Jeff Greason's slide deck for NewSpace 2014

1. Exploration Alternatives
Is human exploration possible
without U.S. government funding it?

2. Earth Departure Stage
• Maximum performance for minimum price
achieved by launching payload first, then a
LOX/LH2 departure stage
– Depots or a “refrigerant module” to avoid boiloff
changes this order
– Can “buddy tank” stages together but $/kg gets
worse
– Can have a lower kick stage (solid or liquid) but
$/kg gets worse

3. Earth escape of existing stages
Launcher
“Stage”
Prop
(kg)
MECO
(kg)
Thrust
(lbf)
Isp
(s)
Propellant
@ zero
payload
(250km)
3.22km/s
dV
payload
No refuel
$/kg to
Escape
(very
rough)
Ariane 5 ECA
“ESC”
14900 5290 14500 446 14900 13820 $26000
Atlas 551
“Centaur”
20830 2550 22300 451 20400 16170 $11000
Delta IV-H
“5m stage”
27200 3490 27200 462 27200 22720 $15500
Falcon 9 90000 4900 180000 342 13150 3230 $27000

4. Light capsule
• Scaling human missions down to
current launchers requires
minimum crew (2), and a
minimal capsule
• Helps a lot if there are several
docking adapters to allow
flexibility in stacking
• Mars Curiosity heat shield 4.6m
diameter with 2600kg entry
mass to Earth or Mars
• Can be used in several beyond
earth missions; has volume for 2
crew on lunar missions without
additional hab

5. LEO to L1/Lunar Surface
• Long term LOX/LH2 storage needed
– 10-20kWe, plus radiators of similar capacity
– Sunshade to keep tanks shaded reduces this
– Docking ports for Centaurs plus reusable landers
based on Centaur tanks plus to hab
• Hab module based on BEAM or Cygnus
• Can put both hab and refrigerator at L1 with Atlas
551. Second 551 to place a lander.
• Including boiloff allowance full Centaur puts 10800kg
to L1, so can push an F9 payload
• Light capsule/heatshield allows crew to be brought
up with enough propellant for a lunar round trip on
single Atlas 551
• Total launch cost: $720M for first mission, $270M
$450M for each following manned round trip, or
$450M to soft land 5t cargo with one way lander:
can just keep going for more missions

6. LEO to Mars Flyby
• Inspiration Mars studies suggest 12000-13000
kg for hab, consumables, and SM
• With light capsule, perhaps 15000-16000kg
• Delta IV Heavy upper stage pushes ~18500kg
from LEO to 3.6 km/s delta V
• Remaining 1.3 km/s provided by electric
propulsion, <100kw concentrator array (SLA),
12 XIPS, 670 kg Xenon
• Atlas 551 for payload, then F9/Dragon for crew
& cargo, then Delta IV Heavy for departure
stage: $620M launch
• This all depends on tolerable GCR hazard; but
we’d find out

7. Dealing with Earth-Orbit Rendezvous
• Challenge for these missions is having time to assemble it all in LEO
• Centaur can loiter a while, Delta IV cannot
• 3 launches: A551, F9/Dragon, Delta IV. Reasonable assumption is 80-90%
likelihood of getting all 3 off succesfully if have time for the launches
• Launch hab/sm/return capsule first; needs to stay on orbit for a while –
say a month.
– Precession of orbit axis reduces launch opportunities to roughly 3-day window
for alignment to mars departure trajectory
• Launch crew on F9/Dragon early; carry extra
consumables– leave excess in the Dragon
• Gives ~3 days to get the Delta off with launch
opportunity every day. Launch to first-orbit-
rendezvous, dock using Dragon SM for the
maneuvering stage, jettison Dragon and TMI
burn on the next orbit
• Stationkeeping solution for plane change would
be a valuable technology

8. L1 Staging
• If need longer launch windows, and if Lunar missions have
already been done, can stage at L1 for Mars missions
– Use refrigerator node to store departure centaurs and transfer
propellant from one to another
– Full Centaur can push 36000 kg from L1 to Mars Flyby with Earth
swingby and burn at perigee; adds 2 Van Allen passes
– Austere Mars Flyby doesn’t need full Centaur; can do a 3-
Centaur mission; two to push the hardware and propellant
uphill, one to push the crew and some propellant uphill, then
consolidate propellant and go with one of the Centaurs. That’s
3 A551 plus 3F9 ($810M launch)
• Of course, once missions are departing from L1, whether to
Moon or Mars, opportunity for adding Lunar propellant
exists to reduce cost further

9. Deimos landing & Return
• Outer edge of current launchers; requires austere hab of Mars Flyby
to work
• Put small propulsion module on capsule, 1100m/s storable
• Max out weight with electric propellant
• First mission sends hab + SM and electric propulsion to spiral down
to Deimos
• Second mission sends crew & hab: capsule and propulsion
aerobrakes to Mars aerocapture then biellptic transfer to Deimos.
Hab spirals in to Deimos orbit electrically over months
• For departure, bring crew & additional Xenon from Deimos up to
hab in capsule, then spiral out & return to Earth
• Launch cost $1840M
• Again, this is pushed right to the edge but if it works – no limit to
how many supplies you can cache at Deimos (or Phobos) this way.
In principle, could keep going until had a Mars crewed lander there

10. Upgrades
• By far, biggest improvement is a bigger but high mass ratio, high Isp earth
departure stage with provision for low boiloff and/or plug-in to
refrigeration module
• The ACES stage on ULA’s roadmap is just right for this; they’ve estimated
$1B for that and a funded customer would be a big help
• That would give about 1.6x increase in mass for Delta 4 heavy EDS launch
and would not take many missions to pay off investment
• Centaur-derived EDS can be aerobraked back to LEO for reuse, which
reduces cost of each subsequent Lunar mission

11. Financing
• Hardware and NRE cost not discussed; wide variety in
what that takes from team to team; swag total mission
at 2-3x launch costs
• Like any exploration mission in history, hard to get
initial expeditions to be profitable
• But <$2B price tags within philanthropic reach – and, if
done intelligently, the ventures that do them now have
a paid off Lunar/Mars transportation system with
possible other customers
• Can we really not figure out how to make back the
media sales of Avatar or Star Wars or E.T from going to
real life other planets, if privately run and no limit on
media rights or sponsorship?

12. Cosmic Radiation Limits
• Austere Mars missions assume GCR hazard is tolerable. This is not
currently known to be true or false
• If GCR is hazardous (high REM/RAD), then what do we do?
• Takes about 0.2m of polyethylene to reduce GCR to type of radiation we
are familiar with and is within our experience base for chronic exposures
on Earth
• If starting from Cygnus module, as with mars flyby, that cuts habitable
volume to 450 ft3 and adds 11000 kg of shielding which can be modular
• Two such modules could then provide long term habitat for two people
plus life support; and if 20m cable connects them, can do Mars gravity at
6RPM
• 5 Atlas + 5 F9 to put this in L1 and electric boost to mars cycler trajectory.
$1350M launch cost for each cycler
• May require two cyclers to take advantage of each Mars opportunity

13. Cycler-based missions
• No longer need to carry hab on each mission
– just capsule and consumables and fuel
• S1L1-B cycler (M-E-E) offers low delta V:
– 4.7km/s excess leaving earth, 5-6.5km/s mars
– Real orbital mechanics takes some hab maneuvers
– Delta-V for hab at most 1 km/s, most orbits nothing
• 2 A551 to bring up 2 centaurs to go to hab, lofting crew capsule,
consumables to come on two F9, and about 2.8 km/s storable stage for
capsule: $540M launch cost
• Centaurs carry payloads to hab, capsule aerobrakes at mars, bielliptic to
Deimos, then biellptic to go from Deimos down and out to cycler for
return, aerobrakes at Earth
• Again, right on the edge – expensive, but doable. Still, if doing this
mission would certainly pay to fund ACES stage
• A successful Mars flyby showing GCR hazards tolerable has huge value

14. References
• Delta IV Launch Services User’s Guide, ULA, June 2013, retrieved from
http://www.ulalaunch.com/uploads/docs/Launch_Vehicles/Delta_IV_Users_Guide_June_2013.pdf
• D. Tito et. al., “Feasibility Analysis for a Manned Mars Free-Return Mission in 2018”, 2013 IEEE
Aerospace Conference, March 2013
• K. Edquist et. al., “Aerothermodynamic Design of the Mars Science Laboratory Backshell and
Parachute Cone”, and “Aerothermodynamic Design of the Mars Science Laboratory Heatshield”,
AIAA Paper 2009-4075
• M O’Neill et. al., “Stretched Lens Array Squarerigger (SLASR): A Unique High-Power Solar Array for
Exploration Missions”, IAC-05-C3.2.01, International Astronautical Congress, 2005. Retrieved from
http://www.markoneill.com/IAC-SLASR-2005.pdf
• M. Lara, “Repeat Ground Track Orbits of the Earth Tesseral Problem as Bifurcations of the Equatorial
Family of Periodic Orbits”, Celestial Mechanics and Dynamical Astronomy, v 86 p 143-162, 2003
• J. Hopkins, W. Pratt, “Comparison of Deimos and Phobos as Destinations for Human Exploration and
Identification of Preferred Landing Sites”, AIAA SPACE 2001, AIAA 2011-7140, September 2011
• T. McConaghy, J. Longuski, D. Byrnes, “Analysis of a Class of Earth-Mars Cycler Trajectories”, Journal
of Spacecraft and Rockets, v 41 n4, July-August 2004
• T. McConaghy, J. Longuski, D. Byrnes, “Analysis of a Broad Class of Earth-Mars Cycler Trajectories”,
AIAA/AAS Astrodynamics Specialist Conference, Monterey, CA, Aug 2002, retrieved from http://trs-
new.jpl.nasa.gov/dspace/bitstream/2014/8886/1/02-1454.pdf
• B. Birckenstaedt, B. Kutter, F. Zegler, “Centaur Application to Robotic and Crewed Lunar Lander
Evolution”, STAIF 2007, AIP conference proceedings, v 880 p 779ff, 2007
• W Chen et. al., “Effects of Cobalt-60 Exposure on Health of Taiwan Residents Suggest New
Approach Needed in Radiation Protection”, Dose-Response, v 5, p 63-75, 2007