gas turbine project

[aukap]

Im a senior year student and having studied a little of gas turbines wished to do a project on this. I was studying combustors and any ideas on this will be appreciated

[racket]

Hi Aukap

Welcome to the Group :-)

Combustors can be as simple or as complicated as you want , what particular aspects of combustors/combustion are you interested in ??

Cheers
John

[aukap]

i did a little reading myself and i found that a combustion chamber is something that has to be designed befitting the requirements . I read about the hole areas in the primary ,secondary and tertiary zone being in the ration of 3:2:5 and since ill have to document the project i was wondering about the theoretical background of this result.I did find some stuff online but i found it too complicated to understand . Any help on this will be of immense benefit.
thank you

[racket]

Hi Aukap

Yep , most scientific papers on the subject are a bit complicated .

The 30/20/50 ratios has to do with burning/finishing off burning/ dilution , the "30%" Primary air for burning should produce complete combustion of enough fuel to produce the required turbine inlet temperature , the "20%" secondary air should complete any unfinished combustion and commence dilution and cooling of the gases whilst not cooling the gases too quickly to prevent combustion , the 50% finishes of the dilution bringing the combustion gas temperatures down to the required turbine inlet temperature ( T I T ).

As for having the total flametube hole area equal to the compressor wheel inducer area , that seems to be conventional ratio for our low compression engines as it produces an acceptable pressure drop across the flametube wall whilst producing the required jets of air entering the flametube.

Cheers
John

[mitch]

Just gonna throw this question in here because it seems like it could fit in with the current discussion, is there a general relationship between turbine inlet temperatures and turbine outlet temperatures?

Also racket, your above explanation of the flame tube and dilution zones was great! That helps me visualize what is occurring during combustion in there.

[racket]

Hi Mitch

Yep , it depends on the temperature rise in the compressor , assume you're idling at 5psi P2 ( 1.34 pressure ratio) and the comp is running at 73% efficiency on a 15 deg C day , there will be a ~35 deg C temp rise in the comp for a T2 exiting the comp of 50 deg C .

We now take that 35 degree rise , multiply by 0.24 and divide by 0.275 , these are the specific heat values for air and hot gases , this will require a temp drop of ~30 degrees across the turb stage to power the comp to produce the 35 deg rise.

If you have a Turbine Outlet Temperature ( TOT) of say 500 C then the T I T will be 530 C .

At a 40 psi P2 ( 3.72 PR) at say 70% effic the temp rise will be ~188 C ( 203 C T2) , multiply by 0.24, divide by 0.275 =164 deg C drop across the turb stage , ...........so if you have 650 deg C for the T O T , the T I T will be 814 deg C

This is the reason why we can generally get away with high TOTs at spoolup as the T I T and T O T are basically the same as theres little compression , but at high power settings the drop in temp across the turb stage is many times greater so we have to be careful .

The required Pressure Ratio (PR) across the turb stage varies with turb efficiency to get the required temp drop , a high effic turb stage will require less PR to achieve the required temp drop than a low effic turb stage , both T I T and T O T will be the same but there'll be less pressure energy downstream for making thrust or shaft horsepower with the low effic turb stage .

Cheers
John

[mitch]

Oh okay I think I understand what you are saying, so basically the temp difference between tit and tot is based on compressor efficiency. The temp drop is based on the amount of energy the turbine wheel sucks out of the fluid due to expansion, to power the compressor.

[racket]

Hi Mitch

Nope , the temperature difference across the turb stage is based on the temp rise in compression due to both the pressure ratio obtained as well as the comp efficiency , the turbine stage is just the reverse of the compressor stage , its just that the hot gases are "more energetic" so there doesn't need to be as big a temp drop as the temp rise in compression to balance out the power required vs the power produced.

The turb temp drop is a direct outcome of the energy extracted ...................using our 164 C degree drop from the example I gave in my last email , we multiply the 164 by our 0.275 to get 45.1 Centigrade Heat Units ( CHUs) , they're the centigrade equivalent of BTUs , we multiply the 45.1 by 1400 ( constant , as 1 CHU = 1400 ft pounds of energy) , then divide by 550 to get the horsepower ( 550 ft lbs/sec =1 hp ) , so 45.1 X 1400 = 63140 ft lbs/sec , div by 550 = 114.8 horsepower per pound of gases flowing through the turbine wheel per second .

Hope this makes things a little clearer :-)

Cheers
John

[mitch]

racket,
this is starting to help! I am trying to learn more about these engines, and you are definitely helping. Where did you learn all of this stuff!? I assume you were an engineer for someone like rolls royce or GE at some point?!
Thanks!

[racket]

Hi Mitch

LOL............no, I'm just a back yard tinkerer , no formal training whatsoever other than what I've taught myself from books.

It ain't that complicated thankfully

Cheers
John

[aukap]

I went through some stuff and apart from hole diameters one needs to put swirlers in order to get the "magic circles" ,how do you design the swirlers?

[racket]

Hi Aukap

The swirler needs to flow ~10% of the total airflow , how you design them is up to you , there are "radial" as well as "axial" swirl vanes used surrounding the fuel nozzle , there can also be louvers in the flametube end cap to produce counter rotating flows .

The simplest is probably a large washer that the fuel nozzle just fits inside off, the washer is then cut radially inwards almost to the centre hole and the metal between the cuts is twisted to form a fan , the extremities/tips of the "fan" are welded to the flametube endcap , you could always use the metal fan off a small electric motor.

The main thing is to have the designed flow area at ~10% of total flametube hole area , the swirler flow area is part of the total area , the other 20% is spread between the endcap (maybe 5%) and the side wall ( 15%)

Cheers
John

[aukap]

So I have decided to base the project on the combustion chamber. My intention is to vary the hole pattern and hole size from an initial design.
I require a little help on the hole sizes for the zones, I did check jet specs and although it gives you the total area of holes but the size of the holes is pretty much a decision one has to make, so what is a good start point

[racket]

Hi Aukap

Combustion chambers design can be a very complex issue with a multitude of issues to consider , I'd strongly suggest you get a copy of this book to start with , them perhaps ask pertinent questions

www.amazon.com/GAS-Turbine-Combustion-Second-International/dp/1560326735/ref=sr_1_2?ie=UTF8&qid=1421787306&sr=8-2&keywords=gas+turbine+combustion&pebp=1421787310461&peasin=1560326735

Jetspecs produces a basic working combustor , but thats all , if you want to experiment with flametube design you'll be needing a lot more info , combustion was one of the hardest parts of getting airborne gas turbines to work due to the level of combustion intensity required .

Cheers
John

[mitch]

Here is a pretty good video related to the subject:


www.youtube.com/watch?v=3Y-U7FT7AU4