Apollo 11 REDUX

Void · 2018-10-15 11:30:37

Well, I took the time to try to get what you wanted Nitrogen for. I may make an As* out of myself here but;

My line of work was process control, Metrology, and programming. More or less.

So, I see a process which I think might fit, and avoid wastage of energy, and yet fulfill needs.

First a containment of water. This could be variable in nature. A mechanical canister, or a ice covered reservoir. Plenty of variations imaginable. I will for this post imagine using an ice covered reservoir of relatively cold or cool water, probably more towards fresh than brine.

To have a water "Sheet" in a water column which least requires compression of the ambient Martian atmosphere, we would compress raw Martian atmosphere to just under the ice. If the ice were 3.2 feet thick, then compressed to ~34 mb+.

And then some means to produce Hydrogen to also inject. Presumably water split into Hydrogen, and Oxygen (For humans to breath, and also for a BFS device to use as Oxidizer).

There are microbes that will happily exist in water, eating Hydrogen, and breathing CO2. A fuel could be produced from this. Depending on the microbe, it could be Methane, or Alcohol, or whatever.

This could be somewhat of a batch process. You would inject large amounts of CO2, and Hydrogen, and then starve the process of them when you wanted to "Harvest" the results, which could be primarily a mixture of Methane, Nitrogen, and Argon.

To harvest, then you must degas the water with a vacuum and also compress the results. You then must do a separation which I presume would be cryogenic in nature (Most likely). But every molecule compressed and treated would be of some value you would most likely not cast anything off as waste.

Obviously the Methane would be able to feed a "Plastics" industry, and fuel something like a BFS. The Nitrogen would be what you have wanted most in this thread. The Argon would be a nice electric rocket propellant. I believe that Kdb512 had talked about that recently.

Well, that's were I would attempt to start, in hopes of getting the "Mostest for the Leasest".

And latched into that could be additional insitu microbial chemical activity. Dune materials that "Rust" in water and make Hydrogen, and perhaps clay.
Perchlorate Salts that could be Oxidizers for the Microbes.

Done.

Last edited by Void (2018-10-15 11:52:30)

GW Johnson · 2018-10-15 11:58:39

By the way, the pressurized thin-wall stainless-steel tanks on the Centaur are exactly the same technology used in the original LOX-kerosene Atlas.

300-series stainless steel are a whole lot less likely to embrittle enough to crack upon being exposed to cryogenic cold, than any other known material. Bar none. That's why it is preferred hands-down for Earthly cryogen storage tanks.

Plus, 300-series stainless steel is easily formed. Titanium is not: you would have to carve your tank from an ingot. The two beta-phase titaniums that were formable, would age at room temperature in a matter of months. The rest are non-formable alpha-phase.

I think you will find that aluminum, even aluminum-lithium, has a short fatigue life, especially when exposed to extremes of temperature. Spacex is talking about flying Falcon stages made of aluminum-lithium maybe 10 times, not hundreds or thousands of times. That's why.

GW

Belter · 2018-10-15 12:03:08

I keep going back to the idea that any equipment for processing fuel should be built into unmanned BFS units that could set down, unload mining robots and water tankers that would go out, retrieve water, bring it back and then the ship would fill its internal tanks with methane and O2. If the equipment could be made small enough, it could and should be standard issue on all of the BFS units so that they could go most anywhere in the solar system and make fuel to come back. Ceres. A trip Europa and back, just drop a straw with a heated tip. On to Titan and Saturn's rings for some methane and/or H2, and O2. Triton. Pluto. But even if not, they could be the processing and storage facility for future manned missions.

Last edited by Belter (2018-10-15 12:21:18)

Oldfart1939 · 2018-10-15 13:00:05

I originally became fascinated by the hydrazine type propellants for long term, deep space applications. The lack of any insulation/pressurization during storage coupled with the hypergolic properties make them attractive for multiyear missions.

spacetechsforum · 2018-10-15 13:33:17

GW, that was also my thinking on aluminium use (still not the main reason to discard), but data provided by kbd512 suggest otherwise:

Here is the exact link:
http://www.dtic.mil/dtic/tr/fulltext/u2/429244.pdf

If I understand correctly tested aluminium alloy in low temperature seems to be doing quite well, The titanium suggestion is also from this paper's summary.
All in all I am still unable to prove that "LH2 fuel depot" is a wrong idea. (I have no reference for active cooling system design, but this is what probably kills the project in NASA analysis).

GW Johnson · 2018-10-15 18:44:04

The link on cryo materials is to a paper from 1964. It supported in part the selection of a straight beta-phase titanium for SR-71 skins, something not in Mil Handbook 5 as a well-understood material. Those had serious cracking problems, operating at up to 750 F, far beyond what turned out to be the aging temperature, something not covered in the 1964 report.

In about the 1988-1990 time frame, we at what was once Rocketdyne Solid Rocket division quantified the aging temperature for the other (of the two) beta phase Ti alloys, as 77 F. It would suffer grain growth, strength loss, and elongation loss, just sitting in an office at 77 F. What was a rocket motor case good to about 2500 psig hydrotest, failed brittle-ly at 200 psig after 6 months room temperature aging.

That kind of thing would probably do well as a cryo tank, but only as long as you NEVER let it warm to 77 F. Bit of a practicality problem there.

Only the mixed beta-alpha phase Ti alloys have any real formability, per Mil Handbook 5, and then not much. The rest are alpha phase,
available in forged or cast massive stock, but not sheet. They cannot be rolled to sheet or drawn to wire. Simple as that. And the straight beta-phase Ti alloys are still not in Mil Handbook 5, based on the SR-71 skins experiences and the rocket motor case experience we had.

Don't ask for more detail than that, I wasn't a materials specialist. But I knew the guy who cut open the specimens from that motor case that failed. He was a real metallurgist.

GW

Last edited by GW Johnson (2018-10-15 18:46:32)

SpaceNut · 2018-10-15 20:58:07

Oldfart1939 wrote:

I originally became fascinated by the hydrazine type propellants for long term, deep space applications. The lack of any insulation/pressurization during storage coupled with the hypergolic properties make them attractive for multiyear missions.

With next up are the heavy gasses used for ION drive engines for the long term multi year missions as well.

So until the mass issues are resolved for LH2/Lox we will be stuck with pump and burn quickly so as to keep as much as possible to get the speed we need to be able to coast the distance to mars.

Matching speed with an ION drive with storables would only work to refuel the section if its a dual fuel capable engine system as well as tanks that can tolerate the different fuels before getting to Mars but this is not something needed for the moon.

For the moon we need more thermal insulation and shielding from the heat of space.

GW Johnson · 2018-10-16 14:02:35

Somebody on these forums mentioned an iodine electric thruster. That sounded really promising. Is anybody really working on one?

GW

kbd512 · 2018-10-16 14:15:59

GW,

Quaoar posted about it. I think I may have posted a few links to it, but here's a presentation from the company working on it:

https://www.youtube.com/watch?v=uKwjk6gbMf0

Edit: It's Busek Space Propulsion and Systems

http://www.busek.com

Last edited by kbd512 (2018-10-16 14:17:16)

GW Johnson · 2018-10-16 14:57:36

Thanks Kbd512:

Seems tailored to cubesats. I wonder if this could be scaled up.

His BIT-3 produced about 1.2 milli thrust at 65 W electric power, and 2800 sec Isp. Supposedly pushes a 12 kg cubesat. That would be roughly 0.01 gee acceleration capability on that cubesat. 3-12 VDC input power at 65 W sounds like modest solar.

I very much like the near-vacuum (or maybe vacuum) storage of subliming iodine fuel. Heat to 80 C and get the right flow for both the thruster and the neutralizer, if I understood him correctly, although he didn't say exactly how that worked. Really simple.

Power/thrust = 65W/1.2 mN. I have no way to judge how good or bad that is.

GW

SpaceNut · 2018-10-16 17:49:57

Here are a few more light reading items.

Iodine Propellant Space Propulsion

Iodine RF Ion Thruster Development - NASA

http://www.busek.com/index_htm_files/Io … UL2017.pdf

https://iepc2017.org/sites/default/file … say_v3.pdf

Mission and System Advantages of Iodine Hall Thruster

kbd512 · 2018-10-17 01:49:46

GW,

Aerojet-Rocketdyne / NASA's X3 ion engine produced 5.4N using 102kWe of input power. There's no reason it can't run on Argon, Xenon, Krypton, Nitrogen (it was actually tested / calibrated using N2 and then the expensive Xe), or Oxygen. I suggest we use Argon since that's available in the Martian atmosphere in significant quantities and the only other use for the gas is welding.

Tons of Argon will be collected if the atmosphere is compressed to obtain enough LOX/LCH4 to leave the surface, so why not use the stuff for in-space propulsion? So what if the input power requirement is 30% greater? It's cheap, available, and works just as well as Xenon.

X3 testing results document:

Implementation and Initial Validation of a 100-kW Class Nested-channel Hall Thruster

From the testing document:

The thruster telemetry for each test condition is presented in the Appendix in table 2. As the ionization gauge is calibrated for N2, the pressures in the table have been corrected for xenon.
As more channels are turned on, the total mass flow rate drops below the superposition value of the individual channel flow rates. When running all three channels (test conditions Xe7 and Xe14), the total mass flow rates are respectively 74% and 82% of the summed single channel values for each condition (Xe1- Xe3 and Xe8-Xe10, respectively). This is the same propellant cross-utilization phenomenon seen previously in the krypton tests. Further exploration and characterization of this significant savings in propellant is a point of particular interest for future work.

X3: 102,000W / 5,400mN = 18.88W/mN

Busek Iodine Thruster: 65W / 1.2mN = 54.16W/mN

That means roughly 2.88 times as much input power per mN of thrust produced. In other words, using Iodine technology mandates a rough tripling of the solar array area / mass and PMAD mass to handle the additional power processing capability. I don't see this as a major problem, let alone a show-stopper, using currently available electronics components.

1.1kWe/kg using Ascent Solar's thin film means 3t of sheet for 3MWe. 250We/kg using Orbital ATK's MegaFlex means 12t. The critical capability of the thin film is 270V operation. Current systems are 100V. The jump to 300V saves 30% to 40% on PMAD mass. High voltage operation is critical to mass reduction of the wiring and PMAD. Thin film saves an enormous amount of wiring cell interconnection mass and hand-assembly cost associated with MegaFlex. MegaFlex and MegaROSA still work, they're just not ideal in terms of mass and cost.

How does higher voltage operation save PMAD mass?

PMAD is basically a giant power transformer (two coils that step voltage and current up or down, as desired, as determined by the number of turns of wiring in the primary and secondary coils) that also includes power conditioning to smooth input (no damaging voltage spikes) and switching (now solid state) technology to drive various loads.

Solar Electric Propulsion (SEP) Tug Power System Considerations

BFS has lots of empty space in its rear end. Why not put something useful in the voids for in-space propulsion?

50%+ payload mass fraction in a purpose-built ITV (not BFS), but BFS could still benefit greatly from using SEP to establish a stable orbit around Mars or Earth following an injection burn using chemical propulsion, then reenter at sane and controllable orbital velocities. So what if it takes a few weeks to spiral in? Is this not about getting everyone to their destination in one piece? How is this not looking good to everyone?

elderflower · 2018-10-17 05:39:43

On Mars atmospheric resources:
On earth, Argon is a by-product of Air separation plants. It tends to come mixed with the liquid Oxygen fraction due to its very similar boiling point. It is abundant and can be discarded after use as a blanketing gas.
When we oxidise the CO and extract the CO2 from the Mars atmosphere N2 and Ar will be the main components of the waste stream. The LOX produced on Mars will be from chemical reduction of CO2 and from electrolysis of water, neither of which will include any Argon.
It is possible to separate N2 and Ar due to their different boiling points or by exploiting their different attraction to absorbent materials, but if we want to use them in an ion engine, maybe we don't need to separate them. Their ionisation energies are not a lot different. It may be that designing for one or the other would be more efficient, but for simplicity of production on Mars the reduction may be OK.

SpaceNut · 2018-11-29 21:31:39

We now have a target year in 2028 moon mission pitched at US National Space Council meeting

Oldfart1939 · 2018-11-30 12:42:38

Hrumpf! Subject to standard change of mission with next administration--no doubt. Give the money to private contractors; somehow, SpaceX comes to mind.

kbd512 · 2018-11-30 13:47:10

We can do lunar missions with Dragon 2 or Orion and Falcon Heavy with IVF-enabled LOX/LH2 upper stages. We don't have to wait for giant rockets to start building the hardware to take us back to the moon. It's just not necessary. If the goal is indeed to go back to the moon, rather than just spend money and spin our wheels, then apart from the lander we already have the hardware required to do it. Incidentally, a LOX/LH2 upper stage can also be the lander if we decide to use it that way. That's exceptionally minimal technology development to achieve something we haven't had the capability to do for almost half a century.

RobertDyck · 2018-11-30 18:31:21

There's the Moon idea I pitched. Two launches of Falcon Heavy, one for the LM, one for Dragon. Reusable LM so subsequent missions just need one launch.

SpaceNut · 2018-11-30 19:26:38

Well first up is getting partners with pockets with cash and hardware to put towards the project. Construction of Russian Lunar Orbital Station May Be Launched in 2025

kbd512 · 2018-11-30 21:10:12

Robert,

Sounds great, then. Let's flesh out your idea and start doing some rudimentary math using what we already have. We're using an all TRL-8 or TRL-9 technology architecture. Falcon 9 Heavy is the baseline. If Vulcan Heavy and New Glenn become operational in the next 2 to 3 years, then we'll consider those rockets after they've successfully flown at least 3 times. We're not even going to consider PowerPoint-based super heavy lift rockets, such as SLS or BFR. Dragon 2, Starliner, and Orion are all TRL-8, but they'll be operational next year. We have AJ10-118K's, Merlin Vac's, and RL-10C's for upper stage and lander propulsion. The cryogenic propellant combinations are limited operational lifetime, meaning a month or so at most. For habitation, we have ISS node modules and BEAM for habitation. ISRU/ISPP will be included in the mission architecture the moment we prove the technology to extract ice deposits from the lunar poles, but not a moment before.

1. crew per mission?

2. surface stay duration?

3. payload capability?

4. missions per year?

5. architecture evolution schedule and plans (inclusion of ISRU/ISPP, testing of new space suits, radiation shielding, etc)?

RobertDyck · 2018-12-01 04:26:04

kbd512,

Ok. I posted about it a few times. Here's my idea.

4 crew per mission
surface stay, hours. Give the LM enough air for landing, take-off, and at least 2 depressurize/repressurize cycles. Possibly more. Spacesuits with 8 hours surface time, without LM capacity to refill suit O2.
payload capacity: absolutely no payload outside the pressure hull. Some capacity for science instruments inside the pressure hull with astronauts. How much? Also payload capacity for samples during ascent. Again, how much?
missions per year? Multiple
architecture evolution plans: this is where it gets interesting. I propose a Mars Direct habitat sent to the Moon. I get this from Robert Zubrin's plan. He proposed a Mars Direct ERV launch directly from the surface of the Moon, with the ERV landed fully fuelled. I point out that's NASA's moon plan prior to 1961, known as Direct Launch. It required a launch vehicle even larger than Saturn V. We only have SLS block 2B, so that won't work. I propose instead using an Apollo architecture: Lunar Orbit Rendezvous. That means LM and a Dragon capsule. However, the Moon base will be a Mars Direct hab, complete with rover and inflatable greenhouse. Radiation shielding: sand bags filled with Lunar regolith piled on the roof. Again, from Mars Direct.

More details: Falcon Heavy will have to be man-rated, but no modification. My calculations showed the side boosters could be recovered on ships (barges), but the central core booster and upper stage would have to be expended. Falcon Heavy upper stage would be existing Falcon 9 upper stage without modification. It would be the TLI stage. RP-1/LOX propellant.

LM: pressure hull built with isogrid aluminum alloy. Apollo LM had a pressure hull made of aluminum alloy foil so thin it was 5 times the thickness of kitchen aluminum foil. Apollo used a grid sandwich, like corrugated cardboard but made of aluminum alloy. This was very light-weight but only able to withstand 5 pressure cycles. I propose a reusable LM, so the hull will have to be more durable. I propose an isogrid, which is the current pressure hull used by Dragon. The LM will have a different shape, but built with the same technology. This is the same technology currently used by Dragon, Orion, and Boeing's CST-100 Starliner.

LM would not have 2 stages like the Apollo LM, instead designed to use a crasher stage for de-orbit. LM would have one engine (or one set of engines) only. One set of propellant tanks. Everything that lands, would take off. Leave nothing behind, other than stuff carried inside.

This would require a new expendable upper stage. On top of Falcon Heavy upper stage. The new stage would use carbon fibre composite propellant tanks, carbon fibre composite structure. LOX/LCH4 propellant. LM would also use LOX/LCH4 propellant. New stage would have 3 jobs: Lunar Orbit Insertion (LOI) with Dragon attached, propellant transfer to LM in Lunar Orbit, de-orbit with LM attached. Crasher stage would also carry breathing oxygen for life support, replenish O2 for LM life support.

Dragon: convert trunk into a service module. Start with trunk of Dragon v1, with flat solar panels that track the Sun. Add carbon fibre composite propellant tanks filled with LCH4/LOX. This service module will be used for Trans-Earth Injection (TEI).

Note: BFS is intended to use carbon fibre composite tanks, LCH4/LOX, and on-orbit propellant transfer. This Lunar architecture will test/demonstrate technologies needed for BFS, testing on a craft a lot smaller.

SpaceNut · 2018-12-01 09:32:20

Have reposted conversation of kbd512 and Robert Dyck on use of Falcon 9 Heavy for lunar mission as they relate to the cover page article by Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years thou this topic is short of the base part of discussion.

As for the Dragon version 1 its got now human internal life support so we will need to go with Version 2 or crewed which is coming soon.

Moon Landers are a new design that must fit the payload restriction of the launch vehicle.

Launch vehicle Falcon 9 Heavy we are assuming will be man rated but if its not then we need a plan B.

RobertDyck · 2018-12-01 14:38:39

The point with Dragon version 1 was the trunk. Not the capsule, the trunk. Yes, Dragon version 2 has life support and launch escape, while version 1 does not. Fine, use the v2 capsule. My comment was to use the v1 trunk as basis to make a service module.

Dragon version 1:

Dragon version 2, used for pad abort test, no solar panels:

Notice the v2 trunk has fins, and curved solar panels that don't track the Sun. Curved cells are more expensive to make. And when solar cells are applied to a curved surface, only those facing the Sun will generate power. This requires several times the number of solar cells to produce the same power. Basically pi (3.14159) times as many solar panels, if the trunk is perpendicular to the Sun. If the axis of the vehicle is not perpendicular, not only are cells around the curve not oriented correctly, but the second dimension isn't oriented either. So basically you need 4 times as many solar cells. This increases mass and cost. V2 trunk doesn't require fairings, but solar cells are far more expensive than fairings.

Once the trunk has propellant tanks and engines required for TEI, the mass will be so great that the service module cannot be dragged along during abort. So abort has to be changed anyway, may as well reduce mass and cost at the same time. Abort control system will have to be changed to permit the capsule to abort without the finned trunk. Once done, we can base the service module on the v1 trunk.

SpaceNut · 2018-12-01 21:21:55

So we do not even have a ready to use means for moon missions without changes to what the crew would use as well.
Even the mini bfr would still not be what we would need.

Then again starliner

starliner_hero_new_1280x720.0.jpg

moon missions dreamchaser is a taxi to orbit.

Cygnus sure in the various ISS configurations would be useful but even these need modifications.

So realistically all we have is 63 T to orbit with Falcon 9 heavy

RobertDyck · 2018-12-01 21:38:41

Starliner needs a larger service module, for TEI. And since its SM is its abort system, you have to ensure engines are powerful enough to quickly lift the capsule & SM in Earth gravity.

Orion can go to the Moon without modification, but it's so heavy it requires SLS. Even SLS block 1 isn't enough to launch it into lunar orbit, without a LM. That takes block 1B.

Required modifications for Dragon are not show-stoppers.

Last edited by RobertDyck (2018-12-01 21:39:57)

SpaceNut · 2018-12-01 23:14:46

https://en.wikipedia.org/wiki/Falcon_9_ … st#Block_5

How much fuel is remaining in the second stage of the Falcon 9 Heavy launch for a Dragon of either version?

Is that enough to send a capsule to the moon?

https://en.wikipedia.org/wiki/SpaceX_Dragon

Falcon 9 block 5 with crewed flight capsule

https://en.wikipedia.org/wiki/Dragon_2

https://en.wikipedia.org/wiki/Cargo_spacecraft
https://en.wikipedia.org/wiki/Cygnus_(spacecraft)
https://en.wikipedia.org/wiki/H-II_Transfer_Vehicle

Modules
https://en.wikipedia.org/wiki/Internati … ce_Station

So we were able to get to orbit now we need to send the assembly on its way or even by seperate pieces which seems like a waste but that may not prove out.

https://en.wikipedia.org/wiki/Parking_orbit

https://en.wikipedia.org/wiki/Free-return_trajectory
https://en.wikipedia.org/wiki/Circumlunar_trajectory

https://en.wikipedia.org/wiki/Trans-lunar_injection

https://en.wikipedia.org/wiki/Orbital_maneuver

We have sent a number of probes but only a few manned landings to the moon and we will be reinventing the wheel once more.
Then we have to slow into orbit and then begin to land the stuff to the surface.

https://en.wikipedia.org/wiki/Moon_landing

Last mission was https://en.wikipedia.org/wiki/Apollo_17 for just 3 days time on the surface

In a 2 stage none reuseable lander
https://en.wikipedia.org/wiki/Apollo_Lunar_Module

We will want to do better with this aspect of stay, reusability and mass to the surface.

Just glad that all worked correctly as the chance to get back to orbit if it failed was grim at best.
https://en.wikipedia.org/wiki/Lunar_Escape_Systems

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Re: Apollo 11 REDUX

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