CA1111774A - Composite metal-ceramic turbine nozzle - Google Patents
Composite metal-ceramic turbine nozzleInfo
- Publication number
- CA1111774A CA1111774A CA327,058A CA327058A CA1111774A CA 1111774 A CA1111774 A CA 1111774A CA 327058 A CA327058 A CA 327058A CA 1111774 A CA1111774 A CA 1111774A
- Authority
- CA
- Canada
- Prior art keywords
- vanes
- metal
- turbine nozzle
- ceramic
- nozzle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Inventor: Theodore Ivanko Assignee: Avco Corp.
Atty. Docket: 74-15 COMPOSITE METAL-CERAMIC TURBINE NOZZLE
Abstract of the Disclosure A turbine nozzle has a metal inner shroud that is joined to a metal outer shroud by a limited number of metal blades for structural purposes, and includes a plurality of air flow guidance vanes made of ceramic material and connected to the shrouds by spring clips. The composite structure allows the turbine nozzle to operate at higher temperatures with less cooling.
Atty. Docket: 74-15 COMPOSITE METAL-CERAMIC TURBINE NOZZLE
Abstract of the Disclosure A turbine nozzle has a metal inner shroud that is joined to a metal outer shroud by a limited number of metal blades for structural purposes, and includes a plurality of air flow guidance vanes made of ceramic material and connected to the shrouds by spring clips. The composite structure allows the turbine nozzle to operate at higher temperatures with less cooling.
Description
7~
COMPOSITE METAL-CER~lIC TURBINE NOZZLE
E~k~o~ d of thu l~v~n=:~A
The present invention relates to a turbine nozzle which is located between the combustion chamber and the turbine wheel of a gas turbine, and more particularly a new and improved composite turbine nozzle having metallic structural blades and ceramic airflow vanes so as to re-duce the cooling air requirements for the turbine nozzle as well as to allow the turbine nozzle to operate at higher temperatures with less cooling air.
In the well known type of gas turbine commonly re--ferred to as a turbojet, air is compressed in a rotating compressor and heated in the combustion chamber, and then expanded through a gas turbine in which no excess power (above that required to drive the compressor) is supplied by the turbine. To increase the available energy in the temperature cycle, and hence the thrust and efficiency of the engine, designers have conventionally attempted to increase the turbine inlet temperature, since useful turbine engine power is directly related to turbine inle-t tempexature, as is turbine engine efficiency. In order to achieve a more efficient turbojet operation, i.e., higher cycle temperatures and hence higher thrust values for a given engine siæe, it has been proposed to use more sophisticated and advanced turbine airfoil cooling .
~ ~ ' .
... .-~' ' ' '' ., . ~ ' , ' :
.
~:
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techni~ues to permit higher turbine inlet temperatures.
With such techniques, the turbine vane and rotor blade temperatures may be brought within the capability of ex isting heat or oxidation resistance metallic materials (metals). Lacking the complete availability of such cooling methods or techni~ues, however, recourse is made to irnproved blade or vane materials and construction methods.
Heretofore, it has been known to use ceramic mat-erials in conventional hlade and vane designs, since ceramic materials have the ability to withstand signi- .
ficantly higher temperatures than the known refractory alloys. In particular, i.t has been known to use ceramic materials as the leading edge of the turbine airfoil where the temperature is always highest and where cooling :
is most difficult, since the heat input is highest at :~
the leading edge also. On the other hand, ceramic mat :~
erials present certain problems in their use, such as the fact that ceramic materials do not have the tensile strength of metallic materials, and also ceramic materials usually have relatively low ductility and thus have a tendency to crack under the impact of severe or suddenly applled thermal shock or stresse~ as may occur in gas turbines. . :
, .~2-~ ~ .
.
' , . ,~ ` ' , ` ` ~ ' 7~
Prior art attempts to utilize both metallic and cer amic materials in the construction of a gas turbine nozzle have employed t.le concept wherein a portio~ of each air-foil surface i5 formed of a ceramic material, while the remaining structural portion is formed of a metal:Lic material. This is exemplified b~ the teachings of ~. S.
Letters Patent No. 3,758,233 wherein a ceramic coating is applied to an airfoil shaped element for use in a gas turbine. Reference is also made to a publication article "Ceramic Gas Turbine Has A Promising Future." Iron Age, pages 37, 38 an~ 39, March 1, 1976 which suggests a ceramic turbine wheel stator made by injection molding the ceramic, then reaction sintering.
None of the prior art teachings, however, is a practical solution to the advantages of marrying ceramic kechnology and the present state of the art of manufactur-ing metallic shrouds and airfoil elements to allow them to exist in harmony, and to achieve maximum advantages of both materials, without compromising either the struc-tural or heat-dissipating characteristics thereof. The present invention overcomes the shortcomings of the prior art systems, and provides a new and improved composite ~urbine nozzle for use in a gas turbine device which is capable of reducing cooling air re~uirements and allowing higher temperature- with les~ cooling air.
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7'7~
Br.ief Summa3~ nvention It is therefore a primary object of the subject in-vention to provide an improved turbine nozz:le made of a composite metal and ceramic turbine element:s~ It is also an object of the invention to achieve a recluction of costs by the reduction of the number of cooled vanes needed in the turbine nozzle, while at the same time maintaini.ng dimensional tolerance control by using sufficient metal lic vanes between inner and outer metal shroud structures not otherwise connected. It is also an object of the subject inven-tion to further provide a new and improved turbine nozzle wherein metal vanes or a part of them are located in lower temperature regions of the turbine noz~
zle inlet flow.
The foregoing objects, as well as others, which will become apparent from the following detailed descript.ion, are accomplished by the present invention. In one form thereof, the invention provides a composite turbine noz- .
zle for use in a gas turbine device and having an outer metallic shroud or casing, an inner metallic shroud or casing, said casings being interconnected by a pluralit~
of metal vanes so as to integrate said outer and inner casings~ A plurality of ceramic vanes are intersperse~d with the metallic vanes and extend between said outer .
~4 "~ :
:- ' ' , .
7~
casin~ and said inner casing. By this arrangement, the cooling air requlrements for the hiyh inlek temperature airflow provided to the gas turbine can be reduced, or higher inlet temperatures can be accommodat:ed without additional cooling air. Cost bene~its are achieved as the number of cool vanes per engine is substantially re-duced, and the structural metal vanes may be selectively located in the lower temperature regions of the inlet to the turbine nozzle. E'urthermore, the use of metallic structural vanes maintains the structural integrity of the metal-to-metal control of the inner and outer shroud relationships, and the ceramic vanes are secured to the shrouds by spring clips whereby the ceramic vanes are not required to absorb axial excursions of the shrouds, and thus the ceramic vanes can ~e made of a simple one-piece design.
Detailed Description of -the Drawings For purpose of facilitating the description of this invention, the appended drawings are incorporated, wherein:
FIG. 1 is a cross--section of a turbine nozzle por-tion of a gas turbine engine of the prior art' FIG. 2 is a partial isometric of a cross-section of a turbine nozzle according to the subject invention; ; -~5~
- ~ . -: .
- . - . : , . .. : . . .
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.. .. : : . . .
7~
FIG. 3 is a par-tial isometric view of a stator vane segment for a turbine nozzle according to the subject in-vention; and FIG. 4 is a cross-sectional view of the metal vane in the turbine nozzle according -to the subject invention.
Detailed Description FIG. 1 illustrates the prior art structure of United States Letters Patent No. 3,619,077 wherein a cerarnic leading edge 10 is pxovided for a body structure 12 of an airfoil shaped stator vane between a liner 14 of an outer casing 16 and an inner liner 18 for an inner casing 20 of a gas turbine. The outer liner 16 and inner liner 18 form a turbine nozzle for a comhustion chamber (not shown) opening towards a turbine wheel, blade por-tion 22, being shown. In order to effect the desired force, the stator vane comprised of the leading edge 10 and body 1~ turn khe flow from the nozzle so as -to im-part a rotational force on the blades 22 of the tuxbine wheel.
It may thus be apprecia-ted that these stator vanes are exposed to very high ternperature gases. In order to obtain adequate life cycle, the all-metal blades were proposed to be changed to a composite of ceramic metal.
~6~
.
. . . . . . . . . : : - . . :
. . . . . .. . .
Also, the metal body was to be open to receive cooling air from passage 24 and exhaust same to passage 26 wherefrom it was generally directed via the turbine wheel section to exhaust in the gas turbine exhaust flow therefrom. It should be noted that this prior art also taught that the entire vane may be made of ceramic wafers with shims therebetween to effect the desired stator vane profile.
Such was located between the nozzle liners by a hollow pin to permit cooling air to flow through such ceramic structure as it was to flow in the metal body.
Contrasted to this is the new and improved composite turbine nozzla of the subject invention as illustrated in FIGS. 2-4. As shown in FIG. 2, the turbine nozzle of the invention includes a plurality of blades 30 formed from a ceramic composition to either side of an all metal blade 32 between outer metallic casing 34 and an inner metallic liner 36 connected C~Y means not shown~ to an inner casing 38~ The latter includes a pluxality of drilled passages 40 to conduct cooling air into the metal blade 32 wherefrom it is exhausted via surface openings 44 thereof. As seen in FIG. 4, each metal blade 32 also has trailing edge openings 48 ~or exhaust of cooling air~
As se n in FIG. 3 a spring clip SQ is attached to casing liner 36, and locates the ceramlc vane 30. Similar means axe suggested for the outer casing vane juncture.
'-:
.
: . . .. ,: . , .. : , .: ..
: . , - . . - . . . .
,: ~ - . - ~ . . : : ~ . :
7~
~ ccordingly, the subject invention provides a compo-site metal and ceramic turbine nozzle which reduces cool-ing air requirements and allows higher temperatllre~ with less cooling air. This will allow also the reduction of costs by the reduction oE the number of cooled me-tallic vanes needed in the nozzle, while at the same time ~ain-taining dimensional tolerance control by using su~ficient metal vanes between inner and outer metal structures, not otherwise connected. It is also possible to further improve this by locating the metal vanes or a part of them in lower temperature regions of the nozzle inlet flow from the combustor. It is also possible with the subject invention to realize a more advantageous employ-ment of inner and outer metal shroud flexibilities in diminishing shroud imposed vane loads. Finally, with the metal vanes 32 maintaining the axial posltion of the shrouds, the ceramic vanes 30 are not required to absorb axial excursions. This means that, in contrast to the prior art use of wafers to construct such ceramic vanes, the ceramic vanes of the subject invention can be o~ simple one piece design.
It will be understood that such other embodiments or obvious changes that are within the spirit and scope of the invention, as disclosed herein, are intended also : to be covered by the claims appended thexeto.
'.
COMPOSITE METAL-CER~lIC TURBINE NOZZLE
E~k~o~ d of thu l~v~n=:~A
The present invention relates to a turbine nozzle which is located between the combustion chamber and the turbine wheel of a gas turbine, and more particularly a new and improved composite turbine nozzle having metallic structural blades and ceramic airflow vanes so as to re-duce the cooling air requirements for the turbine nozzle as well as to allow the turbine nozzle to operate at higher temperatures with less cooling air.
In the well known type of gas turbine commonly re--ferred to as a turbojet, air is compressed in a rotating compressor and heated in the combustion chamber, and then expanded through a gas turbine in which no excess power (above that required to drive the compressor) is supplied by the turbine. To increase the available energy in the temperature cycle, and hence the thrust and efficiency of the engine, designers have conventionally attempted to increase the turbine inlet temperature, since useful turbine engine power is directly related to turbine inle-t tempexature, as is turbine engine efficiency. In order to achieve a more efficient turbojet operation, i.e., higher cycle temperatures and hence higher thrust values for a given engine siæe, it has been proposed to use more sophisticated and advanced turbine airfoil cooling .
~ ~ ' .
... .-~' ' ' '' ., . ~ ' , ' :
.
~:
7~ ~
techni~ues to permit higher turbine inlet temperatures.
With such techniques, the turbine vane and rotor blade temperatures may be brought within the capability of ex isting heat or oxidation resistance metallic materials (metals). Lacking the complete availability of such cooling methods or techni~ues, however, recourse is made to irnproved blade or vane materials and construction methods.
Heretofore, it has been known to use ceramic mat-erials in conventional hlade and vane designs, since ceramic materials have the ability to withstand signi- .
ficantly higher temperatures than the known refractory alloys. In particular, i.t has been known to use ceramic materials as the leading edge of the turbine airfoil where the temperature is always highest and where cooling :
is most difficult, since the heat input is highest at :~
the leading edge also. On the other hand, ceramic mat :~
erials present certain problems in their use, such as the fact that ceramic materials do not have the tensile strength of metallic materials, and also ceramic materials usually have relatively low ductility and thus have a tendency to crack under the impact of severe or suddenly applled thermal shock or stresse~ as may occur in gas turbines. . :
, .~2-~ ~ .
.
' , . ,~ ` ' , ` ` ~ ' 7~
Prior art attempts to utilize both metallic and cer amic materials in the construction of a gas turbine nozzle have employed t.le concept wherein a portio~ of each air-foil surface i5 formed of a ceramic material, while the remaining structural portion is formed of a metal:Lic material. This is exemplified b~ the teachings of ~. S.
Letters Patent No. 3,758,233 wherein a ceramic coating is applied to an airfoil shaped element for use in a gas turbine. Reference is also made to a publication article "Ceramic Gas Turbine Has A Promising Future." Iron Age, pages 37, 38 an~ 39, March 1, 1976 which suggests a ceramic turbine wheel stator made by injection molding the ceramic, then reaction sintering.
None of the prior art teachings, however, is a practical solution to the advantages of marrying ceramic kechnology and the present state of the art of manufactur-ing metallic shrouds and airfoil elements to allow them to exist in harmony, and to achieve maximum advantages of both materials, without compromising either the struc-tural or heat-dissipating characteristics thereof. The present invention overcomes the shortcomings of the prior art systems, and provides a new and improved composite ~urbine nozzle for use in a gas turbine device which is capable of reducing cooling air re~uirements and allowing higher temperature- with les~ cooling air.
-3~
.. .. .
.
: :. . . . . .
: . - ~ . . .
: ;. ~ ~, ' . ~. , ., ,, . . ,:
. , , . .. - :
7'7~
Br.ief Summa3~ nvention It is therefore a primary object of the subject in-vention to provide an improved turbine nozz:le made of a composite metal and ceramic turbine element:s~ It is also an object of the invention to achieve a recluction of costs by the reduction of the number of cooled vanes needed in the turbine nozzle, while at the same time maintaini.ng dimensional tolerance control by using sufficient metal lic vanes between inner and outer metal shroud structures not otherwise connected. It is also an object of the subject inven-tion to further provide a new and improved turbine nozzle wherein metal vanes or a part of them are located in lower temperature regions of the turbine noz~
zle inlet flow.
The foregoing objects, as well as others, which will become apparent from the following detailed descript.ion, are accomplished by the present invention. In one form thereof, the invention provides a composite turbine noz- .
zle for use in a gas turbine device and having an outer metallic shroud or casing, an inner metallic shroud or casing, said casings being interconnected by a pluralit~
of metal vanes so as to integrate said outer and inner casings~ A plurality of ceramic vanes are intersperse~d with the metallic vanes and extend between said outer .
~4 "~ :
:- ' ' , .
7~
casin~ and said inner casing. By this arrangement, the cooling air requlrements for the hiyh inlek temperature airflow provided to the gas turbine can be reduced, or higher inlet temperatures can be accommodat:ed without additional cooling air. Cost bene~its are achieved as the number of cool vanes per engine is substantially re-duced, and the structural metal vanes may be selectively located in the lower temperature regions of the inlet to the turbine nozzle. E'urthermore, the use of metallic structural vanes maintains the structural integrity of the metal-to-metal control of the inner and outer shroud relationships, and the ceramic vanes are secured to the shrouds by spring clips whereby the ceramic vanes are not required to absorb axial excursions of the shrouds, and thus the ceramic vanes can ~e made of a simple one-piece design.
Detailed Description of -the Drawings For purpose of facilitating the description of this invention, the appended drawings are incorporated, wherein:
FIG. 1 is a cross--section of a turbine nozzle por-tion of a gas turbine engine of the prior art' FIG. 2 is a partial isometric of a cross-section of a turbine nozzle according to the subject invention; ; -~5~
- ~ . -: .
- . - . : , . .. : . . .
. . . , . .. , . , . .: .
. . . ,. , . ,, , :
.. .. : : . . .
7~
FIG. 3 is a par-tial isometric view of a stator vane segment for a turbine nozzle according to the subject in-vention; and FIG. 4 is a cross-sectional view of the metal vane in the turbine nozzle according -to the subject invention.
Detailed Description FIG. 1 illustrates the prior art structure of United States Letters Patent No. 3,619,077 wherein a cerarnic leading edge 10 is pxovided for a body structure 12 of an airfoil shaped stator vane between a liner 14 of an outer casing 16 and an inner liner 18 for an inner casing 20 of a gas turbine. The outer liner 16 and inner liner 18 form a turbine nozzle for a comhustion chamber (not shown) opening towards a turbine wheel, blade por-tion 22, being shown. In order to effect the desired force, the stator vane comprised of the leading edge 10 and body 1~ turn khe flow from the nozzle so as -to im-part a rotational force on the blades 22 of the tuxbine wheel.
It may thus be apprecia-ted that these stator vanes are exposed to very high ternperature gases. In order to obtain adequate life cycle, the all-metal blades were proposed to be changed to a composite of ceramic metal.
~6~
.
. . . . . . . . . : : - . . :
. . . . . .. . .
Also, the metal body was to be open to receive cooling air from passage 24 and exhaust same to passage 26 wherefrom it was generally directed via the turbine wheel section to exhaust in the gas turbine exhaust flow therefrom. It should be noted that this prior art also taught that the entire vane may be made of ceramic wafers with shims therebetween to effect the desired stator vane profile.
Such was located between the nozzle liners by a hollow pin to permit cooling air to flow through such ceramic structure as it was to flow in the metal body.
Contrasted to this is the new and improved composite turbine nozzla of the subject invention as illustrated in FIGS. 2-4. As shown in FIG. 2, the turbine nozzle of the invention includes a plurality of blades 30 formed from a ceramic composition to either side of an all metal blade 32 between outer metallic casing 34 and an inner metallic liner 36 connected C~Y means not shown~ to an inner casing 38~ The latter includes a pluxality of drilled passages 40 to conduct cooling air into the metal blade 32 wherefrom it is exhausted via surface openings 44 thereof. As seen in FIG. 4, each metal blade 32 also has trailing edge openings 48 ~or exhaust of cooling air~
As se n in FIG. 3 a spring clip SQ is attached to casing liner 36, and locates the ceramlc vane 30. Similar means axe suggested for the outer casing vane juncture.
'-:
.
: . . .. ,: . , .. : , .: ..
: . , - . . - . . . .
,: ~ - . - ~ . . : : ~ . :
7~
~ ccordingly, the subject invention provides a compo-site metal and ceramic turbine nozzle which reduces cool-ing air requirements and allows higher temperatllre~ with less cooling air. This will allow also the reduction of costs by the reduction oE the number of cooled me-tallic vanes needed in the nozzle, while at the same time ~ain-taining dimensional tolerance control by using su~ficient metal vanes between inner and outer metal structures, not otherwise connected. It is also possible to further improve this by locating the metal vanes or a part of them in lower temperature regions of the nozzle inlet flow from the combustor. It is also possible with the subject invention to realize a more advantageous employ-ment of inner and outer metal shroud flexibilities in diminishing shroud imposed vane loads. Finally, with the metal vanes 32 maintaining the axial posltion of the shrouds, the ceramic vanes 30 are not required to absorb axial excursions. This means that, in contrast to the prior art use of wafers to construct such ceramic vanes, the ceramic vanes of the subject invention can be o~ simple one piece design.
It will be understood that such other embodiments or obvious changes that are within the spirit and scope of the invention, as disclosed herein, are intended also : to be covered by the claims appended thexeto.
'.
Claims (10)
1. In a gas turbine device a turbine nozzle between combustion chamber and a turbine wheel, said nozzle being characterized by:
an outer casing;
an inner casing;
a plurality of ceramic vanes between said outer casing and said inner casing; and a plurality of metal vanes interspersed with said ceramic vanes and integrating said outer casing and said inner casing.
an outer casing;
an inner casing;
a plurality of ceramic vanes between said outer casing and said inner casing; and a plurality of metal vanes interspersed with said ceramic vanes and integrating said outer casing and said inner casing.
2. The turbine nozzle of claim 1 and further char-acterized by means to supply cooling fluid to said metal vanes.
3. The turbine nozzle of claim 2 wherein the means to supply cooling fluid is a passage in the inner casing leading to a chamber open to a hollow interior of said metal vanes.
4. The turbine nozzle of claim 3 wherein the vanes have passages from the hollow interior through their sur-face in the nozzle to exhaust the cooling fluid thereinto.
5. A turbine nozzle comprising:
an outer casing, said outer casing being of a predetermined metal composition with means to affix its location;
an inner casing of similar metal, said inner casing having a passageway therethrough and also having means to affix its location;
a plurality of vanes also of similar metal as said outer casing, said vanes connecting said inner and outer casing and having a hollowed in-terior open to the passage of the inner casing; and a plurality of one-piece, i.e., solid, ceramic vanes between said outer casing and inner casing and located thereby so as to permit axial excursions of the outer casing and inner casing during flow of high temperature fluid in the turbine nozzle, said ceramic vanes being inter-leaved with the metal vanes.
an outer casing, said outer casing being of a predetermined metal composition with means to affix its location;
an inner casing of similar metal, said inner casing having a passageway therethrough and also having means to affix its location;
a plurality of vanes also of similar metal as said outer casing, said vanes connecting said inner and outer casing and having a hollowed in-terior open to the passage of the inner casing; and a plurality of one-piece, i.e., solid, ceramic vanes between said outer casing and inner casing and located thereby so as to permit axial excursions of the outer casing and inner casing during flow of high temperature fluid in the turbine nozzle, said ceramic vanes being inter-leaved with the metal vanes.
6. The turbine nozzle of claim 5 wherein the metal vanes have passages to exhaust cooling air therefrom
7. A turbine nozzle for high temperature gas flow control, said nozzle comprising:
a metal structure consisting of outer and inner casings connected by a limited number of cooled metal vanes; and a plurality of ceramic vanes interleaved with the metal vanes without being solidly inte-grated to said metal structure.
a metal structure consisting of outer and inner casings connected by a limited number of cooled metal vanes; and a plurality of ceramic vanes interleaved with the metal vanes without being solidly inte-grated to said metal structure.
8. The turbine nozzle of claim 7 wherein the ceramic vanes project through both outer and inner casings.
9. The turbine nozzle of claim 8 wherein spring clips affixed to the outer and inner casing frictionally locate the vanes without integration so as to load the vanes with forces created in the metal structure.
10. The nozzle of claim 9 wherein the metal structure has provisions to receive a cooling fluid flow and into the vanes and exhaust same therefrom.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA327,058A CA1111774A (en) | 1979-05-07 | 1979-05-07 | Composite metal-ceramic turbine nozzle |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA327,058A CA1111774A (en) | 1979-05-07 | 1979-05-07 | Composite metal-ceramic turbine nozzle |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1111774A true CA1111774A (en) | 1981-11-03 |
Family
ID=4114145
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA327,058A Expired CA1111774A (en) | 1979-05-07 | 1979-05-07 | Composite metal-ceramic turbine nozzle |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA1111774A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011082243A1 (en) | 2009-12-31 | 2011-07-07 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and frame |
-
1979
- 1979-05-07 CA CA327,058A patent/CA1111774A/en not_active Expired
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011082243A1 (en) | 2009-12-31 | 2011-07-07 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and frame |
| EP2519726A4 (en) * | 2009-12-31 | 2015-02-18 | Rolls Royce Nam Tech Inc | GAS TURBINE ENGINE AND STRUCTURE |
| US9284887B2 (en) | 2009-12-31 | 2016-03-15 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine and frame |
| US10151219B2 (en) | 2009-12-31 | 2018-12-11 | Rolls-Royce North American Technologies Inc. | Gas turbine engine and frame |
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