New Mars Forums

Josh, the alleged 1.6% gold in the LCROSS ejecta is only one of several branches this thread has split into.

I'm at a disadvantage here. I will offer links and math to back up my assertions as well as drawings to convey an idea. You as well as others just write from your gut. You proclaim the math exercises not worth your time. So you can pop out posts in 3 minutes when my posts can consume 15 minutes or 20 minutes. Time is a precious commodity for me and this lopsided exchange is annoying.

Anyway...

Saturn II C-5A

Gross mass: 384,057 kg
Dry mass: 31,740 kg
Thrust: 4446.65 kN
Specific Impulse sea level 200 s.

You can try inputting these quantities to Schilling's launch vehicle calculator

And you continue to ignore that an SSTO RLV would also have to endure re-entry. Adding TPS and other accommodations for re-entry will also harm mass fraction.

I just suggested this to someone knowledgeable in the field. His response wasn't encouraging:

"As for Earth LOS comm, that's difficult due to the 5.5 degree latitude
libration -- you need about several hundred meters of tower to clear the
limb for each degree of elevation/latitude off the pole, so a tower that
encompasses all librational conditions based exactly at the pole would be
well over 2 km high.  Possible, but not in any early stage of lunar
development.  The preferred solution for initial lunar presence is a
constellation of relay satellites."

So back to Molniya relay satellites or pehaps constellations in halos about L1 or L2.

As I mentioned to Josh, there's a trade. You can improve T/W and thus reduce gravity loss by adding more engines. But this increases your dry mass and harms mass ratio.

Yes, a good resource. I've linked to Schilling's calculator several times in this thread. As you mentioned it warns you if your T/W is less than one.

Schilling also posted some explanation of his calculator: Launch Methodology

He talks about the benefit of higher thrust during early ascent. He writes, in part "If dissimilar staging is allowed, the optimal design is generally one which uses a high-thrust first stage (often augmented by strap-on boosters) to rapidly escape the early high-loss flight regime, and a relatively low thrust but high Isp upper stage to most efficiently add delta-V once most loss terms have become irrelevant."

That is why I like two or even three stages to orbit. The lower stage can be high thrust kerosene to get out of the early high loss regime faster. Then the upper stages that don't need high T/W ratios can use better ISP stages. Besides having better ISP, the upper stage can have fewer rocket engines, thus helping with the mass ratio.

The major obstacle for reusing the upper stages is getting past the 8 km/s re-entry. As I mentioned earlier, propellant in LEO could provide a means for upper stages to shed re-entry velocity and thus avoid the temperature and stress of 8 km/s re-entry.

Another possibility is momentum exchange tethers.

The tether foot in the above illustration moves 6.9 km/s at 300 km altitude. Shaving 1 km/s not only helps with the mass fraction but lessens the re-entry abuse.

Mass ratio and T/W are a trade space.

For a given payload you can have better T/W by having more rocket engines, but this harms your mass ratio.

You can also improve T/W by using lower ISP but higher thrust engines. However lower ISP will also hurt your mass ratio.

Another trade space is comfort and spaciousness vs good mileage for motor vehicles. You can have the good mileage of a moped or the comfort of a limo but not both.

Saying a Centaur has the mass ratio to be a SSTO is like like saying a limo get 90 miles to a gallon. A limo doesn't get a moped's mileage. And with a very poor thrust to weight ratio, a Centaur's delta V budget for reaching LEO will not be 9.4 km/s.

The notion that a Centaur has the mass ratio to be an SSTO is not only arguable, it's flat out wrong.

If you are a retailer who benefits from holiday shopping, December isn't an inconsequential month. And December is only 1/12 (8.33%) of the year.

If only 1/12 of the metallic asteroids remain intact, there could be substantial ore bodies on the moon.

The small number of Apollo samples is a lot less than 9%. Yet from these vanishingly minute data points you seem convinced there are no lunar ore bodies.

No, we don't.

I don't agree 9% = virtually no impacts.

If a LEO is coplanar with the moon's orbit, A Hohmann window opens about every 90 minutes.

However a LEO coplanar with the moon's orbit will not remain so. The earth's influence will see to that.

Given a LEO orbit that is inclined wrt moon's orbit, you want to launch when the space craft crosses the line of nodes (the intersection of the low earth orbit plane with the lunar orbit plane). Else you pay a delta V penalty.

A Hohmann launch window opens at a line of nodes about every two weeks. So a launch window from a given LEO opens about every two weeks. This is true of the moon, EML1, EML1 and LLO.

You were talking about mass ratio to function as a SSTO vehicle.

If your T/W is low, you will incur excessive gravity loss and your needed delta V will be higher than 9.4 km/s.

So no, the Centaur IV does not have the mass ratio to function as a SSTO vehicle.

For reference, here is one flavor of Mars Semi-Direct: A Practical Architecture for Exploration-Focused Manned Mars Missions Using Chemical Propulsion, Solar Power Generation and In-Situ Resource Utilisation by Willson and Clarke

The author Clarke I believe is the same Jon Clarke participating in this thread.

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A few notes on the Mars Semi-Direct in Willson & Clarke's pdf:

694 tonnes to LEO (on the page labeled 202)

ERV propellant isn't deep cryogenic nor is the ERV NTR. The itemized mass list includes 44.22 tonnes of liquid oxygen and liquid methane (on the page labeled 200).

Lifting an ERV off Mars' surface and sending on a Trans Earth Injection (TEI) is very challenging. To crack enough propellant, you'd need a very robust plant and power source. ISRU to provide enough propellant to get an MAV to Mars orbit is more doable. In this regard, I regard Mars Semi-Direct more plausible than Mars Direct.

Willson & Clarke have under-estimated the delta V to get from low Mars orbit to Phobos. Thus their mission plan has insufficient propellant to make their side trip to Phobos and achieve TEI. Their plan would kill the astronauts. However if they cut out the side trip to Phobos, their propellant is sufficient.

There is some controversy on the radiation shielding need to protect against GCR (galactic cosmic rays). In my opinion Willson and Clarke's radiation shielding is insufficient.

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I haven't seen an itemized mass list of Mars Direct payloads as Willson and Clarke have done for Mars semi-direct. Given that propellant for the return trip would be from Mars, you can take 40 tonnes off the MTV as well as the ERV. But  the ERV and MTV are two separate vehicles. The MTV and hab would be the same payload in MD. I would think you'd need to land a cargo vehicle just as W & C call for in MSD. And you'd need to have Trans Mars Stages (TMS) to send the MTV, ERV and cargo vehicles on their ways.

So I believe MD would also require multiple Ares V launches (or multiple SLS launches, since Constellation bit the dust).

Valeri Polyakov demonstrated 8 months of weightlessness isn't a show stopper. So I don't regard MSD sans bolo MTV as a major flaw.

The mass ratio is only part of what I said. You need a large ratio and having a spacecraft strong and temperature resistant enough to endure re-entry.

Whether that's possible, I don't know. It has yet to be demonstrated.

Needless to say, I don't regard your math as consistently sound. I would need to see a cite and the work behind your calculations.

I been over this a 100 times with Louis. Supply and Demand. Once rare objects becoming abundant. A niche market spending millions suddenly becoming wealthy enough to spend many billions.

He is deaf to all these arguments.