One way that you could increase J-130's performance by a coupla mT - have specific 3-engine & 4-engine thrust structures & piping.J-130 programme should not be affected in either cost or schedule, but there would be additional costs for phase 2 / J-24x development - substantial costs, I'd guess.
I'll take an operational engine over a "paper" engine any day.
John & mars,I really don't feel comfortable discussing that in public without seeking permission from the panel first.Lets just say that they have been asking questions, we are preparing data for them and some of the team have made contact directly. And the contacts have all been good so far.We have decided to leave it entirely to the panel themselves to control the release of all such materials and discussions for themselves according to their own policies.Ross.
During our available time, we strongly made the point that, being an integrated architecture, the most significant driver that sizes much of the architecture is lunar global access. This is by far the most dominant driver in how much mass must be delivered to translunar injection. Indeed, with our present baseline, the size of the rocket and lander alone do not enable global lunar access - to attempt that would result in a rocket that is too large to reasonably build. That's just the physics of the problem. So we utilize all of the parameters at our disposal (lander propulsion load, loiter time in lunar orbit, etc) to open the hardest to reach places to exploration with what we considered a reasonable heavy lift launch vehicle to lift the necessary mass out of Earth's gravity well - and sure enough, several of the most interesting sites are in such locations. We also pointed out that we had scaled back our Ares V and Altair assumptions to supporting only equatorial and polar landings (while still protecting a reasonable level of cargo delivery capability for establishing an outpost) and the cost variance was only roughly 10 percent cheaper.
Much attention has been focused on the probability of loss of crew (pLOC) as a figure of merit in determining the crew launch aspect of the architecture, and we expressed that the ESAS pLOC numbers were all using the same methodology and that the value was in the comparative results and not in the absolute numbers. Very simply, Ares' clear advantage is in the comparative simplicity of its first stage (the shuttle SRM) and use of a single gas generator cycle upper stage engine. These two attributes alone provide substantial robustness over, for example, a more complex liquid pump fed first stage and a multiengine upper stage - simply put, they are more complex with more moving parts. What Ares affords us, in accordance with the findings of the CAIB, is a crew launch system that has the potential to achieve unmatched safety in human spaceflight history. And this is not just a Constellation 'claim' as some would suggest, but has been validated by independent experts in the field of physics based probabilistic risk assessment. There will be much more provided on this topic as well.
On the topic of 'human rating', it is clear that the panel will want to hear more on this topic as well. The term gets thrown around in the community without a consistent understanding of what 'human rating' means. NASA's human rating 'policy' is clearly documented, but Constellation is the first program to really attempt to apply to a design in a practical manner. Our overall approach to human rating has been briefed to the ASAP, as has our program-wide approach to risk-based design that chooses robustness over blind fault tolerance in engineering these systems. All of our external review has largely validated this approach to date.
I watched a fascinating Apollo documentary last night, and interesting to see comments re the violence of takeoff on Apollo 8 - the guys actually thought they'd impacted the pad. How does that compare to expected levels of Ares I TO, if it happens?
Even with the mitigation efforts, TO on Ares-I is expected to still be able to impart up to +/- 2.0g of vibrations on the Crew Module, although seat isolators are hoped to reduce that for the crew themselves.Ross.
Quote from: Lobo on 06/24/2009 06:41 PMSince the Direct baseline architecture is 2 J-246's, with two JUS's, one fully fueled, one fueled just enough for LEO of CEV/LSAM.Could both JUS's be partially fueled, and have a rendevous in LE1 or LLO? You have to expend two anyway, couldn't the docking procedures be simplified that way? Also, in that vein, if a Centaur or D4US (are they the same? Don't know much about them, if they are two terms for the same upper stage or not) can take Orion around the moon for an Apollo 8 type flyby, could one be used to take the LSAM there too, again for a LE1 or LLO rendevous?I know fueled LSAM is heavier than Orion, so I don't know.But now I'm curious.You've just stumbled upon what, in my view, is the biggest weakness of the EOR-LOR mission profile. Try as you might, you can't distribute the total payload evenly among multiple identical launch vehicles without on-orbit propellant transfer. If the DIRECT ethos can be distilled to "one kind of vehicle launched multiple times", then EOR-LOR is a questionable mission profile, a holdover from the 1.5-launch approach that doesn't make as much sense in a 2-launch architecture.The DIRECT 3.0 architecture calls for one launch of about 100 mT and another of about 70 mT. It's not really a 2-launch architecture, hence the heavily-offloaded J-24x CLV and the barely-viable J-130 CLV alternative.With LOR-LOR, L1R-L1R, or (especially) L2R-L2R, the CEV and LSAM each have their own upper stage for TLI and insert themselves into rendezvous orbit separately.The key thing to understand is that the LSAM actually masses less than the CEV when it separates for lunar descent. The LSAM is only heavier than the CEV at liftoff because it does the LOI burn for itself and the attached CEV.Remember, the Apollo CSM was much more massive than the LM, mostly because the CSM did the LOI burn for the combined mass. Whichever spacecraft does LOI becomes much bigger than the other.But if both spacecraft do LOI and their rendezvous masses are similar, then their TLI masses are similar, and therefore their LEO requirements are similar, and they can be lofted on identical launch vehicles.With LOR-LOR, the rendezvous mass of the CEV is notably higher than the LSAM, because now it has to do its own LOI instead of relying on LSAM. So the CEV drives launch vehicle requirements.But with L1R-L1R or (especially) L2R-L2R, the CEV takes a cheaper round trip to the rim of the moon's gravity well. This increases LSAM mass, but it decreases CEV mass by a much greater amount, and the combined effect helps even out the rendezvous masses.With L2R-L2R, CEV liftoff mass is roughly the same as with EOR-LOR (depending on trajectory), even though it does its own LOI, and LSAM liftoff mass is dramatically reduced to about 20 mT, not much less than the CEV.Either J-130 or Not Shuttle-C could lift a 25 mT spacecraft with a 45 mT EDS to put it through TLI. With a 2-launch L2R-L2R profile, this is enough for the baseline lunar mission.Additionally, this same 5m Centaur-derived EDS could double as the new upper stage for EELV, and it would only make sense for ULA to lead the development, rather than NASA/MSFC.This mission profile allows for the development of a significantly smaller upper stage that's much more versatile and would see higher flight rates. It also allows for the development of a significantly smaller LSAM descent stage with a lower center of gravity for landing stability.The number of SSMEs expended per mission is reduced to six, global access to the lunar surface without expensive plane-change maneuvers, and global communications relay to earth via CEV at EML2.If you try to replicate a 1.5-launch EOR-LOR architecture with one kind of launch vehicle, the closest you can get is DIRECT. But for a true 2-launch architecture, L2R-L2R makes more sense.
Since the Direct baseline architecture is 2 J-246's, with two JUS's, one fully fueled, one fueled just enough for LEO of CEV/LSAM.Could both JUS's be partially fueled, and have a rendevous in LE1 or LLO? You have to expend two anyway, couldn't the docking procedures be simplified that way? Also, in that vein, if a Centaur or D4US (are they the same? Don't know much about them, if they are two terms for the same upper stage or not) can take Orion around the moon for an Apollo 8 type flyby, could one be used to take the LSAM there too, again for a LE1 or LLO rendevous?I know fueled LSAM is heavier than Orion, so I don't know.But now I'm curious.
Quote from: kraisee on 06/25/2009 05:47 PMJohn & mars,I really don't feel comfortable discussing that in public without seeking permission from the panel first.Lets just say that they have been asking questions, we are preparing data for them and some of the team have made contact directly. And the contacts have all been good so far.We have decided to leave it entirely to the panel themselves to control the release of all such materials and discussions for themselves according to their own policies.Ross.Ross:You and the team probably do need to be very careful in what you say about your interactions with the commission. I think you did that above. I wouldn't give any specifics about anything though.However, good to know they are following up.It's amazing to me how quickly the commission is to conduct itself. basically 2 to 3 months to do the whole thing. Very impressive. I wish you the best of luck.