EP3084173A2

EP3084173A2 - Geared turbofan with three turbine sections

Info

Publication number: EP3084173A2
Application number: EP14880333.1A
Authority: EP
Inventors: Frederick M. Schwarz; Joseph Brent Staubach
Original assignee: United Technologies Corp
Current assignee: RTX Corp
Priority date: 2013-12-16
Filing date: 2014-11-07
Publication date: 2016-10-26
Also published as: WO2015112231A3; WO2015112231A2; EP3084173A4

Abstract

A gas turbine engine comprises a fan rotor configured to be driven by a fan drive turbine through a first shaft and a gear reduction. The fan rotor is configured to deliver air into a bypass duct as bypass air and to deliver core air flow into a core engine where it reaches an upstream compressor rotor. The upstream compressor rotor is configured to be driven through a second shaft by an intermediate turbine rotor. A downstream compressor rotor is configured to be driven by an upstream turbine rotor through a third shaft. An overall pressure ratio across the upstream and downstream compressor rotors is greater than or equal to about 35.0 and less than or equal to about 75.0.

Description

GEARED TURBOFAN WITH THREE TURBINE SECTIONS

CROSS-REFERENCE TO RELATED APPLICATION

[0001] This application claims priority to United States Patent Application No. 14/107,169, filed 16 December 2013.

BACKGROUND OF THE INVENTION

[0002] Gas turbine engines are known and typically include a fan delivering air as bypass flow for propulsion within a nacelle. In addition, the fan often delivers core air flow into a core engine where it is delivered into an upstream compressor rotor. The air may be compressed at the upstream compressor rotor and then delivered into a downstream compressor rotor where it may be further compressed to higher pressures. The air from the downstream compressor rotor may be delivered into a combustion section where may be is mixed with fuel and ignited. Products of this combustion can pass downstream over turbine rotors driving them to rotate. The turbine rotors, in turn, can rotate the compressor rotors and the fan rotor.

[0003] In one known type of gas turbine engine, there are three turbine rotors. A first turbine rotor drives the downstream or higher pressure compressor rotor. An intermediate turbine rotor drives an upstream or lower pressure compressor rotor. A third turbine rotor is a fan drive turbine and drives the fan rotor. All of these turbine rotors can have one stage or more than one stage, except for the fan drive turbine which for reasons of efficiency and in order to provide a reasonable fan tip speed have conventionally had many stages. Historically, the fan drive turbine has driven the fan rotor at a common speed.

[0004] It would be desirable to increase the overall pressure ratio across the upstream and downstream compressor rotors. However, there are challenges relating to the strength of available materials and the capability to withstand the stress from thermal gradients arising from engine transients and steady stress arising from centrifugal pull and the material property degradation with increasing temperature. One particular challenge is that the temperature and speed of the downstream compressor rotor limits the overall pressure ratio which can be achieved. The stresses and challenges are greatest at a location in the last disk of this compressor and the hub aft of the downstream compressor rotor. SUMMARY OF THE INVENTION

[0005] In a featured embodiment, a gas turbine engine comprises a fan rotor configured to be driven by a fan drive turbine through a first shaft and a gear reduction. The fan rotor is configured to deliver air into a bypass duct as bypass air and to deliver core air flow into a core engine where it reaches an upstream compressor rotor. The upstream compressor rotor is configured to be driven through a second shaft by an intermediate turbine rotor. A downstream compressor rotor is configured to be driven by an upstream turbine rotor through a third shaft. An overall pressure ratio across the upstream and downstream compressor rotors is greater than or equal to about 35.0 and less than or equal to about 75.0.

[0006] In another embodiment according to the previous embodiment, a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6.

[0007] In another embodiment according to any of the previous embodiments, a pressure ratio across the upstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 12.0.

[0008] In another embodiment according to any of the previous embodiments, a pressure ratio across the downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.

[0009] In another embodiment according to any of the previous embodiments, a gear ratio for the gear reduction is greater than or equal to about 2.6.

[0010] In another embodiment according to any of the previous embodiments, the intermediate turbine rotor has more than one stage.

[0011] In another embodiment according to any of the previous embodiments, the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages, and a ratio of the first number to the second number is greater than or equal to about 1.20.

[0012] In another embodiment according to any of the previous embodiments, a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine. The bypass ratio is greater than or equal to about 6.0.

[0013] In another embodiment according to any of the previous embodiments, a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6. [0014] In another embodiment according to any of the previous embodiments, a pressure ratio across the downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.

[0015] In another embodiment according to any of the previous embodiments, a gear ratio for the gear reduction is greater than or equal to about 2.6.

[0016] In another embodiment according to any of the previous embodiments, the intermediate turbine rotor has more than one stage.

[0017] In another embodiment according to any of the previous embodiments, the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages. A ratio of the first number to the second number is greater than or equal to about 1.2.

[0018] In another embodiment according to any of the previous embodiments, a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine. The bypass ratio is greater than or equal to about 6.0.

[0019] In another embodiment according to any of the previous embodiments, a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine. The bypass ratio is greater than or equal to about 6.0.

[0020] In another embodiment according to any of the previous embodiments, the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages. A ratio of the first number to the second number is greater than or equal to about 1.2.

[0021] In another embodiment according to any of the previous embodiments, the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages. A ratio of the first number to the second number is greater than or equal to about 1.2.

[0022] In another embodiment according to any of the previous embodiments, the intermediate turbine rotor has more than one stage.

[0023] In another embodiment according to any of the previous embodiments, a gear ratio for the gear reduction is greater than or equal to about 2.6. [0024] In another embodiment according to any of the previous embodiments, the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages. A ratio of the first number to the second number is greater than or equal to about 1.2.

[0025] These and other features may be best understood from the following drawing and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 schematically shows a gas turbine engine.

DETAILED DESCRIPTION

[0027] Gas turbine engine 120 is illustrated in Figure 1 having a fan rotor 122 driven through a gear reduction 124 by a fan drive turbine 128. The fan drive turbine 128 drives a shaft 126 that drives the gear reduction 124 to, in turn, drive the fan rotor 122 at a reduced speed. In embodiments, a gear ratio of the gear reduction is greater than or equal to about 2.6.

[0028] Fan rotor 122 delivers bypass air B within a nacelle 130 and core air C within a core housing 132. The air in the core housing 132 enters into an upstream or lower pressure compressor rotor 134. Air compressed by first or upstream compressor rotor 134 is then delivered into a second (higher pressure) or downstream compressor rotor 140. The compressed air downstream of the compressor rotor 140 is delivered into a combustion section 146 where it is mixed with fuel and ignited.

[0029] Products of this combustion pass downstream over an upstream or high pressure turbine rotor 142. As shown, a shaft 144 is driven by turbine rotor 142 to, in turn, drive the downstream compressor rotor 140. Downstream of the turbine rotor 142, the products of combustion pass over a second (intermediate pressure) turbine rotor 136. Turbine rotor 136 is shown driving a shaft 138 that is, in turn, connected to drive the upstream compressor rotor 134. The products of combustion downstream of turbine rotor 136 pass over the fan drive turbine 128.

[0030] An overall pressure ratio is defined across the compressor rotors 134 and

140. In addition, the average fan pressure ratio may be lower than about 1.6 and, in certain embodiments, be between about 1.2 and about 1.45. Applicant has discovered that by shifting more of the work burden to the lower pressure compressor rotor 134, the speed of the downstream compressor rotor 140 may be reduced even though it is generally seen as highly desirable from an aerodynamic efficiency standpoint, and a compressor stability standpoint and in order to reduce the number of compressor stages, for a compressor to be designed with as much speed capability as possible. This is because even though there are benefits to increased compressor speed, there are also countering trends that can sharply reduce the gains from endlessly increasing the speed of the downstream compressor.

[0031] One countering trend involves the mass of the disks throughout the downstream compressor, particularly in their bore areas to hold together at the high speeds and high temperatures. This disk mass in turn adds substantially to a transient thermal mismatch between the disks and engine casing during engine acceleration and deceleration. Rotor tip clearances become particularly large relative to the compressor blade height at the last stage of this compressor owing to high temperature excursions there. The tip clearance there is especially critical because that stage is the smallest in the engine and the open clearance represents a large percentage of the total flowpath annulus and therefore an inordinately large efficiency loss due to air returning to the upstream flowpath by going around the tip.

[0032] A slower speed in the last rotor of the downstream compressor allows an achievement of higher overall pressure ratio, by shifting speed (and stages) to the upstream compressor. At the same time, the speed of the downstream compressor rotor 140 has historically been a limit on overall achievable pressure ratio of about 50 for long range, twin aisle aircraft where the use of takeoff power is infrequent in the overall duty cycle of the engine. For shorter range, single aisle aircraft (often termed "regional jets"), the use of takeoff power is more frequent so the OPR might be limited to 40 or even less owing to durability concerns for the upstream turbine stages. This downstream location sees very high stress levels, especially at take-off and climb conditions, but the absolute level of stress that can be tolerated is increased if, at the elevated compressor exit temperature, and OPR, the speed is reduced. So reduced downstream compressor speed has at least two benefits: (a) reduced clearances; and (b) increased overall pressure ratio.

[0033] Naturally, as speed relationships between the upstream and downstream compressors are revised, the pressure rise across the compressors must also be revised. In addition, there are improvements to be made in turbine durability which, for any characteristic material temperature capability is also a function of the overall pressure ratio. A high overall pressure ratio is desired for better thermal efficiency for a long range engine application such as the use of an engine on a twin aisle aircraft with average flight times of six hours. This engine has fewer takeoff s relative to the time at cruise, so the high overall pressure ratio, impacting both the turbine gas path air and the turbine cooling air can still result in acceptable turbine life. This is in contrast to a single aisle aircraft application.

[0034] In contrast, the lower overall pressure ratio would result in optimum commercial engine economics for a shorter mission aircraft, most likely with a single aisle and medium passenger numbers below 240. This is so because even though the overall pressure ratio is somewhat lower (thereby increasing fuel burn), the extended turbine life benefits the aircraft economics to such an extent that it makes up for the less remarkable fuel burn.

[0035] Thus, in embodiments of the engine 120, an overall pressure ratio across the compressor rotors 134 and 140 may be greater than or equal to about 35.0 and less than or equal to about 75.

[0036] The lower pressure compressor rotor 134 may have a pressure ratio greater than or equal to about 4.0 and less than or equal to about 12.0. The higher or downstream compressor rotor 140 may have a pressure ratio of greater than or equal to about 4.0 and less than or equal to about 8.0.

[0037] The intermediate turbine rotor 136 may have more than one stage if the pressure rise across the first compressor is selected to be high in relation to the second compressor in the ranges mentioned earlier.

[0038] The fan drive turbine should have at least three stages.

[0039] In one non- limiting example, the upstream compressor rotor 134 may have at least about 1.20 times as many stages as the downstream compressor rotor 140.

[0040] Similarly, the fan may have a fan pressure ratio greater than or equal to about 1.1 and less than or equal to about 1.6.

[0041] A bypass ratio may be defined as the ratio of the volume of air delivered as bypass flow B compared to the volume of air delivered as core air flow C. The bypass ratio may be greater than or equal to about 6.0.

[0042] 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 about 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. In one example, the bypass ratio was greater than or equal to about 10.0.

[0043] 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:

a fan rotor configured to be driven by a fan drive turbine through a first shaft and a gear reduction;

said fan rotor being configured to deliver air into a bypass duct as bypass air and to deliver core air flow into a core engine where it reaches an upstream compressor rotor;

said upstream compressor rotor configured to be driven through a second shaft by an intermediate turbine rotor;

a downstream compressor rotor configured to be driven by an upstream turbine rotor through a third shaft; and

an overall pressure ratio across said upstream and downstream compressor rotors being greater than or equal to about 35.0 and less than or equal to about 75.0.

2. The gas turbine engine as set forth in claim 1, wherein a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6.

3. The gas turbine engine as set forth in claim 1, wherein a pressure ratio across said upstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 12.0.

4. The gas turbine engine as set forth in claim 3, wherein a pressure ratio across said downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.

5. The gas turbine engine as set forth in claim 4, wherein a gear ratio for said gear reduction is greater than or equal to about 2.6.

6. The gas turbine engine as set forth in claim 5, wherein said intermediate turbine rotor has more than one stage.

7. The gas turbine engine as set forth in claim 6, wherein said upstream compressor rotor has a first number of stages and said downstream compressor rotor has a second number of stages, and a ratio of said first number to said second number is greater than or equal to about 1.20.

8. The gas turbine engine as set forth in claim 7, wherein a bypass ratio is defined as the volume of air delivered as bypass flow into said bypass duct compared to the volume of air delivered as core flow into said core engine and said bypass ratio being greater than or equal to about 6.0.

9. The gas turbine engine as set forth in claim 4, wherein a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6.

10. The gas turbine engine as set forth in claim 1, wherein a pressure ratio across said downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.

11. The gas turbine engine as set forth in claim 10, wherein a gear ratio for said gear reduction is greater than or equal to about 2.6.

12. The gas turbine engine as set forth in claim 11, wherein said intermediate turbine rotor has more than one stage.

13. The gas turbine engine as set forth in claim 12, wherein said upstream compressor rotor has a first number of stages and said downstream compressor rotor has a second number of stages, and a ratio of said first number to said second number is greater than or equal to about 1.2.

14. The gas turbine engine as set forth in claim 13, wherein a bypass ratio is defined as the volume of air delivered as bypass flow into said bypass duct compared to the volume of air delivered as core flow into said core engine and said bypass ratio being greater than or equal to about 6.0.

15. The gas turbine engine as set forth in claim 1, wherein a bypass ratio is defined as the volume of air delivered as bypass flow into said bypass duct compared to the volume of air delivered as core flow into said core engine and said bypass ratio being greater than or equal to about 6.0.

16. The gas turbine engine as set forth in claim 15, wherein said upstream compressor rotor has a first number of stages and said downstream compressor rotor has a second number of stages, and a ratio of said first number to said second number is greater than or equal to about 1.2.

17. The gas turbine engine as set forth in claim 1, wherein said upstream compressor rotor has a first number of stages and said downstream compressor rotor has a second number of stages, and a ratio of said first number to said second number is greater than or equal to about 1.2.

18. The gas turbine engine as set forth in claim 1, wherein said intermediate turbine rotor has more than one stage.

19. The gas turbine engine as set forth in claim 1, wherein a gear ratio for said gear reduction is greater than or equal to about 2.6.

20. The gas turbine engine as set forth in claim 19, wherein said upstream compressor rotor has a first number of stages and said downstream compressor rotor has a second number of stages, and a ratio of said first number to said second number is greater than or equal to about 1.2.