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LOX/LCH4 vs LOX/LCO Propellant Masses and Volumes for NASA's Mars DRM 5.0 1.5-Stage-to-Orbit MAV
Go to Page #114 for an estimated mass breakdown for a LOX/LCH4 powered Mars Ascent Vehicle (MAV):
Human Exploration of Mars Design Reference Architecture 5.0 Addendum
Inert Vehicle Mass: 13,944kg
32,498kg LOX/LCH4 propellant at 369s Isp delivers 4,353.8m/s Delta-V
47,304kg LOX/LCO propellant at 300s Isp delivers 4,353.8m/s Delta-V
14,806kg propellant mass increase for LOX/LCO, as compared to LOX/LCH4
LOX/LCO Mixture Ratio is 0.57143:1
Oxidizer mass = Mixture Ratio * Fuel mass
20,273kg LOX (17.768m^3); 27,031kg LCO (25.695m^3); 43.463m^3 total propellant volume
24,956kg LOX (21.872m^3); 7,542kg LCH4 (17.838m^3); 39.71m^3 total propellant volume
3.753m^3 total propellant volume increase for LOX/LCO, as compared to LOX/LCH4, for equal Delta-V, and 9.45% volume increase over LOX/LCH4
According to Dr Robert Zubrin, liquid CO2 can be obtained from the Martian atmosphere for an energy cost of ~84kWh/1,000kg.
According to the NIST document below:
Bond Dissociation Energies in Simple Molecules
The energy cost associated with breaking one of the double-bonds between Carbon and one of the two Oxygen atoms in a CO2 molecule is 532.2kJ/mol, which yields Carbon Monoxide and an Oxygen atom.
I recall very little of my basic chemistry, so take what follows with a truck load of salt.
CO2 Molar Mass: 44.01g/mol
CO Molar Mass: 28.01g/mol
1kg of CO2 is 22.72 mol
1kg of CO is 35.70 mol
17.85 moles of CO2 is required to obtain 1kg of CO.
We need 27,301kg / 974,645.7 moles of CO, so we need 620,309.6 moles / 27,302.4kg of CO2.
620,309.6 mol of CO2 * 532,200J/mol = 330,128,769,120J / 91,702,436Wh of dissociation energy for 27,301kg of CO
84,000Wh/t for CO2 liquefaction * 27.302t = 2,293,368Wh of energy to obtain LCO2 from atmospheric CO2
From 298.15K we must remove ~191.5kJ/kg of LOX, so 3,882,279,500J / 1,078,411Wh for 20,273kg of LOX.
From 298.15K we must remove ~225.15kJ/kg of LCO, so 6,086,029,650J / 1,690,564Wh for 27,031kg of LCO.
Carnot efficiency of modern heat pumps is about 0.5, so 2,156,822Wh to liquefy 20,273kg of O2 and 3,381,128Wh to liquefy 27,031kg of CO.
This presumes we start the liquefaction process at room temperature. Thermal dissociation of CO from CO2 typically occurs around 700C, but there are also room temperature methods involving catalysts, which is what we'd likely use in a practical Mars ISRU system to split CO2.
Energy Inputs Summary:
2,293,368Wh to obtain 27,302kg of LCO2
91,702,436Wh using a thermal dissociation method (our fallback if the room temp catalyst doesn't last long enough)
2,156,822Wh for liquefaction of 20,273kg of O2
3,381,128Wh for liquefaction of 27,031kg of CO
99,533,754Wh of total input energy to obtain 47,304kg of LOX/LCO propellant for a 6 person MAV.
Even if we had to double the input energy requirement, we're not talking about that much additional energy. Over the course of 365 days, 99,533,754Wh works out to a constant input power of 11,362.3 Watts. That's roughly what a 10kWe KiloPower fission reactor delivers. I'm generally in favor of "the nuclear option" because we only need to dig a hole, place the reactor in the hole, connect the power cable, remotely command the withdrawal of its singular control rod, and then leave it the hell alone while it delivers the power.
Since the western world seems very anti-nuclear right now, we do have an alternative. It's more complex than the nuclear option because it operates intermittently and is subject to diurnal cycles / seasonality / dust storms, but it could still work for an exploration class mission where money is no object. My personal contention is that we should take both options, in case one or the other fails. The most significant penalty is the increased weight associated with a battery subsystem to store and buffer the power being generated, into the ISRU equipment. This solution won't scale-up very well to provide power to a colony, unless made from ISRU materials. That matters less here where we're conducting a science experiment in the name of "pushing boundaries", and the personnel undertaking the mission have already volunteered to be human guinea pigs for science's sake.
In the summer time, in an equatorial region of Mars, we get about 586W/m^2 for 6 hours per day. There will be some power produced outside of peak hours, which ought to be used to top-up the battery to keep it warm at night. We have 29.5% BOL efficiency triple-junction photovoltaics provided by Northrop-Grumman (previously Orbital ATK). Some of the newest cells are better than this, but 29.5% efficient cells are legacy tech with flight heritage in the actual environment where they will be used. In other words, it's proven to work across multiple missions, and we like things that have exhaustive testing and real world use to their credit. The integrated system is better known as the MegaFlex foldable / deployable photovoltaic array. MegaFlex was successfully used to power our Mars InSight Lander.
That means we should expect to convert 172.87Wh/m^2 * 6hrs = 1,037.22Wh per day under "full Sun" conditions on the surface of Mars. 250m^2 of array surface area therefore nets 259,305Wh per day, so we can make the LOX/LCO propellant over roughly 384 days under ideal power generating conditions. Our surface stay is almost 2 years, so we have some buffer to work with if things don't go precisely to plan.
Methane is a better rocket fuel than Carbon Monoxide, especially as the Delta-V requirement increases, but 100MWh per year is not asking for an inordinate amount of power. The technological hurdle to clear for an atmospheric-only ISPP solution set is far lower than it is for Methane production, which includes a Sabatier reactor and shallow surface mining machines to scoop up Martian regolith and extract whatever water it contains. We should locate a good water supply first, and then figure out how much equipment and money is required to convert that water into rocket fuel. Even if we can do that without undue effort, I will opine that washing, cooking, cleaning, and flushing are higher priorities than obtaining premium rocket fuels when a cheaper / easier to produce fuel is available.
Exploration mission 2 or 3 should introduce the Sabatier reactor and mining equipment. The first mission should simply prove that we can do atmospheric ISPP and extract as well as purify usable quantities of liquid water from the frozen Martian regolith. It's probable that the additional power and equipment mass required to make Methane is not energy or mass favorable. Liquid CO2 is easy to get, whereas liquid water is not. Whatever water a Mars colony does get is going to have a rather lengthy list of higher priority uses.
Here on Earth, apart from electric power generation, Methane is the base stock for modern synthetic lubricants, plastics and rubbers, synthetic fibers, refrigerants, and industrial chemicals such as solvents. The Hydrogen in water will also be required to make Sulfuric acid. Mars has abundant surface deposits of Sulfur, and Sulfuric acid is the basis of most industrial chemical processes. Even if Mars has Methane reservoirs buried in the ground, those "other uses" are more essential to colonization efforts than rocket fuel. We're setting up an entire human civilization from scratch, from the starting point of relative energy poverty since absolutely nothing is "human ready" on Mars. Therefore, economizing on energy and materials inputs, even for rocket fuel, is essential.
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This post is reserved for an index to posts that may be contributed by NewMars members over time.
Update after scanning the document kbd512 cited: This paper from 2009 reflects work by a great number of NASA employees and contractors in an effort to understand and to document the Mars exploration scenario as it was foreseen at that time. A detail that caught my eye is the use of a two stage vehicle to ascend from Mars. As regular NewMars readers will know, GW Johnson created multiple sets of designs for vehicles for a variety of purposes for the Mars adventure. None of those designs used more than one stage. The structure of the NewMars forum does not facilitate easy discovery of all that work. I am hoping we (admins and members) might be able to find ways to improve that weakness. I have asked GW Johnson to see if he can find his work in this particular area.
I would note that my pdf player in the browser displayed "pages" in a panel to the left of the text. Those "pages" did not match the page numbers in the document. Document page 114 is shown as 130 (or so) in the side bar. I assume that pages of the pdf that were not numbered explain the difference.
Index:
kbd512:Update to NASA document: 2019 updated from 2009
http://newmars.com/forums/viewtopic.php … 25#p231625
(th)
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My PDF player displays correctly. For NASA DRM v5 Addendum, page 1 starts with the Introduction, the first page of text. Not included in page numbers are the cover page, NASA program office info, second cover page, forward, table of contents, index of figures, index of tables. The table of contents and indexes have Roman numerals. However, the PDF player simply numbers pages from the start, so document page 1 is PDF page 21.
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What I did was show that a single stage vehicle of very modest payload fraction (near 5%) could indeed make the round trip between the Martian surface, and low Mars orbit, with any of the common propellants that we know and use. And that includes the storables: NTO and any of the hydrazines.
The higher your Isp, the larger your payload fraction can be, but single stage it is still modest even with LOX-LH2. It's just very small with the storables. Over 5% with LOX-LH2, under 5% with storables.
The only reason I did this as a single stage item was to look at reusability/long service life. If you stage the lander, you cannot re-use the first stage because you jettison it. And on a wild and uninhabited planet, recovering such a jettisoned stage is unlikely in the extreme. But, the overall payload fraction of a two-stage vehicle is very much larger. It's a tradeoff: bigger payloads vs one-shot throwaway first stages.
Whether based on the surface, or based from orbit, this single stage lander is possible. Surface based, the numbers are a bit different from orbit-based, but they are still comparable. However, things sent to low orbit around Mars inherently involve a higher dV for the trip from (and back to) Earth, because of what's needed to decelerate into (and accelerate back out of) low Mars orbit, not to mention rendezvous budgets. Direct landings do not have that dV requirement, because hypersonic atmospheric drag deceleration can do the majority of that job, even in Mars's thin atmosphere. That's part of another tradeoff you have to make: reusable orbital transports plus reusable landers and their logistics, vs throwaway one-shot direct landers.
Although, at higher ballistic coefficient, you are looking at a powered landing of much larger dV than you might otherwise think, so the direct-landing dV advantage is nowhere near as attractive as most people seem to think it is!
That advantage almost disappears when you are too large and come out of hypersonics too low to use parachutes! Smaller unmanned probes (at or under 100 kg/sq.m ballistic coefficient) can use chutes, but manned vehicles or items with significant payload mass (300-1000+ kg/sq.m) cannot use chutes, because they come of hypersonics at or under 5 km altitudes, while the small probes come out at or above 20 km altitudes.
If for the sake of illustrative argument you assume end of hypersonics at 0.7 km/s on a straight path angled 45 degrees down, then at 5 km altitude, you are only 10 sec from an un-decelerated impact. If instead you come out of hypersonics at 20 km, same speed and angle, you are almost some 50 sec from an un-decelerated impact. 0.7 km/s is still too fast to deploy a chute, and it takes about 5 sec to deploy one. And it still takes more time (at least 20-30 sec) to slow you down further by any significant amount! See the difference altitude makes?
Even SpaceX's "Starship" must face this issue: I really do not believe they can execute the near-surface pull-up deceleration to subsonic without rocket thrust. The reliable lift is just not there. The extra landing propellant is going to displace payload, even if they make the vehicle work as an Earth orbital transport, which they still have yet to do. Plus, they still have to make tanker refueling work! And, in my humble opinion, very few of these tall narrow vehicles will survive touching down on a rough landing surface that is also soft. SpaceX has never made a dirt landing yet, even with its Falcons.
I'm not saying that this can't be done! I'm saying that doing it "right" is a very difficult thing. And getting those answers is way-to-hell-and-gone more complicated than just running a rocket equation calculation or two.
GW
Last edited by GW Johnson (Yesterday 09:11:15)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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NASA's updated / more detailed MAV design, circa 2019:
Update to Mars Ascent Vehicle Design for Human Exploration
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Using NASA's 226:1 "gear ratio" for propellant / engines / structures masses required to land 1kg on the surface of Mars, we can see that the 8,892kg of LCH4 fuel from the "5 sols" 2019 MAV design refinement, brought all the way from Earth to the surface of Mars, results in the requirement for 2,009,592kg of (mostly propellant) mass to deliver that LCH4 to the surface of Mars. In reality, it's actually worse than that, because additional LOX/LCH4 engines which are not required for Mars ascent are used during Mars EDL to land the MAV. Some of this additional engine power is undoubtedly a result of leaving the heat shield affixed to the lander.
If I understood the document correctly, 8X 100kN LOX/LCH4 engines in total are used by their 5 sols MAV design, yet only 4 are required for ascent. The rest of the engines are not required after EDL. They show engine mass optimization charts wherein more powerful main engines save 1,000kg to 2,000kg of weight in "Figure 12". I think that's a necessary optimization, because it saves 452,000kg of launch weight. If the exploration campaign is only 6 missions, then 2,712t of launch mass is eliminated, which is creeping up on the 3,300t GLOW of a Falcon Heavy heavy lift vehicle, which costs NASA about $95M.
Keeping those LCH4 fuel tanks cold for up to 5 sols, and leak-free, is a fairly tall order. That's another advantage of making all ascent propellant on the surface of Mars, using the Martian atmosphere, at an energy cost on-par with their LOX-only ISPP solution. If you're already making a similar tonnage of LOX from atmospheric CO2, then you may as well collect and liquefy the Carbon Monoxide produced by that process as well.
The 3X 10kWe KiloPower fission reactors NASA proposes to use for their surface exploration campaign can provide 262.8MWh worth of power per year, which is more than sufficient to make enough LOX/LCO propellant from the Martian atmosphere for LOX/LCH4 equivalent Delta-V, and the MAV can be launched and landed using EDL-only propellant in its tanks, when LCO becomes the fuel of choice. The only thing we're changing is adding the energy cost of liquefaction of the Carbon Monoxide instead of dumping it overboard after we get the LOX. Either way, we have to process about as much CO2 because the LOX/LCH4 MAV requires 29,142kg of LOX.
There's not enough LOX/LCO propellant mass increase, relative to LOX/LCH4, for a low Delta-V requirement vehicle (less than half of what is required to achieve orbit here on Earth), to justify the expense and complexity of either making Methane propellant from Martian ground water or shipping Methane to Mars from Earth and keeping it cold for up to 5 years. If you're going to do ISPP for the MAV, then max out that "gear ratio" benefit through simplification of the natural resource processing requirements to only Martian atmosphere processing.
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Spectrolab, a Boeing company. Space Photovoltaics
Datasheet XTE-SF: 32.2% XTE-SF (Standard Fluence)
Based on 20+ years of heritage 3J devices
Fully qualified under AIAA-S111 2014 Standard
Targeting LEO to GEO mission fluences
Best in class 32.2% BOL efficiency
27.9% EOL, 1E15 1MeV electron**
Multiple Sizes Available (<85-cm2)
Currently in Production
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RobertDyck,
Thanks for that contribution. That bumps our power output to 103,308,870Wh per sol per 250m^2 array. From doing the math on their square 27cm^2 cells, 1m^2 of photovoltaic cells weighs 839.16g, the lowest practical weight for a T1200 composite and honeycomb backer board (2 ply per face sheet, around the Kevlar honeycomb sandwich core) is about 1kg per square meter, and the wiring, if CNT / Copper composite (~50% more electrically conductive than Copper alone and less than half the weight), would add another 1kg to get the power off the panel. A small dual-axis stand / Sun tracker for the panel would weigh another 1kg, so 4kg per 1m^2 panel, or 1,000kg per 250m^2 array.
Output will drop during the winter months, but overall the array and battery bank is about equal in weight to a 10kWe KiloPower reactor. We cannot operate the ISPP equipment without a battery bank. Heating up and cooling down MOXIE once per day is a better than average way to crack the ceramic in the solid oxide fuel cell that produces CO and O2.
From the DRM 5 addendum document provided by NASA, their state-of-the-art cryocooler consumes 1,200W of power to provide 150W of cooling capacity to remove heat from the gases undergoing liquefaction. We need to remove 2,768,975Wh of thermal energy from 47,304kg of LOX/LCO propellant to lower their temperatures from 298.15K to just below their respective boiling points, which implies 22,151,800Wh of input electrical power to the cryocoolers, for a constant power draw of 2,528.74W.
Revised Energy Inputs Total:
116,147,604Wh to obtain 47,304kg of LOX/LCO propellant.
We would thus require 2X 250m^2 arrays of those Spectrolabs cells to ensure our ISPP processing time remains below 1 year.
Each LEU-fueled 10kWe KiloPower unit is expected to weigh at least 2,000kg and NASA intended to bring 3 of them, so 6,000kg in total, which provides 262,800,000Wh over a sol. If we have a 6,000kg mass budget to work with for the power generating equipment, then I will opt for 3X of those 250m^2 arrays and 3,000kg worth of batteries in properly protected enclosures.
Amprius offers 500Wh/kg and 1,300Wh/L prismatic cell Lithium-ion batteries, rated for 1,300 full charge / discharge cycles before significant degradation occurs, for commercial sale. NASA presently uses GS Yuasa prismatic cells aboard ISS, so they're familiar with how to design enclosures for the prismatic variety. Presuming our battery pack enclosure roughly halves the stated Wh/kg value of the cells themselves, a 3,000kg battery bank can store 750kWh worth of power (46,875 Watts per hour over 16 hours, if fully discharged), which should be enough to keep that solid oxide fuel cell very very hot and the propellant very very cold during the night.
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We should locate a good water supply first, and then figure out how much equipment and money is required to convert that water into rocket fuel.
I've said before, we need to send a rover with a multi-segment core drill to the frozen pack-ice of Cerberus Fossae. That's 5° north of the Mars equator, and in the dried-up ocean basin so low-altitude. Altitude is important for additional atmosphere for radiation shielding. In a vast region known as Elysium Planitia. About 20° latitude or about 1,000km south of the volcano Elysium Mons. Maybe 1,200km, but considering the pack ice is 800km x 900km, it depends where you measure from.
Back in 2005, the head of the Canadian Space Agency proposed a rover about the size of Spirit/Opportunity with such a drill. They developed a prototype drill, dry drilling using an electric motor. Each segment 1 metre long. So a rover the size of Curiosity isn't necessary. I could cite the work by the European Space Agency that has a lot of evidence to conclude it's ice, but we need "ground truth".
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RobertDyck,
I also think we should be looking for a suitable water source during the very first mission, but "betting the farm" on finding water that we can readily use for propellant production is a bad idea. We have no clue what Martian water is mixed with because we've never even attempted to drill through the ground to obtain samples. If we knew exactly where to stick our soda straw, because we knew how far down the water table is, what was dissolved into that ice, and how much was there, then let's devise a way to obtain some of that water so we can make the best rocket fuel we can get. We need to actually do that at least once before we stake human lives on obtaining local water.
Why did we not undertake that effort already if Mars was the ultimate exploration target, as NASA has claimed since the Apollo program?
I have no idea, but it clearly wasn't a priority in the past. We need more than lip service paid to the idea.
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A mission the size of Mars Exploration Rover (Spirit/Opportunity), could be easily launched. I consider that low risk. The Canadian proposal for a rover would include sample handling and multiple instruments on its back to analyze ice cores. Analysis by ESA claim the pack ice is 45 metres deep, based on craters that punch all the way through to ground beneath the ice. 45m = 147.64 feet, round to 2 significant digits.
The ice is believed to have formed when the volcano Elysium Mons erupted between 2 million and 20 million years ago; most likely 7 to 10 million years ago. Frozen from the ocean floor (frozen wet mud) melted, releasing water. Since the ocean evaporated, it wouldn't just be ocean water, it would be extra salty. The water ran down hill to a low area and pooled. The large lake then froze. When a body of water that large freezes, it forces out salt, creating fresh water ice on top. Eventually the water is so super-saturated that it can't force out the salt, and salt water freezes. So the ice should have differing salinity: fresher on top, salter deeper down.
It would have been muddy water, such as lake water. Could it have sat motionless long enough for silt to settle out, or would it be turbid? That raises the question whether mud is dissolved in the ice. Is it similar to current Mars surface, or more like the ancient Mars ocean floor? Or a mix of both?
Again, a rover with a core drill could answer these questions.
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The more I study NASA's MAV design, the more it looks like pointless complexification to keep engineers and vendors busy, or to prove that they can do some specific set of technology demonstrations that don't function in a way that minimizes the mass, cost, and development timeline of the design. I'm going to work on something that doesn't add complexity and mass for no mission benefit.
If we're going to bring the fuel all the way from Earth, then we're going to bring NTO/MMH. At least that requires no ISPP mass, no brand new engine development, and no cryocooling until ascent from Mars. We get 6,000kg of mass back by not having to bring the reactors, plus whatever mass the support equipment requires to robotically emplace the reactors without any humans present, plus the entire ISPP plant mass.
For a 5,314kg dry vehicle mass and 19,612kg wet mass at ignition, I only need 2X Aestus-II / RS-72 pump-fed NTO/MMH engines, which are used by the Ariane rocket's upper stage. That's quite literally less than half the mass of NASA's ridiculous vehicle design which still uses smaller NTO/MMH engines on the cruise stage used to deliver the MAV to Mars. If we're gonna load Hydrazine anyway, then we may as well cross-out all this pointless complexification associated wtih keeping the Methane liquid from the point in time that the vehicle is launched from Earth to the point in time when someone hits the button on Mars, up to 5 years later, robotic emplacement of the KiloPower reactors, making all the propellant before the crew lands, etc. The entire point of this mission is to explore Mars, not babysit the MAV.
9,368kg (7.807m^3) NTO, 4,930kg (5.634m^3) MMH, 13.441m^3 total propellant volume for the same Delta-V that the LOX/LCH4 provides.
That is less total propellant volume than the LCH4 tank volume for NASA's MAV, and it absolutely will weigh less than all the equipment that would otherwise be required. If we do ISPP at all for the first mission, LOX/LCO all the way. LCH4 makes no sense. Otherwise, NTO/MMH for the win.
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