SpaceX Dragon spacecraft for low cost trips to the Moon.

On Oct 6, 5:03*am, Robert Clark wrote:
...
* An even lower cost possibility for the capsule and lander might be
one proposed by the University of Maryland aerospace engineering
department:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration
Vehicle.http://www.nianet.org/rascal/forum20..._umd_paper.pdf

*As with the Orion CEV, this Phoenix spacecraft was intended to be
used in conjunction with a separate lander for lunar missions.
However, by using it both for the trip from LEO and as the lander you
get great savings in cost.
On page 3 of the report is given a breakdown of the weights of the
various subsystems. By removing the propulsion system as I suggested
for the Dragon for this purpose, the mass with crew would be about
half that of the Dragon, at about 2,000 kg.
Then assuming again 10 to 1 mass ratios for two Centaur style stages
for propulsion, we would need about half the propellant load as for
the Dragon, about 20,000 kg, which could be lofted by a single launch
of the current largest launchers.
Then the cost of lofting this propellant load to LEO would be about
$100 million. And if a new heavy lift launcher could get a $2,400 per
kg launch price, it would only be in the range of $50 million.
This would increase even further the market for such low cost lunar
missions.


Especially innovative about this design is the "parashield" thermal
protection. Not only is this lightweight but another advantage is that
it has a higher protective area so that you can use a larger volume
cylindrical structure rather than the usual conical structure for the
capsule. From the report "Phoenix: A Low-Cost Commercial Approach to
the Crew Exploration":

"Figure 5.9-1: Phoenix ParaShield in stowed and deployed
configurations."
http://oi51.tinypic.com/14e9vd4.jpg

Also, with a 20,000 kg mass for the propulsion system and 2,000 kg
mass for the capsule. The total 22,000 kg mass might be launchable by
a single large class launcher, in the $100-$140 million cost range.
The crew would be launched on a separate high reliability man-rated
launcher such as the Soyuz to link up with the vehicle in orbit.

Bob Clark

On Oct 16, 4:51*pm, Robert Clark wrote::
...
* An even lower cost possibility for the capsule and lander might be
one proposed by the University of Maryland aerospace engineering
department:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration
Vehicle.http://www.nianet.org/rascal/forum20..._umd_paper.pdf

*As with the Orion CEV, this Phoenix spacecraft was intended to be
used in conjunction with a separate lander for lunar missions.
However, by using it both for the trip from LEO and as the lander you
get great savings in cost.
On page 3 of the report is given a breakdown of the weights of the
various subsystems. By removing the propulsion system as I suggested
for the Dragon for this purpose, the mass with crew would be about
half that of the Dragon, at about 2,000 kg.
Then assuming again 10 to 1 mass ratios for two Centaur style stages
for propulsion, we would need about half the propellant load as for
the Dragon, about 20,000 kg, which could be lofted by a single launch
of the current largest launchers.
Then the cost of lofting this propellant load to LEO would be about
$100 million. And if a new heavy lift launcher could get a $2,400 per
kg launch price, it would only be in the range of $50 million.
This would increase even further the market for such low cost lunar
missions.

*Especially innovative about this design is the "parashield" thermal
protection. Not only is this lightweight but another advantage is that
it has a higher protective area so that you can use a larger volume
cylindrical structure rather than the usual conical structure for the
capsule. From the report "Phoenix: A Low-Cost Commercial Approach to
the Crew Exploration":

"Figure 5.9-1: Phoenix ParaShield in stowed and deployed
configurations."http://oi51.tinypic.com/14e9vd4.jpg

*Also, with a 20,000 kg mass for the propulsion system and 2,000 kg
mass for the capsule. The total 22,000 kg mass might be launchable by
a single large class launcher, in the $100-$140 million cost range.
The crew would be launched on a separate high reliability man-rated
launcher such as the Soyuz to link up with the vehicle in orbit.



The price for these commercial lunar flights could be cut
dramatically if instead of hauling the fuel from the Earth, it could
be obtained from the Moon. This would require automated systems to
produce propellant from the materials on the Moon.
Then as a precursor to show this is feasible it would be necessary to
do a smaller unmanned lunar lander mission that demonstrates ISRU
propellant production. We will want to do a reusable, round trip
mission to also show the feasibility of the manned missions. However,
as a low cost first step we'll only do an expendable one-way lander
that drops off an electrolysis station to produce hydrogen/oxygen from
the water found by LCROSS to be near surface in the polar regions.
To keep costs low we'll use the Russian Dnepr rocket:

Dnepr specifications.
http://www.spaceandtech.com/spacedat...pr_specs.shtml

According to this page, the price is $10-$13 million for up to 4,500
kg to LEO. So we'll need to keep the total mass for the lander and the
propulsion system under 4,500 kg.
One possibility for the propulsion might be the solid motor "Star"
series, but multiply staged. Find the specifications for the Star 48
version he

Star 48 - Specifications.
http://www.spaceandtech.com/spacedat...48_specs.shtml

They have a good mass ratio at around 18 or 19 to 1. And a moderate
Isp, from 286 s to 292 s. However, it should be noted that the low dry
mass indicated, which results in the high mass ratio, is coming from
the fact this is only considering the nozzle and casing. Reaction
control thrusters and the avionics assemblies are not included in this
dry mass.
A more accurate accounting for the dry mass for this upper stage might
be he

PAM-S.
"Solid propellant rocket stage. Loaded/empty mass 2,182/220 kg. Thrust
66.60 kN. Vacuum specific impulse 288 seconds.
Cost $ : 4.060 million."
http://www.astronautix.com/stages/pams.htm

Note this page, with the higher dry mass, indicates this upper stage
with the Star 48 engine does also use reaction control thrusters. The
extra mass was about 100 kg added onto the 111 kg Star 48 bare mass.
I'll reserve 100 kg for the RCS and avionics within the mass of the
payload, and use the bare masses for the Star engines in the delta-V
calculations. The final, smallest stage will have slightly more
powerful RCS than needed and for the lower stages I'll rely on spin-
stabilization and the upper stage RCS for stability while the lower
stage motors are firing.
Let's calculate how much payload we could deliver to the Moon's
surface. This page gives the delta-V requirements in the Earth-Moon
system:

Delta-v budget.
Earth–Moon space.
http://en.wikipedia.org/wiki/Delta-v...0.93Moon_space

To get to the lunar surface from LEO would require a delta-V of 5.93
km/s. The stages used will be the Star 48B:

STAR 48B - Short Nozzle PAM STS.
"Effective Isp (vacuum): 286.0 sec
Motor Loaded Mass: 4705.4 lb, 2134.3 kg
Motor Burnout Mass: 245.4 lb, 111.3 kg"
http://www.spaceandtech.com/spacedat...48_specs.shtml ,

the Star 37FM:

STAR 37FM.
"Effective Isp (vacuum): 289.8 sec
Motor Loaded Mass: 2530.8 lb, 1148.0 kg
Motor Burnout Mass: 162.5 lb, 73.7 kg"
http://www.spaceandtech.com/spacedat...37_specs.shtml ,

and the Star 30:

Star 30.
"Gross mass: 492 kg (1,084 lb).
Unfuelled mass: 28 kg (61 lb).
Diameter: 0.76 m (2.50 ft).
Specific impulse: 293 s."
http://www.astronautix.com/engines/star30.htm

Estimate the payload to the Moon as 400 kg. The delta-V needed for
Trans Lunar Injection will be in the range of 3.05 to 3.25 km/s:

Trans Lunar Injection.
History.
http://en.wikipedia.org/wiki/Trans_L...ection#History

The delta-V you could get from the Star 48 first stage would be:
286*9.8ln((2134.3+1148+492+400)/(111.3+1148+492+400)) = 1,857 m/s.
The delta-V you get from the Star 37FM second stage will be:
289.9*9.8ln((1,148+492+400)/(73.7+492+400)) = 2,125 m/s. The two lower
stages give you a total of 3,982 m/s, sufficient for TLI.
You need now 5,930 - 3,982 = 1,948 m/s additional delta-V to complete
the landing. The delta-V you get from the Star 30 will be:
293*9.8ln((492+400)/(28+400)) = 2,109 m/s, sufficient for the landing.
The total gross mass of the 3 stages plus payload will be
2,134.3+1,148+492+400 = 4,174.3 kg, within the lift capacity of the
Dnepr 1. The cost of the Dnepr 1 might be $13 million. The costs of
the upper stages? The Astronautix page on the PAM-S powered by the
Star 48 motor gives the price as $4.06 million. The Star 37 is smaller
by half, and the Star 30 is smaller by an additional factor of one-
half. Then we might estimate their prices as $2 million and $1 million
respectively, for a total cost of these upper stages of $7 million.
Then the total launch cost might be $20 million.
We would have to add onto that the cost of the avionics and the cost
of the lander.


Bob Clark

On Sun, 14 Nov 2010 04:39:49 -0800 (PST), Robert Clark wrote:

On Oct 16, 4:51*pm, Robert Clark wrote::
...
* An even lower cost possibility for the capsule and lander might be
one proposed by the University of Maryland aerospace engineering
department:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration
Vehicle.http://www.nianet.org/rascal/forum20..._umd_paper.pdf

*As with the Orion CEV, this Phoenix spacecraft was intended to be
used in conjunction with a separate lander for lunar missions.
However, by using it both for the trip from LEO and as the lander you
get great savings in cost.
On page 3 of the report is given a breakdown of the weights of the
various subsystems. By removing the propulsion system as I suggested
for the Dragon for this purpose, the mass with crew would be about
half that of the Dragon, at about 2,000 kg.
Then assuming again 10 to 1 mass ratios for two Centaur style stages
for propulsion, we would need about half the propellant load as for
the Dragon, about 20,000 kg, which could be lofted by a single launch
of the current largest launchers.
Then the cost of lofting this propellant load to LEO would be about
$100 million. And if a new heavy lift launcher could get a $2,400 per
kg launch price, it would only be in the range of $50 million.
This would increase even further the market for such low cost lunar
missions.

*Especially innovative about this design is the "parashield" thermal
protection. Not only is this lightweight but another advantage is that
it has a higher protective area so that you can use a larger volume
cylindrical structure rather than the usual conical structure for the
capsule. From the report "Phoenix: A Low-Cost Commercial Approach to
the Crew Exploration":

"Figure 5.9-1: Phoenix ParaShield in stowed and deployed
configurations."http://oi51.tinypic.com/14e9vd4.jpg

*Also, with a 20,000 kg mass for the propulsion system and 2,000 kg
mass for the capsule. The total 22,000 kg mass might be launchable by
a single large class launcher, in the $100-$140 million cost range.
The crew would be launched on a separate high reliability man-rated
launcher such as the Soyuz to link up with the vehicle in orbit.



The price for these commercial lunar flights could be cut
dramatically if instead of hauling the fuel from the Earth, it could
be obtained from the Moon. This would require automated systems to
produce propellant from the materials on the Moon.
Then as a precursor to show this is feasible it would be necessary to
do a smaller unmanned lunar lander mission that demonstrates ISRU
propellant production. We will want to do a reusable, round trip
mission to also show the feasibility of the manned missions. However,
as a low cost first step we'll only do an expendable one-way lander
that drops off an electrolysis station to produce hydrogen/oxygen from
the water found by LCROSS to be near surface in the polar regions.
To keep costs low we'll use the Russian Dnepr rocket:

Dnepr specifications.
http://www.spaceandtech.com/spacedat...pr_specs.shtml

According to this page, the price is $10-$13 million for up to 4,500
kg to LEO. So we'll need to keep the total mass for the lander and the
propulsion system under 4,500 kg.
One possibility for the propulsion might be the solid motor "Star"
series, but multiply staged. Find the specifications for the Star 48
version he

Star 48 - Specifications.
http://www.spaceandtech.com/spacedat...48_specs.shtml

They have a good mass ratio at around 18 or 19 to 1. And a moderate
Isp, from 286 s to 292 s. However, it should be noted that the low dry
mass indicated, which results in the high mass ratio, is coming from
the fact this is only considering the nozzle and casing. Reaction
control thrusters and the avionics assemblies are not included in this
dry mass.
A more accurate accounting for the dry mass for this upper stage might
be he

PAM-S.
"Solid propellant rocket stage. Loaded/empty mass 2,182/220 kg. Thrust
66.60 kN. Vacuum specific impulse 288 seconds.
Cost $ : 4.060 million."
http://www.astronautix.com/stages/pams.htm

Note this page, with the higher dry mass, indicates this upper stage
with the Star 48 engine does also use reaction control thrusters. The
extra mass was about 100 kg added onto the 111 kg Star 48 bare mass.
I'll reserve 100 kg for the RCS and avionics within the mass of the
payload, and use the bare masses for the Star engines in the delta-V
calculations. The final, smallest stage will have slightly more
powerful RCS than needed and for the lower stages I'll rely on spin-
stabilization and the upper stage RCS for stability while the lower
stage motors are firing.
Let's calculate how much payload we could deliver to the Moon's
surface. This page gives the delta-V requirements in the Earth-Moon
system:

Delta-v budget.
Earth–Moon space.
http://en.wikipedia.org/wiki/Delta-v...0.93Moon_space

To get to the lunar surface from LEO would require a delta-V of 5.93
km/s. The stages used will be the Star 48B:

STAR 48B - Short Nozzle PAM STS.
"Effective Isp (vacuum): 286.0 sec
Motor Loaded Mass: 4705.4 lb, 2134.3 kg
Motor Burnout Mass: 245.4 lb, 111.3 kg"
http://www.spaceandtech.com/spacedat...48_specs.shtml ,

the Star 37FM:

STAR 37FM.
"Effective Isp (vacuum): 289.8 sec
Motor Loaded Mass: 2530.8 lb, 1148.0 kg
Motor Burnout Mass: 162.5 lb, 73.7 kg"
http://www.spaceandtech.com/spacedat...37_specs.shtml ,

and the Star 30:

Star 30.
"Gross mass: 492 kg (1,084 lb).
Unfuelled mass: 28 kg (61 lb).
Diameter: 0.76 m (2.50 ft).
Specific impulse: 293 s."
http://www.astronautix.com/engines/star30.htm

Estimate the payload to the Moon as 400 kg. The delta-V needed for
Trans Lunar Injection will be in the range of 3.05 to 3.25 km/s:

Trans Lunar Injection.
History.
http://en.wikipedia.org/wiki/Trans_L...ection#History

The delta-V you could get from the Star 48 first stage would be:
286*9.8ln((2134.3+1148+492+400)/(111.3+1148+492+400)) = 1,857 m/s.
The delta-V you get from the Star 37FM second stage will be:
289.9*9.8ln((1,148+492+400)/(73.7+492+400)) = 2,125 m/s. The two lower
stages give you a total of 3,982 m/s, sufficient for TLI.
You need now 5,930 - 3,982 = 1,948 m/s additional delta-V to complete
the landing. The delta-V you get from the Star 30 will be:
293*9.8ln((492+400)/(28+400)) = 2,109 m/s, sufficient for the landing.
The total gross mass of the 3 stages plus payload will be
2,134.3+1,148+492+400 = 4,174.3 kg, within the lift capacity of the
Dnepr 1. The cost of the Dnepr 1 might be $13 million. The costs of
the upper stages? The Astronautix page on the PAM-S powered by the
Star 48 motor gives the price as $4.06 million. The Star 37 is smaller
by half, and the Star 30 is smaller by an additional factor of one-
half. Then we might estimate their prices as $2 million and $1 million
respectively, for a total cost of these upper stages of $7 million.
Then the total launch cost might be $20 million.
We would have to add onto that the cost of the avionics and the cost
of the lander.


Bob Clark

Other mathematical exercises to consider:
1. How much metal would it take to build ships for the exodus that is
the stuff of science fiction?
2. What percentage of the GPP (gross planetary product) can be devoted
to building interstellar craft?
3. How much fuel can we devote to shipping people off-world?
4. At the current rate of population growth, how long does it take for
the earth's population to double?
5. How many ships would it take to transport 7 billion people to other
planets?
6. To how many destinations would they be sent to avoid overpopulating
those planets in a short time?
7. How many ships would we need to launch each day to keep this
planet's population constant at its current level?
8. How many people would it take to "overpopulate" a newly settled
planet starting from zero infrastructure?
9. Given the opportunity to stay on Earth or to make a hazardous,
one-way, years-long journey to a world where there is no
infrastructure, how many would sign up to leave?
10. If only the cream of the crop is shipped off-world, how long can
the exodus continue before the population left behind collapses and no
more ships are sent?

Not a mathematical calculation, but a question about the societies
portrayed in science fiction: If only the "cream crop" is sent, who
will clean toilets on the settled worlds? Did you ever see anyone
cleaning toilets on the Starship Enterprise?

Give a listen to "The Intergalactic Laxative" by Donovan:
http://www.youtube.com/watch?v=ZCpnwJQhoYY

On Nov 14, 7:39*am, Robert Clark wrote:
...
One possibility for the propulsion might be the solid motor "Star"
series, but multiply staged. Find the specifications for the Star 48
version he

Star 48 - Specifications.http://www.spaceandtech.com/spacedat...48_specs.shtml

They have a good mass ratio at around 18 or 19 to 1. And a moderate
Isp, from 286 s to 292 s. However, it should be noted that the low dry
mass indicated, which results in the high mass ratio, is coming from
the fact this is only considering the nozzle and casing. Reaction
control thrusters and the avionics assemblies are not included in this
dry mass.
A more accurate accounting for the dry mass for this upper stage might
be he

PAM-S.
"Solid propellant rocket stage. Loaded/empty mass 2,182/220 kg. Thrust
66.60 kN. Vacuum specific impulse 288 seconds.
Cost $ : 4.060 million."http://www.astronautix.com/stages/pams.htm

Note this page, with the higher dry mass, indicates this upper stage
with the Star 48 engine does also use reaction control thrusters. The
extra mass was about 100 kg added onto the 111 kg Star 48 bare mass.
I'll reserve 100 kg for the RCS and avionics within the mass of the
payload, and use the bare masses for the Star engines in the delta-V
calculations. The final, smallest stage will have slightly more
powerful RCS than needed and for the lower stages I'll rely on spin-
stabilization and the upper stage RCS for stability while the lower
stage motors are firing.
Let's calculate how much payload we could deliver to the Moon's
surface. This page gives the delta-V requirements in the Earth-Moon
system:

Delta-v budget.
Earth–Moon space.http://en.wikipedia.org/wiki/Delta-v...0.93Moon_space

*To get to the lunar surface from LEO would require a delta-V of 5.93
km/s. The stages used will be the Star 48B:

STAR 48B - Short Nozzle PAM STS.
"Effective Isp (vacuum): *286.0 sec
Motor Loaded Mass: *4705.4 lb, 2134.3 kg
Motor Burnout Mass: *245.4 lb, 111.3 kg"http://www.spaceandtech.com/spacedata/motors/star48_specs.shtml,

the Star 37FM:

STAR 37FM.
"Effective Isp (vacuum): *289.8 sec
Motor Loaded Mass: *2530.8 lb, 1148.0 kg
Motor Burnout Mass: *162.5 lb, 73.7 kg"http://www.spaceandtech.com/spacedata/motors/star37_specs.shtml,

and the Star 30:

Star 30.
"Gross mass: 492 kg (1,084 lb).
Unfuelled mass: 28 kg (61 lb).
Diameter: 0.76 m (2.50 ft).
Specific impulse: 293 s."http://www.astronautix.com/engines/star30.htm

*Estimate the payload to the Moon as 400 kg. The delta-V needed for
Trans Lunar Injection will be in the range of 3.05 to 3.25 km/s:

Trans Lunar Injection.
History.http://en.wikipedia.org/wiki/Trans_L...ection#History

*The delta-V you could get from the Star 48 first stage would be:
286*9.8ln((2134.3+1148+492+400)/(111.3+1148+492+400)) = 1,857 m/s.
The delta-V you get from the Star 37FM second stage will be:
289.9*9.8ln((1,148+492+400)/(73.7+492+400)) = 2,125 m/s. The two lower
stages give you a total of 3,982 m/s, sufficient for TLI.
You need now 5,930 - 3,982 = 1,948 m/s additional delta-V to complete
the landing. The delta-V you get from the Star 30 will be:
293*9.8ln((492+400)/(28+400)) = 2,109 m/s, sufficient for the landing.
The total gross mass of the 3 stages plus payload will be
2,134.3+1,148+492+400 = 4,174.3 kg, within the lift capacity of the
Dnepr 1. The cost of the Dnepr 1 might be $13 million. The costs of
the upper stages? The Astronautix page on the PAM-S powered by the
Star 48 motor gives the price as $4.06 million. The Star 37 is smaller
by half, and the Star 30 is smaller by an additional factor of one-
half. Then we might estimate their prices as $2 million and $1 million
respectively, for a total cost of these upper stages of $7 million.
Then the total launch cost might be $20 million.
We would have to add onto that the cost of the avionics and the cost
of the lander.


As a point of comparison the Dnepr has been studied to be used to
launch a 500 kg payload to GEO by using two Star solid motors and
lunar gravity assist:

Dnepr (R-36M2).
"The Dnepr launch vehicle does not have the capability to deploy
payloads directiy into GTO. However, Kosmotras has studied a technique
to deliver small spacecraft to GEO using the gravity of the Moon to
provide the plane change and perigee raising. In this scenario, the
spacecraft is attached to Star 48A and Star 27 solid motors, supplied
separately by ATK Thiokol. The Star48A would send the spacecraft to
the Moon, where a gravity slingshot maneuver would lower the transfer
orbit inclination from 50.5 deg to 0 deg, and raise the orbit perigee
to geostationary altitude. When the spacecraft reaches perigee of the
new transfer orbit, the Star 27 motor would fire to circularize the
orbit at GEO. Using this method, a 500 kg (1100 lbm) spacecraft could
be delivered to GEO."
http://www.b14643.de/Spacerockets_1/...tion/Frame.htm


Bob Clark

On Nov 14, 7:39*am, Robert Clark wrote:
...
*To get to the lunar surface from LEO would require a delta-V of 5.93
km/s. The stages used will be the Star 48B:

STAR 48B - Short Nozzle PAM STS.
"Effective Isp (vacuum): *286.0 sec
Motor Loaded Mass: *4705.4 lb, 2134.3 kg
Motor Burnout Mass: *245.4 lb, 111.3 kg"http://www.spaceandtech.com/spacedata/motors/star48_specs.shtml,

the Star 37FM:

STAR 37FM.
"Effective Isp (vacuum): *289.8 sec
Motor Loaded Mass: *2530.8 lb, 1148.0 kg
Motor Burnout Mass: *162.5 lb, 73.7 kg"http://www.spaceandtech.com/spacedata/motors/star37_specs.shtml,

and the Star 30:

Star 30.
"Gross mass: 492 kg (1,084 lb).
Unfuelled mass: 28 kg (61 lb).
Diameter: 0.76 m (2.50 ft).
Specific impulse: 293 s."http://www.astronautix.com/engines/star30.htm

*Estimate the payload to the Moon as 400 kg. The delta-V needed for
Trans Lunar Injection will be in the range of 3.05 to 3.25 km/s:

Trans Lunar Injection.
History.http://en.wikipedia.org/wiki/Trans_L...ection#History

*The delta-V you could get from the Star 48 first stage would be:
286*9.8ln((2134.3+1148+492+400)/(111.3+1148+492+400)) = 1,857 m/s.
The delta-V you get from the Star 37FM second stage will be:
289.9*9.8ln((1,148+492+400)/(73.7+492+400)) = 2,125 m/s. The two lower
stages give you a total of 3,982 m/s, sufficient for TLI.
You need now 5,930 - 3,982 = 1,948 m/s additional delta-V to complete
the landing. The delta-V you get from the Star 30 will be:
293*9.8ln((492+400)/(28+400)) = 2,109 m/s, sufficient for the landing.
The total gross mass of the 3 stages plus payload will be
2,134.3+1,148+492+400 = 4,174.3 kg, within the lift capacity of the
Dnepr 1. The cost of the Dnepr 1 might be $13 million. The costs of
the upper stages? The Astronautix page on the PAM-S powered by the
Star 48 motor gives the price as $4.06 million. The Star 37 is smaller
by half, and the Star 30 is smaller by an additional factor of one-
half. Then we might estimate their prices as $2 million and $1 million
respectively, for a total cost of these upper stages of $7 million.
Then the total launch cost might be $20 million.
We would have to add onto that the cost of the avionics and the cost
of the lander.


As a point of comparison in regards to the feasibility of using solid
motor upper stages for the purpose, the Dnepr launcher has been
studied to be used to launch a 500 kg payload to GEO by using two Star
solid motor upper stages and lunar gravity assist:

Dnepr (R-36M2)
"The Dnepr launch vehicie does not have the capability to deploy
payloads directly into GTO. However, Kosmotras has studied a technique
to deliver small spacecraft to GEO using the gravity of the Moon to
provide the plane change and perigee raising. In this scenario, the
spacecraft is attached to Star 48A and Star 27 solid motors, supplied
separately by ATK Thiokol. The Star48A would send the spacecraft to
the Moon, where a gravity slingshot maneuver would lower the transfer
orbit inclination from 50.5 deg to 0 deg, and raise the orbit perigee
to geostationary altitude. When the spacecraft reaches perigee of the
new transfer orbit, the Star 27 motor would fire to circularize the
orbit at GEO. Using this method, a 500 kg (1100 lbm) spacecraft could
be delivered to GEO."
http://www.b14643.de/Spacerockets_1/...tion/Frame.htm


Bob Clark

Dear hall...:

On Nov 23, 6:52*am, " wrote:
....
The US is BROKE!!! lets repeat that WERE BROKE

We dont have the $$$ to fund social security, day to
day operations, let alone moon operations

Let alone the NEXT terrorist hit here will collapse what
remains of our economy.......

under such grim [fiancials] it will be hard to impossible
to raise funds for any out of this world projects.......

We are broke, because we have no frontier... it seems to me. Anybody
with a new idea gets thrown in jail, rather than powering expansion.
90+% of my state's budget is spent providing / operating prisons,
either for children while their parents work, or those found guilty of
some crime.

We have no money for social security, because those with money have
other needs than supporting old people. We've stopped taking care of
our parents (or those folks that decided it was inconvenient to have
kids), so I guess we have to implement "Logan's Run" or "Solylent
Green". Because the politicians borrowed that money to fund day-to-
day operations, replacing it with IOUs for decades.

Terrorists won't collapse the economy, the politicians will do it for
them, right afterwards. Watch, they'll trigger an EMP blast into our
luggage next, to make sure any electronics there is disabled
permanently.

We either invest in the future (and a space plane that does more
damage to the ozone layer than the Concorde, isn't that), or we won't
have one.

I disagree with Robert's (apparently) favorite method to getting to
space, but we need something. The money spent never leaves the
planet. It only makes high-tech jobs. Jobs that aren't dedicated
towards reducing the surface population one explosion at a time.

David A. Smith

And... is it supported in Human experiments like throwing women
passengers out at the end of the trip because they could not withstand
accelerations, like in that advertisement of a shooting star that
looked quite like my girlfriend of the Berlioz family I could not hold
on to the laptop/card documenting the pic in it because the laptop was
stolen? I suspect this of more than one such spatial ventu not
enough funding for biophysical research or modeling being offset by
direct Human experimentation...

Danilo J Bonsignore

On 14 nov, 14:50, Matt wrote:
Not a mathematical calculation, but a question about the societies
portrayed in science fiction: If only the "cream crop" is sent, who
will clean toilets on the settled worlds? Did you ever see anyone
cleaning toilets on the Starship Enterprise?

According to Douglas Adams in "The restaurant at the end of the
Universe", Earth was actually colonized by the crew of a spaceship
composed entirely of middle-man, including toilet cleaners and hair
stylists. To answer your other question, still according to the book,
the original population of the planet died after a pestilent infection
spread because there was no one to clean telephones anymore...

Adams saw farther than most...