Rethinking the orbital bus plan | Mission Planning Workgroup | Forum

Rocket-To-The-Moon said:

RP-1 has more energy per volume, but LH2 has more energy per unit mass.

I suppose this will need some careful consideration and Actual Numbers, then.  Our goal should be to reduce overall mass to make the launch vehicle's job easier.  H2 will use less mass of actual fuel but will use larger (hence heavier) tanks with insulation that will have non-zero mass.  RP-1 will use more mass of actual fuel but the tanks will be smaller and non-insulated, hence lighter.  I guess the material we make our tanks from will end up being an important factor in this choice.

See this post for some of my ideas.

Post edited 9:58 am – January 8, 2010 by Luke Maurits

Thanks for taking the time to comment in here, Gary.

Some quick numbers for everyone and then another mission plan modification idea.

Gary seems to think that with regenerative and radiative cooling we could achieve burn times for a liquid rocket of hours if we really wanted to, so there is no fundamental flaw with mission plan bravo as discussed so far.  This is good news!

The delta-v figures on Gary's website are rather different from those I have been using for all my calculations so far (where I copied the delta-vs for Apollo).

Apollo figures from Wikipedia gave me a lunar ascent delta-v of 2220 m/s and a descent delta-v of 2470 m/s, with lunar capture and escape about 1300 m/s.  TLI was about 3100 m/s.  Applying the rocket equation to our current mission plan using these figures for delta-v gives me a total mission stack mass of 8668 kg using LOX/H2 with 450s Isp.  Switching to LOX/RP-1 with 350 Isp brings everything up to 12951 kg, which is a little too heavy for a Falcon 9, but I suspect that our rough estimates for CM, extension and fuel module masses that I used for this are a little on the heavy side – it is feasible that we could bring a LOX/RP-1 mission plan down in mass to Falcon 9 levels.

Using Gary's figures for delta-v, which are lower, we get a total stack mass of 9546 kg using LOX/RP-1 and just 7149 using LOX/H2.  I don't know what the source of the discrepancy between the delta-v values is, but if Gary's figures are good then we are in for an easier trip than planned so far – RP-1 is absolutely feasible, which means reduced explosion risk during handling, which is good.  Methane would also be absolutely considerable too.

Note also that all the fuel mass calculations on the CLLARE by numbers Wiki page are a little wrong due to me using Isp values incorrectly – but not very wrong, they were always within 10% and get the gist across fine.  I will update them soon.  I may also upload a spreadsheet I made and have been using to the Google Code repository soon.

That's it for numbers, now a suggestion.  There were two inspirations for the idea I'm about to pitch.  First of all was this comment from Gary to me in an email (quoted text is mine):

> In particular we are interested in an option where *all* the
> propulsion for the mission is provided by the lander's engine -
> basically, the standard mode of propulsion for our mission would be
> the emergency mode of propulsion that Apollo 13 used to get itself on
> a free-return trajectory, with the LM engine applying a delta-v to the
> entire LM-CSM stack.  The appeal of this would be that we would only
> need to design, test and build the one engine.

I do like this, but building multiple copies of the same engine is pretty
cheap once you have a working design. you could cluster or just use the
same one in several different vehicles.

Which stands very much to reason: the big appeal of the "one engine" approach is saved design and testing time, not necessarily huge manufacturing costs.

The second source of inspiration was this diagram of a planned Gemini lunar mission (I've tried to embed the image in this post but I've noticed that astronautix.com sometimes doesn't work well for this, so if you don't see anything, the direct URL is http://astronautix.com/graphic…..mlmlan.gif and the referencing page is http://astronautix.com/craft/g…..mnilor.htm.

Basically you have the standard Gemini capsule, docked at the nose to  big booster (right of the diagram) and with a small and simple lunar lander docked in an open "shed module" attached to the back, after the equipment module (left of diagram).  The "shed module" is open at the bottom so the lander can just drop out and head for the moon.

What if we did something like this (not the nose docking booster, the "shed module" with engines on it), with the engines at the leftmost of the diagram being identical to the engine on the lander?  We could cluster 4 of the lander engines in a cross configuration at the end of the shed module.

This approach keeps the savings of design and testing time by using the one rocket in multiple places.  It costs us whatever the cost of manufaturing the individual rockets are and the extra mass of those rockets (but remember if Gary's delta-vs are good we have quite a bit of spare mass before we are using a full Faclon 9 payload), but extra advantages above and beyond the current plan are:

• A cluster of 4 engines at the back of the shed modue gives us better acceleration than just the one lander engine, reducing gravity loss.
• A cross shaped cluster of 4 engines at the back of the shed module gives us the ability to steer the stack a little bit with differential thrusting – this is good because the RCS in the nose of the CM would be pretty bad at steering the entire mass of the stack.
• We completely lose the difficult engineering problem of docking a light and open lander with the rest of the stack in such a way that thrust can be safely passed through it.
• We completely lose the difficult engineering problem of needing the lander to be able to switch between two separate fuel sources.
• The lander could be redocked into the shed after lunar ascent to make transfer back into the CM easy, but then it could be dropped again, unmanned, before lunar escape – reducing the mass we need to apply our lunar escape delta-v to, hence reducing the amount of fuel we need for that burn, which has good flow on effects – less total mass we need to do lunar capture on, so less fuel for lunar capture and hence less total mass we need to do TLI on, so less fuel for TLI, meaning less total stack mass.

I'm quite fond of this plan, it feels like it maintains most of the simplifying flavour of the lander-engine-does-it-all plan but gets rid of so many difficult problems.  What do people think?