Removable film for airfoil surfaces

Abstract

A fan section to be incorporated into a gas turbine engine has a rotor and a plurality of fan blades. The fan blades deliver air into a bypass duct defined inwardly of the nacelle and into a core engine. There are static vanes inward of the nacelle. A surface of the fan section is provided with a removable film material. A gas turbine engine and a method of refurbishing a surface are also disclosed.

Claims

1. A method of protecting and refurbishing a surface in a fan section for a gas turbine engine comprising steps of: peeling a protective outer removable layer from a surface in the fan section and leaving an underlying protective removable layer, such that said outermost surface can be removed along with accumulated debris; and further including a step of subsequently peeling away said underlying protective removable layer.

2. A fan section to be incorporated into a gas turbine engine comprising: a rotor and a plurality of fan blades, said fan blades for delivering air into a bypass duct defined inwardly of a nacelle and for delivering air into a core engine, and there being guide vanes inward of said nacelle; a surface of said fan section being provided with a removable film material; wherein said removable film material includes a plurality of layers of removable material, such that an outer layer may be peeled away from an underlying layer; and wherein said underlying layer may be peeled away from a subsequent layer underling said underlying layer.

3. The fan section as set forth in claim 1, wherein said surface of said fan section is on said fan blade.

4. The fan section as set forth in claim 1, wherein said surface of said fan section is on said nacelle.

5. The fan section as set forth in claim 1, wherein said surface of said fan section is on said guide vane.

6. The fan section as set forth in claim 1, wherein said surface of said fan section is on a core engine cowl.

7. The fan section as set forth in claim 1, wherein said surface of said fan section is on a variable area nozzle.

8. The fan section as set forth in claim 1, wherein said removable film material has at least one of hydrophobic or icephobic properties.

9. The fan section as set forth in claim 1, wherein said removable film material has a relatively sticky underside and a relatively less sticky outer side.

10. A gas turbine engine comprising: a fan section for delivering air into a bypass duct with a nacelle and for delivering air into a compressor; said fan section having a rotor and a plurality of fan blades, and there being static vanes inward of said nacelle; a surface of said fan section being provided with a removable film material; wherein said removable film material includes a plurality of layers of removable material, such that an outer layer may be peeled away from an underlying layer; and wherein said underlying layer may be peeled away from a subsequent layer underling said underlying layer.

11. The fan section as set forth in claim 10, wherein said surface of said fan section is on said fan blade.

12. The fan section as set forth in claim 10, wherein said surface of said fan section is on said nacelle.

13. The fan section as set forth in claim 10, wherein said surface of said fan section is on said guide vane.

14. The fan section as set forth in claim 10, wherein said surface of said fan section is on a core engine cowl.

15. The fan section as set forth in claim 10, wherein said surface of said fan section is on a variable area nozzle.

16. The gas turbine engine as set forth in claim 10, wherein said removable film material has at least one of hydrophobic or icephobic properties.

17. The gas turbine engine as set forth in claim 10, wherein said removable film material has a relatively sticky underside and a relatively less sticky outer side.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) FIG. 1 schematically shows a gas turbine engine.

(2) FIG. 2A shows a fan blade in a gas turbine engine.

(3) FIG. 2B shows a novel fan blade.

(4) FIG. 2C shows details of a fan section.

(5) FIG. 3A is a cross-sectional view of the novel fan blade.

(6) FIG. 3B shows another embodiment.

(7) FIG. 3C shows yet another embodiment.

(8) FIG. 3D shows another embodiment.

(9) FIG. 3E shows yet another embodiment.

(10) FIG. 4 shows a method step in refurbishing a surface of a fan blade.

(11) FIG. 5 is a simplified cross-sectional view of a multi-layer adhesive film.

DETAILED DESCRIPTION

(12) FIG. 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.

(13) 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.

(14) 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. A fan exit guide vane 11 is shown downstream of the fan 42. 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.

(15) 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.

(16) 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 five (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 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 2.5:1. 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 gas turbine engines including direct drive turbofans.

(17) 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 conditiontypically 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 consumptionalso known as bucket cruise Thrust Specific Fuel Consumption (TSFC)is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf 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)].sup.0.5. The Low corrected fan tip speed as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

(18) FIG. 2A shows a fan blade 120 which may be incorporated into an engine such as shown in FIG. 1. A dovetail 124 is received within a rotor that rotates to drive a plurality of the blades 120. Airfoil 126 extends radially outwardly of the dovetail 124. Areas of accumulated debris 150 are shown schematically on the airfoil 126. It can be appreciated that the fan blade is in an area where it is subjected to a good deal of airborne debris and as this debris builds up, the aerodynamic qualities deteriorate. The same is true for the surfaces inwardly of the nacelle 15 as shown in FIG. 1, static vanes, or an inlet into the nacelle 15 as shown in FIG. 1.

(19) FIG. 2B shows an improved fan blade 120 wherein a surface 160 may be removed, as will be explained below, from at least the airfoil section 126.

(20) FIG. 2C shows a number of details in the fan section 600. As shown, a rotor 601 rotates blades 120, and sits within a nacelle 15. The exit guide vanes 11 are downstream of the blades 120. A variable area nozzle 13 is positioned at a downstream end of the nacelle 15, and can move to change the area of the bypass duct, as known. Further, an engine core cowl 14 has an outer surface which also defines a portion of the bypass duct. Any one of the illustrated components in the fan section 600 may receive a similar removable surface. While the engine core cowl 14 and/or variable area nozzle 13 may not be considered part of a fan section in some definitions, for purposes of this application, they are included in the definition of fan section.

(21) As shown in FIG. 3A, the surface 160 may be formed on one side of the airfoil 126, and another removable surface 162 may be formed on an opposed side.

(22) Although the variable area nozzle 13 and exit guide vanes 11 may not typically be called part of the fan section, they are part of an airflow through the bypass duct, and will benefit from the removable surface. Thus, for purposes of this application, they are part of the fan section 600.

(23) FIG. 3B shows a nacelle 400 having a removable material 402. FIG. 3C shows a static exit guide vane 410 having a removable material 412.

(24) FIG. 3D shows an engine core cowl 514 having the removable material 516 on a radially outer surface. FIG. 3E shows a variable area nozzle 511 having the removable material 512 on a radially inner surface.

(25) For purposes of this application, the fan blade 120, the nacelle 400, the static fan exit guide vane 410, the variable area nozzle 13, the engine core cowl 14 and any other related structure are collectively part of the fan section.

(26) FIG. 4 shows a method according to this application. As shown, debris 150 has built up on the outer surfaces 160 and 162 of the fan blade 120. As shown at 170, an outer layer is peeled away removing the debris 150 from the side 162. An underlying film layer 172 remains and is clean. A similar step would next occur on the side 160. Such film layers have a relatively sticky underside 300, and a relatively less sticky outer side 301.

(27) As shown in FIG. 5, the surfaces/materials 160, 162, 402 and 412 may all be formed of a plurality of layers 180. Appropriate materials may be similar to those utilized on race car applications.

(28) However, since the materials for this application need not be optically clear, more freedom in materials may be provided and materials that are more resistant to dirt accumulation or pitting may be utilized. Further, film materials that have hydrophobic or icephobic properties may also be utilized to repel water and/or limit ice accumulation.

(29) Example materials may be publicly available from windshield tear-off materials. As an example, materials available under the trade names Pro-Shield or Racing Optics from Pro-Tint of Kannapolis, N.C.

(30) 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.