Here’s a quick thought on a way to use EELVs for launching Orion that I was thinking about tonight.Â Â One of the reasons why the CEV is so big that it’s hard to launch on an existing EELV is because of the amount of service module propellant.Â Basically, the CEV is sized to provide nearly 1800m/s of delta-V.Â Now, you don’t actually need most of that for an ISS mission (except they use it as a contingency to provide enough delta-V during an abort over the North Atlantic in winter to make sure you can put the CEV down in a safer location), but you do for a lunar mission.
I don’t have the latest numbers, but some numbers I’ve seen for the CEV (from a presentation Pete Worden gave a few months back about using Orion for an NEO mission) put it at about 24.7klb dry and 20.5klb propellant.Â In other words almost half of the CEV launch mass is NTO/MMH for the service module.Â Once you factor in the much bigger tanks and engines for such a service module, the CEV isn’t really all that much heavier than a Dragon capsule for instance.Â But once you include all that lunar return propellant, you now have a capsule that’s so big that it’s hard to loft on all but the biggest existing launchers.Â In fact, due probably to the switch back to hypergols from the LOX/CH4 suggested in ESAS, right now in order to fit its mass targets, Orion is having to shed a lot of the redundancy and functionality it needs in order to perform its mission safely.
Right now, as I understand it, the current concept of operations is that the Ares-I puts the orion into a suborbital trajectory with most of the velocity needed for orbit.Â The Orion then provides the circularization burn to put itself in orbit.
The question I had was, if you are already using the Orion somewhat like a third stage, what if you actually did use it as a third stage?Â While you couldn’t launch a fully-fueled Orion into LEO on anything other than one of the EELV Heavies, transfer of hypergolic propellants is now a demonstrated capability, even in the US!Â So, I was curious how small of an EELV you could use and still get Orion into orbit (with enough propellant left for rendezvous and docking maneuvers), if you assumed that for lunar missions you could tank-up the CEV on orbit.
I happened to have some mass numbers from previous conversations about the Atlas V Phase 1 and 2 concepts that ULA did, so I put together a spreadsheet. First, I took the payload numbers from ULA’s site, and the Centaur and CCB mass numbers I had sitting around and estimated the total delta-V to LEO for the stack. Then, I took the rough numbers for Orion I had above, and treated it like a third stage. The Orion payload adapter was added as a dry weight to the Centaur stage, and the LAS was added as a dry weight to the CCB stage (since it would likely be tossed immediately after stage separation). As you can see, a 1.5x Phase 1 Atlas V could do the job, leaving a bit of performance for margin and maneuvering delta-V. Of course, if someone were serious about doing this, you would likely oversize the Phase 1 upper stage a bit to provide extra margin, to relax the constraints on Orion. Call it a 1.6x or 1.65x (referring to 1.6 or 1.65 times the propellant load of an existing Centaur).
For ISS missions, you wouldn’t even need to top the propellant off, as you’d have enough leftover performance after arrival for rendezvous, docking, and deorbiting. For a lunar mission, you would need most of a second launch worth of propellant, but that propellant would be a very low cost cargo (and could also reasonably be launched by other lower-cost commercial suppliers). By using the service module as an upper stage, you would get to test out the engine thoroughly on the way to orbit, which might reduce your risks of unforeseen problems cropping up on your way home from the Moon. Also, if for some reason the upper stage fails (though it has engine-out capability now unlike the Ares-1 US), you still have the extra propellant for contingency maneuvers to avoid landing in the North Atlantic.
The single stick Atlas-Vs meet almost all of the old NASA human rating requirements (and NASA had to lower their standards enough for Ares-I to pass that most of the few human rating requirements that Atlas-V didn’t already meet are no longer there), and most of the remaining requirements are improvements that ULA wanted to do anyway for their Bigelow collaboration. The nice thing about a Phase-1 Atlas V is that the CCB is unchanged from the existing Atlas-V, and even the upper stage changes all have prior design heritage. The Titan Centaurs used the wider body diameter that the Phase 1 design would use, most earlier Centaurs were dual-engine, and between Lockheed and Boeing’s contributions to ULA, the friction stir welding techniques that would be needed are mostly developed and qualified already. And the engine that would be used, the RL-10 is an existing one with excellent heritage, and very benign operating characteristics.
Anyway, I just thought this analysis was kind of interesting. Such a system would still be able to loft a standard Orion, it wouldn’t require a ton of new development work (a little stage work, but a close derivative of an existing stage done by a team that has a proven track record from doing several such stages within the past decade or two), and since it would be a single-stick system, it would likely compare very favorably (LOM/LOC-wise) with the Ares-I.
Just a random thought.
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Couldn’t this approach be used with an Ares I launcher as well? It would be funny (or at least highly ironical) if NASA ended up doing orbital refueling/gasteroids in order to save Ares I.
Happy New Year and thanks to you and everybody here for a great blog ^_^
Yeah, in theory you could do this with Ares-I. But a lot of your excuses for doing Ares-I have been shredded.
Did you look at what happens if you do this with the Delta 4?
the D4 has a bit more performance then the A5, so, you
can avoid some of the stage work, and the D4 has plenty of
heritage, and means less design work.
Personally i think NASA should be stripped of the
ISS launch mission, and the USAF should take over
that mission. The USAF should be tasked to get
X tons to ISS every year, and Y Astronauts, and
NASA should just designate the passengers and cargo.
Let NASA keep funding COTS and work on Research.
The challenge, of course, is that any chosen architecture must address political issues that have little to do with pure engineering or even accomplishing objectives in space. For example, perhaps ATK “deserves” to be excluded from future VSE contracts but do the Atlas V plus propellant depot advocates have the votes in Congress to accomplish that?
I’ll betcha a dollar Dr. Zubrin will have 200 Mars Society members walking the halls of Congress in late July vocally opposing an all-EELV VSE. Are there “ANY” known Congress-critters who have gone on the record as being willing to introduce and support legislation to adopt an Atlas V + propellant depot architecture in lieu of everything else?
With the DIRECT people now calling for propellant depots to further leverage their architecture perhaps its time to ask a “target fixation” question concerning EELVs and propellant depots and acknowledge the political wisdom of Team DIRECT’s commitment to “leave no contractor behind” and work to get funding for propellant depot research and prototype deployment incorporated into the 2009 Congressional legislation, as part of a transition to DIRECT.
I think that for all the talk of political “realities”, people forget that politics does change. Contractors do indeed get “left behind” all the time. The complaints of losing jobs in Utah isn’t going to have quite the impact with a Democrat majority Congress as it did with a Republican one. Likewise, jobs in Florida counties that Obama lost anyway, Texas, Alabama, Mississippi…Do you really think that keeping those people employed is going to take precedence over other fiscal pressures in an economic environment like what we have today? It’s possible, but I wouldn’t bet on it.
Not saying that I think Congress is going to all of the sudden start funding NASA missions on any less corrupt of a basis. Just saying that I wouldn’t be surprised if relative levels of clout changed drastically in the new political environment.
Cool concept. With 10,400 m/s DV it would make orbit with a cool 1,000 m/s to spare.
“Likewise, jobs in Florida counties that Obama lost anyway, ”
Counties aren’t winner-take-all like states. Losing a county by less than you otherwise would have means you don’t have to win other areas by as big a margin.
Within a state’s boundaries, votes are fungible.
True, but there was no evidence that any of Obama’s space positions did him any better in that part of Florida than he would’ve done had he said nothing about Space. I think a lot of people really want to believe in some magic tooth fairy called “political realities” that means they don’t have to actually convince people that their approach is really any better or more efficient. Political realities change all the time, and I think that there are some people that are still thinking this is 2005.
RE: chicken vs egg; the NASA version
If Ares I is dropped, perhaps NASA might also revisit the Orion specifications too, because sizing the Orion spacecraft for six crew makes no sense. If the lunar plan is for four crew than the Orion should also be sized for four crew, not six. Such a reduction in size would ease up much of the margin needed by the crew-launch vehicle and the cargo-launch Ares V (VI?).
In fact the only way the CEV specification makes sense is if the mass was deliberatly scaled-up by NASA to match the anticipated payload of Ares I. And it’s very clear much of the development costs of Ares V are piggybacked on the cost of Ares I. Hence the way NASA makes everything revolve around the desire for the heavy-lift Ares V which is a bass-ackwards method of coming up with a reasonable cost-effective architecture.
I don’t buy the NASA excuse that the six-man Orion capsule is needed for potential Mars misssions because I don’t see how a six man crew is a neccessity for such a hypothetical mission and much more importantly an Apollo moldline reentry vehicle is grossly inadequate for typical reentry speeds from a Mars mission. Reentry speeds from Mars are more like 14 km/s, much higher than the lunar reentry speeds of 11 km/s for which the Apollo shape was selected.
RE: Orbital refuelling of the Orion service module
An excellent idea, and one that had occured to me too as I fooled around with alternative lunar architecture concepts. And this made me consider another wild idea which I have yet to crunch the numbers on, what if the Altair’s ascent module (which is planned to use the same engine as the Orion) is also fuelled in orbit?
Looking at the Altair’s ascent module and it’s requirements and basic specifications made me consider another interesting possibility. Perhaps that ascent module could be so designed as to serve different jobs. Not only could it be used as the Altair ascent module, but also as a mission-module for the Orion (such as for NEO missions) and also as a reusable orbital tug (such as the Russian proposed Parom tug). In fact the configuration of the Parom tug might be a usefull jumping off point to design such a multi-purpose spacecraft. (I’d almost say such a spacecraft could serve as the Orion service module, but that is impractical for several reasons)
Please stop making so much sense. Heaven forbid anybody figure out a way to close the gap and save money over Ares I.
In all seriousness, it’s a pretty sound approach. Ironically, it fulfills the flawed ESAS finding that EELV’s would need a bigger upper stage to pull off the crew launch mission. It also dispenses with the SRB’s which hurt your LOM numbers. My past experience with spreadsheets to do this kind of estimation has been iffy, but I got the impression that Atlas V would really benefit with a larger upper stage relative to the first stage. Due to the higher Isp for Stage 2, a bigger Centaur would result in a more optimal delta-V breakdown between the two stages.
What kind of thrust/weight do you have with that much mass added for a LEO mission on a single-stick Atlas.
The first stage T/W ratio wouldn’t be impacted that much. You’re adding something like 40klb on takeoff to a vehicle stack that already weighs something over 700klb. That might impact gravity losses a bit, but I’m not sure how much. As for the Centaur stage, the one I used in this BOTE analysis was a dual-engine Centaur, so I don’t think you’re in trouble there either.
As it is though, this is just a BOTE analysis showing that something like this *could* work. It would obviously need some actual trajectory modeling to verify the various remaining issues. It may turn out that there’s some reason it wouldn’t work, but at least on a preliminary basis, it appears feasible.
Just looked a bit further (I don’t have all the numbers, since I’m at work), and it looks like it would be down around 1.1 T/W ratio, compared to the current 1.15. Could be an issue, maybe not. Would need a more detailed analysis to be sure either way. The other option is that a Phase 1 Atlas V single-stick has just about enough capacity to launch the Orion dry without using it as an upper stage, so there may be some intermediate plan that works while keeping liftoff T/W reasonable (ie a partial load of the orion to balance offloading some of the delta-V performance to that stage, while still keeping GLOW within reasonable bounds). Good news is that either way, Max-Q would be very low compared to der Griffenschaft.
Iâ€™ll betcha a dollar Dr. Zubrin will have 200 Mars Society members walking the halls of Congress in late July vocally opposing an all-EELV VSE.
Taken. The Mars Society is more diverse than you imagine. The unifying factor is wanting to go to Mars, with at least some govt. facilitation and funding. Belief in a particular architecture is not a sine qua non of membership.
Has anyone seen this?
“military rockets may be cheaper and ready sooner than the space agencyâ€™s planned launch vehicle”
“””””””Just looked a bit further (I donâ€™t have all the numbers, since Iâ€™m at work), and it looks like it would be down around 1.1 T/W ratio, compared to the current 1.15.”””””””
We’re talking Atlas V here, right? For what it’s worth or even relevant, the current Atlas V takes a long time to get off the pad, so lowering the T/W ratio may be a problem. As it is, it takes 11 seconds to clear its launch tower, and, according to one source anyway, probably consumes over 50klb of propellant in its first 20 seconds to get to about 80 mph.
Not sure if that’s a big deal or not. It’s actually about 55klb in 20s by my count. Out of a 625klb propellant load. Which means that all told, that stage is actually developing a *lot* of delta-V. Numbers out of context can sometimes be decieving. For instance, 20s of gravity losses is only about 200m/s of delta-V (about 1/7 of the typical gravity losses I’ve heard for stages). While a lower initial acceleration will definitely increase gravity losses, the question is by how much? And the answer is–hard to tell without a good 3DOF.
Jon, thanks for the follow-up.
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