speed”, described herein in a non-limiting embodiment, is less than about 1150 ft/second (350 m/s).
    [00054] The exemplary gas turbine engine includes fan 42 comprising in a non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than about 18 fan blades. Furthermore, in a described embodiment, the low pressure turbine 46 includes no more than about 6 turbine stages schematically indicated by 34. In another exemplary non-limiting embodiment, the low pressure turbine 46 includes about 3 or more stages. of turbine. A ratio of the number of fan blades 42 to the number of low pressure turbine stages is between about 2.5 and about 8.5. The exemplary low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine stages 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an exemplary gas turbine engine 20 with increased power transfer efficiency.
    [00055] Greater power transfer efficiency is provided in part because of the greater use of turbine blade materials and improved manufacturing methods, such as casting with directional solidification, and monocrystalline materials that allow for higher turbine speeds and reduced numbers. of internships. In addition, the exemplary low pressure turbine 46 includes improved turbine disc configurations that additionally allow for desired durability at the highest turbine speeds.
    [00056] Referring to figures 2 and 3, an exemplary described speed change device is an epicyclic gearbox of a planetary type, where the input is at the central sun gear 62. Planetary gears 64 (only one shown ) around sun gear 62 rotate and are spaced apart by a bracket 68 which rotates in a common direction with sun gear 62. ), contains the entire gear set. Fan 42 is attached and driven by bracket 68 such that the direction of rotation of fan 42 is the same direction of rotation of bracket 68 which in turn is the same direction of rotation of input sun gear 62.
    [00057] In the following figures, nomenclature is used to define the relative rotations between the various sections of the gas turbine engine 20. The fan section is shown with a “+” sign indicating rotation in a first direction. Rotations relative to fan section 22 of other gas turbine engine features are additionally indicated by the use of either a “+” sign or a sign Sign indicates a rotation that is contrary to that of any component indicated with a sign
    [00058] Furthermore, the expression fan drive turbine is used to indicate the turbine that provides the driving power to rotate the blades 42 of the fan section 22. Additionally, the expression “second turbine” is used to indicate the turbine before the fan drive turbine which is not used to drive the fan 42. In this described example, the fan drive turbine is the low pressure turbine 46, and the second turbine is the high pressure turbine 54. It is understood that other turbine section configurations that include more than the high and low pressure turbines 54, 46 shown are within the scope of this description. For example, a three-coil motor configuration may include an intermediate turbine (not shown) used to drive fan section 22 and is within the scope of this description.
    [00059] In an exemplary embodiment described (Figure 2) the fan driving turbine is the low pressure turbine 46 and therefore the fan section 22 and the low pressure turbine 46 rotate in a common direction indicated by the common sign “+” indicating rotation of both the fan 42 and the low pressure turbine 46. Also, in this example, the high pressure turbine 54 or second turbine rotates in a common direction with the fan driving turbine 46. In another example shown in figure 3, the high pressure turbine 54 or second turbine rotated in a direction opposite the fan driving turbine (low pressure turbine 46) and fan 42.
    [00060] Counterrotation of the low pressure compressor 44 and the low pressure turbine 46 relative to the high pressure compressor 52 and the high pressure turbine 54 provides certain efficient aerodynamic conditions in the turbine section 28 as the exhaust gas flow generated high-speed turbine switches from the high-pressure turbine 54 to the low-pressure turbine 46. The relative rotations in the compressor and turbine sections provide approximately the desired airflow angles between the sections, which improves overall efficiency in the turbine section. turbine 28, and providing a reduction in the overall weight of the turbine section 28 by reducing or eliminating the airfoils or an entire row of blades. -
    [00061] Referring to Figures 4 and 5, another exemplary described speed change device is an epicyclic gearbox referred to as a star-type gearbox, where the input is at the central "sun" gear 62. Star Gears 65 (only one shown) around the sun gear 62 rotate in a fixed position around the sun gear and are spaced by a bracket 68 which is fixed to a static case 36 (best shown in Figure 1). A ring gear 66 that is free to rotate contains the entire gear assembly. The fan 42 is attached and driven by the ring gear 66 such that the direction of rotation of the fan 42 is opposite to the direction of rotation of the input sun gear 62. In this way, the low pressure compressor 44 and the low pressure turbine pressure 46 rotate in a direction opposite to the rotation of fan 42.
    [00062] In an exemplary embodiment shown in Figure 4, the fan driving turbine is the low pressure turbine 46 and therefore the fan 42 rotates in a direction opposite to that of the low pressure turbine 46 and the low pressure compressor. Also, in this example, the high coil 32 including the high pressure turbine 54 and the high pressure compressor 52 rotates in a direction opposite to that of the fan 42 and common with the low coil 30 including the low pressure compressor 44 and fan drive turbine 46.
    [00063] In another exemplary gas turbine engine shown in Figure 5, the high pressure turbine or second turbine 54 rotates in a common direction with the fan 42 and against the low coil 30 including the low pressure compressor 44 and the fan drive turbine 46.
    [00064] Referring to Figure 6, the bearing assemblies near the front end of the shafts on the motor at locations 70 and 72, whose bearings support rotation of the inner shaft 40 and outer shaft 50, oppose net thrust forces in a direction parallel to the axis A that are generated by the backward loading of the low pressure turbine 46 and the high pressure turbine 54, minus the high pressure compressor 52 and the low pressure compressor 44, which also contribute to the thrust forces which act on the corresponding low coil 30 and high coil 32.
    [00065] In this exemplary embodiment, a first front bearing assembly 70 is supported on a portion of the static frame shown schematically at 36 and supports a front end of the inner shaft 40. The first exemplary front bearing assembly 70 is a thrust bearing and controls the movement of the inner shaft 40 and thereby the coil 30 lowers in an axial direction. A second front bearing assembly 72 is supported by the static frame 36 to support rotation of the high coil 32 and substantially prevent movement along an axial direction of the outer shaft 50. The first front bearing assembly 70 is mounted to support the inner shaft 40 at a point ahead of a connection 88 of a low pressure compressor rotor 90. The second front bearing assembly 72 is mounted ahead of a connection referred to as a hub 92 between a high pressure compressor rotor 94 and the shaft A first tail bearing assembly 74 supports the tail portion of the inner shaft 40. The first tail bearing assembly 74 is a roller bearing and supports rotation but does not provide resistance to movement of the shaft 40 in the axial direction. Instead, tail bearing 74 allows shaft 40 to thermally expand between its location and bearing 72. First exemplary tail bearing assembly 74 is disposed behind a connecting hub 80 between a low pressure turbine rotor 78 and the inner shaft 40. A second tail bearing assembly 76 supports the tail portion of the outer shaft 50. The exemplary second tail bearing assembly 76 is a roller bearing and is supported by a corresponding static frame 36 through the intermediate turbine frame 58 which transfers the radial load from the shaft through the turbine flow path to the ground 36. The second rear bearing assembly 76 supports the outer shaft 50 and thereby the high coil 32 at a point behind a connecting hub 84 between a high pressure turbine rotor 82 and the outer shaft 50.
    [00066] In this described example, the first and second sets of front bearings 70, 72 and the first and second sets of rear bearings 74, 76 are supported on the outside of any of the corresponding compressor or turbine connection hubs 80, 88 to provide a prominent support configuration of the corresponding inner shaft 40 and outer shaft 50. The prominent support of the inner shaft 40 and outer shaft 50 provides a desired support and rigidity for operation of the gas turbine engine 20.
    [00067] Referring to Figure 7, another exemplary axle support configuration includes first and second sets of front bearings 70, 72 arranged to support the front portion of corresponding inner shaft 40 and outer shaft 50. The first tail bearing 74 is disposed behind the connection 80 between the rotor 78 and the inner shaft 40. The first tail bearing 74 is a roller bearing and supports the inner shaft 40 in a prominent configuration. The prominent configuration may require additional length of the inner shaft 40 and therefore an alternative configuration referred to as a suspended configuration may be utilized. In this example, the outer shaft 50 is supported by the second rear bearing assembly 76 which is arranged in front of the connection 84 between the high pressure turbine rotor 82 and the outer shaft 50. In this way, the connecting hub 84 of the rotor of the high pressure turbine 82 on the outer shaft 50 is suspended behind the bearing assembly 76. This positioning of the second rear bearing 76 in an overhead orientation potentially provides a reduced length of the outer shaft 50.
    [00068] In addition, the positioning of the tail bearing 76 can also eliminate the need for other support structures such as the intermediate turbine frame 58 as both the high pressure turbine 54 is supported on the bearing assembly 76 and the turbine Low pressure 46 is supported by bearing assembly 74. Optionally, intermediate turbine frame stanchion 58 may provide an optional roller bearing 74A which may be added to reduce inner shaft 40 vibratory modes.
    [00069] Referring to Figures 8A and 8B, another exemplary axle support configuration includes first and second sets of front bearings 70, 72 arranged to support corresponding front portions of each of the inner axle 40 and the outer axle 50. The first rear bearing 74 provides support for the outer shaft 40 at a location behind the fitting 80 in a prominent mounting configuration. In this example, the rear portion of outer shaft 50 is supported by a roller bearing assembly 86 supported in a space 96 defined between an outer surface of inner shaft 40 and an inner surface of outer shaft 50.
    [00070] Bearing bearing assembly 86 supports the rear portion of outer shaft 50 on inner shaft 40. The use of bearing bearing assembly 86 to support outer shaft 50 eliminates the requirements for support structures that lead back to the static frame 36 through intermediate turbine frame 58. In addition, exemplary bearing assembly 86 can provide both reduced shaft length and outer shaft support 50 in a position substantially in axial alignment with connecting hub 84 to the rotor. of the high pressure turbine 82 and the outer shaft 50. As can be seen, the bearing assembly 86 is positioned behind the hub 82 and is supported through the rearmost section of the shaft 50. Referring to Figure 9, another Exemplary axle support configuration includes first and second sets of front bearings 70, 72 arranged to support corresponding front portions of each of inner axle 40 and outer axle 50. First rear bearing assembly 74 is supported at a point along inner shaft 40 forward of connection 80 between low pressure turbine rotor 78 and inner shaft 40.
    [00071] Positioning the first rear bearing assembly 74 in front of the connection 80 can be used to reduce the overall length of the motor 20. In addition, positioning the first rear bearing assembly 74 in front of the connection 80 provides support through the frame of the motor. intermediate turbine 58 to static frame 36. Also, in this example, second rear bearing assembly 76 is deployed in a prominent mounting configuration behind connection 84 between outer shaft 50 and rotor 82. Thus, in this example , both the first and second sets of after bearings 74, 76 share a common support structure with the static outer structure 36. As can be seen, a common support feature such as this provides a less complex motor construction along with reduced the overall length of the engine. In addition, the required reduction or support structures will reduce overall weight to provide a further improvement in aircraft fuel-burning efficiency.
    [00072] Referring to Figure 10, a portion of the exemplary turbine section 28 is shown and includes the low-pressure turbine 46 and the high-pressure turbine 54 with the intermediate turbine frame 58 disposed between an outlet of the high-pressure turbine. pressure and the low pressure turbine. Intermediate turbine frame 58 and vane 60 are positioned to be upstream of the first stage 98 of low pressure turbine 46. Although a single vane 60 is illustrated, it should be understood that these could be several circumferentially spaced vanes 60. The vane 60 redirects the flow downstream of the high pressure turbine 54 as it approaches the first stage 98 of the low pressure turbine 46. As can be seen, it is desirable to improve the efficiency to have flow between the high pressure 54 and low pressure turbine 46 redirected by vane 60 such that the expanding gas stream is aligned in the desired manner as it enters low pressure turbine 46. Therefore vane 60 can be a real cambered airfoil and rotation, which aligns the airflow in the desired manner to the low pressure turbine 46.
    [00073] By incorporating a true air rotation blade 60 in the intermediate turbine frame 58, instead of an aerodynamic strut and a row of stator blade after the strut, the overall length and volume of the combined turbine sections 46, 54 are reduced by virtue of the blade 60 serving a number of functions including upgrading the intermediate turbine frame 58, protecting any static structure and any oil tubes serving a bearing assembly from heat exposure, and rotating the flow entering the turbine from heat. low pressure 46 such that it enters the rotating airfoil 100 at a desired flow angle. Additionally, by incorporating these features together, the overall turbine section 28 assembly and arrangement is reduced in volume.
    [00074] The cited features achieve a more or less compact turbine section volume compared to previous technology including both high and low pressure turbines 54, 46. Furthermore, in one example, the materials to form the low pressure turbine pressure 46 can be improved to provide reduced volume. Such materials may include, for example, materials with greater thermal and mechanical capabilities to accommodate potentially greater stresses induced by operating the low pressure turbine 46 at the highest speed. In addition, the higher speeds and higher operating temperatures at the inlet of the low pressure turbine 46 allow the low pressure turbine 46 to transfer a greater amount of energy more efficiently to drive both a larger diameter fan 42 through the geared architecture 48 and an increase in compressor work done by the low pressure compressor 44.
    [00075] Alternatively, cheaper materials may be used in combination with cooling features that compensate for higher temperatures within the low pressure turbine 46. In three exemplary embodiments, a first rotating blade 100 of the low pressure turbine 46 may be a cast blade. unidirectionally solidified, a monocrystalline cast blade or an internally cooled hollow blade. The improved material and better thermal properties of the exemplary turbine blade material provide for operation at higher temperatures and speeds, which, in turn, provide greater efficiencies at each stage, which, thereby, allow the use of a reduced number of turbine stages. low pressure turbine. The reduced number of low pressure turbine stages in turn provides an overall turbine volume that is reduced, and which accommodates desired increases in low pressure turbine speed.
    [00076] The reduced stages and reduced volume provide greater engine efficiency and aircraft fuel burn due to the lower overall weight. What's more, as there are fewer rows of blades, there are: fewer leak paths at the tips of the blades; fewer leak paths in the air seals inside the pallets; and reduced losses across the rotor stages.
    [00077] The exemplary described compact turbine section includes a power density, which can be defined as thrust in pounds force (Ibf) produced divided by the volume of the entire turbine section 28. The volume of the turbine section 28 can be defined through an inlet 102 from a first turbine blade 104 on the high pressure turbine 54 to the outlet 106 of the last rotating airfoil 108 on the low pressure turbine 46, and may be expressed in cubic inches. The static thrust at the rated flat sea level take-off condition of the engine divided by a turbine section volume is defined as the power density, and a higher power density may be desirable for reduced engine weight. The rated static lift-off thrust at sea level can be defined in pounds-force (Ibf), while the volume can be the volume from the annular inlet 102 of the first turbine blade 104 in the high pressure turbine 54 to the annular outlet 106 from the downstream end of the last airfoil 108 on the low pressure turbine 46. The maximum thrust may be takeoff thrust at sea level “SLTO thrust” which is normally defined as the flat rated static thrust produced by the turbofan at sea level.
    [00078] The volume V of the turbine section can be better understood from Figure 10. As shown, the intermediate turbine frame 58 is arranged between the high pressure turbine 54 and the low pressure turbine 46. The volume V is illustrated by a dashed line, and extends from an inner periphery I to an outer periphery O. The inner periphery is defined by the flow path of the rotors, but also by the flow paths of an inner pallet platform. The outer periphery is defined by the stator vanes and external air sealing structures along the flow path. The volume extends from a most upstream end of the vane 104, typically its leading edge, and to the most downstream edge of the last rotating airfoil 108 in the low pressure turbine section 46. Typically this will be the trailing edge of the airfoil 108.
    [00079] The power density in the described gas turbine engine is much higher than in the prior art. Eight exemplary engines are shown below that incorporate turbine sections and general engine drive systems and architectures presented in this order, and can be found in Table I below:

    [00080] Thus, in exemplary embodiments, the power density would be greater than or equal to about 1.5 lbf/in (41,520 kg/m ). More strictly, the power density would be greater than or equal to about 2.0 lbf/inJ (55,360 kg/m3). Even more strictly, the power density would be greater than or equal to about 3.0 lbf/in (83,040 kg/m ). More strictly, the power density is greater than or equal to about 4.0 lbf/in (110,720 kg/m ). Also, in embodiments, the power density is less than or equal to about 5.5 lbf/in3 (152,240 kg/m3).
    [00081] Engines made with the architecture described, and including turbine sections presented in this application, and with modifications within the scope of this description, thus provide very high efficiency operation, and greater fuel efficiency and low weight in relation to their carrying capacity. buoyancy.
    [00082] An outlet area 112 is defined at the outlet location for the high pressure turbine 54 and an outlet area 110 is defined at the outlet 106 of the low pressure turbine 46. Gear reduction 48 (shown in Figure 1) provides a range of different rotational speeds of the fan driving turbine, which, in this exemplary embodiment, is the low pressure turbine 46, and the fan 42 (Figure 1). In this way, the low pressure turbine 46 and thereby the low coil 30 including the low pressure compressor 44 can rotate at a very high speed. The operation of the low pressure turbine 46 and the high pressure turbine 54 can be evaluated towards a performance value which is the output area for the respective turbine section multiplied by their respective velocity squared. This performance value (“PQ”) is defined as: Equation 1: PQitp = (Aipt x Vipt2) Equation 2: PQhpt = (Ahpt x Vhpt2) where A|pt is the area 110 of the low pressure turbine 46 at outlet 106 , Vipt is the speed of the low pressure turbine section; Ahpt is the area of high pressure turbine 54 at output 114, and where Vhpt is the speed of high pressure turbine 54.
    [00083] Thus, a ratio of the performance value for the low pressure turbine 46 to the performance value for the high pressure turbine 54 is: Equation 3: (A)pt x Vlpt2)/(Ahpt x Vhpt2) = PQitp /PQhpt
    [00084] In a turbine modality made according to the previous design, the areas of the low and high pressure turbines 46, 54 are 3599.3 cm (557.9 in ) and 584.9 cm" (90.67 in ), respectively. Additionally, the low- and high-pressure turbine speeds 46, 54 are 10,179 rpm and 24,346 rpm, respectively. Thus, using Equations 1 and 2 above, the performance values for the exemplary low- and high-pressure turbines 46.54 are: Equation 1: PQitp = (A)ptx Vlpl2) = (557.9 in2)( 10,179 rpm)2 = 57,805,157,673.9 in2rpm2 (lin2= 6.45 cm2) Equation 2: PQhpt = (Ahpt x Vhpt2) = (90.67 in2)(24,346 rpm)2 = 53,742,622,009.72 in2 rpm2 (1 in2 = 6.45 cm2) and, using Equation 3 above, the ratio for the low pressure turbine section for the high pressure turbine section is: Ratio = PQitp/PQhpt= 57,805,157,673.9 in2rpm"/ 53,742,622,009.72 in2rpm2 = 1.075 (1 in2 = 6.45 cm2)
    [00085] In another embodiment, the ratio is greater than about 0.5, and in another embodiment, the ratio is greater than about 0.8. With PQitp/PQhpt ratios in the range of 0.5 to 1.5, an overall very efficient gas turbine engine is obtained. More strictly, PQilp/PQhpt ratios greater than or equal to about 0.8 provide greater overall gas turbine efficiency. Even more strictly, PQitp/PQhpt ratios greater than or equal to 1.0 are even more thermodynamically efficient and provide a reduction in weight that improves the aircraft's fuel-burning efficiency. As a result of these PQitp/PQhpt ratios, in particular, the turbine section 28 can be made much smaller than in the prior art, both in diameter and axial length. What's more, the overall engine efficiency is greatly increased.
    [00086] Referring to Figure 11, portions of the low pressure compressor 44 and the low pressure turbine 46 of the low spool 30 are shown schematically and include rotors 116 of the low pressure turbine 46 and rotors 132 of the low pressure compressor 44 Each of the rotors 116 includes a radius of hole 122, a radius of live disk 124 and a width of hole 126 in a direction parallel to axis A. The rotor 116 supports turbine blades 118 which rotate with respect to the turbine blades. 120. Low pressure compressor 44 includes rotors 132 including a bore radius 134, a live disk radius 136 and a bore width 138. The rotor 132 supports compressor blades 128 which rotate with respect to vanes 130.
    [00087] The radius of hole 122 is the radius between an innermost surface of the hole and the geometry axis. The radius of the live disk 124 is the radial distance from the axis of rotation A and a portion of the rotor that supports the airfoil blades. The width of rotor hole 126 in this example is the largest width of the rotor and is arranged at a radially spaced distance from the axis A determined to provide desired physical performance properties.
    [00088] The rotors for each of the low pressure compressor 44 and the low pressure turbine 46 rotate at a higher speed compared to prior art low spool configurations. The geometric shape including hole radius, live disk radius and hole width are determined to provide desired rotor performance in view of the selected mechanical and thermal stresses to be imposed during operation. Referring to Figure 12, continuing with reference to Figure 11, a turbine rotor 116 is shown to further illustrate the relationship between the radius of bore 126 and the radius of live disk 124. In addition, the relationships described are provided in from a known range of materials normally used for the construction of each of the rotors.
    [00089] In this way, the greatest attributes of performance and performance are provided by the desirable combinations of the described features of the various components of the described and described gas turbine engine modalities.
    [00090] Although an exemplary embodiment has been described, those skilled in the art should realize that certain modifications would fall within the scope of this description. For this reason, the following claims should be studied to determine the scope and content of this description.
    权利要求:
    Claims (15)
    [0001]
    1. A gas turbine engine (20), comprising: a compressor section (24); a combustor (26) in fluid communication with the compressor section (24); a turbine section (28) in fluid communication with the combustor (26), the turbine section (28) including a fan drive turbine (46) and a second turbine (54), the fan drive turbine (46) including a plurality of turbine stages (34); a fan (42) including a plurality of blades rotating about an axis (A); and a speed change system (48) driven by the fan drive turbine (46) to rotate the fan (42) about the axis (A); wherein the fan drive turbine (46) includes a first rotor tail (78) attached to a first shaft (40) and the second turbine (54) includes a second tail rotor (82) attached to a second shaft (50), and a first tail bearing assembly (74) is arranged axially rearwardly. of a first connection (80) between the first tail rotor (78) and the first tail shaft (40), and a second tail bearing assembly (76) is arranged axially forward of a second connection (84) between the second rear rotor (82) and the second axis (50); characterized by the fact that: a ratio between the number of fan blades and the number of fan drive turbine stages (34) is between 2.5 and 8, 5; and the second turbine (54) includes at least two stages and works at a first pressure and the fan drive turbine (46) includes more than two stages and works at a second pressure lower than the first pressure.
    [0002]
    2. Engine (20) according to claim 1, characterized in that the first rear bearing assembly (74) and the second rear bearing assembly (76) comprise roller bearings.
    [0003]
    3. Engine (20) according to any one of claims 1 or 2, characterized in that the compressor section (24) comprises a first compressor (44) driven by the fan drive turbine (46) through the first shaft (40) and a second compressor driven by the second turbine (54) through the second shaft (50), wherein the first rear bearing (74) supports a rear portion of the first shaft (40) and the second rear bearing (76) supports a rear portion of the second axis (50).
    [0004]
    4. Engine (20) according to claim 3, characterized in that a front portion of each of the first and second axles (40, 50) is supported by a thrust bearing assembly (70, 72).
    [0005]
    5. Engine (20) according to any one of claims 1 to 4, characterized in that the fan drive turbine (46) has a first outlet area (110) at a first outlet point (106) and routed at a first speed, the second turbine (54) has a second exit area (112) at a second exit point and rotates at a second speed which is greater than the first speed, wherein a first performance value is defined as the product of the first velocity squared by the first area (110), a second performance value is defined as the product of the second velocity squared by the second area (112), and a performance ratio of the first performance value to the second performance value is between 0.5 and 1.5.
    [0006]
    6. Engine (20) according to claim 5, characterized in that the performance ratio is greater than or equal to 0.8.
    [0007]
    7. Engine (20) according to any one of claims 5 or 6, characterized in that the first performance value is greater than or equal to 4.
    [0008]
    8. Engine (20) according to any one of claims 1 to 7, characterized in that the speed change system (48) comprises a gearbox, and in which: the fan (42) and the turbine fan drives (46) both rotate in a first direction around the axis (A) and the second turbine section (54) rotates in a second direction opposite to the first direction; or the fan (42), the fan driving turbine (46), and the second turbine (54) all rotate in a first direction about the axis (A); or the fan (42) and the second turbine (54) both rotate in a first direction about axis (A) and the fan driving turbine rotates in a second direction opposite to the first direction; or the fan (42) is rotatable in a first direction and the fan driving turbine (46) and the second turbine (54) rotate in a second direction opposite to the first direction about axis (A).
    [0009]
    9. Engine (20) according to any one of claims 1 to 8, characterized in that the speed change system (48) comprises a gear reduction with a gear ratio greater than about 2.3.
    [0010]
    10. Engine (20) according to any one of claims 1 to 9, characterized in that the fan (42) dispenses a portion of air in a diversion duct, and a diversion ratio being defined as the air portion dispensed into the bypass duct divided by the amount of air dispensed into the compressor section (24), with the bypass ratio being greater than 6.0 or greater than 10.0.
    [0011]
    11. Motor (20) according to any one of claims 1 to 10, characterized in that a fan pressure ratio across the fan (42) is less than 1.5.
    [0012]
    12. Motor (20) according to any one of claims 1 to 11, characterized in that the fan (42) has 26 or less blades.
    [0013]
    13. Engine (20) according to any one of claims 1 to 12, characterized in that said first fan drive turbine section (46) has between 3 and 6 stages.
    [0014]
    14. Engine (20) according to any one of claims 1 to 13, characterized in that the pressure ratio across the fan drive turbine (46) is greater than 5:1.
    [0015]
    15. Engine (20) according to any one of claims 1 to 14, characterized in that it includes a power density greater than 1.5 lbf/in3 (41,520 kg/m3) and less than or equal to 5.5 lbf/in3 (41,520 kg/m3) in3 (152,240 kg/m3).
    类似技术:
    公开号 | 公开日 | 专利标题
    BR112014016273B1|2022-02-01|gas turbine engine
    BR112014016277B1|2022-02-01|gas turbine engine
    BR112014016303B1|2022-01-25|gas turbine engine
    BR112014016305B1|2022-01-25|gas turbine engine
    BR112014016281B1|2022-02-01|gas turbine engine
    US9739206B2|2017-08-22|Geared turbofan gas turbine engine architecture
    US10794292B2|2020-10-06|Geared turbofan gas turbine engine architecture
    CA2856912C|2017-04-25|Geared turbofan gas turbine engine architecture
    BR102016005926A2|2016-10-25|gas turbine engine
    US9222417B2|2015-12-29|Geared turbofan gas turbine engine architecture
    BR102016005922A2|2016-10-25|gas turbine engine
    BR102016015269A2|2017-01-03|GAS TURBINE ENGINE
    同族专利:
    公开号 | 公开日
    US20130192258A1|2013-08-01|
    BR112014016273A2|2017-06-13|
    CA2857351A1|2013-10-24|
    SG11201402823TA|2014-09-26|
    RU2633218C2|2017-10-11|
    EP2809922A4|2015-11-18|
    EP2809922A1|2014-12-10|
    BR112014016273A8|2017-07-04|
    EP2809922B2|2022-01-26|
    RU2014134426A|2016-03-27|
    EP2809922B1|2018-06-20|
    EP3312404A1|2018-04-25|
    WO2013158191A1|2013-10-24|
    CA2857351C|2016-06-28|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3747343A|1972-02-10|1973-07-24|United Aircraft Corp|Low noise prop-fan|
    US4611464A|1984-05-02|1986-09-16|United Technologies Corporation|Rotor assembly for a gas turbine engine and method of disassembly|
    GB8630754D0|1986-12-23|1987-02-04|Rolls Royce Plc|Turbofan gas turbine engine|
    US5433674A|1994-04-12|1995-07-18|United Technologies Corporation|Coupling system for a planetary gear train|
    DE19828562B4|1998-06-26|2005-09-08|Mtu Aero Engines Gmbh|Engine with counter-rotating rotors|
    USH2032H1|1999-10-01|2002-07-02|The United States Of America As Represented By The Secretary Of The Air Force|Integrated fan-core twin spool counter-rotating turbofan gas turbine engine|
    US6619030B1|2002-03-01|2003-09-16|General Electric Company|Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors|
    US7594388B2|2005-06-06|2009-09-29|General Electric Company|Counterrotating turbofan engine|
    US7966806B2|2006-10-31|2011-06-28|General Electric Company|Turbofan engine assembly and method of assembling same|
    US7721549B2|2007-02-08|2010-05-25|United Technologies Corporation|Fan variable area nozzle for a gas turbine engine fan nacelle with cam drive ring actuation system|
    US8256707B2|2007-08-01|2012-09-04|United Technologies Corporation|Engine mounting configuration for a turbofan gas turbine engine|
    US8590286B2|2007-12-05|2013-11-26|United Technologies Corp.|Gas turbine engine systems involving tip fans|
    US7762086B2|2008-03-12|2010-07-27|United Technologies Corporation|Nozzle extension assembly for ground and flight testing|
    DE102008023990A1|2008-05-16|2009-11-19|Rolls-Royce Deutschland Ltd & Co Kg|Two-shaft engine for an aircraft gas turbine|
    US8695920B2|2008-06-02|2014-04-15|United Technologies Corporation|Gas turbine engine with low stage count low pressure turbine|EP2610463B1|2011-12-30|2016-08-03|United Technologies Corporation|Gas turbine engine gear train|
    EP3165754A1|2015-11-03|2017-05-10|United Technologies Corporation|Gas turbine engine with high speed low pressure turbine section and bearing support features|
    EP3163033A1|2015-10-26|2017-05-03|United Technologies Corporation|Gas turbine engine with high speed low pressure turbine section and bearing support features|
    US10309232B2|2012-02-29|2019-06-04|United Technologies Corporation|Gas turbine engine with stage dependent material selection for blades and disk|
    WO2015088619A2|2013-10-16|2015-06-18|United Technologies Corporation|Geared turbofan engine with targeted modular efficiency|
    EP3084173A4|2013-12-16|2016-12-28|United Technologies Corp|Geared turbofan with three turbine sections|
    EP3108117B1|2014-02-19|2020-02-05|United Technologies Corporation|Gas turbine engine airfoil|
    EP3108107A4|2014-02-19|2017-03-08|United Technologies Corporation|Gas turbine engine airfoil|
    US9869190B2|2014-05-30|2018-01-16|General Electric Company|Variable-pitch rotor with remote counterweights|
    US10287976B2|2014-07-15|2019-05-14|United Technologies Corporation|Split gear system for a gas turbine engine|
    US10072510B2|2014-11-21|2018-09-11|General Electric Company|Variable pitch fan for gas turbine engine and method of assembling the same|
    US10731560B2|2015-02-12|2020-08-04|Raytheon Technologies Corporation|Intercooled cooling air|
    US20160237908A1|2015-02-12|2016-08-18|United Technologies Corporation|Intercooled cooling air using existing heat exchanger|
    US20160245184A1|2015-02-19|2016-08-25|United Technologies Corporation|Geared turbine engine|
    US11225913B2|2015-02-19|2022-01-18|Raytheon Technologies Corporation|Method of providing turbine engines with different thrust ratings|
    US20160333786A1|2015-05-13|2016-11-17|General Electric Company|System for supporting rotor shafts of an indirect drive turbofan engine|
    US9909453B2|2015-05-19|2018-03-06|General Electric Company|Lubrication system for a turbine engine|
    US10415429B2|2015-09-25|2019-09-17|General Electric Company|Planet gearbox with cylindrical roller bearing with high density roller packing|
    US10100653B2|2015-10-08|2018-10-16|General Electric Company|Variable pitch fan blade retention system|
    US10234018B2|2015-10-19|2019-03-19|General Electric Company|Planet gearbox with cylindrical roller bearing with under race lube scheme|
    ITUB20156062A1|2015-12-01|2017-06-01|Gen Electric|HOUSING FOR USE IN A MOTOR-DRIVEN ENGINE AND WASHING PROCESS OF FLUID FROM IT.|
    US11242770B2|2020-04-02|2022-02-08|General Electric Company|Turbine center frame and method|
    法律状态:
    2018-12-04| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2020-02-27| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|
    2021-06-01| B350| Update of information on the portal [chapter 15.35 patent gazette]|
    2021-12-07| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2022-02-01| 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 30/01/2013, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    US13/363,154|US20130192196A1|2012-01-31|2012-01-31|Gas turbine engine with high speed low pressure turbine section|
    US13/363154|2012-01-31|
    US201261653780P| true| 2012-05-31|2012-05-31|
    US61/653780|2012-05-31|
    US13/645,665|US20130192258A1|2012-01-31|2012-10-05|Geared turbofan gas turbine engine architecture|
    US13/645665|2012-10-05|
    PCT/US2013/023715|WO2013158191A1|2012-01-31|2013-01-30|Geared turbofan gas turbine engine architecture|
    [返回顶部]