Henry:
This is interesting; could you be explicit about how you calculate
effective Isp for rentry system mass?
Bill
On Sun, Feb 11, 2018 at 8:04 PM Henry Spencer <hspencer@xxxxxxxxxxxxx>
wrote:
On Sat, 10 Feb 2018, Ed LeBouthillier wrote:
The biggest obstacle I see remaining, is the ability to safely and
reliably recover an orbital stage without excessive TPS mass.
Do you think that something like in-orbit refueling of these stages
might mitigate the entry heat problem? ...just think if you're in orbit
at ~8 km/s and you can drop your entry velocity to 1.5 to 2 km/s before
entering the atmosphere...
Even setting aside questions of launch abort and missed rendezvous --
assuming everything always works perfectly -- how do you propose to get
the fuel *up there*? The fuel needed to reduce entry velocity that much
will weigh (at least) several times the upper stage's payload capacity.
And if you've got some cheaper way of getting fuel up there, why do you
need to recover rocket stages?
If you can do significant delta V deceleration on the way down, you
don't have to rely so heavily on TPS.
The key word is "if".
Maybe propellant might be cheaper than TPS mass and refurbishing.
TPS produces a certain delta-V (of deceleration) for a spacecraft of a
given mass at the cost of a certain TPS mass, which means you can assign
it an effective Isp. The Apollo spacecraft's heatshield had an effective
Isp of about 8000s. And that was an overbuilt heatshield; a less
conservative design with more modern materials might get about 20000s.
Which is to say, each kilogram of heatshield is worth 10-100kg of
propellant. Nothing short of advanced nuclear propulsion will make
propellant competitive with that.
Henry