Just food for thought.

Let’s turn to the Augustine Report itself for some information. On pages 83 and 84 they discuss implementing the Program of Record on entirely unconstrained budgets–ie if we gave the program the full funding it needs to execute, and allot it to move at the full pace it can realistically move at, what do we get?

• A $145B pricetag over the 2010-2020 timeframe, which doesn’t even get us to the point of having Ares V and the LSAM ready for operations, much less a moonbase.  This would require almost $5B extra per year–ie a 25% increase in NASA’s topline budget.
• An international space station deorbited within 5 years of its completion, during which time the only method of access would be by paying the Russian government for flights.
• A crew launch vehicle that becomes available two years after its first destination is deorbited, and whose operational costs have to be carried for over half a decade until we have any of the tools that would be necessary to actually use it for anything.  But don’t worry, we can spend $2B+ per year to send even fewer astronauts flying in even more useless circles.
• A seven plus year manned orbital spaceflight gap in the US.
• Almost no investment in long-term technology development (not much more than the current SBIR budget, and entirely focused on short-term Constellation needs, not on making future missions safer, more affordable, and more valuable).
• No stimulation of commercial industry beyond the CRS contracts which wouldn’t be extended since the ISS would be gone by 2016.  No investment or early market for commercial crew delivery
• No money to actually develop hardware for actually doing anything on the Moon, since almost all of the money will go to figuring out how to go there while maximizing employment in Shelbyville.
• No more robotic orbiters or landers for years to follow-up on the work LCROSS did.

But hey, at least if we do it this way, sometime 15+ years from now, we’ll have the ability to send 8 people to the moon every year at the cost of an “exploration” program that costs almost as much per year as NASA’s entire current budget!

If you assume that there are parts of NASA outside of Huntsville that actually matter (ie that NASA != Northern Alabama Space Administration), the situation gets even worse.  In order to fund Constellation at full speed without splashing the space station almost as soon as it’s completed, you would need $159B over that timeframe, which constitutes a $7B per year increase for NASA.  That increase still:

• Gets you a space station you can’t access without the Russians for most of its operational lifetime (why does Congress trust Russian commercial space more than American commercial space, btw?).
• Gets you no real investment in long-term technologies, ensuring that the cost, safety, and efficiency of manned spaceflight will be stagnant for another couple decades.
• Gets you no real investment or encouragement of the commercial industry (in direct contravention of the laws of the land and NASA’s charter I might mention).
• Gets you no more robotic follow-ons for LRO and LCROSS for over 15 years.

Compare this with the Flexible Path option that Mark likes to mock so much.  For less than half as much of an increase per year, you get:

• Robust ISS utilization through 2020, with multiple methods of providing crew and cargo delivery that aren’t all dependent on Russia
• Investments in commercial space that can help keep the US in the forefront of space technology and utilization
• Robust investments in high-payoff medium-term technologies like propellant depots, space radiation, space nuclear power, aerocapture and other EDL techniques, ISRU, and other high-payoff technologies that can vastly lower the cost of future exploration missions, allowing us to accomplish more for less and at lower risk.
• A manned lunar landing program that at most is only 3-4 years behind the current PoR, but when it gets there, it provides a much more affordable, more commercially and internationally interesting program, and has much greater capabilities once you get there.
• A manned spaceflight program that is much more capable of exploring the whole inner solar system, and not just doing a few flags and footprints landing on the Moon.
• A manned spaceflight program that builds on and leverages our impressive achievements in robotic space exploration.
• A program that in spite of doing a lot more looking, also allows a lot more touching of new destinations like NEOs and Phobos/Deimos, all on about the same timeframe that the PoR would at best be going for its first lunar landings.

Where I come from, we tend to think that getting a heck of a lot less while paying a heck of a lot more is usually the sign of a sucker.  I just wish that a few space pundits and public figures didn’t keep enabling Senator Shelby and his ilk from hijacking NASA’s budget to enrich his campaign contributors at the rest of our expense.

[Note: As an aside, am I the only one who finds Shelby's latest childish tantrum accusing the Augustine Committee of being compromised by biased by evil commercial lobbyists to be richly and hilariously ironic?  When it comes to lecturing people about the evils of lobbyists corrupting the political process for their own personal gain, Senator Shelby has about as much moral standing as Tiger Woods does when it comes to lecturing people about marital fidelity.]

To me one of the more interesting points is found at the top of page 11.  There were several misleading statements made by several people today about the relative safety of Ares-I compared to commercial crew vehicles.  As Brett put it (my emphasis added):

Second, some have claimed that NASA’s Exploration Systems Architecture Study (ESAS) shows that the current exploration vehicles are safer than commercial crew vehicles. In actuality, commercial crew vehicles were never even analyzed in the ESAS report – the ESAS report only looked at vehicles large enough to carry Orion, such as Ares I and variants of the triple-core Delta IV Heavy, and did not examine the smaller, simple, single-core vehicles, such as Atlas V Medium and Falcon 9 Medium that are sufficiently sized for commercial crew missions.  Moreover, even if ESAS had compared exploration vehicles to commercial crew-sized vehicles, the comparisons would be “apples vs. oranges,” because of the dramatically different tasks of these two types of vehicles.

When Jeff Hanley talks about how the Great Oz and supercomputers at NASA show that Ares-I is 3x safer than commercial launch vehicles, I wonder if he’s ever going to release their analyses for actually commercial crew vehicles, or if he’s being accidentally or intentionally dishonest. Because so far we haven’t been shown any data about the safety of actual commercial crew launchers. So far we have lots of data shown for the risks of using existing or modified commercial launch vehicles for launching a massive spacecraft designed to go to and return from the moon, including significant plane change maneuvers to allow anytime returns (ie Orion). It’s interesting to note that over half of the mass on Orion is the oversized launch escape system needed to get away from an SRB you can’t shutoff, and enough propellant for about 1500m/s of maneuvering to reach orbit and then to do in-space ops. That’s above and beyond the RCS propellant on the CM itself.

Most of the stuff that make Orion so massive are flat-out completely unnecessary for an earth-to-LEO crew capsule. You don’t need those kinds of delta-V capabilities. You don’t need as roomy of facilities, since by definition the flight times should be a lot shorter. Etc. There’s a reason why almost all of the proposed commercial crew systems are able to utilize single-stick launchers like Atlas V or Falcon 9–for an actual earth-to-LEO capsule you really don’t need anything bigger.

This realization that earth to LEO capsules can be much smaller than Orion leads to at least two important corollaries that I can think of:

1. Smaller capsules mean higher structural margins.  One of the existing vehicles most often suggested for commercial crew, Atlas V, was designed for the worst-case loading environment of any of its configurations (in this case I believe that would be the Atlas V 551 or 552).  The Atlas V 552 sees much higher max-Q’s than the 401/402 do, and has a much heavier payload on top, which exerts much larger structural loads on the Centaur stage than are seen in the 401/402 configuration.  While the Centaur structures may not meet the 1.4 magic number NASA likes in some of the bigger configurations, as I understand it, it actually exceeds that number in the 401/402 config most likely used for commercial applications.  The Falcon 9 was designed from the start to meet NASA structural margin specs.
2. No need for strapons.  Only one of the commercial crew ideas I’ve seen so far used a vehicle with strapons for crew launch (Dreamchaser).  This alone should make a huge difference in launcher reliability, since there are less things that can go wrong, less staging events, etc.  Most of the commercial launcher ideas they mentioned in ESAS assumed multi-core configurations.

There’s also the possiblity on the Atlas-V of using a dual-engine Centaur configuration to allow for some upper stage engine-out capability, or running the RL-10 at a derated performance level (not sure if that’s something it can do automatically, or if you’d have to make modifications–if you have to modify it it probably isn’t worth it).  With the much lower max-Q, and the ability to shut off the booster engine in case of an abort, I have a hard time believing that Ares-I is really that much more reliable than an actual commercial crew capsule launched on a commercial launch vehicle that has dozens of flights under its belt.

Food for thought.

Since I’ve been doing a lot of rocket plumbing, let me use an analogy.  Imagine you have a small rocket engine, like our igniter.  It’s got a tiny fuel orifice that’s only a tiny fraction of an inch in diameter.   Now, if the solenoid valve upstream of that orifice is small enough, it can meaningfully decrease the overall flow.  But after you reach a certain valve size, the valve size ceases to be relevant to the overall flow rate.  You could put a 1″ full port ball valve leading up to that orifice, and the flow difference between that an a solenoid valve with a 1/32″ port is going to be round-off error.  Once the size difference between the most constricting component and the rest of the components gets past a certain point, you can pretty much ignore them.

There are tons of other analogies from electronics, manufacturing, structures, etc.  Basically, in almost any system, you can improve one component only so far before you hit rapidly diminishing returns.  A wise engineer will spend his resources in a way to maximize the overall system reliability, not just one small sub-component.

Unfortunately, that’s exactly the mistake that the CxP guys have been making with Ares-I.  Even if, in spite of all evidence so far, and in spite of all historical precedent, Ares-I really is as reliable as their Probibalistic Risk Assesments suggest, it still isn’t a wise investment of capital when you look at the overall exploration mission.

Let’s just go back to the math again.

According to ESAS, the predicted probability of losing a crew on a lunar mission was something like 1.6% (or about 1/60).  Of that, only about 1/2000 (or .05%) came from ascent risk–or about 3% of the overall crew risk for the entire mission.  A mission that used the worst numbers they came up with for existing EELVs with an LAS attached estimated a 1/600 probability of losing a crew (or about .16% chance).  That would increase the probability of losing a crew to 1.7% or about 1/59…

Investing tens of billions of dollars to reduce the probability of losing a crew from 1.7% to 1.6% only makes sense if there are no better safety investments out there.  With a 1/600 ascent safety rating, about 90% of the danger to the crew is coming from other phases of the mission–which strongly suggests that the best safety return on investment is not in overoptimizing the ascent reliability.

Now, to be fair, Shuttle has a safety rating estimated to be around 1/100.  At that reliability rating for crew launch, it would be one of the dominant crew safety risks in a lunar mission.  Spending money to get it up past the 1/500 range is a good investment.  But after a point, if your goal is to improve the overall probability of getting a crew safely to the moon and back, you’re best off finding other areas to invest your money.

On a related note, the self-righteous attitude that many CxP people like Griffin and Hanley take that the Augustine Committee is ignoring crew safety is kind of farsically hypocritical at best.  Blowing so much of your budget on ascent risks, when they aren’t the dominant risk is actually making the overall mission less safe, not more.  Wisely spending that money instead on mitigating the risks in lunar landing, ascent, surface ops, and earth return would result in a higher probability of not killing astronauts on a given mission.  If astronaut safety is really so important to CxP, why do they only seem to care about the first 5-10 minutes of the mission, when it’s the rest of the mission that accounts for 90-97% of the danger?

Anyhow, Jeff managed in 25 minutes to address human rating, depots, whether or not we need heavy lift, technology maturation and R&T investment, and the need for NASA to find new ways to interact with business. I don’t think he could have hit more of my hot-button issues in 25 minutes if he had tried.

Anyhow, I hear that the HSF committee will have video of today’s proceedings up online soon (possibly tonight) for those who didn’t get up at 5:30am PDT to watch. I’ll comment more later.

Whew! I haven’t had this much hope for this nation’s space program in years!

[Update:  Here's the link to the subgroup's presentation (warning, it's a 14MB powerpoint presentation).  All of it is interesting, but Jeff's part starts on page 76 and goes through page 89.  Chris Chyba's section was the last three pages.]

To quote the former senior NASA official who was poo-pooing depots in the article I linked to earlier

Rocket malfunctions are not uncommon, and the more launches are needed for each moon mission, the more likely it is that something will go wrong, a former senior NASA official told New Scientist.

ESAS had a whole section slamming lunar architectures that used more than two launches as having too high of a probability of losing the mission. The ESAS study pointed out that the odds of not losing any of the launches is much lower than the odds of not losing any one given launch. The probability of successfully doing N launches without any failures is the probability of success for a single launch raised to the Nth power (ie 6 launches with a 98% reliable vehicle only has 88.5% chance of not losing one vehicle). They also went into rendezvous reliability, launcher availability, boiloff losses etc…

…but for some reason that doesn’t matter for Mars missions. Why are large numbers of launches, rendezvous, and long duration cryo storage considered perfectly acceptable for Mars, but completely unnacceptable for the Moon?

The mission uses the following tricks to make things work:

• Dual Engine Centaur for this mission is stretched by 50% and includes an “Extended Mission Kit” to allow for it to function for the ~5 days necessary for the mission (normal DEC dry mass is ~5400lb, and the EMK is ~1750lb and includes stuff like extra hydrazine bottles, more batteries, deep space navigation upgrades to avionics, sunshields, etc)
• Command module does a powered lunar swingby to go to L2, thus cutting down on overall dV requirements (~750m/s total required, 335m/s per leg), thus allowing for a much smaller CSM (possibly with the service module integrated into the command module).
• The Stretched Centaur and the Lander break into lunar orbit and descend to the surface instead of continuing to L2.  I’m not positive if this allows you to land anywhere on the lunar surface or not (this is one of the few big questions for this mission mode).  This avoids the extra dV requirements you normally get for stopping everything at L2 first.
• Upper stage performs part of the landing burn (between LOI and the descent burn it provides about 1950m/s out of the total 3050m/s needed for LOI and landing).
• RS-68A Upgraded Delta-IVH. This upgrade is already in engine testing and is badly needed by the DoD, so there’s a good chance this will work out.  Expected payload capacity I’ve heard is 27mT for the system.
• Instead of carrying a second stretched Centaur as a payload on one of the flights, the Atlas V 552 uses the stretched Centaur as its upper stage.  In order to tank up the LH2, it carries an LH2 drop tank between the lander and the command module.  It gets transfered right after reaching orbit, and gets dumped shortly before TLI.

Here are the major components of the system:

• Command Module: This module is based on the Apollo outer mold line, but only carries two people, and enough life support consumables for the mission.  I budgetted 11,000lb dry and 3250lb of propellant for the capsule (not including RCS propellants).  I assumed hypergols for the stage, with a crappy 314s Isp.  The Apollo CM wet mass was 12.8klb, and the SM weighed 54klb wet, 13.5klb dry.  However, most of the SM mass was due to the CSM performing the LOI burn for the Apollo Stack.  About half of the dry mass of the CSM was the huge main engine, and a good chunk of the remaining mass was electrical equipment and the huge tanks for the 40klb of propellant.  With modern materials, electronics, a smaller crew, solar panels instead of fuel cells, and the much lower propulsive requirements for the Command Module in this architecture, I think 11klb is actually pretty conservative for such a system.  For another comparison the latest CEV numbers I’ve heard (which are pretty far out of date) were ~18klb for a four person capsule.
• Stretched Centaur Lunar Transfer/Crasher Stage:  As mentioned above, this is a dual engine centaur using two RL10A-4-2 engines, but with a 50% barrel stretch to the tanks.  The tanks are actually less than 40% of the dry mass of a centaur stage, but you also need more helium for pressurization of the larger stage…assuming that the 50% greater propellant load requires a 50% higher dry mass should be a conservative estimate.  The idea of a stretched Centaur shouldn’t be too crazy when you realize how many iterations General Dynamics, Martin Marietta, and Lockheed Martin have done on the Centaur just in the past 20 years (including 5m diameter Centaurs for use on Titan IV among other things).  The 1750lb for the extended mission kit is also based on numbers from previous papers LM/ULA has published about converting their stages over for longer-duration missions.  Total dry mass I assumed was 9850lb.  Note that the Atlas V 552 performance numbers also include 5400lb worth of Centaur burnout weight, so you only have to provide ~4450lb worth of “payload” for the Stretched Centaur.  Also note, that if you tank the stretched Centaur up all the way for launch, it should probably increase the payload capacity of the Atlas V 552 a little compared to a normal Centaur, but for purposes of this analysis we’re assuming only the nominal payload of a normal Atlas V 552, to be conservative.
• Single Stage Lunar Lander/Ascender: This stage takes the crew the rest of the way to the lunar surface after the Centaur has provided the first part of the descent burn, and then provides the ascent burn, and the burn to take the crew to the L2 staging point to rendezvous with the Command Module.  I budgetted 1100m/s for its portion of the descent burn, 100m/s to allow for a 90s hover to find the best landing spot, 2650m/s for the lunar surface to L2 burn, and about 50m/s more for contingencies.  This is probably the most aggressive part of the mission.  For this vehicle, I’m assuming a piston-pump-fed LOX/CH4 stage, based off of the piston pump and LOX/Methane engine work XCOR has done  (possibly combined with stuff that we at Masten have done that they haven’t like gimbals, throttling, etc).  The piston pump requires very low net peak suction head, which allows for very low pressure tanks, that can be made of  the LOX/Cryo-compatible Nonburnite composites that XCOR has been devleoping.  XCOR developed the piston pump and Nonburnite composites explicity for making propellant tanks out of shapes that aren’t typical for propellant tanks (in their cases to make the CG numbers work, they wanted to do LOX-filled “wet wings”).  Using this technology, instead of heavy pressure fed tanks and heavy helium tanks, you have lightweight composite tanks that can actually form part of the load-bearing structure of the vehicle.  As I understand it, based on my recollection of their public statements, the piston pumps they’re looking at using scale to about enough flow for a 2500lbf engine in a single pump.  By combining them with the 7500lbf engine XCOR developed (with a nozzle extension of course), you have significantly more thrust than you need for landing.  More importantly, you can possibly make the three pumps operate in a redundant fashion, so the loss of one pump can be tolerated at any point in the mission, and the loss of a second pump can be tolerated through most of the mission.  If done right, the pumps could be “armored” as XCOR calls it, but placed in such a way that they have removable manways between them and the main compartment that would allow for shirtsleeve troubleshooting/repair (the pump compartments would need to be done in a manner that if something went horribly wrong, that any debris/blast would be directed away from the crew cabin…but I can imagine a few ways that could be done).  All told, I’m assuming a 4350lb dry weight, a 9000lb propellant weight, 500lb worth of hardware to be left on the moon, and a 360s Isp.  The LM ascent stage was 4200lb, but held only 65% of the propellant mass, and only about half the propellant volume of this lander, and didn’t have to do landings, and didn’t have to support the crew for as long (about 3 days vs. the target 9 days to give you a week on the surface and 2 days in transity to L2).  But as mentioned above, it used pressure fed tanks, with the mass of a helium blowdown system, had to provide significant RCS capabilities since the stage did not have a gimballed main engine, was using crappy 60s era electronics and electrical systems, and had tanks that were entirely non structural, and also didn’t have access to modern materials like lithium-aluminum or modern composites.  However, the 13,850lb total mass for the lander actually compares pretty well with the 13,510lb currently assumed for the pressure-fed, hypergol-fueld Altair Ascent stage (from this document), which carries 4 crew for the same mission duration.
• Pre-Depot LOX Tank: This ~2.2klb Tank holds ~57.1klb of LOX for the Stretched Centaur.  It includes a docking port (possibly using LIDS technology?), a sunshield, and a Centuar-derived LOX tank.  It gets launched as the sole payload for the Delta-IVH, using up all but about 200lb of its capacity.  But since it is so dense, it might be able to get away with using a shorter (and lighter weight) fairing than is typical for Delta-IVH if that wouldn’t require lots of expensive aero analysis.  This tank, if launched with the LOX pre-chilled can hang out for over a month waiting for the Atlas V 552 launch.
• LH2 Drop Tank: This ~62.5 m^3 tank weighs about 2000lb (with another 2000lb budgetted for connecting structures between the various parts of the launch stack).  It would be housed between the Lander and the Command Module on the Atlas V 552 launch.  It would possibly use 5m tankage derived from the Delta-IV US.  After reaching orbit, the LH2 from this tank would be transfered (using propulsive settling) into the Stretched Centaur.  After the Command Module docks with the Pre-Depot LOX tank, and has transferred all the propellants from that (and discarded the pre-depot LOX tank), the CM and empty LH2 drop tank would separate from the stack, the drop tank would be discarded, and the CM would reattach to the lander much like was done on the Apollo Missions.

Now, this mission model isn’t perfect.  It uses most of the capabilities of the two launchers without a huge amount of margin (except in the fact that the Atlas V 552 with stretched Centaur probably has some margin built in that isn’t being explicitly called out).  And I’m not a fan of launching the crew on an EELV with 5 solid strapons.  It would be a lot easier if you assumed the development of something like the Common Upper Stage that ULA has been talking about recently.  With that, you would have tons more margin (since a CUS would add nearly 7mT of capacity to the DIVH, and probably at least 5mT to the Atlas V 552–possibly enough to go with less or no strapons on the crew launcher).  But it demonstrates that a 2-launch EELV mission using almost no modifications to existing launch vehicles (beyond the Centaur mods) is within feasibility.

The system also has several good things going for it.  First off, it can deliver lunar crew to the surface without a depot.  It doesn’t need Autonomous Rendezvous and Docking (since the rendezvous and docking can be piloted), or tankers to be developed.  It doesn’t need HLVs or 10m fairings (everything can fit within a stock Atlas V fairing).  It doesn’t need really long term LH2 storage in orbit.  It only requires two launches for the mission, and doesn’t put anywhere near as much launch timing constraints as the ESAS architecture does.  It can provide for cargo missions (~19klb delivered mass to the surface assuming that 2klb of the lander stage is in the form of a removable crew cabin, which just happens to be enough to land a Bigelow Module).

And most importantly, if depots do come into existence, it can immediately take advantage of them.  With just an LEO depot, you can both cut down on the number of EELV launches to just one (and use lower-cost systems like Falcon 9’s, Zenits, Ariane-Vs, Soyuzes, future commercial RLVs, etc to launch the remaining propellant).  Also by getting rid of the huge LH2 drop tank, you simplify the stack, remove about 15klb worth of hardware from the Atlas stack , dropping it to the point where it can possibly be launched by a 502 launch instead of a 552 launch (since the stretched Centaur provides almost as much propellant as a Phase 1 Atlas, which was supposed to boost the LEO capacity of the single-stick Atlas to almost 30klb).  Or you could use that saved mass to beef up the lander and/or command module for more capable missions.

If you have both a LEO and an L1 or L2 depot, the Centaur can top itself up again that depot, and provide a much larger chunk of the descent burn to the lander stack.  With enough propellant left over to return to LLO then to L1/L2 after separating from the lander, allowing the Stretched Centaur to be reused multiple times.  With such a system you could actually soft-land bigger payloads than the Altair cargo lander…and you’d have the capability of making the lander and transfer stage fully reusable.  The transfer stage, since it wouldn’t see atmospheric flight, reentry, lunar dust, or even particularly bad thermal environments should actually be reusable for several flights–the RL10 is after all rated for 200 relights.  The lander may be tougher, but by the time you have an L1/L2 depot, you’ve probably had enough time (and enough surface infrastructure built up) that you can work that out to.

Ok, so maybe it’s not so bad of an idea after all.

Well, Mark’s at it again.  Commenting on some interesting questions raised by the Vision Restoration blog, where the optimal crew size question is raised again, Mark repeats his old argument:

A blog calling itself “Vision Restoriation” has some questions it would like to pose to the Augustine Commission. But the first question made me roll my eyes and wonder whether the blog ought to have been named “Vision Gutting.”

…
One hardly knows where to begin. Shrinking the crew to two roughly halves what one can do on the Moon for not, I would think, a lot of savings. And let’s just imagine the reaction of the Vision’s stakeholders, including Congress, the scientific community, and the new space entrepeneurs.

On the other hand, maybe we can shrink the crew to zero, make a movie about returning to the Moon, and save some real money. And people wonder why I can’t take these internet rocketeers seriously.

The first and most important flaw in this argument is the assumption that shrinking the crew per landing would not amount to a lot of savings.  It’s a nice opinion, but shows a complete and utter lack of understanding of the physics of lunar transportation.  The lander and capsule masses drive the IMLEO requirements for a lunar mission.  Halving the crew requirement would reduce the IMLEO requirement substantially.  Maybe not quite by half, but probably by at least 40%.  More importantly, doing it that way eliminates the need for big new boosters which are slated to use up something like 2/3 of NASA’s “exploration” budget over the next 10-15 years.

More importantly, the current ESAS-derived architecture is already presenting us with the “Incredible Shrinking Moon Program” that Mark bewailed back when I first raised this point 3 years ago.  Since that time, Orion and LSAM’s capabilities have been cut back substantially, Orion has gone from 6 crew to 4 to the ISS, and Ares V still doesn’t close performance-wise, and it’s already getting to the limits of its growth capacity.  If we continue down our current ill-thought-out path, there’s a very real chance that we’ll end up with a 2 or 3 person crew, just at the cost of a 4-person mission.

I’m not positive that a two-person mission is the right way to go, but handwaiving it away seems kind of lame.  With the kind of analysis that the Augustine Panel is doing, this is the kind of question that they should be asking.  Even if they come to the conclusion that 4 people is about right, that’s something that should be investigated, not just assumed.

Now, I have a lot of respect for the DIRECT guys, and have to give them kudos for even mentioning depots, but the fact is that this argument for heavy lift isn’t anywhere near as solid as it appears on the surface.  The fact is that there are many ways you could use dry-launch/propellant depot techniques to do ESAS-sized missions without the need of a big fairing or an HLV.  It may be true that you can’t cram the current Altair conceptual design into a Delta-IV fairing, however there are several legitimate alternatives out there to the current Altair concept, that can do the same job without requiring anything bigger than the already massive 5m fairings that come with existing EELVs.  And it’s important to remember that Altair is still in the early conceptual phase, where even fairly significant changes don’t really cost that much yet.

I want to keep this post brief, so I’ll just list a few options out there:

1. Horizontal landers: I don’t have the latest numbers, but the most recent numbers I’ve found show the LSAM descent stage holding only a little more propellant than a Centaur stage.  A horizontal lander of the type described by ULA in their papers could easily fit within the fairing of an EELV.  Even more so if based on 5m diameter tanking.
2. Crasher stage landers: It’s possible to split a lander up such that the descent burn is mostly done by a “crasher stage”, which is dropped shortly before the final landing burn (with shortly possibly being over a minute before).  This means that your actual lander stage can be a lot smaller and more compact.
3. EDS TLI burn: There’s nothing that says the LOI burn has to be done by the LSAM.  An EDS that’s big enough to do the TLI burn can still be fit within an EELV fairing, especially if using a 5m diameter stage like the Common Upper Stage.  With the lander descent stage only having to do the actual descent burn, it can be a lot more tightly packed.
4. LLO depots: If you have a depot in lunar orbit that is regularly topped off, you can tank up the lander stage after doing the LOI burn and before doing the landing.  This allows for less tankage, since you don’t have to size the lander for both LOI and lunar descent.  Alternately, if you have a reusable lunar lander (ie an SSTO lander designed to work with depots), you can send that lander independently from LEO to to the lunar vicinity.
5. L2 rendezvous: Having the CEV separate from the lunar stack prior to LOI, and then perform its own powered swingby maneuver greatly reduces the size of the lunar stack that needs to perform the LOI burn.  At this point having either the EDS or the lander do the burn allows for a much smaller lander.

And the list could go on and on.  Basically, so long as you don’t stick to “black aluminum” lander and transfer stage strategies, you can actually use depots to enable ESAS-equivalent landings without needing HLVs or big payload fairings.  There may be other arguments for big fairings (Mars reentry shields if we can’t get hypercones or rocket decelerators to do the trick, other large payloads, who knows), but this isn’t it.