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General Discussion => Q&A Section => Topic started by: JosephB on 03/19/2009 01:53 am

Title: Rocket Engine Q&A
Post by: JosephB on 03/19/2009 01:53 am
I'm curious to know how far the RS-84 program got. Did it end at the power point presentation stage or was there serious design work?
I can't help but envisage the Ares V using expendable kerolox engines with 5 seg boosters. Thanks.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/19/2009 01:41 pm
Never mind, after a little digging I see they did a PDR and tested a subscale preburner. Talk about a project worthy of finishing. Especially with shuttle retirement on the horizon.

Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/21/2009 06:41 pm
Anyone?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/21/2009 06:44 pm
Never mind, after a little digging I see they did a PDR and tested a subscale preburner. Talk about a project worthy of finishing. Especially with shuttle retirement on the horizon.

Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?

Because RS-68 and SSME already exist
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/21/2009 06:58 pm
Wow, that's a fast response. Thanks.
From what I read on the web isn't a modern kerolox supposed to be more reliable & less complex than LH2? For instance wouldn't 3 or 4 RS84's be better than 6 RS-68's? Not to mention less expensive (development costs aside)?

Or is the consensus that is room for improvement down the road?
The reason I ask is the extra RS68 & 1/2 segment on each side on the latest A5 could be remedied by the RS84 it seems.
Title: Re: Rocket Engine Q&A
Post by: Antares on 03/21/2009 08:18 pm
Development costs and expertise.  Too much and too little.
Title: Re: Rocket Engine Q&A
Post by: TrueGrit on 03/21/2009 09:47 pm
Merits of LO2/Kero over LO2/LH2 aside do how long it would take to develop a new engine?  Or cost?  Just as an example...  The RS-68/EELV inital contract was awarded in 1995 and first flight wasn't until 2002.  7 yrs with a price tag well under $1bil...  But that was a commerical structured development.  History has shown a NASA directed activity would take much longer and cost much more.  Heck the J-2X is costing more and taking long than the built from scratch RS-68...  7 yrs and $1.2bil.  My guess extraplotaed off this experience is that RS-84 would take ~10yrs to develop and cost ~$3bil.  Compare that to ~$200mil and ~2yrs to modify RS-68.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/22/2009 01:53 am
Good points. I have to say I keep becoming more and more impressed with the designers & engineers from the 60's & 70's. What they accomplished with the tools, money & materials available in comparison to today...

As far as expertise goes, it seemed like the MSFC/Rocketdyne team had a pretty good handle on the project before it was shut down. Or is there more to it?Although it does sound like development is a BIT on the spendy side, it would be such an incredible 1st stage engine it seems worth it to swallow the up front costs to get on with moon/mars down the road. But as Jim said, if you can put enough 68's on it and get the job done...

With Goverment the way it is cost is relative. For the price of a couple B2 bombers we could have an engine to push us to the planets! I'd check a box on my 1040 to donate $3 to Ares V instead of presidential candidates.
Does anyone know how much shuttle engines cost to develop?
Title: Re: Rocket Engine Q&A
Post by: Patchouli on 03/22/2009 03:35 am
Merits of LO2/Kero over LO2/LH2 aside do how long it would take to develop a new engine?  Or cost?  Just as an example...  The RS-68/EELV inital contract was awarded in 1995 and first flight wasn't until 2002.  7 yrs with a price tag well under $1bil...  But that was a commerical structured development.  History has shown a NASA directed activity would take much longer and cost much more.  Heck the J-2X is costing more and taking long than the built from scratch RS-68...  7 yrs and $1.2bil.  My guess extraplotaed off this experience is that RS-84 would take ~10yrs to develop and cost ~$3bil.  Compare that to ~$200mil and ~2yrs to modify RS-68.

Instead of building the J2X in house they should have just handed a set of performance requirements to various engine manufactures and held a competition.

Title: Re: Rocket Engine Q&A
Post by: William Barton on 03/22/2009 08:11 am
Merits of LO2/Kero over LO2/LH2 aside do how long it would take to develop a new engine?  Or cost?  Just as an example...  The RS-68/EELV inital contract was awarded in 1995 and first flight wasn't until 2002.  7 yrs with a price tag well under $1bil...  But that was a commerical structured development.  History has shown a NASA directed activity would take much longer and cost much more.  Heck the J-2X is costing more and taking long than the built from scratch RS-68...  7 yrs and $1.2bil.  My guess extraplotaed off this experience is that RS-84 would take ~10yrs to develop and cost ~$3bil.  Compare that to ~$200mil and ~2yrs to modify RS-68.

Instead of building the J2X in house they should have just handed a set of performance requirements to various engine manufactures and held a competition.



In order for that to work there would have to be "various" engine manufacturers working in the technology and size class (in the US, for LH2/LOX there's PWR and who else?), and those various engine manufacturers would have to have and be willing to invest their own capital. Otherewise, if you hand out taxpayer money to various engine manufacturers in order to see who comes up with an engine first, it's not a competition, it's corporate welfare. Give me a billion dollars up front and I'll join any competition you want.
Title: Re: Rocket Engine Q&A
Post by: William Barton on 03/22/2009 08:18 am
Merits of LO2/Kero over LO2/LH2 aside do how long it would take to develop a new engine?  Or cost?  Just as an example...  The RS-68/EELV inital contract was awarded in 1995 and first flight wasn't until 2002.  7 yrs with a price tag well under $1bil...  But that was a commerical structured development.  History has shown a NASA directed activity would take much longer and cost much more.  Heck the J-2X is costing more and taking long than the built from scratch RS-68...  7 yrs and $1.2bil.  My guess extraplotaed off this experience is that RS-84 would take ~10yrs to develop and cost ~$3bil.  Compare that to ~$200mil and ~2yrs to modify RS-68.

One of the commonplace ideas that gets put forward from time to time is to allow taxpayers to designate how their taxes are spent, at the federal level. Put a form in with the 1040 that allows you to mark government agencies with a percentage of your tax outlay, is one version of it. You like welfare, pay for welfare. You like the Navy, pay for the Navy. You like art or basic science, your taxes turn to grant money. All the objections to it are trivial, along the lines of, what if nobody wanted to pay for defense (or whatever the objector thinks is important)? But the real reason it gets no traction is because the government already gets to take our money at gunpoint, and we get no real say in how it's spent. Voting only gives you a very tiny input. My favorite line from "Charlie Wilson's War" is too chauvinist to quote in full here, but it begins, "Honey, I'm not elected by voters..."
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/23/2009 01:24 am
It's frustrating to see all these vehicle & technology development programs never leaving the laboratory to get practically applied in space. It's unfortunate the Nasa paradigm is captive to each successive administration like it is & left chasing the next big thing without finishing the last every few years or so.

Welcome to the real world I guess.

We should be thankful to have shuttle & station.
And here I was all pumped to be be "taking the next step"....
Silly man.
Title: Re: Rocket Engine Q&A
Post by: edkyle99 on 03/23/2009 03:24 am
Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?

It wouldn't work as a straight replacement for RS-68 engines on a long core. 

Although four RS-84 engines would produce roughly the same total liftoff thrust as six RS-68 engines, the substantial specific impulse advantage of the RS-68 LH2 engines provide much more payload.  The RS-84 powered core boosted by two five-segment boosters topped by a J-2X powered Earth Departure Stage would only boost about 3/4ths as much payload to TLI as the RS-68 powered core version.

An RS-84 engine would work better on a different launch vehicle design - probably something more like Saturn V with no solid boosters at all, a kerosene first stage, a big hydrogen second stage, and a smaller, optimized TLI third stage.  Or, alternatively, a hydrogen core lifted by RS-84 powered liquid boosters. 

 - Ed Kyle
Title: Re: Rocket Engine Q&A
Post by: Jorge on 03/23/2009 03:40 am
Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?

It wouldn't work as a straight replacement for RS-68 engines on a long core. 

Although four RS-84 engines would produce roughly the same total liftoff thrust as six RS-68 engines, the substantial specific impulse advantage of the RS-68 LH2 engines provide much more payload.  The RS-84 powered core boosted by two five-segment boosters topped by a J-2X powered Earth Departure Stage would only boost about 3/4ths as much payload to TLI as the RS-68 powered core version.

Since the mixture ratios are different, the tanks would have to be resized as well.
Title: Re: Rocket Engine Q&A
Post by: edkyle99 on 03/23/2009 03:46 am
Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?

It wouldn't work as a straight replacement for RS-68 engines on a long core. 

Although four RS-84 engines would produce roughly the same total liftoff thrust as six RS-68 engines, the substantial specific impulse advantage of the RS-68 LH2 engines provide much more payload.  The RS-84 powered core boosted by two five-segment boosters topped by a J-2X powered Earth Departure Stage would only boost about 3/4ths as much payload to TLI as the RS-68 powered core version.

Since the mixture ratios are different, the tanks would have to be resized as well.


Right.  And they would be much smaller in overall length than the current core, which would complicate  the booster attachment design (it currently uses a thrust beam running through the core inter-tank).

 - Ed Kyle
Title: Re: Rocket Engine Q&A
Post by: kraisee on 03/23/2009 04:02 am
Good points. I have to say I keep becoming more and more impressed with the designers & engineers from the 60's & 70's. What they accomplished with the tools, money & materials available in comparison to today...

They did have one big advantage.   They managed to do pretty-much all of their designs before they lost their first crew. The loss of the Apollo-1 crew changed a lot of things. *Everything* had to be scrutinized a lot more after that, from initial designs to prototype hardware, to production processes and everything else all the way to launch -- and even after. Congress forced an enormous amount of additional oversight at every level in order to prevent such another international embarrassment from occurring again.

The loss of Challenger and Columbia both also forced considerable increases in the work needed behind the scenes too.

One example:   Take the re-design work which had to go into the foam Ice Frost Ramps on the External Tank, or the removal of the Pal Ramps.   When you look at them, the changes are pretty simple jobs to accomplish physically -- essentially just shave some 'excess' foam off the tank -- but the process required to quantify the aerodynamic effects of those changes took *years* of effort to allow the Shuttle Program to actually certify the changes.

Back in the early days, such a change as that might very well have been decided in a two hour long meeting and a pad rat would be sent out to the Pad to shave the offending article off the vehicle within a few hours of that!   That sort of 'fix' did happen from time to time back then.

The agency today is a world apart from its younger self in terms of being risk averse.   Some say that's a good thing. Others say its gone too far.   As is so often the case, the truth is probably somewhere in between the two extremes.

But regardless, the effects are very tangible:   Human rated hardware for NASA is a lot more expensive and takes a lot more time to prepare than hardware for almost any other human endeavour.   Because when someone screws up, people die in an extraordinarily visible way and these days all the world watches the USA screw-up live on CNN.

That's the biggest difference between then and now.

Ross.
Title: Re: Rocket Engine Q&A
Post by: gospacex on 03/23/2009 04:21 am
Any thoughts as to why a kerolox RS-84 would/wouldn't work with the 10m Ares V and 5 seg boosters?

It wouldn't work as a straight replacement for RS-68 engines on a long core. 

Although four RS-84 engines would produce roughly the same total liftoff thrust as six RS-68 engines, the substantial specific impulse advantage of the RS-68 LH2 engines provide much more payload.  The RS-84 powered core boosted by two five-segment boosters topped by a J-2X powered Earth Departure Stage would only boost about 3/4ths as much payload to TLI as the RS-68 powered core version.

Well, "the same total liftoff thrust" means that your core stage is much smaller, because RP-1 is so much more dense. If you retain general dimensions of tankage, LOX/RP-1 based booster would weigh much more, need more engines, but also would have much bigger performance.

Anyway, as resident gurus say, "rockets aren't lego blocks", swapping LH engines for RP-1 ones cannot be done like that, serious redesign of the stage is inevitable.
Title: Re: Rocket Engine Q&A
Post by: gospacex on 03/23/2009 04:27 am
The agency today is a world apart from its younger self in terms of being risk averse.   Some say that's a good thing. Others say its gone too far.   As is so often the case, the truth is probably somewhere in between the two extremes.

But regardless, the effects are very tangible:   Human rated hardware for NASA is a lot more expensive and takes a lot more time to prepare than hardware for almost any other human endeavour.   Because when someone screws up, people die in an extraordinarily visible way and these days all the world watches the USA screw-up live on CNN.

That's the biggest difference between then and now.

Ross.

Do you have an explanation why "risk-averse" NASA's designs are complex and work-intensive? This doesn't look logical.
Title: Re: Rocket Engine Q&A
Post by: MKremer on 03/23/2009 07:32 am
Do you have an explanation why "risk-averse" NASA's designs are complex and work-intensive? This doesn't look logical.

Imagine, starting tomorrow, everything you do once you wake up must be documented - not only on paper (with details and samples tested) but with photographs during and after, including your morning ablutions, dressing, travel, workday, lunch, and going home, etc., etc.

Things you use must have extensive documents as well - for example, your toothbrush must have documents detailing how it was made, the materials (sources and processing), how it was inspected and packaged (with its own documented evidence), when and how it was purchased, and when it was opened and additional inspections at that time as well.
(and that's just for one item!)

That's what's meant by "complex and work-intensive". Everything dealing with manned spaceflight hardware, down to individual components, must have a documentation and testing trail that follows it through creation, assembly, and testing. Doesn't matter if it's part of the spacecraft, the booster, or anything else being 'flown' (including crew items), it has to have extensive and expensive testing and documentation, which costs lots and lots of extra $$$.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/23/2009 01:20 pm
Man you guys are good! Thanks for all the detailed info and thanks for taking the time to explain. Reading the more technical members posts is the main reason I joined L2. What appears to make sense to an interested observer is very deceiving. It IS rocket science after all! Also, very interesting commentary on the old/new Nasa.

I have a bit of a moon/Mars bias so reading about heavy lift becoming a reality raises an eyebrow or two. What likelihood would you guys give Ares V actually seeing a launch pad?
Title: Re: Rocket Engine Q&A
Post by: edkyle99 on 03/23/2009 03:42 pm

 If you retain general dimensions of tankage, LOX/RP-1 based booster would weigh much more, need more engines, but also would have much bigger performance.

That would be a beyond-monster rocket!  It would need at least 10 RS-84 engines on the core (if they could be made to fit).  The core and twin boosters would produce roughly 18 million pounds of thrust at liftoff.  The rocket would weigh something like 14.5 million pounds at liftoff.  That's roughly equivalent to 2.4 Saturn V rockets, or about 1.8 Ares Vs, or 14.5 fully loaded A380s going straight up.  :)

House foundations would crack in Titusville!  Car alarms would go off in Orlando!  All that for only a factor of 1.2 improvement in payload. 

"Woof" nonetheless!

 - Ed Kyle
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/24/2009 01:02 am
Not to beat a dead horse on Ares I (but I will anyway), after reading this and other threads wouldn't a better 1st stage be a single core with a cluster of 4 RS-68A's with a J2X upper (manned rating aside) maybe GEM capable? Then you could have a bigger service module, more capable Orion, etc.

I can see why they went with the solid for safety reasons but still. And a five seg solid still deserves devlopment just for the A5 in it's own right (imho).
Good, bad or ugly?

Title: Re: Rocket Engine Q&A
Post by: Jim on 03/24/2009 01:08 am
(manned rating aside) maybe GEM capable?

GEMs are a no no wrt safety/reliability
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/24/2009 01:17 am
I didn't know GEM's were that sketchy. Although they would be on my mind if I were riding on top! What about a 4 engine core? Sufficient?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/24/2009 01:20 am
I didn't know GEM's were that sketchy. Although they would be on my mind if I were riding 'on top! What about a 4 engine core? Sufficient?

It isn't the GEM's themselves, it is the additional parts/systems which reduces "reliability".  That is the line that gave us the SRB first stage
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/24/2009 01:45 am
I see your point, but then the russian launchers come to mind with all the engines & plumbing...

Solids may be a different animal entirely in that respect but the SRB sure limits what can be done. One good thing is that it's in the here & now.
Title: Re: Rocket Engine Q&A
Post by: edkyle99 on 03/24/2009 02:09 am
Not to beat a dead horse on Ares I (but I will anyway), after reading this and other threads wouldn't a better 1st stage be a single core with a cluster of 4 RS-68A's with a J2X upper (manned rating aside) maybe GEM capable? Then you could have a bigger service module, more capable Orion, etc.

I can see why they went with the solid for safety reasons but still. And a five seg solid still deserves devlopment just for the A5 in it's own right (imho).
Good, bad or ugly?

MSFC proposed a similar idea back in 2004-2005, an ET-based core launcher that would serve as a building block toward a larger vehicle.  The problem was that the core didn't serve either purpose (crew launch or cargo launch) very well.  It needs to be smaller for crew and larger for cargo.

 - Ed Kyle
Title: Re: Rocket Engine Q&A
Post by: Proponent on 03/24/2009 02:11 am
GEMs are a no no wrt safety/reliability

A couple of questions about solid reliability, if I may:

* Do solids ever fail to ignite these days, either on the pad or at staging?
* Are there any solid or solid-boosted LVs that are held down until proper ignition is verified?
* What proportion of solid failures are catastrophic (i.e., of the "boom" variety rather than of, say, the loss-of-thrust variety)?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/24/2009 02:14 am

* Are there any solid or solid-boosted LVs that are held down until proper ignition is verified?

that would be none I believe
Title: Re: Rocket Engine Q&A
Post by: GI-Thruster on 03/24/2009 03:49 pm
Anyone heard anything about aerospike in the last 6 months?

http://www.csulb.edu/colleges/coe/mae/views/projects/rocket/news_2008/aerospike06252008.shtml

Anyone heard anything about pulsed detonation research or continuous detonation research in the last year or is continuous still a pipe dream?

Anyone know how Qinetiq's ion drive on GOCE compares to NASA's ion thrusters?  With 40 years of ion research it would be embarrassing if ESA is ahead of NASA and if they're not one wonders why they didn't ask to fly one of ours. . .

http://www.esa.int/SPECIALS/GOCE/index.html
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/25/2009 02:59 am
Ed, very interesting. I didn't know that was ever proposed. I have no idea about performance numbers but something 1/2 to 3/4 the size of the ET would would seem perfect for what Orion is trying to accomplish (with plenty of margin).

Would anyone have a good link to find out the latest progress on the J2X?
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/25/2009 11:50 am
I have no idea about performance numbers but something 1/2 to 3/4 the size of the ET would would seem perfect

LOL That didn't sound to good did it?
Title: Re: Rocket Engine Q&A
Post by: Antares on 03/26/2009 02:17 am
Would anyone have a good link to find out the latest progress on the J2X?

Good news is usually provided here (http://www.nasa.gov/mission_pages/constellation/ares/ares_weekly.html).  Bad news is usually provided (at least the first anyone outside of NASA hears about it) here (http://forum.nasaspaceflight.com/index.php?board=29.0).
Title: Re: Rocket Engine Q&A
Post by: duane on 03/27/2009 06:27 pm
Your "bad news" link is broken, or is that a forum on the pay side of NSF ?
Title: Re: Rocket Engine Q&A
Post by: kch on 03/27/2009 06:55 pm
Your "bad news" link is broken, or is that a forum on the pay side of NSF ?

That's the "Constellation" section of L2.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 03/31/2009 03:03 am
Not to beat a dead horse on Ares I (but I will anyway), after reading this and other threads wouldn't a better 1st stage be a single core with a cluster of 4 RS-68A's with a J2X upper (manned rating aside) maybe GEM capable? Then you could have a bigger service module, more capable Orion, etc.

I can see why they went with the solid for safety reasons but still. And a five seg solid still deserves devlopment just for the A5 in it's own right (imho).
Good, bad or ugly?



Looks like other folks were thinking along the same lines well before me (imagine that). See attached...
An 8 meter CBC could lift the "Apollo on steroids" I bet. Even the larger more capable versions of Orion.
Title: Re: Rocket Engine Q&A
Post by: kkattula on 03/31/2009 06:17 am
Anyone know what the performance of a single core, 7 or 8 meter CBC with 2, 3 or 4 RS-68A's and an improved US would be?

Although at 8m and 4 engines, it's getting close to an ET with engines underneath and no SRB's.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 04/14/2009 07:07 pm
Question for the more technical members...
Since A5 is set for 5 1/2 segment SRB's, wouldn't A1 gain from using it as well?
I tried searching but to no avail.
Title: Re: Rocket Engine Q&A
Post by: jabe on 04/14/2009 09:55 pm
A real newbie question on this..
why is a LO2/LH2  engine more expensive to design then a LO2/Kero? Is it the temperature involved the main problem?  I think I'm missing the obvious on this.
jb
Title: Re: Rocket Engine Q&A
Post by: JosephB on 04/15/2009 01:34 am
Question for the more technical members...
Since A5 is set for 5 1/2 segment SRB's, wouldn't A1 gain from using it as well?
I tried searching but to no avail.

OK, how about any member?
Does that extra weight push it in the realm of diminishing returns?
Title: Re: Rocket Engine Q&A
Post by: JosephB on 04/15/2009 01:41 am
Found the thread but if anyone wants to comment please do. Thanks.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 04/15/2009 03:46 am
Looks like there is a good chance that both Ares I & V will be using expendable 5 segment HTPB composite boosters down the road eventually. Very nice payload improvement (See page 9).

Reasonable assessment?

Title: Re: Rocket Engine Q&A
Post by: Antares on 04/15/2009 03:47 am
5.5 seg SRB makes TO even worse.  According to one of Ross's posts, somewhere above 5.5 but less than 6 actually has worse performance for Ares 1 or 5 or both.


I'm not sure if it can blanket be said that a H2 engine is more expensive to design than kero.  However, H2 is certainly harder in operations.  You have to use helium purges and pressurant instead of N2 since N2 would freeze at LH2 temperatures.  H2 also freezes O2, so common bulkheads are much more difficult to design and build unless you're experienced at it (ahem).

As for the engine, since LH2 has such low density, it requires more pump stages in a turbopump.  That's hard by itself and really hard for a single shaft.

H2 burns cleaner, so combustion chamber design is easier.  Mostly I think it's just the temperature of LH2 that makes it harder, if that can be said universally.
Title: Re: Rocket Engine Q&A
Post by: Antares on 04/15/2009 03:52 am
Reasonable assessment?

No
1) Ares V is so out of bed on cost, nothing should be considered even remotely certain.

2) Composite and HTPB lose all semblance of human rated heritage.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 04/15/2009 12:56 pm
I originally was trying to find out how much extra performance an Ares I with a 5.5 segment booster would gain but found out that the higher alt. & heating on the casing makes reusability a big problem.

Which leads to a second question. Development cost & man rating aside, which would be cheaper to fly, a composite 5 seg HTPB booster or a liquid stage?
Title: Re: Rocket Engine Q&A
Post by: Antares on 04/15/2009 06:56 pm
Any idea why ATK was not selected in the EELV competition back in 1996?  That might be the easiest place to find an answer.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 04/27/2009 01:39 am
Okay heres one that has been rattling arround in my brain a little and could make an interesting discussion. We know solids give a very rough ride. I was wondering what liquid fuel gives the smoothest ride. (Asuming a properly designed POGO system).

I think I know the answer. LH/LOX because of how of it burns and the low molecular weight. But is that true?

Considering Kero/LOX seems to have the most violent combustion, I do not think it is it.
The first rockets used Alcohol/LOX because it was easy to work with. I wonder, lower ISP, but would the ride have been smoother?

Since Hypergols burn on contact do you get more complete combustion and a smoother burn?

I am sure there are other things we can throw into the mix, pressure fed verses turbo fed. How well the combustion chamber is designed.
Title: Re: Rocket Engine Q&A
Post by: Jim on 04/27/2009 01:46 am
I don't believe there is a difference in the ride due to propellants
Title: Re: Rocket Engine Q&A
Post by: Patchouli on 04/27/2009 02:46 am
5.5 seg SRB makes TO even worse.  According to one of Ross's posts, somewhere above 5.5 but less than 6 actually has worse performance for Ares 1 or 5 or both.


I'm not sure if it can blanket be said that a H2 engine is more expensive to design than kero.  However, H2 is certainly harder in operations.  You have to use helium purges and pressurant instead of N2 since N2 would freeze at LH2 temperatures.  H2 also freezes O2, so common bulkheads are much more difficult to design and build unless you're experienced at it (ahem).

As for the engine, since LH2 has such low density, it requires more pump stages in a turbopump.  That's hard by itself and really hard for a single shaft.

H2 burns cleaner, so combustion chamber design is easier.  Mostly I think it's just the temperature of LH2 that makes it harder, if that can be said universally.

I once read that the RL-10 would be a very cheap engine if it were mass produced due to it's simplicity.

You're probably right it's likely the handling of the fuel that makes it so costly.

Then there are nasty things like hydrogen embrittlement not sure how bad an issue this is for pad operations but maybe a shuttle or delta person fill that in.
Title: Re: Rocket Engine Q&A
Post by: butters on 04/27/2009 03:40 am
As a rocket tech newbie, I'm trying to get my head around the design trade-offs between LOX/LH2 and LOX/RP-1 (or some shorter-chain hydrocarbon) for the first stage of a heavy launch vehicle.

Obviously LH2 has a specific impulse advantage, greater cooling capacity, and cleaner combustion.  But everything else seems to favor RP-1: cost, density, pumping, plumbing, lubrication, insulation, tank pressure, boil-off, mass ratio, production, storage, logistics, etc.

What are the best arguments for why LH2 is less desirable for first stages than for upper stages?  Is there still a reason to use RP-1 first stages if LH2 is more appropriate for the upper stages?

How much room is there for specific impulse improvement, for example, with a LOX/LCH4 full flow staged combustion cycle engine with a chamber pressure of about 4000 psia?
Title: Re: Rocket Engine Q&A
Post by: meiza on 04/27/2009 10:15 am
Having worse ISP in the first stage makes the first stage bigger in wet weight. (It still might be lighter in dry weight, like Atlas is lighter than Delta.)
Having worse ISP in the second stage makes the second stage bigger in wet weight AND thus the first stage bigger in TOTAL (since it carries the wet second stage).

Now you can see that it doesn't matter that much if a first stage is a bit heavy when wet if it can compensate that with good thrust and low dry weight (which kerolox does). But if a second stage is heavy when wet, it must not only thrust more itself, but the first stage must deliver more impulse (more thrust, more fuel, ie a bigger stage). Having a good thrust or a good mass fraction don't help in the second stage as much as in the first stage, since the penalty of high wet weight is so drastic.

Hence it's smarter to spend more effort/money on wet mass optimization in the top part of the rocket than at the bottom. And hydrogen gives lower wet mass while kerosene gives lower dry mass.
Title: Re: Rocket Engine Q&A
Post by: gospacex on 04/27/2009 02:41 pm
Having worse ISP in the first stage makes the first stage bigger in wet weight. (It still might be lighter in dry weight, like Atlas is lighter than Delta.)
Having worse ISP in the second stage makes the second stage bigger in wet weight AND thus the first stage bigger in TOTAL (since it carries the wet second stage).

Now you can see that it doesn't matter that much if a first stage is a bit heavy when wet if it can compensate that with good thrust and low dry weight (which kerolox does). But if a second stage is heavy when wet, it must not only thrust more itself, but the first stage must deliver more impulse (more thrust, more fuel, ie a bigger stage). Having a good thrust or a good mass fraction don't help in the second stage as much as in the first stage, since the penalty of high wet weight is so drastic.

Hence it's smarter to spend more effort/money on wet mass optimization in the top part of the rocket than at the bottom. And hydrogen gives lower wet mass while kerosene gives lower dry mass.

Another consideration is that not only dry mass of kerolox stage is smaller, but the stage is smaller itself. When you build a big rocket and/or you need to adapt an existing infrastructure, this becomes a factor too.

An example: VAB has limited dimensions. Biggest possible LH/LOX rocket that fits in the VAB is going to have lower performance that biggest possible kerolox one. (Assuming only 1st stage is kerolox)
Title: Re: Rocket Engine Q&A
Post by: edkyle99 on 04/27/2009 03:21 pm
As a real world example, the AtlasV provides better performance than DeltaIV. 

This isn't all due to kerosene, etc.  If Delta IV Medium had a Centaur second stage, it would nearly match Atlas V 401 performance to GTO.  The only limiting factor would be the liftoff thrust to weight ratio.  RS-68A will solve that problem. 

Delta IV has a heavy upper stage compared to Centaur, but it avoids the balloon tanks.

The real difference between these two launch vehicles is cost.  Delta IV is a bigger rocket.  It has to be to carry all of that low-density first stage propellant.  That means bigger ground processing and launch facilities, which means more dollars.  (Although Atlas has "smarter" launch facilities too.  If Delta used a mobile platform rather than a massive mobile service tower, it's facilities would probably have cost less than they did.)

 - Ed Kyle
Title: Re: Rocket Engine Q&A
Post by: Antares on 04/27/2009 04:35 pm
I was wondering what liquid fuel gives the smoothest ride.

The difference is really only in near-pad acoustics and engine section acoustics.  It's very small compared to the difference between solid and liquid.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 05/01/2009 09:23 pm
Came across an interesting link in another thread...
http://www.californiaspaceauthority.org/html/press-releasesandletters/pr060718-1.html

I'm sure IPD is old news to veteran posters but for those on the learning curve it seems a fair amount of effort has been put into reusable engines just a few years back. RS-84 kerolox and then the IPD demonstrator (250,000 lb class).

Were these to be used in flyback first and/or second stages?
What applications did they have in mind?

EDIT: Oooops! Never mind. Read up on SLI history.
Yet another "late to the party" post.
In this case, by how many years?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 05/22/2009 05:22 pm
Is NASA Glenn still involved in spacecraft propulsion research or has that all been moved to MSFC?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 08/17/2009 02:03 am
How are the code names for US rocket engines (RS-68, RS-84 etc) assigned? Is there some government agency that maintains these names or are they just chosen by their manufacturers??
Title: Re: Rocket Engine Q&A
Post by: gin455res on 08/17/2009 07:34 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines. Engines with different propellants, e.g. 6 merlins and 1 SSME, so that you could gradually throttle down the merlins to approximate the function of a true tri-propellant engine.

Has this idea ever been tried, and if not why not?

thanks   
Title: Re: Rocket Engine Q&A
Post by: ugordan on 08/17/2009 07:41 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines.

Energia used 4 kerosene boosters (Zenit) and a hydrogen core so that's that.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 08/17/2009 08:07 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines.

Energia used 4 kerosene boosters (Zenit) and a hydrogen core so that's that.

I meant on a single stage together as a unit.
Title: Re: Rocket Engine Q&A
Post by: Jim on 08/17/2009 08:37 pm
How are the code names for US rocket engines (RS-68, RS-84 etc) assigned? Is there some government agency that maintains these names or are they just chosen by their manufacturers??

Both
In the early days, the USAF gave designations, now it is manufacturers
Title: Re: Rocket Engine Q&A
Post by: Jim on 08/17/2009 08:38 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines. Engines with different propellants, e.g. 6 merlins and 1 SSME, so that you could gradually throttle down the merlins to approximate the function of a true tri-propellant engine.


Tank design and the requirements of the different engines
Title: Re: Rocket Engine Q&A
Post by: ugordan on 08/17/2009 08:44 pm
On conceivable way to use two different propellants on a single stage, but with the same engine type would be say a hydrogen engine with a high expansion ratio, thrust augmented nozzle burning kerosene for extra thrust low in the atmosphere. But that was so far only demonstrated on a small scale.
Title: Re: Rocket Engine Q&A
Post by: rklaehn on 08/17/2009 08:47 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines. Engines with different propellants, e.g. 6 merlins and 1 SSME, so that you could gradually throttle down the merlins to approximate the function of a true tri-propellant engine.

Has this idea ever been tried, and if not why not?

I have seen several proposals for something like this. It makes a lot of sense for reusable single stage vehicles, but it is not as useful for expendable and/or multi-stage vehicles.

I think the shuttle can fire the OMS engines during ascent (OMS assist burn). That would be the only example I can think of where this is used in an existing vehicle.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 08/17/2009 08:54 pm
With some first stages having as many as 9 separate engines, is there any reason why you couldn't  have a mixture of two types of engines. Engines with different propellants, e.g. 6 merlins and 1 SSME, so that you could gradually throttle down the merlins to approximate the function of a true tri-propellant engine.


Tank design and the requirements of the different engines

Thanks for  answering.

Is that for the specific engine examples i gave, or for all possible versions of the idea?

I imagine from your answer that I need two separate oxygen tanks because the different engines need different input pressure requirements. Is this what you are alluding too, or is it something else?
Title: Re: Rocket Engine Q&A
Post by: gin455res on 08/17/2009 09:07 pm
On conceivable way to use two different propellants on a single stage, but with the same engine type would be say a hydrogen engine with a high expansion ratio, thrust augmented nozzle burning kerosene for extra thrust low in the atmosphere. But that was so far only demonstrated on a small scale.

thanks.
Title: Re: Rocket Engine Q&A
Post by: Jim on 08/17/2009 09:08 pm

I imagine from your answer that I need two separate oxygen tanks because the different engines need different input pressure requirements. Is this what you are alluding too,

Yes
Title: Re: Rocket Engine Q&A
Post by: gin455res on 08/17/2009 09:24 pm

I imagine from your answer that I need two separate oxygen tanks because the different engines need different input pressure requirements. Is this what you are alluding too,

Yes
And that this is; a) completely unavoidable, b) not possible with any existing extant set of engines, or c) just impossible with the specific example engines I used to help illustrate the question?
Title: Re: Rocket Engine Q&A
Post by: hop on 08/17/2009 10:22 pm
And that this is; a) completely unavoidable, b) not possible with any existing extant set of engines, or c) just impossible with the specific example engines I used to help illustrate the question?
The RD-701 (http://www.buran.ru/htm/rd-701.htm) might be relevant to this discussion.

I have difficulty thinking of a situation where using multiple engine types would be a win. If you are staging anyway, you are much better off dumping your high thrust engines and associated tankage as soon as you are done with them. If you are building an SSTO, the penalty for lugging the high thrust engines all the way to orbit will be prohibitive.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/04/2009 03:21 pm
Merits of LO2/Kero over LO2/LH2 aside do how long it would take to develop a new engine?  Or cost?  Just as an example...  The RS-68/EELV inital contract was awarded in 1995 and first flight wasn't until 2002.  7 yrs with a price tag well under $1bil...  But that was a commerical structured development.  History has shown a NASA directed activity would take much longer and cost much more.  Heck the J-2X is costing more and taking long than the built from scratch RS-68...  7 yrs and $1.2bil.  My guess extraplotaed off this experience is that RS-84 would take ~10yrs to develop and cost ~$3bil.  Compare that to ~$200mil and ~2yrs to modify RS-68.

Instead of building the J2X in house they should have just handed a set of performance requirements to various engine manufactures and held a competition.



The J-2X engine was contracted to Pratt & Whitney Rocketdyne who has the sole responsibility to design & develop the engine per NASA spec.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/04/2009 03:33 pm
How are the code names for US rocket engines (RS-68, RS-84 etc) assigned? Is there some government agency that maintains these names or are they just chosen by their manufacturers??

Both
In the early days, the USAF gave designations, now it is manufacturers

RS = Rocket System

It's a Pratt & Whitney Rocketdyne internal designation.

Title: Re: Rocket Engine Q&A
Post by: JosephB on 09/06/2009 01:50 am
I came across the RL-10 based CECE and wondered if it could eventually be tested on the X-37B or is that way to big of an engine for possible testing?


CECE - check out the video!
http://www.nasa.gov/mission_pages/constellation/news/cece.html

X-37B's AR2-3
http://www.astronautix.com/engines/ar23.htm
Title: Re: Rocket Engine Q&A
Post by: Jim on 09/06/2009 02:49 pm
I came across the RL-10 based CECE and wondered if it could eventually be tested on the X-37B or is that way to big of an engine for possible testing?


CECE - check out the video!
http://www.nasa.gov/mission_pages/constellation/news/cece.html

X-37B's AR2-3
http://www.astronautix.com/engines/ar23.htm

Too big.  X-37 is just a spacecraft
Title: Re: Rocket Engine Q&A
Post by: strangequark on 09/06/2009 05:17 pm
I came across the RL-10 based CECE and wondered if it could eventually be tested on the X-37B or is that way to big of an engine for possible testing?


CECE - check out the video!
http://www.nasa.gov/mission_pages/constellation/news/cece.html

X-37B's AR2-3
http://www.astronautix.com/engines/ar23.htm

Speaking of the CECE, does anyone know the specifics as to the methane option on it? Is it the exact same engine, turbopumps and all, do they use a different turbopump for the methane (and gear ratio...), or are the modifications even greater?
Title: Re: Rocket Engine Q&A
Post by: JosephB on 09/07/2009 09:47 pm
I tried searching for the answer to this question but really didn't find what I was looking for (although I'm sure it's been covered somewhere, probably numerous times.)

If Ares I had an airstart SSME it wouldn't be taking the heat is is. Fair assessment?

If so, how much of a showstopper was this really? Is it technically extremely difficult? Too expensive? Or both? If Nasa wanted Ares I bad enough I would think they would push forward to a solution.

What specifically is the sticking point in the engine that stopped progress?
I know it's a fine tuned hotrod but I would think they could come up with something between J2X and airstart SSME?

Also, could anyone suggest an excellent modern book on engines for the layperson?
Title: Re: Rocket Engine Q&A
Post by: Jorge on 09/07/2009 10:08 pm
I tried searching for the answer to this question but really didn't find what I was looking for (although I'm sure it's been covered somewhere, probably numerous times.)

If Ares I had an airstart SSME it wouldn't be taking the heat is is. Fair assessment?

Beware the "grass is always greener" fallacy. Ares I with SSME would likely still be getting heat, just for different reasons.

Quote
If so, how much of a showstopper was this really? Is it technically extremely difficult? Too expensive? Or both?

It was never an absolute showstopper. It was discovered to be more difficult than first thought and that would mean a big hit to development cost and schedule. The 5-seg+J-2X option scored just below the 4-seg+SSME option in the ESAS evaluations and the additional problems with airstart SSME tipped the balance.

Quote
If Nasa wanted Ares I bad enough I would think they would push forward to a solution.

NASA wants Ares V just as much. Another argument advanced at the time in favor of 5-seg+J-2X was greater commonality with Ares V and therefore lower operational costs. That was, of course, later negated due to Ares V going 5.5 seg and mods required for *restartable* J-2X, but it was the argument at the time.

Quote
What specifically is the sticking point in the engine that stopped progress?
I know it's a fine tuned hotrod but I would think they could come up with something between J2X and airstart SSME?

"Something in between" means a whole new engine and a lengthier and more expensive development process. Rocket engines aren't Legos; you can't just mix and match.

Quote
Also, could anyone suggest an excellent modern book on engines for the layperson?

I'm not aware of any for a lay audience. Sutton's Rocket Propulsion Elements is the best for a technical audience.
Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 09/08/2009 01:02 am
I tried searching for the answer to this question but really didn't find what I was looking for (although I'm sure it's been covered somewhere, probably numerous times.)

If Ares I had an airstart SSME it wouldn't be taking the heat is is. Fair assessment?

If so, how much of a showstopper was this really? Is it technically extremely difficult? Too expensive? Or both? If Nasa wanted Ares I bad enough I would think they would push forward to a solution.

What specifically is the sticking point in the engine that stopped progress?
I know it's a fine tuned hotrod but I would think they could come up with something between J2X and airstart SSME?

Also, could anyone suggest an excellent modern book on engines for the layperson?

It would still have big problems with thrust oscillation and SRB debris.  Acoustics could probably be fixed by lowering max q due to the increased performance. 

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence.  I am also certain this was well known at the time of ESAS but was used to shove a flawed design into the system.

Can someone find and post copies of the documents mentioned in Appendix 6 of ESAS stating the SSME could be redisgned for airstart?
Thanks in advance to anyone that can.

Danny Deger
Title: Re: Rocket Engine Q&A
Post by: JosephB on 09/08/2009 02:09 am
Thanks for the excellent input guys. I'll put Sutton's on my X-mas list and hope I won't need a Doctorate in Physics to glean a few nuggets.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 09/08/2009 02:18 pm
I'm not aware of any for a lay audience. Sutton's Rocket Propulsion Elements is the best for a technical audience.

People have said the 7th edition is full of errors. Is there a reliable list of errata out there somewhere or should I go for the 6th edition?
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/08/2009 03:30 pm
I'd go for the 6th.

JosephB, you might also try Huzel and Huang.  It's somewhat more accessible than Sutton.  The 1971 version is available online
http://www.spl.ch/publication/sp125.html

The current version can be purchased from AIAA.
Title: Re: Rocket Engine Q&A
Post by: kneecaps on 09/08/2009 04:06 pm
Thanks for the excellent input guys. I'll put Sutton's on my X-mas list and hope I won't need a Doctorate in Physics to glean a few nuggets.

Plenty of good information in Sutton even if you don't understand all the math, you'll certainly gain insight into things.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/08/2009 05:45 pm
I came across the RL-10 based CECE and wondered if it could eventually be tested on the X-37B or is that way to big of an engine for possible testing?


CECE - check out the video!
http://www.nasa.gov/mission_pages/constellation/news/cece.html

X-37B's AR2-3
http://www.astronautix.com/engines/ar23.htm

Too big.  X-37 is just a spacecraft

Also, X-37 no longer uses the AR2-3.

Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/08/2009 06:08 pm
If Ares I had an airstart SSME it wouldn't be taking the heat is is. Fair assessment?

If so, how much of a showstopper was this really? Is it technically extremely difficult? Too expensive? Or both? If Nasa wanted Ares I bad enough I would think they would push forward to a solution.

Ares I would taken less heat if it had an airstart SSME.

For one, it would not needed to spend $1B for a brand new J-2X, plus it wouldn't had to spend an additional $1B to put an extra seg.ment (20%) on the existing Shuttle SRB.  That's a wopping $2B impact that Mike Griffin had not counted on. 

Also, a 4-segment SRB would have less of Thrust Oscillation issue (different structural natural frequency plus a lower magnitude of oscillation). 

One had to ask NASA if it was worth to spend the extra money, plus the engineering effort (time & money) on TO, just to avoid an SSME 'upgrade' to an expendable air-startable RS-25

But the contractors are not asking any questions as they are happy.  Rocketdyne got a new $1B engine development contract.  ATK got a $1B SRB development contract, plus $X hundreds of Millions on TO problem solving and isolator development.  Both Boeing & LM are happy because they get 'extension' on their deliverable schedule on Ares I Upper Stage and the Orion.

How much does it take to upgrade SSME to an air-startable RS-25?  Can NASA use that extra $2.X Billion budget if it didn't need to develop a J-2X and a FSB?  It is certainly interesting to ask if, had someone done some upfront trade study & cost analysis, would MSFC still gets an black eye on the Ares I design? 

Quote
What specifically is the sticking point in the engine that stopped progress?
I know it's a fine tuned hotrod but I would think they could come up with something between J2X and airstart SSME?

Also, could anyone suggest an excellent modern book on engines for the layperson?

That would be the "Modern Design of Liquid Rocket Engines" by Huang & Huzel from AIAA.  Both authors worked at Rocketdyne for sometime and it provides great insights on the design of pump-fed liquid rocket engines.



Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/08/2009 06:12 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 09/08/2009 06:13 pm
On Airstarting an SSME, ESAS mentions a NAS8- paper that I have found.  It also mentions a 2004 Marshal paper on looking at airstarting an SSME.  Can someone find and post?

Danny Deger
Title: Re: Rocket Engine Q&A
Post by: Jorge on 09/08/2009 06:25 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.

Title: Re: Rocket Engine Q&A
Post by: Jim on 09/08/2009 06:26 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.


and the SSME can't be restarted and therefore is not a EDS candidate
Title: Re: Rocket Engine Q&A
Post by: Jorge on 09/08/2009 06:27 pm
If Ares I had an airstart SSME it wouldn't be taking the heat is is. Fair assessment?

If so, how much of a showstopper was this really? Is it technically extremely difficult? Too expensive? Or both? If Nasa wanted Ares I bad enough I would think they would push forward to a solution.

Ares I would taken less heat if it had an airstart SSME.

For one, it would not needed to spend $1B for a brand new J-2X, plus it wouldn't had to spend an additional $1B to put an extra seg.ment (20%) on the existing Shuttle SRB.  That's a wopping $2B impact that Mike Griffin had not counted on. 

But both those items were needed for Ares V anyway. So it would just delay the impact, not eliminate it. Whereas switching to J-2X *eliminated* the impact of airstart SSME.
Title: Re: Rocket Engine Q&A
Post by: Jorge on 09/08/2009 06:28 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.


and the SSME can't be restarted and therefore is not a EDS candidate

Right, so you need the J-2X (or some other restartable engine) anyway.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/08/2009 09:20 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.


and the SSME can't be restarted and therefore is not a EDS candidate

Right, so you need the J-2X (or some other restartable engine) anyway.

Surely for $2 Billon dollars, we can design an air-startable, re-startable, staged combustion cycle engine?

Call it a derived SSME, RS-83, IPD, Russian RD-0120... whatever heritage you'd like
Title: Re: Rocket Engine Q&A
Post by: cleo on 09/08/2009 09:34 pm
For some a (slight) notion of the complexity take a look at the ground support for a launch start of SSMEs  Now pack it all up and do same things up in the air /vacuum (spin up of the turbine rotor at such rpms where a tiny spec will rip a blade apart)
Title: Re: Rocket Engine Q&A
Post by: Jorge on 09/08/2009 09:57 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.


and the SSME can't be restarted and therefore is not a EDS candidate

Right, so you need the J-2X (or some other restartable engine) anyway.

Surely for $2 Billon dollars, we can design an air-startable, re-startable, staged combustion cycle engine?

It will cost more than modifying an existing engine like the J-2. J-2X budget was $1.2 billion back in 2006. Don't know if it's gone up since then. And don't call me Shirley.
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/08/2009 10:25 pm
Gents, isn't this getting more into a discussion that should be in the CxP threads?

Another resource for JosephB: http://www.pwrengineering.com/data.htm
There are some papers in there that aren't so technical.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 09/09/2009 04:35 am
Great points from all. Thank you. Adding all the above mentioned books to the wish list. I have to say that having the air-startable, re-startable, staged combustion cycle engine (Call it a derived SSME, RS-83, IPD) that Propforce mentioned would be wise to have. In hindsight I'll bet Nasa now wishes they would have gone down that path?

On a side note, what does the X-37 use now if not the AR2-3?
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/09/2009 04:03 pm

I think it would be VERY hard to change the SSME to airstart.  I am certain the design relies heavily on head pressure at the pumps due to gravity during the start sequence......

Is it $1 BILLION dollar hard?   ::)

No. That's probably a lowball. Airstart SSME is harder than J-2X.


and the SSME can't be restarted and therefore is not a EDS candidate

Right, so you need the J-2X (or some other restartable engine) anyway.

Surely for $2 Billon dollars, we can design an air-startable, re-startable, staged combustion cycle engine?

It will cost more than modifying an existing engine like the J-2. J-2X budget was $1.2 billion back in 2006. Don't know if it's gone up since then. And don't call me Shirley.

If you say so, Georgi  ;D

BTW, the J-2X looks NOTHING like the J-2.  Just about the only thing 'derived' is the name.

I recalled some discussions on airstart RS-25 back in 2005 and associated NRE & RE costs. 
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/10/2009 05:29 pm
Hyper vs cryo question:

Is it more common to express run conditions on a hypergol engine in Pc vs MR, whereas on a cryogen engine it's P vs T for the inlet conditions of the propellants?
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/11/2009 03:26 am
Hyper vs cryo question:

Is it more common to express run conditions on a hypergol engine in Pc vs MR, whereas on a cryogen engine it's P vs T for the inlet conditions of the propellants?

I think it's the other way around.  In the US, the yhpergols tend to be smaller pressure-fed engines so inlet pressure (& temp) is important, thought the Russians & Chinese are very good at pump-fed hypergol engines.  Cryos tend to be pump-fed engines therefore we usually specify it in Pc & MR.  Inlet pressure is often expressed in terms of NPSP (net positive suction pressure = P-inlet minus P-vapor) requirement.  Engine guys will then design their pump performance characteristic based on the requirements (NPSP, Pc, MR, cycle, & engine balance) given.  At the end, it's iterations & compromises between the engine guys & the vehicle guys.
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/18/2009 04:48 pm
Reposting an old request I made in the video forum, here in the "engine shop":

In the late 90s there was a video on the web, in the expert.??.purdue.edu domain IIRC, of a couple of enterprising aerospace students putting liquid hydrocarbon fuel in a 2-liter bottle attached to a skateboard and lighting it off in their kitchen.

I'd be in debt to anyone who could find and post it here.  Watching the flame front move through the bottle in slo-mo was pretty neat.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 09/18/2009 07:20 pm
OK, now that sounded too neat not to investigate. I didn't find the Purdue rocketboard but did find something equally entertaining....

http://www.davesdaily.com/videoclips/94-rocketpoweredskateboard.htm

And just for kicks...  (add a "W" to www)

ww.youtube.com/watch?v=LpqwDcM8zcs&feature=channel

Edit: And as Alex noted: Watch out for TO!
Title: Re: Rocket Engine Q&A
Post by: AlexInOklahoma on 09/19/2009 01:15 am
The skateboard motor should not exceed a *four* seg SRB as 5 or more segments just might shake/throw the rider to the ground (vibrations, ya know!)  ;-) 

(Mods: I just HAD to say that, please, please forgive me.....)

Alex
Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 10/03/2009 02:41 am
Has anyone ever tried gasoline as a fuel for a large engine?  It might coke less than kerosene and not have the huge tank size needed for hydrogen.  Or even ethanol might be a good rocket fuel.  It should have very little coking.

Danny Deger
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/03/2009 02:44 am
Or even ethanol might be a good rocket fuel.

Wernher von Braun thought so.  ;D
Title: Re: Rocket Engine Q&A
Post by: strangequark on 10/03/2009 03:20 pm
Has anyone ever tried gasoline as a fuel for a large engine?  It might coke less than kerosene and not have the huge tank size needed for hydrogen.  Or even ethanol might be a good rocket fuel.  It should have very little coking.

Danny Deger

Gasoline makes a terrible regenerative coolant, as the volatiles will try to boil off. RP-1 is all around 12 carbons for that reason. Also, sulphur content in gasoline is not tightly controlled like RP-1. Sulphur does horrendous things to your oxidation properties. Look up the SSME problem they had recently that was linked to a sulphur-containing tacky tape.
Title: Re: Rocket Engine Q&A
Post by: hop on 10/03/2009 08:57 pm
Kerosene also has a density advantage.
Title: Re: Rocket Engine Q&A
Post by: William Barton on 10/03/2009 09:11 pm
Or even ethanol might be a good rocket fuel.

Wernher von Braun thought so.  ;D

In the 1950s "Tom Corbett - Space Cadet" TV series and juvenile novels, which had Willy Ley as scientific advisor, the bad guys are repesented as drinking rocket fuel as an intoxicant. At the time I read them, the only rocket fuel I ever heard of was kerosene, drinking which, I knew would kill you. Eventually, I read Ley's "Rockets, Missile, and men in Space," which supplied the necessary info about V-2 fuel (which was watered down enough it probably was like a particularly horrendous vodka)! Eventually, a chemistry-major friend came up with some pure ethanol, which we sampled, and then agreed Tom Corbett's space pirates must have been some pretty tough hombres...
Title: Re: Rocket Engine Q&A
Post by: William Barton on 10/03/2009 09:13 pm
Has anyone ever tried gasoline as a fuel for a large engine?  It might coke less than kerosene and not have the huge tank size needed for hydrogen.  Or even ethanol might be a good rocket fuel.  It should have very little coking.

Danny Deger

Goddard, apparently...

http://science.howstuffworks.com/rocket5.htm
Title: Re: Rocket Engine Q&A
Post by: DiggyCoxwell on 10/06/2009 08:30 pm
   A friend from Colorado told me that those rockets that use
turbopumps to pump propellant into the engine(s) since 1942
use either an integral H2O2 steam generating system (like that in the V2)
or use pyrotechnics to build up the turbopump's RPM before
siphoned hot engine exhaust takes over the task.
 (1) Is he correct?
 
(2) Or is there a third and maybe even a fourth method of operating the turbopumps?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 10/06/2009 08:54 pm
   A friend from Colorado told me that those rockets that use
turbopumps to pump propellant into the engine(s) since 1942
use either an integral H2O2 steam generating system (like that in the V2)
or use pyrotechnics to build up the turbopump's RPM before
siphoned hot engine exhaust takes over the task.
 (1) Is he correct?
 
(2) Or is there a third and maybe even a fourth method of operating the turbopumps?

(1) Not really, see (2)

(2) Startup in large, groundstarted engines is usually achieved with cold, inert gas (N2, He). Upper-stage expander engines (RL-10, RL-60, Vinci, LE-5) start off of residual heat in the nozzle and chamber. Not sure on all the gas generator upper stage engines, J-2 used a solid propellant (not pyrotechnic) to turn up the turbines. The two methods you mentioned have been used, but I wouldn't call them standard these days. As far as running the engines during flight, there are three common cycles used to run turbopumps when the engines are up and running. The bleed cycle (siphoned, hot engine exhaust) is uncommon, though I believe it was used on the J-2S engine. The classical cycles are:

Expander: Cryogenic fuel is used to cool the thrust chamber and nozzle, then the warm, gasified propellant is run through a turbine, and finally injected into the main combustion chamber

Gas Generator: A small amount of fuel and oxidizer are burned in a separate chamber, the exhaust runs through a high pressure-ratio turbine, and is vented overboard, usually through a secondary nozzle, to provide some thrust.

Staged Combustion: All of one propellant is burned with some of another (to achieve low temperature) in a separate chamber. The exhaust runs through a low-pressure turbine, and is then injected into the main chamber with the rest of the other propellant.

There are variations on these, of course. For instance, there is a Japanese expander engine which only uses some of the hydrogen fuel, then vents it overboard after going through the turbine. There have also been proposals for afterburning gas generators, full flow staged combustion, etc. However, the above represents the vast majority of liquid engines.
Title: Re: Rocket Engine Q&A
Post by: Jim on 10/06/2009 10:46 pm
[Startup in large, groundstarted engines is usually achieved with cold, inert gas (N2, He).

This is the exception vs the rule.

 F-1, H-2, MB-3, RS-27, MA-5, use head start.  These engines used the main propellants in the gas generators to turn the turbines

SSME, RD-180 use head start. These engines used the main propellants in the preburners  to turn the turbines

Titan first and second stages, Agena  used start cartridges. These engines used the main propellants in the gas generators to turn the turbines

The RS-68 and Merlin are the ones that use an inert gas for spin up.

Title: Re: Rocket Engine Q&A
Post by: Jim on 10/06/2009 10:49 pm

Expander:

Gas Generator:

Staged Combustion:

You forgot tapoff and the gas generator with separate mono or bi props
Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 10/07/2009 01:33 pm

Expander:

Gas Generator:

Staged Combustion:

You forgot tapoff and the gas generator with separate mono or bi props

The summary is -- there is more than one way to get the turbines spinning.

Danny Deger
Title: Re: Rocket Engine Q&A
Post by: meiza on 10/07/2009 01:44 pm
Soyuz is an exception. It uses hydrogen peroxide gas generators to turn the pumps. There are separate hydrogen peroxide tanks just for that purpose.

It is a very old fashioned system that has seemed to work just fine and reliably for more than the last 50 years.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/07/2009 01:48 pm
The summary is -- there is more than one way to get the turbines spinning.

What methods are used to spin up turbines for aircraft engines? You seem exactly the person to ask.  :P
Title: Re: Rocket Engine Q&A
Post by: Jim on 10/07/2009 02:17 pm
What methods are used to spin up turbines for aircraft engines? You seem exactly the person to ask.  :P

direct mechanical torque supplied by a ground cart or an onboard electric motor
start cartridges
high pressure gas supplied by a flight or ground APU
Title: Re: Rocket Engine Q&A
Post by: strangequark on 10/07/2009 02:19 pm

Expander:

Gas Generator:

Staged Combustion:

You forgot tapoff and the gas generator with separate mono or bi props

Tapoff! That's the word I was trying to think of. Thought that "bleed cycle" was the wrong term. You are talking about the J-2S cycle, right? Never seen gas generator with separate biprops (maybe I haven't been looking hard enough ;-)), which engines? Thanks for the corrections and additional information.
Title: Re: Rocket Engine Q&A
Post by: DiggyCoxwell on 10/07/2009 08:18 pm
   A friend from Colorado told me that those rockets that use
turbopumps to pump propellant into the engine(s) since 1942
use either an integral H2O2 steam generating system (like that in the V2)
or use pyrotechnics to build up the turbopump's RPM before
siphoned hot engine exhaust takes over the task.
 (1) Is he correct?
 
(2) Or is there a third and maybe even a fourth method of operating the turbopumps?

(1) Not really, see (2)

(2) Startup in large, groundstarted engines is usually achieved with cold, inert gas (N2, He). Upper-stage expander engines (RL-10, RL-60, Vinci, LE-5) start off of residual heat in the nozzle and chamber. Not sure on all the gas generator upper stage engines, J-2 used a solid propellant (not pyrotechnic) to turn up the turbines. The two methods you mentioned have been used, but I wouldn't call them standard these days. As far as running the engines during flight, there are three common cycles used to run turbopumps when the engines are up and running. The bleed cycle (siphoned, hot engine exhaust) is uncommon, though I believe it was used on the J-2S engine. The classical cycles are:

Expander:

Gas Generator:

Staged Combustion: .

   Interesting. Thanks.
Between you and Jim you revealed some interesting variety of methods to do the task.

 BTW, with some commercial H202 propulsion systems in the works,
would H202 react/decompose prematurely when whipped around
by turbine blades? And would that present more of a problem than
a benefit?
 
Title: Re: Rocket Engine Q&A
Post by: strangequark on 10/07/2009 08:24 pm
   A friend from Colorado told me that those rockets that use
turbopumps to pump propellant into the engine(s) since 1942
use either an integral H2O2 steam generating system (like that in the V2)
or use pyrotechnics to build up the turbopump's RPM before
siphoned hot engine exhaust takes over the task.
 (1) Is he correct?
 
(2) Or is there a third and maybe even a fourth method of operating the turbopumps?

(1) Not really, see (2)

(2) Startup in large, groundstarted engines is usually achieved with cold, inert gas (N2, He). Upper-stage expander engines (RL-10, RL-60, Vinci, LE-5) start off of residual heat in the nozzle and chamber. Not sure on all the gas generator upper stage engines, J-2 used a solid propellant (not pyrotechnic) to turn up the turbines. The two methods you mentioned have been used, but I wouldn't call them standard these days. As far as running the engines during flight, there are three common cycles used to run turbopumps when the engines are up and running. The bleed cycle (siphoned, hot engine exhaust) is uncommon, though I believe it was used on the J-2S engine. The classical cycles are:

Expander:

Gas Generator:

Staged Combustion: .

   Interesting. Thanks.
Between you and Jim you revealed some interesting variety of methods to do the task.

 BTW, with some commercial H202 propulsion systems in the works,
would H202 react/decompose prematurely when whipped around
by turbine blades? And would that present more of a problem than
a benefit?
 

If you're actually talking turbines, then the H2O2 would never see a turbine blade. It'll be decomposed into H2O and O2 by then. If you mean the pump inducer or impeller, I would imagine that cavitation would be the primary concern. The pumping action itself would be fine, as long as sufficient margin existed between the lowest system pressure and the peroxide vapor pressure.
Title: Re: Rocket Engine Q&A
Post by: hop on 10/07/2009 11:17 pm
BTW, with some commercial H202 propulsion systems in the works would H202 react/decompose prematurely when whipped around by turbine blades? And would that present more of a problem than a benefit?
The Soyuz boosters and first stage core still use H202 to drive the turbopumps. Seems to work OK ;)

You can find some pretty detailed information on this site: http://www.lpre.de/energomash/RD-107/index.htm (in russian, but google translate will get you the gist. Thanks to pm1823 for posting that http://forum.nasaspaceflight.com/index.php?topic=18646.msg472072#msg472072 )
Title: Re: Rocket Engine Q&A
Post by: Antares on 10/11/2009 04:38 pm
newbie question regarding engine tests..
after the test fires do they "clean" the engine at all after the test or can they take the engines straight to the cape as is.

Generally all engines have some sort of post hot-fire inspection and processing regime.  This may involve cleaning, especially on a hydrocarbon engine - usually isopropyl alcohol or a similar solvent as a wipe.  There's probably something in the SSME bible on L2.
Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 10/11/2009 07:07 pm
What methods are used to spin up turbines for aircraft engines? You seem exactly the person to ask.  :P

direct mechanical torque supplied by a ground cart or an onboard electric motor
start cartridges
high pressure gas supplied by a flight or ground APU

If air borne the wind blowing through the engine will spin it up enough to start.

Danny Deger

Edit: For an amusing story on using start cartridges in starting a large number of F-4Es training for "Toe to toe nuclear combat with the Ruskies" see page 58 of this download www.lulu.com/dannydeger

The Wing Commander's airplane gets burned up in the process.   Needless to say, he was not happy about this situation.  Of all the airplanes on the ramp you want to burn up, the Wing Commander's personal airplane is NOT the one to choose  :o
Title: Re: Rocket Engine Q&A
Post by: kkattula on 10/12/2009 07:28 am
XCOR aren't using a turbine at all.  They use a piston pump powered by, IIRC,  compressed gas.

Flowmetric have developed a pistonless pump system, again powered by compressed gas, or one of the 'traditional' cycles.

These pumps have lower power densities than turbo-pumps, but are cheaper, simpler and more robust.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 10/15/2009 02:33 pm
Has anyone ever tried gasoline as a fuel for a large engine?  It might coke less than kerosene and not have the huge tank size needed for hydrogen.  Or even ethanol might be a good rocket fuel.  It should have very little coking.

Danny Deger

Late reply, sorry if someone else have already answer this question.

An interesting fact not well known is that kerosene engines only cokes if the chamber pressure is less than 2,000 psia.  Not sure if I got that from Huzel & Huang, or from talking to one of engine manufacturers.  But perhaps someone here can tell us if the RD-180 cokes in the chamber?

Kerosene (RP-1) is a hydrocarbon mixture, higher density than the regular gasoline (Octane, C8H18).  But all hydrocarbon fuels burn, down to methane (CH4) with only 1 carbon.  The higher the hydrogen-to-carbon ration (H/C ratio), the less likelihood to coke. RP-1 averages about 2:1 for H/C ratio.

The disadvantage of lower carbon fuel such as methane is lower density which means a bigger tank.  As you know, tanks dominates the size of rocket and cost money to build.  That's why vehicle manufacturers stay with RP-1 for its higher bulk density property.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 10/15/2009 02:54 pm
Has anyone ever tried gasoline as a fuel for a large engine?  It might coke less than kerosene and not have the huge tank size needed for hydrogen.  Or even ethanol might be a good rocket fuel.  It should have very little coking.

Danny Deger

Gasoline makes a terrible regenerative coolant, as the volatiles will try to boil off. RP-1 is all around 12 carbons for that reason. Also, sulphur content in gasoline is not tightly controlled like RP-1. Sulphur does horrendous things to your oxidation properties. Look up the SSME problem they had recently that was linked to a sulphur-containing tacky tape.

Well, our refining technology has gotten much much better, particularly in removing sulfur.  For example, commercial refineries produce much CLEANER RP-1 than the current MIL-SPEC specifies.

Another interesting phenomenon is that, up until 15 yrs ago, not many people were interested in getting hydrocarbon fuel heated up.  In fact, the Air Force fuel handbook did not show jet fuel thermal stability temperature above 300 deg. F.  Since then, much work has been done to understand fuel thermal stability and AFRL has published AIAA papers showing one can heat jet fuel up past 800 deg. F.



Title: Re: Rocket Engine Q&A
Post by: Propforce on 10/15/2009 03:00 pm

Gas Generator: A small amount of fuel and oxidizer are burned in a separate chamber, the exhaust runs through a high pressure-ratio turbine, and is vented overboard, usually through a secondary nozzle, to provide some thrust.

There are two different issues here, start-up the turbopump and during the regular run (steady-state) condition.  What you've stated above is true for the run condition. 

For start-up, take the J-2/ J-2X for example, they use(d) an inert gas, helium in this case, to spin start the pumps.  On the J-2, they refill the helium bottle with gaseous hydrogen and use it for the 2nd start (in a 2-burn mission).
Title: Re: Rocket Engine Q&A
Post by: Propforce on 10/15/2009 03:02 pm
What methods are used to spin up turbines for aircraft engines? You seem exactly the person to ask.  :P

direct mechanical torque supplied by a ground cart or an onboard electric motor
start cartridges
high pressure gas supplied by a flight or ground APU

If air borne the wind blowing through the engine will spin it up enough to start.

Danny Deger

I guess there is no issue on air-start !!!  ;D

Kinda like old cars, just push it down the hill and start the engine.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 11/12/2009 12:37 am
Are Flyback Booster concepts such as these:
say, using Hydrogen RS-83's or Kerolox RS-84's just not competitive cost wise with Shuttle derived solid rocket motors?

I know flight rate factors into it, but let's just use a rate of 12 per year for example.

Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 11/12/2009 03:14 am
Are Flyback Booster concepts such as these:
say, using Hydrogen RS-83's or Kerolox RS-84's just not competitive cost wise with Shuttle derived solid rocket motors?

I know flight rate factors into it, but let's just use a rate of 12 per year for example.



My guess is solid with parachute recovery is cheaper than this design.  Also a problem is where to land two flybacks and have a runway open for a Return to Landing Site abort.  Last point is lots of empty mass for all the flyback and landing hardware. 

This extra mass is a killer on a rocket.

Danny Deger
Title: Re: Rocket Engine Q&A
Post by: sdsds on 11/12/2009 03:59 am
Are Flyback Booster concepts such as these:
say, using Hydrogen RS-83's or Kerolox RS-84's just not competitive cost wise with Shuttle derived solid rocket motors?

I know flight rate factors into it, but let's just use a rate of 12 per year for example.

I'm curious what people imagine the turn-around time on liquid fly-back boosters might be?  If all the boosters in a set could be processed in parallel, and the fleet included two sets, is the 12-per-year flight rate mentioned above realistic?  That is, could a given set of boosters fly every 60 days?
Title: Re: Rocket Engine Q&A
Post by: deltaV on 11/12/2009 02:34 pm
1) What is the minimum altitude that a RL-10B-2 is designed for use at? For example could a vacuum RL-10B-2 be lit at 60 kft? Is the extendable nozzle designed for altitude compensation? Do other vacuum engines have a similar minimum altitude?

The following links seem to suggest that a RL10B2 at 60 kft would be quite overexpanded.
http://www.aerospaceweb.org/design/aerospike/compensation.shtml
http://books.google.com/books?id=s1C9Oo2I4VYC&lpg=PA93&ots=eM5dM5hvLV&dq=rocket%20overexpansion&pg=PA93#v=onepage&q=rocket%20overexpansion&f=false
http://www.spaceandtech.com/spacedata/engines/rl10_specs.shtml

2) Would it be hard to modify a vacuum engine to allow it to operate at sea level during an abort? No need for good ISP, just produce some thrust and don't cause too much damage.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 11/12/2009 02:39 pm
I don't know how long SSME's take to refurb but I think we'd be crazy to let the SRB capability/technology we have now fall to the wayside. Yet another reason the Ares V is a good idea (IMHO).

With RS-68A's to be the mainstay in US engines... Ares V.
I'm curious once the shuttle goes away will SSME's go the way of F1? Too expensive? It will be interesting to see what happens. Guess it all depends on what we as a nation want to commit to.
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/12/2009 04:11 pm
RL10 questions:

Do you know how to do basic nozzle expansion calculations?  This is pretty easy.  You can use
http://www.aoe.vt.edu/~devenpor/aoe3114/calc.html
You can get the area ratio and chamber pressure (p0) off the net.  Input area ratio and get out Mach and p/p0 at the exit plane.  Use that Mach on the normal shock calculator to get a pressure ratio (p2/p1) if there's a normal shock at the exit.  Multiply chamber pressure by p/p0 and by p2/p1 to get the pressure downstream of the shock.  Use that pressure in a standard atmosphere calculator to figure the altitude
http://www.digitaldutch.com/atmoscalc/

The altitude I get is 10671 meters.  But that's only the minimum altitude (not to mention isentropic) that doesn't tear apart the nozzle.  Your thrust would suck due to the shock.  Which leads to...

In general, even without any horrible shock losses, upper stage engines have insufficient thrust for an abort motor.  You wouldn't want that much thrust on a nominal mission, and please don't mention something designed for both modes.
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/12/2009 04:11 pm
This is interesting:

http://gravityloss.wordpress.com/2009/11/12/dual-propellant-expander/
Title: Re: Rocket Engine Q&A
Post by: yinzer on 11/12/2009 07:11 pm
This is interesting:

http://gravityloss.wordpress.com/2009/11/12/dual-propellant-expander/

Hard to make out the details of their picture, but they appear to be showing a LOX pump on the same shaft as a GH2 turbine.  I don't think that's how the RL10 works - doesn't it have the LH2 pump on the same shaft as the turbine and a gear-driven LOX pump?
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/12/2009 11:46 pm
It's a LH2 pump on a GH2 turbine and a LOX pump on a GOX turbine.  That way you don't have to have an interpropellant seal.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 11/13/2009 01:15 am

Thanks, those tools are quite helpful. A follow-up question: can an RL10B2 operate with the nozzle extension retracted? If so would it be possible (and easy) to lower the extension while the engine is operating? I'm not asking for continuous control of the area ratio, just a sudden increase from area ratio 130 to 285 at the appropriate altitude, which if my calculations are right is around 120 kft.
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/13/2009 05:11 pm
Yes

No
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/14/2009 06:19 am
No

Would the answer be different for an ablatively cooled nozzle?
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/14/2009 03:27 pm
No, the deployment system can't overcome the pressure of the expanding exhaust.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/15/2009 07:27 am
As the exhaust gases in a rocket cool down as they pass through the nozzle, is there ever a risk of cooling them down so much that they condense? That sounds like a bad thing, but maybe it's impossible.
Title: Re: Rocket Engine Q&A
Post by: William Barton on 11/15/2009 09:18 am
As the exhaust gases in a rocket cool down as they pass through the nozzle, is there ever a risk of cooling them down so much that they condense? That sounds like a bad thing, but maybe it's impossible.

Somewhere (I don't recall where), there is an image of an LH2/LOX rocket engine running with ice forming at the rim of the engine bell. I think it's from atmospheric moisture, but I'm not sure. Running in an atmosphere, I don't think hydrogen and oxygen could cool enough to liquify, and in vacuum, I don't think it would matter whether the rocket exhaust was gas, vapor, or ice crystals, or bricks of frozen whatever, for that matter.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/15/2009 09:32 am
That was probably a test firing of the CECE, I was thinking of that too. Don't fire extinguishers work by cooling through a nozzle? Is the white stuff that comes out solid CO2 or is it just frozen water from the atmosphere?
Title: Re: Rocket Engine Q&A
Post by: William Barton on 11/15/2009 09:49 am
That was probably a test firing of the CECE, I was thinking of that too. Don't fire extinguishers work by cooling through a nozzle? Is the white stuff that comes out solid CO2 or is it just frozen water from the atmosphere?

In dry chemical fire extinguishers (I keep one each at home and in my office), the white stuff is a bicarbonate powder/foam that works with the decomposition-generated CO2 to smother the fire. Most of my home/office fires have been either grease of electrical. I've never used one of those pure-CO2 restaurant extinguishers, so I don't know if they make any white stuff (I doubt it). I've also seen one big IT server room halogen extinguisher in action. I was too busy getting out of the room to notice if there was any vapor!
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/15/2009 09:58 am
I was too busy getting out of the room to notice if there was any vapor!

Damn it Barton, pay attention next time. This stuff is important. If you won't do it for me, do it for science!  ;D
Title: Re: Rocket Engine Q&A
Post by: nacnud on 11/15/2009 03:11 pm
I've used pressurised liquid co2 released into a container to make dry ice before. So a pure co2 extinguisher could defiantly be used to make dry ice too.
Title: Re: Rocket Engine Q&A
Post by: kch on 11/15/2009 03:22 pm
I've used pressurised liquid co2 released into a container to make dry ice before. So a pure co2 extinguisher could defiantly be used to make dry ice too.

How defiant would one have to be to make it work?  ;)
Title: Re: Rocket Engine Q&A
Post by: Danny Dot on 11/15/2009 03:23 pm
I've used pressurised liquid co2 released into a container to make dry ice before. So a pure co2 extinguisher could defiantly be used to make dry ice too.

And all this dry ice could be use to make some very cold martinis after the fire is out  8)

Seriously now.  I know a divergent nozzle with supersonic flow accelerates the flow while a subsonic divergent nozzle has the flow slowing down.  I never have understood why in detail.  It must have to do with the fact any disturbance is not going to be transmitted forward into the supersonic flow.  Can anyone put words to this effect?

Anyway, this is how a rocket throat/nozzle works.  The flow is Mach 1 at the throat, then accelerates to high Mach numbers (14 or so) as it flows out the divergent nozzle.

Danny Deger

P.S.  Good luck to the crew and the entire NASA team tomorrow.  Scortch was kind of a pain in the rear to train at times, but is certainly one of the best commanders in the history of the shuttle -- in my opinion anyway.
Title: Re: Rocket Engine Q&A
Post by: Antares on 11/15/2009 03:53 pm
It's the back pressure / area ratio combination that drives whether the nozzle is supersonic or subsonic.  In space, every nozzle is supersonic.  There's a good picture in Anderson that describes this.  Edit: Chapter 4 in Fundamentals, Sections 5.4 and 5.5 and Figure 5.24 in MCF
Title: Re: Rocket Engine Q&A
Post by: meiza on 11/15/2009 07:06 pm

Thanks, those tools are quite helpful. A follow-up question: can an RL10B2 operate with the nozzle extension retracted? If so would it be possible (and easy) to lower the extension while the engine is operating? I'm not asking for continuous control of the area ratio, just a sudden increase from area ratio 130 to 285 at the appropriate altitude, which if my calculations are right is around 120 kft.

Extending the nozzle while an engine is firing has been demonstrated on a test stand. Even when it was gimbaling at the same time. So it is certainly doable in a broader sense even if a current design doesn't do it.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 11/16/2009 12:39 pm
That was probably a test firing of the CECE, I was thinking of that too. Don't fire extinguishers work by cooling through a nozzle? Is the white stuff that comes out solid CO2 or is it just frozen water from the atmosphere?

To use Jim's (tm) reply style,

Yes http://science.nasa.gov/headlines/y2009/15jan_cece.htm?list1294036
Title: Re: Rocket Engine Q&A
Post by: Propforce on 11/16/2009 02:25 pm

Thanks, those tools are quite helpful. A follow-up question: can an RL10B2 operate with the nozzle extension retracted? If so would it be possible (and easy) to lower the extension while the engine is operating? I'm not asking for continuous control of the area ratio, just a sudden increase from area ratio 130 to 285 at the appropriate altitude, which if my calculations are right is around 120 kft.

Extending the nozzle while an engine is firing has been demonstrated on a test stand. Even when it was gimbaling at the same time. So it is certainly doable in a broader sense even if a current design doesn't do it.

Yes it has.  I think Antares was addressing specifically if the B-2 engine can do this, e.g., retracting the nozzle while engine is firing.  The B-2 NEDS is not designed for the stiffness and the forces to either extend or retract nozzle extension during hot fire.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 11/16/2009 02:33 pm

Thanks, those tools are quite helpful. A follow-up question: can an RL10B2 operate with the nozzle extension retracted? If so would it be possible (and easy) to lower the extension while the engine is operating? I'm not asking for continuous control of the area ratio, just a sudden increase from area ratio 130 to 285 at the appropriate altitude, which if my calculations are right is around 120 kft.

I think what you're trying to achieve is the concept of "altitude compensating nozzle" by mechanically deploy nozzle extension as the rocket climbs in altitude.  Recognizing this has been done, via hot fire demonstration on the ground, the question is if this is worthwhile trading overall system benefit with the mechanical complexity (& weight) of a deployment system. 


Title: Re: Rocket Engine Q&A
Post by: Propforce on 11/16/2009 02:34 pm
I was too busy getting out of the room to notice if there was any vapor!

Damn it Barton, pay attention next time. This stuff is important. If you won't do it for me, do it for science!  ;D

I think this will be a good CFD thesis for someone  ;D
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/17/2009 04:24 pm
Seriously now.  I know a divergent nozzle with supersonic flow accelerates the flow while a subsonic divergent nozzle has the flow slowing down.  I never have understood why in detail.  It must have to do with the fact any disturbance is not going to be transmitted forward into the supersonic flow.  Can anyone put words to this effect?

See this link, which derives this mystery from the mystery of isentropic flow: Nozzle Design (http://exploration.grc.nasa.gov/education/rocket/nozzle.html).
Title: Re: Rocket Engine Q&A
Post by: Propforce on 11/18/2009 08:49 am
I've used pressurised liquid co2 released into a container to make dry ice before. So a pure co2 extinguisher could defiantly be used to make dry ice too.

And all this dry ice could be use to make some very cold martinis after the fire is out  8)

Seriously now.  I know a divergent nozzle with supersonic flow accelerates the flow while a subsonic divergent nozzle has the flow slowing down.  I never have understood why in detail.  It must have to do with the fact any disturbance is not going to be transmitted forward into the supersonic flow.  Can anyone put words to this effect?

Danny, it's the same reason why heat flows from hot to cold, why a boat is just a money pit, and why you should never lick a frozen telephone pole.  Just accept it because it is.

There's a (P/Pt M-dot) relationship in the explanation somewhere, and you can trace the fanno line from the method of characteristic.  Now do that calc by hand like any self respecting engineer would.  But damn it, nobody quote Shapiro anymore!  That was the bible where I learned compressible flow.  ;D
Title: Re: Rocket Engine Q&A
Post by: butters on 12/12/2009 06:27 pm
Has anyone studied the concept of compressed gas propellant feed pumps as a compromise between pressure-fed and pump-fed liquid rocket engines?

The basic idea is to use a convenient gas (nitrogen?) as the pressure medium instead of the propellants themselves, and use the expanding gas to drive pneumatic pumps to feed the propellants into the nozzle.

This avoids high-pressure propellant tanks.  Perhaps the convenient gas would require much smaller, lighter, and cheaper tanks to contain the necessary pressure/mass?

If high enough chamber pressure could be produced in this manner, then the specific impulse could be on par with staged combustion engines with less complexity than even gas generator engines.

Silly idea?
Title: Re: Rocket Engine Q&A
Post by: jongoff on 12/12/2009 07:08 pm
Butters,
Google "flometrics pistonless pump"

~Jon
Title: Re: Rocket Engine Q&A
Post by: butters on 12/12/2009 07:32 pm
Butters,
Google "flometrics pistonless pump"

~Jon

Thanks, Jon!
Title: Re: Rocket Engine Q&A
Post by: jongoff on 12/12/2009 11:10 pm
Butters,
Google "flometrics pistonless pump"

~Jon

Thanks, Jon!

They've done a decent amount of consulting with us (including doing some work on a pistonless pump that we had to put on hold back in '06).  I hope at some point in the next year or so that we can have enough money to have them finish up the work they had started.  Of all the pump-feed concepts out there, this is probably tied for my favorite.

~Jon
Title: Re: Rocket Engine Q&A
Post by: butters on 12/20/2009 04:47 am
Of all the pump-feed concepts out there, this is probably tied for my favorite.

~Jon

OK, I'm too curious not to ask... what is your other favorite pump-feed concept?

I also have a general launch vehicle design question (for anyone who would like to answer):

What are the trade considerations for LOX/RP-1 rocket stages between fuel on the bottom (e.g. Atlas V, Falcon 1/9) or fuel on the top (e.g. Delta II)?
Title: Re: Rocket Engine Q&A
Post by: jongoff on 12/20/2009 05:14 am
Of all the pump-feed concepts out there, this is probably tied for my favorite.

~Jon

OK, I'm too curious not to ask... what is your other favorite pump-feed concept?

One I can't really talk about yet due to an NDA (and yeah, I probably shouldn't have been a tease with something I couldn't talk about yet).  If I can talk the guy who came up with most of the idea into it, and if I can talk the session chair into allowing a very late entry, I may try putting together an AIAA paper for the Joint Propulsion Conference this fall.  Alternately, if the guy is able to file a provisional patent, I'll blog the concept.  It's a *really* cool idea that has the potential to be a gamechanger in pump-fed rocket design.

It's not as simple as pistonless pumps, and while it does have higher performance, it's not clear yet even if it works that it would always be better.  Which is why I say it's a tie at the moment.

~Jon
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/20/2009 09:39 am

What are the trade considerations for LOX/RP-1 rocket stages between fuel on the bottom (e.g. Atlas V, Falcon 1/9) or fuel on the top (e.g. Delta II)?

CG, feed duct length
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 12/20/2009 04:13 pm
feed duct length

Can you say more? Doesn't one duct get longer and the other shorter? Is length more important for the fuel than for the oxidiser or the other way round?
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/20/2009 04:42 pm
feed duct length

Can you say more? Doesn't one duct get longer and the other shorter? Is length more important for the fuel than for the oxidiser or the other way round?

One of the ducts handles a cryogenic fluid and has to deal with insulation/heat transfer issues
Title: Re: Rocket Engine Q&A
Post by: yinzer on 12/20/2009 05:02 pm
feed duct length

Can you say more? Doesn't one duct get longer and the other shorter? Is length more important for the fuel than for the oxidiser or the other way round?

Both ducts have to go from the bottom of the bottom tank to the engine. One duct has to go from the top of the bottom tank to the bottom of the bottom tank. With the shorter tank on bottom, total duct length is reduced.
Title: Re: Rocket Engine Q&A
Post by: Antares on 12/21/2009 11:40 pm
Closely related is whether to plumb the top tank through the bottom tank or around the bottom tank and in the airflow.
Title: Re: Rocket Engine Q&A
Post by: Art LeBrun on 12/22/2009 02:24 am
Closely related is whether to plumb the top tank through the bottom tank or around the bottom tank and in the airflow.

Besides Atlas throughout the ages and N-1 what other vehicles have used external feed or fill/feed lines?
Title: Re: Rocket Engine Q&A
Post by: kraisee on 12/22/2009 05:25 am
Shuttle ET, Saturn S-II, Saturn S-IVB, Delta-IV.   The list goes on...

Most vehicles use external feedlines because they are less complicated and thus less expensive than internal ones.   The aerodynamic penalty is often less of a concern than the economically painful alternative.

Ross.
Title: Re: Rocket Engine Q&A
Post by: sdsds on 12/22/2009 06:56 am
Shuttle ET, Saturn S-II, Saturn S-IVB, Delta-IV.   The list goes on...

Most vehicles use external feedlines because they are less complicated and thus less expensive than internal ones.   The aerodynamic penalty is often less of a concern than the economically painful alternative.

Ross.

The aerodynamic drag penalty for external feedlines must surely be small? 

So the reasons for internal feedlines are:
* Less complex aerodynamic behavior allows simpler (or no) active flight control.
* Leaves room for more SRBs.  See e.g. Delta II with nine!
* Less vulnerable to damage (?)
* Aesthetics.

Any others?
Title: Re: Rocket Engine Q&A
Post by: ugordan on 12/22/2009 08:19 am
* Less complex aerodynamic behavior allows simpler (or no) active flight control.

No active control? That won't get you far in the real world...
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/22/2009 12:07 pm
The aerodynamic drag penalty for external feedlines must surely be small? 

So the reasons for internal feedlines are:
* Less complex aerodynamic behavior allows simpler (or no) active flight control.
* Leaves room for more SRBs.  See e.g. Delta II with nine!
* Less vulnerable to damage (?)
* Aesthetics.

Any others?

Less turns and bends in the lines.

Major drawbacks are assembly complexity and insulation issues.
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/22/2009 12:09 pm

So the reasons for internal feedlines are:
* Less complex aerodynamic behavior allows simpler (or no) active flight control.


That isn't a reason. All launch vehicles require active control.
Title: Re: Rocket Engine Q&A
Post by: Antares on 12/22/2009 09:35 pm
The "aerodynamic" effect is direct frictional and convective heating of the external surface, which then conducts inward to the propellant flowing through the duct.

Both configurations could likely create operational requirements for keeping the propellant in the line conditioned prior to engine start... like the "LOX drainback" call in a Shuttle count.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 12/22/2009 09:46 pm
Are noncryogenic propellants immune from this issue?
Title: Re: Rocket Engine Q&A
Post by: sdsds on 12/23/2009 12:09 am

So the reasons for internal feedlines are:
* Less complex aerodynamic behavior allows simpler (or no) active flight control.


That isn't a reason. All launch vehicles require active control.

Understood, thanks.  The aerodynamic issues would only apply to rockets used for other purposes, though even lowly scuds have active TVC that could probably compensate for external feedlines. 

Admittedly I was thinking mainly of hobby rocketry where fins play a large role in stabilization.  The issue with developing active guidance for hobbiests might be more regulatory than technical.  (See comment above about scuds.) 

Does anything preclude a liquid-propellant sounding rocket from using only fin and spin stabilization?  If not, that could be a commercial vehicle where external feedlines might be an aerodynamic issue....
Title: Re: Rocket Engine Q&A
Post by: kraisee on 12/23/2009 02:00 am
You still need pitch, yaw and roll control outside the atmosphere.

Ross.
Title: Re: Rocket Engine Q&A
Post by: Antares on 12/24/2009 12:19 am
Are noncryogenic propellants immune from this issue?

It's a question of meeting the qualified pump inlet conditions of the engine no matter what the propellants are.  One could design a pump to run on gaseous propellants, but engine performance would suck.
Title: Re: Rocket Engine Q&A
Post by: butters on 12/24/2009 10:15 am
So if the upper tank feed line is to be external, then the insulation and pumping considerations would favor RP-1 on the top.  This would be the simplest design, both for the LOX system and the upper tank feed line.
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/24/2009 12:07 pm
So if the upper tank feed line is to be external, then the insulation and pumping considerations would favor RP-1 on the top.  This would be the simplest design, both for the LOX system and the upper tank feed line.

But there are control considerations to account for.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/15/2010 04:26 am
Ran across this dated J-2X info on the Volvo Aero site:

"More facts on the J-2X nozzle
The J-2X nozzle has two components, an upper regenerative cooled nozzle and a lower film cooled extension.
For the upper part Volvo Aero will propose a sandwich design, the same technology that was chosen by Pratt & Whitney in 2001 for the RL60 engine. In the RL60 program Volvo Aero managed to concurrently design and build a sandwich demon nozzle in the record time of 18 months. The sandwich technology, patented by Volvo Aero, has also been selected by ESA for a full scale demonstration on the Vulcain 2 engine late 2007.
The lower J-2X nozzle extension will be cooled with a supersonic film injection of turbine exhaust gases. This is a technology Volvo Aero developed with great success for the Vulcain 2 engine, and is the only flight proven super sonic film cooling technology system in the world. This experience puts Volvo Aero in a unique position and enables a straight forward development of a similar system for the J-2X engine."

http://www.volvogroup.com/group/global/en-gb/newsmedia/pressreleases/previous/_layouts/CWP.Internet.VolvoCom/NewsItem.aspx?News.ItemId=14597&News.Language=en-gb

A few questions if I may:
Is PWR still going forward with this nozzle from Volvo Aero for J-2X?

Also, if a decision was made to go RS-68 Regen, would this type of nozzle be applicable? Is there a certain type of nozzle PWR is looking at for the 68 regen option?

There was also an interesting "sandwich nozzles guide" here if anyone was interested:
http://www.volvoaero.com/VOLVOAERO/GLOBAL/EN-GB/PRODUCTS/SPACE%20PROPULSION/NOZZLES/Pages/Nozzles.aspx
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/15/2010 09:14 pm

A quote from this blog:
http://chairforceengineer.blogspot.com/2009/08/cool-engines.html

"But as I've argued in my previous post, SSME's that are designed for low production costs will be very different from the baseline SSME."


Question: Is it fair to say the SSME route isn't as rosy as envisaged? Especially given the performance of the 68R / J-2X EDS combo?
Title: Re: Rocket Engine Q&A
Post by: jongoff on 01/15/2010 11:51 pm

A quote from this blog:
http://chairforceengineer.blogspot.com/2009/08/cool-engines.html

"But as I've argued in my previous post, SSME's that are designed for low production costs will be very different from the baseline SSME."


Question: Is it fair to say the SSME route isn't as rosy as envisaged? Especially given the performance of the 68R / J-2X EDS combo?

Honestly I think it's something that you can have honest differences of opinions.  If you assume human rating has to be really expensive and hard, then taking an already "human rated engine" might be cheaper.  I'm skeptical that you can really make a 3000psi LOX/LH2 rocket engine of that complexity "cheap" compared to designing a new nozzle for something designed with a much simpler cycle and half the chamber pressure. But what would I know--I just build regen cooled engines for a living...  much smaller ones that don't use LH2, but still regen cooled.

~Jon
Title: Re: Rocket Engine Q&A
Post by: jongoff on 01/15/2010 11:58 pm
Ran across this dated J-2X info on the Volvo Aero site:

<snip stuff about their sandwich nozzle concept>

I'm a big fan of the sandwich wall concept--it gives you lots of benefits compared to traditional regen cooled nozzle concepts.  I've actually been doing some low-level work on a similar concept based on our aluminum-chamber work at Masten, but have been too busy with other stuff to get funding to do the actual manufacturing development work (my Bachelors was in Manufacturing Engineering, so I like coming up with crazy new ways of building stuff).

~Jon
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/16/2010 02:14 am
Thanks for the comments Jon. PWR must have a few concepts for a 68R but I sure haven't found articles from them. (Maybe I'm not looking in the right places?) Since J-2X is using this type of nozzle, a layperson like myself naturally wonders if it could be used for the 68R as well?
Title: Re: Rocket Engine Q&A
Post by: Antares on 01/16/2010 04:25 am
1) why would regen RS-68 be made differently from the part of the chamber/nozzle that is already regen?

2) questions/discussions about specific engines could threaten to derail this heretofore mostly generic thread.  I ask my fellow rocket scientists to consider taking long and specific discussions to other threads.   
Title: Re: Rocket Engine Q&A
Post by: kraisee on 01/16/2010 08:06 am
Thanks for the comments Jon. PWR must have a few concepts for a 68R but I sure haven't found articles from them. (Maybe I'm not looking in the right places?) Since J-2X is using this type of nozzle, a layperson like myself naturally wonders if it could be used for the 68R as well?

Are they still using that style of nozzle for J-2X?   The Volvo site appears to have been thoroughly purged of most mentions related to J-2X (except in a small number of documents/pages which are no longer "in date", like those mentions above).

I have had a ^&%$%$ of a time trying to get an answer whether Volvo are still involved in that project or not -- nobody seems to know for sure.   Given the benefit of publicity from such involvement on Volvo's part, I would expect that they would be making a lot more of this cooperation if they were still involved.

Ross.
Title: Re: Rocket Engine Q&A
Post by: clongton on 01/16/2010 08:40 am
Question: Is it fair to say the SSME route isn't as rosy as envisaged? Especially given the performance of the 68R / J-2X EDS combo?

... If you assume human rating has to be really expensive and hard, then taking an already "human rated engine" might be cheaper.  I'm skeptical that you can really make a 3000psi LOX/LH2 rocket engine of that complexity "cheap" compared to designing a new nozzle for something designed with a much simpler cycle and half the chamber pressure. But what would I know--I just build regen cooled engines for a living...  much smaller ones that don't use LH2, but still regen cooled.

~Jon

Jon, I'm surprised at that from you. Your answer is incomplete.

It's one thing to design a human rated regeneratively cooled engine from the start, which is not overly expensive to do. You know that because that's what you do for a living (except the human rating part), as well as what I used to do for a living years ago when I worked on the F-1A. It's quite another thing to take an existing non-human rated ablatively cooled engine that was *never* even remotely envisioned to be human rated *and* cooled in any anther way and *CONVERT* it to be human rated *and* regeneratively cooled. That process is far more expensive. Changing an engine like the RS-68 from non-human rated and ablative cooling to human rated and regenerative cooling is a lot harder and a lot more expensive than just designing an engine from scratch to be a human rated regeneratively cooled engine of the same performance level. It's not just a nozzle swap, as your post implies. Changing the nozzle from ablative to regen is the least expensive part of the conversion. So much of the engine internals need changing out and complete redesign that it becomes for all intents and purposes an entirely new engine. In addition to what has to happen inside the turbine to create pressures that even make regen possible, so many of the things like valves and actuators among other things, which are currently single fault systems, now need to be replaced to be dual fault tolerant for the human rating process but with the added constraint that it has to stay an RS-68 variant. The difficulties of staying within that constraint drive the cost up far beyond what it would otherwise cost to just do a clean sheet design to get the same performance from a human rated regen engine. So why didn't NASA just do a new design? Because then the whole cost would be theirs alone while if it stays an RS-68, even if it costs more, they can get the Air Force to share the cost, thus keeping their own costs down. Politically smart - technically stupid.

Think of it this way: The RS-68 regen is to the RS-68 as the J-2X is to the J-2. And we all know how vast the difference is between the J-2X and the J-2 - it's enormous. It would have been less expensive to just do a clean sheet new design. Heck, mounting a single pair of existing J-2's, a *very* reliable engine combination, would have produced far more performance at less than 1/4 the cost of the expense of designing and building the J-2X.

I could go over with you almost to the individual part level what has to be changed inside the RS-68 engine to enable the use of a regen nozzle as well as make the engine human rated but that would take this discussion far off topic.

The bottom line, and to bring this post back to being on topic for this thread - Clean sheet designs are usually far less expensive than any major conversions of existing engines for the same end result performance. They always have been and they always will be.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 01/16/2010 03:04 pm
A quote from this blog:
http://chairforceengineer.blogspot.com/2009/08/cool-engines.html

"But as I've argued in my previous post, SSME's that are designed for low production costs will be very different from the baseline SSME."


Question: Is it fair to say the SSME route isn't as rosy as envisaged? Especially given the performance of the 68R / J-2X EDS combo?

An HR'd RS-68R will also be very different from the currently-available RS-68, and won't be available for quite a few years.

But the block II SSME can be used today, with guaranteed performance, reliability, environmental requirements, acoustic output, etc. I'd guess this makes it easier to design the new vehicle.

Perhaps it's simplest to think of the SSME > RS-25e route as one of simplifying & cost saving, and the RS-68a > RS-68r route as higher turbopump pressure, performance upgrade & Human Rating. Both involve a channel-wall nozzle as stage 1.

BTW, DIRECT 1.0 used the ~430s Isp variant discussed at the very end of the post. They describe it as a high cost, long timeline programme.

I don't think anyone is downplaying the size of the RS-25e development programme, but it can be taken at whatever pace fits the budget. With RS-68r, nothing will fly until the programme is complete.

Shuttle was already moving towards a channel-wall nozzle, so this is an understood first major step to simplification.

cheers, Martin

PS IIRC DIRECT claim NASA identified the base heating issue on Ares V, not themselves.

Quote
Yet the findings of the DIRECT team prior to unveiling DIRECT 3.0 may put a wrench in NASA's plans. Their studies (and apparently NASA's internal studies confirm this) indicate the current RS-68 engines will not survive the extreme heating environment nestled between two SRB's.
Title: Re: Rocket Engine Q&A
Post by: kraisee on 01/16/2010 07:20 pm
PS IIRC DIRECT claim NASA identified the base heating issue on Ares V, not themselves.

That is correct.

Ross.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/17/2010 04:59 am
In reading a little about the history of channel wall here:
http://www.astronautix.com/engines/rd0120.htm

and here (could only get 1st page from google unfortunately):
http://pdf.aiaa.org/preview/CDReadyMJPC2005_1177/PV2005_4306.pdf

Question:
What makes Volvo's sandwich nozzle different from the channel/jacket developed in russia? Is it that it's laser welded?
Title: Re: Rocket Engine Q&A
Post by: Patchouli on 01/17/2010 09:16 pm
Following the discussion I guess the reason why regen on the Merlin 1C was so easy for Spacex was they must have had it designed into their engine from the start.

Or is Merlin 1C a completely different engine from Merlin 1A and 1B?

But I have to agree with clongton on the J2X I always wondered why they didn't just use two J2S engines.

But I guess Ares I  and the limited space inside it's inter-stage is why the J2X exists which it's self was originally supposed to use a SSME.

This makes the J2X kinda a band aid solution for the lack of an airstarted SSME.

I guess it could be like hopping up a VW type II engine to make it work in a Cessna 172.
Sure it would work but it's a poor replacement for a Lycoming O-360.
Title: Re: Rocket Engine Q&A
Post by: William Barton on 01/17/2010 09:33 pm
Something I'd still like to know is, what became of all the SSMEs that were built over the years? There are something like 17 flight-ready engines in inventory, and of course 6 have been destroyed in accidents. Were the other 40 or so scrapped, or do they still exist somewhere?
Title: Re: Rocket Engine Q&A
Post by: clongton on 01/17/2010 10:52 pm
Something I'd still like to know is, what became of all the SSMEs that were built over the years? There are something like 17 flight-ready engines in inventory, and of course 6 have been destroyed in accidents. Were the other 40 or so scrapped, or do they still exist somewhere?

They were recycled.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/18/2010 03:02 am
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

Is it fair to say that whatever engine has to be designed for the new HLV (RS-25E or RS-68R) they both have considerable hurdles? And are BOTH basically new engines?

RS-25E - after "simplifying" what performance will it really have and at what cost?
RS-68R - how much commonality will it really have and for how much?

Without some official documentation from PWR to see whats really possible it all seems like speculation & educated guesswork.
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/18/2010 03:10 am
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

It would reduce ISP
Title: Re: Rocket Engine Q&A
Post by: Propforce on 01/19/2010 03:39 pm
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

It would reduce ISP

How's so?
Title: Re: Rocket Engine Q&A
Post by: Downix on 01/19/2010 04:04 pm
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

Is it fair to say that whatever engine has to be designed for the new HLV (RS-25E or RS-68R) they both have considerable hurdles? And are BOTH basically new engines?

RS-25E - after "simplifying" what performance will it really have and at what cost?
RS-68R - how much commonality will it really have and for how much?

Without some official documentation from PWR to see whats really possible it all seems like speculation & educated guesswork.

There are a few known steps for the RS-25e.  Closing off access points for refurbishment/not adding them in the first place is one.  Replacing the current labor-intensive bell design with a newer/cheaper bell design is another.  These two changes would not change the performance of the system so I can see.

For the RS-68R however, you need to redesign the bell entirely, redesign the turbopump and fuel flow system entirely.  This adds in new variables into the equasion.
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/19/2010 05:19 pm
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

It would reduce ISP

How's so?

More use of fuel in a non propulsive manner.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/19/2010 07:07 pm
That brings up another question I had. I've seen the term "useable propellant" before. Percentage wise, roughly how much residual O2 & H2 are left in the tanks of modern hydrogen fueled rockets?
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/19/2010 07:17 pm
That brings up another question I had. I've seen the term "useable propellant" before. Percentage wise, roughly how much residual O2 & H2 are left in the tanks of modern hydrogen fueled rockets?

the LOX engine cutoff sensors for the Shuttle ET and Delta IV are in the LOX feedline coming down from the LOX in the top.

As for residuals, I believe 2% is too high
Title: Re: Rocket Engine Q&A
Post by: sdsds on 01/19/2010 07:42 pm
And what about incorporating supersonic film cooling to the RS-68 like the vulcain 2?

It would reduce ISP

How's so?

More use of fuel in a non propulsive manner.

The RS-68 is open cycle, so  isn't the propellant in question (which powered the turbomachinery) either exhausted externally or exhausted into the nozzle?  Doesn't exhausting into the nozzle improve performance?
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/19/2010 10:49 pm

The RS-68 is open cycle, so  isn't the propellant in question (which powered the turbomachinery) either exhausted externally or exhausted into the nozzle?  Doesn't exhausting into the nozzle improve performance?

A little but it is not being burned.  The fluid for film cooling is parasitic and reduces overall ISP.
Title: Re: Rocket Engine Q&A
Post by: Antares on 01/20/2010 03:22 am
My esteemed colleague, that is not universally true.  With the extreme low mass of H2, film cooling or running rich can increase Isp though it might decrease thrust.  The exact answer is dependent on MR (Tc), Pc and design altitude (area ratio).
Title: Re: Rocket Engine Q&A
Post by: Propforce on 01/20/2010 04:09 pm
My esteemed colleague, that is not universally true.  With the extreme low mass of H2, film cooling or running rich can increase Isp though it might decrease thrust.  The exact answer is dependent on MR (Tc), Pc and design altitude (area ratio).

Agree.  But in the RS-68 case, the MR shift in the MCC does not gain you much of Isp increase due to the relatively low nozzle expansion ratio.  It will be significantly better for upper stage engines like the RL10s and J-2X, etc.

For booster engines, one needs to review the thrust accounting of RS-68 vs. the F-1 and compare the thrust gain from GG-exhaust vs. injected back into nozzle.   Personally, I think we get a better thrust gain if re-injected back in the main nozzle.  But the current RS-68 uses the GG exhaust as roll-control which is useful for single engine application.  Not sure if RCN is needed for multi-engine application as in the case of Ares V.

For ablative nozzles such as the RS-68, the film cooling also slows down the nozzle erosion which further benefits the Isp.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/20/2010 07:22 pm
Appreciate the great comments. Could film cooling on ablative nozzles be one possible way to address base heating on Ares V? Or in combination with vanes/ducting/insulation mentioned in other threads?
Title: Re: Rocket Engine Q&A
Post by: clongton on 01/20/2010 07:27 pm
Appreciate the great comments. Could film cooling on ablative nozzles be one possible way to address base heating on Ares V? Or in combination with vanes/ducting/insulation mentioned in other threads?

No. The thermal environment is far too severe. That would be like putting a band-aid on IED wound.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 01/20/2010 08:40 pm
Appreciate the great comments. Could film cooling on ablative nozzles be one possible way to address base heating on Ares V? Or in combination with vanes/ducting/insulation mentioned in other threads?

No. The thermal environment is far too severe. That would be like putting a band-aid on IED wound.

Yup.  It's like drinking from a firehose.  The 'cheapest' solution is to spread ablative insulation on outside of RS-68.  The penalty is weight.  Those insulation can add up to hundreds of pound.

Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/21/2010 02:15 am
Would that be hundreds of pounds per engine? Even if it was say 1200 lbs for all engines isn't that relatively a small % of total payload? Or would it be way more than 1200?

Also, I'm surprised PWR hasn't come out with any kind of press release or document saying "here is what's possible engine wise for the HLV. We can do this for the 68 regen and this for the 25E."

They seem like the conspicuous absentee in the public debate on engine selection.
Or am I just super naive and they would never do this as it's highly proprietary?
They must have some excellent ideas for both approaches.
Title: Re: Rocket Engine Q&A
Post by: kraisee on 01/21/2010 02:30 am
The ablative nozzle on the RS-68 accounts for more than 65% of the total engine mass already.

Don't forget too, that if you make the nozzle significantly heavier, will have to re-design its mounting/supports or you will not longer have the required margins in the design.

Ross.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 01/21/2010 05:39 pm

Also, I'm surprised PWR hasn't come out with any kind of press release or document saying "here is what's possible engine wise for the HLV. We can do this for the 68 regen and this for the 25E."

They seem like the conspicuous absentee in the public debate on engine selection.
Or am I just super naive and they would never do this as it's highly proprietary?
They must have some excellent ideas for both approaches.

The bottom line is that NASA has not formerly define the "new" RS-68 requirements for Ares V/ HLVs.  The follow-up issue would be the cost & risk associated with a new RS-68"X" development.

The Air Force's RS-68"A" upgrade only goes so far to meet Air Force's mission requirements but not including NASA's new HLV requirements.  Wish list is long for the upgrade, but the budget is tight. 

Title: Re: Rocket Engine Q&A
Post by: kraisee on 01/21/2010 08:48 pm
They seem like the conspicuous absentee in the public debate on engine selection.
Or am I just super naive and they would never do this as it's highly proprietary?
They must have some excellent ideas for both approaches.

They make money both ways.

Why rock the boat by stirring up this debate any more than is already happening?

Ross.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/22/2010 01:51 am

What are the trade considerations for LOX/RP-1 rocket stages between fuel on the bottom (e.g. Atlas V, Falcon 1/9) or fuel on the top (e.g. Delta II)?

CG, feed duct length

What about loads?  If the heavier tank is further aft, then can't one get by with a somewhat flimsier and lighter structure?
Title: Re: Rocket Engine Q&A
Post by: JosephB on 01/22/2010 07:04 am
The bottom line is that NASA has not formerly define the "new" RS-68 requirements for Ares V/ HLVs.  The follow-up issue would be the cost & risk associated with a new RS-68"X" development.

The Air Force's RS-68"A" upgrade only goes so far to meet Air Force's mission requirements but not including NASA's new HLV requirements.  Wish list is long for the upgrade, but the budget is tight. 


So, if I read you correctly, be patient, see what the current administration has in mind and go from there? Should be interesting.

Edit: for those interested, TrueGrit also has some really informative posts on RS-68 & makes for good reading.

Title: Re: Rocket Engine Q&A
Post by: Propforce on 01/25/2010 04:54 pm

What are the trade considerations for LOX/RP-1 rocket stages between fuel on the bottom (e.g. Atlas V, Falcon 1/9) or fuel on the top (e.g. Delta II)?

CG, feed duct length

What about loads?  If the heavier tank is further aft, then can't one get by with a somewhat flimsier and lighter structure?

The loads you need to look at is the vehicle bending load throughout the flight, which will most likely be the primary design driver.  Having the Cg closer to the center of vehicle helps the vehicle bending moment, a combination of thrust and aero & vibroacoustic loads, thereby reduces the need to thickens structures for the primary load path.

Title: Re: Rocket Engine Q&A
Post by: DiggyCoxwell on 05/08/2010 07:19 pm
   
If an amateur rocketeer wants to refine
a cheap (retail supply) bulk source of H2O2,
that has a concentration of only 3 percent, can he or she
use the small difference between water and H2O2 freezing
points, and the large difference between their pure liquid densities,
to freeze separate the two substances and increase the final H202 concentration to between 50 - 70 percent?
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 05/08/2010 07:51 pm
   
If an amateur rocketeer wants to refine
a cheap (retail supply) bulk source of H2O2,
that has a concentration of only 3 percent, can he or she
use the small difference between water and H2O2 freezing
points, and the large difference between their pure liquid densities,
to freeze separate the two substances and increase the final H202 concentration to between 50 - 70 percent?
You can do that. There are places where you can get 50% H2O2, though sometimes you have to lie about what you use it for (assuming you're using it for rocketry), since there's been some ridiculous law suits that somehow managed to blame the H2O2 provider for the mistakes of the person experimenting with it.

The perfect place to find out more information on this is ARocket, an active email list for amateur/experimental/entrepreneurial rocketry, frequented by the likes of Armadillo, Masten, Unreasonable Rocket (Paul Breed), XCor, and lots of amateurs:
http://www.arocketry.net/forum.html

Look in the archives or just ask the list.
Title: Re: Rocket Engine Q&A
Post by: hop on 05/08/2010 09:49 pm
Worth pointing out that a fair number of people have lost property, limbs and lives attempting this kind of thing. Peroxide can be handled safely, but it needs to be treated with respect.

Since the "liquid bomber" plot, this sort of activity might also attract attention from the authorities.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 05/09/2010 12:17 am
Also, you will need A LOT of 3% peroxide to create a small amount of 50% peroxide.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 05/09/2010 06:04 am
In general, the Peroxide Religion loses most of its converts within a few years. On the surface, it seems like a great oxidizer, but in practice, selecting this "solution" tends to kill your project.

Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 05/09/2010 07:23 am
In general, the Peroxide Religion loses most of its converts within a few years. On the surface, it seems like a great oxidizer, but in practice, selecting this "solution" tends to kill your project.



Liquid Oxygen is also really cheap, and I would say more readily available than High Test Peroxide. Besides having higher performance.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 05/09/2010 05:13 pm
Also, concentrated peroxide is both explosive and toxic. LOX is neither.
Title: Re: Rocket Engine Q&A
Post by: kraisee on 05/09/2010 08:49 pm
Not as clean burning as LOX either.

Ross.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 05/13/2010 05:20 pm
With the interest in a reusable booster:
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2010/04/19/AW_04_19_2010_p30-219818.xml&headline=USAF%20Plans%20For%20Reusable%20Booster%20Development&channel=defense

I was wondering about the possibilities of a winged RS-84 booster.
http://www.astronautix.com/engines/rs84.htm

In particular, three variants:
A. One engine
B. Two engine
C. Three engine cluster (like shuttle)

The powerful RS-84 would help offset the penalties of wings, wheels. canards, robust framework, etc.

Would an interstage be practical for an in-line application (with say a J-2X upper) without interfering with any needed TPS on the nose?

For a twin booster config with a liquid core, would there be the same base heating environment as with shuttle solids? (assuming RS-68A core & 3 engine boosters)

And what about solid motors for approach & landing?

The price of engines aren't going to get cheaper with time. Just wondering. Thanks.
Title: Re: Rocket Engine Q&A
Post by: Antares on 05/13/2010 05:48 pm
It has to be an integrated look at life cycle costs.  SSMEs are too intricate to be cost effective, and an RS-84 probably would be too.  RS-68s are were very cheap for their capability because they were designed to be easy to put together, at least compared to other existing engines.  Engine prices could go down, especially relative to performance, if innovations in manufacturability are achieved.  That's where we should be focusing engine development work, IMEO.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 05/14/2010 05:43 am
Peroxide does have its advantages too. It is noncryogenic, very dense, can be used as a monopropellant, is excellent for gas generators and with a catalyst it is hypergolic with RP-1. It also isn't very toxic, you obviously shouldn't drink it but it isn't an environmental threat and you don't need fume hoods or anything. It biodegrades very quickly and it is even used to clean up after certain toxic spills.

Like everything else it is not a panacea, but it does have its supporters.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 05/14/2010 02:08 pm
http://en.wikipedia.org/wiki/Russian_submarine_Kursk_explosion (http://en.wikipedia.org/wiki/Russian_submarine_Kursk_explosion)

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Propforce on 05/14/2010 05:47 pm
It has to be an integrated look at life cycle costs.  SSMEs are too intricate to be cost effective, and an RS-84 probably would be too.  RS-68s are were very cheap for their capability because they were designed to be easy to put together, at least compared to other existing engines.  Engine prices could go down, especially relative to performance, if innovations in manufacturability are achieved.  That's where we should be focusing engine development work, IMEO.

Not sure the folks picked up the subtle hint on the RS-68 cost.  Why is this a surprise to the 'market place' when you consolidate two rocket factories into one?  The third one is acting as importer only these days.

The word "cost effective" is also highly subjective.  If you are a commercial launch provider, then it is the "effectiveness" of the ratio of your projected revenues over projected cost.  But it is not that obvious if you're government where the metrics are not so straight forward.

As far as hoping for engine price to go down, remember; selecting an engine for your launch vehicle is like a marriage.  The weight never goes down after a marriage, the requirements always creep up and so is the per unit price that's going to cost you.

Changing engine selection afterward is even more costly and complicated.  But I'd assume no elaboration here is necessary.

Title: Re: Rocket Engine Q&A
Post by: JosephB on 05/15/2010 04:40 am
It has to be an integrated look at life cycle costs.  SSMEs are too intricate to be cost effective, and an RS-84 probably would be too.  RS-68s are were very cheap for their capability because they were designed to be easy to put together, at least compared to other existing engines.  Engine prices could go down, especially relative to performance, if innovations in manufacturability are achieved.  That's where we should be focusing engine development work, IMEO.

Not sure the folks picked up the subtle hint on the RS-68 cost.  Why is this a surprise to the 'market place' when you consolidate two rocket factories into one?  The third one is acting as importer only these days.

The word "cost effective" is also highly subjective.  If you are a commercial launch provider, then it is the "effectiveness" of the ratio of your projected revenues over projected cost.  But it is not that obvious if you're government where the metrics are not so straight forward.

As far as hoping for engine price to go down, remember; selecting an engine for your launch vehicle is like a marriage.  The weight never goes down after a marriage, the requirements always creep up and so is the per unit price that's going to cost you.

Changing engine selection afterward is even more costly and complicated.  But I'd assume no elaboration here is necessary.



Thank you for the replies. The point to always strive for innovations in manufacturing was well taken. However, in the era of CAD/CAM together with decades of global engine manufacturing experience, I get the impression (from my laypersons armchair) that major innovations to drastically reduce engine cost are less likely than innovations to make them more serviceable for reuse (from a refurb cost perspective) especially with kerolox engines.

At the risk of further talking out of my keester (mainly to draw some expert thoughts on the recent interest in reusable boosters) I’d say that in 20 years or so the U.S. cost of professional & technical labor will make the idea of engine reuse even more desirable.
Granted the flyback booster structure will also take some labor, but in comparison to say three RS-84 engines that go inside? Life cycle cost estimates are helpful but in the end are still a bit of a crap shoot (as they are for any expensive long duration govt. program.)

The RS-84 seemed to make a good candidate for a few classes of LFBB’s based on engine count. Shoot it down or worth investigating?
Title: Re: Rocket Engine Q&A
Post by: Antares on 05/15/2010 06:19 am
As far as hoping for engine price to go down, remember; selecting an engine for your launch vehicle is like a marriage.  The weight never goes down after a marriage, the requirements always creep up and so is the per unit price that's going to cost you.

I don't care how much DARPA paid: that's the best thing ever posted on the internet.  A bargain at any price, doubly good for rocket scientists.

Looks like the DoD was getting NA$A help on both liquids and solids and didn't know it.  ('Course, NASA didn't know it either.)  Sigh, makes the solution a mite closer to intractable.  Requires more thought.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 05/16/2010 02:02 am
It has to be an integrated look at life cycle costs.  SSMEs are too intricate to be cost effective, and an RS-84 probably would be too.  RS-68s are were very cheap for their capability because they were designed to be easy to put together, at least compared to other existing engines.  Engine prices could go down, especially relative to performance, if innovations in manufacturability are achieved.  That's where we should be focusing engine development work, IMEO.

Not sure the folks picked up the subtle hint on the RS-68 cost.  Why is this a surprise to the 'market place' when you consolidate two rocket factories into one?  The third one is acting as importer only these days.

The word "cost effective" is also highly subjective.  If you are a commercial launch provider, then it is the "effectiveness" of the ratio of your projected revenues over projected cost.  But it is not that obvious if you're government where the metrics are not so straight forward.

As far as hoping for engine price to go down, remember; selecting an engine for your launch vehicle is like a marriage.  The weight never goes down after a marriage, the requirements always creep up and so is the per unit price that's going to cost you.

Changing engine selection afterward is even more costly and complicated.  But I'd assume no elaboration here is necessary.



Exception: if you can achieve economies of scale. This is doubtful in almost all cases, but applies to the Falcon 9.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 05/17/2010 06:39 pm
With the interest in a reusable booster:
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2010/04/19/AW_04_19_2010_p30-219818.xml&headline=USAF%20Plans%20For%20Reusable%20Booster%20Development&channel=defense

I was wondering about the possibilities of a winged RS-84 booster.
http://www.astronautix.com/engines/rs84.htm

In particular, three variants:
A. One engine
B. Two engine
C. Three engine cluster (like shuttle)

The powerful RS-84 would help offset the penalties of wings, wheels. canards, robust framework, etc.

Would an interstage be practical for an in-line application (with say a J-2X upper) without interfering with any needed TPS on the nose?

For a twin booster config with a liquid core, would there be the same base heating environment as with shuttle solids? (assuming RS-68A core & 3 engine boosters)

And what about solid motors for approach & landing?

The price of engines aren't going to get cheaper with time. Just wondering. Thanks.


Another question along these lines, could carbon fiber make up most of the non-tankage portion of such a booster & still be able to attach any needed TPS without much issue?

EDIT: Ran across something fairly recent on Hydrocarbon Boost:
http://www.space-travel.com/reports/Aerojet_And_Florida_Turbine_Technologies_To_Develop_NASA_New_Rocket_Engines_999.html
Also:
Title: Re: Rocket Engine Q&A
Post by: strangequark on 05/17/2010 07:35 pm
Not as clean burning as LOX either.

Ross.

Peroxide is cleaner burning than oxygen. The bulk of your exhaust product is water vapor.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 05/17/2010 10:15 pm
Not as clean burning as LOX either.

Ross.

Peroxide is cleaner burning than oxygen. The bulk of your exhaust product is water vapor.

I guess Ross is talking about burning a fuel with LOX or peroxide.

Don't you need to be using peroxide as a monopropellant to get mostly-water exhaust?

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Propforce on 05/18/2010 01:11 am

Looks like the DoD was getting NA$A help on both liquids and solids and didn't know it.  ('Course, NASA didn't know it either.)  Sigh, makes the solution a mite closer to intractable.  Requires more thought.

Well... NA$A is pretty clear on reasons why for the new LOX-Rich stage combustion cycle kerosene engine.  The only "heavy lift propulsion" missions identified so far are only the DoD and Commercial.

Speaking of which, Elon Musk was at MSFC the other day discussing about this topic.  I guess he too want to replace those 27 Merlin engines one day.
Title: Re: Rocket Engine Q&A
Post by: Antares on 05/18/2010 02:54 am
Well... NA$A is pretty clear on reasons why for the new LOX-Rich stage combustion cycle kerosene engine.  The only "heavy lift propulsion" missions identified so far are only the DoD and Commercial.

You did note that our friends in Utah/Minneapolis succeeded in having all references to liquid stricken from the RFI?  Yay.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 05/18/2010 02:48 pm


I guess Ross is talking about burning a fuel with LOX or peroxide.

Don't you need to be using peroxide as a monopropellant to get mostly-water exhaust?

cheers, Martin

Burning with RP-1 is still almost 60% water. Versus around 30% for Kerolox. Otherwise, the combustion products are pretty much identical. Peroxide plus organic gives a lower amount of CO2, CO, and anything else objectional. I mean, peroxide has it's issues, but cleanliness of its exhaust products would definitely not be one of them.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 05/18/2010 02:52 pm
Another advantage of peroxide I've seen mentioned is that it lends itself very well to full flow staged combustion when combined with catalytic decomposition. The decomposition temperature is sufficiently low that the oxygen-rich gas doesn't cause problems.

Peroxide can also apparently be used for regenerative cooling and cannot produce coking in that case since it doesn't contain any carbon.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 05/18/2010 04:30 pm
Another advantage of peroxide I've seen mentioned is that it lends itself very well to full flow staged combustion when combined with catalytic decomposition. The decomposition temperature is sufficiently low that the oxygen-rich gas doesn't cause problems.

The decomposition temperature for HTP is around 1200K, which should be comparable to ox-rich Kerolox engine preburners. The temp on those is a pretty controllable factor, too. You can make simpler staged combustion engines with peroxide, because you only have to run one propellant up to preburner pressures, but full flow would negate that. Unless you're using the term "full flow" in a manner with which I'm unfamiliar.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 05/18/2010 04:36 pm
The decomposition temperature for HTP is around 1200K, which should be comparable to ox-rich Kerolox engine preburners. The temp on those is a pretty controllable factor, too.

Maybe it's the absence of hot streaks or the presence of large quantities of water vapour and absence of carbon to mess things up.

Quote
You can make simpler staged combustion engines with peroxide, because you only have to run one propellant up to preburner pressures, but full flow would negate that. Unless you're using the term "full flow" in a manner with which I'm unfamiliar.

If so I'm probably using it incorrectly or even Incorrectly. :) I thought staged combustion was called full flow if all propellant was preburned or decomposed to drive the turbines and if all propellant was vaporised before entering the combustion chamber.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 05/18/2010 04:46 pm
Another potential advantage of peroxide that I've seen mentioned is that you can run a peroxide bipropellant engine in monopropellant mode to start up, thus reducing the risk of hard starts. If you google "Mark Ventura peroxide" you'll find lots of articles extolling the virtues of peroxide. It's good to remember there's no such thing as a panacea.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 05/18/2010 08:03 pm
The decomposition temperature for HTP is around 1200K, which should be comparable to ox-rich Kerolox engine preburners. The temp on those is a pretty controllable factor, too.

Maybe it's the absence of hot streaks or the presence of large quantities of water vapour and absence of carbon to mess things up.

Quote
You can make simpler staged combustion engines with peroxide, because you only have to run one propellant up to preburner pressures, but full flow would negate that. Unless you're using the term "full flow" in a manner with which I'm unfamiliar.

If so I'm probably using it incorrectly or even Incorrectly. :) I thought staged combustion was called full flow if all propellant was preburned or decomposed to drive the turbines and if all propellant was vaporised before entering the combustion chamber.

Full flow staged combustion means that there are 2 preburners, one fuel-rich and one oxidizer-rich.
Title: Re: Rocket Engine Q&A
Post by: 2552 on 05/28/2010 06:58 pm
What is the engine being test-fired in this picture? It's the picture on page 7 of the FY2011 Budget Overview PDF (http://www.nasa.gov/pdf/420990main_FY_201_%20Budget_Overview_1_Feb_2010.pdf).

Title: Re: Rocket Engine Q&A
Post by: ugordan on 05/28/2010 07:02 pm
RS-68.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/05/2010 08:16 am
If most hydrocarbon engines are running fuel rich, why isn't methanol (or hydrous ethanol, ammonia, or perhaps even water) blended into the fuel so that the disassociation products of the excess fuel (excess to stoichemetric) have lower molecular mass and ?presumably? higher isp?
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/05/2010 01:28 pm
If most hydrocarbon engines are running fuel rich, why isn't methanol (or hydrous ethanol, ammonia, or perhaps even water) blended into the fuel so that the disassociation products of the excess fuel (excess to stoichemetric) have lower molecular mass and ?presumably? higher isp?

Advantage is more likely offset by additional tankage and complexity
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/05/2010 07:08 pm
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.
Title: Re: Rocket Engine Q&A
Post by: clongton on 07/05/2010 07:28 pm
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.

Aerojet has done the work to demonstrate the feasibility of providing a TAN for the RS-68. I am not at liberty to discuss results except to say categorically that the results were very, very good. The application would be to have a LH2/LOX core, ground ignited, with RP-1/LOX burning in the nozzle to provide the same advantage as having side-mounted RP-1/LOX LRB's, however without the additional mass of the RP-1 engines. The RP-1 would be carried in side mounted tanks with the LOX being provided by the core LOX tank. When the RP-1 is depleted, the side-mounted tanks would be jettisoned just as if they were actual LRB stages when in reality they are only drop tanks. Meanwhile the LH2/LOX core continues upward on its own. Very efficient.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/05/2010 07:38 pm
Looking at the OP, the question was about blending fuels which would avoid plumbing complexity, provided the fuels will actually blend. Water will not blend with kerosene, but maybe anhydrous methanol would since it can also be blended with gasoline. I don't understand why this is though, since gasoline is nonpolar and methanol is polar.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 07/05/2010 09:28 pm
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.

Aerojet has done the work to demonstrate the feasibility of providing a TAN for the RS-68. I am not at liberty to discuss results except to say categorically that the results were very, very good. The application would be to have a LH2/LOX core, ground ignited, with RP-1/LOX burning in the nozzle to provide the same advantage as having side-mounted RP-1/LOX LRB's, however without the additional mass of the RP-1 engines. The RP-1 would be carried in side mounted tanks with the LOX being provided by the core LOX tank. When the RP-1 is depleted, the side-mounted tanks would be jettisoned just as if they were actual LRB stages when in reality they are only drop tanks. Meanwhile the LH2/LOX core continues upward on its own. Very efficient.

Does that also imply that TAN might be equally applicable to SSME?

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: beb on 07/05/2010 10:07 pm
Looking at the OP, the question was about blending fuels which would avoid plumbing complexity, provided the fuels will actually blend. Water will not blend with kerosene, but maybe anhydrous methanol would since it can also be blended with gasoline. I don't understand why this is though, since gasoline is nonpolar and methanol is polar.
Ethanol can be mixed with gas up to 15% Ethanol (or 85% or greater) Methanol probably has a lower limit on blending.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/05/2010 10:19 pm
I thought it was the other way round: M85 is 85% methanol with 15% gasoline. But you probably do have to have significantly more of one than the other, which is not a problem for the application we're talking about.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 07/06/2010 12:52 am
OK very basic question.
What does TAN stand for?
It's not in the acronyms menu. Thanks.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/06/2010 12:53 am
Thrust Augmented Nozzle.
Title: Re: Rocket Engine Q&A
Post by: clongton on 07/06/2010 01:28 am
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.

Aerojet has done the work to demonstrate the feasibility of providing a TAN for the RS-68. I am not at liberty to discuss results except to say categorically that the results were very, very good. The application would be to have a LH2/LOX core, ground ignited, with RP-1/LOX burning in the nozzle to provide the same advantage as having side-mounted RP-1/LOX LRB's, however without the additional mass of the RP-1 engines. The RP-1 would be carried in side mounted tanks with the LOX being provided by the core LOX tank. When the RP-1 is depleted, the side-mounted tanks would be jettisoned just as if they were actual LRB stages when in reality they are only drop tanks. Meanwhile the LH2/LOX core continues upward on its own. Very efficient.

Does that also imply that TAN might be equally applicable to SSME?

cheers, Martin

Theoretically yes, but that would be expensive to implement.
IMO it would probably cost less to create an engine from scratch than to modify the SSME for TAN.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 07/06/2010 01:51 pm
Ah, OK. Thanks.

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/06/2010 02:16 pm
Looking at the OP, the question was about blending fuels which would avoid plumbing complexity, provided the fuels will actually blend. Water will not blend with kerosene, but maybe anhydrous methanol would since it can also be blended with gasoline. I don't understand why this is though, since gasoline is nonpolar and methanol is polar.
Ethanol can be mixed with gas up to 15% Ethanol (or 85% or greater) Methanol probably has a lower limit on blending.

If the following (http://en.wikipedia.org/wiki/Rocket_propellant) I found on wikipedia is correct:

Quote
LOX/hydrocarbon rockets are run only somewhat rich (O/F mass ratio of 3  rather than stoichiometric of 3.4 to 4), because the  energy release per unit mass drops off quickly as the mixture ratio  deviates from stoichiometric. LOX/LH2 rockets are run very  rich (O/F mass ratio of 4 rather than stoichiometric

then F/O ratio stoichiometric is between .294 and .25  whereas fuel rich it is .33,  which means of those .33 between .08 and .04 is excess. which is roughly somewhere between 12.5% and 25% of excess. So maybe based on this v. simplistic reasoning, E10-E30 might work. 

Would it be a better coolant?

Title: Re: Rocket Engine Q&A
Post by: MP99 on 07/06/2010 04:34 pm
SSME has a ratio of 6 & J-2X 5.5, I believe.

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/07/2010 01:59 am
Quote
Quote from: beb on 07/05/2010 02:07 PM
Quote
LOX/hydrocarbon rockets are run only somewhat rich (O/F mass ratio of 3  rather than stoichiometric of 3.4 to 4), because the  energy release per unit mass drops off quickly as the mixture ratio  deviates from stoichiometric. LOX/LH2 rockets are run very  rich (O/F mass ratio of 4 rather than stoichiometric

then F/O ratio stoichiometric is between .294 and .25  whereas fuel rich it is .33,  which means of those .33 between .08 and .04 is excess. which is roughly somewhere between 12.5% and 25% of excess. So maybe based on this v. simplistic reasoning, E10-E30 might work. 

Would it be a better coolant?

Now look here, you're in the Rocket world here now.  When in Rome, do as the Romans do.  You'll have to forget about that airbreathing jargon and pick up new vocabulary. 

Starting with O/F ratio, not F/O ratio.  It's a simple conversion, sure, but it convey you know what you're talking about.  Just like when draw an airplane, it always fly from right to left.  Likewise, launch vehicle always fly from the bottom to top. 

I had to politely correct an Air Force Captain on her power point chart that, if her objective it to get to space, it doesn't help if her rocket is flying sideways on her charts!!  This way, it will NEVER get to space !  :)
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/07/2010 08:26 am
Quote
Quote from: beb on 07/05/2010 02:07 PM
Quote
LOX/hydrocarbon rockets are run only somewhat rich (O/F mass ratio of 3  rather than stoichiometric of 3.4 to 4), because the  energy release per unit mass drops off quickly as the mixture ratio  deviates from stoichiometric. LOX/LH2 rockets are run very  rich (O/F mass ratio of 4 rather than stoichiometric

then F/O ratio stoichiometric is between .294 and .25  whereas fuel rich it is .33,  which means of those .33 between .08 and .04 is excess. which is roughly somewhere between 12.5% and 25% of excess. So maybe based on this v. simplistic reasoning, E10-E30 might work. 

Would it be a better coolant?

Now look here, you're in the Rocket world here now.  When in Rome, do as the Romans do.  You'll have to forget about that airbreathing jargon and pick up new vocabulary. 

Starting with O/F ratio, not F/O ratio.  It's a simple conversion, sure, but it convey you know what you're talking about.  Just like when draw an airplane, it always fly from right to left.  Likewise, launch vehicle always fly from the bottom to top. 

I had to politely correct an Air Force Captain on her power point chart that, if her objective it to get to space, it doesn't help if her rocket is flying sideways on her charts!!  This way, it will NEVER get to space !  :)


I suspect this is not supposed to be taken too literally.. but the  straight roads, sanitation, hypocaust, use of concrete and arches, and  that a man (.. or woman!) could walk home safely on a dark night, I  like; the ritual slaughter of men (.. or women!) for entertainment, not so  much; but the use of R. numerals, or the ban on a temporary switch from  using 1/F to F for back of the wax seal calculations, is taking it too  far!  :)
 
 Perhaps the loss of energy taken by the disassociation of the alcohol  and/or reforming (assuming hydrous ethanol) of the fuels, more than  offsets any benefit of the increase (if this IS what would result?) in  the number of exhaust molecules (higher exhaust pressure?), and reduction in exhaust molecular weight?
 
 Any one know if this is the case?
 
Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/07/2010 01:30 pm
If most hydrocarbon engines are running fuel rich, why isn't methanol (or hydrous ethanol, ammonia, or perhaps even water) blended into the fuel so that the disassociation products of the excess fuel (excess to stoichemetric) have lower molecular mass and ?presumably? higher isp?

Rocket engine likes it simple but powerful, adding all these "extras" only complicates the engine design conditions.  Designing a rocket engine is a multidisciplinary process that must satisfy a wide array of design requirements, with safe & reliable operation among the prime objective.

The fuel used in rocket engine, either LH2 or HC, are used to cool the main combustion chamber and a part (or whole) of nozzle.  Those heating rate can be extremely high, with combustion temperature in excess of 6,000 deg. F and peak heat flux in the range of 100 Btu/sec-in^2.  The fuel must be able to cool this heat flux across a 0.1 inch thin layer of copper alloy and takes away that heat before the metal melts and the entire engine burns up.

With HC fuel, it is important that these fuel has a very high thermal stability properties and will not decompose, vaporize, gum up or coke up in the coolant channels.  Because once it does, a run-away heat transfer problem occur and we are looking at a rapid disintegration of rocket engine in front of our eyes ! Fuel is also used inside of combustion chamber as coolant via film cooling or transpiration cooling techniques, hence similar requirements exist.

All these "additives" you mentioned above have much lower thermal stability limits than the RP-1, our current HC propellant of choice.  As such, it will severely impact the way engine operates and its performance.

There are other ways of improving engine Isp performance such as improving engine cycles, etc. 
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 06:27 pm
Using a third cooling fluid (say water) would avoid those problems. In addition this would be much easier to pump than LH2, and much easier on the turbines.

From a recent NATO paper on High Speed Propulsion Cycles (http://ftp.rta.nato.int/public//PubFullText/RTO/EN/RTO-EN-AVT-150///EN-AVT-150-02.pdf):

Quote
Major benefits of the TFC technology applied to the LOX/LH2 engines are significantly higher (50-65%)
engine thrust-to-weight ratio and feasibility of higher combustor pressures. A preliminary comparison with
SSME shows that at the same combustor pressure, 34% of the structural weight saving can be expected.
Due to significantly lower cycle pressures, TFC technology may be a key to rocket engine reusability.
Other significant advantages include the fact that compared to the staged combustion cycles of the SSME
type, which utilize 3 combustion devices (2 preburners and a main combustor), the TFC eliminates 2 of
these 3 combustion devices with an accompanying weight savings and no loss in performance. As a
result, development cost and time savings can be expected. This is also true for the LOX/HC rocket
engines.
The TFC configuration is a promising choice for the LOX/HC engine because it permits:
• The elimination of the preburner and associated systems
• Low turbine temperature
• Low maximum cycle pressure (on the level of the chamber pressure).
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/07/2010 07:11 pm
Using a third cooling fluid (say water) would avoid those problems. In addition this would be much easier to pump than LH2, and much easier on the turbines.


XCOR has a patent on this.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 07:16 pm
XCOR has a patent on this.

Interesting. Do you have a link for this?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/07/2010 07:23 pm

Interesting. Do you have a link for this?

Hmm, I did, once upon a time read a paper on it. Can't seem to find it right now. Will dig later on. They do mention it briefly here  (http://www.xcor.com/products/pumps/index.html) at the bottom of the page:

Quote
XCOR’s proprietary piston pumps can pump like  turbopumps but are able to stop and start quickly.  They are also less expensive. A benefit of a piston-style design is that the pump is capable of pumping more fuel at a higher operational speed. Using this innovative pump design, drive gas to operate the pumps can be delivered by any of the three classical methods: staged combustion, gas generator, or expander. We have chosen a fourth, however, which is a proprietary thermodynamic cycle that is most similar to the expander. XCOR has patented this cycle, which has the advantage of not lowering the engine specific impulse, as a gas generator would.

They also mention it here (http://www.xcor.com/products/engines/2P1_N2O_ethane_rocket_engine.html):

Quote
The Tea Cart engine continues to support unique research.  In early 2006, XCOR and ATK GASL received a DARPA contract to "investigate, develop, and demonstrate a novel configuration for a liquid rocket engine, namely a Third Fluid Cooled (TFC) liquid rocket engine," for which we used the Tea Cart to generate superheated steam suitable for driving a Rankine cycle. This steam cycle will allow a turbopump system to develop high chamber pressure in a more efficient way than staged combustion cycles, thus improving the performance and durability for boost propulsion, as well as orbital transfer applications.  The Tea Cart engine was pressed well beyond its original design parameters to higher temperatures and pressures than it had ever seen before, and provided cost-effective proof that this new technology is feasible.

I'd bet your paper (by Dr. Balepin of ATK GASL) is related to the DARPA project mentioned, since ATK GASL was involved in that.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 07/07/2010 07:25 pm
In the TFC case why not use LOX instead?

1. It will already be available in almost any conceivable liquid engine negating the need for a third fluid.
2. It is chilled to a much lower temp than H2O meaning you do not have to waste as much mass to get the same cooling.
3. Besides most liquid engines burn fuel rich, meaning you can dump the coolant into the nozzle and get higher thrust boost than you would with plain H2O.

No need for a "third" liquid.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 07:27 pm
The third liquid would run in a closed cycle.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 07/07/2010 07:29 pm
The third liquid would run in a closed cycle.

So you need an additional heat exchanger... Interesting
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/07/2010 07:29 pm
The third liquid would run in a closed cycle.

And it allows you to create something like a Kerolox expander engine. Since you can choose your coolant pretty arbitrarily, and kerosene is way easier to pump than hydrogen, I bet you could push 150 klbs with that (educated guess only).
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 07:30 pm
I do remember reading something on NTRS about LOX cooling which is apparently not as fiendishly difficult as you might think. Quite an old paper I think.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 07:37 pm
I'd bet your paper (by Dr. Balepin of ATK GASL) is related to the DARPA project mentioned, since ATK GASL was involved in that.

Sounds like it! I found TFC interesting, I find XCOR interesting and it is even more interesting to see the connection...
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 07:43 pm
1. It will already be available in almost any conceivable liquid engine negating the need for a third fluid.

It would be interesting for bipropellant engines with noncryogenic oxidisers, maybe peroxide, maybe HAN.
Title: Re: Rocket Engine Q&A
Post by: alexw on 07/07/2010 08:09 pm
In the TFC case why not use LOX instead?
1. It will already be available in almost any conceivable liquid engine negating the need for a third fluid.
2. It is chilled to a much lower temp than H2O meaning you do not have to waste as much mass to get the same cooling.
     The specific heat capacity of LOX is much less than water; will the total heat transfer capacity really be greater given the lower temperature? Might need to know the design temperatures and pressures inside the nozzle and integrate the heat capacity curves over the full temperature range.
        -Alex
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 08:20 pm
The LOX would likely still be used for cooling the water in the heat exchanger.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/07/2010 10:10 pm
If most hydrocarbon engines are running fuel rich, why isn't methanol (or hydrous ethanol, ammonia, or perhaps even water) blended into the fuel so that the disassociation products of the excess fuel (excess to stoichemetric) have lower molecular mass and ?presumably? higher isp?

Rocket engine likes it simple but powerful, adding all these "extras" only complicates the engine design conditions.  Designing a rocket engine is a multidisciplinary process that must satisfy a wide array of design requirements, with safe & reliable operation among the prime objective.

The fuel used in rocket engine, either LH2 or HC, are used to cool the main combustion chamber and a part (or whole) of nozzle.  Those heating rate can be extremely high, with combustion temperature in excess of 6,000 deg. F and peak heat flux in the range of 100 Btu/sec-in^2.  The fuel must be able to cool this heat flux across a 0.1 inch thin layer of copper alloy and takes away that heat before the metal melts and the entire engine burns up.

With HC fuel, it is important that these fuel has a very high thermal stability properties and will not decompose, vaporize, gum up or coke up in the coolant channels.  Because once it does, a run-away heat transfer problem occur and we are looking at a rapid disintegration of rocket engine in front of our eyes ! Fuel is also used inside of combustion chamber as coolant via film cooling or transpiration cooling techniques, hence similar requirements exist.

All these "additives" you mentioned above have much lower thermal stability limits than the RP-1, our current HC propellant of choice.  As such, it will severely impact the way engine operates and its performance.

There are other ways of improving engine Isp performance such as improving engine cycles, etc. 

Thanks for explaining this. I read from this that its the phase change resulting from the disassociation of the additives that occurs in the cooling channels that would be the main problem?

Why is bubbly coolant a problem, will it lead to uneven fuel flow or uneven cooling, or something else?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/07/2010 10:51 pm
If the bubbles are small (nucleate boiling) it's not a problem at all, this is the most effective cooling. Large bubbles are a problem, except with hydrogen which is a good coolant even in gaseous form.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/07/2010 11:44 pm
If most hydrocarbon engines are running fuel rich, why isn't methanol (or hydrous ethanol, ammonia, or perhaps even water) blended into the fuel so that the disassociation products of the excess fuel (excess to stoichemetric) have lower molecular mass and ?presumably? higher isp?

Rocket engine likes it simple but powerful, adding all these "extras" only complicates the engine design conditions.  Designing a rocket engine is a multidisciplinary process that must satisfy a wide array of design requirements, with safe & reliable operation among the prime objective.

The fuel used in rocket engine, either LH2 or HC, are used to cool the main combustion chamber and a part (or whole) of nozzle.  Those heating rate can be extremely high, with combustion temperature in excess of 6,000 deg. F and peak heat flux in the range of 100 Btu/sec-in^2.  The fuel must be able to cool this heat flux across a 0.1 inch thin layer of copper alloy and takes away that heat before the metal melts and the entire engine burns up.

With HC fuel, it is important that these fuel has a very high thermal stability properties and will not decompose, vaporize, gum up or coke up in the coolant channels.  Because once it does, a run-away heat transfer problem occur and we are looking at a rapid disintegration of rocket engine in front of our eyes ! Fuel is also used inside of combustion chamber as coolant via film cooling or transpiration cooling techniques, hence similar requirements exist.

All these "additives" you mentioned above have much lower thermal stability limits than the RP-1, our current HC propellant of choice.  As such, it will severely impact the way engine operates and its performance.

There are other ways of improving engine Isp performance such as improving engine cycles, etc. 

Thanks for explaining this. I read from this that its the phase change resulting from the disassociation of the additives that occurs in the cooling channels that would be the main problem?

Yes it would be.  Another consideration is that the dissociation of both fuel & additive may promote free radical and, when combined with the wall catalytic effect, they would further accelerate the decomposition reaction.

Another big issue with the RP-1, being a refined fuel from crude oil, is the existence of contaminants in the refined fuel mixture itself.  Sulfur is the biggest issue.  Thanks to the modern day refinery technology, its content has significantly dropped.  This is one few area where commercial industry has out-paced the government specification.  They actually deliver CLEANER fuel than what the MIL-SPEC specifies. 


Quote
Why is bubbly coolant a problem, will it lead to uneven fuel flow or uneven cooling, or something else?

Bubbly coolant will effectively form an insulation layer of "film" thereby blocking the heat transfer between the coolant and the wall.  Without the forced convection cooling from the coolant drawing heat away from the wall, the wall temperature will climb in a matter of seconds, to effectively melt away the wall (since it is only ~0.1 inch thick copper alloy).  You can tell this by the greenish flame coming out of combustor.  It's pretty to see, but it doesn't make the management happy.

IF you are brave and are able to control the degree of boiling to within a range called "subcooled nucleate boiling", the it is like watching a pot of water boil.  You'll first see little bubbles formed at the wall, it rises then is quickly cooled by the bulk temperature of water in the pot.  But if you are not careful, with continue increasing in heating, these bubbles quickly overwhelm the liquids and soon you'll have a "film" of gas bubbles on the wall.  At the high heat flux regime, you'll burn up the wall in no time.

Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/08/2010 08:31 am
If the bubbles are small (nucleate boiling) it's not a problem at all, this is the most effective cooling. Large bubbles are a problem, except with hydrogen which is a good coolant even in gaseous form.
Again thanks for taking the time to explain this.

Here is the link to a relevant section (http://en.wikipedia.org/wiki/Regenerative_cooling_%28rocket%29#Heat_flow_and_temperature) of a wikipedia page.

From this, is it correct to assume that coolant boiling and disassociation is more of a problem in fuels that are mixtures of substances with different boiling points because the fraction with the lower boiling point can remain much longer in a gaseous phase while still in thermal equilibrium with other substances in the mixture which are in liquid phase. Whereas, a bubble of steam or ammonia surrounded by water or liquid ammonia will continue to cool as it vaporizes the surrounding liquid until it finally collapses back into the liquid phase, and this is why it is okay to use water in third fluid cooling systems or pure ammonia as a coolant in an engine like the xlr99, which was used in the x-15.
Title: Re: Rocket Engine Q&A
Post by: JosephB on 07/16/2010 07:04 pm
I saw this article on the fourth & final test of CECE:
http://www.nasa.gov/centers/marshall/news/news/releases/2010/10-081.html

104% down to 5.9%? Sounds like she's ready for a mission!
I was wondering how that compared to the LEM?


EDIT: found it:
http://en.wikipedia.org/wiki/Lunar_Excursion_Module

To the CECE folks: Well done, great work!
Title: Re: Rocket Engine Q&A
Post by: jongoff on 07/16/2010 08:33 pm
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.

Aerojet has done the work to demonstrate the feasibility of providing a TAN for the RS-68. I am not at liberty to discuss results except to say categorically that the results were very, very good. The application would be to have a LH2/LOX core, ground ignited, with RP-1/LOX burning in the nozzle to provide the same advantage as having side-mounted RP-1/LOX LRB's, however without the additional mass of the RP-1 engines. The RP-1 would be carried in side mounted tanks with the LOX being provided by the core LOX tank. When the RP-1 is depleted, the side-mounted tanks would be jettisoned just as if they were actual LRB stages when in reality they are only drop tanks. Meanwhile the LH2/LOX core continues upward on its own. Very efficient.

Another variation on the theme that I remember hearing about was to take the intertank volume on the ET and convert it to a Kerosene tank, with the LOX tank getting a stretch.  You wouldn't need to actually pressurize the kero tank area, because the lower suction head requirements for the kero (IIRC it was only something like 3psi for the turbopumps on the NK-33) could be entirely provided by the gravity head from the tank down to the engines.  Gives you both LOX/Kero and LOX/LH2 in the same stage.  Really high thrust for liftoff, while allowing really good vacuum Isp once the afterburning is off.

I haven't run this specific scenario using my afterburning rocket calculation tool, but my guess is that you could get a vacuum Isp in the ~440s range (with a larger expansion bell than the existing RS-68), while still having the liftoff nozzle exit pressure in happy ranges, and thrusts in the 1.5-2Mlbf range.  Of course, you'd probably have to beef the engine and thrust structure up a bit if you did that.  But my guess is in the end you'd have an engine with a T/W north of 100:1......

I'm tempted to run the numbers just for pure nerdlichkeit.

You have any idea how much fun you could have with a rocket engine like that?


~Jon
Title: Re: Rocket Engine Q&A
Post by: alexw on 07/16/2010 09:14 pm
Since RS-68 TAN seems unlikely if we now build an SSME SDHLV, what about its use on an upper stage? J-2X is less well suited for TLI/TMI burn on account of the dry mass, but RL-10 calls for a 4 to 7 engine cluster to reduce gravity losses during the burn to LEO. Can you speak about the prospect of TAN'ing RL-10?
      -Alex
Title: Re: Rocket Engine Q&A
Post by: jongoff on 07/16/2010 10:36 pm
   Jon's point is well taken. "Everybody" loves the idea of TAN (see for example all the dreamy statements about TAN+SSME) yet it also seems quite popular around here that we're going to eviscerate the proposed kerolox engine applied research, and that we need to "stop waiting" and "get on with flying". People need to acknowledge that we are making tradeoffs, and those have consequences.

That was the point I was trying to make.  There are real costs to all decisions, and often the most painful costs aren't readily visible because they're the opportunity costs of futures we've foregone.  The good news is that even if NASA ignores technologies like this (which isn't actually a 100% certainty), DoD is working on projects like RBS which benefit even more from TAN than a normal booster would.   

Quote
Jon, to return to topic: since RS-68 TAN seems unlikely if we now build an SSME SDHLV, what about its use on an upper stage? J-2X is less well suited for TLI/TMI burn on account of the dry mass, but RL-10 calls for a 4 to 7 engine cluster to reduce gravity losses during the burn to LEO. Can you speak about the prospect of TAN'ing RL-10?

I've run some numbers.  You end up taking a decent Isp hit while running the TAN system, which means that you'd want to mostly use it for parts of the burn where the gain in gravity losses avoided outweighs the penalty in less efficient combustion.  FWIW, I was getting an RL-10B2 running at around 50klbf and ~435s vacuum Isp.  You could actually boost the thrust quite a bit more than that, since the RL-10B2's nozzle is so long, but the Isp penalties start getting high.  Might make sense for an US working with propellant depots.  Use the really high thrust and vehicle T/W on ascent to LEO to minimize gravity losses.  Once you've refueled though, run the thing with the TAN flow off for in-space burns where the Isp matters more than the T/W ratio. 

Not as big of a slam-dunk though as it is for booster engines.  Like most good things in engineering, you can overapply it...

~Jon
Title: Re: Rocket Engine Q&A
Post by: MP99 on 07/17/2010 09:52 am
since RS-68 TAN seems unlikely if we now build an SSME SDHLV, what about its use on an upper stage? J-2X is less well suited for TLI/TMI burn on account of the dry mass, but RL-10 calls for a 4 to 7 engine cluster to reduce gravity losses during the burn to LEO. Can you speak about the prospect of TAN'ing RL-10?

I've run some numbers.  You end up taking a decent Isp hit while running the TAN system, which means that you'd want to mostly use it for parts of the burn where the gain in gravity losses avoided outweighs the penalty in less efficient combustion.  FWIW, I was getting an RL-10B2 running at around 50klbf and ~435s vacuum Isp.  You could actually boost the thrust quite a bit more than that, since the RL-10B2's nozzle is so long, but the Isp penalties start getting high.  Might make sense for an US working with propellant depots.  Use the really high thrust and vehicle T/W on ascent to LEO to minimize gravity losses.  Once you've refueled though, run the thing with the TAN flow off for in-space burns where the Isp matters more than the T/W ratio. 

Not as big of a slam-dunk though as it is for booster engines.  Like most good things in engineering, you can overapply it...

DEC increases Atlas payload to LEO by overcoming gravity losses, while SEC maximises GTO payload.

It seems that a TAN'd SEC could have the thrust to equal DEC performance to LEO, balancing lower Isp against lower burnout mass. Indeed, you'd only want to apply TAN for the early part of the ascent, switching to a pure-SEC, high Isp config to maximise payload to orbit.

I even wonder if this would use less prop achieving orbit, so increasing SEC's GTO performance. For DEC-type LEO missions, you'd probably need to apply TAN for longer to overcome the greater gravity losses.

Would the TAN on an RL-10 run with H2/O2, or does it also require kero as per the booster concept?

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: jongoff on 07/17/2010 07:58 pm
DEC increases Atlas payload to LEO by overcoming gravity losses, while SEC maximises GTO payload.

It seems that a TAN'd SEC could have the thrust to equal DEC performance to LEO, balancing lower Isp against lower burnout mass. Indeed, you'd only want to apply TAN for the early part of the ascent, switching to a pure-SEC, high Isp config to maximise payload to orbit.

I even wonder if this would use less prop achieving orbit, so increasing SEC's GTO performance. For DEC-type LEO missions, you'd probably need to apply TAN for longer to overcome the greater gravity losses.

Quite possibly.  Without a good trajectory optimizer (I'll get one yet), it's hard to do much more than handwave.  But it sounds like doing an RL-10 equivalent engine with thrust augmentation for the first part of the burn might give you a net performance boost for GTO missions.  But it would probably have a bigger impact for LEO missions.

Quote
Would the TAN on an RL-10 run with H2/O2, or does it also require kero as per the booster concept?

TAN can run with any propellant combination that burns well.  It turns out LOX/Kero is one of the hardest TAN propellant combos to make work.  LOX/LH2 (or even better yet LOX/GH2 or supercritical H2) is much easier to mix and combust.  I think Aerojet did GOX/GH2 for their first TAN engine, and LOX/Kero for their second.  IIRC.

~Jon
Title: Re: Rocket Engine Q&A
Post by: clongton on 07/18/2010 11:59 am
I haven't run this specific scenario using my afterburning rocket calculation tool, but my guess is that you could get a vacuum Isp in the ~440s range (with a larger expansion bell than the existing RS-68), while still having the liftoff nozzle exit pressure in happy ranges, and thrusts in the 1.5-2Mlbf range.  Of course, you'd probably have to beef the engine and thrust structure up a bit if you did that.  But my guess is in the end you'd have an engine with a T/W north of 100:1......

I'm tempted to run the numbers just for pure nerdlichkeit.

You have any idea how much fun you could have with a rocket engine like that?


~Jon

NerdAlert! (me of course :) )
I would dearly love to see the RS-68 with TAN nozzle on the Delta with Kero drop tanks. O-M-G!!
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/18/2010 10:39 pm
I'm wondering about how much more expensive expander cycle engines are than pressure-fed engines and why that is. To be concrete compare a LOX/LCH4 expander cycle engine similar to the RL-10 to a pressure-fed engine with similar thrust. Two questions:

1. IIRC RL-10s supposedly cost a few million dollars a piece. How much cheaper would a similar pressure-fed engine be? I'm looking for a ballpark figure, i.e. is an expander cycle closer to 30% or 1000% more expensive than pressure-fed?

2. How much do each of the following factors account for the increased cost of expander cycle engines?
* The Turbopump
* Increased use of regenerative cooling
* Increased working pressure

I'm looking for ballpark but quantitative information.

Thanks.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/19/2010 01:59 am
Jon, to return to topic: since RS-68 TAN seems unlikely if we now build an SSME SDHLV, what about its use on an upper stage? J-2X is less well suited for TLI/TMI burn on account of the dry mass, but RL-10 calls for a 4 to 7 engine cluster to reduce gravity losses during the burn to LEO. Can you speak about the prospect of TAN'ing RL-10?

I've run some numbers.  You end up taking a decent Isp hit while running the TAN system, which means that you'd want to mostly use it for parts of the burn where the gain in gravity losses avoided outweighs the penalty in less efficient combustion.  FWIW, I was getting an RL-10B2 running at around 50klbf and ~435s vacuum Isp.  ......
~Jon

You managed to take a 462 sec Isp engine and drop it down to ~435 sec with TAN concept.  The RL10B-2 is already designed for optimum vacuum performance, you will not gain much appreciable Isp benefit with a bigger nozzle.

Just remember, thrust is not king and neither is engine Isp.  It is the overall integrated mission performance for the cost of rocket.  Delta IV upper stage with RL10B-2 engine was designed to optimize performance to GTO and not to LEO, but it will do LEO mission just the same.

The real benefit of TAN concept is altitude compensation with afterburner boost capability.  When you are in an Upper Stage flight regime, there's really isn't much back pressure to worry about altitude compensation, so the concept is really self-defeating for an Upper Stage engine.

Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/19/2010 02:15 am
I'm wondering about how much more expensive expander cycle engines are than pressure-fed engines and why that is. To be concrete compare a LOX/LCH4 expander cycle engine similar to the RL-10 to a pressure-fed engine with similar thrust. Two questions:

1. IIRC RL-10s supposedly cost a few million dollars a piece. How much cheaper would a similar pressure-fed engine be? I'm looking for a ballpark figure, i.e. is an expander cycle closer to 30% or 1000% more expensive than pressure-fed?

2. How much do each of the following factors account for the increased cost of expander cycle engines?
* The Turbopump
* Increased use of regenerative cooling
* Increased working pressure

I'm looking for ballpark but quantitative information.

Thanks.

You are not approaching the issue correctly.  You are focusing on the manufacturing cost of engine itself, but the comparison really goes to your vehicle/engine as a "system". 

RL10 engine has a chamber pressure (Pc) at ~600 psia.  To obtain the same pressure for your equivalent pressure-fed engine, you must add

1) Additional pressure for pressure drop because of regenerative cooling in engine nozzle and combustion chamber

2) Additional pressure for pressure drop for across the injector plate

3) Additional pressure for pressure drop for line losses from the tank to the engine

So this puts your tank pressure at around 750~800 psia.  Now calculate the additional wall thickness to be able to withstand a 800 psia pressure.  Now add a factor of safety of additional 20%...

Now as you "fly" this stage with pressure-fed engine, watch your stage drops like a rock into the ocean because your tank is so heavy that your engine can not support the weight which it needs to fly....

Tell me how much $ are you saving by dropping payload in the ocean?




Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/19/2010 02:19 am
I haven't run this specific scenario using my afterburning rocket calculation tool, but my guess is that you could get a vacuum Isp in the ~440s range (with a larger expansion bell than the existing RS-68), while still having the liftoff nozzle exit pressure in happy ranges, and thrusts in the 1.5-2Mlbf range.  Of course, you'd probably have to beef the engine and thrust structure up a bit if you did that.  But my guess is in the end you'd have an engine with a T/W north of 100:1......

I'm tempted to run the numbers just for pure nerdlichkeit.

You have any idea how much fun you could have with a rocket engine like that?


~Jon

NerdAlert! (me of course :) )
I would dearly love to see the RS-68 with TAN nozzle on the Delta with Kero drop tanks. O-M-G!!

O-M-G is right as you see the Delta IV engulfed by smokes but can't take off from the ground.
Title: Re: Rocket Engine Q&A
Post by: jongoff on 07/19/2010 02:03 pm
Jon, to return to topic: since RS-68 TAN seems unlikely if we now build an SSME SDHLV, what about its use on an upper stage? J-2X is less well suited for TLI/TMI burn on account of the dry mass, but RL-10 calls for a 4 to 7 engine cluster to reduce gravity losses during the burn to LEO. Can you speak about the prospect of TAN'ing RL-10?

I've run some numbers.  You end up taking a decent Isp hit while running the TAN system, which means that you'd want to mostly use it for parts of the burn where the gain in gravity losses avoided outweighs the penalty in less efficient combustion.  FWIW, I was getting an RL-10B2 running at around 50klbf and ~435s vacuum Isp.  ......
~Jon

You managed to take a 462 sec Isp engine and drop it down to ~435 sec with TAN concept.  The RL10B-2 is already designed for optimum vacuum performance, you will not gain much appreciable Isp benefit with a bigger nozzle.

Just remember, thrust is not king and neither is engine Isp.  It is the overall integrated mission performance for the cost of rocket.  Delta IV upper stage with RL10B-2 engine was designed to optimize performance to GTO and not to LEO, but it will do LEO mission just the same.

The real benefit of TAN concept is altitude compensation with afterburner boost capability.  When you are in an Upper Stage flight regime, there's really isn't much back pressure to worry about altitude compensation, so the concept is really self-defeating for an Upper Stage engine.

Propforce, yeah TAN is really a much better deal for first stage engines.  That said, it's still possible, especially for very low T/W ratio upper stages, that at certain parts of the burn, having more thrust might outweigh the losses.  Hard to tell without an actual trajectory simulator.  But it is clear, that it's nowhere near as good of a fit for upper stage applications than first stage ones.

~Jon
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/20/2010 02:58 am
You are not approaching the issue correctly.  You are focusing on the manufacturing cost of engine itself, but the comparison really goes to your vehicle/engine as a "system". 

RL10 engine has a chamber pressure (Pc) at ~600 psia.  To obtain the same pressure for your equivalent pressure-fed engine, you must add

1) Additional pressure for pressure drop because of regenerative cooling in engine nozzle and combustion chamber

2) Additional pressure for pressure drop for across the injector plate

3) Additional pressure for pressure drop for line losses from the tank to the engine

So this puts your tank pressure at around 750~800 psia.  Now calculate the additional wall thickness to be able to withstand a 800 psia pressure.  Now add a factor of safety of additional 20%...

Now as you "fly" this stage with pressure-fed engine, watch your stage drops like a rock into the ocean because your tank is so heavy that your engine can not support the weight which it needs to fly....

Tell me how much $ are you saving by dropping payload in the ocean?

Thanks for your reply.

Sorry I didn't ask my question clearly enough. I didn't mean that the hypothetical pressure-fed engine should be identical to the RL-10 except for the pressurization method. The hypothetical pressure-fed engine should have parameters typical of optimized pressure-fed engines, such as a more modest chamber pressure of 150 psi or so, and somewhat lower performance. The only commonality is the propellants and thrust.

BTW I am not proposing that the RL-10 be replaced one for one with a pressure-fed engine. What I'm trying to do is get a data point to aid my quantitative understanding of the cost side of rocketry. The internet is loaded with information on thrust, mass and ISP for various engines but cost data is very sparse, hence my asking.
Title: Re: Rocket Engine Q&A
Post by: Antares on 07/20/2010 05:57 am
You're still going to have to increase the pressure in the tank by a factor of 5 or 6.

Cost is driven by touch-labor - i.e. design/manufacturing complexity - and flight rate - how many flights you amortize fixed costs over.  Today's ELV fixed costs are the same or lower than they used to be, but there are fewer flights.  Things look more expensive per flight, but they kind of aren't.  The biggest cost on older engines is the hand-work on the combustion chamber tubes.  That is, happily, going extinct.

Sorry, but it's really hard to parameterize costs in low-production systems.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/20/2010 08:39 pm
Sorry, but it's really hard to parameterize costs in low-production systems.
Oh well. Thanks for your help.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 07/20/2010 08:51 pm
You are not approaching the issue correctly.  You are focusing on the manufacturing cost of engine itself, but the comparison really goes to your vehicle/engine as a "system". 

RL10 engine has a chamber pressure (Pc) at ~600 psia.  To obtain the same pressure for your equivalent pressure-fed engine, you must add

1) Additional pressure for pressure drop because of regenerative cooling in engine nozzle and combustion chamber

2) Additional pressure for pressure drop for across the injector plate

3) Additional pressure for pressure drop for line losses from the tank to the engine

So this puts your tank pressure at around 750~800 psia.  Now calculate the additional wall thickness to be able to withstand a 800 psia pressure.  Now add a factor of safety of additional 20%...

Now as you "fly" this stage with pressure-fed engine, watch your stage drops like a rock into the ocean because your tank is so heavy that your engine can not support the weight which it needs to fly....

Tell me how much $ are you saving by dropping payload in the ocean?

Thanks for your reply.

Sorry I didn't ask my question clearly enough. I didn't mean that the hypothetical pressure-fed engine should be identical to the RL-10 except for the pressurization method. The hypothetical pressure-fed engine should have parameters typical of optimized pressure-fed engines, such as a more modest chamber pressure of 150 psi or so, and somewhat lower performance. The only commonality is the propellants and thrust.

BTW I am not proposing that the RL-10 be replaced one for one with a pressure-fed engine. What I'm trying to do is get a data point to aid my quantitative understanding of the cost side of rocketry. The internet is loaded with information on thrust, mass and ISP for various engines but cost data is very sparse, hence my asking.


Did you change your original question?
Quote
Quote from: deltaV on 07/18/2010 02:39 PM

    I'm wondering about how much more expensive expander cycle engines are than pressure-fed engines and why that is. To be concrete compare a LOX/LCH4 expander cycle engine similar to the RL-10 to a pressure-fed engine with similar thrust. Two questions:

    1. IIRC RL-10s supposedly cost a few million dollars a piece. How much cheaper would a similar pressure-fed engine be? I'm looking for a ballpark figure, i.e. is an expander cycle closer to 30% or 1000% more expensive than pressure-fed?

    2. How much do each of the following factors account for the increased cost of expander cycle engines?
    * The Turbopump
    * Increased use of regenerative cooling
    * Increased working pressure

    I'm looking for ballpark but quantitative information.

    Thanks.

If you just need a pressure-fed engine, by itself, at a Pc ~ 150 psia.  Now that's a much simplified engine.

* You don't need REGEN cooling because the heat flux is very low

* You can use ablative/ heat sink for chamber/ throat/ nozzle cooling

But then again,

1) What is the thrust level requirement?
2) What application do you want to you them in?
3) Does it have to be flight-weight?
4) What comparable vehicle will this engine fly on?

Depending on your answers, the cost can go all over the map.

For starter, you may want to visit Microcosm's scorpious <sp?> launch vehicle design.  It is an all pressure-fed propulsion system
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/21/2010 01:29 am
Did you change your original question?

My original question was ambiguous so it's hard to say whether or not I changed it.

Quote
1) What is the thrust level requirement?
2) What application do you want to you them in?
3) Does it have to be flight-weight?
4) What comparable vehicle will this engine fly on?

Depending on your answers, the cost can go all over the map.

For starter, you may want to visit Microcosm's scorpious <sp?> launch vehicle design.  It is an all pressure-fed propulsion system

Thanks for your pointers.

Answer to your questions about my question:
1) Thurst level: in the 20 klbf class, like RL-10.
2) A two-stage to orbit LEO vehicle designed for low-value splittable cargo such as space station consumables and LOX for prop depots. Second stage would be expendable and use the engine in question. The first stage may be air launched out the back of a C-17 and might be reusable.
3) Yes.
4) How is this question different from (2)?

Thanks for the pointer to Scorpius. For other people reading this, here's some info I found on that vehicle:
http://www.smad.com/scorpius/index.html
http://wikileaks.org/wiki/Development_of_the_Scorpius_Rocket_Engine,_2005
.

Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/22/2010 02:47 pm
Here's another (unrelated) question. How obnoxious is high temperature and pressure gaseous oxygen? In the neighborhood of 40 atmospheres, 600 Kelvin. Is it easy to make lightweight valves, heat exchangers, and so on that don't burn in that environment? (For example compare to the difficulty of making LH2-tolerant items.) Or is such an oxidizing environment best avoided?

Would a say 50/50 mix (by moles) of oxygen and steam be significantly easier to design for?
Title: Re: Rocket Engine Q&A
Post by: TyMoore on 07/22/2010 03:03 pm
Ultra-high chromium stainless steel (0.3-0.4) ought to do nicely in a high pressure oxygen environment, as long as you 'pickle' it first by pressure aging it in presence of oxygen...

The same kinds of alloys are used in various internal engine components of the Energomash RD-170/180/190 series engines.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/22/2010 11:18 pm
Ultra-high chromium stainless steel (0.3-0.4) ought to do nicely in a high pressure oxygen environment, as long as you 'pickle' it first by pressure aging it in presence of oxygen...

The same kinds of alloys are used in various internal engine components of the Energomash RD-170/180/190 series engines.

Thanks. Too bad steel's strength-to-weight ratio leaves something to be desired.
Title: Re: Rocket Engine Q&A
Post by: TyMoore on 07/23/2010 12:34 am
If it's thick enough, it should be strong enough. The high thrust to weight ratio of high pressure kerosene staged combustion rocket engines makes up for the thickness of the pressure lines and combustion chambers, I would think.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/23/2010 02:29 am
If it's thick enough, it should be strong enough. The high thrust to weight ratio of high pressure kerosene staged combustion rocket engines makes up for the thickness of the pressure lines and combustion chambers, I would think.

I wonder if there's some easy way to coat a stronger metal such as titanium with a thin oxygen-compatible layer. Something vaguely similar to galvanization of steel to resist corrosion, but with the coating acting as a physical barrier rather than a sacrificial anode.

Title: Re: Rocket Engine Q&A
Post by: TyMoore on 07/23/2010 03:25 am
The trouble with doing this, especially with titanium, is that high pressure oxygen tends to diffuse through a thin barrier layer. Titanium can actually burn in the presence of pure oxygen, especially at high pressure. It is even pyrophoric--which means it can spontaneously combust with oxygen.

Small dings or even erosion can cause that thin protective layer to fail.

This is why using the high chrome stainless steel is essential. Chromium forms an oxide matrix which only allows oxygen to penetrate only so far, and no farther. It is also self-healing: erosion or dings only exposes fresh chromium, which instantly oxidizes, and then stops. A thin layer will not do this.

Title: Re: Rocket Engine Q&A
Post by: deltaV on 07/28/2010 04:30 am
The trouble with doing this, especially with titanium, is that high pressure oxygen tends to diffuse through a thin barrier layer. Titanium can actually burn in the presence of pure oxygen, especially at high pressure. It is even pyrophoric--which means it can spontaneously combust with oxygen.

Small dings or even erosion can cause that thin protective layer to fail.

This is why using the high chrome stainless steel is essential. Chromium forms an oxide matrix which only allows oxygen to penetrate only so far, and no farther. It is also self-healing: erosion or dings only exposes fresh chromium, which instantly oxidizes, and then stops. A thin layer will not do this.

Good point about dings. Something that burns up if scratched is the enemy of low costs. Thanks for your help.
Title: Re: Rocket Engine Q&A
Post by: Hells on 08/02/2010 08:42 pm
Even though one of RP's advantage is that it doesn't need deep cryogenic temperatures, is there any point in freezing it to get it just above the melting point to reduce thermal expansion?

Also what kind of pressures does RP tanks run at?
Title: Re: Rocket Engine Q&A
Post by: Antares on 08/03/2010 12:26 am
It gets thicker, so you'd have to design the engine pump for that or put a heater on the inlet.  RP tank pressures are an optimization of keeping the pump from cavitating (higher better), vehicle stiffness (higher better) and tank weight (lower better).
Title: Re: Rocket Engine Q&A
Post by: Propforce on 08/03/2010 01:01 am
Even though one of RP's advantage is that it doesn't need deep cryogenic temperatures, is there any point in freezing it to get it just above the melting point to reduce thermal expansion?

Also what kind of pressures does RP tanks run at?

It is a concept called "densified propellant" by cooling the propellant further thereby increase its density. This way you can pack more propellant in a given volume tank.  This concept works for room temperature RP, but also works for cryogenic hydrogen (yes, it can be chilled to an even-lower temperature, just more expensive to do it).

Antares is correct, when you change propellant properties (such as density) at the engine inlet, you need to either change it back (warm it up - but it takes energy) or re-design engine pump to take the higher density (hence higher mass flow).  Most engine suppliers (who are we kidding, there're only two in the U.S.) are not willing to do that without getting paid. So unless a new engine is designed to accept densified propellants, this pretty much is a "non-starter" for engine upgrades.

There are launch vehicles exist today that take densified propellants, but I think who that is proprietary information.  :)
Title: Re: Rocket Engine Q&A
Post by: Jim on 08/03/2010 01:06 am
Even though one of RP's advantage is that it doesn't need deep cryogenic temperatures, is there any point in freezing it to get it just above the melting point to reduce thermal expansion?

Delta actually does the opposite and heats the RP-1 before loading to increase its energy content
Title: Re: Rocket Engine Q&A
Post by: TyMoore on 08/03/2010 04:11 pm
Huh, I didn't know that Jim. Thanks.

Of course, we're only talking about 1-2% energy content variations, I would think.

Apparently warm RP-1 is a slight performance advantage.

Lower viscosity too ---> lower pump power needed, and lower injector pressure drop. Interesting...
Title: Re: Rocket Engine Q&A
Post by: Propforce on 08/03/2010 06:21 pm
Huh, I didn't know that Jim. Thanks.

Of course, we're only talking about 1-2% energy content variations, I would think.

Apparently warm RP-1 is a slight performance advantage.

Lower viscosity too ---> lower pump power needed, and lower injector pressure drop. Interesting...


No single factor has a clear advantage.  One must look at how vehicle and engine match up. 

Title: Re: Rocket Engine Q&A
Post by: sdsds on 08/17/2010 03:48 am
Simple version of my question:  "For how long can a liquid rocket engine fire?"

Background:  I'm interested in an approach which uses a relatively low thrust storable bipropellant engine to propel a spacecraft from LEO onto an Earth-escape trajectory.  The Ariane 5 ES upper stage engine is advertised as having a "Nominal single firing" duration of 1,100 seconds.  If I'm doing the math right that's 10.39 t of propellant, which is plenty!

But a variation of that engine that adds a turbpump and has a chamber pressure of 60 bar rather than 11 bar is only advertised as having a "Mission duty" of 5 starts/600 seconds.  600 seconds is not enough!  What's the physics that explains this difference?  And, how easy is it to improve this aspect of a design?
Title: Re: Rocket Engine Q&A
Post by: Antares on 08/17/2010 05:09 am
Limits I can think of:
tank size vs mass flow rate
heat transfer in the combustion chamber (some chambers have cracks after exceeding a single mission duration, others could go forever)
fatigue in turbopump blades and bearings
if deeply throttled, combustion instability

Your question needs more constraints.
Title: Re: Rocket Engine Q&A
Post by: TyMoore on 08/25/2010 02:02 pm
There is also the problem of erosion on throat surfaces. This is the most severely stressed (thermally and mechanically) component after turbopump impellers and turbines (and their bearings,) I think.


Also, wear seals in turbopumps. Some are faced with lead to provide a good seal, but their design lifetime is on the order of 100-200 seconds.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 08/30/2010 01:41 pm
Pictures of amateur or New Space rocket engines show yellow and blue discolourations of the aluminium chambers after they've been fired. What's the explanation for these colours? Are they metal oxides? I thought aluminium was covered in a thin layer of oxide anyway, but maybe other oxides are formed at higher temperatures?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 08/30/2010 04:30 pm
Pictures of amateur or New Space rocket engines show yellow and blue discolourations of the aluminium chambers after they've been fired. What's the explanation for these colours? Are they metal oxides? I thought aluminium was covered in a thin layer of oxide anyway, but maybe other oxides are formed at higher temperatures?

Could be contamination because of oxidation of the alloying elements. Doesn't need to be a lot either to cause visible color. I know trace amounts of lithium will cause aluminum to turn blue under temperature/humidity/altitude testing. Also, AL7075 has a lot of copper in it. I wouldn't be surprised if that would give you a blue or yellow discoloration after high heat. I'm not an M&P guy, buy figured I'd toss in a semi-educated guess.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 08/30/2010 04:32 pm
Thanks!
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 09/20/2010 01:15 pm
Would it be possible/worthwhile to add a form of afterburning to a gas generator exhaust to get performance closer to that of staged combustion without the full complexity? I suppose not or people would be doing it. If so can anyone explain why not?
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 09/20/2010 01:29 pm

I thought that was the whole point of TAN (Thrust Augmented Nozzle). Might one of the issues be whole having the turbopumps switching from a high flow low pressure output to a low flow high pressure output when you turn it back off? Assuming you have a flight profile that requires switching from high thrust to high ISP.
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/21/2010 03:52 pm
GG exhaust thrust is negligible.  The added complexity to duct it over wouldn't be worth the benefit of more energetic GG exhaust.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 09/21/2010 04:36 pm
What is the point of staged combustion then?
Title: Re: Rocket Engine Q&A
Post by: ugordan on 09/21/2010 04:45 pm
Higher turbine power which enables higher chamber pressures and higher performance.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 09/21/2010 04:50 pm
Do you mean higher pressure without the losses of obtaining those pressures with a gas generator?
Title: Re: Rocket Engine Q&A
Post by: ugordan on 09/21/2010 04:54 pm
I suppose that can be one way of looking at it. Typically most/all of one propellant goes through the preburner and a fraction of another so there's can be a higher mass flow through the turbine without sacrificing it by later dumping it overboard as is the case with GG engines.
Title: Re: Rocket Engine Q&A
Post by: Antares on 09/21/2010 05:10 pm
Right, the point of staged combustion is higher Isp.  100% of the mass flow comes out the one and only nozzle, optimized for thrust.  A GG exhaust has to worry about pluming other parts of the engine, among other things.
Title: Re: Rocket Engine Q&A
Post by: yinzer on 09/22/2010 04:37 am
Would it be possible/worthwhile to add a form of afterburning to a gas generator exhaust to get performance closer to that of staged combustion without the full complexity? I suppose not or people would be doing it. If so can anyone explain why not?

This is probably wrong in some subtle way, but here goes:

The gas in the chamber of a rocket is hot, at high pressure, and dense.  At the nozzle exit it is cooler, lower pressure, less dense, and moving very very fast.  Energy is conserved, and the kinetic energy of the exhaust had to come from somewhere.

If the pressure of the gas coming out the nozzle is too low you get backpressure losses, flow separation, and a gigantic nozzle.  This puts a limit onto how much energy you can extract from the gases in the chamber.  And the propellants have to be pumped up to higher than chamber pressure to get in, and the chamber has to not melt, which puts a limit onto how much energy you can get into the gases in the chamber.

The turbine for the pump works more or less the same way - high energy gases upstream, low energy gases downstream, but in this case work comes out of the drive shaft instead of as kinetic energy of the exhaust.  But the same principles apply - the gas at the exit of the turbine has to be higher pressure than where it's going, propellants have to be pumped up to higher that turbine inlet pressure, and the turbine inlet can't melt.

In a staged combustion engine they get pumped WAAAAY above chamber pressure, burned a little to add energy, expanded down to a little above chamber pressure to get pump drive energy, and then fed into the chamber to be burned the rest of the way.

In a gas generator engine, the propellants get pumped up to a bit higher than chamber pressure, burned a little to add energy, and then expanded down to a low pressure to turn their energy into shaft power for the main pumps.  They aren't at high enough pressure to really do anything useful as a rocket engine - I found something saying the gas after the turbine on the Vulcain is at ~150 psi - although there's usually enough left over to generate thrust for roll control.
Title: Re: Rocket Engine Q&A
Post by: Propforce on 09/28/2010 12:32 am
Would it be possible/worthwhile to add a form of afterburning to a gas generator exhaust to get performance closer to that of staged combustion without the full complexity? I suppose not or people would be doing it. If so can anyone explain why not?

This is probably wrong in some subtle way, but here goes:

In a staged combustion engine they get pumped WAAAAY above chamber pressure, burned a little to add energy, expanded down to a little above chamber pressure to get pump drive energy, and then fed into the chamber to be burned the rest of the way.

In a gas generator engine, the propellants get pumped up to a bit higher than chamber pressure, burned a little to add energy, and then expanded down to a low pressure to turn their energy into shaft power for the main pumps.  They aren't at high enough pressure to really do anything useful as a rocket engine - I found something saying the gas after the turbine on the Vulcain is at ~150 psi - although there's usually enough left over to generate thrust for roll control.

Good enough for government work !  :)

In short, Gas Generator (GG) cycle does not pump the pressure up high enough.  Most GG cycle engine chamber pressure (Pc) is around 1,500 ~ 2,000 psia max.  But the Staged Combustion (SC) cycle can go upward to 6,000 ~ 7,000 psia and the Pc for the SSME is around 3,000 psia. 

Because the SC can pump up to high pressure (6,000 ~ 7,000 psia), the "fuel" has enough pressure after cooling the nozzle & chamber to inject into chamber to burn with "oxidizer".

Title: Re: Rocket Engine Q&A
Post by: Antares on 09/28/2010 12:54 am
There's no causation yet mentioned here.  All we're talking about is turbopump power balance, which is designable.  A GG could be designed to pump pressure up high enough.

You could have a low pressure SC or a high pressure GG.  The question is, why wouldn't you?
Title: Re: Rocket Engine Q&A
Post by: yinzer on 09/28/2010 02:27 am
There's no causation yet mentioned here.  All we're talking about is turbopump power balance, which is designable.  A GG could be designed to pump pressure up high enough.

You could have a low pressure SC or a high pressure GG.  The question is, why wouldn't you?

You wouldn't have a high-pressure GG engine because in order to get the energy to drive the pumps without melting the turbine you need to send too much mass through the gas generator and overboard without generating thrust.  That part's easy.

As to why you wouldn't have a low-pressure staged combustion engine... staged combustion requires that propellants be pumped up significantly higher than chamber pressure and that the turbine exhaust be suitable for injecting into a rocket chamber.  I could imagine the turbine exhaust requirement causing trouble for kerosene engines where the exhaust is either sooty or oxygen-rich.  There's also the matter of staged combustion engines needing to handle a lot of high-temperature gas outside the chamber while gas generator engines can handle most of their propellants as liquids.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 09/28/2010 07:45 am
You wouldn't have a high-pressure GG engine because in order to get the energy to drive the pumps without melting the turbine you need to send too much mass through the gas generator and overboard without generating thrust.  That part's easy.

Couldn't you fix that by combusting the gas generator exhaust further? On reflection, that would itself be a form of staged combustion. Could there be applications where this has an advantage over a normal GG or SC engine?

It would be a bit like an air turborocket, where the gas generator exhausts into an afterburner, only in this case you probably wouldn't use atmospheric air. Or could it be useful to have an airbreathing gas generator for first stages?
Title: Re: Rocket Engine Q&A
Post by: yinzer on 09/28/2010 06:33 pm
You wouldn't have a high-pressure GG engine because in order to get the energy to drive the pumps without melting the turbine you need to send too much mass through the gas generator and overboard without generating thrust.  That part's easy.

Couldn't you fix that by combusting the gas generator exhaust further? On reflection, that would itself be a form of staged combustion. Could there be applications where this has an advantage over a normal GG or SC engine?

Combusting the gas generator exhaust turns it into a smaller staged combustion rocket engine.  But in order to run a rocket engine the chamber pressure has to be much higher than ambient, and there was already a big pressure drop across the GG turbine.  To use the exhaust as a rocket you have to really increase the inlet pressure and thus pump power, and just eating the performance losses of dumping the exhaust overboard and making a slightly bigger rocket can be significantly cheaper.

Quote
It would be a bit like an air turborocket, where the gas generator exhausts into an afterburner, only in this case you probably wouldn't use atmospheric air. Or could it be useful to have an airbreathing gas generator for first stages?

Even a first stage spends very little time in dense atmosphere.  You have to run the numbers to see whether it makes sense, but almost all the time it doesn't.
Title: Re: Rocket Engine Q&A
Post by: AnalogMan on 10/01/2010 05:06 pm
Thought this might be of general interest to readers of this thread.

Liquid Rocket Engine Testing - Historical Lecture: Simulated Altitude Testing at AEDC*
N S Dougherty (July 27, 2010) 46 pages

The span of history covered is from 1958 to the present. The outline of this lecture draws from historical examples of liquid propulsion testing done at AEDC primarily for NASA's Marshall Space Flight Center (NASA/MSFC) in the Saturn/Apollo Program and for USAF Space and Missile Systems dual-use customers. NASA has made dual use of Air Force launch vehicles, Test Ranges and Tracking Systems, and liquid rocket altitude test chambers / facilities. Examples are drawn from the Apollo/ Saturn vehicles and the testing of their liquid propulsion systems. Other examples are given to extend to the family of the current ELVs and Evolved ELVs (EELVs), in this case, primarily to their Upper Stages. The outline begins with tests of the XLR 99 Engine for the X-15 aircraft, tests for vehicle / engine induced environments during flight in the atmosphere and in Space, and vehicle staging at high altitude. The discussion is from the author's perspective and background in developmental testing.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100032986_2010034410.pdf (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100032986_2010034410.pdf)  [2.65 MB]

*AEDC - Arnold Engineering Development Center
Title: Re: Rocket Engine Q&A
Post by: JosephB on 10/01/2010 05:29 pm
Good reading & nice find.

Edit: that page 8 staging photo is phenominal. Gotta try to find a better resolution version. Thanks for posting this!
Title: Re: Rocket Engine Q&A
Post by: Antares on 10/01/2010 07:09 pm
Vacuum pumps that had been brought over from Peenemunde were just retired at AEDC a few years ago.  Phenomenal.
Title: Re: Rocket Engine Q&A
Post by: PahTo on 10/01/2010 07:23 pm

I have a question--hope this is the correct place:

How is horsepower measured for rocket engines?

Or more directly, why is it that many sources say the RS-25 (SSME) produces 12 million horsepower (approx 400,000 pound thrust) while the recent 5-seg DM tests say "22 million horsepower" (approx 3.5 million pounds thrust)?  Those numbers just don't seem to jibe.
Perhaps the horsepower is total energy for the entire burn time (SSME is measured over the 8.5 minute burn, SRB for 2 minute)?

Title: Re: Rocket Engine Q&A
Post by: JosephB on 10/01/2010 07:23 pm
Peenemunde? Thats too funny.
At least the price was right.
Title: Re: Rocket Engine Q&A
Post by: Antares on 10/01/2010 08:06 pm
One might do Thrust times exhaust velocity to get a power.

Horsepower is power (energy divided by time) not energy.

One might also do power of the turbopumps to get horsepower for a liquid engine.

Power of a solid motor doesn't make any sense to a technical audience.  They just needed something for a press release.  The general public knows something about power because they drive cars.  They know nothing about thrust and total impulse.
Title: Re: Rocket Engine Q&A
Post by: PahTo on 10/01/2010 08:11 pm

Re: horsepower equivalent for rocket engines:

Thanks Antares!
Title: Re: Rocket Engine Q&A
Post by: simonbp on 10/01/2010 10:51 pm
Vacuum pumps that had been brought over from Peenemunde were just retired at AEDC a few years ago.  Phenomenal.

Though that was as much a case of procrastination as endurance. The icing tunnel at Glenn is just now getting its circa-1944 compressors replaced.

Ironically, AEDC nearly went to Huntsville over Tullahoma. Putting the ABMA at Redstone was considered a consolation prize...
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/12/2010 10:23 am
Another wild idea, this time about hybrids. As I understand it there are two types of hybrids, the ordinary kind where liquid oxidiser is injected into the hollow core(s) of a solid fuel and the reverse hybrid where liquid fuel is injected into the core(s) of a solid oxidiser. The other day I was reading a presentation about air turborockets with a solid gas generator and I wondered if a variant of this could be applied to hybrids. What if you didn't inject liquid at the top, but used a very fuel-rich solid as a giant gas generator that injects hot fuel-rich gas into a combustion chamber at the bottom? If you did things this way the solid might not need as much thermal protection and the final combustion chamber could be liquid cooled. You'd also get excellent mixing. I've never read anything about this. Does this idea have any merit?
Title: Re: Rocket Engine Q&A
Post by: 93143 on 10/12/2010 08:28 pm
Yes, it does.  You aren't the first to think of it.  It solves a number of problems with conventional hybrids, and promises to be very useful for a small, low-cost expendable LV...

Power of a solid motor doesn't make any sense to a technical audience.

Sure it does.  As you said, 1/2 * thrust * exhaust velocity = jet power.

A frustrating metric, to be sure, since you have to combine it with another parameter to calculate anything you actually care about...
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/12/2010 08:38 pm
Thanks! I didn't think I was the first to think of it, but it's nice to know I found something that's not completely crazy.  :)
Title: Re: Rocket Engine Q&A
Post by: Proponent on 10/13/2010 05:41 am
A frustrating metric, to be sure, since you have to combine it with another parameter to calculate anything you actually care about...

Indeed, it's not of much use for propulsion, but it's more relevant in the context of rocket-driven MHD.
Title: Re: Rocket Engine Q&A
Post by: butters on 10/25/2010 07:06 am
Another wild idea, this time about hybrids. <snip> What if you didn't inject liquid at the top, but used a very fuel-rich solid as a giant gas generator that injects hot fuel-rich gas into a combustion chamber at the bottom? If you did things this way the solid might not need as much thermal protection and the final combustion chamber could be liquid cooled. You'd also get excellent mixing. I've never read anything about this. Does this idea have any merit?

Interesting. Chamber cooling with oxidizer might be a metallurgical problem.
Title: Re: Rocket Engine Q&A
Post by: 93143 on 10/25/2010 07:39 am
Reaction Engines needs LOX cooling for Skylon because the hydrogen is otherwise occupied.  They don't seem to consider it a major potential showstopper.

I seem to recall some NewSpace company accidentally plumbing a rocket backwards and finding out that it is, in fact, possible to film-cool a rocket with LOX...
Title: Re: Rocket Engine Q&A
Post by: butters on 10/25/2010 08:02 am
Why do R-7/Soyuz booster and core stages use LOX with an H2O2 gas generator?  Why didn't they inject the decomposed H2O2 into the combustion chambers and eliminate the LOX system? 

The H2O2 turbine would just pump the RP-1 through the cooling jacket and into combustion chamber, where it should be hypergolic with the decomposed H2O2.  No cryogenics, no igniters, higher density, and storable propellants for their ICBM.

They already worked out the gas generator and turbine technology, and presumably H2O2 production wouldn't have been a major problem if they could depend on it in somewhat lower quantities.

So why didn't it work out that way, and why hasn't closed-cycle H2O2/RP-1 really caught on for booster rockets (besides Black Arrow)?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/25/2010 11:13 am
Interesting. Chamber cooling with oxidizer might be a metallurgical problem.

It can be done with peroxide and I read on NTRS that there were successful experiments with LOX cooling too, which is apparently not as difficult as it seems, even with the occasional leak. The Reaction Engines people want to use LOX cooling for Skylon and they are planning to do some tests with it, or maybe they have done a few already.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/25/2010 11:17 am
So why didn't it work out that way, and why hasn't closed-cycle H2O2/RP-1 really caught on for booster rockets (besides Black Arrow)?

I'm curious about that too. Maybe it's a case of if it ain't broke, don't fix it. H2O2 with RP-1 or even jet fuel is fine, but so is LOX/RP-1. Large peroxide catalyst packs may also be very heavy and inconvenient. On the other hand, I read that Aerojet tested a tri-fluid injector which used a relatively small stream of decomposed peroxide to ignite larger streams of liquid peroxide and liquid fuel. I assume this was done to avoid the use of a very large catalyst pack.

Beal also tried peroxide, but in a pressure-fed system. In that case the high density of peroxide and the high O/F ratio help keep down the mass penalty.
Title: Re: Rocket Engine Q&A
Post by: yinzer on 10/25/2010 06:04 pm
Why do R-7/Soyuz booster and core stages use LOX with an H2O2 gas generator?  Why didn't they inject the decomposed H2O2 into the combustion chambers and eliminate the LOX system? 

The H2O2 turbine would just pump the RP-1 through the cooling jacket and into combustion chamber, where it should be hypergolic with the decomposed H2O2.  No cryogenics, no igniters, higher density, and storable propellants for their ICBM.

They already worked out the gas generator and turbine technology, and presumably H2O2 production wouldn't have been a major problem if they could depend on it in somewhat lower quantities.

So why didn't it work out that way, and why hasn't closed-cycle H2O2/RP-1 really caught on for booster rockets (besides Black Arrow)?

Significantly better performance with straight LOX than with H2O2.  H2O2 is already partially burned, in a manner of speaking.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/25/2010 06:11 pm
You have lower GLOW with LOX, but not better mass fractions, because of the higher overall density of the propellant.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 10/27/2010 01:32 am
The Russians had plans for a tripropellant engine that would do this with hydrogen, but in their case the main purpose was to start in kerolox mode and then switch to hydrolox at high altitude. If you have the plumbing anyway, you might as well have an intermediate mode in which all three propellants are used.

Aerojet has done the work to demonstrate the feasibility of providing a TAN for the RS-68. I am not at liberty to discuss results except to say categorically that the results were very, very good. The application would be to have a LH2/LOX core, ground ignited, with RP-1/LOX burning in the nozzle to provide the same advantage as having side-mounted RP-1/LOX LRB's, however without the additional mass of the RP-1 engines. The RP-1 would be carried in side mounted tanks with the LOX being provided by the core LOX tank. When the RP-1 is depleted, the side-mounted tanks would be jettisoned just as if they were actual LRB stages when in reality they are only drop tanks. Meanwhile the LH2/LOX core continues upward on its own. Very efficient.

Does that also imply that TAN might be equally applicable to SSME?

cheers, Martin

Theoretically yes, but that would be expensive to implement.
IMO it would probably cost less to create an engine from scratch than to modify the SSME for TAN.

Not meaning any disrespect, but why?
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 10/27/2010 01:36 am
If the bubbles are small (nucleate boiling) it's not a problem at all, this is the most effective cooling. Large bubbles are a problem, except with hydrogen which is a good coolant even in gaseous form.

Hmm. If they are using water, then most of the cooling would come from the boiling, right?
Title: Re: Rocket Engine Q&A
Post by: Antares on 10/27/2010 05:49 am
Incropera and Dewitt is your friend.  But reread the subject post anyway.  It was propellant cooling a combustion chamber.  Water is not exactly in the running for that application.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 10/27/2010 08:30 am
Incropera and Dewitt is your friend.

Thanks for that reference.

Quote
But reread the subject post anyway.  It was propellant cooling a combustion chamber.  Water is not exactly in the running for that application.

Well, we were earlier talking about the third fluid cooling concept, which could use water for cooling. Still requires the propellant as a heatsink of course.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 10/31/2010 01:29 am
Incropera and Dewitt is your friend.  But reread the subject post anyway.  It was propellant cooling a combustion chamber.  Water is not exactly in the running for that application.

What are Incropera and Dewitt?
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 10/31/2010 01:31 am
Sorry for asking-found on Amazon
Title: Re: Rocket Engine Q&A
Post by: scienceguy on 11/03/2010 12:01 am
How would something like polywell fusion help a launch? How much of say the space shuttle is used for power generation? Doesn't all the thrust come from the chemical rockets?
Title: Re: Rocket Engine Q&A
Post by: 93143 on 11/03/2010 09:26 am
The Space Shuttle Main Engines put out around 5 GW each in a vacuum, and each solid booster puts out about 18 GW.  This is pure chemical rocketry; power in this case refers to jet power: 0.5 x thrust x exhaust velocity.

Electrical power for the Shuttle is provided by three hydrogen/oxygen fuel cells weighing a total of about 350 kg.  At peak power, they can temporarily supply 36 kW, or about a millionth as much power as the solid rocket boosters.

...

Propulsion schemes using Polywell involve application of the high-voltage electricity generated by the reactor to heat and/or accelerate propellant, which can be atmospheric air, internally-carried propellant, or both.  The idea is that a substantial fraction of the engine power is provided by the Polywell's electrical output rather than by chemical reactions, thus allowing much higher specific impulse than pure chemical rockets (and/or better performance than chemically-fueled scramjets) and making a reusable, economical SSTO possible.  In all cases, extensive engine research and development would be required.

Very high power-to-weight ratios are required for fusion-powered launch vehicle concepts.  According to my calculations, reactor performance in the range of 10 kW/kg including radiation shielding might enable an airbreathing SSTO, while 100 kW/kg (which may be difficult if not impossible due to anticipated shielding requirements) could potentially enable an all-rocket SSTO, if the engines can be designed to allow substantial LOX augmentation...

Aneutronic p-¹¹B operation with direct electrical conversion is probably required, since a neutronic reaction like D-T would require either a thermal generating plant (prohibitively heavy) or direct use of the thermal energy (which doesn't work very well due to low peak temperatures; fission thermal would be better).
Title: Re: Rocket Engine Q&A
Post by: scienceguy on 11/03/2010 07:55 pm
Thanks
Title: Re: Rocket Engine Q&A
Post by: gin455res on 11/25/2010 08:12 am
Are there any designs for tri-propellants that use two different oxidisers and one fuel? 

e.g. n2o, lox and lh2.

The motivation being to improve t/w at lift-off, as opposed to trying to improve isp.  It seems to me that since most propellant combinations have much larger fractions of oxidiser, varying the composition of the oxidiser is the most effective way to increase mass flow and consequently improve t/w.
Title: Re: Rocket Engine Q&A
Post by: madscientist197 on 11/25/2010 09:51 am
Perhaps a better combination might be initial use of (high thrust density) H2O2 followed by later use of O2 (higher thrust per mass). I'm not sure whether there are any fuel specific trade-offs in these sorts of cases -- perhaps someone else knows?
Title: Re: Rocket Engine Q&A
Post by: Hop_David on 12/12/2010 03:49 am
On page 196 of A Practical Architecture for Exploration-Based Mars Missions (http://www.marsdrive.com/Libraries/Downloads/A_Practical_Architecture_for_Exploration-Focused_Manned_Mars.sflb.ashx) it talks about an MAV with methane burning RL10s. Can RL10s burn methane? If so, do they need to be modified substantially from the hydrogen burning version?
Title: Re: Rocket Engine Q&A
Post by: AZ-JRB on 12/12/2010 08:08 pm
The RL10 running on methane was essentially proving, successfully, that one could use methane in an expander cycle engine. Because of the significantly different density, the hydrogen component flow areas had to be significantly reduced [e.g. the fuel pump, turbine, chamber/nozzle coolant tubes] when methane was used.
Title: Re: Rocket Engine Q&A
Post by: Scotty on 12/12/2010 10:17 pm
Yes, a modifled RL-10 was run on Methane during ground testing.
Title: Re: Rocket Engine Q&A
Post by: tnphysics on 12/12/2010 11:15 pm
How expensive, time-consuming, etc were the mods?
Title: Re: Rocket Engine Q&A
Post by: Hop_David on 12/14/2010 03:32 am
How expensive, time-consuming, etc were the mods?

That's the question I was going to ask.

I had imagined a vehicle getting propellent from various depots, sometimes hydrogen and oxygen, other times methane and oxygen.

However, if the needed modifications are extensive, it would be best to limit a given vehicle to one type of fuel.
Title: Re: Rocket Engine Q&A
Post by: HMXHMX on 12/14/2010 07:09 am
Yes, a modifled RL-10 was run on Methane during ground testing.

I used to have the report, but that was in the early 1980s, and it is long gone.  There were a number of issues, and I recall seeing damage to the engine.  But they were running an essentially unmodified engine as I recall.
Title: Re: Rocket Engine Q&A
Post by: AZ-JRB on 12/14/2010 03:52 pm

I had imagined a vehicle getting propellent from various depots, sometimes hydrogen and oxygen, other times methane and oxygen.

However, if the needed modifications are extensive, it would be best to limit a given vehicle to one type of fuel.
[/quote]

For the Methane demonstrator tests, the fuel pump impellor blades were cut back, the turbine flow area was reduced and the chamber/nozzle coolant flow passages had inserts inside to reduce the area and therefore keep the fluid velocity up. Because of the density difference, it is unlikely to be able to use both methane and hydrogen as fuel in a "simple" engine. BTW, the demonstrator engine used FLOX rather than Oxygen, however no modification to the engine hardware was required [other than passivation of the Lox components] since the densities are about the same.
Title: Re: Rocket Engine Q&A
Post by: Scotty on 12/14/2010 08:30 pm
The engine would be configured to run on one fuel.
It would be very difficult to set the engine up to run on either fuel.
Do some reading:

http://www.pw.utc.com/StaticFiles/Pratt%20&%20Whitney%20New/Media%20Center/Press%20Kit/1%20Static%20Files/pwr_cece.pdf

http://www.dlr.de/sart/publications/pdf/0095-0212prop.pdf

Title: Re: Rocket Engine Q&A
Post by: Hop_David on 12/23/2010 02:36 pm
http://www.pw.utc.com/StaticFiles/Pratt%20&%20Whitney%20New/Media%20Center/Press%20Kit/1%20Static%20Files/pwr_cece.pdf


Thank you. The methane ISP in the above pdf says >350 sec.

Is 386 seconds an unrealistic expectation for an RL10 methane engine?

http://www.dlr.de/sart/publications/pdf/0095-0212prop.pdf

This one will take some time to digest.
Title: Re: Rocket Engine Q&A
Post by: joertexas on 12/24/2010 05:24 am

http://www.pw.utc.com/StaticFiles/Pratt%20&%20Whitney%20New/Media%20Center/Press%20Kit/1%20Static%20Files/pwr_cece.pdf


Just out of curiousity, how much does one of these engines cost?

JR
Title: Re: Rocket Engine Q&A
Post by: deltaV on 01/14/2011 06:08 pm
Has anyone considered using Aluminum/Lithium alloy as rocket propellant for an ion engine or other low-thrust engine? Aluminum/Lithium alloy is available in space for free in large quantities in the form of spent propellant tanks. Both aluminum and lithium also have fairly low ionization energies.

I'm guessing that the hard part would be converting the AlLi tanks into some usable form (e.g. AlLi dust) without heavy or expensive equipment and without releasing oodles of space debris in the process.
Title: Re: Rocket Engine Q&A
Post by: Nomadd on 01/14/2011 06:23 pm
Has anyone considered using Aluminum/Lithium alloy as rocket propellant for an ion engine or other low-thrust engine? Aluminum/Lithium alloy is available in space for free in large quantities in the form of spent propellant tanks. Both aluminum and lithium also have fairly low ionization energies.

I'm guessing that the hard part would be converting the AlLi tanks into some usable form (e.g. AlLi dust) without heavy or expensive equipment and without releasing oodles of space debris in the process.

 With a nick like that, you should have an idea of the delta-v that would be needed to scoot around scavenging old fuel tanks in space. You'd need to use chemical rockets for that unless you wanted to take 20 years doing it. It would probably be about a hundred times as efficient just to replace the fuel you'd use picking up the tanks with whatever your ion engine uses in the first place.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 01/14/2011 07:02 pm
With a nick like that, you should have an idea of the delta-v that would be needed to scoot around scavenging old fuel tanks in space. You'd need to use chemical rockets for that unless you wanted to take 20 years doing it. It would probably be about a hundred times as efficient just to replace the fuel you'd use picking up the tanks with whatever your ion engine uses in the first place.

I didn't mean to suggest scavenging old fuel tanks from various orbits, which as you say would be rather ridiculous (except possibly for a spacecraft whose job is space debris removal and has to rendezvous with space junk anyway). Fortunately there's no need to change orbits to get junk propellant tanks since every spacecraft enters its life with at least one propellant tank in precisely the same orbit, namely the tanks in the upper stage that launched it. (Of course upper stages are usually purposefully left in unstable orbits for disposal, but it doesn't take much delta-vee to get from there to a stable orbit.) One example would be an ISS reboost engine using tanks from launch vehicles servicing the ISS.
Title: Re: Rocket Engine Q&A
Post by: Nomadd on 01/14/2011 07:42 pm
 Gotcha. But I'm still betting that it wouldn't be practical with any technology we have today. I'm not sure if using solid particles instead of nobel gass for ion has ever been seriously considered. The processing of the metal and less than optimum design and efficiency from not using the best fuel would probably make it a whole lot better to just take more argon or xenon with you.
 Those accelerator grids are pretty tight. Increasing the spacing even a tiny bit to use solid fuel would drop their efficiency a bunch.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 01/14/2011 09:36 pm
Gotcha. But I'm still betting that it wouldn't be practical with any technology we have today. I'm not sure if using solid particles instead of nobel gass for ion has ever been seriously considered. The processing of the metal and less than optimum design and efficiency from not using the best fuel would probably make it a whole lot better to just take more argon or xenon with you.
 Those accelerator grids are pretty tight. Increasing the spacing even a tiny bit to use solid fuel would drop their efficiency a bunch.

There is certainly precedent for using metals in high-isp engines, including cesium, bismuth and lithium. Apparently for some sorts of engines lithium is a top performer: http://en.wikipedia.org/wiki/Magnetoplasmadynamic_thruster.
I get the impression that the metals are vaporized (and/or plasma-ized) before entering the engines. Of course AlLi alloys are mostly aluminum so the main point is that some metals are usable in some engines, not that lithium in particular is. I haven't found any mention of using aluminum as a propellant (in non-chemical rockets).

Not all low-thrust engines have accelerator grids.
Title: Re: Rocket Engine Q&A
Post by: Malderi on 01/14/2011 09:49 pm
Has anybody studied that particular case, I.e., the feasibility of scavenging upper stages and using them for the fuel to reach the next one? If the engine could be made to work, then that"solves" the big delta v problem of debris removal. Obviously, that is a big if.
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/14/2011 10:18 pm
Has anybody studied that particular case, I.e., the feasibility of scavenging upper stages and using them for the fuel to reach the next one?

There isn't any left over propellant.  All upperstages are passivated by depletion burns and blow downs.
Title: Re: Rocket Engine Q&A
Post by: deltaV on 01/14/2011 11:32 pm
I just discovered a couple of studies from the early 1980s proposing various uses of shuttle external tanks, including as rocket propellant:
* http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19940004970_1994004970.pdf
* http://www.freemars.org/studies/et/

Propellant possibilities mentioned in those studies include:
* Use of residual LOX/LH2 propellants in a traditional chemical rocket
* Combine residual LOX/LH2 with powdered aluminum from the tank for a tripropellant chemical rocket
* Use aluminum in some sort of mass driver, rail gun or coil gun.

I didn't see any mention of ion engines in those studies, perhaps because those weren't in common use in the west when these studies were written.
Title: Re: Rocket Engine Q&A
Post by: gospacex on 01/15/2011 01:06 am
Mass driver (linear motor) may be a better way to accelerate small bits of Al to large velocities (compared to ion engine). Aluminum with its high conductivity is a good projectile material.
Title: Re: Rocket Engine Q&A
Post by: Antares on 01/16/2011 02:08 am
Can we take this la-la land stuff to one of the advanced propulsion threads?
Title: Re: Rocket Engine Q&A
Post by: WulfTheSaxon on 04/07/2011 06:13 pm
Astronautix lists the RS-56-OBA (http://www.astronautix.com/engines/rs56oba.htm) as 805 kg with 1046.8 kN. That can’t be right. Is that not including pumps on the MA-5A or something?  ???

Several charts (http://www.sei.aero/downloads/SEI_TW_Trends_022801.pdf) list it as having a T/W of around 120-130 (for some reason it manages to vary), higher than even the NK-33.
Title: Re: Rocket Engine Q&A
Post by: Antares on 05/20/2011 08:19 pm
Been thinking/wondering: While I realize that Space-X is looking into building a LH2/LOX engine what kind of work would have to be done to turn something like the Kestral (or Merlin for that matter since it's a re-gen already) LOX/Cryo-Propane, or LOX/Propylene engine for higher performance?
http://www.dunnspace.com/alternate_ssto_propellants.htm

All things being equal (which I grant they are not of course but in a general sense) the increased delta-v kick alone would make it seem a worthwhile thing to look into. (If I read it right there is almost a 40% increase in payload with Propylene)

Most likely a completely new turbopump and a completely new combustion chamber.  Since those are pretty much the two most important parts of an engine, it's basically a new engine.  Unless... somehow one is adaptable as-is and the other is changed.  The needs for new pieces are driven by the vast differences in the fluid characteristics of the different fuels.
Title: Re: Rocket Engine Q&A
Post by: RanulfC on 05/24/2011 12:40 pm
Been thinking/wondering: While I realize that Space-X is looking into building a LH2/LOX engine what kind of work would have to be done to turn something like the Kestral (or Merlin for that matter since it's a re-gen already) LOX/Cryo-Propane, or LOX/Propylene engine for higher performance?
http://www.dunnspace.com/alternate_ssto_propellants.htm

All things being equal (which I grant they are not of course but in a general sense) the increased delta-v kick alone would make it seem a worthwhile thing to look into. (If I read it right there is almost a 40% increase in payload with Propylene)

Most likely a completely new turbopump and a completely new combustion chamber.  Since those are pretty much the two most important parts of an engine, it's basically a new engine.  Unless... somehow one is adaptable as-is and the other is changed.  The needs for new pieces are driven by the vast differences in the fluid characteristics of the different fuels.
One of the reasons I'd asked in the Space-X section in the first place was because they ARE developing a LH2/LOX upper stage engine which would also be Cryo-Propane usable (probably).

I found an answer (somewhat) to the LOX/Propylene question in information from Garvey Aerospace; Seems NO-ONE has done much research on Propylene as a fuel GA is static testing a motor at the moment and the preliminary results look good but they are in unknown territory since the research on the combination is so thin.

Randy
Title: Re: Rocket Engine Q&A
Post by: fatjohn1408 on 06/10/2011 02:36 pm
I have a question regarding Hybrid rocket engines.

How does a thrust profile of hybrids look like and on what do these profiles depend? Can these engines be throttled also like liquids?

I'm mainly interested into whether or not they can for instance power a launch vehicle at a constant load of for example 3g and if they cannot, how close can they get to such a profile?

Thank you in advance.
Title: Re: Rocket Engine Q&A
Post by: fatjohn1408 on 06/14/2011 02:21 pm
Hi, I have another question

Has anyone ever studied the thrust density of chemical rocket engines?
Regarding electrical propulsion, thrust density is the thrust divided by the area and is kinda important.

I think it is also important for chemical engines, certainly if you want to design a slender high thrust system.

My initial idea would be that low density fuels would still have a lower density at the nozzle exit and that this would not be counteracted by their slightly higher Isp, therefore they would also lead to low thrust density engines.

This might be a partial reason why space shuttle, buran and ariane used a combination of solid an LH2.

But then again Atlas does it all with LH2, so it's clearly possible.
Anybody knows alot more about this topic?
Title: Re: Rocket Engine Q&A
Post by: gospacex on 06/14/2011 02:58 pm
But then again Atlas does it all with LH2, so it's clearly possible.
Anybody knows alot more about this topic?

Atlas does not. Delta-IV does.
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 06/14/2011 05:27 pm
Hi, I have another question

Has anyone ever studied the thrust density of chemical rocket engines?
Regarding electrical propulsion, thrust density is the thrust divided by the area and is kinda important.
...
That's merely the thrust coefficient (Cf) times pressure. Of course it's been "studied." Look in your Rocket Propulsion Elements by Sutton (Chapter 3, I believe).
Title: Re: Rocket Engine Q&A
Post by: Downix on 06/28/2011 06:29 am
In a discussion about deeply throttleable engines, someone remarked that an engine's combustion chamber and throat were the same, which limited throttling.  So, I studied a few 60's and 70's era engines and ran across a design draft, which I then made a rough schematic of, which actually in part addresses that issue.  Wide pintle, which can operate in one of two modes, low-thrust mode with the pintle fully extended, or high-performance mode with it retracted.  It used the fuel itself as the hydraulic fluid to extend the pintle, with a simple valve to guide which way the system would run.  As you can see, when extended it basically puts a new throat into the design, and reduces the combustion chamber size dramatically. Excuse the crudeness of the drawing.
Title: Re: Rocket Engine Q&A
Post by: clongton on 06/28/2011 10:39 pm
In a discussion about deeply throttleable engines, someone remarked that an engine's combustion chamber and throat were the same, which limited throttling.  So, I studied a few 60's and 70's era engines and ran across a design draft, which I then made a rough schematic of, which actually in part addresses that issue.  Wide pintle, which can operate in one of two modes, low-thrust mode with the pintle fully extended, or high-performance mode with it retracted.  It used the fuel itself as the hydraulic fluid to extend the pintle, with a simple valve to guide which way the system would run.  As you can see, when extended it basically puts a new throat into the design, and reduces the combustion chamber size dramatically. Excuse the crudeness of the drawing.

Thanks Nate. This is exactly the kind of thing which is very appropriate for this thread. Much appreciated.
Title: Re: Rocket Engine Q&A
Post by: HMXHMX on 07/06/2011 05:42 pm
In a discussion about deeply throttleable engines, someone remarked that an engine's combustion chamber and throat were the same, which limited throttling.  So, I studied a few 60's and 70's era engines and ran across a design draft, which I then made a rough schematic of, which actually in part addresses that issue.  Wide pintle, which can operate in one of two modes, low-thrust mode with the pintle fully extended, or high-performance mode with it retracted.  It used the fuel itself as the hydraulic fluid to extend the pintle, with a simple valve to guide which way the system would run.  As you can see, when extended it basically puts a new throat into the design, and reduces the combustion chamber size dramatically. Excuse the crudeness of the drawing.

Do you recall where you saw this?  It seems similar to a recent, attached, TRW expansion-deflection nozzle patent.
Title: Re: Rocket Engine Q&A
Post by: Downix on 07/07/2011 01:55 am
In a discussion about deeply throttleable engines, someone remarked that an engine's combustion chamber and throat were the same, which limited throttling.  So, I studied a few 60's and 70's era engines and ran across a design draft, which I then made a rough schematic of, which actually in part addresses that issue.  Wide pintle, which can operate in one of two modes, low-thrust mode with the pintle fully extended, or high-performance mode with it retracted.  It used the fuel itself as the hydraulic fluid to extend the pintle, with a simple valve to guide which way the system would run.  As you can see, when extended it basically puts a new throat into the design, and reduces the combustion chamber size dramatically. Excuse the crudeness of the drawing.

Do you recall where you saw this?  It seems similar to a recent, attached, TRW expansion-deflection nozzle patent.
it was while I was researching the low cost pressure fed booster concepts for the TAOS shuttle, I think it was by either TRW or Boeing.
Title: Re: Rocket Engine Q&A
Post by: HMXHMX on 07/07/2011 03:06 am
In a discussion about deeply throttleable engines, someone remarked that an engine's combustion chamber and throat were the same, which limited throttling.  So, I studied a few 60's and 70's era engines and ran across a design draft, which I then made a rough schematic of, which actually in part addresses that issue.  Wide pintle, which can operate in one of two modes, low-thrust mode with the pintle fully extended, or high-performance mode with it retracted.  It used the fuel itself as the hydraulic fluid to extend the pintle, with a simple valve to guide which way the system would run.  As you can see, when extended it basically puts a new throat into the design, and reduces the combustion chamber size dramatically. Excuse the crudeness of the drawing.

Do you recall where you saw this?  It seems similar to a recent, attached, TRW expansion-deflection nozzle patent.
it was while I was researching the low cost pressure fed booster concepts for the TAOS shuttle, I think it was by either TRW or Boeing.

Thanks.
Title: Re: Rocket Engine Q&A
Post by: RanulfC on 07/07/2011 12:50 pm
I have a question regarding Hybrid rocket engines.

How does a thrust profile of hybrids look like and on what do these profiles depend? Can these engines be throttled also like liquids?

I'm mainly interested into whether or not they can for instance power a launch vehicle at a constant load of for example 3g and if they cannot, how close can they get to such a profile?

Thank you in advance.
As far as I can figure it really depends on the hybrid, what type of fuel/oxidizer combo used and the propellant properties themselves.

There is some good information here that might help:
http://aa.stanford.edu/events/50thAnniversary/media/Karabeyoglu.pdf (http://aa.stanford.edu/events/50thAnniversary/media/Karabeyoglu.pdf)

I've still got high hopes for the Paraffin hybrids :o)

Randy
Title: Re: Rocket Engine Q&A
Post by: RanulfC on 07/07/2011 12:54 pm
It was while I was researching the low cost pressure fed booster concepts for the TAOS shuttle, I think it was by either TRW or Boeing.
OT a bit but "TAOS shuttle"? I did a search and didn't find anything but the acronym in other posts?

Randy
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/07/2011 01:02 pm
It was while I was researching the low cost pressure fed booster concepts for the TAOS shuttle, I think it was by either TRW or Boeing.
OT a bit but "TAOS shuttle"? I did a search and didn't find anything but the acronym in other posts?

Randy

Thrust Assisted Orbiter Shuttle (TAOS).  This is what the configuration of the shuttle was called after "deletion" of the flyback booster.
Title: Re: Rocket Engine Q&A
Post by: Downix on 07/07/2011 04:06 pm
It was while I was researching the low cost pressure fed booster concepts for the TAOS shuttle, I think it was by either TRW or Boeing.
OT a bit but "TAOS shuttle"? I did a search and didn't find anything but the acronym in other posts?

Randy
Jim put down the anacronym's meaning, but basically TAOS is the concept which lead to the shuttle we have today.
Title: Re: Rocket Engine Q&A
Post by: RanulfC on 07/07/2011 04:46 pm
Thank you both :)

Randy
Title: Re: Rocket Engine Q&A
Post by: Jason1701 on 02/25/2012 06:49 pm
What is the temperature of the exhaust gases at the nozzle exit of a hydrolox engine such as the RL-10?
Title: Re: Rocket Engine Q&A
Post by: Antares on 02/25/2012 09:48 pm
NEI.  You need to specify expansion ratio and mixture ratio.  There are flow solvers on the web to determine the exit temperature.  There are probably combustion solvers to determine the initial temperature.
Title: Re: Rocket Engine Q&A
Post by: Jason1701 on 02/26/2012 12:25 am
There are flow solvers on the web to determine the exit temperature.

Thanks, do you recommend any?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 02/26/2012 07:23 am
There are flow solvers on the web to determine the exit temperature.

Thanks, do you recommend any?

CEA (http://www.grc.nasa.gov/WWW/CEAWeb/) is pretty easy to use, you will need to know some essentials of the engine, but they should be easy to find. The online version is a bit buggy, but it will probably work for your purpose.

To answer you more directly. An RL-10A-4-2, the engine on the Atlas V upper state, operating at a mixture ratio of 5.5, will be a little above 1000 degrees Fahrenheit at the exit. That is as opposed to 6000-7000 in the chamber.
Title: Re: Rocket Engine Q&A
Post by: Jason1701 on 02/26/2012 07:43 am
 
There are flow solvers on the web to determine the exit temperature.

Thanks, do you recommend any?

CEA (http://www.grc.nasa.gov/WWW/CEAWeb/) is pretty easy to use, you will need to know some essentials of the engine, but they should be easy to find. The online version is a bit buggy, but it will probably work for your purpose.

To answer you more directly. An RL-10A-4-2, the engine on the Atlas V upper state, operating at a mixture ratio of 5.5, will be a little above 1000 degrees Fahrenheit at the exit. That is as opposed to 6000-7000 in the chamber.

Thanks. Using the calculator, I found that the RL-10B has an exhaust temperature of about 550 C.
Title: Re: Rocket Engine Q&A
Post by: Antares on 02/26/2012 06:52 pm
Keep in mind that's isenthalpic.  There was a video a few years ago of some engine with icicles hanging off of the nozzle due to the regenerative cooling while it was firing.
Title: Re: Rocket Engine Q&A
Post by: ugordan on 02/26/2012 07:02 pm
There was a video a few years ago of some engine with icicles hanging off of the nozzle due to the regenerative cooling while it was firing.

CECE.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 02/27/2012 04:51 am
Keep in mind that's isenthalpic.  There was a video a few years ago of some engine with icicles hanging off of the nozzle due to the regenerative cooling while it was firing.

Well, and quasi-1D. Probably still close for what the core static temperature is in the exhaust jet though.
Title: Re: Rocket Engine Q&A
Post by: ARD on 03/05/2012 12:34 am
Has there ever been a rocket engine burning hydrazine (N2H4, no methyl groups) in N2O4?  If so, which was it, what was its specific impulse, and what was the oxidizer:fuel ratio?  Every other engine seems to be Aerozine 50 or UDMH or MMH, but I haven't found any bipropellant engine that burns only hydrazine. 
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/05/2012 01:46 am
Has there ever been a rocket engine burning hydrazine (N2H4, no methyl groups) in N2O4?  If so, which was it, what was its specific impulse, and what was the oxidizer:fuel ratio?  Every other engine seems to be Aerozine 50 or UDMH or MMH, but I haven't found any bipropellant engine that burns only hydrazine. 

Juno
Title: Re: Rocket Engine Q&A
Post by: Proponent on 03/05/2012 02:21 am
I thought that was hydyne (https://en.wikipedia.org/wiki/Hydyne) (with lox).
Title: Re: Rocket Engine Q&A
Post by: ARD on 03/05/2012 02:28 am
Has there ever been a rocket engine burning hydrazine (N2H4, no methyl groups) in N2O4?  If so, which was it, what was its specific impulse, and what was the oxidizer:fuel ratio?  Every other engine seems to be Aerozine 50 or UDMH or MMH, but I haven't found any bipropellant engine that burns only hydrazine. 

Juno

Thank you.  I take it you refer to the Leros 1b engine on the Juno spacecraft.  I initially confused it for the satellite launcher.  Once again, thank you for your answer. 
Title: Re: Rocket Engine Q&A
Post by: strangequark on 03/11/2012 11:04 pm
Has there ever been a rocket engine burning hydrazine (N2H4, no methyl groups) in N2O4?  If so, which was it, what was its specific impulse, and what was the oxidizer:fuel ratio?  Every other engine seems to be Aerozine 50 or UDMH or MMH, but I haven't found any bipropellant engine that burns only hydrazine. 

The HiPAT DM (http://www.astronautix.com/graphics/h/hipatdua.jpg) is another example. Straight hydrazine's a little more finicky, but it's nice when your monoprop thrusters can pull from the same fuel supply.

Also, I can’t remember the name of the engine, but TRW had a true dual-mode engine, capable of operating in monoprop mode on hydrazine, or biprop mode with added NTO/MON.
Title: Re: Rocket Engine Q&A
Post by: alirezame on 03/13/2012 08:25 am
hi
I'm new in this forum
I'm wonder in starting the liquid rocket engines in vac, i'm searching on the net and don't find anything useful if anyone have searched or worked on this  ,please answers me or give me useful references.
thanks
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/13/2012 09:36 am
There are many methods depending on propellents.
H2 O2 use spark plug type devices
RP-1 uses hypergolic fluid, start cartridges, spark plugs, etc
Hypergolic needs nothing

Same applies for booster stages



The sparklers seen before shuttle launches do not start the engines
Title: Re: Rocket Engine Q&A
Post by: alirezame on 03/13/2012 09:45 am
 Thanks a a lot Jim, for your attention
i'm searching on this items, i have an other question in this filed
what do you think about ullage solid motors to make a high pressure gas for pressurization in tanks on vaccum condition, is it used in launch vehicles or that's concept only?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/13/2012 09:52 am
The Russians use a little of each hypergolic component to pressurize the other components tank.

A.  How would you gather the gas , it is needed for propulsion
B.  ullage rockets burn a sort time, pressuziation is needed mostly in the latter portion of the stage bun.
C.  Many stages don't use ullage rockets.  I believe no major US vehicle does at this time
Title: Re: Rocket Engine Q&A
Post by: alirezame on 03/13/2012 10:04 am
I think method of using a hypergolic component for pressurization had named chemical  pressurization, which never flown , it means you have combustion on the head of your propellant
I convince that ullage motors used for providing artificial gravity on vacuum condition to make  a minimum static head for injection the medium to the inlet of pump, what do you think ?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/13/2012 01:02 pm
1.  Yes, chemical pressization has flown
2. No, using thrusters or settling jets (they are not ullage) is not feasible.  How would you tap off of it?  Also, its products might not be compatable with the propellants.
Title: Re: Rocket Engine Q&A
Post by: spacecane on 05/02/2012 01:11 am
What is the "thermal efficiency" of a RP-1/LOX rocket engine at sea level?  Basically, what percentage of the energy in the fuel is turned into usable thrust?  I think for the SSME's I've seen something crazy like 99% (different fuel of course).
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 05/02/2012 01:20 am
What is the "thermal efficiency" of a RP-1/LOX rocket engine at sea level?  Basically, what percentage of the energy in the fuel is turned into usable thrust?  I think for the SSME's I've seen something crazy like 99% (different fuel of course).
You mean the efficiency of turning the energy in the propellant into directed kinetic energy of the exhaust, I believe.
Title: Re: Rocket Engine Q&A
Post by: spacecane on 05/02/2012 01:27 am
What is the "thermal efficiency" of a RP-1/LOX rocket engine at sea level?  Basically, what percentage of the energy in the fuel is turned into usable thrust?  I think for the SSME's I've seen something crazy like 99% (different fuel of course).
You mean the efficiency of turning the energy in the propellant into directed kinetic energy of the exhaust, I believe.

Yes, exactly.
Title: Re: Rocket Engine Q&A
Post by: spacecane on 05/06/2012 06:09 pm
Another question, what causes the appearance of the exhaust for various engines?

I've read that the blue cone on the SSME is caused by the extremely hot steam going into a plasma state.  Somebody on the forum told me that the RS-68 has the color it does due to the ablative nozzle.  What causes the specific appearance in other types of engines?

I assume that the incredibly bright and long Solid Rocket exaust plume is from extremely hot glowing particles of aluminum oxide?
Title: Re: Rocket Engine Q&A
Post by: gin455res on 05/13/2012 07:13 am
Are there any designs for micro nuclear-thermal rockets, using radio-isotopes in place of reactor cores?

Title: Re: Rocket Engine Q&A
Post by: Jim on 05/13/2012 12:00 pm
Are there any designs for micro nuclear-thermal rockets, using radio-isotopes in place of reactor cores?



not hot enough
Title: Re: Rocket Engine Q&A
Post by: Proponent on 05/13/2012 01:03 pm
And even if you could make the RTG hot enough, the specific power is way too low to be attractive.  Consider a 1-kg kerolox engine.  At typical kerolox thrust-to-weight ratios, it could easily generate 1000 N of thrust.  If the exhaust velocity is 3000 m/s, then the mass flow rate through the engine is (1000 N)/(3000 m/s) = 0.333 kg/s.  The mechanical power in the exhaust is then 0.5*(0.333 kg/s)*(3000 m/s)^2 = 1.5 MW, giving a power per unit engine mass of 1.5 MW/kg.  RTGs might produce something like 1 kW of heat per kilogram of plutonium oxide.  So, even if you ignore all the other mass that would be required, the specific power of the isotope-driven engine would be a thousand times less than for a conventional chemical engine.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 05/13/2012 03:20 pm
And even if you could make the RTG hot enough, the specific power is way too low to be attractive.  Consider a 1-kg kerolox engine.  At typical kerolox thrust-to-weight ratios, it could easily generate 1000 N of thrust.  If the exhaust velocity is 3000 m/s, then the mass flow rate through the engine is (1000 N)/(3000 m/s) = 0.333 kg/s.  The mechanical power in the exhaust is then 0.5*(0.333 kg/s)*(3000 m/s)^2 = 1.5 MW, giving a power per unit engine mass of 1.5 MW/kg.  RTGs might produce something like 1 kW of heat per kilogram of plutonium oxide.  So, even if you ignore all the other mass that would be required, the specific power of the isotope-driven engine would be a thousand times less than for a conventional chemical engine.

Thanks for the analysis.

I suspected it would be very low thrust.  And your analysis assumes 3000m/s, which understates the problem, as this concept would only make any sense if it was more fuel efficient than chemical propulsion, which would suggest higher exhaust velocities. And you also lose the oberth effect unless it is pulsed at perigee.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 05/20/2012 02:55 pm
Does the output of the gas Generator chokes? I.e. does the gas generator have a throat, or is it sort of "straight" output to the tubopump? I thought that if you could choke you could have less problems with back pressure, right?
Title: Re: Rocket Engine Q&A
Post by: STS-200 on 05/21/2012 09:50 am
The gas generator itself doesn't have a throat. However, a throat is almost invariably formed by the first stage stator vanes of the turbine, which twist and accelerate the gas to supersonic speeds.

In most designs, the turbine rotor blades are not operating in a supersonic regime, as the speed of the blades is sufficient to lower the relative motion to typically Mach 0.8-0.9. Supersonic turbine designs do exist, but they tend to be lower efficiency.

Title: Re: Rocket Engine Q&A
Post by: baldusi on 05/21/2012 05:39 pm
The gas generator itself doesn't have a throat. However, a throat is almost invariably formed by the first stage stator vanes of the turbine, which twist and accelerate the gas to supersonic speeds.

In most designs, the turbine rotor blades are not operating in a supersonic regime, as the speed of the blades is sufficient to lower the relative motion to typically Mach 0.8-0.9. Supersonic turbine designs do exist, but they tend to be lower efficiency.
So the flux on the GG interacts directly with the first stator of the TP, right? Wouldn't a throat help avoiding this interaction?
Title: Re: Rocket Engine Q&A
Post by: DMeader on 05/21/2012 06:51 pm
The purpose of the stator blades (nozzles) before the first turbine stage is to control and direct the flow from the gas generator into the turbine blades at the optimum angle, and since they form convergent ducts, they increase the velocity of the flow. I don't think you'd want any constriction between the gas generator and the turbine.
Title: Re: Rocket Engine Q&A
Post by: STS-200 on 05/22/2012 08:35 am
The purpose of the stator blades (nozzles) before the first turbine stage is to control and direct the flow from the gas generator into the turbine blades at the optimum angle, and since they form convergent ducts, they increase the velocity of the flow. I don't think you'd want any constriction between the gas generator and the turbine.

Yes, thats right.  A choke (or any major constriction like a sharp bend) between the GG and the turbine would only cause shocks or turbulence which would reduce the pressure of the flow. 
Title: Re: Rocket Engine Q&A
Post by: baldusi on 05/23/2012 11:52 am
The other question is if the Gas Generator is regeneratively cooled. Would cooling the Gas Generator lower its pressure? I guess you could decrease the ID of the tube as you cool and keep the pressure? Or is is straight? I'm trying to get a grasp of the trade offs.
The Gas Generator generates power, but you have to be able to run it through the turbopump without undue degradation of the turbo blades, so you end up being temperature limited.
Let's say that the maximum temperature allowable by the turbopump is Km. And let's say that you chose a O/F that gives you that Km temperature. So, if you use a new mixture that give Kl>Km, but you lower the temperature of the GG output by way of passing cryo fluid around the output jacket until it's Km, would the pressure be the same for both outputs? Would you be using more or less mass? In other words, would there be any advantage? Or the GG->Turbo is simply a game of materials?
Title: Re: Rocket Engine Q&A
Post by: STS-200 on 05/23/2012 02:19 pm
There is little point in cooling the GG, as you cannot use temperatures which are high enough to need it. GG temperature is 600-1000K, which can be handled without active cooling.

Regen cooling will have little effect on gas temperature. Varying mixture ratio varies the gas temperature, and the GG will run at whatever pressure the pumps can supply. Higher pressures and/or temperatures allow you to extract more work from the gas.

As you say, the limiting factor is the strength of the turbine blades, which will be nearly as hot as the gas flowing over them.

Jet engines get round this by using various types of hollow blades with coolant (air) running through them and out of them. In a rocket turbine this isn't possible as the blades are much more heavily loaded - they have to be stronger and therefore can't be hollow. Blade materials are also stronger when cool. For these reasons rocket GGs have to run at lower temperature than jet engine CCs.

Title: Re: Rocket Engine Q&A
Post by: joertexas on 06/14/2012 05:50 pm
I was curious about the CECE's suitability as a space vehicle engine, such as a earth-moon orbital transfer vehicle. The literature says the engine is good for 10,000 seconds and 50 restarts. How difficult would it be - for that engine, or one with similar specifications, to extend it's service life? Also, what components would be the most impacted by such a change?

Thanks

JR
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/15/2012 12:44 pm
Another question about Gas Generator. If you expand the output, would it lower the temperature at a higher volume and thus allow more mass flow at the limiting temperature, or would you be loosing that at the camber side of the equation and thus not earn much overall performance? I should have studied turbo machinery  :-[
Title: Re: Rocket Engine Q&A
Post by: STS-200 on 06/15/2012 03:25 pm
Another question about Gas Generator. If you expand the output, would it lower the temperature at a higher volume and thus allow more mass flow at the limiting temperature, or would you be loosing that at the camber side of the equation and thus not earn much overall performance? I should have studied turbo machinery  :-[

Sorry, I'm not sure I understand your question - at what stage are you "expanding the output"?

General answer:
- Expanding the gas other than in the turbine is a waste.
- If you want more mass flow, you need to either raise the pressure or make the unit bigger.

If you are mathematically minded this book is quite good:
"Jet Propulsion" ISBN-13: 978-0521541442

I think there is also a chapter on it in "Rocket Propulsion Elements"
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/15/2012 03:44 pm
Another question about Gas Generator. If you expand the output, would it lower the temperature at a higher volume and thus allow more mass flow at the limiting temperature, or would you be loosing that at the camber side of the equation and thus not earn much overall performance? I should have studied turbo machinery  :-[

Sorry, I'm not sure I understand your question - at what stage are you "expanding the output"?

General answer:
- Expanding the gas other than in the turbine is a waste.
- If you want more mass flow, you need to either raise the pressure or make the unit bigger.

If you are mathematically minded this book is quite good:
"Jet Propulsion" ISBN-13: 978-0521541442

I think there is also a chapter on it in "Rocket Propulsion Elements"

I thought increasing the ID of the GG output would cool the gases due to adiabatic expansion. But I guess you can do that with a bigger stator on the turbine, right?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 06/15/2012 09:44 pm
I thought increasing the ID of the GG output would cool the gases due to adiabatic expansion. But I guess you can do that with a bigger stator on the turbine, right?

No actually. First off, there's a distinction you need to know:

Stagnation temperature: The temperature of the flow if it were brought from current velocity down to zero velocity (stagnation) isentropically.

Static temperature: The temperature of the flow while in motion, as defined by the average molecular velocity but excluding the bulk velocity of the fluid.

For an isentropic subsonic flow, expanding the flow area decreases the velocity, and raises the static temperature. Stagnation temperature is the same, because you are trading kinetic energy for thermal energy.

With that said, to a first order, the components exposed to the hot flow "see" the stagnation temperature. Local flow velocity approaches zero in the fluid layer close to a hard surface, and so the local temperature approaches the stagnation temperature (the reason this is not totally true is because the real flow isn't adiabatic, especially at the chamber/casing walls). So, what really impacts your heat transfer to the turbomachinery is the initial combustion temperature (which is basically the stagnation temperature). That can't be changed with fluid mechanic trickery, only by transferring heat out of the system, which means less available energy in your flow.

Now, on a tangent, there have been proposed cycles where you burn, run through a single turbine stage (which removes energy, dropping stagnation temperature), then burn again (raising stagnation temperature back to allowable limits), and go through the next turbine stage. However, there are weight and complexity issues with that kind of configuration.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/15/2012 10:19 pm
So it's the only way to allow higher temperature (and power) is, basically, by using more resistant materials, right? So the trickery is on the layering of the turbopump blades? If you could, for example add a layer of a very good heat conductor to avoid heat spikes on certain zones, or putting a stronger but less hear resistant material on the core and adding a heat insulator and a heat conductor on top of that? Probably composite turbine blades are one f the holy grails of turbo machinery, right? Or do you see the future on exoeskeletal engines?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 06/15/2012 10:36 pm
So it's the only way to allow higher temperature (and power) is, basically, by using more resistant materials, right? So the trickery is on the layering of the turbopump blades? If you could, for example add a layer of a very good heat conductor to avoid heat spikes on certain zones, or putting a stronger but less hear resistant material on the core and adding a heat insulator and a heat conductor on top of that? Probably composite turbine blades are one f the holy grails of turbo machinery, right? Or do you see the future on exoeskeletal engines?

Oh, that is outside my paygrade ;). But yes, materials science is definitely the limiting factor (and has been). I know that on the jet engine side, there is considerable interest in ceramics, with the problem being that they are less resistant to the mechanical environments. There's also a point of diminishing returns. Turbomachinery manufacturing is already a complex art, and a major driver of expense. Do you really want to increase manufacturing complexity by an order of magnitude to squeeze out a mild performance benefit? In rocket engines, there are schemes that offer more benefit for less crazy development (full flow staged combustion, for instance).

From a systems level too, you might be better off investing your money in improved structures to reduce vehicle inert mass (i.e. the Boeing 787's fuel efficiency is mostly due to structure).
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/16/2012 05:51 pm
I part yes. But I still believe that the secret is to invest in the crazy technologies that would make that manufacturing an order of magnitude cheaper. Like the EBM that allowed new structures manufacturing (still not cheap, though).
The day somebody develops a DMLS that can work around small ceramic/carbon blocks, making crazy stiff turbo machinery might get ridiculously cheap. And that's something that every turbo machine can take advantage of.
In any case, I find it interesting that there's still some very obvious architectural improvements to rocket engines (like full stage), that haven't be tried, since is still cheaper to improve other parts. In that sense, it would seem that the low scale of the market has left a lot of margin to grow.
Title: Re: Rocket Engine Q&A
Post by: e of pi on 06/21/2012 03:25 pm
I have a question about how rocket engine chamber pressure improvments relate to potential ISp and thrust improvements. Looking at it, both the compressible and incompressible forms of Bernoulli's princible seem to suggest that exhaust velocity should rise in proportion to the square root of the pressure increase (i.e. if the pressure ratio is 1.5, then the exhaust velocity and ISp should see a ratio of 1.22).

However, looking at some instances of this in the real world, like the Merlin 1C-1D which is supposed to go from 6.77 MPa to 9.7 MPa, which is a ratio of 1.43. My understanding is that this would be an increase of 19% in ISp, but the increase is more like 2%: 304 to 310s vacuum. I see something similar looking at Astronautix numbers for the change from H-1 to RS-27, so clearly I'm not fully accounting for something critical. Is this that I've mis-understood or mis-applied the theory, or are there real world complications that the theory isn't accounting for? 
Title: Re: Rocket Engine Q&A
Post by: Antares on 06/22/2012 03:35 am
Isp is a weak but non-zero function of chamber pressure.  Combustion temperature is a stronger contributor to Isp.  I'll see if I can find the equations at work tomorrow and edit this.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 06/22/2012 01:10 pm
(i.e. the Boeing 787's fuel efficiency is mostly due to structure).
On the 787, a big chunk of those efficiency improvements are the engines. It is the engines under performing that have caused the efficiency headaches, not the plane being overweight. The composite structure just reduced mass, allowing higher fuel loads and/or larger payloads.

Just look at the 747-8's and the jump it took with the same engines.

(btw. This discussion can easily spin out of control, since the 787 was really optimized for long Pacific routes and doesn't offer as great a leap forward (compared to the 767) on shorter routes. Though to be fair, ANA's only complaint to date is the window tint is not dark enough).

Of course, like everyone else I can't wait to see the 787-9 with it's laminar tail :-)
Title: Re: Rocket Engine Q&A
Post by: sdsds on 07/02/2012 12:37 am
Recently staged-combustion methane has received considerable attention, but for upper stage use how would an expander cycle methane engine compare? Expander cycle works for RL10, arguably the best existing upper stage engine. Is there some property of methane (vs. hydrogen) that makes it less suited for an expander cycle?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/02/2012 05:07 am
Recently staged-combustion methane has received considerable attention, but for upper stage use how would an expander cycle methane engine compare? Expander cycle works for RL10, arguably the best existing upper stage engine. Is there some property of methane (vs. hydrogen) that makes it less suited for an expander cycle?

An expander will have a lower T/W. The available power is limited, which means expanders have lower chamber pressures (this is part of why they are largely relegated to upper stage use). Methane has a lower specific heat than hydrogen, which makes it less effective at heat transfer. However, it's still been done. Admittedly, I haven't run through the calcs, but I think you'd have to run at a lower pressure, or maybe a faster flow (higher pressure drop) in the nozzle/chamber regen section.

The benefit of expander is that it's simple, and the turbine drive gases are as gentle as a kitten on the turbomachinery (RL-10 isn't far from room temperature hydrogen). The downside is you will have a much larger engine for the same thrust level.

Now, deviating on that topic somewhat, there has been some work done on a combined expander/staged cycle. The idea is that you still get a substantial amount of your energy from the regen cooling of the nozzle and chamber, but then you dump a little bit of oxidizer into the flow to bump up the temperature to maybe 400-500 Kelvin (As opposed to 900-1200K).
Title: Re: Rocket Engine Q&A
Post by: baldusi on 07/02/2012 11:41 am
The Japanese have their bleed expander cycle, which is replacing the GG of the open cycle with the expander, but still dumping the gases after they've passed through the turbine, in an open cycle. Apparently they can take it to something like 1,500kN. I guess it won't have good T/W, then?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/02/2012 04:13 pm
The Japanese have their bleed expander cycle, which is replacing the GG of the open cycle with the expander, but still dumping the gases after they've passed through the turbine, in an open cycle. Apparently they can take it to something like 1,500kN. I guess it won't have good T/W, then?

Do you have a reference for that? I know about the LE-5B, but that's RL-10 class. I have heard about expander bleeds hypothetically going to higher thrust (which makes sense, you can have a much higher pressure ratio across the turbine), but I'm not aware of any high thrust production engines that use expander bleed.

As for T/W, LE-5B is about 50, which is in line with RL-10.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 07/02/2012 04:32 pm
Can't really remember where did I found this papers. I think it was Mitsubishi Heavy Industries Annual Technology discussion, or something like that. Please bear in mind that this is open cycle expander, not close cycle.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/29/2012 08:32 pm
How much of an engine's throttlability is due to nozzle constraints and how much is due to the pump?

And can multi-chambered engines with a single pump, stop the flow to individual chambers to increase throttlability (assuming a pump has a significantly wider power range than a single nozzle can except without the flow separating)?
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 07/30/2012 07:40 pm
It's usually the injector design and the chamber design that affect throttleability. The nozzle doesn't have much to do with it.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 07/30/2012 09:09 pm
Not over-expansion?

cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Fequalsma on 08/01/2012 01:37 am
Here's a talk by Al Miller, Boeing director for  787 Technology Integration
http://www.uwtv.org/video/player.aspx?mediaid=16215102

Most of the 787 performance improvement is due to the powerplant, much more than the structures.


From a systems level too, you might be better off investing your money in improved structures to reduce vehicle inert mass (i.e. the Boeing 787's fuel efficiency is mostly due to structure).

Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 08/03/2012 05:06 pm
Not over-expansion?

cheers, Martin
Ah, true.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 10/23/2012 09:44 pm
If the problem of using cooling on the turbo blades is that they suffer too much stress to tolerate the channels. As the size growths, does the law of surface to mass makes it possible to use channels? Or are the effect of strength non scalability bigger?
Title: Re: Rocket Engine Q&A
Post by: Antares on 10/24/2012 12:52 am
Resonant frequencies in the turbopumps can also affect throttleability.

Nozzle expansion plays very little into throttleability.  You want the engine at full blast anywhere that would be a problem.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 11/30/2012 07:21 am
When designing a catalyst pack for decomposing peroxide or hydrazine, is there any attention given to the aerodynamics of the flow?

For example, would a tapering, or waisted, pack act like a venturi. That is, would it reduce the pressure in the flow. Reducing pressure in a reaction thats products have more mols (and consequently more volume) than the starting reactants is an additional mechanism (additional to the catalysis on the pack surface) to drive the reaction in a forwards direction.

And, can we even assume that mixed phase flows obey the Bernoulli equation that explained the venturi effect?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 11/30/2012 07:25 pm
When designing a catalyst pack for decomposing peroxide or hydrazine, is there any attention given to the aerodynamics of the flow?

For example, would a tapering, or waisted, pack act like a venturi. That is, would it reduce the pressure in the flow. Reducing pressure in a reaction thats products have more mols (and consequently more volume) than the starting reactants is an additional mechanism (additional to the catalysis on the pack surface) to drive the reaction in a forwards direction.

And, can we even assume that mixed phase flows obey the Bernoulli equation that explained the venturi effect?

That's undesirable. Reduced pressure reduces the reaction rate, and increased velocity decreases the residence time. You'd have to have a much longer catalyst bed. There's also a loss of total pressure with higher flow rates.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 11/30/2012 09:04 pm
Why isn't oxidizer (LOX) never used to regeneratively cool the camber? It has a good deal more density to absorb the heat. Is it because it has bigger pressure losses? Because the risk of phase change?
The vortex method does use it, obviously. But has none of those issues.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/30/2012 09:08 pm
LOX cooling has been used experimentally, and apparently it's not as dangerous as it sounds, even if there are cracks. I came across a paper describing this on NTRS once. Skylon's SABRE is also meant to use LOX cooling.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 11/30/2012 09:17 pm
LOX cooling has been used experimentally, and apparently it's not as dangerous as it sounds, even if there are cracks. I came across a paper describing this on NTRS once. Skylon's SABRE is also meant to use LOX cooling.
I thought they use a He loop.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/30/2012 09:25 pm
That's for cooling the incoming air, it would use LOX for cooling the nozzle, as the LH2 is already used to cool the helium.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 11/30/2012 09:32 pm
Why isn't oxidizer (LOX) never used to regeneratively cool the camber? It has a good deal more density to absorb the heat. Is it because it has bigger pressure losses? Because the risk of phase change?
The vortex method does use it, obviously. But has none of those issues.

Specific heat capacity is more important. Fuels, and LH2 especially, tend to have better specific heats. LH2 is something like 10 kJ/kg-K, whereas LOX is more like 1. So, for an engine with a 5.5 O/F ratio, the much smaller quantity of hydrogen can handle about twice the heat flux as the oxygen for the same temperature change. That, coupled with the additional aggressiveness of LOX/GOX toward your materials. Not impractical though, if there's enough benefits. Been proposed for plenty of engines, though they still tend to use the fuel first, and oxidizer as a supplement (i.e. fuel cooled combustion chamber, oxygen cooled nozzle).
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/01/2013 04:41 pm
Are there many examples of engines that were designed to burn hypergolic propellants that were converted to burn cryogenic prop instead (like LOX/Kerosine)? Is there an improvement in ISP burning LOX instead of Nitric Acid or N2O4?
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 03/04/2013 01:02 pm
Well, Titan went the other direction ... I believe both used variants of the LR-87/LR-91 for the first and second stages.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/11/2013 05:17 pm
For a vehicle with a single turbopump powered engine, do some open cycle systems use the exhaust of the turbopump for roll control or steering?
 
Title: Re: Rocket Engine Q&A
Post by: baldusi on 03/11/2013 05:27 pm
For a vehicle with a single turbopump powered engine, do some open cycle systems use the exhaust of the turbopump for roll control or steering?
 
You mean besides the Falcon 1/9?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/11/2013 06:16 pm
For a vehicle with a single turbopump powered engine, do some open cycle systems use the exhaust of the turbopump for roll control or steering?
 
Delta IV CBC, Titan second stage, Jupiter missile....
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/11/2013 06:21 pm
What are the requirements/constraints for using turbopump exhaust for roll control or steering - are there some systems that do not do this, and why not?

Is the issue that some turbopumps do not produce sufficient exhaust to perform the steering function?
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/11/2013 06:27 pm
What are the requirements/constraints for using turbopump exhaust for roll control or steering - are there some systems that do not do this, and why not?

Is the issue that some turbopumps do not produce sufficient exhaust to perform the steering function?


It depends on many things, like moments of inertia of the vehicle, the aero disturbances it will see, the engine will have some effects, etc.  Delta IV with SRM's need the SRM's to have gimballing, since slight thrust misalignments will over power the roll nozzle.
Title: Re: Rocket Engine Q&A
Post by: a_langwich on 03/11/2013 08:41 pm
When designing a catalyst pack for decomposing peroxide or hydrazine, is there any attention given to the aerodynamics of the flow?

For example, would a tapering, or waisted, pack act like a venturi. That is, would it reduce the pressure in the flow. Reducing pressure in a reaction thats products have more mols (and consequently more volume) than the starting reactants is an additional mechanism (additional to the catalysis on the pack surface) to drive the reaction in a forwards direction.

And, can we even assume that mixed phase flows obey the Bernoulli equation that explained the venturi effect?

That's undesirable. Reduced pressure reduces the reaction rate, and increased velocity decreases the residence time. You'd have to have a much longer catalyst bed. There's also a loss of total pressure with higher flow rates.

I believe gin455res was correct:  reduced pressure can increase the reaction rate, for the kinds of reactions that were specified.

But the increased reaction rate will likely not be enough to compensate for the increased velocity, so the rest of your response applies.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/12/2013 01:49 am
What are the requirements/constraints for using turbopump exhaust for roll control or steering - are there some systems that do not do this, and why not?

Is the issue that some turbopumps do not produce sufficient exhaust to perform the steering function?


It depends on many things, like moments of inertia of the vehicle, the aero disturbances it will see, the engine will have some effects, etc.  Delta IV with SRM's need the SRM's to have gimballing, since slight thrust misalignments will over power the roll nozzle.

In the earliest days of liquid fueled rocketry, rocket engines used vanes and the like for steering, rather than the approach of using the turbopump exhaust, at least until the Viking program; was this because turbopump exhaust steering hadn't been invented yet, or because the turbopumps of the day were not sufficiently powerful, or at least their exhaust was not sufficiently powerful, to function as part of the guidance system?

What was the first major rocket system to use turbopump exhaust for steering?

Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 03/12/2013 07:17 am
Rocket vanes are pretty easy to make. Easier than gimbaling a nozzle of any type.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 03/12/2013 02:18 pm
There was recently two threads on the history of gimballing:

http://forum.nasaspaceflight.com/index.php?topic=31053.0

http://forum.nasaspaceflight.com/index.php?topic=31059.0

Note worthy quote from Jim: http://forum.nasaspaceflight.com/index.php?topic=31059.msg1010377#msg1010377
Quote
Verniers came about for Thor and Atlas because even with gimbaling, single engine vehicles need a method of roll control.
Which implies that turbo-pump roll control had not been developed yet.
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/12/2013 02:49 pm
There was recently two threads on the history of gimballing:

http://forum.nasaspaceflight.com/index.php?topic=31053.0

http://forum.nasaspaceflight.com/index.php?topic=31059.0

Note worthy quote from Jim: http://forum.nasaspaceflight.com/index.php?topic=31059.msg1010377#msg1010377
Quote
Verniers came about for Thor and Atlas because even with gimbaling, single engine vehicles need a method of roll control.
Which implies that turbo-pump roll control had not been developed yet.

See Jupiter ;-).  For Thor and Atlas, the term verniers was applicable since they did provide final adjustment to the vehicle for targeting.

As a cost savings, Atlas II eliminated them and went to a hydrazine thruster system for roll control during the sustainer phase.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 03/12/2013 04:45 pm
Really... That's a very interesting, so I assume switching to hydrazine thrusters led to a less complex and cheaper system. I assume this means that all three engines gimbaled on the Atlas II.

Ed would know, but considering how successful the Atlas II was, I would assume they never caused a LOM.
Title: Re: Rocket Engine Q&A
Post by: yinzer on 03/13/2013 12:45 am
I think all three engines gimbaled on all Atlas variants.

The Atlas ICBM needed to precisely control burnout velocity so it wanted verniers anyway, and once you have them you may as well use them for roll control too.

Once you have a second stage you can use that to precisely control velocity so you don't need the verniers any more.  But if you are going to fly with only one engine you need something for roll control and keeping the verniers around is easier than deleting them.

Replacing the gimbaled biprop verniers with some fixed hydrazine monoprop thrusters should simplify and reduce recurring costs, so it makes sense that they got around to it.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/13/2013 02:05 pm
There was recently two threads on the history of gimballing:

http://forum.nasaspaceflight.com/index.php?topic=31053.0

http://forum.nasaspaceflight.com/index.php?topic=31059.0

Note worthy quote from Jim: http://forum.nasaspaceflight.com/index.php?topic=31059.msg1010377#msg1010377
Quote
Verniers came about for Thor and Atlas because even with gimbaling, single engine vehicles need a method of roll control.
Which implies that turbo-pump roll control had not been developed yet.

I assume that on those older vehicles, the turbopump exhaust was simply vented downwards, in parallel with the bell nozzle exhaust; this begs the question as to whether the turbopump exhaust provides any useful thrust in this mode, or is this like asking whether the tailpipe on a car provides any useful forward thrust?

Title: Re: Rocket Engine Q&A
Post by: Jim on 03/13/2013 02:26 pm

I assume that on those older vehicles, the turbopump exhaust was simply vented downwards, in parallel with the bell nozzle exhaust; this begs the question as to whether the turbopump exhaust provides any useful thrust in this mode, or is this like asking whether the tailpipe on a car provides any useful forward thrust?


yes, but on a relative scale.   The F-1 turbopump produced 30klb.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/13/2013 05:17 pm

I assume that on those older vehicles, the turbopump exhaust was simply vented downwards, in parallel with the bell nozzle exhaust; this begs the question as to whether the turbopump exhaust provides any useful thrust in this mode, or is this like asking whether the tailpipe on a car provides any useful forward thrust?


yes, but on a relative scale.   The F-1 turbopump produced 30klb.

ie 2% of the engine thrust was produced by the turbopump exhaust.

My dim recollection is that using vanes in the exhaust for steering costs 5% of overall engine performance, so using turbopump exhaust might not have been useful in replacing vanes for vehicles like Redstone - not quite enough output from the turbopump.
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/13/2013 05:26 pm

My dim recollection is that using vanes in the exhaust for steering costs 5% of overall engine performance, so using turbopump exhaust might not have been useful in replacing vanes for vehicles like Redstone - not quite enough output from the turbopump.


Redstone used the vanes for pitch and yaw too.  Also, the Redstone turbopump was powered by H2O2. 
Title: Re: Rocket Engine Q&A
Post by: baldusi on 03/13/2013 07:48 pm

My dim recollection is that using vanes in the exhaust for steering costs 5% of overall engine performance, so using turbopump exhaust might not have been useful in replacing vanes for vehicles like Redstone - not quite enough output from the turbopump.


Redstone used the vanes for pitch and yaw too.  Also, the Redstone turbopump was powered by H2O2. 
So is the Soyuz's btw.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/14/2013 05:24 pm
I gather from this that only turbopumps powered by propellant products can be useful for roll control, and that turbopump exhaust cannot be used for pitch or yaw control under any condition.
Title: Re: Rocket Engine Q&A
Post by: Jim on 03/14/2013 05:53 pm
I gather from this that only turbopumps powered by propellant products can be useful for roll control, and that turbopump exhaust cannot be used for pitch or yaw control under any condition.


Depends on the control authority required.  From history, it doesn't look like it can
Title: Re: Rocket Engine Q&A
Post by: baldusi on 03/14/2013 05:55 pm
I gather from this that only turbopumps powered by propellant products can be useful for roll control, and that turbopump exhaust cannot be used for pitch or yaw control under any condition.
I meant that the RD-107/8 also used H2O2 to drive the turbopump. The steering engines appear to have their own injectors for RG-1/LOX, but are also fed from the same turbopump.
Look here:
(http://www.lpre.de/energomash/RD-107/img/flow_diagram.jpg)
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/16/2013 09:50 pm
Are there many examples of engines that were designed to burn hypergolic propellants that were converted to burn cryogenic prop instead (like LOX/Kerosine)? Is there an improvement in ISP burning LOX instead of Nitric Acid or N2O4?


I found the answer to my own question - the German Wasserfall engine burned storable components, but was converted to cryogenic oxidizer and alcohol (or maybe Kerosene) for the Viking rocket program, and then on to Vanguard.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 03/17/2013 03:20 am
You know what is more interesting? You started a thread on using a scud (vehicle derived from the Wasserfall) as the first stage of an orbital launchers. It's been done, Vanguard ...
Title: Re: Rocket Engine Q&A
Post by: Danderman on 03/17/2013 05:00 am
You know what is more interesting? You started a thread on using a scud (vehicle derived from the Wasserfall) as the first stage of an orbital launchers. It's been done, Vanguard ...

As I am finding out.

But, Vanguard was a derivative of a derivative of a Scud ancestor (Wasserfall) and my issue in the other thread is whether an actual Scud could be used as a satellite launcher.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 03/22/2013 07:12 am
Terminology question: is the injector of the main combustion chamber considered part of the powerhead of an engine?
Title: Re: Rocket Engine Q&A
Post by: Antares on 04/05/2013 06:55 pm
No.  The powerhead is the turbopump(s) and gas generator or preburner.
Title: Re: Rocket Engine Q&A
Post by: kalif3000 on 06/04/2013 08:50 am
Hello everybody,

I am Faissal and new in the forum. I hope that is sufficient as intro :) .
I find this forum the best place for up to date info on current spaceflight topics, a really great source for news. I think many members here are experts in many aerospace related flields which is why I think I can finally find an answer to a (I think basic) question that is bothering me for a while here.

I am currenty studying the design of rocket engines and after scouting many books and other sources I hit a roadblock.
I noticed that all the equations I have seen so far that assess the performance of rocket engines use chamber pressure as input variable. However, those equations and the theory I have seen so far does not continue to elaborate on how the desired chamber pressure dictates the geometry and design of the rocket combustion chamber (or, the other way around, how to arrive at the chamber pressure from given chamber geometry and gas property)

I have seen some claimes here and there that the chamber pressure is equal to the pump discharge pressure minus the losses in the pipes (unfortunately not mentioned specifically in the books and good sources I have). Is that true?
If yes, then it seems counterintuitive to me that the size of the combustion chamber doesn't matter or am I missing something here?

I have recently whatched a short presentation where Tom Mueller explained how the Merlin engine works. There he explained that the turbopump discharges the propallent at 1400 and 1500 psi while the chamber pressure is 1000 psi. Where is this difference then coming from. It can't be all losses in the piping, or is it?

http://moonandback.com/2013/03/31/the-merlin-engine-presented-by-spacexs-tom-mueller/

Thank you for any advice and I apologize for any errors in the text. I am german, and not even a pro writing texts in my own language.


Faissal
Title: Re: Rocket Engine Q&A
Post by: strangequark on 06/04/2013 10:47 pm
Hello everybody,

I am Faissal and new in the forum. I hope that is sufficient as intro :) .
I find this forum the best place for up to date info on current spaceflight topics, a really great source for news. I think many members here are experts in many aerospace related flields which is why I think I can finally find an answer to a (I think basic) question that is bothering me for a while here.

I am currenty studying the design of rocket engines and after scouting many books and other sources I hit a roadblock.
I noticed that all the equations I have seen so far that assess the performance of rocket engines use chamber pressure as input variable. However, those equations and the theory I have seen so far does not continue to elaborate on how the desired chamber pressure dictates the geometry and design of the rocket combustion chamber (or, the other way around, how to arrive at the chamber pressure from given chamber geometry and gas property)

I have seen some claimes here and there that the chamber pressure is equal to the pump discharge pressure minus the losses in the pipes (unfortunately not mentioned specifically in the books and good sources I have). Is that true?
If yes, then it seems counterintuitive to me that the size of the combustion chamber doesn't matter or am I missing something here?

I have recently whatched a short presentation where Tom Mueller explained how the Merlin engine works. There he explained that the turbopump discharges the propallent at 1400 and 1500 psi while the chamber pressure is 1000 psi. Where is this difference then coming from. It can't be all losses in the piping, or is it?

http://moonandback.com/2013/03/31/the-merlin-engine-presented-by-spacexs-tom-mueller/

Thank you for any advice and I apologize for any errors in the text. I am german, and not even a pro writing texts in my own language.


Faissal

Hi Faissal,

Chamber pressure is equal to pump discharge pressure minus all down stream losses, not just in "pipes" (propellant feed lines). The other big loss is across the injector. Injector pressure drop is usually 15-25%. Most of this pressure gets converted to propellant injection velocity, which helps break up propellant streams into drops and mix them well. It also keeps the chamber from "chugging". Merlin for example might have 5% loss in lines, then 25% through injector. 0.7*1500=1050psi.

"Chugging" happens when the injector pressure drop is too low. A small, random increase in chamber pressure will cause a large decrease in propellant flow. This will cause a large decrease in pressure, which causes a large increase in propellant flow, and so on in a cycle.

Pressure does not depend on chamber size and shape. Throat size is set directly by mass flow requirement. Chamber diameter is set as a multiple of throat diameter, based on experience with a given thrust. The designer must balance having a short convergent nozzle section, with keeping the chamber a reasonable diameter. You want neither a "pancake" or a "stovepipe", but there is some flexibility. Length is set to allow enough time for reaction, while keeping the chamber as light as possible. This is often set by defining a characteristic length (L*), which is Volume/Throat Area, and using a value, based on experience with a given propellant combination. It could also be set from first principles by looking at reaction time compared to residence time.

Chamber pressure is set to a desired value. Throat area, injector area, and pump output are selected for this value. You do not design the throat, injector, and pump and then calculate pressure, which is a misconception I remember having when I was in school.

Hope this helps, and welcome to the forum.
Title: Re: Rocket Engine Q&A
Post by: kalif3000 on 06/05/2013 07:41 am
Hi strangequark,

Yes, that was very helpful. I was planning to follow up on my question with one about the injector but you answered it all. Thanks!  :)
Title: Re: Rocket Engine Q&A
Post by: Oli on 07/27/2013 09:43 am

I have a question. Scaling up rocket engines seeems to be difficult/costly. I'm not talking about changing the cycle or increasing chamber pressure, only about increasing thrust. I wonder why this is the case and to what extent higher thrust engines are more expensive in development as well as in production.
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/27/2013 09:46 am

I have a question. Scaling up rocket engines seeems to be difficult/costly. I'm not talking about changing the cycle or increasing chamber pressure, only about increasing thrust. I wonder why this is the case and to what extent higher thrust engines are more expensive in development as well as in production.

production and testing facilities have to be larger, hence most cost
Title: Re: Rocket Engine Q&A
Post by: Oli on 07/27/2013 10:05 am
^

Ok, so if you had to give development and production cost (seperate) as a function of thrust, how would it look like? Log, fractional, linear, quadratic, exponential?

Edit: Well I guess its not possible to describe it with a simple function.

Anyway, I've got another question  :)

Edit2: Is there a measure for engine efficiency independent of nozzle expansion ratio? Usually vacuum ISP for a given expansion ratio is given for an engine.

Title: Re: Rocket Engine Q&A
Post by: Proponent on 07/29/2013 08:13 am
Well, there's c*, which is defined as chamber pressure times throat area divided by mass flow rate and is measured in velocity units.  When multiplied by the thrust coefficient, CF, a dimensionless factor defined as thrust divided by the product of chamber pressure and throat area, it gives the exhaust velocity: c = CFc*.

You can think of c* as a a measure of what happens in the combustion chamber and nozzle inlet and of CF as a measure of the efficiency of the expanding part of the nozzle.
Title: Re: Rocket Engine Q&A
Post by: R7 on 07/29/2013 10:53 am
Edit2: Is there a measure for engine efficiency independent of nozzle expansion ratio? Usually vacuum ISP for a given expansion ratio is given for an engine.

IIRC you asked for theoretical efficiency with infinite expansion too before edited the question? Equation for that is

vexhaust = sqrt( 2*k*R*T/(k - 1) )

where k is ratio of specific heats, R specific gas constant (universal gas constant divided by average molecular weight of exhaust) and T combustion chamber temperature. Divide the result by g to get theoretical Isp in seconds. Note how result is independent of initial combustion chamber pressure and finite despite area ratio grown to infinite.

Title: Re: Rocket Engine Q&A
Post by: Oli on 08/02/2013 12:58 pm
Quote from: R7
Note how result is independent of initial combustion chamber pressure and finite despite area ratio grown to infinite.

Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero? I.e. the bigger P_e in the term below, the more P_c matters.

1 - (P_e / P_c)^((k-1) / k)

So basically at sea level you want engines with high chamber pressure because P_e is high.

Consequently in vacuum higher chamber pressure makes only sense if it reduces nozzle size/mass (or is q somehow related to P_c?).

(variables from http://www.braeunig.us/space/propuls.htm)
Title: Re: Rocket Engine Q&A
Post by: R7 on 08/02/2013 03:32 pm
Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero? I.e. the bigger P_e in the term below, the more P_c matters.

1 - (P_e / P_c)^((k-1) / k)

So basically at sea level you want engines with high chamber pressure because P_e is high.

Yes, yes and yes. You want the P_e/P_c ratio to be as low as possible, zero P_e drives it to zero as long as P_c is non-zero to not to throw division by zero error.

In real world you have lower limit for P_e, even in vacuum, because temperatures get so low that your non-ideal exhaust gases stop behaving ideally and eventually condense.

Quote
Consequently in vacuum higher chamber pressure makes only sense if it reduces nozzle size/mass (or is q somehow related to P_c?).

(variables from http://www.braeunig.us/space/propuls.htm)

Things get quite unpractical if you drop P_c so much that required upper stage vacuum nozzle diameter exceeds that of LV's (like Sea Dragon concept).
mass flux q is directly related to P_c due to the good old PV = nRT. See equation 1.26 on your link (great page btw!). Equation solved for q(dot m) is;

(http://s21.postimg.org/l044vxokn/mdot.jpg)

(Sutton notation, t = throat, 1 = combustion chamber)
Title: Re: Rocket Engine Q&A
Post by: Oli on 08/02/2013 03:56 pm
Quote from: R7
mass flux q is directly related to P_c due to the good old PV = nRT. See equation 1.26 on your link (great page btw!). Equation solved for q(dot m) is;

So if you want to increase q, and thus thrust, you must increase A_t (and hence the size of the combustion chamber as well as the nozzle, I guess) and/or P_c.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 08/02/2013 04:09 pm
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output. Basically, allow for expanding the thrust envelope of the expander cycle. I'm wondering if it would give better performance in the limited case. The tanking would be really bad, but the pressurization system would be "free". And if you direct the TP out gas to the nozzle the molecular weight should give some extra isp, right?
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 08/02/2013 11:23 pm
Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero?...
Yes...
Besides the impracticality of the bell becoming ridiculously big (which you mentioned), you also have the limit of condensing (or freezing) the exhaust as the temperature drops which limits performance.
Title: Re: Rocket Engine Q&A
Post by: R7 on 08/03/2013 09:26 am
So if you want to increase q, and thus thrust, you must increase A_t (and hence the size of the combustion chamber as well as the nozzle, I guess) and/or P_c.

Correct. Note that Isp is very non-linear function of area ratio so if engine diameter is restricted a quick'n'dirty solution is to just increase A_t (and CC) for larger thrust increase but minor Isp decrease. Attached example diagram for NK-33 (actual area ratio is 28).  The k value also affects thrust so propellant choice matters.

Besides the impracticality of the bell becoming ridiculously big (which you mentioned), you also have the limit of condensing (or freezing) the exhaust as the temperature drops which limits performance.

You appear to have skipped the next paragraph I wrote but no worries, the matter is important and deserves to be mentioned twice  :D
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 08/03/2013 04:24 pm
Correct you are!  :)
Title: Re: Rocket Engine Q&A
Post by: Remes on 08/04/2013 03:39 pm
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output. Basically, allow for expanding the thrust envelope of the expander cycle. I'm wondering if it would give better performance in the limited case. The tanking would be really bad, but the pressurization system would be "free". And if you direct the TP out gas to the nozzle the molecular weight should give some extra isp, right?

I think this doesn't make sense. You have to vaporize the fuel and oxydizer anyway. Be it in the thrust chamber or, with a expander cycle, in the nozzle channels. But the energy must be invested anyway.

If you take this energy now to heat up helium, you still have to vaporize Fuel/Oxidizer, so you didn't save anything. Furthermore you have a lot of hot helium which can't be used for anything (Edit: after it served its purpose to spin the TP), which doesn't provide thrust (as it can't be convinced to do any useful chemical reaction) and which was nevertheless heavy.
Title: Re: Rocket Engine Q&A
Post by: Remes on 08/04/2013 03:51 pm
Scaling up rocket engines seeems to be difficult/costly. I'm not talking about changing the cycle or increasing chamber pressure, only about increasing thrust. I wonder why this is the case and to what extent higher thrust engines are more expensive in development as well as in production.

More thrust@same pressure->more fuel/oxidizer:
- bigger thrust chamber, bigger injector, completely new geometry
- TP/GG redesign
- bigger fuel/oxidizer lines (perhaps throughout the whole stage)
- valves needs to be bigger, have higher flow forces, needs stronger actuation, different actution->maybe different power stages
- different volumes/flows/geometry->different dynamics, perhaps different oscillation behaviour
- Cooling needs to be redesigned (different surface/volume ratio)
- All the control loops need to be investigated tuned (if it is not open loop control, in that case all the orifices&Co needs to be tuned)
- Some adaptation in software
- ...

Actually you are building a new engine and you have to recertify it. Some basic things like materials used, supplier, etc. remain the same, but actually most of the machining/tooling/parts/... changes.

Edit:
aaaaand:
- higher thrust-> do you have to change the thrust transfer structure of the LV?
- Do the tanks have to modified?
- do you need new/stronger TVC-actuators?
- what about the engine gimbal?
- Does the ascent profile have some undesired side effects? (I was thinking of the first DIV and the cavitation problem. This was not really a engine redesign and fits only partially to this topic. But it's a good example for sideffects).

Some of these points are not that critical if you improve the thrust by a few percent. At higher improvement rates the points get more and more critical.
Title: Re: Rocket Engine Q&A
Post by: R7 on 08/05/2013 07:43 am
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output.

Strangequark has some nice relevant posts on expander/cooling issues

http://forum.nasaspaceflight.com/index.php?topic=16279.msg986192#msg986192

http://forum.nasaspaceflight.com/index.php?topic=30910.msg1006206#msg1006206

http://forum.nasaspaceflight.com/index.php?topic=31040.msg1009889#msg1009889

In short, specific heat capacity (Cp) matters more and hydrogen is in it's own league. Also remember the He is getting more and more expensive, also difficult to store efficiently (AIUI rockets don't store it liquefied but compressed and chilled in supercritical form).
Title: Re: Rocket Engine Q&A
Post by: baldusi on 08/05/2013 12:46 pm
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output.

Strangequark has some nice relevant posts on expander/cooling issues

http://forum.nasaspaceflight.com/index.php?topic=16279.msg986192#msg986192

http://forum.nasaspaceflight.com/index.php?topic=30910.msg1006206#msg1006206

http://forum.nasaspaceflight.com/index.php?topic=31040.msg1009889#msg1009889

In short, specific heat capacity (Cp) matters more and hydrogen is in it's own league. Also remember the He is getting more and more expensive, also difficult to store efficiently (AIUI rockets don't store it liquefied but compressed and chilled in supercritical form).
I had confused heat of vaporization with heat capacity. Sorry.
Title: Re: Rocket Engine Q&A
Post by: fatjohn1408 on 08/05/2013 01:39 pm
Hello I am wondering the following:

Does the difference in Isp for an engine under vacuum conditions and under sea level conditions solely arise from the pressure related term (Pe-Pa)*Ae?

If so, does that mean that the mass flow of an engine is constant throughout flight? Just to refresh my memory.

Regards
Title: Re: Rocket Engine Q&A
Post by: MP99 on 08/05/2013 06:08 pm
Hello I am wondering the following:

Does the difference in Isp for an engine under vacuum conditions and under sea level conditions solely arise from the pressure related term (Pe-Pa)*Ae?

If so, does that mean that the mass flow of an engine is constant throughout flight? Just to refresh my memory.

Regards

I assume the generator on a GG engine would also be affected same way, IE pumping efficiency is reduced. I guess that would also reduce efficiency (assuming the GG compensates to keep Pc the same).

Cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Galactic Penguin SST on 08/15/2013 06:44 pm
What rocket engines using hydrogen peroxide/kerosene are there in current service? How good would they perform as a restartable upper stage engine?

(I'm asking because China is apparently making a new multi-restartable upper stage based on a 35 kN thrust H2O2/RP-1 closed cycle engine and plans to put it into service next year)
Title: Re: Rocket Engine Q&A
Post by: baldusi on 08/15/2013 07:08 pm
What rocket engines using hydrogen peroxide/kerosene are there in current service? How good would they perform as a restartable upper stage engine?

(I'm asking because China is apparently making a new multi-restartable upper stage based on a 35 kN thrust H2O2/RP-1 closed cycle engine and plans to put it into service next year)
The Soyuz's first stage use the combination to run the turbopump.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 09/05/2013 01:30 pm
Am I right in my understanding that in the expander cycle rockets is purely liquid? In other words, the expander part works by upping the pressure of the liquid but avoiding the phase change? Thus, the injector is liquid-liquid (as the Russians say)?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 09/05/2013 07:12 pm
Am I right in my understanding that in the expander cycle rockets is purely liquid? In other words, the expander part works by upping the pressure of the liquid but avoiding the phase change? Thus, the injector is liquid-liquid (as the Russians say)?

I'm stretching a bit out of my expertise, but I believe expanders typically run supercritical (http://en.wikipedia.org/wiki/Supercritical_fluid). In short, once it gets to the injector, the fuel will be a "gas", but the process to get there does not involve a sharp phase change (like boiling). As for the pressure, it will actually drop through the cooling jacket. The energy increase is expressed in the temperature of the fluid, not the pressure.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 09/05/2013 08:22 pm
Am I right in my understanding that in the expander cycle rockets is purely liquid? In other words, the expander part works by upping the pressure of the liquid but avoiding the phase change? Thus, the injector is liquid-liquid (as the Russians say)?

I'm stretching a bit out of my expertise, but I believe expanders typically run supercritical (http://en.wikipedia.org/wiki/Supercritical_fluid). In short, once it gets to the injector, the fuel will be a "gas", but the process to get there does not involve a sharp phase change (like boiling). As for the pressure, it will actually drop through the cooling jacket. The energy increase is expressed in the temperature of the fluid, not the pressure.
I forgot, you have pressure drop and temperature goes up. Silly me. So, if critical point of H2 is (33K,1.3MPa), as long as you keep it above 1.3MPa, you don't have nasty things like bubbles/cavitation, right? My question is, then, by the time it arrives at the turbine, how's energy transferred? I can understand that a gas expands and since the other side is incompressible you trade movement (which in a gas is temperature or pressure?) for potential energy (higher pressure) on the pump side. In the case of super critical H2, does it behave like a gas by the time it reaches the turbine?
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 09/05/2013 08:25 pm
Supercritical works like a gas. It is compressible.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 09/05/2013 11:45 pm
I forgot, you have pressure drop and temperature goes up. Silly me. So, if critical point of H2 is (33K,1.3MPa), as long as you keep it above 1.3MPa, you don't have nasty things like bubbles/cavitation, right? My question is, then, by the time it arrives at the turbine, how's energy transferred? I can understand that a gas expands and since the other side is incompressible you trade movement (which in a gas is temperature or pressure?) for potential energy (higher pressure) on the pump side. In the case of super critical H2, does it behave like a gas by the time it reaches the turbine?

Right, no bubbles, boiling or any of that mess. So, turbine power is mass flow*specific heat*T_inlet*(1-[pin/pout]^((1-k)/k).

So, what's that mean? If the pressure drop across the turbine is held constant, then for a given amount of gas, the hotter it is, the more power you get out of it. This is how the expander cycle works. The energy is in the temperature, rather than the pressure, so to speak.

This is true for all cycles, actually. Pressure will never be higher than it is at pump outlet, but temperature will increase substantially.
Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 09/06/2013 10:52 pm
Does anybody know how much the ISP of a rocket engine changes throughout its atmospheric flight? Is there any quick way to calculate this? For example, if a rocket engine is sea-level optimized, what would how much would its ISP change at 10 km, 25 km, etc.? Is there a table for this?
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 09/06/2013 11:19 pm
Does anybody know how much the ISP of a rocket engine changes throughout its atmospheric flight? Is there any quick way to calculate this? For example, if a rocket engine is sea-level optimized, what would how much would its ISP change at 10 km, 25 km, etc.? Is there a table for this?
Rocket Propulsion Elements, Sutton.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/10/2014 05:19 pm
Below I list data of some 1st stage liquid hydrogen engine:

                                           combustion pressure                                area ratio

vulcain 1                               100bar                                                        45

vulcain 2                                117bar                                                       58

le-7                                         127bar                                                       52

le-7a                                      120bar                                                       52

rs-68                                        97bar                                                       21.5

Here comes my question, how comes that similar engines have so different area ratios, I mean,is rs-68 properly designed?

Thanks in advance.

BTW,  why is there not a engine section for such a specialized forum?



Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 10/10/2014 05:42 pm
Remember the RS-68 was designed for cost. Hence the use of an ablative nozzle instead of a regeneratively cooled nozzle. I wonder if it had something to do with manufacturing costs.

If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here.

Could it be the weight of a larger nozzle would have negated the ISP gain?
Title: Re: Rocket Engine Q&A
Post by: sdsds on 10/10/2014 05:50 pm
Just a guess? Thrust. The low area ratio implies a big throat. Have you compared the propellant mass flow of the engines?
Title: Re: Rocket Engine Q&A
Post by: R7 on 10/10/2014 05:54 pm
If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here.

Look beyond the engines to the LVs which use them. Ariane and H-II have solid boosters which do most of the lifting at low altitudes. Thus you don't have to worry about core engine sea level Isp so much. Situation is different with Delta IV, good sea level performance is required from RS-68.

sea level Isps:

Vulcain 1: 326s
Vulcain 2: 318s
LE-7: 349s
LE-7A: 338s
RS-68: 365s
Title: Re: Rocket Engine Q&A
Post by: baldusi on 10/11/2014 12:08 am
Vulcain 1/2 and LE-7/A are, basically, sustainer engines. When you use a booster augmented h2/lox core (like Ariane 5, H-II or STS did), you want to optimize for vacuum performance. When you need an engine to propel the whole stack alone (like RS-68 might do) you want to focus more on sea level performance.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 10/11/2014 12:26 am
Vulcain 1/2 and LE-7/A are, basically, sustainer engines. When you use a booster augmented h2/lox core (like Ariane 5, H-II or STS did), you want to optimize for vacuum performance. When you need an engine to propel the whole stack alone (like RS-68 might do) you want to focus more on sea level performance.

Yeah, it's all about the operating envelope. According to the Ariane V user's manual (http://www.arianespace.com/launch-services-ariane5/Ariane5_users_manual_Issue5_July2011.pdf), the core engine cutoff is at an altitude of about 250km, and burns for over 500 seconds for a typical GTO trajectory. It needs to do well in vacuum. The Delta IV (http://www.ulalaunch.com/uploads/docs/Launch_Vehicles/Delta_IV_Users_Guide_June_2013.pdf) on the other hand has MECO occur at 121km, after 245 seconds, and a third of that time is below 10km altitude. The time averaged altitude is much lower for the RS-68.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 01:36 am
If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here.

Look beyond the engines to the LVs which use them. Ariane and H-II have solid boosters which do most of the lifting at low altitudes. Thus you don't have to worry about core engine sea level Isp so much. Situation is different with Delta IV, good sea level performance is required from RS-68.

sea level Isps:

Vulcain 1: 326s
Vulcain 2: 318s
LE-7: 349s
LE-7A: 338s
RS-68: 365s

So your opinoin is that the other 4 engines are optimized for vacuum performance?

It makes sense.

Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 02:27 am
Just a guess? Thrust. The low area ratio implies a big throat. Have you compared the propellant mass flow of the engines?


For a given propellant combination, the theoretical area ratio is generally related to combustion pressure.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 02:48 am
Remember the RS-68 was designed for cost. Hence the use of an ablative nozzle instead of a regeneratively cooled nozzle. I wonder if it had something to do with manufacturing costs.

If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here.

Could it be the weight of a larger nozzle would have negated the ISP gain?

Weight or t/w has no influence on thrust, inproper area ratio does.

Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 09:12 am
As we all know, most liquid hydrogen rocket engines have extraordinary Isp compared to other chemical rocket engines. I think the reason is generally the low average molecular mass of the resultants of combustion. The resulant water vapor has small molecular mass compared to CO2 or CO, also fuel rich combustion allows plenty of  hydrogen molecules to be left over, which results in  hot mixture of low average molecular mass, usually 12-13, compared to over 18 for cases of hydrocabon fuels.

Now the rocket engine function

c=ve+(Pe-Pa)Ae/q

The low average molecular mass of the resultant mixture contributes to both "ve" and "(Pe-Pa)/q" as follows:

For the "ve" part,

(http://images-mediawiki-sites.thefullwiki.org/06/1/5/1/35312061362500042.png)

Low average molecular mass means high ve, which is the case for flights in designed ambient pressure.
But even for a vacuum optimized liquid hydrogen engine burning in the vacuum, this part of Isp can hardly make it over 420s, which you can deduce yourself.

For the "(Pe-Pa)Ae/q" part, low average molecular mass means larger exit area even for the  same thrust and optimzed for the same ambient pressure, which can be seen as in the case of rd-0120 and rd-170.


(http://www.buran.fr/energia/img/vulkan.jpg)

Larger exit area means larger value for the  "(Pe-Pa)Ae/q" part of Isp.
I also conclude that Isp part over 400s are mainly the result of larger exit area induced by the low average molecular mass of resultant of combustion.

What do you think of my understanding? Welcome to reply.


 
Title: Re: Rocket Engine Q&A
Post by: smoliarm on 10/11/2014 10:14 am
>>mocular mass
??
It should molecular mass
Or, better yet - molar mass.

The answer to your question is in chemical thermodynamics. Which is bad, because in general opinion this is very complex matter. But the good news is that here you need only the simplest part of all the thermodynamics - its First Law and the term Enthalpy.

The answer - in a short -
LH2 has very high combustion Enthalpy (which is here energy per mole) AND the lowest possible molar mass.
This gives to the LH2-LOX pair the highest specific combustion energy -- per gram of propellant - among all the things which can burn (in the Universe:)
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 10:33 am
>>mocular mass
??
It should molecular mass
Or, better yet - molar mass.

The answer to your question is in chemical thermodynamics. Which is bad, because in general opinion this is very complex matter. But the good news is that here you need only the simplest part of all the thermodynamics - its First Law and the term Enthalpy.

The answer - in a short -
LH2 has very high combustion Enthalpy (which is here energy per mole) AND the lowest possible molar mass.
This gives to the LH2-LOX pair the highest specific combustion energy -- per gram of propellant - among all the things which can burn (in the Universe:)

I Did mean molecular. Thanks for bringing it out.

As for the thermodynamics, one doesn't have to understand thermodynamics thoroughly to design rocket engines, what he need to do is learn to read relative charts.

Title: Re: Rocket Engine Q&A
Post by: R7 on 10/11/2014 01:49 pm
Larger exit area means larger value for the  "(Pe-Pa)Ae/q" part of Isp.

Too simplified statement. For instance engine optimized for vacuum should have as low Pe as possible. Play with engine simulation tools, the greater the area ratio and the lower the Pe the closer optimum expansion and vacuum Isps are, meaning pressure thrust component shrinks.

Optimum would be not to have the pressure thrust component at all but in practice you always have some. At sea level it is usually negative, meaning optimum expansion doesn't happen at sea level but at higher altitude (~4-5km for many kerolox, even higher for hydrolox core engines you listed (and SSME)).
Title: Re: Rocket Engine Q&A
Post by: Jim on 10/11/2014 01:57 pm

As for the thermodynamics, one doesn't have to understand thermodynamics thoroughly to design rocket engines, what he need to do is learn to read relative charts.


Quite wrong.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 02:11 pm
Larger exit area means larger value for the  "(Pe-Pa)Ae/q" part of Isp.

Too simplified statement. For instance engine optimized for vacuum should have as low Pe as possible. Play with engine simulation tools, the greater the area ratio and the lower the Pe the closer optimum expansion and vacuum Isps are, meaning pressure thrust component shrinks.

Optimum would be not to have the pressure thrust component at all but in practice you always have some. At sea level it is usually negative, meaning optimum expansion doesn't happen at sea level but at higher altitude (~4-5km for many kerolox, even higher for hydrolox core engines you listed (and SSME)).

I can't say you are wrong, you just missed the point.

I was not talking about optimizing a liquid hydrogen engine, but comparing liquid hydrogen engines with other chemical rocket engines.
Title: Re: Rocket Engine Q&A
Post by: R7 on 10/11/2014 02:45 pm
I can't say you are wrong, you just missed the point.

I was not talking about optimizing a liquid hydrogen engine, but comparing liquid hydrogen engines with other chemical rocket engines.

The point was exit areas / pressures as such aren't reason why hydrolox excels. Smoliarm already gave you the short answer.


Furthermore statements like this...

For the "(Pe-Pa)Ae/q" part, low average molecular mass means larger exit area even for the  same thrust and optimzed for the same ambient pressure, which can be seen as in the case of rd-0120 and rd-170.
 

... are baffling at best because RD-0120 and RD-170 are neither same thrust nor optimized for same ambient pressure.
Title: Re: Rocket Engine Q&A
Post by: R7 on 10/11/2014 03:01 pm
At lease one thing is for sure, the optimized ambient pressure for rd-0120 and rd170 is approximately sea level pressure.

Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 03:20 pm
At lease one thing is for sure, the optimized ambient pressure for rd-0120 and rd170 is approximately sea level pressure.

Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.

Can't believe that.
Perhaps you are right.
And as for the thrust, I meant thrust of a chamber of RD-170 and that of rd-0120.


Title: Re: Rocket Engine Q&A
Post by: R7 on 10/11/2014 03:49 pm
Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.

Can't believe that.

It's not a matter of belief, that's how it is.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/11/2014 04:24 pm
Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.

Can't believe that.

It's not a matter of belief, that's how it is.

OK,and thanks for sharing these sceenshots.

It seems that you have very good simulation software.

Can you also provide the exit area of a chamber of rd-170 and that of rd-0120 optimized for a common  ambient pressure, say 0.5atm?


Title: Re: Rocket Engine Q&A
Post by: Jim on 10/11/2014 05:44 pm

Can't believe that.


An understanding of thermodynamics would open your eyes.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/12/2014 01:47 am

Can't believe that.


An understanding of thermodynamics would open your eyes.

Of course.
And I do know a little.
Title: Re: Rocket Engine Q&A
Post by: R7 on 10/12/2014 10:14 am
OK,and thanks for sharing these sceenshots.

It seems that you have very good simulation software.

Can you also provide the exit area of a chamber of rd-170 and that of rd-0120 optimized for a common  ambient pressure, say 0.5atm?

You are welcome. It is a good software indeed, developed by a real Russian rocket scientist Alexander Ponomarenko. Go ahead and try it yourself (http://www.propulsion-analysis.com/index.htm). RD-170 configuration file is included in the software examples, RD-0120 you can quickly build from public information but I'll attach configuration file for that to speed things up. Remove the ".txt" from the filename, forum software requires it.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/12/2014 12:39 pm
OK,and thanks for sharing these sceenshots.

It seems that you have very good simulation software.

Can you also provide the exit area of a chamber of rd-170 and that of rd-0120 optimized for a common  ambient pressure, say 0.5atm?

You are welcome. It is a good software indeed, developed by a real Russian rocket scientist Alexander Ponomarenko. Go ahead and try it yourself (http://www.propulsion-analysis.com/index.htm). RD-170 configuration file is included in the software examples, RD-0120 you can quickly build from public information but I'll attach configuration file for that to speed things up. Remove the ".txt" from the filename, forum software requires it.

It is a file of a not usual format, it will take some time to read it.

And I have a new question, is a combustion chamber temperature of over 3800K normal?I haven't seen so high a value before.
Title: Re: Rocket Engine Q&A
Post by: R7 on 10/12/2014 04:44 pm
@nimbostratus

You are not supposed to read the attached file but load it into the software I provided link for :)

Chamber temperature is connected to chamber pressure. RD-170 Pc is very high, therefore the Tc is high too. IIRC only RD-180 has higher Pc among kerolox engines, the RPA tool calculates Tc nearly 3900K for it.
Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/13/2014 01:50 am
@nimbostratus

You are not supposed to read the attached file but load it into the software I provided link for :)

Chamber temperature is connected to chamber pressure. RD-170 Pc is very high, therefore the Tc is high too. IIRC only RD-180 has higher Pc among kerolox engines, the RPA tool calculates Tc nearly 3900K for it.

Oh, I didn't realize that.

I visited that site several times, but never found that software.

As for chamber temperature, as far as I know, the chamber termparature can hardly exceed 3700K. I guess the software has problem with the temperature part.

Title: Re: Rocket Engine Q&A
Post by: nimbostratus on 10/13/2014 03:56 am
@nimbostratus

You are not supposed to read the attached file but load it into the software I provided link for :)

Chamber temperature is connected to chamber pressure. RD-170 Pc is very high, therefore the Tc is high too. IIRC only RD-180 has higher Pc among kerolox engines, the RPA tool calculates Tc nearly 3900K for it.

I run several cases on this software, it is really good for selecting the proper area ratio.

After runnig the software,I found that flow seperation stops 1st stage Hydrolox engines from adopting too big an area ratio, so both rs-25 and rd-0120 adopt the highest area ratio possible in which case flow seperation doesn't occur.

And thank you for sharing the software
Title: Re: Rocket Engine Q&A
Post by: MarsMethanogen on 10/30/2014 12:47 pm
While is may have already been addressed somewhere in this thread, I was hoping that someone could educate me on how hypergols are used in the Antares rocket engine of launch vehicle.  My question originates from the recent launch failure and the post-event warning for civilians not to enter the area given the remnanats of hypergols, which I presume was something like a hydrazine derivative and/or N2O4.  I am aware that the first stage propellants are RP-1 and LOX, and I am aware that the second stage uses solid propellants.  So then I thought that maybe the first stage turbopumps are powered by these hypergols.  But then I found out that the first stage engines are staged combustion engines.  So hypergols wouldn't be involved here, as the gas driving the turbines are generated by using the kerolox propellants in the pre-burner.  So where are these hypergols used in this launch vehicle?  Is there an APU, conceivably used to provide hydraulic power to what I preseme might by TVC?  Please educate me here.  Thank you.
Title: Re: Rocket Engine Q&A
Post by: kch on 10/30/2014 12:53 pm
While is may have already been addressed somewhere in this thread, I was hoping that someone could educate me on how hypergols are used in the Antares rocket engine of launch vehicle.  My question originates from the recent launch failure and the post-event warning for civilians not to enter the area given the remnanats of hypergols, which I presume was something like a hydrazine derivative and/or N2O4.  I am aware that the first stage propellants are RP-1 and LOX, and I am aware that the second stage uses solid propellants.  So then I thought that maybe the first stage turbopumps are powered by these hypergols.  But then I found out that the first stage engines are staged combustion engines.  So hypergols wouldn't be involved here, as the gas driving the turbines are generated by using the kerolox propellants in the pre-burner.  So where are these hypergols used in this launch vehicle?  Is there an APU, conceivably used to provide hydraulic power to what I preseme might by TVC?  Please educate me here.  Thank you.

As far as I know, they're used in the Cygnus cargo ship.
Title: Re: Rocket Engine Q&A
Post by: MarsMethanogen on 10/30/2014 01:00 pm
Of course!  You're right!  I went out and read about the cargo craft and it uses hypergols in the "propellant system".  I feel so silly for limiting my thinking to the launch vehicle and not the cargo vehicle.  The use of hypergols as the propellant for the engines and the RCS is an industry standard that goes way back. 
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 10/30/2014 02:38 pm
While I don't know the answer for Antares.

I do know that other launch vehicles have in the past used Hypergols in the first stage. The typical use would be roll control. Atlas IIAS, and many solids have used them for that reason.

It is also not uncommon to use them on the second stage for not only roll control, but also keeping the second stage at the correct attitude. The Antares second stage has a coast period before the solid fires. It is possible they might be used on the second stage to keep it properly oriented during the coast period.

So while Cygnus is a no brainier, they may be used in other parts on the Antares.

Also, nasty chemicals like TEA/TEB are often used to "light" the engine. While it is usually provided by the ground  support equipment, there may be some sort of tank and contaminated plumbing in what remains of the engines.

Does anyone know how toxic the battery chemistry  in launch vehicles is? The remains could easily be another source of nasty chemicals that could harm the careless.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/02/2014 03:00 pm
Would a compressed air / helium powered turbopump be a reasonable alternative to a Flometrics-style pistonless pump?
Title: Re: Rocket Engine Q&A
Post by: Damon Hill on 11/05/2014 01:22 am
Silver-zinc batteries used to be popular for rocket systems; the technology may be leaning towards lithium-ion.  Eagle-Pischer has been a provider for decades.  As a rule, I don't think batteries are too much of a concern, compared to hydrazines and nitrogen tetroxide, which are the most commonly used hypergols.  There are other possible chemistries, but they are far less common.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/06/2014 02:07 pm
The other day I found a partial answer to my question about compressed air turbopumps. It turns out that as early as 1941 the Russians were experimenting with compressed air turbopumps. It was only a temporary solution, and heavy, but it did work.

Rocketing Into the Future: The History and Technology of Rocket Planes (http://books.google.nl/books?id=4M9i-FXVKckC&pg=PA127&lpg=PA127#v=onepage&q&f=false)

In that case, I wonder why you would ever prefer a valve-based Flometrics style pistonless pump to a compressed air turbopump, which doesn't need very reliable valves that are cycled many times. The energy density of compressed gas is similar to that of lead acid batteries, which suggests that lithium-ion or silver-zinc powered electric pumps might be even better. For high gravimetric density you could use dielectric immersion cooling, which is a mature technology and might already be in use in the automotive sector.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 11/06/2014 03:45 pm
I fail to understand the conceptual difference between piston-less compressed air valve and normal pressure fed rockets.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/06/2014 04:28 pm
Only a small pressure vessel needs to be at high pressure, the propellant tanks themselves can be kept at low pressure. A small tank is filled with propellant from the main propellant tanks at low pressure, then the feed valve is closed, the pressurisation valve is opened, once the pressure has equalised it is closed again and then the discharge valve is opened and just before it's empty, it is closed again, and then the cycle repeats. It's a bit more complicated than that, but that's the general idea.

How the pistonless pump operates (http://www.rocketfuelpump.com/technology/operation/)
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 11/06/2014 05:39 pm
One thing you could do to keep the weight of a pistonless system down is to generate the gas from an LN or LOX or LHe by running it through the cooling jacket of the engine. Storing it as a liquid and converting it into a gas would save a fair bit of weight on the container... Being warmed, it will also save on the amount of gas that needs to be dragged up hill.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 11/07/2014 07:38 am
Only a small pressure vessel needs to be at high pressure, the propellant tanks themselves can be kept at low pressure. A small tank is filled with propellant from the main propellant tanks at low pressure, then the feed valve is closed, the pressurisation valve is opened, once the pressure has equalised it is closed again and then the discharge valve is opened and just before it's empty, it is closed again, and then the cycle repeats. It's a bit more complicated than that, but that's the general idea.

How the pistonless pump operates (http://www.rocketfuelpump.com/technology/operation/)
Why is the pump gas vented overboard instead of being used to pressurise the main tank?

High pressure gas should be able to be vented into the main tank until it reaches a safe max pressure, then the rest vented overboard.

Cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/07/2014 09:12 am
The diagram shows the two pumping cylinders being the same size, but I've seen Flowmetrics diagrams in which one cylinder is about a quarter of the size of the other.  I see only disadvantages of differently-sized cylinders:

1. The need to manufacture more types of components, and
2. With the two cylinders having different cycle durations, the system generates a wider spectrum of mechanical forcings and has a higher probability of suffering a problematic resonance.

What advantage(s) of differently-sized cylinders am I missing?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/07/2014 09:42 am
I believe the main advantage is that the system will be lighter and more compact as a result.
Title: Re: Rocket Engine Q&A
Post by: R7 on 11/07/2014 10:26 am
I believe the main advantage is that the system will be lighter and more compact as a result.

Yes.

http://www.rocketfuelpump.com/wp-content/uploads/2013/12/RocketPumpJPC2003.pdf

Quote
Instead of two similar pump chambers, it uses
one main chamber which supplies fuel for most
of the time and an auxiliary chamber which
supplies fuel for the rest of the time. The main
chamber is placed inside the tank, and it is filled
through a number of check valves so that it can
be filled quickly, thereby reducing the size of the
auxiliary chamber, which is typically one fourth
the size of the main chamber. The optimized
design offers a substantial weight savings over
the basic design, in that it uses one primary
pumping chamber and one auxiliary chamber
instead of two pumping chambers.

Wondering if the volume asymmetry also helps with avoiding bad resonances because the valving sort of "gallops" instead of steady cycling.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/07/2014 02:02 pm
Here's a video of the pistonless pump in action. I don't know what to call the white things going up and down inside a cylinder, but apparently it's not 'piston'. :-)

https://www.youtube.com/watch?v=z0Zk2ANHlLE

And another one:

https://www.youtube.com/watch?v=7Fa7SCB2Jd4

I love the steampunk sound of this thing!
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 11/07/2014 03:51 pm
The "disk" might be to prevent sloshing, ingestion of the gas, and loss of working fluid when the cylinder is vented. It may also not have a very tight fit and just float on the working fluid.

I wonder if the Flometrics definition of "piston" is it requires a rod to push it.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/07/2014 04:07 pm
I believe the main advantage is that the system will be lighter and more compact as a result.

Yes.

http://www.rocketfuelpump.com/wp-content/uploads/2013/12/RocketPumpJPC2003.pdf

Quote
Instead of two similar pump chambers, it uses
one main chamber which supplies fuel for most
of the time and an auxiliary chamber which
supplies fuel for the rest of the time. The main
chamber is placed inside the tank, and it is filled
through a number of check valves so that it can
be filled quickly, thereby reducing the size of the
auxiliary chamber, which is typically one fourth
the size of the main chamber. The optimized
design offers a substantial weight savings over
the basic design, in that it uses one primary
pumping chamber and one auxiliary chamber
instead of two pumping chambers.

Wondering if the volume asymmetry also helps with avoiding bad resonances because the valving sort of "gallops" instead of steady cycling.

Why not have two equally-sized cylinders, each of which is filled by multiple check valves?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/07/2014 04:22 pm
I get the impression that the advantages over a pressure-fed system are much smaller if you don't use a gas generator / heat exchanger to supply the pressurant gas.

If I understand Wikipedia (https://en.wikipedia.org/wiki/Pressure_vessel) correctly the mass fraction of a pressure vessel does not improve with pressure, so I'm guessing the pressurant tank would be about as heavy as a high-pressure propellant tank, just smaller. Using a pistonless pump might reduce the mass of the propellant tank from 1M to 0.1M, but the pressurant tank would still be at 1M. So tankage in total would go from 2M to 1.1M, which is still significant of course, but only by a factor of about 2, rather than a factor of 10.

You can probably use composites for the pressurant tank however, even if the propellant isn't compatible with composites, so that might give additional efficiency.
Title: Re: Rocket Engine Q&A
Post by: Specifically-Impulsive on 11/08/2014 10:59 pm
Would a compressed air / helium powered turbopump be a reasonable alternative to a Flometrics-style pistonless pump?

One of the upgrades being studied for the STS was to replace the hydrazine APUs in the SRBs with a high pressure helium system that blew down through a turbine.  Don't know how far it got but it was being given serious consideration at one time (prior to STS-107).
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/09/2014 09:49 am
I get the impression that the advantages over a pressure-fed system are much smaller if you don't use a gas generator / heat exchanger to supply the pressurant gas.

If I understand Wikipedia (https://en.wikipedia.org/wiki/Pressure_vessel) correctly the mass fraction of a pressure vessel does not improve with pressure, so I'm guessing the pressurant tank would be about as heavy as a high-pressure propellant tank, just smaller. Using a pistonless pump might reduce the mass of the propellant tank from 1M to 0.1M, but the pressurant tank would still be at 1M. So tankage in total would go from 2M to 1.1M, which is still significant of course, but only by a factor of about 2, rather than a factor of 10.

You can probably use composites for the pressurant tank however, even if the propellant isn't compatible with composites, so that might give additional efficiency.

That makes a lot of sense.  Since the pressurant tank is much smaller than the propellant tank, it's probably easier for it to have an optimal (spherical) shape, improving the mass efficiency a little further.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/09/2014 09:55 am
Would a compressed air / helium powered turbopump be a reasonable alternative to a Flometrics-style pistonless pump?

One of the upgrades being studied for the STS was to replace the hydrazine APUs in the SRBs with a high pressure helium system that blew down through a turbine.  Don't know how far it got but it was being given serious consideration at one time (prior to STS-107).

It seems to me that the pistonless pump is thermodynamically quite efficient.  I would think a compressed-gas-powered turbopump would be more complicated and less efficient, since the energy would have to be converted from one form to another.

For an APU, on the other hand, a turbine is necessary, since the whole point is to spin a generator.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/09/2014 12:21 pm
It seems to me that the pistonless pump is thermodynamically quite efficient.  I would think a compressed-gas-powered turbopump would be more complicated and less efficient, since the energy would have to be converted from one form to another.

Interesting point, it would be good to see some analysis. I don't think a compressed-gas turbine is more complicated however. Or it might be a bit more complicated, but also a lot more mature, since there are many engineering firms that design them for industrial applications. You can more or less order them off the shelf or go through a routine procurement process for a custom solution. It only gets tricky when you have a high temperature gas generator driving the turbine. I wonder how difficult a closed cycle steam turbine would be compared to the alternatives. More complex perhaps, but again very mature technology. Like a pistonless pump with a gas generator powered by a heat exchanger and like XCOR's piston-pump it would be somewhat similar to a traditional expander cycle engine.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/09/2014 05:43 pm
Only a small pressure vessel needs to be at high pressure, the propellant tanks themselves can be kept at low pressure. A small tank is filled with propellant from the main propellant tanks at low pressure, then the feed valve is closed, the pressurisation valve is opened, once the pressure has equalised it is closed again and then the discharge valve is opened and just before it's empty, it is closed again, and then the cycle repeats. It's a bit more complicated than that, but that's the general idea.

How the pistonless pump operates (http://www.rocketfuelpump.com/technology/operation/)
Why is the pump gas vented overboard instead of being used to pressurise the main tank?

High pressure gas should be able to be vented into the main tank until it reaches a safe max pressure, then the rest vented overboard.

The diagram shows the propellant being pumped to 4 MPa while the ullage pressure in the tank is 300 kPa -- just 7.5% as much.  To a first approximation, that should mean that exhausting the cylinder to the tank would save less than 7.5% of the gas, since gas will still have to be exhausted to the atmosphere to get the cylinder pressure below the tank pressure for the start of a new cycle.  At the same time, exhausting to the tank at 300 kPa instead of the atmosphere at 100 kPa probably slows things down.  So, my guess is that saving a few percent of the gas isn't worth the extra time needed, since a slower cycle corresponds to a bigger, heavier cylinder.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/09/2014 09:24 pm
Is it really true that the pistonless pump is thermodynamically more efficient? It's true that you're not converting energy twice, but the gas that exits a turbine after doing its work can be at very low pressure, while the gas that exits the pistonless pump is still at very high pressure. From that perspective, it looks as if the turbine pump would actually be more efficient.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 11/09/2014 11:08 pm

Is it really true that the pistonless pump is thermodynamically more efficient? It's true that you're not converting energy twice, but the gas that exits a turbine after doing its work can be at very low pressure, while the gas that exits the pistonless pump is still at very high pressure. From that perspective, it looks as if the turbine pump would actually be more efficient.
Plus, the turbine is a thermal machine. If it is fed by the regen system or heat exchanger, it actually works like an energy recovering system.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/10/2014 12:16 am
Plus, the turbine is a thermal machine. If it is fed by the regen system or heat exchanger, it actually works like an energy recovering system.

True, but that also applies to a pistonless pump using a heat exchanger. For an apples to apples comparison you should compare the two systems when combined with the same pressurisation system, whether that is stored gas or something coming out of a heat exchanger. Based on my current understanding I'd say you would likely prefer a turbine pump over the equivalent pistonless pump in both cases.
Title: Re: Rocket Engine Q&A
Post by: R7 on 11/11/2014 10:22 am
The "disk" might be to prevent sloshing, ingestion of the gas, and loss of working fluid when the cylinder is vented. It may also not have a very tight fit and just float on the working fluid.

I wonder if the Flometrics definition of "piston" is it requires a rod to push it.

Also can act as a thermal barrier if needed; if you pump cryogenic fluids you want to reduce the pressurant gas cooling and condensing.

Some of the exhausted pressurant probably would be used to pressurize the main tank ullage. The mass savings vs conventional pressure fed comes from reduced main tank ullage pressure (tank mass directly proportional to the pressure) and venting most of the pressurant on the way up.

Have been wondering why this technology hasn't found its way to any of the newspace nano/microlaunchers yet.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/11/2014 06:31 pm
Would an expander bleed cycle that uses a hydrogen expander to pump a hydrocarbon and oxygen into a combustion chamber make for a moderate-thrust rugged (reusable)  engine?

No, the advantage of expander bleed is that you have the hydrogen anyway. Storing a, relatively, small amount of a deep cryogen just to run the turbine would be overly complicated.

I've read that high temperatures are what makes turbopumps difficult, so maybe using water as a third fluid to power a steam turbine that drives a centrifugal pump might make sense for a cheaper pump-driven kerolox engine that offers better performance than a compressed gas turbopump or an electrical pump.

Edit: heh, I'm getting old, we've discussed some of this before:

http://forum.nasaspaceflight.com/index.php?topic=16279.msg615769#msg615769 (http://forum.nasaspaceflight.com/index.php?topic=16279.msg615769#msg615769)
Title: Re: Rocket Engine Q&A
Post by: gin455res on 11/11/2014 07:09 pm
I was looking at the wikipedia article on hydrazine. It stated that when hydrazine is used as a monopropellant it can reach temperatures of 800degC (approx 1100degK) and gets about 220isp.

If you burnt it with a little of oxygen to reach 1300degC (1600degK), what would the isp be?

From the depths of my memory I thought v squared was proportional to t

so v should be proportional to root t
hence:
v2/v1=root(t2/t1)
hence:
v2=220xroot(1600/1100) =265

and since 1300degC is below some turbine inlet temperatures, could one build a simple rocket engine that was essentially just a full-flow turbo-pump with no afterburner. And would this be useful intermediate between a pressure fed monopropellant and a (conventional) pump-fed bi-propellant?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 11/12/2014 11:56 pm
Have been wondering why this technology hasn't found its way to any of the newspace nano/microlaunchers yet.

I think a compressed air turbine or an electrical motor powering a centrifugal pump would be a better next step for Masten, with the compressed air motor probably being cheaper to buy / develop, and the electrical motor having better performance. The compressed air turbine would have the advantage of venting its working fluid continuously while battery mass is constant during flight, but the battery would have higher energy density and the electric motor would have better efficiency.

The compressed air turbine would also be a stepping stone towards an expander-like steam turbine with even better performance.
Title: Re: Rocket Engine Q&A
Post by: Phillip Clark on 11/29/2014 04:54 pm
I have been trawling the internet to find details of the spacecraft engines designed by Khimmash, which started out as being Isayev's OKB-2.

This link gives a lot of designators for the engines and details of their applications, but it does not give the engine performance data (thrusts, Isp, design lifetimes, O/F ratio, number of restarts, physical dimensions, dry masses - yes, my "dream wish list"!): http://www.b14643.de/Spacerockets_1/Diverse/KB-Isayev_KDUs/

There used to be a site which seemed to list data at FSU rocket/spacecraft propulsion systems: Venik's Aviation which was based at http://www.aeronautics.ru, but that seems to have disappeared (although I have archived the pages).   That included some numerical data but is was incomplete and by today's standard probably dated.

Can anyone please point to a site which gives detailed numerical performance data for the propulsion systems produced by Isayev/Khimmash, please?   Of course, the simple answer might be that the data are simply unavailable .......

Title: Re: Rocket Engine Q&A
Post by: Danderman on 11/29/2014 06:34 pm
http://novosti-kosmonavtiki.ru/forum/forum9/topic9898/

NK forum thread about Isaev engines.
Title: Re: Rocket Engine Q&A
Post by: Phillip Clark on 11/29/2014 09:28 pm
Thank you Danderman!   I think Google translate will be rather busy as I explore the discussion.   There are what appear to be listings of the Isayev engines and numerical data (I love numbers! - they are fun to play with) which I will download and then I hope i can make them ;larger so I can read them.
Title: Re: Rocket Engine Q&A
Post by: Zero-G on 11/29/2014 09:36 pm
Here is some data about Isayev engines. Is this what you are looking for?
http://www.kbhmisaeva.ru/main.php?id=31
http://www.lpre.de/kbhm/index.htm
Title: Re: Rocket Engine Q&A
Post by: Phillip Clark on 11/29/2014 09:50 pm
Thank you Zero-G!   The second link is to a better quality and earier-to-read version of the tables which are copied to the NK forum which Danderman guided me to.

Simply to ensure accuracy, is it possible to translate the various column headings into English please?   Google Translate is fine for a rough translation but is not good for accurate translations!
Title: Re: Rocket Engine Q&A
Post by: Zero-G on 11/29/2014 10:40 pm
You are welcome, Phillip! I can't help you with the translation, since I don't speak Russian, sorry.
Title: Re: Rocket Engine Q&A
Post by: Danderman on 11/30/2014 01:54 am
Thank you Zero-G!   The second link is to a better quality and earier-to-read version of the tables which are copied to the NK forum which Danderman guided me to.

Simply to ensure accuracy, is it possible to translate the various column headings into English please?   Google Translate is fine for a rough translation but is not good for accurate translations!

Okay, dokey, here's what I could figure out:

Engine
Development year
Propellant components
Ignition type
Thrust, tons
Specific impulse, seconds
Chamber pressure, atmospheres
?
?
Burn time?
Mass, kg
Height, mm/Diameter, mm
Notes
Title: Re: Rocket Engine Q&A
Post by: baldusi on 11/30/2014 01:04 pm
The first unkown column looks like main combustion chambers.
Title: Re: Rocket Engine Q&A
Post by: Kerbonaut on 11/30/2014 02:11 pm
Simple Question: Does anybody know good books about space-rockets and/or propulsion? Please only in English/German.

Thanks!  ;D

Kerbonaut
Title: Re: Rocket Engine Q&A
Post by: Krevsin on 11/30/2014 03:01 pm
Well, I got my introduction to rocketry in the form of The Problem Of Space Travel: The Rocket Engine (http://books.google.si/books?id=te15mpHmmTwC&printsec=frontcover#v=onepage&q&f=false) by Hermann 'Noordung' Potočnik.
I have found it to be a very comprehensible and well-illustrated introduction to rocket science. I am unsure as to the quality of the translation, since I read it in slovene, but I highly recommend this book.

Similarily, I have found the non-fiction section of the Atomic Rockets' Suggested Reading page (http://www.projectrho.com/public_html/rocket/reading.php#id--Non-Fiction) to be most helpful.

I also recommend reading the works by Hermann Oberth (https://www.google.si/search?tbo=p&tbm=bks&q=inauthor:%22Hermann+Oberth%22&gws_rd=ssl) if you're interested in an overview of the basics of rocket science.
Title: Re: Rocket Engine Q&A
Post by: msat on 12/08/2014 05:24 am
I've been wondering about the logic behind the the canting of the NF-104's rocket engine. I can't find enough relevant specs on this aircraft, but I suppose the images provide enough information to allow us to make some guesses. I've taken 3 images and drew lines from the nozzle end of the rocket motor towards the fuselage (see attached images).

I think the image with red lines is the most accurate, but what is the thrust line intersecting with? CL? CG? What is the possible reasoning behind it? Is it to provide a horizontal thrust component for the zoom climb, or just stability reasons (in the case of it intersecting with CG)?
Title: Re: Rocket Engine Q&A
Post by: Helodriver on 12/08/2014 06:27 am
The engine, a liquid fueled rocket, was canted a few degrees to make up for the fact it was mounted above the aircraft centerline, that space being taken by the J-79 jet engine. The angle, directed through the aircraft center of mass,  allowed the rocket to be fired without causing a pitching moment that could not be overcome with aircraft control authority available at high altitude from the control surfaces or RCS jets. The Space Shuttle Main Engines were mounted with a similar angle for similar reasons, due to the mass of the external tank below the orbiter.
Title: Re: Rocket Engine Q&A
Post by: msat on 12/08/2014 06:32 am
Awesome! Thanks for the info!
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 12/08/2014 02:36 pm
A small nit on the Space Shuttle ET, most of the mass was in the LOX tank, which was at the top of the tank. So the ET center of mass was technically above, or to be more correct forward of the shuttle...

I do believe  the shuttle also flew underneath the tank during powered flight.

On the NF-104 was the line directed through the center of mass or center of pressure?
Title: Re: Rocket Engine Q&A
Post by: Helodriver on 12/08/2014 04:02 pm
A small nit on the Space Shuttle ET, most of the mass was in the LOX tank, which was at the top of the tank. So the ET center of mass was technically above, or to be more correct forward of the shuttle...

I do believe  the shuttle also flew underneath the tank during powered flight.

On the NF-104 was the line directed through the center of mass or center of pressure?

As far as shuttle goes, underneath, above, it does not really matter the shuttle was capable of flying in either heads up or heads down attitude. The key, which I could have been clearer on, was that the main engines were canted to balance thrust due to the asymmetrical mass distribution of the stack and the distance away from centerline of the engine mountings.

NF-104 thrust went through center of mass, as the CoP shifted and then became less of an issue rapidly with the drop in atmospheric density during the near vertical climb.
Title: Re: Rocket Engine Q&A
Post by: IslandPlaya on 12/08/2014 04:14 pm
Forgive me for maybe asking a question that has maybe been answered many times before...
If the shuttle could fly in any roll orientation then why did we NASA choose the shuttle down orientation it did?
Surely it could have gone uphill 'sideways' and it wouldn't matter?
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/08/2014 04:38 pm
Forgive me for maybe asking a question that has maybe been answered many times before...
If the shuttle could fly in any roll orientation then why did we NASA choose the shuttle down orientation it did?
Surely it could have gone uphill 'sideways' and it wouldn't matter?


no, it couldn't go side way.  It flew inverted to reduce the aero loads on the wings.  Also to give the crew a view of the horizon.  It roll to the out side later in the mission to provide antenna views to TDRSS



There is also SRB separation to consider.
Title: Re: Rocket Engine Q&A
Post by: IslandPlaya on 12/08/2014 04:44 pm
Thanks Jim. Why would the orientation matter to loads on the wings except the 1g downwards (Maybe answered my own question here.)
Other points... Yes.
SRB sep. Is it because the thin atmos at that height still provides lift?
Title: Re: Rocket Engine Q&A
Post by: indycar82 on 12/08/2014 04:52 pm
I have a couple questions in regards to the Orion launch last Friday. I am new to the forum, so hopefully this question is being asked at the right thread.

Question 1: During the initial stages of launch friday there were some very bright flashes coming from the exhaust of the engines. was this from materials coming off the vehicle itself or was it coming from the RS-68 engines?

Question 2: The exhaust plume from the RS-68 are much brighter and have more color than the SSME  which were alot lighter, even colorless. What is the reason for this?

Apologies if this has already been answered. I looked about in the vast NSF library, but I could not find anything.

Thanks
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/08/2014 04:54 pm
I have a couple questions in regards to the Orion launch last Friday. I am new to the forum, so hopefully this question is being asked at the right thread.

Question 1: During the initial stages of launch friday there were some very bright flashes coming from the exhaust of the engines. was this from materials coming off the vehicle itself or was it coming from the RS-68 engines?

Question 2: The exhaust plume from the RS-68 are much brighter and have more color than the SSME  which were alot lighter, even colorless. What is the reason for this?

Apologies if this has already been answered. I looked about in the vast NSF library, but I could not find anything.

Thanks

Both are from the ablative nozzles
Title: Re: Rocket Engine Q&A
Post by: indycar82 on 12/08/2014 04:57 pm
Thanks Jim !
Title: Re: Rocket Engine Q&A
Post by: AS-503 on 12/08/2014 04:57 pm
I have a couple questions in regards to the Orion launch last Friday. I am new to the forum, so hopefully this question is being asked at the right thread.

Question 1: During the initial stages of launch friday there were some very bright flashes coming from the exhaust of the engines. was this from materials coming off the vehicle itself or was it coming from the RS-68 engines?

Question 2: The exhaust plume from the RS-68 are much brighter and have more color than the SSME  which were alot lighter, even colorless. What is the reason for this?

Apologies if this has already been answered. I looked about in the vast NSF library, but I could not find anything.

Thanks

Hello and welcome to the forum.

The combined (combusted) hydrogen and oxygen exhaust from the SSME is a clean semi-transparent blue color.
The SSME has a regenerative cooled nozzle. The RS-68 burns the same fuel and oxidizer but it has an ablative nozzle which chars and burns off in order to shed heat. This debris is what is causing the exhaust not to be like that of the SSME. 


Edit: Jim beat me to it!
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/08/2014 04:58 pm
Thanks Jim. Why would the orientation matter to loads on the wings except the 1g downwards (Maybe answered my own question here.)
Other points... Yes.
SRB sep. Is it because the thin atmos at that height still provides lift?


Aero loads and not gravity loads.

It flies at a negative angle of attack and that can't be done in any other attitude.

And it is loads on the orbiter -  ET attachment.
Title: Re: Rocket Engine Q&A
Post by: IslandPlaya on 12/08/2014 05:11 pm
Thanks for your perseverance Jim... But...
Aero loads wouldn't matter with the roll orientation if gravity doesn't matter. So any roll would do.
-ve AOA gotcha.
Please expand on ET attachment loads... am not clever enough to read between the lines.
Many Thanks!
Title: Re: Rocket Engine Q&A
Post by: Jim on 12/08/2014 06:01 pm

Aero loads wouldn't matter with the roll orientation if gravity doesn't matter. So any roll would do.
-ve AOA gotcha.


wrong, aeroloads are depended on flight path and vehicle attitude with respect to the attitude.   Because of the location of the SSME's, the shuttle is aways flying at an angle of attack. 
Title: Re: Rocket Engine Q&A
Post by: IslandPlaya on 12/08/2014 06:11 pm
ok, gotcha
Title: Re: Rocket Engine Q&A
Post by: pericynthion on 12/08/2014 10:27 pm
As far as shuttle goes, underneath, above, it does not really matter the shuttle was capable of flying in either heads up or heads down attitude.

Jim has already mentioned that the shuttle can only fly a 'heads-down' ascent trajectory due to aero loads; Carl Ehrlich's wonderfully titled and quite accessible paper (https://www.aiaa.org/uploadedFiles/About-AIAA/History_and_Heritage/Final_Space_Shuttle_Launches/Why_the_Wings_Stay_On.pdf) gives some more details.
Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 01/05/2015 04:46 pm
Hello all!

I was wondering if someone could explain to me the simplest method to approximate chamber pressure, optimum convergent/divergent rocket nozzle angles, and optimum throat diameter all for a solid rocket motor. I understand this may not be an easy task, but I would appreciate if someone could help me learn and/or point me in the right direction to learn such calculations. I'm patient and eager to learn!

A few questions off the bat though...

Does the chamber pressure of a rocket motor rely heavily on the volume freed up by expelled propellant or can this be ignored for simple approximations? In other words, does the changing non-propellant volume in a rocket motor profoundly affect calculations regarding chamber pressure? Thinking about it, it seems that the mass flow rate would be the dominant factor.

In core burning solid rocket motors, does the propellant insulate itself from the motor casing?

If no space is left between the top of the motor casing and the propellant, is the propellant grain being forced downward? I think not, but figured I'd ask.

If anyone could help me out with my questions and calculations, that would be great! :)
Title: Re: Rocket Engine Q&A
Post by: spacecane on 01/06/2015 01:28 pm
What is the typical efficiency of an RP-1 rocket engine?  Essentially, how much of the energy stored in the fuel is output as thrust?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/06/2015 02:22 pm
What is the typical efficiency of an RP-1 rocket engine?  Essentially, how much of the energy stored in the fuel is output as thrust?

That's a tricky question to answer, because thrust is not an energy or an energy flow rate (power), so it can't be directly compared to chemical energy.  We can ask how much of the chemical energy in the propellants winds up as kinetic energy of the rocket, but then the answer depends on how fast the rocket is moving.

You can calculate the rate at which chemical energy is provided (flow rate of the fuel multiplied by its heat of combustion, say) and divide that by the power of the exhaust jet (0.5 times the mass flow rate times the square of the effective exhaust speed).  That will probably give you a pretty good efficiency -- like 70% or better.  Maybe even around 90% if we're talking about a high-expansion hydrogen-fueled engine in a vacuum.

But then the question is how fast is the kinetic energy of the rocket itself increasing.  The rate at which it's kinetic energy -- ½mv² -- is increasing is mva, where a is the acceleration.

This makes some intuitive sense.  If I use a rocket engine with an exhaust velocity of 3000 m/s to accelerate my car to a speed of 3 m/s, obviously most of the energy winds is in the exhaust jet and very little is in my car.  On the other hand, if I see a rocket flying past of 3000 m/s with its engine on, then from my point of view the exhaust stream is at rest and carries away no residual energy, and all of the energy goes into the rocket.

Just as an aside, let me mention that the standard de Laval (converging-diverging) nozzle is an amazing thing.  It's rare that it's possible to convert heat into another form of energy (kinetic energy of the exhaust gases) so efficiently.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 01/06/2015 05:33 pm
When designing a catalyst pack for decomposing peroxide or hydrazine, is there any attention given to the aerodynamics of the flow?

For example, would a tapering, or waisted, pack act like a venturi. That is, would it reduce the pressure in the flow. Reducing pressure in a reaction thats products have more mols (and consequently more volume) than the starting reactants is an additional mechanism (additional to the catalysis on the pack surface) to drive the reaction in a forwards direction.

And, can we even assume that mixed phase flows obey the Bernoulli equation that explained the venturi effect?

That's undesirable. Reduced pressure reduces the reaction rate, and increased velocity decreases the residence time. You'd have to have a much longer catalyst bed. There's also a loss of total pressure with higher flow rates.

Are we sure that reduced pressure reduces reaction rate in a decomposition reaction?

( http://en.wikipedia.org/wiki/Haber_process#Reaction_rate_and_equilibrium (http://en.wikipedia.org/wiki/Haber_process#Reaction_rate_and_equilibrium) reduced no .of moles in this example favours high pressure - decomposition would be opposite so shouldn't it favour low pressure?)

I thought energy of collision was only dependent on temperature?


h2o2(l) --> h2(g) + o2(g)   1)

h2o2(g) --> h2(g) + o2(g)  2)

In both cases volume increases.  In case 2) the activation energy is lower because the latent heat of vaporization is 0?

Sure a longer catalyst bed would be needed. 
Title: Re: Rocket Engine Q&A
Post by: gin455res on 01/06/2015 05:43 pm
When designing a catalyst pack for decomposing peroxide or hydrazine, is there any attention given to the aerodynamics of the flow?

For example, would a tapering, or waisted, pack act like a venturi. That is, would it reduce the pressure in the flow. Reducing pressure in a reaction thats products have more mols (and consequently more volume) than the starting reactants is an additional mechanism (additional to the catalysis on the pack surface) to drive the reaction in a forwards direction.

And, can we even assume that mixed phase flows obey the Bernoulli equation that explained the venturi effect?

That's undesirable. Reduced pressure reduces the reaction rate, and increased velocity decreases the residence time. You'd have to have a much longer catalyst bed. There's also a loss of total pressure with higher flow rates.

Are we sure that reduced pressure reduces reaction rate in a decomposition reaction?

( http://en.wikipedia.org/wiki/Haber_process#Reaction_rate_and_equilibrium (http://en.wikipedia.org/wiki/Haber_process#Reaction_rate_and_equilibrium) reduced no .of moles in this example favours high pressure - decomposition would be opposite so shouldn't it favour low pressure?)

I thought energy of collision was only dependent on temperature?


h2o2(l) --> h2(g) + o2(g)   1)

h2o2(g) --> h2(g) + o2(g)  2)

In both cases volume increases.  In case 2) the activation energy is lower because the latent heat of vaporization is 0?

Sure a longer catalyst bed would be needed. 

need more sleep sorry

2h2o2(l) --> 2h2o(g) + o2(g)   1)

2h2o2(g) --> 2h2o(g) + o2(g)  2)
Title: Re: Rocket Engine Q&A
Post by: pagheca on 01/09/2015 03:45 pm
Question: does somebody know which are the analytical functions approximating a "typical" engine bell - at least as theoretically derived from "basic" principles?

Do there is one?

Thanks in advance.
Title: Re: Rocket Engine Q&A
Post by: quanthasaquality on 01/09/2015 07:31 pm
For the RD-180, kerosene/lox oxidizer rich staged combustion engine, we've heard from various people that:
Staged combustion in general is hard.
metallurgy for oxidizer rich kerosene/lox staged combustion is difficult.

The Soviet Union developed lox/kerosene, and n2o4/udmh oxidizer rich staged combustion engines in the 1960s. How hard is oxidizer rich staged combustion for N2O4/UDMH, compared to lox/kerosene?

I suppose it will come down to how corrosive supercritical oxygen is, versus supercritical n2o4.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/10/2015 11:16 am
Question: does somebody know which are the analytical functions approximating a "typical" engine bell - at least as theoretically derived from "basic" principles?

I don't think there is any simple analytical function giving the exact optimal shape, but according to what I've read a circular profile at the throat smoothly joined to a parabolic profile further downstream is apparently near optimal.  A textbook like Sutton's Rocket Propulsion Elements provides graphs of the optimal initial and final opening angles of the parabola as a function of nozzle length and expansion ratio.
Title: Re: Rocket Engine Q&A
Post by: MP99 on 01/11/2015 11:30 am
What is the typical efficiency of an RP-1 rocket engine?  Essentially, how much of the energy stored in the fuel is output as thrust?

That's a tricky question to answer, because thrust is not an energy or an energy flow rate (power), so it can't be directly compared to chemical energy.  We can ask how much of the chemical energy in the propellants winds up as kinetic energy of the rocket, but then the answer depends on how fast the rocket is moving.

You can calculate the rate at which chemical energy is provided (flow rate of the fuel multiplied by its heat of combustion, say) and divide that by the power of the exhaust jet (0.5 times the mass flow rate times the square of the effective exhaust speed).  That will probably give you a pretty good efficiency -- like 70% or better.  Maybe even around 90% if we're talking about a high-expansion hydrogen-fueled engine in a vacuum.

But then the question is how fast is the kinetic energy of the rocket itself increasing.

ISTM the answer to the question is the kinetic energy of the exhaust, vs the potential energy of the propellants.

At SL, that's the same, even when test firing on a stand.

Looks like you've answered it nicely, above.

Cheers, Martin
Title: Re: Rocket Engine Q&A
Post by: Raj2014 on 01/13/2015 02:13 pm
Has NASA thought of using Aerospike engines to reach orbit and use bell nozzle engines in space only on their rockets e.g SLS?   
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/13/2015 02:44 pm
Has NASA thought of using Aerospike engines to reach orbit and use bell nozzle engines in space only on their rockets e.g SLS?   

No, because SLS is only using shuttle components and NASA has no other rockets.   The rest of US fleet of rockets is owned by companies such as ULA, Spacex or OSC. 
Title: Re: Rocket Engine Q&A
Post by: Raj2014 on 01/13/2015 03:05 pm
Has NASA thought of using Aerospike engines to reach orbit and use bell nozzle engines in space only on their rockets e.g SLS?   

No, because SLS is only using shuttle components and NASA has no other rockets.   The rest of US fleet of rockets is owned by companies such as ULA, Spacex or OSC.

Yes I understand that, what I mean is, since bell nozzle engines are more efficient in space and aerospike engines are apparently efficient at all altitudes. It make sense to use both engines. Could this also bring down costs and weight?
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/13/2015 03:15 pm

Yes I understand that, what I mean is, since bell nozzle engines are more efficient in space and aerospike engines are apparently efficient at all altitudes. It make sense to use both engines. Could this also bring down costs and weight?

If you understand "that", then why are you asking the question?  "Sense" doesn't enter the picture,  SLS is using shuttle components: efficiency, costs and weight don't matter.  NASA doesn't have other rockets to use aerospikes on. 

Aerospike engines need more research before they can be used operationally
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 01/13/2015 06:02 pm
I thought aerospace where also heavier than bell nozzles, so you are trading that efficiency gain for a mass penalty.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 01/13/2015 07:54 pm
There must be a reason why nobody is using aerospikes operationally. May be they need further research. But they may also have issues like worse T/W, and worse performance after effectively being in vacuum, which is still a lot of the booster's time. The optimization might even depend on the propellant.
There's also the Thrust Augmented Nozzle that Aerojet developed, but never implemented. It might be a lack of funding (they don't usually move a finger without a government contract for main propulsion), or just that is not that great. As always, res non verba. If nobody of the actual rocket scientists use it, it might not be such a huge breakthrough.
Title: Re: Rocket Engine Q&A
Post by: TrevorMonty on 01/13/2015 10:02 pm
There must be a reason why nobody is using aerospikes operationally. May be they need further research. But they may also have issues like worse T/W, and worse performance after effectively being in vacuum, which is still a lot of the booster's time. The optimization might even depend on the propellant.
There's also the Thrust Augmented Nozzle that Aerojet developed, but never implemented. It might be a lack of funding (they don't usually move a finger without a government contract for main propulsion), or just that is not that great. As always, res non verba. If nobody of the actual rocket scientists use it, it might not be such a huge breakthrough.
On paper the aerospike has better performance but it costs a lot to get it from paper to flying HW and there are a lot of unknowns.

There is a possible fullback option for Firefly if they can't make it work. Go to a F9/Electron configuration using the upperstage engine they are developing.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 01/18/2015 03:01 pm
I'm still trying to figure out the efficiency of a Flometrics style pistonless pump vs a dead simple compressed air turbopump. In the process I'm upgrading my still very limited understanding of elementary thermodynamics.

I have a simple spreadsheet that models the work I could ideally get from a given amount of compressed air at a specified temperature and pressure using a perfect pneumatic motor, assuming no friction and adiabatic expansion. The higher the pressure ratio, the better the work, but the discharge temperature also drops precipitately, and we can't go below the boiling point of air. On the other hand, we also want to use high pressure in order to keep the tanks compact.

So now I'm imagining we have a high pressure (say 200 bar) tank discharging into a medium pressure (say 10 bar) tank through a pressure regulator, with the medium pressure tank feeding the pneumatic motor. How do I correctly model the discharge through the regulator? Can I simply assume a pressure ratio of 10:1 over the motor and constant temperature and mass in the medium pressure tank as long as the high pressure tank is capable of keeping the pressure in the medium pressure tank at 10 bar? Or would the high pressure tank affect the temperature in the medium pressure tank? To keep it simple I'm not interested in transients, Joule Thomson cooling etc.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/18/2015 04:54 pm
I'm no expert, but it seems to me that a regulator is simply a throttled expansion.  Throttled expansion leaves enthalpy unchanged.  The enthalpy of an ideal gas solely a function of temperature.  Hence, constant temperature across the regulator is correct for an ideal gas.

I'll be interested in your calculations, by the way.
Title: Re: Rocket Engine Q&A
Post by: ZachS09 on 01/22/2015 08:20 pm
Since the canceled Ares V rocket would have used five to six RS-68B engines, I have a question about the ignition sequence. Had Ares V not been canceled, and it still used the RS-68s, would a huge fireball erupt upwards at ignition like the Delta IV?
Title: Re: Rocket Engine Q&A
Post by: Chris Bergin on 01/23/2015 02:04 pm
Since the canceled Ares V rocket would have used five to six RS-68B engines, I have a question about the ignition sequence. Had Ares V not been canceled, and it still used the RS-68s, would a huge fireball erupt upwards at ignition like the Delta IV?

Yep!
Title: Re: Rocket Engine Q&A
Post by: baldusi on 01/24/2015 02:20 am
Since the canceled Ares V rocket would have used five to six RS-68B engines, I have a question about the ignition sequence. Had Ares V not been canceled, and it still used the RS-68s, would a huge fireball erupt upwards at ignition like the Delta IV?

Yep!
Isn't the self inmolation a feature of pad design (namely lack of burners)
Title: Re: Rocket Engine Q&A
Post by: DaveS on 01/24/2015 02:31 am
Since the canceled Ares V rocket would have used five to six RS-68B engines, I have a question about the ignition sequence. Had Ares V not been canceled, and it still used the RS-68s, would a huge fireball erupt upwards at ignition like the Delta IV?

Yep!
Isn't the self inmolation a feature of pad design (namely lack of burners)
No. Both SLC-37B at CCAFS and SLC-6 at VAFB feature the Radially Outward Firing Ignitors (ROFIs).
Title: Re: Rocket Engine Q&A
Post by: CorvusCorax on 01/24/2015 07:51 pm
Hi.

A thrust augmentor is a hollow tube (nozzle shaped) attached right behind the main nozzle of a thrust producing engine. It's effect is that it takes excess energy from a high velocity volume exiting the main engine to accelerate surrounding fluid that also enters it in the same direction and creates a lower velocity higher volume with the same impulse but effectively higher thrust force.

They are regularly used by those freaks (in the positive sense) who build pulse jets in their garage, then strip them to their bikes and post footage on youtube ;)

Since those have HOT exhausts, the heat of the primary exhaust causes the sucked in extra air to expand, turning the augmentor work effectively like a ramjet (fueled by the excess-heat of the primary engine exhaust) which increases the efficiency significantly.

The same would go for rocket, just even more so (higher exhaust velocities, higher temperatures) I was wondering if a rocket could be equipped with a thrust augmentor, like those guys use on their pulse jets, since especially during launch and early ascend it could save huge amounts of propellant (like 50% less fuel used to get to Mach 1)

All I could find was some general research work on thrust augmentors for generic jets done by NACA - as well as (interestingly) and more modern patent applications for using them - also for pulse jet engines - for vtol drone (which is a joke, since youtube is full of prior art, but well - patents ... *snicker*)

But I found not a single mentioning of a rocket engine augmenting its thrust while in atmosphere.

In retrospect it almost seems a no-brainer, so why hasn't it been done?

Or has it been done and I missed it?  Is there any obvious reason why it would fail or be a bad idea?

Just in case I wrote a pdf document laying out the concept, attached. I hope this is the right forum to discuss such stuff ;)

Title: Re: Rocket Engine Q&A
Post by: baldusi on 01/25/2015 04:10 am
Wikipedia (http://en.m.wikipedia.org/wiki/Air-augmented_rocket) has an article with a good summary of the issues. LV actually spend little time where there's enough air to make a difference and after that it's just too heavy. Oldest developed model was Soviet ICBM Gnom.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/25/2015 08:21 am
On a very small scale, I have encountered the term "augmenter tube" in connection with Jetex solid-propellant motors for model aircraft (http://jetex.org).
Title: Re: Rocket Engine Q&A
Post by: CorvusCorax on 01/25/2015 11:43 am
Wikipedia (http://en.m.wikipedia.org/wiki/Air-augmented_rocket) has an article with a good summary of the issues. LV actually spend little time where there's enough air to make a difference and after that it's just too heavy. Oldest developed model was Soviet ICBM Gnom.

That's interesting, that would indicate that the concept in theory works even better than I expected : All the way through the atmosphere, even in supersonic regimes.

I guess however that also explains the problems encountered. The (non jetisonable) high weight and complex inlet design come with the problems designing an inlet that works through this wide realm of speed regimes.

In subsonic flight you can simply just stick a pipe past the exhaust to get a significant boost but when reaching sonic velocities you'd likely need  (do you?) an integrated design that leads air past the engine and compresses it prior to heating in a supersonic RAM layout.  At late ascend you'd even need scramjet like aerodynamics, with the airflow through the engine being supersonic.  (Arguably without the main issue of scramjets - keeping the flame lit)

But that attempt to make an airbreathing engine all the way until it exits the atmosphere seem to be overly complicated. It wastes the large potential even a very simple solution would offer at/around launch, while a hypersonic design would likely only increase performance marginally when still flying subsonic (and, as mentioned in the article, adds lots of non-jetisonable weight)

So maybe NASA's SSTO design is simply too ambitious. By trying to make such an engine they make the problem "hard" and the obvious benefits vanish.  That sounds to me a bit like the story of the shuttle. They tried so hard to make a reuseable design that it ended to be more expensive than a simple throwaway rocket.

The russian story is just sad, Its one of many where they had one contructor follow an idea, it worked well, and just died with him when he did. Didn't the same happen with their ecranoplane and so many other projects?
Title: Re: Rocket Engine Q&A
Post by: CorvusCorax on 01/25/2015 12:03 pm
I found only this document on the NASA GTX project referenced from the wiki article

http://trajectory.grc.nasa.gov/aboutus/papers/NASA-TM-2003-212315.pdf

as far as I read that, it's not an augmented rocket design, but an actual air-breathing rocket using atmospheric oxygen instead of carryon oxydizer for combustion in a traditional RAM/SCRAMJET configuration (at least in the $300M technology demonstrator)

(nevertheless pretty cool!!!)
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 01/25/2015 12:53 pm
I'm no expert, but it seems to me that a regulator is simply a throttled expansion.  Throttled expansion leaves enthalpy unchanged.  The enthalpy of an ideal gas solely a function of temperature.  Hence, constant temperature across the regulator is correct for an ideal gas.

Hmm, it looks as if the medium pressure tank isn't even needed, you could connect a turbine directly to the regulator. You say temperature is constant across the valve, which I assume also means that the temperature of the remaining gas in the tank remains constant. All the work and cooling down would then be done in the turbine itself, after the gas has passed through the valve. The gas passing through the valve would also cool down a little bit from converting some microscopic thermal energy into macroscopic kinetic energy. If true, that would be very convenient, but I don't think that's realistic for real tanks. For instance, doesn't the tank of a propane burner cool as it slowly empties? If so, is that exclusively due to Joule Thomson cooling?

The question I'm trying to clear up is if you can simply model the system as adiabatic expansion from 10 bar to atmospheric pressure without worrying about the 100 bar tank cooling down / losing too much pressure.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 01/25/2015 01:28 pm
Never mind the propane, I just realised that there is liquid propane inside the tank, which would inevitably cool down as it evaporated.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/25/2015 07:51 pm
There is no change in enthalpy in the throttling process, but I think the gas in the tank is doing work on the gas flowing into the regulator and will therefore tend to cool, even if ideal.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 01/25/2015 08:50 pm
There is no change in enthalpy in the throttling process, but I think the gas in the tank is doing work on the gas flowing into the regulator and will therefore tend to cool, even if ideal.

So how much work is the gas inside the high pressure tank doing and what kind of work precisely? In the end all the energy that the gas loses on the whole path from tank to turbine outlet is due to the work done during the expansion in the turbine, isn't it? Of course, in the real world there will always be friction etc.

EDIT: as for ideal gases, doesn't the derivation of the ideal gas law assume that there is no interaction between the particles? If so, can a parcel of ideal gas ever do work on another parcel of ideal gas?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/26/2015 10:11 am
All good questions, which bring us back to re-examine the fundamentals.  After thinking about it a bit, I believe that what "non-interacting" must mean is that there are no forces between molecules unless they are very close to one another, "very close" meaning separated by a distance much less than the average distance between molecules.  This condition is necessary so that the energy of a molecule is simply the sum of the kinetic energy arising from the motion of its center of mass and its internal degrees of freedom.

To look at it another way, if "non-interacting" meant collisionless, we wouldn't be able to approximate air at STP as an ideal gas, because the mean free path is very short, like a micron or less.  The fact that the path is so short means that we can think of a parcel of gas as doing work on a neighboring parcel.  In the present system, we could imagine a thin piston placed in the duct between the high-pressure tank and the regulator.  As the piston moves along the duct, work is done by the gas upstream of the piston on the gas downstream.  Take the piston away, and the same work is performed.

As an example, the textbook I referred to in refreshing my memory about throttled expansions derives the conservation of enthalpy as follows.  Consider a volume V1 of gas at pressure p1 just upstream from the throttle.  As it passes into the throttle, work p1V1 is done on it by the gas further upstream.  As the parcel emerges from the throttle with volume V2 and pressure p2, it does work p2V2 on the gas further downstream.  If there is no heat transfer during the throttling, then the conservation of energy requires that E1 + p1V1 = E2 + p2V2, where E is the internal energy of the gas.  Of course, enthalpy is defined precisely as E + pV, so throttling conserves enthalpy.
Title: Re: Rocket Engine Q&A
Post by: Raj2014 on 01/26/2015 09:47 pm
Which rockets will NASA use to send Astronauts to and from Mars?   
Title: Re: Rocket Engine Q&A
Post by: Jim on 01/27/2015 12:29 am
Which rockets will NASA use to send Astronauts to and from Mars?   

There is only one NASA rocket, SLS.  At this time, there is no program to go to mars.  The white house and congress have yet to direct and fund NASA for such an endeavor.
Title: Re: Rocket Engine Q&A
Post by: scienceguy on 02/24/2015 04:36 pm
What are the chemicals in kerosene? Methane is CH4, gasoline is octane, or C8H18. Is kerosene a bunch of different chemicals?
Title: Re: Rocket Engine Q&A
Post by: beb on 02/24/2015 04:53 pm
What are the chemicals in kerosene? Methane is CH4, gasoline is octane, or C8H18. Is kerosene a bunch of different chemicals?
While I can't answer your question, I can tell you that gasoline is a mixture of hydrocarbons and not simply Octane. Kerosene is undoubtedly also a mixture of heavier hydrocarbons.  RP-1 is a refined version of kerosene but that's all I know.
Title: Re: Rocket Engine Q&A
Post by: PahTo on 02/24/2015 04:57 pm

I would add that gasoline is a by-product of the refinement/production of kerosene from crude oil.  Way back when, it was considered a waste product of said process, which is in part what led to the gasoline fueled combustion engine and our beloved automobiles.
Title: Re: Rocket Engine Q&A
Post by: Jim on 02/24/2015 05:05 pm
C12H26
Title: Re: Rocket Engine Q&A
Post by: scienceguy on 02/24/2015 06:16 pm
thanks Jim
Title: Re: Rocket Engine Q&A
Post by: deltaV on 02/25/2015 04:05 am
I'm pretty sure Jim's formula is only an average and RP-1 is made up of a variety of hydrocarbons.
Title: Re: Rocket Engine Q&A
Post by: R7 on 02/25/2015 10:35 am
I'm pretty sure Jim's formula is only an average and RP-1 is made up of a variety of hydrocarbons.

True. (http://forum.nasaspaceflight.com/index.php?topic=32112.msg1063854#msg1063854)

There is no single exact chemical composition for kerosene because its a common term for various fuel products and even individual fuel product composition varies some from batch to batch.

Google is your friend, googling "kerosene composition" gives more, for instance http://www.atsdr.cdc.gov/toxprofiles/tp121-c3.pdf which has JP-8 chemical breakdown. Jim's dodecane is most abundant in it, but still only 22.5wt%.

Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 04/04/2015 07:29 pm
The Centaur IVF talk has me wondering something.

Specifically a turbo pump has a thermal efficiency of ~60%, on IVF using a piston engine to generate electricity is maybe 30%, switching to fuel cells jumps that to 80%, maybe 85%.

So I am wondering, would switching to large fuel cells and electric propellant pumps result in an engine that is simpler while still maintaining the high ISP of expander and staged combustion engines?

Thoughts?
Would the extra mass of the fuel cells and motors weigh more than a traditional engine?
Or is it to far out in left feels?
Title: Re: Rocket Engine Q&A
Post by: Jim on 04/04/2015 07:48 pm
The Centaur IVF talk has me wondering something.

Specifically a turbo pump has a thermal efficiency of ~60%, on IVF using a piston engine to generate electricity is maybe 30%, switching to fuel cells jumps that to 80%, maybe 85%.

So I am wondering, would switching to large fuel cells and electric propellant pumps result in an engine that is simpler while still maintaining the high ISP of expander and staged combustion engines?

Thoughts?
Would the extra mass of the fuel cells and motors weigh more than a traditional engine?
Or is it to far out in left feels?

I think you have to look at power density.
Title: Re: Rocket Engine Q&A
Post by: R7 on 04/05/2015 03:56 pm
The Centaur IVF talk has me wondering something.

Specifically a turbo pump has a thermal efficiency of ~60%, on IVF using a piston engine to generate electricity is maybe 30%, switching to fuel cells jumps that to 80%, maybe 85%.

The 60% sounds like turbine efficiency, not thermal.

You forget the heat from IVF ICE is not waste, it does useful work creating ullage pressurant gasses and the exhaust provides settling thrust. Read the links in the ULA IVF thread, there are explanations why fuel cell is not a good choice here.
Title: Re: Rocket Engine Q&A
Post by: Oli on 04/09/2015 09:58 am

If I'm not mistaken bigger turbo pumps tend to be more efficient. Anybody knows what the influence on engine isp might be?

I'm asking because I wonder whether one can expect better isp from bigger engines, all other things equal.
Title: Re: Rocket Engine Q&A
Post by: Damon Hill on 04/09/2015 10:57 am

If I'm not mistaken bigger turbo pumps tend to be more efficient. Anybody knows what the influence on engine isp might be?

I'm asking because I wonder whether one can expect better isp from bigger engines, all other things equal.

The answer seems to be complicated:

I think average molecular weight of the exhaust dominates, but expansion ratio of the nozzle and chamber pressure/temperature have a say in the matter.  Then there's a major difference in how much of the propellants go through the nozzle and how much is dumped out of the gas generators if it's an open cycle design as most liquid rocket engines are.  That's (partly) why the RL10, which is a closed cycle expander with no hot gas generator has a much better Isp than the far larger RS-68 (in a vacuum, of course).

The high pressure staged combustion SSME, with a fully expanded nozzle, could wipe the floor with the RL10, but the nozzle would be enormous and couldn't be safely used in the atmosphere because of turbulent flow separation.  The RS-68 would be significantly improved, but the same limits apply.  Certain solutions that can adjust expansion ratio like plug and aerospike nozzles are possible but not used in practice; some use of extendable nozzles is made but partly to keep the interstage length down until after staging.  One version of the RL10 does use an extendable nozzle and gets the highest Isp.  The SSME is slightly overexpanded as it is, but the general approach has been to accept the compromise when the engine operates the full range of atmospheric pressure, and let the upper stage engines benefit from operating only in a vacuum providing a fully expanded nozzle can be accommodated.

The mixture ratio could be varied too, but that's not commonly done.  Deep throttling of thrust can impair Isp a bit; go too deep and the engine gets into problems.

I think sheer power of the turbopumps tends to trail the pack, so to speak because of other factors mentioned.  As with just about anything, it's all about compromises.
Title: Re: Rocket Engine Q&A
Post by: kevin-rf on 04/14/2015 02:38 pm
Interesting, looks like someone did the trades and decided pumps powered Brushless DC motors with LiPo's was worth the effort.

http://www.rocketlabusa.com/about-us/propulsion/rutherford/

It's a small motor for a small rocket, but interesting none the less.
Title: Re: Rocket Engine Q&A
Post by: QuantumG on 04/14/2015 10:30 pm
I remember Peter Beck asking about electric turbopumps on some space forum about five years ago. I don't think anyone thought he'd actually go do it. A different group - Ventions (http://ventions.com/2014-news/) - flew an electric pump fed rocket at FAR on July 7, 2013 and they received a patent on their technique last year.
Title: Re: Rocket Engine Q&A
Post by: Damon Hill on 04/15/2015 02:47 am
Interesting, looks like someone did the trades and decided pumps powered Brushless DC motors with LiPo's was worth the effort.

http://www.rocketlabusa.com/about-us/propulsion/rutherford/

It's a small motor for a small rocket, but interesting none the less.

50 horsepower is about 37 kW for several minutes; Tesla does this all the time, but launch is weight critical.  It'd certainly make for a simpler rocket engine not having to manage a hot gas generator.  50 hp from a beer can motor is impressive.

Think I'll keep an eye on them.
Title: Re: Rocket Engine Q&A
Post by: TrevorMonty on 04/15/2015 04:46 am
37kw motor is going produce a significant amount of heat, but nothing a ready supply of LOX couldn't solve.
Title: Re: Rocket Engine Q&A
Post by: Mini_Elon on 05/23/2015 04:27 pm
Hello can anyone help me figure out how a Turbo-Pump works. I know some of the basics of how it works but I trying to figure out how it uses Bernoulli's principle and How does this Principle help during engine start up.
Title: Re: Rocket Engine Q&A
Post by: QuantumG on 05/23/2015 11:53 pm
Hello can anyone help me figure out how a Turbo-Pump works. I know some of the basics of how it works but I trying to figure out how it uses Bernoulli's principle and How does this Principle help during engine start up.

Welcome to the forum! You may find this helpful:

http://www.engineeringtoolbox.com/centrifugal-pumps-d_54.html

A turbopump is just a centrifugal pump driven by turbines. Design of the turbines is considered the more challenging part over the design of the pump.
Title: Re: Rocket Engine Q&A
Post by: Mini_Elon on 05/24/2015 02:40 am
Just what I was looking for
Title: Re: Rocket Engine Q&A
Post by: a_langwich on 06/11/2015 08:01 pm
I'm asking because I wonder whether one can expect better isp from bigger engines, all other things equal.


Close to what I wanted to ask...

Is a 450s Isp engine LH2/LOX engine possible at 1 k lbf ?  what about 100 lbf?  what about 10 lbf?

In-space only.  We've got the LH2/LOX sitting around, so that's a fixed part of the problem.

Would piston-pumps be more efficient at this thrust range?  Could you push the O/F mix closer to stoichiometric?
Title: Re: Rocket Engine Q&A
Post by: msat on 06/14/2015 06:12 pm
Hello can anyone help me figure out how a Turbo-Pump works. I know some of the basics of how it works but I trying to figure out how it uses Bernoulli's principle and How does this Principle help during engine start up.

Welcome to the forum! You may find this helpful:

http://www.engineeringtoolbox.com/centrifugal-pumps-d_54.html

A turbopump is just a centrifugal pump driven by turbines. Design of the turbines is considered the more challenging part over the design of the pump.

I vaguely remember seeing a cutaway image of a Russian unit, and the pump was a screw type rather than centrifugal. I don't know how common such a pump was, or if the drawing was even accurate.
Title: Re: Rocket Engine Q&A
Post by: R7 on 06/16/2015 10:08 am
I vaguely remember seeing a cutaway image of a Russian unit, and the pump was a screw type rather than centrifugal. I don't know how common such a pump was, or if the drawing was even accurate.

High performance turbopumps usually come with centrifugal impellers and 'screws' called inducers. Inducer's job is to raise propellant pressure gently so that the flow won't cavitate when it hits the inducer blade nor when it hits the impeller vanes after the inducer.

In really high pressure engines there are also separate boost pumps which raise the propellant pressure to intermediate level before entering the main turbopump. They can have inducer-like screws too.

lpre.de has good imagery and google translate is your friend.

http://lpre.de/energomash/RD-170/index.htm
http://lpre.de/sntk/NK-33/index.htm
Title: Re: Rocket Engine Q&A
Post by: msat on 06/16/2015 02:39 pm
I vaguely remember seeing a cutaway image of a Russian unit, and the pump was a screw type rather than centrifugal. I don't know how common such a pump was, or if the drawing was even accurate.

High performance turbopumps usually come with centrifugal impellers and 'screws' called inducers. Inducer's job is to raise propellant pressure gently so that the flow won't cavitate when it hits the inducer blade nor when it hits the impeller vanes after the inducer.

In really high pressure engines there are also separate boost pumps which raise the propellant pressure to intermediate level before entering the main turbopump. They can have inducer-like screws too.

lpre.de has good imagery and google translate is your friend.

http://lpre.de/energomash/RD-170/index.htm
http://lpre.de/sntk/NK-33/index.htm

Ah, thanks! Excellent content in those links. I was familiar with inducers being used on atmospheric air centrifugal compressors, but of course they look quite a bit different. I guess I never took a real close look at actual units before. Is this style of inducer also common on american turbopumps?
Title: Re: Rocket Engine Q&A
Post by: R7 on 06/17/2015 05:22 am
Is this style of inducer also common on american turbopumps?

Yup, it's pretty much a must if you want light weight TP achieving very high pressures.

This page has several nice cutaway images plus the related science for light bedtime reading ;) ; http://www.k-makris.gr/RocketTechnology/TurboPumps/turbopumps_for_liquid_rocket_eng.htm



Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 06/25/2015 10:30 pm
I think I've confused myself on the definition (or how to calculate) Kn (the ratio of surface burning area to nozzle throat cross-sectional area) in solid rocket engines. I was hoping someone could help me and clarify if I am calculating Kn correctly.

The engine is a monolithic, hollow-cylindrical grain engine, and its grain is 2 inches in diameter and 12 inches long. The hollow-cylindrical bored grain is 1 inch in diameter. The nozzle throat diameter is .5 inches. I also realize these specs may not be realistic, but for purposes of calculating, I think they'll work.

So the minimum Kn would be when the engine is first ignited, so the surface area of the exposed grain at ignition would be:

pi * (diameter of bore) * (height of the grain) = pi * 1 inch * 12 inches = 37.6991118 inches squared of burning propellant

... divided by the nozzle's cross-sectional area which would be pi * (nozzle's radius squared) = pi * .25 inches * .25 inches = .19634954 inches squared

so then the minimum Kn would (surface area of propellant at ignition) / (nozzle cross-sectional area) = 37.6991118 inches squared / .19634954 inches squared = 192.

I'm pretty sure that is correct. I did not include the bottom of the "cylinder" of the grain, because from what I understand, a rocket engine can be made where the bottom of the grain is not an exposed burning surface. I am assuming the maximum Kn would be the outer most surface area of the propellant (assuming the grain is burning from the inside out) divided by the same nozzle cross-sectional area? In addition, in large segmented rockets that are constructed, do all the "cylinders" of grains burn inside out? Because they are segmented and the grains are inserted into the casing (does this happen?), do the grains also burn on the top and bottom of their "cylinder" because they are not tightly packed? Do segmented grains sometimes burn on their top, bottom, inside and outside of their "cylinder"? If someone could please clarify anything here for me, I would greatly appreciate it!

Clayton
Title: Re: Rocket Engine Q&A
Post by: msat on 07/12/2015 06:51 pm
Could water be used as the coolant and turbine working fluid in an expander-bleed cycle engine?

It has pretty high heat capacity, is very dense, and is commonly used as the working fluid in power generation. Is its relatively high boiling point (relative to H2/methane/propane) problematic? What are the factors preventing it from being able to be used in such an engine cycle?
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/12/2015 08:38 pm
Weight.  It is non propulsive.  Also, how would it be cooled
Title: Re: Rocket Engine Q&A
Post by: msat on 07/12/2015 09:49 pm
Weight.  It is non propulsive.  Also, how would it be cooled

Hmm?

I was asking about it used in an expander-bleed engine like the LE-5A/B, in which case the cooling fluid is dumped after passing through the turbine so it being non-propulsive isn't an issue.

As for weight, if water has double the density of another fluid, but the same heat capacity, wouldn't it need only half the flow rate for the same rate of cooling?
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/12/2015 10:08 pm
[

I was asking about it used in an expander-bleed engine like the LE-5A/B, in which case the cooling fluid is dumped after passing through the turbine so it being non-propulsive isn't an issue.

As for weight, if water has double the density of another fluid, but the same heat capacity, wouldn't it need only half the flow rate for the same rate of cooling?

It is an issue, it is an expended mass that doesn't contribute propulsively and hence, would reduce ISP.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/12/2015 10:31 pm


I was asking about it used in an expander-bleed engine like the LE-5A/B, in which case the cooling fluid is dumped after passing through the turbine so it being non-propulsive isn't an issue.

As for weight, if water has double the density of another fluid, but the same heat capacity, wouldn't it need only half the flow rate for the same rate of cooling?

It is an issue, it is an expended mass that doesn't contribute propulsively and hence, would reduce ISP.

I wasn't asking about its affect on ISP, though. The most common gas generator cycle also loses ISP by having a portion of the propellants bypass the CC.  That doesn't preclude them and existing expander-bleed engines from being very useful engines.

What I want to know is if water would work reasonably well in the method I described, and not if it gets a crappier ISP than an H2/LOX full-flow whatever.
Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 07/13/2015 02:22 am
I wasn't asking about its affect on ISP, though. The most common gas generator cycle also loses ISP by having a portion of the propellants bypass the CC.  That doesn't preclude them and existing expander-bleed engines from being very useful engines.

What I want to know is if water would work reasonably well in the method I described, and not if it gets a crappier ISP than an H2/LOX full-flow whatever.

It would not work reasonably well.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/13/2015 03:15 am
I wasn't asking about its affect on ISP, though. The most common gas generator cycle also loses ISP by having a portion of the propellants bypass the CC.  That doesn't preclude them and existing expander-bleed engines from being very useful engines.

What I want to know is if water would work reasonably well in the method I described, and not if it gets a crappier ISP than an H2/LOX full-flow whatever.

It would not work reasonably well.

Could you please elaborate a bit?

The RL10 had worked with methane, which has a lower heat capacity than water. I know the RL10 is just an expander cycle, but that means methane would also work in an expander-bleed engine. So what is it about water that would prevent it from working?
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/13/2015 03:50 am
I wasn't asking about its affect on ISP, though. The most common gas generator cycle also loses ISP by having a portion of the propellants bypass the CC.  That doesn't preclude them and existing expander-bleed engines from being very useful engines.

What I want to know is if water would work reasonably well in the method I described, and not if it gets a crappier ISP than an H2/LOX full-flow whatever.

It would not work reasonably well.

Could you please elaborate a bit?

The RL10 had worked with methane, which has a lower heat capacity than water. I know the RL10 is just an expander cycle, but that means methane would also work in an expander-bleed engine. So what is it about water that would prevent it from working?

So, first off, you could absolutely design a working expander bleed engine using water as the working fluid.

However, there's a number of reasons not to do this, and not any truly compelling reasons to do so.

Water has only about a quarter of the heat capacity as hydrogen. It needs to be at high pressure, so you need a separate pump. It also needs dedicated tankage. Much additional complexity and weight for no obvious benefit.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/13/2015 01:55 pm


So, first off, you could absolutely design a working expander bleed engine using water as the working fluid.

However, there's a number of reasons not to do this, and not any truly compelling reasons to do so.

Water has only about a quarter of the heat capacity as hydrogen. It needs to be at high pressure, so you need a separate pump. It also needs dedicated tankage. Much additional complexity and weight for no obvious benefit.

Thanks, strangequark.

I know hydrogen is the "mother of all" coolants, or even fuels for that matter, but not all rockets use it. However, while water might have 1/4 the heat capacity of H2, it has more than 10x the density. To make this easier to visualize, lets say we stored the hydrogen used as coolant in a separate tank from the H2 used for propellant; if I'm not mistaking, replacing the H2 with H20 would reduce the coolant tank volume by 60%. This may or may not be negligible depending on the coolant flow rate in an expander-bleed cycle (numbers for this cycle have evaded me).

As you say, it certainly adds complexity, but all engineering is about trade-offs. Is the added complexity a worthy trade-off? I don't know. For an ELV, I'd definitely say "no", but for an RLV I can't be as certain.

Thanks for your thoughts.
Title: Re: Rocket Engine Q&A
Post by: R7 on 07/13/2015 02:41 pm
I know hydrogen is the "mother of all" coolants, or even fuels for that matter, but not all rockets use it. However, while water might have 1/4 the heat capacity of H2, it has more than 10x the density. To make this easier to visualize, lets say we stored the hydrogen used as coolant in a separate tank from the H2 used for propellant; if I'm not mistaking, replacing the H2 with H20 would reduce the coolant tank volume by 60%.

Too simple analysis, does not account for the usable temperature range of the coolant which also dictates how much you need it and can extract work from it. With liquid hydrogen you start at 20K and work your way up to the upper temperature limit of coolant channel alloy. Liquid water starts at 273K, same upper limit. Also heat of vaporization matters to certain extent.

As you say, it certainly adds complexity, but all engineering is about trade-offs. Is the added complexity a worthy trade-off? I don't know. For an ELV, I'd definitely say "no", but for an RLV I can't be as certain.

Added complexity + poorer performance coolant = added mass --> it makes even less sense in RLVs which have more severe mass-ratio requirements than ELVs.
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/13/2015 02:51 pm
I wasn't asking about its affect on ISP, though. The most common gas generator cycle also loses ISP by having a portion of the propellants bypass the CC.  That doesn't preclude them and existing expander-bleed engines from being very useful engines.

What I want to know is if water would work reasonably well in the method I described, and not if it gets a crappier ISP than an H2/LOX full-flow whatever.

Getting crappier ISP with no additional benefit, it is not "working reasonably well"

Just saw the above post:  Added complexity + poorer performance coolant = added mass is not  "working reasonably well"

Also, common gas generator cycle engine still produce some thrust with turbo pump exhaust, which also can be used for cooling nozzle components.  Additionally, they don't require additional tankage.
Title: Re: Rocket Engine Q&A
Post by: strangequark on 07/13/2015 03:31 pm


So, first off, you could absolutely design a working expander bleed engine using water as the working fluid.

However, there's a number of reasons not to do this, and not any truly compelling reasons to do so.

Water has only about a quarter of the heat capacity as hydrogen. It needs to be at high pressure, so you need a separate pump. It also needs dedicated tankage. Much additional complexity and weight for no obvious benefit.

Thanks, strangequark.

I know hydrogen is the "mother of all" coolants, or even fuels for that matter, but not all rockets use it. However, while water might have 1/4 the heat capacity of H2, it has more than 10x the density. To make this easier to visualize, lets say we stored the hydrogen used as coolant in a separate tank from the H2 used for propellant; if I'm not mistaking, replacing the H2 with H20 would reduce the coolant tank volume by 60%. This may or may not be negligible depending on the coolant flow rate in an expander-bleed cycle (numbers for this cycle have evaded me).

As you say, it certainly adds complexity, but all engineering is about trade-offs. Is the added complexity a worthy trade-off? I don't know. For an ELV, I'd definitely say "no", but for an RLV I can't be as certain.

Thanks for your thoughts.

So, I kind of already said this, but to trade there must be competing benefits. For this, you have a lot of cons, and no pros. There are no trades. Not saying you couldn't do it, from a physics perspective, just shouldn't. What do you think you would gain?
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/13/2015 04:39 pm
However, there's a number of reasons not to do this, and not any truly compelling reasons to do so.

Water has only about a quarter of the heat capacity as hydrogen. It needs to be at high pressure, so you need a separate pump. It also needs dedicated tankage. Much additional complexity and weight for no obvious benefit.

The original poster may be interested in Balepin's somewhat related Third Fluid Coolant concept, which applies this idea to staged combustion engines with several advantages.

But I'm wondering if this might be interesting for amateur / semi-professional / New Space groups like Masten, Armadillo or Copenhagen Suborbitals if they want an expander-like system with noncryogenic hydrocarbon / alcohol fuels.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/13/2015 05:43 pm


Too simple analysis, does not account for the usable temperature range of the coolant which also dictates how much you need it and can extract work from it. With liquid hydrogen you start at 20K and work your way up to the upper temperature limit of coolant channel alloy. Liquid water starts at 273K, same upper limit. Also heat of vaporization matters to certain extent.



Thanks, R7. These are some of the things I'm trying to understand. I was thinking that since H2 has a much lower critical temperature than water, that would possibly offset the higher initial temperature of the water. I was also working under the assumption that it is desirable to avoid the coolant flow being in a supercritical state, but it seems that that may not be true for the cryo fuels. Is that true?



So, I kind of already said this, but to trade there must be competing benefits. For this, you have a lot of cons, and no pros. There are no trades. Not saying you couldn't do it, from a physics perspective, just shouldn't. What do you think you would gain?

I was just trying to get a better understanding of why it would be a bad idea. I tend to prefer more in-depth answers than someone simply telling me "It's a bad idea". Hope I can't be faulted for that  :)


The possible pros I had in mind were A) potentially [slightly?] smaller LV, and B) reduction in propellant costs (only a concern for heavily reusable LVs)

One could easily make the argument that the Falcon9 is overly complicated, with the first stage alone having 9 CCs, turbopumps, and all the necessary support hardware, yet despite all those drawbacks, it's a cost-effective platform relative to its competitors. So, complexity on it's own isn't necessarily a deal breaker.

I'd actually like to avoid H2 in this discussion, as I'm interested in some of the alkanes (methane, propane).

Ok.... Finding some more information. Around 100K, methane and propane has roughly similar density to water. Heat capacity for propane is about 100 J/mol K, water is 75, and methane 50. Propane's critical temp is also almost 200C higher than methane, with water being almost 300C higher than propane.

So I guess right off the bat propane is a better coolant than water, and perhaps even more so if it can operate in a supercritical state. If it can't, then water has a tiny bit more headroom if I'm not mistaking.

I think I can see where this idea starts to fall apart. Any thoughts on this besides "I told you so"?
Title: Re: Rocket Engine Q&A
Post by: msat on 07/13/2015 06:18 pm


The original poster may be interested in Balepin's somewhat related Third Fluid Coolant concept, which applies this idea to staged combustion engines with several advantages.

But I'm wondering if this might be interesting for amateur / semi-professional / New Space groups like Masten, Armadillo or Copenhagen Suborbitals if they want an expander-like system with noncryogenic hydrocarbon / alcohol fuels.

Thanks for the heads up about the "Third Fluid Coolant" work. I had not heard about it.

As for who might be interested in it, I couldn't say. I'd imagine anyone who could handle LOX could also handle chilled propane. If they go with a non-cryo oxidizer such as H2O2 or N2O, then it's probably just easier to decompose those to drive the turbine and let the fuels or maybe H2O2 do the cooling. On the upside, water is cheaper and easier to get than either of those oxidizers (especially high-test peroxide), so maybe there is value to it. Though I'd say the added weight of an additional pump may be more of a concern on a smaller rocket.
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/13/2015 06:23 pm

The possible pros I had in mind were A) potentially [slightly?] smaller LV, and B) reduction in propellant costs (only a concern for heavily reusable LVs)


Neither are true.  Adding another tank would go against reducing main tank size. 
Saving a few hundreds of gallons of prop really isn't much cost in light of the additional maintenance for the water system
Title: Re: Rocket Engine Q&A
Post by: R7 on 07/13/2015 07:01 pm
I was also working under the assumption that it is desirable to avoid the coolant flow being in a supercritical state

It is the opposite, smooth predictable and stable changes in supercritical region are preferred over distinct erratic phase change in the coolant channels.

http://forum.nasaspaceflight.com/index.php?topic=30910.msg1006206#msg1006206
Title: Re: Rocket Engine Q&A
Post by: msat on 07/14/2015 12:52 am

The possible pros I had in mind were A) potentially [slightly?] smaller LV, and B) reduction in propellant costs (only a concern for heavily reusable LVs)


Neither are true.  Adding another tank would go against reducing main tank size. 
Saving a few hundreds of gallons of prop really isn't much cost in light of the additional maintenance for the water system

How much space does a common bulkhead take up? Maybe a few inches depending on the amount of temperature isolation? You may have a point regarding maintenance, but that largely depends on the design and performance of the hardware.



I was also working under the assumption that it is desirable to avoid the coolant flow being in a supercritical state

It is the opposite, smooth predictable and stable changes in supercritical region are preferred over distinct erratic phase change in the coolant channels.

http://forum.nasaspaceflight.com/index.php?topic=30910.msg1006206#msg1006206

I get that. I just thought that the coolant was operating below its critical temperature, but apparently that's often(?) not the case. Actually, what commonly used fuels is that the case with? Just hydrogen? Or also kerosene, alcohol, and the hypergolics?
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/14/2015 02:14 am
How much space does a common bulkhead take up? ?

Not feasible with a cryogenic propellants.  The insulation requirement negate any benefits.  It wouldn't be done even for storables.  Would cause too much issues with pressure management and propellant outlets
Title: Re: Rocket Engine Q&A
Post by: msat on 07/14/2015 02:47 am
How much space does a common bulkhead take up? ?

Not feasible with a cryogenic propellants.  The insulation requirement negate any benefits.  It wouldn't be done even for storables.  Would cause too much issues with pressure management and propellant outlets

Ah. I completely neglected the plumbing. Though non-standard rocket layouts could potentially alleviate some of issues.
Title: Re: Rocket Engine Q&A
Post by: R7 on 07/14/2015 07:29 am
I just thought that the coolant was operating below its critical temperature, but apparently that's often(?) not the case. Actually, what commonly used fuels is that the case with? Just hydrogen? Or also kerosene, alcohol, and the hypergolics?

Hydrogen critical point is just 33K / 13bar so it effectively always becomes supercritical in coolant channels.

RP-1 is 662K / 22bar (source (http://www.dlr.de/Portaldata/55/Resources/dokumente/sart/0095-0212prop.pdf)). The temperature approaches unsafe coking temperatures so can't say for sure without actual RP-1 engine flow diagram showing pressures and temperatures. I recall seeing such diagram for RD-180 or something similar, can't find it right now. Anyone?

Alcohols ... nobody uses them anymore (in rocketry ;) ) except a few hobbyists. Copenhagen Suborbitals works with isopropyl alcohol, critical point at 509K / 54bar. Temperature might go above with pure IPA (some tests have run with 25% water mixture) but pressure is far below supercritical.

Hydrazine 653K / 147bar. May be a close call as with RP-1 but without the coking issue. Hard to say without proper large hypergolic engine fluid diagram. A kingdom for RD-253/-275 specs!
Title: Re: Rocket Engine Q&A
Post by: msat on 07/14/2015 06:57 pm
Thanks for the info, R7. Besides H2 then, we can't be sure if the other common (present or past) coolants are above their critical points.




The original poster may be interested in Balepin's somewhat related Third Fluid Coolant concept, which applies this idea to staged combustion engines with several advantages.

But I'm wondering if this might be interesting for amateur / semi-professional / New Space groups like Masten, Armadillo or Copenhagen Suborbitals if they want an expander-like system with noncryogenic hydrocarbon / alcohol fuels.

I didn't really understand the Third Fluid Coolant system because the main paper is behind an AIAA paywall, and what I initially saw was some three propellant systems. However, besides finding a few posts discussing it on this forum (what's new to me, is old to the forum  ;) ) I found some interesting details and simpler diagrams in:

http://ftp.rta.nato.int/public/PubFullText/RTO/EN/RTO-EN-AVT-150/EN-AVT-150-02.pdf

Most of the focus is on TFC and MIPCC (which I find particularly interesting), which is not surprising because he was quite involved in both. I really like the information on the testing rig for the MIPCC. Very cool!

For the TFC, the thing that stuck out most to me is his choice of coolants. From page 17:

"In the TFC engine, the nozzle and combustor assembly 4 is cooled by a circulating coolant such as water,
methanol, ethanol, or liquid having equivalent properties, or mixtures thereof. "

Propane would seem to make a better coolant, and while static pressures in the system would be higher than the coolants listed above, it might not be a significant issue compared to the pressures the system would be exposed to when operating. Perhaps propane couldn't condense when it expands in the turbine section like water or the other coolants listed would, requiring a larger heat exchanger?

In that paper, Balepin talks about H2 being used as the fuel in his examples, but I'm not sure if there's really any benefit of using a third fluid for coolant if the engine is using cryo fuel. For example, the Closed Split Expander basically solves the same issues with the standard expander cycle, but doesn't require a heat exchanger

(http://blogs.nasa.gov/J2X/wp-content/uploads/sites/212/2014/03/closedsplitexpander.jpg)
https://blogs.nasa.gov/J2X/2014/03/24/inside-the-leo-doghouse-the-art-of-expander-cycle-engines/

But as you note, it may very well be of interest to those that would prefer to have an expander cycle work with non-cryo fuels. Though as I stated before, if they can work with LOX, chilled propane should be within reach as well.

I really like the TFC concept, but I'm having a hard time seeing where it would be practical outside of RP1.


Edit:
I am so behind the times! Apparently there was already such a discussion here. Sorry  :-[
https://forum.nasaspaceflight.com/index.php?topic=23623.0
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/14/2015 08:58 pm
The reason I thought a steam turbine could be interesting to Masten / Copenhagen Suborbitals / Armadillo is that it might be easier to develop than a gas generator engine.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/14/2015 11:08 pm
For the small rockets Armadillo and Masten have flown, I don't think turbomachinery would be an option. Though there may be a particular size of rocket where COTS turbocharger turbines could possibly be used, but designing the impellers might be tricky. I suppose that could be contracted out to an experienced company. The heat exchanger could "simply" be a coil running through the bottom of the fuel or oxidizer tank (have to make sure the coolant can't possibly freeze - another point in favor of propane). 

One thing I've been thinking about is the upper limit of thrust an expander cycle engine could generate without making the combustion chamber really long. I wonder if an aerospike nozzle would be a better candidate as the nozzle surface is supposedly exposed to a higher heat flux than a bell (though I'd imagine the total surface area is actually smaller). I don't actually understand it, but that's what all the pros say, so who am I to argue. Anyway, the aerospike may provide the much needed heat to drive a more powerful expander cycle pump. Add in the benefit of reasonably good altitude compensation, and it seems like a win. 
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/14/2015 11:26 pm
For the small rockets Armadillo and Masten have flown, I don't think turbomachinery would be an option.

So far they've stuck with pressure-fed systems, but I suspect they'd love to move beyond that. Steam turbines are an old technology and probably easier than gas generator systems, although CS is moving straight to a gas generator system. Electric pumps could also be an option.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/15/2015 12:06 am
A steam turbine shouldn't inherently be different than any other kind of turbine for a given size and pressure ratio. The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.   
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 07/15/2015 07:23 am
The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.

Yeah, that's what I meant. You also don't need to design a gas generator. Not show stoppers, I just imagine it make things easier while you're still learning how to do it.
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 07/15/2015 08:39 am
A steam turbine shouldn't inherently be different than any other kind of turbine for a given size and pressure ratio. The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.
You're right. In fact the first steam turbines were running in the 1900's, the first gas turbines in the 1930's.

IIRC  the team under Eugene Sanger in Germany in WWII planned a steam turbine drive for the turbo pumps for the "Silver Bird" concept. I'm not sure if they actually built the whole engine (they seem to have managed a 100 tonne combustion chamber) at least.

There is a translation of their report done by IIRC the US Navy, complete with (very bad) photographs.

Note that the water used on steam plants is not straight from the tap water. For this application you'd probably want to go with deionized water with no scale forming impurities in it running in a sealed system.

Note that the work of the Whitehead group at LLNL indicates turbines scale down badly below about 5000 lb thrust. The turbines are small, which means the clearances are very tight and the surface finish has to be very good, as the boundary layer will be proportionately thinner.

OTOH a positive displacement reciprocating  pump using say corrugated pistons would be easier to fabricate, eliminate rotating seals and bearings and deliver complete separation of propellant from driving fluid. Also the pump drive fluid is relatively inert (super heated steam should be treated with considerable respect.  :( ).

The problem with all separate pump driver fluid concepts is now you need two heat exchangers, the CC wall and whatever you're dumping the waste head the was not used in the pump drives. This might not be the problem people think it is if the engine stays in the atmosphere. Different goals, different trades.

In vacuum you're looking at either radiators (which get big at low temperature differences, even with a carbon fibre "carpet" around them to enhance emission) or dumping it to the propellant flow IE a 2nd HX.

Historically people have said "No, weight penalty not worth the flexibility of being able to select optimum drive fluid properties"

OTOH amateur developers, or people wanting lower maintenance reusability might prioritize things differently.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/15/2015 06:54 pm

Yeah, that's what I meant. You also don't need to design a gas generator. Not show stoppers, I just imagine it make things easier while you're still learning how to do it.

Possibly. I couldn't say with any kind of certainty. I will however counter Jim's assertion that TCF is too complicated. Perhaps more complex than a "standard" expander like the RL10, but not more so than the gas generator cycle and the variations thereof.




Note that the water used on steam plants is not straight from the tap water. For this application you'd probably want to go with deionized water with no scale forming impurities in it running in a sealed system.

Note that the work of the Whitehead group at LLNL indicates turbines scale down badly below about 5000 lb thrust. The turbines are small, which means the clearances are very tight and the surface finish has to be very good, as the boundary layer will be proportionately thinner.

OTOH a positive displacement reciprocating  pump using say corrugated pistons would be easier to fabricate, eliminate rotating seals and bearings and deliver complete separation of propellant from driving fluid. Also the pump drive fluid is relatively inert (super heated steam should be treated with considerable respect.  :( ).

The problem with all separate pump driver fluid concepts is now you need two heat exchangers, the CC wall and whatever you're dumping the waste head the was not used in the pump drives. This might not be the problem people think it is if the engine stays in the atmosphere. Different goals, different trades.

In vacuum you're looking at either radiators (which get big at low temperature differences, even with a carbon fibre "carpet" around them to enhance emission) or dumping it to the propellant flow IE a 2nd HX.

Historically people have said "No, weight penalty not worth the flexibility of being able to select optimum drive fluid properties"

OTOH amateur developers, or people wanting lower maintenance reusability might prioritize things differently.


Of course you wouldn't fill up the rocket with water straight from the river! Distilled, and maybe deionized for sure. You don't want crud buildup or all the various other things that could happen to occur. Eliminating impurities in a third coolant fluid is no less important than doing it for your propellant.

I know that jet engines become less and less efficient as they're scaled down for numerous reasons. The thing that surprises me is the amount of power rocket engine turbines generate relative to their size. Any inefficiencies may be acceptable in light of the other options. But when it comes to rockets, there seems to be a point where active pumps just aren't worthwhile, and pressure fed is more suitable. Out of curiosity, I looked up some turbine efficiency figures which are annoyingly hard to come across: for turbochargers, I found that one of Garrett's largest units has a turbine efficiency of 82%, while GE's GE90 turbofan is 93%. I couldn't find anything about the F1, but if I'm not mistaking, according to http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850010710.pdf the RL10 turbine has an efficiency of 95%. I wouldn't rule out turbocharger turbines for the next step up from pressure-fed designs in amateur rocketry.

Positive displacement pumps are an option, as Xcor has proven, but they pose some of their own challenges. Pumping isn't steady state, even with multiple pistons. If you're pumping gasses, this isn't much of an issue as you can feed it to a pressure vessel and maintain output with a pressure regulator (at the expense of reduced output pressure). You can do something similar with liquids I suppose, but it's a bit trickier. I'm not entirely convinced that the challenges and potential mass penalties outweigh the use of turbopumps. Even if it's easier to fabricate, it's still more complex.

In the case of the Third Fluid Coolant concept, the idea is to dump the heat into the propellants rather than the atmosphere or space. It's certainly the most compact option. Plus it's desirable to pre-heat the propellants in certain cases.

I agree that maximal efficiency isn't always the be-all-end-all of rocketry, this is especially true of teams with limited experience and budgets. There's so many possible solutions to a given problem, with some of those solutions being better for a team than others. I do appreciate how some very clever solutions are born out of necessity due to limited budgets. 
Title: Re: Rocket Engine Q&A
Post by: msat on 07/16/2015 07:27 pm
I had a thought about a variation of the TFC concept. Well, it's not actually TFC, but it is an expander cycle and uses heat exchangers like one. The fuel should be suitable for an expander cycle and probably subcooled (such as propane at LOX temps). The idea is that the fuel exiting the compressor is split off with one path going to the combustion chamber and the other going to the cooling channels. After the cooling channels, it is expanded and drives the turbine and then routed through a heat exchanger (possibly integrated as part of the injector head?) where it's condensed and then dumped back somewhere upstream of the compressor inlet.

Pros:
1) No need for an additional compressor as a standard TFC requires
2) Higher pressure drop across the turbine like a TFC than possible with a standard expander cycle

Cons:
1) More limited in the fuels that can be used in comparison to a TFC. No H2 either, probably
2) There will be some pressure losses for the propellants entering the combustion chamber as they have to pass through a heat exchanger like a TFC.


It would seem that the most important factor in making the TFC or variants practical is the efficiency and weight of the necessary heat exchangers. I tried finding figures that might give an indication of what's possible (even looking for data on power plants) but came up empty handed. 
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/16/2015 07:32 pm
I still have to ask why?   Your list of cons is missing many items.
Title: Re: Rocket Engine Q&A
Post by: msat on 07/16/2015 08:36 pm
I still have to ask why?   Your list of cons is missing many items.

Then help populate it.

Why? Why do anything? Why make rocket engines or anything at all? If there was a single type of engine that is better than every other one in every way, then why are there so many different ones with various different cycles currently in use? It would seem that by your logic, there's only one metric of importance.

I didn't think I needed to state some of the more obvious pros, but I will:

0) No secondary combustion chambers
1) Closed cycle
2) Higher chamber pressure than a standard expander
3) Lower temps and higher pressure drops across turbines, and possibly higher chamber pressures compared to the various staged combustion cycles (not to mention less corrosive than the hot oxidizer rich ones) while being no more complex, and likely simpler.

Those sound like some decent pros to me.
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/16/2015 09:14 pm
1.  If there was a single type of engine that is better than every other one in every way, then why are there so many different ones with various different cycles currently in use? It would seem that by your logic, there's only one metric of importance.

I didn't think I needed to state some of the more obvious pros, but I will:

0) No secondary combustion chambers
1) Closed cycle
2) Higher chamber pressure than a standard expander
3) Lower temps and higher pressure drops across turbines, and possibly higher chamber pressures compared to the various staged combustion cycles (not to mention less corrosive than the hot oxidizer rich ones) while being no more complex, and likely simpler.

Those sound like some decent pros to me.

1. All the goods ones are in use or have been used.  This one is on sideline because it isn't good.
0) why is that good?

The complexity, extra mass, and losses in the heat exchangers are greater drawbacks.  Not to mention the inability to start such an engine. 
Title: Re: Rocket Engine Q&A
Post by: msat on 07/16/2015 10:44 pm

1. All the goods ones are in use or have been used.  This one is on sideline because it isn't good.
0) why is that good?

The complexity, extra mass, and losses in the heat exchangers are greater drawbacks.  Not to mention the inability to start such an engine.

It may not be good, and perhaps I'm overlooking something, but saying "all the good ones have been used" is a poor argument and akin to saying everything that could possibly be invented already has. What happened the last time someone made that very claim?

Complexity compared to what? Are you saying staged combustion is less complex? If we want minimal complexity, then every rocket would be pressure fed. After all, pumps are complex. Extra mass compared to what? Pressure fed rockets, GG combustion chambers? I'll agree that there's no free lunch. When it comes to rocket motors, if you remove mass by eliminating one component, chances are you'll have to add mass by adding another. The biggest drawback is definitely the heat exchanger, and I've admitted as much. I also don't know what is to be expected in terms of mass and efficiency. If you have any insight, then please share it. Otherwise how can you claim it's a deal-breaker outright? Every method will have losses somewhere, that's why there's no single engine that rules them all in every way. You have pressure losses prior to [some of] the propellants entering the combustion chamber even in the mighty staged combustion engines, because they're pushed through the cooling system prior to entering the gas generator and turbine (so even more losses!).

It wouldn't be any good if it couldn't be started, right? This isn't significantly different than starting many other types of rockets. The RL10 starts just fine. Vacuum starting would be the most trivial. One option is to have an external pressurant and associated valving ahead of the turbine, as well as a valve behind the turbine that vents to the atmosphere. This pressure difference gets the turbopump started until it's self-sustaining. It's not all that different from many other engines.

From the sound of it, you are involved in the field, but I have no idea what your specialty is. In that light, whose thoughts on the matter of using heat exchangers on a rocket should I value more: yours, or Vladimir Balepin's - who has actual and verifiable professional experience on the subject?

In fact, there has at least been some testing done (and please don't say that just because we haven't heard anything since means it was a failure. Many workable ideas don't get pursued for various reasons as you know)

http://aerospace.xcor.com/rocket-engines/research-development/

Quote
In early 2006, XCOR and ATK GASL received a DARPA contract to "investigate, develop, and demonstrate a novel configuration for a liquid rocket engine, namely a Third Fluid Cooled (TFC) liquid rocket engine," for which we used the Tea Cart to generate superheated steam suitable for driving a Rankine cycle. This steam cycle will allow a turbopump system to develop high chamber pressure in a more efficient way than staged combustion cycles, thus improving the performance and durability for boost propulsion, as well as orbital transfer applications. 
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/16/2015 11:45 pm
DARPA contracts are not relevant indicators for viability of technology.  More like throwing stuff at a wall and see what sticks.
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/16/2015 11:50 pm
The RL10 starts just fine.

It just needs valves opened to start. 

Existing engines don't have a closed system driving the turbo pump so back pressure is not a concern
Title: Re: Rocket Engine Q&A
Post by: msat on 07/17/2015 02:42 am
DARPA contracts are not relevant indicators for viability of technology.  More like throwing stuff at a wall and see what sticks.

Fairly true, but generally you need some preliminary research before you get funding. I'd imagine many of these questions have been addressed to some degree in Balepin's AIAA paper written on ATK's dime.


The RL10 starts just fine.

It just needs valves opened to start. 

Essentially the same for this

Quote
Existing engines don't have a closed system driving the turbo pump so back pressure is not a concern

Keep in mind, I'm not talking about TFC. The coolant system isn't a closed loop. Instead, once the coolant (which is the same as the fuel) passes through the heat exchanger, it is dumped back into the fuel supply upstream of the compressor.

Pretty much all rocket engines besides pressure-fed/open cycle GGs with ablative thrust chambers & nozzles have some pressure drops relative to the pump output at the very least due to losses from the cooling channels. In the case of staged combustion, you have similar issues as an expander cycle where you have to provide a suitable pressure drop across the turbine to generate sufficient power for the pumps, further lowering the pressure to the thrust chamber relative to pump output. To avoid these pressure drops so to generate more thrust (at the expense of ISP), Mitsubishi built the LE-5A/B as an expander bleed cycle ensuring the engine saw maximum pump pressure. Can't argue that there's no logic behind it.
Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 08/06/2015 12:13 am
So I have questions regarding erosive burning that solid rocket engines experience. This next explanation is from braeunig.us (http://www.braeunig.us/space/propuls.htm).

Quote
For most propellants, certain levels of local combustion gas velocity (or mass flux) flowing parallel to the burning surface leads to an increased burning rate. This "augmentation" of burn rate is referred to as erosive burning, with the extent varying with propellant type and chamber pressure. For many propellants, a threshold flow velocity exists. Below this flow level, either no augmentation occurs, or a decrease in burn rate is experienced (negative erosive burning).
The effects of erosive burning can be minimized by designing the motor with a sufficiently large port-to-throat area ratio (Aport/At). The port area is the cross-section area of the flow channel in a motor. For a hollow-cylindrical grain, this is the cross-section area of the core. As a rule of thumb, the ratio should be a minimum of 2 for a grain L/D ratio of 6. A greater Aport/At ratio should be used for grains with larger L/D ratios.

So if I understand correctly, if we're using a cylindrical core burning solid rocket engine with a core diameter of 3 inches and an outer grain diameter of 6 inches with a throat diameter of 2 inches, the port-to-throat area ration would be:

3 inches / 2 = 1.5 inch radius

1.5 inches * 1.5 inches * pi = 7.07 inches squared for the port area. And then the throat area would be:

2 inches / 2 = 1 inch radius

1 inch * 1 inch * pi = 3.14 inches squared for the throat area. So then the port-to-throat ratio would be:

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right). Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches? Or does L/D ratio refer to the rocket engine length to the core diameter? So the max length of my solid rocket engine could only be around:

3 inch diameter * 6 = 18 inches? Can someone help by clarifying the approximations for what the dimensions of a solid rocket engine should be to minimize erosive burning?
Title: Re: Rocket Engine Q&A
Post by: R7 on 08/06/2015 08:46 am

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right).

All good so far but note that you can simplify the math by just squaring the ratio of diameters (or radii, does not matter which). Area is proportional to square of length, that's enough knowledge so no need to calculate the actual areas.

(3/2)2 = 2.25

Quote
Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches?

It's the grain L/D, L is grain length and D is outside diameter. I don't know how the thumb rule 2 Ap/At ratio scales with bigger L/Ds.

The combustion area in your simple cylinderical core desing would double at the end of the burn, thrust also. Eventually you want to seek out more complex port geometries for more even thrust.

Have you read Richard Nakka's pages? A lot of information for solid rocket hobbyists there.

http://www.nakka-rocketry.net/th_grain.html
Title: Re: Rocket Engine Q&A
Post by: ClaytonBirchenough on 08/06/2015 12:34 pm

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right).

All good so far but note that you can simplify the math by just squaring the ratio of diameters (or radii, does not matter which). Area is proportional to square of length, that's enough knowledge so no need to calculate the actual areas.

(3/2)2 = 2.25

Quote
Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches?

It's the grain L/D, L is grain length and D is outside diameter. I don't know how the thumb rule 2 Ap/At ratio scales with bigger L/Ds.

The combustion area in your simple cylinderical core desing would double at the end of the burn, thrust also. Eventually you want to seek out more complex port geometries for more even thrust.

Have you read Richard Nakka's pages? A lot of information for solid rocket hobbyists there.

http://www.nakka-rocketry.net/th_grain.html

Thanks for the response!

I understand everything you said, and what you explained is actually what I tried to communicate I was doing haha. In the math I did, I showed that the max length of a 6 inch diameter grain was 36 inches (again, using the 6 L/D ratio where I don't exactly know comes from). This *would be right?

I'm also aware that the grain surface area would double in my design. Not optimal, but makes for a easier fabrication. Also, it may actually be a good thing to minimize erosive burning because the mass flow rate and chamber pressure at ignition happens with a smaller port. In addition, I have read a lot of Nakka's website. He's the man.

But I do get confused on how L/D can become an approximation of the grain length? Is it maybe because they take into account chamber pressure and propellant surface area and assume your Kn (propellant surface area to throat area) is within reason? Thanks R7 for your reply!
Title: Re: Rocket Engine Q&A
Post by: ZachS09 on 11/21/2015 11:36 pm
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?
Title: Re: Rocket Engine Q&A
Post by: Newton_V on 11/21/2015 11:41 pm
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?

OSIRIS-REX and possibly all Dual Engine Centaurs.
Title: Re: Rocket Engine Q&A
Post by: ZachS09 on 11/21/2015 11:57 pm
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?

OSIRIS-REX and possibly all Dual Engine Centaurs.

Thank you.
Title: Re: Rocket Engine Q&A
Post by: Tommy OSullivan on 01/03/2016 08:37 pm
Hey, I was wondering if anyone could help me out with this project I've started working on. I want to put together a full flow methane mockup engine with a fully fleshed out interior and exterior, I have made a first rendition but I want the mark.2 to be more realistic.

Link to the full post is here http://forum.nasaspaceflight.com/index.php?topic=39200.msg1468297#msg1468297 (http://forum.nasaspaceflight.com/index.php?topic=39200.msg1468297#msg1468297) (sorry to be advertising it).

But I was just wondering if anyone could help me with how the turbopumps and preburners work internally and how they operate with one another in full flow engines like the SSME.

Thanks!
Title: Re: Rocket Engine Q&A
Post by: gin455res on 05/13/2016 06:12 am
Are there any piston pump designs based on 2-stroke geometries, specifically stepped piston designs where the crank case compressor volume is multiple  times as large as the power cylinder and massive short circuiting is promoted?

Two of these massively short circuiting engines, One running oxidizer rich , and one running fuel rich  might be equivalent to a full flow staged combustion turbo pump .
Title: Re: Rocket Engine Q&A
Post by: gin455res on 06/01/2016 10:29 pm
Have any pumps been built that can switch between combustion chambers?
(like below - is it feasible)

(http://webpidgin.co.uk/svg/trees/Eagle5-MultiChamberRaptor-Rocket2.jpg)
And what configurations might be most interesting?
below:
(http://webpidgin.co.uk/svg/trees/OtherConfigurations.jpg)
Title: Re: Rocket Engine Q&A
Post by: Jim on 06/02/2016 03:14 am
Have any pumps been built that can switch between combustion chambers?


no
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/02/2016 04:53 pm
I remember an NPO Energomash paper stating that the RD-170 could (not sure if can) shutoff individual chambers. And then you have cases like the RD-0110 that feeds the main chambers and the vernier chambers from the same TP. But generally speaking the combustion chamber and nozzle are really heavy and duplicating the mass is simply not worth it for the possible increase in isp.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 06/02/2016 06:48 pm
I remember an NPO Energomash paper stating that the RD-170 could (not sure if can) shutoff individual chambers. And then you have cases like the RD-0110 that feeds the main chambers and the vernier chambers from the same TP. But generally speaking the combustion chamber and nozzle are really heavy and duplicating the mass is simply not worth it for the possible increase in isp.

Any differences between a normal expendable booster thrust profile, and an F9 style RTLS profile.  I'm wondering if improved isp on the boost back  improves or exacerbates the chamber and nozzle mass issue?
Title: Re: Rocket Engine Q&A
Post by: baldusi on 06/03/2016 08:36 pm
Given how little time is needed for the boostback, I seriously doubt that it would ever be worth it. In any case, you could work on an extendable and retractable nozzle. But again, I seriously doubt it is worth it.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 06/04/2016 10:24 pm
I guess I'd need to model it to truly convince myself. I was imagining a system where the second stage is smaller; payload smaller; and staging at a much higher speed.

And consequently the rocket would  spend more time throttled back on two high isp chamber/nozzles on the way up and on the deceleration and re-entry burn.

Say, staging at mach 14? then decelerating to mach 5? for re-entry, before  a barge landing. With a small cheap pressure-fed disposable upperstage.
Title: Re: Rocket Engine Q&A
Post by: nicp on 07/01/2016 11:58 am
I have an unusual and pointless question.
What does a liquid fueled rocket engine sound like? And I don't mean the hypersonic exhaust hitting the atmosphere, the huge roar we are all familiar with (at least second hand).

Mullane in 'Riding Rockets' describes the SSMEs ('three Rocketdyne beauties..') as smooth as glass (or similar, it's been a while since I've read it) after the SRBs have separated. Now there's some distance between the flight deck and the engines and probably lots of sound insulation.  (I'm not sure he actually describes them as quiet, perhaps only smooth).

So.. subtracting anythng from the engine bell, would they sound like a jet engine perhaps?
Title: Re: Rocket Engine Q&A
Post by: PahTo on 07/01/2016 05:54 pm
I have an unusual and pointless question.
What does a liquid fueled rocket engine sound like? And I don't mean the hypersonic exhaust hitting the atmosphere, the huge roar we are all familiar with (at least second hand).

Mullane in 'Riding Rockets' describes the SSMEs ('three Rocketdyne beauties..') as smooth as glass (or similar, it's been a while since I've read it) after the SRBs have separated. Now there's some distance between the flight deck and the engines and probably lots of sound insulation.  (I'm not sure he actually describes them as quiet, perhaps only smooth).

So.. subtracting anythng from the engine bell, would they sound like a jet engine perhaps?


Correct--smooth, not quiet.  Remember the HPFT is spinning at (36,000?) rpm and HPOT at a good clip too.  One cannot have that much compression/impeller force without plenty of sound.
Title: Re: Rocket Engine Q&A
Post by: the_other_Doug on 07/01/2016 06:21 pm
I have an unusual and pointless question.
What does a liquid fueled rocket engine sound like? And I don't mean the hypersonic exhaust hitting the atmosphere, the huge roar we are all familiar with (at least second hand).

Mullane in 'Riding Rockets' describes the SSMEs ('three Rocketdyne beauties..') as smooth as glass (or similar, it's been a while since I've read it) after the SRBs have separated. Now there's some distance between the flight deck and the engines and probably lots of sound insulation.  (I'm not sure he actually describes them as quiet, perhaps only smooth).

So.. subtracting anythng from the engine bell, would they sound like a jet engine perhaps?


Correct--smooth, not quiet.  Remember the HPFT is spinning at (36,000?) rpm and HPOT at a good clip too.  One cannot have that much compression/impeller force without plenty of sound.

Yeah, the engines definitely made some noise, which was transmitted through the structure of the orbiter.  After going supersonic, none of the sound the engines make in the air gets up to the cabin, of course, and then shortly thereafter there wasn't enough air to carry a lot of sound, anyway.  And consider that the crews regularly reported the OMS firings as sounding like howitzers going off, you have the think the SSMEs generated sound that carried through the structure, as well.

The only description I've ever heard of the sound of an engine firing while the spacecraft was in space was of the LM ascent engine.  It literally sat about two feet behind the crew station, the guts of the engine and the combustion chamber all inside a little can in the back.  With that small engine, the crews reported it sounded like the wind blowing in the distance -- very little sound, almost no vibration.  So, while an engine is going to generate some noise in a vacuum, it won't necessarily be a big, loud noise.  It all depends on the type of engine -- the SSMEs were pump-fed, so the pumps would generate more noise than a pressure-fed engine.

Of course, most of the classic sound of a rocket engine we hear on Earth is caused by the rapid displacement of air by the exhaust.  That sound doesn't exist in space, but most definitely, sound travels through the spacecraft structure.

I know that every simulation of the RCS engines firing in Apollo I have ever heard, both in the LM and the CSM, sounded for all the world like someone opened up a high-pressure steam valve for just a moment, so perhaps some of the engine sounds you'd encounter in a spacecraft would sound more like that, and less like the roar we're used to...
Title: Re: Rocket Engine Q&A
Post by: Proponent on 07/02/2016 11:09 am
I recall a presentation by Firestar, the company that was behind the NOFBX monopropellant, that said non-hypergolic bi-propellant engines tend to be noisier than hypergols which in turn tend to be noisier than monoprops.  I believe the claim was that the smoothness of mixing of fuel and oxidizer was a factor in noisiness.
Title: Re: Rocket Engine Q&A
Post by: PahTo on 07/02/2016 07:37 pm
I recall a presentation by Firestar, the company that was behind the NOFBX monopropellant, that said non-hypergolic bi-propellant engines tend to be noisier than hypergols which in turn tend to be noisier than monoprops.  I believe the claim was that the smoothness of mixing of fuel and oxidizer was a factor in noisiness.

Thanks for that, Proponent.  I was pondering the mass of material passing through multiple pintles, and figured that would itself generate significant vibration (sounds), and then of course there's the actual "intermixing"...
Title: Re: Rocket Engine Q&A
Post by: the_other_Doug on 07/03/2016 02:50 am
ISTR that Mike Collins described the sound and feel of the S-IVB on Apollo 11 as "crisp and rattly," though how much of that was a vibration felt more than heard is hard to say.  I'd say that most of the astronauts who have ridden rockets into space would be hard pressed to separate their perceptions of sound vs. those of vibrations felt through the seat of their couch...
Title: Re: Rocket Engine Q&A
Post by: Proponent on 07/04/2016 06:00 pm
By the way, we also have the negative evidence from Gemini 8.  That it took the crew a while to realize that the gyrations the spacecraft was undergoing were caused by a thruster that was stuck open indicates they couldn't hear thrusters in operation.  I do recall being told in another thread, however, that the opening and closing propellant valves on the Shuttle was very noisy. 
Title: Re: Rocket Engine Q&A
Post by: envy887 on 07/13/2016 05:26 pm
I guess I'd need to model it to truly convince myself. I was imagining a system where the second stage is smaller; payload smaller; and staging at a much higher speed.

And consequently the rocket would  spend more time throttled back on two high isp chamber/nozzles on the way up and on the deceleration and re-entry burn.

Say, staging at mach 14? then decelerating to mach 5? for re-entry, before  a barge landing. With a small cheap pressure-fed disposable upperstage.

Trading thrust for ISP is generally a bad idea on first stages. Every second the 1st stage has to burn is almost 10 m/s of gravity losses. Over 30% of the F9's first stage performance is just fighting gravity losses.

And the F9 can't get to Mach 14 at MECO with ANY reserves for boostback, no matter what ISP gains you theorize. It simply doesn't have the mass fraction after gravity losses (and reducing thrust will make those worse).
Title: Re: Rocket Engine Q&A
Post by: gin455res on 07/13/2016 09:28 pm
I guess I'd need to model it to truly convince myself. I was imagining a system where the second stage is smaller; payload smaller; and staging at a much higher speed.

And consequently the rocket would  spend more time throttled back on two high isp chamber/nozzles on the way up and on the deceleration and re-entry burn.

Say, staging at mach 14? then decelerating to mach 5? for re-entry, before  a barge landing. With a small cheap pressure-fed disposable upperstage.

Trading thrust for ISP is generally a bad idea on first stages. Every second the 1st stage has to burn is almost 10 m/s of gravity losses. Over 30% of the F9's first stage performance is just fighting gravity losses.

And the F9 can't get to Mach 14 at MECO with ANY reserves for boostback, no matter what ISP gains you theorize. It simply doesn't have the mass fraction after gravity losses (and reducing thrust will make those worse).


Yes, I was surprised when I first read of the t/w of rockets at take-off. 


I would still have to model it to believe it, not saying you are wrong, but that it is not self evident. Those mach numbers were plucked out my b@tt, they are not definitive. If those numbers are unrealistic, this doesn't mean others are not (or for that matter are). I would have to model it. I don't have the time or inclination right now (I need to get on with my phonics apps).


I would model a tri-core falcon 5 with a modified central core that has 3 sea level falcons and start with a single falcon feeding 6 chambers for higher ISP (this might be too heavy, and illustrate the need for a new smaller pump, or just kill the concept entirely). This is complicated and we would have to map the whole space of solutions before declaring there isn't one.


Additionally, it is not just about payload fractions. There are unknowns such as how much cheaper re-usability might make the ratio between the cost of the upper expendable stage and the marginal cost of the tri-core. This is not so easy to analyse/predict/guess.
Title: Re: Rocket Engine Q&A
Post by: Nicolas PILLET on 07/17/2016 12:54 pm
Little historical question...
S1.5400, built in OKB-1 for Molniya launch vehicle, was the first staged combustion engine.
But what was the first American staged combustion engine ? SSME ?
Thanks !
Title: Re: Rocket Engine Q&A
Post by: baldusi on 07/18/2016 01:48 pm
Little historical question...
S1.5400, built in OKB-1 for Molniya launch vehicle, was the first staged combustion engine.
But what was the first American staged combustion engine ? SSME ?
Thanks !
Non classified and production, yes. We don't know about all military engines, and regarding experiments. But yes, the Americans needed 21 more years to fly a staged combustion engine than the Russians. And the Russians started with an oxidizer rich version. Granted, SSME is clearly the most difficult, complex, expensive and sophisticated of the two debuts.
Title: Re: Rocket Engine Q&A
Post by: Nicolas PILLET on 07/18/2016 04:07 pm
the Americans needed 21 more years to fly a staged combustion engine than the Russians

But SSME was a bit more reliable than S1.5400, which is a good thing for shuttle crews ! :D
Title: Re: Rocket Engine Q&A
Post by: Sesquipedalian on 07/19/2016 03:42 pm
Why are engines usually "chilled down" prior to startup?  This strikes me as exactly the wrong thing to do.  When the engine actually starts, it immediately gets extremely hot, going through a substantial change in temperature in a fraction of a second.  Chilling down the engine would seem to drastically increase the amount of temperature change it undergoes.  Most materials don't like wild and abrupt temperature swings; they will often fracture or deform or otherwise suffer substantial degradation.  Anyone who has placed a hot glass on a cold counter and seen it shatter will be familiar with this.

Obviously, since engines don't fracture upon startup from the thermal shock, there's a link in the sequence that I'm missing.  Perhaps "chilldown" doesn't mean chill down in the usual sense?
Title: Re: Rocket Engine Q&A
Post by: PahTo on 07/19/2016 04:23 pm
Why are engines usually "chilled down" prior to startup?  This strikes me as exactly the wrong thing to do.  When the engine actually starts, it immediately gets extremely hot, going through a substantial change in temperature in a fraction of a second.  Chilling down the engine would seem to drastically increase the amount of temperature change it undergoes.  Most materials don't like wild and abrupt temperature swings; they will often fracture or deform or otherwise suffer substantial degradation.  Anyone who has placed a hot glass on a cold counter and seen it shatter will be familiar with this.

Obviously, since engines don't fracture upon startup from the thermal shock, there's a link in the sequence that I'm missing.  Perhaps "chilldown" doesn't mean chill down in the usual sense?


Hi Ses,
You actually answered your own question.  Materials don't like abrupt change, hence a controlled (slow) introduction of cryos to condition the engine components to the temps at which they'll be operating (read:  passing massive quantities of propellants at start up/through run cycle).
Title: Re: Rocket Engine Q&A
Post by: Sesquipedalian on 07/20/2016 08:09 pm
I don't follow.  Rockets are fueled over a long period of time (30 minutes or so, at least, based on the latest SpaceX launch), which means that the cryogenic components should reach cryogenic temperature gradually.  But rocket engines go from ambient temperature to red-hot in a matter of seconds.
Title: Re: Rocket Engine Q&A
Post by: PahTo on 07/20/2016 08:27 pm
I don't follow.  Rockets are fueled over a long period of time (30 minutes or so, at least, based on the latest SpaceX launch), which means that the cryogenic components should reach cryogenic temperature gradually.  But rocket engines go from ambient temperature to red-hot in a matter of seconds.

The only portions of a rocket engine that get hot are (mostly) the combustion chamber and nozzle, and preburners as applicable (frictional heating of bearings for the turbo machinery, etc. should be considered separate of this conversation).  Each of these hot components usually cycle cryos around them to keep them from melting (this also serves to pre-heat said props for (in part) helping to drive turbo machinery).  Chill down is for piping, stators and vanes, other turbo hardware, valves, valve housings, etc.    There are some excellent write ups on SSME/RS25 (or used to be, haven't been there in years) on NASA sites.  I'm not a rocket scientist, and have learned most of what I know from googling and reading manuals and such.  This site helps too!  :) 
Title: Re: Rocket Engine Q&A
Post by: nicp on 07/27/2016 07:03 pm
A couple of book I've read explain the cancellation of the E-1 engine due to it being a 'dead end'. That description never seems to be backed up. A dead end because LR-87 was closer to ready? Or - as I sometimes read it - not big enough with the F-1 on the horizon?

Perhaps there just were no stages that needed it, but I see no great reason why E-1 was fundamentally a dead end. A shame it didn't fly.

Any opinions ?
Title: Re: Rocket Engine Q&A
Post by: Jim on 07/27/2016 07:14 pm
It was OBE.  With the "re engining" of the Saturn with the H-1 precursor and success of the LR-87, there was no need for it.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 09/23/2016 03:35 pm
Highly technical question, but here it goes: should steam generators like RD-107/8 be considered different than traditional gas generators? On the same line of thought, Gamma rocket should be considered staged combustion?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 09/24/2016 12:51 pm
Well, I remember reading Wernher von Braun's popular book Space Frontier long ago.  IIRC, the four cycles he described were pressure-fed, gas-generator (by which he meant monoprop gas-generator), bi-propellant (i.e., what we call GG around here), and topping (regenerative).  So he seemed to regard it as a separate cycle.
Title: Re: Rocket Engine Q&A
Post by: baldusi on 09/24/2016 06:37 pm
Well, I remember reading Wernher von Braun's popular book Space Frontier long ago.  IIRC, the four cycles he described were pressure-fed, gas-generator (by which he meant monoprop gas-generator), bi-propellant (i.e., what we call GG around here), and topping (regenerative).  So he seemed to regard it as a separate cycle.
I haven't been able to read that book. I don't quite understand what does he refers as to with topping. And science has really advanced a lot since then. You have closed expander, bleed expander, staged combustion, full flow staged combustion, tap off, electrically pumped and any combination of the above.
With a modern view, the difference is quite minimal. You still take some of the mass of the liquids on board and make hot gas to drive the turbopumps. I'm wondering if writing a whole article for it or just an entry on the Gas Generator cycle on Wikipedia. But I haven't been able to find exactly the quote that considers them different. Now, if I only had access to a copy of that book. NASA perhaps?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 09/24/2016 11:10 pm
"Topping cycle" was his term for closed expander.  Now that you mention it, he also described the tap-off cycle, referring to it as "thrust-chamber bleed" (mind you, I'm recalling something I read years ago as a child, and could be making a mistake).  The book was a compilation of articles that von Braun had written for the magazine Popular Science.

The ancient (1961) Handbook of Astronautics identifies precisely the same four cycles: bipropellant, monopropellant, topping and thrust-chamber bleed.

By the way, I see that Sutton & Bilbarz (8th ed.) use "topping cycle" as a synonym for "closed cycle."   On gas generators, unlike von Braun, they regard (p. 225) the monoprop cycle as simply an obsolete variant of the gas-generator cycle.
Title: Re: Rocket Engine Q&A
Post by: Prober on 10/01/2016 05:12 pm
Rocket Engine Plumbing


https://www.youtube.com/watch?v=4QXZ2RzN_Oo (https://www.youtube.com/watch?v=4QXZ2RzN_Oo)
Title: Re: Rocket Engine Q&A
Post by: haywoodfloyd on 02/26/2017 11:54 am
I heard a report on Daily Planet on the Discovery Channel that said that CO2 was produced along with water vapour when the RS-25 was test fired.
Is this true?
Title: Re: Rocket Engine Q&A
Post by: Jim on 02/26/2017 12:04 pm
I heard a report on Daily Planet on the Discovery Channel that said that CO2 was produced along with water vapour when the RS-25 was test fired.
Is this true?


no
Title: Re: Rocket Engine Q&A
Post by: nacnud on 02/26/2017 12:06 pm
RS-25 burns Oxygen and Hydrogen.

Where would the carbon come from?
Title: Re: Rocket Engine Q&A
Post by: haywoodfloyd on 02/26/2017 12:17 pm
RS-25 burns Oxygen and Hydrogen.

Where would the carbon come from?

My thoughts exactly.
Title: Re: Rocket Engine Q&A
Post by: nacnud on 02/26/2017 12:40 pm
If they are thinking about pollutants maybe there is some NOx produced?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 02/26/2017 05:45 pm
Since the exhaust is hydrogen rich and hot, I imagine a number of compounds involving H, N and O would form, possibly including some nitric acid (HNO3).
Title: Re: Rocket Engine Q&A
Post by: haywoodfloyd on 02/26/2017 07:03 pm
Since the exhaust is hydrogen rich and hot, I imagine a number of compounds involving H, N and O would form, possibly including some nitric acid (HNO3).

But not CO2, right?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 02/27/2017 08:34 am
No -- like nacnud said a few posts up, where would the carbon come from?
Title: Re: Rocket Engine Q&A
Post by: gospacex on 02/27/2017 02:14 pm
Since the exhaust is hydrogen rich and hot, I imagine a number of compounds involving H, N and O would form, possibly including some nitric acid (HNO3).

Exhaust is not hot. It's actually circa ~100 Celsius - because a well-designed engine converts almost all thermal energy (random motion) into energy of the *directed* stream of gas.

On impact into "motionless" atmosphere, some reheating occurs, but I don't know how efficient it is. It might be too low to dissociate N2.
Title: Re: Rocket Engine Q&A
Post by: R7 on 02/27/2017 02:47 pm
Exhaust is not hot. It's actually circa ~100 Celsius - because a well-designed engine converts almost all thermal energy (random motion) into energy of the *directed* stream of gas.

RS-25 exhaust is about 1150K because the area ratio is still limited by sea level ambient pressure. Still below thermal NOx level which is about 1800K. Reheat happens in the shock diamonds, but AFAIK at that point there's not much mixing between plume and ambient air so there's no or just little N present.
Title: Re: Rocket Engine Q&A
Post by: Robert Willis on 08/03/2017 03:26 pm
Engines designed to burn liquid hydrogen, such as RD-0120 & RD-0146 have been extensively test fired running on liquid methane with little modification. RD-701 was actually capable of switching back & forth from kerosene to hydrogen in flight! Seeing as Raptor was originally planned to burn LH2, how difficult would it be to produce such an engine with a high degree of component commonality with the CH4 burning model currently under development? NASA buying a few dozen of these for an improved SLS at a fraction of what AR charges per unit for RS-25 would be a helpful source of funding for SpaceX. Please correct me if I'm wrong, but I would guess that an LH2 fueled Raptor would have lower thrust, but higher ISP than the CH4 powered Raptor baseline. Can anyone out there do some rough calculations/estimates?

Doubtless the Raptor will drastically less expensive than the RS-25; NASA is doling out one point six billion for a mere six new engines to Aerojet-Rocketdyne. Would a hydrogen burning raptor not make a drastically more cost effective RS-25 replacement for SLS applications?
Title: Re: Rocket Engine Q&A
Post by: nicp on 09/03/2017 11:04 am
I've just read the Wikipedia article on the J-2. It mentions that the J-2S would have used a de Laval nozzle.

For quite some time I had assumed all rocket engines used de Laval nozzles, though on a few occasions (looking at photographs of an F-1 for example) I did wonder, but put the seeming lack of convergent/divergent form to camera angle or perspective.

The implication is that some rocket engines do not use a de Laval nozzle, and that the (more efficient) J-2S would have.

So my questions are...

Why use - or not use - a de Laval nozzle?
Can there be a disadvantage in using one?
Does a non-de Laval nozzle achieve choked flow (surely it must?)


Title: Re: Rocket Engine Q&A
Post by: robert_d on 09/03/2017 12:48 pm
What in general is required to make an engine air-startable and restartable?
What makes a design such as the SSME harder to accomplish this? Hope to find some commonality to similar questions regarding the Raptor development in that thread. Thanks.
Title: Re: Rocket Engine Q&A
Post by: brickmack on 09/03/2017 02:43 pm
I've just read the Wikipedia article on the J-2. It mentions that the J-2S would have used a de Laval nozzle.

J-2 definitely had a de Laval nozzle. Wikipedia just says that to distinguish J-2S and J-2T, since the latter used an aerospike
Title: Re: Rocket Engine Q&A
Post by: Jim on 09/05/2017 08:22 pm
What in general is required to make an engine air-startable and restartable?
What makes a design such as the SSME harder to accomplish this? Hope to find some commonality to similar questions regarding the Raptor development in that thread. Thanks.

The ability to get the engine parameters into the start box and provide energy to start the engine.

SSME was a head start engine, it relied on the pressure generated by the weight of the propellants.  It also was constantly conditioned by ground sources.  Both of these are hard to do in a free fall at over 100kft.

Don't need to worrying about the Raptor, it will be designed for ir-startable and restartable, just like the Merlin.  There is nothing special that needs to be done.
Title: Re: Rocket Engine Q&A
Post by: brickmack on 09/06/2017 01:48 am
Addressing *re*-startability, regardless of location (ie, if you've got an engine that stays on the ground and want to restart it), since Jim already covered the main points of air start. A lot of engines, particularly older ones, did things to themselves during startup and shutdown that would make it very difficult to restart them without serious maintenance. Valves would be opened with pyrotechnics, and then closed again at the end of the burn in the same manner. Thermal stresses could also seal valves in one position. Pumps, in engines that had them, might be spun up by small solid rockets. Ignition might use hypergolic or pyrophoric injection (often in burst discs rather than normal plumbing), or pyrotechnics/small solids. All of these issues would require at least replacing several easily-accessible parts between firings (some such engines could actually fire multiple times, but usually only a very small number), if not a significant disassembly.

So, main points anyway: pneumatic or electromechanical valves, temperature-safe parts, hypergolic/pyrophoric ignitor fluid injected through normal reusable plumbing (or, if feasible, use electrical ignition and avoid the problem completely), and use compressed gas/cryogenic fluid expansion/electromechanical means to spin up turbopumps
Title: Re: Rocket Engine Q&A
Post by: darkenfast on 09/06/2017 06:14 am
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage?  Didn't the people who pushed the SSME for the Upper Stage know this?  How were they going to deal with this?
Title: Re: Rocket Engine Q&A
Post by: PahTo on 09/06/2017 02:35 pm
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage?  Didn't the people who pushed the SSME for the Upper Stage know this?  How were they going to deal with this?

I answered earlier with a snarky "money" post, then had a sip of coffee and my brain kicked in.  I believe the J-2X was slated for the upper stage, not SSME.  Even still, there was a ton of money on an engine that now will likely never see in-space action.
Title: Re: Rocket Engine Q&A
Post by: Welsh Dragon on 09/08/2017 08:45 am
Nope, the original concept had an airstart SSME.
Title: Re: Rocket Engine Q&A
Post by: wolfpack on 09/09/2017 01:14 am
What is the cleaning procedure for RP-1 fueled engines between static fires and flights/re-flights? Is it still a trichloroethylene flush?
Title: Re: Rocket Engine Q&A
Post by: JAFO on 01/07/2018 09:41 pm
Ok, bear with me, and I'm probably going to set the record for the dumbest question in the history of the forum, but here's a question I missed from Freshman Rocketry 1.

When I was building Estes rockets a long time ago I accepted the premise that you press the button, the solid ignites, and the force of the escaping gases exiting the motor produce an opposite force against the "top" of the engine, making the rocket go up. For a solid motor, it seems pretty straightforward.

But what about a liquid fueled engine? Hot gases go out the back, but what are they "pushing" against? The engine bell? The thrust chamber? Whatever they're pushing against, how is that force transmitted to the vehicle? In the case of the Falcon, how complicated is it to transfer the thrust of 9 engines to the first stage? If they're pushing against the engine bell/nozzle, how strong is that thing? We all rest our models on them, but I doubt they're strong enough to hold the mass of even an unfueled vehicle in real life.


Thanks for your patience,
Title: Re: Rocket Engine Q&A
Post by: Bernie Roehl on 01/07/2018 10:04 pm
Rocket engines actually don't "push" against anything. It's all about conservation of momentum.

The rocket+propellant are initially at rest. Expelling propellant in one direction means the rest of the system has to move in the other direction so that the net momentum remains at zero. Momentum is mass times velocity, so the more mass you expel and the faster you expel it, the faster you go in the opposite direction.

The classical rocket equation is Vrocket = Vexhaust x ln(M/m), where M is the mass of the rocket+propellant and m is the "dry weight" or the mass of the rocket by itself without the propellant.

Notice that I say "propellent", which is not quite the same as "fuel". Fuel is what provides the energy, propellant is the reaction mass. In a chemical rocket, those two are basically the same thing -- the reaction mass is the byproduct of burning the fuel (with an oxidizer).

In a nuclear thermal rocket, the propellant is usually hydrogen and the energy comes from a reactor.

In an ion engine, the reaction mass is something like Xenon and the energy comes from a solar or nuclear electric source which accelerates the propellant ions using electric fields.

And so on.

Title: Re: Rocket Engine Q&A
Post by: Jim on 01/09/2018 12:21 am
Rocket engines actually don't "push" against anything. It's all about conservation of momentum.


Wrong, they do,  the combustion chamber
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/09/2018 04:04 pm
Rocket engines actually don't "push" against anything. It's all about conservation of momentum.


Wrong, they do,  the combustion chamber

Two sides of the same coin.  The usual expression for the thrust generated by a rocket engine is

      qc + (pe - pa)Ae ,

where q is the mass flow rate, c is the exhaust velocity, pe is the pressure at the nozzle exit, and pa is the ambient pressure.  This is usually derived by appealing to conservation of momentum for the first term and then, rather unconvincingly in my opinion, adding the second term to account for pressure differences.  But it is also possible to derive the whole expression by considering nothing but the pressures, both internal and ambient, on the combustion chamber and nozzle.  In this case the pressure term arises quite naturally.
Title: Re: Rocket Engine Q&A
Post by: the_other_Doug on 01/09/2018 06:27 pm
I think we're getting lost in semantics, here.  The original question, I believe, was asking where in the rocket is the force applied by the engine's escaping gasses -- commonly just called thrust -- transferred into the structure of the rocket as a whole.

Jim's right -- the escaping gasses apply force against the side of the combustion chamber opposite from the hole in the chamber that lets the gasses escape.  That force (thrust) is transferred into the body of the rocket by firmly attaching the combustion chambers to a thrust bulkhead, so to speak -- a portion of the structure of the rocket built heavily enough to withstand the force applied through the backsides of the combustion chambers and safely, with no structural failures, transfer that force through to the entire rocket's mass.

Without a strong enough thrust bulkhead, a rocket would act like a model rocket were the builder forgot to glue the engine stop ring into the engine mount.  The engine would, while it was thrusting, just pass up through the rocket's structure and out the top...  :o
Title: Re: Rocket Engine Q&A
Post by: Lars-J on 01/09/2018 08:13 pm
I think we're getting lost in semantics, here.  The original question, I believe, was asking where in the rocket is the force applied by the engine's escaping gasses -- commonly just called thrust -- transferred into the structure of the rocket as a whole.

Jim's right -- the escaping gasses apply force against the side of the combustion chamber opposite from the hole in the chamber that lets the gasses escape.  That force (thrust) is transferred into the body of the rocket by firmly attaching the combustion chambers to a thrust bulkhead, so to speak -- a portion of the structure of the rocket built heavily enough to withstand the force applied through the backsides of the combustion chambers and safely, with no structural failures, transfer that force through to the entire rocket's mass.

Without a strong enough thrust bulkhead, a rocket would act like a model rocket were the builder forgot to glue the engine stop ring into the engine mount.  The engine would, while it was thrusting, just pass up through the rocket's structure and out the top...  :o

Exactly... In image 1 (Merlin 1C for Falcon 1), you can see such a thrust structure. The four beams in this case transfers the load from the engine to the tanks and rocket structure.

Image 2 is the Merlin 1D, which just uses a plate. (top of engine and thrust vector actuators connect to the plate) That plate is bolted to the octaweb, which transfers the combined load to the rocket.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 01/10/2018 02:05 am
... the escaping gasses apply force against the side of the combustion chamber opposite from the hole in the chamber that lets the gasses escape.

At points within the chamber or nozzle where the internal pressure exceeds the ambient pressure, diverging sections (including the top of the combustion chamber) contribute to thrust while converging sections detract.  Where ambient pressure exceeds internal pressure, as is the case in the tail end of an over-expanded nozzle, it's the other way around.
Title: Re: Rocket Engine Q&A
Post by: nicp on 01/30/2018 07:50 am
I was wondering recently what kind of injector Raptor might have, and vaguely assumed it might be another pintle.
But then I remembered (I think this is right) that the fuel and oxidizer are both fully in the gaseous phase before hitting the main combustion chamber. This may also be true of BE-4.

As I recall one advantage of pintle injectors is you can get 'free' film cooling from fuel hitting the chamber walls. Which presumably is not going to happen with fully gaseous propellants.

But then (presumably) even a classic waterfall injector is going to have to be designed a little differently for gaseous propellants.

So what does a gaseous injector look like?
Title: Re: Rocket Engine Q&A
Post by: DaveS on 12/12/2019 12:56 am
Does anyone know how much of the RL-10 its thermal shield blankets covered the actual engine? I can't find any actual photos of it, just partial views from Rocketcams from the various Atlas/Delta flights that used them.
Title: Re: Rocket Engine Q&A
Post by: fl1034 on 12/25/2019 12:43 pm
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.
Title: Re: Rocket Engine Q&A
Post by: Unrulycow on 12/25/2019 01:20 pm
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.
It has. The AR1 is an ORSC engine proposed for use in Vulcan.  They choose BE4 instead and AR1 may potentially be used in Firefly Beta in the future.
https://en.m.wikipedia.org/wiki/AR1
Title: Re: Rocket Engine Q&A
Post by: fl1034 on 12/28/2019 11:29 am
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.
It has. The AR1 is an ORSC engine proposed for use in Vulcan.  They choose BE4 instead and AR1 may potentially be used in Firefly Beta in the future.
https://en.m.wikipedia.org/wiki/AR1

Well, why not LOX cooling? Splitting the usual coaxial shaft pump of the kerolox engine into 2, use LOX and LOX only for regen cooling, adjust the fuel side preburner temperature, then you get completely independent mixture ratio control, no coking to worry about, much easier pumps to design, and the ability to use from methane all the way to kerosene or even diesel for fuel. Just saying.
Title: Re: Rocket Engine Q&A
Post by: Redclaws on 12/28/2019 12:37 pm
The obvious concern - which may not be relevant, but still - is well, it’s LOX.  It’s comparatively *very* nasty, even cold, and is going to introduce new material requirements in the cooling channels.  I would be interested to hear livingjw on this topic.
Title: Re: Rocket Engine Q&A
Post by: mmeijeri on 12/28/2019 01:02 pm
I recall reading an old ('70s) study on LOX cooling on NTRS, which found that, surprisingly, this was not as dangerous as you might imagine, even if the cooling channels sprung a leak.
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 12/28/2019 01:12 pm
Well, why not LOX cooling? Splitting the usual coaxial shaft pump of the kerolox engine into 2, use LOX and LOX only for regen cooling, adjust the fuel side preburner temperature, then you get completely independent mixture ratio control, no coking to worry about, much easier pumps to design, and the ability to use from methane all the way to kerosene or even diesel for fuel. Just saying.
That is a very good question.

It also means you don't have to fiddle with the cooling circuit if you change fuels, but not oxidizer (and given LOX is about the best available in terms of performance and cost why would you?)

In fact it has been used in at least one engine (the rocketdyne plug nozzle test bed for the USAFRL in the mid 70's used a dual expander cycle with LOX cooling and LH2 cooling of the modular combustion chambers to drive the separate pumps, eliminating interpropellant seals and the criticality 1 failure modes. 

Likewise Rotary Rocket tested LOX cooling in the early 90s. NASA ran tests on a 40 000 lb pressure fed test engine (late 80's, early 90'x) which included deliberate  leaks into the chamber. Nothing bad actually happened.

Here's the thing.  The bulk of rocket engineering was done in the 50's and 60's. It was done at breakneck speed (by modern standards) because of the Cold War.  If something looked too difficult to deliver in the short term it was discarded.

And TBH many of those decisions have never been reviewed or retested.

So generations of folklore and myth have accumulated (passed on by lecturers who've either never worked with it or never been asked to question what they have been told by their lecturers) that "Oh noes. LOX will burn anything"
It's quite interesting that AFAIK the propulsion team at Rotary had not gone through the conventional aerospace engineering education process and so had not received the stories, myths and old wives tales that have built up. 

The other classic meme is that HTP in unstable, despite it being specified for "Super performance" booster engines on early jet aircraft and supplying on orbit station keeping for a comms sat for 6 years in the early 60's.

IRL LOX cooling is likely to cool the area around the hole and reduce the risk of ignition. A big enough leak in the chamber cooling system is a serious issue regardless of what the coolant is.
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 12/28/2019 01:18 pm
The obvious concern - which may not be relevant, but still - is well, it’s LOX.  It’s comparatively *very* nasty, even cold,
Welcome to the site.

Actually it's nasty mostly because it is cold.  If it leaks it warms up, vaporizes and disperses. Contrast that with NTO, which is seriously toxic as a liquid and makes a quite good WMD if it vaporizes. It has also been know to explosively decompose.
Quote from: Redclaws
and is going to introduce new material requirements in the cooling channels.  I would be interested to hear livingjw on this topic.
I'd be interested in hearing more from HMX on the subject, given his company actually built LOX cooled combustion chambers.

Or you could just look up the NASA papers on www.sti.nasa.gov for the subject.
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 12/28/2019 01:26 pm
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(
Title: Re: Rocket Engine Q&A
Post by: envy887 on 12/30/2019 03:39 pm
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(

Wasn't the main issue lack of pressure head in freefall after staging? I'm curious why they didn't opt for large solid ullage motors, or hot staging, as those would provide the acceleration needed for a pressure head.
Title: Re: Rocket Engine Q&A
Post by: DaveS on 12/30/2019 05:50 pm
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(

Wasn't the main issue lack of pressure head in freefall after staging? I'm curious why they didn't opt for large solid ullage motors, or hot staging, as those would provide the acceleration needed for a pressure head.
Airstarting the SSME after staging wasn't the problem, it was re-starting it for the TLI burn. NASA wanted a common engine for both the Ares I and Ares V upper stages so when they ran into the problem of getting a good re-start of the SSME on the Ares V upper stage, they decided it wasn't worth expending the resources into making it re-startable in flight and went with the J-2X as the upper stage engine on both LVs. This had the cascade effect of making the Ares I upper stage unable to insert the CEV into orbit which forced them to not only upsizing the upper stage but the booster stage as well (was derivative of the STS SRB but became a brand new five segment version with no flight history).
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 12/30/2019 08:57 pm
Airstarting the SSME after staging wasn't the problem, it was re-starting it for the TLI burn. NASA wanted a common engine for both the Ares I and Ares V upper stages so when they ran into the problem of getting a good re-start of the SSME on the Ares V upper stage, they decided it wasn't worth expending the resources into making it re-startable in flight and went with the J-2X as the upper stage engine on both LVs. This had the cascade effect of making the Ares I upper stage unable to insert the CEV into orbit which forced them to not only upsizing the upper stage but the booster stage as well (was derivative of the STS SRB but became a brand new five segment version with no flight history).
Wow.

That is a truly impressive lack of anything like foresight.
Title: Re: Rocket Engine Q&A
Post by: nicp on 01/03/2020 09:26 pm
What is the big deal about throttling liquid fueled rocket engines? This was definitely considered a big deal for the LEM descent engine, and even that (which clearly worked well) could get erosion at certain throttle settings.
As I understand it _that_ engine displaced some of the propellants with inert helium.

But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
Title: Re: Rocket Engine Q&A
Post by: john smith 19 on 01/06/2020 08:41 pm
But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
It isn't.

F9 takes off on 9 engines then lands on 1.

That alone deals with most of the "deep" throttling problem.

A Merlin at 50% thrust (a fairly shallow throttle providing your nozzle expansion ratio is not extreme) has 1/18 of an F9's take off thrust.
Title: Re: Rocket Engine Q&A
Post by: walkermo on 01/13/2020 09:13 pm
What are the challenges around designing an electric pump engine similar to the Rutherford or Stealth Space/Astra’s engine? How much complexity & cost does it add over a pressure fed system?

Asking as a member of a student team that has successfully built and tested a N2O/Kerosene pressure fed engine.
Title: Re: Rocket Engine Q&A
Post by: TrevorMonty on 01/13/2020 10:55 pm
What are the challenges around designing an electric pump engine similar to the Rutherford or Stealth Space/Astra’s engine? How much complexity &amp; cost does it add over a pressure fed system?

Asking as a member of a student team that has successfully built and tested a N2O/Kerosene pressure fed engine.

Design suitable battery, power electronics, electric motor and pumps. Maybe able to source motor off shelf but battery pack would be custom build. Keeping everything cool at these high power levels would also add to design challenges.

Would be good project for student rocket but would take few years and multi discipline team. Great to have on team members' CV even if not successful as all these disciplines are in high demand across lots of industries.
Title: Re: Rocket Engine Q&A
Post by: Hog on 01/14/2020 05:14 pm
But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
It isn't.

F9 takes off on 9 engines then lands on 1.

That alone deals with most of the "deep" throttling problem.

A Merlin at 50% thrust (a fairly shallow throttle providing your nozzle expansion ratio is not extreme) has 1/18 of an F9's take off thrust.
I remember the paper dealing with the RS-25 17%RPL 24%RPL 40%RPL deep throttle testing.  It's amazing how complicated things cab get.
Title: Re: Rocket Engine Q&A
Post by: robert_d on 06/04/2020 08:42 pm
 I had a question that is sort of related to the development of the Super Heavy booster by SpaceX. I thought, especially before the engine count reduction to thirty-one, that stuffing them all into nine meters was impossible so that they would need a larger ‘skirt’ or shroud to shield the sides of the outer ring. It appears to be a good idea to keep that ring as small as possible. I also know that certain existing rocket engines such as Atlas 5’s RD-180 have one set of turbo machinery feeding two engine bells. So is the reverse possible? Could multiple engine sets feed into one bell to save space? What I conceived was a compound shaped bell with three lobes that had an exit area close to the exit area of three standard engines. This would need to assume that the turbo-machinery could coexist nearer to each other than any limitation of the outer bell diameter. Is this at least theoretically possible? I have attached a picture of the conceptual shape I had in mind. The outer triangle depicts a hemispherical divide where the normal engine bells would just touch. The inner triangle defines the position where the total exit area of the lobed shape would approximate the actual area of three standard engine bells.
Anyone care to comment?
Title: Re: Rocket Engine Q&A
Post by: robert_d on 06/04/2020 08:57 pm
... Could multiple engine sets feed into one bell to save space? What I conceived was a compound shaped bell with three lobes that had an exit area close to the exit area of three standard engines. ...
Clarification: This would all occur AFTER the actual combustion chambers and hopefully only deal with the geometry of the final exhaust and cooling thereof.

edit: "and cooling thereof" isn't clear but I was trying to acknowledge that the extra-ordinary shape might cause complications in any regenerative cooling system.

edit 2: note on outer circle in the diagram - it isn't anything except the it was the only way I could find to hold the proper geometric center of the lobed shape in my inexpensive CAD program.
Title: Re: Rocket Engine Q&A
Post by: Crispy on 06/04/2020 09:45 pm
If one of your three combustion chambers suffers a failure, then you have severely unbalanced pressure inside the bell. It would likely deform or suffer from flow separation on the faulty side. You'd have to shut all three chambers down to avoid further damage. This reduces engine-out capability

You have less surface area to absorb the same amount of energy via bell cooling. This reduces efficiency (heat is "wasted" by ejecting a hotter exhaust)

If you really want packing efficiency, you might consider hexagonal bells, but no pressure vessel likes being far from a circle without getting heavy (a downside that especially applies to the lobed shape)
Title: Re: Rocket Engine Q&A
Post by: robert_d on 06/07/2020 12:24 am
If one of your three combustion chambers suffers a failure, then you have severely unbalanced pressure inside the bell. It would likely deform or suffer from flow separation on the faulty side. You'd have to shut all three chambers down to avoid further damage. This reduces engine-out capability


Thanks for your thoughtful response.
I can understand the first objection but wonder if the fact that since a standard engine can throttle down to 66% or more this might not be a deal breaker. All changes would be reductions in pressure and the flow would not be affected until some point beyond the combustion chamber. So can I hold to a faint hope that it wouldn't be as bad as you think? 
Title: Re: Rocket Engine Q&A
Post by: robert_d on 06/07/2020 12:29 am

You have less surface area to absorb the same amount of energy via bell cooling. This reduces efficiency (heat is "wasted" by ejecting a hotter exhaust)
I don't understand this at all, sorry. I always thought having the most energetic exhaust was the 'holy grail' of rocket engines, but obviously I am totally off on this. Can you point me to something that would explain?
Title: Re: Rocket Engine Q&A
Post by: Bureaucromancer on 06/17/2020 05:49 am
I'm wondering if there's anyone out there who's particularly familiar with the M-1 can give me a sense of how it would have likely compared to the RS-68... I mean I can read the nominal thrust numbers for the 68 and what Astronautix has on the M-1, but how realistic are those M-1 numbers, and how far could it have been M-1 really be pushed?

Reading the numbers as published makes it seem like the M-1 wasn't really all that much more powerful, especially in terms of what I'm contemplating, which is all hydrolox first stages, but I'm vaguely suspicious of concluding the performance gain is so small between the size difference and fifty years...

This is all by way of me considering the potential for all hydrolox boosters the super heavy and medium lift ranges. Mostly just musing on what something Falcon/New Glenn like, or SuperHeavy/Saturn V like is would be like with an all hydrolox first stage, and most especially what kind of engine it would take to roughly equal an F-1's sea level performance with hydrolox.
Title: Re: Rocket Engine Q&A
Post by: robert_d on 06/18/2020 01:10 pm
Regarding my questions regarding stuffing rocket engines closer together:

If you really want packing efficiency, you might consider hexagonal bells, but no pressure vessel likes being far from a circle without getting heavy (a downside that especially applies to the lobed shape)

I do not see how the lobed shape is heavier. In fact just the opposite (or I would not have even asked). At the lower end it would have 2/3rd the surface area. Working back toward the linked combustion chambers, it would have to transition back to the original shape, but I can't conceive of additional structure beyond that. At no point can I see where the pressure is higher (or flow forcefully closer than the three standard bells) that would require beefier structure. I haven't been able to make a good diagram yet showing how to get from the combustion chambers to the bell openings, so I may be way off base.
   
Title: Re: Rocket Engine Q&A
Post by: gin455res on 06/29/2020 06:31 am
Are there any gas-generator rocket engines that pre=heat the fuel via a heat exchanger in the the gas-generator combustor, and would this reduce sooting and improve efficiency by allowing the gas-generator to run less rich?


And how bad are methane gas-generators for sooting?


I read that the ariane 4 engine used water injection, can this be used in hydrocarbon gas generators to reduce sooting?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 06/29/2020 01:48 pm
And how bad are methane gas-generators for sooting?

Well, in first attachment to this post (https://forum.nasaspaceflight.com/index.php?topic=31040.msg1310156#msg1310156) you can see the chemical results of burning a very rich lox-methane mixture, O/F = 0.23 by mass (1/9 by mole fraction).  The result is about 8% carbon by mass, which sounds a bit sooty to me, at 800 K.
Title: Re: Rocket Engine Q&A
Post by: panckage on 07/04/2020 11:38 pm
Why does the combustion from a turbopump chamber point towards the top of the rocket? It seems that if it was pointed down, it would contribute to the engine thrust. Does it have something to do with maximizing pressure through the turbopump?
Title: Re: Rocket Engine Q&A
Post by: edzieba on 08/28/2020 04:02 pm
More of a fun question: How many, if any a craft have reached or exceeded their engine's exhaust velocity, such that a 'stationary observer' would see a vehicle fly past at exorbitant speed, depositing a 'stationary' cloud of reaction products as it goes?

Inevitable caveats:
- 'Stationary' being relative, this would encompass three categories: stationary relative to the Sun, stationary relative to the Earth, and stationary relative to the nearest object in stable orbit (I was going to go for 'planetary body' but asteroids and comets can count too).
- Only partial credit given if the craft has undergone gravity boosts to reach that velocity subsequent to final engine firing (e.g. probes with a non-detachable kick stage). Full marks only if it reached that velocity while the engine was operating.

My suspicion is that only cold-gas RCS systems on hyperbolic-trajectory probes may manage to qualify due to their low exhaust velocity.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 08/28/2020 04:15 pm
Neglecting losses, all it would take is a mass ratio of at least e=2.718.... 
Title: Re: Rocket Engine Q&A
Post by: whitelancer64 on 08/28/2020 04:46 pm
More of a fun question: How many, if any a craft have reached or exceeded their engine's exhaust velocity, such that a 'stationary observer' would see a vehicle fly past at exorbitant speed, depositing a 'stationary' cloud of reaction products as it goes?

Inevitable caveats:
- 'Stationary' being relative, this would encompass three categories: stationary relative to the Sun, stationary relative to the Earth, and stationary relative to the nearest object in stable orbit (I was going to go for 'planetary body' but asteroids and comets can count too).
- Only partial credit given if the craft has undergone gravity boosts to reach that velocity subsequent to final engine firing (e.g. probes with a non-detachable kick stage). Full marks only if it reached that velocity while the engine was operating.

My suspicion is that only cold-gas RCS systems on hyperbolic-trajectory probes may manage to qualify due to their low exhaust velocity.

This is possible, but AFAIK it hasn't been done in space.

Mythbusters gave a practical demonstration of this by firing a ball out the back of a moving vehicle at the same speed as the vehicle; from the perspective of the camera at the firing point, the ball dropped straight down.

It looks very strange because it defies our expectations of what happens when a cannon fires a ball.

https://youtu.be/BLuI118nhzc
Title: Re: Rocket Engine Q&A
Post by: 1 on 08/28/2020 06:01 pm
More of a fun question: How many, if any a craft have reached or exceeded their engine's exhaust velocity, such that a 'stationary observer' would see a vehicle fly past at exorbitant speed, depositing a 'stationary' cloud of reaction products as it goes?

All of them.

Fundamentally because a reference frame can always be defined where that exhaust velocity is zero (which you allude to in your 'stationary caveats') but more generally because exhaust velocities are not very high; topping out around 4.5 km/s for high isp (i.e. hydrolox) engines. Since LEO requires an orbital velocity a bit below 8 km/s, all vehicles will pass through a point where it appears to an earth-bound observer that the exhaust products are not moving.

And, of course, beyond that point is the regime where exhaust products are moving forward relative to an observer on Earth.
Title: Re: Rocket Engine Q&A
Post by: whitelancer64 on 09/05/2020 02:06 am
Offhand question, how big would the bell of a vacuum optimized F-1 engine be?
Title: Re: Rocket Engine Q&A
Post by: Damon Hill on 09/05/2020 02:17 am
Offhand question, how big would the bell of a vacuum optimized F-1 engine be?

See images of M-1, a hydro-lox F-1 equivalent.  Really huge.   Looks like the Russian practice of multiple smaller engines and nozzles is a bit more practical.
Title: Re: Rocket Engine Q&A
Post by: Proponent on 09/05/2020 04:49 pm
Offhand question, how big would the bell of a vacuum optimized F-1 engine be?

Since some versions of the proposed Nova launch vehicle had one or two F-1's in the second stage, there might actually be documentation on a vacuum version of the F-1 lying around somewhere.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 11/08/2020 06:15 am
Is there a minimum pressure that a FFSC/ORCS engine will work at?
[and could this be lowered using catalytic combustion]


I imagine that there is a minimum temperature in the pre=burner which may set a minimum power release that will get converted* into pressure.


*how much scope is there to vary the conversion efficiency of the pump
Title: Re: Rocket Engine Q&A
Post by: Okie_Steve on 02/21/2021 10:11 pm
This is a general engine question, not about Raptor, but a quick search did not turn up anything and the people here seem like the best collective memory to ask  :)

Has anyone ever tried to use a turbo pump assembly essentially as a magneto to supply electrical power for the engine while it was running? I realize it would change the stresses etc and not be a magic drop in, just wondering if it was ever tried for TVC or whatever in the past.  :o
Title: Re: Rocket Engine Q&A
Post by: chopsticks on 02/22/2021 04:30 am
Really interesting question, I'd like to know too. Electric TVC is kind of the holy grail as I understand it.

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Title: Re: Rocket Engine Q&A
Post by: Jim on 02/23/2021 01:03 pm
Really interesting question, I'd like to know too. Electric TVC is kind of the holy grail as I understand it.

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pressure fed engines use electric TVC
Title: Re: Rocket Engine Q&A
Post by: Hog on 02/24/2021 12:05 am
Really interesting question, I'd like to know too. Electric TVC is kind of the holy grail as I understand it.

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Advanced Boosters for SLS are to drop the hydrazine and use eTVC.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 10/06/2021 01:18 am
Are there any designs for engines that are a mixture of expansion deflection and aerospike?


I imagine a two concentric aerospikes with the inner aerospike angled  outwards and the outer one angled inwards.  The spike would be annular and perhaps expander-cycle.
Title: mach diamonds more common than you think!
Post by: brettly2021 on 10/29/2021 01:25 pm
you dont need to be a rocket scientist to produce mach diamonds, in fact people are probably doing it when they use some spray cans and dont even realise it. I was really surprised to find this out, some spray cans the gas exiting is faster then speed of sound an mach diamonds are produced ( not usueally visible to naked eye but non ther less they are there)
nice video gives the info:
https://www.youtube.com/watch?v=-eGTuNfc_kk
Title: Re: Rocket Engine Q&A
Post by: Robotbeat on 11/02/2021 08:33 pm
Yeah, took some pictures once of some Mach diamonds from the tip of an air compressor nozzle. Any time you get a sizable pressure drop, you pretty much get Mach diamonds. Unless you’re trying not to.
Title: Re: Rocket Engine Q&A
Post by: JetProp on 11/24/2021 06:50 am
I'm looking for article LEMAITRE, A. ─ MARCIQUET, C.: Propellant Electric Pump for low thrust cryogenic propulsive systems. 4th European Conference for Aerospace Sciences (EUCASS), Saint Petersburg, Russia, 2011
Sometime the article was available on the site: http://eucass2011.conferencecenter.ru/cs/upload/gF76bMq/papers/papers/978-1505-1-RV.pdf
but this site is dead.
Please, help find this article...

Sorry my bad English.
Title: Re: Rocket Engine Q&A
Post by: gin455res on 11/27/2021 12:09 am
If one had a 1.3m diameter rocket nozzle and combustion chamber but could not 3d print it because it was too big, but could 3d print 4-5  60-65cm diameter nozzles and combustion chambers and connect them together as a multi-chambered engine; which engine would have fewer parts, and which would be cheaper to build?
 
i.e. Does 3d-printing favour/afford multiple chamber rocket engines?
Title: Re: Rocket Engine Q&A
Post by: Proponent on 11/27/2021 12:54 pm
Though I can't say much about 3-D printing, it's certainly possible to build multi-chamber engines.  That's what the famed RD-107/108 is, not to mention the RD-180, which, in turn, is just half of an RD-170.

To a first approximation, a nozzle is a pressure vessel, so its mass scales as the product of pressure and volume.  For a constant nozzle shape and chamber pressure, then, the thrust-to-weight ratio would actually tend to improve with increasing numbers of smaller nozzles.  And the nozzles will be shorter than a single monolithic nozzle.

On the other hand, while the "good" things about rocket plumbing -- flow rates and cooling requirements -- tend to improve with increasing cross-sectional area, the "bad" things -- viscous losses and leakage -- tend to scale with diameter, i.e., more slowly with increasing size.