Up front comments:
This article is an earlier,
smaller effort, aimed at
identifying and characterizing the 3-phase process required to plant colonies
off-Earth. It examines the effects of
the process upon mission plans and the requirements upon the appropriate
vehicle designs. I plan to supersede it
with a longer article or articles, which
will include some vehicle rough-sizing results.
There is a corresponding slide show to this shorter
article, that could be given in a 30-45
minute window. It and myself are
available to speak on this topic at meetings,
preferably (but not exclusively) local to me here in central Texas.
--------
This article is about a reliable process for getting from
initial explorations on Mars, to
actually being able to reliably plant a permanent settlement there, without killing a lot of people. That process is defined by the experiences of
the cross-ocean voyages from Europe,
starting about 500 years ago, but
with due consideration for what they did wrong back then.
The
Lesson of History
Based on what Europeans did, establishing colonies in the New World and the
far Pacific, there are definitely 3
phases. They didn’t get it “right” much
of the time: the Roanoke colony in North
America disappeared entirely in rather short order. The Jamestown colony almost disappeared but
for knowledge obtained from the hostile local Indians. The Plymouth Rock colony would have failed, but for direct aid (plus useful knowledge
obtained) from friendly local Indians.
But when they did do it “right”, it worked rather well, such as in Indonesia, and with the later colonies in North America
after it had become widely known how to “live off the land” there. The proper process is illustrated in Figure
1, complete with the necessary
phases, and with the objectives, characteristics, and who usually does the funding, listed for each phase.
Figure 1 – The Lesson of History: 3 Phases Ending in a Settlement
Phases
Set the Missions
The same 3 phases apply to colonizing Mars (or anywhere
else, but Mars is the example
here). Different needs in the different
phases result in different missions being necessary during each of the 3
phases. Note that the Mars analog to
multiple sites explored in the first mission requires basing out of low Mars
orbit to visit multiple sites in the one mission to Mars! There is no way around that, precisely because there will be no long-range
surface transport on Mars during that first exploratory mission!
Other sites cannot be visited from a direct surface landing at one
site!
It’s
either visit multiple sites in the one mission,
or else mount a mission to each and every site of possible interest, or else bet lives on remote sensing results
(which you should never do)!
But done “right” by visiting multiple sites in the one mission, there will only be the one exploratory
mission! This is actually a good
outcome, considering the high costs of
mounting any sorts of missions to Mars.
See Figure 2.
Figure 2 – The Phases Set Different Mission During the
Process at Mars
Different
Mission Requirements and Vehicles
The different phases have different mission
requirements, and they in turn require
different vehicles. There may be significant
vehicle overlap between the first 2 phases,
but not very much at all with the third.
Note in Figure 3 that one required outcome of the experimental
base phase is hard-surfaced,
large-and-level landing pads, and
another is in-situ propellant manufacture at full scale. Those enable completely different vehicles to
serve more efficiently later in the phase.
Therefore, the mix of
vehicles used in the experimental base phase is going to change as that phase
proceeds.
Bear in mind that these mission approaches and vehicle
concepts are all “clean sheet of paper” designs! This is what could be done, if we could get away from a space program
micromanaged by Congress to only maximize the political return from pork-barrel
and corporate-welfare projects in powerful Senator’s districts. Privatization may help some with that, but it also brings other resource allocation
problems associated with an oligarchy of the rich and powerful.
Figure 3 – Different Vehicles Are Appropriate in the
Different Phases, at Mars
Typical
Transfer Velocity Requirements
These numbers reported in Figure 4 for the
interplanetary transfers are rough, but “well
inside the ballpark”, good enough to get
started. One should obtain better
estimates before actually sizing vehicles,
because of the exponential nature of the rocket equation. One should also use actual engine ballistics estimates, not handbook specific impulse values, to size appropriate specific impulses for use
in the rocket equation. The remaining
uncertainties will lie in the inert mass fractions for the weight statements of
the vehicles, and the resulting mass
ratios.
The Hohmann min-energy transfer is for “average planetary
distances from the sun”. There’s not
much effect of the Earth’s low eccentricity on this, but there is,
for Mars’s more-eccentric orbit.
However, these average values are
quite representative values for initial sizing purposes.
The same is true of the “fast trajectory” shown. This is an ellipse with an exactly-2-year-period, so that it could also serve as an abort
orbit. That way, Earth is there at perihelion, when the craft arrives at perihelion after a
single two-year circuit about the ellipse.
Slightly-different velocity requirements obtain, for more extremized planetary distances about
the sun. But that is a smaller
effect, so these are good “ballpark”
numbers for getting started.
Be aware that the near-field encounter velocities shown are
corrected from the 2-body solar orbit values, by the third-body gravitational attraction of
Mars (or Earth), as the distance closes
between Mars and the spacecraft, or
opens between Earth and spacecraft. The
far-field “encounter” velocities computed from simple 2-body equation models of
orbits about the sun are lower, but
unrealistic! Budgets for two course
corrections are also estimated in the figure.
One of these is to be done about mid-way, the other takes place as the craft approaches
Mars close-up.
Figure 4 – Rough Figures for Transfer Trajectory Velocity
Requirements
Typical
Local Mission Velocity Requirements at Mars
The numbers indicated in Figure 5 are fairly
reasonable, but that ignores thrust and
acceleration-level issues, which affect
engine inert weights, as well as the numbers
of engines vs thrust turndown ratios needed.
One must actually do the Mars entry ballistics and the final descent and
landing estimates, in order to firm up lander
vehicle thrust/weight requirements!
Entry, descent, and landing on Mars is both similar and
dissimilar to that same process on Earth.
The Mars atmosphere is thick enough to use entry aerobraking to “kill”
most of the close approach velocity, but
it is also so thin that the end-of-entry-hypersonics altitudes are very
much lower, and also much more scattered
with varying vehicle masses.
Almost regardless of size,
at Earth the end-of-hypersonics altitudes are above 40 km, and the atmosphere below that is thick enough
to enable the effective use of parachutes or wings to conduct landings without
any rocket braking. Mars is quite
different: even at smaller
sizes, vehicles come out of the entry
hypersonics at rather low altitudes, and
even lower still at higher vehicle mass and higher entry speeds. Impacting the surface still-hypersonic is a very
real risk!
Terminal velocities on parachutes at Mars are just barely
subsonic, so that terminal rocket
braking is absolutely required, even at
only 1-ton-or-smaller vehicle masses. At
higher masses, there is just not time to
deploy such a chute at all, before
surface impact, much less have it
decelerate you from high supersonic. Either
way, that Mars landing scenario requires
significant, even major, amounts of terminal rocket braking, in order to achieve a survivable touchdown at
all!
And while the velocity to “kill” is not all that large at only
0.7 km/s, you have a rough-field
obstacle problem to design for. You must
essentially hover and divert to avoid fatal obstacles or hazards on the
surface. That dominates over gravity and
drag loss effects, so that you need to
use a factor of somewhere between 1.5 and 2.0,
applied to the 0.7 km/s velocity-to-kill, for estimating the lander braking-rocket
velocity requirement, as near 1.0 to 1.5
km/s.
Beyond that, there is
also the wildly-varying thrust-to-local-weight deceleration requirement: near 4+ gees for braking-to-zero before
impact, versus only about 0.382 gees for
hover-and-divert. These are NOT easy
design requirements to satisfy, but
they must be satisfied, for all
lander designs at Mars! Rocket
engines, even today, do NOT have that kind of turndown ratio (near
11).
Figure 5 – Local Entry,
Descent, and Landing Velocity
Requirements at Mars
Rough/soft
field requirements drive exploration and experimental-base designs
The rough/soft field issues will drive vehicle designs in
both of the first two phases, because
hard, level, smooth landing pads do not yet exist! Some design criteria shown are shown in Figure
6.
There are fundamentally 3 problems to address: (1) static stability vs overturn on rough
ground, (2) sinking into the surface at
too high a dynamic or static bearing pressure upon soft ground, and (3) touching down at non-zero horizontal
speed, causing the leading-side landing
pads to “dig in” and “trip” the vehicle dynamically.
There is a rule of thumb used successfully for many decades
for landers on the moon, Mars, and elsewhere. There is a minimum lander pad footprint
dimension, as indicated in Figure 6. That dimension needs to exceed the height of
the vehicle center of gravity above the surface. This criterion simply rules out the safe
touchdown of tall, narrow vehicles on
rough ground! It is based on high
school physics: when the weight vector
points outside the landing pad footprint at its minimum dimension, the vehicle WILL topple over!
Sinking into the regolith happens when the landing pad
bearing pressure exceeds the ultimate failure pressure of the soil. Murphy’s Law says this will always occur
unevenly, leading to the craft being at
an angle, even on level ground. Too much, and it topples over! Even if it does not topple, pads buried in the regolith accumulate loads
of soil that must be removed before a takeoff can be attempted. One must design for landing pads large enough
to reduce the soil bearing pressure below that ultimate failure pressure! That is true dynamically at landing, and statically at takeoff.
99% of Mars’s surface corresponds to Earthly “soft, dry,
fine sand”, whether in dunes or
in plains with a loose rock content.
Such loose rocks cannot add strength until their spacing is essentially
zero, which is rare on Mars. The civil engineering handbooks have values
for the “safe” or “allowable” soil bearing pressures for a variety of
soils, up to and including “hard rock
ledge”. These allowable values are lower
than ultimate, to prevent soil settling
in the long-term foundation design problem.
The ratio of ultimate to allowable is usually about 2, sometimes 2.5.
As for the residual horizontal velocity problem, there is a mechanical energy criterion for
that. There is a radius from the center
of gravity to the pad or pads that dig in.
Dug in, the craft rotates about
that dig-in point, raising its center of
gravity. If the kinetic energy of the
horizontal velocity exceeds the potential energy change of the
center-of-gravity rise, then the vehicle
WILL topple over! This criterion also
pretty much eliminates landing tall,
narrow vehicles on rough ground.
Figure 6 – Rough/Soft Field Lander Design Requirements
There are 3 different vehicles required at Mars during this
phase, as listed in Figure 7. The direct 1-way cargo shots can be sent
prior to the manned mission. It is
presumed that a few of these need to arrive fairly quickly, although Hohmann min energy transfer should
be adequate for most. The manned
orbit-to-orbit transport will need to cross the Van Allen belts quickly both
outbound and on return for re-use. The
landers and their propellant supplies (plus propellants for the manned transport
return) can be sent ahead unmanned, and slowly, by electric propulsion. The space tug assist concept can be used to
reduce departure velocity requirements from Earth orbit.
Figure 7 – Recommended Vehicle Concepts for Exploration
Phase
Experimental
Base Phase Vehicles
Although they don’t have to be, the same mix of 3 vehicles can be used to
support much of the experimental base phase.
Note the additional requirement to have nothing jettisoned
before, during, or after Mars entry for the 1-way direct
cargo vehicles. This is to avoid
falling debris hazards to people and things already on the surface. All of this is listed in Figure 8.
The right time to apply the debris requirement is during the
exploration phase, so that no design
changes are needed when the phase changes to experimental base. Bear in mind that during this phase, the mission is still entirely supplied by
Earth, until and unless there is full success
in living off the land. The 1-way cargo
flight rate only decreases when success obtains in living off the land.
Again, the space tug
concept can be used to reduce departure velocity requirements from Earth.
Figure 8 – Recommended Vehicle Concepts for Experimental
Base Phase
Permanent
Settlement Phase Vehicles
This phase can only happen once all the “living off the
land” experiments succeed reliably in the experimental base phase, otherwise lots of people will die! That includes both in-situ sustainable life
support and in-situ propellant production, plus the construction of large, flat,
level, hard-surfaced landing
pads. The infrastructure for in-situ
production of large amounts of electricity is implied. See Figure 9.
The mix of vehicles is quite different: there can be both orbit-to-orbit and
direct-landing transports, and there need
be no further 1-way direct cargo flights,
alleviating that hazard to people and things on the ground at the
selected site. The “lighter” is a much
larger 2-way 1-stage surface-to-LMO-to-surface vehicle, with a larger payload fraction, based on the surface, and using higher-energy in-situ propellants
and the appropriate engines. It
functions to load and unload orbit-to-orbit transports, of both cargo and people.
And as with the other two phases, Earth departure velocity requirements can be
reduced by using the tug-assisted departure concept.
Figure 9 – Recommended Vehicle Concepts for Permanent
Settlement Phase
Conclusions
There is overlap among vehicle designs for phases 1 and
2, but not much with phase 3, as indicated in Figure 10. Rough/soft field landing is the driving
vehicle design requirement for both phase 1 and the first part of phase 2. Having such a rough field capability as an
abort capability would be wise even in later phase 2, and in phase 3. Each vehicle design is worthy of its own
vehicle design study. Such studies are
not included here!
The manned vehicle designs are the most demanding, because of the needs to provide not only life
support over months-to-years in space,
but also radiation protection,
and protection against microgravity diseases. Those are all worthy topics in and of
themselves, not covered here!
Figure 10 – Overall Conclusions
Final
Comments
Perhaps
the most important finding here is also quite divergent from most other mission
concepts for Mars! That is the need to visit multiple
sites in the one exploration mission,
driven by two things.
First,
the huge difficulty and expense of mounting any sort of mission to Mas
at this time in history. Second, the need to definitively-determine real ground
truth (including deep underground) at each candidate site, in order to reliably select the “best one”.
This
drives one to orbit-to-orbit manned transports with landers, instead of direct manned landings!
That is
true precisely because it is not just unwise to bet lives on possibly-wrong remote-sensing
results, it is actually immoral and
unethical to do so off Earth! Why? Because even today, there are still (more often than not) small
but significant disparities between remote sensing results and real ground
truth. Such is likely lethal, in a hostile lethal environment!
