Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

kbd512 · 2017-03-06 00:19:33

SpaceNut,

The article says they're using MR-80B (3100N) and MR-107U (275N) thrusters. MR-80C (3100N) and MR-104D (440N) are thrusters I made up. They're notional variants of the MR-80B (3100N) and MR-104C (440N), both of which are real flight hardware, that have been modified to use materials compatible with the slightly acidic HAN (AF-M315E) monopropellant.

* HAN monopropellant thrusters do not use double valves as do thrusters using N2H4 / hydrazine (MR-80B and MR-104C are examples thereof) use. HAN requires high temperature catalyst beds (200C to 300C) compared to hydrazine (100C to 150C), and thus more electrical power, to decompose the propellant. Hydrazine will start to decompose purely as a function of contacting the catalyst bed, although Isp will be lower without pre-heating. The switch to single valves for HAN-fueled thrusters is principally a characteristic of the high heat requirement to facilitate propellant decomposition. Simply put, propellant leaking through the feed valve won't fire the thruster without a hot catalyst bed. HAN is also less likely to decompose in hot propellant feed lines connected directly to the thruster, compared to hydrazine. This simplifies the propellant feed system.

* HAN does not require continuous heating of the propellant tanks and feed lines. It has a glass transition temperature, but doesn't freeze like hydrazine will if it gets too cold during the deep space transit. Basically, you heat the tanks and feed lines when you intend to use the thrusters.

* AF-M315E (a trade name for a specific HydroxlAmmonium Nitrate or HAN blend / formulation developed by Air Force Research Laboratory or AFRL and AeroJet-Rocketdyne) monopropellant is comparatively non-toxic to humans, in relation to hydrazine. That is another reason for the simplified propellant feed system. Spills or leaks are not events that pose immediate danger to human health during propellant handling. For comparison, a gasoline spill is more hazardous to your health and a hydrazine spill would seriously injure or kill you without protection.

kbd512 · 2017-03-06 02:55:33

Further notes:

16. The Cygnus PCM has 4 variable length (extendible) landing legs attached to steel wire wheels that can caster. Each wheel hub has a small electric motor so the PCM can move at low speeds. This is how the Cygnus PCM is mated to the Earth return stage. The PCM drives over to the Earth return stage after landing, extends the legs to raise the PCM above the upper stage, and then rotates about its axis to lock the PCM in place.
17. The empty RCS module stays attached to Cygnus during ascent because it is later required during Earth return to stabilize the Cygnus PCM during Earth reentry. After the Cygnus PCM ascends to Mars orbit using the ascent stage, they have a series of three spacewalks to perform to prep their Cygnus PCM for Earth return.
Space Walk I - Transfer food and water from the Earth Return Stage to the Cygnus PCM
Space Walk II - Refill the empty HAN and GN2 tanks using tanks integrated into the Earth Return Stage
Space Walk III - Attach a new HIAD to the Cygnus PCM for Earth reentry
18. The Earth reentry HIAD is specifically designed for that purpose. All HIAD's have to be constructed specifically for their intended payload mass. It's the same basic technology set and materials, but different donut sizes, material thickness, or material layers. An Earth reentry HIAD may have five layers of thermally protective fabric whereas a Mars reentry HIAD may only have three layers.
19. If SLS and SEP can deliver Lockheed-Martin's Mars orbital station, better known as Mars Base Camp, to LMO, then the spacewalks are performed while the crew stay aboard the space station in Mars orbit. There will be 2 to 4 crews (4 to 8 astronauts) loading supplies at the same time, so an ISS node module with 4 docking ports and an ISS MPLM module to store consumables is the bare minimum required to service 4 Cygnus PCM's. An ISS node module and ISS Destiny (lab habitat) module is more realistic.
20. 1g artificial gravity is provided on the departure flight by tethering off to the expended upper stage. Artificial gravity may or may not be practical for Earth return. The PCM really needs to stay solidly attached to the Earth return stage because that stage is responsible for slowing the PCM just prior to reentry. If that fails, you get BBQ'd during reentry and that's not a good way to end an otherwise successful mission.
21. If NASA is willing to give up 30 to 60 days of surface time by spiraling into LMO using SEP and/or spiraling into LEO using SEP, then reentry velocities drop substantially. If that Mars orbital station ever materializes, it's a good place to inspect PCM's prior to committing to a reentry. If I was an astronaut, I'd want to determine whether or not my PCM, RCS, and HIAD modules were in good working order after my trip through interplanetary space.
22. Using SEP to spiral into both LMO and LEO, I eliminate the requirement to attach a second HIAD to the Cygnus PCM. The RCS tank still requires a refill. The food and water still require replenishment prior to Earth return. Upon Earth return, the PCM's dock at ISS for samples offload, crew de-briefing, and return to their respective countries along with astronauts from ISS. The Americans and Canadians return to America aboard Dragon V2, the Asians return to China aboard Shenzhou (assuming they participate; they seem to be more interested in the moon), and the Europeans and Russians return to Kazakhstan aboard PTK NP. There is no need to store NTO/MMH in the PCM service module or the Earth Return Stage, either. HAN and SEP are sufficient for all mission requirements except the Mars ascent. NTO/MMH, as dangerous as they are, are required by the Mars Ascent Stage to achieve sufficient Isp so that a single Falcon Heavy can deliver the stage to Mars.

GW Johnson · 2017-03-06 12:15:12

I wish NASA would get on with the war testing these extendible / inflatable heat shield things. The sooner these are flying routinely back from orbit, the sooner we can trust them to land on Mars. What I see is a lot of papers and meeting presentations, and precious few exploratory (!) experiments.

GW

Oldfart1939 · 2017-03-06 12:45:57

GW-

You are absolutely correct: Theory Guides, but Experiment Decides...a truism for a long time.

GW Johnson · 2017-03-06 17:02:22

Same applies to the new less-toxic propellants. I sure do wish somebody was trying to fly these things.

GW

SpaceNut · 2017-03-06 19:01:47

Here is some data for the HIAD testing....

NASA to Hold Briefing to Discuss Status of 'Flying Saucer' Test

https://www.nasa.gov/ldsd/

https://www.nasa.gov/langley/nasa-langl … craft-test

https://www.nasa.gov/pdf/737628main_Fin … -26-13.pdf

https://www.nasa.gov/sites/default/file … sd2015.pdf

Comparison of Analysis with Test for Static Loading of Two Hypersonic Inflatable Aerodynamic Decelerator Concepts

kbd512 · 2017-03-07 09:29:40

The AF-M315E demonstrator is scheduled to fly this year. HEART could fly next year if NASA devoted serious funding to the project.

I don't see us going to Mars for any reasonable cost without these new technologies:

Radiation Shielding
BNNT (boron nitride nanotube)

NASA | Radiation Shielding Materials Containing Hydrogen, Boron, and Nitrogen

638828main_Thibeault_Presentation.pdf

716082main_Thibeault_2011_PhI_Radiation_Protection.pdf

Boron Nitride Nanotube: Synthesis and Applications

Structural Boron Nitride Nanotube Composite Development

Simulation of Hydrogenated Boron Nitride Nanotube’s Mechanical Properties for Radiation Shielding Applications

Life Support
CAMRAS (amine bed atmospheric scrubber) - far lighter, more compact, and less electrical power than current ISS atmospheric scrubbers
MOXIE (O2 from Martian CO2) - sucks CO2 out of the Martian atmosphere and turns it into O2
IWP (ionomer membrane waste water processor) - recovers 90%+, as high as 98%, of water from waste water and turns it into potable water with very little electrical power

We go from multiple ISS racks of equipment to a half rack of equipment with these technologies. The IWP uses lightweight replaceable plastic bags that can be discarded after several uses. Combined with improved waste heat management, total power consumption for life support should be less than 1kWe for 2 astronauts.

Expandable Heat Shield
PICA (any payload mass range) - phenolic impregnated carbon ablator for reentry heat shields
HIAD (10t-20t) - stacked donuts nitrogen inflatable application of commercial fabrics for reentry heat shields
ADEPT (10t-75t) - umbrella expandable application of commercial fabrics for reentry heat shields

Propulsion
SEP (solar electric propulsion) - required to perform high-dV maneuvers when mass is at a premium
HAN (AF-M315E monopropellant) - hydrazine replacement that can't freeze, readily decompose without high heat, or poison you
Composite Tanks - we now have robots that wrap composite tanks that are 30% to 40% lighter than traditional aluminum can cost 30% to 30% less than traditional aluminum alloy tanks since the production process is faster, requires fewer personnel, and no autoclaving is required

Oldfart1939 · 2017-03-19 16:43:55

We've now had another "Lunar Distraction" from the far more important goal of Mars. I really believe that putting together a complete inventory of available space vehicles, not only the ones made here in the USA, would allow us to conceptually select some of the other modules for building a Mars mission system.

GW Johnson · 2017-03-20 10:22:52

Kdb512's AF-M315E "green" propellant is actually a catalyst-decomposed monopropellant, not a fuel to be used with an oxidizer. Not yet, anyway.

It is the non-toxic analog to monopropellant hydrazine catalytic decomposition thrusters. Its performance is said to be "10% higher than hydrazine", which as a monopropellant, is around 250 sec Isp.

None of the hydrazine thruster hardware can be directly adapted, because the catalyst is different and hotter, and the decomposition chamber temperature is very much higher at 1800 C, instead of 800 C with hydrazine.

As near as I can tell looking about on the internet and in AIAA's magazine that ran an article on it in the latest issue, no one has yet tried it as the fuel in a bipropellant engine. The application is monopropellant thrusters for satellite attitude control and maneuvering. At least so far.

GW

Last edited by GW Johnson (2017-03-20 10:28:11)

SpaceNut · 2017-03-20 20:34:57

I did remember a monpropellant vehicle topic Armadillo VTOVL hop - In the spirit of DC-X it was post #4 that talked about silver-catalyzed monopropellant (Hydrogen-peroxide) engine..
Its a bit dated as its back to the VSE of 2004 plan but I would think that something could be found on the NASA web sites for technical info....

elderflower · 2017-03-21 04:23:46

The UK rocket program used silver catalysed peroxide with 85% peroxide as oxidiser. Higher percentage peroxide results in a discharge too hot for silver and you have to move to platinum catalyst.
85% peroxide freezes at -20C and has better stability than higher concentrations. However Isp is reduced, both as monopropellant and bipropellant due to the added water. I suspect that they will have used stabilisers as well.
In more than twenty launches there were no incidents on the pad, no fires, no explosions, nothing. One failure was due to a guidance fault and a second was due to a fuel leak. Peroxide was shipped from Sydney to Woomera , also without incident.
The fuel was injected downstream of the catalyst (they used jet fuel, described as paraffin). Isp was reported to reach 250-260.
The original rocket was used for testing re-entry vehicles for nuclear warheads, in conjunction with the US. This was called Black Knight. A satellite launcher was then developed called Black Arrow. After a few launches the government, in the usual short sighted way of governments in the UK, cancelled the program having launched one satellite.

Oldfart1939 · 2017-03-21 10:04:40

I haven't seen any data regarding any Isp data for the couples of MMH/LOX or ADMH/Lox. It seems that there isn't too much of a problem at deep space temperatures for keeping LOX around; both of these hydrazines have pretty low freezing points, so less supplementary heat would need be applied their tankage for keeping them available in the liquid form. All the gaseous combustion products would be low molecular weight compounds which allow for higher exhaust velocities which relate back to higher Isp values.

SpaceNut · 2017-03-21 21:18:02

This topic http://www.thespacerace.com/forum/index … pic=2583.0 has a table with possible ISP values.

This is a lengthy pdf but possibly may have what you are looking for.
ADVANCED HIGH PRESSURE ENGINE STUDY

Oldfart1939 · 2017-03-21 21:59:10

SpaceNut-

Great tables in the first article; took some time to dig out the appropriate numbers, but seems that Hydrazine comes out second only to LH2 w/r Isp, and much better Id, when combined with LOX. Better than LCH4, and when density is considered, may be the best overall non-cryogenic fuel/oxidizer combination. UDMH isn't too shabby, either.

Last edited by Oldfart1939 (2017-03-22 09:03:59)

JohnX · 2017-03-22 19:08:21

From what I've read (not having much background in rocketry) LH2 has difficulties in storage, and needs big tanks because it's not so dense.
LCH4 can be manufactured on Mars (as can LH2 obviously) as long as there's water-ice and power. And certain catalysts. But you all know that already. You're looking at other fuels such as hydrazine for a descent stage & attitude control, right? Can hydrazine be simply manufactured ISRU-wise?

Oldfart1939 · 2017-03-23 21:14:48

http://www.thespacerace.com/forum/index … pic=2583.0

Tables contained in this forum indicate that for getting a rocket off the ground with minimum vehicular weight, the fuel/oxidizer combination most efficient seems to be hydrazine/LOX. There are definitely more energetic propellant combinations, but they would unleash clouds of toxic combustion products that are totally unacceptable. Basing my conclusions on both the Isp and the Id values of the fuel. Id = (Isp)x (density ). MMH is just a tad lower, and UDMH slightly less. For some comparisons: CH4/LOX; Isp = 299, Id = 235; N2H4; Isp = 303. Id = 321; UDMH; Isp = 297, Id = 286; MMH; Isp = 300, Id = 298. Now compare these numbers for RP-1/LOX: Isp = 289, Id = 294.

From this data analysis, it would appear that a Falcon 9 first stage powered by N2H4/LOX might be able to carry more hydrazine in the existing tanks and get a performance enhancement for stage recovery. All numbers stated here are for sea level performance. Interestingly, the combination of N2H4/NTO is even higher at Isp = 286, Id = 342. The Russians are using MMH/NTO in all 4 stages of their Proton M space vehicles.

As I've stated before, I'm not a Rocket Scientist as is GW, but I am a chemical thermodynamacist by training.

elderflower · 2017-03-24 09:16:21

Whats the freezing point of NTO?

GW Johnson · 2017-03-24 16:42:03

My antique Pratt & Whitney Aeronautical Vest-Pocket Handbook says the freezepoint of NTO is 11.8 F, which is fairly insensitive to pressure level, and its boiling point is 70.1 F at 1 std atmosphere, which means that you have to put pressure on it to go hotter. Vapor pressure is 17.2 psia at 77 F, and 111 psia at 160 F. Those are the min inert-gas pressurant levels required to prevent immediate active boiling with bubble formation.

At thermal equilibrium, the vapor pressure plus your initial pressurant level will be the total pressure in the tank. At 77 F liquid pool temperature with 17.2 psia initial dry nitrogen (or helium) pressure, the final equilibrium will be 34.4 psia (half NTO vapor, half inert gas). At 160 F liquid pool temperature with 111 psia initial dry nitrogen (or helium) pressure, the final equilibrium pressure will be 222 psia (111 vapor plus 111 initial inert pressurant).

As you withdraw propellant from the tank, pressures fall rapidly unless you add makeup pressurant gas. After the withdrawal, given time to reach vapor equilibrium, the NTO vapor partial pressure will rise to the equilibrium value for that liquid pool temperature, increasing the total tank pressure, after the makeup inert gas addition. If you vent to maintain a pressure level, you lose both NTO vapor and inert gas. More NTO will evaporate to make up the equilibrium vapor pressure shortfall at the liquid pool temperature, but without violent boiling. However, NTO vapor is "lost" to you for purposes of liquid NTO withdrawals from the tank.

Sorry, it just ain't simple. But, it is exactly like handling anhydrous ammonia out on the farm. If farmers can do it, so can we.

But you need to understand what the vapor pressure curves really mean in terms of the how and why of phase change, and in terms of partial pressures. This is what so many misunderstand about the stability of liquid water (briny or fresh) at Martian conditions.

GW

Last edited by GW Johnson (2017-03-24 16:54:35)

Oldfart1939 · 2017-03-24 18:07:50

Looking at the possible fuels in more detail this afternoon; hydrazine is maybe too unstable for use as a fuel, but MMH and UDMH are a lot better; the blend of UDMH with hydrazine called Aerozine 50 looks very useable, as does MMH.

elderflower · 2017-03-25 04:13:44

GW, is there a flexible material which could be used to make a bladder in the tank to reduce demand for inert gas? Or would the candidates all be oxidised in short order? Would it be possible to breakdown NTO in small controlled quantities and feed the resulting gas back to the tank, so pressurizing the tank without using any inert gas.
After we have removed CO2 and CO from the martian atmosphere what is left is about 50 % N2. How easy is it to make NTO from this? Can it be done using a small automatic plant and is the product stable and storable over the long term?
The last question would also apply to hydrazine and it's derivatives.
I am still looking for a fuel/oxidiser combination for use both in a hopper to give crew mobility on the surface, and in an ascent vehicle.

Oldfart1939 · 2017-03-25 09:57:01

Looking at just the reactions for manufacturing UDMH, there is a possibility of producing this on Mars once we develop a source of Hydrogen, and to a lesser extent, Chlorine.

Reaction for production of UDMH: (CH3)2NH + H2NCl -----> (CH3)2-NH2 (Hydrochloride) ------------------> UDMH
Neutralize, Purify

NTO Might also be possible in the future, as well. Meanwhile, UDMH is also a great fuel for LOX--which is possible the easiest thing besides Methane for Mars production by a fully automated system.

These chemical products would definitely be high on any list for potential ISPU manufacture, as they could be utilized during exploration and development of an asteroid-based mining economy. This product---UDMH-- is non-cryogenic, a major advantage for storage thereof.

These product could at some time be manufactured by fully automated plant facilities, but not initially--let's go with the Zubrin "Kiss" philosophy. Oxygen is the #1 item for ISRU; followed by Methane. These. Will. Work.

As an afterthought, this sort of production will be pretty energy intensive, and will definitely require Nuclear power plant, not Solar...

Last edited by Oldfart1939 (2017-03-25 15:30:13)

SpaceNut · 2017-03-25 11:16:40

Yes priorities are for insitu use is for life support (oxygen & water) and fuel (methane & oxygen)...It appears that we do have the materials in the soils of mars to make lots of the alternative product that we need.

aHR0cDovL3d3dy5zcGFjZS5jb20vaW1hZ2VzL2kvMDAwLzAyOS84MjAvb3JpZ2luYWwvbWFycy1hc3Ryb25hdXQtaWxsdXN0cmF0aW9uLmpwZw==.jpeg

Curiosity Finds Water And Poison In Martian Soil

about 835 degrees Celsius (about 1,535 degrees Fahrenheit). At that temperature, the minerals in the soil break down and release volatile gases. A team of international scientists found water vapor, sulfur dioxide, carbon dioxide and oxygen in the sample, in that order of abundance.
a good amount of the sample—about 1.5 percent to 3 percent by weight—was water. a square foot of this—or, a cubic foot

NASA Curiosity rover digs Mars, finds sulfur, chlorine and organic traces of unknown origin

Toxic Mars: Astronauts Must Deal with Perchlorate on the Red Planet

NASA's Viking Mars landers in 1976 measured signatures of perchlorates, in the form of chlorinated hydrocarbons. Other U.S. Mars robots — the Sojourner, Spirit and Opportunity — detected elemental chlorine. Moreover, orbital measurements taken by the Mars Odyssey spacecraft show that chlorine is globally distributed. NASA's Curiosity rover found perchlorates within Gale Crater, where it landed in August 2012.
perchlorate is widespread in Martian soils at concentrations of between 0.5 to 1 percent. There are dual implications of calcium perchlorate on Mars. On one hand, at such concentrations, perchlorate could be an important source of oxygen.
microbes on Earth use perchlorate for an energy source.

GW Johnson · 2017-03-25 15:26:21

Elderflower:

Using polymer bladders is long practice in airplanes and missiles. In airplanes, they most commonly contain kerosenes and gasolines with urethane, a few with neoprene. Both work fine. When they start getting stiff, it is time to replace.

I dunno what might hold NTO, or IRFNA, or the hydrazines, or some of the other harsh or corrosive propellants. That's beyond my experience. To the best of my knowledge, bladders are not used with cryogens, due to catastrophic embrittlement cold.

The urethanes used in airplanes are destroyed by ethanol, which is why the airplane STC's for autogas are no longer desired. Neoprene is fine with ethanol, but it's nowhere near as common in airplanes.

Doing biodiesel blends with jet fuels, I found that the biodiesel rejuvenated age-stiffened fuel bladders without softening them beyond their original new-part durometer readings and sandpaper scuff resistance. Surprised the hell out of me (and the FAA), but very pleasing, too. You never know what will happen till you try.

I did an STC to put E-95 ethanol in an old Piper Pawnee cropduster. That stock bladder must have been neoprene, because I never had a problem running straight ethanol in it. The only incompatibility I ran into was the plastic blister over the fuel quantity indicator, which was also the tank vapor vent collection point.

Ethanol vapor destroys Lexan and Plexiglas. Liquid contact is OK, vapor is deadly. That's why the STC for ethanol in the Pawnee has you replace that blister with a steel flange and a mason jar from the grocery store. If it sucks corn squeezins, then there just has to be a mason jar somewhere. The FAA agreed with me, too.

Real life experience is not only stranger than any imaginable fiction, it can be funnier, too!

GW

RobertDyck · 2017-03-25 15:46:02

Elderflower:
Thruster quads of the Apollo Service Module used MMH as fuel, and oxidizer was N2O4 aka NTO. They contained both MMH and N2O4 in tanks with a rubber bladder liner. That way the fuel and oxidizer could be "squeezed out" to feed the thrusters. The space between bladder and metal tank was pressurized with helium. Because helium is inert and helium is very light.

The main engine of the SM used aerozine-50 as fuel, and N2O4 as oxidizer. However thruster quads used MMH and N2O4. Tanks for the main engine did not have bladders, but tanks for thruster quads sure did. Not sure how they managed that. Third stage of Saturn V was called S-IVB, and it had ullage motors. S-IVB used small solid rockets as ullage motors, which propelled the stage forward with just enough acceleration to cause propellant to settle at the bottom of the tank. Then propellant could pass through feed lines to turbo-pumps, and feed the main engine. Once the main engine ignited, it provided acceleration so ullage motors weren't needed any more. So the ullage motors only had to fire until the main engine ignited. Did the Apollo SM use the aft-pointing thruster of each thruster quad as ullage motors? Technically they wouldn't need all 4 to fire, just a pair on opposite sides of the SM.

Bottom line to your question is "yes". And Apollo already did it. I'm not sure but I think they used silicone rubber, because it doesn't react with either MMH or N2O4.

Last edited by RobertDyck (2017-03-25 15:47:02)

kbd512 · 2017-03-25 15:57:17

Oldfart1939,

If LOX/HAN produces an Isp similar to LOX/MMH or LOX and Hydrazine, then the ascent stage could propulsively land using HAN alone. The LOX tanks would be empty and the ascent stage mass would be similar to the Cygnus PCM-derived habitat. Once on the surface, MOXIE would use power from solar panels and batteries to produce and store LOX for the ascent stage, which means a reliable cryocooler is required. The similar landed masses would seem to indicate that a single HIAD design could be used for both Cygnus and the ascent stage. The TMI mass for the ascent stage would be comfortably below Falcon Heavy's 13.6t limit. LOX/HAN density impulse may not end up being quite as good as NTO and Hydrazine, dependent upon the optimal O/F ratio, but it should be similar to LOX/MMH or LOX and Hydrazine.

This seems like a reasonably good low-risk, high-reward ISRU / ISPP project that combines the Isp advantages of cryogenic oxidizers with low-toxicity storable fuels that don't require electrical power for thermal stabilization during storage periods.

If toxicity is not the major problem everyone has made it out to be, and since we're talking about NTO/MMH propellants I don't think it is, then LF2 combined with HAN should yield a fantastically high specific impulse, something near or above 420s. The impulse density should approach that of common solid fuels. The LF2 requires a cryocooler, but the mass of the fluorine should be quite manageable. LF2 and HAN should be hypergolic, so no ignition system is required and thus no fundamental design changes to the rocket engine are required. However, it's all but certain that materials changes are required, as was the case for the switch from Hydrazine to HAN.

There are problems with both approaches, but LOX/LCH4 or LOX/LH2 both increase the power requirements for the active cooling required. Strangely enough, in the late 70's F-based oxidizers were considered by NASA to meet STS in-space propulsion requirements.

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Re: Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

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