FISO Telecon Lecture on LEO Propellant Depots for Interplanetary Smallsat Launch

Posted on November 28, 2018 by Jonathan Goff

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)

Posted on September 17, 2018 by Jonathan Goff

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)

Posted on September 15, 2018 by Jonathan Goff

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

Posted on July 21, 2018 by Jonathan Goff

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

Posted on July 18, 2018 by Chris Stelter

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

Posted on July 17, 2018 by Jonathan Goff

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

Plan D for space settlement

Posted on July 15, 2018 by Chris Stelter

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)

Posted on July 5, 2018 by johnhare

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)

Posted on May 5, 2018 by Chris Stelter

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

Posted on March 18, 2018 by johnhare

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