Phobos First!

[j05h]

We been touching on this subject in "Mars 9 tons at a time", and it seems time to bring the old subject back up. Should we go to Mars' moons first? Accessing Phobos and Deimos is definitely easier than the surface of Mars and cheaper in delta-V than getting to the lunar surface. <br /><br />I covered this argument extensively in spring '06 in the thread "Private Mars Missions" and it's also been touched on lately in "Mars 9 tons..."<br /><br />The argument breaks down to one thing: volatiles. If Phobos or Deimos has any amount of accessible water or other hydrates, these bodies will immediately become competitive as a target for industrial development. We know there is water on Mars, but it requires surface-access. Lunar volatiles, if they exist, are in dark polar craters - the hydrogen readings from our sensors have not confirmed water, it could be locked up in rocks. We already know how to function in freefall and some of the modules for this are off-the-shelf (Russian FGB), helping the utilization of these mini-moons. The real question: where's the water?<br /><br />Another aspect is building a "base camp" in Martian orbit. Stockpiling supplies and locally-produced volatiles (propellant, potable water and breathing atmosphere) in Mars orbit provides strong efficiencies for rocket transport and supply production heading to cis-lunar space. It also provides huge safety margins for any Mars surface activity. Instead of thinking of a single Mars base, supplies on-orbit can provide global surface access. <br /><br />So, what are the advantages and disadvantages of using and exploring the Martian Moons? What timeframe can they be accessed in? <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[thereiwas]

To make methane you also need carbon. Carbon is readily available on Mars since the atmosphere is nearly all CO2. On the moons the carbon would have to be extracted from the rocks.<br /><br />On the plus side, the larger cruise habitats from the Earth-Mars transit, that would not be brought to the surface, could be collected at Phobos to form a base there.

 

[docm]

IMO just going to Phobos after going to all the way to Mars space is like getting into bed with the woman of your dreams and not having sex <img src="/images/icons/tongue.gif" /> <div class="Discussion_UserSignature"> </div>

 

[spacester]

Phobos First! . . . as a means to explore the Martian Surface.<br /><br />If there's water . . . if not . . . Phobos Later.<br /><br />Send a prospector and find out and then we can decide.<br /><br />In the meantime, it's well worth talking about.<br /><br /><font color="yellow">On the moons the carbon would have to be extracted from the rocks.</font><br /><br />Or brought up as compressed atmosphere on the back-haul from taking stuff down to human habs.<br /><br /> <div class="Discussion_UserSignature"> </div>

 

[keermalec]

<blockquote><font class="small">In reply to:</font><hr /><p>IMO just going to Phobos after going to all the way to Mars space is like getting into bed with the woman of your dreams and not having sex <p><hr /></p></p></blockquote>Docm you are of course right but getting to and from Mars is MUCH more expensive in terms of delta-v than getting to and from Phobos.<br /> <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[keermalec]

<blockquote><font class="small">In reply to:</font><hr /><p>To make methane you also need carbon<p><hr /></p></p></blockquote><br />Why make methane? A LOX/CH4 engine has an ISP of "only" 379s. A LOX/LH2 engine has an ISP of 465s. We have mastered the technique of storing LH and LOX for long periods of time (years) using active cooling. Zubrin's LOX/CH4 idea was valid before actived cooling became reality. <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[keermalec]

OK here are some basic delta-vs assuming all circular orbits and Hohmann transfers:<br /><br /><br />Phobos<br />------<br />Departure from LEO: 3.9 km/s incl. gravity losses<br />Mid-course corrections: 0.1 km/s<br />Arrival aerobrake delta-v: 1.42 km/s (entry velocity 5.61 km/s)<br />Orbit circularisation at 5988 x 5988 km: 0.57 km/s<br />Land on Phobos: 0.02 km/s<br />Total propulsive delta-v: 4.59 km/s<br /><br /><br />Deimos<br />------<br />Departure from LEO: 3.9 km/s incl. gravity losses<br />Mid-course corrections: 0.1 km/s<br />Arrival aerobrake delta-v: 1.02 km/s (entry velocity 5.61 km/s)<br />Orbit circularisation at 20070 x 20070 km: 0.66 km/s<br />Land on Deimos: 0.02 km/s<br />Total propulsive delta-v: 4.68 km/s<br /><br /><br />Josh you are right the difference between landing on Deimos and landing on Phobos is negligeable.<br /><br />I see two possible ship configurations to get there:<br /><br /><br />1. The fast ship<br />----------------<br />- Expendable launch stage to boost onto TMI<br />- Aeroshell for capture at Mars<br />- Total delta-v to Phobos = 4.59 km/s<br />- Payload as a percentage of inital mass in LEO: 18%<br />- Time of flight: about 9 months<br /><br /><br />2. The slow ship<br />----------------<br />- One piece ship with ion drive<br />- Multiple burns in Earth Orbit to reach interplanetary space<br />- Continuous thrust from Earth to Mars<br />- Multiple aerobraking at Mars without an aeroshell<br />- Multiple burns in Mars orbit to reach the orbit of Deimos or Phobos<br />- Total delta-v to Phobos = 10.12 km/s<br />- Payload as a percentage of inital mass in LEO: 50% (at 0.00002 Gs acceleration)<br />- Time of flight: about 24 months<br /> <br />The above assumes circular earth and Mars orbits, which is not correct. As the orbit of mars is elliptic, depndeing on the conjuntion trip tiomes could be longer or shorter. The above indicates average trip time and delta-v. <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[gunsandrockets]

<Why make methane? A LOX/CH4 engine has an ISP of "only" 379s. A LOX/LH2 engine has an ISP of 465s. We have mastered the technique of storing LH and LOX for long periods of time (years) using active cooling. Zubrin's LOX/CH4 idea was valid before actived cooling became reality.> <br /><br /><br />You are both right and wrong.<br /><br />For a rocket in Mars orbit using water mined from Deimos/Phobos, a LOX/LH2 rocket engine using propellant active cooling is perfectly reasonable.<br /><br />But even so there is nothing obsolete about a Zubrin style hydrocarbon propellant rocket for launch from the surface of Mars. Hydrogen is much easier to keep liquified in space than on the surface of Mars, so active cooling could be much more difficult on Mars. <br /><br />Secondly the Sabatier Reactor and/or Reverse Water Gas Shift for propellant production on Mars depends on a supply of carbon dioxide gas which is fairly easy to gather on Mars by taking it from the Martian atmosphere. Unless honest to god ice is found at fairly shallow depths beneath the surface of Phobos/Deimos, deriving volitiles from the moons of Mars is going to be much more difficult than from the surface of Mars. Of course balanced against that is the tremendous potential energy advantage the moons gain from orbital position.<br /><br />Even if the moons of Mars are exploited for propellant production, it's still easier to leave the surface of Mars and gain low Mars orbit using propellant produced on the surface of Mars. So there would be a logical division of propellant production in a fully developed flight architecture -- Deimos/Phobos propellant for leaving Mars orbit to either land on Mars or depart towards Earth, and Mars propellant for launching off of Mars into low Mars orbit.

 

[gunsandrockets]

<IMO just going to Phobos after going to all the way to Mars space is like getting into bed with the woman of your dreams and not having sex><br /><br />I imagine the crews of Apollo 8 and Apollo 10 felt the same way. Yet a logical system of transportation development requires intermediate steps to most quickly and successfully achieve the final goal, the goal in this case the landing of men on Mars.<br /><br />So if you want to get to Mars, choose your poison...<br /><br />1) Nuclear rocket engines?<br /><br />2) Propellant manufacture at Phobos/Deimos?<br /><br />3) Titanic sized spacecraft with chemical rocket engines?<br /><br />4) Extremely risky flyby-rendezvous mission architecture?<br /><br />Any of these will work, but they all have obvious drawbacks. <br />

 

[j05h]

<i>> IMO just going to Phobos after going to all the way to Mars space is like getting into bed with the woman of your dreams and not having sex</i><br /><br />No, the plan would be to go to Phobos (or D), mine volatiles and provide early tele-operation for robotic Mars outposts. The idea is that a well-financed consoritia in "Private Mars Missions" would set up shop around Mars perhaps a decade before any landing architecture. People could base out of existing or near-term station modules almost as soon as they could be purchased. It would provide years of leverage in orbital Mars ops before any reasonable Mars Direct or DRM, then provide "base camp" support to surface expeditions. There are other uses for the moons, too. <br /><br />Josh <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[keermalec]

<blockquote><font class="small">In reply to:</font><hr /><p>Even if the moons of Mars are exploited for propellant production, it's still easier to leave the surface of Mars and gain low Mars orbit using propellant produced on the surface of Mars. So there would be a logical division of propellant production in a fully developed flight architecture -- Deimos/Phobos propellant for leaving Mars orbit to either land on Mars or depart towards Earth, and Mars propellant for launching off of Mars into low Mars orbit. <br /><p><hr /></p></p></blockquote><br />Actually there is a good reason to use LOX/LH2 based engines to lift from Mars surface, instead of LOX/CH4: useful payload to LMO.<br /><br />Let me explain, because of its higher isp, a LOX/LH2 lander/lifter will lift twice as much payload to LMO than a LOX/CH4 rocket.<br /><br />Given the availlability of water on the martian surface, and the higher ease with wich it can be transformed into LH and LOX (simple hydrolysis), as compared to CH4 and LOX (Sabatier reactor depending on imported H2), a LH2/LOX rocket is a better choice to me. <br /><br />Also, I see no reason why active cooling should be easier to carry out in orbit than on the surface. Active cooling is just 2 inches of insulation around the tank surrounded by tubes of cooling fluid connected to a Brayton reactor. If it works on Earth at 288 K it should work on Mars at 234 K. <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[thereiwas]

Sabatier reactions do not need imported H2. Local H2 extracted from that same ice will do.<br /><br />How much energy does it take to keep LH2 from boiling here on Earth? LH2 boils at 21 Kelvin. Mars is colder than Earth (night time low 184 Kelvin) so the energy would be less, but it isn't <i>that</i> much colder, compared to space. <br /><br />CH4 boils at 111.5 Kelvin.

 

[gunsandrockets]

<Let me explain, because of its higher isp, a LOX/LH2 lander/lifter will lift twice as much payload to LMO than a LOX/CH4 rocket.><br /><br />It's not quite that simple.<br /><br />Let us compare a variety of Mars ascent vehicles, the only difference of which is the ascent propulsion. These ascent vehicles would have the same empty weight on the surface of Mars before they fuel-up for blast-off.<br /><br />Because empty masses are the same, the vehicles would also share the same sized propellant tank volume. So a primary limitation on any vehicles performance would be the limited volume of propellant any vehicle could carry as well as the ISP of the rocket engines. <br /><br />This problem in fact is very similar to that faced by developers of Single-Stage-To-Orbit (SSTO) reusable launch vehicles, in quest of the holy grail of cheaper transportation to space. Dunnspace has an invaluable chart for examining different types of propulsion for exactly this situation, that of SSTO performance to LEO. You should check it out...<br /><br />Alternate SSTO Propellants<br /><br />... noteworthy is the low performance level (as expressed in tons of payload delivered to LEO) of the classic LOX plus liquid hydrogen propellant combination.<br /><br /> <br />

 

[gunsandrockets]

<Also, I see no reason why active cooling should be easier to carry out in orbit than on the surface. Active cooling is just 2 inches of insulation around the tank surrounded by tubes of cooling fluid connected to a Brayton reactor. If it works on Earth at 288 K it should work on Mars at 234 K.><br /><br />Did that experiment in active cooling take place in an environment of sea-level-pressure room-temperature-warm air? Could a lightweight active cooling system cope with such conditions? Doubtful.<br /><br />Cryogenic hydrogen storage is an interesting subject which came up before in this fascinating post... <br /><br /><br />http://uplink.space.com/showthreaded.php?Cat=&Board=missions&Number=583189&page=0&view=collapsed&sb=5&o=0<br />

 

[gunsandrockets]

Conference to Discuss Exploration of the two Moons of Mars

 

[gunsandrockets]

<br />SCOUT for Phobos/Deimos EVA.

 

[keermalec]

Precisely ThereIwas, if you are hydrolizing ice to get Hydrogen, you are obtaining Oxygen for free. Why then bother using the Hydrogen in a Sabatier reactor to produce CH4 if you already have a better fuel?<br /><br />Concerning zero boilloff, this is a proven technique carried out at standard temperature and pressure. Therefore it can only be easier at the colder temperatures on Mars. For information, in NASA's DRM-3 revision 1, active cooling is used to cool the liquid Hydrogen tanks and cancel boilloff for years. <br /><br /><blockquote><font class="small">In reply to:</font><hr /><p><Also, I see no reason why active cooling should be easier to carry out in orbit than on the surface. Active cooling is just 2 inches of insulation around the tank surrounded by tubes of cooling fluid connected to a Brayton reactor. If it works on Earth at 288 K it should work on Mars at 234 K.> <br /><br />Did that experiment in active cooling take place in an environment of sea-level-pressure room-temperature-warm air? Could a lightweight active cooling system cope with such conditions? Doubtful.<p><hr /></p></p></blockquote><br />For NASA at least, a lightweight active cooling system is a reality. I dont have the link here as I am travelling but I posted it earlier. Google "zero boiloff" to find it. <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[keermalec]

<blockquote><font class="small">In reply to:</font><hr /><p>Let us compare a variety of Mars ascent vehicles, the only difference of which is the ascent propulsion. These ascent vehicles would have the same empty weight on the surface of Mars before they fuel-up for blast-off. <br /><br />Because empty masses are the same, the vehicles would also share the same sized propellant tank volume. So a primary limitation on any vehicles performance would be the limited volume of propellant any vehicle could carry as well as the ISP of the rocket engines. <br /><br />This problem in fact is very similar to that faced by developers of Single-Stage-To-Orbit (SSTO) reusable launch vehicles, in quest of the holy grail of cheaper transportation to space. Dunnspace has an invaluable chart for examining different types of propulsion for exactly this situation, that of SSTO performance to LEO. You should check it out... <br /><br />Alternate SSTO Propellants <br /><br />... noteworthy is the low performance level (as expressed in tons of payload delivered to LEO) of the classic LOX plus liquid hydrogen propellant combination. <p><hr /></p></p></blockquote><br />Gunsandrockets, I believe this article is misleading. Rockets with different propellant types are not designed in the same way (with the same tank volume for instance) and therefore propellant preformance cannot be compared by volume. They must be compared by mass, and in this case LOX/LH2 is always the winner.<br /><br />You will note that all of our existing large payload-to-LEO launchers use LOX/LH2: The Space Shuttle, the Saturn V, Energiya. The reason for this is because ISP performance dwarfs any other consideration when attempting to reach the largest possible delta-v.<br /><br />The Dunnspace article makes some strange assumptions, worthy of note:<br /><br />1. It assumes tank and engine size are roughly the same, irrespective of the propellant type and density.<br /><br />2. It assumes target delta-v is 9.2 km/s "only", when all existing rocket <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[gunsandrockets]

I am replying on the fly so here is a partial response, more later:<br /><br /><You will note that all of our existing large payload-to-LEO launchers use LOX/LH2: The Space Shuttle, the Saturn V, Energiya. The reason for this is because ISP performance dwarfs any other consideration when attempting to reach the largest possible delta-v. /><br /><br />Not true. The examples you provided proove that ISP does NOT dwarf all other considerations. <br /><br />The only modern booster which uses pure LOX/LH2 propulsion is the Delta IV (and even then only in some variations such as the Delta IV heavy). All the other rockets and all of the examples you provided, either don't use hydrogen at all in the first stage such as the Atlas V and Saturn V, or use hydrocarbon fueled liquid boosters or solid rocket boosters to supplement a hydrogen fueled sustainer. The ISP advantage of hydrogen really comes into it's own for upper stage propulsion when factors such as drag and thrust to weight are no longer as important as they are during the early stages of flight.<br /><br /><The Dunnspace article makes some strange assumptions /><br /><br />Not really.<br /><br /><, worthy of note: 1. It assumes tank and engine size are roughly the same, irrespective of the propellant type and density. /><br /><br />The study does not assume engines of the same size, it assumes 2% of propellant weight for hydrocarbon engines and 3% of propellant weight for hydrogen engines. And because the point of the study is to compare vehicles of indentical dry mass instead of indentical gross mass, the vehicles tank size must be the same. If anything assuming identical tank design provides an unfair advantage to low temperature propellants such as liquid hydrogen which requre heavier than normal tanks. <br /><br />The DunnSpace study is especially usefull for analyzing Mars Ascent Vehicle performance (which exploits ISRU) because the dry mass of a MAV that can be delivered to the surface of Mars is a key

 

[gunsandrockets]

<Concerning zero boilloff, this is a proven technique carried out at standard temperature and pressure. Therefore it can only be easier at the colder temperatures on Mars... For NASA at least, a lightweight active cooling system is a reality. I dont have the link here as I am travelling but I posted it earlier. Google "zero boiloff" to find it.><br /><br />I searched and found the original link you posted about the NASA zero boiloff experiment...<br /><br />http://www.grc.nasa.gov/WWW/RT1998/5000/5870plachta.html<br /><br />... and as I thought the experiment was conducted in a vacuum chamber to simulate a space environment, it was not conducted in a room temperature/pressure environment.

 

[gunsandrockets]

... followup continued.<br /><br /><As can be seen, even though the tank mass of the LOX/CH4 launcher is half that of the LOX/LH2 launcher, because of the lower density of LOX/LH2, the final payload in LMO is still larger for LOX/LH2 because its isp is better. /><br /><br />Well your delta-V seems a little high, and the engine masses seems grossly oversized, but you are right that when comparing two different Mars Ascent Vehicles of the same liftoff mass of 10 tonnes the hydrogen fuelled MAV would outperform the methane fuelled MAV. But your analysis contains a fatal assumption.<br /><br />A truly valid comparison would not compare vehicles of the same liftoff mass. Comparing the same liftoff mass would only be valid if both vehicles were landed on Mars already fuelled for leaving Mars again. But instead the vehicles we are debating about would be empty when they land on Mars, only to be re-fuelled later from resources derived from the Martian environment.<br /><br />So different MAV landed on Mars would still have the same empty mass, which means the different propellant tanks would still have the same volume available. Depending on the type of propellant used, different MAV could have radically different fully fuelled weights and therefore different payloads they could deliver to orbit. That is why propellant density is so important to MAV performance and why ISP is not the only factor.<br /><br />The DunnSpace SSTO propellant study fairly conclusively shows that for vehicles of the same dry mass, hydrogen fuel delivers less payload than hydrocarbon fuels. In this case hydrogen's higher ISP does not compensate for hydrogen's rotten density. <br /><br />

 

[keermalec]

Good point Gunsandrockets, if the performance criterium is payload to LMO per dry surface mass, using LOX/CH4 looks better than in a wet mass comparison. However, LOX/LH2 is still better. Using the figures from my earlier calculations:<br /><br /><br />LOX/LH2 launcher:<br />-----------------------<br />Dry mass: 2.4 tons<br />Payload to LMO: 2.4 tons<br />Performance: 100%<br /><br /><br />LOX/LCH4 launcher:<br />------------------------<br />Dry mass: 2.14 tons<br />Payload to LMO: 1.91 tons<br />Performance: 89%<br /><br /><br />All in all, one still gets over 10% more payload to LMO using LOX/LH2 than LOX/LCH4, for the same dry mass.<br /><br />Another advantage of LOX/LH2: all ressources are realtively easilly availlable on Mars. Producing CH4 requires a source of H2: either imported or produced from local water, but in that case the process is obviously more complex than for producing LOX/LH2 (Hydrolysis + Sabatier reactor versus Hydrolysis only). <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[gunsandrockets]

<However, LOX/LH2 is still better. Using the figures from my earlier calculations: ... ><br /><br />Gives a flawed result. <br /><br />The numbers you use to describe both vehicle are off the mark. So of course your numbers give your hydrogen vehicle the edge -- considering your hydrogen vehicle devotes 24.5% of it's dry mass to propellant tanks and your methane vehicle only devotes 15.9% of it's dry mass to propellant tanks! Your methane vehicle suffers a 1/3 handicap. <br /><br />Vehicles of the same dry mass will have propellant tanks of the same volume (ignoring cryogenic storage issues which make it even worse for hydrogen). But rather than laboriously grind through every propellant combination for tanks of equal volume, it's simpler to refer to the DunnSpace SSTO propellant comparison tables to find results. Hydrogen and methane both come at the lower end of the volumetric performance scale, and hydrogen is even lower than methane.<br /><br />Perhaps though I am jumping to an incorrect conclusion about your methodology. Maybe you did compare equal volumes of propellant. So are you claiming that an equal volume of hydrogen/LOX propellant would outperform an equal volume of methane/LOX propellant? Which of course would mean the DunnSpace SSTO propellant study is full of false results.<br />

 

[gunsandrockets]

<Another advantage of LOX/LH2: all ressources are realtively easilly availlable on Mars. Producing CH4 requires a source of H2: either imported or produced from local water, but in that case the process is obviously more complex than for producing LOX/LH2 (Hydrolysis + Sabatier reactor versus Hydrolysis only).><br /><br />Hydrolysis of water is simpler than the Sabatier process, yes. But that doesn't mean acquiring adequate LOX/LH2 propellant would be easier than acquiring LOX/hydrocarbon propellants.<br /><br />Practical LOX/LH2 propellant production means harvesting ground ice, which means your LOX/LH2 shuttle is only practical flying from a fixed base location. Wheras atmospheric harvesting, which hydrocarbon fuels such as methane would use, could be built into a shuttle allowing flight operations to most places on Mars.<br /><br />

 

[j05h]

Methane can be made with biologics. That's a tougher nut to crack for harvesting hydrogen (but also possible). For Mars, methane may be the "natural" fuel of choice because it is simpler to store or make.<br /><br />One interesting possibility would be to manufacture or ferment methane, then compress it into methane hydrate ices and bound in tough plastic compression membranes for long-term storage. These kind of ices can also feed a certain type of marine worm, but that seems counterproductive. Any thoughts on that form of energy storage? Also of banks of small flywheels for baseline power?<br /><br />Interesting point about atmospheric harvesting. For low-use/emergency ISRU, the equipment is surprisingly light-weight. Not keep-in-your-hopper light, but enough for something like Zubrin's proposed NIMF craft.<br /><br />I'm not sure why you guys are fighting about tank size. The volume needed for hydrogen and methane are almost inverse with their LoX. In a given rocket (if the pumps and engines could do it) the LoX would be used in either tank, flying with LH in the larger tank with less LoX. When using CH4, the LoX would occupy the larger tank because methane is denser than hydrogen. This is vaguely remembered from a while ago, but would provide an interesting engine in a wider space economy. <br /><br />Another point on ISRU - there have been demonstrations of micromachined plates "pressing" liquid O out of a gas (not sure if it was compressed air or what), which could be applied on Mars. Nuclear reactors make harvesting water much, much easier, but it can be done with solar reflectors and sterling engines as well. <br /><br />Josh <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>