Mars 9 tons at a time.

[keermalec]

Zero boiloff of cryogenic propellants is now possible:<br /><br />http://www.grc.nasa.gov/WWW/RT1998/5000/5870plachta.html<br /><br />Therefore LOX and LH can be stored for long periods of time. That means liquid methane can be stored for even longer than liquid hydrogen, as its boiling point is much higher. <br /><br />What do you mean by:<blockquote><font class="small">In reply to:</font><hr /><p>I doubt a single engine could run on both <p><hr /></p></p></blockquote> <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]

I was assuming that since different chemicals will have different mixing ratios and produce different chamber pressures that a single engine could not be made fuel-flexible; it would require different injectors, or something. If the differences can just be handled by adjusting external valves I would be glad to hear it.<br />

 

[gunsandrockets]

<Zero boiloff of cryogenic propellants is now possible:><br /><br />I was excited at first to read this, but then I saw your link was to an old story about a storage experiment that used active cooling. No one is shocked that active cooling could produce zero boiloff.<br /><br />Now if you could get zero boiloff with a purely passive lightweight system, that would be impressive.<br />

 

[keermalec]

According to Borowski, the refrigeration system weighs 4.57 kg per Watt of refrigeration or per 75 kWe (it takes 75 Watts of electricity to produce 1 Watt of refrigeration).<br /><br />51 tons of LH2 requires 15 kWe of power to prevent boiloff. Therefore the refrigeration system weighs 0.91 tons. Solar panels for 15 kWe at 1.52 AU from the sun will weigh 0.39 tons, considering Solar concentrator arrays as in Deep Space 1.<br /><br />Therefore preventing boiloff from an LH2 tank adds only 2.5% to the propellant mass, to which must be added about 2.3% of additional insulation and micro-meteoroid shielding for long-term storage. That is a total of 4.8% or 48kg per ton of liquid Hydrogen. Whoever would want to do without? <br /><br />I strongly believe that zero boiloff with purely passive techniques is impossible. You can only slow down heat transfer; you can't stop 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>

 

[thereiwas]

I added a calculation of the maximum aerodynamic force against the ballute. For a nominal reentry it comes to 229 N per square meter, or just under 5 lbf per square foot. The trick is that the ballute has to support that force over a diameter of nearly 100 feet, without substantially changing shape. Perhaps the totally enclosed ballute with internal pressurization is a good idea after all.

 

[j05h]

The self-pressurized designs allow for much more pressure to be applied to the reentry "face" of the ballute. <br /><br />Something like this shape could be pressurized an arbitrary amount to maintain or adjust it's shape. Perhaps it could be tailored to inflate to the right amount for that landing's atmospheric conditions. <br /><br />http://www.ilcdover.com/products/aerospace_defense/entrylanding/brakes.htm<br /><br />Also, considering how light ballute material is, the added "flap" looping back to the central module doesn't add much mass. <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[gunsandrockets]

<The Delta IV Heavy can put 17,600 lbs into Trans Mars Injection.><br /><br />Could you please post a link to the data? I'm having a little trouble finding it.

 

[j05h]

Numbers are on Page 52:<br /><br />http://www.boeing.com/defense-space/space/delta/kits/d310_d4heavy_demo.pdf<br /><br />I linked to this about 10 pages back in the thread. 8000kg or 17600lbs to TMI<br /><br />Josh <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[gunsandrockets]

Thanx!

 

[keermalec]

Interesting illustration on page 51 of the document u linked to Josh:<br /><br />With a modified launch pad, an upgraded Delta-IV Heavy could lift as much as 47 tons to LEO, that would be about 14.5 tons on a trans-Mars injection course (or 18 on a trans-lunar), with a fairing diameter of 6.5 m.<br /><br />With a new pad/infrastructure the new Delta (V?) could lift up to 95 tons to LEO, or 29 tons to TMI (or 36 tons on a trans-lunar) with a fairing diameter of 6.5 m.<br /><br />This is almost the planned performance of the Ares V, which has a 130 ton to LEO capacity (40 tons TMI, 53 tons TLI). <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]

Yes, Delta or Atlas upgrades are very interesting. They are achievable, relatively near-term modifications to existing rockets. They provide a growth path to heavy-lift that doesn't fly like Frankensteins' monster. <br /><br />Just the increase in fairing size would allow for expanded types of missions. <br /><br />But... none of these modifications will happen "only" for Mars-lift. There needs to be a larger market that needs 50t to LEO to make that happen. As it is, I think we can do a lot with what we've got. <br /><br />Josh <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[thereiwas]

The required internal pressure for the ballute varies quite a bit over time. This graph shows the pressure in Pascals for the ballute to keep its shape over the time of re-entry. Horizontal scale is in seconds from simulation start. The general increase to the right represents the need to balance the increasing outside air pressure so it does not deflate on the way down. (Surface pressure is about 660 Pa). The lump between 600-700 seconds represents the overpressure required to resist the aerodynamic forces during the major deceleration that happens between 50km and 100km in altitude. The vehicle goes subsonic at around 700 sec.<br /><br />The discontinuity near the end is when the ballute is jettisoned at engine start, at around 1km alt.<br />

 

[keermalec]

You're right, Josh, I don't think there is a market for 47 tons to LEO just yet.<br /><br />A lunar base though would certainly create that need.<br /><br />Concerning our initial subject though, of "Mars 9 tons at a time", sending 14.5 tons to Mars makes a big difference as compared with 8 tons. This definetely allows for a crew of 2-3 persons with life support etc. <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]

<The Delta IV Heavy can put 17,600 lbs into Trans Mars Injection [TMI].> <br /><br />Therefore the first thing to consider are simple ways to increase that payload.<br /><br />In that spirit I played around with some numbers to see how a solar-electric-propulsion (SEP) system third-stage rocket could improve that payload. So instead of the LOX/LH2 Delta IV upperstage blasting an 8,000 kg payload directly towards Mars, it places a third-stage of 23,800 kg into LEO.<br /><br />Inspired by SMART-1 , I equipped my third-stage with with 4 Hall-effect plasma thrusters running at a low ISP of about 1,600 seconds. The T-220 thruster seemed appropriate for the job and each one needs up to 20,000 watts. I estimate the combined mass of the third-stage thrusters at about 250 kg.<br /><br /><br /><br />To provide 80 kWe for the thrusters, I add solar panels based on the SCARLET II technology used by Deep Space 1 . I conservatively estimate the mass of the third-stage Solar-Power-System (SPS) to be about 2,000 kg.<br /><br />The final result of all this is a TMI payload of about 17,000 kg (not including the still useful SPS), which is more than twice the payload of a regular Delta IV heavy. I'm kind of surprised the SEP advantage isn't even greater. Despite the crudeness of my calculations, the payload number is relatively insensitive to significant changes of delta-V and ISP; the T-220 can operate at considerably higher ISP than the number I used (at the penalty of lower thrust to

 

[usn_skwerl]

It's probably a stupid idea, and I don't know if its been mentioned, I didnt see it in a search. <br /><br />SRB weighs about 1.3 million lbs loaded, makes "thrust" of 2.8 million lbs. So T-to-W ratio of ~2:1 and increasing over 2 mins.<br /><br />Why not send a 2-pack, or 4-pack of SRB's up? they're already somewhat equipped with onboard (GPS) navigation. Would it be complex for 2 (or more) boosters to fire from the pad, launch 4 (or more)boosters strapped onto it, but not firing? maybe some segments? And equip it with a star stracker system?<br /><br />That's existing technology, they're reliable, but I dont know if SRB's can be fired in vacuum. If they can't, forget I mentioned it, but with some amount of modification, without spending a ridiculous amount, could they realistically be used to send the payload to mars?<br /><br />I'm seeing something like that used with the D-IV H config, with unfired strap-ons encircling the active ones, using only the actives to get the bundle into orbit. A rondezvous with the SRB's, an adapter, and the mars lander/equipment would result in a good bit of cargo able to go to Mars. As the boosters are used and depleted, they could be dumped, or used for something else with purpose...i dunno, just an idea. <div class="Discussion_UserSignature"> </div>

 

[keermalec]

usn_skwerl, SRBs are cheap and their thrust is high, but their ISP is low (something like 250s). In space, ISP counts more than thrust.<br /><br />Your basic idea is good: launch payload into LEO, then launch the propulsion unit. Thats 23.8 tons of payload and 23.8 tons for the propulsion unit. I would like to jump onto your idea but optimise it somewhat. SRBs have such a low ISP that u would need 93 tons of these to accelerate 23.8 tons to TMI...<br /><br />However, 23.8 tons of propellant and thruster COULD definetely get your 23.8 ton payload to Mars if:<br /><br />- You use LOX/LH2, not SRBs<br />- You launch from a 30'000 x 30'000 km orbit to reduce departure delta-v. Launching from LEO requires more than 23.8 tons for the propulsion unit. <br /><br />Therefore you need a way of lifting the ship to 30'000 km first (ion thruster?).<br /><br />However, if we assume part of the payload 23.8 tons is propellant and the total ship masses 47.2 (composed of two parts), the useful payload sent to Mars using LOX/LH2 from LEO is about 13.8 tons.<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]

Gunsandrockets, I get less than 17 tons to TMI using a 4'000s ISP NEXT ion thruster (an evolution of the NSTAR drive used on DeepSpace 1).<br /><br />I believe the reason my result is different is because I use a different delta-v calculation:<br /><br />For continuous thrust the delta-v is more than for high thrust. According to Stockermans, the delta-v for continuous low thrust can be summarised as the difference in orbital velocities between departure and arrival orbits.<br /><br />For example, to go from LEO to infinity (for interplanetary transfer) requires V(LEO) - 0 = Velocity at LEO = 7.79 km/s using continuous low thrust, and 3.5 km/s when using high thrust.<br /><br />To go from Earth to Mars requires V(Earth) - V(Mars) = 29.78 - 24.15 = 5.63 Km/s instead of 2.94 + 2.63 = 5.57 km/s as in the case of a high thrust Hohmann transfer.<br /><br />So total delta-v accoplished by ion drive would be 7.79 (escape Earth) + 5.63 (transfer to Mars) = 13.42 km/s.<br /><br />To this must be added 3.43 km/s for slow-down to Low Mars Orbit (an aerobrake shield would weigh 3.4 tons and the extra fuel to do a propulsive slowdown only 1.4 tons).<br /><br />Total delta-v therefore = 16.85 km/s for continuous low-thrust from LEO to LMO.<br /><br />Using the rocket equation, 23.8 tons in LEO becomes 15.5 tons in LMO.<br /><br />Duration of the transfer depends on thrust. Assuming a thrust of 0.00001 Gs, the trip would take 1990 days, or over 5 years. <br /><br />Mass breakdown could be approximately:<br /><br /><br />Propellant: 8.75 tons <br />Structure: 0.95 tons <br />Tanks: 0.82 tons <br />Thrusters: 0.26 tons <br />Solar panels: 0.76 tons <br />RCS + RCS proopellant: 0.48 tons <br />Avionics: 0.24 tons <br />Payload: 11.54 tons <br />Total mass: 23.80 tons <br /> <br /> <br />Time of flight: 1990 days <br /><br /><br />If we were to incr <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]

K- <br />Those are very interesting numbers for low-thrust flights. It brings up two issues. First is building gear that can survive that long in space (esp. Van Allen Belts) and still function on Mars. Second is that building a low-thrust, high-mass cargo tug means fully developing a second transit system, essentially doubling those development costs. 5 years in transit just seems long, too.<br /><br />Thoughts, esp. on it's practicality? <br /><br />josh <div class="Discussion_UserSignature"> <div align="center"><em>We need a first generation of pioneers.</em><br /></div> </div>

 

[thereiwas]

"an aerobrake shield would weigh 3.4 tons"<br /><br />That surely depends on what it is made out of.

 

[gunsandrockets]

Wait a moment, slow down.<br /><br />I never said the payload was to LMO. The 17,000 kg payload I quoted for an SEP third-stage is for Trans-Mars-Injection only ; for the purpose of direct comparison to the Trans-Mars-Injection payload of a standard Delta IV heavy rocket.<br /><br />And are your delta-V numbers accurate? For example the 5.57 km/s figure you quote for high thrust propulsion transfer to Mars does not match the numbers I find. Zubrin's book says it's only about 4 km/s.<br /><br />Now I freely admit my math skills are terrible and I imagine you derived your delta-V results by cranking the numbers through your reference equations. But do me a favor as a check on your technique, does a normal Delta IV heavy still have payload of 8,000 kg TMI when using your methodology? If not there may be something wrong with your results.<br /><br />And I wonder about the claim that low-thrust propulsion is so inefficient. I've asked around and the best answer I have found so far is that low-thrust does not suffer an extraordinary penalty in delta-V. One factor that might make a difference is your link assumes constant thrust, but low-thrust missions I've read of only thrust about half the time and at orbit periapsis. But there is a way you could double check conflicting assumptions.<br /><br />Does your methodolgy derive results that matches the specifications of the low-thrust DAWN mission? That is another way to double check the assumptions. Dawn carries a total of 450 kg of propellent for the three NSTAR ion engines, and has a loaded mass of 1,200 kg. So DAWN has a mass ratio of 1.6 and an exhaust velocity of 30 km/s, and NASA claims a total delta-V capability of 10 km/s. Does that agree with your calculations?<br /><br />If you want to go even further you can check your method vs the numbers published for the SMART-1 lunar orbiter mission too. The SMART-1 traveled from GTO to LLO using it's low-thrust electric propulsion engine. The mass ratio of the SMART-1 was

 

[keermalec]

Hi Gunsandrockets, first of all I must say I also find close to 17 tons at end of TMI using continuous low-thrust, like you do. 15.5 tons in LMO means 17 tons at Mars SOI (the limit between the interplanetary and orbital portions of the journey).<br /><br />I did the following check though as you suggested.<br /><br /><blockquote><font class="small">In reply to:</font><hr /><p>And are your delta-V numbers accurate? For example the 5.57 km/s figure you quote for high thrust propulsion transfer to Mars does not match the numbers I find. Zubrin's book says it's only about 4 km/s.<br /><p><hr /></p></p></blockquote><br />5.57 km/s includes departure delta-v (2.94 km/s) and arrival delta-v (2.63 km/s) for a Hohmann transfer from Earth (1 AU) to Mars (1.52 AU), assuming circular orbits.<br /><br />In reality, as the orbit of Mars has a relatively high eccentricity, delta-v may be higher or lower, depending on the conjunction. But the average delta-v will still be 2.94 + 2.63 = 5.57 km/s.<br /><br />2.94 = difference between Earth orbital velocity and orbital velocity of transfer ellipse at periapse = 32.72 - 29.78 km/s.<br /><br />2.63 = difference between Mars orbital velocity and orbital velocity of transfer ellipse at apoapse = 26.78 - 24.15 km/s.<br /><br />To attain an interplanetary velocity of 2.94 km/s one must leave LEO at 11.4 km/s (see equation below). As orbital velocity in LEO is already 7.8 km/s, one only requires an additional 3.6 km/s to attain the 11.4 needed to launch onto a TMI. From the hyperbolic trajectory equations:<br /><br /><br />sm = -gp / (vi ^2)<br /><br />sm = semi-major axis of hyperbolic trajectory (km)<br />gp = gravitational parameter for Earth (398600 km2/s3)<br />vi= required velocity at infinity (2.94 km/s)<br /><br />vp = sqrt(2 * <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>Those are very interesting numbers for low-thrust flights. It brings up two issues. First is building gear that can survive that long in space (esp. Van Allen Belts) and still function on Mars. Second is that building a low-thrust, high-mass cargo tug means fully developing a second transit system, essentially doubling those development costs. 5 years in transit just seems long, too. <br /><p><hr /></p></p></blockquote><br />Spending time in space is not necessarily a problem for inert cargo: note how deep space probes like Cassini spent 7 years in space before functioning perfectly on arrival. The Van allen Belts are in medium Earth orbit so going through them should take 1-2% of the total trip time at most. Provided the equipement is designed to withstand such exposure, it should be ok. Once again, all ion-driven deep space probes have passed the van allen belts without trouble.<br /><br />Concerning developement costs, the ion drive is an existing technology. I read somewhere that building a new NSTAR drive should cost "only" about 2-4 million.USD. <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>"an aerobrake shield would weigh 3.4 tons" <br /><br />That surely depends on what it is made out of. <p><hr /></p></p></blockquote><br />I am basing myself on the Viking aerobrake shield (25% of total mass) and Borowski's shield mass formula, which would make it around 17% for a 60-ton craft. I just assume our 17-ton craft is somewhere in between the two, at 20%. <br /><br />Of course, a ballute would be lighter, but I do not have enough information on its mass and storage requirements to factor it in. If it massed less than 1.4 tons then of course it would be preferable to propulsive capture.<br /><br />Actually, now that I think of it, an aeroshell is not necessary. The flight being non-inhabited, adding in several months for a slow, multipass aerobrake is perfectly feasible, and this requires neither aeroshell nor ballute.<br /><br />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.<br /><br />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! <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]

So my favorite Mars Mission scenario, using only Delta-IV Heavy launchers would be:<br /><br /><br />Inert cargo to LMO by ion drive (6 launches, 24-28 month trip):<br />-------------------------------------------------------------------------------<br />21.8-ton First Mars Lander/Lifter (carries 9.6 tons of surface payload + 4.2 tons of ascent vehicle + 3-ton aeroshell) + 21.8-ton ion thruster stage <br />21.8-ton Second Mars Lander (carries 13.8 tons of surface payload including room for 3 astronauts + 3-ton aeroshell) + 21.8-ton ion thruster stage<br />21.8-ton fully-fuelled LOX/LH2 Return Trip Thruster + 21.8-ton ion thruster stage<br /><br /><br />Inhabited trip to Mars (3 launches, 4.9 month trip):<br />----------------------------------------------------------------<br />4 Delta-IV Heavy launches to LEO. Ship is composed of one 21.8-ton habitat with aerobrake shield and a 65.4-ton thruster stage (3 Delta-IV Heavy launches assembled in LEO). Total Delta-v using LOX/LH2 = 4.2 km/s. Net payload after subtracting aerobrake shield and others = 16 tons = habitat for 3 persons. C3 = 17.6, Earth-Mars angle = 118°. Departure from a 200 x 200 km orbit. Arrival at a 250 x 33793 km orbit at Mars.<br /><br /><br />Return trip from Mars (5.2 month trip):<br />-------------------------------------------------<br />Ship is composed of the initial 21.8-ton habitat with aerobrake shield and the 21.8-ton Return Thruster Stage. Total Delta-v using LOX/LH2 = 2.2 km/s. C3 = 22 km2/s2, Earth-Mars angle = 90°. Departure from a 250 x 33793 km orbit.<br /><br /><br />The inert cargo modules are launched on a 24-month trip to Mars. The Return Trip Thruster Module aerobrakes doing multiple passes lasting 4 extra months. The First Mars Lander/Lifter module aerobrakes and lands directly, using a parachute. The Second Mars Lander aerobrakes and waits in a 250 x 33793 parking orbit (period = 1 martian day) for the inhabited mission.<br /><br />Once the First Mars Lander/Lifter has landed it is remotely contro <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]

<Departure delta-v is therefore 3.6 km/s, to which must be added about 0.3 km/s in gravity losses, and mid-course corrections of about 0.1 km/s. The arrival delta-v is zero if aerocapturing directly into the martian atmosphere. This is Zubrin's Mars Direct scheme, and that is why he indicates 4.0 km/s.><br /><br />Now that you have gone into all the nitty-gritty details of the trajectory, the delta-V numbers you used before are more sensable. In your earlier post I thought you were claiming that LEO to Mars departure delta-V was 5.57 km/s! -- which sounded way too high.<br /><br />Despite my math skills I don't have too much trouble dealing with simpler things such as the rocket equation. However, what does throw me is orbital mechanics which I am just beginning to learn. Which is why I was so concerned with the very large delta-V numbers you were tossing around regarding low-thrust propulsion trajectories.<br /><br />That's why I suggested checking your numbers versus some real low-thrust missions. It's easy enough to figure out the maximum delta-V capacity of low-thrust spacecraft such as SMART-1 or DAWN using the rocket equation. The question is does your method of calculating delta-V needs for a low-thrust trajectory mission show results which match the actual missions these spacecraft were built for?<br /><br />In other words does your trajectory calculations match the actual mission that SMART-1 flew? Or do your trajectory calculations show a much larger delta-V requirement than the actual SMART-1 spacecraft was capable of performing? <br /><br />