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  #51  
Old January 8th 05, 07:52 AM
Henry Spencer
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In article ,
Paul Hovnanian P.E. wrote:
As far as introducing some sort of coolant into the skin, you'd have to
determine the amount of heat generated. A good estimate is that all of
the shuttle's kinetic energy is converted into heat at the skin
interface...


No, that's a lousy estimate. There is more than enough kinetic energy in
a vehicle reentering from orbit to vaporize the entire vehicle, no matter
*what* it's made of. The only reason reentry is practical at all is that
with careful design (notably, a very blunt leading surface) only a *tiny*
fraction of the heat actually reaches the skin.

Then, given the specific
heat of various coolants, calculate how many tons of coolant you'd have
to haul up at launch and throughout the mission in order to cool the
skin.


The fast answer will be: too much. The shuttle makes a prolonged reentry
with quite high total heat loads, which would require an awful lot of
expendable coolant. Expendable coolants work much better with Apollo-style
lifting-capsule reentries, which are short and sharp, with higher peak
heating rates but much lower total heat loads. That's what an ablative
heatshield is: solid expendable coolant.

The shuttle is cooled by transferring the heat of friction to the
surrounding atmosphere which carries it away.


No, the air is in general hotter than the surface. Almost all of the heat
is spread into the air and never reaches the surface. What does reach the
surface is to some small extent soaked up (heat continues to soak through
the tiles for quite a while after reentry -- the cargo-bay temperature
actually peaks *after* landing), but mostly radiated. Conveniently, even
fairly hot air is essentially transparent to radiated heat. But that does
require a very hot surface -- radiated heat flux scales with the fourth
power of temperature -- with very good insulation behind it.
--
"Think outside the box -- the box isn't our friend." | Henry Spencer
-- George Herbert |
  #52  
Old January 13th 05, 02:33 AM
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In article ,
(Henry Spencer) writes:
In article PzNsd.141967$V41.95361@attbi_s52,
Tommy wrote:
a better question is why didn't they make the shuttle out of titanium
rather than aluminium.


NASA did in fact consider it, with the higher cost and greater weight of
titanium structure almost completely balanced out by the ability to relax
the requirements on the tiles a little. The decision was ultimately made
on secondary issues, notably the US's limited titanium supply.

It would not have made any great difference to Columbia. Titanium is not
*that* much better; the conditions in Columbia's wing were far beyond the
working limits of *any* reasonable structural metal.


Thanks for an answer to one of my pet post-Columbia questions.
Recognizing that a complete titanium rework was out of the question,
I was wondering about using titanium for just the leading structural
member. Comparing Al to Ti the melting points are 669C vs 1660C,
quite a bit different. I guess that says nothing of how much strength
either loses as they approach the melting point. How hot do they
estimate it was inside the leading edge?

The second part of my pet post-Columbia idea was to flow air through
the leading-edge space. Arrange some sort of exhaust port at the
wingtip to create some negative pressure. Open an airflow from the
leading-edge wing-root into the cargo bay. Open some sort of positive-
pressure inlet from the coldest (I know no point is really cold.)
exterior into the cargo bay. The net airflow would be from cold-side,
probably somewhere on top near the back, through the cargo bay, and
out the leading edges. The idea was to dilute hot leaks from the
leading edge, itself. For small leaks I suspect it might help, but
for one the size on Columbia, I don't know if anything could have.

Dale Pontius
  #53  
Old January 13th 05, 05:02 PM
Henry Spencer
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In article ,
wrote:
It would not have made any great difference to Columbia. Titanium is not
*that* much better; the conditions in Columbia's wing were far beyond the
working limits of *any* reasonable structural metal.


I was wondering about using titanium for just the leading structural
member. Comparing Al to Ti the melting points are 669C vs 1660C,
quite a bit different. I guess that says nothing of how much strength
either loses as they approach the melting point.


Exactly. The maximum *working* temperatures -- temperatures at which they
still dependably retain most of their strength -- are much, much lower,
and if I recall correctly, only a couple of hundred degrees apart. This
is significant for aircraft but a very minor advantage for space reentry.

If memory serves, the strength loss in aluminum is fast enough that
there's little advantage in just making it thicker to reduce the working
stresses. With titanium you can go a bit higher by doing that, but there
is of course a mass penalty.

How hot do they estimate it was inside the leading edge?


Don't remember the accident-report numbers, but RCC panels are used only
where temperatures exceed about 1250degC, and the stagnation points -- the
worst case -- on nose and leading edge are at about 1650degC.
--
"Think outside the box -- the box isn't our friend." | Henry Spencer
-- George Herbert |
  #54  
Old January 14th 05, 01:46 AM
Paul F. Dietz
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Henry Spencer wrote:

Don't remember the accident-report numbers, but RCC panels are used only
where temperatures exceed about 1250degC, and the stagnation points -- the
worst case -- on nose and leading edge are at about 1650degC.


The air entering the hole in the RCC was much hotter. The hottest
air at the shock is normally kept some distance from the vehicle,
but the hole let that hot air flow directly into the wing.

I've seen a figure of 8000 F, enough to erode even the RCC
to a thin edge at the boundaries of the enlarged hole.

Paul
  #55  
Old January 14th 05, 02:01 PM
Jan Vorbrüggen
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I was wondering about using titanium for just the leading structural
member. Comparing Al to Ti the melting points are 669C vs 1660C,
quite a bit different. I guess that says nothing of how much strength
either loses as they approach the melting point.


Exactly. The maximum *working* temperatures -- temperatures at which they
still dependably retain most of their strength -- are much, much lower,
and if I recall correctly, only a couple of hundred degrees apart.


Additionally, this was a dynamic event. AIUI, Ti has lower thermal
conductivity, so in such a case the Al would have the advantage of
being able to "cool" the impacted parts of the structure to the rest
of it, while the local temperature rise would be faster for Ti, further
reducing its advantage.

Jan
  #56  
Old January 31st 05, 03:43 PM
Andrew Nowicki
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Peter Fairbrother wrote:

Who needs Al or Ti? Carbon fibre is stronger than
either weight for weight. Let's make the entire
wing structure out of large closed cell CF/phenolic,
or RCC foam - or at least fill the spaces with
broken aerogel


Your idea makes sense, but it would be rather
expensive to retrofit the existing shuttle fleet
with new wings.

NASA improved insulation on the outside of the
external cryogenic tank, and they made uncertified
repair kit to fix the lost refractory tiles in orbit.
I believe that it would be good idea to protect
the fragile refractory tiles on the leading edges
with disposable plastic foam. A backup water cooling
system (like the fire sprinklers) would also help.
 




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