TR Propulsion Systems

Windchime

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TS1400 turbo shaft turbine inlet temperatures should not be too high. LHTEC CTS800 engine (ts1400 was based on the same radial flow design) has a TIT of 1050 degrees. When TS1400 was first made, to stress on the single crystals used, it was mentioned that the TIT could go as high as 1350 degrees C. - not necessarily the TIT value of TS1400. I don’t have the exact TIT values for TS1400. If you do I would appreciate it if you could share.

TF6000/10000 turbofan engines are being produced primarily as test bed engines for the KAAN’s prospective TF35000 engine. This engine is to contain technologies and specifics that would lead to an engine that can produce around 26000lbf dry and 35000lbf wet thrust similar to F119. Prof Aksit himself claimed that at TEI they could produce such an engine. Now proof of the pudding is in the eating.
To be used as test bed, TF6K is to contain third generation single Crystal blades, Blisk fans and compressors. Plus many parts being produced with additive technologies. F119 has a TIT value of 1649 degrees C and uses third generation single Crystal blades with Blisk fan and compressor sections. So TF6000 is being produced with similar tech on board. So I think it would be wrong to draw analogies between TF6K and TS1400. Next few months will be very enlightening for all.
I see, so I was wrong about the technological interconnect between the TS1400 and TF-6000. The TIT figure for the TS1400, which I was informed about to be around 1600K at emergency power output, was something I've read on this very forum, possibly on this thread.

So it should either be that I've misremembered the details (mixing up possible maximum temperature of TS1400 turbine blades as the actual maximum operational temperature of the engine) or the person who've published said figures was citing wrong information, be it intentional or not.

Nevertheless, I am still of the opinion that TF-6000 combustor exit temperature/turbine inlet temperature would at least be around the same ballpark of F404, thus TIT wouldn't be the major factor when it comes to the reason the two engines have such different thrust figures. I still think that the lower thrust is due to smaller engine core and thus lower MFR, as well as lower OPR.

Now going back to the part of the Defence Turkey article you have mentioned on your previous post, I actually find it funny what is conveyed in that paragraph.
by replacing the HP fan and LP turbine and rescaling the compressor, combustion chamber, and HP turbine.
Yeah, that's what we usually call a "new design". As much as it would be great to be able to simply "rescale" the engine and that it just works, it isn't as simple as that in reality. You are basically changing the entierty of the engine core at that point. I should say that there's quite an exaggeration at play on that assertion.

Talking about the level of technology applied on the TF-6000 as well as its role as a precursor, or 'demonstrator' if I can put it that way, for the TF-35000, it surely is very encouraging that there are said progresses being made, but there are obviously many more important aspects to an advanced low-bypass gas turbine if you want to compare it to an engine as advanced as F119, although it is basically 3 decades old technology at this point.

I get to see the "3rd generation single crystal blades" "blisks" "additive manufacturing" being repeteadly mentioned to highlight Turkish progress in the field of gas turbine manufacturing, but there are much more to the advanced manufacturing techniques and more importantly, design of the operationally capable and cost-effective low-bypass gas turbine. I'm sure you are already aware of such @Yasar , but I just wanted to stress this again, so bear with me. So, although TF-6000 will be able to demonstrate at least part of the basic Turkish capabilities in designing and manufacturing indigenous turbofans, which is an important stepping stone for the TF-35000, what's probably more important about the TF-6000 is the effect of the program itself to the entire gas turbine industry and R&D ecosystem. It will basically introduce Turkiye to basic capabilities and infrastructures, an overall structure to develop more advanced and sophisticated gas turbines in the future. We are already seeing new facilities being built for TRMotor, for example.

Yes, we'll get to know more details about the TF-6000 and where Turkiye is currently at in terms of gas turbine technologies, but if we are talking about TF-35000, its still a very long shot.
 

Yasar_TR

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Nevertheless, I am still of the opinion that TF-6000 combustor exit temperature/turbine inlet temperature would at least be around the same ballpark of F404, thus TIT wouldn't be the major factor when it comes to the reason the two engines have such different thrust figures.
First F404 engines used in Hornets were tested by NASA and the paperwork mentions 770-800degrees Centigrade TIT values. If I remember correctly they did not have FADEC either and had a thrust level of just over 16000lbf. Since then they have undergone many modifications whereby thrust levels have gone over 19000lbf now. This has been achieved by more modern parts being used and higher temperatures at the exit of combustion chamber thanks to single crystals etc.
Now going back to the part of the Defence Turkey article you have mentioned on your previous post, I actually find it funny what is conveyed in that paragraph.
That is a bit funny actually. You are right. They might as well have said produce a new engine. F404 has 3 LP and 7 HP stages in its compressor. I thought the author was trying to describe an engine similar to it with the amount of new stages he was suggesting.

Regarding the TF35K engine, in between the lines of various interviews it was mentioned that the engine is ready in digital domain. Now it has to take physical form. Before attacking such a project they have decided to produce a smaller scale model in the form of TF6K.
I agree with you; this is a gargantuan project to tackle. prof Aksit had said that his team is being asked to achieve the impossible. He had said before that It takes minimum of 14 years from scratch for a company like TEI to fly an engine like TF35K. He says his team knows how to produce and have produced f110 class engines. But this is a 5th generation engine they are being asked to produce in such a short time. (they have produced 50% parts of f110-129, f110-132 and f118 engines- including fans, compressors and combustion chambers. And assembled these engines). The teams for the engine design, have been working for over 4 years now. They are supposed to fly an engine by 2028. Personally I find it very difficult without outside help.
After TF6K we will see where we stand. Like you said; more than being capable of building such an engine, the more important point is having the right “eco-system” in place.
Again let us wait and see!
 

Nilgiri

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There are issues with scaling up a core past the flow aerodynamics and temperatures.

There are serious issues with things like vibrations and bearing fatigue as well.

Give multi-variate multi-dimensional transfer functions involved, it can be modelled with monte carlo simulations to some degree....to give sense of where human resources are best arranged and used in 5 - 10 year time chunks. No idea if TEI has set up an experienced team to do that kind of thing, what is the overall computing power they have steady access to and process flow and handling for this. Lot of it comes only with time and some level of mistakes made, there is lot current PW builds upon from earlier PW team foundations etc and all documentation related to (real or likely) dead ends and why.

There is an issue with lot of montecarlo sims too the more complex they become....tied to things like the sample error reducing by a magnitude (10) only if you increase the sample by twice that magnitude (100)...and what this cascades to for computing (to direct and optimise lab work and engineering later).

Large part of my current job it to revisit a number of old code we have running reliably to sandbox and optimize them better (especially some of its random number generators in the old FORTRAN that need updating now) because processing time is always at premium and given way montecarlo and other stochastic sims can cascade wildly to begin with, its important RNG intrinsic to their inputs are actually as random as possible.
 

Windchime

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First F404 engines used in Hornets were tested by NASA and the paperwork mentions 770-800degrees Centigrade TIT values. If I remember correctly they did not have FADEC either and had a thrust level of just over 16000lbf. Since then they have undergone many modifications whereby thrust levels have gone over 19000lbf now. This has been achieved by more modern parts being used and higher temperatures
770~800℃ is way too low, and is also quite a bit far off from the temperature range I am aware of. It is probably measured on other parts of the engine. The actual figure known is 1660K TIT for GE-400 variant. Add another 70K to that figure for GE-402 and RM12 variants.
 

Yasar_TR

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770~800℃ is way too low, and is also quite a bit far off from the temperature range I am aware of. It is probably measured on other parts of the engine. The actual figure known is 1660K TIT for GE-400 variant. Add another 70K to that figure for GE-402 and RM12 variants.
You are correct. My bad. I was checking US documents. Mostly in Fahrenheit. Reading 1390 to 1450 degrees as Fahrenheit was my mistake. They were in fact in Celsius. They correspond with your figures.
Recheked lhtec800 TIT ; It is 1007 degrees Celsius max.
 

Zafer

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What should we expect as TIT for the TEI TF-6000 engine? About 1600-1650 Celsius?
1400 °C is achievable with metal without cooling, you put gains with cooling and coating on top of this. There is no telling how much we gained by such improvements.
 

Afif

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What should we expect as TIT for the TEI TF-6000 engine? About 1600-1650 Celsius?

Should be around 1300-1400.
Slightly higher than other engines in class. (I am just speculating given TF-6000 will achieve same thrust with lower SFC, logically combustion temperature would be little higher)

Iirc, according to Mr. Aksit current Turkish single crystal blades can withstand up to 1500c. (1800k)

(It was posted in this thread)

To my very basic understanding, usually highest turbine inlet temperature is 100-150 degrees lower then the highest temperature blades can withstand. (For safety reasons and endurance) I could be wrong though, maybe @Yasar and @Nilgiri could tell us more.
 

Yasar_TR

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What should we expect as TIT for the TEI TF-6000 engine? About 1600-1650 Celsius?
It is possible with some “higher” bypass engines to achieve higher thrust levels than their low by pass counterparts whilst keeping TIT the same. Or alternatively with lower TIT at the same thrust level.

For a jet engine ; the higher the energy you throw out of the nozzle the more thrust you achieve.

So one way to do this is by increasing TIT and hence increase the energy thrown out of the nozzle.
The other way is to increase the amount of bypass stream whilst keeping TIT at a lower more manageable level. (By the way these effect the engines performance differently at different altitudes)

It seems that we may be adopting the latter.

F119 engine has a bypass ratio of 0.3:1 with a TIT of 1649 degrees C. It is clear that TEI is not building an engine that resembles F119. But whatever the final TF35K is going to be; This TF6K is going to resemble that. TF6K is going to have a bypass ratio of 1.08:1. That means a lot of thrust will be generated by the bypass stream. (1.08 kg mass of air is being thrown out for every 1kg mass of air flowing through the engine core) . It was mentioned that , of the 6000lbf thrust , approximately 2000lbf of it will be provided by the bypass stream.

If our 3rd generation single crystals are as efficient and good as the ones used in f119, they should be operable up to 1700 degrees C. But the size of blades and the rotational centrifugal forces they are subjected to play a part as well. Prof Aksit mentioned that the 3rd generation blades they built had 20-25% higher tensile strength when compared to foreign made counterparts. That is good but that doesn’t give you the performance levels of the blade. Making a Crystal means, aligning metal molecules in straight line as well as you can. That increases tensile strength. But also makes the material brittle and more susceptible to breakages. So how well our 3rd generation blades will perform and under what TIT conditions are still to be seen. At the moment TIT of these engines, is anybody’s guess.
 

TheInsider

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BPR of TF-6k is not a problem because it has a very efficient fan and compressor section that can push more air into the engine so TEI can design a higher bypass engine without losing too much performance. 1.08 is a sweet spot for TF-6k.
 

Nilgiri

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Should be around 1300-1400.
Slightly higher than other engines in class. (I am just speculating given TF-6000 will achieve same thrust with lower SFC, logically combustion temperature would be little higher)

Iirc, according to Mr. Aksit current Turkish single crystal blades can withstand up to 1500c. (1800k)

(It was posted in this thread)

To my very basic understanding, usually highest turbine inlet temperature is 100-150 degrees lower then the highest temperature blades can withstand. (For safety reasons and endurance) I could be wrong though, maybe @Yasar and @Nilgiri could tell us more.

It depends what you mean by "withstand up to".

Using Mr. Aksit's figure of 1500C for example (under some setting/context and overall tradeoff design elected for maybe the earlier TS-1400 prototype testing), the blade doesn't suddenly melt or lose cohesion at higher temperatures than this.

i.e the largest driving bedrock technologies in blade survival are things like compressor bleed air cooling and thermal barrier coatings (i.e things that allowed turbines to achieve temperature profiles that exceeded the melting point of Nickel superalloys to begin with by drastically reducing what the actual blade alloy itself experiences along with the alloy being picked to have that melting temp as high as possible while retaining all mechanical advantages metals bring).

Single Crystal technology maturation simply helped extend lifetime of the turbine blades (ceterus paribus) by ~ 3 - 5 times (compared to traditionally cast and early directional solidification processes) at the higher end operating temperature range we are talking about to give sense of the buffer "on top" of raw survival addressed by the base mitigation technologies.

The 1500C associates with some picked E(X) expected lifetime of the single crystal blade. You reduced the E(X), you can increase that withstand temperature and vice versa.

Refer to my older post: https://defencehub.live/threads/tf-...iner-aircraft-projects.5/page-187#post-203889

So there is no hard and easy way to say what the TIT will be in the end from the statements released so far.

The larger underlying tradeoff driving what Yasar mentions in his latest post has to do with there being an "ideal" compression ratio for every TIT w.r.t maximising heat engine work area (i.e available work capacity per mass flow rate).

i.e it is easy to understand why the first large increases in compression ratio (for every TIT, i.e heat engine max) benefit what can be extracted by the turbine....but once the compression ratios get really large this actually starts to decline given the working fluid is heated by compression and starts to diminish what the combustor can "add" (i.e combustion does much better thermal performance with cold air feed than hot air)....though thermal efficiency continues to increase with higher compression ratios (and this often gives extra options for pure turbojet designs that turbofans do not have, past the issue of fan tip velocity restricting fan diameter that way).

In essence you want to both limit how much you rely on compressor bleed air for turbine cooling (i.e allocating 10% for newer design compared to 20% of an older design is a big advantage) and there is also a limit to how big/long you can have the compressor once you have selected a max TIT (or at least a range to study its sensitivity impact on rest of engine).

This has ramifications for a turbofan more than for a turbojet (plain core only) given the fan ideally wants as high work capacity as possible from the core to run off for bypass. Or conversely the ability/tradeoff in doing this is what limits the bypass range on offer...i.e why we see lot of fairly low bypass ranges (i.e augmented turbojet) elected for in many designs well before the fan diameter issue can kick in... given the supersonic profile design driver in small volume problem that lends itself to turbojet as the natural basis.

i.e there are lot of things going on with feedback loops (on rest of engine) w.r.t TIT and turbine design drivers overall. These are only some of them from the thermodynamic profile.

From there its a great many wiggle rooms on offer as the larger design matures, so there is no way to tell at this early stage from statements as the (especially finalised) context is not really available for it from what I can see.
 

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Kale Arge prepares for KTJ-3700 and Turbofan surprise


We have three different engines that are currently qualified. These are the KTJ-3200, KTJ-1750 and KTJ-3700 engines. We have developed these engines primarily to respond to the needs of our country. We have now started mass production of these engines, and we have an intensive mass production process ahead of us, again primarily in line with the needs of our country.

The design of the KTJ-3700 engine was actually completed in 2020. In 2021, upon Roketsan's request for the engine needed for the ÇAKIR Missile, we quickly developed and prioritised the KTJ-1750 engine based on the KTJ-3700 design. As you know, our KTJ-1750 engine successfully completed its first flight test by flying the ÇAKIR Missile.

Subsequently, the engine requirement of the KARA ATMACA Missile emerged, and we quickly converted the KTJ-3700 design into a prototype engine. Since the KTJ-1750 and KTJ-3700 engines have largely the same design, the success we achieved in the KTJ-1750 guarantees that the KTJ-3700 will also be successful. As a matter of fact, in the first prototype engine tests we conducted this year, success was achieved immediately and we saw that the desired performance levels were achieved.

We are now producing the delivery engines required by Roketsan. We plan to deliver more than one KTJ-3700 to Roketsan this year.

We continue to implement the engine development roadmap agreed between Kale Arge and SSB in 2012, in line with the needs of the country. This roadmap started with the turbojet engine family and included other engine architectures, including turbofan. At the point we have reached today, we have created a modern turbojet engine family for missiles, and have largely started mass production. In line with the low fuel consumption engine requirements for new generation unmanned aerial vehicles, we have started small turbofan development studies in the background. With the progress of the studies, we will continue to share information about Turbofan.
 

zio

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I think there is some difficulties about implentation of KTJ 3200 to roketsan Atmaca missiles
 

uçuyorum

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It is so weird that Çakır has been tested with KTJ1750 but we haven't seen SOM or ATMACA with 3200 despite the fact that serial production began. Maybe in their design update for 3700 and 1750 they figured new things which didn't make their way to 3200 so they are having trouble now? And if 3700 has lower SFC with same diameter why not use that instead? I really wish someone with info could tell us what is wrong with Kale
 

Trakya_forever

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It is so weird that Çakır has been tested with KTJ1750 but we haven't seen SOM or ATMACA with 3200 despite the fact that serial production began. Maybe in their design update for 3700 and 1750 they figured new things which didn't make their way to 3200 so they are having trouble now? And if 3700 has lower SFC with same diameter why not use that instead? I really wish someone with info could tell us what is wrong with Kale
There is not a problem. Original project was KTJ 3500. However, at the first stage Kale couldn't reach 3,5 N. KTJ 3200 was accepted and the project went on. But in time, they did it and got 3,7 N from an engine with the same weight and diameter.
 

MADDOG

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Stating that accessories, cabling and pipes will be added to the prototype, Akşit stated, "Our team is working as fast as possible with serious dedication. It took its final form the day before it came to the fair, with the work carried out day and night. Our entire team also has calendar pressure. Our state needs it. I hope it works out. Hopefully we will also see it fly." he said.
Explaining that they follow the design, plastic model creation, exhibition model creation, real prototype creation, real engine production and flight processes for each new engine, Akşit said that the TF6000 engine is also in the early stages of the development process.
Asked when the engine sound will be heard, Akşit said, "I won't give a date. We are working as fast as possible. From the last TEKNOFEST to this TEKNOFEST, in a few months, it has turned into an engine that has been made from a few parts into flesh and blood. Hopefully, it will not take long. Our whole team will make sure that this time will not be too long. We hope that it works with all its might." he replied.


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Via @Avionot , @DrErincErdem1 , @gdhdefence
 
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