2-Man EELV Based Mission Numbers

One of the commenters on last night’s post wanted to see some more information on how I came up with the results I’ve been discussing. Well, I finally broke down and figured out how to use Google Spreadsheets, and here is the spreadsheet with the details.

The surprising thing I found was that if you assume LOX/LH2 for the lander using an RL10 as a baseline engine (since they are off-the-shelf, and have been used for throttling and VTVL applications before), that you can actually do a two-man sortie using only two EELV launches. I usually have done most of my analysis assuming the landers use LOX/Kerosene or LOX/Propane, as those are closer to my experience base with LOX/alcohol. However, I was trying to come up with an architecture that used existing boosters and stages as much as possible, and I found that LOX/LH2 for the landers actually makes that feasible.

The other thing that made this architecture possible was using Weak Stability Boundary trajectories for the unmanned landings. I first heard of WSB trajectories from the t/Space CE&R report. These rather complicated trajectories use multiple lunar encounters over a 3 month period to more or less eliminate the need for the Lunar Orbit Insertion burn. The required Delta-V from LEO to LUNO drops from ~4.2km/s using a normal 3-day Hohmann transfer to ~3.2km/s using a WSB transfer. While 3 Months in deep space is a long time, the Centaur guys at Lockheed have design ideas for keeping boiloff reasonable over that long of a period. This enables the cargo launches to also be two-vehicle missions.

The basic architectural blocks for this system are:

  • The Single Engine Centaur: This is a stock SEC as used on the Atlas V launch vehicle, with a lunar-mission kit attached. This mostly consists of things like solar cells, long-duration navigation hardware (stuff like star-trackers), extra MLI on the tanks to reduce boiloff, and things like that. All of these have been under development there at LM.
  • The Single-Stage Multi-Purpose Lander: This is a light, LOX/LH2 lander based around a similar construction technique to that proposed for the Wide Body Centaur. The lander can carry either a 5klb crew capsule on a round trip, drop-off a 15.5klb cargo pallet on the lunar surface (with the lander returning to orbit), or haul a full 19.9klb Sundancer module all the way to the lunar surface (with the lander staying on the surface either to be cannabalized, or to be refueled from a cargo landing). The lander itself is reusable, so eventually you only have to send propellants for it instead of having to ship the lander ever time.
  • The Crew Capsule: This is a light, 5klb crew capsule. The Lunar Module Ascent stage weighed less than 4500lb, including the weight of the propellant and pressurization tanks, the main ascent engine, and a bunch of 60s era electronics and batteries. Using modern materials and electronics, this capsule can probably be quite a bit roomier than the LM was, though they will likely still be a bit cramped. This capsule includes some RCS engines, a reentry heat shield, and parachutes.

For each of the three mission sorts (manned, Sundancer landing, or cargo dropoff), a single Delta IVH places the fully fueled Centaur stage into orbit, and then an Atlas V launch (a man-rated 401 for crew, or a 531 for either Sundancer or cargo delivery) places the rest of the mission payload into orbit, where they rendezvous and dock. The Centaur then performs the TLI burn. In the case of the manned missions, the Centaur also performs the LOI burn, and the TEI burn after the crew returns from the surface. All of the missions were designed assuming at least some boiloff, and including some extra Delta-V to make up for contingencies.

This isn’t the absolute most perfect architecture in the world, but it’s an inexpensive and simple one that gives a lot of flexibility. The ideal architecture would probably use fully reusable landers that had multiply redundant engines, that used WBC derived reusable transfer stages, and that used both on-orbit propellant transfer and lunar ISRU derived propellants to continually drive costs down while driving performance up. But this is a simple, straightforward, low-risk, near-term feasible technical solution.

Between existing Atlas and Delta IVH capabilities, up to 10-12 missions per year could be accomplished without adding new pads or infrastructure. At that flight rate, a lunar sortie would cost about $300M for a manned or a cargo mission, and probably an extra $100-150M for a Sundancer mission. And that’s assuming no Wide Body Centaurs, no on-orbit propellant transfer, and no new lower-cost launchers. If on-orbit propellant transfer comes on-line, you can get rid of the Delta-IVH launches and extra Centaur stages and just refuel the Centaur used for launching the capsule and lander. That would cut the mission costs about in half to $160M for the manned and cargo missions and $260-310M for the Sundancer mission.

So, ignoring the on-orbit propellant transfer for a second, we find that for just the fixed annual costs of the ESAS architecture, you could put:

  • Two Sundancer modules down at 1-2 locations
  • Four Two-man teams at 1-4 outpost/sortie locations
  • Three 7mT cargo pallets down at 1-3 different outpost/sortie locations

The coolest thing is that you only have two pieces of hardware that need to be developed for this architecture: the capsule and the lander. While they are a bit sophisticated, neither of them ought to take more than 2-3 years to field. Which means that you could start having lunar missions again this decade.

Mark Whittington likes to poke fun at a 2-man architecture calling it the “Incredible Shrinking Moon Mission”, but I kind of prefer it to the “Rapidly Receding Moon Mission Schedule” which promises Apollo-on-Steroids but keeps pushing the landing date out a year for every year of real time. When Bush announced it, the first landing was supposed to be in 2015. Then it was 2018. Now it’s possibly as late as 2020. So what started as an 11 year schedule when Bush proposed it is now a 14 year schedule in spite of the fact that it’s now two years later…Me, I’m a fan of Lunar Sooner (and Better).

Anyhow, whoever the anonymous commenter was, I hope he enjoys the spreadsheet.

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Jonathan Goff

Jonathan Goff

President/CEO at Altius Space Machines
Jonathan Goff is a space technologist, inventor, and serial space entrepreneur who created the Selenian Boondocks blog. Jon was a co-founder of Masten Space Systems, and is the founder and CEO of Altius Space Machines, a space robotics startup in Broomfield, CO. His family includes his wife, Tiffany, and five boys: Jarom (deceased), Jonathan, James, Peter, and Andrew. Jon has a BS in Manufacturing Engineering (1999) and an MS in Mechanical Engineering (2007) from Brigham Young University, and served an LDS proselytizing mission in Olongapo, Philippines from 2000-2002.
Jonathan Goff

About Jonathan Goff

Jonathan Goff is a space technologist, inventor, and serial space entrepreneur who created the Selenian Boondocks blog. Jon was a co-founder of Masten Space Systems, and is the founder and CEO of Altius Space Machines, a space robotics startup in Broomfield, CO. His family includes his wife, Tiffany, and five boys: Jarom (deceased), Jonathan, James, Peter, and Andrew. Jon has a BS in Manufacturing Engineering (1999) and an MS in Mechanical Engineering (2007) from Brigham Young University, and served an LDS proselytizing mission in Olongapo, Philippines from 2000-2002.
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27 Responses to 2-Man EELV Based Mission Numbers

  1. Brad says:

    “The Crew Capsule: This is a light, 5klb crew capsule. The Lunar Module Ascent stage weighed less than 4500lb, including the weight of the propellant and pressurization tanks, the main ascent engine, and a bunch of 60s era electronics and batteries. Using modern materials and electronics, this capsule can probably be quite a bit roomier than the LM was, though they will likely still be a bit cramped. This capsule includes some RCS engines, a reentry heat shield, and parachutes.”

    I hate to be the bearer of bad news, but I think many of your capsule assumptions need work.

    1st off, the Grumman LM ascent stage had a mass of 4,500 kilograms, not 4,500 pounds. (according to astronautix.com)

    Secondly, a better analog for your 2-man multi-purpose capsule is the Gemini spacecraft. And even there the Gemini reentry module alone weighs more than 4,300 pounds. To that should be added all the mass needed for power and life support for 14 days. For that the reentry capsule needed the equipment modules which upped the total Gemini mass to 3,851 kg.

    Cutting 40% from the mass of the Gemini spacecraft to fit the 5,000 weight limit would be very very tough. You could get rid of much of the heavy electronic equipment, much of the OMS propellant and much of the associated structure. But something has to provide power and life support. So I don’t see any practical cutting elsewhere. The Gemini was a pretty lightweight spacecraft using a fuel cell power system (which produced water for the crew too) and a titanium capsule structure.

    An umbilical from your lunar lander to your capsule could provide power and life-support, though those functions would cut into the mass budget of the lander. But what about the return trip when the capsule transfers from the lander to the centaur stage? Could the capsule contain enough supplies for the 3 day journey home? Or would the capsule have to establish a docked supply link to the centaur stage?

  2. Jon Goff says:

    Brad,
    Read more carefully. That 4500kg was the *wet* mass of the ascent stage. Once you subtract off the propellants (a hurking big about of N2O4, and Hydrazine), you’re around 5000lb. There’s a more detailed mass breakdown at Wikipedia: http://en.wikipedia.org/wiki/Lunar_module

    Yes a Gemini capsule was a little bit heavier all things told than what I’m talking about, but yes it’s still perfectly feasible. Electronics, materials, and power sources (think solar panels with some batteries) have gotten much, much better over the years. I know someone who thinks that 1000lb/person is a reasonable goal. 2.5klb per person is perfectly attainable.

    The lander is just a lander that that the capsule docks to. But the lander, the capsule, and the Centaur stage will all likely have on-board solar power, RCS capabilities, etc.

    As I said, it’s a bit aggressive, but probably doable. If the LM could support 3 people for 6-7 days (Apollo 13) for only about 5000lb, I think with a more modern system we can do better.

    ~Jon

  3. habitat Hermit says:

    I thought the anonymous post was sort of trollish but who cares when the reply is as interesting as this? Way to go Jon! 🙂

  4. Bill White says:

    Jon, what is your assumed retail price point for a Delta IVH?

    Re-run the cost analysis using Proton or Falcon 9 and see how low you can go.

  5. Mark says:

    The last paragraph was cute, but also false. The planned first mission was always “before 2020”, with guesses of it happening anywhere from 2016 to 2019.

  6. jv says:

    Mark, to quote the VSE speech:
    – Our second goal is to develop and test a new spacecraft, the Crew Exploration Vehicle, by 2008, and to conduct the first manned mission no later than 2014.
    – Our third goal is to return to the moon by 2020, as the launching point for missions beyond. Beginning no later than 2008, we will send a series of robotic missions to the lunar surface to research and prepare for future human exploration. Using the Crew Exploration Vehicle, we will undertake extended human missions to the moon as early as 2015, with the goal of living and working there for increasingly extended periods.

  7. Gaetano Marano says:

    .

    A size-reduced lunar hardware is the BEST way to accomplish the VSE moon plan FASTER and CHEAPER. That can be done also with low cost and ready available Shuttle-derived hardware as I’ve xplained seven months ago in my article “Single Launch Vehicle NOW! And “On The Moon” in 2012 !” here: http://www.gaetanomarano.it/articles/005_SLVnow.html

    .

  8. Jon Goff says:

    Mark,
    I’m just going off of what Bush said in his VSE speach:
    “Using the Crew Exploration Vehicle, we will undertake extended human missions to the moon as early as 2015, with the goal of living and working there for increasingly extended periods.”
    http://www.whitehouse.gov/news/releases/2004/01/20040114-3.html

    ~Jon

  9. Jon Goff says:

    Jv,
    Oops, you beat me to the punch.
    ~Jon

  10. Jon Goff says:

    Bill,
    I was assuming that you’d be flying a few of these per year (at leat four or five missions), and that thus it would be getting back closer to the original $170M range.

    Falcon IX could do the crew/cargo launch, and that would drop the per mission pricetag by about 10% overall.

    Proton doesn’t quite have the LEO mass to put the transfer stage up–it’s about 5000lb shy. You could use it for the unmanned missions, but once again, it wouldn’t save you very much. If it could displace the Delta IVH, it could cut things down a bit, but it’s about 10% short of the performance needed.

    Now, if you can do on-orbit propellant transfer, things change quite a bit (as I mentioned in the post).

    ~Jon

  11. Will McLean says:

    Jon writes:

    As I said, it’s a bit aggressive, but probably doable. If the LM could support 3 people for 6-7 days (Apollo 13) for only about 5000lb, I think with a more modern system we can do better.

    Except the LM ascent stage couldn’t. It could only support two men for a few hours. Supporting two men for three days required power, oxygen and cooling water from the descent stage. Stretching it further for Apollo 13 required the same, plus stringent rationing of power that would never have been attempted if there was any other choice.

    Removing the ascent stage engine doesn’t save you as much as you might think (about 100 kg.) Neither does downsizing the propellant tanks (you still need some mass for RCS tankage).

    Reality check: If you can build a 5,000 lb capsule that can provide life support to the moon and back for two people, then you should be able to build a four person capsule to do the same job for less than 10,000 lbs. The larger capsule should mass less per seat, since some of your components, such as avionics and hatches, will mass the same the same even though you are carrying twice as many crew.

    This is about half what NASA is estimating for an Orion capsule to carry four to the moon and back. For a fair comparison, you need to apply the same optimism to both plans.

    By the way, do you realize that you have essentially reinvented General Dynamics’ Early Lunar Access from 1993, but with even more optimistic mass estimates?

  12. Jon Goff says:

    Will,
    Reality check: If you can build a 5,000 lb capsule that can provide life support to the moon and back for two people, then you should be able to build a four person capsule to do the same job for less than 10,000 lbs. The larger capsule should mass less per seat, since some of your components, such as avionics and hatches, will mass the same the same even though you are carrying twice as many crew.

    T/Space was claming ~9000lb for a 4-person CXV capsule. Lockheed was talking about an 8 person module for 20klb. It’s aggressive, but I’m pretty sure it’s doable. Electronics, power systems, etc are getting lighter and lighter every year. Manufacturing processes and materials are better now than previously. Etc.

    This is about half what NASA is estimating for an Orion capsule to carry four to the moon and back. For a fair comparison, you need to apply the same optimism to both plans.

    Yup, and Orion is extremely bloated, and much larger than it really needs to be. NASA is intentionally making it more spacious and less cramped, which means that no, you shouldn’t apply the same “optimism” for both plans. There’s a big difference between between design optimism and designing with different requirements in mind.

    They’re talking about using CEV for long-duration exploration to places like NEAs and such. That’s a lot more than you need for a cramped, bare-bones lunar transportation system.

    By the way, do you realize that you have essentially reinvented General Dynamics’ Early Lunar Access from 1993, but with even more optimistic mass estimates?

    Well, that’s good. Its about time that aerospace, like every other industry on the planet actually advances in capability over time. It’s 13 years since 1993, one would really hope that the state of practice has improved a bit since then. But it’s also important once again to remember, that just because GD said one mass back in 1993, doesn’t mean that that was the absolute lightest possible, or even that it was being designed to be the lightest possible.

    Just comparing designs to what was done years and decades ago doesn’t really prove or disprove the feasibility of new designs.

    ~Jon

  13. Big D says:

    Why fly a heat shield to the moon?

    Wouldn’t it be more efficient to dock back at a station and take a dedicated E-LEO capsule back down, and re-use the capsule/LM stack each time?

    Or does that go back to avoiding refueling or any other infrastructure for the purposes of this exercise?

  14. Peter says:

    Leaving the heat shield and some of your fuel in lunar orbit or L1 then picking them up on the way home saves the fuel needed to land the mass on the moon then lift it again, at the expense of operational complexity and risk. For higher flight rate and longer term operations, having 1 vehicle for LEO-L1 transfer and another for LI-Lunar surface becomes attractive.

  15. Brad says:

    Hey Will,

    Thanx for the tip about the 1993 ‘early lunar access’ plan by GD. I hadn’t heard of it before, very interesting, and the plan provides another data point for a 2 man lunar mission Earth reentry capsule.

    The GD capsule, based on a full sized Apollo capsule but designed for two men and using up to date equipment and materials, had a gross mass of 3,688 kg.

    http://www.abo.fi/~mlindroo/Station/Slides/sld051e.htm

  16. Brad says:

    Jon,

    “Read more carefully. That 4500kg was the *wet* mass of the ascent stage. Once you subtract off the propellants (a hurking big about of N2O4, and Hydrazine), you’re around 5000lb.”

    I did read carefully. I considered that maybe you were refering to dry weight, in spite of what you wrote.

    “The Lunar Module Ascent stage weighed less than 4500lb, including the weight of the propellant and pressurization tanks, the main ascent engine, and a bunch of 60s era electronics and batteries.”

    So just in case, I examined the possibility you meant dry mass. I calculated the dry mass of the ascent module by subtracting the mass of the propellant and came up with weight over 4,800 pounds, while you wrote of a weight of less than 4,500 pounds. I concluded you merely made a simple misreading mistake that slipped by you.

    But what you intended to say about the ascent module mass is not important. The main case I was originally trying to make is that the lunar ascent module is a bad example for comparison to your lunar capsule. Other than crew size there is hardly anything comparable to the job of a reentry capsule.

    “Yes a Gemini capsule was a little bit heavier all things told than what I’m talking about, but yes it’s still perfectly feasible.”

    8,472 pounds is more than a bit heavier than 5,000 pounds. That’s 69% heavier!

    “I know someone who thinks that 1000lb/person is a reasonable goal.”

    For a lunar mission? I would like to hear more about that. Enough life support plus a heatshield that copes with return from luna? And that scales down for a 2,000 pound capsule? I’m very skeptical.

    “…2.5klb per person is perfectly attainable…T/Space was claming ~9000lb for a 4-person CXV capsule. Lockheed was talking about an 8 person module for 20klb.”

    Yes and the SpaceX capsule is less than 20,000 pounds for 7 people. But those spacecraft are all designed for short duration LEO missions. I doubt they are built for 11 km/s reentry speeds or have 7+ days of life support. And even for LEO missions those capsules might turn out considerably less capable than the current claims, with weight growth forcing crew reduction.

    Looking over the small scale 2 man lunar direct missions from the Apollo era, it’s interesting that both the lunar version of the Gemini capsule and the small scale 2-man Apollo capsule massed over 5,000 pounds. And that’s not even counting the mass of the power system and life support provided by their service modules.

    ————————————————————————-

    Now here is a vital issue…why should your lunar architecture use the reentry capsule on the lunar lander?

    If I understand your plan correctly, the manned lander only transports the crew because there is no margin in the architecture for anything greater. Which is why surface rendezvous with the sundancer habitat is key to the plan. Right?

    So that being the case, why haul down to the lunar surface (and then back up to LLO) the mass of the reentry capsule? The key factor is – get to the surface and then transfer to the sundancer. So a small rover carried down by the lander would be more important to LSR success than a capsule would, even if the crew had to fly the lander with no more protection than EVA suits.

    The only sensible reason I can see for hauling a reentry capsule down to the lunar surface is if the lunar lander will fly a direct return to Earth. Just as the various lunar direct architectures do. But your plan uses LOR, right?

    I think a more practical plan is to leave the heavy reentry module in lunar orbit still attached to the centaur stage, and use a simple lightweight pilot cabin on the manned lander which is not much more than a pressurized bubble tent. Maybe even an inflatable cabin.

  17. Will McLean says:

    Jon writes:

    They’re talking about using CEV for long-duration exploration to places like NEAs and such. That’s a lot more than you need for a cramped, bare-bones lunar transportation system.

    That wasn’t a driver for the design, which was based on a crew of four to the moon and back. Longer missions, such as Mars, assumed that the capsule would be docked to a larger spacecraft, and only had to be designed for habitability for the brief period from undocking to reentry.

    Your requirements may actually be more demanding, since the Orion CEV doesn’t need to provide volume to don moonsuits in gravity.

  18. Jon Goff says:

    Brad,
    Oops, I meant to say RCS propellant, not the ascent propellant (which wouldn’t be needed for this design). I think I had taken out the RCS propellant too for that less than 4500lb number. Adding that back in, you’re up to 4800lb, adding the 300lb of water from the descent stage and say 50lb of oxygen and other consumables, you’re up to 5150lb. Subtracting off the ascent engine gets you back to 4900lb, and subtracting off the main propellant tanks probably gets you back down into the 4700-4800lb range. So yeah, with 60s era technology, and big power and cooling hungry 60s era computers, the ascent stage isn’t a *great* analogy.

    As for your point that LEO spacecraft don’t carrey enough consumables for a long duration trip…I’m getting about 540lb of consumables for a 2 people over the course of 8 days. Most of that is sanitary water, which might be reduceable. 540lb is a lot, but a good chunk of that would be needed for a short-stay LEO vehicle. The Dragon capsule for instance was sized for 30 man-days, which comes out to about 4.3 days per person. Adding another 4 days per person would add about 125lb per person….not that much really.

    Cooling water and power have both been drastically reduced by modern electronics, and at with a LOX/LH2 stage on the ground with you, you might be able to use the boiloff LH2 to cool the electronics.

    ——————

    As for the idea of using a much lighter structure for the landing cabin…that’s not necessarily a bad idea at all…Let me look into that. If say we could cut it down to say 1600lb, that would drastically reduce the propellant needed for the manned lander substantially….

    Would a 8400lb 2-man capsule with a 1600lb Lander cabin sound better? I’m getting the numbers to work. Check out the new spreadsheet at:
    http://spreadsheets.google.com/pub?key=pD9i_cIb_0jC2LlfgSOQJAQ

    Anyhow, thanks for the suggestion, I think it has some real merit. Sorry for being a bit on the confrontational side.

    ~Jon

  19. Anonymous says:

    Jon

    Good work.

    I think that the whole point of your analysis is that a lot of time and money can be saved if a 2-man lunar architecture is utilized, and if existing technologies are utilized.

    You should drop development of that 5,000 lb capsule, and use an existing reentry capsule that weighs 6,000 lb to 10,000 lb, if lowering development costs and timeline are your goals. You should also drop the Earth Orbit Rendezvous (EOR) portion of your architecture, and just do one launch on a 35 ton to 45 ton to LEO capable launch vehicle if lowering costs and complexity is your goal. Lockheed Martin, Arianespace, Boeing, and SpaceX have all said that they will offer a 35 ton to 45 ton LEO capability on their existing launch vehicles by 2010 as soon as a customer pays for that mission. The only development costs and risk within your architecture should be the creation of an RL10-based single stage lunar lander that can handle Lunar Orbit Insertion, Lunar Descent, and Lunar Ascent. I think total development costs would drop to $100 Million and the cost per moon mission could eventually drop to as low as $100 Million per mission if you dropped the development of that 5,000 lb capsule and the requirement to launch on 2 vehicles with an EOR. The Europeans, the Russians, the Chinese, the Indians, the Japanese, and American Commercial companies like SpaceX/Bigelow/Armadillo would copy your architecture if you made these changes.

    The existing Soyuz capsule weighs 6,000 lbs and was designed (and successfully tested once in the late 1960’s) to return 2 cosmonauts from the moon safely. The Chinese Shenzou capsule, the new Indian manned capsule, the SpaceX Dragon capsule, and the new Soyuz-K capsule for the European/Japanese ACTS manned spacecraft should all weigh between 6,000 lbs and 10,000 lbs for the reentry capsule portion. The most important thing is that these reentry capsules will all be developed with other people’s money, so you do not have to include them within your own development costs.

    Lockheed Martin has a good paper on their web site of a single stage lunar lander that handles lunar insertion burn of the entire stack, including the 20 ton CEV, descent to the lunar surface with 4 astronauts, and ascent back to the CEV, all on one RL10 CECE engine. This lander would weigh in at about 40 tons using this one RL10 engine. If you resize this LM lunar lander for a 10-ton SpaceX Dragon or 10-ton Russian/European/Japanese Soyuz-K space vehicle, then you would get a 20-ton lunar lander that looks a lot like your lunar lander concept. Carmak and Armadillo Aerospace could probably purchase an RL10 from Pratt&Whitney for $5 million and they could use their existing software and avionics to build a mock-up of this RL10-based lunar lander for demo at the next X-prize cup within 11 months.

    If you search the Internet, you will see that the Ariane V is supposed to be upgraded to a 36 ton capability with a new restartable LH2 upper stage engine (this might be approved at the ESA ministers conference in December 2006), and that Lockheed, Boeing, and SpaceX all claim that they can easily do 40 tons to 50 tons to LEO with upgrades to the number of RL10 engines used on upper stage of the heavy versions of their rockets. Elon Musk has been quoted many times in saying that when he finds a customer that he will do a 100,000 lb to LEO version of his Falcon IX parallel stage vehicle that uses RL10’s on the upper-stage.

    Your architecture should involve a 10-ton manned space vehicle, the 20 ton lunar lander that I discussed above, a larger RL10-based upper stage (or the new European restartable LH2 upperstage for Ariane V flights), and an evolved heavy lift launch vehicle with around 40 ton capability to LEO. This 40-ton to LEO vehicle will lift the 10-ton space vehicle, the 20-ton lunar lander, and extra propellant on this larger upper stage (some like to call this the simple upper stage the “earth departure stage” but it really is just a standard upper stage with a higher propellant loading and a restart capability for the RL10 that could be a month into the future instead of the usual maximum time between first and last restart of 8 hours).

    You only have to do an analysis now of the delta-V requirements and propellant loadings to make this work. I do not know if you want to put more propellant in the upper stage or in the lunar lander to make all of the burns that you need.

    All of the development work with the above architecture surrounds the “what ifs” of what could be done with an RL10 engine. SpaceX, Bigelow, and Armadillo Aerospace are already investigating the “what ifs” surrounding RL10 upper stages and RL10 lunar lander designs so I would say that you could sell your analysis of this architecture to them.

    If Armadillo built an earth-bound mock-up of this lunar lander that used the RL10 engine, then they could probably start selling the lunar-lander version of this mock-up to any company or country that is interested in manned moon landings.

    Remember that Space Adventures is looking for $200 million to send 2 people around the moon, so there is a potential commercial market for this. If SpaceX could conduct a mission to the Moon using their $78 million Falcon IX heavy, their Dragon space vehicle, and RL10 upper stage, and this lunar lander, then they could probably profit on a $200 million price from Space Adventures. Bigelow could probably profit if their Sundancer module was used within the same architecture for later commercial time shares of “lunar condos” given to foreign nations at $100 million per astronaut.

    I think this all works if you just focus on what can be done with an RL10, drop the development of your 5,000 lb reentry capsule, and drop the EOR portion of your architecture using 2 separate launch vehicles. If you do these things you probably should be able to find a paying customer for your work. Great work.

  20. SMetch says:

    Jon, in early 2005 post Mike Griffin we detailed a similar approach to NASA management using upgraded EELV’s (30-35mT) with 2 launches EOR (EDS/Everything Else) approach. We also utilized an EELV upper stage as the LOA & PDI stage for the lunar landing separating short of landing for final descent using the CEV’s TEI engine. Our CM and SM were close in weight/dimensions to Apollo for a mission of two astronauts.

    We also showed the approach we detailed in the AIAA paper

    Here;

    http://www.teamvisioninc.com/services-consulting-space-exploration-optimization.htm

    The first all EELV approach was rejected because it would basically shut down major portions of NASA as we know it. When I suggested the China/Russia could pull this off in the near future using this approach they arrogantly discounted their ability to do this. I still think a Chinese/Russian Apollo 8 for 2008 is in work.

    The second approach (i.e. AIAA paper) was rejected because we used ELV’s for the CEV testing and ISS mission and later EML1 missions. As such the expense of the shaft was avoided and applied to transforming the SSTS to a HLV much sooner, i.e. Ross’s Direct approach which he has done a great job of detailing.

    Here;

    http://www.directlauncher.com/

    We are working on a new AIAA paper which will combine some of the best feedback we have received for the next Space 2007 conference. If only NASA could be this open to constructive criticism and alternate approaches as many in these forums are.

    I thought the last century firmly established how dictatorial and insular organizational systems were not as effective as the opposite.

    P.S. Jon double check your TLI numbers the lunar DV looks high for a 3 day transit.

  21. Brad says:

    Jon,

    Thanx for the updated spreadsheet, very interesting, you even managed to squeeze in a rover!

    And you are absolutely correct about the oversized Orion capsule crew module. Looking over the other 4 man architectures proposed by the various aerospace companies makes it obvious just how elephantine the NASA 5.5m (now 5m) diameter capsule is for the job.

    And apology accepted. With all the trolls about it’s all too easy to become hypervigilant about unfair attacks. And it’s natural to charge to the defense of the children of your mind. My intentions have always been constructive and I think you do solid and imaginative work. Sometimes it’s hard to provide the right tone with the printed word and I tried with my crappy writing skills to critique your proposal and avoid personal offense.

  22. Jon Goff says:

    Brad,
    And apology accepted. With all the trolls about it’s all too easy to become hypervigilant about unfair attacks. And it’s natural to charge to the defense of the children of your mind.

    Yeah. It’s never good when one gets one’s hackles up. You had a good idea, and I’m glad I wasn’t too pigheaded to see it.

    My intentions have always been constructive and I think you do solid and imaginative work. Sometimes it’s hard to provide the right tone with the printed word and I tried with my crappy writing skills to critique your proposal and avoid personal offense.

    Oh, I can empathize with the difficulty of trying to convey things without being offensive.

    Keep the comments coming though. You’ve also got good ideas.

    ~Jon

  23. Jon Goff says:

    Anonymous,
    I think I prefer trying to developing a new capsule over trying to field a new commercial heavy lift launch vehicle. Either is tough, but trying to field new super Atlases or super Deltas or super Falcons just sounds like a recipe for trouble. Sure, you could fit it all in one launch, but I just don’t really see the benefit.

    ~Jon

  24. Ed says:

    Sure, you could fit it all in one launch, but I just don’t really see the benefit.

    Well, that depends on what your goals are. What would you do if you were in charge of NASA and wanted to maintain space as your own personal fiefdom?

  25. Anonymous says:

    Jon

    The Soyuz capsule is 6,000 lbs, and was already tested for a manned flight around the Moon in the 1960’s. You can probably buy a Soyuz capsule from the Russians for $10 million and elimiante your largest development cost.

  26. Jon Goff says:

    Anonymous,
    Regarding the Soyuz, the capsule itself may only weigh in the ~6000lb range (Mark Wade says 6500lb), but that’s only part of the Soyuz vehicle. The full Soyuz vehicle is about 15000lb. You might be able to eliminate the PAO module in the back and just use the RCS on the lander and the Centaur for steering things around…and then the numbers for the TMA suggest a fairly large amount of cargo in LEO that wouldn’t be needed. And it is designed for three people and 14 days…

    It isn’t crazy, I just don’t have enough detail to know if it would work. Maybe I should put it up as a main post to see if other people have thoughts.

    If you could do a Soyuz for the crew capsule, then that leaves only the lander (and the lunar mission kit for the Centaur) that would need work…

    ~Jon

    ~Jon

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