Propellant depots for interplanetary flight.

[RGClark]

Key points about a propellant depot based Mars architecture, once the propellant depots are in place at both departure and arrival points:

1.)One single medium-lift booster first stage, Falcon 9, Atlas V, Delta IV, etc., delivered empty to orbit can then do ALL the propulsion from LEO departure, to Mars orbit insertion, to Mars landing, to Mars liftoff, to return to Earth.
No Saturn V, Constellation, Ares V, SLS, Mars Colonial Transport, or even Falcon Heavy required. The required boosters are already existing IF those propellant depots are already in place.

2.)SpaceX has shown that you can do reentry burns in the hypersonic airstream with the F9 first stage reuse tests. Then the problem of landing large masses on Mars is solved by doing a fully propulsive burn to Mars landing once that one, single stage is refueled in Mars orbit.

3.)That one single mid-lift stage could also be used to make an approx. 30 day flight to Mars. No VASIMR, solar electric or nuclear propulsion required. However, very high reentry velocity heat shields, ca. 20 km/s instead of ca. 6 km/s, would need to be developed for this.

4.)The most important point of all: getting the propellant depots to cislunar orbit is easy using near Earth asteroids. You don't need to use the Moon's proposed water ice deposits or develop a manned lunar base. This was the most surprising calculation of all: a single Centaur upper stage, of ca. 20 mT gross mass, could drag a 500 metric ton asteroid to cislunar space.

See:

Propellant depots for interplanetary flight.
http://exoscientist.blogspot.com/2015/08/propellant-depots-for-interplanetary.html


Bob Clark

 

[RGClark]

What makes this doable is the surprisingly low delta-v requirements for the closest near Earth asteroids. See this NASA page for asteroid close approaches:

http://neo.jpl.nasa.gov/cgi-bin/neo_ca

Here are the orbital specifications for the 2008HU4 asteroid I mentioned in the blog post as having especially low delta-v requirements to drag into cislunar space:

2008 HU4

http://ssd.jpl.nasa.gov/sbdb.cgi?sstr=2008 HU4;orb=1

How long would it take a chemical propulsion system, as opposed to slow SEP, to drag it to say L2?

Bob Clark

 

[RGClark]

The idea behind the asteroid-derived propellant depots was to get the large amount of propellant required in orbit at low cost. But what if we have low launch cost due to reusability? Surprisingly in that case we might be able to do a manned Mars mission for costs in the hundreds of millions of dollars range, not tens of billions, simply by launching directly from Earth.

In the post "Reusable Falcon Heavy" I estimated a reusable Falcon Heavy might be able to bring costs down to the ca. $300 per kilo range.

Now note I estimated that a Delta IV Heavy's upper stage (U/S) at a ca. 30 mT gross mass might be able to get a ca. 10 mT Bigelow Sundancer habitat to Mars in 48 days:

Math needed for 5-week flight from Earth to Mars.
http://orbiter-forum.com/showthread.php?p=512693&postcount=80

That's 40 mT for the trip to mars, and another 40 mT for the trip back to Earth. What about the landing? Take for the landing crew module 2 mT, then a ca. 20 mT gross mass Centaur will suffice to take the crew to and back from the Mars surface to Mars orbit. That's ca. 100 mT total so far for the gross mass.

However, since in this scenario we are launching from Earth orbit rather than cislunar space we need an additional 3.1 m/s delta-v to get to escape velocity. A rule of thumb is using a hydrolox stages you get the same mass to escape as the propellant size, so double the total size of the mission mass to ca. 200 mT. But at a launch cost of $300 per kilo, this will be only $60 million in launch costs!

There is though the cost of the in-space hydrolox stages. The Centaur at 20 mT gross costs $30 million. So estimate the cost of the Delta IV Heavy U/S at 50% higher gross mass at a 50% higher cost so at $45 million.

We need two 20 mT Centaurs, or a 40 mT hydrolox stage, to get the 40 mT Delta IV U/S plus habitat to escape. So this half of the trip will cost 2*30 million + 45 million = $105 million. The return flight stages will cost the same at $105 million. The landing Centaur stage costs $30 million. That's $240 million for the in-space stages. Then with the reusable Falcon Heavy costs, the total comes to $300 million.

I did not include the costs of the two habitats or the landing crew module. Considering they will be reusable the costs for all three will likely still cost in the few hundred million dollar range.

Bob Clark

 

[MaverickSawyer]

The idea behind the asteroid-derived propellant depots was to get the large amount of propellant required in orbit at low cost. But what if we have low launch cost due to reusability? Surprisingly in that case we might be able to do a manned Mars mission for costs in the hundreds of millions of dollars range, not tens of billions, simply by launching directly from Earth.

In the post "Reusable Falcon Heavy" I estimated a reusable Falcon Heavy might be able to bring costs down to the ca. $300 per kilo range.

Now note I estimated that a Delta IV Heavy's upper stage (U/S) at a ca. 30 mT gross mass might be able to get a ca. 10 mT Bigelow Sundancer habitat to Mars in 48 days:

Math needed for 5-week flight from Earth to Mars.
http://orbiter-forum.com/showthread.php?p=512693&postcount=80

That's 40 mT for the trip to mars, and another 40 mT for the trip back to Earth. What about the landing? Take for the landing crew module 2 mT, then a ca. 20 mT gross mass Centaur will suffice to take the crew to and back from the Mars surface to Mars orbit. That's ca. 100 mT total so far for the gross mass.

However, since in this scenario we are launching from Earth orbit rather than cislunar space we need an additional 3.1 m/s delta-v to get to escape velocity. A rule of thumb is using a hydrolox stages you get the same mass to escape as the propellant size, so double the total size of the mission mass to ca. 200 mT. But at a launch cost of $300 per kilo, this will be only $60 million in launch costs!

There is though the cost of the in-space hydrolox stages. The Centaur at 20 mT gross costs $30 million. So estimate the cost of the Delta IV Heavy U/S at 50% higher gross mass at a 50% higher cost so at $45 million.

We need two 20 mT Centaurs, or a 40 mT hydrolox stage, to get the 40 mT Delta IV U/S plus habitat to escape. So this half of the trip will cost 2*30 million + 45 million = $105 million. The return flight stages will cost the same at $105 million. The landing Centaur stage costs $30 million. That's $240 million for the in-space stages. Then with the reusable Falcon Heavy costs, the total comes to $300 million.

I did not include the costs of the two habitats or the landing crew module. Considering they will be reusable the costs for all three will likely still cost in the few hundred million dollar range.

Bob Clark

So, you're saying we need the upper stage of the Block I SLS for this mission?

 

[Dantassii]

Let's take this solution 1 step further. Instead of having refueling depots all the way to Mars (and back if you want to come back), how about instead having a chain of specialized vehicles each one specifically designed for 1 phase of the mission, and each vehicle is put in place and shuttles people (and cargo) along its specialized route.

An analogy would be this:

You don't bring a bus and/or train along with you every time you go on a trip from the USA to Europe to visit a few countries and then come back to the USA. No, the train and/or buses are already in Europe before you even leave the USA and are ready to be used by you and anyone else who wants to use them. You take a specially designed vehicle (airplane) to get you to Europe. You then use the trains and buses already in Europe to travel around Europe. But then you don't use those trains and buses to get back to the USA. No, you use another airplane.

So design a series of vehicles, each one efficiently designed to get its payload part of the way, to one of those propellant depots for instance. As the only thing that MUST be transported from Earth are people, you'd take a people transport from the Earth to LEO. Then you'd take a specially built vehicle designed to transport people and cargo from LEO to Lunar Orbit (say the L1 point). You can then take a space elevator from the L1 point to the lunar surface. You then take a maglev vehicle around the equator of the moon to the far side where you take another space elevator to the L2 point. Once you are there, you hop into a specially built vehicle that takes you from the L2 depot to a GMO (Geosynchronous Mars Orbit) depot. From there you take a space elevator to the Martian surface. Once there, you can take any one of a number of transport options to take you to your final Martian destination.

In each case, a specially built vehicle (be it spacecraft, space elevator, or surface transport) takes you part of the way. Each part of this trip can be built up over time. Maybe the part that gets you to the Lunar Surface gets built first, then the Lunar Maglev gets built, then the L2 depot, and so on. Along the way you learn better ways of doing things, so the final design of the Martian space elevator will be built upon the knowledge gained building the 2 Lunar space elevators.

Oh, and I don't count on building an Earth-based space elevator any time soon. LEO is so junked up that any Earth to GEO space elevator can't possibly be built until that is all cleaned up and a permanent moratorium on LEO space vehicles that leave debris is in place.

Just my $0.02 worth

Dantassii
HUMONGOUS IMS shipbuilder

 

[Col_Klonk]

My Mars plan that I started in KSP..

1) Besides Earth/Mars LEO supplies, 'deposit' about 8x fuel/supplies depot tankers at 180/190 million Km sun orbit.
2) With Earth/Mars to/fro trips these can be used to either resupply or deposit excess fuel/supplies.
3) Once Mars mining and fuel manufacturing is established.. re-supplies can be run from here to the tankers.

The idea is to run a ship with enough fuel supplies to get to mars itself and the extra gas needed for landing/taking off from mars is taken from the depot tankers.

Now the problem is asteroid fields and comets with their accompanying fragment belts that are floating around the equatorial plane, so one might have to have the tanker orbit slightly off this plane, hopefully ensuring longer survival.

Just thinking..
:hello:

 

[Lmoy]

What are the advantages of going from LEO to lunar orbit and then to Mars, rather than just from LEO to Mars?

 

[Col_Klonk]

I would imagine refueling would be it... Less Dv to exit a lunar orbit, means you can have a smaller ship, or carry a bigger payload and more fuel.

Depends on the mission objective.. I'd think.

 

[meson800]

Based on my handy subway map of the solar system hanging in front of me, isn't it harder to go to low Moon orbit to refuel then escape, instead of just escaping Earth?

From LEO to escape, it appears to be 3210 m/s dV, whereas it takes 3940 m/s dV to go from LEO to LMO. Of course some (most?) of that dV to go from LEO to LMO is effectively applied in advance for escape, but still.

(I haven't done the math or tried it in Orbiter yet)

 

[Col_Klonk]

I stand corrected, but i think it depends on how fast you want to get there - the moon that is.

From what i understand is that you can get to the moon in a few hours, or a few days.
A few hours means you're really shunting it.. and will need a lot of Dv to 'stop', as compared to coming in at a slower velocity - which will cost less Dv for orbit insertion. Also you orbit altitude determines how much Dv you use.

Should be fun doing the math optimisation.. later on that though.

 

[RGClark]

Based on my handy subway map of the solar system hanging in front of me, isn't it harder to go to low Moon orbit to refuel then escape, instead of just escaping Earth?
From LEO to escape, it appears to be 3210 m/s dV, whereas it takes 3940 m/s dV to go from LEO to LMO. Of course some (most?) of that dV to go from LEO to LMO is effectively applied in advance for escape, but still.
(I haven't done the math or tried it in Orbiter yet)

Most suggestions for propellant depots have them stationed at L2. Also, when you have large amounts of propellant at L2, either asteroid or lunar derived, then you should bring also propellant in to depots at LEO. Then assuming your rocket is space-restartable, you could have a stage arrive empty at LEO, refuel, and go to L2 or directly to Mars.


Bob Clark

 

[Dantassii]

What are the advantages of going from LEO to lunar orbit and then to Mars, rather than just from LEO to Mars?

Delta V from Lunar Orbit could be partially (fully?) done using a mass driver sort of sling shot. LEO is so cluttered up with junk and requires nearly constant delta V to keep anything there (due to atmospheric drag) that it's super expensive to keep anything permanent in LEO. Now, go to the L2 point in the Earth-Moon system, and the amount of delta V required to keep a station there is trivial.

Remember, keeping something in orbit requires a constant stream of resources from somewhere else, be it propellant, raw materials, or life support substances (food, water, air). Putting a station someplace that requires LESS support resources would be an economic thing to do. Once you get something somewhere that requires less support, then you build on that.

Dantassii
HUMONGOUS IMS shipbuilder

 

[RGClark]

On another forum someone noted that delta-v from Phobos or Deimos is lower than from the Moon and therefore it would be easier to get propellant from the moons of Mars than from the Moon. It is indeed the case that at least for Phobos its low density is believed to be due to large amounts of water/ice content.
Looking at this chart of delta-v's to the Mars system I was surprised how low was the delta-v to get from Phobos or Deimos to Earth assuming you use full aerobraking on arrival at Earth:

You see to lift-off from Deimos, exit out of Deimos orbit around Mars, and be put on a transfer trajectory towards Earth would require delta-v's of .7 + .2 +.9 km/s = 1.8 km/s. And for Phobos, it would be .5 + .3 + .2 + .9 km/s = 1.9 km/s. After this you assume you do aerobraking on return at Earth. This less than that of the Moon.

Also quite interesting is if you add up the required delta-v's to get to the moons of Mars and then return, the total is also surprisingly low, 7.4 km/s for Deimos and 7.5 km/s for Phobos. This means it would be quite easy in delta-v terms to do a sample return mission from them. It could be launched by the Falcon 9 using existing upper stages to serve as the in-space stages. On an up coming blog post I'll describe this.

About the return flight though, when delta-v's to and from Mars are quoted it's usually for a Hohmann transfer trajectory for when Mars and Earth are at closest approach. So when the sample return mission was to head back to Earth it would have to wait two years to have the low delta-v requirements of the close approach to recur. But perhaps someone on Orbiter could do the calculation for the required delta-v if you departed back to Earth soon after arrival. It would be larger, but likely still doable by a Falcon 9 launch.


Bob Clark