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Table of content Page no.
1. Abstract …………………………………………………………………………… 02
2. Introduction ………………………………………………………………………. 02
3 .Overall structures and design …………………………………………………….. 03
4. Nozzle selection ………………………………………………………………….. 07
5. Trajectory and launch window for mission ………………………………………
08
6. Capsule design and module planning ……………………………………………. 14
7. Total mission planning …………………………………………………………… 15
8. Total mass analysis ……………………………………………………………….. 16
9. Re-entry speed and heat shield …………………………………………………… 19
10. Several subsystems ……………………………………………………………… 22
11. Radiation protection and safety system ………………………………………….
25
12. Detailed system safety design …………………………………………………… 28
13. Cost analysis. …………………………………………………………………… 33
14. System and mission engineering and landing ……………………………………
36
15. Choices for a Mars Orbit Base Location …………………………………………
38
16. Descent Technologies …………………………………………………………… 40
17. Re-entry and finally to earth …………………………………………………….. 45
18. Conclusion ………………………………………………………………………. 48
19. References ……………………………………………………………………….. 48
20. Team profile ……………………………………………………………………… 49
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Abstract:
This paper investigates means for achieving human expeditions to
Mars utilizing existing or
near-term technology. Mission plans are described here.Those are
accomplished with tandem
direct launches of payloads to Mars using the upper stages of
the heavy lift booster used to lift
the payloads to orbit. No on-orbit assembly of large
interplanetary spacecraft is required. In
situ-propellant production of hydrogen propellant on the Martian
surface is used to reduce return
propellant and surface consumable requirements, and thus total
mission mass and cost.
Chemical combustion powered ground vehicles are employed to
afford the surface mission
with the high degree of mobility required for an effective
exploration program. Data is
presented showing why medium-energy conjunction class
trajectories are optimal for piloted
missions, and mission analysis is given showing what
technologies are optimal for each of the
missions primary maneuvers. The optimal crew size and
composition for initial piloted
Mars missions is presented, along with a proposed surface
systems payload manifest. The
back-up plans and abort philosophy of the mission plans are
described. We have chosen radiation
magnetic protection system. The most significant idea of our
concept is that we have tried not to
show “reusable” the usual meaning. We will show even mars ice
can be used as hydrogen fuel
for mission cost effectiveness. It’s concluded that the plans
offer viable options for robust
piloted Mars missions employing near-term technology.
Introduction:
Getting to Mars with a Lunar Class Transportation System:
Figure 1: spacecraft interior(developed by catia)
Our dreamt requirements for this mission in a nutshell given
below-
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1. Propulsion:
Abort Motor, Attitude Control Motor, and High Burn Rate
Propellant for Solid Rocket Motors,
chemical hydrogen based propellant
2. Navigation:
Atmospheric Skip Entry, Autonomous Rendezvous and Docking, Fast
Acquisition GPS
Receiver, High Density Camera Sensors
3. Avionics:
Algorithmic AutoCAD Generation, ARINC-653 / DO-178 Standard
Operating System,
Baseband Processor, High Speed/High Density Memory Devices, and
Honeywell HX5000 North
star ASIC
4. Communication:
C3I - Standard Communications, Communication Network Router
Card, Digital Video Recorder
5. Power:
Column Grid Array Packaging (CGA), Direct Energy Power Transfer
System
6. Thermal Protection System:
Ablative Heat shield with Composite Carrier Structure
7. Life Support & Safety:
Backup and Survival Systems, Closed Loop Life Support,
Contingency Land Landing, Enhanced
Waste Management, Environmental Control, Hazard Detection,
Isolation and Recovery
8. Structures:
Composite Spacecraft Structures, Human Rated Spacecraft Primary
Structures Development
Overall structures and design:
Interdependency of Manned Mars Entry Vehicle
Types with Booster Diameter
• You cannot assemble a re-usable entry vehicle with an integral
aero-
Shell in Earth orbit (no factory equipment and no manpower), so
such
vehicles must be launched intact from the surface.
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• Two types of Mars Entry Vehicle concepts exist:
– 1. Wide base – Blunt body (capsule shaped)
– 2. Narrow Body – lifting body or cylindrically shaped
• Wide body (up to 15 meters wide at the base) landers are more
stable
And can carry more cargo since they need less fuel due to entry
drag,
but they need an HLV with a 10 meter or wider diameter.
• Narrow body landers carry less cargo but they can be launched
on
Some currently projected HLV boosters with a 7-8 meter
diameter.
• The booster’s launch cost must be affordable for dozens of
launches
Per year to support a continuing Mars exploration program.
• We can choose a vehicle design based on the booster available
OR we
Can pick a booster design to FIT the needs of the payload (the
lander).
Capsule shaped blunt body lander:
Figure 2: carbon –carbon tile configuration for 10 and 15 meter
diameter vehicle
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Types of Re-usable first Stages for HLVs ordered by increasing
development cost:
Cluster of Boosters which separate and are recovered
Individually from the water.
• Cluster of Boosters where each one separates and
Individually flies back to a landing strip.
• Single Large Rocket-powered airframe which flies back to a
Landing strip with jet engines.
• Single very Large Rocket-powered cone-shaped airframe
Which lands vertically on its own rockets.
• Single fly-back rocket powered vehicle which captures its
own LOX supply during flight & for the second stage
engine.
• Fully air-breathing (Hypersonic) Booster which flies
itself
back to a landing strip with scramjets.
• Highest development cost = lowest operating cost.
• Operating costs usually far exceed development costs.
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Cluster base power booster planning:
Figure 3: phase 3B super HLV (54-140) mt Leo
Expected near term HLV feature:
• Re-usable first stage or first stage segments (required).
• Air breathing engine to increase payload mass.
• Minimizing refurbishment to recovered stages, such as a stage
that
flies back and lands like an airplane.
• Flexible payload mass/size if a cluster.
• Very Wide payload capability to accommodate wide aero-shells,
re-entry shields and vehicles
(minimum 33 feet (10 meters) wide or more,up to 15 meters).
Wider payloads can be launched
with an inverted conical fairing, creating a “hammerhead”
payload configuration, up to 50%
wider diameter as the booster.7 meter (23 ft.) wide booster can
launch a 10 meter wide payload
8.4 meter booster (ET) can launch a 12.6 meter wide payload 10
meter wide (33 ft.) booster can
launch a 15 meter (49 ft.) wide payload
• Large payload shroud volume to hold large integral structures
with low
density like habs.
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• Ability to recover and re-use the second stage if
possible.
Nozzle selection:
Two-Step Nozzles. Several modifications of a bell-shaped nozzle
have evolved that allow
full or almost complete altitude compensation; that is, they
achieve maximum performance
at more than a single altitude. Figure 3-15 shows three concepts
for a two-step nozzle,
one that has an initial low area ratio AZ/At for operation at or
near the earth's surface
and a larger second area ratio that improves performance at high
altitudes. See Ref.3-5. The
extendible nozzle requires actuators, a power supply, mechanisms
for moving the
extension into position during flight, fastening and sealing
devices. It has successfully
flown in several solid rocket motor nozzles and in a few liquid
engine applications,
where it was deployed prior to ignition. Although only two steps
are shown, there have
been versions with three steps; one is shown in Fig. As yet it
has not made the change
in area ratio during rocket firing. The principal concerns are a
reliable rugged mechanism
to move the extension into position, the hot gas seal between
the nozzle sections, and the
extra weight involved. The droppable insert concept avoids the
moving mechanism and
gas seal but has a potential stagnation temperature problem at
the joint. It requires a reli-
able release mechanism, and the ejected insert creates flying
debris. To date it has little
actual test experience. The dual bell nozzle concept uses two
shortened bell nozzles
combined into one with a bump or inflection point between them,
as shown in Fig. 3-15.
During ascent it functions first at the lower area ratio, with
separation occur- ring at the
inflection point. As altitude increases and the gas expands
further, the flow attaches itself
downstream of this point, with the flow filling the full nozzle
exit section and operating
with the higher area ratio at higher performance. There is a
small performance penalty
for a compromised bell nozzle contour with a circular bump. To
date there has been little
experience with this concept.
Figure: two step dual bell nozzle
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A CAD model of dual bell nozzle given below:
Trajectory and launch window for mission
Planning suitable trajectory and launch window for a space
mission specially, when it is an
interplanetary mission is very important task. It is one of the
biggest challenges for a mission
design. For manned mission in mars this task must be done
perfectly to make the mission safe
and cost effective. While designing trajectory for manned
mission in mars some factors are
arisen are choosing suitable launch window, position of earth
and mars in their orbits, Δv budget,
time of flight, departure energy etc. Here some idea for
designing a suitable trajectory for
manned mission in mars in 2018 is given.
Why 2018
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The planets don’t follow circular orbits around the sun. They’re
actually travelling in ellipses.
Sometimes they’re at the closest point to the Sun (called
perihelion), and other times they’re at
the furthest point from the Sun (known as aphelion). To get the
closest point between Earth and
Mars, it is needed to imagine a situation where Earth and Mars
are located on the same side of
the Sun. Furthermore it is a situation where Earth is at
aphelion, at its most distant point from the
Sun, and Mars is at perihelion, the closest point to the Sun.
When Earth and Mars reach their
closest point, this is known as opposition. In the fig-1 the
position of earth and mars at
opposition is shown.
And theoretically at this point, Mars and Earth will be only
54.6 million kilometers from each
other. But here’s the thing, this is just theoretical, since the
two planets haven’t been this close to one
another in recorded history. The last known closest approach was
back in 2003, when Earth and Mars
were only 56 million kilometers apart [1]. And this was the
closest they’d been in 50,000 years.
Here’s a list of Mars Oppositions from 2007-2020[2]
Dec. 24, 2007 – 88.2 million km (54.8 million miles) Jan. 29,
2010 – 99.3 million km (61.7 million miles) Mar. 03, 2012 – 100.7
million km (62.6 million miles) Apr. 08, 2014 – 92.4 million
km (57.4 million miles) May. 22, 2016 – 75.3 million
km (46.8 million miles) Jul. 27. 2018 – 57.6 million
km (35.8 million miles) Oct. 13, 2020 – 62.1 million
km (38.6 million miles)
Figure 4: Position of earth and mars at
opposition
So 2018 is more suitable for manned mission in mars cause, mars
will be nearer to earth and the
date is Jul.27.2018.
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Trajectory plan
Two major types of human Mars missions have been proposed based
upon their
interplanetary trajectories and associated Mars stay times. They
are known as conjunction-class
(or long-stay) and opposition-class (or short-stay) missions.1
Conjunction-class missions are
characterized by long stay times on Mars (order of 400 to 600
days), short in-space durations
(approximately one year total for the Earth-Mars and Mars-Earth
legs), and relatively
small propulsive requirements. Opposition-class missions have
significantly shorter Mars stay
times but longer stay time in space. Now which will be the best
choice for this mission is the
question. Since this is manned mission safety issue for
astronauts is very important. Longer stay
in space will cause longer exposure of
spacecraft in direct radiation from sun or
other celestial bodies. So conjunction
class trajectories will be better option for
the mission. [3] Figure-2 shows a simple
representation of conjunction class
trajectory
Figure 5: Conjunction class trajectory.
There are two types of conjunction class trajectories. They are
type I and type II. If the
spacecraft travels less than a 180° true anomaly, the trajectory
is termed type I. If the spacecraft
travels more than 180° and less than 360°, then it is a type II
transfer. Type I trajectories are
preferable for manned mars mission.
Another important consideration for trajectory design is C3
(departure energy). C3 is not
constant. Variations in C3s can be due to many causes: the
relative positions of the planets, the
plane change required into the transfer orbit, the velocities of
the planets, and the eccentricities
of the orbits. However, this relies on the superposition of two
synodic variations. The first
synodic period occurs every 2.14 years, or 25.6 months, and
refers to the angular positions of the
two planets. The second is due to the eccentricity of Mars orbit
(e = 0.093). The planets nearly
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return to their original relative heliocentric position every 7–
8 oppositions, or every 15–17
years. [4]
Figure 6: Piloted optimal departure energies, 2009– 2024.
Figure-3 shows the piloted departure C3s for minimum initial
departure mass in LEO (Low Earth
Orbit) for 180-day outbound mission flights. [4]
Now finding the optimized trajectory can be done using software
like Mission Analysis
Environment (MANE) software tool, STK Astrogator by AGI,
trajectory planner, MATLAB
based programs etc. Here using trajectory browser website
developed by NASA some
trajectories are found for manned mission in mars. Figure-4
represents graphical representation
of trajectories for round-trip rendezvous mission in mars for
departure from earth in 2018.
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Figure 7: Trajectories for manned rendezvous mission.
In figure-4 the right column shows total Δv in km/s, horizontal
line represents earth departure
and left side vertical line shows duration of the mission (Earth
departure to earth reentry) in
years.
Figure 8: Details of the trajectories for manned rendezvous
mission in mars.
Figure-5 gives the detail information of possible trajectories
for the mission. In the figure
injection C3 is the departure energy for earth departure, Abs
DLA is declination of the launching
asymptote, injection Δv is the change in velocity required for
earth departure from 200 km LEO
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(Low Earth Orbit). Analyzing the information given in figure-5
two trajectories can be chosen.
These two trajectories are indicated by green rectangle (say
trajectory-1) and yellow rectangle
(say trajectory-2). Departure C3s of trajectory-1 and
trajectory-2 are same and smaller than
others and injection Δvs are also smaller. So less propulsive
thrust is necessary that is
economically important for the mission. Considering the post
injection Δv, trajectory-1 is better
than trajectory-2 and earth reentry velocity of trajectory-1 is
less than trajectory-2. But the
mission duration of trajectory-1 is more than trajectory-2.
Trajectry-1 allows the crew 48 days
more stay on mars but 32 days less stay in space. Stay more in
space causes more radiation
exposure in space and more stay in mars surface allows the crew
more exploration time. The
earth reentry velocity of trajectory-1 is less than that of
trajectory-2. So trajectory-1 can be
selected for the mission. The earth departure date for this
trajectory is May 10 2018 and arrival
date to mars is December 04 2018. The mars departure date is
August 01 2019 and earth reentry
date is September 20 2020.
The orbital view of trajectory-1 is shown in figure-6.
The above discussion shows not a final trajectory for the
mission but it gives an idea to find a
perfect trajectory for rendezvous mission to mars in 2018 with
an example. There are so many
trajectories for mission to mars in a finite interval of launch
window that can be found using
various trajectory finding software (some of them are mentioned
before) and various
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computational method. From these trajectories the final one can
be selected according to the
requirements of the mission such as Δv budget, reentry velocity,
mission duration, mars stay
time, departure energy, safety of the crew, total cost of the
mission etc. Keeping the cost of the
mission minimum is a challenge for the mission. An
interplanetary mission requires burning of
enormous amount of fuel in less than one hour to give necessary
velocity to the spacecraft. So Δv
should be minimum as much as it possible. Departure energy C3
should also be kept minimum.
For this mission window is an important consideration. It is
also necessary to keep in mind that
the reentry velocity to earth atmosphere or entry velocity to
the mars atmosphere should not be
so high that causes overheating than the limit of the
spacecraft. Total mission duration should
also be considered. Longer mission duration more effort to
prevent radiation is necessary that
causes more cost. All these consideration must be focused to
plan the final trajectory.
Capsule design and module planning:
Figure 10: module planning
In capsule we will use the usual planning but we are choosing
the Orion planning. We are
planning for some lunar module. That’s picture is given
below:
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Figure 11: looner module
Total mission planning:
Figure 12: mars concept of operation
The Concept of Operations (ConOps) resulting from the hybrid
architecture solution. The IM
Vehicle Stack would be lifted into LEO by the SLS rocket with
the DUUS, the only launch
vehicle with enough throw mass for the job. Then the crew would
rocket into orbit using the
services of a Commercial Crew provider. Once on orbit the
Commercial Crew vehicle would
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dock with the IM Vehicle Stack, transfer the crew and depart.
Afterwards, the DUUS would
perform the trans-Mars injection burn to send the spacecraft and
crew on their way to
Mars.
The Martian flyby alters the trajectory of the IM Vehicle Stack
sending it on a path to intercept
Earth. Just prior to entry interface the crew would transfer
over to the ERP and separate from the
Habitat Module. The ERP would then carry out the Entry, Descent
and Landing portion of the
mission ending with a splashdown in the Pacific Ocean.
Total mass analysis:
In calculating the throw mass: Using the throw mass numbers we
generated a curve of C3 as a
function of throw mass. We first estimate the excess ∆V from a
200 km parking orbit to each of
the trajectories described in TABLE below:
Figure: ∆V EXCESS FROM 200 KM LEO
Then we can feet a curve on this data. That is given below.
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Figure: LEO ∆V Excess vs. Payload Mass
For a C3 of 38.8 km2/sec2, the ∆V excess is 4.86 km/sec.From
these curves, we could further
generate a payload mass vs. C3 curve for the Falcon Heavy, shown
in Fig. This curve was
generated by adding the excess velocity from table to the state
of a 250 km LEO in
STK/Astrogator, and then calculating the resultant C3 for each
point. From this curve you can
see that the estimated mass to a C3 of 38.8 km2/sec2is 9,800
kg.We rounded this to 10,000 kg
for the purposes of this study.
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Figure: Falcon Heavy Payload Mass vs. C3
For TMI finite burn modeling:
We use estimated parameters for the Falcon 9 second stage based
on a Falcon Data Sheet [13]
(This is the same stage used for the Falcon Heavy.) This stage,
the V1.1 Stage 2, will use the
Merlin 1D engine, which is estimated to have a vacuum thrust of
400,000 N, a vacuum ISP of
340 seconds, and a dry mass of 4.7 tonnes. To estimate the fuel
usage for stage 2above the 200
km circular LEO baseline, we made an estimate of what fuel would
be required for the
appropriate ∆V. using our STK/Astrogator-based numerical
integration, we were able to produce
the data in TABLE VI. For the finite maneuver, we then
calculated the burn duration to be 420
seconds. We estimated 15 tonnes of payload (including about
4.7tonnes for stage 2 dry mass)
and derived the fuel load for a finite TMI maneuver to require
roughly 50 tonnes of fuel beyond
LEO. This compares to the reference data[13] of a73.4-tonne
total fuel load for the stage. Note
that while the SpaceX references claim that the Falcon Heavy can
lift 53 tonnes to LEO, this
refers to 53 tonnes of payload mass to a 200 km circular LEO. In
our case of targeting a
Mars Escape C3 of 38.8 km2/sec2, we are only assuming a payload
mass of 10 tonnes. This
leaves excess propellant in stage 2 in LEO for our case. TABLE
VI shows an estimated 50
tonnes excess mass in LEO that we consider as excess propellant
for the TMI maneuver. (We
assume that there is zero propellant excess when inserting 53
tonnes into LEO.)
Another table is given below:
Launch vehicle and payload value:
To determine if a launch vehicle can throw a space capsule to an
Earth departure C3 of 38.835
km2/sec2we need a value for the payload mass. In this study we
use a value of 10,000 kg, which
is based on an estimated SpaceX Dragon that has a dry mass of
4,200 kg and a cargo mass of
6,000 kg [7].For a representative launch vehicle we chose to use
the Falcon Heavy. According to
a press release by SpaceX [8] in April, 2011, the first Falcon
Heavy launch should be in 2013
or2014. According to version 7 of the SpaceX brochure, the
Falcon Heavy will be able to
deliver (from Cape Canaveral)29,610 kg to a 28.5 degree
inclination 185 km circular Low Earth
Orbit (LEO), and 15,010 kg to a 185 x 35,788 km .Geostationary
transfer orbit
(GTO)[7]. Version 12 of this document[9] states the Falcon Heavy
will be able to
deliver53,000 kg to a 28.5 degree inclination orbit (no altitude
mentioned), and 19,000 kg to
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a 28.5 degree inclined GTO orbit(no specific orbit dimensions
are given.) Another reference,
however, on the SpaceX website itself, says that the Falcon
Heavy will be able to deliver
53,000 kg to a 28.5 degree inclination LEO, and 12,000 kg to a
27 degree
inclinationGTO[10].
Re-entry speed and heat shield:
By targeting the altitude of perigee to be in the range from
56.5 km to 61.5 km. The numbers in
the figure refer to the perigee altitude that caused that
trajectory. A perigee of 62km or greater
escapes, and 56 km or less performs a direct reentry. We used
the 1976 US Standard
Atmospheric Model(developed jointly by NOAA, NASA, and the USAF)
and a five-tonne mass
with a heat shield area about 10.2 m. We modeled simple
atmospheric drag without lift, with
a constant Coefficient of Drag (Cd) of 1.3 in the numerical
integrator. (This value is similar to
Apollo.) Fig.7 shows that the velocity (dashed line, with
right-hand axis) during reentry peaks at
about 14.2 km/sec velocity. This could possibly be reduced by
changing the launch trajectory or
Mars flyby conditions, which we will look at in future
studies.
Figure: Aero brake and Reentry Trajectories for Various Perigee
altitude
Again,
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Figure: Altitude and Velocity during Direct Re-entry
The purpose of the aero capture before the reentry is to reduce
the g-loads on the capsule and the
crew. Studies looking at reentry trajectories from Mars have
attempted to define acceptable
effects on the crew, and in particular, the g force astronauts
can withstand after long duration
spaceflight. [14] Next figure shows the g force due to
atmospheric drag during aero capture for
the range of perigee altitude from 56.5 km to62 km. The lowest
perigee has the highest g force,
and the peak load ranges from just under 6 g’s to just over 9
g’s. Of course, an aero capture
increases the length of the mission, up to an additional 10 days
or so if the apogee reaches to
lunar orbit. Because the service module of the capsule would be
released before aero braking, the
power system of the capsule would likely rely on batteries. The
post-aero braking orbit must be
optimized with these considerations, as well as with the reentry
conditions.
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Figure: G-force during aero braking
After jettisoning structures that must be released before
reentry, the spacecraft is estimated to be
5,000 kg. The reentry of a 5,000 kg, 3.6 m diameter spacecraft
into Earth’s atmosphere
present some challenges from an aerodynamics,
aerothermodynamics, and thermal protection
system (TPS)perspective. The mission calls for both an aero
capture maneuver and a
reentry at Earth. To date, no aero capture maneuver of this type
has been attempted either at
Earth or other planetary destinations. The atmospheric entry
speed forth aero capture is estimated
at 14.2 km/sec, which would make it the fastest reentry of any
manned vehicle by far. The
fastest, successful reentry of a man-made, but unmanned, vehicle
to date was the sample return
capsule for the NASA Stardust mission, which reentered at 12.6
km/s with an estimated total
stagnation heat flux of 1,200 W/cm2 [14, 15, and 16]. Assuming
that a baseline vehicle
architecture for reentry Isa Dragon capsule, several TPS related
issues that require more study.
First, we need to predict with higher fidelity the reentry
trajectory and aero thermal loads. The
details of the reentry will have a significant impact on the
thermal loads. The total heat load
during Earth reentry will determine how thick the TPS material
will need to be to protect the
underlying structure. The peak heat flux will drive the
selection of the actual TPS material.
The Dragon capsule heat shield uses a variant of Phenolic
Impregnated Carbon Ablator (PICA),
which has been demonstrated in actual flight environments up to
1,200 W/cm2[17]. Some
very preliminary aero thermal predictions of the Mars return
aero capture maneuver show peak
stagnation heating of over 3,500 W/c, dominated by shock layer
radiation heating. To date there
has not been any ground testing of PICA at those heating heat
flux levels. However, there is no
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data indicating that PICA would not perform as predicted under
those conditions, using
existing analytical models. Ground testing, such as arc-jet
testing, would need to be conducted to
verify PICA's performance. Second, currently operational manned
spacecraft have been designed
to operate in LEO conditions for missions that are nominally up
to six months to a year in
duration. The micrometeoroid and orbital debris (MMOD) and
thermal environments for these
missions are well characterized and systems are in place to
mitigate or control their effects. Some
investigation of the space environments during the 500-dayMars
mission would be required in
order to predict the impacts that these environments may have on
the TPS and determine any
necessary design changes. Third, TPS performance during reentry
is influenced by two primary
factors: peak heat flux and total heat load, as well as other
factors. Trade studies will need to be
performed to determine if using an aero capture maneuver to
spread out the heat load is an
optimal approach. The longer soak times associated with the aero
capture may lead to
even more stringent TPS requirements than a direct reentry.
Using a direct skipping entry
trajectory with banking maneuvers to bleed off energy, may work
just as well as the aero
capture/reentry combination. This study would include trajectory
optimization, predictions of
aero thermal environments, and TPS sizing calculations. And
lastly, the PICA TPS material has
not been used or tested at the high heat fluxes that can be
expected on the Earth reentry from
Mars. As mentioned earlier, ground based arc-jet testing would
need to be performed to
verify that PICA performs acceptably at those conditions. This
test data would also be used to
validate the material response models used its sizing analyses.
TPS thickness is the primary
adjustable design parameter available to satisfy TPS
requirements. Higher heating conditions
usually lead to thicker TPS. Due to its manufacturing processes,
PICA can be fabricated in
thicknesses up to about ten inches. If sizing predictions
indicate that the required PICA thickness
is more than that, another TPS option may be needed. In
addition, the thicker the material gets,
the more difficult it is to accommodate mechanical loads. PICA
is not a mechanically strong
material, and induced strains in the material due to flexure of
the underlying structure or
differential thermal expansion, may lead to mechanical failure
(fracture) of the material.
In some cases these mechanical loads on the TPS material can be
minimized by using a strain
isolation pad or by increasing the stiffness of the heat shield
structure.
Several subsystems:
ECLSS is made up of several subsystems: air, water, food,
thermal, waste and human
accommodations (often referred toes crew systems). Human
accommodations takes into
consideration the crew’s personal, exercise and medical
provisions, radiation protection, how
much free volume should be required for each person, and
pressure suit needs. Figure below
shows primary ECLSS subsystems and their interfaces.
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“Barsoom Inception: A mission for humanity”
23
Figure: integrated class interface
The cabin atmosphere assumes 101kPA (14.7psi) with adnominal
makeup of 78% Nitrogen
(N2), 21% Oxygen (O2), and1% other. Subsystem components and
stored gasses are
included for makeup gasses for leakage and gas equivalent
storage for one cabin
repressurization, though no pressure suits). We reviewed
numerous technology options to
revitalize the cabin air, and determined that an electrolyzer
and Sabatierwork in tandem to
provide necessary air revitalization needs. The electrolyzer
produces O2 and H2 from water. The
SabatierprocessesCO2, but needs H2 to do so, which it receives
from the electrolyzer. The
amount of water needed to fully replenish all O2 requirements
was included in the water
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24
subsystem calculations, and the difference in H2 production and
H2requirements are
balanced to determine the stored H2 required for this subsystem.
The water subsystem assumes
basic metabolic requirements for drinking, hygiene, and food
processing. Water is added to
accommodate O2 production. Various configurations include no
recycling, recycling
condensation from the atmosphere, and recycling water from the
biological waste. Hygiene water
is a variable that can be used to reduce mass for this
subsystem, and water recycling is
assumed for reduced launch mass, discussed further in section 8,
discussing ECLSS sizing.
The SOA technology for condensate recovery utilizes
multifiltration, ion exchange, and catalytic
oxidation. Food estimates include packaging and storage factors,
as well as a galley and food
processing equipment such as heating, preparing, and disposing
of waste. While it is probable
that all food is consumed, factors to account for food adhesion
to the packages are included. The
thermal system is based on a redundant loop radiator, and is
sized for the ECLSS power as
estimated for the different configurations, plus an assumed 2kW
of heat coming from air-cooled
components such as avionics. The waste system provides the
interface for collecting and recycle
water from the liquid waste are included in the water system as
noted above. Personal provisions
are limited to items such as clothing and hygiene products. Free
volume estimations are based on
minimum NASA requirements, and are only used to calculate makeup
gasses that are required
for the full duration. No consideration of privacy, or separate
sleeping quarters was contemplated
for this study. The mass and volume associated with free volume
need not included in the
ECLSS mass and volume, as we assume that they will be carried at
the overall vehicle
level, as is typical. Long term exposure to microgravity has
demonstrated that deconditioning
must be counteracted with routine vigorous exercise so we assume
the use of resistive
exercise equipment that provides full range of motion exercise
of all primary muscle
groups. Medical equipment and supplies are provided to address
emergencies. In addition to
spacecraft materials of construction, radiation protection is
provided by a water shield made up
of water in storage for other subsystems. Other options reviewed
were Hydrogen-Impregnated
Carbon Nan-fibers, and a Liquid Hydrogen Shell. Further study
needs to be done to find creative
solutions for radiation protection, including the amount of
radiation and the level of risk of a
high radiation event deemed to be acceptable. Evaluating these
risk factors will include an
exploration of crew age, gender, and various exposure types; and
will affect future ECLSS trade
studies.
Class sizing:
The crew’s metabolic rates drive parameters such as O2
consumption and CO2 production.
TABLE below shows interface values for each subsystem based on
an average crewmember
(CM) of 70 kg and metabolic rate of 11.82 MJ/d (17).As
demonstrated, the most significant
consumable is water at kg/CM-d of dry food. We are using a Water
Lean approach, with a
reduced water requirement of 1.2 kg/CM-d as shown in TABLE:
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“Barsoom Inception: A mission for humanity”
25
We can also share a master equipment list. Given below:
Radiation protection and safety system:
HT mission design needs to be carried out with the aim of
reducing radiation exposures in line
with ALARA. The application of ALARA requires that the space
radiation environment is well
known – outside and inside a spacecraft. This requires knowledge
of the external radiation
environment with its various components of the
spacecraft/habitat construction, and of the
results of transport calculations modelling the internal
radiation environment. The design
of the spacecraft/habitat requires the use of radiation
transport codes to compute dose
equivalents. As described in Chapter 5 the computer codes may be
one- or three
dimensional, deterministic or based on Monte Carlo (MC) methods.
The construction of a
spacecraft/habitat should include areas where the dose rates are
lower than elsewhere in the
spacecraft. There should be area monitors with visual displays
of dose rates.
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“Barsoom Inception: A mission for humanity”
26
To reduce uncertainties further improvements are needed in the
models of the galactic cosmic
radiation, the solar energetic particles, and the trapped
radiation to allow the accurate forecasting
of the fully integrated model of the radiation environment
incident on the spacecraft/ habitat.
Models have been developed for each of the radiation components.
These models suffer several
shortcomings: (i) the GCR models inadequately characterize the
solar cycle dependency and the
scaling with heliocentric distance; (ii) the SPE models have an
incomplete understanding of
the acceleration mechanism of the transport through the
Heliosphere and a lack of prediction
capability; (iii) the radiation belt models no longer reflect
the current state of the Earth’s
magnetosphere and lack the ability to properly describe the
dynamic behavior of the trapped
particles. The forecast models require an improved understanding
of the physical processes on
the Sun; the transport and acceleration of the solar wind
through the heliosphere; the
processes in the magnetosphere (wave-particle interactions,
source and loss processes, and
acceleration mechanisms). The space environment is highly
variable on very different time
scales as a result of the variability of the Sun. In general all
aspects of the space environment are
affected, but SPEs and CMEs are the most dramatic radiation
events and may constitute for
several missions a serious hazard. All the radiation components
(including GCR and
trapped) are also modulated by SPEs(For bush decreases in the
GCR fluences, for example).
An accurate prediction of SPEs and CMEs would allow for a more
effective approach in
the shielding strategy. Forecasting through real time
observation and propagation modelling
should be improved. Astronauts are particularly vulnerable
during EVAs, when they should be
monitored with active dosimeters. Real-time space weather
predictions and remote satellite and
areas Instrumentation will assist in EVA activity. The real-time
measurements will provide
guidance, and can suggest changes in mission scheduling to
maintain the total risk below
predefined limits. The development of shielding requirements and
strategies is important
for the achievement of ALARA. The reduction in exposure can be
made by reducing the
exposure time or by passive shielding. Passive shielding may
cause an increased risk by
increasing the dose equivalent from any generated secondary
particles, projectile and target
fragments (including neutrons). For shielding effectiveness, the
use of a shielding material with a
low mean atomic mass is generally better. Information about
radiation transport codes is
important and the strengths and weaknesses of the codes should
be investigated in detail
via benchmarking procedures against experimental data, including
data obtained with
advanced anthropomorphic phantoms exposed at accelerators. The
physics at the basis of the
particle transport and cross sectional data tables must also be
improved to further develop the
codes.
Area monitoring:
Area monitors at well selected locations in the spacecraft can
determine the environmental
conditions and are appropriate for an immediate warning about
changing exposure
conditions. Instruments are required to determine the radiation
environment in terms of particle
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“Barsoom Inception: A mission for humanity”
27
type, fluency rate, energy, and direction distributions and, in
some instances, dose quantities.
Dose quantities used to assess doses to astronauts and to
monitor radiation at a number of
locations should give values of the dose rate. These data can be
used to implement ALARA.
Area monitors at well selected locations in the spacecraft can
be appropriate for immediate
warning about changing exposure conditions. This can be of
importance before or during SPEs,
electron belt enhancements, and EVA. Real-time calibration of
instruments should be explored.
If appropriately designed and accurately calibrated instruments
are used, it may be that a
quantity measured in fixed position in a spacecraft can, along
with appropriate occupancy
data, provide the basis for an adequate assessment of doses to
an astronaut or of doses to the
local skin or the extremities. While in principle this procedure
may be applicable to astronauts in
space, the large variation of the radiation field in intensity
and composition of radiation types
inside a spacecraft, and its variation with time together with
the flexibility of the astronaut´s
position, has the consequence that area monitoring is not
sufficient to completely
substitute individual monitoring, especially considering the
high individual doses to astronauts
and the interest in providing a basis for individual risk
estimates.
Individual monitoring
The assessment of organ and tissue absorbed doses, together with
radiation quality factors, of
individual astronauts can be accomplished by calculations using
anthropomorphic phantoms or
by measurements using personal dosimeters.
One method of calculation of organ and tissue absorbed doses and
radiation quality factors does
so directly for a standard male or female phantom for various
locations in a spacecraft
with appropriate shielding. The phantoms can be adjusted to
approximate a particular
astronaut. The results are normalized using readings of area
monitors and personal dosimeters.
Another method requires knowledge of particle fluence and
applies conversion coefficients
from particle type, energy and direction distribution of fluence
to organ and tissue absorbed
doses and corresponding radiation quality factors for uniform
irradiation of an astronaut.
Individual monitoring is mostly performed using personal
dosimeters worn at the surface
of the body. The personal dosimeter serves as the dosimeter of
record. A single dosimeter
system is, however, not sufficient to provide an assessment of
the absorbed dose at the surface of
the body weighted by radiation quality. The broad range of
different types of particle requires a
minimum two detectors, one sensitive to low-LET radiation and
the other to high-LET radiation.
Because of a possible anisotropy of the exposure in the
spacecraft due to variations of shielding
properties, it may be useful to wear more than one dosimeter.
Also care needs to be taken
regarding low-energy electrons and particles which are stopped
in the skin and, therefore,
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“Barsoom Inception: A mission for humanity”
28
contribute only marginally to organ doses other than the skin
dose, but may induce a large signal
in an external dosimeter.
The use of adequate active personal radiation detectors would
enable improved 4036
characterization (input energy, nuclear abundance, fluence rate,
direction) of the radiation
field both on the body of the astronaut as well as in the
environment. The measurement of
dose-rate can contribute directly to ALARA.
The results of bio-marker measurements can be additionally used
to estimate individual
radiation exposure. The determinations can be collaborative and
provide all the
experimental radiation information and relative codes needed to
achieve an efficient risk
assessment, minimizing the uncertainties in the final risk
estimates.
Dose recording
Astronauts in space are exceptionally exposed and the assessment
of their individual doses
should be part of the radiation protection programmer for space
flights. Astronauts should
be informed of their doses and risk assessments as soon as
possible. Their doses should
be regularly registered and a long term registry for all
missions should be maintained. The dose
record is the formal statement of the crew member’s exposure and
should be kept as a
confidential medical record. The record should contain the
history of the exposure and all
the calculation and experimental results, including all
information on the particle type
energy and direction distributions of fluence; computer codes;
conversion coefficients and
weighting factors; area monitor, personal dosimeter, and
biomarker results.
Detailed system safety design:
The system safety activities needed to support detailed system
design are presented in Figure
This phase is different from the concept development and early
design phase because the length
of time required to complete the detailed design makes it likely
that new risk issues
requiring resolution will surface as design details are being
developed. These new risks often
involve the necessity of having to reallocate subsystem masses
while ensuring that
allocated safety requirements, codes, and standards continue to
be satisfied. The lower level
safety objectives for this phase consist of a combination of
objectives listed in Figure of S under
the phrases “Design the system to ...” and “Build the system to
...” This results in an
expansion lower-level objectives under the ASARP objective to
five, as shown in Figure
The broadening of the early design safety objectives reflects
the fact that during detailed
design, new information, both positive and negative in
character, is being generated.
Sometimes this information takes the form of new risks and
challenges. Other times the new
information takes the form of new opportunities to improve
safety.
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“Barsoom Inception: A mission for humanity”
29
Figure: Principal System Safety Activities and Related Processes
during Detailed System
Design, and their Interfaces with the Safety Objectives
Radiation protection (passive protection):
The first protection is constituted by the mechanical structure
of the spacecraft. This, in general,
will be very elective for cosmic rays with no more than about 30
MeV/nucleon of E, For
example, the average thickness on the MIR space station is1
g/cm2of equivalent aluminum,
reaching 7 g/cm2in particular places of the spacecraft. For a
more generalized example, let us
consider a spacecraft whose external dimensions match the
Shuttle capability, i.e. with a radius
of 2 m and several meters in length. Its external vessel should
have a mass of1020 kg for each 3
mm of equivalent aluminum and each 10 m in length. This
thickness of aluminum can stop, by
ionization losses, all protons with less than 25 MeV of E,
Adding a further absorbing material
around a more restricted volume, where the astronauts are
supposed to spend most of their time,
would result in the addition of a significant mass to the
spacecraft. The protection of a section of
our hypothetical cylindrical spacecraft, 1.5 m in radius and 3 m
long, requires 115kgfor each mm
of aluminum, i.e. 3450 kg for stopping protons up to 100 MeV. We
could do better by
concentrating the absorber around a smaller volume, where the
astronauts could shelter in case of
radiation alarm'. The mass of aluminum around a cylindrical
volume 0.5 m in radius and 2 m
long would be 12 kg for each mm of thickness, i.e. 360 kg for
stopping protons up to 100MeV,
which rises up to 1160 kg for stopping protons up to 200
MeV.
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30
A magnetic lens for shielding the spacecraft from the solar part
of the cosmic ray
radiation:
The good directionality of the solar cosmic rays allows one to
conceive a cylindrical magnetic
lens for defocusing all the positively charged258 P. Spillantini
et al. /Nuclear Instruments and
Methods in Physics Research A 443 (2000) 254}2631The industrial
production of NbTi
superconducting wires is already at a critical current density
of J#"5000A/mm2 atB"2.8 T and
4.2 K, so we consider to work at 70% of the critical current.
Particles away from the axis of the
lens. The need to minimize the mass limits the choice for the
magnetic field production: it must
be realized by superconducting coils, and without the use of
iron yokes. This implies a
refrigeration system to keep the superconductor at the low
temperature necessary for its
operation. For what concerns the techniques that can be
envisaged for the coil, in ten years or
more from now, it is probable that the NbTi material presently
used for the industrially produced
superconducting cables could be substituted by theNbSn material
that is already used in magnets
for high field spectroscopy and, since a few years, is also used
by industry for construction of
large coils (for fusion research and specialized projects).The
use of NbSn material will certainly
reduce the mass of superconductor material needed forth coils
and increase both the field
intensity and the operating temperature while also increasing
testability limits. Actually, beyond
NbTi and NbSn, one can envisage the use of ceramic high-Tc
superconductors (HTS), like Ag-
stabilizedBi2223tapes;in fact, within 2 or 3 years the
electiveness of such material in real size
coils may be proved industry for big projects (SMES and other
electric applications). The fast
progress of cry generators for space and other applications
allows one treasonably envisage that
in the future their reliability, mass and power consumption will
match the request for their use in
interplanetary sights, thus reducing drastically the volume and
mass of the helium supply. Surely,
it should be much better to use cry refrigerating systems
capable of maintaining the temperature
of the coils at a few Kelvin without requiring liquid helium.
Presently, the needed mass and
power consumption rule out this solution. However, also for this
kind of device, the technical
progress promises to be quick and a suitable R&D activity
devoted to their use in long duration
space sights could bring their parameters in the useful range.
The following evaluation for the
magnetic system will be based on the use of NbTi superconducting
material whose performances
are well known especially for the intensity of the field and the
needed refrigeration power. No
special hypotheses will be made for the cooling system at this
stage of the discussion. A
magnetic lens can be obtained by a toroidal configuration: an
electrical current running along a
cylindrical conductor, parallel to its axis, with the electric
current circuit closed in a return
circuit. The simplest scheme forth return circuit consists of a
cylindrical conductor at a larger
radius carrying the return current, connected by radial
conductors to the inner cylinder. Because
at this stage is not worthwhile to go into detail for the
geometry of the lens, we will assume that
the innermost and outermost cylinders, as well as the connecting
radial connectors, are
continuous surfaces, and the total cross-section of the
conductor is constant along the whole
circuit. The diameter of the inner cylinder is assumed to be
0.5m and its length 1 m. The radius
of the external cylinder can be as large as we wish, since its
mass does not change with the
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“Barsoom Inception: A mission for humanity”
31
radius, and only the mass of the connecting radial connector’s
changes. However, since the
magnetic lens must be transported to space in an already
assembled state, its external diameter
cannot exceed that of a big rocket or that of cargo bay of the
shuttle vehicle (the diameter of
cargo bay of the present shuttles is 4.6 m). We will assume 4 m
for the external diameter of the
lens, and we will maintain the length of 1 m also outside (see
Fig. 5). The region inside the inner
cylinder could be well protected against the incoming radiation
by the mass of the technical
devices servicing the coil, adding a further absorber only in
the parts not well protected by them.
For an approximate evaluation, let us assume that we deviate by
453thecosmic rays arriving
parallel to the coil axis at a radius slightly exceeding 0.25 m.
For 500 MeV kinetic energy
protons, slightly less than 3 T] nonintegrated magnetic field is
then required. The total current
required to obtain such a field is about 3.5MA. We assume to
have a current density of3500
A/mm2 in the inner cylindrical conductor, which is a safe
working point for a NbTi
superconductor.1This choice brings to an average current density
of 700 A/mm2 #owing in the
whole coil. Hence, the mean cross section of the s.c. coil turns
out to be 50 cm2, accounting for a
total volume of the s.c. of 50 cm2]560 cm"27 840 cm3 and a mass
of 103 kg. The average
thickness of the s.c. in the inner cylinder is 3.2 mm and in the
outer one is0.4 mm. With this coil
the resulting trajectories of500 MeV protons arriving at some
selected distances from the axis of
the lens are reported in Fig. 6.A look at this sketch of
trajectories indicates that the deflection
power of the lens should be reinforced around 1 m of radius in
order to maximize the volume of
the fully protected part of the space craft. A circular-shaped
coil can obtain this maximization.
Such a choice represents also a better geometrical situation for
the mechanical structure that
should support the magnetic stresses produced by the high field
inside the lens. With this choice
for the shape of the coil and assuming the same parameters as
before, with the same maximum
field atR"0.25 m, the trajectories of 500 MeV protons are those
sketched in Fig. 7. The fully
protected volume significantly increases in comparison to the
case of the rectangular-shaped coil
of Fig. 6. Furthermore, if the coil could be realized by a
nearly continuous surface, the coil could
be nearly self-supporting against the magnetic stresses, in
spite of its thin average thickness: in
fact, the magnetic stresses are directed outside the volume of
the coil and nearly perpendicular to
the conductor, and their elect is equivalent to an internal
pressure that probably requires only
relatively light external structures. Therefore, the total mass
of the magnetic lens will require the
addition of the only masses of the thermal shields and of the
cooling system. Another important
advantage of constructing a nearly continuous surface coil is
that the magnetic field will be fully
contained inside the volume of the coil, with no
external-fringing field.
Comparative discussion of passive and active protection from the
cosmic ray radiation:
We begin by comparing the mass of the magnetic system based on
the circularly shaped coil of
Fig with the mass of the absorber needed to have the same
protection. Also, the absorber can
protect from the directionality of the arriving cosmic rays, and
should consist therefore of a disk
of the same diameter as the external diameter of the coil. The
internal part of the disk (up to 0.25
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“Barsoom Inception: A mission for humanity”
32
m in radius) is common to the two devices and is kept out of the
comparison. The surface of the
disk will be indeed12.5 m2. The total mass of the magnetic lens
has tube evaluated adding to the
mass of the coil the masses of the thermal shields, of the
cooling system and of the supporting
structures. In "rst approximation, these masses can be
considered proportional to the mass of the
coil itself. The proportionality factor is completely arbitrary
without the consideration of a
specific design, and it will be considered arbitrarily equal2to
1. However, it is reasonable to
assume that all the elements servicing the coil cannot be
lighter than some minimum mass that
will be arbitrarily3assumed to be 50 kg.The results of the
comparison are reported in Table2 and
in Fig. 8 for several values of the maximum kinetic energy of
the particles. It is clear that as we
move up from the proton energy region of a few tens of MeV, the
magnetic system is much
superior.
Figure: Comparison between the directional magnetic lens and an
equivalent Al absorber
Figure: Weights of active and passive shield vs. cutoff energy
triangles: passive shield;
diamonds: magnet mass of the active shield; squares: total mass
of the active shield
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“Barsoom Inception: A mission for humanity”
33
Figure: Monitored cosmic rays 1967}1972 by IMP IV and IMP V
experiments. The differential
fluxes are measured in the energy range between 20 and 80 MeV.
In the same energy the galactic
cosmic ray flux is less than 10~4(cm2s sr MeV) ~1
COST ANALYSIS:
1. Cost Breakdown Structure
2. Lander module
3. Earth return vehicle
4. Mars habitation module
5. Mars Surface Elements
6. Earth Ground Infrastructure
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34
7. Prime Contractor Programme Level
8. Space Agency Programme Level
8. Operations
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36
System and mission engineering and landing:
Rationale for using Low Earth Orbit Propellant Depots with
HLVs:
Most sources now show that Mars landers will need to have very
wide diameters or bases: 10
meters (33 feet) or more. To use a lander on Mars, FIRST it
needs to get to LEO. Current ELV
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“Barsoom Inception: A mission for humanity”
37
(small-diameter) launchers cannot launch such wide payloads into
LEO. We need Mars landers
that can carry very large payloads to the surface - protected
from re-entry heating. We can launch
much larger landers DRY than when WET. If we launch them dry, we
need orbital Propellant
Depots to accumulate propellants for Transit and Mars landing
(EDL).Without Depots,
cryogenic propellants will sometimes boil off before a crew can
reach the vehicle to use it .With
Depots, the propellant in the first vehicle is not lost.
Building Propellant Depots should NOT be
used as a rationale for not building large Diameter HLV’s.
Buildup of Mars Fleet in LEO using
HLVs and Use of Propellant Depot as a ‘Vehicle Accumulator
‘Vehicles and equipment are
launched dry to LEO. Mars Ferries may be able act as a second
stage and put themselves in orbit
if fully fueled. Space tugs move orbiting vehicles and cargo to
depot area and dock at adjacent
assembly base. External (non-integral) aero-capture shields are
attached to all Mars bound
Transiting vehicles such as cargo carriers, crew Earth return
vehicles and two Depots: LEO to
LMO (Low Mars orbit).Ferries have integral aero-shells (as part
of their vehicle structure).Cargo
is loaded aboard cargo transit vehicles. All Transit vehicles
are fueled from large LEO Depot.
Sun-Shielded, Insulated Propellant Depot:
VASIMR – a possible Transit Alternative:
VASIMR is a very efficient plasma rocket system that, given a
200 Megawatt electrical power
source, could reduce Mars Transit time to ~50 days from 6-8
months, greatly reducing crew
exposure to solar radiation. This would save a huge mass of Mars
Transit propellants.VASIMR
could widen windows for Mars Transit missions. It reduces
mission risk from damage to liquid
fuel engines and loss of liquid propellant accidents. It could
also move Mars bound vehicles and
cargo from LEO to GEO before departure from GEO and maintain a
Mars base orbit. A current
version is rated in tests at 250 KW. For Mars Transit purposes,
development should start now on
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“Barsoom Inception: A mission for humanity”
38
a much larger and light weight space rated power supply needed
to power VASIMR, such as a
compact nuclear reactor, ultra-light solar panels, etc.
PROVISONS:
This System is NOT useful for Mars without the large power
supply. The use of chemical
(cryogenic) propellants for Mars transit missions is practical
without waiting for a compact
VASIMR power source.
Using aero capture at Mars:
Aero capture uses a very large diameter rigid aero-shield or
integral aero-shell to slow all
spacecraft arriving at Mars by flying through the atmosphere
once to brake down to orbital
speed. This saves a huge mass of liquid propellant.
Non-Cryogenic OMS Propellants may be
used for the orbit trim maneuver and rendezvous with the Low
Mars Orbit base. An Aero-shield
is not integral to the spacecraft, but is much wider than it and
partly surrounds it during the aero
capture maneuver. Aero-shields may be able to be assembled from
sections in LEO for use, (so
they will fit in an HLV cargo space). They can be kept extended
(out of the way) on a boom in
front of Mars- bound vehicles until arrival at Mars. The
Aero-shield is retracted and locked
before arrival. The Aero-shield can also be used during return
to Earth. Aero-capture may not be
compatible with use of VASIMR propulsion, which could eventually
replace it.Mars Ferry
vehicles would use an integral aero-shell which is part of the
vehicle’s structure to accomplish
the aero-capture.
Choices for a Mars Orbit Base Location:
The base should be high enough to avoid frequent orbit re-boosts
which use up fuel (about 400
km high or more – Low Mars Orbit .If the orbit is elliptical, it
will reduce the number of
opportunities for landings and takeoffs compared to a circular
orbit. A High Mars Orbit (HMO)
would require a lot more fuel to reach from the surface and for
landings than a LMO.A near-
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“Barsoom Inception: A mission for humanity”
39
equatorial orbit will maximize the equatorial eastward speed of
240 meters per second to reduce
fuel use for landings and takeoffs. This is 6% of takeoff
delta-V and about 25% or more of
powered landing delta-V requirements. A high inclination or
polar orbit would increase the
propellant mass needed for landings and takeoffs considerably.
It would also reduce the number
of opportunities for landings and takeoffs compared to a
near-equatorial orbit.
Other Equipment Needed in LMO Base:
2 or more Independent Crew Habitats with Solar Radiation Refuges
(same as in Transit
Vehicles), each capable of supporting all crew members until
return to Earth. Multiple spare
replaceable equipment modules and parts. Redundant food, water
and consumables for the
crew.2 Cryogenic Propellant Depots (shielded & active
cooling) Sets of Mars cargo landers (one-
way) and Mars Ferries.Intra-system crew vehicles or tugs to
explore Phobos etc.Tele-operation
equipment to run surface robots. All equipment needs to be
optimized to minimize crew time to
operate and maintain it, including accessibility and
modularization. This is a major lesson
learned from the space station.
Supersonic Decelerators:
Ballutesand Hyper cones a Balluteis like an inflated semi-rigid
parachute used to decelerate at
the end of re-entry. Work is beginning on a Ballute-like system
called a hyper cone that would
deploy after speed dropped below about 3250 mph. For a 60 ton
vehicle, the wide end of the
hyperconewould be about 100 feet across. The hyperconefabric
still has to be able to withstand
heating caused by friction with the air. The Hyperconeis hard to
deploy and control.Ballutes and
Hyperconesare expendable only.
Combinations are required if Supersonic Decelerators are
used.
Once speed drops below Mach 1, a subsonic parachute could be
deployed.
Below about 1 kilometer, landing rocket engines would be needed
to set the lander down gently.
This design requires 2 expendable systems, the hyper cone and
the subsonic parachute. There is
no way to recover and prevent a high speed crash if either the
Hyperconeor the parachute fails to
open properly. A failed chute could also endanger separation of
a crew cabin/escape capsule for
abort to surface mode.
Supersonic Retro Propulsion
This method is called Supersonic Retro Propulsion (SRP).It is
now being taken seriously. It
requires rocket thrust firing directly through the heat shield
and against the supersonic flow of air
pressing against the base of the vehicle as it decelerates. The
rocket engines must be fixed in
position with the nozzle ends flush with and sealed to the heat
shield and thus they cannot gimbal
for steering. Steering during re-entry and landing must be done
by varying the thrust of a set of
engines or by using side-mounted small Vernier engines.
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Descent Technologies:
Deceleration for Mars EDL from 400 km orbit.
We should use:
Banish the “Expendable Mentality “Think Re-Usable’’Current
scenarios for Manned Mars
landings envision a very large lander which has, inside it ,
just like a nested Russian doll
(Matryoshka), another entire vehicle for the ascent with its own
engines, tanks, controls,
structure, etc. This means that every trip to the surface
requires an entire additional pair of
vehicles with all of the descent propellant brought from Earth.
It also wastes all of the perfectly
good equipment in the descent vehicle or lander. This kind of
architecture is only good for the
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41
kind of Mars expedition we do not want: the “Flags and
Footprints “style mission, which is
financially unsustainable, and leads to the “one-way “Mars trips
currently being proposed by
those desperate to see any kind of Manned Mars Mission during
their lifetime.
Designing a Re-Usable Mars Ferry:
Re-usable vehicles using SRP can use their engines for:(1)
Initial de-orbit burn: (7560 mph –
7400mph) ~ Mach 13 (1) “Passive “Atmospheric Entry -Deceleration
(Mach 13 –Mach 6) (2)
Supersonic Retro-propulsion: (~Mach 6 -Mach 0.9) (3) Subsonic
Deceleration Mach 0.9 –Mach
0.2 (4) Final descent and landing: (Mach 0.2 / 100 mph to
surface) (5) Ascent back into Low
Mars Orbit. With a fully re-usable vehicle, nothing is thrown
away. Hydrogen and oxygen
propellants can be created from Mars ice and volatiles, using a
nuclear power source and carried
back to orbit for use on the next trip down, since there is
little cargo other than propellant that
needs to go up. Wide base vehicles with lower density slow down
more during re-entry and thus
need less propellant to land than a narrow base vehicle. All of
the propellants needed for Crew
and Cargo Descent can be carried UP on ascent. 5 tons of extra
propellant can be loaded back
into the Propellant Depot in Low Mars Orbit from each Cargo
Ferry trip UP for use in Earth
Return
A Single Stage to Orbit and Back Vehicle (SSTOAB) for Mars
The Mars Ferry is essentially an SSTO for Mars. Mars gravity is
0.38% of Earth’s, so achieving
low orbit is much easier than on earth -about 2.5miles per
second. This takes only about ¼of the
energy to reach L.E.O. Mars has 1/10 Earth mass, and 8 times
lunar mass.If there is no staging,
then there is no first stage to recover –the entire vehicle goes
to the Orbital “base”and back to the
surface base -intact. Much less fuel is needed to land than take
off to orbit since normal
atmospheric re-entry sheds up to the first 4310 mph (1.7-2.4
km/sec) of speed. A cargo ferry
would carry 25 tons of modules and equipment down to the surface
and 20 tons of propellants
back to orbit (15 tons to use for descent and 5 tons for the
Depot). Acrew ferry (with its 5 ton
crew cabin) would carry a crew with 20 tons of cargo down to the
surface, and a crew and 15
tons of propellant back to orbit.
Discovery of Widespread Sub-Surface Ice on Mars:
makes propellant production a Non-Exotic operation 20 years ago,
we had no knowledge of the
widespread existence of Wateron Mars in the Form of Sub-surface
ice deposits (ice regolith or
permafrost), some of them fairly close to the equator. Mars
Direct (1989) and related concepts
assumed we would bring the hydrogen to make methane fuel from
all the way from Earth. Now
the hydrogen can be obtained from Mars ice in large enough
quantities to use as a fuel directly.
Producing propellants at a Mars surface base is NOT an exotic
zero-gravity technology that still
needs to be developed. All of the individual steps are already
performed on Earth every day. This
avoids the need to first develop a technology to do it in
micro-gravity, and allows a direct
process of developing the hardware to do it under Mars
conditions. The extraction equipment
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“Barsoom Inception: A mission for humanity”
42
and cryogenic storage system would be built in “package
plant”modules so that it could be set up
easily by tele-operated robots, before crew members descend to
the surface. Any Surface Base
Site would be influenced by the availability of near-surface ice
posits that can be mined with
simple excavation equipment.
Descent Engine Placement for Lander –similar to Ferry using
S.R.P.:
Mars Ferry: Launch Weight - 125 Tonnes:
Both Cargo & Crew designs mass ~ 70 mt at descent and 125 mt
at lift off, with 20 tons total
ascent payload and 25 tons for descent. Both vehicles have a
base heat shield diameter 14 meters
wide. Both Ferry versions have a cargo bay to hold up to 25 tons
of large cargo with cargo bay
doors that open on the side below the fuel tanks. Both Ferry
versions have oversize tanks that can
hold 95 tons of propellant (about 75 tons for ascent, about 15
tons for the subsequent descent,
and about 5 tons extra for deposit in the propellant depot.
(Cargo ferry only)
CREW AND CARGO FERRY DIFFERENCES:
The Crew Ferry has a 5 ton crew cabin / escape capsule on top.
It would never separate from the
Ferry except in an emergency. The cargo Ferry has no equivalent
capsule on top. The Crew Ferry
would carry 15 tons of propellants UP as payload in its tanks
during an ascent mission, and 20
tons of cargo DOWN during a descent mission. The Cargo Ferry
carries (as payload) 20 tons
(fuel) UP in its tanks and 25 tons (cargo) DOWN in the cargo
bay. Total Ascent and Descent
Masses are the same for both Ferri.
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45
For safe landing and better descending we will use sky crane and
supersonic parasuit.
Figure: sky crane
Re-entry and finally to earth:
Most Mars Sample Return scenario involve two launches
One with an orbiter One with a lander, providing a rocket with a
sample storage container.
The rocket performs a rendezvous with the orbiter, then the
orbiter returns toward Earth.
This type of mission is both costly (two launches) and risky
(Mars orbit rendezvous), hence the
interest of single launch. Single launch mission profiles
Electric propulsion is a key enabler
for single launch missions. Two options are presented: A short
term mission with a
rendezvous in Mars orbit, A more advanced mission with EP,
Cryogenic propulsion, ZBO and
a heavy lander with direct return toward Earth.
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46
Most Mars Sample Return scenario involve two launches:
One with an orbiter One with a lander, providing a rocket with a
sample storage container.
The rocket performs a rendezvous with the orbiter, then the
orbiter returns toward
Earth. This type of mission is both costly (two launches) and
risky (Mars orbit rendezvous),
hence the interest of single launch. Single launch mission
profiles Electric propulsion is a key
enabler for single launch missions. Two options are presented :
A short term mission with a
rendezvous in Mars orbit, A more advanced mission with EP,
Cryogenic propulsion, ZBO and
a heavy lander with direct return toward Earth.
We will return using ep. process given below:
Mission profile:
Orbiter is separated from the main spacecraft (lander and MAV)
after Mars orbit insertion, EP
is used to circularize orbit.
The orbiter waits for rendezvous with the MAV capsule.
Then orbiter EP is used to increase orbit apoapsis and
eventually inject orbiter into
interplanetary orbit. EP is used to reach Mars – Earth transfer
orbit.
Launch mass 3905 kg (18% L/V mass margin)
Includes EP orbiter + DM + MAV + Carrier
EP orbiter wet mass: 625 kg.
Total xenon used 354 kg, EP Orbiter S/C dry mass 271 kg
Mission duration: 57.5 months (4.8 yrs.)
Operational mission at Mars: 107 days
Descent to nominal orbit 690 days and return up to escape 396
days
Delayed transfers very well suited for dual (simultaneous)
launch of electric low thrust and
chemical vehicles.
1PPS®1350 operating at any time (2 for the required
lifetime)
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47
Decisions:
EP orbiter yields higher mass margins
Longer mission duration allows for significant reduction of
∆V needs ( 10500 hrs. of operations. > 7000 ON-OFF cycles
(incl. 50 cold)
Nominal operating point
1500 W / 4,28 A / 350 V
88 mN / 1650 s / 50% total eff.
4,4 kg
Variable power feature demonstrated on
Smart-1
Can allow for cluster operations with
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48
Thrust steering capability.
Conclusion:
This is a dreamt mission. A mission for humanity. If there would
be no humanity it would be a
failure though it may have so much success. Because making earth
livable is more important
than making mars livable.
References:
1. R. Zubrin and D. Baker, "Humans to Mars in 1999,"Aerospace
America, August 1990.
2. D. Baker and R. Zubrin, "Mars Direct: Combining Near-Term
Technologies to Achieve a
Two-Launch Manned Mars Mission," JBIS, Nov. 1990.
3. R. Zubrin, D. Baker, and O. Gwynne, "Mars Direct: A Simple,
Robust, and Cost-Effective
Architecture for the Space Exploration Initiative," AIAA
91-0326, 29th Aerospace Science
Conf., Reno NV Jan.1991.
4. M. Duke, Private communication, Oct. 1992.
5. D. Weaver, "Mars Study Team Reference Mission Overview,"
Presented at Mars Working
Group conf. NASA Ames Research Center, May, 1993.
6. T. Stafford et-al, “America at the Threshold: Report of the
Synthesis Group on America’s
Space Exploration Initiative," U.S. Gov. Print. Off. May
1991
7. K. Joosten, Bret Drake, D. Weaver, and John Soldner, "Mission
Design Strategies for the
Human Exploration of Mars," IAF-91-336, 41st Congress of the
International Astronautically
Federation, Montreal, Canada, Oct. 1991.
8. R. Zubrin and T. Sulmeisters, "The Application of Nuclear
Power and Propulsion for Space
Exploration Missions,” AIAA-92-3778, AIAA/ASME Joint Propulsion
Conference, Nashville,
TN, July 1992.
9. A. Cohen et al "the 90 Day study on the Human Exploration of
the Moon and Mars,” U.S.
Government Printing Office, 1989.
10. A.Gonzales, L.Harper, E.Dunsky, and B.Roberts, “Mars Surface
Mission Life support
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49
Team profile:
Team leader: Nazmul Ahsan Saikat; Email:
[email protected]
Co- Team leader: Yeasir Mohammad Akib; Email:
[email protected]
Other team member: Md. Nur-e-alam Siddiky; Email:
[email protected]
Every one of our team member is 3rd year graduate student.
University: Rajshahi University of Engineering & Technology
(RUET), Rajshahi, Bangladesh.
mailto:[email protected]