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#776 Re: Interplanetary transportation » Mars Semi-Direct with Falcon » 2012-04-12 00:01:56
3.8 km/sec and 6.4 km/sec are the numbers for Hohmann (minimum energy) transfers, which take 8-9 months each way and are not free return trajectories. If you go a bit faster (4.3 km/sec outbound and about 7 km/sec inbound) you make the trip in 6 months instead and you put yourself into a 24-month solar orbit, so if you miss Mars on the way out (say, because of damage to the craft) you come back to Earth anyway. You also cut down on consumables.
Zubrin calculates that a Falcon Heavy putting 53 metric tonnes in LEO, including a hydrogen-oxygen stage, can push 17 tonnes to trans-Mars injection, put 14 metric tonnes into orbit, and 11 metric tonnes on the surface. I suppose those numbers represent 3 metric tonnes of heat shield and retrorocket fuel and 3 tonnes of TMI staging. If you can protect the stage with your heat shield, you can probably land it on Mars as well, but that probably means a bigger heat shield and more landing fuel, so your 11 tonnes of cargo diminishes to maybe 10 or 9. As we have noted elsewhere, the "Red Falcon" project manages to put even less on the surface, probably because they are using bigger margins and RP1/LOX propellant all the way.
Nine or 10 tonnes may be enough for a 1-way trip for 1 or 2 people, but it'd be hard to include enough other stuff to make it a round trip. If you already have solar panels and a robotically drilled water well, it might be barely enough, but I doubt it. You'd do better by launching two Falcon Heavies, one with a dedicated LH2/LOX stage to which the earlier launch would dock. Both Falcon launches would need to include propulsion stages, but the one with the cargo would have a much smaller stage and your payload mass doubles; a bit more than doubles, probably. If you can land 22 or even 24 tonnes on the surface, you're getting close to the Mars Direct ERV (28.6 tonnes).
Thanks for the info. My plan certainly has slim margins.
My numbers for landing on Mars are actually similar to Zubrin's if you use the larger numbers, for instance 4.3 km/s for outbound, for the delta-V for the shorter transit times. You would also get about 17 mT, instead of 20, for the TMI mass for my architecture using the larger delta-V requirement. Now consider that I am fully using aerobraking for landing on the surface so no or minimal propellant is used for landing after arrival at Mars.
I also believe that production of hydrolox propellant on Mars will be easier than expected due to nearly pure water ice close to the surface on Mars. Separating hydrogen and oxygen by electrolysis is a simple process. It even appears in children's science toys.
Some exciting recent discoveries show that water ice is close to the surface even in the mid-latitudes on Mars:
Water Ice Exposed in Mars Craters.
by Andrea Thompson
Date: 24 September 2009 Time: 02:18 PM ET
Byrne told SPACE.com that it was surprising to the team to find the bluish ice, though "in retrospect maybe it shouldn't have been." Scientists knew of the existence of underground ice and had been monitoring craters as they formed, but "I guess we didn't put the two together," he said.
Several of the craters were also near the landing site of the Viking Lander 2. Viking also looked for water ice on Mars, but was only able to dig down about 6 inches (15 cm) below the surface ? about 4 inches (10 cm) shy of where Byrne and his colleagues think the ice table sits.
"It's a shame that didn't happen," Byrne said. "You might have been having this conversation 30 years ago."
http://www.space.com/7333-water-ice-exp … aters.html
Water ice seen in fresh craters on Mars.
DR EMILY BALDWIN & KEITH COOPER
ASTRONOMY NOW
Posted: September 24, 2009
“The scientifically heartbreaking aspect of this work is that these craters are located very close [350 miles] to the Viking 2 lander site, which landed on Mars in 1976, and dug a trench about 4-6 inches deep,” says Cull. “What this new study is telling us is that if Viking 2 had been able to dig down a few more inches, it would have hit ice. That would have been a major discovery for our understanding of Mars, and it was literally inches away from our robotic fingertips.”
http://www.astronomynow.com/news/n0909/24mars/
What I would like though is a method where you could just dig down the required shallow depth and remain stationary there to process the water ice there rather than have a rover constantly digging up fresh ice. Still thinking about that one.
My architecture has the advantage of simplicity in only requiring a single stage for the space trip. It's in fact SSTO at least from the Mars surface.
Bob Clark
#777 Re: Interplanetary transportation » Mars Semi-Direct with Falcon » 2012-04-11 13:01:18
It may be possible to do the manned mission with a single Falcon Heavy based on fully using aerocapture and fully using in situ production of the return propellant.
Recall the calculations that fully used aerocapture to eliminate propellant burns at Mars and at Earth:
Index » Human missions » Developing the cis-Lunar economy and infrastructure
http://www.newmars.com/forums/viewtopic … 11#p111711
It was only 10.2 km/s total delta-V, 3.8 km/s for the outbound trip from LEO to Mars, and 6.4 km/s for the return trip from Mars to LEO. Let's suppose we need 20 mT payload. I'm considering something like the Bigelow Sundancer as the hab module. This was to have a pressurized volume of 180 m^3 and mass of 8,618.4 kg for a crew of 3. The original NASA Transhab might also work. It was to have a pressurized volume of 339.8 cubic meters at a mass of 13.2 mT. The Bigelow 330 derived from Transhab would probably be too heavy since even though the internal volume was the same, the mass was up to 20 mT.
For the outbound trip we would need a Centaur style upper stage with a 465.5s Isp at 30 mT propellant load and 3 mT dry mass: then we would get a 3.8 km/s delta-V: 465.5*9.81ln(1 + 30/(3 +20)) = 3,812 m/s.
I'm assuming all the consumables, aerobrake, thermal protection, power systems, would be contained within the left-over mass allowance from 20 mT once you subtract off the mass of the hab.
But the problem is how do you get the 6.4 km/s delta-V for the return trip? The only thing I could think was to carry some empty drop tanks along from Earth. Then on Mars fill up these tanks as well. Remember the entire propellant load for the return trip is taken at Mars. I calculated you would need to take on an additional 50 mT of propellant in these drop tanks. But then you also have to figure in the mass of the empty tanks and subtract this off from the payload allowance. Probably with lightweight composites you can get the mass of these under 1 mT, even considering that hydrolox tanks are comparatively heavy.
Bob Clark
#778 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-04-10 14:32:45
...
But this slingshot effect is potentially quite large though according to the Wikipedia page. It can be as high as twice the speed of the planet around the Sun. Since for the Earth this is about 30 km/s, this means you can get a boost of about 60 km/s (!) How fast can you be going beforehand and still get swung around by the planet to still get the boost?
The gravitational slingshot won't work from Earth since the spacecraft even if you launch from the Moon is still moving in the same direction as Earth with respect to the Sun.
It might work from Venus. I remember reading some of the plans to reduce the return time from a Mars mission is to do a swingby of Venus. As I recall though, the reduction in time was not that dramatic as to reduce the trip time to days instead of months, so likely the same would be true for using a Venus swingby for the outbound trip.
I think we could use the Oberth effect though.
Bob Clark
#779 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-04-10 05:19:11
Note that all the components for such a mission already exist, the launcher, the spacecraft, and the rover. All that is required is to mate them together. On that basis such a mission probably could be launched within a year. Note also all of the U.S., Russia, and Europe have the required 20 mT launcher, and the upper or space stage capable of the space traverse. And China will also with the introduction of the Long March 5 in 2014. Then the question arises who will be first?
A common charge leveled at the space program is what is it good for? If the U.S. government fully financed the mining operation then based on an estimated $20 trillion value for the minerals on a single asteroid, this would have enough value to retire the entire U.S. debt(!) Preferably though the U.S. would only be a partial investor to retain the costs savings of a privately financed venture. Even then as a minority investor, the return in value to the U.S. government could be in the trillions.
...
Astrobotic Technology Reveals Design for Robot to Prospect at Moon’s Poles.
April 3, 2012 – 8:00 pm|Releases
PITTSBURGH, PA – April 3, 2012 – Astrobotic Technology unveiled
its new Polaris lunar rover design, which will prospect for
potentially rich deposits of water ice, methane and other resources at
the moon’s north pole in three years.
A powerful Falcon 9 rocket from SpaceX will launch Polaris from Cape
Canaveral in late October 2015. Four days later Polaris will land
during north pole summer, when patches of ground that are in cold
shadow most of the year get brief illumination. This is where ice will
be found closest to the surface, and when a solar-powered robot will
get the sunlight needed to sustain exploration. Polaris will search
for ice for the next 12 days until sundown in early November.
http://astrobotic.net/2012/04/03/astrob … ons-poles/
Such missions are very important to further quantify the amount of
water and other minerals suggested by orbiter missions to lie within
the polar regions of the Moon. They are a key first step, though there
is nothing better than having the actual sample in your hands for
geologists to assay and determine if there really are such valuable
minerals as gold and silver tentatively identified by the LCROSS
mission.
Bob Clark
#780 Re: Interplanetary transportation » Cycling Spaceship » 2012-04-09 10:22:16
By repeatedly making use of the gravitational assist of swinging by the planets it appears the cyclers can be made to travel at arbitrarily high speed between the planets. Then in fact the time of travel can be days instead of months.
The problem for manned trips though is how do you get the crew up to these high speeds? The cyclers can still be useful for delivering supplies on standby as it were if there is an emergency where the crew on one of the planets needs a replacement engine, power supply, habitat, etc.
Another way they could be useful for minimizing spaceship size, is that the supplies, habs, propellants, etc. could use this method to get up to high speed, but the crew only liftoff in a small capsule so that a much smaller vehicle could get them up to high speed. Then the small capsule could link up with the much larger craft carrying all the supplies, habs, propellants, that reached its high speed by the cycler method.Bob Clark
However, see here for some problems with this:
Developing the cis-Lunar economy and infrastructure.
http://newmars.com/forums/viewtopic.php … 53#p112153
Bob Clark
#781 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-04-09 09:59:26
...
It involves orbital mechanics of spaceflight. There are a few different components to the question. First, if our Mars rocket departed from the Moon or a Lagrange point propellant depot fully fueled towards Earth at, say, 11 km or more, so it's moving at speeds beyond Earth's escape velocity, then in just passing by the Earth it should pick up additional speed equal to Earth's escape velocity about 11 km/s. So at least temporarily it should have a speed of 22 km/s. But the problem is that it still be slowed down by the Earth as it proceeds to Mars, so it will lose some of this speed. How much speed will it lose?
What I want to do is leave Earth's vicinity at such high speed so that you don't have the long travel times of the Hohmann orbit, and in fact so that the trajectory approximates a straight-line path and if you do it at closest approach of Mars then the travel time could be say 60,000,000 km/22 km/s = 2,700,000 s, about 31 days. (You would have the problem of aerocapture at such highly elevated speeds but I'll leave that to another discussion.) So another question I have is at what high speed would you need so that the path is approximately straight-line?
This is just using Earth flyby. Could we in addition also use a Venus flyby? You would need an orbital arrangement where both Venus and Mars are near the Earth at the same time. Say you are now traveling at 22 km/s towards Venus, minus the amount you're slowed by leaving the Earth. You can likewise pick up about 11 km/s additional speed by just passing by Venus on the way to Mars, perhaps arranging it so that the path is bent by Venus to aim the craft towards Mars. So you could conceivably be traveling now at 33 km/s, again though I need to know how much speed you would lose in leaving Venus. You would also have to factor in the additional time it takes to get to Venus and the longer straight-line distance to Mars from Venus. Also, in being within Venus's orbit around the Sun, the greater gravitational effects of the Sun will have a greater effect to curve the trajectory.
Finally, could we use repeatedly the gravitational boosts of Earth and Venus? Suppose we are now at 33 km/s, more or less, after leaving Venus but we arrange it so our path is bent completely around to head back towards Earth. Could we once more get an additional 11 km/s to bring our velocity to 44 km/s? Could we do this repeatedly to get arbitrarily high speeds?
I'm having trouble disentangling the gravitational slingshot effect and the Oberth effect.
By the Oberth effect I can get greater velocity if I apply my rocket burn when I'm closest to the planet. Plugging in some speeds into the equation on the Wikipedia page I am able to get an additional boost about that of Earth's or Venus' escape speed if I make the rocket burn high, say, 10 km/s or above. The problem is this page seems to be suggesting to get the gravity boost, I need to apply a rocket burn but I wanted to get the gravity boost without having to apply an additional rocket burn.
On the other hand the Wikipage on the gravitational slingshot effect suggests I can get an additional boost without having to supply an additional burn. But the problem here is I want to get these additional boosts while my craft is already moving at high speed, to boost even higher, but if I'm going too fast I won't swing around the planet but instead go right by it without getting the slingshot effect.
But this slingshot effect is potentially quite large though according to the Wikipedia page. It can be as high as twice the speed of the planet around the Sun. Since for the Earth this is about 30 km/s, this means you can get a boost of about 60 km/s (!) How fast can you be going beforehand and still get swung around by the planet to still get the boost?
Bob Clark
#782 Re: Interplanetary transportation » Cycling Spaceship » 2012-04-08 07:20:16
By repeatedly making use of the gravitational assist of swinging by the planets it appears the cyclers can be made to travel at arbitrarily high speed between the planets. Then in fact the time of travel can be days instead of months.
The problem for manned trips though is how do you get the crew up to these high speeds? The cyclers can still be useful for delivering supplies on standby as it were if there is an emergency where the crew on one of the planets needs a replacement engine, power supply, habitat, etc.
Another way they could be useful for minimizing spaceship size, is that the supplies, habs, propellants, etc. could use this method to get up to high speed, but the crew only liftoff in a small capsule so that a much smaller vehicle could get them up to high speed. Then the small capsule could link up with the much larger craft carrying all the supplies, habs, propellants, that reached its high speed by the cycler method.
Bob Clark
#783 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-04-08 01:41:49
...
Duke's proposal is perhaps the simplest and most elegant proposal for sending astronauts back to the moon that has been make in recent years. Three launches would establish a reusable transportation system to the moon.
He estimated that an eight-tonne fuel-making system on the moon would allow one or two manned flights to the moon per year. The lunar-based vehicles would be designed for ten uses before a solar electric vehicle would have to bring them back to low earth orbit for refurbishment (if a facility for such refurbishment were built).
Thanks for that. Do you have a link to the full paper?
Bob Clark
#784 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-04-07 09:54:43
I mentioned I had a question for Hop in this thread:
Robotic Mining.
http://newmars.com/forums/viewtopic.php … 92#p111392
It involves orbital mechanics of spaceflight. There are a few different components to the question. First, if our Mars rocket departed from the Moon or a Lagrange point propellant depot fully fueled towards Earth at, say, 11 km or more, so it's moving at speeds beyond Earth's escape velocity, then in just passing by the Earth it should pick up additional speed equal to Earth's escape velocity about 11 km/s. So at least temporarily it should have a speed of 22 km/s. But the problem is that it still be slowed down by the Earth as it proceeds to Mars, so it will lose some of this speed. How much speed will it lose?
What I want to do is leave Earth's vicinity at such high speed so that you don't have the long travel times of the Hohmann orbit, and in fact so that the trajectory approximates a straight-line path and if you do it at closest approach of Mars then the travel time could be say 60,000,000 km/22 km/s = 2,700,000 s, about 31 days. (You would have the problem of aerocapture at such highly elevated speeds but I'll leave that to another discussion.) So another question I have is at what high speed would you need so that the path is approximately straight-line?
This is just using Earth flyby. Could we in addition also use a Venus flyby? You would need an orbital arrangement where both Venus and Mars are near the Earth at the same time. Say you are now traveling at 22 km/s towards Venus, minus the amount you're slowed by leaving the Earth. You can likewise pick up about 11 km/s additional speed by just passing by Venus on the way to Mars, perhaps arranging it so that the path is bent by Venus to aim the craft towards Mars. So you could conceivably be traveling now at 33 km/s, again though I need to know how much speed you would lose in leaving Venus. You would also have to factor in the additional time it takes to get to Venus and the longer straight-line distance to Mars from Venus. Also, in being within Venus's orbit around the Sun, the greater gravitational effects of the Sun will have a greater effect to curve the trajectory.
Finally, could we use repeatedly the gravitational boosts of Earth and Venus? Suppose we are now at 33 km/s, more or less, after leaving Venus but we arrange it so our path is bent completely around to head back towards Earth. Could we once more get an additional 11 km/s to bring our velocity to 44 km/s? Could we do this repeatedly to get arbitrarily high speeds?
Bob Clark
#785 Re: Human missions » Landing on Mars » 2012-04-06 10:39:11
"Lower heating if decelerating at higher up" is really lower peak skin temperatures at lower ballistic coefficient. The total integrated BTU's (KW-hr) absorbed over the trajectory is actually higher in that case, but that's an easier problem to solve than high skin temperatures. Control of survivable gees is the real key to how this trajectory is selected.
For a winged vehicle here on Earth, a "ballistic coefficient" is better expressed as "wing loading", which is vehicle weight divided by wing planform area. To obtain the benefit of earlier deceleration higher up, this needs to look more like a small light aircraft (10-20 lb/sq.ft) than the shuttle or a jet fighter (100-200 lb/sq.ft). Peak skin temperatures are under 2000 F, which ceramics, or very, very, very heavily-cooled Inconel-X, can survive. Better odds on the ceramics.
For a space capsule here on Earth, it is vehicle weight divided by heat shield broadside area. This has typically been about 4 times the shuttle/jet fighter range (400-800 lb/sq.ft). These vehicles punch very deep into the air before they slow significantly. Peak skin temperatures (stagnation point, leading edges) is around 3500-4000 F. That's why Shuttle had ablative carbon-carbon leading edge and nose cap pieces, and why Mercury, Gemini, Apollo, and Dragon had/have ablative phenolic composite heat shields.
PICA-X on Dragon (and the other new capsules) is actually nothing but a modernized version of the 1960-ish vintage silica-phenolic, which is really nothing more than what the old capsules used in the 1960's and 1970's.
Think about it: if you can actually achieve such a low wing loading/ballistic coefficient (around 10-20 lb/sq.ft), you could survive re-entry here on Earth (from orbit) with a steel truss "airframe" and a ceramic fire-curtain cloth skin, as long as there was some sort of insulating standoff between the frame and the skin, and some sort of ventilation inside to absorb the BTU's. In other words, a ceramic fabric-skinned variant of a Piper Cub might actually be a feasible re-entry vehicle from LEO.
Anything that might work here would work on Mars. The heating there is less demanding. I see some experiments that need to be done here!!!!
GW
Thanks for that. Your suggestion of getting a lightweight structure with high wing area, or equivalent, reminded me of the Lockheed X-33:
X-33.
http://www.astronautix.com/lvs/x33.htm
It was only to weigh 63,000 lbs dry, compared to the space shuttle orbiter at ca. 200,000 lbs. Note that the problem of the composite propellant tanks not holding up when pressurized would not be a problem here since we would remove them for this purpose. In fact this would make the weight even lighter. This page gives the total propellant tank weight as 15,200 lbs:
Marshall Space Flight Center
Lockheed Martin Skunk Works
Sept. 28, 1999
X-33 Program in the Midst of Final Testing and Validation of Key Components.
http://www.xs4all.nl/~carlkop/x33.html
So conceivably the weight might be as low as 48,000 lbs, though actually the propellant tanks provided structural rigidity. So we would have to add some strengthening members if they were removed, but the extra weight would quite likely be much less than the 15,200 lbs of the tanks. The reference area for the X-33 is in the range of 1600 sq.ft. So if not too much added weight for strength was required for the tankless version, we might get a wing loading of 48,000 lbs/1600 sq.ft. = 30 lbs/sq.ft.
The volume you would get by removing the tanks on the X-33 is about 300 m^3, which is about the volume of the payload bay on the shuttle orbiter. So it could contain about the same volume as the Bigelow BA 330 inflatable habitat proposed for Mars missions or orbital or surface habitats.
BTW, since the X-33 could weigh 1/3rd to 1/4th the weight of the space shuttle orbiter, I suggested that a plan that was announced last year for a commercial, next-generation shuttle system:
Next Gen Shuttle-Capable vehicle interest as secret effort to save orbiters ends.
December 19th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/12/ … ters-ends/
should use instead the X-33 in place of the shuttle orbiter.
This would mean you could carry more payload for this STS 2.0, if you will, than the original shuttle system. Ironically, the fully orbital version of the suborbital X-33 was considered to be a "next-generation" shuttle:
Lockheed Vying to Design Replacement for Shuttle : Space: Skunk Works in Palmdale competes against McDonnell Douglas Aerospace and Rockwell International.
March 14, 1995|JEFF SCHNAUFER | SPECIAL TO THE TIMES
http://articles.latimes.com/1995-03-14/ … ll-douglas
However, strictly speaking in being single stage to orbit, this orbital version would have been quite different than the space shuttle orbiter. But having the X-33 in place of the shuttle orbiter in a multi-stage system would be more in keeping with the idea of a "next generation" shuttle.
Bob Clark
#786 Re: Life support systems » Wind power : possible ? » 2012-04-04 13:09:58
The solar may be easier than we think by time we get to go with these types of advances.
Spray-on solar may be future for green energy
Mitsubishi Chemical is the first company to create prototype spray-on solar cells, which at present have a practical conversion level of 10.1 percent of light energy into electricity.
The new solar cells utilize carbon compounds which, when dried and solidified, act as semiconductors and generate electricity in reaction to being exposed to light.
Thanks for that, Space. Given that they expect to reach the same efficiency level as usual solar cells, this will mean that the mass of the solar cells needed for a certain power level will be about 1/10th of that currently needed. Since current cells are about 100W to 200W per kilo, we might get up to the 2,000 W per kilo range.
This page suggests a lunar base might only need 100 kW to 1 MW power and a Mars base 1 MW:
Lunar Production of Solar Cells.
http://www.asi.org/adb/02/08/solar-cell-production.html
Then this might be only 50 kg to 500 kg that needed to be transported to the Moon or Mars for the power for the base. It would also be easy to transport since it could be sent in liquid form.
Also, in regards to the VASIMR plasma propulsion system Robert Zubrin has raised some points to question its feasibility since it requires high power at lightweight, which was proposed to be nuclear. Zubrin argued nuclear plants at the power level required, in the range of 1,000 watts per kilo, were not likely to be achieved in the near future:
VASIMR and a new war of the currents.
by Chuck Black
Monday, August 1, 2011
http://www.thespacereview.com/article/1896/1
But these new solar cells might be able to provide the power at the lightweight required.
Bob Clark
#787 Re: Meta New Mars » Increasing use » 2012-04-04 12:16:48
Unfortunatey I miss those many meaningfully made posts (5000 +) as many contained lots of useful data for getting mission to succeed for mars....
About another 1000 were lost on the 2 crashes that MarsDrive had and possibly another 1000 + on Red Colony....
As for welcomed Ideas I have felt that its the topics that one has of interest that makes or breaks how you feel on any forum...and how one puts thoughts from one to the other for continued discussion in a topic...
I saw you had over 5,600 posts Space dude, dating back all the way to '04. I think most of these older posts dating back to before the crash have been recovered based on the fact that I was able to view several of yours going back that far by clicking the link under your profile for your prior posts.
NewMars was indeed a great repository for information about achieving manned Mars missions, and a relatively well frequented one based on the number of posts. I think it can be well frequented again because the interest in achieving manned Mars missions, I believe, will again be renewed by Elon Musk making it a key objective for his company over the next decade.
It will be realized it is an achievable goal when SpaceX achieves first the goal of cutting the costs to LEO by two orders of magnitude by reusability. I believe such cuts in the cost to space are indeed reachable. Then when this is reached it will be understood that the cost for a Mars mission will be similarly cut, bringing it well within a range that we can afford.
Bob Clark
#788 Re: Human missions » Landing on Mars » 2012-04-03 11:54:36
There is a link to report on "Red Dragon" that discusses this question of landing on Mars in this thread on NasaSpaceFlight:
Re: Red Dragon
« Reply #297 on: 12/12/2011 10:19 PM »
http://forum.nasaspaceflight.com/index. … #msg838572
For landings that use aerobraking, rather than retro-rockets, the problem is one of the "ballistic coefficient", as discussed in this article:
Mars Exploration Entry, Descent and Landing Challenges.
http://www.4frontierscorp.com/dev/asset … rs_EDL.pdf
It's proportional to the ratio of landed mass to area of the air brake. The problem is because of Mars thin atmosphere the air brake's surface area has to be large for a large landed mass which would mean an even heavier mass. (Note: in some sources the ballistic coefficient is defined in terms of the ratio of the weight, in Newtons, to the surface area.)
So the air brake would have to look like it does here:
How to Go to Mars--Right Now!
Human exploration of Mars doesn't need to wait for advanced rockets, giant spaceships, or lunar base stations.
By Robert Zubrin / June 2009
http://spectrum.ieee.org/aerospace/spac … ight-now/1
and here:
Aerobrake.
http://smpritchard.deviantart.com/art/A … -276129026
But it would have to be of a lightweight material to be this large without incurring too large a weight penalty while at the same time having great heat resistance for reentry. One solution that has been proposed is a ballute. The SpaceX PICA-X material used for the Dragon heat shield might also work since it appeared to undergo minimal degradation on Earth reentry so it's possible it could be made thinner to get a larger heat shield at low weight.
Bob Clark
#789 Human missions » Elon Musk: ticket to Mars for $500,000. » 2012-03-22 07:12:04
- RGClark
- Replies: 40
Elon Musk was interviewed on the U.S. news program "60 minutes" on Sunday:
SpaceX: Entrepreneur's race to space.
March 18, 2012 4:44 PM
From PayPal to electric cars to rockets, billionaire entrepreneur Elon Musk wants his company, SpaceX, to build America's next manned spacecraft. Scott Pelley reports.
http://www.cbsnews.com/video/watch/?id=50121782n
He was also interviewed on a BBC radio program where he states that eventually, after perhaps a decade of regular flights, the price for a round trip ticket to Mars might be down to $500,000 per person:
20 March 2012 Last updated at 19:25 ET
Jonathan Amos, Science correspondent.
Mars for the 'average person'.
http://www.bbc.co.uk/news/health-17439490
Bob Clark
#790 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-03-03 14:28:58
Hop, I wanted to see what kind of payload we could get to Mars for a sample return mission using one of the current 20 mT payload launchers or the 53 mT payload Falcon Heavy. If you use aerobraking both for landing at Mars and for Earth orbit insertion, what would be the total delta-V to go from LEO to the Martian surface and back to LEO?
From this diagram I get a surprisingly low 10.2 km/s:
Delta-v budget.
Delta-vs between Earth, Moon and Mars.
<image>
http://en.wikipedia.org/wiki/Delta-v_bu … n_and_Mars
LEO to GTO: 2.5 km/s
GTO to Earth C3: .7 km/s
Earth C3 to Mars transfer: .6 km/s
Now notice for the delta-v's after this leading into Mars they all have red arrows indicating this part of the trip can be done by aerobraking. So this portion leading into Mars orbit and landing on the surface is only 3.8 km/s.
Then for the return trip:
Mars(surface) to low Mars orbit: 4.1 km/s
low Mars orbit to Phobos transfer: .9 km/s
Phobos transfer to Deimos transfer: .3 km/s
Deimos transfer to Mars C3: .2 km/s
Mars C3 to Mars transfer: .9 km/s
Now the delta-v's after this leading into Earth all have red arrows indicating this part of the trip can be done by aerobraking. So the return part of the trip can amount to only 6.4 km/s, for a total of 10.2 km/s for the round trip.
As for the heat shield for these Mars return velocities notice that the SpaceX Dragon's PICA-X heat shield was designed to withstand such velocities. It reportedly weighs only half of Apollo era heat shields which would put it at about 8% of the landed mass.
Using a 465.5 s Isp for the LH2/LOX engines and one of the later Centaurs with a ca. 20 mT propellant load and ca. 2 mT dry mass then with a single stage you get a payload of .39 mT: 465.5*9.8ln(1 + 20/(2 + .39)) = 10,206 m/s. The problem is this payload mass also has to account for the heat shield mass for the aerobraking/aerocapture.
We can do better than this using staging, but the problem then is the smaller LH2/LOX stages do not have as good a mass/ratio. For instance see the gross mass/dry mass ratio of the Ariane upper stages here:
http://www.astronautix.com/props/loxlh2.htm
We may suppose that with using lightweight composites we can get a smaller stage with a high mass ratio, say 10 mT propellant load and 1 mT dry mass. Then with staging of two of these we can get a payload of 1.3 mT:
465.5*9.8ln(1 + 10/(1 + 11 + 1.3)) + 465.5*9.8ln(1 + 10/(1 + 1.3)) = 10,206 m/s.
Bob Clark
#791 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-02-29 02:45:35
Funny how this topic is being written of.
The cislunar econosphere (part 2)The article points to Lunar Oxygen
The to keep Nasascientists busy
Then again another topic of our discusions are 3D printing technology
Tieing this all up is a story on how Mining on the moon: gold, fuel, and Canada's possible role in a new space race
Thanks for those links, all very informative. I like the fact in that last one gold was mentioned on the Moon. LCROSS may have indeed detected gold but this detection is controversial. I wonder if the author was informed of some further information to firm up that detection.
Bob Clark
#792 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-28 07:14:08
There will be a media demonstration of the Scarab lunar rover using a new fuel-cell technology on Wednesday, Feb. 29th at the NASA Glenn center:
Media Invited to NASA Glenn to See New Fuel Cell Demonstration on Mobile Rover.
Source: Glenn Research Center
Posted Thursday, February 23, 2012
CLEVELAND - A demonstration of a fuel cell that will allow rovers on extraterrestrial surfaces to go farther and last longer will be conducted at NASA's Glenn Research Center on Feb. 29 at 11 a.m.
The new type of fuel cell will extend the range of surface operations for rovers that will explore new worlds as part of future NASA missions. Unlike a conventional fuel cell that needs a pump to remove the water produced inside the device, this non-flow-through fuel cell uses capillary action to wick away the water. By eliminating the pump, a non-flow-through fuel cell is simpler, lighter, and more reliable.
The rover that will demonstrate the fuel cell in Glenn's Simulated Lunar Operations (SLOPE) facility is called SCARAB. It was developed by Carnegie Mellon Robotics Institute, Pittsburgh, under a grant from Glenn, and is regularly used for Human Robotic systems project mobility research in SLOPE.
http://www.spaceref.com/news/viewpr.html?pid=36206
Perhaps one of the reporters will inquire when the test vehicle can be turned into a flight ready version.
Bob Clark
#793 Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-02-23 07:05:31
Hop, I wanted to see what kind of payload we could get to Mars for a sample return mission using one of the current 20 mT payload launchers or the 53 mT payload Falcon Heavy. If you use aerobraking both for landing at Mars and for Earth orbit insertion, what would be the total delta-V to go from LEO to the Martian surface and back to LEO?
From this diagram I get a surprisingly low 10.2 km/s:
Delta-v budget.
Delta-vs between Earth, Moon and Mars.
http://en.wikipedia.org/wiki/Delta-v_bu … n_and_Mars
LEO to GTO: 2.5 km/s
GTO to Earth C3: .7 km/s
Earth C3 to Mars transfer: .6 km/s
Now notice for the delta-v's after this leading into Mars they all have red arrows indicating this part of the trip can be done by aerobraking. So this portion leading into Mars orbit and landing on the surface is only 3.8 km/s.
Then for the return trip:
Mars(surface) to low Mars orbit: 4.1 km/s
low Mars orbit to Phobos transfer: .9 km/s
Phobos transfer to Deimos transfer: .3 km/s
Deimos transfer to Mars C3: .2 km/s
Mars C3 to Mars transfer: .9 km/s
Now the delta-v's after this leading into Earth all have red arrows indicating this part of the trip can be done by aerobraking. So the return part of the trip can amount to only 6.4 km/s, for a total of 10.2 km/s for the round trip.
As for the heat shield for these Mars return velocities notice that the SpaceX Dragon's PICA-X heat shield was designed to withstand such velocities. It reportedly weighs only half of Apollo era heat shields which would put it at about 8% of the landed mass.
Bob Clark
#794 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-22 02:08:38
More on the low cost approach to the LCROSS mission from program manager Dan Andrews:
Going Lunar for Less.
By Melissa Salpietra
Posted 04.01.09
NOVA scienceNOW
DAN ANDREWS: The whole key with LCROSS is to use what exists. You take things that are available, you glue them together, you attach them in as simple a way as you can. You’re not doing a bunch of custom designs and development. You’re leveraging everywhere you can. And that’s a really smart way to get the most out of the money that you're given.
http://www.pbs.org/wgbh/nova/space/moon-for-less.html
LCROSS won Popular Mechanics' Breakthrough Award in 2010:
NASA's LCROSS Wins 2010 Popular Mechanics Breakthrough Award.
The individual LCROSS 2010 Breakthrough Award recipients are:
Daniel Andrews, LCROSS project manager at Ames
Anthony Colaprete, LCROSS project scientist and principal investigator at Ames
Stephen Carman, LCROSS spacecraft project manager at Northrop Grumman
Craig Elder, LCROSS spacecraft manager at Northrop Grumman
"We are honored to win this award," said Steve Hixson, vice president of Advanced Concepts - Space and Directed Energy Systems for Northrop Grumman Aerospace Systems in Redondo Beach, Calif. "It is a significant acknowledgement of the high caliber of our engineering skills and our close partnership with Ames, which developed the LCROSS payload and conducted mission operations. It also validates our ability to build small, inexpensive spacecraft with high science value very quickly, awakening the industry and the nation to the viability of this mission class."
http://www.nasa.gov/centers/ames/news/r … -86AR.html
Program manager Dan Andrews also was awarded NASA's Systems Engineering Excellence award for the mission:
2010 Systems Engineering Excellence Award.
01.06.10
http://www.nasa.gov/centers/ames/news/f … award.html
Bob Clark
#795 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-21 08:46:57
...
As I said to keep costs low these missions should be privately financed. NASA is planning to launch an asteroid sample return mission in 2016. This would not return the samples though until 2023 and is budgeted at $800 million without even launch costs:NASA to Launch Asteroid-Sampling Spacecraft in 2016.
Mike Wall, SPACE.com Senior WriterDate: 25 May 2011 Time: 07:10 PM ET
http://www.space.com/11788-nasa-asteroi … -rq36.htmlWhen you add on launch costs and considering the usual NASA cost overruns this will probably wind up being a billion dollar mission. Also, since some proposed human missions to asteroids would have a duration of 5 to 6 months, these sample return missions could return their samples in months rather than the seven years planned for the NASA mission.
Note that all the components for such a mission already exist, the launcher, the spacecraft, and the rover. All that is required is to mate them together. On that basis such a mission probably could be launched within a year. Note also all of the U.S., Russia, and Europe have the required 20 mT launcher, and the upper or space stage capable of the space traverse. And China will also with the introduction of the Long March 5 in 2014. Then the question arises who will be first?
A common charge leveled at the space program is what is it good for? If the U.S. government fully financed the mining operation then based on an estimated $20 trillion value for the minerals on a single asteroid, this would have enough value to retire the entire U.S. debt(!) Preferably though the U.S. would only be a partial investor to retain the costs savings of a privately financed venture. Even then as a minority investor, the return in value to the U.S. government could be in the trillions.
However, it may indeed be possible that a fully NASA financed venture could maintain the low costs of a privately financed one - with the right management. I consider the LCROSS lunar impactor to be the perfect NASA mission because it returned such profoundly important results and at low cost, only $79 million without launch costs, which is like pocket change for planetary missions:
Inside NASA's Plan to Bomb the Moon and Find Water.
By Michael Milstein
October 1, 2009 12:00 AM
Typically, 10 to 15 percent of a spacecraft's budget goes into instruments; on LCROSS, it's roughly 3 percent, or $2 million. When Anthony Colaprete, NASA's lead scientist for the mission, went to big aerospace companies for instruments, they laughed at his budget. So he turned to small outfits instead. He bought near-infrared spectrometers from a company that makes them for breweries to test the alcohol content of beer on assembly lines. He resisted agency reviewers who wanted him to put an anodized coating on the aluminum storage boxes. "One of their arguments was, `It's not very expensive--just do it,'" he says. "I'm like, `Well, I want to save that $1000. I'm very cheap.'"
http://www.popularmechanics.com/science/space/4277592
LCROSS: A HIGH-RETURN, SMALL SATELLITE MISSION.
Daniel Andrews, LCROSS PM
NASA-Ames Research Center, MS 240-3, Moffett Field, CA 94035, USA.
http://ntrs.nasa.gov/archive/nasa/casi. … 030093.pdf
Academy of Program/Project & Engineering Leadership.
Lunar CRater Observation and Sensing Satellite (LCROSS).
The Good Enough Spacecraft.
From Andrews‘s perspective, the LCROSS spacecraft had to be ―faster, good enough, cheaper.‖ He made clear to his team from the beginning that LCROSS was not about maximum performance. ―It was about cost containment,‖ Andrews said. ―LCROSS was not about pushing the technical envelope. It was about keeping it simple – keeping it good enough.‖
The LCROSS team had 29 months and $79 million to build a Class D mission spacecraft. (See below for a brief explanation of NASA mission risk classifications.) The low-cost, high-risk tolerance nature of the project led to a design based on heritage hardware, parts from LRO, and commercial-off-the-shelf components.
http://www.nasa.gov/pdf/474589main_LCRO … _23_10.pdf
LCROSS rode piggyback on the LRO mission so did not have to pay for the Centaur space stage, but even if you include this that would only be an additional $30 million or so.
LCROSS Program Manager Daniel Andrews and lead scientist Anthony Colaprete deserve major kudos for using innovative methods to accomplish such a successful mission under cost saving constraints. If we were to have NASA financed asteroidal and lunar prospector landers then they would be my choice to manage those missions.
Note now that if NASA funded these exploratory lander missions that proved definitively that asteroids or even the Moon contained such extraordinary mineral wealth, then under the principle that the government has the authority to grant mining rights to private companies, the U.S. government could sell these rights for a total of, say, $1 trillion, while only having to have spent ca. $200 million for the lander missions.
Bob Clark
#796 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-16 03:01:43
...
On this listing of space vehicles you can find that the later versions of the Centaur upper stage have a mass ratio of about 10 to 1:http://www.friends-partners.org/partner … pndexc.htm
The Isp's given for the RL-10A engines used on these stages are around 450 s, but an updated version with a longer, extensible nozzle has an Isp of 465.5 s:
RL10B-2.
http://www.pw.utc.com/products/pwr/asse … l10b-2.pdfThis page gives the delta-V's needed for trips within the Earth-Moon system:
Delta-V budget.
Earth–Moon space.
http://en.wikipedia.org/wiki/Delta-v_bu … Moon_space...
The RL-10 engine was proven to be reusable for multiple uses with quick turnaround time on the DC-X. The total propellant load of 40,000 kg could be lofted by two 20,000+ kg payload capacity launchers, such as the Atlas V, Delta IV Heavy, Ariane 5, and Proton.
The price for these launchers is in the range of $100-140 million according to the specifications on this page:Expendable Launch Vehicles.
http://www.spaceandtech.com/spacedata/elvs/elvs.shtml
The original architecture was to use two of the 20 mT to LEO launchers currently available with two Centaur upper stages to get a 4 mT Dragon to the Moon and back.
What can we do with a single one of these launchers currently available? Using a single one of these launchers to carry a single Centaur upper stage we could carry about 1 mT to the Moon and back:
From the delta-V table, you need 4.04 km/s to go from LEO to low lunar orbit, 1.87 km/s to go from low lunar orbit to the lunar surface, and 2.74 km/s with aerobraking to go from the lunar surface back to LEO for a total of 8.65 km/s delta-V for a single stage making the round-trip.
Then with a 465.5 s Isp, 20 mT total mass including payload, 2 mT dry mass, and 1 mT payload we get: 465.5*9.8ln(20,000/(2000 + 1000)) = 8,650 m/s, sufficient for the round-trip.
This would suffice to carry a lunar rover to operate in the permanently shadowed regions of the lunar poles or for an NEO asteroid:
Lunar Prospecting Robot To Be Field Tested On Hawaii's Mauna Kea
ScienceDaily (Oct. 14, 2008)
http://www.sciencedaily.com/releases/20 … 134111.htm
This university developed robot probably cost no more than a few million dollars. The single Centaur upper stage costs in the range of $30 million. And the 20 mT to LEO launchers cost in the range of $100-140 million, according to the Spaceandtech.com site estimates, for a total in the range of $200 million. This is a fraction of the amount spent by mining interests on exploration:
Explore Mining.
World non-ferrous expenditures for all exploration in 2007 are estimated to be about $10.4 Billion dollars.
http://www.holden.house.gov/comm/explor … ploration/
This same site also indicates that mining exploration is by nature high risk:
So just what is exploration?
It’s the collection of processes that gather information about the presence or absence of mineral deposits
The over-riding goal of exploration is to find deposits that can be worked as profitable mining operations.
It is a time-consuming, multi-stage investment in information different gathering processes.
It’s also an expensive, high-risk investment, unlike ordinary businesses investments.
Depending on the literature source, the success rate for finding profitable mining operations (when weighed against the total number of mineral properties examined by a company) have ranges from a high of 4 in 100 (that’s a 4% success rate!), to less than 1 in 100 and as low as 1 in 1000 (that’s a .1% success rate!).
For any investment venture a cost/risk/benefit analysis has to be made. Compared to the cost already spent by mining interests yearly the cost is relatively low especially for a consortium of mining interests funding the mission together.
The risk is composed of the risk of the mission failing and of it not finding the high amounts of precious minerals. At least for the asteroid missions the risk of it not finding the high value minerals is low as there are several independent lines of evidence that precious metals are located uniformly on asteroids. So that leaves the risk of the mission failing. Considering the amount of U.S. experience with planetary missions, this risk is considerably better than the 1 in 1,000 chance of success some estimates put on Earth bound mining exploration.
However, quite important when measuring cost and risk, are the benefits to justify them. The possible benefits are more mineral wealth in a single asteroid than all that mined in all of human history.
Indeed the likelihood of the high amounts of precious minerals is so good, and the benefits of success are so extraordinarily high, that it would pay to do several missions if there are failures.
That is for the asteroid missions. However, if such asteroid mining missions are to be profitable then it would be much cheaper if the large amount of propellant needed to carry out the transport could be obtained from the Moon rather than by lofting it from Earth's deep gravity well. Then to insure that propellant could be obtained from the Moon's polar regions sample return missions to the lunar poles would have to be mounted as well. The nice thing about these missions is that the same rovers and spacecraft could be used for the asteroid sample return missions. Then these lunar sample return missions could be regarded as test missions to give further assurance of the technology for returning the samples from asteroids. And if the lunar polar samples show the high precious metal amounts tentatively detected by LCROSS then so much the better.
As I said to keep costs low these missions should be privately financed. NASA is planning to launch an asteroid sample return mission in 2016. This would not return the samples though until 2023 and is budgeted at $800 million without even launch costs:
NASA to Launch Asteroid-Sampling Spacecraft in 2016.
Mike Wall, SPACE.com Senior WriterDate: 25 May 2011 Time: 07:10 PM ET
http://www.space.com/11788-nasa-asteroi … -rq36.html
When you add on launch costs and considering the usual NASA cost overruns this will probably wind up being a billion dollar mission. Also, since some proposed human missions to asteroids would have a duration of 5 to 6 months, these sample return missions could return their samples in months rather than the seven years planned for the NASA mission.
Bob Clark
#797 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-14 01:09:20
Back to using Falcons and Dragons to go to the moon. You will need a lander of some tonnage, probably not unlike the old Apollo lander. You might even build it out of a Dragon with extra tanks, or maybe a from-scratch design. Whatever.
But, I don't see why you need a Centaur or any other departure stage. You don't need any more engines, you just need delta-vee. The Dragon will have the new Super-Draco thrusters on it with 120,000 lb of axial thrust, according to their website. There's eight of them, plenty of redundancy there. Just add a big dumb propellant tank, and plumb it up to the Dragon's system. Use the Dracos for all the delta-vee from LEO to lunar orbit, and back, sucking from the big dumb tank.
Launch the lander on one Falcon-Heavy, launch the Dragon and the big dumb tank on the other. Rendezvous in LEO, and dock the big dumb tank to the rear end of the Dragon, and the lander to its nose. Make up your plumbing connections. Then go to the moon.
Lots of minor details to work out, but I don't see any show stoppers here.
GW
Two Falcon Heavy's would be 106 mT to LEO which is getting into Saturn V class, and SpaceX already said it would take two Falcon Heavy's to send a manned mission to the Moon.
Could you get your plan to work with a single Falcon Heavy?
Bob Clark
#798 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-12 17:59:57
RGClark wrote:Just saw this discussed on Nasaspaceflight.com
Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PMThe key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."http://www.popularmechanics.com/science … ns-6653023
Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.Bob Clark
Well I can't enough of Musk personally...he seems to hit the nail bang centre on the head. He knows exactly what needs to be done to revolutionise space travel and he's doing it.
Bob, "time is money" as they say. I don't think he's worried about scheduling - it's the labour input that drives up the cost of recovery from sea. A return to a spaceport is what is required. Rest assured he will have done all the calculations. I think he's right.
6MT is still a huge amount to get up there.
I am presuming that for the Mars transit he must be thinking in terms of orbital assembly.
Still I'd like to see the trade study.
Bob Clark
#799 Re: Interplanetary transportation » Reusable Rockets to Orbit » 2012-02-12 07:39:14
While the experiment did not fully work it did take a big step towards proving that its possible...
http://www.universetoday.com/wp-content … allute.jpg
Armadillo Launches a STIG-A Rocket; Captures Awesome Image of ‘Ballute’
Thank for that. I didn't know they they had gotten that high at 82 km or had tested this recovery system, the 'ballute'.
Bob Clark
#800 Re: Interplanetary transportation » SpaceX Dragon spacecraft for low cost trips to the Moon. » 2012-02-12 02:17:23
Just saw this discussed on Nasaspaceflight.com
Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PM
The key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."
http://www.popularmechanics.com/science … ns-6653023
Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.
Bob Clark