For serious interplanetary operations we need fusion propulsion – plain nukes aren’t much better than chemical rockets performance wise. Outer Planet access with trip times under a year are probably vital on biomedical grounds due to the nastiness of high-energy Cosmic-rays. Thus the necessity of fusion propulsion.

But before we shoot off to Jupiter, what can we do about Mars and a little bit beyond?

Assume three FH Tankers (52 tonnes fuel, 3 tonnes dry-mass) and a payload massing 55 tonnes. Arrange two Tankers as First Stage and one as the Second Stage to push the payload. What delta-vee do we get? Over-all mass ratio is (220/(220-104))*(110/(110-52)) = 3.6, thus with the Merlin Vacuum engine we get 1.28 x 342s x 9.80665 = 4,293 m/s – enough to put our cargo on a Hohmann transfer to Mars, with a bit of a reserve.

For unmanned vehicles carrying cargo the 258 day Hohmann orbit is preferrable, but punitive for a manned mission. With a bit of extra delta-vee – such as the above figure – a manned mission can save on supplies and cosmic-ray exposure. Gerald Nordley discusses the issue in his on-line essay…

…indicating trip-times of 130-180 days are reasonably feasible. Thus crew can travel quicker than freight. The canonical Mars Semi-Direct would require delivery to Mars of a Habitat, and Earth Return Vehicle and a Mars Ascent Vehicle, all in the roughly 55-60 tonne mass range. Thus a total of 12 Falcon Heavy launches to deliver a crew of six to Mars. A launch cost of just $1.5 billion for a Mars mission is a *dream*! But eminently practical with Falcon Heavies available.

Going to Mars lets us save propellant via aerobraking – aerocapture into a highly elliptical Mars orbit – which isn’t available if we go beyond Mars to the Asteroid Belt. Trip-times rapidly go up as we move further away from the Sun, especially for tricky fuel-saving orbits with higher aphelia than the destination. Another speed-bump is the non-zero inclinations of the asteroids, which makes them even trickier to reach.

So what do we do? Personally I think this is where we have to start getting out of the rocket straight-jacket and start getting serious about solar-sails – as recently successfully demonstrated by IKAROS and Nanosail-D. There’s a certain elegance – and zero-fuel budget – which has an immense appeal.

After the third wine now so please excuse the slurred speech.

The more I think about it, the more I’m in favour of solar thermal, if Isp of 900s is achievable with H2 propellant, transit times of less than 100 days using aerobraking at Mars should be achievable with quite reasonable propellant mass fractions. Then if we’ve got the H2 + O2 to get back up to orbit already produced on Mars by an unmanned nuclear fission powered probe using on planet indigenous resources, the total Earth departure mass could be impressively modest.

Aerobraking counts for a lot if we want shorter transit times, so Nautilus-x just went down in my stock (at least for Mars missions).

I don’t see solar sails as being as effective as you, here’s my maths:

At 1 AU from the sun light pressure is about 10N / km^2 on a mirrored surface. If we have a sail with a mass of 3g/m^2 (current best) we get a maximum acceleration of 3mm/s^2, ie. three ten thousandth of a g.

If we attach a ship that has about twice the mass of the sail we’re down to about one ten thousandths of a g acceleration, at that rate of acceleration we get to 3.46 km/s after 40 days, and that’s with the sail perpendicular to the sun.

For a trip as challenging as an expedition to Mars, I’d go for a larger crew and more than one ship.

Here’s my architecture:

8 large aerobraking capable capsules; 3 of them manned (each capable of supporting 6 people under normal operation), 4 of them fuelers, 1 a nuclear reactor.

Fuelers each:

Dry mass 20t

H2 capacity 80t (1200m^3)

O2 capacity 120t (100m^3)

Manned:

solar concentrator 10t

H2 20t (300m^3)

O2 120t (1200m^3)

expendables 12t

dry mass 40t

Reactor:

62t 10mW electrical generation with H2O electrolysis converter.

20t H2

OMS system included

They’re mated in pairs when traveling interplanetary, one fueler plus a manned unit or the reactor. Each mated pair departs Earth using solar thermal rockets Isp900s, total paired mass 82 tonnes dry, 10t O2, 100t H2; delta V over 6 km/sec. Solar concentrator stored after acceleration period.

During transit pairs can be tethered to provide 0.38g gravity.

At arrival at Mars each capsules aerobrakes into initial parking orbit, then they land at the same location where water has been previously determined to be extractable.

The reactor, once set up starts filling the tanks of the fuelers and manned units. The length of stay on Mars is to be 150 days, 380 tonnes of H2 and 840 tonnes of O2 needs to be produced in that time which requires continuous 4.2 mW of power other that period plus loses.

At departure the 3 manned units and the 4 fuelers launch to orbit using H2 + O2 rockets, burning 20t H2 and 120t O2 with total liftoff masses around 220t they have delta v’s at 450s Isp of about 4.5 km/s, easily enough to achieve orbit.

The fuelers are each able to lift 60 tonnes of H2 to orbit, thats 80 tonnes to power the solar thermal rockets for each of the 3 paired units heading home.

Oh, the total Earth departure mass is 768 tonnes, which should be able to be lifted in 16 Falcon H flights, the physical dimensions of the units could require some thought to minimize drag at launch while still being a suitable shape as re-entry capsules.

Ahh, in my 5:40 comment a small error “O2 120t (1200m^3)” under “manned” should be “O2 120t (100m^3)”

Note that these figures refer only to capacity, not Earth departure mass.