Post by Count Zero on Jul 26, 2005 3:29:35 GMT -4
Yeah, I was pretty ticked when that board went away. Fortunately, I saved some of the work I did for that thread. The challenge was to figure out a lunar landing mission that A.) used Apollo hardware (perhaps heavily modified), and 2.) Used Saturn IB (or a variant) rockets. Without a 2/3rds fueled S-IVB in orbit, I couldn't get an EOR mission to work. For posterity's sake, here's what I came up with (the quotes were from an earlier post of mine):
Using known delta-v requirements and spacecraft masses listed at astronautix.com, I've generated some rough numbers by working backwards through the rocket equation.
Delta-v assumptions:
TLI: 3.2km/sec
LOI: .55km/sec
Descent: 3.1km/sec (d-v designed into actual Apollo LM descent stage)
Ascent: 2.2km/sec (d-v designed into actual Apollo LM ascent stage)
TEI: .55km/sec
This first step is the most difficult part.
If the DSM is to do the LOI, the ascent stage needs to lose ~half its weight. If we cut 1000kg from its structure (out of an original 2189kg), then we can also eliminate almost 1100kg of ASM propellant, and still keep its 2.2km/sec d-v. This feeds back in a positive way: less required propellant means the tanks (and their associated structure) can be smaller.
The big problem is leaving Earth orbit. Boosting the LM/Centaur-C stack to LEO uses over 95% of the S-IVB. Assuming it can restart with nearly empty tanks,
the S-IVB/Centaur combination still falls short of the required d-v for TLI by over .5km/sec. To make up for the shortfall, we must not only strip-down the LM even more, we've got to burn the DSM, also. If we can reduce the structural mass of the descent stage by 700kg (out of an original 1984kg) without reducing the fuel load, we can achieve trans-lunar injection velocity. [what's the emoticon for Tom Kelly spinning in his grave?]
This could have worked. The S-IVB would provide ~.9km/sec, and the Centaur-C would add another 1.23km/sec. Then the fully fueled CSM would undock with the Centaur, turn around and burn roughly half of its SPS propellant to complete TLI. It would have sufficient delta-v left over for LOI, TEI and ~.2km/sec for rendezvous with the LM and any course corrections.
The only option I can see would be to send the CSM & LM to the moon separately, using Centaur upper stages to augment the S-IVBs for TLI. The spacecraft would dock in lunar orbit.
Using known delta-v requirements and spacecraft masses listed at astronautix.com, I've generated some rough numbers by working backwards through the rocket equation.
Delta-v assumptions:
TLI: 3.2km/sec
LOI: .55km/sec
Descent: 3.1km/sec (d-v designed into actual Apollo LM descent stage)
Ascent: 2.2km/sec (d-v designed into actual Apollo LM ascent stage)
TEI: .55km/sec
Launch 1: Saturn-IB-D launches LM with Centaur. S-IVB & Centaur perform TLI. LM is severely stripped-down to allow descent engine to perform LOI.
This first step is the most difficult part.
If the DSM is to do the LOI, the ascent stage needs to lose ~half its weight. If we cut 1000kg from its structure (out of an original 2189kg), then we can also eliminate almost 1100kg of ASM propellant, and still keep its 2.2km/sec d-v. This feeds back in a positive way: less required propellant means the tanks (and their associated structure) can be smaller.
The big problem is leaving Earth orbit. Boosting the LM/Centaur-C stack to LEO uses over 95% of the S-IVB. Assuming it can restart with nearly empty tanks,
the S-IVB/Centaur combination still falls short of the required d-v for TLI by over .5km/sec. To make up for the shortfall, we must not only strip-down the LM even more, we've got to burn the DSM, also. If we can reduce the structural mass of the descent stage by 700kg (out of an original 1984kg) without reducing the fuel load, we can achieve trans-lunar injection velocity. [what's the emoticon for Tom Kelly spinning in his grave?] Launch 2: Saturn-IB-D launches fueled CSM to earth orbit.
Launch 3: Saturn-IB-D launches Centaur with docking adaptor.
The CSM has 6 hours to rendezvous & dock with the S-IVB & Centaur stack before the cryogenics lose pressure. S-IVB & Centaur perform TLI (with the astronauts facing backwards). Since the CSM doesn't have the LM for LOI, we can save weight on SPS propellants. Alternatively, the SPS may use some fuel for TLI.
Launch 3: Saturn-IB-D launches Centaur with docking adaptor.
The CSM has 6 hours to rendezvous & dock with the S-IVB & Centaur stack before the cryogenics lose pressure. S-IVB & Centaur perform TLI (with the astronauts facing backwards). Since the CSM doesn't have the LM for LOI, we can save weight on SPS propellants. Alternatively, the SPS may use some fuel for TLI.
This could have worked. The S-IVB would provide ~.9km/sec, and the Centaur-C would add another 1.23km/sec. Then the fully fueled CSM would undock with the Centaur, turn around and burn roughly half of its SPS propellant to complete TLI. It would have sufficient delta-v left over for LOI, TEI and ~.2km/sec for rendezvous with the LM and any course corrections.