Selenian Boondocks

Ok, two weeks ago, I mentioned that the “pre-depot” dual EELV launch concept my friend had passed to me could be adapted to do manned lunar missions.  Lunar missions are a lot easier to close if you assume a depot in LEO (and even easier if there’s also a small depot at L2).  But it turns out that if you use a couple of tricks, you can actually make a pre-depot concept close as well.  This wouldn’t be my optimal approach, but it at least illustrates the point.

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.

I recently found a fun presentation on lunar excavation technologies that I thought deserved a bit wider circulation.  I’ve actually been interested in lunar excavation for over a decade now (in fact, it played a role in leading me to my thesis topic, but that’s a post for another day), and I think that this presentation summed up a lot of my thinking on the topic better than I could.

The presentation was done by a company called Honeybee Robotics, that has been doing space robotics systems for over 20 years.  I found out about them through a joint project they had been doing with one of our new neighbors at the Mojave Spaceport, Firestar Engineering.  They had worked together on using the mixed nitrous monoprop that Firestar is developing as a gas generator for a lunar pneumatic excavation idea (which is discussed in this presentation, and also in this paper on using it for a Mars or Lunar sample return mission).  I’ve been interested in the concept of pneumatic excavation ever since I read an article titled “Foundation Slab for Lunar Base Construction” that was presented at the fourth ASCE conference on Engineering, Construction, and Operations in Space back in 1994.  The concept seemed to show a way of producing very large enclosed subselenian areas for relatively small initial investments in hardware and materials.  Anyhow, my curiosity on the topic led me to do some searches to see if I could dig up any more information about their Lunar Pneumatic Excavator concept, and that led me to the presentation I wanted to write about in this post.

I’d strongly suggest reading the whole thing, but here are my notes on the paper:

• Lunar regolith is highly compacted, abrasive, high surface friction, and sticks together very strongly.  All of these things drive up excavation forces.
• Most terrestrial excavation equipment uses the weight of the vehicle and the friction forces in its wheels/tracks to react against the excavation forces.
• Lower lunar gravity means that for typical lunar excavators, you might actually need several times heavier equipment to do the same job.
• Methods that lower the excavation force required can greatly reduce the mass requirement for the excavation equipment.
• Percussive/Vibratory excavation systems can lower the force and excavator mass requirements dramatically for such situations (up to a 10x mass reduction).
• Pneumatic excavation can be done using a coaxial tube setup, where the inner tube blows low-pressure gas at the regolith, and the slightly longer outer tube provides a return path for the gas vent.  Gas doing the turn from the inner tube to the outer annulus imparts momentum into the regolith, and then entrains the regolith particles in the gas forcing it up the outer tube (see illustration below)

• In an experiment performed on a vomit comet, they showed that a lunar pneumatic excavator system operating at 7psia, could give a regolith mass excavated to gas expended ratio of over 3000:1 in a 1/6g environment (ie each gram of gas could move over 3kg of regolith)
• Such excavation techniques can also be used for sample return or prospecting missions, using relatively simple hardware with almost no moving parts.

They didn’t go into it much, but if you could store the gas as a liquid (either inert, or as rocket propellant), and then vaporize or combust it before shooting it down the nozzle, you could excavate a pretty sizable amount of regolith using a pretty small volume of liquid.   That means that a tank about the same size as the propellant tanks we’re using on XA-0.2 (36in spherical tanks) could hold enough liquid CO2 to excavate enough regolith to bury a Bigelow Nautilus module.  Of course, to move the regolith back, you’d need another tank that big.  Call that about 700lb of CO2 and about 100lb of tankage.  Not bad if those numbers are accurate.

Anyhow, read the presentation and that other article, and post your thoughts in comments.

In case some of you are wondering why I’ve managed to blog four times today, I dropped Tiff and the boys off at the airport yesterday to go spend a month with her family. I’ll be batching it for the next three weeks, then joining them either right before or right after NewSpace 2009.

My usual sequence when batching it goes something like this:

1. Delusions of grandeur about how much I’m going to be able to accomplish without having to worry about taking care of kids and losing sleep from little kids waking me up 2-3 times every night.
2. Depression when I remember that I really love those little squirts and that I’m not going to get to rough-house with them for a long time.
3. Boredom as I realize that without someone else here to motivate me that I easily get stuck in ruts just surfing the web or playing computer games.
4. Mild insanity from missing said squirts (and Tiff too of course!)
5. Blogging quantity goes rapidly up, while quality rapidly drops
6. Workaholic phase as I no longer have someone reminding me to go home at night
7. Sleep deprivation and lack of personal hygiene kick in as I don’t have anyone to remind me to clean up the apartment, or to be reasonable with how late I stay up.
8. Oversleeping phase where my body starts forcing me to make up for me staying up too late in the previous phase.
9. Wake-up phase when I realize that I need to get the house and everything back in order before they get home.
10. Euphoria at finally having them all home again.

I think I’ve set a new record–it actually took over a day this time to get to step 5!

For your amusement, Ben explains:

I’ve found that the difference between “screwing around with heavy equipment” and “science” is whether or not you record the results with a high speed camera. This principal has been repeatedly demonstrated by the Mythbusters.

FWIW, I just noticed today that last week (on the 16th) was the four year anniversary of me starting this blog.  I can’t say my readership is that high compared to some, but nearly 300,000 hits and over 600 published articles over the past four years is still pretty impressive to me.  I’m not as prolific as many, but I’d like to think that my attempts at putting out some original thought on space technology and policy have earned me a happy niche in the blogosphere.

Here’s looking forward to another good year.

A few years ago, I asked the question of “how many crew do you really need for a lunar program?“  The conclusion was that if you could reduce the crew requirements (at least initially), it might allow for a much more capable, affordable, and flexible architecture.  And you’d eventually be back up to 4 or even 6 or 8 person crews as more infrastructure gets set up and in place.  The idea is unorthodox, but worth serious contemplation.  However, Mark Whittington pooh-poohed the idea back then as “the incredible shrinking moon program”.  His theory was that since we had already said we’d do 4 people, that if we switched to two people, that the program would lose face and risk being canceled.  Even if the two-person architecture actually allowed us to do more for less money, and sooner.

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.

With yet another episode of “let’s-just-quote-Jorge-Frank-because-he-puts-it-so-much-better-than-I-could”.  In response to a comment about how the problem with Shuttle was that it tried to be everything to everyone, Jorge said:

That was an effect, not a cause. The cause was the decision to make the shuttle an operational, rather than an experimental vehicle. This was key. An experimental vehicle would have been much smaller and would have had a cockpit and an instrument bay rather than a crew cabin and a payload bay.

The decision to concentrate on existing governmental (civilian and military) and commercial customer requirements was subsequent to that. At the time, the entire US launch market was around 50 per year. So in order to be economical, the shuttle had to be capable of meeting the requirements of all those customers. That led to the “too many cooks” problem you mention. But note that the problem could have been averted at *two* prior points: the decision for an operational vehicle (rather than experimental) and the decision to target existing markets (rather than postulate that the presence of an ultra-low-cost launcher on the market, even if that launcher could only carry small payloads, might drive the market toward smaller payloads).

Monte Davis, and several others, have made the first point–that the shuttle should’ve been an experimental vehicle, not an operational one.  But I think Jorge’s second point is even more interesting–that even though they made that first mistake, that they didn’t have to compound it by going for something at the first try that could meet all the nation’s *existing* launch needs.  Even if that would’ve meant a lot less political support (and hence money for development), if they had gone with something small, and focused on trying to enable new markets instead of trying to replace existing ones, I think things could’ve turned out a lot differently.  If they had only had 1/4 the development budget, but build a vehicle 1/10th as big, I think they could’ve actually delivered a fully reusable vehicle.  And even if it was a crappy, first generation reusable vehicle, that had lots of flaws, there’s still a good chance it would have been good enough to make a huge difference.

The DC-3 was revolutionary for its time even though it is far inferior to modern jet liners.  Even a Falcon-1 class RLV that could only fly twice a month per airframe would revolutionize the industry.

Just a thought.

guest blogger john hare

The multiple problems of solid rocket first stages would lead one to believe that the people that specified them would change their minds after they sobered up. That not being the case, it is somewhat interesting to think of ways to make it work anyway. The primary problems seem to be, excessive vibration, catastrophic failure modes requiring robust escape systems, control on all 3 axis, and poor scaling.

Apparently the upper stages and interstages are going to be tasked to handle most of these shortcomings. This Orion II idea is in response to people commenting on the Aries II idea. I agree that Direct, EELV, Falcon 9, depots, and a hundred other approaches are better than der Griffenschaft, but it is possible that the country will remain stuck with it. The context here is not supposed to be a better vehicle than the competition, but rather a way to get some lemon juice if even lemonade is not available.

Place an over sized H2/O2 tank on top of Aries with low pressure gas generator driven pumps pushing both propellants into the small 3rd stage tanks on Orion. The Orion rests on shock absorbers between the large tank and its’ RL10 clusters. Outboard of the stack are two or more large H2/O2 engines that light on the ground before Aries ignition. The large engines have sufficient thrust to keep the shock absorbers in tension throughout the Aries burn. The secondary function is  enough thrust to make up the performance shortfalls of the solid first stage.

 Three axis control is also supplied by these engines both during Aries first stage and their independent second stage burn after Aries drop off. In case of abort, the second stage tank becomes the shrapnel absorber while the Orion accelerates away on the large engines using the propellant that would have been used for the RL10s in a nominal mission. On a nominal mission, the large H2/O2 engines are dropped when the second stage tank is depleted, leaving the third stage with a full propellant load for its’ onboard RL10s.

The large propellant tank provides one layer of shock absorption from the solid rocket shaking. The large H2/O2 engines have enough thrust to keep the Orion clear of solid contact with the tank during the solids burn for a second layer of protection. The tensile shock absorbers between the tank and the Orion are the third layer of protection and should be capable of eliminating all but the most violent shaking.

The large H2O2 engines suffer considerable performance losses due to being canted outboard. These Isp losses on the order of 5% may be not too high a price to pay if it is necessary to save the concept. Gimbaling these engines can control the whole stack without much problem considering the relative sizes. During the first stage burn, they make up for the possible under performance of the Aries by providing several hundred thousand pounds of higher Isp thrust that would not normally be available to the concept. After Aries separation, they have enough thrust to carry the partially depleted second stage tank and the Orion spacecraft to near orbital velocity.

In case of abort, the second stage tank remains attached to the Aries to absorb shrapnel, while the large engines accelerate the bare spacecraft away using the onboard propellant that would have been used for the RL10s during a nominal mission. The large engines have enough thrust to escape at 4+G because they are sized to keep Orion clear of hard contact with the lower stages and carry the partially full second stage tank during a nominal mission.

On a nominal mission, the large engines are dropped with the second stage tank leaving the spacecraft to continue with the reliable vacuum rated RL10s. Internal tankage and number of RL10s would depend on final design details. One possibility is that the whole stack could be optimised for a lunar mission with abort to orbit on RL10s possible.

With the extra performance and escape routes available, this upper stage/escape system/spacecraft could be placed on the Atlas or Delta launchers with sufficient confidence to guarantee that either of them could perform the missions. With the safety aspects of the inherent escape system, relatively early flights of the Falcon 9 could do the job. Up and coming launch providers could be used as available and desirable for unmanned resupply at fairly early stages in development.

guest blogger john hare

I have been reading about some of the financial  numbers and technical problems of the Aries I for the last couple of weeks with more interest than usual. The idea that $8B has been spent already, and that total cost to first flight is projected to be $35B just boggles my mind. The laundry list of technical problems from under performance to extreme vibration should have stopped the program even before this price tag became known to NASA management.

If we accept as given that a shuttle derived vehicle with solid rockets was the only politically acceptable choice at the time, was it possible that something better could have been chosen? Something that would keep much of the shuttle work force and contracting organization together as possible? As a basher of der Griffenschaft, it seems possible to me that a somewhat similar architecture could have done better.

Suppose that instead of stretching and totally redesigning the SRB, they had shifted to a two SRB first stage with the exact same units as used for the last hundred and something shuttle launches. These rockets are fully developed and tested with an extensive flight history of operating in pairs over the last three decades. The purchase costs, handling , and performance are known quantities. Development consists of building an attachment structure, upper stage adapter, and vibration dampening gear. With the considerably more lift performance available from eight segments compared to five in Aries I, the problem fixing payload hits could be absorbed without sacrificing the flavor de jour safety systems NASA would like to have. They wouldn’t even have to game the requirements to match the competition from ULA, Direct, and various upstarts. While it’s possible that this would cost as much as the projected Aries I, it shouldn’t, and if it did, it would be for a system nearly twice as capable. 

The Aries I upper stage has another interesting budget feature in development of the J2S. This was supposed to be a simple upgrade of the flight proven J2 of Saturn fame. Now it seems to be a major development project in its’ own right. If the two SRBs were used instead, the upper stage would be much larger, perhaps too large for even the upgraded (read new engine) J2S. SSMEs, for all their expense, have a lot more flight history than the original J2 ever had. Modifying an SSME for second stage use would involve a new ignition system that would work in vacuum, and a much expanded nozzle for improved vacuum performance. Though SSMEs are expensive, a lot of them could be bought for the development price of the J2S, and the payload to orbit would about double.

I noticed during the DIRECT presentation at today’s HSF public meeting, that they were asked why they would need an HLV if they had depots.  Now, I didn’t hear the exact question, since I had a phone call come in just a few minutes earlier (ironically enough from one of my friends in the depot community), but I think the questioner was asking about drylaunching a lunar stack on an EELV and then tanking up at a depot.  The reply given was that while an EELV could loft the mass, it couldn’t handle the volume.

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.