Research Papers I Wish I Could Con Someone Into Writing Part I: Lunar ISRU in the Age of RLVs

One of the things I’d love to do if I were successful enough at Altius to afford it would be to sponsor graduate-level research into space technology, business, economics, and policy topics that I’m interested in. Not just because I don’t have time to dig into these topics as deeply myself as I would like, but also because frankly there are lots of graduate students out there who have better analytical tools they could bring to bear than the crude ones I could come up with informally. I decided to share some of these ideas via a blog post in the hope that maybe I could either inspire someone in grad school who is looking for a research topic, or if not I could at least plant the seed for conversation on this blog. If someone is interested in doing one of these research topics, I’d love to do a review of the final paper when it comes out.

Topic One: Lunar ISRU Economics In The Age of RLVs
This is one that I often discuss with coblogger Chris Stelter on Twitter. There have been a lot of papers over the years looking at ISRU economics, but the vast majority, if not all of them, have made the assumption that launch costs are more or less static. I think I understand the usual reasons for doing so–either a) these papers are trying to recommend a policy change, and therefore are being compared against the status quo approaches of say government exploration missions using entirely earth-launched propellants, or b) at the time of the papers, RLVs weren’t taken very seriously, and the last thing they wanted to do was to make ISRU look less respectable by making it look like it depended on RLVs.

But now is probably a good time to start looking into what lunar ISRU economics look like if you assume RLVs can be successful in driving down launch costs in the foreseeable future. I’ve seen a lot of SpaceX fans recently who have made the argument that lunar or NEO ISRU is totally irrelevant because BFR costs are guaranteed to be so cheap that there’s no way lunar ISRU could possibly compete with it. I think this is… mildly overoptimistic, but one result of lunar ISRU studies that assume status quo earth-to-orbit launch costs (both for launching ISRU infrastructure, and as competition) is that the lunar ISRU price points they quote really do seem kind of high compared to potential RLV price points. I personally don’t think lunar ISRU is in as much trouble as all that, but I do think that since it is more likely that we’re at the dawn of the age of the RLV, that those interested in lunar ISRU economics should at least start looking as RLVs become available.

Some thoughts on approaches:

  • Launch costs are going to vary over time–even if gas and go RLVs happen in the foreseeable future, it’ll still take time to get there. So instead of treating launch costs as static, make a few scenarios where you make different assumptions about the shape of the launch cost vs. time “S-Curve”. How long does it take for significant reductions in $/kg to start appearing? How low can they realistically get before hitting diminishing returns? How steep is the slope of $/kg over time once that initial decrease starts creating new demand that creates virtuous cycles? The nice thing is that you can probably characterize these S-Curves with only a few parameters, and then you can come up with say at least three scenarios–a pessimistic one where RLVs are only mildly successful, and launch costs decrease slowly, hitting diminishing returns at a moderate price point, an optimistic one, where RLVs are very successful, and the transition is fast, with the point of diminishing returns being dramatically lower than current prices, and then a middle of the road S-Curve shape.
  • Assume that lunar ISRU developers are smart enough to leverage RLVs as they become available, so that launch of ISRU hardware can take advantage of the decreasing costs over time. For example–George Sowers was mentioning a recent CO School of Mines analysis that showed it was possible to extract water from the lunar poles for $500/kg of extracted water on the lunar surface. But he was assume a $35,000/kg delivery cost to the lunar surface for all the infrastructure.
  • It would be interesting to see analyses that reflect the idea that lunar ISRU developers might be able to leverage decreasing launch costs to also lower the exploration and development costs of their lunar ISRU capabilities.
  • It would be good to include scenarios for how hard lunar ISRU ends up being, ranging from scenarios where trying to crack oxygen out of the regolith is the best we can do, through lunar polar ice being legit, all the way through Warren Platts’ lunar aquifers scenario. My guess is that this could also be modeled by some sort of S-Curve as well, as there’s going to be a learning curve for developing lunar mining, that eventually snowballs, but then hits diminishing returns, but the timing, depth, and steepness of the curve could vary.
  • It would be cool to see analyses that assume different cislunar transportation architectures for getting lunar ISRU propellants back to LEO. Not just rocket only, but also architectures that use propellantless launch options (see my unfinished “Slings and Arrows” series), aerocapture, SEP transfer, nodes at different cislunar orbital locations (LEO, EML1/2, LLO, etc).
  • It would be interesting to see with these analyses where the equillibrium point ends up being for lunar ISRU vs RLV-earth-launched propellants under different assumptions. I could see some cases (optimistic RLVs, pessimistic lunar resource difficulty, lame approaches to cislunar transportation) where lunar ISRU isn’t even competitive on the lunar surface, while there may be other scenarios, where lunar ISRU wins hands down even in LEO. But it would be interesting to see patterns and what assumptions lead to which outcomes.

Anyhow, I just wanted to seed the thought. I’ll probably turn this into a series for other research topics I’d like to see others write, but I wanted to throw this one out there.

Posted in BFR, ISRU, Lunar Commerce, Lunar Exploration and Development, Research Papers I'd Like to Con Others Into Writing, Space Development | 21 Comments

Flyback Shrouds

Seeing the picture of the SpaceX shroud floating in the water, it struck me how much it resembled the bottom half of certain lifting bodies. Then it struck me that this thought had been around before, whether it had been a passing thought, or a hint of memory of a past conversation.

My current thought is that the shroud inflates an upper body after separation and deploys a couple of vertical tails. Very fluffy flying reentry for a controlled glide to recovery. It seems possible that a very light propulsion unit of about 100-200 hp could extend the glide to an RTLS. Or possibly a aerial tow RTLS.

Does anyone know if this is a new (if not unique) thought or just a forgotten memory of a previous discussion?

Posted in Uncategorized | 5 Comments

Random Thoughts: A Now Rather Cold Take on BFR

When Elon gave his update on BFR at the IAC conference in Australia, I was originally going to post some thoughts right away1. But with Falcon Heavy’s maiden launch attempt coming up tomorrow, I realized I still hadn’t collected my thoughts about BFR into one place, instead of leaving them scattered over two dozen twitter arguments with Chris and others. I’ve had a lot of thoughts since the original announcement, but I wanted to share five thoughts that I had at the time that I still feel pretty strongly about:

  1. More Reasonable Size (Though Still Probably Too Big): I liked that Elon shrunk the size of BFR to something slightly less insane than ITS (a 50% drop from 300mT to 150mT). I still think he’s going way too big for any realistic markets near or medium term markets, but it’s a step in the right direction. I’m not convinced you need anything bigger than ~30-40mT to LEO to do Mars exploration and settlement, and you definitely don’t need anything that big to service near-term and future markets.
  2. Replacing Falcon 9, Falcon Heavy and Dragon with BFR/BFS: This was actually my favorite part of the plan in theory. In theory replacing the semi-expendable Falcon 9 and Falcon Heavy with a fully-reusable, and in-space refuelable launch vehicle would be a great idea. Especially one that was a single-stick, not crazy high aspect-ratio vehicle. And once you have that, and have the upper stage reliability up high enough, having an integral crew/passenger capability without needing a separate capsule could be a really powerful combination. Getting to high flight rate reusability is far more important for affordable deep space transport than getting to gargantuan rocket size. Something more modestly sized (say in the 30-40mT fully reusable range) would’ve been a much smaller leap, and I think would’ve much better taken advantage of the best part of Elon’s updated plan.
  3. BFR Leaves Open Room for Competition:At 150mT, BFR would be flying mostly empty on most flights for the foreseeable future. It would only really replace Falcon 9, Falcon Heavy, and Dragon if nobody else succeeds at doing a fully-reusable vehicle in a more sane scale. While it may be possible that a BFR sized fully-reusable launch vehicle might have much lower $/kg when flying completely full than a smaller sized fully-reusable vehicle using similar design architecture and technology choices, if BFR is flying mostly empty for most realistic near-term missions (satellite launch, ISS crew/cargo, etc), the actual cost to fly a realistic payload will probably be cheaper on a more right-sized vehicle. Personally, I think there’s a huge potential here for someone who wants to make a 1-10mT to LEO full RLV. While the $/kg might not be as good as a fully-loaded BFR or fully-reusable New Glenn/New Armstrong, the $/mission for most realistic near-term missions would likely be lower. I really hope someone else is able to raise money and execute on a small to medium RLV, I really don’t want to have to go back to launch vehicles for my next startup.
  4. Skeptical about Suborbital Point to Point: If you project BFR economics out to the point where it really hits some low multiple of the propellant costs, it theoretically could be competitive for some long-range travel. I just have a hard time seeing a rocket-based system with that high of performance and that razor thin of margins ever getting within spitting distance of the reliability of jet aircraft, especially within the foreseeable future. There are just so many technical and non-technical challenges for this market to make sense, and I think a lot of them are exacerbated by how big BFR is.
  5. What About Space Tourism? While I’m really skeptical about how realistic the suborbital point-to-point market for BFR, I’m kind of surprised Elon didn’t propose space tourism as a market. After all, if BFR can really keep 100+ people comfortable for a 6+ month Mars mission, you’d think they could easily handle 100 people for a 1-2 week stay in LEO. Even without space hotels as a destination, if he can really get down to a $200k/person Mars ticket using 5 launches, he should be able to get down to a $40k/person ticket for a two week space trip. If he was going to a space hotel and could pack people in as tightly as they were suggesting for BFR point-to-point suborbital flights, he could probably get the price for a LEO vacation down below $10k. While there are legitimate questions about how much market there is for space tourism at $20M+ per seat, is there really any doubt that there’d be a market for space tourism if it really cost only $10-20k per person for a 1-2 week LEO cruise?

Those were the five things that hit me the most. While I think BFR is an improvement over the original ITS plan, I think it still leaves a big opening for a serious competitor that didn’t feel the need to get into rocket size competitions with Elon and Jeff2.

Posted in BFR, Blue Origin, Random Thoughts, SpaceX | 49 Comments

AAS Paper Review: Practical Methodologies For Low Delta-V Penalty, On-Time Departures To Arbitrary Interplanetary Destinations From A Medium-Inclination Low-Earth Orbit Depot

I’d like to share a technical paper about propellant depots and interplanetary mission orbital dynamics that I helped co-author this past year, with the help of two of my astrogator friends1, Mike Loucks and John Carrico of Space Exploration Engineering and The Astrogators Guild blog.

By way of preface, this is a paper that Mike and I have been meaning to write for almost five years now2 as a rebuttal to some anti-LEO depot arguments that had started to come out back in the 2010 timeframe. You see, back when the FY2011 NASA Budget came out, those of us who had been advocating LEO propellant depots as a source-agnostic way of driving innovations in low-cost launch and Lunar/NEO ISRU thought we had finally won the day. Constellation had been cancelled. The president was proposing having NASA invest heavily in technology demonstrations for reducing to practice ideas like depots.

But then the idea of LEO depots started taking a lot of flack from many directions. Probably one of the most effective critiques of LEO propellant depots came from a NASA Flight Dynamics Office out of JSC who pointed out some orbital dynamics challenges of using LEO depots for doing interplanetary missions to places like asteroids. While I think the issues were raised in good faith–there are some legitimate challenges that LEO depots need to overcome–groups at NASA that didn’t want competition for their Monster Rocket used these arguments to “prove” that LEO depots weren’t really that useful after all, because you see, they weren’t useful for performing the asteroid missions NASA was now planning. It probably didn’t help that a certain one of its parent companies informed ULA’s depot advocates in no uncertain terms that depot was now a four-letter word that could be severely career limiting to use3.

The AAS Paper
Before I get into the orbital dynamics issue that was raised, and the solution we present in the paper, why don’t I share a link to the paper itself for those of you who would like to cut to the chase: AAS 17-696. This was presented at the AAS/AIAA Astrodynamics Specialist Conference which was held August 20-24, 2017 in Stevenson, Washington, U.S.A. If you’re interested, hard copies of this and the other presentations will soon be available in Volume 162 of Advances in the Astronautical Sciences. Fortunately AAS was fine with me sharing a copy of this on the blog so long as I gave proper credit.

The “Show Stopper”
So, unless you’re a real space nerd and happen to already know the answer, you’re probably wondering what the orbital dynamics issue was that NASA used as an excuse to ignore depots over the past five years. The tl;dr version of the issue is that while you can launch a LEO propellant depot in a way that lines up well for one specific interplanetary departure opportunity, it’s almost guaranteed to not be aligned well for most subsequent interplanetary missions you’d like to perform.

The longer version is driven by the concept of nodal precession. You see, because our planet spins, it’s a little… round about the middle. This bulge causes a “J2 perturbation” to orbits4, that basically causes the orbit plane to slowly precess around the earth’s rotational axis. I can’t remember exact numbers off the top of my head, but I think we’re talking 5-7 degrees per day for an ISS-like depot orbit. I could geek out on more facts about nodal precession, but here’s why that matters–once you launch a depot into LEO, you’ve established an orbital plane for that depot. That plane will precess over time in a very predictable way. The problem is that for a given interplanetary departure window, you need to leave earth on a specific departure vector (called the departure asymptote)5, and if the plane of your LEO parking orbit doesn’t intercept that departure asymptote vector at your departure date, you can’t do a coplanar one-burn departure, which means you have to deal with a very painful plane-change delta-V penalty that gets rapidly more painful the further the misalignment between your departure vector and your orbital plane. Another problem is that for weird destinations like some NEOs, the required departure asymptote is pointing off in a direction far from the equator–far enough that it might be higher than the inclination of a reasonable LEO depot, which means that you would never get a one-burn coplanar departure opportunity that didn’t require a debilitating plane change maneuver.

The good news is that since nodal precession is pretty easy to model, you can definitely place your depot into an orbital plane that will drift into alignment with a single departure opportunity. The problem is that the odds of it lining up with any given future arbitrary departure opportunity is pretty poor. Depending on how short the window is, and it tends to be really short for many NEOs, your odds might be less than 25% of having your depot in the right place at the right time. The odds of alignment are better if you’re talking planetary departures, as those tend to have longer windows at lower declinations (the angle from the equatorial plane to the departure asymptote vector), but still nowhere near the 100% you’d probably prefer for a critical piece of space infrastructure.

Ground launches don’t have the same problem because unlike an orbital plane that precesses very slowly, a ground launch site rotates with the earth 360 degrees every 24hrs. And a ground site can easily launch into a parking orbit with any inclination higher than the latitude of the site. Basically, you can pick the parking orbit of your departure to line up with your target every time.

I’m glossing over some details, but it all seems pretty damning when you think of it like that. Why build an expensive piece of infrastructure that you can’t count on being usable for any given mission?

Previous Proposed Solutions
Fortunately, there are some potential solutions that have been proposed in the past that still allow you to use LEO depots but get around this problem6. One option would be to have multiple depots instead of just one. If you have somewhere around 4-6 evenly-spaced depots, you should be able to always have a depot that will be aligned with a given destination at the right time, and you’ll know far enough in advance that you can logistically schedule different missions out of different depots. The problem is that this probably only makes sense if you have a really high flight rate. Another option proposed is to phase the depot orbit between missions–you can change your nodal precession rate by raising your apogee or lowering your perigee. If you know where you need to be far enough in advance, you don’t have to tweak your orbit too far to have it either precess faster or slower so it ends up in the right place at the right time. The problem with this approach is that it only works with really low flight rates, since it can take a long time to phase into the right orbit, especially if radiation or drag limits prevent you from changing your apogee or perigee too much. And neither of these solutions solve the problem if your departure declination is higher than the inclination of of your orbit, because the only solution at that point would be to do a plane change for your depot to get into a higher inclination orbit–and those are really painful maneuvers from a delta-V standpoint.

3-Burn Departures
While both of those solutions kind of work, Mike Loucks and I came up with a potentially better solution based on a MXER tether paper by someone who many Selenian Boondocks readers might recognize. You see, the best place to put a MXER tether is in equatorial orbit, because that maximizes the frequency of launch opportunities you have. But as I mentioned above, if the declination of the departure vector is higher than the inclination of your orbit, you’ll never have a coplanar one-burn departure. So what was Kirk’s solution? Use a 3-burn trajectory!

But first I need to explain something about departure orbits. Interplanetary departure orbits are by definition hyperbolic trajectories. And hyperbolic trajectories don’t behave quite like circular or elliptical ones. With a circular orbit, if you do an instantaneous (impulsive) burn, that point in your orbit where you did the burn becomes the perigee of an eliptical orbit, and the apogee will occur on a line that goes from that perigee through the center of the planet (that line is called the line of apsides if you’re wondering). So long as you’re still in an elliptical orbit this will always be the case. If you hit exactly escape velocity (i.e. have no excess velocity above escape), you’re in a parabolic orbit, which will point asymptotically as it approaches infinity in that same direction (i.e. parallel with the line between your perigee and the center of the earth). But as soon as you go a little faster, you’re in a hyperbolic trajectory, and the trajectory you end up asymptotically approach at infinity no longer aligns with that same line–the angle that a hyperbolic orbit has relative to that line between the perigee and the departure asymptote is I believe called the turning angle. And the turning angle gets bigger and bigger the faster you’re going relative to escape velocity. Combine that with the fact that any plane that is coplanar with the desired departure asymptote can allow a hyperbolic departure, and you get a famous drawing like that in the astrodynamics book by Bate, Mueller, and White (Figure 2 from both of the referenced papers) shown below:

Hyperbolic Departures (credit: Bate, Mueller, and White)

What you may notice is that there’s actually a ring, or “locus” of injection points (aka locus of periapses) that all allow you to get to the same departure asymptote. This ring, which is centered on the axis formed by the desired departure asymptote gets wider as the required injection velocity increases (higher C3), and narrower as it decreases.

Ok, so how does this enable a 3-burn departure that gets around the constraints described in the earlier section?

In order to do a coplanar departure without penalties, you need your perigee to be on that locus of injection points at the time of the final departure burn, you need the plane to be coplanar with the departure asymptote, and you need to be going in the right direction in your orbit. You can meet those three criteria if your LEO depot orbit plane happens to line up with the departure asymptote at the right time, but it turns out there’s a 3-burn trick you can do that allows you to meet those criteria for your final burn so long as your depot’s orbital plane ever crosses through any point on that locus of periapses at a point prior to your desired launch window.

Here’s how:

  1. At any time your orbit crosses through that locus of injection points, you do a large apogee raising burn that has an orbital period timed so that you return to perigee at the exact time you want to do the interplanetary departure burn7. This solves having your perigee on the right locus of injection points at the right time even if your LEO depot orbit has long since precessed out of optimal alignment. This burn also is not wasted as all of the energy you put into raising your apogee has come back as kinetic energy when you’re back at perigee for the final burn.
  2. Once you’re at apogee you’ll probably need to do a plane change maneuver to rotate your plane so that when you get back down to perigee your plane is coplanar with the desired departure asymptote, and you’re headed in the right direction. Since plane change costs are proportional to your velocity at the point in your orbit that you do them, they’re cheapest at apogee, especially for a high apogee near escape velocity.
  3. When you drop down to the perigee, you’re now lined up for your third and final burn which sends you to your destination.

If you’re having a hard time visualizing it, our AAS paper has a bunch of illustrations of what the trajectory looks like, including some practical mission examples, including an excellent acid test case provided by Josh Hopkins of Lockheed Martin (a mission to NEO 2007 XB23, which has a really crazy departure declination of -72 degrees).

Now there are all sorts of subtle nuances that can complicate this or make things better. For instance, because of how vectors add, you almost certainly want to split your plane change up between burns one, two, and three, because that’ll lower the overall penalty. Also, if your apogee is too high, you might start running into lunar perturbations that would have to be compensated for. Also, you might want to lower your perigee when you’re at apogee, to get a little more boost from the Oberth effect. Also, while your nodal precession rate drops off dramatically for a highly elliptical orbit, it’s not exactly zero. But all of these perturbations can be modeled and planned for when designing your 3-burn departure trajectory.

It’s worth mentioning that you can do this process in reverse to rendezvous with a LEO depot when coming in from an arbitrary interplanetary trajectory8.

While this is more complex than a single-burn departure, look what this does for you:

  1. You now can always hit your desired departure window even if your depot orbit itself is very misaligned with the asymptote.
  2. You can hit a departure asymptote even if the declination is higher than your inclination–it now just has to be lower than the sum of your inclination and turning angle, meaning that the higher the required departure velocity, the lower the inclination of your depot can be to hit a given departure declination. If your departure C3 is >16km^2/s^2, you can hit any departure declination from an ISS-like 51.6 degree depot orbit.
  3. It can also allow you to do missions that would require more propellant than your depot can handle. Basically, you can launch one tanker into this highly elliptical parking orbit that has your perigee on the desired injection point, long in advance of your desired departure window. The nodal precession rate of this highly elliptical orbit will be almost zero, so the depot orbit will rotate into alignment with it approximately once every 2-2.5 months. Each time it aligns with this parking orbit with the tanker in it, you can launch another tanker to rendezvous with, and add propellant to the first one, until it’s all the way full. Once that’s done, you can expend, or better yet aerobrake these depleted tankers back into LEO for reuse. Then when you’re at your last perigee before departure, you can launch the actual mission stack into the highly elliptical phasing orbit, rendezvous with the now full tanker, transfer propellant from the tanker, and then do your plane change burn at apogee, and your departure burn when you’re back down to perigee. In this way, or with variations on the theme, you can do really impressive missions using relatively modest sized depots9.
  4. And you get all of these benefits without having to do large numbers of depots and without having to move the depot, so you could theoretically park the depot in something like a resonant orbit that makes refueling logistics much easier10.

There are a few drawbacks or complexities, but they’re mostly minor:

  1. You do get two or more extra passes through the Van Allen Belt per departure mission. Not the best thing in the world, but not the end of the world either.
  2. You do still have some plane change penalty11, unless you can perform the first burn at a time when your LEO depot orbit is coplanar with the the departure asymptote12.
  3. You add a non-trivial amount of time to the mission, possibly on the order of weeks.
  4. You now have to do at least two burns a very long time after the first burn. Most current rocket stages are only designed for mission durations of 14hrs or less. Which means if you don’t have a long-lived stage (like ULA’s planned ACES stage), you probably need to do the second and third burn using something storable, which does cause a slight performance hit. In most cases the first burn is the biggest of the three though, so you still get some benefit from having your higher performance stage even if you have to have a kick stage of sorts for the final departure.
  5. You have more mission complexity and three important burns rather than just one for departure. But the nice thing is if either the first or second burn fails, you can probably abort the mission.

One “turning lemons into lemonade” advantage of this approach is that you get more time to check out your vehicle before its committed to interplanetary space. Once you’ve done that final departure burn, most systems really have no way to abort if something goes wrong. And with the bathtub reliability curve most complex systems have, it might actually be better to have an extra 2-3 weeks of checkout time during the elliptical phasing orbit to make sure everything is really ready for commit, with the option of aborting the mission if its not. Another side benefit of the elliptical orbit is that they have much lower (likely 10x lower) cryogenic boiloff rates, since you spend more time far away from nice warm planetary bodies. And you spend most of your time away from LEO where the MMOD environment is much, much better, which means that if you want to build up a mission stack over time (in a similar manner to the tanker concept discussed above), you can do so with less worry of MMOD damage to your mission stack while it waits than if you built it up in LEO.

While there’s still a lot of work to better evaluate optimizations and tradeoffs, and to explore various architectures enabled by this 3-burn departure (or arrival) method, we were able to identify and demonstrate a method that allows you to reuse a LEO depot for multiple missions, in a way that can always hit the desired departures at minimum penalty, even in spite of the previously raised issues due to LEO depot nodal precession. I wish we had been able to present this four or five years ago when we first discovered the solution. Maybe that would’ve at least dispelled the notion that orbital dynamics are a show-stopper for LEO propellant depots.

Posted in Orbital Dynamics, Propellant Depots, Space Transportation, ULA | 12 Comments

Black Aluminum New Car Tech

Over on Transterrestrial, there was a discussion about self driving cars eliminating traffic jams. The site doesn’t allow comments from my computer calling it through a proxy, so I’ll just throw out a few thoughts here.

The discussion focused mostly on cars accelerating from a red light and the amount of time saved when/if they all accelerated together as soon as the light turned green. Some felt that a lot of time would be saved, and some felt that the computer would have to wait for a safe braking distance just as with a human driver with the only savings being a few seconds here and there.

I saw no discussion of timing the lights that some of us do a bit of already. You see a red light in the middle distance and slow enough that it should turn green by the time you get there. So you might be driving 20 mph in a 30 mph zone by the time you reach the now green light. However, you are doing 20 mph through the light when the hurry up and stop drivers are still accelerating from zero. If done right, it is a faster and smoother trip using less fuel and much less wear and tear on the drive train and brakes. By missing this possibility, many people will be in the same position as the composite people that treat their product as black aluminum and fail to get the potential performance gains possible.

With self driving cars it seems likely that the car will know when the light is going to turn green, and if cars coming the other way are likely to do something stupid. It will also be plugged into the local traffic management systems so as to know the speed the lights are timed for. Many places have had their lights set up in such a way as to allow drivers to hit continuous green lights as long as they maintain some exact speed.

With self driving cars interacting with each other there will also have to be early fault detection systems such that cars with problems can be allowed for and given room to recover, pull over, or get out of the way so that the rest can maintain a steady flow. Follow distance with fault detection could be drafting close even at highway speeds.

So my thought is that a mature system of self driving and interacting cars interfaced with traffic control systems will have a flow pattern occasionally glimpsed but seldom maintained. Cars in city traffic will cluster at a speed that allows them to all get through the green light without stopping at all. I timed a light today that had a minute of green. allowing 2 seconds of margin on each end would give 56 seconds of full flow under the control systems I think likely. 56 seconds of cars with a 30 foot spacing center to center at 30 mph would put over 80 cars per lane through one green light, theoretically of course. That is at least 4 times the number I normally see get through per light change.

I think the focus on red lights missed the possibilities of the green lights. I think many of us are going to get blind sided by, not the technology itself, but the way it changes the way we operate. I tend to believe this extends to space far more than we would like to admit.


Posted in Uncategorized | 13 Comments

The Rivers of Progress

Most if not all people disagreed with my thoughts on bringing in various technologies in the last few posts about the architecture Doug Plata is laying out at On many blogs, and especially with many of the commenters  on those blogs, I could dismiss their concerns almost out of hand and be right most of the time. On this blog, and with the known high quality of the majority of the commenters disagreeing with me, there is necessarily something else going on. Obviously I could be wrong with them pointing it out. Or there could be misunderstanding as I am not a professional writer that is crystal clear in laying out my ideas.

There is another possibility that came up in the email exchanges with Doug. How many of the disagreements spring from variations in  the ideological base of the individuals involved. Life experiences and historical knowledge of space development does not resolve our differences. It can explain some of them so that we can move forward in developing ideas and eventually hardware.

Our viewpoints on space development could be viewed in the way a river develops. Most rivers in the eastern US grow on their way to the ocean. Rivulets, and creeks, and canals, other rivers, and lakes feed them up as they flow such that the river at the ocean is huge compared to its’ humble origins. In the southwestern US, some rivers have some of their vitality tapped off in so many places that the flow at the ocean is a fraction of the size of some upstream locations. Cities and farms and dams can reduce it to nothing in some cases. There are lawsuits about upstream usage before downstream availability.

Some of the discussion about space development mirrors river development. The question being if a given technology or suggestion is a tributary making the river bigger and stronger. Or is it a city or irrigation system draining the vital juices preventing the full flow to the destination. Much of spaceflight history is that of pet projects and congressional set asides draining the river en route such that the salt water flows upstream into the delta regions poisoning the  freshwater plants that depend on the river. SLS is the current flagship for that view with the funds going to it and its’ precursors being more than sufficient for real progress if it had been properly focused. SLS could be seen as a city in the desert that built a dam that keeps the water from flowing to the downstream drought.

The various ideas I throw out could be tributaries or dams depending on the ideological approach involved. If funds are diverted from the main goals for endless toy  development, potential dam. If they must fund their own way to the river, potential tributary. Whether or not the technology of a concept will work is important. Where the funding comes from, and which strings are pulled to get it is critical.

So in the recent discussions, who is right? There are a lot of variables that could make it either, both, or neither.

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Early Testing and Demonstrations on the Depots and Rotovators

Doug Plata has done some detailed comments on the suggestions, ideas, and differences from his ideas that my posts laid out in minimal form. His comments deserve better than I can do in a little comment window. Probably better than I can do in a post, but I won’t admit that part in public. It is my opinion that these technologies would speed up Lunar development, expand the possible scope of operations, and cut the cost in the process. They do carry risk though, which is Dougs’ objection, along with getting permission to use them in the first place.

I did a post on my views on the Lunar hoverslam landings. This one is how I see the depot and rotovator getting underway. I think initial proof of concepts and bringing the TRL up to snuff should be done on company internal resources without fighting the Federal funding battles.

The only real reason depots are not in use now is that there is no real demand for them. A comment on one post noted that propellant in orbit was as valuable as dirt. A bitingly true statement as long as there are limited missions beyond LEO, and few satellites share orbits that would make depots useful. A constant flow of material to and from Lunar orbit changes the situation with many vehicles taking the same path.

I see the initial depot flights as secondary payload technology demonstrators. Often flights to LEO are at less than payload capacity of the launcher. Either the upper stage carries a small second spacecraft, or it makes a rendezvous with another vehicle. The upper stage docks(berths?) with another vehicle and transfers propellant to it. They separate for a while and then hook up and transfer propellant back. Operating as a secondary payload on a stage that is expendable anyway should have the possibility of being a fairly economical mission. This would give a chance to solve propellant settling and transfer in (off?) the real world. Several missions could be flown for relatively low operating costs until the company is comfortable with the transfer techniques and has the boil off data for a few configurations. Then start flying more ambitious missions that do need some help until it is an accepted practice. There is too much information out there on depots to justify me going long on the subject.

Rotovators are far more risky. The payoff is also very high. The Lunar rotovator alone would offer major savings to a serious development operation. The ability to return material from the Lunar surface to an Earth bound trajectory without propellant, engines, or tanks would make it attractive even without the ability to intercept cargos from Earth for   a soft landing without fuel. The TRL is very low for tethers of any kind in space with rotovators having no test data at all.

I suggest the rotovator  demonstration unit be a secondary payload with the minimum mass that can demonstrate the principles.  This mission would be the rotovator itself, whatever auxiliary equipment is needed to make it work, and a bunch of expendable small spacecraft with the only function being thrown and caught.

The rotovator is lightly spun up when orbit is reached testing deployment and system dynamics. The initial target velocity is that which brings the tip thrown  vehicles to a 15 orbit per day instead of 16 of the base vehicle. This brings the small vehicle back to rendezvous in one day if all the calculations and results work out. It is to be expected that most of the little test spacecraft will be missed and lost early on. Perigee would be kept low enough that missed ships would reenter in a matter of days to avoid creating more orbital debris. It would be a risk that there would not be enough of the little ships to establish success and possibly no captures at all on the first rotovator mission. Further rotovators would be sent out as secondaries  until accurate slinging and reliable captures were expected instead of experimental.

After initial proficiency is reached at the 15/16 orbits, velocities are increased to 14/16 and 13/16 until the 1,600 m/s target is reached that would validate a Lunar rotovator. Then one is sent to Lunar orbit as a working system. If the 1,600 /s units were successful enough in Earth orbit some would remain to pick up suborbital ships to sling them most of the way to GTO or TLI with the rockets relighting after the rotovator boost.

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Lunar Hoverslam

I suggested in the middle of a couple of recent posts that the hoverslam techniques pioneered by SpaceX with the Falcon9 be used for Lunar landings. It was a kind of throwaway thought along with several other suggestions. As I think about it though, it seems to me that there might be a serious possible schedule and reliability gain from adapting the technique to Lunar development. That’s why I’m putting it up as a separate post.

I didn’t think hoverslam was a viable technique until it had been demonstrated. I was wrong. Now that it has been demonstrated multiple times, it may be time to see if there are more applications in which it might give an advantage. Lunar landings being the application under discussion recently, I want to lay out a few possibilities.

First thing would be a discussion with the team that is already using the technique in operational vehicles. From the outside looking in, it appears that hoverslam is a software solution to landings that was previously considered a hardware development problem. If this thought is accurate, then it may not be necessary to develop engines and control systems that allow an empty tank vehicle to hover in 1/6 gee. It seems that it is a requirement to bring velocity to zero at the instant that altitude is zero with thrust/weight being far less relevant than most of us previously thought possible. It seems that the SpaceX team is landing with thrust/weight levels of well over two on Earth, which would be well over a dozen at the Lunar surface.

If a Lunar lander is at 20 tons at touchdown, then the hovering that most of us consider a requirement would need engines capable of throttling to  3 tons for a gentle descent at very low velocities. The experience of Apollo 11 finding a clear landing area validates this opinion. This is however, not 1969. The Lunar surface is not only far better known now, but any potential landing sites could be imaged to near centimeter precision at relatively low cost. So hovering while making sure of a clear landing zone may not be a requirement. Navigation to the clear areas is also much less of a challenge than a half century ago. So it may be possible to go straight in to a site on near side without even orbiting first. It may be possible to land that 20 ton vehicle with engines that will only throttle down to 50 or 60 tons.

Doug believes that getting funding authorities to sign off on Lunar hoverslam would be a nonstarter. He is right unless the technique is fully validated just as it was on Earth/barge. I suggest the first step would be an RFI to SpaceX to confirm that it would or would not be possible to use the technique in this manner. If the answer is affirmative, then a test mission could be envisioned. For a test mission, perhaps an upper stage of the Falcon9 could be refueled by a Facon9 tanker in Earth orbit to validate tanker technology as well before sending it on to the Lunar surface.

The Falcon9 upper stage with one refueling should be able to place well over 5 tons on the Lunar surface during the test mission if the concept is valid. Depending on the flight backlog and the interest of both NASA and SpaceX, this could fly by Q4 2018. I doubt any other system could land a comparable payload in anything close to that time frame regardless of interest. Cost would be for two Falcon9s plus payload and Lunar operations. 5 tons in useable condition on the Lunar surface would go a long way towards convincing a funding authority to further use the technique for unmanned payloads.

Central to acceptance of the concept would be the failure modes. Obviously a high enough speed impact would destroy the stage and cargo. Hitting a rock with a landing leg and tipping over could be almost eliminated with a good survey and navigation. A sideways vector on landing that tipped the stage over should not be a factor with the current experience level. The most likely failure modes would seem to be engine failure at altitude from fuel depletion, and excess velocity at touchdown from software or navigation error.

Payloads on the first flight(s) should be very robust as well as being useful so that good work can be done even with a less than successful landing. During an excessive velocity landing, the stage propellant tanks provide a crumple zone if done right. An impact at 100 m/s (200 mph) in the vertical orientation could subject the payload to under 10 gees which is survivable to most hardware. It should be expected that the first payload may have to cut its’ way out of the wreckage before deploying solar panels and starting the primary mission. If the stage soft lands but with a side component that tips it over in the 1/6 gee, the payload should also see less than 10 gees.

The spectacular failures we saw from the early Falcon9 barging attempts were almost all from residual propellant exploding. Though technically not detonations, the burns were fast enough that most of us would call it a good boom. The vertical and horizontal vectors on most of the early Falcon9 barging attempts would have been payload survivable without the propellant reactions on impact. In the vacuum at the Lunar surface there would be no reaction from residual propellants in a crash other than fast evaporation and site contamination. All of those spectacular RUDs on the barge would have been stage lost and payload delivered on the Lunar surface.

I suggest that this concept be considered at some low level to see if there is any merit to it. If there is, it could speed up Lunar development by several years and save a few Dirksens.

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Tesla Semi is reasonable, part 1

Tesla finally unveiled their semi truck:
An electric monster with 4 Model 3 motors, one for each drive wheel, with an astounding 500 mile range (beating most people’s expectations). Fully loaded (i.e. 80,000 lbs total), it can accelerate from 0 to 60mph in 20 seconds. Unburdened, it can do 0-60 in 5s. What’s the point of this kind of performance, other than the ability to go up a 5% grade at 65mph fully loaded? In a traditional vehicle, the conventional wisdom is that an over-powered vehicle will kill your efficiency. And how can you make a 500 mile range electric truck without destroying your payload capacity?

In an electric truck, efficiency and power are related in the opposite way you might expect. The truck shaped a bit like a bullet, with a drag coefficient of 0.36 (like a sports car instead of a “barn”) and an efficiency of better than 2kWh/mile. And the insane performance?

Well, by running more motors in parallel, the effective power loss due to resistance in the coils is reduced since Power loss = I^2*R. Torque is proportional to current and number of motors, so if one motor with resistance R1 needs to generate torque T1, it requires a current I1, and thus a power loss: I1^2*R1.

But if you take the same torque requirement and split it among 4 motors, the current for each motor is now just I1/4, and the power loss for each is: (I1/4)^2*R1 = I1^2*R1/16. All 4 motors in total have a resistive power loss of 4*I1^2*R1/16 = I1^2*R1/4, i.e. a quarter of the value of just a single motor. Additionally, the heat generated would be split among 4 motors, reducing the temperature of the windings and further reducing the coil resistance.

However, there are other loss mechanisms, like friction, “windage” (i.e. aerodynamic drag from the spinning of the motor itself), and eddy currents in the windings and the iron of the magnet. However, these loss mechanisms can be reduced by clever and careful engineering of the motor. Resistive losses are more fundamental and at a certain limit, the only way to reduce the resistive loss is to add more copper (or operate at higher RPM through gearing). Anyway, the point is that having more motors can actually increase the efficiency of the powertrain by reducing the effective resistive losses in the windings. (if you’re interested in how all these loss mechanisms play out in the total efficiency of a motor, here’s a brushless motor efficiency map, showing the maximum efficiency of about 98% at relatively low total torque: motor efficiency map and source:

Another instance in which superlative specifications, which would seem to be detrimental to performance, actually improve efficiency is in the battery size. Batteries are not pure voltage sources but contain “internal resistance” which limits performance and reduces efficiency under load. Doubling the capacity of a battery (by, say, adding an identical battery pack in parallel) thus has the effect of halving the internal resistance and thus halving the power losses under a certain load. Therefore, unlike the conventional wisdom which states that you should use the smallest battery possible, under some metrics adding a larger battery will actually improve efficiency. At proportionally low loads, discharge efficiency can be on the order of 99% or higher (source).
If on the other hand, the smallest motor and the smallest feasible battery were used, energy efficiency would be lower. Motor efficiency, as the above map showed, can drop to mid-80% efficiency or lower. Discharge efficiency for the battery can also drop to 90% or so. Therefore, using the smallest motor and battery (and then relying on, say, a fuel cell or something) would be a false economy. Tesla claims a sub-2kWh/mile figure for the semi, a figure that other manufacturers (such as Nikola Motors) scoffed at.

And let’s look in detail at the <2kWh/mile efficiency claim, starting with drag.

Tesla gave a drag coefficient of 0.36 for the Semi. A typical semi trailer is 96-104 inches wide and 12.5-13.5ft tall. I will pick the smaller end of this range in an attempt to show the minimum possible energy figure. EDIT: Just kidding, I’ll use 104in and 13.5ft because a poster reminded me I had forgotten a factor of 0.5, which makes such optimistic assumptions unnecessary. I will assume the standard 1.225kg/m^3 sea level air density (higher altitudes offer better efficiency! The Tesla Semi seems perfect for mountain routes…). And to start, 60mph:
.5*0.36*104in*13.5ft*1.225kg/m^3*(60mph)^2 in kWh/mile:
0.770843181 kWh / mile

and for kicks, at 55mph:
0.647722395 kWh / mile

So, already, the 500 mile Tesla semi has to have a battery of 385 or 324kWh assuming perfect drivetrain efficiency and no rolling resistance. Oh, let’s calculate rolling resistance:

Rolling resistance, for our purposes, requires a constant amount of energy per mile to move a certain load. The speed doesn’t matter (much…). Whether 1mph or 30mph, if you neglect air resistance, the rolling resistance will require you to expend energy to move stuff on Earth (although less on other planets like Mars…).

This is basically the same thing as coefficient of friction, if you remember your first physics course in high school or college. Crr, the coefficient of rolling resistance, is technically unitless, although sometimes it’s expressed as kg-f/tonne-f (i.e. Crr times 1000), and sometimes it’s expressed not as a coefficient but as quantity with a unit, such as power in Watts (for a given load and speed) or force (for a given load). Unfortunately, this number seems really hard to find for typical commuter car tires (but I digress).

The energy needed per unit length is calculated: gravity*Crr*load. g is 9.80665m/s^2, load is 80,000lbs.

Typical semis have about a 0.006 rolling resistance coefficient, so:*.006*80000lb+in+kWh%2Fmile
0.954495836 kWh / mile

Some good semi tires with low rolling resistance have a coefficient of about 0.0045 (so 0.715871877 kWh / mile). But this is not an ultimate limit. A really, really good road bike tire may have a coefficient of about 0.002 (0.318165279 kWh / mile), which is in the range of rail (which has a rolling resistance–metal wheel on metal rail–of about 0.001 to 0.002 in real life conditions… although fundamentally it can be even lower). If you want to compete with rail, you’re going to have to reduce this as much as you can. But, of course, a semi truck is not a road bike.

But there is one way to get even better rolling resistance: a wider tire. Most semi trucks use dual tires on each of the four drive wheels. But it’s also possible to combine the two tires into one. A “super single” tire can have as low as 0.0034 (0.540880974 kWh / mile ), significantly better than the usual low rolling resistance semi tire. Super singles also tend to last longer, and since they have more contact surface with the road, it’s possible they also have better stopping power and even lower road wear. They also can reduce weight (in combination with alloy wheels) by up to 1000 pounds. And 0.0034 is not a fundamental limit. Physics allows you to do better, as road bike tires prove. So this may be an area that Tesla invested in, and according to some rumors, Tesla is planning to use super singles on the Semi.

So, bringing this all together, with a full 80,000 total load, 95% combined motor and battery discharge efficiency (97% may be possible, but not required… and 95% also allows some parasitic load, such as lights, etc), 0.36 Cd at 60mph and 13.5ft tall and 104in wide trailer with a more conservative Crr of 0.0041, I get an energy-per-mile of:
(0.5*0.36*104in*13.5ft*1.225kg/m^3*(60mph)^2 + 9.80665m/s^2*.0041*80000lb)/(.95) in kWh/mile
1.49798105 kWh / mile

That’s consistent with Tesla’s “<2kWh/mile" and with all the data Tesla has given so far and with fairly optimistic assumptions. But if they're using 55mph as a baseline instead and the more efficient tires and drivetrain with the slightly smaller trailer, then the figure could be as low as 1.16kWh/mile (meaning a 600kWh battery may even be realistic). 1.5kWh/mile I think it’s a pretty realistic figure, and so that’s what I’ll use.

That works out to a 750kWh battery for the 500 mile version, which is $180,000 base. The 300 mile version is $30,000 less at $150,000. Does that mean it's 450kWh and thus the cost per kWh is only $100? I'm not so sure that's what it means, although that’s no show-stopper. I DO think Tesla has the ability to make super cheap batteries, but I think there's more to the story.

Tesla announced a 1 million mile guarantee of sorts on the semi drivetrain. I assume that means 1 million miles before the battery degrades by, say, 20%. For a 500 mile battery, that's 2000 cycles. But a 300 mile battery (450kWh?) at 2000 cycles only lasts 600,000 miles until that 20% degradation. Not only that, but since you're loading the battery more relative to its capacity, it might not even last 2000 cycles. So what I think they'll probably do is put a slightly larger battery in there than you might think. Perhaps they increase the battery size to 500kWh to give more margin, thus reducing the depth of discharge slightly and giving room for a little more degradation before your range is reduced too much. That gives a cost of about $120/kWh, which even gives room for profit (imagine that!). But even that kind of number will leave many analysts incredulous. "Breaks the laws of batteries" they'll say: bloomberg…tesla-s-newest-promises-break-the-laws-of-batteries

And I say balooney. Tesla is selling the Semi with initial release in 2019. But they're going to start with the more-expensive, early-adopter $200,000 "founder's edition" semi. So getting to $120/kWh isn't required right away. Ramp-up may be relatively slow, and so let's say they don't start delivering the regular semis until 2020, maybe 2021. Even GM expects to have battery cell prices down to $100/kWh by 2022 (with pack costs estimated by GM at about 20% more). And a cost of $109/kWh for a battery like the one Tesla is using is consistent with Argonne Lab’s modeling BatPaC 3 for 500,000 units per year of 90kWh each: greentechmedia…How-Soon-Can-Tesla-Get-Battery-Cell-Cost-Below-100-per-Kilowatt-Hour
Since Tesla is also going to be selling a bunch of Model 3s and Model S/Xes and PowerPacks and PowerWalls by then in addition to 5 years of chemistry improvements since that analytical model came out, it’s fully realistic that the learning curve will allow a $120/kWh figure to be achieved. Elon Musk has long predicted a $100/kWh figure once the Gigafactory was fully operational. That’s basically why the Gigafactory was and is being built. I think the biggest reason people have a hard time accepting even a $120/kWh figure (let alone $100) is that it breaks a lot of their dearly held notions about energy. At that price (and combined with solar cells at 19 cents per watt at the spot market), anywhere with sun is going to have cheaper energy than fossil fuels (unless you’re right above a gas field or something). But that’s a topic for another post. I will need to address the 7 cents per kWh that Musk promised.

EDIT: Thanks to comments by James who pointed out I missed the factor of 0.5 in the drag equation. That makes the whole Tesla semi project way easier and should give Tesla a decent profit opportunity. It means only $120/kWh is needed and shrinks the required battery, which reduces the weight considerably and also thus reduces the required charging speed.

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A View on the Alternatives

In the post on the architecture that Doug Plata proposes for the Lunar development, I likely appear as a naysayer as I picked on the sections that I disagreed with or thought that there was a better way forward. Not surprisingly, Doug disagrees with my nitpicking.  There is one point that we will not likely agree on which is government involvement. He believes that a Lunar COTS approach can hold costs to roughly a billion a year. I believe that involving NASA at management level brings in baggage that will blow past a billion a year under any contract method due to congressional involvement. There are a lot of talented people in the agency that answer to political reality. As a single example, commercial crew is $6.8 billion to develop two capsules plus a handful of flights. There is a lot of development for a Lunar program that is far more involved than a couple of capsules. On to the technical stuff that I actually like.

Doug thinks that the technical suggestions that I made would delay Lunar settlement, add risk, and drive up costs. It is a series of reasonable concerns that I will try to address. I suggested Lunar hover slam landings, a rotovator, a low gravity research facility in LEO, and propellant depots. Each of these is questionable without more detail on my meanings, and certainly debatable even after I explain. I will be surprised if one or more of them isn’t honestly trashed in comments.

I was one that didn’t see the hover slam landings working, didn’t expect a high percentage success on the barge, and didn’t believe that landings without deep throttling would be possible. You can probably dig up old posts and comments of mine naysaying on all these things. The fact is that SpaceX has succeeded in reliably landing vehicles vertically with a thrust/weight considerably higher than one. It seems to be a matter of having control systems that can accurately decelerate vertical velocity to zero at a specific altitude. It seems to me that that altitude can be regolith zero just as it can be barge or LZ1 zero. The advantage would be that it would become unnecessary to develop a whole new landing stage for the early equipment and supply deliveries. This would be a way of using the second stage all the way to the Lunar surface. While there is risk, I believe a dedicated Lunar lander will also have some element of risk, just later and more expensive. Later on, dedicated landers will be necessary, especially for human landings.

A rotovator is potentially one of the key pieces for Lunar transportation. A 1,600 m/s unit could catch payloads from TLI and soft land them on the surface. The orbital energy gained could be used to pick up payloads from the surface and sling them to TEI. Dougs’ concern on this one is cost, risk, and schedule. A 1,600 m/s rotovator would seem to have a mass ratio of roughly 25. That is a 25 ton unit could handle 1 ton payloads, in theory. I suggest that the testing and development should take place in LEO. At $2K per kilogram launch costs, a 25 ton rotovator would cost $100M to launch and about a quarter of that for the tether material itself. This would be about an eighth of the first year budget. I would not suggest waiting on it to prove itself before starting the settlement missions. In LEO, the operators train, learn, and develop by picking up payloads from 1,600 m/s below orbital velocity and slinging them to 1,600 m/s above orbital velocity. This would be a long way towards a GTO or TLI. Only after proficiency is reached is a unit sent to Lunar orbit. Assuming it works and survives the clutter of LEO of course. Once in Lunar orbit, it roughly doubles the payload of a vehicle in TLI to the Lunar surface. It also allows material from the Lunar surface to be placed in TEI with no propellant or engines. A third function would be pick up and drop off at various points on the Lunar surface for multi-location prospecting without propellant or engines. The rotovator is a risk, a risk with high payoff. And if it doesn’t work, 1/8 of one years’ budget might be an acceptable loss against the potential benefits.

A low gravity research facility in LEO is another point of disagreement. Doug sees no need for the data until actual settlement generates the information needed. I see it as yet another set of information that should be generated in parallel with development. A 50 ton variable gravity facility could be placed in LEO for a similar cost as the rotovator, as long as feature creep and over engineering is avoided. A simple design that can run unmanned most of the time with occasional visits for maintenance, specimen swap, or clean up. Only after a few generations of rodent and primate trials would it be necessary to send permanent human crew to nail down the data points. It may be that Lunar level gravity is enough to maintain health without centrifuges. Or it may be that it is totally inadequate. It would be good to have some realistic data before shipping large scale equipment to the moon that turns out to be unnecessary on one hand or inadequate on the other. The early animal studies could be started by the time the first humans  are on the surface with the primate studies completed well before adverse reaction might be expected to show up. A possible benefit would be if the settlers could avoid some of the extreme exercise requirements of LEO. This would be another quarter years budget from his program. It just might more than pay for itself in reduced future equipment and exercise requirements.

To me, propellant depots are almost a no brainer. Not everyone shares my opinion. The way I see it, Lunar development would almost require depot type facilities. The first depot could simply be a repurposed upper stage combined with high sortie rates. A relatively high boil off rate could be acceptable if it means being able to launch the large pieces with the equipment you have instead of the equipment you hope will be available eventually. Refuel the upper stage of the FH and send a 60 ton payload to the Lunar surface in one shot. Propellant can be bought from any delivery including from excess capacity in a given launcher. If a vehicle has  the capability of placing 20 tons in LEO and the volume restricted payload is 10 tons, then 10 tons of propellant can be bonus payload. The same applies in Lunar orbit where it would make almost as little sense to send a half loaded vehicle to the surface as it would to send one with insufficient margins. There should be no reason for depots to slow down Lunar development,  and the efficiencies should make the system self funding almost from the start. As long as the over specified and over designed techniques of the past are avoided of course.

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