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SRB CEV Launcher

Posted by: bad_astra - Sun Mar 20, 2005 1:01 am
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SRB CEV Launcher 
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Post SRB CEV Launcher   Posted on: Sun Mar 20, 2005 1:01 am
I haven't seen any detailed information on Thiokol's proposed SRB derived launcher. Is the SRB stage supposed to be recovered as with STS? Seems (to me) that if Thiokol get's shut ouf or any consideration for Project Constellation, they could at the very least build an America's Space Prize launcher as a proof of concept with a fiarly simple upper stage, and they'd be years ahead of any competition. (what a wild ride it would be, though)

If anyone knows about possible recovery, I'd like to know. Thanks.


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Post    Posted on: Sun Mar 20, 2005 11:04 am
I posted this article under the Delta IV for CEV booster thread a while back but havent seen anything since.

http://www.thespacereview.com/article/226/1

It talks about using a H2/LOX upper stage with a J-2 engine (used on Saturn 1B and V), 90,000kg fuel would give a payload of 18,100-22,700kg to LEO.

Although it doesnt specifically say that the SRB will be recovered, it doesnt say that they will not be recovered either. So I guess they will operate like normal SRBs and be recovered from the sea.

I cant see how they could use an SRB without reducing the thrust though as you would subject any CEV mounted on top to quite a few Gs. the article quotes Mach 18 and 20Gs at burn out so this will have to be reduced.

Overall weight would be something like

SRB 590,000kg
upper stage 100,000kg
CEV 20,000kg

Total weight of 710,000kg with a thrust (Vac) of 1,175,000kg.

I got the SRB figures from http://www.astronautix.com/engines/srb.htm.

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Post    Posted on: Sun Mar 20, 2005 11:21 am
Just noticed a link in the above space review article to a report from the planetary society which suggests the use of an SRB, the team who wrote the report were jointly lead by NASA Administrator Elect Michael Griffin.

http://planetary.org/aimformars/study-report.pdf

Makes you think doesnt it? :)

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Post    Posted on: Mon Mar 21, 2005 12:11 pm
Hmm, very interesting study report. I like it alot.

The use of the Shuttle SRB as a launch vehicle is also a very interesting concept. When re-used, it's cost is only $35M (according to the paper I read), which places it a long way below the cost of a Delta-IV or Atlas V. However, looking at the maths, it does have numerous problems also.

I use 35-tonnes, because that is the average weight of a Delta IV or Atlas V upper stage with 12 tonne payload (just approximated).

Code:
Various launch modes

Stage         Gross     Empty    ISP   Thrust       Delta-V
                  
Atlas CCB     306,914   22,461   338   423,386      5914
Delta RS-68   226,400   26,760   420   337,807      5945
Shuttle SRB   589,670   86,183   269   1,174,713    4328


As you can see, the Delta and Atlas are very comparable, whereas the Shuttle SRB is actually much less capable at launching on its own. Also, it's far greater thrust means that the payload would undergo far greater acceleration.

An interesting thing about the SRB though is that the DV it achieves does not particularly depend upon the payload it is carrying (or is affected a lot less than the other two), because its empty mass is so large. As for whether using a very large upper stage would allow larger payloads to be launched, I'm not sure, although it would be a very unusual launch design. I'll try to do some more calculations on it though.

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Post    Posted on: Mon Mar 21, 2005 12:43 pm
Ok, did a little followup considering the upper stages of the rockets, and made a few discoveries. Firstly, the Atlas V and Delta IV are astonishingly similar. I can only suppose that the USAF was very specific in their requirements for the EELVs. Secondly, in order for a SRB to match the launch power of the EELVs, it would essentially have to use a triple upper stage - three times the number of engines, three times the amount of fuel.

Somewhat interestingly, Pratt and Whitney are developing a successor to the RL-10, the RL-60, which offers over double the performance of a single RL-10, which could be quite useful for this application.

The data:
Code:
Stage         Gross     Empty    ISP   Thrust    Payload  Delta-V
                  
Atlas CCB     306,914   22,461   338   423,386   12,000   5,922
Centaur VI    22,825    2,026    451   10,115    12,000   4,024
Atlas V       341,739   36,487                   12,000   9,946
                  
                  
Delta RS-68   226,400   26,760   420   337,807   12,000   5,886
Delta 4-2     24,170    2,850    462   11,222    12,000   4,035
Delta IV      262,570   41,610                   12,000   9,920
                  
                  
Shuttle SRB   58,9670   86,183   269   1,174,713 12,000   3,625
SRB Upper     72,510    8,550    462   33,666    12,000   6,409
SRB CEV       674,180   106,733                  12,000   10,033


The problem with the SRBs is that the upper stage is basically doing all of the work. Especially if you used a J-2 with 200,000 lbs of fuel, you might as well make that the basis of the whole launch vehicle. The actual SRB doesn't help it all that much.

The SRB is never capable of boosting even itself to the velocities of the first stages of the Delta or Atlas rockets (and thats with no payload), which definitely limits its use as a first stage.

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Last edited by Sev on Mon Mar 21, 2005 6:50 pm, edited 1 time in total.



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Post    Posted on: Mon Mar 21, 2005 6:46 pm
I think that a SRB vehicle might be OK to deliver cargo, fuel and food to orbit but it would require some serious modifications to make it suitable for astronauts. 20Gs is far to much, they would need to change the solid propellant so that it burns slower and longer producing less thrust to make it viable for a crew.

Still only having to fuel the upper stage with 90,000kg of LOX/H2 on the launch pad has some advantages for turn around times and speed of launch so if a demo rocket could be produced relatively cheaply then I think it would be worth doing. Perhaps NASA will run a project to build one to see how it performs.

This system could produce substatial savings over an EELV derived vehicle and would utilise some of the existing shuttle infrastructure.

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Post    Posted on: Mon Mar 21, 2005 6:55 pm
Andy Hill wrote:
Still only having to fuel the upper stage with 90,000kg of LOX/H2 on the launch pad has some advantages for turn around times and speed of launch so if a demo rocket could be produced relatively cheaply then I think it would be worth doing. Perhaps NASA will run a project to build one to see how it performs.


Remember that preparing the SRB takes several weeks. True, that can be done on a production line, and then they can be stored. But it's hardly rapid reaction times.

Quote:
This system could produce substatial savings over an EELV derived vehicle and would utilise some of the existing shuttle infrastructure.


I'm really not sure about this. Some savings maybe, but I would be careful about this. The cost for the SRBs on the Shuttle, each flight, is $35M each. Now, that is over 50% more expensive than the engines used in either the Delta or Atlas rockets - and I'm not sure what the cost of their first stages is (in terms of recurring costs), but I imagine it would be very close. The only real advantage is if the core had already been made, it could be fired on relatively short notice. But that seems more like a military advantage than a commercial one :P

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Post    Posted on: Mon Mar 21, 2005 7:09 pm
Since the SRBs are made up of segments, it seems to me they could limit peak acceleration by using fewer SRB segments and making the upper stage fuel tanks larger.


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Post    Posted on: Mon Mar 21, 2005 7:32 pm
campbelp2002 wrote:
Since the SRBs are made up of segments, it seems to me they could limit peak acceleration by using fewer SRB segments and making the upper stage fuel tanks larger.


Throw enough segments away and you have a SSTO vehicle. :)

I think that the SRB should be kept as close as possible to the original design, with the exception of the fuel grain used, to minimise development costs. This should only be attempted if there is a real advantage in either time or costs.

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Post    Posted on: Mon Mar 21, 2005 7:50 pm
I don't think the SRB's could be SSTO with any number of segments. The case is too heavy and the exhaust velocity is too low.

As for minimum change from existing configuration, I would think that a new propellant grain would be a major change but simply stacking fewer segments in the VAB would be a minor change.


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Post    Posted on: Mon Mar 21, 2005 9:25 pm
Ok lets see what we come up with by removing a segment using the numbers I posted earlier.

Assuming that each segment of the booster weighs about 100,000kgs (5 segment SRB 590,000kg) and that the thrust is reduced by a fifth to 940,000kg we end up with a ship weighing about 610,000kg.

Now to put the same payload into orbit we would need to increase the fuel in the second stage lets say by 30% (Big Guess) that gives a craft weighing about 640,000kg.

This would be better for the crew who would be subjected to less Gs but mean that you were designing a bigger upper stage to throw away with each launch while the cost of recovering/refurbishing the SRB will probably not change that much. Begins to get as expensive as using an EELV and probably better to use a conventional booster with strap ons instead.

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Post    Posted on: Mon Mar 21, 2005 9:35 pm
campbelp2002 wrote:
I don't think the SRB's could be SSTO with any number of segments. The case is too heavy and the exhaust velocity is too low.

As for minimum change from existing configuration, I would think that a new propellant grain would be a major change but simply stacking fewer segments in the VAB would be a minor change.


It's a calculable fact - the maximum velocity for a single SRB, with no payload attached, is around 5000 ms-1. It's this very low value that makes it, in my opinion, unrealistic for application outside of simply being a booster. The velocity it would achieve as a first stage is too low for anything short of a three stage to orbit vehicle - making a two-stage to orbit vehicle based on an SRB would quite simply be vastly top-heavy. Due to the large amounts of second-stage fuel required, combined with the high power-density of solid rockets, the second stage would quite literally dwarf the SRB. The little diagram that's been passed around of it (http://www.thespacereview.com/archive/226a.jpg) is completely unrealistic.

There is a good reason that every solid-rocket launch vehicle made thus far has been at least three or four stages, which is simply that the casing to contain solid rocket motors is far too heavy.

There's also a far more down to Earth reason why the SRBs could not be used "as-is" for orbital launches - they were never designed to take the kind of axial loadings which they would undergo through launch. Having to transfer a force of up to 10,000,000 newtons (a 100,000 kg upper stage at 10g's) through it's length is way beyond it's initial design specifications, and to re-manufacture it to take that kind of loading would vastly increase the weight of the casing (which is already too heavy).

The SRB is a brilliant booster. It is not a particularly good first stage.

A much more interesting concept is what kind of benefits would there be from an SRB-assisted EELV. I did some more multi-stage calcs, and a dual SRB-boosted two stage Zenit rocket could lift over 50 tonnes into orbit - not bad from the 20 tonnes it starts at without the booster. (I'm aware the Zenit is not an EELV, but it made a more interesting calculation)

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Post    Posted on: Wed Mar 23, 2005 10:32 pm
Andy Hill wrote:
the article quotes Mach 18 and 20Gs at burn out so this will have to be reduced.
I think that 20Gs was for an empty SRB. Put the 90,000 kg upper stage and the 20,000 kg CEV on top and the burnout acceleration is less than 8Gs I think. Still high compared to the shuttle/SSO 3Gs, but better.


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Post    Posted on: Thu Mar 24, 2005 12:58 pm
And still vastly more than the structure of the current SRBs can take :P

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Post    Posted on: Thu Mar 24, 2005 1:05 pm
And also still vastly more than any sane human being would want to endure for any perceivable length of time.

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