Apollo 11 REDUX

Oldfart1939 · 2018-10-12 22:45:53

Those here who are familiar with my posts regarding the non cryogenic propellants may think me crazy, but known chemical hazards are faced on a daily basis in the process chemical industry. OK, NTO is highly toxic and lethal. Big deal--use proper PPE (Personal Protective Equipment) when working around these substances. My personal favorite hydrazine is Aerozine-50, which offers a high Isp and few handling problems or too high a shock sensitivity. Hydrazine, N2H4, is just too "twitchy" w/r use as a heat jacket coolant or around excessive vibrations. Hydrazine is used a lot in industry, but has disposal problems; normally sent out for incineration, but can also be decomposed to non toxic waste by use of bleach. I would strongly favor a lunar lander powered by Aerozine-50/NTO. The other option is Aerozine-50/LOX, which isn't too shabby with a good Isp. Nobody seems to be bitching about LOX being a problem.

Oldfart1939 · 2018-10-12 23:02:03

As a separate set of comments: all of the proposed propellants when dumped on a technician, astronaut, or whoever--are VERY hazardous. I had a YouTube video bookmarked for presentation whenever the subject of toxicity ever arose; prepared by NASA. Seems that YouTube has "disappeared" this video.

kbd512 · 2018-10-12 23:04:16

LOX / Aerozine-50 would yield Isp every bit as good as LOX/LCH4 and a density impulse approximately 27.5% better than LOX/LCH4.

Edit: Dumping LOX on anything is an extreme hazard, with or without a toxic chemical fuel. I don't take any particular issue with using storable chemical propellants. It's just the high dV requirements that make me think we needed to use LOX/LH2.

Last edited by kbd512 (2018-10-13 01:20:00)

spacetechsforum · 2018-10-13 01:33:58

GW, I do not agree at all. I will respond to your post by paragraph, since it mixes a lot of topics - most that do not apply to the given situation.

1. We are talking zero boil-off tank. The amount of energy removed form tank with active cooling system is equal to heat leak - pressure and temperature stays constant (and hopefully uniform). ZERO boil-off tank - no fuel loss EVER.
2. We are talking tank in space - convective does not apply. Rest addressed below.
3. This is clearly not the design for tank. OK- we have two walls and vacuum (provided by default in space), but material does not matter - the MLI blankets are placed in-between. Heat leak from support structures does not matter - in space the tank is weightless, so most of them can be disconnected after liftoff - a few can be made of material with very bad thermal conductivity and actively cooled on the hot side.
4. Does no apply. MLI in-between.
5. This is totally wrong. The surface on the sun side needs high reflectivity. On The dark side the high emissivity in IR is needed. All else surfaces need just low emissivity - there is no sunlight. Any material can be covered with proper layer to adjust for required properties, but actually only the sunny side needs it.
6. This rule applies to any tank is space - we have those already.
7. Does not apply. At all times the outer wall has equal pressure on both sides. (In atmosphere tank is empty - no insulation needed - no vacuum needed, and you vent it as you go up).
8. Material - does not apply. How big is the cryocoler? - no reference.
9. I cannot agree with this opinion - bad analysis based on wrong assumptions and no calculations.
Rest - does not apply.

I do not know how you actually imagine this tank, but clearly - a short description would be welcome, since is hard to connect the problems you specified with actual design. The designs i checked suggest construction as follows:
1 . Outer wall - aluminium with reflective layer on sunny side, emmisive layer on dark side.Layer Inside - does not care.
2. MLI blankets in vented space.
3. Very few supports made of non-conductive material.
4. Inner wall - material suitable for given substance. Emissivity - does not care.
5. Active cooling system powered with solar panels.

SpaceNut · 2018-10-13 08:51:34

Conductivity, Obsortion, Reflection, Heat sinking, Convection, Cold sinking....all terms for how we see the heat move through an object in space.....

Yes the shade side of the object is cold so making it dark allows for the cold sink effect but the tanks are insulated to space all the way around.

Yes a sun shade effect much like that was used by skylab but thats extra structure that can not deploy until in orbit... so point the reflective side of the structure is equally not a difference maker as the insulate foam would be under and only works if pointed to the sun all the time.

here are some construction images:

inner tanks formed
centaur-tank-construction-centaur-assembly-binder-7-16-59-centaur-ehp2eg.jpg

ribbing applied
centaur-manufacturing-centaur-assembly-binder-9-2-59-bldg-5-centaur-ehp2e4.jpg

top of tank foam applied
centaur-tank-construction-centaur-assembly-binder-10-7-59-centaur-ehp2eb.jpg

outer shell attached and engines built

on trailer being moved

finished foam applied

erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/1993index/IEPC1993-204.pdf
CENTAUR-DERIVED PROPELLANT SUPPLY SYSTEM FOR A SOLAR ELECTRIC ORBIT TRANSFER VEHICLE

sciences.ucf.edu/class/wp-content/uploads/sites/58/2017/02/Kutter-ACES-Space-2015.pdf
ACES Stage Concept: Higher Performance, New Capabilities, at a Lower Recurring Cost

Oldfart1939 · 2018-10-13 10:08:32

The myriad of problems associated with long term storage of cryogenic propellants--specifically LH2--requires that we look further than the attractiveness of high Isp. The density adjusted value, Id, is far more informative. LH2 isn't very dense, so requires much larger tanks for storage, and for missions of any duration, require extensive insulation (additional mass). Hydrazine/NTO combinations all present very good Id with no requirement for excessive insulation. The Russian Proton M is 100% MMH/NTO powered. Then recall, the W.W. II Me163 fighter was powered by a combination of hydrazine and hydrogen peroxide.

kbd512 · 2018-10-13 10:49:22

Oldfart1939,

NASA has already identified the best insulation materials and methods for storing LH2 in a vacuum. It's MLI with a tank design that minimizes radiative and conductive transfer using an intermediate barriers and attachment hardware with poor thermal conductivity, such as a few layers of MLI between the tank and the Sun and certain types of composites. Stainless steel is a good choice for the propellant tank, having relatively poor conductivity compared to Aluminum. The Centaur upper stages that SpaceNut has shown above in various stages of completion are pressurized stainless steel balloon tanks, for example.

The Aerogels, glass beads, etc are all less than optimal for in-space use, but work better than MLI when the LH2 is stored in a place like Earth. Convective heat transfer starts to become a real problem when a rocket is sitting on the pad.

The Me-163 is a bad example to use. Those things had a habit of exploding on landing and during fueling. On the other hand, Proton is a reliable heavy lift launch vehicle. However, Falcon 9 is more than 100t lighter than Proton (694t for Proton vs 550t for Falcon 9) and delivers an equivalent payload on account of the higher-Isp of the propellants used. If we keep the vehicle light, the trade-off in increased propellant mass is probably worth it.

GW Johnson · 2018-10-13 10:57:28

Spacetechsforum didn't like what I posted. But, look at the pictures of the Centaur stage in Spacenut's post 205 just above. That's one of the things I was talking about: a single-wall insulated tank, with a reflective outer surface.

Those flown on Centaur have no cryocooler, and the stage is only rated to store cryogens for several days in space. The trip to Mars is months long. That kind of technology will not do the job going to Mars.

What you do is greatly thicken the insulation, and install a larger cryocooler rig to get months of storage lifetime, some even claim a zero boiloff rate (although I choose to not believe it until I see it flying).

If you did the even heavier Dewar, you could use a smaller cryocooler rig; whether that's worth it, I don't know. But thicker insulation does cost enclosable propellant volume. That might make the weight worth it. But either way, I think you can safely forget about tankage that is 5% hardware and 95% propellant, and be capable of storage for months to years in space.

GW

Last edited by GW Johnson (2018-10-13 11:09:15)

Belter · 2018-10-13 11:17:08

What are you trying to insulate though? If you're talking the sun, a simple shade can do that. Vacuum insulation. If you're trying to keep it from getting so cold it actually freezes, that is a different issue, but it wouldn't be hard to add just enough heat to keep it from cooling completely.

spacetechsforum · 2018-10-13 11:53:39

Belter, update your information on thermal radiation.

GW, i think i made it very clear that I am talking about the tank designed with single purpose of long time propellant storage in space. Centaur is designed to reach orbit and do the burn and be done - the propellant does not need do be stored for very long so I think the boil-off issue is not addressed at all in this design, as long as the boil-off is low.

Still, I do not know on what assumptions you have based your opinion.

The insulation used in space is different that the one used in atmosphere. If you are filling the tank in space you can omit the insulation used on earth, so you are not thickening anything. The data i got on the subject suggest that very good insulation that is used in space weights 2 kg/m2 (only blankets, not walls - so you need to add the standard inner wall and very thin outer wall).
The bigger the tank the better ratio you have for useful propellant/heat leak since volume grows with ^3 and external surface with ^2. From NASA documents gathered i know that heat pump that can remove heat from LH2 tank has COP of 0.008 - but they do not specify mass of the installation, so i was wondering if you did the math with the machinery. Apparently, not.

Kbd512 - stainless steel is used not to insulate the tank, but to prevent corrosion issue.

Last edited by spacetechsforum (2018-10-13 11:59:48)

Oldfart1939 · 2018-10-13 12:26:20

In my mind, the only time that LH2 makes any sense is in a "load it and burn it" design. No long term storage necessary.
Introduction of additional insulation, cryocoolers, etc., defeats the KISS principle, which should underlie any engineering design.
Propellants should be selected for specific missions, not only based on theoretical performance.

kbd512 · 2018-10-13 13:12:22

spacetechsforum,

Thanks for the info. That makes sense. I thought one of the reasons stated for using stainless was that it had lower thermal transfer than Aluminum, about 8.5 times lower than Aluminum, IIRC, and it was lighter because they made it so thin that they had to use tank pressurization to maintain the structural integrity of the tank. It definitely has lower CTE than Aluminum, which might also be a factor for a tank that's pressurized and de-pressurized.

From ULA's documentation:

"The stainless steel thin-walled propellant tanks are pressure stabilized, and give Centaur an excellent energy to weight ratio. The vehicle depends on additional pressure to provide the required strength to react to bending and vibration loads on ascent and throughout the mission. The last Titan Centaur tank incorporates the improved aft bulkhead design used on the Atlas Centaur with welded gimbal mounts and pressure assisted seals. The tanks are manufactured in San Diego in the same facility as the Common Centaur tanks.

Using this lightweight design requires a fluids system that is capable of maintaining the required pressure levels for all phases of flight as well as on the ground. The fluid and propulsion systems consist of a helium supply system, a pressurization system, a vent system, a hydraulic system, two main engines, a reaction control system, and a computer control system. The long TIV missions require radiation shields on the LO2 and LH2 tank. In addition, the LH2 tank uses foam purged on the ground with helium."

Secondary source:

"Centaur boasted a unique tank design that used a common double-bulkhead to separate LOX and LH2 tanks. The two stainless steel skins were separated by a 6.4 mm layer of fiberglass honeycomb. Since LH2 is extremely cold it creates a vacuum within the fiberglass layer, giving the bulkhead low thermal conductivity that prevents heat transfer from relatively warm LOX to LH2."

Tertiary sources:

http://erps.spacegrant.org/uploads/imag … 93-204.pdf

https://ntrs.nasa.gov/archive/nasa/casi … 014252.pdf

http://www.dtic.mil/dtic/tr/fulltext/u2/429244.pdf

Oldfart1939,

LH2 tanks aren't nearly as heavy as storable propellants. I like the storable aspect of NTO/MMH, but what payload penalty are you willing to pay for that?

spacetechsforum · 2018-10-13 14:18:37

Oldfart1939, KISS solutions loose with more efficient ones in long run most of the time. My favourite example is DC electricity transfer (Edison) vs DC to AC conversion and AC transfer (Tesla). On the other hand first attempts with not KISS solutions fail miserably (like first reusable SpaceX rockets?)
Also - I agree that the LH2/LOX is not the best choice in this mission, but it makes NASA do some actually useful designs and research in the future.

Kdb512, does you post includes irony? It is hard to tell, since I am not a native speaker, but the sources provided suggest this. I was not aware that in low temperatures the stainless steel performs better than the aluminium - still the articles i found earlier suggest that the main reason for using stainless steel or other exotic materials to make long term LOX tanks is the corrosion. The information in last link point to titanium as best solution but I think that it is good that the proposed stainless steal is one of the most feasible materials.

Belter · 2018-10-13 14:21:35

spacetechsforum wrote:

Belter, update your information on thermal radiation.

Not sure what you mean. Anything not exposed to thermal radiation or other heat source will cool to 0K over time. So keeping liquid propellant cool in space is pretty easy. Much easier than on Earth.

"The long TIV missions require radiation shields on the LO2 and LH2 tank."

Last edited by Belter (2018-10-13 14:38:04)

SpaceNut · 2018-10-13 16:05:31

https://www.space.com/14719-spacekids-t … space.html

https://sciencing.com/temperatures-oute … 20254.html

If you could travel from world to world, from star to star, out into the gulfs of intergalactic space, you’d move away from the warmth of the stars into the vast and cold depths of the void. Unlike your house, car, or swimming pool, the vacuum of space has no temperature.

The three ways that heat can transfer: conduction, convection and radiation with every object sinking or sources the energy to an equilibium state.

Photons of energy get absorbed by an object, warming it up. Which is what solar panels in space measure in wattage and on orbit of earth its around 1300 W/m^2 and by time we get to Mars the value is about half of the amount. A piece of bare metal in space, under constant sunlight can get as hot as two-hundred-sixty (260) degrees Celsius. And yet, in the shade, an object will cool down to below -100 degrees Celsius. The surface temperature of Pluto can get as low as -240 Celsius, just 33 degrees above absolute zero.

For hydrogen to be in a full liquid state without evaporating at atmospheric pressure, it needs to be cooled to −252.87°C. Which even in the shade around earth will still boil....

SpaceNut · 2018-10-13 16:16:59

Something else to consider in a first stage tank full is falling foam strikes are deadly in aside by side vehicle design.
If we can make ammonia NH₃ on mars insitu we can make Hydrazine N₂H₄ ...and we can also make Dinitrogen tetroxide N₂O₄

kbd512 · 2018-10-13 20:00:15

spacetechsforum,

No irony intended.

My prior understanding, which may not be entirely correct, was that stainless balloon tanks were intended to be lightweight and easy to weld using thin stainless sheet. Back in the late 1950's when Centaur development began, thin Aluminum sheet may have been difficult to reliably weld. Obviously we have better Aluminum welding techniques today, but I assume that stainless was retained as the material of choice because of the decades of experience using it, institutional knowledge of how to fabricate it, the equipment to fabricate it, and the cost penalty associated with a major redesign of the tankage to use Aluminum. The balloon tanks are not as durable as Aluminum tanks because they're so thin that pressurization is required to prevent deformation under load. However, it's been proven to work well enough through dozens of successful launches.

A secondary benefit for brief periods of on-orbit storage, between burns, is that stainless has substantially lower thermal conductivity than Aluminum. That should mean less LH2 boil-off during those storage periods. Anyway, Lockheed-Martin (prime contractor), USAF (the initial project manager), and NASA (mostly NASA since the agency took from the Air Force over 2 years after the General Dynamics / Astronautics Corporation project which resulted in the ARPA project which ultimately resulted in LM's Centaur) did a considerable amount of iterative design and testing work to minimize transfer of heat to the LH2 because it was initially somewhat problematic.

Today, most LOX and LH2 propellant tanks used in rocketry are Aluminum alloy. The LOX tank on Falcon 9 is Aluminum-Lithium alloy and SpaceX intends to use the tanks for up to 10 missions. I would think that if corrosion was a major problem, they'd use something else. The Space Shuttle also successfully stored very large quantities of LOX and LH2 in Aluminum tanks for many decades, albeit for immediate use after loading. Those STS / SLS tanks have to survive for months after several propellant loading and unloading cycles because that's part of the design criteria. Launches get scrubbed for lots of reasons that have nothing to do with the propellant tanks.

For long term storage measured in months to years, I believe you're correct. If memory serves, all the stuff we used in the Navy to store LOX was stainless and those tanks were used for many years.

All,

We basically have two suitable TRL-9 engines for heavy landers, RL-10's and AJ-10's.

Aerojet-Rocketdyne RL-10:
dry mass: 190kg (RL-10C-1); 277kg (RL-10B-2)
propellants: LOX/LH2
thrust: 102kN (RL-10C-1); 110kN (RL-10B-2)
Isp: 450s (RL-10C-1); 462s (RL-10B-2)

Aerojet-Rocketdyne AJ-10-118K:
dry mass: 100kg
propellants: NTO/Aerozine 50
thrust: 43.7kN
Isp: 319s

Any near term lunar lander must use one of those engines. Anything else will necessarily be an engine development program, rather than a lander development program.

Oldfart1939 · 2018-10-14 11:44:18

SpaceNut-

What are actually needed for production of MMH: Methylamine CH3NH2 and Chloramine: ClNH2. Reaction:

Methylamine + ClNH2 -------------> CH3NH-NH2 + HCl.

The HCl could be reclaimed/recycled for production of more Chloramine. The component missing here, in addition to finding the source of Hydrogen, vis-à-vis water, is enough Nitrogen. The Martian atmosphere contains some Nitrogen, which could be concentrated and collected.

kbd512 · 2018-10-14 13:35:28

Oldfart1939,

How much Nitrogen would we need? If we can't get enough, does that mean we have to import it? If we have to do that, why not just import Hydrogen and use it to make LCH4? If there's not a substantial quantity of Nitrogen available, then why not use what we can extract for breathing gases? Is there any Nitrogen on the moon?

Oldfart1939 · 2018-10-14 15:29:54

kbd512-

What we don't really know is whether there are any Nitrogen bearing minerals on Mars that can be easily accessed. I know of no sources of Nitrogen on the Moon.

How much Nitrogen do we need? Let me introduce the concept of the Kilogram Mole (a Kilogram-Molecular weight). A Kilogram Mole of MMH would be 56 kg; the amount of Nitrogen required is 28 Kg, plus 12 kg Carbon, and 6 Kg Hydrogen. Mars has the Carbon (atmospheric CO2) and presumably Hydrogen (from water).

SpaceNut · 2018-10-14 18:59:21

Mars Nitrogen:
https://www.cambridge.org/core/services … 0403001599
https://www.lpi.usra.edu/publications/r … dMars2.pdf
https://www.cambridge.org/core/journals … B5D69870A6

Of course we are most likely bringing some bufffer gasses for or colonist but then again we will get some from processing our human waste as well.

http://canada.marssociety.org/winnipeg/soil.html
https://en.wikipedia.org/wiki/Martian_soil
https://en.wikipedia.org/wiki/Composition_of_Mars

I think the small missions we could achieve the amounts required, its the unknown value of concentration that is the issue for the equipment and energy to actually achieve the goal for larger vehicles.

As spacetechsforum noted the centaura while it is well designed for the limits of its use, the duration that we need is greatly longer, which is why I suggested we might need to go another route for a BFR since we need so many launches to fill just one up on orbit....

As for the engines kbd512 got the version which are being used even for the SLS rather than a J2x or some other version....

kbd512 · 2018-10-14 19:33:24

Oldfart1939,

I'm already familiar with that concept.

What I meant was, how much propellant do you intend to make and where do you think we'll get enough Nitrogen to do that?

Nitrogen is a trace gas on Mars. It exists, but there's not much of it that we know about. Maybe NTO and MMH/UDMH/Aerozine 50 could work for exploration missions, but then what about colonization?

If Nitrogen is a precious commodity required to support human life, why dump it back into the atmosphere after you've collected it?

We're pretty sure there's lots of CO2 and H2O for anyone with the right equipment to make propellant, but that kinda limits what type of fuels and oxidizers could be manufactured to LOX/LH2 or LOX/LCH4. There may be a way to make kerosene, but you still need LOX to burn it.

Oldfart1939 · 2018-10-15 07:59:34

Manufacture of any or all of the previously mentioned propellants and oxidizers would be entirely dependent upon finding some nitrogen bearing minerals. Extraction of Nitrogen from the atmosphere is a losing proposition. Otherwise we stick with methane and LOX.

GW Johnson · 2018-10-15 09:08:35

Just for the record, Aerozine 50 is listed in my ancient edition of the Pratt & Whitney vest-pocket handbook as 50% plain hydrazine, and 50% UDMH (unsymmetrical dimethyl hydrazine). The two most commonly utilized forms of any of the hydrazines are Aerozine 50 and MMH. It's a practical combination of properties that selects them.

GW

RobertDyck · 2018-10-15 10:25:30

I previously posted my idea to harvest nitrogen. It starts with a compressor from Mars ambient to 10 bar. That will require a multi-stage compressor. Do you realize how much power that will take? Why not just use methane?

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Re: Apollo 11 REDUX

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Re: Apollo 11 REDUX

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Re: Apollo 11 REDUX

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