HPC CASE CLEARANCE CONTROL THERMAL CONTROL RING SPOKE SYSTEM

Abstract

A case clearance control system comprising a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate the blade outer air seal support structure; a thermal control ring coupled to the blade outer air seal support structure, the thermal control ring including a plurality of receivers configured to couple with the plurality of protrusions; a thermal break formed between the plurality of protrusions and the plurality of receivers, the thermal break configured to control heat transfer between the blade outer air seal support structure and the thermal control ring; and a plurality of flow passages formed between the blade outer air seal support structure, the thermal control ring and the plurality of protrusions, the plurality of flow passages configured to allow cooling air flow to condition the thermal control ring and maintain thermal control ring dimensions.

Claims

1. A case clearance control system comprising: a blade outer air seal support structure having a plurality of protrusions extending radially from the blade outer air seal support structure opposite a blade outer air seal proximate said blade outer air seal support structure; a thermal control ring coupled to said blade outer air seal support structure, said thermal control ring including a plurality of receivers configured to couple with said plurality of protrusions; a thermal break formed between said plurality of protrusions and said plurality of receivers, said thermal break configured to control heat transfer between said blade outer air seal support structure and said thermal control ring; and a plurality of flow passages formed between said blade outer air seal support structure, said thermal control ring and said plurality of protrusions, said plurality of flow passages configured to allow cooling air flow to condition said thermal control ring and maintain thermal control ring dimensions.

2. The case clearance control system according to claim 1, wherein said thermal control ring comprises a material composition comprising a coefficient of thermal expansion less than said blade outer air seal support structure and said blade outer air seal.

3. The case clearance control system according to claim 1, wherein said protrusions are formed integrally with said blade outer air seal support structure.

4. The case clearance control system according to claim 1, wherein said protrusions comprise a base portion proximate said blade outer air seal support and an end portion radially distal from said base portion.

5. The case clearance control system according to claim 4, wherein said end portion and said receiver are configured in a shape of a dovetail fitting that maintains said thermal break.

6. The case clearance control system according to claim 1, wherein said protrusions are spaced apart with a distance of from not less than 3.5 degrees to not more than 20 degrees.

7. The case clearance control system according to claim 1, wherein said thermal beak is configured to inhibit thermal energy transfer from said blade outer air seal support structure to said thermal control ring.

8. A case clearance control system comprising: a blade outer air seal support structure configured to support a blade outer air seal proximate a tip of a blade of a gas turbine high pressure compressor, said blade outer air seal support structure including a plurality of protrusions projecting radially outward from said blade outer air seal support structure opposite said blade outer air seal; a thermal control ring coupled to said blade outer air seal support structure at said plurality of protrusions, said thermal control ring including a plurality of receivers configured to couple with said plurality of protrusions; a thermal break formed between said plurality of protrusions and said plurality of receivers, said thermal break configured to control heat transfer between said blade outer air seal support structure and said thermal control ring; and a plurality of flow passages formed between said blade outer air seal support structure, said thermal control ring and said plurality of protrusions, said plurality of flow passages configured to allow cooling air flow to pass through and cool said thermal control ring in order to maintain thermal control ring dimensions to control expansion and contraction of said blade outer air seal support structure.

9. The case clearance control system according to claim 8, wherein said protrusions are formed spaced apart around a circumference of the blade outer air seal support structure.

10. The case clearance control system according to claim 8, wherein said thermal control ring is configured to add thermal mass to said blade outer air seal.

11. The case clearance control system according to claim 8, wherein said cooling air flow thermally conditions said thermal control ring.

12. The case clearance control system according to claim 8, wherein said thermal beak is configured to inhibit thermal energy transfer from said blade outer air seal support structure to said thermal control ring.

13. The case clearance control system according to claim 8, wherein said protrusions can be configured at least one of tapped to receive threaded fasteners inserted radially through the thermal control ring or receive pins insertable bayonet style with said thermal control ring.

14. A process for maintaining a tip clearance with a case clearance control system comprising: attaching a thermal control ring to a blade outer air seal support structure, said blade outer air seal support structure configured to support a blade outer air seal proximate a tip of a blade of a gas turbine high pressure compressor, wherein said blade outer air seal support structure includes a plurality of protrusions extending radially from the blade outer air seal support structure opposite said blade outer air seal, said thermal control ring including a plurality of receivers configured to couple with said plurality of protrusions; forming a thermal break between said plurality of protrusions and said plurality of receivers, said thermal break configured to control heat transfer between said blade outer air seal support structure and said thermal control ring; and flowing cooling air through a plurality of flow passages formed between said blade outer air seal support structure, said thermal control ring and said plurality of protrusions, wherein said cooling air cools said thermal control ring for maintaining the thermal control ring dimensions to control expansion and contraction of said blade outer air seal support structure, controlling the tip clearance.

15. The process of claim 14, further comprising: thermally conditioning said thermal control ring such that said thermal control ring thermally decouples from a transient gas turbine engine operation.

16. The process of claim 14, further comprising: inhibiting the thermal energy transferred from a core gas path flow into the blade outer air seal support structure into the thermal control ring, maintaining a stable shape in the thermal control ring in the absence of thermal growth or thermal contraction.

17. The process of claim 14, further comprising: coupling said plurality of receivers with said plurality of protrusions by use of a dovetail fitting.

18. The process of claim 14, further comprising: dimensioning said thermal control ring to cooperate with said flow passages for flowing said cooling air over and between said thermal control ring and said blade outer air seal support structure; and maintaining a structural integrity of said thermal control ring during a gas turbine engine transient response.

19. The process of claim 14, further comprising: flowing said cooling air through said flow passages and said thermal break proximate said at least one protrusion and said at least one receiver.

20. The process of claim 14, wherein said at least one protrusion comprises a base portion proximate said blade outer air seal support and an end portion radially distal from said base portion, said receiver comprises a corresponding shape with said end portion, said receiver comprises a larger dimension than said end portion configured to promote said thermal break.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0029] FIG. 1 is a simplified cross-sectional view of a gas turbine engine.

[0030] FIG. 2 is a simplified cross-sectional view of a gas turbine engine proximate an exemplary high pressure compressor section.

[0031] FIG. 3 is a simplified isometric view of an exemplary blade outer air seal support structure and thermal control ring.

[0032] FIG. 4 is an isometric view of the exemplary blade outer air seal support structure of FIG. 3.

[0033] FIG. 5 is a simplified cross-sectional view of an exemplary blade outer air seal support structure and thermal control ring.

[0034] FIG. 6 is a simplified cross-sectional view of an exemplary blade outer air seal support structure and thermal control ring.

DETAILED DESCRIPTION

[0035] FIG. 1 is a simplified cross-sectional view of a gas turbine engine 10 in accordance with embodiments of the present disclosure. Turbine engine 10 includes fan 12 positioned in bypass duct 14. Turbine engine 10 also includes compressor section 16, combustor (or combustors) 18, and turbine section 20 arranged in a flow series with upstream inlet 22 and downstream exhaust 24. During the operation of turbine engine 10, incoming airflow F.sub.1 enters inlet 22 and divides into core flow F.sub.C and bypass flow F.sub.B, downstream of fan 12. Core flow F.sub.C continues along the core flowpath through compressor section 16, combustor 18, and turbine section 20, and bypass flow F.sub.B proceeds along the bypass flowpath through bypass duct 14.

[0036] Compressor 16 includes stages of compressor vanes 26 and blades 28 arranged in low pressure compressor (LPC) section 30 and high pressure compressor (HPC) section 32. Turbine section 20 includes stages of turbine vanes 34 and turbine blades 36 arranged in high pressure turbine (HPT) section 38 and low pressure turbine (LPT) section 40. HPT section 38 is coupled to HPC section 32 via HPT shaft 42, forming the high pressure spool. LPT section 40 is coupled to LPC section 30 and fan 12 via LPT shaft 44, forming the low pressure spool. HPT shaft 42 and LPT shaft 44 are typically coaxially mounted, with the high and low pressure spools independently rotating about turbine axis (centerline) C.sub.L.

[0037] Combustion gas exits combustor 18 and enters HPT section 38 of turbine 20, encountering turbine vanes 34 and turbines blades 36. Turbine vanes 34 turn and accelerate the flow of combustion gas, and turbine blades 36 generate lift for conversion to rotational energy via HPT shaft 42, driving HPC section 32 of compressor 16. Partially expanded combustion gas flows from HPT section 38 to LPT section 40, driving LPC section 30 and fan 12 via LPT shaft 44. Exhaust flow exits LPT section 40 and turbine engine 10 via exhaust nozzle 24. In this manner, the thermodynamic efficiency of turbine engine 10 is tied to the overall pressure ratio (OPR), as defined between the delivery pressure at inlet 22 and the compressed air pressure entering combustor 18 from compressor section 16. As discussed above, a higher OPR offers increased efficiency and improved performance. It will be appreciated that various other types of turbine engines can be used in accordance with the embodiments of the present disclosure.

[0038] Referring also to FIG. 2, an exemplary portion of the gas turbine high pressure compressor section 32 is shown. The compressor section 32 disclosed can achieve a technical effect through a thermal contraction (or inhibit thermal expansion) of the case 46 and blade outer air seal (BOAS) support structure 48 through the use of a thermal control ring 50 in conjunction with cooling air 52 in cooperation with the blade outer air seal (BOAS) support structure 48.

[0039] The blade outer air seal support structure 48 supports a blade outer air seal 54. The blade outer air seal 54 interacts with the blade 28 to maintain a tip clearance 56 between the blade outer air seal 54 and blade 28. The control of the tip clearance 56 can be accomplished by actively conditioning the thermal control ring 50 in the high pressure compressor section 32 using cooler upstream bleed air, i.e., cooling air 52. The tip clearance control for the high pressure compressor section 32 can be influenced by core gas path 58 heat transfer into the blade outer air seal 54 and support structure 48. During transient engine conditions, such as speed deceleration followed by rapid speed acceleration, the core gas path flow 58 can experience rapid variation in temperature under these transients, such that a mismatch in radial thermal growth between the blade outer air seal 54 and the blade 28 tips occurs. When the mismatch occurs at a re-acceleration condition of maximum radial growth of the blade 28 tips, and minimum radial growth of the blade outer air seal 54, the running tip clearance 56 between the blade outer air seal 54 and blade 28 will become negative and excessive blade outer air seal 54 abradable rub will occur. Prior art designs for this condition by sizing the tip clearance at a large enough value to prevent occurrence, but this also reduces compressor efficiency.

[0040] Referring also to FIG. 3 and FIG. 4, the thermal control ring 50 attached to the blade outer air seal support structure 48 adds circumferential thermal mass to the blade outer air seal support structure 48. Additionally, the thermal control ring 50 is coupled to the blade outer air seal support structure 48 at multiple protrusions or spokes 60. The thermal control ring 50 can be coupled to the blade outer air seal support structure 48 radially opposite the blade outer air seal 54 at the rear 62 of the support structure 48. Coupling 63 between the thermal control ring 50 and the blade outer air seal support structure 48 can be configured as a thermal break 64 to reduce the heat transfer between the thermal control ring 50 and the blade outer air seal support structure 48. The thermal break 64 can include additional lox alpha, low heat transfer materials, coatings, inserts as well as air gaps. The combination of the thermal control ring 50 coupled to the blade outer air seal support structure 48 acts to control the expansion and contraction behavior of the blade outer air seal support structure 48 relative to the tip clearance 56 and other compressor section components subjected to the core gas path flow 58 temperature fluctuations.

[0041] The cooling air 52 thermally conditions the thermal control ring 50 and provides a more consistent thermal profile during engine operations, thermally decoupled from the transient nature of the core gas flow path 58. The combination of the thermal control ring 50 coupled to the blade outer air seal support structure 48 and cooling air 52 provides for a case clearance control system 66.

[0042] In order to control the thermal growth deflection of certain BOAS in the high pressure compressor section 32, such as, rear stage BOAS, and provide the thermal break 64 to minimize the core gas path flow 58 thermal convection and conduction heat transfer from the BOAS support structure 48, the protrusions 60 can be formed integral to and projecting radially outward from the BOAS support structure 48. The protrusions 60 can be formed around the entire circumference of the BOAS support structure 48.

[0043] In an exemplary non-limiting embodiment, the protrusions 60 can have a base portion 68 proximate the BOAS support structure 48 and an end portion 70 extending distally from the BOAS support structure 70. The end portion 70 can be shaped accordingly to provide for a secure coupling between the thermal control ring 50 and the BOAS support structure 48 while allowing for the thermal break 64 between the thermal control ring 50 and the BOAS support structure 48. The thermal break 64 is understood to be configured to inhibit/reduce the heat transfer between the thermal control ring 50 and the BOAS support structure 48. The thermal break 64 inhibits the thermal energy being transferred from the relatively hot core gas path flow 58 through the BOAS support structure 48 to the thermal control ring 50, thus allowing the thermal control ring 50 to maintain a relatively stable shape due to thermal growth/contraction. The thermal control ring 50 can be formed from low alpha materials that have low thermal conductance, as opposed to the BOAS support structure 48 which can be formed of high alpha materials that has good thermal conductive properties.

[0044] In an exemplary embodiment the end portion 70 can couple with the thermal control ring 50 by use of a dovetail fitting 72. The end portion 70 can be shaped as a bulbous form that interlocks with a receiver 74 of similar shape formed in the thermal control ring 50. The receiver 74 can correspond with the shape of the end portion 70 with a larger dimension to allow for a loose fit between the end portion 70 and receiver 74. The loose fit can promote the thermal break 64 in the form of an air gap and allow for the movement of the cooling air 52 to pass over the end portion 70 and through the receiver 74. The receiver 74 and end portion 70 shapes can be optimized to minimize mechanical or thermal stresses based on the materials used in construction.

[0045] The protrusions 60 can be dimensioned so as to form flow passages 76 that allow for the cooling air 52 to flow between each protrusion 60 and between the thermal control ring 50 and BOAS support structure 48, thus maintaining relatively stable dimensions in the thermal control ring 50. The flow passages 76 can be passages proximate the receiver 74 in addition to the thermal break 64.

[0046] The thermal control ring 50 can also be dimensioned to cooperate with the flow passages 76 optimizing the flow of cooling air 52 flowing over and between the thermal control ring 50 and BOAS support structure 48 while maintaining sufficient structural integrity in view of thermal and mechanical stresses to achieve the desired low thermal deflections during transient engine responses. The thermal control ring 50 and BOAS support structure 48 can be configured to expose both the OD and ID surfaces of the thermal control ring 50, to increase heat transfer. The thermal control ring 50 comprises a material composition with a coefficient of thermal expansion that is less than the blade outer air seal 54 or blade outer air seal support structure 48 material. The thermal control ring 50 can be segmented or continuously formed as a hoop. The thermal control ring 50, if formed of a continuous hoop, can be configured to slide axially onto the protrusions 60 to couple with the BOAS support structure 48.

[0047] The protrusions 60 can be spaced apart sufficient to allow for cooling air 52 to pass between. In an exemplary non-limiting embodiment, the protrusions 60 can be spaced apart at a dimension of not less than 3.5 degrees to allow for sufficient cooling air 52 to flow and not more than about 20 degrees to prevent unwanted distortion in the BOAS support structure 48.

[0048] In an exemplary embodiment the protrusions 60 can be configured to receive threaded fasteners and/or posts 78 that can be inserted radially through the thermal control ring 50 into the protrusions 60 at the end portion 70 and configured to secure the thermal control ring 50 to the BOAS support structure 48, as seen in FIG. 5.

[0049] In an exemplary embodiment the protrusions 60 can be configured to receive pins or double shear rivets 80 that can be inserted bayonet style with the thermal control ring 50 into the protrusions 60 proximate the end portion 70 and configured to secure the thermal control ring 50 to the BOAS support structure 48, as seen in FIG. 6.

[0050] The disclosed case clearance control system provides the technical advantage of allowing for the cooling air to dominate the thermal growth and control deflections during transient engine conditions preventing heavy rub conditions.

[0051] The disclosed case clearance control system provides the technical advantage of enabling tighter tip clearances during operating conditions.

[0052] The disclosed case clearance control system provides the technical advantage of controlling thermal growth deflection in the BOAS by use of a thermal control ring having thermal breaks between protrusions integral to the BOAS support structure and cooling air passages between the protrusions, thermal control ring and BOAS support structure.

[0053] The disclosed case clearance control system provides the technical advantage of controlling BOAS contraction during transient engine speed conditions.

[0054] The disclosed case clearance control system provides the technical advantage of decoupling the mass within the compressor by conditioning a thermal control ring and thermally breaking the thermal control ring from the BOAS support structure.

[0055] The disclosed case clearance control system provides the technical advantage of having the thermal control ring to dominate BOAS radial growth instead of the core gas path flow thermal influences.

[0056] There has been provided a tip clearance control system. While the case clearance control system has been described in the context of specific embodiments thereof, other unforeseen alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations which fall within the broad scope of the appended claims.