11institutetext: German Aerospace Center, Institute of Aerospace Medicine,
Radiation Biology Department, Linder Höhe, 51147 Köln, Germany, 11email:
<EMAIL_ADDRESS>
22institutetext: Delft University of Technology, Faculty of Mechanical
Engineering, Mekelweg 2, 2628CD, Delft, The Netherlands
33institutetext: University of Southampton, Faculty of Applied Sciences and
Engineering, SO17 1BJ Southampton, United Kingdom
44institutetext: Vienna University of Technology, Karlsplatz 13, Vienna,
Austria
55institutetext: Department of Experimental Physics, National University of
Ireland Maynooth, Co. Kildare, Ireland
66institutetext: University of Granada, Faculty of Sciences, Avda. Severo
Ochoa s/n, Granada, Spain
77institutetext: University of Cologne, Institute of Geophysics and
Meteorology, Pohligstrasse 3, 50969 Cologne, Germany
88institutetext: University of Stuttgart, Keplerstraße 7, 70174 Stuttgart,
Germany
99institutetext: AIM Paris-Saclay Laboratory, CEA/Irfu Paris Diderot
University CNRS/INSU, 91191 Gif-sur-Yvette, France
1010institutetext: Politehnica University of Bucharest, Faculty of
Electronics, Telecommunications and Information Technology, Splaiul
Independentei 313, Romania
1111institutetext: Institut de Recherche en Astrophysique et Planetologie,
Toulouse University (UPS), CNRS UMR 5277, France

# A Space weather information service

based upon remote and in-situ measurements of coronal mass ejections heading
for Earth
A concept mission consisting of six spacecraft in a heliocentric orbit at 0.72
AU

Birgit Ritter 11 Arjan J. H. Meskers 22 Oscar Miles 33 Michael Rußwurm 44
Stephen Scully 55 Andrés Roldán 66 Oliver Hartkorn 77 Peter Jüstel 88 Victor
Réville 99 Sorina Lupu 1010 Alexis Ruffenach 1111

###### Abstract

The Earth’s magnetosphere is formed as a consequence of interaction between
the planet’s magnetic field and the solar wind, a continuous plasma stream
from the Sun. A number of different solar wind phenomena have been studied
over the past forty years with the intention of understanding and forecasting
solar behavior. One of these phenomena in particular, Earth-bound
interplanetary coronal mass ejections (CMEs), can significantly disturb the
Earth’s magnetosphere for a short time and cause geomagnetic storms. This
publication presents a mission concept consisting of six spacecraft that are
equally spaced in a heliocentric orbit at 0.72 AU. These spacecraft will
monitor the plasma properties, the magnetic field’s orientation and magnitude,
and the 3D-propagation trajectory of CMEs heading for Earth. The primary
objective of this mission is to increase space weather forecasting time by
means of a near real-time information service, that is based upon in-situ and
remote measurements of the aforementioned CME properties. The obtained data
can additionally be used for updating scientific models. This update is the
mission’s secondary objective. In-situ measurements are performed using a
Solar Wind Analyzer instrumentation package and flux gate magnetometers, while
for remote measurements coronagraphs are employed. The proposed instruments
originate from other space missions with the intention to reduce mission costs
and to streamline the mission design process. Communication with the six
identical spacecraft is realized via a deep space network consisting of six
ground stations. They provide an information service that is in uninterrupted
contact with the spacecraft, allowing for continuous space weather monitoring.
A dedicated data processing center will handle all the data, and then forward
the processed data to the SSA Space Weather Coordination Center which will, in
turn, inform the general public through a space weather forecast. The data
processing center will additionally archive the data for the scientific
community. The proposed concept mission allows for major advances in space
weather forecasting time and the scientific modelling of space weather.

###### keywords:

Coronal Mass Ejection (CME) – Remote sensing – In-situ measurement –
Geomagnetic storms – Forecast – Services

##  1 Introduction

The Earth and its near surroundings are affected by space weather, which is
defined as ”the physical and phenomenological state of natural space
environments” (COST 724 final report [2009]). The Sun is the main driver of
space weather phenomena and coronal mass ejections (CMEs) are one of these
phenomena. CMEs are created by plasma eruptions of solar material in a mass
range between $10^{12}$ and $10^{13}$ kg (Meyer-Vernet [2007], COST 724 final
report [2009]). Magnetic reconnection, caused by the twisting and tangling of
magnetic field lines, occurs in the solar corona, where a vast amount of
energy is contained. This process results in an eruption of a magnetic field
structure, potentially leading to a CME (Glaßmeier and Scholer [1991];
Gopalswamy [2003]). However, at present, the trigger of a CME is not fully
understood. The rate of occurrence of CMEs varies roughly according to the
solar cycle, ranging from an average rate of 3 per day at solar maximum to an
average rate of 1 per week at solar minimum (Meyer-Vernet [2007]). Earth-
directed CMEs possibly impacting Earth’s environment exhibit a minimal opening
angle of 60 degrees and their travel velocities range from 50 to 3000 km/s
(Howard et al. [1985]; Gopalswamy et al. [2004]; Gopalswamy [2008]). After its
release, the CME plasma and magnetic field propagate through the
interplanetary medium (IPM) and interact with the ambient solar wind. The
manifestation of a CME in the IPM is sometimes called interplanetary coronal
mass ejection (ICME), whereas the expression ’CME’ is also accepted for both
objects and is used throughout this article. The expansion of a CME is
approximately self-similar within the IPM and its plasma density and magnetic
field strength decrease faster than linearly during its propagation through
the heliosphere (e.g. Gulisano et al. [2012], and references therein).
Moreover, comprehensive studies on CME propagation have been conducted by
Manoharan ([2006]) and Manoharan ([2010]) using remote-sensing observations
indicating that the magnetic energy associated with a CME is responsible for
maintaining the CME structure during its propagation through the IPM. The
shock front develops since the speed differential between the CME plasma and
the usual solar wind plasma often exceeds the magnetosonic wave speed. It is
followed by the shocked solar wind plasma, the sheath that is a turbulent
region, and the driving ejecta of the CME.
The magnetic field and plasma properties of the driving ejecta are crucial for
the process properties of the interaction between a CME and Earth’s plasma
environment. As a matter of fact, CMEs can cause geomagnetic storms on Earth
upon interaction with Earth’s magnetosphere.

Indeed, they are the primary cause for the most severe storms (Gosling et al.
[1990]). Such storms are defined as intervals of time, in which the
magnetospheric ring current is intensified as a result of increased energy and
particle injections from solar wind plasma into Earth’s magnetosphere-
ionosphere system. The enhancement of the ring current generates perturbations
of the geomagnetic field on the ground. Effects of geomagnetic storms on and
near Earth include communication disruptions, current surges in power lines
and radiation hazards to operating astronauts and spacecraft (e.g. COST 724
final report [2009], Reitz [2008]).
Coronagraphs, such as the Large Angle and Spectrometric COronagraph (LASCO,
Brueckner et al. [1995]) instrument onboard the Solar and Heliospheric
Observatory (SOHO, Domingo et al. [1995]), are used to monitor the solar
corona for disturbances such as CMEs (Hundhausen et al. [1984]). CMEs
containing a dominant southward magnetic component are most geoeffective
(Gonzales et al. [1994]), meaning they have a high ability to cause
geomagnetic storms (Gopalswamy [2008]). In this case, magnetic reconnection at
the dayside of the magnetopause (Dungey [1961]) allows for energy of the
ejecta to be transferred into the inner regions of the Earth’s magnetosphere
(Akasofu [1981]), which enhances the ring current intensity. Besides having a
specific magnetic field orientation, a CME’s dimension and velocity are
additional important factors that determine the effectiveness of a magnetic
storm (Gosling et al. [1990]).
CME s can currently be detected in-situ by spacecraft at L1 (Lagrange-1
point), such as the Advanced Composition Explorer (ACE, Smith et al. [1998]),
the Comprehensive Solar Wind Laboratory for Long-Term Solar Wind Measurements
(WIND, Ogilvie et al. [1995]), and SOHO. They allow for an effective warning
time of about one hour before effects of a geomagnetic storm are detectable on
Earth.
Previous mission proposals have considered how to increase this warning time.
Examples include ’Geostorm’ (West [1996], [2004]), which is a solar-sail
mission at 0.98 AU and ’Space Weather Diamond’ (St. Cyr et al. [2000]), which
is a multi-spacraft mission at 0.9 AU. ’Geostorm’ is supposed to be placed in
a rearranged orbit at L1, which is shifted closer to the Sun due to the effect
of a solar sail. ’Space Weather Diamond’ consists of four spacecraft on
eccentric heliocentric orbit seeming to circuit around the Earth 0.1 AU apart
from it. All spacecraft are equally equipped with in-situ measurement
instruments.

This paper presents a new mission proposal to monitor CMEs. The mission
constellation is comprised of six spacecraft placed in heliocentric circular
orbits at 0.72 AU at a separation angle of 60 degrees from each other. All
spacecraft are able to perform remote and in-situ measurements. The mission
serves as a continuous information-service system, providing all necessary
information of the heliosphere that allow for an increased space weather
forecasting time and enhancement of scientific models regarding space weather.
The data will help to forecast if there will be a magnetic storm and determine
its severity, and will also provide information on radiation levels for
astronauts and space missions further away from the Sun. This mission poses a
large advancement in protecting human’s health and technology, and in
preventing technological crashes and negative results for Earth’s eco-system.

In contrast to Geostorm, the proposed mission is based on well-known and
tested technical equipment, ready for the usage as a steadily operating system
within the next decades. ’Space Weather Diamond’ is called a monitoring
system, as well as the here presented mission proposal. However, ’Space
Weather Diamond’ focuses on the reactor of the system, the Earth, whereas the
here presented mission monitors the actual actor, the Sun, including remote
measurements systems allowing a more comprehensive view on the state of space
weather than single point measurements. This mission proposal sets the
spacecraft significantly closer to the Sun, at 0.72 AU and thereby yields a
much longer forecasting time, which is the most important advantage of this
mission.

However, with an increase of forecasting time comes naturally a decrease of
accuracy. Lindsay et al. ([1999]) concluded that in-situ measurements between
1.0 AU and 0.7 AU in the equatorial plane within 10 degrees east to 5 degrees
west of the Earth-Sun line allow substantial space weather forecasts. While
the underlying model of this conclusion is a simple linear model by Burton et
al. ([1975]), the authors of the paper at hand believe that advanced data-
driven CME models (e.g. Luhmann et al. [2004], Tóth et al. [2005]), calibrated
by data of the early mission-phase, will allow for a reliable warning system,
if at least one spacecraft is located within a range $\pm$30 degrees apart
from the Earth-Sun line.
The proposal was originally entitled as the CARETAKER mission, and has been
designed by a group of Bsc. and Msc. students, PhD-candidates and Postdocs
from the ESA member states, as part of the Alpbach Summer
School111http://www.summerschoolalpbach.at/ 2013\. This summer school was
organized and funded by an international cooperation between the European
Space Agency (ESA), the Aeronautics and Space Agency of Austria (part of the
Austrian Research Promotion Agency FFG), the International Space Science
Institute (ISSI), and Austrospace, the association of Austrian space
industries and research institutions.
In this paper the CARETAKER mission proposal is presented. The paper is
structured as follows: in Section 2 the mission overview is given including
the mission statement and requirements, and the operational concept. The
flight segment is described in Section 3, followed by the communication
segment in Section 4 and the operation and ground segment in Section 5. In
Section 6 the data processing is discussed, and Section 7 covers the budgets
of the mission followed by a risk analysis in Section 8, and a cost analysis
in Section 9. The paper ends with a conclusion in Section 10.

##  2 Mission Overview

The mission consists of six identical spacecraft in a heliocentric orbit at
0.72 AU, performing in-situ and remote measurements. The obtained data will be
communicated periodically via a direct link between spacecraft and Earth
through a dedicated deep-space network, while communication with the
spacecraft located behind the Sun is naturally excluded. The deep-space
network consists of six ground stations that are located in a way to ensure
continuous contact with the spacecraft. Two ground stations will operate
simultaneously, each alternating communication between three spacecraft. All
obtained information will be collected by a single mission operations center
that passes it along to a data processing center where the data will be
processed and submitted to the SSA Space Weather Coordination
Center222http://swe.ssa.esa.int/web/guest/ssa-space-weather-activities (SSCC),
and the information will additionally be archived for the scientific
community. It will be the SSCC’s responsibility to send out a warning to
respective users requiring such an alert system in order to trigger their
precautions that need to be undertaken.

###  2.1 Mission Statement and Requirements

The mission states that: _the system consists of a near real-time information-
service, based on the physical properties of solar Coronal Mass Ejections._
The primary objective of the mission is to _obtain information about CMEs as
input for a CME warning system_ , which covers information about:

  1. 1.

the propagation trajectory of CMEs heading for Earth, and

  2. 2.

the physical properties of CMEs.

The secondary mission objective is to _improve CME models by remote and in-
situ multipoint measurements._

Achieving the above stated mission objectives requires that the mission shall:

  1. 1.

monitor the three dimensional trajectory of CMEs that head for Earth;

  2. 2.

measure the CME’s magnetic field orientation and magnitude (related to
geoeffectiveness);

  3. 3.

determine the CME’s propagation envelope within an accuracy of less than 4.3
arcmin (i.e. the angular diameter of the Earth’ s magnetosphere as viewed from
the Sun);

  4. 4.

remotely monitor CMEs between 2 to 15 solar radii (for enhancement of
scientific models, Thernisien et al. [2009]);

  5. 5.

provide a minimum forecasting time of 12 h, measured upon reception of in-situ
measurement values;

  6. 6.

measure the following plasma properties: the 3D velocity distribution of
protons and electrons, and the composition of heavy ions up to 56 amu/q, all
measured with a time resolution of 60 s for the detection of the CME shock
front (Richardson and Cane [2004]);

  7. 7.

measure the low-energy ion particle flux in the range of 0.26 keV/q to 20
keV/q as well as the low-energy electron flux in the range of 1 eV to 5 keV;

  8. 8.

measure the magnetic field with a resolution of 0.1 nT in the range between
-200 nT and 200 nT (Burlaga [2001]);

  9. 9.

provide the SSCC with data, processed according to their standards (allowing
them to construct a space weather forecast for the general public);

  10. 10.

have an operational life time of 5 years (with a possible extension of 5
years);

###  2.2 Operational Concept

The mission combines results from remote stereo images and multiple in-situ
measurements in order to determine the trajectory and physical properties of
CMEs. This is achieved by having six identical spacecraft in a heliocentric
orbit at 0.72 AU, at a separation angle of 60 degrees apart. This separation
angle is driven by the argument that CMEs with an angular extent larger than
60 degrees have a notable impact on Earth (Gopalswamy [2008]). All spacecraft
are equipped with a coronagraph (2-15 solar radii field of view, 5 min
cadence), a Solar Wind Analyzer instrumentation package, and flux gate
magnetometers, both sampling at 1 Hz. The in-situ measurements are performed
continuously by all spacecraft, whereas only two spacecraft will be performing
remote measurements simultaneously, limiting the amount of data communicated
back to Earth. The two spacecraft performing remote measurements will be the
ones having the best point of view for stereoscopic imaging with respect to
Earth. The three spacecraft closest to the Earth (two of which perform remote
sensing as well as in-situ) will downlink their data every 15 minutes whereas
the three distant spacecraft (only making in-situ measurements) will downlink
every 45 minutes. The processing of stereoscopic remote sensing of a CME will
result in information regarding the CME’s trajectory, velocity and size. These
results are processed in a standardized format (in accordance with the SSCC)
and will be passed on to the SSCC within approximately 45 minutes after the
CME’s occurrence at the Sun. The CMEs travelling towards Earth will be
continuously monitored up to 15 solar radii. In-situ measurements of the solar
wind and passing CMEs will be made at 0.72 AU. This information will also be
processed according to standards and made available to the SSCC within 45
minutes after detection. From the moment on in which the in-situ measurement
data has been processed it is possible to determine the geoeffectiveness of
the approaching CME, resulting in a precaution time of at least 12 hours
before the CME reaches Earth’s magnetosphere (estimated for fast CMEs with a
travelling speed of 1000 km/s when arriving at the Earth).

##  3 Flight Segment

###  3.1 Orbit

From the mission requirements it can be concluded that the CARETAKER concept
must consist of six spacecraft at a heliocentric orbit of at most 0.8 AU, each
with a 60 degrees separation. Analysis of the options for realizing this
concept led to the optimal solution of inserting the spacecraft into a Venus
orbit at 0.72 AU using gravity assist maneuvers (GAMs) at Venus to reduce the
required propellant mass. The dependance on the GAM limits the launch window
to a maximum of 3 weeks every 19 months.
Solar cycles 25 and 26 (counting began in 1761, on cycle after the end of the
Maunder minimum) will likely occur between 2020-2031 and 2031-2042,
respectively (DeRosa et al. [2012]). The maximums of those cycles will thus be
around 2026 and 2037, with an enhanced number of solar events expected to
occur. The first spacecraft will be in final orbit in 2027, the last in 2030.
Thus during calibration of the payload, some spacecraft are likely to be
exposed to extreme events. The nominal mission will cover an increasing solar
activity during the 5 years of operations, even if the launch is delayed. With
the extended mission lifetime, an entire solar cycle can be monitored.

###  3.2 Launch

A required characteristic energy of 5.71 km2/s2 was calculated to insert the
spacecraft into the transfer orbit. The calculation is based on the mass of
the spacecraft and its overall trajectory. Table 1 displays the resulting
launcher options investigated for this mission.

Table 1: Launcher tradeoff between Soyuz and Ariane 5. Despite the higher costs, Ariane 5 launchers are considered the better option with the Soyuz being a viable descoping option. Parameter `\` Launcher |  | Soyuz | Ariane 5
---|---|---|---
Performance at Required Escape Speed |  | 1850 kg | 5255 kg
Cost per Launch |  | 75 M EUR | 160 M EUR
Cylindrical Fairing Dimensions (h x $\varnothing$) |  | 5060 x 3860 mm2 | 10039 x 4570 mm2
Number of Launchers Required |  | 3 | 2
Total Mass Delivered to Transfer Orbit |  | 5550 kg | 5255 kg
Total Cost |  | 225 M EUR | 320 M EUR

As the launch window constraint has more significant implications for the
Soyuz launch, it was concluded that the delay on mission operations was severe
enough to justify using two Ariane 5 launchers. The Soyuz launcher is however
a viable descoping option.
Kourou in French Guiana is chosen as the designated spaceport because of its
proximity to the equator.

###  3.3 Orbital Insertion Procedure

The use of parking orbits is essential in order to insert the six spacecraft
into the distinct positions around the Venus orbit separated by 60 degrees and
using two launchers and GAMs. These parking orbits have been designed to
minimize the required propellant mass at minimal cost to the time taken to
insert all spacecraft into the desired orbits.
The first three spacecraft will be launched into a transfer orbit from Earth’s
to Venus’s orbital radius (Figure 1). During the transfer period the
spacecraft will be in safe mode. After six months they will perform the GAM
inserting two spacecraft into parking orbit A and one into parking orbit B.
Parking orbit A is an elliptical orbit with perihelion equal to Venus’s
orbital radius and aphelion greater such that the period is 13/12 of Venus’s
orbital period. Similarly, parking orbit B is also elliptical, however the
aphelion is equal to Venus’s orbit and the perihelion is inside Venus’s orbit
such that the period is 11/12 of Venus’s orbital period. The two spacecraft in
parking orbit A will complete three and five full orbits until they reach
orbital positions five and four, respectively. Similarly for the spacecraft in
parking orbit B, which will complete five full orbits to reach position three.
The last three spacecraft will be launched 19 months after the first three
spacecraft and will follow the same transfer orbit to the gravity assist. Two
spacecraft will be inserted into the parking orbit B and one into parking
orbit A to fill the remaining positions. Once each spacecraft is in its final
destination relative to Venus it can begin the one month commissioning phase;
upon completion, operation can begin. The first two spacecraft will be
operational 32 months after the first launch date, where the mission can
become partially operational. Full operations can begin within less than 47
months after the first launch date.

Figure 1: The six spacecraft (S/C) are launched using two Ariane 5 rockets,
two sets of three spacecraft. After launch the spacecraft are separated and
travel towards Venus for a gravity assist maneuver (GAM). After the GAM the
spacecraft will be parked into two orbits and will from thereon reach their
final orbit, a Venus orbit with 60 degrees separation between each spacecraft.

###  3.4 Spacecraft Design

All six spacecraft will have the same design. Figure 2 outlines the schematic
view of one spacecraft that includes the sensors (indicated in red text color
in the figure) that are described in Section 3.5 and some of the spacecraft
subsystems (indicated in black text color), partly explained in Sections 3.6
to 3.8. The side of the spacecraft facing the Sun has an area of 1.5 $\times$
1.5 m2, the dimensions towards space are 1.7 m.

Figure 2: Spacecraft design for the six identical spacecraft of the dimensions
1.5 m $\times$ 1.5 m $\times$ 1.7 m (yellow) excluding the solar panels (in
blue, Section 3.7). The payload instruments are indicated in red text color
(Section 3.5) and the spacecraft subsystems in black text color.

###  3.5 Payload

The instruments employed for CARETAKER originate from other space missions
with the intention to reduce mission costs and accelerate overall mission
design. In-situ measurements are performed using the Solar Wind Analyzer
instrumentation package from the _Solar Orbiter_ mission (Marsch et al.
[2005]), together with flux gate magnetometers from _Venus Express_ (Titov et
al. [2006]), while remote measurements are performed using coronagraphs from
the _STEREO_ mission (Kaiser et al. [2008]).

####  3.5.1 Coronagraph

Each spacecraft will include a coronagraph. Stereoscopic images of the Sun
will be obtained by pairs of coronagraphs separated by 120 degrees. In order
to obtain satisfactory stereoscopic images, each coronagraph will be switched
off for the period the spacecraft orbits along the Sun-Earth line. The
coronagraphs used are externally occulted Lyot coronagraphs that observe from
2 solar radii to 15 solar radii. This choice of heliocentric distances
provides the opportunity to observe the early stages of a CME. The
coronagraphs derive their heritage from the coronagraph COR2 on-board the
SECCHI suit of instruments of the _STEREO_ mission (Howard et al. [2008]). The
external occultation shields the objective lens from direct sunlight, hence it
enables a low stray light level and makes the observation in this range of
heliocentric distances possible (Thernisien et al. [2009]). The technical
features of the coronagraphs are presented in Table 2, where L0 is the solar
polarization brightness and DN is the measured response of the instrument in
the passband of the observations. The stereoscopic observations, combined with
geometrical models (e.g. ice cream cone model, hollow croissant model) and
reconstruction methods (forward modelling, inversion, and triangulation) allow
to obtain the three-dimensional structure of CMEs and extract their
propagation direction and velocity. The stereoscopic images obtained will
assist to find the propagation direction and velocity of CMEs in the region
2-15 solar radii and hence to acquire an estimation of whether a CME will be
Earth-directed or not.

Table 2: Coronagraph technical details. Field Of View (FOV) | 11.4∘
---|---
Passband | 450 nm – 750 nm
Data rate | 16.7 kbps
Compression factor | 10
Pixel size | 15 arcsec
Exposure time |  $<$4 sec
Image sequence cadence | 5 min
Images per hour | 12
Photometric response (L0/DN) | 1.3$\cdot$10-12

####  3.5.2 Solar Wind Analyzer

The Solar Wind Analyzer (SWA) consists of 3 sensors with a shared processing
unit. The main objective of the SWA will be to characterize CME properties and
to determine their magnetic field structure via comprehensive in-situ
measurements of CMEs. In order to meet all the measurement requirements, the
SWA must be able to measure the three-dimensional velocity distribution
functions of the major solar wind components: protons, electrons and heavy
ions (ESA [2011]).
The Electron Analyzer System (EAS) will make a high temporal resolution
determination of the 3D electron velocity distributions and derive their
moments (density, temperature, bulk velocity, heat flux).
The Proton Alpha Sensor (PAS) contains a top hat electrostatic analyzer that
measures the 3D velocity distribution of protons and alpha particles in the
energy range of 0.2 keV/q to 20 keV/q with a relative accuracy of 8% in energy
and an angular resolution of less than 2 degrees. Unlike the EAS, the PAS
consists of only one device and has therefore a narrower field of view.
The Heavy Ion Sensor (HIS) contains an electrostatic analyzer module and a
time of flight detector. It determines mass, charge, energy distribution, and
direction of incidence of heavy ions up to 56 amu/q.
Table 3 lists the properties of the three sensors and their location onboard
the spacecraft can be seen in Figure 2.

Table 3: Properties of the three sensors (Electron Analyzer System (EAS), Proton Alpha Sensor (PAS), and Heavy Ion Sensor (HIS)) that constitute the Solar Wind Analyzer (SWA). | EAS | PAS | HIS
---|---|---|---
Field of View | 4$\pi$ | -24∘ – 42∘ (Az) | -30∘ – 66∘ (Az)
|  | -22.5∘ – 22.5∘ (El) | -17∘ – 22.5∘ (El)
Particle Species | Electrons | H+, He++ |  3He – Fe
Energy Range | 1 eV – 5 keV | 0.2 keV/q – 20 keV/q | 0.5 keV/q – 100 keV/q (Az)
|  |  | 0.5 keV/q – 16 keV/q (El)
Energy Resolution | 12% | 8% | 6%
Measurement Parameters | flux, velocity, | 3D velocity distribution | energy, charge,
| energy distribution 3D |  | mass, direction 3D
Angular Resolution | 10∘ | 2∘ | 6∘
Cadence | 4 sec | 4 sec | 30 sec

####  3.5.3 Fluxgate Magnetometer

The Fluxgate Magnetometer (FGM) will perform in-situ measurements of the
magnetic field vector. The instrument uses two triaxial fluxgate sensors which
allow separation of stray field effects of the spacecraft from the ambient
magnetic field. One of the sensors is mounted on a boom with a length of 3 m,
while the other sensor is directly attached to the spacecraft. The sampling
rate is $1~{}\mathrm{Hz}$ for normal operation. Values are averaged to one
minute values. The instrument design is based on the Fluxgate Magnetometer
(MAG, Zhang et al. [2007]) on _Venus Express_ , except for increased boom
length. Similar instruments have also been flown on _Rosetta Lander_ (Biele
and Ulamec, [2007]) and the Mir Space station. This instrument is capable of
measuring in interplanetary space as well as inside the magnetic field of
Venus (Balogh [2010]). The range of the outer sensor is $\pm
262~{}\mathrm{nT}$ with an accuracy of $8~{}\mathrm{pT}$ and the on-board
sensor with a range of $\pm 524~{}\mathrm{nT}$, also with an accuracy of
$8~{}\mathrm{pT}$. This theoretically available accuracy exceeds the required
resolution of $100~{}\mathrm{pT}$ (Section 2.1) and is therefore not needed at
such a high level, which would increase mission costs unnecessarily.

###  3.6 Attitude Orbit Control System (AOCS)

All of the spacecraft will be 3-axis stabilized. The alternative (spin
stabilized) maintains a fixed orientation with respect to the stars and is
therefore not suitable for this mission. Each spacecraft must face the Sun at
all times and subsequently must complete one full rotation in sync with its
orbital period. In addition, to transmit to Earth consistently, the antenna
must be steerable for pointing, which is only feasible with 3-axis stabilized
configuration. The control system for stabilization will consist of momentum
wheels333http://www.rockwellcollins.com/sitecore/content/Data/Products/Space_Components/Satellite_
Stabilization_Wheels/RSI_12_Momentum_and_Reaction_Wheels.aspx (accessed July
2013).  in conjunction with
thrusters444http://cs.astrium.eads.net/sp/spacecraft-propulsion/bipropellant-
thrusters/4n-thruster.html (accessed July
2013),555http://cs.astrium.eads.net/sp/spacecraft-propulsion/bipropellant-
thrusters/22n-thruster.html (accessed July 2013) in the reaction control
system (RCS) for de-saturation. The various components are detailed in the
following sections.

####  3.6.1 AOCS Components

In order to optimize stability while accounting for redundancy the reaction
wheels are mounted in a tetrahedral configuration (Wagner et al. [2012]);
attitude control can be achieved with four wheels operating simultaneously
(the nominal operational scenario) or any combination of three wheels (if one
wheel were to fail).

Three star trackers666http://www.selex-es.com/domains/space/attitude-control-
sensors (accessed Feb. 2014) (STRs), one in cold redundancy, determine the
orientation relative to the stars. Two inertial measurement units (IMU) are
mounted on a common platform, providing a finer control of orientation. These
IMUs consist of three Fibre Optic Gyros for rate measurement (Hablani [1994]).
One is required for normal operation and the second is for redundancy. Sun
sensors determine the spacecraft orientation relative to the Sun and allow for
attitude control in safe mode, in which thrusters are used to keep solar cells
aimed at the Sun and avoid damaging any vulnerable imaging instruments. There
are a total of two Sun sensors4 (one redundant) and four reaction wheels (one
redundant).
During science operations, at least two STRs will be used in combination. In
the event of major system anomaly in the spacecraft and consequent loss of
attitude control, dedicated shutters will protect the STR optical paths to
prevent damage due to accidental Sun pointing. Orientation algorithms will be
processed on the spacecraft’s main CPU.

####  3.6.2 Reaction Control System (RCS)

The spacecraft are required to be separated by 60 degrees in their orbit
around the Sun. As discussed in Section 3.1, the period of the transfer orbits
is crucial to allow for efficient insertion into the desired orbit with the
correct phase.

There will be 24 RCS thrusters arranged in groups of three at the corners of
the main cubic structure of each spacecraft. The thrusters of one block form a
triad, thereby producing thrust in the three perpendicular directions. Four
blocks are sufficient to allow the spacecraft to rotate and translate in any
direction. Eight sets of thrusters thus grant full, built-in redundancy. The
RCS momentum dumping will require only short, sporadic pulses. The total burn
time estimated for the RCS thrusters is calculated at 10.2 hours with cycles
per thruster not exceeding 10% of the rated capacity of the thrusters.

Table 4 shows the propellant mass required to maintain the orbit of each
individual spacecraft. Perturbation calculations were completed with respect
to a full solar cycle ($11\times 365$ days).

Table 4: Breakdown of the mass of propellant required for station keeping for each of the 6 spacecraft. The spacecraft are listed in order of heliocentric angular separation from Venus. The associated propellant mass required for station keeping, over the lifetime of the mission, is shown per spacecraft. Spacecraft number | Angular separation from Venus | Propellant mass for station keeping
---|---|---
| $\left[\mathrm{{}^{\circ}deg}\right]$ | $\left[\mathrm{kg}\right]$
1 | 30 | 16.6
2 | 90 | 19.9
3 | 150 | 16.8
4 | 210 | 14.0
5 | 270 | 16.8
6 | 330 | 18.0

####  3.6.3 Propulsion System

The propulsion systems considered for the CARETAKER mission include chemical,
electrical and hybrid options. Electrical propulsion has the potential to save
more than 100 kg per spacecraft with respect to chemical propulsion. However,
electric propulsion has low thrust capabilities and therefore longer transfer
times as well as a high power consumption. It consequently requires extra
solar arrays and the resulting high thermal output has furthermore
implications on the thermal system. Additionally, being a relatively new
technology it harbors an increased risk of failure.
A third option is a hybrid solution. This would involve employing chemical
thrusters to provide the major burns, then utilizing low thrust electric
thrusters to maneuver into the final position. The RCS uses chemical thrusters
therefore, compared to a purely electrical system, the increase in mass of
additional equipment would be minimal. Calculating the optimal chemical and
electric propulsion hybrid along with the subsequent orbital insertion
procedure is beyond the scope of this paper but there is potential to reduce
the overall insertion time and mass.
After considering the above, it was concluded that propulsion system will be
chemical. The total mass of propellant required for the insertion of each
spacecraft into its desired orbit is 310 kg. Each spacecraft will have 22 N
thrusters which allows them to perform the necessary burns for the orbital
maneuvers in about two hours. With an accumulated burn life of 70 hours and
1,000,000 cycles, the chosen thrusters are well within the operational limits.

####  3.6.4 Propellant

RCS and main engine thrusters will be supplied from a common bi-propellant
tank set. Propellant requirements have an optimal mix ratio of 1.4 and the
total mass requirement amounts to 306.2 kg.
Two bladder tanks were selected, allowing for sufficient expansion. The
specific tanks selected were Astriums 198L bladder
tanks777http://cs.astrium.eads.net/sp/spacecraft-propulsion/propellant-
tanks/198-litre-bipropellant-tank.html (accessed Feb. 2014).

###  3.7 Power

The electrical power sub-system provides, stores, distributes and controls the
spacecraft energy. The primary power sources considered are the solar panels;
batteries are the secondary power source.

####  3.7.1 Primary Power Sources – Solar Panels

The solar panels convert the radiation from the Sun into electrical power to
maintain full operation of the sub-systems. Included in the trade-off analysis
for the solar panels are the spacecraft orbit and the power requirements of
the sub-systems.
Hence, the best option in terms of efficiency and dimensions are gallium-
arsenide semiconductor cells which double the efficiency of a silicon solar
cell used in the past. 28% Triple Junction GaAs Solar Cell of Type: TJ Solar
Cell 3G28C888http://www.azurspace.com/images/products/HNR_0002490-00-03.pdf
were considered the best solution for the spacecraft. With respect to the
total amount of power needed to operate each spacecraft, the required area of
the solar panels is approximately 3 m2.

####  3.7.2 Secondary Power Sources – Batteries

Batteries are used as a secondary power system, operating to store energy.
Taking into account the available space-rated batteries and the mission
requirements, the energy is stored in nickel-hydrogen batteries. With space-
based communications requiring large amounts of power and high reliability,
the $\mathrm{NiH}_{2}$ batteries represent a desirable power supply with a
long cycle life (Dermott et al. [1996]). During the launch, the CPU and the
communication requires up to 30 W, thus leading to a discharge of 6% after 2
hours. In safe mode, the batteries will still be charged at 52% after 24
hours.
The secondary power supply will contain 12 cells of $\mathrm{NiH}_{2}$, to
provide approximately 100 Wh per cell.

###  3.8 Thermal Control System

The thermal control system has been designed to ensure that the spacecraft
instrument and component temperatures are always within the operational range
of 298 K $\pm$ 15 K during the orbital periods and operational lifetime. A
total of 2.2 m2 of the deep-space facing sides will be coloured matt-black to
radiate excess heat. The remainder of the spacecraft surface will be covered
in multi-layer insulation (MLI). Fluid pipes will be employed to transport
excess heat from the instruments and the Sun facing side to the radiator
panels. The spacecraft will be actively heated during the cold phases where
the instrument power is not heating the spacecraft (safe mode and transfer
from Earth orbit to Venus orbit). A maximum of 260 W will be required to
maintain the spacecraft within the acceptable temperature range. This power
will be drawn from either the batteries or the solar panels and will be fed to
resistive heaters.

###  3.9 Radiation Environment and Shielding

During the mission all spacecraft encounter the complex and harsh radiation
environment in space composed of galactic cosmic rays (GCR) and solar
energetic particles. All of the spacecraft will be briefly exposed to
significantly increased particle fluxes as they cross Earth’s radiation belts
during the orbital transfer. Electronic equipment is particularly vulnerable
to radiation exposure, therefore the spacecraft will be equipped with
radiation hard electronics where possible. Aluminum shielding will be
implemented to reduce the risk of spacecraft failure. In order to account for
the increased radiation exposure in the space environment, a preliminary
shielding study is performed employing the ’radiation sources and effects’
package in the online tool SPENVIS999http://www.spenvis.oma.be, accessed Sept.
2013 (Space Environment Information System). Solar protons are dominating,
hence only solar energetic particles are considered in this analysis. As an
upper limit for the total ionizing dose (TID) the time period between the
launch date and the end of the five-year mission operation (9 years in total)
is considered at a distance from the Sun of 0.7 AU. Figure 3 shows the dose as
a function of aluminum shielding thickness for nine years and for the extended
mission period of 14 years in total using the SHIELDOSE-2 model within the
’long-term radiation doses’ package. All electronic components shall tolerate
a TID of 20 krad, therefore the dose in silicon is calculated. The shielding
will be designed to yield a TID exposure of 10 krad, providing a factor 2
margin (Wertz and Larson [1999]). This results in an estimated thickness of 9
mm for the nominal mission lifetime. Under the assumption that the spacecraft
itself provides a minimum shielding of 3 mm aluminum, all critical components
are shielded with an individual aluminum envelope of 6 mm thickness. The
estimated shielding mass calculates to 43 kg for each spacecraft (Table 5).

Figure 3: Total ionizing dose in silicon vs. aluminum absorber thickness
inside solid aluminum spheres calculated with SPENVIS. The blue lines indicate
the required limit of 10 krad corresponding to a shielding thickness of 9 mm
aluminum.

Studies of the total non-ionizing dose (TNID) are of special importance for
the degradation and therefore for the performance of the solar panels. The
EQFLUX package within SPENVIS has been used to calculate the damage equivalent
fluences. The thereby obtained results for the mission duration is compared to
the respective solar panel properties (Section 3.7). The solar panels
consequently show an estimated loss of efficiency of less than 2% until the
end of the nominal operational time.

##  4 Communication Segment

###  4.1 Communications Scenarios and Subsystem

There are two communication scenarios, each corresponding to a different
mission phase (see Figure 4).

  1. 1.

Safe mode: launch and early orbit phase (LEOP) and cruise phases
The downlink data consists of housekeeping information and science data.
Communication via two wide beam-width low gain antennas (LGAs) using X-band is
considered as baseline for 7.145-7.190 GHz and 8.40-8.45 GHz for up and
downlink respectively (Fortescue et al. [2011]). The LGA provides omni-
directional coverage and telemetry (TM) up to a distance of 0.8 AU from 15 m
antennas on Earth.

  2. 2.

Science mode: nominal and extended phases
A simple pointing mechanism will be used (Noschese et al. [2010]) on the
deployable high gain antenna (HGA), with 1 m diameter and 1.75 degrees beam-
width.

Figure 4: Schematic overview of the onboard communication systems. Each
spacecraft has a full duplex communication system and consists of two
operational modes, a safe mode using two low gain antennas (LGA) and a science
mode using one high gain antenna (HGA).

Additionally, the communication subsystems will include a hot redundant set of
X/X Band transponders and 60 W traveling wave tube amplifier (TWTA) and a RF
distribution unit (RFDU) with diplexers, 3dB couplers and wave guides to
provide the nominal communication with Earth during the duplex spacecraft-
Earth Station connection phases of the mission (Figure 5).

###  4.2 Earth Stations

Science and housekeeping data is received by two antennas in parallel at the
same time from both sides of the Earth, (Figure 5). The uplink and downlink
will use X-band. It is planned to distribute the stations equally over the
Earth to ensure parallel download from the spacecraft. In order to ascertain a
continuous data stream from the spacecraft, six dedicated 15 m dishes will be
used to download the science and housekeeping data.
Due to high priority data downloading and time consuming antenna pointing
tasks, there is no possibility to reuse the existing ESTRACK Deep Space
Network (DSN)
101010http://www.esa.int/Our_Activities/Operations/Estrack_tracking_stations
(accessed Feb. 2014), NASA’s DSN 111111http://deepspace.jpl.nasa.gov/
(accessed Feb. 2014) or the Indian DSN
121212http://www.isro.org/GroundFacilities/trackingfacility.aspx (accessed
Jun. 2014). Figure 6 shows the proposed longitudinal locations of the Earth
stations that are part of the DSN defined for this mission.

Figure 5: Illustration of Earth–spacecraft communication and instrument
activity. The deep space network at Earth consists of six antennas of which
two (E-Stationi and E-Stationi+3) are operating simultaneously as displayed in
the figure. Each Earth station is intermittently in contact with only three
spacecraft, all spacecraft perform in-situ measurements whereas only two
spacecraft also perform remote measurements (i.e. in-situ + imager), depending
on their location. Figure 6: This overview shows the longitudinal locations of
the six 15 m diameter dishes that are part of the deep space network; these
locations ensure simultaneous operation of two dishes for contacting the
spacecraft.

##  5 Operation and Ground Segment

As discussed in Section 3.1, every 19 months there is a three week window to
launch this mission. The mission timeline is depicted in Figure 7, showing the
time separated launch at a nominal launch date of the spacecraft in two sets
of three spacecraft, the individual cruise duration for each spacecraft,
commissioning and nominal operation in the final orbits. Subsequently the
optional extended mission timeline and decommissioning are displayed. The
spacecraft numbering in the figure refers to the numbering in Figure 1.

Figure 7: Life cycle diagram for all spacecraft, showing that the six
spacecraft are not launched simultaneously but in two sets of three, at
different launch dates. The differences in cruise duration for individual
spacecraft (e.g. Spacecraft 2 compared to Spacecraft 1) is required for the 60
degree spacecraft separation. The spacecraft numbering refers to the numbering
introduced in Figure 1.

The mission operations concept shall minimize the costs both in the area of
ground segment tools and facilities as well as in the sharing of manpower and
expertise in the development and operations teams.
It is important to simultaneously approach the spacecraft system-level testing
between the spacecraft manufacturer and the spacecraft operations team,
maximizing the synergy between spacecraft manufacturer and operators in the
preparation of operational documentation, spacecraft user manual, operations
database etc.
The ground segment will rely on six ground stations all around the world with
two of them continuously communicating with the spacecraft (Section 4.2). The
ground stations must be built for the mission since the capacity of any
existing DSN is insufficient for a continuous link. Figure 6 shows the
proposed locations of such a network. The objective is to build stations able
to command the spacecraft and to receive data. For a given time two of them
will be absolutely dedicated to communication with CARETAKER. However there
will always be four stations available for other purposes. The ground segment
organization including the further data transfer and processing is presented
in Figure 8.

Figure 8: Illustrating data handling at Earth. Raw data is obtained through
six ground stations and is collected by a mission operation center that
forwards the data to a data processing center. This center processes the data
and provides the SSCC Space Weather Coordination Center with data according to
their format, additionally, the data is also archived for the scientific
community. It is eventually the SSCC that provides the public with a forecast.

##  6 Data Processing

###  6.1 Mission Operation Center

The CARETAKER mission operation center (MOC) will be in charge of all
telecommand and telemetry operations of the mission. Communication with the
six spacecraft will be provided by two ground stations at a time through the
CARETAKER Network 24/7. The detailed communication schedule is given in Figure
9. The two antennas that communicate with the spacecraft at a given time
download data from four spacecraft every 15 minutes. For each such interval,
each antenna will download data from two spacecraft, one providing imaging and
in-situ data (C+IS), the second only sending in-situ (IS) data. While data
from the two spacecraft sending imaging (Si+1 and Si-1 in the figure) data and
the spacecraft within the Sun–Earth line (Si) is downloaded in every interval,
the other three spacecraft are not necessarily addressed every 15 minutes.
The operation center will perform 0-level data processing as well as backing
up and storing the previous 30 days of the processed data. Data will be
transmitted continuously to the data processing center. Data processing for
in-situ measurements is likely to take less than one minute so that the
requirements of the SSCC are fulfilled.

Figure 9: Data collection from the spacecraft: Colors stand for the data
collected by each spacecraft communicating with the ground segment. Data from
four spacecraft is downloaded every 15 minutes with two antennas on Earth. For
each such interval the downlink time from spacecraft sending imaging and in-
situ (C+IS) data takes nine minutes. The antennas then switch to the next
spacecraft through a one minute pointing procedure and download only the in-
situ (IS) measurements from the other spacecraft.

###  6.2 Data Processing Center

The data processing center shall perform different tasks:

  1. 1.

Calibration, validation and processing of the data for the early phase of the
mission.

  2. 2.

Provide algorithms, methods and models to a warning unit which shall be able
to give standards compatible data to the SSCC. The time between taking the
measurement and delivering the data products should meet those standards (see
Section 6.3).

  3. 3.

Perform higher level processing on the data and make it available on the
Internet.

  4. 4.

Archive all data for further and long term investigation (over a semi-solar
cycle as it is possible with the minimum life time of the mission).

###  6.3 Data Products

Given the extraordinary position of the spacecraft constellation, the
processed data is to become a new reference for space weather event warning as
well as premium scientific content. For the in-situ measurement, a fifteen
minutes range between measurement and data product will be ensured as required
by the SSCC. For 3D-modelling of coronagraphic data and CMEs, an extended time
line is proposed: those data should be available within a maximum of sixty
minutes after remote measurement. Further processing will be performed later
to make the data fully exploitable by the scientific community in the long
range. The whole content will be available on the internet.

##  7 Budgets

Budgets concerning mass and power of the spacecraft are analyzed considering
the mission profile and the lifetime of the mission. The launcher fairing size
and mass constraints were taken into account and all spacecraft were
investigated individually. Margins for each subsystem are applied following
the ESA Margin Philosophy for Science Assessment Studies (SCI-PA/2007/022,
[2007]). According to this, off-the-shelf components that are implemented
without any changes are considered with a margin of 5%, a 10% margin is added
for off-the-shelf items with minor modifications and a 20% margin is used in
case of major modifications, new designs and new developments. Additional 20%
system margin are added in the end to the complete system. The resulting mass
and power budgets are summarized in Table 5 and Table 6, respectively.

The propellant in the mass budget includes additional fuel for AOCS/RCS and
safe mode recovery as well as a margin of 20%. As this budget refers to a
single spacecraft, the total launch mass within one Ariane 5 rocket adds to
2812 kg, including three spacecraft plus launch adapters.

Table 5: Mass budget for a single spacecraft including the relevant margins. Subsystem | Mass | Margin | Mass total
---|---|---|---
| [kg] |  | [kg]
Power | 43.00 | 2.75 | 45.75
Payload | 31.00 | 2.96 | 33.96
Communications | 28.20 | 5.22 | 33.42
Onboard Data Handling / Avionics | 15.00 | 0.75 | 15.75
AOCS | 92.86 | 2.45 | 95.31
Thermal Control | 24.00 | 1.40 | 35.20
Additional Shielding | 42.67 | 0.00 | 42.67
Chemical Propulsion System (dry mass) | 37.60 | 3.63 | 41.23
Harness (5%) | 15.72 | 0.00 | 15.72
Structure (20% of dry mass) | 66.01 | 0.00 | 66.01
TOTAL (dry, without system margin) | 396.05 | 22.18 | 418.23
System Margin (20%) | 79.21 |  |
TOTAL (dry, with margin) |  |  | 497.44
Propellant |  |  | 306.2
TOTAL (wet mass) |  |  | 803.64
Table 6: Power budget for a single spacecraft including the relevant margins. Subsystem | Power Consumption | Margin | Power Consumption total
---|---|---|---
| [W] |  | [W]
Power | 48.0 | 0.0 | 48.0
Payload | 18.5 | 1.8 | 20.3
Communications | 165.0 | 16.3 | 181.3
Onboard Data Handling / Avionics | 17.0 | 0.85 | 17.85
AOCS | 198.6 | 99.3 | 297.9
Thermal Control | 260 | 13 | 273
Chemical Propulsion System | 5.0 | 1.0 | 6.0
TOTAL (without margin) | 712.1 | 132.2 | 796.3
System Margin (20%) | 142.4 |  |
TOTAL (with margin) |  |  | 986.67

##  8 Risk Analysis

As per ECSS (European Cooperation on Space
Standardization131313http://www.ecss.nl/) standard operating procedures,
outlined in ECSS-M-ST-80 ([2008]), a risk matrix was developed to classify
potential risks for the mission. There are two main sub-classifications under
which each risk will be listed; their likelihood (A [low] to E [high]) and
their severity (1 [low] to 5 [high]). Risk is calculated as a combination of
severity and likelihood; the result ranges from A1 [very low] to E5 [very
high]. The high and very high risks are of the utmost concern and if not
possible to reduce their classification they may represent reasons for mission
postponement until such time as the level can be reduced. In extreme cases it
can also lead to overall mission cancellation. Due to the complexity of this
multi-spacecraft mission Table 7 shows the risk analysis of a subset of cases
and their respective countermeasures.

The dependence of the primary mission objective on the performance of all six
spacecraft produces the most severe risk: the failure of only one spacecraft
compromises the mission, as due to the incomplete coverage of the in-situ
measurements, a CME heading to Earth might be missed. This problem could be
mitigated by employing a larger number of spacecraft, which would increase
costs significantly.

Table 7: Risk examples for the mission classified following ECSS-M-ST-80 ([2008]) for their their likelihood (A [low] to E [high]) and their severity (1 [low] to 5 [high]). Possible Scenario | Risk Index |  Proposed action
---|---|---
Launch fails | A5 | low |  No action - mission failed
Failure of one spacecraft | B5 | medium |  Continue mission with remaining spacecraft, focus on scientific research
Launch misses time window | C2 | low |  Wait 19 months for the next time window
Miss the trajectory for GAM | B5 | medium |  mission failed
Collisions of spacecraft during deployment | B4 | low |  Communicate to spacecraft and try to recover possible orbit problems with propulsion system
Not enough data rate during safe mode to ensure communication between spacecraft and ground station | B3 | low |  Spacecraft are on a circular orbit pointing to the Sun. Wait until spacecraft is closer to the ground station, where higher data rate is possible. Two small emergency antennas (70 m dish ground station)
Damage of sensitive optics of a coronagraph during space flight | B3 | low |  Analyze the influence of the damage and try to reduce it using image processing at the science ground station
Radiation damage of the measurement system caused by GCR background and SEPs | D2 | medium |  Expected damage, which limits the lifetime of the measurement system. Possible to move spacecraft slightly in order to protect them.
If the spacecraft enters safe mode in a 40 degrees range on the backside of the Sun, connection will not be possible in this area | B3 | low |  Ask communication specialist. Communicate with spacecraft, when it leaves the 40 degrees area behind the Sun and recover it
Blackout of measurements device | A3 | low |  Ask payload specialist. Check the whole spacecraft housekeeping system, special communication time to the spacecraft
Explosion of some propulsion tank after deployment | A4 | low |  Check damage, continue mission with remaining spacecraft

##  9 Cost Analysis

In comparison to typical ESA science missions, the CARETAKER mission is
expensive. This is due to the mission objective of creating a forecast system
which is not purely scientific, but shall rather serve as a warning system in
order to prevent possibly catastrophic consequences on Earth, of which some
have been discussed before. Such a warning system however requires that events
with a possible impact can be observed reliably. This can be only achieved
with significantly more resources than are available for purely scientific
missions. A smart resource management is required to reduce the costs of the
mission in a reliable frame. The total costs are estimated to lie between 1.2
billion Euro and 1.4 billion Euro. About 50% of the budget will be spent on
design and construction of spacecraft platform and payload. This would be much
higher if the spacecraft were not designed to be identical. Approximately 30%
of the costs will be taken to provide the infrastructure of the ground segment
and the operation service. This part is relatively large, since the mission
requires its own deep space network. The main uncertainties within the cost
approximation are caused by the ground segment. There is no experience for
multi-spacecraft missions in deep space with a continuous communication
coverage. The architecture of the spacecraft allows to pack the complete set
of spacecraft into two Ariane 5 launchers. The launch takes the smallest
portion of the cost with 20%.
In comparison, the JUICE mission (Grasset et al. [2013]), costs ESA 870
million Euros and Rosetta141414http://sci.esa.int/rosetta/ is estimated to
cost 1.4 billion Euros (according 2014 economic
conditions151515http://www.esa.int/Our_Activities/Space_Science/Rosetta/Frequently_asked_questions,
What is the total mission cost?). CARETAKER can only be funded by an
international collaboration of space agencies and economic companies.

##  10 Conclusion

The presented mission concept consists of six identical spacecraft, located in
a heliocentric orbit at 0.72 AU at a separation angle of 60 degrees. It is
aimed to become fully operational (i.e. with all spacecraft in place) by 2030,
while partial operation with less spacecraft can already begin three years
prior to this. The nominal mission life time with all spacecraft is 5 years
with a possible extension of additional 5 years, covering an entire solar
cycle.
The mission’s primary objective is to provide a near real-time information-
service for space weather, based on the physical properties of solar CMEs. The
information will consist of both in-situ and remote measurements. The mission
is optimized to monitor CMEs with an angular extent of 60 degrees and larger,
since these have the most potential to cause geomagnetic storms upon impact
with Earth. The most important information will contain the propagation
velocity and direction of a CME, as well as the spatial resolution of the
magnetic field orientation and strength. This information helps in predicting
the time of impact at Earth, and the potential severity of a geomagnetic storm
if the CME proves to be geoeffective.
The proposed instruments to be employed on each spacecraft originate from
other space missions with the intention to reduce mission costs and to
accelerate overall the mission design. The spacecraft will perform in-situ
measurements using the Solar Wind Analyzer instrumentation package from the
_Solar Orbiter_ mission, together with flux gate magnetometers from _Venus
Express_ , while performing remote measurements with a coronagraph from
_STEREO_.
The mission’s secondary objective is to provide data helping to improve CME
models. As for the choronagraph the present limitation for the
3D-reconstruction for the CMEs is given by the two _STEREO_ spacecraft and
_SOHO_ which are not always positioned in a useful way to provide the
necessary view angles (Mierla et al. [2010]). For the in-situ measurements
models exist for determining the flux rope geometry, which in the case of one
spacecraft this represents a 1D local cut through a global 3D structure.
Having more than one spacecraft sampling the CME could help to constrain its
global structure (Möstl et al. [2012]).
Communication is realized via a dedicated deep space network consisting of six
ground stations equipped with 15 m dish antennas, with continuous operation of
two ground stations simultaneously, each of them alternating contact between
three spacecraft. Received data will be processed by a data processing center
that forwards the processed data to the Space Weather Coordination Center who
eventually informs the general public through a space weather forecast. The
data processing center will additionally archive the data for the scientific
community, for updating scientific models.

The obtained information will result in major advances in space weather
forecasting time and the scientific modelling of space weather. The data
provided can be combined with data from other missions running at the same
time, which has the potential to further expand the solar physics
understanding. Furthermore, the proposed mission will not only monitor the
solar wind in the Sun – Earth direction, but it also offers the potential for
360 degrees warnings, providing information to a warning system that protects
astronauts in future manned missions (e.g. to Mars). The mission also offers
unprecedented possibilities to study the Sun’s corona and the inner
heliosphere. Any CME with an angular extent of 60 degrees or larger will be
monitored in-situ and remotely, leading to a better understanding of a CME’ s
morphology and its spatial as well as temporal distribution. Our view of the
Sun’s corona will change from 2D to 3D, giving rise to new unexpected
discoveries. Additionally, the gravity assist with Venus presents a case study
on its own and allows for analysis of Venus’s magnetic field through
multipoint in-situ measurements.
This mission will not only provide information vital to protect the health,
economy, and modern technology of the human society. It will also bring space
weather monitoring to a higher level, and propel advancement in the scientific
community.

###### Acknowledgements.

The authors acknowledge the contribution of Daniel Schmid, Ioanna Patsou,
Joseph Nguyen, and Markus Suter, during the Alpbach summer school and are very
grateful for the expertise of Joe Zender and Rumi Nakamura, who proved to be
invaluable during the Alpbach summer school.
Furthermore the authors are greatly indebted for the advice and efforts from
Peter Falkner, Christian Erd, and Markus Hallmann, as with their help this
mission developed to its current shape.
Nikolaos Perakis and Francesco Checco Gini offered very good support regarding
orbital calculations, and Viktor Andersson for helped with thermal aspects.
The reviewing of the article by Günther Reitz and Luciano Rodriguez proved
valuable and is highly appreciated.
Aditionally, the authors are grateful for all the organisational effort of
Michaela Gitsch, who is of great help to young scientists by introducing them
into the (new) field of space science and engineering.
Finally, the authors acknowledge the financial support by the European Space
Agency (ESA), the Aeronautics and Space Agency of Austria FFG, the
International Space Science Institute (ISSI), and Austrospace, the association
of Austrian space industries and research institutions, all of them making the
Alpbach summer school financially possible. In addition, all participants of
the summer school were funded by organisations from their respective member
states.
The editor thanks Norma B. Crosby and an anonymous referee for their
assistance in evaluating this paper.

## References

  * [1981] Akasofu, S. I., Energy coupling between the solar wind and the magnetosphere, Space Sci. Rev., 28 (2), 121-190, 1981.
  * [2010] Balogh, A., Planetary Magnetic Field Measurements: Missions and Instrumentation, Space Sci. Rev., 152 (1-4), 23-97, DOI: 10.1007/s11214-010-9643-1, 2010.
  * [2007] Biele, J., and S. Ulamec, Capabilities of Philae, the Rosetta Lander, Space Sci. Rev., 138 (1-4), 275-289, DOI: 10.1007/s11214-007-9278-z, 2007.
  * [1995] Brueckner, G.E., R.A. Howard, M.J. Koomen, C.M. Korendyke, D.J. Michels, et al., The Large Angle Spectroscopic Coronagraph (LASCO), Solar Physics, 162 (1-2), 357-402, DOI: 10.1007/BF00733434, 1995.
  * [2001] Burlaga, L., Magnetic fields and plasmas in the inner heliosphere: Helios results, Planet. Space Sci., 49 (14-15), 1619-1627, DOI: 10.1016/S0032-0633(01)00098-8, 2001.
  * [1975] Burton, R.K., R.L. McPherron, and C.T. Russell, An empirical relationship between interplanetary conditions and Dst, J. Geophys. Res. 80 (31), 4204-4214, DOI: 10.1029/JA080i031p04204, 1975.
  * [1996] Dermott, J.C., Eagle-Picher Ind. Inc., Joplin, MO, USA, J.A. DeGruson, and L.E. Miller, Large Capacity $NiH_{2}$, battery cells for advanced, high-power spacecraft applications, Energy Conversion Engineering Conference. IECEC 96., Proceedings of the 31st Intersociety, 1, 380-383, DOI: 10.1109/IECEC.1996.552911, 1996.
  * [2012] DeRosa, M.L., A.S. Brun, and J.T. Hoeksema, Solar magnetic field reversals and the role of dynamo families, ApJ, 757, 96-112, DOI: 10.1088/0004-637X/757/1/96, 2012.
  * [1995] Domingo, V., B. Fleck, and A. Poland, SOHO: The Solar and Heliospheric Observatory, Space Sci. Rev., 72 (1-2), 81-84, DOI: 10.1007/BF00733425, 1995.
  * [1961] Dungey, J. W., Interplanetary Magnetic Field and the Auroral Zones, Phys. Rev. Lett., 6 (2), 47-48, DOI: 10.1103/PhysRevLett.6.47, 1961
  * [2011] Fortescue, P., G. Swinerd, and J. Stark, Spacecraft Systems Engineering, 4th Edition, Wiley, ISBN: 978-0-470-75012-4, 2011.
  * [1991] Glaßmeier, K.H., and M. Scholer, Plasmaphysik im Sonnensystem, BI Publisher Wetenschap, Mannheim (Germany), 378 p., ISBN 3-411-15151-X, 1991.
  * [1994] Gonzalez, W.D., J.A. Joselyn, Y. Kamide, H.W. Kroehl, G. Rostoker, B.T. Tsurutani, and V.M. Vasyliunas, What is a geomagnetic storm?, J. Geophys. Res., 99, 5771-5792, DOI: 10.1029/93JA02867, 1994.
  * [2003] Gopalswamy, N., Coronal mass ejections: Initiation and detection, Adv. Space Res., 31 (4), DOI: 10.1016/S0273-1177(02)00888-8, 869-881, 2003.
  * [2008] Gopalswamy, N., Solar connections of geoeffective magnetic structures, J. Atmos. Sol. Terr. Phys., 4th Asia Oceania Geosciences Society meeting, 70 (17), 2078-2100, DOI: 10.1016/j.jastp.2008.06.010, 2008.
  * [2004] Gopalswamy, N., M. Shimojo, W. Lu, S. Yashiro, K. Shibasaki, and R. A. Howard, On coronal streamer changes, Adv. Space Res., 33 (5), 676-680 DOI: 10.1016/S0273-1177(03)00236-9, 2004.
  * [1990] Gosling, J.T., S.J. Bame, D.J. McComas, and J.L. Phillips, Coronal mass ejections and large geomagnetic storms, Geophys. Res. Lett., 17 (7), 901-904, DOI: 10.1029/GL017i007p00901, 1990.
  * [2013] Grasset, O., M.K. Dougherty, A. Coustenis, E.J. Bunce, C. Erd, et al., JUpiter ICy moons Explorer (JUICE): An ESA mission to orbit Ganymede and to characterise the Jupiter system, Planet. Space Sci., 78, 1-21, DOI: 10.1016/j.pss.2012.12.002, 2013.
  * [2012] Gulisano, A. M., P. Démoulin, S. Dasso, L. Rodriguez, Expansion of magnetic clouds in the outer heliosphere, A&A, 543, A107, DOI: 10.1051/0004-6361/201118748
  * [1994] Hablani, H.B., Sun-Tracking Commands and Reaction Wheel Sizing with Configuration Optimization, Journal of Guidance, Control, and Dynamics, 17 (4), 1994.
  * [1985] Howard, R.A., N.R. Sheeley Jr, M.J. Koomen, and D.J. Michels, Coronal Mass Ejections: 1975-1981, J. Geophys. Res., 90, 8173-8191, DOI: 10.1029/JA090iA09p08173, 1985.
  * [2008] Howard, R.A., J.D. Moses, A. Vourlidas, J.S. Newmark, D.G. Socker, et al., Sun Earth Connection Coronal and Heliospheric Investigation (SECCHI), Space Sci. Rev., 136 (1-4), 67-115, DOI:10.1007/s11214-008-9341-4, 2008.
  * [1984] Hundhausen, A.J., C.B. Sawyer, L. House, R.M.E. Illing, and W.J. Wagner, Coronal mass ejections observed during the solar maximum mission - Latitude distribution and rate of occurrence, J. Geophys. Res. (May 1, 1984) (ISSN 0148-0227), 89, 2639-2646, 1984.
  * [2008] Kaiser, M.L., T.A. Kucera, J.M. Davila, O.C. St. Cyr, M. Guhathakurta, and E. Christian, The STEREO Mission: An Introduction, Space Sci. Rev., 136, 5-16, DOI: 10.1007/s11214-007-9277-0, 2008.
  * [1999] Lindsay, G.M., C.T. Russell, and J.G. Luhmann, Predictability of Dst index based upon solar wind conditions monitored inside 1 AU, J. Geophys. Res. A: Space Phys. (1978- 2012), 104, A5, 10335-10344, DOI: 10.1029/1999JA900010, 1999.
  * [2004] Luhmann, J.G., S.C. Solomon, J.A. Linker, J.G. Lyon, Z. Mikic, D. Odstrcil, W. Wang, and M. Wittberger, Coupled model simulation of a Sun-to-Earth space weather event, J. Atmos. Sol. Terr. Phys., 66 (15), 1243-1256, DOI: 10.1016/j.jastp.2004.04.005, 2004.
  * [2006] Manoharan, P.K., Evolution of coronal mass ejections in the inner heliosphere: A study using white-light and scintillation image, Solar Physics, 235, 345-368, DOI: 10.1007/s11207-006-0100-y, 2006.
  * [2010] Manoharan, P.K., Ooty interplanetary scintillation – remote-sensing observations and analysis of coronal mass ejections in the heliosphere, Solar Physics, 265, 137-157, DOI: DOI: 10.1007/s11207-010-9593-5, 2010.
  * [2005] Marsch, E., R. Marsden, R. Harrison, R. Wimmer-Schweingruber, and B. Fleck, Solar Orbiter, mission profile, main goals and present status, Adv. Space Res., 36, 1360-1366, DOI: 10.1016/j.asr.2004.11.012, 2005.
  * [2007] Meyer-Vernet, N., Basics of the Solar Wind, Cambridge Atmospheric and Space Science Series, Cambridge University Press, ISBN-13 978-0-521-81-420-1, 2007.
  * [2010] Mierla, M., B. Inhester, A. Antunes, Y. Bourisier, J.P. Byrne, et al., On the 3-D reconstruction of Coronal Mass Ejections using coronagraph data, Ann. Geophys, 28, 1-13, DOI: 10.5194/angeo-28-203-2010, 2010.
  * [2012] Möstl, C., C.J. Farrugia, E.K.J. Kilpua, L.K. Jian, Y. Liu, et al., Multi-point Shock and Flux Rope Analysis of Multiple Interplanetary Coronal Mass Ejections, around 2010 August 1 in the Inner Heliosphere, ApJ, 758, 18pp, DOI: 10.1088/0004-637X/758/1/10, 2012.
  * [2010] Noschese, P., M. Milano, E.F. Zampolini, G. Mascolo, D. Landi, et al., Bepicolombo Mission to Mercury, Design Status of the High Temperature High Gain Antenna, 32nd ESA Antenna Workshop on Antennas for Space Applications, Noordwijk, The Netherlands, 2010.
  * [1995] Ogilvie, K., D. Chornay, R. Fritzenreiter, F. Hunsaker, J. Keller, et al., SWE, a comprehensive plasma instrument for the WIND spacecraft, Space Sci. Rev., 71, 55-77, DOI: 10.1007/BF00751326, 1995.
  * [2002] Reames, D.V., Magnetic topology of impulsive and gradual solar energetic particle events Astrophys. J., 571, L63-L66, 2002.
  * [2008] Reitz, G., Characteristic of the radiation field in low earth orbit and in deep space Z. Med. Phys., 18 (4), 233-243, DOI: 10.1016/j.zemedi.2008.06.015, 2008.
  * [2004] Richardson, I.G., and H.V. Cane, Identification of interplanetary coronal mass ejections at 1 AU using multiple solar wind plasma composition anomalies, J. Geophys. Res., 109, A09104, DOI: 10.1029/2004JA010598, 2004.
  * [1998] Smith, C.J. L’Heureux, N. Ness, M. Acuña, L. Burlaga, J. Scheifele, C. Russell, R. Mewaldt, and T. Rosenvinge, (Eds.) The Ace Magnetic Fields Experiment, Space Sci. Rev., 86 (1) 613-632, DOI: 10.1023/A:1005092216668, 1998.
  * [2000] St. Cyr, O.C., M.A. Mesarch, H.M. Maldonado, D.C. Folta, A.D. Harper, et al. , Space Weather Diamond: a four spacecraft monitoring system, J. Atmos. Sol. Terr. Phys., 62 (14), 1251-1255, DOI: 10.1016/S1364-6826(00)00069-9, 2000.
  * [2009] Thernisien, A.; A. Vourlidas, and R. Howard, Forward Modeling of Coronal Mass Ejections Using STEREO/SECCHI Data, Sol. Phys., 256 (1-2), 111-130, DOI: 10.1007/s11207-009-9346-5, 2009.
  * [2006] Titov, D.H. Svedhem, D. Koschny, R. Hoofs, S. Barabash, et al., Venus Express science planning, Planet. Space Sci., 54 (13-14), 1279-1297, DOI: 10.1016/j.pss.2006.04.017, 2006.
  * [2005] Tóth, G., I.V. Sokolov, T.I. Gombosi, D.R. Chesney, C.R. Clauer, et al., Space Weather Modeling Framework: A new tool for the space science community, J. Geophys. Res. A: Space Phys., 110, A12, DOI: 10.1029/2005JA011126, 2005.
  * [2012] Wagner, M., S. Airey, G. Piret, and P. Le, New Reaction Wheel Characterisation Test Facility (RCF), 35th Annual AAS Guidance and Control Conference, AAS 12-077, Breckenridge, Colorado, 2012.
  * [1999] Wertz, J.R., and W.J. Larson, Space Mission Analysis and Design, Third Edition, Space Technology Library, ISBN: 978-1881883104, 1999.
  * [1996] West, J.L. ”NOAA/DoD/NASA Geostorm Warning Mission.” Jet Propulsion Lab, California Institute of Technology, JPL D-13986, Pasadena, CA, 1996.
  * [2004] West, J. ”The Geostorm Warning Mission: Enhanced Opportunity Based on New Technology.” 14th AAS/AIAA Space Flight Mechanics Conference, Maui, HI, AAS. Vol. 102. 2004.
  * [2007] Zhang, T.L., G. Berghofer, W. Magnes, M. Delva, W. Baumjohann, et al., MAG: The Fluxgate Magnetometer of Venus Express, ESA Special Publication, SP 1295, 1-10, 2007.
  * [2009] COST 724 final report Developing the scientific basis for monitoring, modelling and predicting SpaceWeather, Introduction to COST 724, p.11, 2009.
  * [2011] CENTRA Technology, OECD/IFP Futures Project on ”Future Global Shocks” , ” Geomagnetic Storms”, CENTRA Technology, Office of Risk Management and Analysis, United States Department of Homeland Security, 2011.
  * [2007] SCI-PA/2007/022, Margin Philosophy for Science Assessment Studies, ESA, ESTEC, 2007.
  * [2011] Solar Orbiter Definition Study Report (Red Book), ESA/SRE(2011)14, Issue 1.0
  * [2008] Space project management – Risk management, Eurpean Cooperation for Space Standardization ECSS, ECSS-M-ST-80, 2008.