US20140010663A1

Movatterモバイル変換

Info

Publication number: US20140010663A1
Application number: US13/536,156
Authority: US
Inventors: Joseph Parkos, JR.; James O. Hansen
Original assignee: Individual
Current assignee: RTX Corp
Priority date: 2012-06-28
Filing date: 2012-06-28
Publication date: 2014-01-09
Also published as: WO2014031195A2; EP2867477A2; WO2014031195A3; EP2867477A4

Abstract

A method of manufacturing a fan blade includes providing a metallic fan blade body, and applying at least a 200 volt potential to the fan blade body in a solution to produce a crystalline oxidation layer. A fan blade for a gas turbine engine includes a metallic fan blade body having a tip with a crystalline oxidation layer. A fan blade for a gas turbine engine includes a metallic fan blade body having a tip with a titanium dioxide layer.

Description

BACKGROUND

This disclosure relates to a fan blade for a gas turbine engine. In more particularly, the disclosure relates to a fan blade having a treated tip.

Composite fan blades have been proposed for gas turbine engines. One example fan blade includes an aluminum substrate or fan blade body having a titanium sheath adhered to a leading edge of the fan blade body. A polyurethane coating is applied over the fan blade body and forms a fan blade contour with the sheath.

During operation of the gas turbine engine, a tip of the fan blade may rub with an adjacent sealing structure. During the rub event, the tip may wear and heat may be generated. To this end, the tip of the fan body has been anodized to create a hardened layer. One typical sealing structure has a hardness greater than that of the tip. Thus, during rub events the tip may still become undesirably worn. In some instances, sufficient heat may be generated to delaminate the coating from the fan blade body.

SUMMARY

In one exemplary embodiment, a method of manufacturing a fan blade includes providing a metallic fan blade body, and applying at least a 200 volt potential to the fan blade body in a solution to produce a crystalline oxidation layer.

In a further embodiment of any of the above, the fan blade body includes a tip. The method includes the step of masking the fan blade body to leave at least the tip exposed. The tip having the crystalline oxidation layer.

In a further embodiment of any of the above, the fan blade body provides an electrode in an alkaline solution.

In a further embodiment of any of the above, the method includes generating a plasma discharge on a surface of the fan blade body.

In a further embodiment of any of the above, the crystalline oxidation layer is aluminum oxide.

In a further embodiment of any of the above, the crystalline oxidation layer has a thickness of at least 0.0005 inch (0.013 mm).

In a further embodiment of any of the above, the crystalline oxidation layer has a hardness of at least 1700 HV.

In a further embodiment of any of the above, the fan blade body is one of a 7000 series and a 2000 series aluminum alloy.

In a further embodiment of any of the above, the method includes the steps of removing the masking and adhering a sheath to a leading edge of the fan blade body.

In a further embodiment of any of the above, the adhering step includes arranging an adhesive-saturated scrim between the sheath and the leading edge.

In a further embodiment of any of the above, the method includes the step of coating the fan blade body with polyurethane to provide a fan blade contour along with the sheath.

In a further embodiment of any of the above, the crystalline oxidation layer is left exposed subsequent to the coating step.

In another exemplary embodiment, a fan blade for a gas turbine engine includes a metallic fan blade body having a tip with a crystalline oxidation layer.

In a further embodiment of any of the above, the crystalline oxidation layer has a thickness of at least 0.005 inch (0.12 mm).

In a further embodiment of any of the above, the fan blade includes a sheath adhered to a leading edge of the fan blade body.

In a further embodiment of any of the above, the fan blade includes a polyurethane coating arranged over the fan blade body and adjoining the sheath to provide a fan blade contour.

In another exemplary embodiment, a fan blade for a gas turbine engine includes a metallic fan blade body having a tip with a titanium dioxide layer.

In a further embodiment of any of the above, the tip includes a plasma spray coating arranged on the titanium dioxide layer, which is provided between the plasma spray coating and the metallic fan blade body.

In a further embodiment of any of the above, the amorphous titanium oxidation layer has a thickness of at least 0.0005 inch (0.13 mm).

In a further embodiment of any of the above, the amorphous titanium oxidation layer has a hardness of at least 640 HV.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2 is a perspective view of an embodiment of a fan blade of the engine shown inFIG. 1.

FIG. 3 is a perspective view of a blade body for the fan blade shown inFIG. 2.

FIG. 4 is a flowchart depicting an example manufacturing process used to produce the fan blade shown inFIG. 2.

FIG. 5 is an end view of the fan blade shown inFIG. 2 produced by the manufacturing process depicted inFIG. 4.

FIG. 6 is a schematic cross-sectional view of another example fan blade produced by another manufacturing process.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an examplegas turbine engine20 that includes afan section22, acompressor section24, acombustor section26 and aturbine section28. Alternative engines might include an augmenter section (not shown) among other systems or features. Thefan section22 drives air along a bypass flow path B while thecompressor section24 draws air in along a core flow path C where air is compressed and communicated to acombustor section26. In thecombustor section26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through theturbine section28 where energy is extracted and utilized to drive thefan section22 and thecompressor section24.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including 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 a 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.

Theexample engine20 generally includes alow speed spool30 and ahigh speed spool32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure36 viaseveral bearing systems38. It should be understood thatvarious bearing systems38 at various locations may alternatively or additionally be provided.

Thelow speed spool30 generally includes aninner shaft40 that connects afan42 and a low pressure (or first)compressor section44 to a low pressure (or first)turbine section46. Theinner shaft40 drives thefan42 through a speed change device, such as a gearedarchitecture48, to drive thefan42 at a lower speed than thelow speed spool30. The high-speed spool32 includes anouter shaft50 that interconnects a high pressure (or second)compressor section52 and a high pressure (or second)turbine section54. Theinner shaft40 and theouter shaft50 are concentric and rotate via thebearing systems38 about the engine central longitudinal axis X.

Acombustor56 is arranged between thehigh pressure compressor52 and thehigh pressure turbine54. In one example, thehigh pressure turbine54 includes at least two stages to provide a double stagehigh pressure turbine54. In another example, thehigh pressure turbine54 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 examplelow pressure turbine46 has a pressure ratio that is greater than about 5. The pressure ratio of the examplelow pressure turbine46 is measured prior to an inlet of thelow pressure turbine46 as related to the pressure measured at the outlet of thelow pressure turbine46 prior to an exhaust nozzle.

Amid-turbine frame57 of the enginestatic structure36 is arranged generally between thehigh pressure turbine54 and thelow pressure turbine46. Themid-turbine frame57 furthersupports bearing systems38 in theturbine section28 as well as setting airflow entering thelow pressure turbine46.

The core airflow C is compressed by thelow pressure compressor44 then by thehigh pressure compressor52 mixed with fuel and ignited in thecombustor56 to produce high speed exhaust gases that are then expanded through thehigh pressure turbine54 andlow pressure turbine46. Themid-turbine frame57 includesvanes59, which are in the core airflow path and function as an inlet guide vane for thelow pressure turbine46. Utilizing thevane59 of themid-turbine frame57 as the inlet guide vane forlow pressure turbine46 decreases the length of thelow pressure turbine46 without increasing the axial length of themid-turbine frame57. Reducing or eliminating the number of vanes in thelow pressure turbine46 shortens the axial length of theturbine section28. Thus, the compactness of thegas turbine engine20 is increased and a higher power density may be achieved.

The disclosedgas turbine engine20 in one example is a high-bypass geared aircraft engine. In a further example, thegas turbine engine20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example gearedarchitecture48 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, thegas turbine engine20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of thelow pressure compressor44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. Thefan section22 of theengine20 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.

“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) 0.5]. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.

The example gas turbine engine includes thefan42 that comprises in one non-limiting embodiment less than about26 fan blades. In another non-limiting embodiment, thefan section22 includes less than about20 fan blades. Moreover, in one disclosed embodiment thelow pressure turbine46 includes no more than about 6 turbine rotors schematically indicated at34. In another non-limiting example embodiment thelow pressure turbine46 includes about 3 turbine rotors. A ratio between the number offan blades42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The examplelow pressure turbine46 provides the driving power to rotate thefan section22 and therefore the relationship between the number ofturbine rotors34 in thelow pressure turbine46 and the number ofblades42 in thefan section22 disclose an examplegas turbine engine20 with increased power transfer efficiency.

Referring toFIGS. 2 and 3, afan blade60 of thefan42 includes aroot62 supporting aplatform64. Anairfoil66 extends from theplatform64 to atip67. Theairfoil66 includes spaced apart leading and trailing

edges

68,70. Pressure and

suction sides

72,74 adjoin the leading and trailing

edges

68,70 to provide afan blade contour86. It should be understood that thefan blade60 is exemplary, and “platformless” fan blades may also be used.

Thetip67 is arranged adjacent to a sealingstructure83, which is typically arranged in relation to thetip67 to provide aclearance84. One example sealing structure may have embedded glass particles with a hardness of 650 HV. During certain engine operating conditions, thetip67 may be prone to rubbing with the sealingstructure83, which can generate heat and undesirably wear thetip67.

Eachfan blade60 includes an aluminumfan blade body80, which may be hollow or solid. Aleading edge sheath76 is applied to aleading edge78 of thefan blade body80. In one example, thefan blade body80 is constructed from a 7000 series aluminum alloy, such as 7255. In another example, a 2000 series aluminum is used.

Thefan blade body80 has aleading edge78. Asheath76 is secured to thefan blade body80 over theedge78 withadhesive82. In one example, thesheath76 and thefan blade body80 are constructed from first and second metals that are different from one another. In one example, thesheath76 is constructed from a titanium alloy. It should be understood that other metals or materials may be used.

The adhesive82 provides a barrier between thefan blade body80 and thesheath76 to prevent galvanic corrosion. Referring toFIG. 5, the adhesive82 includes a scrim88 (e.g., a glass scrim) that carries the adhesive82. Examples of the adhesive82 include a variety of commercially available aerospace-quality metal-bonding adhesives, including several epoxy- and polyurethane-based adhesive films. In some embodiments, the adhesive82 is heat-cured via autoclave or other similar means. Examples of suitable bonding agents include type EA9628 epoxy adhesive available from Henkel Corporation, Hysol Division, Bay Point, Calif. and type AF163K epoxy adhesive available from 3M Adhesives, Coatings & Sealers Division, St. Paul, Minn. The adhesive82 may be cured using a vacuum bag and autoclave.

Certain adhesives

82, including the example film-based adhesives above, are compatible withscrim88.Scrim88 provides dielectric separation betweenblade body80 andsheath76, preventing galvanic corrosion between the two different metal surfaces ofblade body80 andsheath76. Oneexample scrim88 is a flexible nylon-based layer with a thickness between about 0.003 inch (0.08 mm) and about 0.010 inch (0.25 mm) thick. Other examples of the adhesive82 and other aspects of thefan blade60 are set forth in U.S. Patent Application Publication 2011/0211967 to the Applicant, which is incorporated herein by reference in its entirety.

Apolymer coating90 is applied over thefan blade body80 adjacent to thesheath76 to provide afan blade contour86. Thecoating90 is polyurethane in one example.

Anexample method100 of manufacturing thefan blade60 is illustrated in the flow chart shown inFIG. 4. The substrate provided by thefan blade body80 may be masked to leave at least thetip67 exposed, as depicted inblock102. A chemical masking, mechanical masking tape, lacquer, or other painted on coating may be used and temporarily deposited upon the exterior surface of thefan blade body80. Since the oxidation process employed by theexample method100 requires significant voltage, reducing the area to be oxidized greatly reduces the necessary power and cost of oxidizing thetip67.

Thetip67 is oxidized using a high voltage process, such as micro-arc oxidation (MAO), which may also be referred to as plasma electrolytic oxidation (PEO) or plasma arc oxidation (PAO). Such a process employs an at least a 200 volt potential, and as high as 2000 volts. One example process is available from Keronite International Ltd. in the United Kingdom. During the MAO process, a plasma is generated on the surface of the fan blade body, which forms a crystalline aluminum oxide structure having a hardness of at least 1700 HV. A typical anodized aluminum hardness is 450 HV, and the oxidation layer produced by anodizing is non-crystalline. The crystalline oxidation layer92 (FIG. 5) may have a thickness of at least 0.005 inch (0.12 mm), which is significantly thicker than a typical anodized layer thickness of 0.002 inch (0.05 mm). However, the layer may be thinner than 0.005 inch (0.12 mm) since thecrystalline oxidation layer92 is harder than typical anodizing.

Since thecrystalline oxidation layer92 is significantly harder than the sealingstructure83, which has embedded particles around 650 HV in one example, thetip67 more easily cuts through the sealingstructure83 during a rub event, ultimately generating less heat. The reduced heat minimizes the possibility of thecoating90 from delaminating from thefan blade body80. The thicker crystalized oxidation layer also better insulates the substrate of thefan blade body80, reducing the that reaches the interface of thefan blade body80 and thecoating90. The coating is also more wear resistant than a conventional anodize. It also provides superior corrosion protection than a conventional anodize.

The mask may be removed following theoxidation process104. Thefan blade60 may be assembled, as indicated atblock106, by securing thesheath76 over the leadingedge78 using the adhesive82 andscrim88. Thefan blade body80 is covered with thecoating90 to provide thefan blade contour86, as best shown inFIG. 5.

Another example fan blade160 is illustrated inFIG. 6. The fan blade160 includes an aluminum alloy substrate of thefan blade body180. Atitanium dioxide layer192 is deposited on the end of thefan blade body180 using, for example, an EC2 process available from Henkel. The EC2 process applies a current to thefan blade body180, which is immersed in a neutral solution. The EC2 process adds atitanium dioxide layer192 to the aluminum alloy substrate and provides a sufficient bond. Thetitanium dioxide layer192 is 0.0005-0.001 inch (0.013-0.25 mm) thick. If additional wear resistance is desired, a plasma-sprayedcoating110 may be applied to thetitanium dioxide layer192. Example plasma spray coatings are alumina or metal. Thetitanium dioxide layer192 andplasma spray coating110, if used, provides a wearresistant tip167 that also thermally insulates thefan blade body180. The coating is also more wear resistant and provides superior corrosion protection than a conventional anodize.

Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

Claims

What is claimed is:

1. A method of manufacturing a fan blade comprising:

providing a metallic fan blade body; and

applying at least a 200 volt potential to the fan blade body in a solution to produce a crystalline oxidation layer.

2. The method according toclaim 1, wherein the fan blade body includes a tip, and comprising the step of masking the fan blade body to leave at least the tip exposed, the tip having the crystalline oxidation layer.

3. The method according toclaim 1, wherein the fan blade body provides an electrode in an alkaline solution.

4. The method according toclaim 3, comprising generating a plasma discharge on a surface of the fan blade body.

5. The method according toclaim 1, wherein the crystalline oxidation layer is aluminum oxide.

6. The method according toclaim 5, wherein the crystalline oxidation layer has a thickness of at least 0.0005 inch (0.013 mm).

7. The method according toclaim 5, wherein the crystalline oxidation layer has a hardness of at least 1700 HV.

8. The method according toclaim 5, wherein the fan blade body is one of a 7000 series and a 2000 series aluminum alloy.

9. The method according toclaim 2, comprising the steps of removing the masking and adhering a sheath to a leading edge of the fan blade body.

10. The method according toclaim 9, wherein the adhering step includes arranging an adhesive-saturated scrim between the sheath and the leading edge.

11. The method according toclaim 10, comprising the step of coating the fan blade body with polyurethane to provide a fan blade contour along with the sheath.

12. The method according toclaim 11, wherein the crystalline oxidation layer is left exposed subsequent to the coating step.

13. A fan blade for a gas turbine engine comprising:

a metallic fan blade body having a tip with a crystalline oxidation layer.

14. The fan blade according toclaim 13, wherein the crystalline oxidation layer has a thickness of at least 0.005 inch (0.12 mm).

15. The fan blade according toclaim 14, wherein the crystalline oxidation layer has a hardness of at least 1700 HV.

16. The fan blade according toclaim 15, wherein the fan blade body is one of a 7000 series and a 2000 series aluminum alloy.

17. The fan blade according toclaim 16, comprising a sheath adhered to a leading edge of the fan blade body.

18. The fan blade according toclaim 17, comprising a polyurethane coating arranged over the fan blade body and adjoining the sheath to provide a fan blade contour.

19. A fan blade for a gas turbine engine comprising:

a metallic fan blade body having a tip with a titanium dioxide layer.

20. The fan blade according toclaim 19, wherein the tip includes a plasma spray coating arranged on the titanium dioxide layer, which is provided between the plasma spray coating and the metallic fan blade body.

21. The fan blade according toclaim 19, wherein the amorphous titanium oxidation layer has a thickness of at least 0.0005 inch (0.13 mm).

22. The fan blade according toclaim 21, wherein the amorphous titanium oxidation layer has a hardness of at least 640 HV.

23. The fan blade according toclaim 22, wherein the fan blade body is one of a 7000 series and a 2000 series aluminum alloy.

24. The fan blade according toclaim 23, comprising a sheath adhered to a leading edge of the fan blade body.

25. The fan blade according toclaim 24, comprising a polyurethane coating arranged over the fan blade body and adjoining the sheath to provide a fan blade contour.