Space shuttle for space tourism and first stage of a TSTO.

On Jan 19, 5:56*pm, Robert Clark wrote:
...This page gives the specifications of the Ares I:

Space Launch Report - Ares I.http://www.spacelaunchreport.com/ares1.html

*The gross weight including payload is given as 912,660 kg and the
gross weight of the first stage as 732,550 kg. So the gross weight of
the Ares I second stage plus payload is 180,110 kg.
*Then the gross weight for the 55,442 kg dry weight of the
reconfigured shuttle, plus 300,000 kg propellant load, plus 180,110 kg
second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3 NK-33
engines I was suggesting to use only put out a total of 1,018,518 lbs.
of thrust at sea level. For this purpose you would need a fourth
NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg,
1,180,903 lbs., and the sea level thrust of the 4 engines is 1,358,024
lbs.
*Using the average Isp of the NK-33 as the midpoint of the sea level
and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln
(536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent
delta-V, speed + altitude, provided by the Ares I first stage. The
achieved delta-V is actually higher than this since the rocket spends
most of the time at high altitude, where the Isp is closer to the
vacuum value.
Note that if you want to increase the delta-V, the space occupied by
the crew compartment is now empty. This gives an additional 74 cubic
meters that could be used for propellant, which amounts to 74,000 kg
additional lox/kerosene propellant that could be carried.
Then we could still use the planned upper stage of the Ares I while
having a significantly lower development cost and per launch cost of
the now reusable first stage.

If we only instead wanted suborbital tourism for the vehicle then you
would require much less fuel load. Having engine-out capability is a
necessary requirement for manned flights. According to this page, the
shuttle has a max emergency landing weight of 240,000 lbs, 109,000 kg:

Space Transportatin System.
http://science.ksc.nasa.gov/shuttle/...l#launch_sites

In the calculation in the above post, after removing several
subsystems that wouldn't be needed for an unmanned first stage booster
I got a 55,442 kg dry weight using three NK-33 engines. However, for
suborbital tourism we need the crew seats and environmental systems so
I'll add back the 3,250 kg for these to get a dry weight of 58,692 kg.
Now assume we max our fuel load for the suborbital tourism use at
50,308 kg, so our max takeoff weight is 109,000 kg, the max allowed
for the shuttle for landing under abort modes. Then our delta-V
assuming a 315 s average Isp of the NK-33's would be:

315*9.8*ln(109,000/58,692) = 1,911 m/s. This is well above the total
equivalent delta-V, speed + altitude, required for reaching the 100 km
altitude for space tourism. And the achieved delta-V would actually be
higher than this because the actual average Isp is closer to the
vacuum value than to the midpoint value.
Note that in this configuration we even have 2 engine-out capability
since the thrust put out by the NK-33's at sea level is over 300,000
lbs. And since our max weight is at the max allowed for landing the
vehicle could even glide to a landing with a full fuel load if all
three engines failed as long as we did reach high enough initial
velocity for aerodynamic lift to operate. (The orbiter has a
respectable lift/drag ratio of 4.5 at subsonic speeds.)
Keep in mind that not even jet airliners can land safely during an
aborted takeoff if they have all engines out unless they reach
sufficient altitude and velocity for lift to operate.
This is for the case of just carrying 6 passengers in the crew
compartment. The case for carrying a full passenger cabin in the
payload bay and fuel only in the wings is a much more complicated
analysis and probably not susceptible to an elementary analysis.


Bob Clark

Robert Clark wrote:
On Jan 19, 5:56 pm, Robert Clark wrote:
...This page gives the specifications of the Ares I:

Space Launch Report - Ares
I.http://www.spacelaunchreport.com/ares1.html

The gross weight including payload is given as 912,660 kg and the
gross weight of the first stage as 732,550 kg. So the gross weight of
the Ares I second stage plus payload is 180,110 kg.
Then the gross weight for the 55,442 kg dry weight of the
reconfigured shuttle, plus 300,000 kg propellant load, plus 180,110
kg second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3
NK-33 engines I was suggesting to use only put out a total of
1,018,518 lbs. of thrust at sea level. For this purpose you would
need a fourth
NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg,
1,180,903 lbs., and the sea level thrust of the 4 engines is
1,358,024 lbs.
Using the average Isp of the NK-33 as the midpoint of the sea level
and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln
(536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent
delta-V, speed + altitude, provided by the Ares I first stage. The
achieved delta-V is actually higher than this since the rocket spends
most of the time at high altitude, where the Isp is closer to the
vacuum value.
Note that if you want to increase the delta-V, the space occupied by
the crew compartment is now empty. This gives an additional 74 cubic
meters that could be used for propellant, which amounts to 74,000 kg
additional lox/kerosene propellant that could be carried.
Then we could still use the planned upper stage of the Ares I while
having a significantly lower development cost and per launch cost of
the now reusable first stage.

If we only instead wanted suborbital tourism for the vehicle then you
would require much less fuel load. Having engine-out capability is a
necessary requirement for manned flights. According to this page, the
shuttle has a max emergency landing weight of 240,000 lbs, 109,000 kg:

Space Transportatin System.
http://science.ksc.nasa.gov/shuttle/...l#launch_sites

In the calculation in the above post, after removing several
subsystems that wouldn't be needed for an unmanned first stage booster
I got a 55,442 kg dry weight using three NK-33 engines. However, for
suborbital tourism we need the crew seats and environmental systems so
I'll add back the 3,250 kg for these to get a dry weight of 58,692 kg.
Now assume we max our fuel load for the suborbital tourism use at
50,308 kg, so our max takeoff weight is 109,000 kg, the max allowed
for the shuttle for landing under abort modes. Then our delta-V
assuming a 315 s average Isp of the NK-33's would be:

315*9.8*ln(109,000/58,692) = 1,911 m/s. This is well above the total
equivalent delta-V, speed + altitude, required for reaching the 100 km
altitude for space tourism. And the achieved delta-V would actually be
higher than this because the actual average Isp is closer to the
vacuum value than to the midpoint value.
Note that in this configuration we even have 2 engine-out capability
since the thrust put out by the NK-33's at sea level is over 300,000
lbs. And since our max weight is at the max allowed for landing the
vehicle could even glide to a landing with a full fuel load if all
three engines failed as long as we did reach high enough initial
velocity for aerodynamic lift to operate. (The orbiter has a
respectable lift/drag ratio of 4.5 at subsonic speeds.)
Keep in mind that not even jet airliners can land safely during an
aborted takeoff if they have all engines out unless they reach
sufficient altitude and velocity for lift to operate.

If "lift" was not "operating" then an airliner would not have reached _any_
altitude.

cretin.

plonk

On Jan 26, 9:33*am, Robert Clark wrote:
On Jan 19, 5:56*pm, Robert Clark wrote:

...This page gives the specifications of the Ares I:

SpaceLaunch Report - Ares I.http://www.spacelaunchreport.com/ares1.html

*The gross weight including payload is given as 912,660 kg and the
gross weight of the first stage as 732,550 kg. So the gross weight of
the Ares I second stage plus payload is 180,110 kg.
*Then the gross weight for the 55,442 kg dry weight of the
reconfiguredshuttle, plus 300,000 kg propellant load, plus 180,110 kg
second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3 NK-33
engines I was suggesting to use only put out a total of 1,018,518 lbs.
of thrust at sea level. For this purpose you would need a fourth
NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg,
1,180,903 lbs., and the sea level thrust of the 4 engines is 1,358,024
lbs.
*Using the average Isp of the NK-33 as the midpoint of the sea level
and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln
(536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent
delta-V, speed + altitude, provided by the Ares I first stage. The
achieved delta-V is actually higher than this since the rocket spends
most of the time at high altitude, where the Isp is closer to the
vacuum value.
Note that if you want to increase the delta-V, the space occupied by
the crew compartment is now empty. This gives an additional 74 cubic
meters that could be used for propellant, which amounts to 74,000 kg
additional lox/kerosene propellant that could be carried.
Then we could still use the planned upper stage of the Ares I while
having a significantly lower development cost and per launch cost of
the now reusable first stage.


In addition to the Air Force wanting to build reusable first stages
with expendable upper stages to cut launch costs, the Europeans are
also planning such boosters, though sled launched:

Hopper (spacecraft).
http://en.wikipedia.org/wiki/Hopper_%28spacecraft%29

Europe's space shuttle passes early test.
* 16:05 10 May 2004 by Maggie McKee
"Europe took a step towards creating an unmanned space shuttle on
Saturday when a prototype landed autonomously after a test flight in
Sweden.
"The shuttle prototype, called Phoenix, is one of several proposals
for a European reusable launch vehicle (RLV) planned to cheaply ferry
satellites into orbit by 2015."
http://www.newscientist.com/article/dn4975

"An unmanned scale model prototype of the planned European shuttle is
pictured during its first free test flight at the North European
Aerospace Test Range in Kiruna, Sweden 1,230 kilometers (770 miles)
north of Stockholm, Saturday, May 8, 2004. (AP Photo/Peter Degerfeldt,
Blue Sky AB, DLR/EADS)."
http://i.space.com/images/h_phoenix_flight_02.jpg

Reusable space transport systems to reduce cost.
Bremen/Le Bourget, 01 June 2001
"In its HOPPER concept, Astrium has gone for an autonomous transport
system that is noted for its high degree of reusability and
comparatively low mission costs. Although the unmanned HOPPER is very
similar in appearance to the US Space Shuttle there are, however, some
substantial differences: The system will be launched horizontally on a
skid sled running on a four kilometre long track. The vehicle itself
is more compact than the Space Shuttle. The re-entry angle will be
optimised so as to keep the frictional heat at the outer skin
substantially lower than that of the Space Shuttle. This will allow
the sensitive and expensive thermal protection shield to be replaced
by an affordable, low-maintenance heat protection system. Due to these
features, Hopper is intended to transport payloads to orbit at 75%
lower cost than conventional transport systems."
http://www.eads.net/1024/fr/pressdb/...rium_reus.html

This article said the "Hopper" is more "compact" than the shuttle
orbiter, but actually it's about 13 meters longer with a 3 meter wider
wing span. This will allow it to have internal propellant tanks and
also an internal payload bay:

http://www.hobbyspace.com/AAdmin/Ima...RLVConf_md.jpg

The Europeans expect to use upgraded versions of Ariane's hydrogen-
fueled Vulcain engine. The choice of the hydrogen-fueled Vulcain,
rather then a kerosene engine typically used for first stages, perhaps
stems from the desire to have it be sled launched, requiring a lower
GLOW, and the desire to use the ESA's most advanced engine.
Rather than expending such large amounts of money building their own
Hopper vehicle, it might be cheaper as a technology demonstrator for
ESA to adapt the shuttle/Buran for the purpose. NASA seems to want to
give the SSME engines away for free, so the ESA could even get the
needed engines.
The shuttle/Buran is smaller than the proposed Hopper. To get similar
payload to orbit on an expendable upper stage, about 7,500 kg, they
could have it be vertically instead of sled launched to allow the
upper stage to be attached on the vehicle bottom. This would then
allow the full fuselage to be used for propellant.
However, as I have argued by using kerosene instead as the propellant
for this first stage you could have the expendable upper stage be as
heavy as that of the Ares I second stage at 180,000 kg, and with a
much larger payload to orbit than 7,500 kg, at about the size of the
Orion capsule, 20,000 kg to 25,000 kg.


Bob Clark

The real problem at the moment though, is not a shortage of workable ideas,
its a shortage of money to build them really.

Brian

--
Brian Gaff -
Note:- In order to reduce spam, any email without 'Brian Gaff'
in the display name may be lost.
Blind user, so no pictures please!
"Robert Clark" wrote in message
...
On Jan 26, 9:33 am, Robert Clark wrote:
On Jan 19, 5:56 pm, Robert Clark wrote:

...This page gives the specifications of the Ares I:

SpaceLaunch Report - Ares I.http://www.spacelaunchreport.com/ares1.html

The gross weight including payload is given as 912,660 kg and the
gross weight of the first stage as 732,550 kg. So the gross weight of
the Ares I second stage plus payload is 180,110 kg.
Then the gross weight for the 55,442 kg dry weight of the
reconfiguredshuttle, plus 300,000 kg propellant load, plus 180,110 kg
second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3 NK-33
engines I was suggesting to use only put out a total of 1,018,518 lbs.
of thrust at sea level. For this purpose you would need a fourth
NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg,
1,180,903 lbs., and the sea level thrust of the 4 engines is 1,358,024
lbs.
Using the average Isp of the NK-33 as the midpoint of the sea level
and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln
(536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent
delta-V, speed + altitude, provided by the Ares I first stage. The
achieved delta-V is actually higher than this since the rocket spends
most of the time at high altitude, where the Isp is closer to the
vacuum value.
Note that if you want to increase the delta-V, the space occupied by
the crew compartment is now empty. This gives an additional 74 cubic
meters that could be used for propellant, which amounts to 74,000 kg
additional lox/kerosene propellant that could be carried.
Then we could still use the planned upper stage of the Ares I while
having a significantly lower development cost and per launch cost of
the now reusable first stage.


In addition to the Air Force wanting to build reusable first stages
with expendable upper stages to cut launch costs, the Europeans are
also planning such boosters, though sled launched:

Hopper (spacecraft).
http://en.wikipedia.org/wiki/Hopper_%28spacecraft%29

Europe's space shuttle passes early test.
* 16:05 10 May 2004 by Maggie McKee
"Europe took a step towards creating an unmanned space shuttle on
Saturday when a prototype landed autonomously after a test flight in
Sweden.
"The shuttle prototype, called Phoenix, is one of several proposals
for a European reusable launch vehicle (RLV) planned to cheaply ferry
satellites into orbit by 2015."
http://www.newscientist.com/article/dn4975

"An unmanned scale model prototype of the planned European shuttle is
pictured during its first free test flight at the North European
Aerospace Test Range in Kiruna, Sweden 1,230 kilometers (770 miles)
north of Stockholm, Saturday, May 8, 2004. (AP Photo/Peter Degerfeldt,
Blue Sky AB, DLR/EADS)."
http://i.space.com/images/h_phoenix_flight_02.jpg

Reusable space transport systems to reduce cost.
Bremen/Le Bourget, 01 June 2001
"In its HOPPER concept, Astrium has gone for an autonomous transport
system that is noted for its high degree of reusability and
comparatively low mission costs. Although the unmanned HOPPER is very
similar in appearance to the US Space Shuttle there are, however, some
substantial differences: The system will be launched horizontally on a
skid sled running on a four kilometre long track. The vehicle itself
is more compact than the Space Shuttle. The re-entry angle will be
optimised so as to keep the frictional heat at the outer skin
substantially lower than that of the Space Shuttle. This will allow
the sensitive and expensive thermal protection shield to be replaced
by an affordable, low-maintenance heat protection system. Due to these
features, Hopper is intended to transport payloads to orbit at 75%
lower cost than conventional transport systems."
http://www.eads.net/1024/fr/pressdb/...rium_reus.html

This article said the "Hopper" is more "compact" than the shuttle
orbiter, but actually it's about 13 meters longer with a 3 meter wider
wing span. This will allow it to have internal propellant tanks and
also an internal payload bay:

http://www.hobbyspace.com/AAdmin/Ima...RLVConf_md.jpg

The Europeans expect to use upgraded versions of Ariane's hydrogen-
fueled Vulcain engine. The choice of the hydrogen-fueled Vulcain,
rather then a kerosene engine typically used for first stages, perhaps
stems from the desire to have it be sled launched, requiring a lower
GLOW, and the desire to use the ESA's most advanced engine.
Rather than expending such large amounts of money building their own
Hopper vehicle, it might be cheaper as a technology demonstrator for
ESA to adapt the shuttle/Buran for the purpose. NASA seems to want to
give the SSME engines away for free, so the ESA could even get the
needed engines.
The shuttle/Buran is smaller than the proposed Hopper. To get similar
payload to orbit on an expendable upper stage, about 7,500 kg, they
could have it be vertically instead of sled launched to allow the
upper stage to be attached on the vehicle bottom. This would then
allow the full fuselage to be used for propellant.
However, as I have argued by using kerosene instead as the propellant
for this first stage you could have the expendable upper stage be as
heavy as that of the Ares I second stage at 180,000 kg, and with a
much larger payload to orbit than 7,500 kg, at about the size of the
Orion capsule, 20,000 kg to 25,000 kg.


Bob Clark