US20150300264A1

US20150300264A1 - Geared turbofan architecture for regional jet aircraft

Info

Publication number: US20150300264A1
Application number: US14/188,733
Authority: US
Inventors: Frederick M. Schwarz
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 2013-09-30
Filing date: 2014-02-25
Publication date: 2015-10-22
Also published as: US20160195022A1; US10794291B2; EP3052790A1; EP3052790A4; WO2015073076A1

Abstract

A gas turbine engine comprises an upstream compressor rotor and a downstream compressor rotor, each having a first stage blade row and a tip speed defined at the first stage blade row. A ratio of the tip speed of the first stage of the downstream compressor rotor to the tip speed of the first stage in the upstream compressor rotor is equal to or above about 1.18.

Description

RELATED APPLICATION

This application claims priority to U.S. Provisional Application 61/884,230, filed Sep. 30, 2013.

BACKGROUND OF THE INVENTION

This application relates to a geared turbofan engine which may be particularly beneficial for application on regional jet aircraft.
Gas turbine engines are known and, typically, include a fan delivering air into a compressor and into a bypass duct as propulsion air. Air in the compressor is compressed and delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate.
Historically, a turbine rotor drove an upstream compressor rotor and a fan rotor at a single speed.
More recently, it has been proposed to include a gear reduction between the fan rotor and the upstream compressor rotor such that the fan can rotate at slower speeds. This has provided a great deal of freedom to the designer of gas turbine engines.
To date, there has been little activity in tailoring of geared gas turbine engines to the particular application and aircraft which will utilize the gas turbine engine.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises an upstream compressor rotor and a downstream compressor rotor, each having a first stage blade row and a tip speed defined at the first stage blade row. A ratio of the tip speed of the downstream compressor rotor first stage blade row to the tip speed of the upstream compressor rotor first stage blade row is equal to or above about 1.18.
In another embodiment according to the previous embodiment, the upstream compressor rotor has an upstream number of stages and the downstream compressor rotor has a downstream number of stage. A ratio of the downstream number of stages to the upstream number of stages is greater than or equal to about 1.6 and less than or equal to about 3.4.
In another embodiment according to any of the previous embodiments, a compression ratio is defined across the upstream and downstream compressor rotors with the compression ratio being greater than or equal to about 38.
In another embodiment according to any of the previous embodiments, a fan rotor has less than or equal to 26 blades.
In another embodiment according to any of the previous embodiments, the turbine section includes a lower pressure turbine that drives the upstream compressor rotor, and has three or four stages.
In another embodiment according to any of the previous embodiments, the lower pressure turbine also drives the fan rotor.
In another embodiment according to any of the previous embodiments, the lower pressure turbine drives the fan rotor through a gear reduction.
In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to about 2.6.
In another embodiment according to any of the previous embodiments, a pressure ratio across the fan rotor is less than or equal to about 1.45.
In another embodiment according to any of the previous embodiments, the gear reduction is a planet-type epicyclic gearbox.
In another embodiment according to any of the previous embodiments, the gas turbine engine is designed for use on a single aisle aircraft.
In another embodiment according to any of the previous embodiments, a compression ratio is defined across the upstream and downstream compressor rotors with the compression ratio being greater than or equal to about 38.
In another embodiment according to any of the previous embodiments, the gas turbine engine is designed for use on a single aisle aircraft.
In another embodiment according to any of the previous embodiments, the compression ratio is less than or equal to about 55.
In another embodiment according to any of the previous embodiments, the turbine section includes a lower pressure turbine that drives the upstream compressor rotor, and has three or four stages.
In another embodiment according to any of the previous embodiments, the lower pressure turbine also drives the fan rotor.
In another embodiment according to any of the previous embodiments, the fan rotor is driven through a gear reduction.
In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to about 2.6.
In another embodiment according to any of the previous embodiments, the gear ratio of the gear reduction is less than or equal to about 3.3.
In another embodiment according to any of the previous embodiments, a pressure ratio across a fan rotor is less than or equal to about 1.45.
In another embodiment according to any of the previous embodiments, the gas turbine engine is designed for use on a single aisle aircraft.
In another embodiment according to any of the previous embodiments, the ratio of the tip speed of the first stage of the downstream compressor rotor to the tip speed of the first stage in the upstream compressor rotor is equal to or below about 1.43.
In another embodiment according to any of the previous embodiments, a fan rotor has less than or equal to 26 blades.
In another featured embodiment, a gas turbine engine comprises an upstream compressor rotor and a downstream compressor rotor. Each of the compressor rotors has a first stage blade row and a tip speed defined at the first stage blade row. A ratio of the tip speed of the downstream compressor rotor first stage blade row to the tip speed of the upstream compressor rotor first stage blade row is equal to or below about 1.43.
In another embodiment according to the previous embodiment, the upstream compressor rotor has an upstream number of stages and the downstream compressor rotor has a downstream number of stages. A ratio of the downstream number of stages to the upstream number of stages is greater than or equal to about 1.6 and less than or equal to about 3.4.
In another embodiment according to any of the previous embodiments, a compression ratio is defined across the upstream and downstream compressor rotors with the compression ratio being greater than or equal to about 38.
In another embodiment according to any of the previous embodiments, a fan rotor has less than or equal to 26 blades.
In another embodiment according to any of the previous embodiments, the fan rotor is driven through a gear reduction.
In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to about 2.6.
In another embodiment according to any of the previous embodiments, the gear ratio of the gear reduction is less than or equal to about 3.3.
These and other features may be best understood from the following drawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A schematically shows a gas turbine engine.

FIG. 1B schematically shows a regional jet aircraft.

FIG. 2 shows parameters of one example engine.

DETAILED DESCRIPTION

FIG. 1A 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 two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.
The exemplary 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, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as 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 second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 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.
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 C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
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 about 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 five. 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 five 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 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of 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.
FIG. 1B schematically shows what may be called a regional jet aircraft 10 that mounts engines 20. A regional aircraft may be defined in a number of ways, however, one way is that they typically have a single aisle 12 between passenger sections 11. Another way may be a total flight length of less than 3000 nautical miles.
The engines utilized in such an aircraft have a particularly high percentage of time in take-off and climb relative to cruise. That is, compared to jet engines utilized on longer range aircraft, the engines 20 mounted on aircraft 10 will be spending more time at take-off and cruise conditions which are particularly challenging on components within the engine. Thus, there are elevated temperatures adjacent a last stage of the downstream or high pressure compressor 32 and elevated turbine cooling air temperatures utilized to cool components in the turbine sections. This disclosure tailors the gas turbine engine 20 such that it is uniquely structured to address the challenges faced by such an engine.
In such an engine, there may be a lower overall pressure ratio and a lower temperature at a downstream end of the downstream compressor rotor 52. The overall pressure ratio across the upstream compressor rotor 44 and the downstream compressor rotor 52 may, in some embodiments, be greater than or equal to 38, and may, in some embodiments, be less than or equal to 55. Turbine cooling airflow is typically taken from the downstream compressor rotor, and may be at a lower temperature, consistent with a moderate temperature at a downstream end of the downstream compressor rotor. The turbine cooling air temperature will still be high given the frequent occurrence of take-off and climb temperature exposure on turbine airfoils.
A greater portion of the total compression work is shifted to the downstream compressor rotor 52 relative to the upstream compressor rotor 44. Although, the lower pressure spool 30 may be more efficient than the higher pressure spool 32 doing more work with the downstream compressor rotor 52 reduces the temperature of the gas at the exit of the high pressure turbine section 54, thus, reducing the temperature reaching the low pressure turbine 46. The first blade of the low pressure turbine may be cast with a directionally solidified material or even a single crystal material.
Since there is more work done with the downstream compressor rotor 52, a ratio of compressor stages for the downstream compressor rotor 52 compared to the number of stages in the upstream compressor rotor 44 may be greater than or equal to about 1.6 and less than or equal to about 3.4. In one embodiment a ratio of the stages in the downstream compressor rotor 52 to those in the upstream compressor rotor was 2.7. The speed of the low pressure turbine 46 and the fan 42 may be closer than in other engines. Thus, a gear ratio of the gear reduction 48 may be greater than or equal to about 2.6 and less than or equal to about 3.3.
The low pressure turbine 46 may have only three or four stages since it is doing less work.
The gear reduction 48 may be a planet-type epicyclic gearbox where a sun and ring gear revolve around an engine centerline, whereas the intermediate planet gears rotate on stationary axes but do not revolve around the sun.
The downstream compressor rotor 52 speed is higher relative to the upstream compressor 44 speed since there is a greater amount of work performed at the downstream compressor rotor 52. Again, this reduces the temperature reaching the low pressure turbine 46. As an example, a redline tip speed of a first stage rotor in the downstream compressor rotor 52 to the redline tip speed of a first stage rotor in the upstream compressor rotor 44 is greater than or equal to about 1.18 and less than or equal to about 1.43 with the speeds measured in feet/second.
The higher pressure turbine rotor 54 may have two or three turbine sections.
The fan may be greater than about 73 inches in diameter and have fewer than about 26 blades. A pressure ratio across the fan may be less than about 1.45 and in some embodiments may be greater than or equal to about 1.35 and less than or equal to about 1.45.
FIG. 2 tabulates several variables, on one example engine, made according to the teachings of this disclosure.
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 gas turbine engine comprising:

an upstream compressor rotor and a downstream compressor rotor, each of said compressor rotors having a first stage blade row and a tip speed defined at said first stage blade row, with a ratio of said tip speed of said downstream compressor rotor first stage blade row to the tip speed of said upstream compressor rotor first stage blade row being equal to or above about 1.18; and

said gas turbine engine is designed for use on a regional aircraft defined as having at least one of a single aisle between passenger sections or a total flight length of less than 3,000 nautical miles.

2. The gas turbine engine section as set forth in claim 1, wherein said upstream compressor rotor having an upstream number of stages and said downstream compressor rotor having a downstream number of stages with a ratio of said downstream number of stages to said upstream number of stages being greater than or equal to about 1.6 and less than or equal to about 3.4.

3. The gas turbine engine section as set forth in claim 2, wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38.

4. The gas turbine engine as set forth in claim 3, wherein a fan rotor having less than or equal to 26 blades.

5. The gas turbine engine as set forth in claim 3, wherein a turbine section including a lower pressure turbine driving said upstream compressor rotor, said lower pressure turbine having three or four stages.

6. The gas turbine engine as set forth in claim 5, wherein said lower pressure turbine also driving said fan rotor.

7. The gas turbine engine as set forth in claim 6, wherein said lower pressure turbine driving said fan rotor through a gear reduction.

8. The gas turbine engine as set forth in claim 7, wherein a gear ratio of said gear reduction being greater than or equal to about 2.6.

9. The gas turbine engine as set forth in claim 7, wherein a pressure ratio across said fan rotor is less than or equal to about 1.45.

10. The gas turbine engine as set forth in claim 7, wherein said gear reduction is a planet-type epicyclic gearbox.

11. The gas turbine engine as set forth in claim 10, wherein said gas turbine engine is designed for use on a single aisle aircraft.

12. The gas turbine engine as set forth in claim 1, wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38.

13. The gas turbine engine as set forth in claim 12, wherein said gas turbine engine is designed for use on a single aisle aircraft.

14. The gas turbine engine as set forth in claim 12, wherein said compression ratio being less than or equal to about 55.

15. The gas turbine engine as set forth in claim 1, wherein a turbine section including a lower pressure turbine driving said upstream compressor rotor, said lower pressure turbine having three or four stages.

16. The gas turbine engine as set forth in claim 15, wherein said lower pressure turbine also driving a fan rotor.

17. The gas turbine engine as set forth in claim 16, wherein said fan rotor is driven through a gear reduction.

18. The gas turbine engine as set forth in claim 17, wherein a gear ratio of said gear reduction being greater than or equal to about 2.6.

19. The gas turbine engine as set forth in claim 18, wherein the gear ratio of said gear reduction being less than or equal to about 3.3.

20. The gas turbine engine as set forth in claim 17, wherein a pressure ratio across said fan rotor is less than or equal to about 1.45.

21. The gas turbine engine as set forth in claim 1, wherein said gas turbine engine is designed for use on a single aisle aircraft.

22. The gas turbine engine as set forth in claim 1, wherein said ratio of said tip speed of the first stage of the downstream compressor rotor to the tip speed of the first stage in the upstream compressor rotor being and equal to or below about 1.43.

23. The gas turbine engine as set forth in claim 1, wherein a fan rotor has less than or equal to 26 blades.

24. A gas turbine engine comprising:

an upstream compressor rotor and a downstream compressor rotor, each of said compressor rotors having a first stage blade row and a tip speed defined at said first stage blade row, with a ratio of said tip speed of said downstream compressor rotor first stage blade row to the tip speed of said upstream compressor rotor first stage blade row being equal to or below about 1.43, and said ratio being greater than or equal to about 1.18, and a fan rotor is included upstream of said upstream compressor rotor, and said fan rotor being driven by a turbine rotor through a gear reduction.

25. The gas turbine engine section as set forth in claim 24, wherein said upstream compressor rotor having an upstream number of stages and said downstream compressor rotor having a downstream number of stages with a ratio of said downstream number of stages to said upstream number of stages being greater than or equal to about 1.6 and less than or equal to about 3.4.

26. The gas turbine engine section as set forth in claim 25, wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38.

27. The gas turbine engine as set forth in claim 24, wherein the fan rotor has less than or equal to 26 blades.

28. (canceled)

29. The gas turbine engine as set forth in claim 27, wherein a gear ratio of said gear reduction being greater than or equal to about 2.6.

30. The gas turbine engine as set forth in claim 29, wherein the gear ratio of said gear reduction being less than or equal to about 3.3.