Patent Application
20200025106
This application relates to a geared turbofan engine which may be particularly beneficial for application on regional jet engines.
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 geared turbofan engines to particular application in aircraft which will utilize the gas turbine engine.
In a featured embodiment, a gas turbine engine comprises a fan rotor having blades with an outer diameter. The outer diameter is greater than or equal to 77 inches (196 centimeters) and less than or equal to 135 inches (343 centimeter). The fan rotor has less than or equal to 26 fan blades, and is driven by a fan drive turbine through a gear reduction. The gear reduction has a gear ratio of greater than 2.6:1. The fan rotor delivers air into a bypass duct as bypass air, and into a duct leading to a compressor rotor as core air. A ratio of bypass air to the core air is greater than or equal to 12:1. An upstream turbine rotor is upstream of the fan drive turbine and drives a compressor rotor. The upstream turbine rotor has at least two stages, and the fan drive turbine rotor has at least three stages. The turbine blades in at least one stage of the fan drive turbine rotor are provided with a performance enhancing feature. The performance enhancing features is at least one of the blades being manufactured by a directionally solidified blade material. The blades are provided as single crystal blades, and have a radially outer platform provided with scalloping to reduce the weight of the blades. The blades are provided with cooling air.
In another embodiment according to the previous embodiment, the gear ratio is greater than 3.06:1.
In another embodiment according to any of the previous embodiments, the gear ratio is a star gearbox.
In another embodiment according to any of the previous embodiments, an area is defined at a downstream end of the upstream turbine rotor, and a second area is defined at a downstream end of the fan drive turbine rotor. The upstream turbine rotor rotates ng at a first speed and the fan drive turbine rotor rotates at a second speed. A performance quantity of the upstream turbine rotor is defined by the first cross-sectional area multiplied by the first speed squared. A performance quantity of the fan drive turbine rotor is defined by the second cross-sectional area multiplied by the second speed squared. The first performance ratio is less than the second performance ratio.
In another embodiment according to any of the previous embodiments, the second performance quantity is greater than or equal to at least 5.0 times 10 squared (5.0×102).
In another embodiment according to any of the previous embodiments, the blades in the at least one stage are manufactured by a directionally solidified blade material.
In another embodiment according to any of the previous embodiments, the blades in the at least one stage are provided as single crystal blades.
In another embodiment according to any of the previous embodiments, the blades in the at least one stage have a radially outer platform, which is provided with scalloping to reduce the weight of the blades in the at least one stage.
In another embodiment according to any of the previous embodiments, the blades in the at least one stage are provided with cooling air.
In another embodiment according to any of the previous embodiments, an area is defined at a downstream end of the upstream turbine rotor, and a second area is defined at a downstream end of the fan drive turbine rotor. The upstream turbine rotor rotates at a first speed and the fan drive turbine rotor rotates at a second speed. A performance quantity of the upstream turbine rotor is defined by the first cross-sectional area multiplied by the first speed squared and a performance quantity of the fan drive turbine rotor being defined by the second cross-sectional area multiplied by the second speed squared. The first performance ratio is less than the second performance ratio.
In another embodiment according to any of the previous embodiments, the second performance quantity is greater than or equal to at least 5.0 times 10 squared (5.0×102).
In another embodiment according to any of the previous embodiments, the blades in the at least one stage are manufactured by a directionally solidified blade material.
In another embodiment according to any of the previous embodiments, the blades are provided as single crystal blades.
In another embodiment according to any of the previous embodiments, the blades in the at least one stage have a radially outer platform, which is provided with scalloping to reduce the weight of the blades in the at least one stage.
In another embodiment according to any of the previous embodiments, the blades in the at least one blade row are provided with cooling air.
In another embodiment according to any of the previous embodiments, the gas turbine engine is for use on a single aisle aircraft.
These and other features may be best understood from the following drawings and specification.
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 (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), 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 (350.5 meters/second).
The features described below, with regard to turbine sections, may be incorporated into engines such as shown in
In particular, the disclosed gas turbine engine will have a very high bypass ratio. Further, a gear ratio of the gear reduction will be relatively high. To increase the efficiency of the fan drive turbine, the speed of the fan drive turbine will be beneficially increased. This can be best achieved by increasing the gear reduction such that the fan rotor does not also increase in speed.
However, increasing the speed of the fan drive turbine will increase the temperature related challenges on the fan drive turbine. Thus, the following disclosure will provide improvements to the fan drive turbine and the overall turbine section, to increase the ability to withstand high temperatures.
In particular, the bypass ratio is desirably greater than or equal to about 12:1. The gear reduction preferably is a gear ratio of 3.06:1, and is a star gearbox.
A fan diameter is preferably greater than or equal to 77 inches (196 centimeters), but less than or equal to 135 inches (343 centimeters). Preferably, there are less than 26 fan blades.
As shown in
A first area AH is defined at a downstream end of the upstream turbine rotor 96. A second area AL is defined at a downstream end of the fan drive turbine rotor 98. The upstream turbine rotor 96 rotates at a first speed and the fan drive turbine rotor 98 rotates at a second speed. A first performance quantity (AN2) of the upstream turbine rotor 96 is defined by cross-sectional area AH multiplied by its speed squared. A second performance quantity (AN2) of the fan drive turbine rotor 98 is defined by cross-sectional area AL multiplied by its speed squared. The second performance ratio is greater than or equal to at least 5.0 times 10 squared (5.0×102) RPM in2 at engine redline speed, which is slightly higher than takeoff.
The performance quantity of upstream turbine rotor 58 is less than the performance quantity of the fan drive turbine rotor 98.
As shown in
An upstream turbine rotor has two stages, 116A and 116B. The fan drive turbine rotor has three stages, 118A, 118B, and 118C. As shown, blades 120A, 120B, and 120C are associated with the stages 118A, 118B, and 118C, respectively.
The blades 120A, 120B, and 120C preferably have some feature provided to increase their resistance to high temperatures. Such features are disclosed in the following paragraphs and
In one embodiment, the blades may be formed of a directionally solidified blade material, or a single crystal blade.
As shown in
The “scallops” could be described as an interrupted body of revolution. A shroud has two ID corners and two OD corners, and these are interrupted by the scallops and are no longer a two dimensional curve, but rather become a 3 dimensional curve. Similarly, the OD surface of the shroud, adjacent to the knife edges, is conventionally a surface of simple revolution also; adding scallops interrupts that surface and the surface moves toward the centerline locally.
An alternative blade 140 is illustrated in
Short to medium range aircraft spend more time at high stress conditions and, in particular, take-off and climb than do longer range aircraft. As such, the engines are subjected to more stresses and the disclosure of this application will provide valuable benefits which are synergistically realized in such aircraft.
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.
This application is a continuation of U.S. patent application Ser. No. 14/624,668 filed on Feb. 18, 2015.
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14624668 | Feb 2015 | US |
| Child | 16395623 | US |
