Patent Application
20210087948
The disclosure relates generally to gas turbine engines and more specifically to configurations of individual turbine stages.
To facilitate high turbine rotor temperatures and shaft speeds in superalloy turbine rotor and stator systems, toward the goal of improving efficiency and increasing combustion power extraction, additional cooling and other mitigating systems have been developed to achieve incremental gains in efficiency with superalloy-based turbines. However, the financial cost as well as the marginal “cost” in cooling air and other efficiency losses inherent in these mitigating systems can approach any marginal efficiency gain achievable with ever-higher turbine temperatures.
As a next-stage technology, ceramic matrix composites (CMCs) are being adopted into certain parts of the engine hot section to overcome these incremental losses. One of the main attractions of CMCs is the potential gain in thermal resistance over the most recent generations of superalloys. But as with any large step-change, a number of other technical issues arise when implementing CMCs into existing systems. One issue that frequently comes up is the porosity of CMCs, so that under increased pressures and temperatures needed for increased turbine efficiency, several components of the combustion fluid can infiltrate an unprotected composite matrix and severely shorten the useful life of CMC components.
An embodiment of a turbine engine outer case ring assembly includes a substantially circular first main body and a sealant layer covering at least a main gas-facing surface of the inner gas path side, the sealant layer having at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
An embodiment of a turbine engine outer case assembly includes a plurality of first outer case ring assemblies arranged axially along a central longitudinal axis to define at least part of a turbine section case. Each of the plurality of first outer case ring assemblies include a substantially circular first main body having a plurality of first porous ceramic matrix composite (CMC) segments, each first segment including an outer mounting side and an inner gas path side. A sealant layer covers a main gas-facing surface of the inner gas path side, the sealant layer having at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
An embodiment of a turbine module includes a turbine engine outer case assembly including a plurality of first outer case ring assemblies and a plurality of second outer case ring assemblies, each alternating and centered axially along a longitudinal axis to define at least part of a turbine section case. Each of the axially alternating first and second outer case ring assemblies include a substantially circular main body including a plurality of porous ceramic matrix composite (CMC) segments, each segment including an outer mounting side and an inner gas path side. A sealant layer covers a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body. A plurality of turbine rotors are centered and rotatable about the longitudinal axis, each turbine rotor axially aligned with all of the first outer case ring assemblies or all of the second outer case ring assemblies.
Fan section 18, compressor section 20, combustor section 22, and turbine section 24 are arranged sequentially along centerline 12 within engine housing 26. Engine housing 26 includes inner case 28 (e.g., a core case) and outer case 30 (e.g., a fan case). Inner case 28 may house one or more of (or even a portion of) fan section 18, compressor 20, combustor section 22, and turbine section 24 (e.g., an engine core). Outer case 30 may house at least fan section 18. Each of gas turbine engine sections 18, 20A, 20B, 24A and 24B includes respective rotors 32-40. Each of these rotors 32-40 includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s).
Fan rotor 32 is connected to gear train 42, for example, through fan shaft 44. Gear train 42 and LPC rotor 34 are connected to and driven by LPT rotor 40 through low speed shaft 46. Alternatively, LPC rotor 34 may be connected to and driven by gear train 42. The combination of at least LPC rotor 34, LPT rotor 40, and low speed shaft 46 may be referred to as “a low speed spool.” HPC rotor 36 is connected to and driven by HPT rotor 38 through high speed shaft 48. The combination of at least HPC rotor 36, HPT rotor 38, and high speed shaft 48 may be referred to as “a high speed spool.” Shafts 44-48 are rotatably supported by a plurality of bearings 52, which can be rolling element bearings, thrust bearings, or other types of bearings. Each of these bearings 52 is connected to engine housing 26 by at least one stationary structure such as, for example, an annular support strut.
During operation, air enters gas turbine engine 10 through airflow inlet 14. Air is directed through fan section 18 and is then split into either core gas path 54 or bypass gas path 56. Core gas path 54 flows sequentially through fan section 18, compressor section 20, combustor section 22, and turbine section 24. The air within core gas path 54 may be referred to as “core air.” Bypass gas path 56 flows through a duct between inner case 28 and outer case 30. The air within bypass gas path 56 may be referred to as “bypass air.”
The core air is compressed by LPC rotor 34 and HPC rotor 36 and directed into combustion chamber 58 of combustor 60 in combustor section 22. Fuel is injected into combustion chamber 58 and mixed with the core air that has been compressed by compressor section 20 to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof expand and flow through and sequentially cause HPT rotor 38 and LPT rotor 40 to rotate. The rotations of HPT rotor 38 and LPT rotor 40 drive rotation of HPC rotor 36 and LPC rotor 34, respectively and compression of the air received from core gas path 54. The rotation of LPT rotor 40 also drives rotation of fan rotor 32, which propels bypass air through and out of bypass gas path 56.
Some of the main attractions of CMCs are the reduced weight and/or potential gain in thermal resistance over the current and expected advances in superalloys for rotors and stators. But as with any large step-change, a number of other technical issues arise before successfully and cost-effectively implementing CMCs into existing turbine stator systems. One issue with the use of CMCs is its porosity, so that if used untreated or otherwise unprotected under increased pressures and temperatures similar to a pressure vessel—such as in the hot section of a gas turbine engine, several components of the combustion/working fluid can easily infiltrate an unprotected composite matrix and severely shorten the useful life of CMC components.
Nearly all CMCs suitable for use in a hot section (e.g., SiC, carbon/carbon, etc.) have relatively low alpha (thermal expansion) coefficients compared to those of metals. Thus, it becomes desirable to consider, and adapt, structural parts of the hot section to operate as a quasi-static pressure vessel rather than as a highly variable and flexible metallic system that must adapt to temperature variations of 1000° F. or more throughout every flight cycle. By considering the outer structure as a quasi-static pressure vessel and designing the structural material in ways that can minimize or prevent working gases from infiltrating the CMC, life of the outer CMC ring structures can be increased while also taking advantage of the decreased weight, their thermal resistance and low alpha properties.
To this end,
This portion of turbine section 66 includes at least metallic rotor assembly 68 and outer case ring assembly 70 annularly arranged therearound. Assembly 70 includes at least a substantially circular first main body 72 comprising a plurality of first porous (when untreated) ceramic matrix composite (CMC) blade outer air seal (BOAS) segments 74. Each BOAS segment 74 includes outer mounting side 76 and inner gas path side 78. To the end of reducing or eliminating negative effects of this porosity at typical hot section operating temperatures and pressures, at least one external sealant layer 82 covers main gas-facing surface 84 of inner gas path side 78. External sealant layer 82 includes at least one coating sufficient to prevent infiltration of working gas into pores of the main body. In certain embodiments external sealant layer 82 is substantially impermeable to oxidizing gases typically present in the heated combustion/working gas. The sealant layer(s) allow the fixed CMC components to operate as a quasi-static pressure vessel as described above with minimal risk of infiltration and resulting damage.
Turbine section 66 can include sealed stator outer platform segments 88 axially upstream or downstream of BOAS segments 74. In the event that main body 72 makes up stator platforms 88, one or more CMC stator vanes 90 can also extend radially from each of the plurality of stator outer platform segments.
To minimize axial variation in thermal expansion coefficients, it is desired in certain embodiments that both the BOAS and the outer stator platforms are formed from sealed CMC segments such as those described above. To that end, turbine engine outer case assembly 70 can include both a plurality of first outer case ring assemblies arranged axially along a central longitudinal axis A to define at least part of a turbine section case, and a substantially circular second main body comprising a plurality of second porous ceramic matrix composite (CMC) segments.
In the most common configuration, the CMC segments alternate axially between a plurality of blade outer air seal (BOAS) segments and a plurality of stator outer platform segments. CMC segments can further comprise one or more CMC stator vanes extending radially inward from each of the plurality of stator outer platform segments.
Each first and second segment can include an outer mounting side and an inner gas path side. As above, each of the plurality of first outer case ring assemblies includes a substantially circular first main body comprising a plurality of first porous ceramic matrix composite (CMC) segments, each first segment including an outer mounting side and an inner gas path side. A sealant layer covers a main gas-facing surface of the inner gas path side, the sealant layer including at least one external coating sufficient to prevent infiltration of working gas into pores of the main body.
Without surface treatment, the porosity of a CMC substrate is susceptible to infiltration of working gas from core path 54. Some of the combustion byproducts can rapidly wear down the structural materials in the composite, which has hindered broader adoption of more thermal resistant CMC materials. To this end, at least external sealant 82 (also shown in
By way of explanation and enablement, several combinations of CMCs and sealants could be used but not all combinations have been tested for suitability, compatibility, and useful life. It is expected that, generally speaking, the outer case will have the following minimum mechanical and thermal requirements. At a typical turbine flow path pressure of 400-1000 psig (about 2750-6900 kPa) and at least 1500° F. (about 825° C.), the working fluid will not infiltrate substrate 90 due to sealant layer 82 and/or diffusion layer 96. The fibers will have a density and orientation so that a tensile strength of at least 300 ksi (about 207 MPa), and in some embodiments up to or exceeding 500 ksi (about 345 MPa) on a constant basis. And toward the goal of approximating a pressure vessel, the above design conditions will not result in more than 2% radial deflection of the main body or outer ring assembly relative to ambient conditions.
In certain embodiments, an active clearance control (ACC) system, adapted particularly to minimize the efficiency losses due to e.g., excess weight and cooled air usage, can be disposed radially outward of at least one of the plurality of first or second outer case ring assemblies. One such example, using a clearance control ring not directly tied to the engine case or seal supports, is described in U.S. Patent Application Publication 2018/0320542 A1 (Ser. No. 15/589,009) by Suciu et al., which is hereby incorporated by reference in its entirety. The incorporated application is pending and commonly assigned, along with the instant application, to United Technologies Corporation.
The following are non-exclusive descriptions of possible embodiments of the present invention.
An embodiment of a turbine engine outer case ring assembly includes a substantially circular first main body and a sealant layer covering at least a main gas-facing surface of the inner gas path side, the sealant layer having at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
The assembly of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A turbine engine outer case ring assembly according to an exemplary embodiment of this disclosure, among other possible things includes a substantially circular first main body comprising a plurality of first porous ceramic matrix composite (CMC) segments, each first segment including an outer mounting side and an inner gas path side; and a sealant layer covering a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
A further embodiment of the foregoing assembly, wherein additionally and/or alternatively, the plurality of first porous CMC segments comprise a plurality of blade outer air seal (BOAS) segments.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous ceramic matrix composite (CMC) segments comprise carbon fibers and silicon carbide (SiC) fibers, and combinations thereof.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous ceramic matrix composite (CMC) segments comprise silicon carbide (SiC) matrix.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the sealant layer comprises an amorphous material.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous CMC segments comprise a plurality of stator outer platform segments.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous CMC segments further comprise one or more CMC stator vanes extending radially from each of the plurality of stator outer platform segments.
An embodiment of a turbine engine outer case assembly includes a plurality of first outer case ring assemblies arranged axially along a central longitudinal axis to define at least part of a turbine section case. Each of the plurality of first outer case ring assemblies include a substantially circular first main body having a plurality of first porous ceramic matrix composite (CMC) segments, each first segment including an outer mounting side and an inner gas path side. A sealant layer covers a main gas-facing surface of the inner gas path side, the sealant layer having at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
The assembly of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A turbine engine outer case ring assembly according to an exemplary embodiment of this disclosure, among other possible things includes a plurality of first outer case ring assemblies arranged axially along a central longitudinal axis to define at least part of a turbine section case, each of the plurality of first outer case ring assemblies comprising: a substantially circular first main body comprising a plurality of first porous ceramic matrix composite (CMC) segments, each first segment including an outer mounting side and an inner gas path side; and a sealant layer covering a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
A further embodiment of the foregoing assembly, wherein additionally and/or alternatively, the plurality of first porous CMC segments comprise at least one of a plurality of blade outer air seal (BOAS) segments and a plurality of stator outer platform segments.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous CMC segments further comprise one or more CMC stator vanes extending radially from each of the plurality of stator outer platform segments.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively further comprises a substantially circular second main body comprising a plurality of second porous ceramic matrix composite (CMC) segments, each second segment including an outer mounting side and an inner gas path side; and a sealant layer covering a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of second porous CMC segments comprise the other of a plurality of blade outer air seal (BOAS) segments and a plurality of stator outer platform segments.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the plurality of first porous ceramic matrix composite (CMC) segments comprise one or more of carbon fibers and silicon carbide (SiC) fibers in a silicon carbide (SiC) matrix.
A further embodiment of any of the foregoing assemblies, wherein additionally and/or alternatively, the sealant layer comprises an amorphous material.
An embodiment of a turbine module includes a turbine engine outer case assembly including a plurality of first outer case ring assemblies and a plurality of second outer case ring assemblies, each alternating and centered axially along a longitudinal axis to define at least part of a turbine section case. Each of the axially alternating first and second outer case ring assemblies include a substantially circular main body including a plurality of porous ceramic matrix composite (CMC) segments, each segment including an outer mounting side and an inner gas path side. A sealant layer covers a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body. A plurality of turbine rotors are centered and rotatable about the longitudinal axis, each turbine rotor axially aligned with all of the first outer case ring assemblies or all of the second outer case ring assemblies.
The module of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A turbine module according to an exemplary embodiment of this disclosure, among other possible things includes a turbine engine outer case assembly comprising: a plurality of first outer case ring assemblies and a plurality of second outer case ring assemblies, each alternating and centered axially along a longitudinal axis to define at least part of a turbine section case, each of the axially alternating first and second outer case ring assemblies comprising: a substantially circular main body comprising a plurality of porous ceramic matrix composite (CMC) segments, each segment including an outer mounting side and an inner gas path side; and a sealant layer covering a main gas-facing surface of the inner gas path side, the sealant layer comprising at least one coating sufficient to prevent infiltration of working gas into pores of the main body; and a plurality of turbine rotors centered and rotatable about the longitudinal axis, each turbine rotor axially aligned with all of the first outer case ring assemblies or all of the second outer case ring assemblies.
A further embodiment of the foregoing module, wherein additionally and/or alternatively, at least one of the plurality of turbine rotors is metallic, and has a coefficient of thermal expansion greater than a coefficient of thermal expansion of the sealed ceramic matrix composite (CMC) segments.
A further embodiment of any of the foregoing modules, additionally and/or alternatively, further comprises an active clearance control (ACC) system disposed radially outward of at least one of the plurality of first or second outer case ring assemblies.
A further embodiment of any of the foregoing modules, wherein additionally and/or alternatively, the ACC system comprises a first mount feature and a second mount feature such that a clearance control ring of the ACC system can move independently of the first and second outer case ring assemblies.
A further embodiment of any of the foregoing modules, wherein additionally and/or alternatively, the plurality of first porous CMC segments comprise at least one of a plurality of blade outer air seal (BOAS) segments and a plurality of stator outer platform segments, and the plurality of second porous CMC segments comprise the other of a plurality of blade outer air seal (BOAS) segments and a plurality of stator outer platform segments.
A further embodiment of any of the foregoing modules, wherein additionally and/or alternatively, the turbine outer case assembly withstands a working gas pressure of at least 100 psig (690 kPa) at 1500° F. (˜825° C.) without infiltration of the working gas into the segments beyond the sealed gas-facing surfaces.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
This application claims the benefit of U.S. Provisional Application No. 62/903,304 filed Sep. 20, 2019 for “SEALED CMC TURBINE CASE” by G. L. Suciu.
| Number | Date | Country | |
|---|---|---|---|
| 62903304 | Sep 2019 | US |
