Amistics of Human Spaceflight, or How Autonomy and Miniaturization can be the Enemies of Human Spaceflight (Part 1)

File:Lancaster County, Pennsylvania. An Old-Order Amishman working in his repair shop. Good machine sho . . . - NARA - 521078.jpgNeal Stephenson in his novel Seveneves coined the term “Amistics”, deriving from how some Amish people have strong preferences for certain technological paths to achieve the same goal. For instance, these Amish folk swear off modern technology, which for them means electricity. Therefore, they cannot use electric power tools for their furniture-making. Instead, they use just-as-modern air-powered tools. Similar productivity, same result, but they’re able to honor their cultural proclivities. In Seveneves (not to spoil it for Jon), similar proclivities develop in the groups mentioned in the book.

Spaceflight is rife with examples of this. One is the pro-vs-anti hydrogen schools of thought. Dumb, mass-produced expendable vs high tech reusable. But probably the most important for the future of humanity is the amistics of robots vs humans.

It gets started at the beginning of the space race in another example of technology path-dependence. Due to the US’s earlier start, America’s nuclear weapon technology had significantly more advanced miniaturization technology than the Soviets. For reasons I’m not entirely sure of, the US also maintained a very strong advantage in electronics and computerization. Additionally, the US had an advantage in long-range bomber technology. This led to the fact that the Russians focused on ICBMs while the US focused on long range bombers. And secondly, that the first Russian launch vehicles were ENORMOUS in comparison to the US’s. Russia developed the R7 and the Proton in part to be able to lob their nuclear weapons, which (from my limited knowledge) lacked both the miniaturization and precision of their American counterparts. The R7 was so big, that they could use it to launch Sputnik to orbit. And later on, the first crewed launch (Vostok), and eventually even up to 3 people on a single rocket that is used to this day. The US, on the other hand, was caught by surprise by the advanced Soviet ICBMs. Large ICBMs like Proton were not required due to better targeting and miniaturization, thus the US had to develop heavy launch vehicles intently for spaceflight purposes.

And thus the Soviets racked up success after success in the early history of human spaceflight due to the path dependency of tech development. It was only after a concerted, civilian-focused effort of development that the US exceeded the Russians, by an enormous margin.

But the Soviets maintained some of these advantages. They pressed their early leads in human spaceflight and while the US rushed to the Moon, the Soviets developed crewed space stations designed for surveillance. The Almaz program launched Salyut 2, 3, and 5. Soviet military personnel conducted surveillance from orbit in real time. The Americans, for their part, had a similar program, the Manned Orbital Laboratory, or MOL, based on Gemini technology. An uncrewed demo of the capsule was launched, but the program was cancelled soon (in 1969) as it became clear that automatic satellite surveillance was sufficiently advanced that it wasn’t required nor worth the cost. The US’s lead in automation again struck a blow to human spaceflight.

About a decade after (1978), the Soviets came to a similar conclusion and ended their manned orbital surveillance program. But not before advancing their space station technology sufficiently to place them at a Image result for soyuz rocketdramatic advantage over the US in long-duration human spaceflight (as measured by orbital refueling, human spaceflight duration records, etc), an advantage that STILL has not quite yet been eclipsed (although it’s close). And because of the early focus on large launch vehicles and human spaceflight over miniaturization and automation, the Russian human spaceflight program survived the fall of the Soviet Union and to this day US NASA astronauts rely on Russian vehicles to get space.

Now, humans make terrible surveillance satellites, but these historical examples should make us think twice about whether the best way to push for a future where millions are living and working in space is to invest in miniaturization and automation. Because in my opinion, the most likely result is that any useful things a human can do in space will become obsoleted by robotics much faster than otherwise, thus reducing the need for humans in space at all. That’s not a winning strategy, IMHO. So I hope to blog later about how we can use humans in space MORE, in direct contradiction to the current trendy meme of increasing robotic automation in space.
We need to:
1) Find things for HUMANS to do in space.
2) Make it cheaper for humans to go to space.
3) Make it cheaper for humans to live and work in space.

We need a pro-human amistics, not the current pro-automation amistics (even when it doesn’t make sense, like when Elon tried to fully automate Model 3 production and had to switch over to human assembly). We need to engineer systems very close to the humans, including perhaps modifying the human body itself (or at least developing advanced biomedical countermeasures) to make humans more competitive with robotics in space.

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Blue Moon: Is this really it?

By Chris Stelter

Blue Moon, the recent announcement of an uncrewed lander by Blue Origin, had flare and pomp. A starfield surrounded the select audience as they watched Jeff Bezos, the richest man (okay, if you count his family) in the world, deliver an anticipated announcement. They waited patiently as Bezos gave his usual spiel about Earth being the best planet, about the criticality of reusability, about a trillion people living in O’Neil colonies, about moving heavy industry to space. Then Bezos unveiled…

An expendable descent stage with less payload capacity (3.5 tons) than the Apollo LM truck variant (5 tons). Because it uses liquid hydrogen, it’s very tall and therefore needs a sophisticated mechanism for unloading payloads.

It was so anti-climactic. Everyone knew Blue Origin was working on this lander, I was sure it was going to be something more important or at least *innovative*. I’m not sure if the rocket engine is pumpfed or not, but the lander is designed as if it’s pressure-fed, with Apollo-like large round tanks with external structure.

Its example mission is… landing several smaller payloads simultaneously. Basically, competing with all the smaller lunar lander companies out there. Super disappointing there as well.

It’s like a tiny, uncrewed version of Altair with all the drawbacks but without the advantages of a 16 ton payload capacity. And sure, they showed an ascent stage on top of it, but that appears to be provided by NASA.

In fact, let me list off some concepts I think are better:

1) Starship. Obviously. Fully reusable, much larger payload capacity, crew capable, and being crudely prototyped right now in Texas, not just made into a fancy mockup.

2) The reusable Lockheed Martin lander. Dinospace is not supposed to be this much better, but this is a lot more interesting than Blue Moon.

3) ULA/Masten Centaur/Xeus. More payload capacity, still hydrolox, much closer to the ground. Looks to be a more efficient design. Some of the hardware already exists in some form.

4) Altair. At least they were trying for more capability than Apollo.

5) Apollo LM/LMtruck. 5 ton payload capacity, much closer to the ground. Crewed variant was the only one that flew, so it started out crewed.

6) The Soviet LK lander. Crasher stage FTW. Less payload, but the Soviets did a fine job systems engineering a clever way of dealing with the constraints they were given by the much-less-to-TLI N-1 rocket.

7) Various crasher/uncrasher lander concepts, as discussed here.

8) Delta Clipper on the Moon. If Delta Clipper had been successful, there was thought given to variants of it for Moon or Mars. If you have a SSTO VTVL RLV, why not refuel and go to the Moon? Basically, like Starship. Bezos hired a bunch of old DC-X folk. Why such a mundane lunar architecture?

Blue Origin gets like $1 billion per year from Bezos. Couldn’t they come up with something better than Blue Moon? Or at least something that didn’t look designed to squash the other small lunar lander outfits? A reusable upper stage? A reusable lander? Anything?

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SBIR Proposaling Advice

A few weeks back, I asked my followers on Twitter what topics they’d like to see me blog about. One of them suggested that since I do a lot of SBIR proposals at my day job, that maybe I could write an article on advice and lessons-learned for how to write better SBIR proposals.

For those of you not familiar with the term, SBIRs in this case stands for Small Business Innovative Research grants. Basically they’re small R&D contracts that NASA, the DoD and other federal agencies do that only small businesses1 can bid on. Each of these agencies will put out solicitations 1-3 times per year, with a list of topics of interest to the agency. Companies can then submit a proposal for innovative research addressing something in that topic, and if they get selected they’re given typically around $100-150k of money for a six month Phase I feasibility study contract. If that goes well, they can submit a Phase II proposal, and if they win that, it’s a 2-year $750k effort that usually culminates in some level of prototype. To make a long-story short, for bootstrapping aerospace startups, SBIRs if done right, can be a good source of non-dilutive funding, and can help you build up technology expertise as you work to get to a point where you can go after bigger commercial markets.

Anyhow, with that digression aside, I wasn’t sure when I got the advice if I was really competent to write something like this. Sure we’ve written a few and won a few, but we’re not really what I would consider an SBIR farm. I think to-date SBIRs have been less than 1/3 of our total revenue as a company. I ran the numbers tonight, and apparently between NASA and DoD SBIR and STTR2 Phase I proposals, including ones where we proposed as subs, we’re now up to over 43 proposals, with 11 wins. So, I think I can say we no longer totally suck at this, and are in fact probably solidly mediocre.

So… with that stirring endorsement in mind, here’s a few tips I could think of to help you improve your SBIR proposaling fu.

  1. Get to Know Your Customer: One of the best ways to improve your odds of writing a good proposal is to have spoken with the Topic Manager (TM) or Technical Point of Contact (TPOC) first, to learn more about what they’re really looking for, before the solicitation opens3. DoD makes this easy by having a “pre-solicitation” period where they explicitly list the contact info for the TPOC and give you a month to talk with them before the real solicitation opens. NASA makes it harder by not giving out TM contact info. In order to get to the TM your best bet is to look at old SBIR topics from previous years4. Find one that looks aligned with your interests. Look up which center is listed as the lead center for the topic. It usually stays the same over time. Look up that center’s SBIR Program or Tech Transfer Office, and ask them if they can put you in contact with the TM. Most of the time they will. Once you’re in contact, you can now do some customer discovery research. Find out what problems are keeping them up at night. Find out what specific sub-problems are the ones they’re most interested in solving. Find out what types of solutions they’ve looked at, and what their opinions are about the relative merits of those solution types. Find out if they have a transition path in mind–a government program that needs this, or if it’s more just professional curiosity. Find out about their philosophy on SBIRs and how strategic they try to be–are they just seeking fun ideas to fund? Or do they use the topic as a way to get free money to fund research in specific areas they need for an operational program? Find out if they can recommend any NASA researchers in this area that they work with who’ve written papers in this area. If it’s a DoD presolicitation period, I’d also try to dig to find out if the there is a specific technical approach they’re already sold on. On how many Phase I and Phase II awards they expect to give out for the topic, and things like that. Lastly, after you’ve done your customer discovery and know more about the problems they’re trying to solve, that’s when you can start batting ideas around with them to get their feedback. Doing this properly helps you not only by making sure you’re proposing something relevant, but now the people who’ll be reviewing your proposal at least know your name, and you have a better idea of their biases, technologies of interest, and how they see the competition. But best of all, if you’ve done your homework right you at least know if you have a shot at all of being awarded something. There’s nothing more frustrating than wasting several weeks writing something you think they’ll love, only to find that they consider your idea “non-responsive” to the topic. Ever since I found that the DoD has a pre-solicitation period, I’ve submitted way fewer proposals, but my win rate has gone up substantially. It’s much better to have them shoot a hole in your idea when you don’t have multiple weeks into it than to invest what is going to be thousands of dollars worth of your time into something that has zero chance of winning. So, to summarize — fail fast by talking with your customers and really getting to know them before trying to write a proposal.
  2. Sell the Combined Phase I/Phase II Story: This was a piece of advice I got from Greg Mungas, who used to run a propulsion R&D shop5 out in Mojave, and before that used to work at NASA JPL, where he was involved in reviewing SBIR proposals. Basically Greg’s point is that nobody funds SBIR Phase Is just for the Phase I. They fund Phase Is that they think will lead to good Phase IIs. So, when you tell your story in your SBIR, you always want to talk as though you’re proposing both phases. Here’s where we get to in Phase I, which sets us up to do X, Y, and Z in Phase II, getting to this really awesome demo or this point where the technology is ready for flight demonstration, or something like that.
  3. Hook Them With the First Page: Greg also pointed out that you get a lot of different sorts of people reviewing SBIRs, but one thing they have in common is that they almost never have time to thoroughly review everything they’re given. So, it’s really important to quickly hook the reviewer with an exciting first page or two, so they’ll be interested and engaged for the rest of the proposal. My typical formula for the first page includes a) a pretty piece of faux CAD6, b) a paragraph setting up a compelling-sounding problem you’re trying to solve, c) a paragraph introducing your concept in 1-3 sentences as a solution to said problem, d) a short bulleted list with some key highlights of the features/capabilities you think you can achieve with your concept, usually with a 1-3 word summary bolded at the start of the bullet followed by an unbolded one sentence description7, and e) a one paragraph summary at the end talking about what you’ll do in Phase I, what technology readiness level that gets you to, and how that sets you up to do awesome things X, Y, and Z in Phase II that will get you to an ever cooler end-state. Basically it’s almost a mini-proposal on one page that gets them excited to read the rest of your concept, and helps them understand where you’re going from there.
  4. Cite Chapter and Verse: Most SBIR solicitations ask you in your first section, where you’re describing the innovation, to explain why your innovation is relevant to the topic. In these cases, I’ve found it handy to quote the solicitation verbatim, and highlight or bold or italicize the parts of the topic that you specifically hit. Here is also where you can mention insights you gleaned about their needs from your customer discovery interviews.
  5. Quantity Has a Quality All of Its Own: Even if you’ve done your homework well, and are writing an exciting intro, and properly quoting chapter and verse, SBIR solicitations still can be fairly competitive. Your raw odds are typically around 1 in 6 to win something. If you’ve done your homework well, you can get up into the 50/50 range. But that still means you’re best off proposing more topics than you expect to win. We try to shoot for 4-6 proposals each cycle (if we have time), in the hopes of winning around 2 Phase Is, with the hope of converting at least one of those two to a Phase II.
  6. Sometimes Its Them, Not You: Sometimes you’ve done everything right and you still get shot down on a topic. Never fear–you should get a debrief a few months later with reviewer comments. Sometimes you caught someone on a bad day, or they misunderstood something. A lot of times we’ll rebid things a few times in a row, with tweaks and improvements based on the feedback you got, and any customer discovery you can fit in between cycles. Rebids typically don’t take as much time to write, and you have more data going in on what they liked and didn’t like. So don’t be afraid to rebid something if you think you had a good idea that got ignored last time around.
  7. Don’t Waste Time on an Overly Detailed Materials Budget: Most SBIR solicitations require you to provide a fairly detailed cost budget, including labor rates and hours, indirect rates, and “other direct costs” including materials. The annoying thing is that they want detailed quotes for everything you’re going to order in this research project that you haven’t even started yet. You don’t have final designs yet, and don’t really know what things are going to cost, but they want not just numbers, but quotes or at least some other form of cost justification. A lesson we’ve learned is that proposals never get rejected because your cost justification isn’t detailed enough. If you win the proposal, during contract negotiations they’re going to demand you back things up in more detail, but at that point its a winner’s problem. You still need to get in the right ballpark or you could get yourself in trouble8, but a lot of times our cost justification in the original proposal is a handwavy “estimated based on engineering judgement from past projects of similar complexity.” You do have to get them more justified numbers during contract negotiation, but now you’re dealing with a much smaller number of contracts.
  8. Contract Negotiators Hate Travel in Phase I: The people getting you under contract are paid to try and find wasteful proposed expenditures that they can disallow, to cut down a little on how much the government has to pay you for your research. I’m pretty sure they almost always cost the government more than they save, but they’ll go over things with a fine toothed comb. And one of their favorite things to nitpick is travel in Phase I. So it might be worth trying to do as much as possible via video calls and such to avoid giving them a juicy target to go after. I’ve only had my travel request rejected once, but they nitpick it and second guess it every time. In Phase II they’re more understanding, but not in Phase I.
  9. Baking In Commercialization Help in Phase I Might Be a Good Idea: In the Phase II proposals, at least for NASA, the Part 7 on your commercialization plan is a bear. In Phase I, they just want to know what applications you think there are, and your rough development plan for Phase II. But in the Phase II proposal, they want what amounts to a 5-10pg business plan for taking this product to market. It’s a lot of work, especially for technologies that are often at least a few degrees off of your main business plan. So, getting some help on that section could be worthwhile. One thing we’re going to try this time is that there are companies that specialize in “Commercialization Technical Assistance”. In Phase II, you can request $5k of CTA funding that doesn’t come out of your $750k cost cap. The CTA companies then help you with identifying potential Phase III opportunities, commercial customers, and government programs you could infuse your technology into. They won’t give you that bonus $5k for CTA help in Phase I, but there’s no reason you couldn’t include a subcontract to a CTA firm to provide that help during Phase I–so long as you can fit it into the allowed ~33% subcontracting limit. That’s what we’re going to try doing on a few of our Phase Is this time around.
  10. Pay Yourself a Competitive Wage: As a startup founder, it’s often tempting to underpay yourself. The logic goes that doing so means more money going back into the business, which increases your effective runway for the company. And after all, you own a chunk of it, so why not take a lower salary to improve your odds of being successful. Unfortunately, I’ve gotten bit by this several times, and only finally learned my lesson. Basically, since you’re proposing more contracts than you expect to win, you almost always win some combination of contracts that leaves someone important double-booked, where you’re now going to have to hire someone new to fill in the gap. But if you were paying yourself below your replacement cost, that makes it really hard. Either you end up trying to recruit people paying them below market wages, or you end up losing money on the contract. But if you pay yourself at least what it would take to replace yourself as an engineer, then if you have to replace yourself, it’s a lot easier to do so without losing money because you can offer them a competitive salary.
  11. Always Have a Fun Proposal Just in Case: When we do our “gut gate” meeting, where we go over topics we could propose on, to triage things down to a reasonable number of proposals, I always try to leave at least one proposal that’s my “in case we get done with the others early enough” proposal. This is often the fun one that you can’t quite justify prioritizing over the better aligned topics that you feel compelled to do for business alignment reasons. You’ll need to spend a little time up front on long-lead time items like faux CAD, figuring out your teaming strategy, and coming up with a high-level plan for your tech objectives and key tasks. But other than that, you don’t spend a lot of proposal time on that one until you’ve knocked out your high priority ones. But by leaving the fun one for last, a) you’re incentivized to take care of your responsible ones first so you can get to the fun one, and b) once you’ve gotten into the groove of writing proposals, that last one goes a lot faster than you think. Our one Phase II we won, for our Cryo Coupler was just such a “fun” proposal. I had done just enough to keep it alive while focusing on the others, and I freed up the morning the solicitation closed with 8hrs left until proposals were in. After you’ve written a few in the past few weeks, I’ve found it’s often possible to crank out a full Phase I proposal in a single day at the end–so long as you’ve taken care of the long lead time items up front. Even if that last one is skimpier, and not as detailed, it’s still worth sending it in. Even if your odds are only 10% instead of the normal 15%, that’s still one extra shot at a contract award that you wouldn’t have had if you didn’t at least try.
  12. Take Advantage of I-Corps Training If Possible: The NSF SBIR program found that they were getting a lot of PhD researchers who couldn’t commercialize their way out of a wet paper bag. So they created a program to train researcher sorts on how to do customers discovery and commercialization. NASA recently started offering it as an option, and at least when we did it, it both forced you to talk with customers, but also gave you training on how to do it, and at least a small additional travel budget that they don’t nitpick as much to go meet with customers and/or attend conferences or trade shows. It’s a great program, and if you’re at all the typical engineer founder, you probably don’t know the first thing about customer discovery. I-corps won’t make you a master, but they can at least get you from the totally suck level to the solidly mediocre level, if you take it seriously. I really wish this had been available during my first year or two of running Altius. Knowing how to actually talk with customers, and realizing that engineers are really good at creating business hypotheses, but often horrible at actually testing those hypotheses, has been one of the biggest insights I’ve made in the last few years. The sooner you get good at this, the sooner you can get to something scalable and worth investing in.

I could probably go on, but those are the best pieces of advice I could think of in one sitting. I may update this post later if I can think of any more items, but I wanted to at least get these out there, in the hopes of helping at least some startups avoid making all the same mistakes we’ve made. Good luck, and happy proposaling!

Posted in Altius Space Machines, Business, Entrepreneurship, NASA | 2 Comments

Ultimate Off Road Race—-The Lunar 10,000

The arguments about reasons to go to the moon will continue until people start making a profit on site without ambiguity. A profit that doesn’t depend on the taxpayers and their chosen elected changing their minds. Most of us think in terms of “useful stuff”. Water, building materials, oxygen and such. I have noticed (as hard as I try not to) that it is not the necessities that get people pumped up. It is the frills. Sports, jewelry, cruises, and vacationing in general are much higher profile than what I am normally interested in, not to mention, enormously profitable. A wise friend of mine once said “If you want to make a living, give people what they need, but if you want to get rich, give them what they want.”

So an attraction on the moon that is mostly delivered in electrons to the customers, and have them begging for more. There are some sports events that take a few minutes or hours per game, but games are weekly for months. And seasons can trace back decades to over a century of history.  There are “reality” shows that go on season after season. And endless television shows that stay on for over a decade.

I suggest an event that would be unique, challenging, and very hard to predict. Circumnavigating the moon on the ground would be close enough to 10,000 kilometers for purposes of hype. All off road as there are no roads. An east to west course chasing the sun with a start just after sunrise. 28 checkpoints that represent the distance that must be traveled daily (24 hours, earth day) to make it back to the finish line before sundown. Other than the checkpoints, navigation is your problem.

There will have to be some rules such as no suborbital hoping between checkpoints. Other than that though, rules should be as simple and straightforward as possible. Cameras on at all times. No sabotaging competitors. Sportscaster interviews at each check point. And so on.

First race might be a single daredevil proving it could be done, winning the prize money by completing and surviving. Check points could be landers with com and supplies. Or perhaps sportscasters would hop ahead to cover the laps which would be on the order of 400 kilometers each. Get people excited enough to prove it on pay-per-view, and it might just expand into a major annual event.

Byproducts would be better transportation technology on airless/inhospitable planets. Considerable  exploration of the surface in a band perhaps 100 km wide around the whole planet. The possibility of a lot of people getting interested in more than one planet.

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FISO Telecon Lecture on LEO Propellant Depots for Interplanetary Smallsat Launch

Today I gave a lecture on the Future In-Space Operations Telecon Series on the idea of using LEO propellant depots for interplanetary smallsat missions.

Here’s a link to the archive page, which has both the presentation itself and an .mp3 recording of the talk and the associated Q&A/discussions: http://fiso.spiritastro.net/telecon/Goff_11-28-18/

We went over a lot of the same material that I discussed in the previous two posts, but with more illustrations, and some description of what we were doing that hopefully helps make the idea more clear. The main new addition was a “RAAN sweep analysis” we did to quantify the costs of using this 3-burn departure. tl;dr is that it’s not very painful–less than 3% dV hit compared to using a single-burn departure, and if you’re doing a human mission, and launch the crew to rendezvous on the last phasing loop, you can keep the flight-time penalty to <10days. All told, I was really excited to give this talk. It's a neat topic, and I'm becoming more and more convinced that there may be a commercial path forward for propellant depots for providing dedicated smallsat launches to MEO, GEO, and beyond. Way beyond.

Posted in Commercial Space, Lunar Commerce, Mars, Orbital Dynamics, Propellant Depots, Space Transportation, SpaceX, ULA | 6 Comments

AAS Paper Review: RAAN Agnostic 3-Burn Departure Methodology for Deep Space Missions from LEO Depots (Part 2 of 2)

Ok, picking up where the first part left off, we’ve reviewed the background, methodology, and some general observations on the methodology from our RAAN-agnostic 3-burn departure paper. In this part, I want to go over an interplanetary launch campaign concept that demonstrates how such a technique could be used to enable exciting interplanetary exploration missions that would be hard to perform otherwise. To do so, I’ll introduce the Interplanetary Blitz campaign, describe some of the key elements of the campaign, and then go into the results we found. Finally I’ll wrap up with some lessons learned, and areas for future research.

2024 Interplanetary Blitz
In order to illustrate the power of this RAAN-agnostic 3-burn departure methodology, we developed an interplanetary mission campaign that showed how a single ISS-coorbital depot could enable a series of interplanetary “smallsat” missions to four planets1 (Mercury, Venus, Mars, Jupiter), the Moon, and four NEOs2 (2007 XB233, 2008 EV54, 2001 QJ142, and 2009 HC) over an approximately five month period between late August 2024 and late January 2025. The main point of this exercise was to show that a single depot in a convenient location could support a rapid-fire series of interplanetary missions, without having any time to phase its orbit to align for each mission, and without excessive penalties for doing so.

A secondary but very important point of this exercise was to illustrate how a LEO micro-depot could enable dedicated smallsat launch vehicles, which currently can barely deliver payloads to LEO, to send those payloads practically anywhere in the solar system, at price-points that rival even semi-reusable larger launch vehicles.

Key Elements of the Interplanetary Blitz
This interplanetary blitz involves four key elements:

  • The LEO propellant micro-depot
  • One or more dedicated smallsat launchers
  • A long-duration storable bipropellant kick stage
  • A long-duration storable bipropellant lunar lander stage

LEO Micro-Depot
We’ll take them in order. First, for this scenario, we used a LEO micro-depot concept that I’ve been noodling on for some proposals over the past few years. The basic concept is a low-cost, single-launch depot that can be used primarily for refueling dedicated smallsat launcher upper stages and storable bipropellant kick states. This depot is actually capable enough that it could support much larger missions, and ones including LOX/LH2 stages, but its primary near-term mission would be allowing dedicated smallsat launch stages to access much higher energy destinations, with much larger payloads, than they otherwise could.

Artist’s Conception of a LEO Propellant Depot[note]I’m including this picture again, both because it’s awesome, and because I can[/note] (Credit: Brian Versteeg)

We didn’t really go into the technical concept for this depot in the paper, but because many of you may be curious, here’s the thinking behind some key features:

  • The micro-depot is comprised of a depot kit (the conical section and everything to the left) attached to a repurposed Centaur V upper stage. This kit is designed to be able to ride as a secondary payload on an ISS Cygnus mission with approximately 10-12 tonnes of initial propellant on-board.
  • The LOX and LH2 tanks of the Centaur V are repurposed after arriving on orbit–the LOX tank is reused as a primary depot LOX tank, allowing for the storage of over 50 mT of LOX onboard. The LH2 would probably be used to chill the LOX, helium, and other propellants down as far as possible, to extend how long the depot could last with minimal boiloff. In theory, the depot could be used for LOX/LH2 missions, but only if the LH2 was used quickly–in this configuration, and in LEO, it’s not designed to be able to store LH2 without boiloff for super long periods of time.
  • I show a conical sunshield, similar to previous ones studied by ULA, though I’m not positive at 400km if that trades better than just a bunch of layers of MLI. If you’re at too low of an altitude, the earth starts peeking into the conical sunshield, dramatically lowering its effectiveness. I haven’t created a spreadsheet yet that allows you to figure out what altitude you have to be around a given body with a given sunshield half-angle to make sure it works5.
  • The main body of the depot kit is the kerosene tank, which is inside the cylindrical part of the depot kit. In this particular design concept, I’m assuming that the kerosene tank is derived from a Centaur III LOX tank, and that the structures surrounding it are derived from Centaur III and V forward bulkhead elements. Reusing structural elements like that can sometimes enable lower development costs while still providing a reasonably efficient structure. The total capacity of LOX/Kerosene in this depot, assuming 50% boiloff margin on the LOX was somewhere around 40mT, which is enough to support dozens of interplanetary smallsat missions.
  • Hidden inside the conical section are the helium tanks and the plumbing that control the main Centaur V tanks after it has been repurposed as a depot element. The helium tanks would probably be kept right up against the Centaur V hydrogen tank, with insulation isolating them from the warmer parts of the depot, so they can be stored at cryogenic temperatures where the helium is higher density. It may be possible if desired to actually store the helium at low pressure inside the Centaur V hydrogen tank, if all of the LH2 is vented previously, but that depends on the depot CONOPS and how often it receives LH2 deliveries, and how often it has customers that need LH2.
  • The tanks around the outside of the Kerosene tank on the cylindrical section can be used for storable propellants for the kick stages and landers. This could be your traditional storable bipropellants (hydrazine/NTO), or any of the green variants being developed by companies such as Bradford, DSI, and Tesseract. I can’t remember the exact dimensions on these tanks, but there are eight of them, and I think they’re approximately 2-3 cubic meters each, so this may actually be able to support several missions worth of kick stage propellant.
  • The solar arrays are oversized for just powering the main depot functions (robot arms, rendezvous/prox-ops sensors, spacecraft control avionics, so it may be possible to have power available either for running a small cryo-cooler, or for things like converting water electrolytically back to hydrogen and oxygen, or into hydrogen peroxide and hydrogen.
  • There are shown six “Sticky Boom” style capture arms and two refueling arms. The Sticky Boom arms are scaled-up versions of the ones Altius is developing for its Bulldog satellite servicing vehicles, and they use our magnetic grappling technology for grappling the target stages6. These capture arms have reaches in excess of 10m, which can allow a relatively non-agile upper stage to rendezvous close enough to the station to allow the arms to capture the target and damp out relative rates. A trio of arms is used to allow for a parallel robotic structure, which is much stiffer than a series connection. The fueling arm is a more traditional 6DOF arm that can also support depot maintenance. Two sets of capture/fueling arms are provided to enable handling fueling operations with two stages simultaneously.
  • One last detail is that the reason I selected an ISS coorbital location was to take advantage of crew/cargo traffic to the station. Most ISS crew/cargo vehicles launch on vehicles with a lot of excess mass capacity that ends up going to waste. I’m not sure exactly how much wasted propellant there is on say a typical Dragon flight, or on Cygnus flight where you add a few extra solids, but in theory the numbers I’ve run suggest you might be able to get a better deal on this leftover prop than you could by buying a dedicated F9R or FH reusable launch, while still being economically interesting to the ISS crew/cargo launch companies. It is true that NASA isn’t intending to keep the ISS operating indefinitely, but I wouldn’t be surprised if commercial ISS replacements initially start in a similar orbit, which could create a similar dynamic.

Ok, that’s a lot more details on the LEO micro-depot, but I wanted to share some of my thinking, since we didn’t get a chance to go into it much in the paper.

Dedicated Smallsat Launcher Upper Stage
The next element is the smallsat launcher upper stages. For this paper, we focused on Virgin Orbit’s LauncherOne, as it is in the middle of the size range for the more credible smallsat launcher capabilities. But there’s no particular reason you couldn’t use a Rocket Lab Electron7, or a Firefly Alpha, etc. It should be noted that while these stages are not currently designed for rendezvous operations, we think there are a few credible paths forward that can require minimum modifications to the stages themselves. The Rendezvous/Prox-Ops (RPO) sensors would be on the depot itself, with some added avionics to receive commands from the depot and translate them into maneuvers that could be performed either by the kick stage, by upgraded RCS on the stages themselves, or even by using the engine purges as a sort of ghetto cold-gas thruster to augment their 3dof steering RCS thrusters. Additionally the stage would need some DogTag grappling interfaces, and fill/drain ports designed for in-space refueling8.

Since LauncherOne’s full stage performance specs aren’t yet currently available, we derived them from a mix of publicly available data9. Here’s the numbers we used for LauncherOne performance:

  • Propellant mass: 2415kg10
  • Dry mass: 329kg11
  • Stage Isp: 325s12
  • Payload to ISS-like LEO: 475kg13

It should be reiterated that LauncherOne by itself has almost no payload capacity beyond LEO. You could add a large storable kick stage and launch small payloads beyond LEO (likely <100kg net payload), but by refueling the upper stage and a small storable kick stage, the payloads are a lot closer to the full LEO capacity for not dramatically higher costs.

Storable Bipropellant Kick Stage
The next element in the scenario is a long-duration storable bipropellant kick stage. Because most rocket upper stages are not designed for missions much longer than even an hour, we only used the refueled/recharged upper stage to perform the first burn of the 3-burn maneuver–the apogee raise into the near escape-velocity highly elliptical phasing orbit. But there are still at least a pair of burns that need to happen after this–the final interplanetary injection burn that happens at periapsis of the final phasing loop, and any plane change and/or perturbation correction maneuvers that need to happen at apogee of the orbit. For these burns, we assumed the use of a storable kick stage, though in theory this could also be performed by a storable propulsion system integral to the spacecraft. Because the delta-V required for each of these missions is different, we assumed a sort of modular “dial-a-stage” for these calculations, based on Isp and structural fraction estimates from two companies developing storable bipropellant kick stages for smallsat launchers, Tesseract and Deep Space Industries. We took an average Isp, and used the worst number on the provided structural fraction curves, even though for larger stages this is probably excessively conservative. We assumed that the stages were infinitely stretchable with the structural fraction defining how much dry mass was associated with the desired propellant mass.

Here’s the specs we used for the paper:

  • Kick Stage Isp: 310s
  • Kick Stage Structural Fraction: 0.2514

We feel these are pretty darned conservative, and could be readily improved on with additional development. Also for most of these scenarios, we assumed the kick stage would be launched empty, and filled-up in LEO at the depot, though for most of the lower-energy missions launching the kick stage prefueled would probably impact the delivered payload by less than 20%.

Storable Bi-Propellant Lunar Lander
For the one lunar landing mission, we needed specs for a lunar lander, so we just took the same approach as the kick stage, but assumed a worse structural fraction:

  • Lander Isp: 310s
  • Lander Structural Fraction: 0.40

This may be overly pessimistic of a structural fraction, but we ran with it for conservatism sake, to cover things like landing gear, landing sensors, etc. In the lunar scenario, we assumed this was launched dry and filled-up at the depot.

Destinations and Departure Schedules
The following table shows the departure order and departure C3s used for this campaign:

Interplanetary Blitz Targets and Depature Conditions

Trajectory TargetEarth Departure DateDeparture C3 (km²/s²)
Jupiter22 Aug 202486.9
Mercury15 Sep 202454.9
Moon26 Sep 2024-1.99
Mars07 Oct 202411.12
2007 XB2302 Nov 20240.38
2008 EV505 Nov 20242.15
Venus28 Nov 202411.32
2001 QJ14202 Jan 20250.65
2009 HC29 Jan 20250.31

A couple of quick notes on this table before going on to the results:

  • The departure windows weren’t optimized much for arrival C3, and the missions were simulated as though they were flybys15, though as you’ll see from the performance specs, it would be quite possible with most of the payloads to include propulsion or aerocapture capabilities in the payload available to turn these into orbiters or lander missions.
  • All of the trajectories, including the Jupiter and Mercury missions assumed a direct trajectory without using any intermediate gravity assists. As can be seen for Jupiter this is a very high energy mission, requiring more than 6km/s of delta-V from LEO. Going to a Venus or Venus/Earth gravity assist would likely increase the payload to close to same range as is seen for Venus missions, at the cost of added mission time and complexity.
  • For the lunar landing, the C3 is negative, since the TLI burn is lower than escape velocity. For this mission though there is an additional 827m/s for lunar orbit insertion performed by the kick stage, and then 2150m/s of delta-V for the lander stage.
  • It should be noted that the closest two missions are only 3 days apart, which while ambitious should theoretically be possible–LauncherOne was designed for surge capacity, and the depot operations themselves should only take a few hours. But doing back-to-back missions like that would be a sight to behold.
  • It should also be noted that while we picked 2024 for the blitz window, that there’s a similar season where the planets align in a similar manner in 2026. We picked 2024 as that was the soonest we thought a depot capability could realistically be available.

Mission CONOPS
The missions in this campaign all used a fairly similar CONOPS:

  1. The mission stack–launcher upper stage, dry kick stage, dry lander stage (if used), and science payload/spacecraft are launched into LEO in preparation for rendezvousing with the depot. For air-launched missions, this can be readily done as a single-orbit rendezvous mission. For ground launch, you’d still want to limit the number of orbits prior to rendezvous due to the short stage life.
  2. The upper stage would then perform maneuvers to rendezvous with the depot16, which grapples and secures the mission stack.
  3. The depot would then refuel the upper stage with enough propellant for the mission, recharge batteries if necessary, and fuel the kick stage and lander (if present)17.
  4. The launcher upper stage would then depart the depot, and at a specific pre-planned time, would perform an apogee raise maneuver to place the kick stage, lander stage, and payload into a highly-eliptical phasing orbit. The launcher stage would separate as soon as this burn is completed, keeping the amount of time the stage needs to operate at a minimum.
  5. At each apogee during the phasing orbit, the kick stage will perform required plane change and perturbation correction burns. For some of these near-escape orbits, lunar and solar perturbations can move the perigee around that these maneuvers are necessary to avoid either hitting the atmosphere, or having the perigee raised high enough to negatively impact the final burn. These burns tend to be pretty modest in delta-V (typically <100m/s total).
  6. At the end of the final phasing loop, the kick stage would then perform the final injection burn that sends the payload into interplanetary space. In this scenario, for all missions other than the lunar landing mission, we assumed the kick stage would then be jettisoned.
  7. For the lunar landing mission, the kick stage would stay attached to perform the lunar orbit insertion maneuver when the stack arrives at lunar orbit, inserting the lander into a low (250km) lunar polar orbit, at which point the kick stage would be jettisoned. The lander would then perform the descent to the lunar surface.

There are tons of variations on the theme that could’ve been used, and I didn’t have time to create one of those cool mission CONOPS illustrations, but this gives the general idea of the approach used.

Interplanetary Blitz Results
Using the 3-burn methodology described in the previous section, we were able to complete all nine of the missions from an ISS-coorbital depot. The following table provides a summary of key results, including providing some numbers on the delta-V and trip-time penalties incurred for using the 3-burn maneuver, the total delivered payload, and the amount of propellant that would need to be loaded at the LEO micro-depot:

Interplanetary Blitz Payload Results

Destination# of Phasing LoopsTotal ΔV Penalty (m/s)Flight-time Penalty (days)Total Net Payload (kg)LauncherOne Prop (kg)Storable Prop (kg)
Jupiter153.425.9922415769
Mercury181.629.22852300571
Moon2negligible10.71192415671
Mars142.812.24391497108
2007 XB23166.621.3467137323
2008 EV5122.324.8466138628
Venus147.416.14351519119
2001 QJ142116.719.0470136316
2009 HC235.122.8468135021

Here are some key takeaways from these results:

  • As can be seen from this campaign, even though the depot was often pretty poorly aligned at the departure date for a single-burn maneuver, the delta-V penalties for using the approach were very modest–less than 100m/s, and the trip time penalties were all less than one month.
  • Even to an extremely high-C3 trajectory (the Jupiter direct trajectory), this method still provides a pretty substantial net payload. For the high-C3 missions there are several ways that could improve the total net payload, including using flyby trajectories, adding an additional kick stage to split the injection burn into two segments, using a higher-Isp kick stage (like a LOX/Methane storable, or a cryo-cooled LOX/LH2 stage), doing the mission as a interplanetary boost followed by a low-thrust/high-Isp SEP mission, etc. But the fact that using this methodology LauncherOne could send over 90kg on a Jupiter direct trajectory is pretty crazy when you think about it.
  • Even with the really poor structural fraction assumed for the lunar lander, we’re still talking almost 120kg of net payload to the lunar surface using this approach. Which is pretty impressive when you think about it.
  • The Mars payload is around 1/2-2/3 of the injection mass of the Mars Insight lander. Which was launched on an Atlas V launch vehicle, which is around 10x bigger than LauncherOne, and has one of the highest performance upper stages in history.
  • For the lunar mission, you don’t really need to use the 3-burn departure approach unless you’re either a) trying to rendezvous with a depot or other facility in lunar orbit, in which case you could probably increase the net payload to the lunar surface pretty dramatically, or b) if you were trying to land at a very specific local lunar time. Otherwise, the Moon has optimal 1-burn departure opportunities every 7-10 days from an ISS-coorbital depot.
  • Most of the burns had the best payload with a single phasing loop, though for about half of the trajectories, the difference between one loop and three loops was in the noise (<1-2kg). Many of the trajectories had lower delta-V penalties on the 2 or 3 phasing loop options, but had less delta-V provided by the upper stage, which is the most efficient from an Isp and structural fraction standpoint.
  • While we didn’t analyze it, it’s pretty clear that a long-lived, high-Isp/high-pmf stage like ACES or Centaur III with IVF would be pretty amazing, since you could have the stage itself perform all three burns.
  • For most of the asteroid missions, the kick stage is really small, and similar enough in size that you could probably make a “one-size fits all” kick stage for asteroidal missions using this 3-burn departure methodology. And it would be a pretty small stage–less than 50kg wet.
  • Most of the missions didn’t require refilling the LauncherOne stage much more than about halfway. Only the very high C3 missions to Jupiter and the lunar surface needed a full LauncherOne.
  • For Rocketlab Electron payloads my gut suggests that you could multiply these results by ~50% (since it’s about half the payload to LEO but similarish performance), and for Firefly Alpha, multiply approximately 2x (since it’s twice as big). It wouldn’t be hard though to run the numbers for a different launch vehicle using the data in the paper.
  • Performing all nine of these missions would require ~15.6mT of LOX/Kero (not counting boiloff losses) and about 2.3mT of storable bipropellant. This is about the amount of net payload that could be launched on the first depot launch, as a secondary payload to Cygnus, if you added 3-4 solids to the Vulcan/Centaur (call it ~$40M in net depot launch costs if you include a $10M payment to NGIS to entice them to use Vulcan instead of Antares).
  • We didn’t go into the economics of the concept, but if the depot cost $100M to develop, this means you could do the depot development, launch, commissioning, and all nine of these missions for a total cost of less than $250M. If you assume the asteroid missions all use the same spacecraft design/payloads, you could do this complete blitz for less than the cost of a single NASA Discovery mission if you could keep the spacecraft design/fabrication costs below about $40M/each. The previous calculations we did a few years ago suggested that it would be possible to make a decent profit off of 2-4 missions per year, and a price point of around $25M for a dedicated deep space LauncherOne mission, and around $15M for a dedicated deep space Electron mission. This is much cheaper than buying a whole Falcon 9 mission anytime in the foreseeable future, and to many of these destinations, if you want to go at all, you’re unlikely to get many secondary payloads who can use your trajectory, so your main alternative would be buying a full Falcon 9.

Conclusions and Next Steps
I think with this paper, we’ve successfully retired concerns about LEO depots being a viable platform for enabling deep-space missions. While we would still like to run some analyses showing what the worst-case delta-V or trip-time penalties look like even if your depot RAAN has the pessimal alignment for a given mission18, the data from the Interplanetary Blitz campaign preliminarily suggests that the penalties are likely pretty minor. The possibility of enabling 100-400kg class payloads to be sent almost anywhere in the solar system for launch costs in the <$25M range could be a game changer for the interplanetary community. While I don't think this would replace what NASA does with flagship, New Frontiers, or Discovery class missions, lowering the cost of a useful interplanetary mission into the $50-60M range could enable more space agencies and non-space agencies to participate in interplanetary science, enable more frequent visits to destinations that don't get enough love currently (Venus, Neptune/Uranus, etc), enable companies that would like to launch smallsat-class MEO or GEO (or lunar or Martian) telecoms relay constellations.

Posted in Commercial Space, International Space Collaboration, International Space Competition, Launch Vehicles, Lunar Exploration and Development, Mars, NEOs, Orbital Dynamics, Propellant Depots, Space Transportation, Technology, ULA, Venus | 5 Comments

AAS Paper Review: RAAN Agnostic 3-Burn Departure Methodology for Deep Space Missions from LEO Depots (Part 1 of 2)

Artist’s Conception of a LEO Propellant Depot (Credit: Brian Versteeg)

About a year ago, I wrote a review of an AAS conference paper that I coauthored with a few of my astrogator friends, Mike Loucks and John Carrico regarding an mission design tool for enabling the use of LEO depots for deep-space missions. At this year’s AAS/AISS Astrodynamics Specialist Conference in Snowbird, Utah, we did a follow-on paper, with the help of Altius’s Matt Isakowitz Fellow, Brian Hardy, and I wanted to provide a review of this paper, since it was a lot of fun, and I think extremely relevant and timely. As with last time, the paper will be published in a future volume of Advances in the Astronautical Sciences1.

Before I review the paper, here’s a full-text copy for reference: AAS 18-447: RAAN-Agnostic 3-Burn Departure Methodology for Deep Space Missions from LEO Depots

Backstory/Introduction
As a quick reminder of what led us to develop these mission planning techniques, or for those who haven’t had a chance to read the previous blog post, back in 2011 when there was a lot of NASA interest in orbital propellant depots, some flight dynamicists at NASA Johnson Space Center raised a serious concern about the feasibility of using LEO propellant depots for deep space missions. The tl;dr version of this argument is that for any given interplanetary departure, you have to leave along a certain V-infinity vector, and for a reusable LEO depot that wasn’t just launched for this specific mission, the odds that the depot plane would align with that V-infinity vector at the right time was small. You could launch a depot per-aligned for one specific mission, but the odds of it then lining-up correctly for any particular future opportunity was small enough (<25%) to make LEO depots impractical.

What we did was come up with a 3-burn departure that would allow you to leave a LEO depot into a phasing/alignment orbit that would put you back at perigee, in the right place, at the right time, and with the right alignment to do your planetary injection burn, even if the depot’s plane wasn’t aligned with the departure vector at the departure time. In fact, we found that in many cases it was possible to mount a deep space mission from a depot even if the depot plane never intersects with the V-infinity vector (i.e. if the declination2 of the departure asymptote3 is higher than the inclination of your propellant depot’s orbit), so long as it’s close enough. What this means is that you could have a LEO depot that you refill and reuse multiple times for a wide range of missions without having to move the depot around to line things up for a given mission. Which is kind of important for a depot to be economically useful.

In our first AAS paper, we described the genesis of the 3-burn methodology, which was actually a paper by Selenian Boondocks alumni Kirk Sorensen, and showed how it could be used to enable a Mars mission or a mission to a NEO with a very high declination angle (2007 XB23). However, to simplify things for the first paper, we assumed a phasing orbit with a specific apogee altitude, which basically still required you to align the depot plane with that phasing orbit, which kind of defeats the purpose. We knew we could use this technique for enabling departures from a depot regardless of what its RAAN4 was at the time of the departure window by varying the altitude of the phasing loop, but we hadn’t been able to take things that far by the time we had to present last year’s paper.

So the purpose of this paper was to flesh-out the methodology showing how you could use it for missions regardless of where the depot plane was at the desired departure time. Also, to illustrate how powerful this capability was, we illustrated the use of this RAAN-agnostic 3-burn maneuver for enabling a rapid-fire series of deep-space missions from a single LEO depot–4 planets, 1 moon, and 4 NEOs in a 5 month timeframe. Without further ado, I’ll dive into the work we did in this paper.

Methodology Refinement
We described the methodology in a more rigorous manner in the paper, but here’s a quick summary:

  1. Identify the desired departure geometry (C3, declination and RAAN of the departure asymptote, the resulting locus of periapses5, and departure date), and determine the orbital parameters of your depot at around the time of your planned departure.
  2. Check if a simple one-burn departure is possible–the odds aren’t great, but if the plane happens to be lined-up correctly, may as well keep things simple.
  3. Calculate when to enter the phasing orbit–if your depot isn’t aligned with the departure asymptote at the departure date, you need to enter a phasing orbit the last time your depot was optimally aligned. Because your depot plane precesses over time, you can time-step back to the last time you were aligned properly, and have that be the time you do the injection burn to enter your highly elliptical phasing orbit.
  4. Design your phasing orbit–first you calculate how long you need to be in the phasing orbit, and then you can pick a one, two, three, or four loop phasing orbit, with the loops taking some integer fraction of the required phasing time. Lastly, using a high-fidelity simulator you will want to add in required plane changes and/or perturbation correction burns at the apogees of the phasing orbits.
  5. Calculate the final departure burn and tally the required Delta-Vs for each of the maneuvers.

While for the mission simulations we did in the paper we mostly eyeballed several of the steps and then used targeting algorithms to correct for eyeballing-errors, it should be possible to automate these steps6.

In the process of designing this methodology and exercising it, we learned several lessons worth mentioning (in no particular order other than what I could think of when writing this summary):

  1. If the declination of the departure asymptote is lower than your depot inclination, the lowest delta-V departure will happen if you enter your phasing orbit the last time the depot plane intersects the departure asymptote7 prior to the departure date. In this case, you don’t have to do a plane change to align for the departure, just corrections for lunar or solar perturbations.
  2. If the declination is higher than your depot’s inclination, but the angular extent of the locus of periapses8 is larger than the difference between the two (ie if your depot plane at any point crosses through the locus of periapses), you can still use this 3-burn departure methodology, you’ll just have to do a plane change at apogee to align your final departure plane with the departure asymptote. Since this plane change takes place at near escape velocity, the cost of the plane change can be very modest. The delta-V optimal timing for this orbit would be at the last time where the depots orbital plane came closest to intersecting with the departure asymptote9.
  3. The angular extent of the locus of periapses is a function of the injection C3. The faster you have to leave the earth, the wider that locus is. So for a medium-inclination depot (such as one in an ISS-coorbital plane), the only missions you can’t use the 3-burn departure method for are a few NEO missions with high declinations but very low C3. Those are fairly rare, and there may be more complicated departure methodologies that can enable these, but one brute-force solution would be to have a small depot in a near-polar orbit.
  4. For either case, the solution with the lowest total trip time (including phasing orbit) will occur if you enter your phasing orbit the last time the locus of periapses intersects your depot orbital plane prior to your departure date10. In this case you’ll definitely need a plane change at apogee.
  5. As mentioned previously, phasing orbits don’t have to be a single-loop. You can actually go for anywhere from 1-4 orbits while still keeping the orbit elliptical enough to freeze your plane’s orbital precession.
  6. Phasing orbits with several smaller loops tend to be less susceptible to solar or lunar perturbations, which will vary in magnitude depending strongly on where the moon is relative to your departure asymptote and your phasing orbit11. On the other hand, with smaller numbers of phasing loops, more of the departure burn is performed by the refueled upper stage, which typically is higher performance than the kick stage(s). Long story short, you’ll want to check the 1, 2, 3, and 4 phasing loop options to see which is performance optimal for a given mission, because it’ll vary.
  7. Worst case trip-time penalties that we saw were less than 45 days. For a robotic mission, this is probably not an issue, but for a human spaceflight mission, these could be an annoying penalty. One way to solve this would be to use a 3 or 4 loop phasing orbit, and use the depot to fuel and launch everything in the departure stack other than the crew, and then have the crew launched separately only during the last phasing loop, meaning you could keep the trip-time penalty for the crew below ~10 days, and only add two extra Van Allen Belt crossings, at the expense of requiring a launcher that can send the crew capsule into the same highly-elliptical phasing orbit as the mission stack12.

I’m going to take a break at this point to keep the blog post from getting too long. In the second half of this review, I’ll go over the Interplanetary Blitz campaign I mentioned in the introduction.

Posted in Launch Vehicles, NEOs, Orbital Dynamics, Propellant Depots, Space Transportation | 9 Comments

Random Thoughts: Why Cameras Might be Critical to Venus Settlement

It’s been a while since I last posted on the idea of Venus settlement, but the idea came up again on Twitter recently, and it got me thinking about several of the challenges that still need to be resolved to make it a reality. On the technical side, the big ones are still: a) can we extract enough water or hydrogen from the atmosphere to serve as a feedstock for life support needs and plastic production for the habitats, b) can we find a fully-reusable, robust/fault-tolerant way of traveling between cloud cities and orbital facilities, and c) can we realistically get from ISRU feedstocks to practical cloud colony materials that provide the needed functionality while being compatible with the still somewhat harsh environment in the Venusian atmosphere.

But as interesting as those questions are, the question of why someone would want to settle the atmosphere of Venus is probably even more fundamental.1 You’re probably wondering what this has to do with cameras, but I’ll get back to that in a bit. First I want to talk about the economics of settlement.

The Economics of Venus Settlement
I’m not trying to do a detailed treatise on space settlement economics in this short blog post, but I did want to touch on a few ideas I’ve had on the topic.

First, regardless of how good you get at ISRU, you’re almost certainly going to need to import at least some materials. Even if you can get all of your life support materials, and most of your construction materials from the Venusian atmosphere (the Massive, Unitary, and Simple part of Peter Kokh’s MUScle framework for ISRU), you’re still likely going to  be importing electronics and complex equipment for a long time (the complex, lightweight, and expensive parts of the MUScle framework), and until you can get surface mining capabilities, you’ll likely need to import metals, and any biomass items that you can’t get from the atmosphere. If you’re importing stuff from Earth, this implies the need to provide something in return. Now, I don’t want to get into all the complexities of real world trade economics, just suffice it to say a Venus settlement will most likely need at least a few economic drivers.

While there likely are others I’m not thinking of, the three best answers I’ve been able to think of to-date for economic drivers for a Venus settlement are:

  1. Extraction of fusion fuels (Deuterium and maybe Helium-3) from the atmosphere.
  2. Tourism
  3. Immigration to the colony

I could go into it in more detail later, but the first concept is based on the fact that Venus has a Deuterium to Hydrogen ratio that’s over 100x higher than exists on earth. We don’t have fusion reactors yet, but Deuterium is likely to be important, and while we can get some form seawater, if interplanetary transportation became cheap enough, it might be possible to profitably extract Deuterium from the Venusian atmosphere and ship it back to Earth and other places that need it. This seems a lot simpler than concepts of strip mining vast regions of the lunar surface for Helium-3. Speaking of Helium-3, I wasn’t able to find any data on the Helium-3/Helium-4 ratio in Venus’s atmosphere. The Helium concentration on Venus is about 2x that of Earth, but I’m not sure whether we ought to expect it to be concentrated with He-3 (implanted from the solar wind?) or depleted in Helium-3 (since it is light enough to be lost to space like most of Venus’s hydrogen) relative to Earth. If it turns out the He-3/He-4 ratio is enhanced relative to Earth, that could also provide a potential export, and would likely be a lot easier to implement than most lunar He-3 extraction concepts (while also being a lot easier to reach than the outer gas giants).

But really any resource play doesn’t necessarily require a lot of people. It might actually be possible to get some of those materials via orbital atmospheric mining without ever coming down from orbit.

Tourism definitely kind of requires people to be there. That’s kind of the point. And those customers will require people to run the experience. Venus tourism would still involve much longer trips than lunar or LEO orbital tourism, which will likely make it more like going on safari during the 19th century than going to Disneyland. But still, it could be a legitimate economic driver.

Lastly, settlement itself can be an economic driver if people want to immigrate to a place. If people want to move some place to live or retire, they bring their wealth with them, effectively importing or investing that wealth in the Venusian economy in a way that can be used to pay for imports from Earth.

But those last two items strongly depend on something most engineers don’t think much about–the aesthetics of the place. And that’s where cameras come in.

Why Cameras Matter
While Russia did manage to send a pair of balloons to explore the region of the atmosphere we’re interested in, as part of the Vega 1 and 2 missions, neither of those balloons had a camera on board. Some of the landers had cameras, but as far as I can tell, neither balloon had one. They had atmospheric sensors and photometers and a few other sensors, but nothing that could show you what it really looked like inside the cloud layer of Venus.

And frankly, when it comes to tourism and settlement, I wouldn’t be surprised if the look and feel of that region of the Venus atmosphere matters a lot.

For instance, is Venus more Lando Calarisian-esque:

or are we talking more like a Beijing smogfest?

You may think this is a trivial matter, but I think it probably matters a lot. It’s one thing to go to a flying city with breathtaking views and stunning vistas. It’s another to be flying around in pea soup smog so thick that you may as well not even have windows2.

So, this is why I hope we see some balloons visiting the atmosphere of Venus again sometime soon, and this time, I hope they bring cameras. I’m keeping my fingers crossed that the view is amazing.

 

Posted in Random Thoughts, Space Settlement, Venus | 18 Comments

Airbreathing hypersonic travel is less energy efficient over long distances than rocket travel

There’s a certain misunderstanding common in aerospace that rockets are horribly inefficient and that long term we need air breathing ramjets or scramjets to efficiently launch things, with the idea that we can thus avoid accelerating oxygen to flight speed, which is considered wasted energy. “Airbreathing hypersonics are five times as efficient as rockets” they say. This, however, is not so.

The misunderstanding comes in part by considering oxygen as just as costly as fuel. Oxygen is not. It can be condensed out of the atmosphere with little energy and is available by the truckload at $100/ton or less. A dedicated production plant can produce it for as low as $10/ton. That compares to $1200 to $3500 per ton for industrial liquid hydrogen which is often the fuel being compared to.

A stoichiometric rocket burns 8 times as much oxygen as it does hydrogen. So if an airbreather consumes a factor of 5 times less propellant than a rocket, that means it consumes nearly twice the hydrogen!

Hydrogen requires the vast bulk of the energy to produce compared to oxygen, a couple orders of magnitude more energy. So for our purposes we can ignore the energy needed to produce liquid oxygen.

Let’s look at LAPCAT II, and airbreathing hypersonic airline concept capable of traveling to the antipodes of the world at Mach 8.

As a percentage of its gross takeoff weight, 45% is hydrogen fuel and 15% is payload: http://www.icas.org/ICAS_ARCHIVE/ICAS2014/data/papers/2014_0428_paper.pdf

That means each kg of payload requires 3 kg of liquid hydrogen, which has an energy density of 142MJ/kg, giving an energy cost of 426MJ per kilogram of payload.

Hydrogen with variable mixture from oxygen rich to near stoichiometric would be the best fuel to compare with and the most efficient for rockets, but I will use SpaceX’s ITS from 2016 as a comparison point even though it’s less energy efficient.

https://SpaceX.com/Mars

ITS has a payload to LEO of 300 tons (more for the tanker variant), and uses a total of 6700 tons of propellant for the first stage and 1950 tons for the second stage ship (both including landing propellant). Given a O:F weight mixture ratio of 3.9, and a specific energy of 55.5MJ/kg for methane, the cost per kg of payload to orbit is just 330MJ, actually less than the hypersonic airliner in spite of using less efficient methane.

You might as well use rockets for long distance transport at high Mach numbers.

Posted in Uncategorized | 37 Comments

Administrivia

Hey guys, just FYI at Chris’s request (and with some help from Mike Mealling), I updated the website to add SSL encryption and use https instead of http. Apparently, Chrome is about to start giving people warnings if they visit sites that aren’t secured properly.

Also, it’s been a while since I’ve blogged–I’ve been in proposal hell for a while and am only finally coming out of that mode, but I wanted to mention that I’m planning on doing some posts soon reviving the previous thread on Venus ISRU and Settlement issues. More later.

~Jon

Posted in Administrivia | 1 Comment