Patent Application
20150361798
Fan blades of a gas turbine engine are thin and have an aerodynamic shape to reduce weight and minimize thrust specific fuel consumption. However, the fan blades must also meet structural requirements to withstand an event, such as a bird strike.
Internal cavities can be employed to reduce the weight of the fan blade, while still meeting structural requirements. As the cavities are internal, air does not flow through the cavities and affect performance.
A blade of a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a blade portion. The blade portion includes a first portion located radially outwardly that is located in a flow path and a second portion located radially inwardly of the first portion that is not located in the flow path. A plurality of external cavities is in the second portion of the blade portion.
In a further embodiment of any of the foregoing blades the blade is a fan blade of a fan.
In a further embodiment of any of the foregoing blades the blade has a pressure side and a suction side spaced apart by a thickness, and each of the plurality of external cavities extend from the pressure side to the suction side.
In a further embodiment of any of the foregoing blades the blade has a pressure side and a suction side spaced apart by a thickness, and each of the plurality of external cavities extend partially through the thickness of the blade.
In a further embodiment of any of the foregoing blades each of the plurality of external cavities is substantially elliptical in shape.
In a further embodiment of any of the foregoing blades each of the plurality of external cavities extend in a radial direction.
In a further embodiment of any of the foregoing blades includes a plurality of internal cavities in the first portion of the blade, and a cover is secured over the plurality of internal cavities.
In a further embodiment of any of the foregoing blades each of the plurality of external cavities have a substantially triangular shape.
In a further embodiment of any of the foregoing blades includes a root portion radially inward of the blade portion.
A fan of a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a rotor including a plurality of slots, and a plurality of fan blades extending radially about an axial centerline. Each of the plurality of fan blades includes a blade portion and a root portion that is received in one of the plurality of slots of the rotor. The blade portion includes a first portion located radially outwardly that is located in a flow path and a second portion located radially inwardly of the first portion that is not located in the flow path, and the second portion of the blade portion includes a plurality of external cavities.
In a further embodiment of the foregoing fans the plurality of fan blades each have a pressure side and a suction side spaced apart by a thickness, and each of the plurality of external cavities extend from the pressure side to the suction side.
In a further embodiment of the foregoing fans the plurality of fan blades each have a pressure side and a suction side spaced apart by a thickness, and each of the plurality of external cavities extend partially through the thickness of each of the plurality of fan blades.
In a further embodiment of the foregoing fans each of the plurality of external cavities is substantially elliptical in shape.
In a further embodiment of the foregoing fans each of the plurality of external cavities extend in a radial direction.
In a further embodiment of the foregoing fans includes a plurality of internal cavities in the first portion of each of the plurality of fan blades, and a cover is secured over the plurality of internal cavities.
In a further embodiment of the foregoing fans each of the plurality of external cavities have a substantially triangular shape.
In a further embodiment of the foregoing fans the root portion is a dovetail.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
Although the disclosed non-limiting embodiment depicts a geared turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with geared turbofans as the teachings may be applied to other types of traditional turbine engines. For example, the gas turbine engine 20 can have a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive the fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
The example gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about a 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.
The low speed spool 30 generally includes an inner shaft 40 that connects the fan 42 and a low pressure (or first) compressor 44 to a low pressure (or first) turbine 46. The inner shaft 40 drives the fan 42 through a speed change device, such 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 high pressure (or second) compressor 52 and a high pressure (or second) turbine 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the central longitudinal axis A.
A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double 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 a corresponding “low pressure” compressor or turbine.
The example low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
A mid-turbine frame 58 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 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.
The air in the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 58 includes vanes 60, which are in the core flow path C and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture 48 and that the present disclosure is applicable to other gas turbine engines.
A significant amount of thrust is provided by the air in the bypass flow path B due to the high bypass ratio. The fan section 22 of the gas turbine 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 pound-mass (lbm) of fuel per hour being burned divided by pound-force (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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.
The example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about 26 fan blades 62. In another non-limiting embodiment, the fan section 22 includes less than about 20 fan blades 62. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about 3 turbine rotors. A ratio between the number of fan blades 62 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of fan blades 62 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
The second portion 74 of the blade portion 68 of the fan blade 62 includes external cavities 76. In one example, the external cavities 76 extend through an entire thickness T of the second portion 74 of the blade portion 68 of the fan blade 62. That is, the external cavities 76 extend from a pressure side 80 to a suction side 82 of the blade portion 68 of the fan blade 62. The external cavities 76 are exposed to the flow path.
As shown in
The external cavities 76 are located in the second portion 74 of the blade portion 68 of the fan blade 62 below the boundary 84 such that the external cavities 76 are not located within the flow path. In one example, the external cavities 76 are substantially elongated and extend in a radial direction. When the external cavities 76 are oriented in the radial direction, the external cavities 76 are also in a direction of a blade pull load and provide for better stress concentration because both the radius of curvature and the load are radial. In one example, the external cavities 76 have an elliptical shape to further reduce stress concentrations. The external cavities 76 have a length such that the external cavities 76 do not extend above the boundary 84 and into the second portion 74 that is exposed to the flow path. In one example, each of the external cavities 76 is a slot.
The external cavities 76 can be formed by several methods. In one example, the external cavities 76 are machined after the fan blade 62 is created. In another example, the fan blade 62 around the external cavities 76 is formed by an additive manufacture process where the external cavities 76 are formed from a three dimensional model. In another example, the external cavities 76 are formed by near net forging.
The external cavities 76 reduce the weight of the fan blade 62 while still allowing the fan blade 62 to meet structural requirements to allow the fan blade 62 to withstand an event, such as a bird strike or in the event of pull if the low spool 30 runs at a higher than expected speed. However, as the external cavities 76 are located in the second portion 74 of the fan blade 62 and are not located in the flow path, the external cavities 76 do not affect performance or aerodynamic efficiency of the fan 42. As the external cavities 76 do not need to be covered because they are not located in the flow path, the weight can be reduced as a cover is not needed to encase the external cavities 76. As the external cavities 76 reduce the weight of the fan blade 62, the chances of a fan blade out event are also reduced.
In another example shown in
In another example shown in
Although a fan blade 62 has been illustrated and described, the external cavities 76 can be added to a blade that is part of the low pressure compressor 44, the high pressure compressor 52, the high pressure turbine 54, or the low pressure turbine 46.
Although a gas turbine engine 20 with geared architecture 48 is described, the fan blade 62 can be employed in a gas turbine engine without geared architecture.
The foregoing description is only exemplary of the principles of the invention. Many modifications and variations are possible in light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than using the example embodiments which have been specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2014/013341 | 1/28/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61763681 | Feb 2013 | US |
