Publication Year
2016
Acceptance in OA@INAF
2020-06-03T14:13:16Z
Title
Small Mars satellite: a low-cost system for Mars exploration
Authors
Pasolini, Pietro; Aurigemma, Renato; Causa, Flavia; Dell'Aversana, Pasquale; de
la Torre Sangrà, David; et al.
Handle
http://hdl.handle.net/20.500.12386/25908
67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
IAC-16-A3.3A.6
SMALL MARS SATELLITE: A LOW-COST SYSTEM FOR MARS EXPLORATION
Pasolini P.*a, Aurigemma R.b, Causa F.a, Cimminiello N.b, de la Torre Sangrà D.c, Dell’Aversana P.d,
Esposito F.e, Fantino E.c, Gramiccia L.f, Grassi M.a, Lanzante G.a, Molfese C.e, Punzo F.d, Roma I.g,
Savino R.a, Zuppardi G.a
University of Naples “Federico II”, Naples (Italy)
b
Eurosoft srl, Naples (Italy)
c
Space Studies Institute of Catalonia (IEEC), Barcelona (Spain)
d
ALI S.c.a.r.l., Naples (Italy)
e
INAF - Astronomical Observatory of Capodimonte, Naples (Italy)
f
SRS E.D., Naples (Italy)
g
ESA European Space Agency, Nordwijk (The Netherlands)
a
* Corresponding Author
Abstract
The Small Mars Satellite (SMS) is a low-cost mission to Mars, currently under feasibility study funded by the
European Space Agency (ESA). The mission, whose estimated cost is within 120 MEuro, aims at delivering a small
Lander to Mars using an innovative deployable (umbrella-like) heat shield concept, known as IRENE (Italian ReEntry
NacellE), developed and patented by ALI S.c.a.r.l., which is also the project's prime contractor. The Lander includes
two small payloads, i.e., a dust particle analyzer and an aerial drone. The former is based on an instrument, developed
by the Astronomical Observatory of Capodimonte (INAF-OAC), performing in-situ measurements of the size
distribution and abundance of dust particles suspended in the Martian atmosphere. The drone is being designed by the
University of Naples and aims at demonstrating the feasibility of low-altitude flight in the Martian atmosphere. The
project also involves the Space Studies Institute of Catalonia (IEEC), responsible for launch and trajectory design. In
the paper, we illustrate the results of the feasibility study of SMS, including a description of the mission profile, launch
and escape phases, interplanetary trajectory, Mars approach, entry, descent and landing (EDL), and payload
deployment and operations. The current baseline envisages launching to LEO with VEGA. Then, a dedicated
propulsion module will provide a series of apogee raising maneuvers to place the vehicle on the hyperbolic trajectory
to Mars. A targeting maneuver, provided by a cruise stage, will direct the spacecraft to the atmospheric entry point
providing initial conditions suitable for the deployment of the heat shield. This will provide a ballistic coefficient much
lower than in previous Mars missions, thus allowing to reach subsonic conditions without the use of a supersonic
parachute. To demonstrate this, EDL and aero-thermo-dynamic analyses are performed with a 3-DoF model of the
entry trajectory and high fidelity CFD and DSMC analysis tools. Finally, particular attention is devoted to the
description of the deployable shield technology and verification.
Keywords: Mars Exploration, HDAD, Aerial Drone, DPA
Acronyms/Abbreviations
AD
Aerial Drone
CCD
Charge Coupled Device
COTS
Commercial Off The Shelf
CS
Cruise Stage
DHS
Deployable Heat Shield
DPA
Dust Particle Analyzer
EDL
Entry, Descent and Landing
ESA
European Space Agency
FS
Flexible Shield
HDAD
Hypersonic Deployable Aerodynamic
Decelerator
HIAD
Hypersonic Inflatable Aerodynamic
Decelerator
IRENE
Italian REentry NacellE
LEO
Low Earth Orbit
IAC-16- A3.3A.6
MEMS
OBDH
PCP
PM
SMS
SOI
TPS
TRL
UAV
GCS
Micro Electro-Mechanical Systems
On-Board Data Handling
Porkchop Plot
Propulsion Module
Small Mars Satellite
Sphere Of Influence
Thermal Protection System
Technology Readiness Level
Unmanned Aerial Vehicle
Ground Control Station
1. Introduction
Over the past 20 years, great effort has been put on
Mars exploration, with a primary aim at understanding
its surface composition and habitability potential.
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
Since 1976, several missions have successfully landed
on Mars: Viking 1 and 2 [1], Mars Pathfinder [2], the
two Mars Exploration Rovers [3], Phoenix [4] and
Mars Science Laboratory [5]. All Mars landers,
including those of the on-going ESA’s ExoMars
programme [6], are based on similar entry, descent and
landing (EDL) strategies, which employ a fixed
forebody for the lander. Only recent studies have
focused on deployable and inflatable landing systems.
Hypersonic Inflatable Aerodynamic Decelerator
(HIAD) [7-9] and Hypersonic Deployable
Aerodynamic
Decelerator
(HDAD)
[10-13]
technologies represent the new challenge in planetary
exploration.
The aim of this contribution is to present the Small
Mars Satellite (SMS) mission, whose primary goal is
to test a deployable shield for Mars atmospheric entry.
The shield shall provide both thermal protection and
deceleration through the thin Martian atmosphere.
SMS will also carry a technology payload, i.e., an
aerial drone (AD) for planetary surface exploration,
and a science payload, i.e., a dust particle analyzer
(DPA).
Accommodating the shield in folded configuration
at launch and deploying it at Mars entry, offers the
advantage of increasing the mass/volume ratio at
launch, thus widening the choice of the possible
launchers. Moreover, the shield deployment reduces
the ballistic coefficient to much lower values than
previous missions of the same category, thus allowing
a significantly higher deceleration in the upper and
more rarefied region of the atmosphere. This brings
lower thermal and dynamic loads, and the possibility
of reaching subsonic conditions without the use of a
supersonic parachute. These are the core
characteristics of SMS. The present work illustrates in
detail the choices and the methods that led to the
current system’s design.
In particular, Section 2 illustrates the mission and
the system outline, focusing on the system
configuration and the mission profile. Section 3 is
devoted to the description of launch, interplanetary
transfer and Mars approach. Section 4 describes the
EDL phase, followed by the illustration of the design
of the deployable shield. Sections 5 and 6 describe the
two payloads. Conclusions are drawn in Section 7.
Mission and system outline
Core characteristics of SMS are the deployable
shield, the low cost and the small size, features which
stand out with respect to previous Mars Landers. The
overall mission cost, estimated with state-of-the-art
CER models for small missions, is within 120 M€,
including launch and operations. Concerning the
deployable shield technology, an umbrella-like
mechanism will be tested to open a flexible shield
providing thermal protection and aerobraking
functions during Mars entry. The shield is retracted at
launch and during the interplanetary cruise. This
feature allows to adopt a small launcher, gives higher
flexibility in the shield design and sizing, and brings
advantages in the achievable ballistic coefficient at
planetary entry. Indeed, the ballistic coefficient can be
made sufficiently small for the deceleration through
the atmosphere to subsonic Mach numbers to be
achieved without a supersonic parachute, and at
altitudes higher than those of previous Mars missions.
This is discussed in detail in Section 4 which describes
the EDL phase results and the deployable shield
technology. Here, we wish to mention that a lower
ballistic coefficient allows to deliver payloads at
higher elevations on Mars surface, with the possibility
of reaching unexplored sites, of high interest for life
search [14].
As outlined in the previous section, SMS will carry
on board a DPA and an AD. These are designed so to
fit the limited power, mass and volume resources
available on board SMS (see sections 5 and 6).
Nevertheless, they can potentially improve Mars
atmosphere knowledge and modeling, as well as
demonstrate the viability of atmospheric flight on
Mars, issues of high interest to future Mars exploration
missions.
Fig. 1. Overall system configuration at launch (top)
and SMS with the shield opened (bottom).
2.
IAC-16- A3.3A.6
SMS will be launched to Mars by a Vega rocket on
a direct hyperbolic transfer. To this end, an additional
propulsion module (PM), acting as a fifth stage, shall
be introduced because Vega does not perform to Earth
escape. The propulsion module will separate soon after
executing the injection into interplanetary trajectory.
Fig. 1 illustrates the launch setup in the payload fairing
of Vega: SMS (with the shield in retracted
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
configuration) is integrated with the Cruise Stage (CS),
and the ensemble is installed on top of the fifth stage.
The CS provides electric power, telecommunications,
thermal control, and attitude determination and control
during the interplanetary transfer, and is jettisoned
shortly before Mars entry. On approaching Mars, the
CS applies a targeting maneuver to achieve the
required kinematics conditions at atmospheric entry
(i.e., a flight path angle between -14º and -12º).
To host the payloads and carry out the mission,
SMS exploits a modular architecture, which consists of
two main elements: the Lander, including the payload
and the avionics modules (see Fig. 2), and the
deployable shield (see Fig. 1). The payload module
(drone bay in the figure) hosts the AD in folded
configuration, preserving its integrity during all the
mission and allowing its release after landing. The
avionics module contains the DPA and all the
subsystems and components required to carry out the
mission and perform the payload operations on Mars.
A vented airbag is stowed in the hollow nose and in the
upper side of the avionics bay.
Fig. 2. SMS capsule configuration.
As outlined in Fig. 3, at Mars entry the umbrellalike mechanism unfolds the shield. Then, a subsonic
parachute is opened to further reduce the ground
impact speed to values compatible with a passive soft
landing system (i.e., the vented airbag). The
deployable shield is jettisoned soon after parachute
deployment. Once on ground, the cover of the Lander
unfolds (see section 6) exposing the AD. Operations on
the surface of Mars should last from a few days to a
few weeks.
As shown in Table 1, the wet mass is of 304.5 kg.
Upon separation of the CS, the mass of the system
decreases to about 200 kg. After parachute and shield
ejection, the mass further decreases to 132 kg. This
value is the mass at landing. About 15 kg of propellant
are needed for maneuvers during cruise.
IAC-16- A3.3A.6
Figure 3. EDL operation sequence.
Table 1. Overall system mass budget.
Unit
Mass(kg)
Cruise Stage
74.3
Lander
110
Deployable Heat Shield
57.1
System margin (20%)
48.3
Total dry mass
289.7
Total wet mass
304.5
3.
Launch, interplanetary transfer and Mars
approach
SMS will follow a direct trajectory from Earth to
Mars and shall enter the atmosphere of Mars from a
hyperbolic orbit relative to the planet. Environmental
considerations put restrictions on the choice of the
landing site and on the selection of the arrival date.
According to the temperature maps of Fig. 4, the
southern hemisphere is characterized by the warmest
temperatures, and this occurs in winter. However, the
Martian atmospheric conditions are severe at this
epoch because the heat transport causes strong air
currents and winds which raise the dust. The dust
circulation causes devils and may even result in global
storms. Therefore, landing close to the equator (the
mildest region on a yearly basis) before autumn is to
be preferred.
Following the algorithm proposed by [15], the
Earth-to-Mars trajectories have been computed by
solving the two-body Sun-spacecraft Lambert’s
problem between positions obtained by assuming that
the planets are massless bodies revolving on secularlyprecessing Keplerian ellipses.
The transfer opportunities have been determined
with a time resolution of one day at departure and at
arrival. These opportunities are represented in the form
of porkchop plots (PCPs), and their overall cost (sum
of departure and arrival ΔV’s) exhibits a typical pattern
with a periodicity of two years approximately,
resembling the synodical period of the two planets
(2.14 years). Hence, three PCPs are available for the
interval 2020-2024, centered in 2020, 2022 and 2024,
respectively. The opportunities with the minimum cost
at departure are those of the year 2024, in particular
those corresponding to type II trajectories (i.e.,
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
trajectories with transfer angles larger than 180º) (see
Fig. 5).
Fig. 6. Heliocentric plot of the baseline Earth-to-Mars
trajectory (details in the text). The positions of Earth
at arrival and Mars at departure are also indicated.
Fig. 4 - Daily average maximum (top) and minimum
(bottom) atmospheric temperatures close to the
ground over one Martian year and with respect to
geographical latitude. The lowest minima are close to
-130ºC, whereas the highest maxima are at the level
of 20ºC.
Fig. 5. Porkchop plots for the 2024 launch
opportunities: colour maps of departure C3 (km2/s2
top) and arrival v∞ (km/s, bottom).
IAC-16- A3.3A.6
The lowest cost at departure corresponds to a 334days trajectory leaving the Earth on 2 October 2024
and reaching Mars on 1 September 2025 (see Fig. 6).
This solution has an Earth C3 (the square of the
hyperbolic excess speed v∞) of 11.3 km2/s2 and a Mars
v∞ of 2.5 km/s. The arrival date is compliant with the
requirement of landing before the local autumnal
equinox.
The C3 level of the trajectory sets the requirements
on the wet mass of the spacecraft, in a way which
depends on the performance of the launcher. Among
the three European vehicles, i.e., Ariane 5, Soyuz and
Vega, Vega is the cheapest and the smallest, in other
words the most suitable for a small-class mission.
Unfortunately, Vega by itself launches to
geocentric orbit (C3 < 0) only. The performance of
Vega can be enhanced to positive values of C3 by an
additional ad hoc stage assembled with the spacecraft
and stowed in the payload fairing. In particular,
according to a recent study [16], a particular bipropellant engine (i.e., the PM) provides the
performance shown in Fig. 7.
Following insertion into a 300-km LEO lowinclination orbit and separation from the launcher, the
PM can inject a spacecraft of the mass indicated in the
y-axis, given the C3 (x-coordinate) and one of two
options for the size of the PM’s propellant tank, i.e.,
the short (yellow line) or the long (red line). The Lisa
Pathfinder spacecraft [17] was launched to the L1 point
of the Sun-Earth system with a version of the PM: upon
separation from the fourth stage of the rocket, the PM
executed a series of apogee raising maneuvers until
escape was achieved. The baseline launch profile that
has been conceived for SMS is very similar to this.
Adoption of the long propellant tank allows to launch
320 kg of wet mass, margins included. The apogee
raising sequence shown in Fig. 8 is just an exercise
based on the example of Lisa Pathfinder.
Modifications shall be made based on communications
and power requirements. However, although in a
preliminary and simplified way, this apogee raising
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
sequence demonstrates that SMS can be injected into
the right hyperbolic trajectory by means of six perigee
burns of 0.4249 km/s, followed by a final burn of
0.8297 km/s of 22 minutes duration. The maneuvers
performed by Lisa Pathfinder had similar magnitudes
and duration.
Fig. 7. Performance of Vega (launch mass, kg) with
the PM endowed with either of two versions of the
propellant tank (long in red, short in yellow).
4.
The EDL phase and the deployable thermal
protection system (TPS) concept layout
Here we define the SMS capsule’s entry trajectory
and we make a preliminary assessment of the shield
design. Velocity, Mach number and pressure profiles
have been computed over the entry trajectory using 3DOF models and hypersonic Newton theory for entry
trajectories evaluation.
dV
dt
d
V
dt
d
V
dt
dH
dt
d
dt
d
dt
V 2
g sin
2
M2 r cos sin cos cos sin sin
V2
g cos 2M cos cos
r
M2 r cos cos cos sin sin sin
(1)
V2
cos cos tan
r
2M sin tan sin cos M2 r
cos sin cos
cos
V sin
V
cos sin
r
V cos cos
r
cos
q0 1.83 104 V 3
p0 V 2
Fig. 8. A possible Earth escape sequence through
seven consecutive perigee burns executed by the PM.
The selected interplanetary trajectory targets the
center of Mars in a two-body Sun-spacecraft model.
The gravitational influence of the planet gradually
takes over as SMS approaches its destination, and,
according to the patched conics model, dominates the
dynamics of the spacecraft once an ideal border, i.e.,
the surface of the sphere of influence (SOI, 580000 km
radius) has been crossed. The resulting Mars-centered
hyperbola must be appropriately corrected in order to
satisfy the aerodynamic requirement that the flight
path angle be in the range [-14º, -12º] at an altitude of
125 km, representing the upper limit of the Martian
atmosphere. The correction maneuver has been
modelled as a small impulse (33 m/s) applied by the
onboard propulsion system when crossing the surface
of the SOI. Such impulse makes SMS enter the
atmosphere at a speed of 5.5 km/s and land close to the
equator, as desired.
IAC-16- A3.3A.6
(2)
Rc
(3)
In Equation 1 t is the time, V the velocity, H the
altitude, γ the flight path angle, β the ballistic
coefficient, g the gravitational acceleration, r the
radius of curvature of the trajectory, ψ the azimuth
angle, λ the latitude, Λ the longitude and ωM Mars
angular rotation speed. Heat fluxes (𝑞̇ 0 ) have been
estimated along the trajectories using engineering
formulations (eqn. 2) as a function of the speed V, the
Mars atmospheric density ρ and the capsule nose
radius R0 [13]. Equation 3 describes the hypersonic
Newton theory where p0 is the stagnation point
pressure, V the velocity and ρ the Mars atmospheric
density.
Due to the low atmospheric density of Mars, all
past Mars landers have employed complicated and
expensive EDL strategies in order to safely touch down
on the surface [14]. In these designs, the entry capsule
is characterized by a fixed forebody heat shield which
protects the lander in the high aerodynamic heating
portion of the atmospheric entry. Then, when the
capsule reaches supersonic speeds, a parachute system
is deployed to reduce the speed. Once the lander has
slowed down to subsonic speeds, the heat shield is
jettisoned and a subsonic parachute is deployed.
Different touch down methods, active and passive,
using thrusters and airbags respectively, are employed.
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The advantage of using a deployable umbrella-like
configuration is the small attainable ballistic
coefficient, which can be made lower than 20 kg/m2,
less than one third the ballistic coefficient of any
previous Mars lander [18]. Compared with the EDL
methods adopted by past missions, several advantages
are brought by the decrease in the ballistic coefficient.
First of all, the aerodynamic and aero-thermo-dynamic
loads along the hypersonic entry flight path are much
reduced. Besides, a single subsonic parachute may be
sufficient since the capsule decelerates to subsonic
speeds at higher altitudes. Finally, passive landing
systems, such as airbags or crushing systems, can be
used to absorb the impact energy.
A parametric analysis based on varying the ballistic
coefficient and the flight path angle has been carried
out in order to assess the maximum design parameters
(Fig. 9). The simulations have been performed with
actual values for the entry mass (200 kg) and the shield
diameter (3.11 m), and assuming a drag coefficient Cd
equal to 1. Special relevance has been given to the
results corresponding to β = 21 kg/m2 and FPA = -13°.
The relative initial entry velocity and initial entry
altitude have been set to 5.5 km/s and 125 km,
respectively (see Section 3).
Fig. 9. Landing speed (top-left), stagnation point heat
flux (top-right) and pressure (bottom) as functions of
ballistic coefficient (x-axis) and entry FPA (y-axis).
Figure 10 shows a comparison between the velocity
profiles of previous Mars entry trajectories [19] and the
nominal SMS entry path. The lower ballistic
coefficient achieved by SMS causes lower stress in
terms of aerodynamic and aero-thermo-dynamic loads.
As shown in Fig. 10, one of the most important
results achieved with the deployable shield technology
is the possibility of reaching the subsonic regime at
altitudes of approximately 15 km without sophisticated
deceleration systems (such as supersonic parachutes or
thrusters). Hence, the SMS technology is extremely
flexible and the only low-cost system capable of
IAC-16- A3.3A.6
landing at high altitudes (i.e., sites with elevation
higher than MOLA 0).
Fig. 10. Speed, Mach number, stagnation point heat
flux and pressure profiles for SMS and three previous
Mars landers, i.e., Mars Viking, Mars Pathfinder,
Phoenix.
A parachute analysis has been carried out to assess
the performance of SMS in terms of landing capability
using a single subsonic parachute. In order to simulate
the shield ejection in the entry trajectory after the
parachute deployment, an analysis of the variation of
the landing speed as a function of the Lander mass has
been performed. For this analysis, four alternative
values for the diameter have been considered (Fig. 11).
Fig. 11. Landing speed as a function of capsule mass and
parachute diameter.
Finally, in order to evaluate the aerodynamic and
thermal field around the entry capsule and particularly
on the flexible and rigid elements of the thermal
protection system, an aero-thermo-dynamic analysis
on the selected capsule configuration has been
executed. Depending on the flow regime, different
numerical state-of-the-art codes have been applied. In
continuous regime, the solution of the classical NavierStokes and energy equations has been obtained with
the commercial code STAR-CCM+ [20]. The flow
field around the capsule has been assumed laminar.
Due to its chemical composition (95% carbon dioxide),
the Mars atmosphere has been treated as a singlecomponent ideal gas. The numerical simulations have
Page 6 of 11
67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
been performed with a density-based, time implicit
numerical solution scheme through a control volumebased technique. The AUSM (Advanced Upstream
Splitting Method) scheme has been employed for
convective numerical fluxes [21, 22]. In a rarefied
regime, the study of the aerodynamic characteristics
has required Direct Simulation Monte Carlo (DSMC)
methods [23]. In particular, two-dimensional
axisymmetric analyses have allowed to estimate the
distribution of the thermal and mechanical loads on the
surface of the capsule, and to compare the results
obtained in the EDL phase. Three-dimensional
analyses have allowed to calculate the values of
temperature and pressure distributions on the threedimensional geometry of the capsule at the most severe
conditions along the entry trajectory.
made up of a sliding structure, tensioning poles and
threads. Figure 14 shows the shield in retracted (top)
and deployed (bottom) configuration. A rigid nose
cone and the associated support structure are also part
of the TPS.
Fig. 13. Pressure distribution at maximum stagnation point
pressure at 40 km (left) and temperature distribution at
maximum stagnation point heat flux condition at 50 km
(right).
Fig. 12. CFD, DSMC and hypersonic Newtonian theory
(top) and Tauber/Sutton theory (bottom).
Fig. 12 shows good agreement in terms of
stagnation point pressure prediction, meaning that the
Newtonian theory is a valuable tool to predict the
stagnation point pressure for a blunt body in a
hypersonic flow field. It is known that the correct
evaluation of the stagnation point heat flux over the
planetary entry trajectory is a very difficult task to
accomplish. According to Fig. 12, the more rarefied
the flow regime, the closer the DSMC results are to the
curve predicted by the theory. On the other hand, as the
altitude decreases, the CFD code results get closer to
the Tauber/Sutton equation results.
Fig. 13 illustrates the pressure distribution at
maximum stagnation point conditions and the
temperature distribution at maximum stagnation point
heat flux conditions through the capsule entry
trajectory.
The main component of the TPS is a flexible shield
(FS) whose deployment mechanism is essentially
IAC-16- A3.3A.6
Figure 14. Shield layout in retracted (top) and
deployed (bottom) configurations.
The shield is made of woven ceramic fabrics,
which retain strength and flexibility with little
shrinkage at temperatures lower than 1100 °C. The
shield consists of two 0.38-mm layers, each composed
by twelve trapezoidal patches reinforced at the edges
(Fig. 15). When deployed, the shield approximates a
45° conic surface, with a maximum diameter of 3.11 m
and a maximum working temperature of 1300 °C. The
fabric layers are put in tension by 12 poles and are
clamped at the nose cone support structure.
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Fig. 16. Stress map on the Nextel Fabric during entry.
Fig. 15. Layout of the flexible shield.
In order to reduce the overall system’s mass, a
hollow configuration has been used for the nose cone,
which offers also the possibility of accommodating the
airbag. To increase the nose cone strength, a doublelayer solution has been adopted in which the external
layer is made of RESCOR 310M (a silica foam) with a
thickness of 4 mm and a maximum working
temperature of 1650 °C, and the inner layer is made of
a FW12 Oxide/Oxide ceramic matrix composite with a
thickness of 2 mm and a maximum working
temperature of 1300 °C. Although the temperature
reached at the nose during entry is equal to the
maximum working temperature of FW12, the
temperature at the inner layer is lower due to the 4 mm
RESCOR layer. The deployment mechanism, made of
titanium alloy Ti6Al-4V, is composed by a sliding
structure consisting of two rings, connected by four
rods, sliding along the Lander. The twelve tensioning
poles (50 mm diameter, 1 mm thickness) are connected
to the upper ring by means of cylindrical hinges, while
the lower ring supports one end of the twelve lower
threads (1.5 mm diameter). The lower and upper
threads (4 mm diameter) ensure fabric stabilization in
the tensioning phase before the atmospheric entry.
Indeed, under pressure loads the lower threads lose
tension, while the upper threads support the fabric in
sustaining the loads. Two electric actuators complete
the deployment system, which allow the displacement
of the sliding structure during the opening phase. The
shield design has been verified by using a nonlinear FE
model of the flexible shield and of the main structural
parts of the deployment mechanism, considering load
conditions deriving from the tensioning phase and
entry phase (pressure loads). A pressure value of 4.4
kPa on the nose and 3.2 kPa on the fabric have been
used, as predicted by the aero-thermodynamic analysis.
For the flexible shield and the deployment mechanism,
both the pressure and the tensioning loads have been
simulated to test the design. Fig. 16 shows the stress
distribution of the Nextel fabric during entry: the stress
level is lower than the allowable limit (40 MPa) except
in very small areas (i.e., black areas in Fig. 16).
IAC-16- A3.3A.6
5.
The DPA
The monitoring of airborne dust is of great interest
in the study of Martian climatology because dust
affects the atmospheric thermal structure, balance and
dynamics by absorbing and re-emitting solar and
thermal radiation.
The influence of dust is relevant, even with
moderate presence of dust in the atmosphere, but
during regional or global dust storms more than 80%
of the incoming sunlight is absorbed by the dust
causing intense atmospheric heating. Airborne dust is
therefore a crucial climate component on Mars. The
main dust parameters are the size distribution and
abundance. Moreover, wind and windblown dust
represent nowadays the most active processes having
long-term effects on the Martian geology and on the
morphological evolution, such as aeolian erosion, dust
redistribution on the surface and weathering.
The size distribution is so far measured indirectly
through remote sensing data, but these measurements
exhibit poor vertical resolution, since they are related
to the entire atmospheric column, and give little insight
into the properties of the atmospheric layer close to the
surface, where phenomena of dust lifting are present.
The DPA selected for SMS will be able to perform,
actually for the first time, measurements of dust
concentration directly on the Martian soil. The DPA is
based on an optical particle counter which processes
the light scattered from the single dust particle. In this
way, the instrument measures the size and estimates
the abundance of each grain. As a matter of fact,
according to Mie theory, the intensity of the scattered
light is directly related to the grain size. The DPA is
endowed with a pump to obtain a flow of gas with
embedded dust flowing into the instrument, down to an
area where a laser light is focused and the scattered
light is collected by a mirror and finally sent to a
detector. Eventually, the flow of air and dust is ejected
back into the atmosphere. The light which is not
scattered is captured by a light trap to avoid stray light
inside the instrument. The grain radii which can be
measured range from 0.2 to 10 µm, with a factor of
coincidence lower than 4%. Once the particle counting
and size measurements have been performed, the dust
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
number density can be estimated since the volume
sampled by the instrument is known.
The current design of the DPA is derived from the
MEDUSA instrument, selected for the ExoMars
Humboldt payload which reached a TRL of 5.2 in 2009
[24]. Since the mass and power budgets are very
different from those of the ExoMars mission, an
optimized DPA design was needed to fulfill the more
demanding requirements. The DPA has now a total
mass of 600 g (including a margin of 20%) and
requires about 3 W of power, instead of the 3 kg and
21 W of the MEDUSA instrument.
A DPA laboratory breadboard (Fig. 17) was
developed with funds of the Italian Space Agency in
2011 and thanks to INAF financial resources
afterwards. The breadboard was successfully tested
demonstrating the expected performances. The
estimated TRL is between 4 and 5.
Fig. 17. DPA laboratory breadboard.
The DPA will be accommodated preferably in a
fixed position on the lander. The possibility of
embarking the DPA on the drone was discarded
because this solution is more complex and demanding,
and does not ensure a significant increase in the
scientific return. According to the current baseline
design, the DPA will be activated after landing and
shall execute at least four runs per Sol. Further
activations are desired, but they depend on the
availability of energy and data link resources. Longer
and continuous acquisitions from the DPA can allow
the detection of dust devils. Also the possibility to
trigger the DPA acquisition by means of other sensors
(e.g., pressure sensors) can highlight this kind of
phenomena. The possibility to perform measurements
with the DPA during the descent phase of SMS is also
of great interest and will be studied in the next stages
of the project. The DPA interfaces directly with the
On-Board Data Handling (OBDH) by means of power
supply and serial data interfaces.
The OBDH will activate the DPA according to a
predefined mission timeline resident in the OBDH
memory, and will acquire the data from the DPA after
IAC-16- A3.3A.6
each measurement run. Scientific data can either be in
the form of raw data, i.e., the complete waveform
corresponding to the single event (dust grain), or can
be transmitted in the form of histograms, an option
which reduces the data volume.
6.
The AD
Most of Mars surface exploration has been
performed by means of surface rovers [25]. Rovers are
able to collect data near the surface, but their
exploration capability is limited by the obstacles of the
terrain. To overcome these limitations, SMS carries an
aerial drone.
The first idea of sending an Unmanned Aerial
Vehicle (UAV) to Mars dates back to 1978 with the
Mini-sniffer [27-29]. Since then, many UAV proposals
have been developed including airship/balloons [30],
aircraft [25, 32-34] and rotorcraft [35-38].
Among the several concepts, a two-blade coaxial
rotorcraft appears the best solution in terms of
deployment complexity and ability to perform multiple
controlled flights while acquiring high-resolution
images. In addition, this solution is compliant with the
strict mass and volume requirements of the SMS
mission.
The atmospheric density on Mars is about 100
times lower than on Earth, thus making the design of a
Mars rotorcraft very challenging. This type of vehicle
needs high blade velocity and/or radius to generate
sufficient lift [25]. Furthermore, the blade design must
avoid transonic regimes and Reynold numbers lower
than 5104 [35].
With this in mind, the preliminary sizing of the
drone has been carried out so as to fit the space and
mass available in the payload module. This has
resulted in a 7-kg rotorcraft, with a rotor radius of 1.25
m. The main characteristic of the drone are
summarized in Table 2.
The drone is folded inside the Lander’s payload
bay during launch, interplanetary flight and EDL, and
is exposed by the Lander cover opening after landing.
To fulfill the high-resolution imaging goal, the
drone is equipped with a COTS CCD camera, that can
provide centimeter-level resolution images of the
Martian surface [39]. The camera is used also as aiding
sensor for navigation purposes, providing optical flow.
Low-weight MEMS devices compose the drone
avionics. They shall guarantee autonomous navigation
and control on Mars. The drone is assumed to fly a
loiter path (320 m radius, 100 m altitude, 2 km ground
track) around the Lander that acts as GCS and drone
landing site for battery recharging. Thus, during the
flight the drone always keeps the GCS in line of sight.
The multi-mission capability is ensured by four LiPo
rechargeable batteries [25], which allow 7-min drone
autonomous flight. The batteries are recharged in 1
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67th International Astronautical Congress (IAC), Guadalajara, Mexico, 26-30 September 2016.
Copyright ©2016 by the International Astronautical Federation (IAF). All rights reserved.
SOL with the drone standing on a charging pad [40]
embedded in the Lander cover (see Fig. 18).
Table 2. AD Characteristic
Parameter
N. of rotors and blades (per
rotor)
Cruise altitude (m)
Forward velocity (m/s)
Blade radius (m)
Main body volume (m3)
Total power/Energy (W/Wh)
Value
2
100
11.5
1.25
0.15×0.15×0.15
410/60
Fig. 18. Drone lading site on the unfolded lander. A
charging pad is embedded in the Lander cover.
7.
Conclusions
The main results of the phase 0 study of a small,
low-cost system for Mars exploration have been
presented. The core features of the proposed system
are a flexible shield that can be deployed at Mars
arrival using an umbrella-like mechanism (providing
thermal protection and aerobraking functions), a small
scientific payload, for atmospheric dust analysis, and a
small aerial drone aiming at demonstrating
atmospheric flight on Mars.
Our preliminary analyses have proven the
capability of the innovative shield concept to limit
thermal and pressure loads during re-entry, and to
reduce the ballistic coefficient to much lower values
than all previous Mars missions. In this way, the
capsule decelerates to subsonic speeds in the higher
atmosphere, without the aid of a supersonic parachute.
In the paper, the overall system configuration and
mission profile have been presented. A peculiar feature
of the system is the possibility of using a small
launcher, such as Vega, provided its performance is
extended to Earth escape by the addition of an ad hoc
fifth stage.
Finally, the preliminary design of the two payloads
have demonstrated their compliance with such a small
system, opening the way to achieving scientific and
technological results of great value for future Mars
exploration missions.
Acknowledgements
IAC-16- A3.3A.6
The study has been carried out under ESA Contract
No. 4000115306/15/NL/LF/as (11/12/2015).
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