GE9X TAPS3 Combustor
GE9X TAPS3 Combustor
According to the GE video, the new TAPS 3 combustion liner has no air dilution holes like previous generation combustors. So how does it get the film of cooling and dilution air into the combustor to control the EGT at the nozzle guide vanes.
In the old style can type combustion liners (j75) there is a small scoop that takes air form the inner can to the outer can. Does anyone know what that is for?
In the old style can type combustion liners (j75) there is a small scoop that takes air form the inner can to the outer can. Does anyone know what that is for?
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GE is using a new ceramic matrix material for the combustor and turbines in their stage III design. This material can withstand higher temperatures compared to ceramic coated nickel/titanium alloys used previously. Combined with a new swirl pattern and increased air volume through the combustor, this eliminates the need for cooling bleeds - boosting efficiency and simplifying the engine design while reducing maintenance requirements.
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Turbine inlet nozzle guide vanes are internally cooled and film cooled using high pressure compressor bleed air. The bleed air enters an internal passages at the end of the guide vane. Some internal passages have an S-shape and makes several passes from one end of the vane to the other to maximize heat transfer to the cooling air. Some airflow is directed thru holes to provide intensive impingement cooling of internal faces where exterior surface temperatures are highest. The majority of the internal airflow is discharged thru a series of small holes to provide film cooling of the guide vane or blade.
I think the term "film cooling" is slightly misleading since it does fully describe how the approach works. While the boundary flow provided by the air discharged from the holes does indeed result in some cooling of the adjacent metal surface, the boundary airflow also acts as a very effective form of thermal barrier that minimizes heat transfer from the hot core gas flow to the vane surface.
Here is a schematic the shows the cooling airflow to the turbine inlet nozzle guide vanes and HP turbine blades.
Below is an image of the internal airflow in a HP turbine blade:
I think the term "film cooling" is slightly misleading since it does fully describe how the approach works. While the boundary flow provided by the air discharged from the holes does indeed result in some cooling of the adjacent metal surface, the boundary airflow also acts as a very effective form of thermal barrier that minimizes heat transfer from the hot core gas flow to the vane surface.
Here is a schematic the shows the cooling airflow to the turbine inlet nozzle guide vanes and HP turbine blades.
Below is an image of the internal airflow in a HP turbine blade:
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All very interesting RR, but the new GE TAPS III engines no longer require cooling air bleeds on the hot section parts that previous generations did. They have accomplished this by using a new ceramic composite material for the turbines and combustor sections which is able to withstand higher temps. That was I think the news!
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vapilot2004-
Sorry, I read the OP too quickly.
I know the TAPS III combustor for the GE9X will use CMC liner panels. This seems like a safe first step for use of the material in a commercial engine as the liners are not highly loaded, function mostly as a thermal barrier, and would not be sensitive to surface recession.
I also read that GE will be evaluating CMC turbine nozzle guide vanes and HP turbine blades on their GE9X test rig. But I doubt these more demanding applications will make it onto the GE90 engine model for the 777X. The technology is still relatively new, and there still needs to be more work done with regards to durability/reliability, QC in production, in-service NDI, etc.
I also believe your comment about CMC materials eliminating the need for film cooling of turbine inlet nozzle guide vanes and HP turbine blades may not be entirely correct. As I understand it, internal air cooling and film cooling of these CMC components is still necessary, but the amount of bleed air flow is reduced. The reduced demand for cooling bleed air flow allows a greater percentage of the HP compressor flow to be directed through the combustor mixing nozzles. Here is a picture of a TAPS combustor test rig with CMC liners. You can clearly see the air film cooling holes.
Here is a link to a NASA tech report describing some of the issues surrounding the use of CMCs for turbine engine hot section components.
The use of CMC materials does not really allow higher peak cycle temperatures in the engine. CMCs basically improve cycle efficiency by reducing thermal losses. Peak combustion cycle temperatures in commercial turbofan engines are currently limited by NOx formation, and NOx compounds only form when there is excess oxygen present at the flame front with temperatures greater than about 2800degF. The dilution air flow directed into the combustor helps to minimize NOx content of the exhaust gas, but I don't think it is an efficient way to regulate TIT.
Lastly, the biggest benefit of CMCs (as well as CFRP fan blades and Ti-Al HPC & LPT blades) on commercial turbofan engines is the significant weight savings they can potentially provide.
Sorry, I read the OP too quickly.
I know the TAPS III combustor for the GE9X will use CMC liner panels. This seems like a safe first step for use of the material in a commercial engine as the liners are not highly loaded, function mostly as a thermal barrier, and would not be sensitive to surface recession.
I also read that GE will be evaluating CMC turbine nozzle guide vanes and HP turbine blades on their GE9X test rig. But I doubt these more demanding applications will make it onto the GE90 engine model for the 777X. The technology is still relatively new, and there still needs to be more work done with regards to durability/reliability, QC in production, in-service NDI, etc.
I also believe your comment about CMC materials eliminating the need for film cooling of turbine inlet nozzle guide vanes and HP turbine blades may not be entirely correct. As I understand it, internal air cooling and film cooling of these CMC components is still necessary, but the amount of bleed air flow is reduced. The reduced demand for cooling bleed air flow allows a greater percentage of the HP compressor flow to be directed through the combustor mixing nozzles. Here is a picture of a TAPS combustor test rig with CMC liners. You can clearly see the air film cooling holes.
Here is a link to a NASA tech report describing some of the issues surrounding the use of CMCs for turbine engine hot section components.
The use of CMC materials does not really allow higher peak cycle temperatures in the engine. CMCs basically improve cycle efficiency by reducing thermal losses. Peak combustion cycle temperatures in commercial turbofan engines are currently limited by NOx formation, and NOx compounds only form when there is excess oxygen present at the flame front with temperatures greater than about 2800degF. The dilution air flow directed into the combustor helps to minimize NOx content of the exhaust gas, but I don't think it is an efficient way to regulate TIT.
Lastly, the biggest benefit of CMCs (as well as CFRP fan blades and Ti-Al HPC & LPT blades) on commercial turbofan engines is the significant weight savings they can potentially provide.
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I agree, RR, a fascinating bit of technology in materials research. 
According to GE Aviation, testing phase of CMC's in the 9X engine occurred in 2014. The company also says there will be CMC materials in the combustor and turbines on the new GE9X.
From specifications (pre design-lock, but current as of June 2015):
I think they have made a breakthrough, Riff. Quote from GE technical:
The go on to talk about how traditional turbine blades are laser drilled for air cooling passages and how the CMC eliminates this cooling requirement.
Regarding hot section temps, apologies if my comment was misleading. I agree, one firm goal is indeed to reduce overall temperatures for emissions standards, however engine hot section components will operate at much higher temps than ever before thanks to the new materials. That is what I intended to say.
The latest reference I could find on the NASA doc you linked was 2010. Perhaps tech has advanced considerably since publication?
Weight is a big one for sure, RR. The strength requirement of the turbine disk is one area where the lighter CMC blades will save weight and already the composite fan blades and case currently reduce engine weight without sacrificing performance or safety.
I also read that GE will be evaluating CMC turbine nozzle guide vanes and HP turbine blades on their GE9X test rig. But I doubt these more demanding applications will make it onto the GE90 engine model for the 777X.
From specifications (pre design-lock, but current as of June 2015):
Ceramic matrix composite (CMC) material in the combustor and high-pressure turbine.
I also believe your comment about CMC materials eliminating the need for film cooling of turbine inlet nozzle guide vanes and HP turbine blades may not be entirely correct.
The material has two hugely winning attributes for aviation: it’s one- the weight of metal, and it’s also heat-resistant and doesn’t need to be air-cooled.
Regarding hot section temps, apologies if my comment was misleading. I agree, one firm goal is indeed to reduce overall temperatures for emissions standards, however engine hot section components will operate at much higher temps than ever before thanks to the new materials. That is what I intended to say.
The latest reference I could find on the NASA doc you linked was 2010. Perhaps tech has advanced considerably since publication?
Weight is a big one for sure, RR. The strength requirement of the turbine disk is one area where the lighter CMC blades will save weight and already the composite fan blades and case currently reduce engine weight without sacrificing performance or safety.
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By the way, RR, you seem very knowledgeable on the subject of turbine engines. Question for you please - What is surface recession? One of our engineers thought it was a fancy way of saying wear? Is this correct?
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GE definitely seems to be leading the way right now with adoption of CMC components on commercial engines. But other companies like P&W have been evaluating the material for use on military engines for over a decade.
The NASA GRC paper I linked was published in January 2013, so I'm sure the information is mostly up to date. The paper states that the CMC material itself provides an increase in operating temperature of 200-300degF over current superalloy materials and permits a reduction in cooling airflow, but does not totally eliminate the need for cooling airflow. Here's a quote from page 2 of the paper:
"A. CMC Combustor
As a combustor liner material, SiC/SiC composites are an enabling material that can help meet the NOx reduction goals of ERA. Current superalloys require high cooling air flows to keep them below their maximum allowable operating temperatures (up to about 80% of their melting temperature). CMC materials offer operating temperatures that are 200º-300ºF higher than for superalloys. The higher temperature capability and less component cooling requirements allow for a wider combustor design space so that it can be run more efficiently. Less cooling flow to the component allows for more air to be put into the combustion process."
Here's an article from March of this year discussing GE's testing of CMC components on their GE9X rig. The first build, which began ground test in late January, uses CMC for combustor liner panels, HPT stage 1 nozzle shrouds, and HPT stage 2 nozzles. The second build of the test rig will have CMC HPT stage 1 nozzles, but since GE plans to lock down their design configuration by the middle of this year I doubt the CMC HPT stage 1 or stage 2 nozzles will make it into the first production model of the engine for the 777X. The commercial aircraft industry is very conservative when it comes to introducing new technologies like CMC materials in very demanding applications, and they have very legitimate reasons for being this way. If a CMC component were to experience unforeseen problems in service, the cost to the OEM for replacement and compensation paid to the operator for lost revenue could be devastating.
Lastly, about mid way through the article linked there is this quote about cooling air requirements for CMCs versus superalloys:
"More durable than metal, CMC components contain one-third the density of typical metallic parts, making them lighter and longer-lasting. Also more heat-resistant than metals, the material requires 20 percent less cooling air, which improves overall engine efficiency."
Interesting topic of discussion.
The NASA GRC paper I linked was published in January 2013, so I'm sure the information is mostly up to date. The paper states that the CMC material itself provides an increase in operating temperature of 200-300degF over current superalloy materials and permits a reduction in cooling airflow, but does not totally eliminate the need for cooling airflow. Here's a quote from page 2 of the paper:
"A. CMC Combustor
As a combustor liner material, SiC/SiC composites are an enabling material that can help meet the NOx reduction goals of ERA. Current superalloys require high cooling air flows to keep them below their maximum allowable operating temperatures (up to about 80% of their melting temperature). CMC materials offer operating temperatures that are 200º-300ºF higher than for superalloys. The higher temperature capability and less component cooling requirements allow for a wider combustor design space so that it can be run more efficiently. Less cooling flow to the component allows for more air to be put into the combustion process."
Here's an article from March of this year discussing GE's testing of CMC components on their GE9X rig. The first build, which began ground test in late January, uses CMC for combustor liner panels, HPT stage 1 nozzle shrouds, and HPT stage 2 nozzles. The second build of the test rig will have CMC HPT stage 1 nozzles, but since GE plans to lock down their design configuration by the middle of this year I doubt the CMC HPT stage 1 or stage 2 nozzles will make it into the first production model of the engine for the 777X. The commercial aircraft industry is very conservative when it comes to introducing new technologies like CMC materials in very demanding applications, and they have very legitimate reasons for being this way. If a CMC component were to experience unforeseen problems in service, the cost to the OEM for replacement and compensation paid to the operator for lost revenue could be devastating.
Lastly, about mid way through the article linked there is this quote about cooling air requirements for CMCs versus superalloys:
"More durable than metal, CMC components contain one-third the density of typical metallic parts, making them lighter and longer-lasting. Also more heat-resistant than metals, the material requires 20 percent less cooling air, which improves overall engine efficiency."
Interesting topic of discussion.
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vapilot2004,
If you think of ceramics in the same way as metallics, oxidation or corrosion are mechanisms, particularly when combined with heat, which do attack either of the materials. In turbine combustor, CMC applications, one of the environmental attacks comes from water. So in order to protect the base CMC material, a protective layer material is added to prevent either oxidation or corrosion leading to base material component failure. So, it becomes important to know the rate of surface recession or the resistance to environmental attack provided by the protective layer. It isn't so much wear as it is oxidation or corrosion of the component material. Hope this answers your question.
What is surface recession? One of our engineers thought it was a fancy way of saying wear? Is this correct?
Last edited by Turbine D; 29th Jun 2015 at 01:37. Reason: word corrections
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Turbine D, thank you for your answer. I understand perfectly and will share your illuminating response with our engineer. He's an amiable type and won't mind a differing perspective on things theoretical.
Riff Raff, thanks for your ongoing discussion and research. It is an interesting development to be sure and likely something similar will be used by all manufacturers in the future.
On our divergent understanding of CMC turbine blades and cooling, it could be I may be reading the optimistic news, while you are getting conservative input from your sources. GE Aviation claims a no-bleed cooling requirement for the 2nd stage turbine blades made of CMC.
Dan McCormick (GE head military division)
I agree commercial aviation is a conservative business and new materials are often introduced incrementally with a great deal of time and effort spent on proving their suitability for in-service operation.
Riff Raff, thanks for your ongoing discussion and research. It is an interesting development to be sure and likely something similar will be used by all manufacturers in the future.
On our divergent understanding of CMC turbine blades and cooling, it could be I may be reading the optimistic news, while you are getting conservative input from your sources. GE Aviation claims a no-bleed cooling requirement for the 2nd stage turbine blades made of CMC.
With a 27:1 pressure ratio, 20% higher than that in the GENx (in turn 20% higher than in the GE90), the HPC is one of the key changes being introduced in the GE9X, which will exclusively power the 777X.
Others are a “fourth-generation” composite fan with wider, thinner (and fewer) blades, and lightweight, uncooled ceramic matrix composites (CMC) for high-pressure turbine (HPT) nozzles, shrouds and blades.
Others are a “fourth-generation” composite fan with wider, thinner (and fewer) blades, and lightweight, uncooled ceramic matrix composites (CMC) for high-pressure turbine (HPT) nozzles, shrouds and blades.
“The CMC low-pressure turbine blade is about one-third the weight of the metal blade it replaces, and at the second stage, the CMC doesn’t have to be air-cooled. The airfoil can now be more aerodynamically efficient because it does not need all that cooling air pumping through the middle of it.”
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vapilot2004-
Sorry, I missed your post from 2 days ago. But Turbine D provided a good answer. Recession is just a gradual loss of surface material due to oxidation. The surface oxidation is primarily caused by exposure to high temperatures and water vapor.
If you read thru the NASA paper linked above, there is some discussion of oxidation of the CMC material (SiC-SiC) GE is testing for hot section components. From what I could see the upper safe temp limit for the uncoated CMC material in a liner/shroud application is about 2400degF. However, with a suitable environmental barrier coating (EBC) the upper temperature limit can be extended to around 2700degF.
EBCs are beneficial for preventing surface recession of low-stress static components like liners/shrouds/nozzle vanes. But they absolutely critical for reliability of highly-stressed dynamic components like turbine blades. If the EBC on a blade is damaged due to impact from a piece of debris, and surface temperatures are high enough to produce significant oxidation, the blade could suffer catastrophic structural failure. So the EBC needs to be very durable and damage tolerant.
Damage tolerance from debris impacts is an ongoing issue with the ceramic thermal barrier coatings currently used on metal blades/nozzles. Especially with engines operating in dusty/sandy environments. So there probably needs to be much more work done with regards to the long-term durability EBCs before CMC turbine blades are put into service.
Here's a good technical reference regarding oxidation of CMC materials.
Sorry, I missed your post from 2 days ago. But Turbine D provided a good answer. Recession is just a gradual loss of surface material due to oxidation. The surface oxidation is primarily caused by exposure to high temperatures and water vapor.
If you read thru the NASA paper linked above, there is some discussion of oxidation of the CMC material (SiC-SiC) GE is testing for hot section components. From what I could see the upper safe temp limit for the uncoated CMC material in a liner/shroud application is about 2400degF. However, with a suitable environmental barrier coating (EBC) the upper temperature limit can be extended to around 2700degF.
EBCs are beneficial for preventing surface recession of low-stress static components like liners/shrouds/nozzle vanes. But they absolutely critical for reliability of highly-stressed dynamic components like turbine blades. If the EBC on a blade is damaged due to impact from a piece of debris, and surface temperatures are high enough to produce significant oxidation, the blade could suffer catastrophic structural failure. So the EBC needs to be very durable and damage tolerant.
Damage tolerance from debris impacts is an ongoing issue with the ceramic thermal barrier coatings currently used on metal blades/nozzles. Especially with engines operating in dusty/sandy environments. So there probably needs to be much more work done with regards to the long-term durability EBCs before CMC turbine blades are put into service.
Here's a good technical reference regarding oxidation of CMC materials.
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You probably don't need to slog thru that entire paper. The parts relevant to oxidation of CMCs used in high temp conditions are sections 4 and 5.
Just to elaborate the point I made above about damage tolerance of coatings used to protect hot section component surfaces located in the gas flow path, below is a picture of oxidation damage to (metallic) stage 2 nozzle vanes that suffered degradation/loss of their TBC, which ultimately resulted in a burn-through of the case. This was a GP7270 engine on an Emirates A380 that was operated in conditions with sand/dust present. The vane performed fine (for several thousand hours) as long as the TBC was intact. But after loss of the TBC, there was not sufficient air cooling to prevent oxidation damage of the metal structure. This is the same problem EBCs used to protect CMC materials will face.
There is now an FAA AD issued that requires bore scope inspection of the nozzles every 100 cycles, which will be costly to airlines in terms of downtime/loss of revenue service. Replacement cost of the nozzles are approx $500K per engine, which I'm sure airlines would seek reimbursement from the engine OEM for. This is a good example of unforeseen problems that can occur in service with high-performance engine components that operate with razor thin margins of safety.
Just to elaborate the point I made above about damage tolerance of coatings used to protect hot section component surfaces located in the gas flow path, below is a picture of oxidation damage to (metallic) stage 2 nozzle vanes that suffered degradation/loss of their TBC, which ultimately resulted in a burn-through of the case. This was a GP7270 engine on an Emirates A380 that was operated in conditions with sand/dust present. The vane performed fine (for several thousand hours) as long as the TBC was intact. But after loss of the TBC, there was not sufficient air cooling to prevent oxidation damage of the metal structure. This is the same problem EBCs used to protect CMC materials will face.
There is now an FAA AD issued that requires bore scope inspection of the nozzles every 100 cycles, which will be costly to airlines in terms of downtime/loss of revenue service. Replacement cost of the nozzles are approx $500K per engine, which I'm sure airlines would seek reimbursement from the engine OEM for. This is a good example of unforeseen problems that can occur in service with high-performance engine components that operate with razor thin margins of safety.
