Actually, the image I used was that of the latest Shuttle-C concept, presented at the Augustine Commission. As shown here.
ESA ATL Plausibility Checks and Development
- Thread starter Bahamut-255
- Start date
Actually, the image I used was that of the latest Shuttle-C concept, presented at the Augustine Commission. As shown here.
i have seen that one before, i just never heard it called Shuttle-C before.
i wish i could find more pic of it. this is all i have of it.i have seen that one before, i just never heard it called Shuttle-C before.
Well given it's time of development - about mid 1980s - it will likely be known as Shuttle-C. And it's purpose for existence has been explained in earlier posts. And will likely be seeing some other uses as time goes on.
kool, i will be looking forward to the updates.
~just found this pic here, feel free to use.~
keep up the good work
Mars Sample Return Mission - Part II
Obviously, since MSR missions won't happen with great frequency, it will, IMHO, be essential to select the best possible site for landing such a probe as so to secure the highest possible return value from such a mission.
THis would involve a number of precurser orbiters and landers to perform the task of mapping and assessing the Martian Landscape before an MSR mission is conducted. Scientific returns on these precursers would obviously happen as so to justify the expense. That is, the climate, chemistry and geology of Mars would be studied in detail with which a selection of candidate sites would be selected for the MSR mission.
For timescale, Orbiters and Landers can occur during the mid-late '90s with a subsequent MSR mission in the mid-late '00s. Seems more than reasonable to me and perfectly doable.
The question is LV selection. For the smaller orbiters and landers, the Delta 2 series could cut it, as well as the smaller Argo variants, with the larger Argo-HU, or possibly a US EELV for the MSR mission - with the Argo LV in use by ESA, it seems to make sense that NASA would want to replicate its capabilities for 'staying ahead of the competition' purposes, at least to me it does.
But insofar as Mars Exploration is concerned, Manned Landings are extremely unlikely to occur before the 2020s ITTL, so won't really be covered except in design and policy form.
Obviously, since MSR missions won't happen with great frequency, it will, IMHO, be essential to select the best possible site for landing such a probe as so to secure the highest possible return value from such a mission.
THis would involve a number of precurser orbiters and landers to perform the task of mapping and assessing the Martian Landscape before an MSR mission is conducted. Scientific returns on these precursers would obviously happen as so to justify the expense. That is, the climate, chemistry and geology of Mars would be studied in detail with which a selection of candidate sites would be selected for the MSR mission.
For timescale, Orbiters and Landers can occur during the mid-late '90s with a subsequent MSR mission in the mid-late '00s. Seems more than reasonable to me and perfectly doable.
The question is LV selection. For the smaller orbiters and landers, the Delta 2 series could cut it, as well as the smaller Argo variants, with the larger Argo-HU, or possibly a US EELV for the MSR mission - with the Argo LV in use by ESA, it seems to make sense that NASA would want to replicate its capabilities for 'staying ahead of the competition' purposes, at least to me it does.
But insofar as Mars Exploration is concerned, Manned Landings are extremely unlikely to occur before the 2020s ITTL, so won't really be covered except in design and policy form.
This seems like a good mission plan and timescale. I wonder if they might go for the full on MESUR here instead of just Pathfinder. It was supposed to be the "pathfinder" for several additional probes basically identical to it--don't recall, but it was like four to ten. Here, those later probes could follow in groups on some of your larger launchers, carried together through TMI and possibly all the way through cruise before seperating and performing minor course correction burns to target different landing sites.
On the one hand...they have. They have Shuttle-C. There's less political need to have a duplicate of Argo in service if they have one in service that's superior--since ESA may be riding along on a few Shuttle-Cs for station assembly, there's not too huge a loss of prestige (not to mention significant savings) in doing the reverse when it's justified. Spend the money instead on a big Centaur-ish upper stage for Shuttle-C, and the international partnership will have much better LEO and BEO capabilities overall than if NASA tries to duplicate ESA's rocket.
All in all seems reasonable. To be clear--the truss on this extends perpedicular to the axis as with ISS, not the less useful spine-like truss you were talking about on the earlier frikkin huge design?
Particularly with significant ESA involvement (and thus large components coming from Korou), I'd say there's a persuasive argument for a 28.5 degree orbit. It's equally inconvenient for Russia and Europe, and frankly Russia doesn't have an LV to really contribute on the scale of this station (unless they save Energia)--and with ESA as a firm partner, I think there's less justification to jump through hoops to cater to Russia.
Sounds reasonable, though of course if really needed it might be able to get more--ISS is as I understand now good for another 10 years of life on top of its current decade and more of service.
Yes.
I'll see what I can do as far as some 3D renders, though a better idea of trunk layout, launch configuration, and power requirements (compared to ISS as a baseline, perhaps) would be useful and I can't promise fast action.
So, I realize I've been sort of spamming this thread with a backlog of responses, but I saw something today that I thought would fit well here in regards to Shuttle-C. Hence, without further ado:
~~here is a pic i think is relevant to this tl~~
In seriousness, I'm attaching this to emphasize the major advantage of a sidemount SDHLV like Shuttle-C: enormous freaking fairing volume. That image is concept art of the Contellation Altair lander/LSAM being prepped in a 10m fairing. The fairing on your SD-HLV has this diameter, but is a good 3x the length. The scale of the thing is just staggering.
~~here is a pic i think is relevant to this tl
~~
In seriousness, I'm attaching this to emphasize the major advantage of a sidemount SDHLV like Shuttle-C: enormous freaking fairing volume. That image is concept art of the Contellation Altair lander/LSAM being prepped in a 10m fairing. The fairing on your SD-HLV has this diameter, but is a good 3x the length. The scale of the thing is just staggering.
More or less the preliminary mission plans.
I was thinking about USAF, they tend to use their own LVs for their own payloads - especially since Challenger OTL - so something needs to be thought out for them. I'll sort it out when the time's right.
More or less. Maybe a small addition to increase the attachment area to the centre segment by 2-4x, but no more than that. Otherwise it's like the ISS Truss.
Saving Energia is at best, and extremely slim hope. Even though Buran/Energia was a more capable LV than STS, it came with higher costs - in part due to being all-liquid fueled. If they were able to convince enough of the former Soviet Blocks to set up an ESA-esque Space Agency, they should be able to support the smaller Energia-M. It's payload ability, combined with Zenit and Soyuz should keep them in the game somewhat. Though I may have to change Mir somewhat to make it all work.
So it looks like 28.5 degrees may happen, depending on where ESA is willing to let it's capsules land.
Well by 2012 TTL. that Station is going to be showing it's age and a replacement programme would have been under consideration since before then. But stretching the life by another 4-6 years should be doable.
I'll PM when I get it all sorted out.
I've said it before, but you don't need that any more thrust than 2xRL-10 would deliver. This stage is being used for circularization and orbital maneuvering only, right? Neither of those needs particularly high thrust--worst case for even something like TLI is that the added burn time reduces the advantage you can take of the Oberth efect and you add maybe a hundred m/s of delta-v to the required maneuver--but this is on burns that vary by more than that due to orbital positioning even on mission like Apollo where the hardware was largely similar flight-to-flight. It's not worth the extra cost of 2 more RL-10s or an entire new dev. program like an *RL-60.
Except that there are many different types of RL-10, varying greatly in terms of both Isp and Total Thrust. And TLI is going to feature in the later years.
In a vacumn, 462s Isp trumps 448s Isp anyday. And the 470s Isp of the RL-60 is not to be sneezed at.
I'm not disputing your choice of the RL-10B2 for the engine (or at least an alternate development of similar capabilities, since it post-dates your PoD by about a decade--though I wonder whether that variant would bother with the articulated nozzle--it could just be fitted with a fixed high-expansion nozzle, given both payload volume available and that it will only be used in near-vacuum). The Isp is (as you say) too good to ignore, though I think switching from RL-10B2 to the RL-60 for a gain of 8s isn't worth the dev costs. You could likely squeak the same gains out of the RL-10B-2 and save the cost of entirely new turbo-machinery and all that.
What I am saying that you don't need a cluster of 4 of them for your SDHLV orbital maneuvering stage. Unlike SIVB or the Constellation EDS (or the proposed SLS upper stage), this stage has no need to perform any burn during ascent except perhaps a relatively trivial circularization. Thus, you can get by with a much lower T/W than either of those stages. Deleting two RL-10s could save tens of millions of dollars per flight of this orbital maneuvering stage, without a significant impact on performance.
Last edited:
Perhaps I should elaborate on why I think higher total thrust would be needed for an EDS - an OMS can get away with just one RL-10 so we'll skip that. To get 35,000Kg to TLI with Shuttle-C, the EDS must be large. That is because a 77,000Kg parking orbit payload is simply unable to put a 35,000Kg payload to TLI. I did a quick run of the numbers - 35,000KG payload, 42,000Kg EDS, of which 4,000Kg is dry mass, 457s Isp engine of 40,057Kgf (RD-57) - and I fall at least 350m/s short of the required TLI Delta-V, which, in turn, means a need for a larger EDS, first to complete Parking Orbit Insertion, and then to perform TLI.
I could easily be wrong, but I felt it best to explain exactly where it is I'm coming from.
What delta-v are you using? From this data's list of post-burn velocities, and subtracting the orbital velocities for those altitudes, the delta-v for each flight can be found. The range is 3.09-3.14 km/s (space fixed), with an average of 3.13 km/s and a standard deviation of 17 m/s. With a stage mass of 3,750 kg, an Isp of 464 (RL-10B2), and a gross mass of 77,000 kg, this gives a payload through TLI of exactly 34,956 kg. Will you excuse a 0.12% rounding error? (I justify a stage mass of 3,750 kg as being a round number equal to a 9.8% dry-mass-to-fuel-mass percentage.)
I will agree that were you to be trying to get more payload through TLI, you'd need more stage to carry it. I'm not sure at what point your gravity loss increases on the boost and core burn might start eating any payload gains, but you definitely could squeak out a bit more that way. I'm not sure whether it makes $/kg through TLI sense, it'd depend on the cost of adding them (I don't know 1985 RL-10 prices) and the added payload that havin the stage do some POI could do.
Last edited:
I used the Apollo 10 10,970m/s peak velocity minus the orbital velocity of 7,500 m/s (27,000 Kmph) as being the maximum required delta-v needed - although 7,780m/s* has been quoted as orbital velocity as well. This said to me up to 3,470m/s delta-v for TLI, while I couldn't get more than 3,060m/s on the EDS, even with an Isp of 470s and the dry mass cut to 3,400Kg - just over 8%. With 2,940m/s being what I got on the numbers I'd shown.**
37,000Kg TLI, that's possible. 40,000Kg TLI, doable though tricky. 45,000Kg TLI, it stops making sense at about this point.
I think RL-10 engines costed about $2 million each in 1996 dollars, but I suspect those were OTL Titan IV RL-10s IIRC, more efficient ones will almost certainly be more expensive.
But all of this is off the current topic of Unmanned Payloads. It can all be sorted out properly when it's properly worked on.
* I know this is to do with orbital altitude, to the 7,780m/s velocity comes across as being the more likely. The TLI Payload and EDS wont need to stay in LEO for more than a very few days.
** I'm willing to accept the chance that I'd punched in the numbers wrong, giving me incorrect answers. It does happen to me from time to time. Even after I've checked them.
EDIT: I guess this illustrates the merits of writing down the numbers before working with them, huh?
(1) The altitude Apollos parked in was about 333 km according to the Apollo by the Numbers data, however, Wikipedia gives a value of 195 km for Apollo 10-14 and 172 km for 15, 16, and 17. The Apollo Flight Journal agrees, saying 165 km for Apollo 15. According to this calculator, 165 km is 7.809 km/s, 172 km is 7.805 km/s, 195 km is 7.791, and the 333 km average altitude from Apollo by the Numbers is 7.71 km/s. Whatever the source of the difference, the effect on orbital velocity is negligible. To get down to 7.5 km/s, you'd have to park it in 700 km parking orbit.
(2) Strange. I'm still pretty definitely getting ~3.13 km/s. Mass ratio=77000/(35000+3750)=1.987, ln(MR)=0.6867, Delta-v=9.8*464*0.6867=3125.6 m/s, yeah?
Yeah, it's really a question of "how much gravity loss can you add and still come out ahead?" and that kind of optimization is tricky without good numbers and fine pencils--and variables like "cluster of 3" vs "cluster of 4" vs "cluster of 5" vs "cluster of 6" can both add complexity and play merry hell with the "optimal" answer to an easy-to-ask question like "what's the best number of engines?"
Maybe, maybe not. You'll certainly be buying them in some quantity--even 2 flights of a 2xRL-10 Big Centaur for lunar or other BEO would be a noticeable bump in production rate and a probable drop in cost per unit. It may end up being a wash--individually more expensive, but more produced so the fixed costs are split over more units. End result, roughly the same cost.
Err, yeah, it is.
But I don't really know anything about unmanned I haven't already said, certainly not the end of it with alternate mission plans and reasoning why one thing was done instead of another.EDIT: I guess this illustrates the merits of writing down the numbers before working with them, huh?[/QUOTE]
Well I use a different means that I was taught by my Physics Teacher - who knew exactly what he was talking about. First I calculate the burn time (seconds) via: Propellant Mass/Engine Thrust * Specific Impulse (38,000/40,000*457 = 434.15s). After which I cut the propellant mass in half to measure the Average Mass during the burn. This would mean 77,000 - 19,000 = 58,000. After that, I divide the Engine Thrust by the Average Mass then multiply by the Burn Time and 9.8.
40,000/58,000*9.8*434.15 = 2,934.255172 m/s. The absolute maximum it can do based on these numbers.
So...essentially: find average mass, find burn time from propellant mass flow (m_dot), find acceleration at average mass, multiply by time to get a delta-v? The implicit assumption there is that the moment with average mass is also the moment of average acceleration, and thus some simple multiplication and division can stand in for more complex mathematics like the natural log. Unfortunately, that's not a valid assumption.
And I'm sorry, but I think my propulsion professor and text, combined with having done the derivation myself from the basic fluids definition of conservation of momentum for an assignment on no less than two occasions trumps your physics professor, who I suspect was deliberately using an incorrect version of the calculations (though not a severely incorrect one) to avoid having to get into the deeper fluids stuff. If you want to argue with the method I used, you're arguing with every rocket built since Tsiolkovsky (that page has one way of deriving this, by the way, though not precisely the same one I've used--the results are the same, though).