New Mars Forums

Thanks for that. I definitely think your low cost estimates are in the ballpark of what's doable. I think you get how frustrating it is to hear it might take $100 billion dollars and another 20 years to get back to the Moon. When you look at the technical capability we already have available and take into account the fact that government financed programs inflate costs by as much as a factor of 10, you realize a Moon landing program can be accomplished at a fraction of that both in time and cost.

  Bob Clark

The Centaurs have not been used for manned missions but they were considered for the Gemini missions when low cost alternatives to the Apollo lunar architecture were being considered. The Centaurs have been very successfully used for satellite launches for decades. Also this "balloon tank" type lightweight structure was used to launch men into space with the original Atlas rocket in the 60's.
Resting on the Moon for a lightweight structure is actually easier than on Earth of course. And on the Moon the propellant tanks will not be empty since they'll contain the propellant for the return trip and pressurants.
But in any case the most important fact is they don't have to be Centaurs. Two Centaur stages of total gross mass 40 mT would be able to carry a Dragon to the Moon and back. But you have so much leeway with the 53 mT LEO payload capacity of the FH that you could use currently existing LH upper stages with worse mass ratios that don't use "balloon tanks" for the purpose. Take a look at the list of LH2/LOX stages here:

http://www.astronautix.com/props/loxlh2.htm

Scroll down to the "Associated stages" section. You'll see several russian, european, and chinese stages can be used in combination to transport the Dragon to the Moon and back while fitting within the 53 mT LEO capability of the FH.

   Bob Clark

Nice article here:

SpaceX Might Be Able To Teach NASA A Lesson.
May 23, 2011
By Frank Morring, Jr.
Washington

http://www.aviationweek.com/aw/generic/ … 324881.xml

   Bob Clark

Thanks for those links Hop and Adaptation.
Hop, I wanted to pass by you a calculation I did for a relatively low cost Mars mission. It would require though low cost lunar propellant. I'll post it in a day or so.

  Bob Clark

I've seen discussion of using the Falcon Heavy and the Dragon for a Mars sample return mission. Could the Falcon 9 perform such a mission with a smaller capsule than the Dragon?
Consider two different architectures: 1.)the propellant for the return is carried from Earth,
and
2.)the propellant for the return is produced on Mars. For this, note that orbital observations of Mars show there are locations that have high concentrations both of methane and water vapor. Then those locations could be targeted for the lander. These would also be good locations to search for life. Zubrin's Mars Direct proposal was of taking Mars resources either of the surface or the ground to produce propellant.
For our proposal, there would be a few options for making the propellant. You could take the methane from the atmosphere in the high concentration region, or you could make it by hydrolyzing the water vapor to get hydrogen and reacting it with the CO2 in Mars atmosphere to get methane.
A chemistry question: if you supply an electricity source, could you create methane directly from water vapor and CO2 in the Martian atmosphere?
You would need a power source. For Zubrins proposal he suggested nuclear power. But that was for a large manned lander. For our much smaller lander solar power may suffice.

    Bob Clark

Elon Musk has said he wants to cut the costs to space to the $100 to $200 per kg range by reusability. This is about a two order of magnitude reduction in cost. To put this in perspective, this is like a trans-atlantic flight that costs $1,000 suddenly being cut to cost $10 to $20.
Musk has said this transformation of the Falcon 9 to full reusability will be very hard. I don't believe it will be. But first, keep in mind how important that reduction in cost will be if it succeeds. If it succeeds then SpaceX will monopolize the launch business if the other launch companies do not field their own reusable vehicles. So there is a tremendous financial incentive for SpaceX to invest in reusability. Now, most in the industry believe reusability is very difficult for orbital vehicles and not even worth the expense. So if Musk reinforces that idea then he has a better chance at being able to field one without the other launch providers having one. And since they will not have even started to develop one, it will take them some time to catch up. The effect is that Musk will have a monopoly on all launches for at least a few years.
I don't know if that is Musk's intent in saying reusability is very hard. Actually I'm inclined to believe he is just saying what most in the industry believe including his own engineers. But a key reason why reusability is not very hard is because the cost in mass in reentry and landing systems is surprisingly low. In regards to the technical difficulty, there is none. We know how to do it as the shuttle orbiter and the X-37B and Dragon spacecraft has shown. I include the Dragon in the list of reusables because its heat shield showed minimal degradation on return. Musk has said the same heat shield could make hundreds of flights, at least to LEO.
I made an estimate before of about 28% of the landed mass has to go to reentry/landing systems. This was based on estimates of 15% for thermal protection, 10% for wings or for propellant for vertical landing, and 3% for landing gear. However, I said likely with modern materials this could be cut to half that. In fact, it might even be lower than 10%.

1.)Weight of thermal protection.

Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):

Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_l … at_shields

However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:

Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Com … ifications

However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:

Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shut … iderations

Also, for the X-37B the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:

X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/ … b_otv.html

For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:

Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index. … #msg754168

while being able to still survive lunar and even Martian reentry speeds.

SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.

Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as a high temperature-resistant aerospike nozzle to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
Another idea for a vertical landing vehicle would be to reenter head first. This was the preferred method of the Air Force since it provided increased cross-range. In that case you would have the blunt heat shield at the top of each stage. I thought this method would be unstable with the heavy engines now at the top during reentry, but since this was considered for the orbital version of the DC-X presumably this was solved.

2.)Weight of the wings and the landing gear.

For horizontal landing, a common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:

The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downlo … _22-32.pdf

On that same page the landing gear weight is the only 1.5% of the take-off weight.
Then for a vertical take-off vehicle these low weight proportions should apply to the dry, landing weight.

  Bob Clark

Thanks for the detailed response. That's actually a little too much detail for what I need. I read your post on ramjet boosters:

Sunday, August 22, 2010
Two Ramjet Aircraft Booster Studies
http://exrocketman.blogspot.com/2010/08 … e-boe.html

I noted that you were able to get better payload with more shallow launch angle but it created a problem for retrieving the first stage booster, since it went so far downrange. If I'm reading it correctly you were able to double the payload mass with the shallow angle, presumably using aerodynamic lift.
What I'm trying to determine if I can increase my payload just going to the range turbojets can get to, ca. Mach 3+. I intend to use the jets to get to medium altitude for a turbojet, ca. 15,000 m. But I need to get to a good angle as well as reaching its max speed. Another problem is that I don't know if it can get to max speed while climbing.
I looked at the case of the SR-71 and the XB-70 Valkyrie. These had more thrust than I wanted but that added weight because jet engines are so heavy. In any case I noted the climbing rate. From that it seemed doable, considering the high effective Isp, that you could reduce propellant mass that way. The problem is this is for a SSTO application and I can't afford the weight. What I wanted was the engines to put out in the range of 1/7th the vehicle weight to reduce the jet engine mass. What I don't know is how will that effect the climb rate, and will it even be able to reach supersonic now.
Note that an advantage of the SSTO is that you can get the better payload by flying a shallow angle and not have to worry about recovering the booster stage.

   Bob Clark

  Hello again, GW. Perhaps you can help me with a question I'm working on. I want to calculate the fuel burn for an airbreather, such as a turbojet or ramjet, as it accelerating upwards, climbing to altitude.
What is usually given in the specifications of a jet engine is its specific fuel consumption. But I'm fairly sure this is based on the gross thrust. However, to calculate the acceleration of your vehicle you need to subtract off the ram drag to calculate the net thrust. How do you calculate this net thrust?
Also, I believe the specific fuel consumption also changes with altitude and speed. How do you take that into account?

    Bob Clark

In regards to the costs of Mars missions see the thread "Elon Musk wants to put millions of people on Mars".

    Bob Clark

Great to have NewMars back. About increasing traffic the problem is probably due to most people not knowing it's back. I didn't until a couple of days ago.
A few ways to increase traffic:

A strong way would be for the administrators to send out a group email to all the subscribers before the meltdown.
For us members, on the other forums we contribute to we could include pointers to some of ours or of others interesting postings on NewMars.
I also like Jon Clarke's suggestion to mention it on the Mars Society Facebook page.

   Bob Clark

The impetus for this was this proposal by Launchpoint Technologies
to launch small satellites by magnetic fields:

Huge 'launch ring' to fling satellites into orbit
http://technology.newscientist.com/article/dn10180

However, there are many difficulties with getting large mass objects
up to orbital velocity with EM fields alone, discussed in this thread on sci.astro:

Subject: Coilguns and EM launchers.
http://groups.google.com/group/sci.spac … c6417ca8a/

And this article describes research dating back from 1977 able to
get a 3 gm object up to about 6000 m/s, and that record still hasn't
been exceeded for larger mass objects:

For Love of a Gun By Carolyn Meinel
First Published July 2007
The tumultuous history of electromagnetic launch.
http://www.spectrum.ieee.org/jul07/5296

If the launch system is to stay on the ground and for low mass
payloads you can just as well use reaction mass methods, i.e, rockets,
at high ISP to get the craft up to orbit velocity at short distances.
You wouldn't need to have hundreds of kilometers of cable extending
into air trailing from the craft. You could have a cable lying on the
ground and a short length of cable extending from the craft to the
cable on the ground, say 10 to 100 meters long. Keep in mind, just as
for the magnetic launch proposal, the main thing is getting that
horizontal velocity component required for orbit. To get to the
altitude for LEO is just a small proportion of extra velocity and
energy of that required for orbital velocity.
Note that for large launch systems such as the space shuttle a large
amount of thrust is needed just to accelerate that huge mass of fuel
that needs to be carried along. But when the exhaust velocity is much
larger than the ending velocity, say 100,000 m/s compared to 8,000 m/s
then by the rocket equation the mass of the fuel will be about the
same small proportion to the mass of the rocket, 8/100. (The exhaust
velocity being 100,000 m/s for this MPD thruster means the ISP,
specific impulse, actually is a quite high 10,000 s.)
The Launchpoint magnetic launch proposal only talked about launching
small satellites, 10 kilograms or so. Only one of the NASA Glenn
magnetoplasmadynamic (MPD) thrusters would be needed to accelerate a
10 kg mass to 1 g. Five of them could accelerate it to 5 g's at 5
MWatts power.
However, I should say key for this proposal is the idea the MPD
thrusters could be made lightweight. From the descriptions of the mode
of operation, essentially only requiring two electrodes, I'm assuming
this is the case. The images of them shown also suggest they would be
small and light weight.
Assuming that it is indeed the case the weight of the thrusters would
stay low when the thrust is scaled up, this might be used to launch
most satellites and also astronaut passengers. Most satellites are
less than around 1,000 kg. A 1 Gwatt power plant assuming power to
thrust scales up could accelerate this at 10 g's. Transportable gas
turbine electric generators at the 100's of megawatts scale can be
bought in the 10's of millions of dollars range. So 1 Gwatt total
would cost in the range of 100's of millions of dollars.
NASA documents give the human endurance level for acceleration
according to duration, as described here:

G tolerance (Dani Eder; Henry Spencer; Jordin Kare; James Oberg)
http://yarchive.net/space/science/g_tolerance.html

At 9 g's it's about 3 minutes for astronauts lying down in
acceleration seats. The formula for speed v attained at an
acceleration a over distance d is v^2 = 2ad. So for v = 8,000 m/s and
a = 10 g's = 100 m/s^2, d is 320 km. They would have to undergo this
for t =v/a = 80 s.
You could have the craft go in a circle at a smaller radius to reduce
the scale of the distance covered by the cable on the ground, but this
would result in a higher acceleration according to the formula a = v^2/
r. For a radial distances of a few km's you get accelerations at the
1,000's of g's scale, which would greatly reduce the payload and make
it impossible for human passengers.
However, for small satellites, a few kilos, it might be easier to use
such small linear or radial distances of just a few kilometers.

Bob Clark