Mars 9 tons at a time.

[keermalec]

gunsandrockets I am trying to figure precisely that out. The problem with many low thrust misions is they undertake horribly complicated trajectories and I am still figuring out the SMART-1 path.<br /><br />Of course I will be more than happy to discover it takes less delta-v to get somewhere using low thrust. At the moment though I cannot demonstrate anything. I will be back. <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]

<If we base ourselves on the Mars Global Surveyor mission, the aerobraking took a total of 4 months, with about one atmospheric pass per day. The vehicle used neither aeroshell nor ballute and worked down from an orbital period of 45 hours (263x 54'000 km) to 2 hours (250 x 450 km) with very slight upper atmosphere aerobrakes. Using such a technique would add 4 months to our total mission time but would also increase the useful payload in LMO from 11.5 to 12.9 tons!><br /><br />Ah, but does the manned mission being contemplated require braking into low Mars orbit? Why not just stay in the higher orbit? Consider the different mission of the Mars Global Surveyor versus the mission of an ERV.<br /><br />The Mars Global Surveyor needed to get closer to Mars to enhance the quality of its survey. Wheras an ERV is supposed to eventually depart for Earth so a higher parking orbit is better; less delta-V is required for the TEI burn so less ERV propellant is required.<br /><br /><br /><br />[very quick update!]<br /><br /><...Departure from a 250 x 33793 km orbit... /><br /><br />Whoops! Nevermind. It looks like you are already on top of things! <br /><br /><br /><br />

 

[gunsandrockets]

A have a couple initial questions about your mission design.<br /><br />1) What is the Mars-surface mission-duration?<br /><br />2) I assume from your 21.8 tonne Delta IV payload numbers that the Earth Orbital Rendezvous (EOR) of your payloads takes place in an orbit equivalent to the ISS. Unless the ISS has some key role in your EOR, why penalize yourself the 2 tonnes of extra payload per Delta IV launch?

 

[keermalec]

Actually I was imagining multipass aerobraking only for the unmanned missions: as this increases payload mass in LMO but also increases mission duration by 4 months (which we do not want for inhabited missions).<br /><br />The 250 x 33793 parking orbit comes from Borowski's DRM-3 revision. It is a convenient parking orbit because its period is 1 martian day (ie: always above the landing site) and it dramatically reduces arrival AND departure delta-v as you point out. The down side is: delta-v from the surface is increased from 4.74 km/s (to 250 x 250 km orbit) to about 5.94 km/s (to 250 x 33793 orbit). This reduces payload mass from surface from 13% (to 250 x 250 km orbit) to only 4% (to 250 x 33793 orbit) of launch mass.<br /><br />It seems probable in this case that we would need a two-stage rocket to lift the crew from martian surface to parking orbit. <br /><br />If considering a long term stay, with maybe refuelling of the interplanetary vehicle from surface in situ propellant production, it would be adviseable to park in LMO (250 x 250 km) or have a fuel depot here. But this is going much further than a simple Mars ressearch mission. <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]

<If considering a long term stay, with maybe refuelling of the interplanetary vehicle from surface in situ propellant production, it would be adviseable to park in LMO (250 x 250 km) or have a fuel depot here. But this is going much further than a simple Mars ressearch mission. ><br /><br />With any luck refuelling supplies could come from Deimos or Phobos instead of Mars. Of course we don't know enough yet about Deimos or Phobos to tell if this could even be possible; maybe Deimos and Phobos are mere rocks devoid of any useful volatiles.

 

[keermalec]

Well, thanks to your asking I had to work it out ;-):<br /><br /><blockquote><font class="small">In reply to:</font><hr /><p>1) What is the Mars-surface mission-duration? <p><hr /></p></p></blockquote><br />Earth rotates at 0.986°/day, Mars at 0.524°/day. Earth-Mars trip (inhabited mission) covers 118° and takes 147 days. Therefore 3.7 days after arriving at Mars, Earth and Mars are aligned and at their closest. Mars-Earth return trip (inhabited mission) covers 90° and takes 156 days. Therefore Earth has to be 63.8° behind Mars when return mission is launched. Mission time is therefore (360 - 63.8)/(0.986 - 0.524) = 641 days. From this must be subtracted about 2 days for Mars capture and 2 days for hyperbolic escape + whatever time the MIssion requires in parking orbit before landing and after surface take-off.<br /><br />All in all, mission time on surface would be about 21 months. Do note that we are talking approximations here. If mission is launched in the 2014-2018 period, Earth and Mars are even closer due to Mars' elliptic orbit. 20 years later we have a worse case scenario where Earth and Mars are t their furthest. The durations given are averages.<br /><br /><blockquote><font class="small">In reply to:</font><hr /><p>2) I assume from your 21.8 tonne Delta IV payload numbers that the Earth Orbital Rendezvous (EOR) of your payloads takes place in an orbit eqivalent to the ISS. Unless the ISS has some key role in your EOR, why penalize yourself the 2 tonnes of extra payload per Delta IV launch?<p><hr /></p></p></blockquote><br />You are right, we can certainly add those 2 tons to the departure mass. All payload figures given above can therefore be increased by 9% if departing from a 200 x 200 km orbit. <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>With any luck refuelling supplies could come from Deimos or Phobos instead of Mars. Of course we don't know enough yet about Deimos or Phobos to tell if this could even be possible; maybe Deimos and Phobos are mere rocks devoid of any useful volatiles.<p><hr /></p></p></blockquote><br />Totally agree: to me the first Mars mission should be a water prospection mission to Deimos and/or Phobos. Producing propellant there would substantially reduce the inital mass needed in LEO as propellant for the return trip from Mars would no longer have to be lifted from Earth and sent to Mars. <br /><br />If propellant is produced in Mars orbit we can probably save a whole Delta-IV Heavy launch. <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]

<All payload figures given above can therefore be increased by 9% if departing from a 200 x 200 km orbit.><br /><br />Subject heading changed to match!<br /><br /><All in all, mission time on surface would be about 21 months. /><br /><br />Huh, I thought you might have had a short-stay mission in mind. I didn't want to assume that though, which is why I asked. <br /><br />So here is a followup question about your surface operations -- how big is your surface habitat for the three crew? 21 months is a long time on Mars.<br /><br />

 

[keermalec]

Total surface payload from previous calculations is 23.4 tons + 9% if launching from 200 x 200 km orbit = 25.5 tons, but that includes the propellant production facility.<br /><br />If we need more payload on the martian surface, each DUAL Delta-IV Heavy launch can land an extra 15 tons on the surface.<br /><br />According to Borowski a complete space habitat for 6 persons should mass around 30 tons.<br /><br />For information, the SpaceHab module for the ISS masses 5 tons and provides 28 cubic meters of living space, including life support.<br /><br />Living space may be expanded using local materials (martian concrete, martian soil for food + oxygen production) if this is sufficiently ressearched beforehand.<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>

 

[spacester]

Nice work, Keermalec!<br /><br />I would just emphasize something you've pointed out several times: most of these numbers are based on circular planetary orbits, and in reality the numbers shift quite a bit. <br /><br />It gets very tricky to figure out the timing in the real solar system because that nifty equation using degrees movement per day no longer works. What my software does is move forward in time to find the planets in about the right phase angle position, and then recalculates the planetary radius at the actual dates, then the phase angle is found anew and it searches forward in time again to find the more correct launch date. It continues this iteration until the planetary radius error is less than the SOI distance.<br /><br />Anyway, to all, FWIW, I'm keeping an eye on this thread and all of Keermalec's work, and if the numbers look iffy to me, I speak up. I think I've found just one quibble on this entire thread. I'll have to look again to find it.<br /><br />The point of this post is that Keermalec's numbers look accurate to me, but it is important to realize that the true eccentric orbit of Mars changes things quite a bit.<br /><br />Oh I remember the quibble now: the lower-energy trips to Mars run in a cycle of about 11-1/2 years IIRC. A number of 20 years was stated, implying a 40 year cycle, and that for sure is not correct. This cycle is tricky to define, BTW.<br /><br />We've just past the lowest dV opportunities in 2005 and 2007, and they will be low again in ~2016-2018. <div class="Discussion_UserSignature"> </div>

 

[j05h]

<i>to me the first Mars mission should be a water prospection mission to Deimos and/or Phobos. </i><br /><br />"Phobos First" was the impetus to my older long-running thread in this forum, "Private Mars Missions". I would call that thread and this one complimentary - they deal with related but different aspects of settling Mars. <br /><br />I'd be interested in starting a new Phobos First thread if you guys will post to it - but would rather keep discussion of orbital ops outside this thread. (except for staging/aerobraking concepts)<br /><br />Propellant production would be the #2 goal of any commercial Mars operation - right after producing water. Methane can be burned in surface or orbital fuel cells, in rockets with local LOx or as feedstock for making other hydrocarbons. This is crucial on the surface or on the Martian Moons - and we'll simply follow the volatiles to the best source.<br /><br />In terms of launch savings, early ISRU propellant will save 1/4 to 1/2 the outbound mass of a mission. For "Mars 9 tons at a time" I've assumed ISRU propellant from the start - it can't be done in any reasonable manner without local volatiles. <br /><br />Your numbers for a high-energy stage are interesting, thanks for contributing them.<br /><br />Would you post to a new "Phobos First" thread?<br /><br />Josh<br /> <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[keermalec]

Thanks spacester, for pointing out the 11.5 year cycle. I knew there was a cycle but was unaware of the real figure. So 2016-2018 is next best shot? Lets get to work!<br /><br />As you know, I only consider circular orbits because that allows us to more quickly develope a feel for what mass and time requirements may exist for any given scenario. I see my calculations (and calculator) as a handy tool for establishing the basic concept of a mission by non-specialists. Once a concept is validated, then a date can be chosen and the real position of Mars in its elliptical orbit can be taken into consideration for a more detailled mission scenario. <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]

I would certainly post to to a new "Phobos First" thread. But why Phobos? Deimos has a higher orbit and so is more easilly accessible. According to some it has the same chance of holding water as Phobos. More on the new thread... <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>low-thrust missions I've read of only thrust about half the time and at orbit periapsis<p><hr /></p></p></blockquote><br />Many low-thrust missions actually do apply thrust only at periapsis, as you pointed out. This is the case of the SMART-1 mission to the Moon, but not the case of the DAWN mission to the asteroids, which thrusts pretty much most of the time.<br /><br />Both trajectories are quite complicated. IN SMART-1, thrusts are applied only at periapsis for 1.5 to 8 hours at a time. Gravity losses are disregarded but the vehicle does get to the Moon with a delta-v close to that indicated by Stockermanns for continuous low thrust.<br /><br />Total SMART-1 delta-v from an Earth centered 742x36'000 km orbit to a lunar-centered 468x2'878 km orbit: 4.06 km/s, 82.2 kg fuel, 440 days<br /><br />Delta-v needed with high thrust hohmann transfer and lunar capture: 1.73 + 0.66 = 2.4 km/s, 189 kg fuel (assuming storable bipropellant), 4 days.<br /><br /><br />I worked out a different trajectory to get to the same final orbit, but with as much continuous thrust as possible:<br /><br />1. Circularise from 742 x 36016 to 36016 x 36016 by thrusting at apoapsis for 8 hours each time: 1.44 km/s, 31.5 kg fuel, 237 days<br /><br />2. Transfer from 36016 x 36016 to 312000 x 312000 (Moon SOI) by continuous thrust: 1.95 km/s, 38.4 kg fuel, 97 days<br /><br />3. TRansfer from 6000 x 64400 km (lunar capture orbit) to final 468 x 2878 km orbit by 8 hour thrusts at apoapsis: 0.66 km/s, 12 kg fuel, 30 days<br /><br />TOTAL: 4.05 km/s, 81.9 kg fuel, 364 days.<br /><br /><br />IE: exactly the same in terms of delta-v and fuel as the real mission, but I seem to get there 2 months earlier .<br /><br />This would be due to the continuous thrust part, which loses less time simply coasting than the real mission. I do not know why the mission team chose to apply only short thrusts. Maybe the engine was not designed to sustain 97 days of continuous thrust? <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]

Well the SMART-1 numbers certainly check out. Low thrust propulsion does seem to suffer from about a 1/2 delta-V penalty after all. How depressing.<br /><br />That means to achieve the same performance as high-thrust propulsion, low thrust propulsioin needs an ISP about twice as high. And that means doubling the already slow travel times of typical electric propulsion. Ugh.<br /><br />I suppose there isn't a lot of practical difference between unmanned cargo taking four years as opposed to two years to reach Mars, but everytime the cold-equations seems to raise another difficulty to space-travel, the adventure just seems to recede some more.

 

[solarspot]

Given the path this thread has taken since I last logged into SDC, I'm not sure this is still relevant. However I recently checked the earth-escape payload capacities of Atlas V 551 and Ariane V ECB, and both of those vehicles appear to have TMI payloads of just over 5 tons at a time, not the nearly 9 tons reported for the Delta IV. Aswell, I'm fairly sure the Proton 8k82M Briz M has a TMI payload slightly below that, due to it's use of a storable propellant fourth stage.<br /><br />Hopefully this will not be too limiting, seeing as the last 2 pages at least have been about low-thrust propulsion from LEO...

 

[keermalec]

Gunsandrockets, I wouldn't be too depressed. The SMART-1 analysis has opened my eyes on the possibility of using low thrust for multiple-burn orbit elevation, which I thought wouldn't work due to gravity losses.<br /><br />Even though continuous thrust needs to attain a higher delta-v than single-burn (or multiple-burn) high thrust, the final payload delivered is still larger and the technology exists, now.<br /><br />Multiple-burn orbit elevation adds yet another possibility: that of increasing yet some more the useful payload at destination by increasing trip time some more too. As posted in the Phobos First thread, a combined low-thrust approach (multiple-burn orbit elevation to leave Earth, continuous low thrust from Earth SOI to Mars SOI, and multiple-pass aerobraking at Mars) can theoretically deliver 50% of LEO mass to LMO in around 24 months. To me that is quite a feat!<br /><br />Note that continuous low thrust was only science fiction 10 years ago when Zubrin wrote his Mars Direct scheme. Today sending unmanned payloads ahead of time can significantly reduce the cost of getting humans to Mars. <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>

 

[j05h]

<i>> Even though continuous thrust needs to attain a higher delta-v than single-burn (or multiple-burn) high thrust, the final payload delivered is still larger and the technology exists, now.</i><br /><br />The continuous-thrust stage you are talking about doesn't exist yet. The Delta IV H as configured can send a certain amount to Mars - which was the orginal exercise in this thread. It has certain strengths, including the ability to focus only on payload development. The only electric/ion engine that is up to the task of Mars transit would be VASIMR, and that is not a sure thing (funding and power source). <br /><br /><i>> a combined low-thrust approach (multiple-burn orbit elevation to leave Earth, continuous low thrust from Earth SOI to Mars SOI, and multiple-pass aerobraking at Mars) can theoretically deliver 50% of LEO mass to LMO in around 24 months.</i><br /><br />Delta IV H can put about 25t in LEO, which means 12.5t at LMO via ion drive. That is spitting distance from it's direct-throw capability to TMI and LMO. Also, does the ion-engine count against payload or third-stage mass? <br /><br />Direct-throw simplifies things immensely. <br /><br /><i>> Note that continuous low thrust was only science fiction 10 years ago when Zubrin wrote his Mars Direct scheme. Today sending unmanned payloads ahead of time can significantly reduce the cost of getting humans to Mars.</i><br /><br />This whole thread has been about pre-positioning payloads (all my Mars threads have that theme). Automatic site preparation and cached supplies make all kinds of sense - especially if thinking from space-downward creates global Mars access. Updating teh Zubrin's idea of direct-throw payloads with a different launch scheme, modular payloads and a more node/system-based plan still seems a viable approach. Using a common "cargo container" would allow the lesser ELVs and ion-tugs to be part of the fun - especially useful for shipping bulk items like food and space-suits. There is nice sy <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[keermalec]

Josh, I agree with most of your comments except the first one, where you say continuous low thrust does not exist yet.<br /><br />it does: I use only existing technology in my calculations and in this case the engine is the NEXT ion drive, recently developed by NASA: an upgrade of the NSTAR ion drive which flew the DeepSpace 1 probe. NEXT has an ISP of 3940 seconds for a mass of 48 kg and a thrust of 2 Newtons.<br /><br />VASIMR technology is far from being practical yet and I have been able to find no mass figures.<br /><br />The problem with "9 tons at a time" is, as we have seen before, that it allows only sending a single person with life support on a 9 month trip.<br /><br />Using the same Delta IV Heavy, it makes much more sense to send first the inert payloads, using orbital assembly (one 23.8 ton payload module connected to one 23.8 ton ion propulsion module) and, 2 years later, to send the fast inhabited payload (one 23.8 ton habitat for 3 persons + 2 x 23.8 ton propulsion stages) for a fast 5-month transfer. <br /><br />Remember we need to reduce inhabited trip time to reduce the effects of zero-g and galactic cosmic rays on the crew. <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]

All the prepositioned cargo can be sent 9t at a time. The manned flights, years later, could use something bigger. The current lack of whatever the 'something bigger' is need not delay getting the supplies and robot factories on their way now.

 

[gunsandrockets]

<Multiple-burn orbit elevation adds yet another possibility: that of increasing yet some more the useful payload at destination by increasing trip time some more too. As posted in the Phobos First thread, a combined low-thrust approach (multiple-burn orbit elevation to leave Earth, continuous low thrust from Earth SOI to Mars SOI, and multiple-pass aerobraking at Mars) can theoretically deliver 50% of LEO mass to LMO in around 24 months. To me that is quite a feat!><br /><br />Very impressive. But are you sure that such a high efficiency travel method would only take 24 months travel time? I think the MRO spent 6 months alone just to lower it's orbit via aerobraking. <br /><br />Here is an interesting link somewhat on topic...<br /><br />http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/browse.pl?1999/TM-1999-209646.html

 

[publiusr]

9 tons at a time my foot. Lets just get rid of supertankers and ship oil using 10,000 bass boats while you are at it.

 

[nuaetius]

I feel like the dumbest guy in the room in this forum but here is my idea anyway <img src="/images/icons/smile.gif" />. Considering that the warmest daytime temp on Mars is never above water’s freezing point why not build the whole colony by melting holes into a glacier? Here is how I would do it.<br /><br />-Use the current assets in orbit to find glaciers at least 50 meters think, and in a valley to protect if from the worst of the wind storms and radiation from the sides.<br /><br />-1st mission would send a team of rovers to three best locations found by the satellites. These rovers only purpose would be to confirm the depth and “juckiness†of the ice and determine which site would be the best suited to our needs. The two rovers not selected would be leased to Universities for research. The rover chosen would spend the rest of its functional life mapping and imaging ever rock and hill in the chosen area<br />-Next would be a set of cargo drops. One would be a power station, either nuclear if we have be political will, if not solar. The other would be an industrial rover and a melting cap<br /><br />-The Rover would have a crane on top, an ultrasound scanner on the bottom and a blade on the front and be about the size of a car. The rover would use a more powerful ultrasound scanner to confirm that an area has very little rock suspended in the ice below it. It would then clear the area to be melted of rocks and fines to allow the melting cap to do its work.<br />-The melting cap would look like a 3 meter across flattened bad mitten birdie. The “ball†would be a microwave generator and the skirt would be to hold in the heat. The Rover would attach the generator to the melting cap. Once connected the melting cap would slowly increase the air temp under the cap. Slowly the cap would cause the water underneath the cap to melt causing the cap to sink. The water would be released in the form of warm water vapor into the atmosphere (snowing on Mars anyone?)

 

[keermalec]

<blockquote><font class="small">In reply to:</font><hr /><p>are you sure that such a high efficiency travel method would only take 24 months travel time? I think the MRO spent 6 months alone just to lower it's orbit via aerobraking. <br /><p><hr /></p></p></blockquote><br />Once more, gunsandrockets, good point. After careful checking it seems I underestimated the Earth orbit raising segment. SMART-1 leaves from an elliptical GTO orbit and we leave from LEO. The early part of orbit raising takes a long long time:<br /><br />Assumed ion-drive vehicle characteristics:<br /><br />Drive type: NEXT<br />Solar panel type: DS-1 concentrator arrays<br />Average thrust: 0.000020 Gs <br />Largest distance from sun: 1.52 AU <br />Drive isp: 3940 seconds <br />Propellant: Xenon <br />Initial mass in LEO: 23.8 tons<br /><br />Multiple thrust orbit elevation from LEO to SOI: 3.9 km/s, 690 days (based on SMART-1 data, assuming thrust for 1/3rd of orbit period at each periapse)<br /><br />Continuous thrust from Earth SOI to Mars SOI: 5.6 km/s, 330 days (hypothetical, according to Stockmann's equation)<br /><br />Multiple pass orbit lowering at Mars: 1.0 km/s propulsive, 180 days (from Mars Global Surveyor data)<br /><br /><br />** Total propulsive delta-v: 10.5 km/s, 1'200 days or 40 months **<br /><br /><br />Probable ship mass breakdown:<br /><br />Propellant: 5.96 tons <br />Structure: 0.95 tons <br />Tanks: 0.30 tons <br />Thrusters: 0.55 tons <br />Solar panels: 3.26 tons <br />RCS + RCS propellant: 0.48 tons <br />Avionics: 0.24 tons <br />Payload: 12.07 tons (51% of initial mass in LEO)<br />Total mass: 23.80 tons <br /><br /><br />-- />If thruster capacity is increased to 0.00003 Gs, trip time is reduced to 29 months but final payload is reduced to 10.2 tons (43% of IMLEO).<br /><br />-- /> If thruster capacity is increased to 0.00004 Gs, trip time is reduced to 23 months but final payload is reduced to 8.3 tons (35% of IMLEO).<br /><br />-- /> If thruster capacity is increased to 0.00005 Gs, trip time is re <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]

The MRO had no additional structures to aid in aerobraking by dispersing the heat and load. It had to do it gently over a long time so it would not get damaged. With some sort of robust shield, or even a small ballute, you can do it much faster than that.