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Here's a little something I found on the Uselessnet. Responses 1 and 2 are the two halves of the posting.
2 responses total.
From: mlindroo@aton.abo.fi
Newsgroups: sci.space.tech
Subject: (1/2) Soviet manned Mars mission using nuclear/ion propulsion (long)
Date: Thu, 14 May 1998 08:09:29 GMT
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I have finally found some detailed information on Soviet plans
for manned Mars exploration. "Power-Propulsion Systems for
Orbital Nuclear Transfer" by Koroteyev et al. (ACTA
ASTRONAUTICA Vol.24 pp.181,1991). The authors compare the merits
of NERVA type nuclear thermal engines, a hybrid chemical/ion
propulsion system before finally settling for pure nuclear
electric (ion) propulsion. The Russians appear to have
consistently favored this low thrust, high Isp option since the
early 1970s; this in stark contrast to NASA's plans which have
mostly been based on high-thrust chemical or nuclear thermal
propulsion. The Soviet approach to manned Mars exploration is
both interesting and strikingly different. I will first
describe the mission and then discuss the merits and drawbacks
compared with American plans such as Mars Direct.
1.NUCLEAR POWER PROPULSION OPTIONS
----------------------------------
The Soviet option used a huge nuclear reactor generating up to
50,000kW -- 450 times more power than the International
Space Station's solar panels! In addition to producing
electrical power, the reactor's thermal output can be used to
heat liquid hydrogen propellant to 2800K if the nuclear-thermal
rocket (=NERVA) option is used. A 460-day two-way mission to
Mars orbit would have to start from a high 800km Earth parking
orbit due to the radiation hazard. The total mass would be 800
metric tons, including 495t of LH2 fuel, a 150t payload, a
70t reactor/propulsion module (200kW,200kN thrust,917s Isp).
The remainder is presumably LH2 tankage.
---
Option#2 would feature a chemical liquid oxygen/hydrogen
rocket (1.1mN thrust,480s Isp) for quick acceleration through
the Earth's radiation belts. A 260N nuclear-ion engine would
then provide the thrust for the remainder of the journey. Total
spacecraft mass would decrease to 700t (including 310t of
LOX/LH2 propellant plus 147t of lithium fuel for the ion drive)
but the total mission duration would increase to 615 days. The
required delta-Vs from an 800km parking orbit are 3.1km/s for
the chemical rocket plus 25km/s for the low-thrust transfer to
Mars orbit & return to Earth. Both missions (option#1 & 2)
would have taken place in 2018 and permitted a 30 day stay on
Mars.
2.NUCLEAR-ELECTRIC MARS SPACECRAFT
----------------------------------
The option finally chosen was a pure nuclear-electric system
using closed Brayton-cycle gas turbines working at 1800K
temperature for heat-to-electricity conversion.
The total travel time would be reduced to 320 days. In
addition, the spacecraft would weigh less (550 metric tons,
including 300 tonnes of xenon fuel). The delta-V would increase
to 47km/s, however.
---
An ASCII diagram of the spacecraft is shown below.
Mars
landing
vehicle
Radiators __
|_| | |
|_| | |
|_| | |
|> |_| _____ _ _____ |__|
=/|____________||_______________________________|_|/ \/ \/ \\__/
||/_\/_\/_\ooo||oooooooo\/\/\/\/\/\/\/\/\/\/\/\|_| || || ||__|
=\| ||-------------------------------|_|\_____/\_/\_____/ ||
|> |_| ||
|_| /\
|_| ==
|_| Mars
Reactor Ion thrusters Orbital
("o"=propellant tanks) Vehicle
habitation Earth
modules return
capsule
+airlock
--------------------------------------------
WEIGHT BREAKDOWN
--------------------------------------------
2 x 25mW reactors+radiation shield.......14t
2 x turbogenerators......................15t
2 x heat exchangers .....................18t
Radiators................................22t
Control system............................1t
Ion engines .............................10t
Boom (60kg/meter).........................6t
Fuel tanks................................7t
Refrigerator+misc. .......................7t
Xenon fuel..............................300t
Mars Orbital Vehicle (MOV)...............80t
Mars Landing Vehicle (MLV)...............60t
Earth Return Capsule.(ERV)...............10t
--------------------------------------------
550t
An interesting feature of the propulsion system is that the
thrust can be increased (at the expense of a decreased Isp /
higher propellant consumption) for major maneuvers that have
to be carried out quickly. Most important among these is the
initial departure from parking orbit;the spacecraft "spirals
outwards" throught the van Allen belts to an ~185,000km
circular orbit in just 7.5 days. From there on, the engines
operate in a more economical high-Isp / low-thrust mode.
(Ion thrusters are very efficient because their
exhaust velocity is high, so less fuel will have to be carried
on board. The main drawback is the low thrust which means the
engines will have to be fired continuously for days or months.)
---
This "all-up" mission would begin in May 2018 and could
deliver a 4-crew to Mars orbit, where two cosmonauts would
descend to the martian surface in the MLV for 2-3 weeks of
exploration. However, it would be more economical to launch
an UNMANNED CARGO SPACECRAFT on a one-way mission in 2016,
carrying the MLV plus a "temporal research [orbital space-]
station". Payload mass in Mars orbit would be 150t, but the
total mass in LEO would decrease to 280t. This is because
the cargo spacecraft would not have to carry fuel for the
return trip and because a slower, more economical 320-day
trajectory could be used. Only one 25mW reactor would be
required so the mass of the propulsion system is reduced by
almost 50%.
The MANNED SPACECRAFT (sans MLV and carrying 53t less propellant)
would depart from Earth in May 2018. Below is a timetable for
that mission:
Thrust Exhaust Duration Propellant
velocity consumed dV Remarks
(=Isp/0.0098)
-------------------------------------------------------------
3500N 20km/s 7.5d 113.5t 6.0km/s 800km->185,000km
Earth orbit
1000N 90km/s 17.5d 16.5t 4.7km/s ->1.25AU from Sun
(60 day interplanetary coast period;)
1000N 90km/s 30.0d 29.0t 8.9km/s 1.38 to 1.39AU distance
1500N 50km/s 5.0d 13.0t 2.4km/s Mars orbit insertion
(21 days of Mars exploration in October 2018)
1500N 50km/s 5.0d 13.0t 2.4km/s Departure from Mars
3500N 90km/s 50.0d 48.0t 19.0km/s Transfer from 1.38
to 0.78AU
(100 day coast at 0.72 to 0.78AU distance)
1000N 90km/s 12.0d 11.5t 5.2km/s* Earth arrival,March 2019
-------------------------------------------------------------
~320 days 244 tonnes 45km/s total delta-V
3.COMMENTS / POSTSCRIPT
-----------------------
Nuclear-electric propulsion appears to reduce the mass that
has to be launched into low Earth orbit by a factor of 1.5-2
vs. chemical rockets for unmanned cargo missions. This does
not necessarily translate into reduced costs, however -- the
reactor+ion thrusters appear to push the state-of-the-art in
many ways (e.g. power-to-mass ratio) and xenon is both rare
and expensive.
---
The manned example would not be practical using chemical propulsion;
NASA's 15-month "Sprint" proposal from 1986 would have provided
a similar 3-week Mars exploration opportunity but the required
delta-V (7km/s+ to Mars, 4km/s to Earth) was much too high for
a single spacecraft. Like the Soviets, NASA would have launched
part of the unmanned payload in advance, on a more economical
trajectory.
---
In the wake of the Chernobyl disaster, the Soviets wisely
rejected the huge space reactor concept on grounds of political
risk. Consequently it was decided to investigate if large
solar-dynamic & photovoltaic systems could power the ion
thrusters instead. The favored option (described by David
Portree [ http://members.aol.com/dsfportree/explore9.htm ])
featured two giant 40,000 square meter solar panels and
the required technology could have some interesting applications
for solar power satellites. I will discuss this in Part 2
of this article.
MARCU$
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From: mlindroo@aton.abo.fi
Newsgroups: sci.space.tech
Subject: (2/2) SPS & sending a space station to Mars using solar/ion propulsion
(long) Date: Thu, 14 May 1998 08:18:27 GMT Organization: Deja News - The Leader
in Internet Discussion Lines: 183 Approved: sci-space-tech@isu.isunet.edu
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0.INTRODUCTION
--------------
In 1989, the Soviets unveiled a proposal for a manned mission
to Mars using solar-powered electric/ion propulsion (SEPS). In
most respects, the spacecraft is identical to the nuclear-
electric ship described in my previous article. The crucial
difference is that the power-generation system has been
replaced with two giant 40,000 square meter solar panels, each
generating 7600 kilowatts at Earth's distance from the Sun and
3500kW at Mars -- considerably less than the nuclear-electric
option's maximum output of 50,000kW. Consequently the mission
takes longer; two years including a brief 7-day stay on Mars.
Total spacecraft mass is however reduced to 355t, including a
60t Mars orbiting "mothercraft", 10t return vehicle, 60t Mars
lander, 165t of xenon propellant and 40t solar power/ion
propulsion module.
---
The spacecraft departs from a 500km high Earth orbit and slowly
spirals outward. It is 40,000km high after one month. A major
problem thus becomes crew exposure to radiation inside the van
Allen belts:
the Soviets estimate the total mission dose would be 230 REM
with 130 REM received near Earth. However, it should be possible
to launch the crew separately on a small chemical rocket and
have it rendevouz with the Mars spacecraft in geostationary
orbit or beyond. The mission is described in greater detail
on
[ http://solar.rtd.utk.edu:81/~mwade/articles/aelita.htm ]
and
[ http://members.aol.com/dsfportree/explore9.htm ]
1.IS THIS THE "BEST" CANDIDATE FOR A MANNED MARS MISSION?!
------------------------------------------------------
I feel the Russian solar-electric proposal has some interesting
advantages that may yet make it the frontrunner as a possible
follow-on to the International Space Station... History clearly
shows that technology for major manned space projects such as
Shuttle or Station usually gets developed not strictly on
technical merit, but for "political" reasons. By "politics",
I am not only referring to cost considerations or porkbarrel
but also to how appealing the technology is to scientific or
commercial needs. Solar-electric ion propulsion might have
the following advantages:
2.1 USEFUL COMMERCIAL SPINOFF TECHNOLOGIES,
NO NEED TO DEVELOP A VERY LARGE EXPENDABLE LAUNCHER:
For high power solar-electric propulsion,
the engineering challenge is mostly confined to the development
of large but lightweight radiation-hardened solar panels as well
as power-efficient ion thrusters. Previous research has indicated
that solar array production costs would have to be reduced by an
order of magnitude for solar-electric propulsion to be very
competitive with other manned Mars options. Developing these
capabilities might still cost billions even if the goal can be
achieved, but the finished product would be of greater use for
commercial satellite technology and space science than the other
path to Mars: a large expendable heavy-lift launch vehicle plus
nuclear thermal (NERVA) upper stage. Mars Semi-Direct proponents
claim a Saturn-class HLLV would be sufficient, but political &
safety considerations would dictate a larger spacecraft that can
only be launched on a bigger HLLV or the politically unacceptable
NERVA upper stage. The SEPS Mars spacecraft could however be
launched quickly on an uprated EELV, large RLV or
Shuttle-C/Energia class boosters in the 70-90t to LEO class.
2.2. COMPATIBILITY WITH PREVIOUS SPACE STATION WORK (Salyut to ISS)
SEPS would seem to be a perfect match for the current American
and Russian space programs, which have extensive experience from
flying and operating space stations in Earth orbit. For all
intents and purposes, the SEPS Mars spacecraft would be a large
"slowly moving space station" exposed to only very low acceleration,
no aerobraking at Mars etc.. The habitation modules could thus
have a high degree of commonality with previously developed
hardware from the Salyut/Mir & ISS projects.
2.3.REUSABLE INTERPLANETARY SPACECRAFT
A practical Mars mission based on chemical propulsion would
have to use the same throwaway approach as Apollo; the only
"reusable" element would likely be the tiny Earth return
capsule. In comparison, solar-electric propulsion would return
the entire spacecraft to Earth orbit where it could be
refurbished for another mission. The only expendable component
would be a small Mars lander that must be used to ferry
astronauts between the martian surface and the orbiting mother-
craft. In addition to this, the mass that must be lifted to
low Earth orbit would be 1.5 to 2 times greater if chemical
rockets are used for the trans-Mars injection burn. The total
launch cost & assembly time would thus be greater for chemical
propulsion. Nuclear thermal rockets would suffer from the
same problems, although the higher specific impulse would
translate to a slightly lower spacecraft launch weight.
---
Granted, other factors have to be considered too. Some
HLLV/chemical rocket based plans would land most of the payload
on Mars where it could be used again by subsequent expeditions.
This is however no unique advantage for chemical propulsion (an
ion rocket could do this too). The recurring cost per mission
could still be high for solar electric propulsion, if the giant
solar panels and ion thrusters have to be replaced after each
mission. Assembling the huge solar panels in LEO might also be
more complicated than if chemical propulsion modules (which
tend to be fairly compact) are used.
2.4.HEAVY-LIFT MISSIONS TO THE MARTIAN MOONS OR ASTEROIDS
Phobos and Deimos could be useful for Mars-visiting crews
since they could be mined for water and minerals. Propellant
production facilities and other heavy equipment would however
have to be brought in from Earth before these resources can be
exploited. Since the moons are orbiting fairly deep inside Mars'
"gravity well", a spacecraft powered by chemical propulsion
would need to perform an aerobraking maneuver followed by a
rocket burn to circularize its orbit and land on Phobos/Deimos.
Ion propulsion would likely be more efficient than aerobraking
when one factors in the additional mass of an aerobrake+chemical
rocket, and not subject the payload to the risks & deceleration
loads.
---
Solar-electric propulsion would seem to have more potential
for manned missions to near-Earth asteroids as well. These
are difficult to reach using chemical or even nuclear thermal
propulsion since most asteroid orbits are inclined to the
ecliptic plane and because aerocapture can't be used. A very
high delta-V will be required to reduce the travel time to
level acceptable to humans.
3.DRAWBACKS OF SOLAR-ELECTRIC PROPULSION
----------------------------------------
Ion propulsion naturally has some disadvantages too. Perhaps
the biggest drawback is the crew would be stranded in a
heliocentric orbit if the propulsion system fails. This is
because the ion thrusters have to fire continuously for months
as the spacecraft slowly spirals outward towards Mars. However,
any interplanetary spacecraft would have to carry lots of
engines since the thrust of each ion engine is low. The impact
of losing a few engines would thus be small.
---
Another problem is the long "spiral out" time, which forces
the crew to spend up to a month inside the Earth's radiation
belts. This might do considerable damage to the solar panels and
electronics on the spacecraft as well. However, as noted earlier
it would be possible to launch the spacecraft in unmanned mode
from LEO. The crew could later transfer to (for example-)
the L-2 point in a small "taxi" using chemical propulsion.
The taxi (which could be based on existing capsules such as
Soyuz) then rendezvouses & docks with the waiting interplanetary
spacecraft.
---
The Russian 2-year baseline mission might not appear to
offer good value for money since the crew only gets one week of
exploration on Mars. However, there doesn't seem to be any
reason related to the choice of propulsion why the mission
could not be extended to three years.
4.CONCLUSION / SEND THE INTERNATIONAL SPACE STATION TO MARS??
-----------------------------------------------------------
Building & launching a 355 tonne solar-electric spacecraft would
clearly be an enormous undertaking. However, it might be possible
to re-adapt some hardware previously developed for the International
Space Station (ISS). Another fascinating opportunity would be
to test solar power satellite systems. On a logarithmic scale,
the Mars craft (15,200kW;80,000m2 solar panel area) would be
roughly halfway between ISS (76kW; 2415m2 area) and a solar
power satellite (5,000,000kW+; 50,000,000m2+ area) in terms
of technical complexity. Hopefully, lessons learned from previous
projects (ISS->Mars craft->SPS) would make it easier to develop
lighter and more efficient solar panels & assemble them in
orbit.
---
One wonders if it would have been possible to send ISS to Mars
in ~2010, if NASA had made certain design choices when planning
the facility (e.g. radiation-hardened crew modules). Perhaps
not -- the poor station has been bogged down by overly ambitious
user requirements right from the start. It does seem like the
Russian solar-electric propulsion option would allow us to
utilize Mir/ISS hardware to the fullest possible extent, though.
MARCU$
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