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

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.


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


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.


Posted in Administrivia | 1 Comment

Plan D for space settlement

Plan D
There are three companies I take seriously for making true spacefaring (ie including Mars because I’m a Mars Firster) truly accessible: SpaceX, Blue Origin, And Masten Space Systems. I would have taken XCOR seriously, but unfortunately they went bankrupt.
The other three:
1. SpaceX. By far the top of my list. Fast execution, well-capitalized when they need to be, sustainable, good business plan to scale up to $100 billion level, and great architecture. Actually hard to improve on this one. But SpaceX got where it is on the shoulders of Elon Musk and by taking a lot of risks. The flip side of that is one of their bets could go far south, or something happens to Elon. Don’t want to rely on one, particularly risk-taking, company.
2. Blue Origin. Somewhat a mystery, but ridiculously well capitalized. Sustained by brute force money injections, not (much) actual business yet. Similar near-term architecture to SpaceX, but slower & not quite as aggressively low cost. Not easily extendable to other planetary bodies without separate development (which apparently they’re doing with Blue Moon). Moon and free space focused, so I wonder if they’ll even get around to Mars before I’m elderly.
3. Masten Space Systems. Very small, poorly capitalized, but actually pioneered a lot of the reusable tech SpaceX uses. More experience with reusable rocket vehicles than anyone. Was looking like a real possibility for highly reusable launch before Boeing sadly won the XS-1 DARPA bid. Now has pulled back and seems focused on small commercial lunar landers. But unlike XCOR, they’re still in the game.

Plan D? Still thinking about it. But I think a rapid return to launch pad thing like BFR and Masten is a good plan, although ambitious. Fast integration of upper stage is key as well. I like Jon’s idea of an oxygen-rich hydrolox architecture.

Posted in Uncategorized | 11 Comments

Your Sweat Given Rights (Off Topic)

I would like to introduce a possible new term to help counter the seemingly increased use of “God Given Rights” by lawyers, politicians, and various activists that wish latch onto or hand out other peoples’ resources and freedoms. This particular rant was inspired by an ad on the radio today by a lawyer. “You have a God Given Right to a good job, a living wage, a safe living environment, good food, and so on. Followed by a list of possible reasons to call him so he could force them to deliver yours. Included were brags about how much he had won for deserving clients.

I would like to propose the sound bite “Sweat Given Rights” to indicate to people that all these desirable outcomes have to be paid for by the sweat of  someone, often someone else. That good job was created by someone that earned it, often with long hours and personal financial risk. Same with safety, nutrition, medical care, and education among other desirable  services most of us want. Someone has to sweat to provide whatever desirable thing we want and I think it is past time for the word war to start a reasoned counter attack. Reasoned in today’s world mostly isn’t going to be scholarly tomes invoking the Constitution, or interpretations of biblical verse that are obscure to any not already knowledgeable.

I would prefer TANSTAAFL  except that it doesn’t sound bite well enough to get through to many that aren’t going to research issues. Aggressively looking for simple soundbites to counter bad memes might well be one of the important tools to stemming some of the abuses by well meaning people that are following  bad leaders and getting bad advice.

This phrase doesn’t imply throwing people out on the streets to die. It simply reminds some that might listen that TANSTAAFL.


Posted in Uncategorized | 28 Comments

Megacharger costs… (Tesla Semi Part Two)

Elon Musk announced the Tesla Semi months ago, now. Besides the low cost, one of the things people are most incredulous about are the Megacharger costs. Tesla announced just 7 cents per kWh, flat price, to charge at a Megacharger. And that’d be done with solar power, potentially even unhooked from the grid. How is this feasible?

First we must look at existing Tesla Superchargers. Superchargers are capable of up to 145kWh charge rates (internally). Some versions were made by simply ganging up multiple home charger units together to get the required power (like 12 individual 12kW units). The high power was produced by clustering mass-produced smaller units that also are provided with each car. Just a single connector goes to each car, however.

Superchargers use a very large amount of power. A bunch of supercharger stalls being used at once at a Supercharger location can draw Megawatts. Simply installing a megawatt connection to the electric grid is expensive. So Tesla has started installing Powerpacks (the larger versions of the Powerwall designed for utility and commercial installations… around 210kWh apiece… themselves composed of 16 individual battery packs similar to those used for the Model S/X/3 and each with a DC-DC converter, with a DC-AC inverter of between 50 and 500kW of output, depending on configuration). This allows Tesla to add additional supercharging stalls to existing supercharger locations without upgrading the utility connection, and for new locations allows them to avoid utility peak charges which could actually end up being larger than the actual energy usage charges. As a side benefit, it also means that Superchargers have backup power in case of power outages: https://electrek.co/2017/10/30/tesla-supercharger-stays-online-in-power-outage-powerpack-system/

…and in principle, Tesla could also optimize when the Powerpacks are charged to minimize time-of-day charges. That gives Tesla access to sub-7-cents-per-kWh electricity prices already. Industrial electricity prices are around 5 to 8 cents per kWh on average in the US (exception is Northeast, with about 9-12 cents per kWh), with off-peak electricity being about 2 or 4 cents per kWh less.

So Tesla probably ALREADY pays less than 7 cents per kWh for electricity on average. But they also have a significant amount of capital cost in the form of the chargers themselves and the batteries. The Batteries go for about $400/kWh retail, but Tesla’s internal price may be more like $150-250/kWh, especially without the inverter. That’s about 3 cents per kWh if they last for 20-25 years (which isn’t actually too unreasonable, given careful charging and discharging). 4 cents is more reasonable. But long-term, they hope to get cells below $100/kWh, and packs at, say, $120/kWh. (Raw material costs are about $35-45/kWh.) So <2 cents per kWh for the packs themselves is feasible, especially if they can get them to last a while. And ultimately, these packs can be assembled in an automated fashion, like their Model 3 packs. In fact, they could actually use the same battery line. (The biggest argument against automation is, like reuse, that it's not worth it at the volumes Tesla is considering, but if the same line is producing fairly standardized packs for multiple uses, that can dramatically improve the automation business case.)

But if you already are using battery packs for peak power reduction and maybe even time of day shifting, then the idea must occur to you: why not get rid of the utility entirely?

Solar cells are currently as low as 16 cents per Watt on the spot market (average 17.5 cents), without federal subsidies: http://pvinsights.com/. That means 16 cents of cells produces 16 cents of electricity (at 7 cents per kWh) in a SINGLE year in a place like the American Southwest that has a capacity factor of about 26% or better, paying back the cost of the cells. Modules are more, obviously, but still cheap at 27 cents per Watt. If Tesla can automate installation, they may be able to install them and string them together for less than 40 or 50 cents per Watt. That doesn’t pay for a connection to the grid or the inverter. Because Tesla doesn’t need those. In fact, the batteries already contain a DC-DC converter. Careful selection of voltages could allow a nearly direct connection of the solar panels to the batteries, perhaps with a small and cheap “power optimizer” (i.e. DC-DC converter) to improve solar array efficiency. These things only cost a few cents per Watt at utility scale, so let’s call it an even 50 cents per Watt. But Tesla can avoid the grid connection at both the solar array side AND the Supercharger side. And can avoid the cost of the inverters and the inefficiencies/losses from converting to and from AC. That previous 210kWh Powerpack thus has more like 225kWh. And over 25 years, therefore, a solar array that costs just 50 cents per Watt to install means electricity at less than 1 cent per kWh (0.9 cents) in a place with 26 percent capacity factor. However, there are sometimes cloudy days, where solar arrays will be less effective. To counter-act that, we make the solar array about twice the size. Cost per kWh doesn’t quite double, however, as not all the balance-of-system costs double. So let’s say 1.5 cents per kWh. Batteries also need to be about double, so cost of the battery is about 4 cents per kWh total, maybe less. But total cost of electricity is thus just 5.5 cents per kWh, leaving room financing costs. Doable if financing can be kept at low costs (and the solar arrays can actually last a LOT longer, like 50-100 years… and by doubling up the batteries as we did, they can also last a lot longer). And remember, solar costs will continue to decrease, long-term (tariffs notwithstanding). 3 cents per kWh raw solar+battery costs and 5.5 cents per kWh assuming roughly doubly up number of both li-ion and photovoltaic cells.

So that’s how Tesla can offer 7 cents per kWh for the Tesla semi while disconnecting from the grid and using solar power.

Posted in Uncategorized | 7 Comments

Repost of an Idea

Some years ago I did a post on using Lunar fuel to raise a sub-orbital vehicle to Earth orbit. One of the comments by sjv linked to a similar concept that had been done several years earlier but much more professionally by a far more qualified person. The recent FH flight and the more Lunar focused interest at this time makes the idea more relevant than 10 years ago when I blogged it or 14 year ago when Dr. Walthelm published his work.

What makes it more relevant is the possibility of orbiting a hundred tons with one reusable F9, three hundred with one reusable FH, and both without expending an upper stage. The Blue Origin offerings won’t be far behind if the capability comes to pass. BFR payloads to the four digits. The expendable industry could be mostly extinct in a decade or so except for niche one offs. If there is any compelling reason to get tens of thousands of tons into Earth orbit, this could create an early profitable Lunar export.

This is the link to Dr. Walthelms concept. http://www.walthelm.net/inverted-aerobraking/main.htm

This is my original post. https://selenianboondocks.com/2008/11/earth-launch-with-lunar-fuel/  Comments are better than the post.

If I got the link wrong, this is the post.            Earth launch for heavy vehicles currently involves lifting a lot of propellant to lift a lot less vehicle to lift even less payload. One of the frequent criticisms of suborbital flight is that it only uses a small fraction of the energy required to reach orbit. While that argument has some serious validity issues, it would be nice to be able to pop up a suborbital vehicle and hand off a payload to something else that took it to orbit. Mass ratio required would drop by a factor of 5. While it would be nice, it may be a while before some magic tech makes it possible.

Tethers are the frequent solution suggested for making this happen. Unfortunately current tech seems to be that your suborbital vehicle would need to be traveling at mach 15 or so to match velocities with a rotovator. While this will be a major breakthrough when it takes place, it still requires a serious performance vehicle to make the rendezvous. The suborbital vehicle for that mission will have to be fairly aggressively designed if it is a single stage. The mass ratio is about half that of an orbital vehicle with serious TPS still required. It will have to ride the tether around for a few orbits to get home, or land far enough down range that getting home is another logistics problem. As better tethers become available, these problems will slowly get better until a true beanstalk becomes possible.

I’m not aware of any other feasible technologies for doing the job of an advanced rotovator.

With airless aerobraking propulsion, there is a possible solution.  A chunk of lunar LOX launched into a near earth reentry trajectory will be at over 11 km/sec as it makes a near approach 100 miles up. If one pound of this impacted a suborbital pop up vehicle that had no significant horizontal velocity it would deliver an impulse equivalent to an Isp of 1,100+. If that pound vaporized and rebounded at random from a heat shield, it would deliver equivalent of another Isp of 550. So each pound of lunar volatiles would have an effective ‘Isp’ of 1,650 while the suborbital vehicle is motionless earth relative. As the vehicle gathered momentum, ‘Isp’ would drop as a linear function of less impact velocity of each succeeding LLOXball. By the time the suborbital vehicle was pushed to orbit, ‘Isp’ would be down to about 460 or so. Lunar regolith aerogel was suggested for the airless aerobraking. If feasible, that would solve several problems with the concept.

It works out to about 1.5 times as much LLOX as vehicle to make the push to orbit. A one ton upper stage with heat shield would need about one and one half tons of LLOX impact to push it to orbit. The size vehicle it would take to get that one ton inert upper stage into position is in dispute by the various people that build actual hardware. The old V2 would have used 4 tons of vehicle and 9 tons of propellant. There are at least a half a dozen credible newspace companies that believe they can beat that with a vehicle that flies daily or more. The list of less credible is somewhat more extensive. The list of companies that can place a ton in orbit without help is fairly long, and fairly expensive.

If a firm can just match the old tech and get an upper stage boost from the moon, then a ton to orbit will be considerably cheaper than is currently possible. This is a 14 ton earth GLOW and 1.5 lunar volatiles per ton to LEO. Heavy lift is the field that would make this pay. A modern expendable design for this purpose would have a mass ratio of about 2.5 and a dry mass of less than 10%. A 3,000 ton GLOW (Saturn5 class) would get 900 tons in orbit in one shot with help from 1,350 tons of lunar volatiles in intersect trajectory.

If it becomes desirable to get a lot of large payloads from the earth surface into space, this might be one path for doing it. If a suborbital craft can fly often, then it could launch large payloads once a day as it phased with the moon launched trajectory paths. It would be cheaper to use lunar raw material to facilitate earth launch than to manufacture finished components on the moon for the short term of a few decades. If SPS became economically desirable, this is a technique that could help make it possible to launch millions of tons of earth built products into orbit.

Posted in Uncategorized | 8 Comments

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