Exploration Alternatives

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

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 boiloffchanges this order

– Can “buddy tank” stages together but $/kg gets worse

– Can have a lower kick stage (solid or liquid) but $/kg gets worse

Earth escape of existing stages

Launcher “Stage”

Prop (kg)

MECO(kg)

Thrust (lbf)

Isp(s)

Propellant @ zero payload (250km)

3.22km/sdVpayloadNo 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

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

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

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

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

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

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

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

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?

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

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

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