IAC-03-Q-3.b.08
Mars Sample Return as a Micromission
Steve Kemble & Bob Parkinson
Astrium Ltd
Gunnels Wood Road
Stevenage SG1 2AS
bobparkinson@ntlworld.com
Abstract
The return of ~200 gm of surface/sub-surface
material would provide a major improvement
in our understanding of Mars. With such a
small return payload, sample mass is not a
performance driver, and a micro-spacecraft
approach would reduce mission size and cost.
The possibility of such a mission is
investigated using a Soyuz-Fregat launcher.
Key issues are the size of the Mars Ascent
Vehicle, and the V required for entry into
Mars orbit and return to Earth. Optimized
transfers have been developed to minimise V
for the return. The benefits of solar-electric
propulsion (SEP) depend on trade-offs in the
system design, and high specific impulse does
not necessarily yield the greatest vehicle
payload. New operational technologies could
be tested by adapting the vehicle design for a
precursor Lunar Sample Return (e.g. from the
polar Aitken Basin region), providing
additional scientific return.
1. INTRODUCTION
Mars Sample Return is seen as a key mission
in the exploration of Mars. Past studies have
produced complex and expensive missions
demanding a heavy lift launch, or even
multiple launches. But the 1970 Russian
Luna 16 provided a useful scientific return
from just 100 gm of regolith. On Mars,
aeolian transport of surface fines means that
any sample can be expected to contain grains
from a wide – perhaps even planet-wide –
area. The Beagle 2 “Mole” provides a means
of obtaining subsurface material with the
minimum of equipment. As a consequence,
neither the size of the returned sample nor the
equipment required to retrieve it represent
limiting factors in sizing the mission.
As a possible follow-on to Beagle 2, a study
has been made of performing Mars Sample
Return using a Soyuz-Fregat launch. By
deliberately limiting spacecraft size in a
micro-spacecraft mission approach, costs
might be limited by focussing on a single,
scientifically important objective. The study
was intended to identify key technology
objectives required to define a possible
mission rather than to solve all problems.
Key issues are the Vs required for entry into
Mars orbit and later return to Earth, and the
sizing of the Mars Ascent Vehicle.
2. MISSION STRATEGY
a. Mission Outline
While Soyuz-Fregat represents an established,
low cost, medium class launcher, progressive
improvements to its launch capability are
expected. Russian data presented as a part of
the Bepi-Colombo collaboration identified
modifications that would allow Soyuz-Fregat
to inject 1334 kg or 1564 kg onto a Mars
transfer trajectory. For the purposes of this
study the 1334 kg initial mass was taken as
the spacecraft objective, with the 1564 kg
capability providing the system margin.
The sample return mission assumed that, as
with Mars Express/Beagle 2, the launched
assembly would include both an Orbiter and a
Lander, journeying together to Mars, but
separating during the hyperbolic approach.
The Lander would make a direct entry into the
Martian atmosphere and the Orbiter use onboard propulsion to enter a parking orbit
about the planet. The Lander would provide
entry protection and descent and landing
systems around an Ascent Vehicle that – after
having been loaded with the sample of
Martian fines – would return to a low altitude
(350 km) orbit about Mars. The Orbiter
would then rendezvous with the Ascent
Vehicle, transfer the sample container, and at
the appropriate time boost itself onto an Earth
return trajectory. The sample would be
returned to Earth using a small, direct entry
capsule carried by the Orbiter.
b. Transfer Strategy
The design of the sample return mission aims
to maximise payload and sample return mass
with a very limited stay on the Martian
surface. The stay duration is short, typically
10 days. The first problem to solve is an
optimised transfer from Earth to Mars,
followed by an almost immediate optimised
return. The baseline propulsion option system
is chemical, and manoeuvres are near
impulsive.
For an Earth to Mars transfer, the V required
depends on the relative planetary positions at
injection into the inter-planetary transfer
orbit. A minimum energy case exists where
the transfer orbit has an aphelion close to
Mars orbital radius and a perihelion at Earth
radius. This optimum relationship of the
planetary positions arises at a frequency given
by the synodic period. For Earth and Mars,
this period is 785 days. This defines the
interval between minimum energy transfers.
The geometrical repeat period is 7 synodic
periods, at 15 years.
Exactly the same situation exists regarding
optimum transfers from Mars to Earth.
However, the epochs of the optimum return
opportunities do not correspond to those of
the optimum arrival. Compromise is therefore
required to accommodate the short stay
duration. An optimum solution must derive an
Earth departure epoch, Mars arrival epoch and
Earth return epoch that allow minimum
mission V. These solutions were obtained to
form the basis of the mission design. Typical
characteristics of such a transfer are:
Earth Escape velocity
Mars approach velocity
Outward
Transfer
duration
Mars Escape velocity
Earth
approach
velocity
Return
transfer
duration
3.2 km/sec
2.8 km/sec
220 days
5.8 km/sec
4.8 km/sec
520 days
In fact, the outward journey parameters are
nearly optimal for a one-way transfer. The
example is based on a 2003 or 2018 launch,
with a near minimum outward leg
requirement for any launch epoch. The
transfer duration of the return is longer than
those that optimum ‘single leg’ transfers, as
are the escape and approach velocities.
Similar total V results can be obtained for a
range of launch epochs, with a change in the
balance between outward and return legs.
A more efficient return trajectory can be
found by allowing a longer transfer involving
approximately one and a half heliocentric
revolutions. However, the return transfer now
takes over 700 days. Alternatively, shorter
return transfers can be found, but at the
expense of V. Return trip durations of less
than 400 days involve a passage inside 1 AU
(in fact typically 0.8AU).
The problem is illustrated in Fig.1. The sum
of Vinfnities (Mars departure plus Earth
arrival), for a range of Mars departure epochs
and trip times are shown for short duration
stays. Both the optimal, 500 day trip solution
and shorter return trips can be seen. It is
generally of more importance to reduce the
Mars departure Vinfinity and accept a penalty
on the Earth approach speed, if aeroassisted
capture options are used at Earth.
operational orbits. Staging offers the
possibility of substantial net mass savings.
This could be done using two spacecraft
buses, each with its own propulsion. The first
stage would insert the second stage and Mars
descent composite into Martian orbit, and the
second stage used to return to Earth after
rendezvous with the Mars ascent vehicle.
19000-20000
18000-19000
17000-18000
16000-17000
15000-16000
14000-15000
13000-14000
12000-13000
11000-12000
20000
10000-11000
19000
Alternatively, greater net mass saving can be
achieved using a single spacecraft bus with a
separable tank system. This removes the need
for additional thruster units, but retains the
advantage of jettison of redundant tank mass.
Such a system requires a more complex
separation system but improves performance.
18000
17000
16000
15000
Vinfinity Sum
(m /s)
14000
13000
20
12000
100
11000
540
510
480
450
420
390
360
330
300
600
10000
140
570
Stay tim e
(days)
60
Transfer tim e (days)
A further option is to use Solar Electric
Propulsion (SEP) to implement the transfer,
Transfer times of typically 300 to 400 days
each way can be found with limited stay times
at Mars. SEP can be used for orbit insertion
and escape at Mars in addition to the
interplanetary transfer. The high specific
impulse of the low thrust system ensures low
fuel loads, but a key item is the mass required
for the thrusters and power generation. This
later consideration is particularly relevant for
solar electric systems at Mars.
Fig.1: Return leg total Vinfinity sensitivity to
stay time and trip duration
Fig.2 compares the locally optimal, direct
return route with the locally 1.5 revolution
return route. Total Vinfinity requirements are
shown being the sum of Vinfinities for the
outward and return legs. The global minimum
is seen to lie with the 1.5 rev return route for
short stay times. As the stay time extends to
400 days, the situation is reversed
The use of low thrust opens the possibilities
for other interesting transfer techniques. If
longer transfer durations can be accepted, use
of Earth gravity assist can reduce total
mission Vs by >2 km/sec. Transfer times
increase by over a year on the outward leg
Total Hyp Excess(m/s)
25000
20000
15000
10000
d. Earth Entry Vehicle
Direct return
5000
1.5 Rev return
0
0
100
200
300
400
500
600
Stay Time(days)
Fig.2: Total mission Vinfinity requirements vs
stay time for direct and 1.5 rev return types
c. Orbiter Options
The orbiter is required to implement a large
V, injecting to and departing Mars
700
For a minimal mission, direct entry return to
Earth is necessary. The Japanese Muses-C
return capsule, performing a very similar
mission, has a mass of 25 kg [AWST May 19
2003, p. 40]. For the Mars Sample Return a
simpler (but slightly heavier at 40 kg) Earth
Entry Vehicle has been conceived (Fig. 3)
with a low entry ballistic coefficient limiting
the terminal descent through the Earth’s
atmosphere to low speed (~22 m/s) without
the use of a parachute, relying on crushable
protection to cushion the sample at impact.
Fig. 3: Passive Earth Entry Vehicle Concept
3. LANDER/ASCENT VEHICLE
a. Mars Entry, Descent & Landing
The Mars entry heat shield (see Fig.4) has an
identical ballistic coefficient (152 kg/m2) and
geometry (30 sweep angle, 0.417 m nose
radius) to Beagle 2. With an entry mass of
187 kg, the heat shield diameter is 1.480 m.
The leeward aeroshell cone has to be
proportionately taller than Beagle 2 to
accommodate the ascent vehicle, but will still
be entirely within the wake region of the entry
heat shield. After entry into the Martian
atmosphere, the heat shield and entry
aeroshell separate to allow deployment of the
parachutes. For the MSR a cluster of 3
parachutes is proposed to improve packaging
within the entry capsule. Deployment of the
main parachutes takes place at a descent
speed of 89 m/s. The terminal descent speed,
as for Beagle 2, will be about 16 m/s.
Surface impact attenuation assumes the use of
deflating airbags, as opposed to the
“bouncing” airbag system used on Beagle 2.
Release of the system from the parachutes
occurs at ~50 m altitude, with air-bag sizing
and impact velocity designed to avoid the
vehicle toppling even at the maximum
expected lateral drift velocity.
The descent assembly carries equipment for
the following surface operations:
Fig. 4: Stowage of the Ascent Vehicle within
the entry capsule.





A “Mole” for collecting the sub-surface
sample
A robot arm to deploy and retrieve the
“Mole”, and to load the collected sample
into the Ascent Vehicle
A camera
A UHF transceiver for communication
with the Orbiter
An inclinometer, to establish the local
vertical with respect to the ascent vehicle
co-ordinates.
All but the last are directly derived from
Beagle 2. The robot arm has sufficient reach
to extend past the deflated airbags and obtain
a sample from a clear Martian surface.
Power for surface operations is provided by
LiSOCl2 primary batteries, sized to provide
power to operate the Lander for 4 sols after
touch-down. This is judged sufficient to
collect the sample. Use of deployable solar
arrays as in Beagle 2 was ruled out due to
packaging and deployment problems.
b. Ascent Vehicle
The two-stage Ascent Vehicle uses a simple
bi-propellant propulsion system for the
booster stage, and a monopropellant (N2H4)
orbital stage. The task of the Ascent Vehicle
is to propel itself into a low Martian orbit, and
to maintain itself there as a co-operating
target for Orbiter rendezvous.
Mass
performance is critical, since the vehicle must
achieve a V of about 3850 m/s, with a
launch mass of about 91 kg.
Because the Ascent Vehicle launches from the
Martian surface, the booster stage is a singleburn, pressure fed system without the need for
propellant retention devices. The booster
stage uses a fixed 500 N liquid engine and 4 x
22 N RCS steering thrusters. The small size
of the ascent vehicle makes the resulting
vehicle quite agile, with adequate control for
the ascent burn.
The tiny upper stage (~18 kg at separation)
uses 4x10 N thrusters offset at 7 to perform
the orbital insertion. The stage then deploys a
~30 w solar array and maintains a 3-axis
stabilized, sun-pointing attitude until the
Orbiter makes its rendezvous and retrieves the
sample container for return to Earth.
The sensitivity of the mission to Ascent
Vehicle inert mass means that the design must
avoid “fixed mass” items such as bolted
joints, connectors and individual equipment
boxes, which at this small size become
dominant features in the mass budget.
c. Ascent Vehicle/Orbiter Operations
For Lander surface operations, the Orbiter
will act as a communications relay to Earth.
This drives mission design immediately after
arrival. With a low thrust LAE the Orbiter
will not be able to enter a low Mars orbit
efficiently with a single burn. The first burn
will achieve a 400 km x 33753 km altitude
orbit with a second burn one Martian day
(sol) later, placing it in a 400 x 4000 km orbit
with a 3.52 hour period. Communications
with the Lander cannot begin until the start of
sol 2. The objective then is to determine the
sampling site during sol 2, with Mole
operations extending into sol 3, ending with
the stowage of the sample within the Ascent
Vehicle and launch to Mars orbit. Sol 4
provides a margin for surface operations.
Once the Ascent Vehicle has achieved orbit,
the Orbiter must accurately determine the
orbital elements of each vehicle to carry out
the rendezvous and capture. A variety of
possibilities exist for this, including use of the
star trackers with Mars occulation to
determine the orbits of each vehicle, active
laser range-finding by the Orbiter, or even
direct optical acquisition. Fortunately time is
not a critical constraint at this stage. By
starting in an initially elliptical orbit and
descending to the 350 km circular orbit of the
Ascent Vehicle the Orbiter can minimize the
effects of differences in the orbital planes of
the two vehicles by performing combined
burn manoeuvres.
d. Mass Budget
Table 1 shows an initial mass budget for the
mission. The present design lies within the
projected capability of the Soyuz-Fregat
launcher. However, the system is sensitive to
a number of design assumptions. Because the
Lander separates from the Orbiter before
Mars entry, while mass growth factors in the
Lander are significant, mass growth in the
Lander does not consequentially impact the
Orbiter design. Elsewhere, mass growth
factors are less significant than specific
impulse or V requirements.
Mass (kg)
Mass at Launch
1240
Lander at Separation
187
Entry & Surface Element
97
Ascent Vehicle Dry Mass
28
Propellant Loaded
62
Orbiter at Separation
1053
Arrival Module Dry Mass
75
Arrival Propellant
518
Return Module Dry Mass
77
Departure Propellant
343
Earth Entry Vehicle
40
Margin on Soyuz-Fregat
324
Capability
Table 1: Mass Budget for Mars Mission
4. TECHNOLOGY DEVELOPMENT
/ PRECURSOR MISSION
5. CONCLUSION

Use of a micro-spacecraft approach to a Mars
Sample Return mission suggests the
feasibility of returning 200 gm of surface and
sub-surface fines with a spacecraft assembly
within projected Soyuz-Fregat launch
performance. The projected mission would
use direct entry of the Lander into the Martian
atmosphere, following the approach adopted
by Beagle 2, followed by Mars orbit
rendezvous, and direct entry into the Earth’s
atmosphere by the returning vehicle.
A preliminary “demonstration” mission is
therefore advisable. An interesting possibility
is to adapt the systems and equipment for a
lunar sample return mission (possibly in the
lunar polar region), demonstrating the three
critical aspects close at hand where near-realtime communications are possible. A
precursor mission of this sort would clearly
have a scientific value in its own right.
The design studied has a system level margin
of 26% on the maximum projected capability
of Soyuz-Fregat, but has some mass
sensitivities due to the high V requirement
of the returning vehicle. However, the V
requirements are subject to the stay time at
Mars and other mission design features.
The MSR mission involves a number of
critical operations that have not, or will not
have been done before, specifically:
Relaunch of a sample from a planetary
body
 Rendezvous and sample transfer in orbit
 Recovery with a passive re-entry vehicle
at hyperbolic velocities.
The Lander would need to be modified to
descend using rocket braking, replacing the
entry and descent systems with a hydrazine
monopropellant propulsion system (as a cheap
and reliable development).
The Orbiter
would now require only a small V capability
(~800 m/s each for entry and departure),
which could be accommodated with the return
capability of the Mars Orbiter. In this
instance it is not necessary to have a
separating arrival stage. The resulting mass
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
Key operational technologies not yet
demonstrated (or not demonstrated in the
Mars Express/Beagle 2 mission) include relaunch of a sample from a planetary surface,
rendezvous and sample transfer in orbit, and
passive re-entry return to Earth at hyperbolic
velocities. It is suggested that the Mars
Sample Return hardware could be modified
with limited effort to achieve a lunar sample
return mission that would both demonstrate
the requisite capabilities and also fulfil a
valuable scientific objective.