Author Topic: Basic Rocket Science Q & A  (Read 271439 times)

Offline Antares

  • ABO^2
  • Senior Member
  • *****
  • Posts: 5191
  • Done arguing with amateurs
  • Liked: 342
  • Likes Given: 210
Re: Basic Rocket Science Q & A
« Reply #600 on: 05/13/2011 03:16 PM »
Sorry, I forgot to add the word "nozzle" in there, so the nozzle is in the lower stage prop tank.

Its quite common.

Name one.
If I like something on NSF, it's probably because I know it to be accurate.  Every once in a while, it's just something I agree with.  Facts generally receive the former.

Offline douglas100

  • Senior Member
  • *****
  • Posts: 2150
  • Liked: 214
  • Likes Given: 95
Re: Basic Rocket Science Q & A
« Reply #601 on: 05/13/2011 04:16 PM »
Shtil'
Douglas Clark

Offline Danderman

  • Extreme Veteran
  • Senior Member
  • *****
  • Posts: 9565
  • Liked: 345
  • Likes Given: 453
Re: Basic Rocket Science Q & A
« Reply #602 on: 05/13/2011 05:03 PM »
Dnepr either has the 2nd stage nozzle in the lower stage prop tank, or the entire Dnepr 2nd stage engine is in the 2nd stage prop tank.

« Last Edit: 05/13/2011 05:04 PM by Danderman »

Offline LegendCJS

  • Full Member
  • ****
  • Posts: 575
  • Boston, MA
  • Liked: 7
  • Likes Given: 2
Re: Basic Rocket Science Q & A
« Reply #603 on: 05/13/2011 06:06 PM »
For rockets where the 2nd stage engine is housed inside the 1st stage prop tanks, how exactly does the engine leave the tank during staging? Is the prop tank blown up?

Do you know of a rocket where the second stage engine is immersed in the fuel of the first stage tanks before staging?  Don't you mean inter-stage volume?

Sorry, I forgot to add the word "nozzle" in there, so the nozzle is in the lower stage prop tank.

Its quite common.


So the nozzle of the second stage forms part of the pressure structure of the first stage? I don't think I've ever seen such a design...

No, the nozzle extends into the first stage prop tank, saving loads of interstage area.

The second stage nozzle is immersed in the fluid propellent of the upper first stage tank?  Let me see a diagram of that!  Some part of the nozzle throat or combustion chamber or engine would have to be part of the pressure containing boundary of the first stage tank in that case.

Or do you mean that the first stage tank's top is not a dome, but contains a concave region sized for the nozzle of the second stage engine, and there is no fluid communication between the second stage nozzle and the first stage propellant?

A simple diagram (even in paint) would save a lot of confusion for me and other people here it seems.
Remember: if we want this whole space thing to work out we have to optimize for cost!

Offline Danderman

  • Extreme Veteran
  • Senior Member
  • *****
  • Posts: 9565
  • Liked: 345
  • Likes Given: 453
Re: Basic Rocket Science Q & A
« Reply #604 on: 05/13/2011 06:26 PM »


The second stage nozzle is immersed in the fluid propellent of the upper first stage tank?  Let me see a diagram of that!  Some part of the nozzle throat or combustion chamber or engine would have to be part of the pressure containing boundary of the first stage tank in that case.

Or do you mean that the first stage tank's top is not a dome, but contains a concave region sized for the nozzle of the second stage engine, and there is no fluid communication between the second stage nozzle and the first stage propellant?

A simple diagram (even in paint) would save a lot of confusion for me and other people here it seems.

Its not so easy to show in a diagram, but here is a try, a Rokot diagram.

Offline Antares

  • ABO^2
  • Senior Member
  • *****
  • Posts: 5191
  • Done arguing with amateurs
  • Liked: 342
  • Likes Given: 210
Re: Basic Rocket Science Q & A
« Reply #605 on: 05/13/2011 11:42 PM »
.This is one of the coolest things I've learned in the last few years:

Why did they do that? The nuclear arms treaties limited the length of the silos.  So having the upper tank extend up around the second stage nozzle got them a teeny bit more kilotons or a few more miles of range.

There are launch films where the missile is ejected from the silo and THEN the fairing springs up around the warhead, just so they could fit 3 more feet of rocket into the silo.

Brilliant engineering. A very interesting choice in prioritization of design requirements.
If I like something on NSF, it's probably because I know it to be accurate.  Every once in a while, it's just something I agree with.  Facts generally receive the former.

Offline Antares

  • ABO^2
  • Senior Member
  • *****
  • Posts: 5191
  • Done arguing with amateurs
  • Liked: 342
  • Likes Given: 210
Re: Basic Rocket Science Q & A
« Reply #606 on: 05/13/2011 11:43 PM »
For the rest of us, the build and operational complexity of that doesn't make sense for a commercially viable rocket. Just sacrifice some performance for simplicity.
If I like something on NSF, it's probably because I know it to be accurate.  Every once in a while, it's just something I agree with.  Facts generally receive the former.

Offline douglas100

  • Senior Member
  • *****
  • Posts: 2150
  • Liked: 214
  • Likes Given: 95
Re: Basic Rocket Science Q & A
« Reply #607 on: 05/14/2011 11:10 AM »
From Antares:

Quote
Why did they do that? The nuclear arms treaties limited the length of the silos.  So having the upper tank extend up around the second stage nozzle got them a teeny bit more kilotons or a few more miles of range.

And of course the same technical argument applies to Shtil' and other submarine launched liquid propellant missiles.
Douglas Clark

Offline Danderman

  • Extreme Veteran
  • Senior Member
  • *****
  • Posts: 9565
  • Liked: 345
  • Likes Given: 453
Re: Basic Rocket Science Q & A
« Reply #608 on: 05/14/2011 03:22 PM »
For the rest of us, the build and operational complexity of that doesn't make sense for a commercially viable rocket. Just sacrifice some performance for simplicity.

Unless the rocket is being carried inside an aircraft.

Online kevin-rf

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8506
  • Overlooking the path Mary's little Lamb took..
  • Liked: 1022
  • Likes Given: 234
Re: Basic Rocket Science Q & A
« Reply #609 on: 05/16/2011 01:54 PM »

For the rest of us, the build and operational complexity of that doesn't make sense for a commercially viable rocket. Just sacrifice some performance for simplicity.

Does it?

1. You eliminate the weight of the interstage.
2. The pressure in the empty tank when severed from the upper stage provides a separation impulse, eliminating the need and weight of ullage motors and separation pneumatics.
3. (Assuming LOX) Keeping the upper engine immersed in LOX tank will precondition the thermal environment.

It has some merit, If you look at Airlaunch LLC's design you'll notice that the Second stage nozzle (Vac Opt I'm sure) is almost as long as the first stage LOX tank. That is a fair amount of interstage kit you are saving, plus you don't need a stretched C-17 ;)

Of course I can think of more downsides than the three upsides I just mentioned...
If you're happy and you know it,
It's your med's!

Offline baldusi

  • Senior Member
  • *****
  • Posts: 7436
  • Buenos Aires, Argentina
  • Liked: 1434
  • Likes Given: 4475
Re: Basic Rocket Science Q & A
« Reply #610 on: 06/10/2011 10:53 PM »
D. L. Jensen states in this paper that most current US H2 engines are develop with the "wrong", mixture ratios. Given that H2 needs such a huge volumes, it would only seem natural to try to use as little as possible, specially if you get better isp.
But since I know that there's a lot of intelligent people behind current engines, and I've learned better than to blindly trust random papers from the internet, specially when written as angry as this one, I would like to know what he's not saying. In particular, it doesn't makes any sense to me dual clone turbopumps against low pressure and high pressure.
From what I could grasp, the higher the mixture ration, the higher the pressure and temperature that the engine has to handle. And so it becomes more expensive and heavy. A reduction in the proportion of H2 would lower the tank's sizes, so it should offset a bit of the extra weight. And the reduced use of H2 should reduce the turbopump sizes. But there's so much that I guess I'm missing that I would love someone knowledgeable to tell me.

Offline DMeader

  • Full Member
  • ****
  • Posts: 953
  • Liked: 98
  • Likes Given: 47
Re: Basic Rocket Science Q & A
« Reply #611 on: 06/10/2011 11:14 PM »
I'm not a rocket scientist myself (and I didn't stay at a Holiday Inn Express last night either) but even so some things jump out at me in Mr. Jensen's paper. Pardon me if I make some elementary errors.

Regarding his turbopump arrangement. I was under the impression that the SSME used the low pressure fuel and oxidizer turbopumps to bring the propellants to sufficient pressure to eliminate cavitation in the high-pressure turbopumps. Otherwise the external tank would have to supply propellants at much higher pressure itself and would need to be much stronger and heavier.

Regarding his contentions on "wasteful" mixture ratios. Doesn't H2 beyond stoichiometry contribute to thrust thru greater mass flow through the engine, the extra hydrogen being heated and expanded by combustion? Also, isn't a fuel-rich ratio desirable to protect the engine hardware?

I'm amused by people like Mr. Jensen who come out and say that the people that do this for a living (not me of course, I'm just an interested amateur) are obviously wrong about something so fundamental.

Offline Proponent

  • Senior Member
  • *****
  • Posts: 5111
  • Liked: 770
  • Likes Given: 528
Re: Basic Rocket Science Q & A
« Reply #612 on: 06/11/2011 04:13 AM »
Doesn't H2 beyond stoichiometry contribute to thrust thru greater mass flow through the engine, the extra hydrogen being heated and expanded by combustion?

Excess hydrogen improves specific impulse but not by providing additional mass flow.  Rather it increases specific impulse by raising  the efficiency with which the nozzle can convert heat into kinetic  energy.  That in turn is because hydrogen molecules, being simpler than  water molecules, have fewer ways of soaking up energy, so a larger  fraction of the total energy can go into the desired form, kinetic.

Adding mass can improve propulsive efficiency if the added mass need not be carried by the vehicle -- that's a part of the reason that jet engines have much higher specific impulses than rocket engines.
« Last Edit: 06/12/2011 08:43 AM by Proponent »

Offline baldusi

  • Senior Member
  • *****
  • Posts: 7436
  • Buenos Aires, Argentina
  • Liked: 1434
  • Likes Given: 4475
Re: Basic Rocket Science Q & A
« Reply #613 on: 06/11/2011 05:49 PM »
I understand that running the engine fuel rich gives better isp in the H2 case, than a lean mixture. I could also understand that it could get better isp than a stoichiometric ratio, but with an incomplete mixture (i.e. if the engine is run below its optimum pressure and temperature). The graph that he show, is suggestive.
« Last Edit: 06/11/2011 05:53 PM by baldusi »

Offline TyMoore

  • Veteran
  • Full Member
  • ****
  • Posts: 482
  • Eureka, CA, USA
  • Liked: 2
  • Likes Given: 1
Re: Basic Rocket Science Q & A
« Reply #614 on: 06/11/2011 05:54 PM »
D. L. Jensen states in this paper that most current US H2 engines are develop with the "wrong", mixture ratios. Given that H2 needs such a huge volumes, it would only seem natural to try to use as little as possible, specially if you get better isp.
But since I know that there's a lot of intelligent people behind current engines, and I've learned better than to blindly trust random papers from the internet, specially when written as angry as this one, I would like to know what he's not saying. In particular, it doesn't makes any sense to me dual clone turbopumps against low pressure and high pressure.
From what I could grasp, the higher the mixture ration, the higher the pressure and temperature that the engine has to handle. And so it becomes more expensive and heavy. A reduction in the proportion of H2 would lower the tank's sizes, so it should offset a bit of the extra weight. And the reduced use of H2 should reduce the turbopump sizes. But there's so much that I guess I'm missing that I would love someone knowledgeable to tell me.

I read the paper, and one thing that I saw that was wrong right off the bat, was that he was using 43,000 BTU/lb as the enthalpy of combustion for hydrogen. This is wrong: the lower heat of combustion for hydrogen is about 51,000 Btu per pound (non condensing.) [61,000 BTU/lb for higher heat of combustion (condensing)]

The performance of a rocket engine is not just Isp; it is also thrust. You have to weigh both and optimize the engine calculations to perform the mission. This is why there are different mixture ratios. And burning slightly hydrogen rich reduces the average molecular weight of the exhaust, which improves Isp but reduces the overall thrust slightly. Some nuclear thermal rocket engines would actually use a liquid oxygen injection system to boost total thrust by combusting some of the hot hydrogen exiting the nozzle throat. This improves thrust, but reduces overall Isp slightly.

A rocket engine is more than the sum of its "Thrust and Isp." It has to achieve performance goals necessary to achieve the mission.

Offline Gregori

  • Full Member
  • *
  • Posts: 196
  • Liked: 0
  • Likes Given: 0
Re: Basic Rocket Science Q & A
« Reply #615 on: 06/24/2011 03:52 AM »
I've some basic questions i wanted to know about Mars/Moon missions using chemical propulsion.

1) Does chemical propulsion have an inherent speed limited that prevents it from doing a much faster transit to Mars? Is it that you need a large amount of propellant to slow down and circularize the orbit? Could a huge craft with multiple stages assembled in space speed up the mission?

I've heard of NTRs being able to allow faster transits, but I was never sure why they allowed this. I also heard that New Horizons reached the orbit of Mars in two months with chemical. Would this speed be just too fast to be captured by mars without melting the craft or a crash landing?

2) Does the gravity when closely approaching Mars or the Moon help slow down craft significantly and capture them or does this require lots of propellant and using the atmosphere to slow the craft?


Offline Proponent

  • Senior Member
  • *****
  • Posts: 5111
  • Liked: 770
  • Likes Given: 528
Re: Basic Rocket Science Q & A
« Reply #616 on: 06/24/2011 05:20 AM »
1) Does chemical propulsion have an inherent speed limited that prevents it from doing a much faster transit to Mars? Is it that you need a large amount of propellant to slow down and circularize the orbit? Could a huge craft with multiple stages assembled in space speed up the mission?

In principal there's no reason that fast transits can't be done with chemical propulsion, it's just a matter of cost.  Multiple in-space stages could be used with chemical or nuclear rockets.  The Daedalus stellar probe design from the 1970s, for example, calls for a two-stage nuclear-fusion rocket.

Quote
I've heard of NTRs being able to allow faster transits, but I was never sure why they allowed this. I also heard that New Horizons reached the orbit of Mars in two months with chemical. Would this speed be just too fast to be captured by mars without melting the craft or a crash landing?

Yeah, to fast to capture.  Could have been slowed down with rockets -- chemical, nuclear, or otherwise -- but now were talking about hauling lots and lots of propellant all the way out to Mars, so things get really expensive.  Aerocapture could help.

Quote
2) Does the gravity when closely approaching Mars or the Moon help slow down craft significantly and capture them or does this require lots of propellant and using the atmosphere to slow the craft?

Basically, if the spacecraft arrives at the planet above the escape velocity, then without braking of some sort it will not be captured.  Essentially, as seen from the planet, it will speed up as it approaches the planet and then will slow back down again after it swings past.  So, no, the gravity of the planet alone will not permit capture.
« Last Edit: 06/24/2011 05:22 AM by Proponent »

Offline hop

  • Senior Member
  • *****
  • Posts: 3178
  • Liked: 357
  • Likes Given: 693
Re: Basic Rocket Science Q & A
« Reply #617 on: 06/24/2011 05:26 AM »
1) Does chemical propulsion have an inherent speed limited that prevents it from doing a much faster transit to Mars?
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation

Equally valid for nuclear thermal, ion, chemical, throwing rocks off the back...

Offline IsaacKuo

  • Full Member
  • ****
  • Posts: 435
  • Liked: 1
  • Likes Given: 1
Re: Basic Rocket Science Q & A
« Reply #618 on: 06/24/2011 07:24 AM »
I also heard that New Horizons reached the orbit of Mars in two months with chemical. Would this speed be just too fast to be captured by mars without melting the craft or a crash landing?

Mars would not "capture" a space probe on its own regardless of speed.  The space probe would enter at greater than escape speed and leave at the same speed, relative to Mars, on a hyperbolic trajectory.  The probe might use a rocket thrust or the as yet untried technique of aerocapture to slow down into Mars orbit, but Mars's gravity isn't going to do the job by itself.

Alternatively, the space probe could simply directly land on Mars after atmospheric entry to slow it down.  This requires a heavy heat shield and it requires the probe to be designed to survive high accelerations.  But this atmospheric entry would be a cakewalk compared to the Galileo atmospheric entry probe.  That probe slammed into Jupiter's atmosphere at a blistering 48km/s and survived 230 gees of deceleration!!!

Quote
2) Does the gravity when closely approaching Mars or the Moon help slow down craft significantly and capture them or does this require lots of propellant and using the atmosphere to slow the craft?

The gravity actually accelerates the probes, but ironically they also help thanks to the Oberth effect and the fact that you only need to slow down to orbital speed (rather than zero speed).  In fact, deeper stronger gravity wells are better.  The Moon's gravity is almost more of a hindrance than a help, but Mars's gravity well is certainly helpful.

Offline strangequark

  • Full Member
  • ****
  • Posts: 1070
  • Co-Founder, Tesseract Space
  • Mojave, CA
  • Liked: 211
  • Likes Given: 11
Re: Basic Rocket Science Q & A
« Reply #619 on: 06/24/2011 09:21 AM »
D. L. Jensen states in this paper that most current US H2 engines are develop with the "wrong", mixture ratios. Given that H2 needs such a huge volumes, it would only seem natural to try to use as little as possible, specially if you get better isp.
But since I know that there's a lot of intelligent people behind current engines, and I've learned better than to blindly trust random papers from the internet, specially when written as angry as this one, I would like to know what he's not saying. In particular, it doesn't makes any sense to me dual clone turbopumps against low pressure and high pressure.
From what I could grasp, the higher the mixture ration, the higher the pressure and temperature that the engine has to handle. And so it becomes more expensive and heavy. A reduction in the proportion of H2 would lower the tank's sizes, so it should offset a bit of the extra weight. And the reduced use of H2 should reduce the turbopump sizes. But there's so much that I guess I'm missing that I would love someone knowledgeable to tell me.

I have a case of insomnia (my wife goes in for surgery tomorrow, and I am 2400 miles away), so I felt like tackling this.

D.L. Jensen, P.E.'s problem is that he doesn't know what stagnation enthalpy is. I have attached a PDF I worked up with the correct way to get exit velocity from chamber enthalpy (hm in his paper).

Basically though, he gives you an equation for specific impulse (exit velocity, but basically the same thing):

Ve2=2Jghm

It should read:

Ve2=2Jg(hm-he)

Where he is the energy you couldn't recover from the flow. This value is smaller for a fuel-rich hydrogen engine.

This is a derivation that an engineering sophomore should understand, someone with a P.E. (that's "Professional Engineer") after their name has no excuse, particularly when it is in every book that covers the basics of rocket engines. He especially had no right to be a sanctimonious ass and insult an entire industry of people with a vastly better grip on the subject.

The guy also gets some basic numbers wrong. His chart showing a monotonic increase of Isp with O/F ratio has RS-68 as O/F=2.5. It is actually at 6. The RL-10 is 5.88, with Isp of 465. Plot those correctly, and you just have a scattered cluster between 5 and 6, with no correlation.

Also, Proponent, I realize you're dead right from our earlier discussion. Redoing this derivation made it painfully obvious. Static enthalpy goes to zero for an infinite nozzle (because temp goes to zero), so you obviously want to just maximize chamber temp in that case. That's basically what this guy did is assumed an infinite expansion ratio, and then somehow figured that the same logic would hold for a real nozzle.

EDIT: I should note that she is going in for gallbladder removal, which is very minor as surgeries go, before anyone gets too concerned.
« Last Edit: 06/24/2011 08:03 PM by strangequark »
Don't flippantly discount the old rules of this industry. Behind each one lies a painful lesson learned from broken, twisted hardware. Learn those lessons, and respect the knowledge gained from them. Only then, see if you can write new rules that will meet those challenges.