Shuttle maximum altitude
Mike Miller
regulation-limited? I've seen 600 miles, 500 miles (on a NASA page)
and 600 kilometers in a web search.
Mike Miller, Materials Engineer
Hallerb
>What's the US space shuttle's maximum altitude, both hardware- and
>regulation-limited? I've seen 600 miles, 500 miles (on a NASA page)
>and 600 kilometers in
Presently ground level, in Florida thats almost sea level.
Is that what you were looking for?
Hallerb
>What's the US space shuttle's maximum altitude, both hardware- and
>regulation-limited? I've seen 600 miles, 500 miles (on a NASA page)
>and 600 kilometers in a web sear
At ISS inclination its near its top altitude limit with payload.
I think the number varies based on that. With the ideal inclination and NO
payload it would be a lot higher
Dale
>Presently ground level, in Florida thats almost sea level.
>
>Is that what you were looking for?
Hey, it's the new "smart ass" Bob. I like it :)
Dale
Mike Miller
>
> I think the number varies based on that. With the ideal inclination and NO
> payload it would be a lot higher
How about 10000lbs payload and an ideal inclination?
Mike Dicenso
The ideal inclination is 28.45 degrees due east from Kennedy Space Center.
It's easy enough to calculate; just subtract 110 lbs (50 kg)of payload for
every nautical mile gained over 120 nm.
The maximum theoretical shuttle payload with the current surviving
orbiters is 63,500 lbs (28,863.63 kg). Assume a minimum duration mission
and crew, so that payload capacity is not affected by having to carry the
extra mass of crew and consumables, and you have your answer.
-Mike
Brian Thorn
>What's the US space shuttle's maximum altitude, both hardware- and
>regulation-limited? I've seen 600 miles, 500 miles (on a NASA page)
>and 600 kilometers in a web search.
600 miles is the theoretical maximum. Nobody wants to go much higher
than that unless they're going *much* higher than that, because of
trapped radiation in the van Allen belts. Getting there requires
almost no cargo in the payload bay, so the Shuttle hasn't gone above
the 330 miles of Hubble.
Brian
Kim Keller
"Brian Thorn" <btho...@cox.net> wrote in message
news:dd9oqvc7vt79hvfdb...@4ax.com...
> 600 miles is the theoretical maximum. Nobody wants to go much higher
> than that unless they're going *much* higher than that, because of
> trapped radiation in the van Allen belts. Getting there requires
> almost no cargo in the payload bay, so the Shuttle hasn't gone above
> the 330 miles of Hubble.
I've always heard the 600-mile limit was due to reentry limits on the TPS.
True, it's not fun to mess around in the Van Allen Belts, but that's not the
true reason for the altitude limit.
-Kim-
Bob Martin
> True, it's not fun to mess around in the Van Allen Belts, but that's not
the
> true reason for the altitude limit.
It seems to me that OMS fuel would be another limiting factor. The shuttle
doesn't really carry that much.
Jorge R. Frank
news:booi95$fig$1...@news-int.gatech.edu:
Correct. You'll hit the OMS limit first. Example: STS-109 (an HST servicing
mission) carried practically a full OMS load (25,064 lbm), and a light
payload (20,344 lbm). It launched due east from KSC (28.45 deg
inclination), and performed a direct insertion to a 308 n.mi. apogee. It
then spent 12,088 lbm of OMS prop (spread over four rendezvous burns) to
reach HST in a 317 n.mi. circular orbit. The deorbit burn used another
11,538 lbm. With 798 lbm unusable prop trapped in the OMS lines, that left
a whopping 936 lbm to protect for dispersions and contingencies.
This doesn't leave a whole lot of room for improvement. You can eliminate
the payload, top off the OMS tanks, use up your contingency reserve, use a
lighter orbiter (STS-109 was on Columbia, the heaviest orbiter), use up the
RCS propellant set aside for HST rendezvous, and it still won't get you
that much higher.
Of course, if you don't care about coming home, you can turn that deorbit
burn around and raise the orbit instead. Figure about 480 n.mi. if you
circularize, 320x640 n.mi. if you don't.
There was a proposal early in the shuttle program to develop a payload bay
OMS kit, but it never happened. All the orbiters still carry the flight
deck switches to control the kit. Up to three kits could be carried, each
with about half the capacity of the existing OMS tanks. That could get you
up to around 560 n.mi. circular, and still be able to come home.
--
JRF
Reply-to address spam-proofed - to reply by E-mail,
check "Organization" (I am not assimilated) and
think one step ahead of IBM.
Terrence Daniels
"Jorge R. Frank" <jrf...@ibm-pc.borg> wrote in message
news:Xns942FD099...@206.127.4.10...
> There was a proposal early in the shuttle program to develop a payload bay
> OMS kit, but it never happened. All the orbiters still carry the flight
> deck switches to control the kit. Up to three kits could be carried, each
> with about half the capacity of the existing OMS tanks. That could get you
> up to around 560 n.mi. circular, and still be able to come home.
What killed it? Budget constraints, or lack of need? (or both)?
Jorge R. Frank
news:m4Zrb.22742$9M3....@newsread2.news.atl.earthlink.net:
Both.
Michael R. Grabois ... change $ to "s"
wrote:
>"Bob Martin" <rell...@hotmail.com> wrote in
>news:booi95$fig$1...@news-int.gatech.edu:
>
>> It seems to me that OMS fuel would be another limiting factor. The
>> shuttle doesn't really carry that much.
>
>Correct. You'll hit the OMS limit first. Example: STS-109 (an HST servicing
>mission) carried practically a full OMS load (25,064 lbm), and a light
>payload (20,344 lbm). It launched due east from KSC (28.45 deg
>inclination), and performed a direct insertion to a 308 n.mi. apogee. It
>then spent 12,088 lbm of OMS prop (spread over four rendezvous burns) to
>reach HST in a 317 n.mi. circular orbit. The deorbit burn used another
>11,538 lbm. With 798 lbm unusable prop trapped in the OMS lines, that left
>a whopping 936 lbm to protect for dispersions and contingencies.
>
>This doesn't leave a whole lot of room for improvement. You can eliminate
>the payload, top off the OMS tanks, use up your contingency reserve, use a
>lighter orbiter (STS-109 was on Columbia, the heaviest orbiter), use up the
>RCS propellant set aside for HST rendezvous, and it still won't get you
>that much higher.
One other minor factor is that Endeavour, Discovery, and Atlantis have slightly
bigger OMS capabilities than Columbia did. I'm pretty sure the number 25,158 is
burned into some brain cell somewhere. So you have a lighter vehicle with maybe
100 lbs more OMS - but that'll only get about 4 fps extra delta-v or so.
A better example than 109 would be the previous HST mission, STS-103 on
Discovery.
--
Michael R. Grabois # http://chili.cjb.net # http://wizardimps.blogspot.com
"People say losing builds character. That's the stupidest thing I ever
heard. All losing does is suck. " -- Charles Barkley, 9/29/96
Jorge R. Frank
wrote in news:0h71rvk6pohctjlpj...@4ax.com:
> On Mon, 10 Nov 2003 20:30:23 -0600, "Jorge R. Frank"
> <jrf...@ibm-pc.borg> wrote:
>
>>"Bob Martin" <rell...@hotmail.com> wrote in
>>news:booi95$fig$1...@news-int.gatech.edu:
>>
>>> It seems to me that OMS fuel would be another limiting factor. The
>>> shuttle doesn't really carry that much.
>>
>>Correct. You'll hit the OMS limit first. Example: STS-109 (an HST
>>servicing mission) carried practically a full OMS load (25,064 lbm),
>>
>>This doesn't leave a whole lot of room for improvement. You can
>>eliminate the payload, top off the OMS tanks,
>
> One other minor factor is that Endeavour, Discovery, and Atlantis have
> slightly bigger OMS capabilities than Columbia did. I'm pretty sure
> the number 25,158 is burned into some brain cell somewhere.
That's what I meant by "practically" and "top off"... it isn't much, as you
pointed out.
> A better example than 109 would be the previous HST mission, STS-103
> on Discovery.
Right, but I couldn't find the Level A GR&C for that one online... :-)
Abdul Ahad
> >>> It seems to me that OMS fuel would be another limiting factor. The
> >>> shuttle doesn't really carry that much.
Here's a *what if* scenario: suppose you modified the Shuttle design
and filled the entire 65,000-lb cargo bay with fuel instead of
payload... to feed the OMS while in orbit. What would then be the
theoretical max. altitude for the optimum incline, assuming the entire
payload bay was offloaded in OMS burns?
Would that get us to at least half way to the Moon?!
Abdul Ahad
Chris Jones
I doubt it, but it's a pointless thought experiment anyway. The shuttle
is designed for leaving and returning to earth; too much of its mass is
devoted to that to make it at all a logical vehicle for trans-lunar
travel. I also doubt that its thermal protective system can withstand
reentry at coming-back-from-the-moon speed, so you'd need fuel to slow
down first to earth orbit speed, then reentry speed. All in all, a
gigantic bunch of bad engineering.
Jorge R. Frank
news:3416b228.03111...@posting.google.com:
>> >>> It seems to me that OMS fuel would be another limiting factor. The
>> >>> shuttle doesn't really carry that much.
>
> Here's a *what if* scenario: suppose you modified the Shuttle design
> and filled the entire 65,000-lb cargo bay with fuel instead of
> payload... to feed the OMS while in orbit.
OK, in the spirit of the "what if", I'll answer that. But keep in mind that
in the real world, you can't use all 65 klb for propellant; there has to be
some mass for tankage, including the helium tanks needed to pressurize the
OMS propellant. And of course, in the event of an ascent abort, you'd need
to dump most of that propellant to get down to a safe landing weight and
CG. Ick. Plus all the other stuff Chris Jones said. :-)
> What would then be the
> theoretical max. altitude for the optimum incline, assuming the entire
> payload bay was offloaded in OMS burns?
dv = g*Isp*ln(mo/mf) = 32.174*316*ln((200000+65000)/200000) = 2830 fps
Assuming you had already depleted your standard OMS tanks to get to a
320x640 n.mi. orbit, burning at perigee would raise apogee to around 3300
n.mi. You'd probably have to do it in several burns, of course... if you
tried to burn all at once, by the time you finished you'd no longer be at
perigee, and the effect of the burn would be to raise *both* apses, not
just apogee.
> Would that get us to at least half way to the Moon?!
The moon is about 200,000 n.mi. away, so no. Not even close. It's just
enough to get you into the really bad parts of the Van Allen neighborhood,
though.
Think of it this way: the Apollo CSM/LM stack had less than half the mass
of a space shuttle, and it took a Saturn S-IVB stage to send that from LEO
to the moon. And the S-IVB used much more efficient (but less dense)
propellants. If you want to send something the size of a shuttle to the
moon, you need to think bigger than the payload bay. Much bigger.
Abdul Ahad
> Think of it this way: the Apollo CSM/LM stack had less than half the mass
> of a space shuttle, and it took a Saturn S-IVB stage to send that from LEO
> to the moon. And the S-IVB used much more efficient (but less dense)
> propellants. If you want to send something the size of a shuttle to the
> moon, you need to think bigger than the payload bay. Much bigger.
A Shuttle orbiter weighs in the order of 80,000-kg and the payload
capacity is somewhere around 29,000-kg. So I guess that's insufficient
for a full TLI, but I wanted to know just *how* insufficient.
Without going into the full intricacies of the dynamical equations, is
there a very *broad* ratio of spacecraft mass to propellant mass,
assuming an average LOX/RP-1 type fuel for trans-lunar injection? I
read somewhere that a 300-kg mass orbiter in LEO will require 645-kg
of propellant mass (over and above the 300-kg existing spacecraft
mass) to escape on a trajectory of TLI to reach the Moon - assuming it
burns a hypothetical mixture of LOX/RP-1 in an engine adapted for
vacuum that burns in a single impulse at LEO for the TLI. On this
scale, it implies the Shuttle will need 172,000-kg of total propellant
mass to achieve TLI and be certain to reach the Moon!!!
Abdul Ahad
http://uk.geocities.com/aa_spaceagent/
Jorge R. Frank
> A Shuttle orbiter weighs in the order of 80,000-kg and the payload
> capacity is somewhere around 29,000-kg. So I guess that's insufficient
> for a full TLI, but I wanted to know just *how* insufficient.
> Without going into the full intricacies of the dynamical equations, is
> there a very *broad* ratio of spacecraft mass to propellant mass,
> assuming an average LOX/RP-1 type fuel for trans-lunar injection? I
> read somewhere that a 300-kg mass orbiter in LEO will require 645-kg
> of propellant mass (over and above the 300-kg existing spacecraft
> mass) to escape on a trajectory of TLI to reach the Moon - assuming it
> burns a hypothetical mixture of LOX/RP-1 in an engine adapted for
> vacuum that burns in a single impulse at LEO for the TLI.
Roughly speaking, that's correct. That comes out to a mass ratio of 3.15,
which results in a required Isp of about 290 s to achieve TLI from LEO. You
can do better; there are LOX/kerosene engines with vacuum Isps over 350 s.
That would drop the mass ratio to 2.6.