SSAE 99
Stuttgart, November 17 - 19, 1999
AEROTHERMODYNAMICS DEVELOPMENTS FOR PLANETARY MISSIONS
Ph. Reynier1, V. Marieu, L. Marraffa 2
ESA/ESTEC, Aerothermodynamics and Propulsion Division,
Keplerlaan 1, 2200 AG Noordwijk AZ, The Netherlands
1
+ 31 71 565 3161, preynier@estec.esa.nl
2
+ 31 71 565 3907, lmarraff@estec.esa.nl
ABSTRACT
This paper focuses on the aerothermodynamics activities performed at the European Space Agency to prepare sample
return missions to the inner planets of the solar system. In particular, predictions of trajectories, numerical simulations
of the entry in the atmosphere of the explored planet and reentry in Earth atmosphere are discussed. In the framework
of this contribution, computations of Mars entry have been performed and the results compared to those obtained using
another approach. Numerical simulations with and without ablation of a small capsule entry into Earth atmosphere
have been performed for Mercury sample return mission. A connected domain is the study of flows in hypersonic
facilities in order to investigate their capabilities for experimental validation of the numerical results. With this aim in
view, SCIROCCO capabilities have been studied for entry simulations in CO 2 atmospheres.
1. Introduction
In the framework of planetary exploration, the
European Space Agency (ESA) has developed several
studies like Intermarsnet, Marsnet and missions like
Huygens (see Marraffa et al. [1]), involving planetary
entries. ESA has also studied sample return missions to
comets (ROSETTA/CNSR), Venus and Mercury [2, 3].
These past activities benefited from the technological
developments in the wake of the HERMES project.
They induced also specific research in the field of
aerochemistry [4],[5],[6]. In particular in the frame of
ESA’s Technological Research and Development
Programs (TRP), one has been devoted to
aerochemistry and hypersonic flows [7]. The function
of these TRP is to link the fundamental research and
the technological studies. They prepare the ground, at
assessment study and phase A level, for the use of new
data, facilities and technologies in future projects. The
European interest in new space missions triggered the
development of new facilities, databases and computer
programs or the adaptation of existing ones.
The paper focuses on the different aerothermodynamics
activities developed at the European Space Agency to
prepare sample return missions. For Venus the scenario
is complex. A sample return mission to Venus needs
two Ariane 5 launches. The first launch brings to Venus
an orbiter which contains communication capabilities
and an Earth return vehicle. The orbiter stays in
circular orbit around Venus. The second launch brings
to Venus a lander which is slowed down by
aerobraking and then by a parachute. At the surface,
samples are collected and measurements performed.
Then, a balloon takes off with the module and in
altitude, a rocket containing the samples is launched.
The rocket performs a rendezvous with the orbiter, and
finally a return capsule brings the samples back to
Earth. In the case of Venus and Mars, these missions
are complex, they include the entry of a probe in the
CO2 atmosphere of the explored planet and the reentry
of a capsule containing the samples in the Earth
atmosphere. Concerning Earth entry, the presence of
high thermal fluxes induces severe constraints for the
design of the thermal protection system. For a good
prediction of fluxes, some phenomena like nonequilibrium, radiation and ablation of the shield have to
be considered. To ensure the mission success,
simulations of the entry using software and ground tests
are necessary. However, before using a ground test
facility with CO2, preliminary investigations are
required to identify the necessary hardware
modifications, to evaluate its performance envelope
and to ensure its safety during operations.
2.5 - 1 -
functions of temperature. To create a thermodynamic
database for CO2 valid from 300 to 15000 K, the
database of TINA is used as much as possible. When
data were absent from TINA database, those of Esch et
al. [12] have been employed.
2. Scirocco capabilities
This part concerns the investigation on SCIROCCO
(see Figure 1) plasma wind-tunnel capabilities to
simulate a Mars or Venus entry. The first task was to
investigate the species concentration at equilibrium for
a large range of pressure and temperature in the arcchamber. The chemical composition of the mixture has
an incidence on the evolution of the thermodynamic
variables along the nozzle. In addition, a reacting flow
of CO2 could produce toxic chemical species like CO.
Figure 1: Scheme of SCIROCCO high-enthalpy wind
tunnel [8].
2.1 Thermodynamic database
Before performing numerical simulations of the flow
along the nozzle of the facility, the first task is to
develop a thermochemical model for the CO2. Two
software have been selected to investigate the mixture
composition in function of temperature and pressure: a
NASA code developed by Gordon & McBride [9], and
Chemkin [10]. As the temperature encountered during
an atmospheric entry can reach 15000 K, Chemkin and
Gordon & McBride thermodynamic databases are
inadequate. Hence, two other databases have been
tested, one incorporated in the software TINA [11] and
the one proposed by Esch et al. [12]. The database of
Esch et al. [12] is valid until 15000 K. Using these
different databases, the specific heat at constant
pressure Cp, the enthalpy H and the entropy S are
studied for gaseous carbon. Figure 2 shows the curve
fits of these quantities for temperatures between 300
and 15000 K. TINA and Chemkin uses the same
coefficients from 300 to 5000K. This is not the case for
higher temperatures. The database of Esch et al. [12] is
valid until 15000 K. In figure 2, the curves show that
the three databases are compatible between 300 and
5000 K. From 5000 to 15000 K the results obtained
with TINA and Esch et al. [12] databases are in good
agreement. For temperatures above 15000 K, TINA has
coefficients which keep the specific heat constant and
which interpolate enthalpy and entropy as linear
Figure 2: Curve fits of gaseous carbon, using Chemkin,
TINA and Esch database [12]. Cp: specific heat at
constant pressure, H: enthalpy, S: entropy, R:
universal gas constant.
2.2 Approaches used for the predictions
Two methods have been chosen to predict the flow
behaviour along the nozzle. The non-equilibrium flow
solver, TINA, developed by Netterfield et al. [13] and
Roberts et al. [11], is used to perform axisymmetric
simulations. However, a complete investigation on the
arc-heater envelope of SCIROCCO with a NavierStokes solver has not been performed, due to the cost
of such calculations. For this reason, a quasi onedimensional isentropic approach at equilibrium has
been developed. Such an approach does not require
much computer time and allows an extensive
investigation of the arc-heater envelope. This method
described in [14] is coupled with Gordon & McBride
subroutines, for the determination of the equilibrium
composition. The equilibrium approach is valid when
the chemical reaction time is short in comparison with
the time needed by the flow to go through the nozzle.
2.5 -2-
Therefore, equilibrium calculations are not adapted to
high velocities and low pressures. A solution to
improve the reliability of the numerical results is to
freeze the chemistry above a given Mach number. To
achieve this, the isentropic relations for a perfect gas
have been implemented in the code: The chemical
reactions are arbitrarily frozen when the Mach number
reaches the value of 0.7.
composition are compared to those predicted by TINA.
As there are no experimental data available for
SCIROCCO yet, these comparisons provide a first
validation of the numerical results. They also give an
estimate of the non-equilibrium effects on the flow
behaviour. The computed case corresponds to the point
of the envelope (see Figure 4) where the enthalpy is 10
MJ, and the pressure 5 atmospheres. Figure 5 shows the
axial distributions of temperature and velocity, and
Figure 6 of Mach number and density. The curves
presented in these figures bring out the existence of two
regions. Upstream of the throat, the results of the
equilibrium approach are in agreement with those of
TINA for all the quantities. Downstream, the various
predictions show some discrepancies. In this second
region of the flow the differences between the results
obtained with equilibrium and non-equilibrium
predictions are particularly obvious for the temperature.
These results have been obtained for a low pressure and
a high enthalpy in the reservoir. Here, the chemical
reaction time is long compared to the time needed by
the flow to cross the nozzle. The equilibrium
assumption is not any longer valid and the gas can be
considered as frozen. However, even with the frozen
model there is still a large difference between nonequilibrium and equilibrium results.
Figure 3: SCIROCCO nozzle mesh.
Figure 4: SCIROCCO envelope and performance map
for 75 mm throat nozzle. Fluxes calculated at
stagnation point of MARS miniprobe nose ([15]). The
nose radius is 300 mm. Initial composition: 95% CO2,
5% N2.
To take into account the thermochemical nonequilibrium of the flow, numerical simulations have
been performed using TINA. For the numerical
simulations, a five species model CO2, CO, C, O and
O2, which does not generate ions and electrons, is used.
The characteristics of SCIROCCO nozzle geometry
have been given by CIRA [4]. The geometry and the
mesh are shown in Figure 3. The geometry extents from
the reservoir to the test chamber. The mesh shown in
Figure 3 has 139 cells in the axial direction and 44 cells
in the radial direction. The length of a cell at the throat
is 1 mm and its thickness at the wall is 0.1 mm. Then,
the cell size is increased geometrically in each
direction. More refined meshes have been used and the
simulations showed the independence of the results
from the mesh. The flow is axi-symmetric, therefore a
symmetry boundary condition is imposed on the axis.
On the wall, non-catalytic viscous conditions have been
chosen. At the outflow, Neumann conditions are
applied on the mean quantities. More details about
computation conditions like the flow initialisation are
available in [14], [16].
Figure 5: Comparisons between by TINA and quasi
one-dimensional approach at equilibrium: temperature
and velocity axial distributions. Dotted lines
correspond to frozen flow calculations.
To draw the performance map of the facility, the initial
composition has been fixed at 95 % of CO2 and 5 % of
N2. This is an average composition of Venus or Mars
atmospheres. The chemical model used for the
prediction at equilibrium accounts for ionisation. Since
no envelope of the arc-heater is available for CO2, in a
first approximation, the calculations have been
performed with the envelope for air. A study of the
electrical conductivity and radiation losses in the arc
chamber for a CO2 mixture would present a high
interest. It would allow the determination of the energy
that the electric arc could transfer to the plasma, and
therefore a better prediction of the stagnation
2.3 Results
First, the results predicted by the quasi one-dimensional
approach with and without freezing of the mixture
2.5 -3-
conditions. Then, it would be possible to draw the archeater envelope for CO2.
break the molecule of CO2 but not to split it completely
into atoms. For higher enthalpy (45 MJ/kg), the level of
CO decreases, this means that there the enthalpy is high
enough to decompose the carbon dioxide and monoxide
in a mixture of elementary elements. As a consequence,
the carbon monoxide reaches a high level for the mid
part of the enthalpy range. Of course, the prediction of
the mass fraction of carbon monoxide depends on the
chemical model and another one could lead to a
different conclusion. Therefore, subject to the
verification of the chemical model, the level of CO is
high and a reprocessing might be required in the
facility.
Figure 6: Comparisons between TINA and the quasi
one-dimensional approach at equilibrium: Mach
number and density axial distributions. Dotted lines
correspond to frozen flow calculations.
Enthalpy
(MJ/kg)
2
22.5
45
45
25
10.5
10
2
The performance map of SCIROCCO and the
corresponding arc-heater envelope are shown in Figure
4. The performance map has been computed with the
quasi one-dimensional method (without freezing) and
with TINA. Since the computations with TINA take a
lot of computer time, less points have been predicted
than with the quasi one-dimensional approach. To
compute the convective flux, represented in figure 4,
the relation used by Rubio García et al. [15] to estimate
the flux at the stagnation point of Mars miniprobes has
been selected. This relation is given by:
Qconv  1.156 10 5 ( H 0
 Hw )
Pressure
(atm)
1
1
1
1.5
7.5
16.5
4.5
10
Mass-flowrate (kg/s)
0.752
0.171
0.137
0.09
0.684
4.59
1.14
7.47
CO (massfraction)
0.5 10-3
0.58
0.12
0.11
0.58
0.33
0.29
0.16 10-2
Table 1: Mass fraction of CO and mass-flow-rate for
the computed points of SCIROCCO envelope.
3. Aerocapture and aerobraking
To achieve landing on a massive planet surrounded by
an atmosphere, direct high-speed entry should be
avoided. One option is to perform a first partial entry,
transforming the initial hyperbolic trajectory into an
elliptic trajectory around the planet: namely, to perform
aerocapture. The resulting orbit may still be a high
energy one and require further passes into the upper
atmospheric layers, to erode it.
Such a technique has already been successfully applied
to the Russian mission Zond [17] around the Earth,
bringing the velocity of a spacecraft returning from the
Moon from 11 km/s to 7.7 km/s. It has been
demonstrated [18] that aerocapture would provide
important mass savings for large planetary missions,
such as Venus or Mars sample return. The reason is
obvious, when looking at the heat flux evolutions
plotted in Figure 7 for orbital Venus entry and in
Figure 8 for direct entry.
New problems arise, related to guidance, navigation
and control, configuration, thermal control, operations,
atmospheric sciences but also to aerothermodynamics.
The purpose of aerocapture and aerobraking is to
decrease entry loads by dividing the severe direct entry
into multiple, milder decelerations. This enables the use
of lighter materials and concepts for the heat shield.
However, the resulting trajectories are quite different
from those studied so far, and maximum savings can
only be achieved with a very accurate determination of
U2
5
1.013 10 Rn
where, Qconv is the convective heat flux, H0 the total
enthalpy, Hw the enthalpy at the wall, ρ the density, U
the velocity and Rn the probe nose radius.
The performance map obtained has the same shape as
the one for air obtained by CIRA [4]. Since a flow of
CO2 is studied here, the flux and pressure values are
different. The calculations made with the quasi onedimensional program and with TINA gave similar
results, except for low pressure where the effects of
non-equilibrium are more important.
The control of the level of CO during the experiments
is important for the safety of the test. A release of this
gas in large quantities in the atmosphere is not
desirable for the environment. Furthermore, due to its
toxicity a high level of carbon monoxide would be a
danger for the operating staff. The various calculations
performed to draw the performance map of
SCIROCCO allow the prediction of CO mass-fraction
at the nozzle outlet. The results obtained on the nozzle
axis are presented in Table 1. They show a large massfraction of CO (until 58 %) for the points of the
envelope characterised by a medium enthalpy (10 to 25
MJ/kg). In this region, the enthalpy is high enough to
2.5 -4-
the spacecraft environment, in exotic atmospheres and
conditions. In particular, relatively lower entry
velocities for Venus (or Earth) may lead to more
important non-equilibrium effects in radiation.
Figure 9: Comparison between various radiative flux
correlation’s for Venus orbital entry (from [18]).
5. Mars entry
In the framework of the MARS EXPRESS project,
entry computations of a probe in a CO2 atmosphere
have been performed with TINA. This work takes place
in a study on the thermochemical simulation at the
stagnation point for Mars probe made by Kolesnikov &
Marraffa [20].
Figure 7: Venus entry trajectory from orbit (from [18]).
TINA calculations have been performed to verify the
results given by IPM simulations. IPM calculations rely
on the assumption that the flow is at equilibrium except
in the boundary layer.
Case 1
Case 2
Nose Radius (m)
0.8
1.25
P0 (Pa)
5.2
8.27
T0 (K)
147
150
Figure 8: Direct Venus entry trajectory (from [18]).
U0 (m/s)
5524
5163
4. Radiation
Table 2: Nose radii and trajectory characteristics for
the entry of a MARS EXPRESS probe.
Radiation heat fluxes are the dominant contribution to
the total heat flux reaching the heat shield of a
spacecraft entering Venus atmosphere, or Earth
atmosphere during a planetary return. Their
determination represented a significant research effort
in the HUYGENS program [19]. Figure 9 exhibits a
large dispersion of radiative flux predictions for a
Venus entry from orbit. They are in average of the
same order as the convective flux.
Heat fluxes at the nose of the probe are predicted for a
non-catalytic and isothermal wall. Two cases have been
investigated with different nose radii for two points of
the trajectory. The conditions of these cases are listed
in Table 2. For each case, three different calculations
have been performed, for temperatures of 1200 K,
1500 K and 1800 K.
2.5 -5-
mission is the entry in Earth atmosphere of a small
capsule. In particular, an accurate prediction of the heat
flux at the capsule nose during the entry is required to
design the thermal protection system. A preliminary
assessment is obtained using numerical simulations.
Figure 10: Pressure field and flux at the wall. Case (1),
Tw=1500 K (Tw is the temperature at the wall).
Figure 10 represents the pressure field around the probe
and the heat flux at the wall for the case at 1500 K. The
curve of the heat flux at the wall shows a small
decrease at the stagnation point. This result is not
physics and is attributed to the Roe scheme used by
TINA. This scheme is not adapted for viscous flow
simulations at stagnation point.
Tw(K)
1200
1500
1800
Case 1 - TINA
2.
1.9
1.7
Case 1 - IPM
1.
0.88
0.76
Case 2 - TINA
1.05
0.93
0.8
Case 2 - IPM
0.88
0.76
0.64
Figure 11: Earth entry capsule temperature contours.
The top half represents a frozen flow and the bottom
half an equilibrium dissociated and ionized air flow.
Here, the flow around a capsule has been computed
using several thermochemical models. This allows an
estimate of the modelling influence on the temperature
predictions. The capsule has a weight of 20 kg and the
nose radius is 0.4 m. The thermo-physical conditions of
the point of the trajectory studied are listed in Table 4.
Velocity (km/s)
Pressure (Pa)
Density (kg/m3)
Temperature (K)
5
Table 3: Stagnation point heat flux (in 10 W/m2).
Comparison between TINA and IPM calculations for
three isothermal wall temperatures (1200 K, 1500 K
and 1800 K).
14.17
8.7
1.32 10-5
185
Table 4: Upstream flow conditions for the numerical
simulations of Earth entry.
The heat fluxes at the stagnation point predicted by
TINA and calculated at IPM are shown is Table 3. The
results given by the two simulations are quite different
in the case at high velocity and low pressure but closer
in the case at low velocity and high pressure. The last
case is indeed closer to an equilibrium calculation than
the first one.
Case
(1) Frozen model, γ=1.4
(2) Air model
(3) Non-equilibrium
(4) Non-equilibrium with ablation
6. Earth entry
Table 5: Computed test cases for a Mercury sample
return mission.
ESA has recently studied the scenario for a sample
return mission to Mercury [3]. This mission would
consist in a single ARIANE 5 rocket launch, and then
the probe will fly to Mercury by electric propulsion.
Since Mercury has no atmosphere, engines will be used
to ensure the landing. Then a capsule, containing the
samples will be launched for the return flight to Earth.
Concerning aerothermodynamics, the main issue of this
The different computed cases with the corresponding
mesh sizes are summarised in Table 5. First, an
equilibrium approach has been applied to the flow
using the solver BLT [21, 22]. Two models have been
used: a frozen model where the air is a mixture of
perfect gas without chemical effects, and a reacting
model where the dissociation and ionisation processes
are taken into account. Then, the flow has been
2.5 -6-
Code
BLT
BLT
TINA
TINA
Mesh
80x59
80x59
60x80
60x80
simulated using the non-equilibrium solver TINA. The
Case 3 corresponds to a non-equilibrium air model and
the last case accounts for ablation phenomenon.
projects. This work gives some indications concerning
the adequacy of software and ground test facilities to
planetary exploration. Concerning the numerical
simulations, experimental data are needed to improve
the reliability of the predictions. To achieve this,
ground test facilities have to be adapted to simulate
"exotic" atmospheres. Other improvements concerning
in particular severe entry, radiation, microaerothermodynamics and aerobraking aspects are
necessary to ensure the success of future missions. The
technological efforts are within capabilities of Europe,
which is ready to embark in successful exploration
missions.
In Figure 11, the temperature fields simulated with the
two equilibrium models are shown. The temperature is
much lower in case 2 than in case 1. This highlights the
strong effect of dissociation and ionisation on the flow
field. The predicted shock wave is closer to the capsule
in the second case. The temperature fields predicted by
the non-equilibrium solver are shown in Figure 12.
These computations put in evidence the large influence
of non-equilibrium on the results. The maximums of
temperature are in between the perfect gas and the
equilibrium air model predictions. This is consistent
with the fact that in non-equilibrium only a part of the
fluid is dissociated and ionised. Another aspect of nonequilibrium influence is the shock wave location. The
shock wave location is close to the location predicted in
Case 2. The last aspect is the influence of ablation on
results of non-equilibrium simulations. The temperature
predicted in the ablative case is lower than in the nonablative one. The shock wave is closer from the capsule
in the ablative simulation. Therefore, ablation has a
large influence on the non-equilibrium results. This
process has to be carefully studied to design the
thermal protection shield. These first predictions show
that 1/3 of the thermal shield mass (see [3]) would be
lost during an Earth entry.
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5. CONCLUSIONS
In this paper a partial overview of the activities
conducted at ESTEC on aerothermodynamics for
exploration missions has been done. A more systematic
survey was out of the scope of the paper, but would be
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2.5 -8-