|
Post by bazbear on Feb 3, 2010 3:10:58 GMT -4
Is this goal even remotely feasible? I'm guessing it would have to be piggy backed on another commercial mission with it's own extra stage. It's far beyond the loss of money SpaceShipOne entailed.
Of course if I was that rich I'd have already tried the "space flight participant" route lol
ETA- if I had anything near the money needed to go for this prize, I'd try for a ride to the ISS first.
|
|
|
Post by bazbear on Feb 12, 2010 1:53:34 GMT -4
As redundant as I was, I guess no one wants to comment in any case. It wasn't moved, so I guess it fits the reality thread (I guess general discussion would have worked as well). LO, feel free to delete it at your leisure
|
|
|
Post by ka9q on Feb 22, 2010 5:11:33 GMT -4
I agree, it's highly unlikely. The prize certainly won't come close to covering the cost.
I was annoyed by all the "go private enterprise, rah rah rah" cheerleading during the Spaceship One demonstration. I have nothing against private individuals and commercial companies going into space, but I thought they were more than a little disingenuous in implying that they were close to providing commercial space services. They really exploited the fact that few laymen understand that the hard part in making low earth orbit isn't the altitude but the velocity. I figured SpaceShip One had only about 3% or 4% of the total energy needed to make even a minimal (exo-atmospheric) earth orbit.
And it was hardly the first time that a rocket built by a commercial company had entered space. Practically every US rocket was built by a commercial company.
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 22, 2010 10:37:07 GMT -4
A couple years ago I was contacted by a fellow who was considering putting together a Lunar X-Prize team. I tried to help him by answering questions. He had some good ideas but was grossly over optimistic in what it would take to land a package on the moon. We never put together a final budget because it was realized very quickly that the cost was soaring past his expectations for fund raising. It will definitely take some very deep pockets if someone is to pull this off.
The launch is certainly a very big part of the cost. Piggybacking with another payload still isn’t cheap because two payloads mean you need a bigger and more expensive rocket. The only way it could be done on the cheap is it you could get the company launching the primary payload to pick up most of the cost in exchange for the publicity and advertising. Our conversations never matured to the point that we seriously investigated this possibility.
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 22, 2010 10:37:49 GMT -4
And it was hardly the first time that a rocket built by a commercial company had entered space. Practically every US rocket was built by a commercial company. The big difference with SpaceShipOne was that it was privately funded.
|
|
|
Post by ka9q on Feb 22, 2010 14:11:01 GMT -4
The big difference with SpaceShipOne was that it was privately funded. Sure. And if they'd emphasized this point I wouldn't have minded so much.
|
|
|
Post by ka9q on Feb 22, 2010 14:18:12 GMT -4
The launch is certainly a very big part of the cost. Piggybacking with another payload still isn’t cheap Since about 1980 I've been involved with AMSAT, a worldwide group of radio hams who build their own satellites and get them launched as piggyback payloads. The vast majority of amateur satellites are in LEO because that's where the free (or cheap) launch opportunities are. Right now I'm working on one that will be kicked out the airlock of the ISS. Going to the moon obviously takes more than that. At a minimum, you need a launch into the right orbital plane. Then you need enough energy to nearly escape the earth (e ~0.97), plus more energy to brake at the other end (dv ~= 2.38, the moon's escape velocity). Most amateur satellites carry no propulsion, but a few have (Oscars P3A, 10, 13, 40). And I have to tell you, we haven't been terribly good at it.
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 22, 2010 15:25:06 GMT -4
One of the things I worked on when investigating this was estimating the minimum mass needed to accomplish the Lunar X-Prize objectives. I don’t recall the numbers, but by the time I figured in the bare minimum in instrumentation, communications, guidance, attitude control, etc. and then added a main propulsion system and propellant for TLI and landing, we ended up with a pretty good sized payload that had to be put into Earth orbit. It’s not something that can just be tossed in at little additional cost. Adding so much extra payload would likely require upgrading to a bigger class of rocket, which costs millions of dollars.
For instance, suppose we’re launching on an Atlas V. Generally speaking, each additional ton of payload requires the addition of one solid rocket booster. I believe the cost of an Atlas V SRB is somewhere in the neighborhood of $5 million.
Of course piggybacking has its own problems. If someone did allow you to piggyback on their launch vehicle, the timing and orbit of the primary payload is surely going to take precedence. Chances are the launch window and orbit of the primary is not what is required for the secondary payload. You may have to make orbit modifications and/or loiter in orbit for an extended time before proceeding to the moon.
|
|
|
Post by ka9q on Feb 22, 2010 17:38:57 GMT -4
Exactly. Hams have been able to take advantage of piggyback opportunities because we're not trying to provide a 24/7 communications utility. Even LEO is better than nothing, especially since there's more than one satellite.
As I said, most of these satellites have been passive, without propulsion. Many don't even have active attitude control; they use a bar magnet to line up with the local magnetic field.
Four of our satellites were designed for high altitude. They carried propulsion systems intended to get them out of a geostationary transfer orbit and into something resembling a Molniya orbit. All were launched on Ariane.
The first, Phase III-A, went down with the launcher in May 1980.
The second, Phase III-B/Oscar-10, was launched in 1983. The upper stage of the launcher collided with us within a minute after separation because whoever programmed the launcher forgot to repoint it before venting the excess propellants through the engine nozzle. The bad sun angle nearly killed us, but we recovered. A few months later we fired the kick motor for what was supposed to be the first time, but a wiring error in the timer caused it to fire longer than intended. The long burn cooled the helium pressurization tank so much that the seal failed and lost pressurization, so that was the last burn. The computer memory failed from radiation exposure about 3 years later, but it's still in orbit and the transponder still operates from time to time when the sun is at the right angle.
The interesting thing about that one is that it orbits well within the Van Allen belts, so it gets a pretty hefty radiation dose. I estimated ahead of time that it might have a 3 year lifetime based on a 3 kilorad dose limit to the RAMs. The dose per belt crossing is very similar to the dose that the Apollo missions got on each of their belt crossings. And it drives the Apollo hoax believers nuts when I tell them that I have personal experience with a spacecraft that was exposed to radiation in the VAB, and that it got just the amount of radiation that I'd expect Apollo to have gotten.
The third, Phase III-C/Amsat-Oscar-13, was launched in 1988. This time they got the propulsion system working properly. Unfortunately, orbital resonances from the moon and sun caused the perigee to enter the atmosphere in 1996. But we got a good 8 years of operation out of it.
The fourth, Phase III-D/Amsat-Oscar-40, was launched in November 2000. This was a much larger and more ambitious spacecraft, and once again the propulsion gremlins hit us. A vent cap on the bipropellant engine was left in place, and it caused the oxidizer valve to remain open after the first kick motor firing. The N2O4 quickly ate through the hot combustion chamber, entering the insides of the spacecraft. When the oxidizer was depleted, the helium that was pressurizing it followed the N2O4, blowing off the bottom panel and ripping out a wiring harness that ran along it. Amazingly, one transponder continued to function for a couple of years until the whole thing went quiet.
The point of all this is that space is a very unforgiving place, especially where propulsion is concerned. Even the simplest mistakes can be fatal. When you consider all of the things that have to happen to soft-land a spacecraft on the moon -- the pyro firings, the motor burns, etc -- and that every one of them has to happen exactly as planned at the right time, something like Apollo becomes even more impressive.
There's a reason all this stuff is called "rocket science".
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 23, 2010 20:16:12 GMT -4
Assuming the piggyback option doesn't work....
I’ve looked at some of my notes from a couple years when I worked on this issue. I see that my delta-v budget was 6,450 m/s, which included 3,150 m/s for TLI and 3,300 m/s for midcourse corrections, terminal deceleration, and design margin. My assumption is that TLI is performed by a separate stage that is jettisoned after use. The remaining maneuvers are performed by a propulsion system integral to the lander.
Assuming a hypergolic pressure-fed propulsion system, a specific impulse of about 290-320 seconds is typical, so let’s use 310 seconds. I also generally figure a propellant mass fraction of about 0.9 for this type of system, i.e. 90% of the propulsion system mass is propellant.
Based on the above, the following is a mass breakdown:
Mass in Earth Orbit = 1.0000 TLI stage propellant = 0.6452 TLI stage dry mass = 0.0717 Spacecraft propulsion system propellant = 0.1875 Spacecraft propulsion system dry mass = 0.0208 Lunar payload = 0.0748
Therefore, only 7.48% of the mass inserted into Earth orbit is landed on the Moon as useable payload.
One of the preferred partners for the Lunar X-Prize is Space-X, which is offering a 10% discount on its launch services. Space-X’s list prices for low earth orbit (LEO) insertion are as follows:
Falcon 1 $8.9 million for 420 kg into 185 km orbit.
Falcon 1e $10.5 million for 1,010 kg into 185 km orbit.
Falcon 9 $49.5 million for 10,450 kg into 28.5o orbit from Cape Canaveral. $49.5 million for 8,560 kg into polar orbit from Kwajalein. Both of the above are reduced to $44 million if <80% capacity is used.
Taking the 10% discount and using the 0.0748 factor for lunar payload, we have the following (Cape Canaveral option used for Falcon 9):
Falcon 1 can launch 31.4 kg of lunar payload for $8.01M ($255,000 per kg) Falcon 1e can launch 75.5 kg of lunar payload for $9.45M ($125,000 per kg) Falcon 9 can launch <625 kg of lunar payload for $39.60M (>$63,000 per kg) Falcon 9 can launch 782 kg of lunar payload for $44.55M ($57,000 per kg)
I believe Falcon 9 is still in development and hasn't flown yet.
Of course the payload has to include not only the rover required for achieving the X-Prize objectives, but it must also include everything needed for the mission aside from the propulsion systems. This includes subsystems for the spacecraft structure, computer, power, communications, navigation, attitude control, etc. Taking all this into consideration, the 31.4 kg almost certainly isn’t enough capacity and possibly not the 75.5 kg either. Unless someone can figure out a really ingenious way to miniaturize the spacecraft, the cost for a dedicated launch is likely going to be well in excess of $10 million. Note that the first place prize is only $20 million with the possibility of another $5 million in bonuses. Clearly there are launch services other than Space-X, but none of them are going to be cheap.
An alternative to launching into LEO is having the spacecraft injected into a geostationary transfer orbit (GTO). The launch vehicle, therefore, provides most of the delta-v needed to get to the Moon. I calculate that only about 700 m/s is needed for TLI from GTO. This could probably be accomplished with an enlarged spacecraft propulsion system rather than having a separate TLI stage. In this case, the total spacecraft delta-v is increased to 4,000 m/s. The mass breakdown is now,
Mass in GTO = 1.0000 Spacecraft propulsion system propellant = 0.7317 Spacecraft propulsion system dry mass = 0.0813 Lunar payload = 0.1870
Space-X’s pricing for GTO insertion is,
Falcon 9 $49.5 million for 4,540 kg into 28.5o orbit from Cape Canaveral. $49.5 million for 4,680 kg into 9.1o orbit from Kwajalein. Both of the above are reduced to $44 million if payload <3000 kg.
We now have,
Launch <561 kg of lunar payload for $39.60M (>$70,500 per kg) Launch 849 kg of lunar payload from Cape Canaveral for $44.55M ($52,500 per kg) Launch 875 kg of lunar payload from Kwajalein for $44.55M ($51,000 per kg)
Edited to correct TLI from GTO delta-v from 600 m/s to 700 m/s.
|
|
|
Post by LunarOrbit on Feb 24, 2010 0:39:19 GMT -4
|
|
|
Post by ka9q on Feb 24, 2010 4:45:00 GMT -4
An alternative to launching into LEO is having the spacecraft injected into a geostationary transfer orbit (GTO). What about the launch window? When you fly piggyback, beggars don't get to be choosers. GTOs generally put the major axis roughly parallel to the sun line so that a spinning spacecraft in the proper attitude for its apogee kick burn will have the sun normal to the spin axis for proper power and thermal control. (Although most communication satellites are now 3-axis stabilized in operation, they usually spin with their panels folded until after motor firing.) That implies orbital injection (and perigee) at local midnight or local noon, giving two launch windows per day. The night launch is usually preferable to avoid long apogee eclipses, especially in the spring and fall. The argument of perigee is biased so that the earth oblateness perturbations will cause it to reach 180 degrees, with apogee occurring over the equator, at the desired time of kick motor firing. The moon, of course, doesn't enter into any of this. To go to the moon efficiently in a single impulse from GTO requires that the moon be in your orbital plane, along the extension of your major axis past apogee, when you reach lunar altitude after TLI at perigee. How long do you have to wait for this to happen after a randomly timed launch into GTO on, say, Ariane? Is its lower inclination helpful? I discovered long ago (rediscovered, since it was already well known) that it's often cheaper to make a big plane change by first increasing apogee, even if you have to drop it back down later. The higher your apogee the lower the velocity that you have to push in a different direction. Since you're already going to a very high apogee, maybe you could do your "TLI" without the moon in position. (The eccentricity is high, ~0.97, but it's not an earth escape trajectory.) At first apogee, you do a deep space maneuver that sets you up to reach the moon on a subsequent apogee. Does your analysis involve something like this? Or did you assume that the GTO was already chosen with an efficient TLI in mind? If you're the primary customer on a launcher normally used to reach GTO, the easiest thing to do would be to launch at the time of the month when the moon is near zero declination, and at a time of day that puts perigee (and TLI location) on the antipode from where the moon will be when you get there in a few days. You might have to deal with odd sun angles during cruise and/or landing, but that's about it. I found some early Apollo mission planning material with a very good explanation of the rationale for the use of a parking orbit, even though it cut into payload capacity and required the development of a restartable engine for the S-IVB. Basically, it was essential if they were to have practical launch windows from Kennedy's latitude that could get them to a site on the near side of the moon, in local morning, before the end of the 1960s. Planning is a lot easier if you can relax any of those constraints, particularly if you're willing to wait a very long time to get there.
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 24, 2010 9:52:34 GMT -4
What about the launch window? When you fly piggyback, beggars don't get to be choosers. My previous post was assuming we’re not piggybacking. The analysis was based on paying the full launch cost and being the only payload. Given this condition, the launch buyer should be able to have the GTO designed as necessary to meet the TLI requirements. Does your analysis involve something like this? Or did you assume that the GTO was already chosen with an efficient TLI in mind? If you're the primary customer on a launcher normally used to reach GTO, the easiest thing to do would be to launch at the time of the month when the moon is near zero declination, and at a time of day that puts perigee (and TLI location) on the antipode from where the moon will be when you get there in a few days. You might have to deal with odd sun angles during cruise and/or landing, but that's about it. My analysis of the GTO option was just a very quick back of the envelope type of thing. My basic assumption was that the perigee of the GTO is near the antipode* of where the Moon will be at arrival. TLI, therefore, is simply an altitude change to raise the apogee. Going back over my numbers I just discovered I made an error in calculating the TLI delta-v. It should be about 700 m/s instead of 600 m/s. I’m going to edit may previous post to correct for this. (edit) * The perigee should actually be about 10 o east of the antipode because the spacecraft's true anomaly when it intercepts the Moon is about 170 o.
|
|
Bob B.
Bob the Excel Guru?
Posts: 3,072
|
Post by Bob B. on Feb 24, 2010 11:23:46 GMT -4
It's good to read they're getting that close. I had lost track of the time table and really had no idea when the first launch would be.
|
|
|
Post by ka9q on Feb 25, 2010 0:54:05 GMT -4
The perigee should actually be about 10 o east of the antipode because the spacecraft's true anomaly when it intercepts the Moon is about 170 o. Right, though the exact value depends on the eccentricity of the lunar transfer orbit. In a manned mission, there's a tradeoff between the delta-V required for the TLI and lunar capture burns and the time of flight. You can minimize these delta-Vs by setting apogee to just barely reach lunar altitude, but at the expense of a longer trip requiring more consumables (water, food, oxygen, hydrogen for fuel cells). I don't remember the apogee of the typical Apollo transfer ellipse but it was somewhat past the moon for this reason. For a robotic mission you'd probably want to use the lowest apogee that will get you to the moon and so this angle between the TLI midpoint and the antipode would probably be smaller. Piggyback opportunities on GTO launches do exist -- all four AMSAT Phase III spacecraft, for example -- so maybe it would be worth looking at what it would take to get to the moon from an arbitrary GTO without a lunar constraint on the launch window. I think the high altitude maneuver is probably the way to go, not only to reduce the cost of an arbitrary plane change but also to reduce the number of VA belt passes. Ariane GTOs with their especially low inclinations tend to go through the densest parts of the inner belts.
|
|