WO2014204526A2

WO2014204526A2 - Family of geared turbo fan engines

Info

Publication number: WO2014204526A2
Application number: PCT/US2014/015500
Authority: WO
Inventors: Karl HASEL; Jessica TSAY
Original assignee: United Technologies Corporation
Priority date: 2013-02-17
Filing date: 2014-02-10
Publication date: 2014-12-24
Also published as: WO2014204526A3

Abstract

A method of configuring a plurality of substantially mutually distinct gas turbine engines includes the steps of providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors. Each include a fan and a fan drive gear architecture. Each of the engines is provided with respective ones of a plurality of substantially mutually alike gas generators, each with a compressor section, combustor, and a turbine section configured to drive at a common speed. The plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.

Description

FAMILY OF GEARED TURBO FAN ENGINES

CROSS-REFERENCE TO RELATED APPLICATION

[0001] This application claims priority to U.S. Provisional Application No. 61/765,738, filed February 17, 2013.

BACKGROUND OF THE INVENTION

[0002] This application relates to a method of utilizing a single gas turbine engine propulsor core with distinct gear drives and fan diameters.

[0003] Gas turbine engines are known and, when utilized on aircraft, typically, include a fan delivering air into a bypass duct as propulsion air and also delivering air into a compressor.

[0004] In a two spool gas turbine engine, there are typically low and high pressure compressor rotors. Air is compressed in the low pressure compressor rotor, delivered into the high pressure compressor rotor, and then delivered into a combustor. The air is mixed with fuel and ignited in the combustor and products of that combustion pass downstream over a turbine section. A high pressure turbine drives the high pressure compressor. Historically, a low pressure turbine, downstream of the high pressure turbine, had a direct drive connection to both the low pressure compressor and the fan.

[0005] There are trade-offs between the speed of the fan and the speed of the low pressure turbine. For many reasons it would be desirable to have the low pressure turbine rotate at higher speeds, but the fan rotor cannot always rotate at such speeds. This is a limitation on a direct drive gas turbine engine.

[0006] More recently, it has been proposed to include a gear reduction between the low pressure turbine and the fan. The gear reduction allows the fan to rotate at slower speeds than does the low pressure compressor and low pressure turbine. This enables the fan to be designed to have a greater diameter and provides great freedom to a gas turbine engine designer.

[0007] With a direct drive gas turbine engine, the low pressure turbine, low pressure compressor, and fan all needed to be designed together. That is, as the diameter of the fan in a direct drive turbine changed, so too must the design of the low pressure turbine and the low pressure compressor as their speeds were all the same. SUMMARY OF THE INVENTION

[0008] In a featured embodiment, a method of configuring a plurality of substantially mutually distinct gas turbine engines including the steps of providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors, each including a fan and a fan drive gear architecture. Each of the engines with respective ones of a plurality of substantially mutually alike gas generators provide each with a compressor section, combustor, and a turbine section configured to drive at a common speed, the compressor section and the fan drive gear architecture. The plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.

[0009] In another embodiment according to the previous embodiment, a gas generator has a compressor section with first and second compressors. The turbine section has a first and second turbine. The second turbine drives the first compressor and the fan drive gear architecture.

[0010] In another embodiment according to any of the previous embodiments, the plurality of mutually distinct gas turbine engines produce substantially the same thrust.

[0011] In another embodiment according to any of the previous embodiments, the mutually distinct gas turbine engines capable of producing mutually distinct thrust.

[0012] In another embodiment according to any of the previous embodiments, the plurality of substantially mutually distinct gas turbine engines produce a common overall pressure ratio.

[0013] In another featured embodiment, a method of configuring a plurality of substantially mutually distinct gas turbine engines, includes the steps of providing each of the engines with respective ones of a plurality of substantially mutually distinct gas generators, each with a compressor section, combustor, and a turbine section. Each of the engines is provided with respective ones of a plurality of substantially mutually alike propulsors, each including a fan and a fan drive gear architecture. The turbine section is configured to drive the compressor section and the fan drive gear architecture at a common speed.

[0014] In another featured embodiment, a method of providing a family of gas turbine engines includes the steps of designing a gas turbine engine, a fan having a first diameter, a first gear reduction having a first gear ratio, a low pressure compressor, a high pressure compressor, a combustor, a low pressure turbine, and a high pressure turbine. The low pressure turbine is designed to drive the low pressure compressor at a common speed and to drive the fan through the gear reduction. A distinct gas turbine engine is provided by selecting a second gear ratio for a second gear reduction, and a second fan diameter for a fan, and placing the second gear reduction having the second gear ratio and the fan rotor having the second fan diameter into the distinct gas turbine engine, such that the low pressure turbine now drives the second gear reduction having the second gear ratio and, in turn, the fan having the second fan diameter.

[0015] In another embodiment according to the previous embodiment, the second fan is less than the first fan diameter and the second gear ratio is higher than the first gear ratio.

[0016] In another featured embodiment, a family of gas turbine engines has a low pressure compressor to be driven at a common speed with a low pressure turbine and an output shaft also driven by the low pressure turbine. A high pressure compressor is driven by a high pressure turbine, and a combustor. A plurality of gear reductions has different gear ratios to be driven by the output shaft. A plurality of fan rotors has different diameters to be driven by a selected gear reduction. A particular gear reduction is selected in combination with the selection of a particular fan rotor diameter.

[0017] These and other features may be best understood from the following drawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 schematically shows a gas turbine engine.

[0019] Figure 2A shows a first embodiment of a method according to this invention.

[0020] Figure 2B shows a second embodiment.

[0021] Figure 3 shows yet another embodiment.

DETAILED DESCRIPTION

[0022] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section

22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non- limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

[0023] The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

[0024] The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

[0025] The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.

[0026] The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other geared gas turbine engines.

[0027] 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 designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)]^0'5. The "Low corrected fan tip speed" as disclosed herein according to one non- limiting embodiment is less than about 1150 ft / second.

[0028] With the development of the above disclosed geared gas turbine engine, one need not redesign the low pressure turbine 46 and the low pressure compressor 44 to have a distinct fan diameter and other fan structure. Rather, the one configuration of a low pressure turbine 46 and low pressure compressor 44, for example, from the embodiment shown in Figure 1, can be utilized across a family of gas turbine engines having distinct fan diameters and fan structures and thrust levels. This can be accomplished by changing the gear ratio of the gear reduction 48. [0029] Thus, as shown in Figure 2A, a gas turbine engine 100 has a core 102, which includes the low pressure compressor 44, the high pressure compressor 52, the combustor 26, the high pressure turbine 54 and the low pressure turbine 46, such as in the Figure 1 engine. Output shaft 104 is a shaft driven by the low pressure turbine 46. The gear reduction 106 is selected in combination with a diameter Di of a fan 108, such that the rotational speed of the fan 108 is as desired given the input speed of the shaft 104. In general, gear ratios equal to or between about 2.4 to 4.2 can be utilized. Also, bypass ratios greater than or equal to 6, and greater than or equal to 8, and greater than or equal to 10, and with pressure ratios of less than or equal to 1.52 may be achieved. Overall pressure ratios of greater than 30: 1, and even greater than or about 50: 1 may be achieved.

[0030] On the other hand, when a second engine 110 is configured for the same or different thrust, but a different fan diameter and/or fan pressure ratio, the substantially same configuration of the core 102 and the output shaft 104 can be utilized. However, a gear reduction 112 is selected in combination with the diameter D₂ and a similar or different fan pressure ratio of a fan 114 to achieve a speed for the fan which is as desired for better efficiency and noise control. As shown in Figure 2B, the diameter D₂ of the fan 114 is smaller than the diameter of the fan 108, providing a smaller bypass ratio. In such an engine, the gear reduction ratio of gear 112 would generally be less than the gear ratio of the gear reduction 106.

[0031] A disclosed method of configuring a plurality of substantially mutually distinct gas turbine engines 100,110 includes providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors, each including a fan and a fan drive gear architecture 108/106 and 114/112.

[0032] Each of the engines 100, 110 is provided with respective ones of a plurality of substantially mutually alike gas generators 102, each with a compressor section, combustor, and a turbine section configured to drive at a common speed the compressor section and the fan drive gear architecture. The plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.

[0033] The present invention, thus, allows a dramatic reduction in the design, development and test costs and the manufacturing cost for creating a family of gas turbine engines having different fans. [0034] Figure 3 shows another aspect, where an engine 200 can be constructed with a single propulsor 202 having a fan and a fan drive gear architecture, which can be associated with a family of distinct gas generators 204a and 204b. The benefits mentioned above would be similar from this arrangement.

[0035] In addition, the embodiments of Figures 2A/2B and Figure 3 can be utilized to provide engines having a common thrust across the family, or having distinct thrust.

[0036] Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims

1. A method of configuring a plurality of substantially mutually distinct gas turbine engines, comprising the steps of:

providing each of the engines with respective ones of a plurality of substantially mutually distinct propulsors, each including a fan and a fan drive gear architecture;

and

providing each of the engines with respective ones of a plurality of substantially mutually alike gas generators, each with a compressor section, combustor, and a turbine section configured to drive at a common speed the compressor section and the fan drive gear architecture;

wherein the plurality of mutually distinct propulsors have a mutually distinct diameter fan and/or fan drive architecture with distinct gear reduction ratio.

2. The method as set forth in claim 1, wherein said gas generator has a compressor section with first and second compressors, the turbine section has a first and second turbine, and a second turbine drives the first compressor and the fan drive gear architecture.

3. The method as set forth in claim 1, wherein the plurality of mutually distinct gas turbine engines produce substantially the same thrust.

4. The method as set forth in claim 1 , wherein the mutually distinct gas turbine engines capable of producing mutually distinct thrust.

5. The method as set forth in claim 1, wherein the plurality of substantially mutually distinct gas turbine engines produce a common overall pressure ratio.

6. A method of configuring a plurality of substantially mutually distinct gas turbine engines, comprising the steps of:

providing each of the engines with respective ones of a plurality of substantially mutually distinct gas generators, each with a compressor section, combustor, and a turbine section; and

providing each of the engines with respective ones of a plurality of substantially mutually alike propulsors, each including a fan and a fan drive gear architecture, the turbine section configured to drive the compressor section and the fan drive gear architecture at a common speed.

7. A method of providing a family of gas turbine engines comprising the steps of:

designing a gas turbine engine, including designing a fan having a first diameter, a first gear reduction having a first gear ratio, a low pressure compressor, a high pressure compressor, a combustor, a low pressure turbine, and a high pressure turbine, wherein the low pressure turbine is designed to drive the low pressure compressor at a common speed and to drive the fan through the gear reduction; and

providing a distinct gas turbine engine by selecting a second gear ratio for a second gear reduction, and a second fan diameter for a fan, and placing the second gear reduction having the second gear ratio and the fan rotor having the second fan diameter into the distinct gas turbine engine, such that the low pressure turbine now drives the second gear reduction having the second gear ratio and, in turn, the fan having the second fan diameter.

8. The method as set forth in claim 7, wherein if the second fan is less than the first fan diameter and the second gear ratio is higher than the first gear ratio.

9. A family of gas turbine engines comprising:

a low pressure compressor to be driven at a common speed with a low pressure turbine and an output shaft also driven by the low pressure turbine, a high pressure compressor driven by a high pressure turbine, and a combustor; and

a plurality of gear reductions having different gear ratios to be driven by said output shaft and a plurality of fan rotors having different diameters to be driven by a selected one of said gear reductions, with a particular gear reduction selected in combination with the selection of a particular fan rotor diameter.