Mars Mission Outline 2016

Update 9-14-19:  I have updated and refined this mission plan as “Mars Mission Outline 2019”,  dated 14 September 2019.  That update enlarges the manned transport habitation,  updates the reusable Mars landers to handle larger payload and a rendezvous allowance,  and makes the solar electric propulsion “tugs” recoverable and reusable.  The Earth departure stage is still not recovered,  although that could be done.  The 2019 update now supersedes this 2016 article,  which in turn superseded the 2013 version. 

Update 8-17-18:  this article is a rework of the ideas originally posted 12-13-13 as "Mars Mission Study 2013".  What got added here was fundamentally only using electric propulsion to send ahead the unmanned assets.  That got a huge reduction in launched masses.  The same notions of a manned orbit-to-orbit transport and one-stage,  two-way landers were retained.  Some of the details got refined better in this 2016 article.  If you want to see the earlier one,  use the navigation tool on the left.  Click on 2013,  then December.  It is top of the December list for that year's postings.  
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Update 5-30-16:  this article culminates about 6 years' efforts looking at manned Mars mission ideas.  There are several other related earlier articles about men to Mars on this site,  but this one is by far the best.  It does reference the lander study article below:  "Reusable Chemical Mars Landing Boats Are Feasible",  dated 8-31-2013.    

It was from that lander study article that I drew out my results for the NTO-hydrazine blend lander as the one compatible with the same long-term storable fuel used here to push the orbit-to-orbit manned transport.  That allows fuel allocations to fit conditions without changing hardware.  

The big difference with this study is the huge reductions in launched masses and costs that I saw by sending the unmanned stuff "slowboat" ahead with electric propulsion.  But,  if I did that with the manned portion,  it would turn a 2.5 year trip into a 3.3 year trip for the astronauts,  so I did not do that.  

Update 5-31-16:  corrections to the radiation formula and some typos below,  in red.  

Summary

It should be possible to send six people to the surface of Mars with current technology and very-near-term ready-able hardware,  for under $50 billion,  and get them back in good health.  This can be done while recovering the manned spacecraft in Earth orbit for reuse on other missions.  Typical claims by industry giants and NASA itself that this cannot be done for under half a trillion dollars are nonsense.

Mission Architecture

I used the 1950’s orbit-to-orbit transport concept for the manned portion.  The inspiration was an item in the 1957 Disney “Tomorrowland” film “Mars and Beyond”.   That concept was a fleet of nuclear-powered ion thruster-propelled ships built in Earth orbit.  These traveled to Mars orbit,  where they sent chemical rocket-powered landers to the surface.  They returned for recovery (and presumably re-use) in Earth orbit.  There being six ships,  each with a lander rocket,  up to six different sites could be explored at Mars.  See Figure 1 for two illustrations from the Disney film.  

The actual technologies depicted are obsolete or otherwise in error.   But,  the key concepts are still valid,  so I used them.  These were :  (1) the orbit-to-orbit transport,   (2) separate landers,  (3) ability to visit multiple sites,  and (4) recovery and reuse of essential hardware.

Figure 1 –6-Ship Exploration Fleet,  from Disney’s 1957 film “Mars and Beyond”

The orbit-to-orbit transport vehicle serves only to transport the crew safely both ways,  and to house those not on the surface safely while in Mars orbit.  It comprises a spacious “deep space habitat” plus a crew return capsule plus suitable engines that burn long-term storable propellants for arrival at Mars (nitrogen tetroxide and a hydrazine blend).  It departs with a liquid hydrogen/liquid oxygen departure stage,  then spins for artificial gravity during a Hohmann min energy transfer orbit (the minimum practical flight time of 8-9 months at minimum expended energy).  It returns to Earth using all storable propellants sent ahead to Mars orbit.  The crew of six lives in the habitat,  not the capsule,  for this trip. 

Landers,  their propellant supply,  and the Earth return propellants for the manned vehicle are sent ahead to low Mars orbit,  where the manned vehicle must rendezvous with them.  These unmanned vehicles are sent by solar electric propulsion on longer transits that spiral-out and spiral-in.  The electric propulsion does save enormous amounts of weights and costs,  but adds months to the trip time. 

The landers are single stage re-usable chemical vehicles,  whose crewed control cabin doubles as an abort capsule capable of getting the crew to the surface from anywhere along the descent or ascent trajectories,  from orbit,  but not to orbit,  from the surface.  To provide rescue for a stranded crew,  three landers are sent as fully-fueled items,  plus extra lander propellant to support repeat trips.  That gives a rescue capability even if one lander fails,  or is otherwise lost,  for any reason at all.  The current concept uses the same storable propellants as the manned vehicle. 

The lander is the sole item to be developed “from scratch”,  and may require assembly in Earth orbit.  It is “the long pole in the tent”,  and will pace the program.  It has no new technologies,  but this combination has not been tried before.  The rest is a reshaping or re-sizing of things we already build. 

The trip to and from Mars is about 8-9 months long (one-way) for the manned vehicle,  with roughly a year in orbit at Mars.  This is thus about a 2.5 year mission for the crew,  which does not launch until the supplies are verified as sent ahead successfully to Mars orbit.  While in orbit,  three crew take a lander to the surface for up to a 30-day stay,  while the other three crew do science from orbit,  and provide the rescue capability with the other landers.  This process repeats,  alternating surface crews,  for up to eight landings,  without compromising the rescue capability.  Spin gravity is maintained in Mars orbit. 

Earth return propellant is also sent ahead to Mars orbit with solar electric propulsion,  like the landers and their supplies.  The manned vehicle discards empties,  then docks with the return propellant tanks.  Once Mars departure is complete,  those tanks are staged away,  and the vehicle spun-up for gravity during the voyage home.  The last of the propellant is used to enter Earth orbit,  so that the vehicle can be recovered,  refitted,  and reused,  on other missions (unique among the mission plans I have seen in recent decades).  The crew finally returns home in the attached capsule,  which is rated for a full free-return “emergency bailout” reentry,  in the event that vehicle propulsion should fail approaching Earth. 

Assets left in Mars orbit are the three landers,  plus the empty Mars arrival and lander propellant tanks,  and the four depleted solar electric transfer stages (for the Earth return propellants and for the three landers and their propellants).  With a little extra ion propellant,  these electric propulsion stages could be brought home,  too.  It is probably wise to leave the landers in Mars orbit,  for future re-use,  with more propellants either brought from Earth,  or manufactured on Mars,  or both. 

Propellant Selections

For the manned vehicle and the landers,  I chose nitrogen tetroxide and a hydrazine blend at a very conservative 300 sec specific impulse,  precisely because it is very long-term storable in space without any boiloff risks.  This simplifies propellant tank designs,  enabling the practical achievement of very low tank inert mass fractions,  something fundamentally not possible with insulated and sun-shielded Dewars for cryogenics.  I used the same for the landers,  except at a more realistic 317 seconds of specific impulse for a modest chamber pressure and slight cant angle reduction. 

For the electric propulsion stages,  current ion thrusters usually use xenon,  but I would encourage the use of argon instead,  as far more widely available.  Both are noble gases,  and either should work just fine.  The stage inert weights I used include large solar electric panels.  I assumed 5000 sec specific impulse for these,  which should also be conservative. 

For the Earth departure stage (of the manned vehicle),  I chose liquid hydrogen-liquid oxygen at about 465 seconds of specific impulse (rather realistic).  I chose it for the performance in spite of the boiloff risks.  This item is conceived as the last thing to be sent up and docked to the manned vehicle,  at about the same time as the crew arrives.  It may have to be sent up in multiple pieces and docked together in orbit,  but the boiloff risk is only days long,  not months or years. 

Trip Requirements

The delta-vee requirements from Earth orbit to Mars orbit depend upon whether Mars is a little nearer or farther away at opposition.  To be conservative,  I used the maximum Earth departure delta-vee of 3.8 km/sec,  and the maximum Mars arrival delta-vee of 2.4 km/s,  even though these numbers “do not go together”.  That’s 6.2 km/s one-way,  and 12.4 km/s two-way,  even though the largest two-way delta-vee is really just 11.9 km/s.  A “real” design will therefore be much better than depicted here,  especially since I used such a low conservative value for specific impulse,  too.    

For the solar electric propulsion,  as a rule-of-thumb,  I simply doubled the delta-vee requirements to account for the gravity losses of long-duration non-impulsive “burns”.  It wouldn’t make very much difference to triple the delta-vee,  due to the high specific impulse available with ion thrusters. 

The Deep Space Habitat:  Source Technology

For this,  I assumed the characteristics of the Bigelow B330 inflatable module,  which in fact is based upon NASA’s own design characteristics for a deep space habitat.  The habitat uses two of the modified B330 modules.  The difference are:   (1) I assumed heavier mass to account for strengthening the modules for higher applied loads,  and (2) I spin the assembly for artificial gravity (why the extra strength is needed). 

To this I add one custom “hard” module in the center,  that has electrically-powered spin-up flywheels,  an airlock,  and some docking ports.  Any solar panels should attach here as well,  being near the spin center for low acceleration loads.  (We have already built similar modules for the space station.) 

The basic characteristics of the B330 module are given in Figure 2.  This is the item to be launched by Bigelow Aerospace about 2020 via the ULA Atlas-5 booster,  as a zero-gee space station.  It has an enormous interior volume for a very modest mass,  as compared to all the hard modules used to construct the international space station.  These take the form of a folded inflatable packed about a cylindrical hard core,  with docking hardware  at each end.  The inflatable has many layers,  considerable wall thickness (nearly half a meter),  and thus considerable meteoroid-impact protection.  Multiple modules docked together can make a really voluminous space station very quickly.  

Figure 2 – Characteristics of the B330 Inflatable Module,  Current as of May 2016

The Deep Space Habitat:  Size and Spin Rate

It is weightlessness that induces a myriad of microgravity diseases in the human body.  Not everything we need to know is yet known,  but we know enough to know that beyond just a little over a year’s exposure,   the effects become harder and harder to reverse,  if not actually impossible.  That’s not long enough to go anywhere beyond the moon.  Which says you do artificial gravity,  no if’s,  and’s,  or but’s. 

It is reasonable to suppose that reduced gravity might be somewhat therapeutic,  but we absolutely do not know how much is enough,  because those experiments have never been run directly.  So,  for a long mission in the 2-3 year range,  the only ethical decision at this time in history is to provide one full gee of artificial gravity.  The only method known to science is spin gravity:  using centrifugal force as a substitute for the force of gravity.   General relativity tells us they are fundamentally equivalent. 

How much gravity you get depends upon your radius away from the spin enter,  and upon the square of your spin rate.  Science tells us that 56 m radius at 4 rpm is one full gee at that radius.  Assuming your spin center (center of gravity) is in the middle of your design,  then your radius is about half the length of your design,  if you use a baton-shaped design spinning end-over-end.  This configuration is very stable in spin,  as all of us have seen in Friday night football game half-time shows.  You can ratio-and-proportion this from:

     Gees at radius = (your radius, m / 56 m)*(your spin rate, rpm / 4 rpm)²

Now,  untrained civilians easily tolerate 3-4 rpm spin rates without getting sick.  Well-trained crews with plenty of time to acclimate,  can tolerate something in the vicinity of 12 rpm long-term,  without getting sick.  Down around 7-8 rpm,  almost anybody should be trainable,  and acclimatize well,  fairly quickly. 

There is a gravity gradient effect on blood pressure gradients and fainting that is also proportional to spin rate,  but it should “kick in” nearer 16-20 rpm.  That’s no problem,  if we stay under 12 rpm anyway. 

I chose to consider either two or four somewhat-modified B330 modules with a center “hard” module,  for spinning space stations with artificial gravity.  These are illustrated in Figure 3 below,  and have 1 gee at the outer deck,  less inward,  to zero at the spin center.  Dimensions of the center module are not yet “hard”;  that one still needs to be designed.  It has electrically-powered spin-up flywheels,  an airlock,  and more than one docking port,  plus the best locations for attaching solar panels. 

Arranging the Interior

I would point out that the bed rest studies used on Earth as a surrogate for microgravity are as applicable as they are,  precisely because the body does not benefit from gravity while prone sleeping.  What that says is that you put your daily work shift stations at or near the one-gee level in a spin gravity design,  where being seated or standing provides the same beneficial effects as experienced at home. 

You may put sleeping quarters nearest the spin center at very -low-to-zero gee,  as there is no benefit from gravity while sleeping here,  there or at home. 

The off-duty and recreational functions can be put somewhere in between at the reduced gravity levels.  Thus the crew can benefit from whatever benefits there are from partial gravity,  during waking hours,  and under circumstances that actually could be fun. 

These same spinning space station design layout considerations also apply to the manned orbit-to-orbit transport version.  It is essentially a suitably-modified spin gravity space station design.

Figure 3 – Possible Variations on the B330 to Create a Space Station with Artificial Gravity

What we see with the three-module form is a space station of about 700 cubic meters internal volume and about 65 tons mass,  spinning at around 6-7 rpm,  tolerable with modest training.  This thing could be easily constructed by docking things sent to Earth orbit with current launch rockets.  The 700 cubic meters is almost as big as the entire international space station habitable volume,  as we know it today.  Unlike the current station,  this design allows us to directly explore the therapeutic value of partial gee levels.  And it would be far cheaper to build! 

The five-module form would be a little more expensive to build at 115 tons,  but would be far,  far larger than the current station in internal volume at about 1800 cubic meters.  Its lower 5 rpm spin rate is likely tolerable,  even by untrained civilians,  with only a few days’ acclimatization. 

What does artificial gravity buy you for a space station or an interplanetary spacecraft? 

It buys you freedom from the time limited exposure to microgravity,  with respect to microgravity diseases.  These include at least bone loss,  muscle weakness,  heart degradation,  bad vision effects,  and reduced immune system response. 

It buys you a much wider choice of food selection and cooking methods,  which now look just like they do down here.  It buys you food that will last a full 2-3 years,  or more,  in space without deterioration, by means of enabling frozen food and free-surface water-based cooking.  The current freeze-dried items last only a little more than a year!  Astronaut “food” is literally not sufficient for going to Mars!

It buys you “conventional”,  well-understood means of handling water and wastewater processing.  It buys you “conventional” easier-to-use toilets.  It buys you the ability to do laundry (currently impossible in weightlessness).  It buys you the possibility of conventional tub and shower bathing (also currently impossible in weightlessness). 

For missions in the 2-3 year range,  all these functions are essential to keeping a sane,  healthy crew!

We can equip and plan for a few days weightless as things get docked,  and as the vehicle makes maneuvering burns.  Between these short intervals,  it is imperative to spin-up for gravity to maintain health and fitness.  Electrically-powered flywheels would serve that function very well. 

What I chose for the Mars vehicle design was the lighter 3-module form,  to save exponentially on launched mass in Earth orbit,  which in turn seems to drive estimated program cost. 

The Other Crew Health Issues: Radiation

There are two different kinds of space radiation: galactic cosmic rays (GCR) and solar flare events (SFE).   The GCR is a very slow drizzle of extremely high-energy subatomic particles,  which are very difficult to shield.  This is because of “secondary shower” effects:  GCR particles hitting your shielding create a whole shower of other dangerous particles.  The thicker your shielding,  the worse this is,  in any practical geometries.  Near-Earth GCR intensity varies with the solar activity cycle:

    GCR = 42 REM annual + (18 REM annual) sin (time, yrs/ 11 years)

    where time is measured from the time of minimum exposure value

GCR in the inner solar system near Earth peaks at about 60 REM in years when the sun is quiet,  and minimizes near 24 REM in years when the sun is very active with storms,  sunspots,  and flares.  Among other limitations,  NASA limits astronaut radiation exposure to 50 REM max annual,  based on an expected 3% increase in cancer rates late in life.  Bear in mind that this is a sharp line drawn upon an inherently very fuzzy spectrum of effects. 

When you are orbiting lose to an astronomical body,  such as Mars or the Earth,  the planet blocks roughly half the sky at a thickness dimension too big to allow the “secondary shower” effect to occur,  thus cutting exposures roughly in half.   Otherwise,  even a thin spacecraft hull acts to reduce the exposure a little bit,  and also without the “secondary shower” effect,  precisely because it is thin. 

As close as 50 REM annual allowable is to 60 REM annual in a peak year,  I have to conclude that GCR is simply not the threat that some hype it up to be.  Astronauts wanting to go to Mars (and I wanted to be one myself,  45+ years ago) unanimously agree.

Solar flare events (SFE) vary widely in strength and in the actual location they threaten.  These are far lower-energy particulate radiation,  easy to shield,  but are also extremely intense (quantity).  The worst on record happened in 1972 between two Apollo landings on the moon.  Had a crew been out there to be struck by such a thing,  they would have died within hours,  very much like standing outside unprotected in the fallout from a surface nuclear bomb blast.  It takes but 15-20 cm of water to shield against the worst known case,  according to NASA’s own data.  It is the hydrogen in the water (or any other substance) that provides the shielding effect. 

If you have humans on board,  you have life support,  which means you have water and wastewater on board.  If you have spin gravity,  it is very likely that you have frozen food aboard.  All three of these are very effective shielding materials against SFE.  Arrange 15 to 20 cm of these about a designated sheltering space,  and you have a good shelter in which to ride out the SFE,  which is typically only a few hours long.  The smart designer would make this the vehicle’s flight control station (cockpit),  so that critical mission maneuvers could be flown,  regardless of the solar weather.

The Other Crew Health Issue: Living Space

Living space is the other crew health issue often not adequately addressed.  This is far more than just a volume allotment per person,  because real people require spaces in which to be together,  and other spaces in which to be alone.  But volume-per-person is a figure of merit we can compare.  You can ask anyone who has ever served prison time in solitary confinement how important this is.  He will verify that what we are doing now on the space station for 6-month-to-a-year tours is inadequate for multiple years’-long confinement. 

Putting 6 crew into 700 cubic meters is over 100 cubic meters per person in my mission plan.  That’s about the same as the old 1970’s Skylab station,  larger than anything since except the latest ISS.  The trick is allocating this into communal versus private spaces.  I have not addressed that in detail here. 

The 1960’s Gemini 7 mission was “up there” for two weeks “doing it in the suit” from a sewage standpoint,  and without even the ability to straighten the knees in such a cramped capsule!  They (NASA) do not want you to know that the plan was originally a 3 week mission!  This crew was brought home at 2 weeks,  because they were literally cracking-up mentally.  You really have to dig hard to find that fact;  it is something NASA really does not want you to know. 

Why?  Because if you did know,  then you would also know that NASA’s publicity claims that its new Orion capsule is an “interplanetary travel vehicle” are not true.  It is only adequate for short trips,  such as to,  or near,  the moon.  Longer trips simply require a very much larger space in which to live. 

That 1960’s Gemini 7 mission was astronauts Frank Borman and Jim Lovell,  who later flew around the moon on Apollo 8;  and Lovell was commander of Apollo 13.  As Gemini 7,  they rendezvoused in orbit with Gemini 6,  up there only a very few days.  That 2 weeks-in-space demonstration was “enough” for the moon in a cramped capsule,  knowing that sewage bags would be available for waste disposal. 

It is 8 or 9 months one-way to Mars.  See the problem?

The Mission:  Getting Men There Without Landing

What I assumed was the 3-module 65 ton spin-gravity space station as the deep space habitat for my manned orbit-to-orbit transport,  with a crew of 6.  This thing is illustrated in Figure 4.  It depends upon Earth return propellant previously sent separately to Mars orbit.  I assumed 65 tons for the basic station,  plus 50 tons of supplies to support the crew,  plus a ton for some engines for arrival at Mars. It leaves Earth orbit coupled to the Mars-arrival tanks laterally,  and to an Earth orbit departure stage on one end.  After the departure burn,  this departure stage is jettisoned,  and the habitat spun up. 

The water/wastewater tanks are arranged as radiation shielding about one inflatable module,  with the hydrazine/tetroxide engines fitted to the other.  Mars arrival propellant tanks are arranged about both inflatables as additional shielding.  I assumed that 7.5 tons of the 50 tons of supplies would be consumed and jettisoned during transit to Mars.

 Figure 4 – Orbit Transport and Other Assets for Men to Mars Orbit,  Without Addressing the Landing

I further assumed that about 10 tons of the supplies would be consumed and jettisoned during the year-long stay in Mars orbit.  This lightens the mass to be accelerated home out of Mars orbit,  thus lowering return propellant required.  This effect is exponential in its impact (mass ratio in the rocket equation).

I also assumed a further 7.5 tons of life support mass would be consumed and jettisoned during the voyage home,  thus further reducing the dead-head mass to be decelerated into Earth orbit at the conclusion of the mission.  This actually made a big difference to launched propellant tank tonnage,  so fully-closed-cycle life support may actually not be the thing we need to strive for,  at this time in history!   

Perhaps my assumptions are not good enough,  but the beneficial mass-reduction effects are quite real. To summarize my assumptions:  (1) 50 tons of all life support and consumables initially,  (2) half of this mass gets used up and wastes jettisoned during the entire mission,  (3) 40% of the jettisoned waste mass happens during the roughly-a-year in Mars orbit,  and (4) the other 60% is split equally between the two 8-9 month transits.  Half the 50 ton life support mass is still aboard upon arrival at Earth orbit. 

Off-Design Crew Return with Loss of Vehicle

The mass allotment for the orbit-to-orbit transport does include about 6 tons for the crew return capsule.  This would be used to return six people to Earth from Earth orbit,  in the nominal case.  But if there were to be a complete propulsion failure before Earth arrival,  the crew would have to “bail out” in this capsule for a free return entry situation.  What that means is we need a capsule heat shield capable of 50,000+ fps (17 km/s) reentry speeds,  and large enough for a crew of 6. 

The Orion and the Boeing CST-100 Starliner are not capable of 50,000 fps (17 km/s) reentry,  only 36,000 fps  (11 km/s) from the moon for Orion,  and 25,000 fps (8 km/s) from Earth orbit for the CST-100.  Both are capable of 7 crew.  Orion is really heavy,  the CST-100 not so much. 

Only the Spacex Dragon v2 is capable of 7 crew plus 50,000+ fps reentry speeds.  That is the one I selected to use in my model here.  It (and only it) makes a 6 ton allowance “reasonable”,  if we delete the service module (unnecessary for this application). 

Sending Men to Mars Excluding Landers

There are two items that must be sent to accomplish this:  the manned vehicle,  and its return propellant. To do only this implies no landings at all.  As far as I am concerned,  as difficult as it is to get men to Mars and back at this time in history,  not-to -land is equivalent to Columbus looking over the rail at the New World,  and sailing back to Spain without landing.  That would clearly have been the height of stupidity in Columbus’s day,  just as going to Mars without landing is today. 

Regardless,  you send two clusters to Mars orbit in order to send men there,  with the technologies that we have.  You send the manned habitat as quickly as you can,  and you send its return propellant “slowboat” with electric propulsion,  before you ever send the men.  This same approach would work with Venus or any of the asteroids, or even Mercury. 

For Mars and Mercury,  you need landers to complete the task.  Landers for those two destinations are necessarily very different.  The lander for each destination is thus a custom design. 

The results of my analysis are as follows.  The manned vessel at departure from Earth orbit is on the order of 50 m long,  and about 760 metric tons at departure ignition.  It cannot come home unless you send ahead its Earth return propellant.  That’s about 937 tons of storable-propellant tanks,  sent ahead with a solar electric stage for both the departure and arrival “burns”.   That propellant shipment vehicle is somewhere in the vicinity of 1245 metric tons at Earth orbit departure.  None of these are small,  but they are not “battlestar galactica”,  either.

Sending Landers for Mars

In order to make the journey worthwhile at all,  we have to land.  One-way / one-shot landers can do this at much higher payload fraction and lower ignition weight,  but with a very limited number of trips (one per lander).  If instead you design for a two-way,  one-stage trip,  you can actually fly the vehicle multiple times:  several trips per lander. 

Because of the multiple trip effect, in spite of its far lower payload fraction (near 3% max),  with the reusable lander approach you can raise the “bang for the buck” far above anything the one-shot throwaway approach offers.  This is indeed feasible today with a variety of propellant combinations.  I chose the same nitrogen tetroxide-monomethyl hydrazine as I chose for the manned vehicle,  precisely because it is easily storable at low inert mass fraction,  in space,  for years. 

Reusable lander design is another whole design study in itself.  I had previously done one to confirm that a chemically-powered single stage two-way lander was indeed feasible at all,  for four different propellant combinations.  Suffice it to say that my version uses retropropulsive landing after the entry hypersonics at Mars,  to avoid the high-ballistic coefficient difficulties with aero-deceleration by chutes.  The total delta-vee (two way) is on the order of 4.5 km/s.  That’s 3.6 for ascent,  and 0.9 for the retropulsive deceleration-to-touchdown. 

In that study,  my paper lander designs show enough delta-vee to land just over 3 metric tons of astronauts plus gear on Mars’s surface,  for a stay of up to 30 days.  The idea is to live in the lander itself,  but to travel around the landing site in a small,  drill rig-equipped,  rover.  To support longer traverses,  a inflatable 3-man ”pup tent” is assumed  to be carried on that rover car.  In addition,  I included a payload allotment for 3 in-situ resource utilization experiments,  plus three “other” experiments,  for each trip to the surface.  “Camping out” in the lander for only 30 days completely eliminates the need for a large surface habitat on Mars.  That’s a huge launch mass savings,  and a huge entry descent and landing problem avoided. 

My choice of NTO-MMH leads to the largest,  least-capable lander design,  in terms of total delta-vee.  It’s about 90.7 metric tons ignition mass,  fueled for a trip to the surface and back.  Anything we really do should be better than this,  certainly no worse.   If we leave everything on the surface except the astronauts,  their suits,  and their transit consumables,  there is “room” to bring a fair fraction of a ton of surface samples and experiment results back to low Mars orbit. 

These craft take the form of a slope-sided early space capsule with a projecting neck,  scaled way up.  It is about 10 m long,  about 8 m diameter at the heat shield,  and has a 5 m diameter core.  These are approximate,  but in the lander feasibility study,  all 3 "conventional" propellant combinations were just about these sizes and proportions.  Only the hydrogen-oxygen "bird" was different.  

The internal layout has a cylindrical core of a sealed engine room,  plus tall propellant tankage,  surrounded by a conical volume for cargo spaces.  There is a spherical-segment heat shield,  with holes through which the four slightly-canted engines fire.  Because the engine compartment is sealed,  there is no entry plasma flow into the compartment through the holes for the engine plumes. 

It has retractable landing legs,  whose footprint is roughly as wide as the craft is tall.  That provides stability,  even on rough ground.  Shell panels fold down to act as loading ramps.  The cargo spaces can be sealed and pressurized for living space,  once unloaded.  The crew cabin atop the neck provides great piloting visibility.  If also built as an abort capsule,  then the crew of a stricken lander can “bail out” during descent or ascent,  and reach the surface alive to await rescue. 

Having that lander bail-out capability only makes sense if there is also rescue capability available in the mission design.  That is primarily why I send 3 landers,  not just one.  If you have at least 2 operational landers,  then you can impose the wise safety rule that one lander be ready in orbit for a rescue trip while the other is on a mission to the surface.  You do not take the risk of a trip to the surface unless you have that rescue lander available.  But,  there is also always the risk that one “bird” might fail or otherwise become unusable.  In that case,  your surface missions end,  if you have only two.  So,  I send three.  That way the prime mission still proceeds,  even if one lander becomes unusable or is lost. 

You send down a lander with a crew of three,  while the other three do science from orbit,  watch over the surface crew,  and maintain peak fitness in the spin gravity environment.  When the surface crew returns,  they swap roles for the next trip to the surface.  See Figure 5. 

A rescue flight uses a crew of one only,  and does not stay down long.  You handle the weight of 4 people on ascent by offloading supplies and equipment to “lighten ship” for the rescue flight.

My plan here includes 3 landers,  sent fueled,  plus propellant to support two more trips for each lander.  That’s propellant for 9 total surface trips.  However,  only 8 trips can be made if you maintain rescue capability,  because you need that one fueled rescue lander in orbit at all times. 

Even so,  this mission plan and sizing provides up to 8 months on the surface of Mars,  in alternating crews of 3.  This can be at one site or it can be up to 8 sites (8 flights).  I would recommend visiting at least 7 different sites,  so that the “best” can be identified for installation of a base or permanent experiment station on the final sortie.  That’s a whole lot of “bang for the buck”,  more than I have seen in any of the one-site plans. 

These landers,  with their extra propellant tankage,  get sent as individual packages by means of electric propulsion,  as shown in Figure 6.  Of all the packages “flung” to Mars orbit,  these are individually the smallest at under 300 tons for 3 trips per lander.  The empty tanks,  landers,  and SEP stages all get left in Mars orbit at end of stay there.  With a little extra SEP propellant,  the SEP stages could be brought home to Earth orbit for reuse as well. 

There’s also room to add extra lander propellant,  of course.  For example,  using 5 trips per lander,  3 landers,  and the 1 rescue reserve,  that’s 14 trips available.  If spread over 13 months in orbit,  those average something like 27 days each.  It just costs more to launch and send it there.

Figure 5 – Use of Reusable Landers for Rescue Capability and Multi-Site Exploration Capability

 Figure 6 – Sending a Lander Ahead to Mars Orbit,  Along with Extra Propellant for Multiple Trips

Option to Establish a Permanent Automated Facility

There is an additional option here,  with respect to establishing a base or experiment station on that first manned mission to Mars.  The equipment to do this could be shipped inside one of the landers as an overload cargo item (because the space is available).  This is the third,  or back-up,  lander.  If the best site is determined in 7 visits,  and three landers are still flyable, then the ascent propellant can be off-loaded from this third lander,  with several tons of base equipment substituted inside it,  instead.  It can then be sent one-way to the surface at that best site,  using a crew of one,  who can ride with the three surface crew back to orbit in their lander. 

Some sort of homing guidance will be required to safely land two vehicles at one site,  though.  That will need to be part of the weight allowance for the equipment sent to the surface,  and may involve some small navigation aids in orbit as well.  Thus this option requires precision landing. Everything else about this plan is one vehicle at each one site.  The site need only be where the vehicle actually sets down.  

Another possibility is to send that lander down one-way robotically,  after all the crew has returned to orbit.  This would require equipment that operates entirely without human intervention,  and from within the lander’s cargo bay.  But the option inherently exists.  It requires precision landing guidance.

The real point is this:  there are ways to establish that permanent facility on Mars,  during the very first manned mission there,  and not “break the bank” doing it.  The enabling factors that create this opportunity are specifically (1) the on-orbit “basing” assumed in this mission plan,  and (2) the larger reusable lander vehicle design.  Other mission architectures without these specific features simply do not possess this possibility. 

Returning the Crew Safely to Earth

At end of mission on Mars,  with all crew safely aboard the orbit-to-orbit transport vehicle,  preparations begin for the trip home.  The vehicle is de-spun for a few days to dock up with the return propellant supplies. 

There are two groups of these:  (1) the Earth arrival tanks,  which are docked alongside the habitat modules to “fatten” the baton,  and (2) the Mars departure tanks,  docked to the end opposite the engines.  Docking implies propellant connections,  of course.  These connections likely will require astronauts on spacewalks to do them. 

Once ready,  the engines fire to leave Mars orbit.  Once the burn is complete,  those tanks are staged off,  and the vehicle spun-up for artificial gravity during the 8-9 month transit home.  A normal arrival at Earth requires a braking burn with the last of the propellant tanks.  This is depicted in Figure 7.  The crew returns from Earth orbit in the capsule,  at something like a 4 gee ride or thereabouts.  The vehicle itself can be refitted and used again on subsequent missions,  another huge “more bang for the buck” item. 

An abnormal arrival at Earth would occur if vehicle propulsion capability is lost.  In that event,  the crew must “bail out” in the Earth return capsule,  for a free-return entry at above escape speed.  That is likely a 12-15 gee event,  which is why maximum crew health from spin gravity is so essential.   A crew weakened by 2.5 years exposure to zero or low gravity would very likely not survive this much stress. 

That is also why the crew return capsule needs a heat shield capable of free-return entry,  even though such a return is not in the normal mission outline.  It is an emergency contingency only,  but ethics requires we plan for this properly.  

 Figure 7 – Returning a Healthy Crew to Earth

Providing Redundancy for Mission-Critical Propulsion Events

Loss of functionality of a rocket engine is a critical event,  since there are so many propulsive events required to make this trip.  Thus,  I use multiple engines for each propulsion installation.  The lander designs call for four engines in that compartment,  each firing through its own port in the heat shield.  The storable propellant engines on the transit vehicle should also likely number 3 or 4.  There should be at least 2 engines on the hydrogen-oxygen departure stage.  Thus every propulsion installation has multi-engine reliability.  Further,  there is larger lander thrust turndown capability with multiple engines. 

Rough-Estimate Costs

This is very rough,  “shoot-from-the-hip” stuff.  Take it with more than a grain of salt.  But the point is,  these estimates are very much smaller than those for the “90 day report” and “design reference mission” scenarios presented to congress. 

The reason is twofold:  (1) this is a very different mission plan than any of those other proposals,  and (2) with the exception of the lander design,  the hardware proposed here is not very developmental at all.  It is basically only re-shape,  re-size,  and test-to-prove it works,  of stuff we already build.  Only the landers do not fit that description.  They must be developed “from scratch” and tested extensively to trust that they will function.  That’s a lot more expensive. 

For hardware costs,  most items I simply estimated as $4 million per dry ton.  Launch costs assumed current launchers,  flying fully-loaded,  for a unit price to Earth orbit of $6 million per ton delivered.  The new Falcon-Heavy reduces that considerably,  and it should fly for the first time in 2016.  But I did not take advantage of that beneficial effect in my calculations.

For the landers,  I assumed a factor-100 higher cost of $400 million per dry ton,  to represent the extensive engineering development project that this will be.  Accordingly, development costs, not launch costs, dominate the figure I got for adding the landers to the mission.  The manned transit vehicle is dominated by launch costs,  precisely because it has little engineering development needed. 

For propellant costs,  I “shot from the hip” and used $800/ton (about $2.3 per gallon at 6.5 lb/gal) for the ordinary rocket propellants,  and 10 times that at $8000/ton for the electric propulsion propellants.  These may not be right,  but they have to be “in-the-ballpark” enough to be at least somewhat realistic. 

My results are given in Figure 8 below.  Building the manned orbit-to-orbit transport and sending it to Mars and back totaled something in the neighborhood of $15 billion,  dominated by launch costs for which propellant was trivial.  Adding the 3-lander fleet with propellant for 3 flights per lander was surprisingly modest at $26 billion,  dominated almost entirely by the lander development costs.  The total was still under $50 billion to send 6 people to the surface of Mars at up to 8 different sites,  and bring them and their transit vehicle home,  with spin gravity for assured crew health.  The transit vehicle gets used again for other missions. 

In comparison,  the 90 day report and design reference mission scenarios were quoted to congress a few years ago as $450 billion (probably over half a trillion today) to put 2 to 4 people on Mars at one single site.  The entire manned vehicle was thrown away doing this. 

There is no comparison here!  My old 1950’s way,  updated for today’s technologies and hardware,  is just far, far better. 

 Figure 8 – Resulting Very Rough-Estimate Costs

Challenge for Readers

Use my basic orbit-based mission design,  and see if you can add some nuances and thus squeeze even more bang for the buck out of this.  Some of you have better tools than the ones I used:  a simple hand calculator and ordinary spreadsheet software on a laptop. 

Challenge for Spacex

Get with Bigelow Aerospace and start working on the transit vehicle now.  This stuff can certainly be launched with your Falcon-Heavy,  and uses your Dragon v2.  Think men-on-Mars by 2025,  paced by the lander.  You can show the world how this is really properly done. 

Challenge for ???

“Somebody” needs to get started on the reusable lander,  as it is the pacing item preventing launching men to Mars by 2020.  The booster-landing experiences of both Spacex and Blue Origin are very applicable here.  The Dragon v2 propulsive landings will also be highly and directly relevant.  

References:

See also "Reusable Chemical Mars Landing Boats Are Feasible",  dated 8-31-2013,  located elsewhere on this same website:  paper rough-out designs for a one-stage lander,  done for 4 propellant combinations.

See also "Space Travel Radiation Risks",  dated 5-2-2012,  located elsewhere on this same website:  accumulated dosages for a "typical" manned Mars mission,  with either just a shelter space,  or a shelter space plus shielding around sleeping quarters.  

Update 6-14-16:  This one is based on the orbit-to-orbit transport carrying people both ways,  with separate landers staged out of orbit.  This make a lot of sense from a site selection flexibility standpoint,  and it does not depend at all upon our being able to manufacture suitable propellants on Mars.

Such propellant manufacture feasibility has been established;  the question how much / how fast can we make it?  The prime propellant for this activity is LOX-liquid methane.  That is not compatible with the long-term storable NTO-MMH that I selected for this study.  To use it,  the landers would need to be the LOX-methane design.

The cryogenics are also subject to boil-off losses over time.  The Dewar construction and insulation to counter boiloff will add significantly to inert weights.  It would be hard to ship them from Earth because of the 8-9 months of boil-off losses.  But if manufacture on Mars fails for some reason,  any reason at all, then you lose your ability to land,  after going to all the trouble to get there!  I see that as an unacceptably-undesirable outcome,  even if low probability.  You would have to ship enough LOX-methane from Earth to get a baseline mission done,   if only low or no mass quantities are available on the surface.

The counter-argument to that is sending the propellant manufacturing capability ahead as a robot facility.  Don't send men until you know you have enough on-hand to do what you want.  There is an advantage to having lander propellant made at Mars.  You can swap out un-needed descent propellant for extra cargo tonnage,  and fuel-up on the surface for the ascent.  That makes it far easier to land lots of equipment with any lander design,  even this reusable one. Plus,  you can fly suborbitally to explore other sites.

But you must have at least some propellant on orbit,  for the rescue lander.  That will simply have to come from Earth.  If you don't do that,  you do not have a rescue capability for a stranded lander crew.  Period.  Further,  it doesn't make a lot of sense to be shipping up tonnages of propellant from the surface,  because ascent payload fractions will be inherently far smaller.  So proposing to use Mars propellant as crew return propellant is less attractive than you might think at first glance.

All in all,  I still kind of like best the updated version of the 1950's mission design.