I have a question about heat transfer and cooling/heating of sattelites.In many sources, especially in main stream media, it sounds like an unbelievably big problem, that in space, the sun side is so much hotter than the shadow. And they talk about huge figures hundreds of degrees celsius of temperature difference.On the other hand, my common sense tells me, that because of the low pressure there is probably very little convection and therefor there is only radiation for heat transfer.Therefor, i'd think that though the air moluecules in LEO are technically really hot, they transfer very little energy to a spacecraft. So how is the cooling/heating acomplished? I'd guess that making the spacecraft reflective on the outside would block most of the energy coming from the sun, an because of little convection the spacecraft should cool very little when provided with basic insulation against heat transfer to the outer hull (which would remove the energy by way of infrared radiation). Then I'd guess the electronics or other systems (e.g. a human in a spacecraft) would provide enough heat to keep the sacecraft from freezing. So what, would be left would be to provide a way of radiating exactly as much energy out of the spacecraft as needed o keep constant temperature, so how would I do that?
I have a question about the corrosiveness of N2O4. I have read that adding a few percent of NO reduces the corrosiveness and that the resulting mixture is called MON, which stands for mixed oxides of nitrogen. Various mixture ratios are in use.So how much of a difference does this make and how much of a problem is this corrosiveness? I'm particularly interested in the effect on the near-term feasibility of orbital hypergolic propellant depots. How long can you realistically store MMH/MON in space before corrosion renders your depot inoperable?
It isn't a problem. Comsats have serviceable lives of more than 15 years with MON. Cassini is now over ten. MGS was over ten
I've just found a scanned copy of a Rocketdyne nitrogen tetroxide handling manual from 1961 :-)
Propellants MR dp (kg/L) ve (m/s) Id (Ns/L)O2/H2 5.0 0.3251 4455 1448O2/H2 6.0 0.3622 4444 1610O2/H2 7.5 0.4120 4365 1798O2/CH4 3.6 0.8376 3656 3062O2/C2H6 3.2 0.9252 3634 3362O2/C3H8 3.1 0.9304 3613 3362O2/C3H4 2.4 0.9666 3696 3573O2/RP–1 2.8 1.0307 3554 3663O2/C7H8 2.4 1.0954 3628 3974HTP/C3H4 6.5 1.2553 3319 4166HTP/RP–1 7.3 1.3059 3223 4209HTP/C7H8 6.6 1.3496 3288 4437
strangequark, I wrote two papers on this subject which are attached. In a first stage performance is limited by propellant volume. I show in my paper that the criteria for choosing a fuel for a first stage is its impulse density Id, equal to the product of the propellant density (kg/L) and the exhuast speed (m/s). A list of propellants is given belowPropellants MR dp (kg/L) ve (m/s) Id (Ns/L)O2/H2 5.0 0.3251 4455 1448O2/H2 6.0 0.3622 4444 1610O2/H2 7.5 0.4120 4365 1798O2/CH4 3.6 0.8376 3656 3062O2/C2H6 3.2 0.9252 3634 3362O2/C3H8 3.1 0.9304 3613 3362O2/C3H4 2.4 0.9666 3696 3573O2/RP–1 2.8 1.0307 3554 3663O2/C7H8 2.4 1.0954 3628 3974HTP/C3H4 6.5 1.2553 3319 4166HTP/RP–1 7.3 1.3059 3223 4209HTP/C7H8 6.6 1.3496 3288 4437HTP is 98% hydrogen peroxide
Quote from: Steven Pietrobon on 03/19/2009 04:10 amHTP is 98% hydrogen peroxideIt is an explosion hazard, and is more expensive than LOX.
HTP is 98% hydrogen peroxide
Then why all SSTO projects seems to use O2/H2?
Quote from: gospacex on 03/20/2009 10:40 pmQuote from: Steven Pietrobon on 03/19/2009 04:10 amHTP is 98% hydrogen peroxideIt is an explosion hazard, and is more expensive than LOX.So are solid rocket motors, which are widely used.
Provided that proper precautions are followed, HTP is safe to use, as shown by prior British and US experience.
Propellant cost is a small fraction of launch costs. The 15% performance increase more than makes up for any propellant cost increase.