BLADE OUTER AIR SEAL

Abstract

An example blade outer air seal assembly for a gas turbine engine includes a main body portion extending along an axis. The main body portion has a radially inward facing surface and at least one radially outward facing surface. A passage is provided in the main body portion between the at least one radially inward facing surface and the at least one radially outward facing surface. The passage has a first portion and a second portion transverse to the first portion, and at least one rib disposed along the axis radially outward of the second portion.

Description

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This disclosure was made with Government support under N00019-12-D-0002-4Y01 awarded by The United States Navy. The Government has certain rights in this disclosure.

BACKGROUND

This disclosure relates generally to a blade outer air seal and, more particularly, to enhancing the performance of a blade outer air seal.

As known, gas turbine engines, and other turbomachines, include multiple sections, such as a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. Air moves into the engine through the fan section. Airfoil arrays in the compressor section rotate to compress the air, which is then mixed with fuel and combusted in the combustor section. The products of combustion are expanded to rotatably drive airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan and compressor sections.

A blade outer air seal arrangement includes multiple blade outer air seals circumferentially disposed about at least some of the airfoil arrays. The tips of the blades within the airfoil arrays seal against the blade outer air seals during operation. Improving and maintaining the sealing relationship between the blades and the blade outer air seals enhances performance of the turbomachine. As known, the blade outer air seal environment is exposed to temperature extremes and other harsh environmental conditions, both of which can affect the integrity of the blade outer air seal and the sealing relationship.

SUMMARY

An example blade outer air seal assembly for a gas turbine engine includes a main body portion extending along an axis. The main body portion has a radially inward facing surface and at least one radially outward facing surface. A passage is provided in the main body portion between the at least one radially inward facing surface and the at least one radially outward facing surface. The passage has a first portion and a second portion transverse to the first portion, and at least one rib disposed along the axis radially outward of the second portion.

In a further non-limiting embodiment according to the previous assembly, the at least one rib is at least partially in a cavity that opens in a radially outward facing direction. The at least one rib comprises one of the radially outward facing surfaces.

In a further non-limiting embodiment according to any of the previous assemblies, at least one airflow inlet is provided by one of the radially outward facing surfaces.

In a further non-limiting embodiment according to any of the previous assemblies, the at least one airflow inlet includes a first airflow inlet opening to the cavity.

In a further non-limiting embodiment according to any of the previous assemblies, at least one attachment portion is disposed at both a leading edge of the main body portion and a trailing edge of the main body portion. Each of the at least one attachment portions has a flange extending axially aft.

In a further non-limiting embodiment according to any of the previous assemblies, the at least one rib is circumferentially aligned with one of the attachment portions disposed at the leading edge.

In a further non-limiting embodiment according to any of the previous assemblies, at least one fluid outlet is disposed at a circumferential end of the main body portion radially outward of the radially inward facing surface.

In a further non-limiting embodiment according to any of the previous assemblies, at least one airflow inlet is provided by one of the radially outward facing surfaces. The at least a first airflow inlet includes at least a first airflow inlet having a cross-section greater than a second airflow inlet.

In a further non-limiting embodiment according to any of the previous assemblies, the main body portion comprises a single crystal nickel alloy having a sulfur content that is equal to or below 1 part per million.

In a further non-limiting embodiment according to any of the previous assemblies, a thermal barrier coating is disposed adjacent the radially inward facing surface. A thickness of the thermal barrier coating is substantially equal to a thickness of an inner wall of the main body portion defined between a floor of each of the at least one passages and the radially inward facing surface.

In a further non-limiting embodiment according to any of the previous assemblies, the passage is entirely radially outward of the thermal barrier coating.

In a further non-limiting embodiment according to any of the previous assemblies, the second portion is radially inward of the first portion.

In a further non-limiting embodiment according to any of the previous assemblies, at least one of the radially outward facing surfaces is exposed.

An example gas turbine engine assembly includes a plurality of blade outer air seal assemblies each configured to attach to a casing and disposed circumferentially about an engine axis. The blade outer air seal assemblies each have a main body portion with a radially inward facing surface and a radially outward facing surface. At least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface. The main body portion has an inner wall having a first thickness defined between a floor of each of the at least one passages and the radially inward facing surface. A thermal barrier coating adjacent to the radially inward facing surface. The thermal barrier coating has a second thickness that is substantially equal to the first thickness.

In a further non-limiting embodiment according to the previous gas turbine engine assembly, at least one airflow inlet is opening to the radially outward facing surface. The at least one of passage is generally perpendicular to the engine axis. The at least one passage is in fluid communication with the at least one inlet and configured to receive cooling air flow.

In a further non-limiting embodiment according to any of the previous gas turbine engine assemblies, the at least one passage includes a first passage in fluid communication with a first airflow inlet of the at least one airflow inlets having airflow greater than a second passage in fluid communication with the second airflow inlet.

In a further non-limiting embodiment according to any of the previous gas turbine engine assemblies, the first airflow inlet has a cross-section greater than the second airflow inlet.

In a further non-limiting embodiment according to any of the previous gas turbine engine assemblies, an airflow source communicates compressor bleed air to the at least one inlet.

In a further non-limiting embodiment according to any of the previous gas turbine engine assemblies, each of the plurality of blade outer air seal assemblies includes at least one attachment portion disposed at each of a leading edge of the main body portion and a trailing edge of the main body portion. Each of the at least one attachment portions have a flange extending axially aft. Each of the at least one attachment portions aligned with corresponding receiving portion of the casing attaches the plurality of BOAS assemblies to the casing.

In a further non-limiting embodiment according to any of the previous gas turbine engine assemblies, at least at least one rib is disposed along the engine axis radially outward of the second portion.

An example method of forming a blade outer air seal for a gas turbine engine includes providing a main body portion having a radially inward facing surface and a radially outward facing surface that axially extend along an axis. At least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface. The at least one passage has a first portion extending generally radially and a second portion extending generally circumferentially transverse to the first portion. The second portion has a floor. At least one rib is disposed along the axis on the radially outward facing surface. The at least one rib is radially outward of the circumferential second portion. An inner wall defined between the floor and the radially inward facing surface is machined to reduce a thickness of the inner wall. A thermal barrier coating is deposited adjacent the radially inward facing surface. The thermal barrier coating is substantially equal to the reduced thickness of the inner wall.

In a further non-limiting embodiment according to the previous method, at least one airflow inlet opening to the radially outward facing surface is formed. The at least one passage is in fluid communication with the at least one inlet and configured to receive cooling air flow.

These and other features of the disclosed examples can be best understood from the following specification and drawings, the following of which is a brief description.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 shows a cross-section of an example turbomachine.

FIG. 2 shows a cross-section of an example turbine section of the turbomachine of FIG. 1.

FIG. 3 shows a perspective view of a blade outer air seal assembly of the turbine section of FIG. 2.

FIG. 4 shows a top view of the blade outer air seal assembly of the turbine section of FIG. 2.

FIG. 5 shows another perspective view of the blade outer air seal assembly of the turbine section of FIG. 2.

FIG. 6 shows top cross-sectional view of the blade outer air seal assembly of the turbine section of FIG. 2.

FIG. 7 shows method of forming a blade outer air seal assembly.

DETAILED DESCRIPTION

Referring to FIG. 1, a gas turbine engine 20 is schematically illustrated. 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 flowpath while the compressor section 24 drives air along a core flowpath 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, non-geared turbine engines, and land-based turbines.

The engine 20 generally includes a first spool 30 and a second spool 32 mounted for rotation about an engine central 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 first spool 30 generally includes a first shaft 40 that interconnects a fan 42, a first compressor 43 and a first turbine 46. The first shaft 40 is connected to the fan 42 through a gear assembly of a fan drive gear system 48 to drive the fan 42 at a lower speed than the first spool 30. The second spool 32 includes a second shaft 49 that interconnects a second compressor 52 and second turbine 55. The first spool 30 runs at a relatively lower pressure than the second spool 32. It is to be understood that “low pressure” and “high pressure” or variations thereof as used herein are relative terms indicating that the high pressure is greater than the low pressure. An annular combustor 57 is arranged between the second compressor 52 and the second turbine 55. The first shaft 40 and the second shaft 49 are concentric and rotate via bearing systems 38 about the engine central axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the first compressor 43 then the second compressor 52, mixed and burned with fuel in the annular combustor 57, then expanded over the second turbine 55 and first turbine 46. The first turbine 46 and the second turbine 55 rotationally drive, respectively, the first spool 30 and the second spool 32 in response to the expansion.

The engine 20 is a high-bypass geared aircraft engine that has a bypass ratio that is greater than about six (6), with an example embodiment being greater than ten (10), the gear assembly of the fan drive gear system 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 2.3:1 and the first turbine 46 has a pressure ratio that is greater than about 5. The first turbine 46 pressure ratio is pressure measured prior to inlet of first turbine 46 as related to the pressure at the outlet of the first turbine 46 prior to an exhaust nozzle. The first turbine 46 has a maximum rotor diameter and the fan 42 has a fan diameter such that a ratio of the maximum rotor diameter divided by the fan diameter is less than 0.6. It should be understood, however, that the above parameters are only exemplary.

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 0.8 Mach and 35,000 feet, with the engine at its best fuel consumption. To make an accurate comparison of fuel consumption between engines, fuel consumption is reduced to a common denominator, which is applicable to all types and sizes of turbojets and turbofans. The term is thrust specific fuel consumption, or TSFC. This is an engine's fuel consumption in pounds per hour divided by the net thrust. The result is the amount of fuel required to produce one pound of thrust. The TSFC unit is pounds per hour per pounds of thrust (lb/hr/lb Fn). When it is obvious that the reference is to a turbojet or turbofan engine, TSFC is often simply called specific fuel consumption, or SFC. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane 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 feet per second 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 feet per second.

Referring to FIG. 2, an example blade outer air seal (BOAS) assembly 50 is attached to an inner engine case structure 44 of the gas turbine engine 10 by a receiving portion 68 of the inner engine case structure 44. In this example, the BOAS assembly 50 is located within the turbine section 28 of the gas turbine engine 20. The BOAS assembly 50 faces turbine blade 51 to define a radial tip clearance 53 between the turbine blade 51 and the BOAS assembly 50. Although only one BOAS assembly 50 is shown, a number of BOAS assembly 50 are arranged circumferentially about engine axis A to form a shroud. Alternatively, the BOAS assemblies 50 may be formed as a unitary BOAS structure, with the same features described herein.

Referring to FIGS. 3-6, the BOAS assembly 50 includes a main body portion 54 that extends generally axially from a leading edge portion 56 to a trailing edge portion 58 and from a radially outward facing surface 62 to a radially inward facing surface 64. The BOAS assembly 50 also includes at least one leading attachment portion 60a (“attachment portions 60a”) disposed at or near the leading edge portion 56 and at least one trailing attachment portion 60b (“attachment portions 60b”) disposed at or near the trailing edge portion 58. Each of the attachment portions 60a, 60b define a flange 66 extending in an axially aft direction. Each axially extending flange 66 corresponds to the receiving portion 68 of the inner engine case structure 44 to support and attach the BOAS assembly 50 (Shown in FIG. 2). In this example, the attachment portions 60a may be circumferentially offset, circumferentially aligned, or a combination of both, from the attachment portions 60b in response to BOAS assembly 50 parameters.

In this example, the BOAS assembly 50 includes a cavity 70 opening to the radially outward facing surface 62 between the attachment portions 60a and the attachment portions 60b. It is understood that other configurations of cavity 70 are contemplated by this disclosure.

In this example, the main body portion 54 establishes at least one rib 72 circumferentially aligned with corresponding attachment portion 60a and circumferentially offset from attachment portions 60b. The at least one rib 72 is disposed at least partially within the cavity 70 and extends axially from corresponding attachment portion 60a within the cavity 70 to the trailing edge portion 58 adjacent attachment portions 60b. In another example, the at least one rib 72 is circumferentially aligned with corresponding attachment portions 60a and attachment portion 60b and extends axially from attachment portion 60a to corresponding attachment portion 60b. A ratio of the height 74 in the radial direction of the at least one rib 72 to the width 76 in the circumferential direction of the at least one rib 72 is between 1:1 and 1:10.

In this example, a plurality of fluid inlets 80 open to the radially outward facing surface 62 near a first circumferential end 82 and a second circumferential end 84 of the main body portion 54. The fluid inlets 80 may be located in the cavity 70 or alternatively at other positions on the radially outward facing surface 62. The fluid inlets 80 are varied in size based on pre-determined cooling parameters. Fluid inlets 80a with a larger surface area provide a greater amount of cooling airflow than fluid inlets 80b with a smaller surface area, thus providing cooling flow distribution based on thermal load distribution. In this way, portions of the BOAS assembly 50 subject to relatively higher thermal loads compared to other portion of the BOAS assembly 50 receive greater cooling by receiving cooling airflow through relatively larger fluid inlets 80.

In this example, each of the fluid inlets 80 is in fluid communication with a corresponding cooling passage 86 defined in the main body portion 54 radially inwards of fluid inlet 80 and cavity 70. Each cooling passage 86 includes a first portion 86a generally transverse to a second portion 86b. The first portion 86a extends in a generally circumferential direction from the first circumferential end 82 to the second circumferential end 84 of the main body portion 54 to provide cooling. The second portion 86b extends in a generally radial direction to provide fluid communication between inlet 80 and the first portion 86a. In this example, each cooling passage 86 is defined entirely within the main body portion 54.

The BOAS assembly 50 is in fluid communication with an airflow source 90 (shown schematically), such as an upstream compressor 24 or other source, such that fluid inlets 80 receive cooling airflow, such as bleed compressor air. The cooling airflow passes through fluid inlets 80 and is communicated to the first portion 86a of corresponding cooling passage 86 via second portion 86b for cooling the BOAS assembly 50. The cooling airflow passes through the cooling passage 86 from the fluid inlet 80 to a fluid outlet 92 located at the circumferential end 82, 84 of the cooling passage 86 opposite fluid inlet 80. The amount of cooling airflow communicated to each cooling passage 86 is determined by the size of fluid inlet 80. The BOAS assembly 50 also include a pressure gradient which determines the amount of cooling airflow communicated to each cooling passage 86. In this example, the fluid inlets 80 have a cross sectional area 94 between about 0.00028 in.̂2 (about 0.00181 cm̂2) and about 0.0078 in.̂2 (about 0.00503 cm̂2). Cooling airflow is communicated through each cooling passage 86 in a single circumferential direction in this example. Plugs (not shown) are inserted to close the cooling passage 86 at the circumferential end 82, 84 corresponding to the fluid inlet 80 of each passage.

In this example, the BOAS assembly 50 is made of a material having sulfer levels at or below 1 part per million (PPM), such as a single crystal nickel alloy, but other examples may include other types of material. The thermal barrier coating 110 is a metallic or ceramic based material in this example.

In this example, the BOAS assembly 50 includes a thermal barrier coating 110 disposed on the radially inward facing surface 64 of the main body portion 54. The thermal barrier coating 110 includes a thermal layer 112 and a bond layer 114. A thickness 116 of the thermal barrier coating 110 is substantially equal to a thickness 118 of an inner wall 120 of the main body portion 54 defined between a floor 122 of the plurality of cooling passages 86 and the radially inward facing surface 64. The term substantially equal conveys that one of ordinary skill in the art would consider the measurements to be the same within recognized tolerances.

The thickness 116 of the thermal barrier coating 110 and the thickness 118 of the inner wall 120 are between 0.025 (0.0635 cm) and 0.030 inches (0.0762 cm). In this example, the thickness 116 of the thermal barrier coating 110 and the thickness 118 of the inner wall 120 is about 0.027 inches (0.0686 cm).

During gas turbine engine 20 operation, the BOAS assembly 50 is subjected to different thermal loads and environmental conditions. Cooling air flow from the airflow source 90 is provided to the various fluid inlets 80, which communicate the cooling airflow to the cooling passages 86 to provide varying levels of cooling to different areas of the BOAS assembly 50 and effectively communicate thermal energy away from the BOAS assembly 50 and the tip of the rotating blade 51. The thermal barrier coating 110 is provided on the radially inner facing surface 64 of the main body portion 54 to provide additional protection from the thermal loads. The at least one rib 72 provides stability to the BOAS assembly 50 to prevent axial deformation due to the reduction in BOAS assembly 50 material due to the use of cooling features and thermal bond coating, as described in this disclosure.

Referring to FIG. 6, in another example, adjacent first portions 86a of cooling passages 86 provided in the main body portion 54 are connected at either the first circumferential end 82 or the second circumferential end 84 to form a serpentine passage 96 (shown schematically in phantom in FIG. 6) in fluid communication with a single fluid inlet 80 or multiple fluid inlets 80 via one or more second portions 86b (as shown in FIG. 3).

Referring to FIG. 7, with continued reference to FIGS. 2-6, a method of forming a BOAS assembly 200 includes providing a main body portion along an axis A having a radially inward facing surface and a radially outward facing surface that extend axially between a leading edge portion and a trailing edge portion 202. At least one passage is disposed in the main body portion between the radially inward facing surface and the radially outward facing surface 202. The at least one passage has a radial first portion and an axial second portion transverse to the radial first portion. The second portion defines a floor 202. At least one rib is provided along the axis A 202. The at least one rib is radially outward of the circumferential second portion 202. An inner wall defined between the floor and the radially inward facing surface is machined to reduce a thickness of the inner wall 204. A thermal barrier coating is deposited adjacent the radially inward facing surface 206. A thickness of the thermal barrier coating is substantially equal to the reduced thickness of the inner wall 206.

The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.

Claims

1. A blade outer air seal assembly for a gas turbine engine, comprising:

a main body portion extending along an axis, the main body portion having a radially inward facing surface and at least one radially outward facing surface;

a passage provided in the main body portion between the at least one radially inward facing surface and the at least one radially outward facing surface, the passage having a first portion and a second portion transverse to the first portion; and

at least one rib disposed along the axis radially outward of the second portion.

2. The assembly of claim 1, wherein the at least one rib is at least partially in a cavity that opens in a radially outward facing direction, the at least one rib comprising one of the radially outward facing surfaces.

3. The assembly of claim 1, including at least one airflow inlet provided by one of the radially outward facing surfaces.

4. The assembly of claim 3, wherein the at least one airflow inlet includes a first airflow inlet opening to a cavity that opens in a radially outward facing direction.

5. The assembly of claim 1, including at least one attachment portion disposed at both a leading edge of the main body portion and a trailing edge of the main body portion, each of the at least one attachment portions having a flange extending axially aft.

6. The assembly of claim 5, wherein the at least one rib is circumferentially aligned with one of the attachment portions disposed at the leading edge.

7. The assembly of claim 1, wherein at least one fluid outlet is disposed at a circumferential end of the main body portion radially outward of the radially inward facing surface.

8. The assembly of claim 1, including at least one airflow inlet provided by one of the radially outward facing surfaces, wherein the at least one airflow inlet includes at least a first airflow inlet having a cross-section greater than a second airflow inlet.

9. The assembly of claim 1, wherein the main body portion comprises a single crystal nickel alloy having a sulfur content that is equal to or below 1 part per million.

10. The assembly of claim 1, wherein a thermal barrier coating is disposed adjacent the radially inward facing surface, wherein a thickness of the thermal barrier coating is substantially equal to a thickness of an inner wall of the main body portion defined between a floor of the passage and the radially inward facing surface.

11. The assembly of claim 10, wherein the passage is entirely radially outward of the thermal barrier coating.

12. The assembly of claim 1, wherein the second portion is radially inward of the first portion.

13. The assembly of claim 1, wherein at least one of the radially outward facing surfaces is exposed.

14. A gas turbine engine assembly, comprising:

a plurality of blade outer air seal assemblies each configured to attach to a casing and disposed circumferentially about an engine axis, the blade outer air seal assemblies each having a main body portion with a radially inward facing surface and a radially outward facing surface, wherein at least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface, the main body portion having an inner wall having a first thickness defined between a floor of each of the at least one passages and the radially inward facing surface; and

a thermal barrier coating adjacent to the radially inward facing surface, the thermal barrier coating having a second thickness that is substantially equal to the first thickness.

15. The gas turbine engine assembly of claim 14, wherein at least one airflow inlet is opening to the radially outward facing surface, the at least one of passage generally perpendicular to the engine axis, the at least one passage in fluid communication with the at least one inlet and configured to receive cooling air flow.

16. The gas turbine engine assembly of claim 15, wherein the at least one passage includes a first passage in fluid communication with a first airflow inlet of the at least one airflow inlet having airflow greater than a second passage in fluid communication with a second airflow inlet.

17. The gas turbine engine assembly of claim 16, wherein the first airflow inlet has a cross-section greater than the second airflow inlet.

18. The gas turbine engine assembly of claim 15, wherein an airflow source communicates compressor bleed air to the at least one inlet.

19. The gas turbine engine assembly of claim 14, wherein each of the plurality of blade outer air seal assemblies includes at least one attachment portion disposed at each of a leading edge of the main body portion and a trailing edge of the main body portion, each of the at least one attachment portions having a flange extending axially aft, each of the at least one attachment portions aligned with corresponding receiving portion of the casing to attach the plurality of blade outer air seal assemblies to the casing.

20. The gas turbine engine assembly of claim 14, including at least at least one rib is disposed along the engine axis radially outward of the second portion.

21. A method of forming a blade outer air seal for a gas turbine engine comprising:

providing a main body portion having a radially inward facing surface and a radially outward facing surface that axially extend along an axis, wherein at least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface, the at least one passage having a first portion extending generally radially and a second portion extending generally circumferentially transverse to the first portion, wherein the second portion has a floor, wherein at least one rib is disposed along the axis on the radially outward facing surface, the at least one rib radially outward of the circumferential second portion;

machining an inner wall defined between the floor and the radially inward facing surface to reduce a thickness of the inner wall; and

depositing a thermal barrier coating adjacent the radially inward facing surface, wherein the thermal barrier coating is substantially equal to the reduced thickness of the inner wall.

22. The method of claim 21, including the step of forming at least one airflow inlet opening to the radially outward facing surface, wherein the at least one passage is in fluid communication with the at least one inlet and configured to receive cooling air flow.