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                         Mars Sample Return as a Micromission

                                Steve Kemble & Bob Parkinson
                                          Astrium Ltd
                                      Gunnels Wood Road
                                      Stevenage SG1 2AS

                 Abstract                        minimum of equipment. As a consequence,
The return of ~200 gm of surface/sub-surface     neither the size of the returned sample nor the
material would provide a major improvement       equipment required to retrieve it represent
in our understanding of Mars. With such a        limiting factors in sizing the mission.
small return payload, sample mass is not a
performance driver, and a micro-spacecraft       As a possible follow-on to Beagle 2, a study
approach would reduce mission size and cost.     has been made of performing Mars Sample
The possibility of such a mission is             Return using a Soyuz-Fregat launch. By
investigated using a Soyuz-Fregat launcher.      deliberately limiting spacecraft size in a
Key issues are the size of the Mars Ascent       micro-spacecraft mission approach, costs
Vehicle, and the V required for entry into      might be limited by focussing on a single,
Mars orbit and return to Earth. Optimized        scientifically important objective. The study
transfers have been developed to minimise V     was intended to identify key technology
for the return. The benefits of solar-electric   objectives required to define a possible
propulsion (SEP) depend on trade-offs in the     mission rather than to solve all problems.
system design, and high specific impulse does    Key issues are the Vs required for entry into
not necessarily yield the greatest vehicle       Mars orbit and later return to Earth, and the
payload. New operational technologies could      sizing of the Mars Ascent Vehicle.
be tested by adapting the vehicle design for a
precursor Lunar Sample Return (e.g. from the        2. MISSION STRATEGY
polar Aitken Basin region), providing
additional scientific return.                         a. Mission Outline

   1. INTRODUCTION                               While Soyuz-Fregat represents an established,
                                                 low cost, medium class launcher, progressive
Mars Sample Return is seen as a key mission      improvements to its launch capability are
in the exploration of Mars. Past studies have    expected. Russian data presented as a part of
produced complex and expensive missions          the Bepi-Colombo collaboration identified
demanding a heavy lift launch, or even           modifications that would allow Soyuz-Fregat
multiple launches. But the 1970 Russian          to inject 1334 kg or 1564 kg onto a Mars
Luna 16 provided a useful scientific return      transfer trajectory. For the purposes of this
from just 100 gm of regolith. On Mars,           study the 1334 kg initial mass was taken as
aeolian transport of surface fines means that    the spacecraft objective, with the 1564 kg
any sample can be expected to contain grains     capability providing the system margin.
from a wide – perhaps even planet-wide –
area. The Beagle 2 “Mole” provides a means       The sample return mission assumed that, as
of obtaining subsurface material with the        with Mars Express/Beagle 2, the launched
assembly would include both an Orbiter and a      the optimum arrival. Compromise is therefore
Lander, journeying together to Mars, but          required to accommodate the short stay
separating during the hyperbolic approach.        duration. An optimum solution must derive an
The Lander would make a direct entry into the     Earth departure epoch, Mars arrival epoch and
Martian atmosphere and the Orbiter use on-        Earth return epoch that allow minimum
board propulsion to enter a parking orbit         mission V. These solutions were obtained to
about the planet. The Lander would provide        form the basis of the mission design. Typical
entry protection and descent and landing          characteristics of such a transfer are:
systems around an Ascent Vehicle that – after
having been loaded with the sample of                 Earth Escape velocity     3.2 km/sec
Martian fines – would return to a low altitude        Mars approach velocity    2.8 km/sec
(350 km) orbit about Mars. The Orbiter                Outward       Transfer    220 days
would then rendezvous with the Ascent                 duration
Vehicle, transfer the sample container, and at        Mars Escape velocity      5.8 km/sec
the appropriate time boost itself onto an Earth       Earth        approach     4.8 km/sec
return trajectory. The sample would be                velocity
returned to Earth using a small, direct entry         Return         transfer   520 days
capsule carried by the Orbiter.                       duration

     b. Transfer Strategy                         In fact, the outward journey parameters are
                                                  nearly optimal for a one-way transfer. The
The design of the sample return mission aims      example is based on a 2003 or 2018 launch,
to maximise payload and sample return mass        with a near minimum outward leg
with a very limited stay on the Martian           requirement for any launch epoch. The
surface. The stay duration is short, typically    transfer duration of the return is longer than
10 days. The first problem to solve is an         those that optimum ‘single leg’ transfers, as
optimised transfer from Earth to Mars,            are the escape and approach velocities.
followed by an almost immediate optimised         Similar total V results can be obtained for a
return. The baseline propulsion option system     range of launch epochs, with a change in the
is chemical, and manoeuvres are near              balance between outward and return legs.
                                                  A more efficient return trajectory can be
For an Earth to Mars transfer, the V required    found by allowing a longer transfer involving
depends on the relative planetary positions at    approximately one and a half heliocentric
injection into the inter-planetary transfer       revolutions. However, the return transfer now
orbit. A minimum energy case exists where         takes over 700 days. Alternatively, shorter
the transfer orbit has an aphelion close to       return transfers can be found, but at the
Mars orbital radius and a perihelion at Earth     expense of V. Return trip durations of less
radius. This optimum relationship of the          than 400 days involve a passage inside 1 AU
planetary positions arises at a frequency given   (in fact typically 0.8AU).
by the synodic period. For Earth and Mars,
this period is 785 days. This defines the         The problem is illustrated in Fig.1. The sum
interval between minimum energy transfers.        of Vinfnities (Mars departure plus Earth
The geometrical repeat period is 7 synodic        arrival), for a range of Mars departure epochs
periods, at 15 years.                             and trip times are shown for short duration
                                                  stays. Both the optimal, 500 day trip solution
Exactly the same situation exists regarding       and shorter return trips can be seen. It is
optimum transfers from Mars to Earth.             generally of more importance to reduce the
However, the epochs of the optimum return         Mars departure Vinfinity and accept a penalty
opportunities do not correspond to those of
on the Earth approach speed, if aeroassisted                                                                                                          operational orbits. Staging offers the
capture options are used at Earth.                                                                                                                    possibility of substantial net mass savings.
                                                                                                                                                      This could be done using two spacecraft
                                                                                                                                                      buses, each with its own propulsion. The first
                                                                                                                               17000-18000            stage would insert the second stage and Mars
                                                                                                                               15000-16000            descent composite into Martian orbit, and the
                                                                                                                                                      second stage used to return to Earth after
                                                                                                                                                      rendezvous with the Mars ascent vehicle.
                                                                                                                                                      Alternatively, greater net mass saving can be
                                                                                                                      Vinfinity Sum
                                                                                                                                                      achieved using a single spacecraft bus with a
                                                                                                                          (m /s)                      separable tank system. This removes the need
                              20                                                                              13000
                                                                                                                                                      for additional thruster units, but retains the
                               60                                                                             12000
  Stay tim e                   100                                                                            11000                                   advantage of jettison of redundant tank mass.
    (days)                                                                                                    10000
                                                                                                                                                      Such a system requires a more complex

                                                                                                                                                      separation system but improves performance.

                                                                 Transfer tim e (days)

                                                                                                                                                      A further option is to use Solar Electric
                                                                                                                                                      Propulsion (SEP) to implement the transfer,
Fig.1: Return leg total Vinfinity sensitivity to                                                                                                      Transfer times of typically 300 to 400 days
stay time and trip duration                                                                                                                           each way can be found with limited stay times
                                                                                                                                                      at Mars. SEP can be used for orbit insertion
Fig.2 compares the locally optimal, direct                                                                                                            and escape at Mars in addition to the
return route with the locally 1.5 revolution                                                                                                          interplanetary transfer. The high specific
return route. Total Vinfinity requirements are                                                                                                        impulse of the low thrust system ensures low
shown being the sum of Vinfinities for the                                                                                                            fuel loads, but a key item is the mass required
outward and return legs. The global minimum                                                                                                           for the thrusters and power generation. This
is seen to lie with the 1.5 rev return route for                                                                                                      later consideration is particularly relevant for
short stay times. As the stay time extends to                                                                                                         solar electric systems at Mars.
400 days, the situation is reversed
                                                                                                                                                      The use of low thrust opens the possibilities
                                                                                                                                                      for other interesting transfer techniques. If
                           25000                                                                                                                      longer transfer durations can be accepted, use
                                                                                                                                                      of Earth gravity assist can reduce total
                                                                                                                                                      mission Vs by >2 km/sec. Transfer times
   Total Hyp Excess(m/s)


                           15000                                                                                                                      increase by over a year on the outward leg
                                                                                                                               Direct return               d. Earth Entry Vehicle
                                                                                                                               1.5 Rev return
                                                                                                                                                      For a minimal mission, direct entry return to
                                   0             100                200               300               400              500          600       700   Earth is necessary. The Japanese Muses-C
                                                                                     Stay Time(days)                                                  return capsule, performing a very similar
                                                                                                                                                      mission, has a mass of 25 kg [AWST May 19
Fig.2: Total mission Vinfinity requirements vs                                                                                                        2003, p. 40]. For the Mars Sample Return a
stay time for direct and 1.5 rev return types                                                                                                         simpler (but slightly heavier at 40 kg) Earth
                                                                                                                                                      Entry Vehicle has been conceived (Fig. 3)
                            c. Orbiter Options                                                                                                        with a low entry ballistic coefficient limiting
                                                                                                                                                      the terminal descent through the Earth’s
The orbiter is required to implement a large                                                                                                          atmosphere to low speed (~22 m/s) without
V, injecting to and departing Mars
the use of a parachute, relying on crushable
protection to cushion the sample at impact.

Fig. 3: Passive Earth Entry Vehicle Concept
                                                  Fig. 4: Stowage of the Ascent Vehicle within
   3. LANDER/ASCENT VEHICLE                                    the entry capsule.

     a. Mars Entry, Descent & Landing                 A “Mole” for collecting the sub-surface
The Mars entry heat shield (see Fig.4) has an         A robot arm to deploy and retrieve the
identical ballistic coefficient (152 kg/m2) and       “Mole”, and to load the collected sample
geometry (30 sweep angle, 0.417 m nose               into the Ascent Vehicle
radius) to Beagle 2. With an entry mass of            A camera
187 kg, the heat shield diameter is 1.480 m.          A UHF transceiver for communication
The leeward aeroshell cone has to be                  with the Orbiter
proportionately taller than Beagle 2 to               An inclinometer, to establish the local
accommodate the ascent vehicle, but will still        vertical with respect to the ascent vehicle
be entirely within the wake region of the entry       co-ordinates.
heat shield. After entry into the Martian
atmosphere, the heat shield and entry             All but the last are directly derived from
aeroshell separate to allow deployment of the     Beagle 2. The robot arm has sufficient reach
parachutes. For the MSR a cluster of 3            to extend past the deflated airbags and obtain
parachutes is proposed to improve packaging       a sample from a clear Martian surface.
within the entry capsule. Deployment of the
main parachutes takes place at a descent          Power for surface operations is provided by
speed of 89 m/s. The terminal descent speed,      LiSOCl2 primary batteries, sized to provide
as for Beagle 2, will be about 16 m/s.            power to operate the Lander for 4 sols after
                                                  touch-down. This is judged sufficient to
Surface impact attenuation assumes the use of     collect the sample. Use of deployable solar
deflating airbags, as opposed to the              arrays as in Beagle 2 was ruled out due to
“bouncing” airbag system used on Beagle 2.        packaging and deployment problems.
Release of the system from the parachutes
occurs at ~50 m altitude, with air-bag sizing           b. Ascent Vehicle
and impact velocity designed to avoid the
vehicle toppling even at the maximum              The two-stage Ascent Vehicle uses a simple
expected lateral drift velocity.                  bi-propellant propulsion system for the
                                                  booster stage, and a monopropellant (N2H4)
The descent assembly carries equipment for        orbital stage. The task of the Ascent Vehicle
the following surface operations:                 is to propel itself into a low Martian orbit, and
to maintain itself there as a co-operating
target for Orbiter rendezvous.            Mass     Once the Ascent Vehicle has achieved orbit,
performance is critical, since the vehicle must    the Orbiter must accurately determine the
achieve a V of about 3850 m/s, with a             orbital elements of each vehicle to carry out
launch mass of about 91 kg.                        the rendezvous and capture. A variety of
                                                   possibilities exist for this, including use of the
Because the Ascent Vehicle launches from the       star trackers with Mars occulation to
Martian surface, the booster stage is a single-    determine the orbits of each vehicle, active
burn, pressure fed system without the need for     laser range-finding by the Orbiter, or even
propellant retention devices. The booster          direct optical acquisition. Fortunately time is
stage uses a fixed 500 N liquid engine and 4 x     not a critical constraint at this stage. By
22 N RCS steering thrusters. The small size        starting in an initially elliptical orbit and
of the ascent vehicle makes the resulting          descending to the 350 km circular orbit of the
vehicle quite agile, with adequate control for     Ascent Vehicle the Orbiter can minimize the
the ascent burn.                                   effects of differences in the orbital planes of
                                                   the two vehicles by performing combined
The tiny upper stage (~18 kg at separation)        burn manoeuvres.
uses 4x10 N thrusters offset at 7 to perform
the orbital insertion. The stage then deploys a         d. Mass Budget
~30 w solar array and maintains a 3-axis
stabilized, sun-pointing attitude until the        Table 1 shows an initial mass budget for the
Orbiter makes its rendezvous and retrieves the     mission. The present design lies within the
sample container for return to Earth.              projected capability of the Soyuz-Fregat
                                                   launcher. However, the system is sensitive to
The sensitivity of the mission to Ascent           a number of design assumptions. Because the
Vehicle inert mass means that the design must      Lander separates from the Orbiter before
avoid “fixed mass” items such as bolted            Mars entry, while mass growth factors in the
joints, connectors and individual equipment        Lander are significant, mass growth in the
boxes, which at this small size become             Lander does not consequentially impact the
dominant features in the mass budget.              Orbiter design. Elsewhere, mass growth
                                                   factors are less significant than specific
     c. Ascent Vehicle/Orbiter Operations          impulse or V requirements.

For Lander surface operations, the Orbiter                                        Mass (kg)
will act as a communications relay to Earth.        Mass at Launch                      1240
This drives mission design immediately after        Lander at Separation                 187
arrival. With a low thrust LAE the Orbiter            Entry & Surface Element      97
will not be able to enter a low Mars orbit            Ascent Vehicle Dry Mass      28
efficiently with a single burn. The first burn        Propellant Loaded            62
will achieve a 400 km x 33753 km altitude           Orbiter at Separation               1053
orbit with a second burn one Martian day              Arrival Module Dry Mass      75
(sol) later, placing it in a 400 x 4000 km orbit      Arrival Propellant          518
with a 3.52 hour period. Communications               Return Module Dry Mass       77
with the Lander cannot begin until the start of       Departure Propellant        343
sol 2. The objective then is to determine the         Earth Entry Vehicle          40
sampling site during sol 2, with Mole               Margin on Soyuz-Fregat               324
operations extending into sol 3, ending with        Capability
the stowage of the sample within the Ascent
                                                     Table 1: Mass Budget for Mars Mission
Vehicle and launch to Mars orbit. Sol 4
provides a margin for surface operations.
    4. TECHNOLOGY DEVELOPMENT                        5. CONCLUSION
                                                  Use of a micro-spacecraft approach to a Mars
The MSR mission involves a number of              Sample Return mission suggests the
critical operations that have not, or will not    feasibility of returning 200 gm of surface and
have been done before, specifically:              sub-surface fines with a spacecraft assembly
                                                  within projected Soyuz-Fregat launch
  Relaunch of a sample from a planetary          performance. The projected mission would
  body                                            use direct entry of the Lander into the Martian
 Rendezvous and sample transfer in orbit         atmosphere, following the approach adopted
 Recovery with a passive re-entry vehicle        by Beagle 2, followed by Mars orbit
  at hyperbolic velocities.                       rendezvous, and direct entry into the Earth’s
                                                  atmosphere by the returning vehicle.
A preliminary “demonstration” mission is
therefore advisable. An interesting possibility   The design studied has a system level margin
is to adapt the systems and equipment for a       of 26% on the maximum projected capability
lunar sample return mission (possibly in the      of Soyuz-Fregat, but has some mass
lunar polar region), demonstrating the three      sensitivities due to the high V requirement
critical aspects close at hand where near-real-   of the returning vehicle. However, the V
time communications are possible. A               requirements are subject to the stay time at
precursor mission of this sort would clearly      Mars and other mission design features.
have a scientific value in its own right.
                                                  Key operational technologies not yet
The Lander would need to be modified to           demonstrated (or not demonstrated in the
descend using rocket braking, replacing the       Mars Express/Beagle 2 mission) include re-
entry and descent systems with a hydrazine        launch of a sample from a planetary surface,
monopropellant propulsion system (as a cheap      rendezvous and sample transfer in orbit, and
and reliable development).      The Orbiter       passive re-entry return to Earth at hyperbolic
would now require only a small V capability      velocities. It is suggested that the Mars
(~800 m/s each for entry and departure),          Sample Return hardware could be modified
which could be accommodated with the return       with limited effort to achieve a lunar sample
capability of the Mars Orbiter. In this           return mission that would both demonstrate
instance it is not necessary to have a            the requisite capabilities and also fulfil a
separating arrival stage. The resulting mass      valuable scientific objective.
budget for the lunar mission is shown below
in Table 2.

                               Mass (kg)
Mass at launch                         1122
Orbiter at launch                       449
  Orbiter dry mass            154
  Orbiter propellant          295
Lander at Separation                    680
  Descent stage dry mass      130
  Descent propellant          479
  Ascent vehicle dry mass      41
  Ascent propellant            30
Table 2: Mass Budget for Lunar Mission

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