JP6027250B2 - Gas turbine engine having a forward moment arm - Google Patents

Description

  The present invention relates to a gas turbine engine, and more particularly to the center of gravity of a gas turbine engine.

  This application claims priority to US Provisional Patent Application No. 61 / 708,510, filed Oct. 1, 2012. A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. The air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to produce a high velocity exhaust gas stream. The high velocity exhaust gas stream expands through the turbine section and drives the compressor and fan sections. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

  A reduction gear, such as an epicyclic gear assembly, can be utilized to drive the fan section so that the fan section can rotate at a different speed than the turbine section so as to increase the overall propulsion efficiency of the engine. . In such an engine structure, the shaft driven by one of the turbine parts drives the fan part at a reduced speed so that both the turbine part and the fan part can rotate closer to the optimum speed. Provides input to the epicyclic gear assembly.

  The structure of the gas turbine engine contributes to the total weight of the engine and the equilibrium point is defined at the center of gravity. The position of the center of gravity of a gas turbine engine affects how the engine is installed and how the surrounding nacelle structure is constructed. Moving the center of gravity forward along the engine axis increases the internal moment arm, thereby increasing the load on the engine mounting structure. The position of the engine's center of gravity is affected by the choice of material and component configuration.

  While the gear structure has improved propulsion efficiency, turbine engine manufacturers continue to seek further improvements in engine performance, including improvements in thermal efficiency, transmission efficiency, and propulsion efficiency.

  A gas turbine engine according to an exemplary embodiment of the present disclosure includes, among other things, a plurality of fan blades that are rotatable about an axis, each having a plurality of fan blades including a leading edge and a last having a trailing edge. And a gear structure driven by a turbine section for rotating a plurality of fan blades about an axis, wherein the center of gravity of the gas turbine engine has a leading edge of the plurality of fan blades. And a first axial distance from the trailing edge of the rearmost turbine blade that is about 35% to about 75% of the total length between the rearmost blade and the trailing edge of the last turbine blade.

  In a further embodiment of the gas turbine engine, the center of gravity is disposed substantially along the axis.

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity includes the weight of the structure comprising the gas turbine engine that does not include the engine mounting structure, the engine cowling structure, and the nacelle structure. Determined.

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity is determined including the weight of fluid contained within the operating system of the gas turbine engine.

  In a further embodiment of any of the above turbine engine embodiments, the first axial distance is about 40% to the total length between the leading edge of the plurality of fan blades and the trailing edge of the rearmost turbine blade. About 70%.

In a further embodiment of any of the above turbine engine embodiments, the plurality of fan blades are supported on a rotor and the rotor and fan blades are about 260.19 kg / m ³ ( about 0.0094 lb / stand). having a density of square inches) to about 426.27kg / m ³ (about 0.01540 lbs / stand side inch).

In a further embodiment of any of the above turbine engine embodiments, the gear structure is between about 6022.58 kg / m ³ ( about 0.22 lb / in ³ ) and about 8303.97 kg / m ³ ( about 0). Gearbox with a density of .30 lb / in3 ) .

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity is located at the intersection of a vertical line and an axis extending through the lifting point of the gas turbine engine, the axis being perpendicular to the vertical line. is there.

  Another gas turbine engine according to an exemplary embodiment of the present disclosure includes, among other things, a plurality of fan blades that are rotatable about an axis, each including a plurality of fan blades including a leading edge, and a trailing edge. A turbine section including a last rotating turbine blade having a gear structure driven by the turbine section for rotating a plurality of fan blades about an axis, wherein the internal moment arm of the turbofan engine is about 35 A first distance from the center of gravity of the turbofan engine to the trailing edge of the last rotating turbine blade, and the total length between the leading edges of the plurality of fan blades and the trailing edge of the last turbine blade, from about% to about 75% And the ratio.

  In other embodiments of the gas turbine engine, the ratio is between about 40% and 70%.

In a further embodiment of any of the above turbine engine embodiments, the gear structure is between about 6022.58 kg / m ³ ( about 0.22 lb / in ³ ) and about 8303.97 kg / m ³ ( about 0). Gearbox with a density of .30 lb / in3 ) .

In a further embodiment of any of the above turbine engine embodiments, the rotor includes a rotor that supports a plurality of fan blades, wherein the rotor and fan blades are about 260.19 kg / m ³ ( about 0.0094 pounds per cubic inch). ) To about 426.27 kg / m ³ ( about 0.01540 lb / in ³ ) .

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity is determined including the weight of fluid contained within the operating system of the turbofan engine.

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity is located at the intersection of a vertical line and an axis extending through the lifting point of the turbofan engine, and the axis is perpendicular to the vertical line. is there.

  A method of assembling a gas turbine engine according to an exemplary embodiment of the present disclosure includes, inter alia, supporting a plurality of fan blades, each including a leading edge, about a rotational axis and a trailing edge about the rotational axis. Supporting a turbine portion including a rearmost turbine blade having, supporting a gear structure driven by a turbine portion for rotating a plurality of fan blades about an axis, and leading edges of the plurality of fan blades To position the center of gravity of the gas turbine engine at a first axial distance from the trailing edge of the rearmost turbine blade that is about 35% to about 75% of the total length between the rear edge of the rearmost turbine blade and the rearmost turbine blade. Selecting components of the gas turbine engine structure.

A further embodiment of any of the above turbine engine embodiments includes supporting a plurality of fan blades on a rotor, the rotor and fan blades being about 260.19 kg / m ³ ( about 0.0094 pounds). / Cubic inch ) to about 426.27 kg / m ³ (about 0.01540 pounds / cubic inch ) .

In any of the further embodiments of the embodiments of the turbine engine, a gear structure, about 6089.58kg / m ³ (about 0.22 lb / stand side inch) to about 8303.97kg / m ³ (about Assembling as a gearbox having a density of 0.30 pounds / cubic inch ) .

  In a further embodiment of any of the above turbine engine embodiments, the center of gravity is within a range of about 40% to 70% of the total length between the leading edge of the plurality of fan blades and the trailing edge of the rearmost turbine blade. Selecting components of the gas turbine engine for deployment.

  Although different examples have particular components illustrated in the illustration, embodiments of the present disclosure are not limited to those particular combinations. Some of the components or features from one of the examples may be used in combination with the components or features from another of the examples.

  These and other features disclosed herein can be best understood from the following specification and drawings, of which the following is a brief description.

1 is a schematic view of the center of gravity of an example gas turbine engine. FIG. 1 is a schematic view of the center of gravity of an example gas turbine engine. FIG.

  FIG. 1 schematically illustrates an example of a gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28. An alternative engine may include an augmentor (not shown) of other systems or features. While the fan section 22 sends air along the bypass flow path B, the compressor section 24 sucks air along the core flow path C where the air is compressed and communicated to the combustor section 26. In the combustor section 26, the air is mixed with fuel, ignited and produces a high pressure exhaust gas stream that expands through the turbine section 28, where energy is extracted and utilized, the fan section 22 and the compression. The machine part 24 is driven.

  Although the disclosed non-limiting embodiments depict a turbofan gas turbine engine, the concepts described herein can be used with turbofans so that the teachings can be applied to other types of turbine engines. It should be understood that this is not a limitation. For example, the teachings can be applied to a turbine engine that includes a three-shaft structure in which three axes rotate concentrically about a common axis, and such turbine engines provide a fan to a low-pressure turbine via a gear box. A low spool that allows driving, an intermediate spool that allows an intermediate pressure turbine to drive a first compressor of the compressor section, and a high pressure turbine that drives a high pressure compressor of the compressor section. Includes a high spool to allow.

  An example of the engine 20 generally includes a low speed spool 30 attached for rotation about the engine's longitudinal central axis A to the stationary structure 36 of the engine via several bearing systems 38. A high speed spool 32 is included. It should be understood that various bearing systems 38 at various locations may be provided alternatively or additionally.

  The low speed spool 30 generally includes an inner shaft 40 that connects the fan 42 and the low pressure (or first) compressor section 44 to the low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 via a transmission such as the gear structure 48, and drives the fan 42 at a speed slower than that of the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate through a bearing system 38 about a longitudinal central axis A of the engine.

  The combustor 56 is disposed between the high-pressure compressor 52 and the high-pressure turbine 54. In one example, high pressure turbine 54 includes at least two stages for providing a two-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 a higher pressure than the corresponding “low pressure” compressor or turbine.

  The example low pressure turbine 46 has a pressure ratio greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured before the inlet of the low pressure turbine 46 in relation to the pressure measured at the outlet of the low pressure turbine 46 before the exhaust nozzle.

  The intermediate turbine frame 58 of the stationary engine structure 36 is generally disposed between the high pressure turbine 54 and the low pressure turbine 46. The intermediate turbine frame 58 sets the air flow entering the low pressure turbine 46 and further supports the bearing system 38 within the turbine section 28.

  The air flow through the core channel C is compressed by the low pressure compressor 44 and then by the high pressure compressor 52, mixed with fuel, ignited in the combustor 56, and then through the high pressure turbine 54 and the low pressure turbine 46. Generates high-speed exhaust gas that is expanded. The intermediate turbine frame 58 includes a vane 60 that is in the core air flow path and functions as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the intermediate turbine frame 58 as an inlet guide vane for the low pressure turbine 46 reduces the length of the low pressure turbine 46 without increasing the axial length of the intermediate turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 reduces the axial length of the turbine section 28. Therefore, the gas turbine engine 20 can be miniaturized and a higher power density can be achieved.

  The disclosed gas turbine engine 20 in one example is an aircraft engine with a high bypass gear. In a further example, the gas turbine engine 20 includes a bypass ratio that is greater than about 6 (6), and in an example embodiment is greater than about 10 (10). An example of gear structure 48 is an epicyclic gear train, such as a planetary gear system, a star gear system, or other known gear system, having a gear reduction ratio greater than about 2.3.

  In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about 10 (10: 1) and the fan diameter is significantly larger than the outer diameter of the low pressure compressor 44. However, it should be understood that the above parameters are merely exemplary of one embodiment of a gas turbine engine that includes a gear structure, and that the present disclosure is applicable to other gas turbine engines.

  A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is typically designed for specific flight conditions of about 0.8 Mach and about 35,000 feet of cruise. The engine's highest fuel economy at 0.8 Mach and 35,000 feet flight conditions, also known as "Fuel consumption per bucket cruise thrust" ("TSFC"), is the pound of thrust that the engine produces at its lowest point. The industry standard parameter of pounds of fuel (lbm) per hour burned divided by force (lbf).

  “Low fan pressure ratio” is the pressure ratio before and after a single fan blade without a fan outlet guide vane (“FEGV”) system. The low fan pressure ratio as disclosed herein in accordance with one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment, the low fan pressure ratio is less than about 1.45.

  “Low correction fan tip speed” is the actual fan tip speed in feet / second divided by the industry standard temperature correction value of [(Tram ° R) / (518.7 ° R)] 0.5. The “low correction fan tip speed” as disclosed herein in accordance with one non-limiting embodiment is less than about 1150 feet / second.

  An example gas turbine engine includes a fan 42 that in one non-limiting embodiment includes less than about 26 fan blades. In another non-limiting embodiment, the fan portion 22 includes less than about 20 fan blades. Also, in one disclosed embodiment, the low pressure turbine 46 includes only about six turbine rotors, indicated generally at 34. In another non-limiting example embodiment, the low pressure turbine 46 includes approximately three turbine rotors. The ratio of the number of fan blades 42 to the number of low pressure turbine rotors is about 3.3 to about 8.6. The example of the low pressure turbine 46 provides the driving force to rotate the fan section 22 and thus the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 has been increased. An example of a gas turbine engine 20 having power transmission efficiency is disclosed.

With continued reference to FIG. 1 and with reference to FIG. 2, the center of gravity indicated at 62 has been moved axially to the engine centerline located along axis A. The center of gravity 62 is located at a first distance _XCG between the front end and the rear end of the engine 20. In the disclosed embodiment, the forward end is the leading edge 66 of the fan blade 42 and the trailing end is the trailing edge 68 of the rearmost rotating turbine blade 70. A total length X _L is defined between the leading edge 66 and the trailing edge 68.

The position of the center of gravity 62 along axis A is located at a first distance _XCG and affects the configuration for supporting the engine on the fuselage and is therefore of interest to engine manufacturers and aircraft designers. .

  In the disclosed embodiment, a mounting structure shown schematically at 64 supports an example of a gas turbine engine 20 on a fuselage (not shown). It should be understood that the position of the mounting structure 64 may vary for each engine application, and such variations are within the scope of the present disclosure.

  The example gas turbine engine 20 includes a gear structure 48 for driving the fan portion 22 at a different speed than the fan driven turbine. In this example, the fan driven turbine is a low pressure turbine 46.

  The selection of material and component configuration determines the total weight of the engine 20 along with its weight distribution and determines the position of the center of gravity 62. Structures such as fan section 22 and gear structure section 48 together with compressor section 24 and turbine section 28 define a weight distribution that determines the position of center of gravity 62 as well as the total weight of engine 20.

  The gear structure 48 is a structure located in front of the engine 20 together with the fan part 22. Therefore, the selection of materials and the configuration of the structure influence the position of the center of gravity 62. Also, many structures within the gas turbine engine structure take into account and partially determine the location and position of the center of gravity 62.

In this disclosed example, components of the gear structure 48 such as, for example, journal bearings, lubrication jets, and gutta around the epicyclic component define the total weight of the gear structure 48 and thereby the center of gravity. It is chosen to provide light weight and axial miniaturization taking into account 62 definitions. In one disclosed example, the weight of the gear structure is described as a density to relate the overall size or volume of the gear structure to the weight. In this example, the example gear structure has a density of about 6089.58 kg / m ³ ( about 0.22 lb / in ³ ) to about 8303.97 kg / m ³ ( about 0.30 lb / in ³ ). It is a gear box. Gearbox part density is one consideration that affects the position of the center of gravity 62.

  Further, the fan section 22 is one of the foremost components of the engine example, and therefore also has a significant impact on the final position of the center of gravity 62 in the finished engine 22. Therefore, selection of the parts constituting the fan unit 22 is considered not only from the viewpoint of propulsion efficiency but also with respect to the overall structure of the gas turbine engine 20.

  In the disclosed example, fan portion 22 includes a rotor 72 that supports a plurality of fan blades 42 for rotation about axis A. The arrangement of the rotor 72 radially folded around the bearing 38 provides a weight advantage, thereby providing another means for modifying and positioning the engine's center of gravity 62. The bearing 38 may also include a tapered roller roller bearing that further provides a beneficial effect on the structure of the fan portion without adding additional weight to further affect the position of the center of gravity 62.

The density of the fan rotor may be further selected to utilize a lightweight structure that further corresponds to and achieves the position of the center of gravity 62. In this example, the fan unit 22 and a plurality of fan blades, comprising a rotor 72, together about 260.19kg / m ³ (about 0.0094 lbs / square inch) to about 426.27kg / m ³ (about 0. providing density in the range of 01,540 lbs / cubic inch).

  Furthermore, the fan housing case 16 needs to house the blades 42 and is a choice that is manufactured from composite material to reduce weight and determine the position of the center of gravity 62.

  Utilizing these and other configuration parameters and material selection options, the example geared turbofan engine has an internal moment arm that is a measure of the position of the engine's center of gravity 62. Examples of internal moment arms are in the range of about 40% to 70% of the length of the engine 20 according to the following relationship:

  The internal moment arm MA is defined as follows.

In this example, a MA = X _CG / X _L.

Thus, in a geared turbofan gas turbine engine, the position of the center of gravity 62 is related to the moment arm according to the above equation and the total length X between the leading edge 66 of the fan blade 42 and the trailing edge 68 of the rearmost rotating turbine blade 70. Within about 35% to about 75% of _L. In another disclosed embodiment, an example of a moment arm MA for example of the engine 20 is about 40% of the total length X _L between the trailing edge 68 of the leading edge 66 and rearwardmost rotary turbine blades 70 of the fan blades 42 In the range of ~ 70%.

The position of the center of gravity 62 in the disclosed geared gas turbine engine is approximately 30% longer than in prior art direct drive turbine engines. Table 1 below includes examples of further disclosed embodiments of the moment arm MA of a geared turbofan engine.

The disclosed moment arm MA is a percentage of the engine length X _L (FIG. 1) measured between the leading edge 66 of the fan blade 72 and the trailing edge 68 of the last rotating turbine blade 70 in the low pressure turbine 46. Represents the position of the center of gravity 62.

  Referring to FIG. 2, complex calculations can be used to determine the physical location of the center of gravity 62 of any engine, but there are other means of determining this location in any engine. When the engine centerline or axis A is perpendicular to a vertical line 76 that intersects the axis A through the lift point 74, the lift point 74 located on the engine 20 is positioned at the center of gravity 62 in the axial direction. . Thus, the center of gravity 62 can be determined, as an example, as the point where the vertical line 76 intersects the engine centerline or axis A at right angles when supported by a single lifting point 74.

  The actual center of gravity 62 may be slightly offset from axis A due to locally attached accessories, but an example of engine center of gravity 62 is considered along the engine centerline or axis A.

  The disclosed center of gravity 62 includes fluid contained within the operating system of the turbofan engine 20. However, examples of centroids do not include some structures such as those from a fuselage that would have little effect on the position of gravity 62, such as typical tubes, brackets, and harnesses. Also, the example of the center of gravity 62 is determined only for the bare engine that does not include the engine mount, fan cowl, reverse thrust device, inlet, nozzle, and plug. In other words, the example of the center of gravity 62 is not determined including the weight of the structure constituting the gas turbine engine such as the engine mounting structure, the engine cowling structure, and the nacelle structure.

  Thus, through selection of structural materials and designs such as fan and gear structure, the center of gravity 62 can be located at a structurally desirable location to improve propulsion efficiency and reduce mounting structure requirements.

  While one embodiment of the invention has been disclosed, those skilled in the art will recognize that certain modifications will fall within the scope of the disclosure. As such, the following claims should be studied to determine the scope and content of this disclosure.

Claims (12)

A gas turbine engine,
A plurality of fan blades rotatable about an axis, each of which includes a leading edge;
A turbine section including a rearmost turbine blade having a trailing edge;
A gear structure that is driven by the turbine to rotate the plurality of fan blades about the axis, and
A center of gravity of the gas turbine engine is located at a first axial distance from the trailing edge of the rearmost turbine blade, the first axial distance being determined by the front edge and the rearmost edge of the plurality of fan blades. are three 5% to 7 5% der of the total length between the trailing edge of the turbine blade,
The plurality of fan blades are supported on a rotor, the rotor and the fan blades having a density of 260.19 kg / m ^{3 to} 426.27 kg / m ³ ;
The gear structure comprises a gearbox having a density of ^{6089.58kg / m 3 ~8303.97kg / m 3} , the gas turbine engine.   The gas turbine engine of claim 1, wherein the center of gravity is disposed substantially along the axis.   The gas turbine engine according to claim 1, wherein the center of gravity is determined including a weight of a structure constituting the gas turbine engine not including an engine mounting structure, an engine cowling structure, and a nacelle structure.   The gas turbine engine of claim 3, wherein the center of gravity is determined including a weight of fluid contained within an operating system of the gas turbine engine. The first axial distance is 40 % to 70 % of the total length between the leading edge of the plurality of fan blades and the trailing edge of the rearmost turbine blade. Gas turbine engine.   The gas turbine engine of claim 1, wherein the center of gravity is located at the intersection of a vertical line extending through a lifting point of the gas turbine engine and the axis, the axis being perpendicular to the vertical line. A turbofan engine,
A plurality of fan blades rotatable about an axis, each of which includes a leading edge;
A turbine section including a rearmost rotating turbine blade having a trailing edge;
A gear structure that is driven by the turbine to rotate the plurality of fan blades about the axis, and
An internal moment arm of the turbofan engine from a center of gravity of the turbofan engine relative to a total length between the leading edge of the plurality of fan blades and the trailing edge of the rearmost turbine blade; is the ratio of the first distance to the trailing edge, Ri this ratio is 35% to 7 5% der,
The gear structure comprises a gear box having a density of 6089.58 kg / m ^{3 to} 8303.97 kg / m ³ ;
A turbofan engine comprising a rotor supporting the plurality of fan blades, wherein the rotor and fan blades have a density of 260.19 kg / m ^{3 to} 426.27 kg / m ³ . The turbofan engine according to claim 7 , wherein the ratio is 40 % to 70%. The turbofan engine according to claim 7 , wherein the center of gravity is determined including a weight of fluid contained in an operating system of the turbofan engine. The turbofan engine according to claim 7 , wherein the center of gravity is located at an intersection of a vertical line extending through a lifting point of the turbofan engine and the axis, and the axis is perpendicular to the vertical line. A method of assembling a gas turbine engine comprising:
Supporting a plurality of fan blades each including a leading edge about a rotational axis;
Supporting a turbine portion including a rearmost turbine blade having a trailing edge about the rotational axis;
Supporting a gear structure driven by the turbine to rotate the plurality of fan blades about the axis;
Selecting a component of the gas turbine engine structure to place a center of gravity of the gas turbine engine at a first axial distance from the trailing edge of the rearmost turbine blade; and axial distance, are three 5% to 7 5% der of the total length between the trailing edge of the said leading edge of said plurality of fan blades rearmost turbine blades,
Supporting the plurality of fan blades on a rotor, the rotor and fan blades having a density of 260.19 kg / m ^{3 to} 426.27 kg / m ³ ;
Assembling the gear structure as a gear box having a density of 6089.58 kg / m ^{3 to} 8303.97 kg / m ³ . The gas turbine engine is configured to place the center of gravity within a range of 40 % to 70 % of a total length between the leading edge of the plurality of fan blades and the trailing edge of the rearmost turbine blade. The method of claim 11 , comprising selecting a part.

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