Starliner beyond LEO

[ncb1397]

To me it seems by far the simplest solution is just to build a bigger service module.  Sure it would be heavier due to the extra fuel, but you would also have more propellant available for launch abort, so it would seem workable just by adding additional abort engines.

The rocket equation makes things difficult, because delta-v is proportional to the logarithm of the wet:dry mass ratio. According to otlski, the starliner masses are:

SM dry mass = 10000 lbs.
SM wet mass = 15000 lbs
CM dry mass = 14000 lbs
CM wet mass = 18000 lbs

That means the wet mass is 16.5 tons and the dry mass (with the CM still 18000 lbs) is 14 tons for a ratio of 1.14. The CST-100 SM provides 1200 m/s delta-v. Orion's SM provides 1500 m/s, or 1.25x more delta-v. So your ratio needs to go up by 1.18. This is a minimum of about 1 ton more propellent (not including mass of stretched SM, which is significant). For a spacecraft with 2.5 tons of propellent, stretching the SM propellent tanks by 40% is nontrival.

How do you get 1200 m/s? I get about 500 m/s assuming a 300 second specific impulse using the numbers you posted. To get 1200, you would need a specific impulse of around 750 (so, nuclear thermal propulsion or better).

[Jimmy_C]

To me it seems by far the simplest solution is just to build a bigger service module.  Sure it would be heavier due to the extra fuel, but you would also have more propellant available for launch abort, so it would seem workable just by adding additional abort engines.

The rocket equation makes things difficult, because delta-v is proportional to the logarithm of the wet:dry mass ratio. According to otlski, the starliner masses are:

SM dry mass = 10000 lbs.
SM wet mass = 15000 lbs
CM dry mass = 14000 lbs
CM wet mass = 18000 lbs

That means the wet mass is 16.5 tons and the dry mass (with the CM still 18000 lbs) is 14 tons for a ratio of 1.14. The CST-100 SM provides 1200 m/s delta-v. Orion's SM provides 1500 m/s, or 1.25x more delta-v. So your ratio needs to go up by 1.18. This is a minimum of about 1 ton more propellent (not including mass of stretched SM, which is significant). For a spacecraft with 2.5 tons of propellent, stretching the SM propellent tanks by 40% is nontrival.

How do you get 1200 m/s? I get about 500 m/s assuming a 300 second specific impulse using the numbers you posted. To get 1200, you would need a specific impulse of around 750 (so, nuclear thermal propulsion or better).

I probably misread Steve Pietrobon's post. My apologies, you are right. In that case, the stretch is a lot more!

https://forum.nasaspaceflight.com/index.php?topic=52056.msg2140370#msg2140370

[Pueo]

To me it seems by far the simplest solution is just to build a bigger service module.  Sure it would be heavier due to the extra fuel, but you would also have more propellant available for launch abort, so it would seem workable just by adding additional abort engines.

The rocket equation makes things difficult, because delta-v is proportional to the logarithm of the wet:dry mass ratio. According to otlski, the starliner masses are:

SM dry mass = 10000 lbs.
SM wet mass = 15000 lbs
CM dry mass = 14000 lbs
CM wet mass = 18000 lbs

That means the wet mass is 16.5 tons and the dry mass (with the CM still 18000 lbs) is 14 tons for a ratio of 1.14. The CST-100 SM provides 1200 m/s delta-v. Orion's SM provides 1500 m/s, or 1.25x more delta-v. So your ratio needs to go up by 1.18. This is a minimum of about 1 ton more propellent (not including mass of stretched SM, which is significant). For a spacecraft with 2.5 tons of propellent, stretching the SM propellent tanks by 40% is nontrival.

The Atlas V 551 config can put 22.6 US short tons in LEO (185 km x 28.5 deg).
So an Atlas V 551 could put up a Crew Module and Service Module and have 6.1 tons spare.  Using that mass as kick stage with 19% structural mass fraction and assuming 300 s ISP for both kick stage and service module we get:

725 m/s = 300 *9.8 *ln (22.6 / (16.5 + 0.19 * 6.1))

Then we stage and fire the service module using the LAS fuel instead of wasting it:

483 m/s = 300 * 9.8 * ln (16.5 / 14)

This gives us a total of 1208 m/s, equivalent to the ESM.  We don't have to worry about propellant line separation during abort, only additional life-support line separation.

Upsides: 
No need to re-develop the launch abort system.
Launcher agnostic: Can work with both Atlas V 551 and an expendable Falcon 9.

Downsides:
High G loading during ascent.
You have to cram all the extra life support equipment into the 1.159 tons of dry mass in the kick stage.
Your mission needs to only require the firing of the service module at a point where the SM and CM carry enough life support to get you home.
To reach orbit the Atlas V 551 needs to fly its more efficient lofted trajectory rather than the flat OFT trajectory.  This potentially creates black zones during the first two stages of flight where failure results in loss of crew.
Oh, and you still need to dock to an ICPS in orbit or you aren't getting to the Moon anyways.

[ranma]

Maybe this should wait until the "Starliner" has reached the International Space Station with a crew? I keep hoping for a win for Boeing but it does not seem to be possible.