Definitive new fuel mass estimates | Mission Planning Workgroup | Forum

Post edited 7:10 am – January 15, 2010 by Luke Maurits

Okay, so I've written a small program (will push it to my Hg clone later tonight) that makes it a piece of cake to compute total mass figures using the rocket equation.

The graph below shows total launch mass (y-axis) vs fuel Isp (x-axis) for 5 different stack masses: super light (750 kg CM, other structures similarly unrealistically light) all the way up to super heavy (1750 kg CM, other structures similarly unrealistically heavy).  The medium option is quite sensible but still I think still a little conservative: 1250 kg CM, 400 kg lander, 500 kg PM, etc.  We should expect to manage medium, the other options are just there to show how much spare mass we have before things stop being feasible and to show us how much mass we'd need to shed to be able to user lower Isp fuels than LOX/LH2.

The delta-v budget here is 3100 m/s for TLI, 1000 m/s for LOI, 3400 m/s for the lander and 700 m/s for TEI, with a 10% safety margin on top of each burn's budget.

The horizontal brown line shows the Falcon 9 maximum LEO payload.  For each of the stack mass options, if you find the point where that stack's line intercepts the brown horizontal and then mentally draw a vertical line down to the x-axis, it shows you the minimum Isp required to get that stack to fly in a Falcon 9.

The heavy and super heavy options will never fly on a Falcon 9, so we absolutely must beat those estimates.

The medium option, which is a sensible, slightly conservative estimate of masses, requires an Isp of about 410s.  LOX/LH2 can easily manage this.  This means that the overall mission is absolutely feasible using a Falcon 9, which is good news.

The light option, which would require hard work to keep mass down but which is not outside the realm of possibility, requires an Isp of just less than 350s, which means we could use LOX/LCH4 (LCH4 is safer, easier to store and more dense than LH2).  We should strive for this due to the advantages of LCH4, but probably not expect to make it.

Note that 350s seems to be about the average Isp of all of the LOX/LCH4 engines on this page, but some individual engines go as high as 380.  If we can get decent LOX/LCH4 engines with these higher Isps, we wouldn't need to do quite as good a job as the light option requires.

The super light option requires 300s, which would make LOX/RP-1 feasible, but I would be amazed if we could get things that light without resorting to super expensive/exotic materials everywhere possible.

Summary: CLLARE on a Falcon 9 is definitely possible using realistic mass estimates if we can get working LOX/LH2 engines.  It may well be possible to use LOX/LCH4 engines if we do a really excellent job of keeping mass down everywhere.

I'll put this graph or one very like it in the next version of the CLLARE Project Overview Document.

I have uploaded a new version of the plot which (i) includes an OSM mass (ranging from 100 kg in super light to 300 kg in super heavy) and (ii) includes a Falcon 9 Heavy line, like Rocket's modification of the original.

Since I uploaded it over the old version rather than as a new file on the Wiki, the image at the top of this thread should now be the correct one.  Just note that the commentary below it in terms of Isp requirements is no longer quite right.

The changes are relatively minor and this entire prospect is still feasible.