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Help With Project On Aircraft Engines  
User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Posted (9 years 2 months 4 days 4 hours ago) and read 7635 times:

Hi there,

I am currently a senior on Aerospace Engineering at my University. Instead of being a sensible person and sticking to the final projects in the list I decided to propose one myself and am now paying the penalty...

My project involves the preliminary sizing of a jet engine for a supersonic business jet, which isn't easy when you have to make up most of the specs and do all the design decisions yourself.

The main software I am using is a pre-historic version (v6.0) of GasTurb. It has an optimization function which I have used but unfortunately it only allows 5 variables to be optimized at once. Design condition is M1.8 at 55,000 ft.

Here is the main problem: After establishing most data (efficiencies, mass-flow, etc.) I was left with 5 variables (btw, it is a turbofan)
- Inner fan pressure ratio
- Outer fan pressure ratio
- HP compressor pressure ratio
- Burner exit temperature
- By-pass ratio
For some unknown reason the program uses inner and outer fan pressure ratio (inner being fan x LP compressor) instead of the such simpler fan and booster pressure ratio.

I have tried everything, from changing the initial design point to practically all other input data and the end result is always in the same form: an engine with an extremely high pressure ratio in the LP compressor and a ridiculously low HP compressor pressure ratio (around 2).
As much as I have tried to understand the reasons for it a 2 or 3 stage HP compressor and a HPT pressure drop of just 1.7 seems inplausible to me but that is the optimized result the software presents.

I imagine there are a lot of people on this forum that know a lot about engines so I was wondering if someone could please help me out. I know LightSaber used to design combustors for P&W, can anyone make sense of it?

Thanks in advance,

Miguel


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26 replies: All unread, showing first 25:
 
User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 1, posted (9 years 2 months 4 days 3 hours ago) and read 7616 times:

As a sanity check, throw your results from GasTurb into the NASA EngineSim tool found at:

http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html



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User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 2, posted (9 years 2 months 3 days 7 hours ago) and read 7549 times:

I just checked out that programme, it isn't exactly what I need but the results I get are approximatelly the same (it does not do optimization, though). But thanks anyway.


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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 3, posted (9 years 2 months 3 days 4 hours ago) and read 7516 times:

Can you lock up HPC pressure ratio at a reasonable value and let the program optimize the other four?

What is your specified airflow at sea level / static / standard day?

What component efficiencies are you assuming (and feeding to the program)?



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User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 4, posted (9 years 2 months 2 days 4 hours ago) and read 7439 times:

Hi there,

Yes I can do it, I can even specify a minimum pressure drop in the HPT.
I have an airflow in standard sea-level conditions of 230 kg/s (fan diameter of 1.3m). Fan and LPC efficiency is 0.85 (tip speed 500 m/s) and HPC efficiency is 0.87 (350 m/s), while both turbines efficiencies are 0.89 . I have varied tip speeds but still the same strange thing happens.

Thanks for all the help,

Miguel

P.S. - the booster pressure ratio is always near the maximum I allow. If I allow it it will go even further, where HPC pressure ratio will remain small



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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 5, posted (9 years 2 months 2 days 3 hours ago) and read 7418 times:

Quoting Pyrex (Reply 4):
LPC efficiency is 0.85 (tip speed 500 m/s) and HPC efficiency is 0.87 (350 m/s)

I'm a little suspicious of your LPC & HPC tip speeds. Is this a two-spool or three-spool machine? If two spools, is the LPC on the LP shaft?



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User currently offlineLightsaber From United States of America, joined Jan 2005, 13422 posts, RR: 100
Reply 6, posted (9 years 2 months 1 day 10 hours ago) and read 7353 times:
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Quoting Pyrex (Thread starter):
I know LightSaber used to design combustors for P&W, can anyone make sense of it?

I finally checked tech ops.  spin 

Ok, this looks very suspicious.
Since you properly have HPC efficiency > LPC efficiency The engine should naturally balance to having a greater pressure ration in the HPC. That is unless because of the fixed diameter and other inputs you're forcing the fan to turn at an unrealistic RPM.

First do you have inputs correct?
at 55,000 ft
Outside temperature is -69.7F, 1.323 psia*
Isentropic equations at Mach 1.8
Pt/Ps=5.745
Tt/Ts=1.648

Now, what is the burner exit temperature (Turbine inlet temperature)? This is normally a fixed point in a design (material limit), but its really going to impact your design.

Could you also post your resulting Bypass ratio? One does not expect a high bypass ratio in a supersonic design. A bypass ratio of 1.0 would not surprise me.

One question is on the tip speeds. Why is the HPC tip speed so much lower than the Fan/LPC tip speed? Due to the higher temperatures in the HPC, the internal mach number is lower (for the same RPM) so that the RPM can be pretty high as airfoils have a nearly constant optimum mach number. I would expect these tip speeds to be higher. F14D forever has a good point here.

If you have more questions, instant message me or I'll check this forum. However, I'm on 3rd shift this week (including the weekend), so my hours will be odd, but I will respond.

I apologize for having more questions than answers, but there seems to be an input glitch. Also, what is your fan pressure ratio?

Lightsaber
*From my Pratt&Whitney Aeronautical Vestpocket Handbook, 1986. The so called "Pratt little black book."



Societies that achieve a critical mass of ideas achieve self sustaining growth; others stagnate.
User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 7, posted (9 years 2 months 1 day 9 hours ago) and read 7348 times:

Ok, I'll try to respond to all of your questions

It is a two-spool machine, with the LPC on the LP shaft. With such a high pressure ratio on the LP shaft I naturally have a low booster surge margin.

The fan is turning at around 7380 r.p.m. Is that an unrealistic value? I honestly don't know.

In the software you just have to input Mach number and altitude, it calculates all relevan parameters automatically.

The optimized by-pass ratio is 1.42 .

The burner exit temperature in sustained supersonic cruise is 1600K (I didn't want to go much higher because of durability concerns, it is a civilian and not military engine). Do you think this is too low? (I have 5% cooling air for the NGV and the HPT rotor).
If I increase temperature to 1700K the best HPC pressure ratio increases to 2.2-2.4 and the higher specific thrust leads to a by-pass ratio of 1.8

Actually the HPC tip speed was suggested by the professor, but if I increase it the only thing that happens is that the HP spool speed (currently near 20,000 r.p.m.) increases. Pressure at entry into the HPC, as it stands, is very high, so compressor diameter is low.

Thanks for all your trouble,

Miguel

P.S. - by the way, FPR is 1.9



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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 8, posted (9 years 2 months 1 day ago) and read 7326 times:

Quoting Pyrex (Reply 7):
The fan is turning at around 7380 r.p.m. Is that an unrealistic value?

It's a viable physical speed, but at that altitude we need to know corrected speed: 7380/sqrt(390/518.67 deg R) = 8510. Could be on the ragged edge. What's your fan tip radius?



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User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 9, posted (9 years 2 months 23 hours ago) and read 7323 times:

Hi,

My fan tip radius is 0.65 metres (fan tip diameter of 1.3 metres).
What will be the main limiting factor in this case? Mechanical resistance?

Thanks in advance



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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 10, posted (9 years 1 month 4 weeks 1 day 2 hours ago) and read 7257 times:

Quoting Pyrex (Reply 7):
FPR is 1.9

I'm not a compressor aero guy, but that seems high for a single stage and low for a two-stage design. For high altitude supersonic application, I wonder if you need a two- or three-stage fan.

Regarding the fan tip radius and speed, the usual considerations apply: tip Mach number, blade integrity, and blade retention. Those all relate to physical speed, by the way. (Sorry, I mislead you with the question about corrected speed, which relates more to its ability to pump air relative to its sea level / static airflow.)



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User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 11, posted (9 years 1 month 4 weeks 1 day ago) and read 7236 times:

Hi,

The upper limit on FPR was set by me as I had some data to suggest it could be achieved with a single-stage fan (for simplicity reasons). I don't know at what altitude the graph I showed was.
If I don't limit it the FPR goes up to 2.05 (clearly 2-stage). Maybe I should do that...

Fan tip mach number is 1.35 ... I know a fan can achieve supersonic speeds but I don't know if it can go that high.

Thanks,

Miguel



Read this very carefully, I shall write this only once!
User currently offlineLightsaber From United States of America, joined Jan 2005, 13422 posts, RR: 100
Reply 12, posted (9 years 1 month 4 weeks 15 hours ago) and read 7226 times:
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Quoting Pyrex (Reply 7):
The fan is turning at around 7380 r.p.m. Is that an unrealistic value? I honestly don't know.

Its high, but you're working with a small diameter engine. I'm used to ~1400 rpm on a 100" fan (2.54 meter). You're at a very high RPM, but this is a supercruise engine... see my comment below on Fan tip speed.

Quoting Pyrex (Reply 7):
The optimized by-pass ratio is 1.42 .

A little high, but ok for a FPR of 1.9. Note, a FPR of 1.9 seems low for supercruise.

Quoting Pyrex (Reply 7):
The burner exit temperature in sustained supersonic cruise is 1600K

A bit low. Modern commercial engines run closer to 2800F! (~1800K) Cooling has improved dramatically compared to the assumptions you seem to be making. THIS is probably limiting the HPC PR. For sustained flight consider 1700k.

Quoting Pyrex (Reply 7):
(currently near 20,000 r.p.m.)

Most commercial engines of LARGER diameter are at 15,000 rpm, so your number is reasonable.

Quoting F14D4ever (Reply 10):
I'm not a compressor aero guy, but that seems high for a single stage and low for a two-stage design. For high altitude supersonic application, I wonder if you need a two- or three-stage fan.

The only supercruise engines I'm aware of (F119 for the F-22 fighter, Mig-31 proposed, etc.) all have tripple stage fans (Typical for a military engine, FYI.)
a FPR of 1.9 is aggressive, but possible. I wouldn't want to consider the surge margine though!  scratchchin 

Ok Pyrix, I hope this advice isn't too late (I lost my iternet connection over the weekend, mea culpa.)

Quoting Pyrex (Reply 11):

Fan tip mach number is 1.35 ... I know a fan can achieve supersonic speeds but I don't know if it can go that high.

That's too high. Shock waves would be killing fan efficiency. 1.2 is a realistic maximum for a supercruising engine. Most commercial engines aim for Mach 1.1 to 1.15. But 1.2 with a curved blade fan isn't a stretch at all.

Summing up my recommendations:
1. I strongly recommend going to a two or three stage fan. Since the F22 supercruises at Mach 1.7, I would seriously consider emulating its fan. (That engine is very optimized for supercruise, hint! Sorry, I'm limited to telling you about public numbers...) FPR per stage should probably be backed off a bit too, say to 1.75. A FPR of 1.9 per stage isn't that realistic. Your engine would surge more often than a pw4062!  duck  A FPR of 1.9 would have trouble overcomming the nacelle drag in the real world (for supercruise, obviously lower FPR's work subsonic).

2. Up the Turbine inlet temperature. 1600k is too conservative. If you're doing a supercruise engine, you're not going to skimp. Go with 1800k limit for TO, 1700k for cruise.

3. Don't worry about tip speed so much. I've always designed to mach number. But then again, structures was always another engineer's problem! Tip mach numbers should be Mach 1.1 to 1.2.

Hope this helps. Sorry for the delay.
Lightsaber



Societies that achieve a critical mass of ideas achieve self sustaining growth; others stagnate.
User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 13, posted (9 years 1 month 4 weeks 9 hours ago) and read 7195 times:

Quoting Lightsaber (Reply 12):
The only supercruise engines I'm aware of (F119 for the F-22 fighter, Mig-31 proposed, etc.) all have tripple stage fans (Typical for a military engine, FYI.)

The GE F110-400, which on at least one occasion* has powered the F-14D in supercruise, also has a three-stage fan.

* Sept. 28, 1989; F-14D sans chin pod and glove pylons



"He is risen, as He said."
User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 14, posted (9 years 1 month 4 weeks 7 hours ago) and read 7181 times:

Hey,

Thanks for all the input. I think I have discovered the reason for the low HPC pressure ratio (compressor hub/tip radius was too low - 0.4) so I will be sure to take into consideration all your help. I may stick to a two-stage fan, though (probably cheaper).
About the by-pass ratio: I am not doing detailed acoustic studies but I don't want to go much lower than what I have at the moment because of take-off noise. After all this is a civilian engine.
I am using standard component maps but what is giving a low surge margin is the booster. At the moment LP spool speed cannot go below 75% nominal value...

I'll keep you posted of all developments. Thanks for your help,

Miguel

P.S. - Lightsaber, and now a question for your specialty: according to the papers I have read the objective, emission-wise, would be for less than 5 g of NOx for every kg of fuel in cruise. Do you see that happening in the near future?



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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 15, posted (9 years 1 month 4 weeks 6 hours ago) and read 7170 times:

Quoting Pyrex (Reply 14):
I am using standard component maps but what is giving a low surge margin is the booster. At the moment LP spool speed cannot go below 75% nominal value...

You need a variable bleed valve in the transition duct from booster exit to core inlet to 'burp' the booster (by dumping air to the bypass stream) at low/part power. Or else go to a three-stage fan and delete the booster. Picture a two-dimensional map of all possible fan diameter & FPR combinations. Along some locus of points (combinations) you'll cross over from wanting a booster to not wanting one.

Is this a separate flow or mixed flow engine? I'm thinking for this application you'll go with mixed. Is it fixed or variable exhaust nozzle?

If fixed, does the software model the effect of exhaust nozzle size on noise? Does it give you any points for using a Chevron nozzle (search for pix of the CF34-10 Chevron core nozzle and the new GE90 testbed with Chevron fan nozzle).



"He is risen, as He said."
User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 16, posted (9 years 1 month 4 weeks 4 hours ago) and read 7158 times:

Hi,

It is a mixed-flow turbofan and the idea is to have a variable-geometry exhaust nozzle. With the work I have at the moment, though, I will not go into detailed nozzle design. By the way, this version of the software has absolutely no code for acoustics, so I cannot model exhaust noise.

I tried to eliminate the booster but the problem is, with this software, you cannot make inner fan PR equal to outer fan PR automatically, which means that when I eliminate the booster I cannot optimize FPR automatically (only HPC pressure ratio). I tried that configuration for a number of manually-imposed FPRs and the results I obtained are worse than the ones with the booster, so I will stick with it for the moment.

At the moment this is what I have: a 2-stage fan with a booster (similar to the JT8D) and a HPC. With temperature upped to 1700K and booster pressure ratio limited to about 5 I have:
2.05 fan pressure ratio
4.8 LPC pressure ratio
~5 HPC pressure ratio
I also limited fan tip speed so that the relative Mach number was between 0.9 and 1.2 and oddly enough the optimized result was in the low-end of that scale (around M1.0), so fan speed has reduced to around 4500 r.p.m.



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User currently offlineLightsaber From United States of America, joined Jan 2005, 13422 posts, RR: 100
Reply 17, posted (9 years 1 month 3 weeks 6 days 21 hours ago) and read 7140 times:
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Quoting Pyrex (Reply 14):
: according to the papers I have read the objective, emission-wise, would be for less than 5 g of NOx for every kg of fuel in cruise. Do you see that happening in the near future?

Tough. The problem is that there is a direct trade between fuel economy and NOx. If you desire better fuel economy, the engine must run hotter. With oil at $65/bbl, that is the priority. Oh, I've worked on some WICKED combustor systems that can get emissions down, but once you take into account all of the safety requirements (high altitude relight, burner stability, turbine pattern factor) and other emissions requirements (Carbon monoxide, smoke) its not going to be easy.

GE loves lean burn, which has lower cruise NOx, but EVERY lean burn system has had to be retrofitted to a high NOx configuration due to their inability to maintain light during flight.  Sad Not to mention some of the wicked accoustics! (Airlines dislike pulling apart engines on an annual basis to replace cracked combustor liners.) As TAPS in the GenX will be GE's 3rd attempt at a low NOx lean burn combustor... lets see if they get it right this time. (The CFM-56-5 and GE-90A were the first two attempts that both failed.) Only the Pratt pw4162? (might be 4158) opperated by UPS and a few pw4168's with their Talon II's are in service with "low NOx" combustors. Today, "low NOx" means a mere 20% cut in emissions. At least its not the pw4090 which has 99.2% of the allowed NOx!  duck 

But there thing to note: Right now Cruise NOx preferentially attacks CFC's (vs. attacking ozone). But in a few years (maybe as few as 5) there will be enough traffic and not enough CFC so the NOx will go after the O3...

So obviously more government money needs to be thrown at the problem!  bigthumbsup   spin  It can be done, but a lot more work needs to be done.

Quoting Pyrex (Reply 16):
ariable-geometry exhaust nozzle.

Good catch. Required for noise control and power optimization for sub and super sonic flight.

Quoting Pyrex (Reply 16):
At the moment this is what I have: a 2-stage fan with a booster (similar to the JT8D) and a HPC. With temperature upped to 1700K and booster pressure ratio limited to about 5 I have:
2.05 fan pressure ratio
4.8 LPC pressure ratio
~5 HPC pressure ratio
I also limited fan tip speed so that the relative Mach number was between 0.9 and 1.2 and oddly enough the optimized result was in the low-end of that scale (around M1.0), so fan speed has reduced to around 4500 r.p.m

A much more practical arrangement. Well done.

And I happen to know there is a business jet venture is looking at using JT8D's for Mach 1.2 supercruise. (Yes, if you put the OLD pre-200's on an aerodynamic airframe, the old beasts can supercruise!) Why? Cost. Not to discourage you, but a business jet generally cannot budget millions per engine. Utilizing used engines saves a bundle! (But for supercruise, new nacelle designs are required.) I wonder if it will happen. Its a paper project right now.

Lightsaber

Monitors: please have the spell check fixed!



Societies that achieve a critical mass of ideas achieve self sustaining growth; others stagnate.
User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 18, posted (9 years 1 month 3 weeks 6 days 18 hours ago) and read 7129 times:

Quoting Lightsaber (Reply 17):
there is a business jet venture is looking at using JT8D's for Mach 1.2 supercruise.

Yes, the Aerion. Looks roughly like a Douglas X-3 Stiletto. It would be a thrill to see that fly, but at $80 million a copy, I think they need to sharpen their pencils some more. And unless they do some cutting-edge work on sonic boom attenuation, they'll be hampered by the same operational restrictions in U.S. airspace as the Concorde.



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User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 19, posted (9 years 1 month 3 weeks 6 days 9 hours ago) and read 7110 times:

Yeah, I've heard about the Aerion but, according to a Flight International I read, they do not pretend to have the ability to do supersonic flights overland. I thought they pretended to do M1.6, though. Remember, the JT8D has flown supersonically before, albeit with an afterburner (the Saab Viggen had a Volvo derivative of that engine, if I am not mistaken).

One of the trickiest parts of the design of any engine for this application will be the inlet design but for that you need to know how the entire aircraft will be (because of shock wave interactions with the fuselage and the like) so luckily that reduces my workload a lot.

Actually, after even further changes here is how it looks:
FPR 2.03
BPR 1.59 (quite high)
Booster: 3.88 PR
HPC: 3.80 PR
Fan tip mach number remains at about 1.0

I do not need to go to 1800K on take-off as I already have much more thrust than I need, even in ISA+15K. That has probably to do with the specifications I had to create.

Regarding your question about engine costs: I know practically no engine has been designed from a clean sheet of paper (the costs are just enormous) and initially my idea was to also study adaptations of existing engines (probably in the CFM56/IAE V2500 class but maybe I would have to go higher than that). However, as is quite understandable, engine companies do not provide enough data to do that. I am pretty sure that the first SSBJ to fly (it is only a matter of time, I hope) will have a derivative engine of some sort (probably civilian but possibly also military).



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User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 20, posted (9 years 1 month 3 weeks 6 days 5 hours ago) and read 7100 times:

Quoting Pyrex (Reply 19):
Actually, after even further changes here is how it looks:
FPR 2.03
BPR 1.59 (quite high)
Booster: 3.88 PR
HPC: 3.80 PR

It still looks like that booster just doesn't belong in this picture. Maybe I'm trapped in a paradigm, but that core PR still looks awfully low. What is the core PR for the JT8D?

Does the program have a different 'environment' (at minimum a different suite of compression maps) wherein you can design a military engine, and forego the booster?

Moderator: we're still getting a proxy error with the spell checker.



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User currently offlineLightsaber From United States of America, joined Jan 2005, 13422 posts, RR: 100
Reply 21, posted (9 years 1 month 3 weeks 6 days 2 hours ago) and read 7096 times:
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Quoting F14D4ever (Reply 18):
And unless they do some cutting-edge work on sonic boom attenuation, they'll be hampered by the same operational restrictions in U.S. airspace as the Concorde.

Lockheed skunk works claims thay can allow a mach 1.1 to 1.3 airframe with their boom attenuation. I'm not sold yet.

Quoting F14D4ever (Reply 20):
Maybe I'm trapped in a paradigm, but that core PR still looks awfully low. What is the core PR for the JT8D?

2.03*3.88*3.80=29.9 This is a reasonable cruise pressure ratio. This is about the OPR on a JT8D at SLTO. That said, maximum efficiency is at high pressure ratios. Pyrex is being a little conservative with compressor and turbine efficiencies, but that's ok. (I'm not allowed to state current state of the art. NDA's can be a pesky thing...) The lower the component efficiency, the lower the optimal OPR.

Quoting Pyrex (Reply 19):
Regarding your question about engine costs: I know practically no engine has been designed from a clean sheet of paper (the costs are just enormous) and initially my idea was to also study adaptations of existing engines (probably in the CFM56/IAE V2500 class but maybe I would have to go higher than that). However, as is quite understandable, engine companies do not provide enough data to do that.

Almost true. Now with computers, engines are designed as derivatives of existing designs fairly quickly. However, due to $65/bbl oil, the next engines are radical improvements. And you're correct, engine companies are pretty tight with detail technical data. What was funny is when we had an ex-CFM-56 engineer interviewing with Pratt, he had to stop his interview presentation; the Pratt engineers rattled off the engine stats he couldn't.

Lightsaber



Societies that achieve a critical mass of ideas achieve self sustaining growth; others stagnate.
User currently offlinePyrex From Portugal, joined Aug 2005, 4035 posts, RR: 28
Reply 22, posted (9 years 1 month 3 weeks 6 days 1 hour ago) and read 7092 times:

I don't know the core (meaning HP system) pressure ratio for the JT8D but with a two-stage fan and a booster I assume it isn't much higher than that.

The programme has several possible engines (turbojets, mixed and unmixed turbofans, ramjets, turboprops and turboshafts). You can forego the booster but the problem is, when you start optimization there is no way of guaranteeing inner fan PR equal to outer fan PR (i.e.,, no booster) unless they are not varied as design parameters. I actually tried to do that once (fix FPR to the value I have now and optimize only HPC PR) but the results were worst than what I have now. I guess I'll have to control booster surge margin using FADEC, bleed-air and variable geometry guide vanes.

About the compressor maps: it has one defined for each component but it allows you to make your own. I wouldn't want to go down that road, though...

I'm glad I am being a bit conservative, this way I can have a trump up my sleeve during the presentation... 11th October is coming up awfully fast.

By the way: I didn't really mention it but all compression ratio values are at supersonic cruise (the design condition). In take-off OPR is higher than what I posted.

Quoting Lightsaber (Reply 21):
Lockheed skunk works claims thay can allow a mach 1.1 to 1.3 airframe with their boom attenuation. I'm not sold yet.

Neither is the FAA, last I heard. They would have to build one and prove it to them so they could lift the waiver, and that would cost an awful lot of money.



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User currently offlineLightsaber From United States of America, joined Jan 2005, 13422 posts, RR: 100
Reply 23, posted (9 years 1 month 3 weeks 5 days 23 hours ago) and read 7090 times:
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Quoting Pyrex (Reply 22):
11th October is coming up awfully fast.

Good luck! Hope this helped.

Lightsaber

ps
Still getting the Proxy error on Spell check.



Societies that achieve a critical mass of ideas achieve self sustaining growth; others stagnate.
User currently offlineF14D4ever From United States of America, joined May 2005, 319 posts, RR: 4
Reply 24, posted (9 years 1 month 3 weeks 5 days 19 hours ago) and read 7077 times:

If you have time, tell me a little more about GasTurb. It varies outer (fan tip?) pr, inner (fan hub + booster) pr, core pr, bpr, and t4 to match what? A specified value of cruise thrust? TKOF thrust? Top of climb thrust? Core speed?

Also, did you end up with a booster bleed valve to help iron out speed-speed mismatches (low booster margin)?



"He is risen, as He said."
25 Pyrex : Thanks Lightsaber, it helped a lot. GasTurb is a software designed by Joachim Kurzke, an employee (or former employee, I'm not sure) of MTU Aero Engin
26 F14D4ever : Additionally, the F110-129 supercruised aboard an F-16XL during a NASA test flight. The second aircraft (a two seater) had been retrofitted with a Ge
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