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I'm putting the Raptor's SL TWR at 200:1, which makes its mass 1.55 tonnes. There are nine of these on the ITS Tanker and Spaceship, but I'll go ahead and kick the engine mass up to a total of 15.5 tonnes to account for the mass of the vacuum nozzle extensions.

The dry mass of the ITS tanker is projected at 90 tonnes, which includes the monocoque tanks/body, the intertanks, the RCS thrusters, the TPS, the engines, and the landing legs. Subtract the engine mass, and that's 74.5 tonnes for the tanker's 2,500 tonnes of propellant, or a mass ratio of 33.6:1.

To account for some square-cube losses, I projected my structural mass ratio at 29:1 to get a base structural mass of 4.88 tonnes for the 141 tonnes of propellant I wanted. I added back 1.28 tonnes for the two dev Raptor engines and 405 kg for the eight landing thrusters.

I didn't include payload encapsulation in my overall dry mass because I'd want a monocoque pressure vessel for the crewed version. The payload fairings on the Falcon 9 mass 1.75 tonnes total so that's a good estimate for the cargo bay mass. The actual downmass to Mars on a FH-expendable direct ascent is probably closer to 13 tonnes, then. That's reserving 2.9 tonnes of propellant for the landing burn, which gives about 448 m/s of dV. That should be enough, I think; if I need more it would decrease the downmass slightly but not by much.

This makes sense, except that you're comparing to the ITS tanker. As I pointed out already, the tanker isn't designed to fly to Mars. It most likely doesn't have TPS for an interplanetary entry, which will probably mass about twice as much as TPS for Earth LEO entry. It probably doesn't have significant solar power, operating mostly off batteries for the few hours it's in LEO. And it probably doesn't have active or passive cooling to keep the landing fuel from boiling off during a 6 month transit, which is how long a mass-optimized transfer to Mars takes. I'd figure dry mass of at least 5% of wet, before payload.
See, this is just the sort of engineering counter-analysis I was hoping for with this thread! Thanks. The points regarding high-velocity TPS, cooling, and power are good ones. I hadn't considered that. However, even if there is some dry mass growth on my upper stage, it's not catastrophic. Payload with full reuse is already very high, and this is not a SSTO where a 10% increase in dry mass results in negative payload fraction.

It was difficult to get a good estimate of tankage/structural fraction on the ITS Spaceship, because I don't know how much to subtract for the crew cabin and cargo hold. But after some consideration, I think we have a way we can at least place an upper bound on it.

The ITS Tanker and ITS Spaceship have identical external form lines, so we can say with relative confidence that if the tanker DID have all the TPS and cooling and power systems of the spaceship, it would still have a lower dry mass than the spaceship, since the crew cabin and ECLSS and cargo hold are heavier than monocoque tanks of equivalent size. So we can divide the structural dry mass of the ITS Spaceship (150 tonnes minus 15 tonnes of engine) by the propellant capacity of the ITS Tanker (2,500 tonnes) to get a structural/tankage/body fraction of 5.4%. This, I think, is generous and conservative. Even if there are square-cube losses to factor in, they are greatly outweighed by the fact that we are including the crew cabin and cargo hold and ECLSS dry mass, which we definitely don't need. And the ITS Spaceship dry mass already includes the additional mass of landing legs, control surfaces, RCS, and additional outer mold lines, so I don't need to add anything in to compensate for that.

Add it up, and the dry mass of my second-stage vehicle climbs from 6.56 tonnes to 9.32 tonnes, which means payload is simply reduced by a little under three tonnes. However, my reusable upper stage already beats the Falcon 9 upper stage for all LEO payloads by more than 3 tonnes when lofted on F9FT and it beats the Falcon 9 upper stage for all payloads period when lofted on Falcon Heavy...and that, again, is with full reuse. So it is by no means a dealbreaker.

I would certainly support refueling. But there is really very little additional structure or propulsion required for horizontal landing. Composites have better biaxial rigidity than aluminum, you need biaxial strength for the biconic re-entry anyway, and you need the thrusters regardless because of TWR issues.

You might still need an arm or crane or some other mechanism for getting the payload out, but it is a LOT easier than trying to do it with a capsule.
What Isp and TWR are you assuming for the thrusters? I haven't seen any reliable info on them other than fuel and thrust. I'd be surprised if it's over 340 sec for a low pressure methalox RCS thruster.

Vertical landing allows the more efficient Raptors to null 99% of the velocity with thrusters handling only terminal thrust. And you only need about 1 g worth of terminal thrust instead of 3 to 4 gees. So with vertical landing you some structure mass, some thruster mass, and some fuel mass thanks to being lighter AND more efficient.
Ah, but here's the problem: we're dealing with an upper stage, not a first stage; we can't use vacuum-expanded dev Raptors for landing at sea level. So for Earth entry we must have the thrusters handle the landing burn.

Not to mention that terminal velocity on Earth will be far lower for a blunt lifting body than for a tail-first descent, and the structural mass really doesn't go up because, again, you're already using a biconic re-entry.

Good catch about the isp of the landing engines. They are ten-metric-tonne thrusters, which is overpowered for RCS on a craft of this size, but great for landing. I did already factor in a decreased thrust due to SL expansion. Ten metric tonnes of vacuum thrust is 98.1 kN, but since these are going to need to be sea-level-expanded, underexpansion will decrease vacuum thrust to around 92 kN and SL thrust will be around 86 kN. 

I had been projecting vacuum isp for the landing engines at 360 seconds and SL isp at 330 seconds, forgetting that pressure-feeding decreases their efficiency. But by how much? The pressure-fed methalox thruster tested by NASA recently had a projected vacuum specific impulse of 348 seconds; for SL-expanded thrusters this translates to a vacuum isp of around 326 s and a SL isp of 305 s.

So that's going to increase landing fuel somewhat. But again, I have a LOT of margin, so it's not catastrophic.
Why does the second stage need to have 2 engines? One single Vacuum Raptor-dev (assuming it is at 1/3rd of the thrust, and doesn't have much less TWR compared to the full sized one) could push the full second stage well enough, you don't need a 1 TWR (on earth) for a second stage or even a mars/moon lander.
I tried it with a single dev Raptor first, but I didn't like it. TWR would be under half a gee for crewed missions, which is not really ideal; you end up needing an aggressively lofted trajectory, which really hurts your ability to safely abort. And FFSC gives you a really good TWR; removing one of the dev Raptors would only save you 640 kg of dry mass while cutting your vehicle TWR in half.

And it will need at least 1000 m/s to land. Terminal velocity is going to be Mach 2-ish.

I may be wrong but since it is both a lifting body and it is quite light (<15 tons), i've read that in this case mars EDL delta v could be as low as 800 m/s.
You need 1000 m/s for a capsule, but not for a biconic lifting body. Even 800 m/s is an overestimate, I think. The Mid-L/D MAV concept allocates 650 m/s for the landing burn, and my upper vehicle would be much fluffier (higher-drag, lower-mass) than the Mid-L/D. Speaking of which, the re-entry profile is quite similar:

I designed this from the ground up as an Earth-based reusable upper stage, but the adjustments to make it work equally well for Mars or the moon are really, really minor. The few tweaks merely make it a more capable spacecraft for Earth missions.

How does the lander get vertical again, for takeoff? Once fueled the landing engines will not be enough to get it off the ground will they?
Q&A Section / Re: Thorium Fuel Engine
« Last post by sanman on Today at 10:36 AM »
I think you would be right about that except for the peculiarities of the cycles. the thorium/u233 one can be done on the fly (I think.) I think i recall that the Thorium U233 cycle is safer environmentally and also from a proliferation perspective because you could make U233 at the burn rate for the reactor so there never was much of it at any one time and the uranium was unsuitable for practically sized fission or fusion bombs.

That's wrong - there would be more than enough U233 to build a bomb. Reactors have hundreds of tons of fuel in them. 1% burnup of thorium in such a reactor would yield some ~1 ton of U233 - enough for a hundred nukes.

The problem (for bomb makers, that is) is that thorium reactors also generate U232, which has very undesirable radiological properties, contaminating that U233.

U233 is a strong gamma-radiator, which makes it too deadly for anyone to isolate into a bomb and carry around.

Likewise, you wouldn't want to be enriching it aboard a spacecraft either, without a lot of shielding and distance for safety.

Bur Carlos Rubbia's idea for an Accelerator-Driven Reactor system might be useful for nuclear propulsion one day, if the mass requirements can be kept down. Using fuels like Thorium or U238 which are merely fertile not fissile would at least avoid any danger of runaway chain-reaction meltdown. As soon as the accelerator beam is cut, the nuclear reaction is immediately quenched or shut off.
Space Science Coverage / Re: NASA - Cassini updates
« Last post by Bubbinski on Today at 10:33 AM »
That top that the polar hurricane in the middle of the hexagon?
Why is there such a fuss about this "record"?   It pales into insignificance when compared with the Russian duration record of 804 days by Gennady Padalka.
Has there ever been indication how much propellant can be carried on board Tianzhou for transfer to Tiangong or in future Tianhe 1?   Is this mass in addition to the 5.5 tonnes of supplies which has been quoted?
I think you're still going to have the same old problem of changing suit volume during joint flexion.
Blue Origin / Re: Blue Origin Update and Discussion Thread
« Last post by Chasm on Today at 10:02 AM »
Another talk, this time by A.C. Charania at the Caltech Space Challenge

Skipping through the talk
16:14 has rough pricing for the previously announced experiment lockers.
19:37 this slide now includes a half cut of NG.
22:25 Blue Moon wall of text
In the meantime, apart from testing the RF amplifier, I have been testing some loops (picture) with a special cavity for that purpose and the Windfreak SynthNV signal source/network analyzer. I measure the transmission (S21), so I make them in pairs. Next week I'll try the loops as Zhang (2013) [1]described, see last figure.
Great device, this SynthNV, example of a measurement in the next posting where it is compared with the a measurement with a professional network analyzer.
The copperplates for the first frustum arrived, I have to buy some tools still... The electronics of the torsion balance isn't working well yet. Still lots of things to do. Could be a matter of weeks before everything is working.
I haven't bought any batteries yet. And I haven't decided/figured out how to do the communication. If the balance is opereational, I will test what the influence of 3 tiny copper wires near the torsion wire is. The USB connection will only need these wires. If there is any disturbance of the balance, I will have to use a wireless option.

[1] Hai Zhang et al., Research on Novel Loop Antenna in Microwave Cavity Measurement of Permittivity, Int. J. of Information and Electronics Engineering, Vol. 3, No. 4, July 2013.

Peter, since you're experimenting with antennas (and btw cavities), check out the attached PDF

Tianzhou-1 to complete first propellant refueling mission

China's first space cargo ship, the Tianzhou-1, is about to refuel the Tiangong-2 space lab for the first time, using "in-flight refueling" techniques.
The spacecraft blasted off from the Wenchang Launch Center on April 20 and successfully docked with the space lab two days later. During its two-month orbit, the cargo ship is expected to refuel the space lab three times. The refueling procedure takes 29 steps and lasts for several days each time. If the Tianzhou-1 completes its objective, China will become the third country, after Russia and the United States, to master the technique of refueling in space. Chinese engineers consider it a major breakthrough for future space station operations.
Image showing umbilical cables as I install the Windfreak NV software on the torsional pendulum control computer. 12V 8.5A power supply, HDMI, USB spectrum analyser, and USB data.

Sounds like you cleaned things up quite a lot :) ! This may help speeding up the characterization/calibration and reducing the noise and also help removing the (darn) ground loops; way to go !
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