Analysts of various SRB CEV launcher configurations

[dwightlooi]

The following is a tabulated analyst of the payload capabilities of the various SRB launcher configuration that is likely to be pursued by NASA.<br /><br />If I have to take a guess, I'll say that NASA will go with a 4-segment SRB with a J2S or RLX powered upper stage in the 100 ton class. Using the SSME does not yield much payoff unless the upper stage is in the 200 ton class. The problem with that is that a 200 ton upper stage will be just about as big and as long as a Delta IV booster core ( which is 229 tons). This will yield a VERY long rocket -- imagine a Delta IV stacked on top of the SRB and you get the picture.<br /><br />I hope they go with the RLX. The future of the space program can really use a 300,000 lbs expander cycle hydrogen engine. And, unlike pressing the J2S into service we will actually be moving forward. An expander engine is inherently safe and receptive to air start and multiple restarts.<br /><br />http://img301.imageshack.us/img301/859/srbsevchart5jq.gif<br /><br />In anycase, this is the table:-

 

[dwightlooi]

The chart is now visible

 

[tap_sa]

Nice work! Did you choose the gross deltaV to be 9000m/s and work other parameters from that? It's a bit on the low side.

 

[dwightlooi]

The idea is to achieve 9000 m/s worth of gross delta v. I think that it is reasonable estimate for netting 7800 m/s of net delta v. Overall I think it is a good estimate because the result for the 4-segment SRB + J2S configuration is consistent with ATK Thiokol's claims.<br /><br />The loaded and empty weight for the SRB is from actual statistics for the shuttle application.<br /><br />Loaded upper stage mass is set 100, 200 or 300 tons for comparison purposes. The empty mass is a structural mass of 0.0892 of the loaded mass (the same as a delta IV booster core's), plus engine mass from available stats.<br /><br />delta v for each stage is approximated using the rough formula of 9.81 x IpSec x LN(initial mass/final mass)<br /><br />It is possible that changing the expansion ratio of the nozzle will improve performance. But I don't think it will affect it a great deal. The J2S is an upper stage engine to start with. The SSME is really designed as more like an upperstage engine than a sea level one -- it is SERIOUSLY over expanded at lift off. Besides IpSec in vaccuum doesn't get that much higher than 453 secs anyway (about 470 is the maximum I think one can expect) and a nozzle redesign is not a trivia, especially for a regeneratively cooled engine like the SSME. The RS-68 will probably benefit from a better nozzle, but it's biggest flaw is the hefty mass and a larger nozzle is going to make it even heavier. IpSec for a gas generator, ablative cooled engine like the RS-68 is unlikely to get beyond the 430s anyway regardless of the kind of nozzle you bolt on.<br />

 

[spacester]

That is a beeee-oooot-iful chart. Exciting times we live in, don't you think?<br /><br />IMO 7800 m/s is not enough orbital energy. The 1.2 km/s total grav and drag losses seems quite reasonable according to my research. Sorry, but I feel like posting some equations tonight.<br /><br />Using stuff from a spreadsheet I made 5 years ago . . .<br /><br />E = Orbital Energy = GMm (1/(Re) - 1/(2*Rc))<br />G = grav constant = 6.673e-11 N * m^2 / kg^2<br />M = mass of Earth = 5.975E+24 kg<br />m = spacecraft mass <== ignore by using specific energy = E / m<br />Re = Radius of Earth = 6377.83 km = 6377830 m<br />Rc = Equivalent Circular Radius of orbit = semi-major axis = Re + equivalent orbital altitude<br />and<br /><br />dv = sqrt(2 * E / m)<br />dv = Net velocity imparted to spacecraft after all corrections<br /><br />An approx correction formula for latitude bonus is<br /><br />Vr = w * Re * cos("some Angle")<br />But if due east lauch it becomes exact<br />Vr = w * Re * cos(latitude)<br />w = 7.2921152e-05 rad/sec (Earth’s eastward rotation rate)<br />KSC is at 28.465 deg<br />Vr = 0.41 km/s<br /><br />No elevation bonus at KSC. The formula is<br />v = sqrt((GM)*((1 / Re) - (1 / (Re + elevation))))<br /><br />So corrections would amount to something close to <br />+ 0.41 Latitude bonus<br />- 1.2 Gravity losses<br />- 0.41 Drag losses<br />___<br />-1.2 km/s total corrections (pretty conservative, we prolly have some margin here for later)<br /><br />The numbers say that the orbital radius is way too low for an orbit above the atmosphere with only 7.8 km/s , in fact, at Rc = Re , zero orbital altitude, I get 7.91 km/s<br /><br />If we wanted to use 8.8 km/s - which puts us at a nice even 10.0 km/s with corrections - we'd have a way too high orbit of 2000 km<br /><br />But at 8.2 km/s, we get equivalent orbital altitude = 525 km - or 8.15 km/s at 425 km<br /><br />So I would LOVE to see that chart with gross delta V of 9.35 km/s, net delta V of 8.15 km/s or something like that <br /><br />You explained your <div class="Discussion_UserSignature"> </div>

 

[dwightlooi]

<i>If we wanted to use 8.8 km/s - which puts us at a nice even 10.0 km/s with corrections - we'd have a way too high orbit of 2000 km<br /><br />But at 8.2 km/s, we get equivalent orbital altitude = 525 km - or 8.15 km/s at 425 km<br /><br />So I would LOVE to see that chart with gross delta V of 9.35 km/s, net delta V of 8.15 km/s or something like that </i><br /> <br />First of all, I wasn't shooting for a 500km orbit. It is common for rocket manufacturers to "cheat" and state their payload ratings as what the rocket will lift to a 185km orbit at a low inclination (20-something degrees being common). Practically nothing goes to a 185 km orbit -- not space stations, not satelites, just about nothing -- but it is the low ball altitude that is recognized as LEO. Since a lot of publically available figures for payload capacity of various launchers use the 185km orbit as the benchmark, I was aiming for to make comparisons to easier.<br /><br />Another thing to consider is that once you are coasting through vacuum and microgravity in a 185km pseudo orbit, changing orbital altitude really doesn't take a lot of fuel or thrust. It is safe to assume that the CEV or CXV will have some kind of orbital maneuver engine (OME) and can effect that altitude change on its own. An AJ10-190 engine with a modest loading of hypergolic fuel will likely be sufficient for this. Whatever the OME configuration, it wouldn't have to lug around 10 to 33 tons worth of 2nd stage dead weight.<br /><br />In anycase, here is the chart done with 9,350 m/s as the target gross delta v.<br /><br />http://img374.imageshack.us/img374/5826/comparison93500ip.gif

 

[spacester]

Cool! Too bad the image is not yet approved, and I gotta hit the sack. no biggie<br /><br />FWIW 185 km orbit is 8.02 km/s orbital energy so 8.02 + 1.2 = 9.22 km/s might be the magic number<br /><br />Or we can assume 8.02 - 9.35 = -1.33 total corrections<br /><br />Sorry for the nit-picking, but I want to use this chart for stuff in the future. Good resource, well done, thanks again! <div class="Discussion_UserSignature"> </div>

 

[dwightlooi]

<i>"It is possible that changing the expansion ratio of the nozzle will improve performance"<br /><br />No it is a fact. The closer the exit pressure is to ambient the better the performance. </i><br /><br />Of course it does. What I mean is that the additional performance may not be worth the extra cost, bulk and weight.<br /><br />Take the RL10 for example. The RL10-A4 weighs 168 kg and gets an IpSec of 449 secs. The RL10-B2 has a much larger nozzle -- the largest drop in place carbon nozzle for that matter -- and attains 465 secs. This also increased the weight of the RL-10B2 to 302kg. A huge chunk of the weight of rocket engines is nozzle weight, and when you go to a huge nozzle this can skyrocket very quickly and you reach a point of diminishing returns quite quckly.<br /><br />IF we increase the IpSec rating of the SSME to 470 secs (up from 453) with everything else being constant payload rating goes up from 19.3 to 21.3 tons. That is a good 2 tons. But how much is the new giant nozzle SSME going to weigh? If weight is gained at the same ratio as we see going from an RL-10A4 to an RL-10B2 then the SSME will weight 2534 kg more (5711 kg), completely negating the payload increase and in fact reducing payload by 500kg. Even if that does not happen... lets say the weight gain is only half that -- 1250 kg. That still leaves a mere 750kg increase in payload; a 750kg increase over 19,300kg. One then has to ask whether that is worth redesigning the SSME, retooling for its manufacturer and recertifiying it for manned flights.

 

[teije]

Hey Dwight!<br />Would it be rude to actually ask you the .xls version of that pic?<br />That would be great. <br />Thx adv!<br />Teije

 

[dwightlooi]

<i>Would it be rude to actually ask you the .xls version of that pic? </i><br /><br />I don't mind, but I can neither upload the xls file to space.com nor host it on an image hosting site. So... I'll send it to you as a file attachment if you'll give me your email address.

 

[gunsandrockets]

"The following is a tabulated analyst of the payload capabilities of the various SRB launcher configuration that is likely to be pursued by NASA. "<br /><br />The latest word is the SRB derived CEV launcher will deliver 25mt to orbit (because NASA expects the CEV to mass at least 23mt). Presumably that is a 51 degree inclined orbit, the same orbit as the ISS.<br /><br />I also believe NASA would want to keep the second stage to as close to 100mt as possible, that way the same stage could be used as an upper stage for the SDHLV, and as a moon mission EDS.

 

[dwightlooi]

I tyhink they can do it with a 5-segment SRB, and a 125 ton upper stage mounting a single SSME. that will drive a 25 ton payload to about 9100 m/s (gross) -- not quite enough for a 500km orbit, but the OME on the CEV should be able to take care of the rest.