**Solar Orbiter**

![](media/image1.jpeg){width="5.450694444444444in"
height="5.7756944444444445in"}**A High-Resolution Mission to the Sun and
Inner Heliosphere**

Assessment Study Report

**July 2000**

SCI(2000)6

# Foreword {#foreword .Chapter}

The Solar Orbiter Mission was proposed to ESA by an international team
of scientists led by E. Marsch in response to the call for mission
proposals for two flexible missions (F2 and F3) and submitted on January
27, 2000. In March 2000 ESA\'s Space Science Advisory Committee
recommended the Solar Orbiter for further study based on a
pre-assessment study performed in 1999 in the framework of ESA's Solar
Physics Planning Group activities.

This document is the result of the "delta" assessment study, carried out
by representatives of the European solar and heliospheric physics
communities, in close co-operation with ESA staff.

Study team members:

**E. Marsch**, Max-Planck-Institut für Aeronomie, Katlenburg-Lindau,
Germany

**E. Antonucci**, Osservatorio Astronomico di Torino, Pino Torinese,
Italy

**P. Bochsler**, University of Bern, Switzerland

**J.-L. Bougeret**, Observatoire de Paris, Meudon, France

**R. Harrison**, Rutherford Appleton Laboratory, Chilton, UK

**R. Schwenn**, Max-Planck-Institut für Aeronomie, Katlenburg-Lindau,
Germany

**J.-C. Vial**, Institut d\'Astrophysique Spatiale, CNRS-Université de
Paris-Sud, France

ESA staff members:

Study Scientists: **B. Fleck**, ESA/GSFC, Greenbelt, Maryland, USA

**R. Marsden**, ESA/ESTEC, Noordwijk, The Netherlands

Study Manager: **O. Pace**, ESA/ESTEC, Noordwijk, The Netherlands

Solar System Mission Coordinator: **M. Coradini**, ESA HQ, Paris, France

**Y. Langevin,** Institut d\'Astrophysique Spatiale, Orsay, France has
provided important assistance on the Solar Orbiter mission design and
analysis.

Valuable contributions and comments on the science objectives and
instrumentation were provided by:

**T. Appourchaux**, ESA Space Science Department, ESTEC, Noordwijk, The
Netherlands

**R. Bruno**, Istituto Fisica Spazio Interplanetario del CNR, Roma,
Italy

**M. Decaudin**, Institut d\'Astrophysique Spatiale, CNRS-Université de
Paris-Sud, France

**M. Collados**, Instituto de Astrofisica de Canarias, Spain

**A. Fedorov,** Institute for Space Research, Moscow, Russia

**S. Fineschi**, Osservatorio Astronomico di Torino, Pino Torinese,
Italy

**C. Fröhlich**, Physikalisch-Metorologisches Observatorium, Davos,
Switzerland

**L. Gizon**, Stanford University, California, USA

**K. Goetz,** University of Minnesota, USA

**M. Hilchenbach**, Max-Planck-Institut für Aeronomie, Lindau, Germany

**J.-F. Hochedez**, Observatoire Royal de Belgique, Brussels, Belgium

**P. Lemaire**, Institut d\'Astrophysique Spatiale, CNRS-Université de
Paris-Sud, France

**F. Lefeuvre,** LPCE, Orleans, France

**B. Lites**, High Altitude Observatory, Boulder, USA

**O. von der Lühe**, Kiepenheuer Institut für Sonnenphysik, Freiburg,
Germany

**V. Martinez Pillet**, Instituto de Astrofisica de Canarias, Spain

**R.B. McKibben**, University of Chicago, USA

**G. Naletto**, Università di Padova, Padova, Italy

**E. Priest**, University of St. Andrews, UK

**M. Romoli**, Universita\` di Firenze, Firenze, Italy

**H. Rucker,** Universtität Graz, Austria

**E. Sawyer**, Rutherford Appleton Laboratory, Chilton, UK

**W. Schmidt**, Kiepenheuer Institut für Sonnenphysik, Freiburg, Germany

**S. Solanki**, Max-Planck-Institut für Aeronomie, Lindau, Germany

**P. Wurz**, University of Bern, Switzerland

#   {#section .Chapter .unnumbered}

# Table of Contents {#table-of-contents .Chapter}

Foreword [i](#foreword)

Table of Contents [iii](#table-of-contents)

1 EXECUTIVE SUMMARY [1](#executive-summary)

2 Scientific Rationale [4](#scientific-rationale)

2.1 INTRODUCTION [4](#introduction)

2.2 THE SUN\'S MAGNETISED PLASMA: CLOSE-UP OBSERVATIONS OF THE SOLAR
ATMOSPHERE
[4](#the-suns-magnetised-plasma-close-up-observations-of-the-solar-atmosphere)

2.2.1 Photospheric magnetic flux elements
[5](#photospheric-magnetic-flux-elements)

2.2.2 Basic processes and fundamental scales in the magnetised solar
atmosphere
[5](#basic-processes-and-fundamental-scales-in-the-magnetised-solar-atmosphere)

2.2.3 Waves in the corona [6](#waves-in-the-corona)

2.2.4 Coronal magnetic field [6](#coronal-magnetic-field)

2.3 LINKING THE PHOTOSPHERE AND CORONA TO THE HELIOSPHERE: CO-ROTATION
OBSERVATIONS
[7](#linking-the-photosphere-and-corona-to-the-heliosphere-co-rotation-observations)

2.3.1 Global solar corona and solar wind
[7](#global-solar-corona-and-solar-wind)

2.3.2 Global coronal sources of the solar wind
[8](#global-coronal-sources-of-the-solar-wind)

2.3.3 Magnetic network [9](#magnetic-network)

2.3.4 Boundaries and fine structures in the corona and solar wind
[9](#boundaries-and-fine-structures-in-the-corona-and-solar-wind)

2.3.5 Connections between the internal plasma states of the solar corona
and the solar wind
[10](#connections-between-the-internal-plasma-states-of-the-solar-corona-and-the-solar-wind)

2.3.6 Coronal and solar wind abundances and fractionation effects
[10](#coronal-and-solar-wind-abundances-and-fractionation-effects)

2.3.7 Coronal transients [11](#coronal-transients)

2.4 PARTICLES AND FIELDS: *IN-SITU* MEASUREMENTS IN THE INNER
HELIOSPHERE
[12](#particles-and-fields-in-situ-measurements-in-the-inner-heliosphere)

2.4.1 Microstate of the interplanetary solar wind
[12](#microstate-of-the-interplanetary-solar-wind)

2.4.2 Solar wind ions as tracers of coronal structures
[13](#solar-wind-ions-as-tracers-of-coronal-structures)

2.4.3 Magnetohydrodynamic turbulence
[13](#magnetohydrodynamic-turbulence)

2.4.4 Acceleration and transport of solar energetic particles
[14](#acceleration-and-transport-of-solar-energetic-particles)

2.4.5 Neutral particles from the Sun
[15](#neutral-particles-from-the-sun)

2.4.6 Circumsolar and interplanetary dust
[16](#circumsolar-and-interplanetary-dust)

2.4.7 Solar neutrons [17](#solar-neutrons)

2.4.8 Solar radio emission [17](#solar-radio-emission)

2.5 THE SUN\'S POLAR REGIONS AND EQUATORIAL CORONA: EXCURSION OUT OF THE
ECLIPTIC
[17](#the-suns-polar-regions-and-equatorial-corona-excursion-out-of-the-ecliptic)

2.5.1 The Sun\'s polar magnetic field and the dynamo
[18](#the-suns-polar-magnetic-field-and-the-dynamo)

2.5.2 Polar rotation and internal flows
[18](#polar-rotation-and-internal-flows)

2.5.3 Polar observations of CMEs [19](#polar-observations-of-cmes)

2.5.4 Solar luminosity variations [20](#solar-luminosity-variations)

2.5.5 Polar observation of the fast solar wind
[20](#polar-observation-of-the-fast-solar-wind)

2.5.6 Secular variation of the heliospheric magnetic flux
[20](#secular-variation-of-the-heliospheric-magnetic-flux)

3 SCIENTIFIC PAYLOAD [22](#scientific-payload)

3.1 MEASUREMENT REQUIREMENTS [22](#measurement-requirements)

*3.2* *In-situ* HELIOSPHERIC INSTRUMENTS
[22](#in-situ-heliospheric-instruments)

3.2.1 Solar Wind Plasma Analyser (SWA)
[22](#solar-wind-plasma-analyser-swa)

3.2.2 Radio and Plasma Waves Analyser (RPW)
[25](#radio-and-plasma-waves-analyser-rpw)

3.2.3 Coronal Radio Sounding (CRS) [29](#coronal-radio-sounding-crs)

3.2.4 Magnetometer (MAG) [30](#magnetometer-mag)

3.2.5 Energetic Particle Detector (EPD)
[30](#energetic-particle-detector-epd)

3.2.6 Dust Detector (DUD) [32](#dust-detector-dud)

3.2.7 Neutral Particle Detector (NPD)
[33](#neutral-particle-detector-npd)

3.2.8 Neutron Detector (NED) [34](#neutron-detector-ned)

3.2.9 Heliospheric instruments summary
[36](#heliospheric-instruments-summary)

3.3 REMOTE SENSING SOLAR INSTRUMENTS
[36](#remote-sensing-solar-instruments)

3.3.1 General considerations [36](#general-considerations)

3.3.2 Visible-Light Imager and Magnetograph (VIM)
[36](#visible-light-imager-and-magnetograph-vim)

3.3.3 EUV spectrometer (EUS) [42](#euv-spectrometer-eus)

3.3.4 EUV imager (EXI) [47](#euv-imager-exi)

3.3.5 Ultraviolet and Visible-light Coronagraph (UVC)
[51](#__RefHeading___Toc489691200)

3.3.6 Radiometer (RAD) [54](#radiometer-rad)

3.3.7 Other potential solar instruments
[55](#other-potential-solar-instruments)

3.3.8 Solar remote-sensing instruments summary
[56](#solar-remote-sensing-instruments-summary)

4 Mission profile [57](#__RefHeading___Toc489691204)

5 System design [60](#system-design)

5.1 Design requirements and drivers
[60](#design-requirements-and-drivers)

5.2 Main system design features [60](#main-system-design-features)

5.2.1 Requirements and constraints [61](#requirements-and-constraints)

5.2.2 Spacecraft baseline design [61](#spacecraft-baseline-design)

5.2.3 Thermal Design [64](#thermal-design)

5.2.4 Power Supply [65](#power-supply)

5.3 Data collection and transmission
[66](#data-collection-and-transmission)

5.4 Radiation [68](#radiation)

5.5 Mass Budgets [68](#mass-budgets)

TOTAL LAUNCH MASS [68](#__RefHeading___Toc489691215)

6 Subsystem design [69](#__RefHeading___Toc489691216)

6.1 Attitude and orbit control [69](#attitude-and-orbit-control)

6.2 Reaction Control System [70](#reaction-control-system)

6.3 Propulsion [70](#propulsion)

6.4 Structure [71](#structure)

6.5 Thermal Control [71](#thermal-control)

7 Ground systems and operations [73](#ground-systems-and-operations)

7.1 Ground segment facilities and services
[73](#ground-segment-facilities-and-services)

7.2 Ground stations and communications network
[73](#ground-stations-and-communications-network)

7.3 Mission control centre [74](#mission-control-centre)

7.4 Computer facilities and network
[74](#computer-facilities-and-network)

7.5 flight control software system [74](#flight-control-software-system)

7.6 Mission operations concept [75](#mission-operations-concept)

7.7 Science instrument operations and team structure
[75](#science-instrument-operations-and-team-structure)

8 Communication and outreach [77](#communication-and-outreach)

9 Management [78](#management)

9.1 F2/F3 procurement approach [78](#f2f3-procurement-approach)

9.2 Procurement philosophy [78](#procurement-philosophy)

9.3 Scientific management and payload selection
[78](#scientific-management-and-payload-selection)

9.4 Industrial management [79](#industrial-management)

9.5 Development philosophy and schedule
[79](#development-philosophy-and-schedule)

9.5.1 General [79](#general)

9.5.2 Payload model philosophy [79](#payload-model-philosophy)

9.5.3 Service module model philosophy
[79](#service-module-model-philosophy)

9.5.4 Schedule [80](#schedule)

10 List of Acronyms [81](#list-of-acronyms)

# EXECUTIVE SUMMARY 

The Sun\'s atmosphere and the heliosphere represent uniquely accessible
domains of space, where fundamental physical processes common to solar,
astrophysical and laboratory plasmas can be studied in detail and under
conditions impossible to reproduce on Earth or to study from
astronomical distances.

The results from missions such as Helios, Ulysses, Yohkoh, SOHO, and
TRACE have advanced enormously our understanding of the solar corona,
the associated solar wind and the three-dimensional heliosphere.
However, we have reached the point where further *in-situ* measurements,
now much closer to the Sun, together with high-resolution imaging and
spectroscopy from a near-Sun and out-of-ecliptic perspective, promise to
bring about major breakthroughs in solar and heliospheric physics.

The Solar Orbiter will, through a novel orbital design and its state-of
the-art instruments, provide exactly the required observations.

The Solar Orbiter will for the first time

-   explore the uncharted innermost regions of our solar system,

-   study the Sun from close-up (45 solar radii, or 0.21 AU),

-   fly by the Sun, tuned to its rotation and examine the solar surface
    and the space above from a co-rotating vantage point,

-   provide images of the Sun\'s polar regions from heliographic
    latitudes as high as 38°.

The scientific goals of the Solar Orbiter are

-   to determine *in-situ* the properties and dynamics of plasma, fields
    and particles in the near-Sun heliosphere,

-   to investigate the fine-scale structure and dynamics of the Sun's
    magnetised atmosphere, using close-up, high-resolution remote
    sensing,

-   to identify the links between activity on the Sun\'s surface and the
    resulting evolution of the corona and inner heliosphere, using solar
    co-rotation passes,

-   to observe and fully characterise the Sun\'s polar regions and
    equatorial corona from high latitudes.

The underlying basic questions which are relevant to astrophysics in
general are

-   Why does the Sun vary and how does the solar dynamo work?

-   What are the fundamental physical processes at work in the solar
    atmosphere and in the heliosphere?

-   What are the links between the magnetic field dominated regime in
    the solar corona and the particle dominated regime in the
    heliosphere?

In particular, we want

-   to unravel the detailed working of the solar magnetic field as a key
    to understanding stellar magnetism and variability,

-   to map and describe the rotation, meridional flows, and magnetic
    topology near the Sun\'s poles, in order to understand the solar
    dynamo,

-   to investigate the variability of the solar radiation from the far
    side of the Sun and over the poles,

-   to reveal the flow of energy through the coupled layers of the solar
    atmosphere, e.g. to identify the small-scale sources of coronal
    heating and solar wind acceleration,

-   to analyse fluctuations and wave-particle interactions in the solar
    wind, in order to understand the fundamental processes related to
    turbulence at all relevant scales in a tenuous magnetofluid,

-   to understand the Sun as a prolific and variable particle
    accelerator,

-   to study the nature and the global dynamics of solar eruptive events
    (flares, coronal mass ejections, etc.) and their effects on the
    heliosphere ("space weather and space climate").

The near-Sun interplanetary measurements together with simultaneous
remote sensing observations of the Sun will permit us to disentangle
spatial and temporal variations during the co-rotational phases. They
will allow us to understand the characteristics of the solar wind and
energetic particles in close linkage with the plasma conditions in their
source regions on the Sun. By approaching as close as 45 solar radii,
the Solar Orbiter will view the solar atmosphere with unprecedented
spatial resolution (35 km pixel size, equivalent to 0.05 arcsec from
Earth). Over extended periods the Solar Orbiter will deliver images and
data of the polar regions and the side of the Sun not visible from
Earth.

The Solar Orbiter will achieve its wide-ranging aims with a suite of
sophisticated instruments. Note that due to the Orbiter's proximity to
the Sun the instruments can be fairly small, compared to instrumentation
required at the Earth's orbit.

The payload includes two instrument packages, optimised to meet the
solar and heliospheric science objectives:

-   Heliospheric *in-situ* instruments**:** solar wind analyser, radio
    and plasma wave analyser, magnetometer, energetic particle
    detectors, interplanetary dust detector, neutral particle detector,
    solar neutron detector.

-   Solar remote sensing instruments: Extreme ultraviolet (EUV) full-Sun
    and high resolution imager, high-resolution EUV spectrometer,
    high-resolution visible-light telescope and magnetograph, EUV and
    visible-light coronagraph, radiometer.

The Solar Orbiter spacecraft benefits from technology developed for the
Mercury Cornerstone mission. This allows such an ambitious mission to be
carried out within the frame of an F mission.

Using solar electric propulsion (SEP) in conjunction with multiple
planetary swing-by manoeuvres, it will take the Solar Orbiter only two
years to reach a perihelion of 45 solar radii at an orbital period of
149 days. Within the nominal 5 year mission phase, the Solar Orbiter
will perform several swing-by manoeuvres at Venus, in order to increase
the inclination of the orbital plane to 30^o^ with respect to the solar
equator. During an extended mission phase of about two years the
inclination will be further increased to 38°.

The spacecraft will be 3-axis stabilised and always Sun-pointed. Given
the extreme thermal conditions at 45 solar radii (25 solar constants),
the thermal design of the spacecraft has been considered in detail
during the assessment study and viable solutions have been identified.
Telemetry will be handled via X-band low-gain antennae, and by a 2-axis
steerable Ka-band high-gain antenna.

The total mass of the Solar Orbiter is compatible with a Soyuz-Fregat
launch from Baikonur.

![](media/image2.png){width="5.903472222222222in"
height="3.3916666666666666in"}

Figure 1.1: a) Ecliptic projection of the Solar Orbiter trajectory.
Blue: Solar Orbiter. Pink: Earth Orbit. Red: Venus orbit. Green: SEP
Thruster firings. b) Perihelion distance of the Solar Orbiter as a
function of time. c) Spacecraft latitude with respect to the Sun\'s
equator as a function of time.

+-----------+----------------------------------------------------------+
| **Mission | View the Sun from **near-Sun** and **out-of-ecliptic**   |
| Concept** | perspectives and perform                                 |
|           |                                                          |
|           | **·** spectroscopy and imaging at **high resolution**,   |
|           |                                                          |
|           | **· co-rotational** *in-situ* sampling of particles and  |
|           | fields,                                                  |
|           |                                                          |
|           | **·** remote-sensing of the **polar regions** of the Sun |
+-----------+----------------------------------------------------------+
| **        | Mass: 130 kg                                             |
| Payload** |                                                          |
|           | Power: 125 W                                             |
|           |                                                          |
|           | Telemetry: 75 kb/s                                       |
|           |                                                          |
|           | **Heliospheric instrumentation:**                        |
|           |                                                          |
|           | **·** Solar Wind Plasma Analyser                         |
|           |                                                          |
|           | **·** Radio and Plasma Wave Analyser                     |
|           |                                                          |
|           | **·** Magnetometer                                       |
|           |                                                          |
|           | **·** Energetic Particles Detector                       |
|           |                                                          |
|           | **·** Neutral Particle Detector                          |
|           |                                                          |
|           | **·** Dust Detector                                      |
|           |                                                          |
|           | **·** Neutron Detector                                   |
|           |                                                          |
|           | **·** Coronal Radio Sounding                             |
|           |                                                          |
|           | **Solar remote sensing instrumentation:**                |
|           |                                                          |
|           | **·** Visible-light Imager and Magnetograph              |
|           |                                                          |
|           | **·** EUV Spectrometer                                   |
|           |                                                          |
|           | **·** EUV Imager                                         |
|           |                                                          |
|           | **·** Ultraviolet and Visible Light Coronagraph          |
|           |                                                          |
|           | **·** Radiometer                                         |
+-----------+----------------------------------------------------------+
| **Spa     | **·** Design lifetime = 5 years                          |
| cecraft** |                                                          |
|           | **·** Consumables sized for 7 years                      |
|           |                                                          |
|           | **·** Total mass = 1308 kg                               |
|           |                                                          |
|           | **·** Dimensions: 3 m x 1.2 m x 1.6 m                    |
|           |                                                          |
|           | **·** 3-axis stabilised, Sun-pointing                    |
|           |                                                          |
|           | **·** Pointing stability better than 3 arcsec/15min      |
|           |                                                          |
|           | **·** Solar electric propulsion system: 4 x 0.15 N       |
|           | plasma thrusters                                         |
|           |                                                          |
|           | **·** Cruise solar arrays (28 m²) jettisoned after last  |
|           | SEP thrusting                                            |
|           |                                                          |
|           | **·** Orbit solar arrays (10 m^2^): tiltable             |
|           |                                                          |
|           | **·** X-band low-gain antennae, omni coverage            |
|           |                                                          |
|           | **·** Ka-band high-gain antenna, 1.5m diameter           |
+-----------+----------------------------------------------------------+
| **L       | Dedicated launch with **Soyuz-LV Fregat** from Baikonur. |
| auncher** |                                                          |
+-----------+----------------------------------------------------------+

# Scientific Rationale

## INTRODUCTION

The Sun\'s atmosphere and heliosphere represent uniquely accessible
domains of space, where fundamental physical processes common to solar,
astrophysical and laboratory plasmas can be studied under conditions
impossible to duplicate on Earth and in detail not possible to achieve
at astronomical distances.

Through a novel orbital design, the Solar Orbiter will explore *in-situ*
the innermost heliosphere and will scrutinise, through high-resolution
remote-sensing, the Sun and its atmosphere, while flying closer to the
Sun than any other spacecraft and out of the ecliptic to higher
latitudes. From there, the Solar Orbiter will provide the first
observations of the polar regions of the Sun and the whole equatorial
corona.

The results from missions such as Helios, Ulysses, Yohkoh, SOHO and
TRACE have advanced enormously our understanding of the solar corona and
the associated solar wind and three-dimensional heliosphere. However, we
have reached the point where further *in-situ* measurements, now much
closer to the Sun, together with high-resolution imaging and
spectroscopy from a near-Sun and out-of-ecliptic perspective, promise to
bring about major breakthroughs in solar and heliospheric physics.

There are four totally novel aspects to the Solar Orbiter mission, which
allow unique science investigations to be performed. These are:

-   Close-up remote sensing observations of the magnetised solar
    atmosphere providing unprecedented high-resolution;

-   Unique heliosynchronous observations enabling us to understand the
    links between solar and heliospheric processes;

-   *In-situ* measurements of the unexplored inner heliosphere;

-   First out-of-ecliptic imaging and spectroscopic observations of the
    solar poles and equatorial corona from high latitudes.

The next four sections describe the new science and advances which will
be made in these four areas.

## THE SUN\'S MAGNETISED PLASMA: CLOSE-UP OBSERVATIONS OF THE SOLAR ATMOSPHERE 

![](media/image3.jpeg){width="5.888888888888889in"
height="1.5777777777777777in"}The high-resolution imaging of the solar
atmosphere will represent a major step forward by providing an order of
magnitude improvement in spatial resolution over past missions. The
Solar Orbiter instruments will, in concert, enable us to analyse
thoroughly the time-variability, evolution and fine-scale structure of
the dynamic chromosphere, transition region and corona, to study the
Sun\'s magnetic activity on multiple scales, to investigate energetic
particle acceleration, confinement and release, and to reveal plasma and
radiation processes underlying the heating of the chromosphere and
corona. The Sun is the only star that can be resolved at the level at
which the physical processes responsible for magnetic activity take
place.

**Figure 2.1:** High-resolution field of view of the Solar Orbiter
(right) as compared with EIT on SOHO (left) and TRACE (middle). Pixel
size on the Sun of EIT and TRACE are 1850 and 350 km, respectively. The
EUV imager on the Solar Orbiter will have a pixel size of 35 km on the
Sun.

### Photospheric magnetic flux elements 

While our basic understanding of the equilibrium properties of the Sun
has been validated by helioseismology, most notably by the impressive
results from SOHO, our understanding of the non-equilibrium properties
-- primarily associated with magnetic fields -- remains poor. Magnetic
fields play a crucial role throughout astrophysics, ranging from the
formation of stars to the extraction of energy from supermassive black
holes in galactic nuclei. The Sun provides a natural laboratory for the
study of cosmic magnetism under conditions not accessible on Earth and
on scales not resolvable in distant astronomical objects.

A key goal of the Solar Orbiter is to advance our understanding of solar
magnetism by measuring the structure and dynamics of the magnetic field
at the solar surface down to its fundamental length scale. The major
part of the magnetic flux permeating the solar photosphere outside
sunspots is concentrated in small flux tubes of kilo-Gauss field
strength. The scale of these magnetic flux tubes is believed to be
determined by the pressure scale height, which is about 100 km. This is
also near the typical photon mean free path at the photosphere whose
dynamics is controlled largely by radiation processes at this scale.

The structure and dynamics of these fundamental elements of the
near-surface magnetic field has profound implications for a number of
basic questions:

-   How do magnetic foot-point motion, wave excitation, flux
    cancellation and reconnection contribute to the flux of mechanical
    energy into the corona?

-   In what way do the emergence, evolution and removal of magnetic flux
    elements determine the magnetic flux budget of the Sun? Is there a
    local dynamo operating on the scale of granulation?

-   What is the origin of the facular contribution to the variability of
    the solar constant?

-   What is the physics of the interaction between convection and
    magnetism?

Answers to these questions require the study of magnetic flux elements
on their intrinsic spatial scale (\<100 km). The Solar Orbiter will
allow us to achieve this resolution, providing observations of magnetic
field emergence, dynamics, twist, shearing, mutual interactions and
possible coalescence and subduction in order to follow the evolution and
understand the life cycles of magnetic flux elements.

### Basic processes and fundamental scales in the magnetised solar atmosphere 

We wish to understand the basic processes on all scales in the solar
atmosphere, e.g. magnetic processes leading to particle acceleration or
heating. It has been realised recently that very small-scale processes
are at work in the solar atmosphere. For example, theoreticians have
pointed out the difficulty of having both large-scale and rapid magnetic
reconnection. This led to the concept of heating through numerous,
globally distributed, small-scale "events" for which observers have
candidates (e.g. SOHO observations of explosive events). Numerical
simulations have been performed in the frame of Self-Organised
Criticality or Cell Automation models whose outputs, in terms of
temporal fluctuations and statistical properties, compare well with
observations. Moreover, it seems that the observed distribution laws
determined from SOHO data have self-similarity properties which point at
sub-resolution processes. The nanoflare model of Parker is just an
example of such models.

Whatever the scale, magnetic reconnection leads to particle dissipative
heating and acceleration, and wave generation. Particle acceleration and
wave dissipation have the net effect on the lower solar atmosphere of a
local kinetic energy increase which can be revealed through high
resolution extreme ultraviolet (EUV) imaging and spectroscopy. Wave
propagation can be traced from the source site to the region of
dissipation through observations of EUV-line broadening and Doppler
shifts.

Between the very small scales implied by the above processes and the
sizes of the smallest observed features, there are intermediate plasma
scales. Indications of these intermediate scales are provided by the
very small values of the filling factor derived from observations made
of a large variety of solar structures. Thus, the transition region and
inner corona are highly structured at scales even below those presently
resolvable by SOHO and TRACE, and involve hot and cold plasmas at all
temperatures between 10^4^ K and 10^7^ K. Widely differing temperature
plasmas often co-exist side by side in hierarchies of filamentary
structures, channelled by the magnetic field, which together form the
solar atmosphere.

Given the high spatial resolution due to the spacecraft's proximity to
the Sun, combined with high temporal and spectral resolution, and
multi-wavelength coverage, the Solar Orbiter will measure the signatures
of the basic processes and the elementary structures of the solar
atmosphere, providing plasma diagnostic information over the full
temperature range existing in the solar atmosphere. This will enable us
to understand their nature and importance in the mass and energy budget
of the solar atmosphere. For example, the Solar Orbiter will be used to
measure

-   the outflow and accelerated particles from magnetic reconnection
    sites on all scales,

-   the proper motion, line broadening and Doppler shifts associated
    with wave propagation,

-   the scale properties of MHD turbulence,

-   the scales and evolution of the elementary structures.

### Waves in the corona

Magnetohydrodynamic waves generated in the photosphere by convective
motion are primarily of low frequency. Small-scale magnetic activity is
expected to continually produce waves. Their dissipation could involve
cyclotron damping, which is observed to operate in the solar wind. The
Solar Orbiter will make the first observations of such plasma processes
from a near-Sun perspective and thus address the extended heating of the
outer corona. In the small-scale magnetic structures of the strongly
inhomogeneous network fields higher-frequency waves could be excited up
to the kilo-Hertz range. Such waves would transfer their energy, e.g.
into the transverse kinetic degrees of freedom of the protons, and
particularly the heavy ions, and thereby heat the particles very
effectively to high kinetic temperatures, a process for which the UVCS
and SUMER instruments on SOHO have recently confirmed evidence in the
strong Doppler broadening of emission lines.

Due to its proximity to the heating sites in the corona and the high
sensitivity and spatial resolution of its instrumentation, the Solar
Orbiter will for the first time be able to

-   see very dim emissions, which are not visible from 1 AU due to the
    low density of the gas and the high contrast in emissivity, and long
    lines of sight; these difficulties are not encountered when
    observing plasma confined in small loops,

-   resolve and diagnose dilute plasma on open fields and the outflow in
    coronal funnels,

-   search for evidence of high-frequency waves in the corona (with
    reasonable chances of success due to the high temporal and spatial
    resolving capabilities of the dedicated instrumentation).

### Coronal magnetic field 

A key scientific objective of the Solar Orbiter is to study the
emergence and the cancellation of photospheric magnetic flux (the latter
is the disappearance of opposite polarity regions in close contact), and
to investigate the consequences of such processes for the overlying
global coronal magnetic loops and for the chromospheric and transition
region magnetic network. Flux cancellations are known to be the origin
of various active phenomena, such as filament formation and eruption,
evolution of small points of emission in radio, ultraviolet or X rays,
or the occurrence of flares.

Magnetograms combined with EUV images as well as EUV spectra at high
spatial resolution are the key data necessary to understand the
influence that small-scale magnetic activity has on the corona. Such
co-ordinated observations should reveal basic processes at work
throughout the solar atmosphere which are drivers of the solar wind
acceleration or coronal heating. Thus, the detailed understanding of the
fine-scale processes in the magnetic fields at the base of the solar
atmosphere is an essential goal of this mission.

On a larger scale, coronal loops are the very building blocks of the
outer solar atmosphere, ranging from small loops in active regions,
through loop arcades to active region interconnecting loops and helmet
streamers. Observations from Yohkoh, SOHO and TRACE in extreme
ultraviolet and X-ray wavelengths have revealed a truly complex, highly
dynamic environment with loops confining plasmas at widely varying
temperatures even in adjacent loops. It is within such loop systems that
we witness the most dramatic of solar phenomena, the solar flares and
mass ejections, as magnetic loop systems interact or become unstable.
The study of the interaction and evolution of loop systems, through
high-resolution EUV mapping combined with fine-scale diagnostic analysis
using high-resolution spectroscopic observations, will be a major
activity for Solar Orbiter scientists in particular in an effort to
understand processes leading to mass ejection and flaring.

The TRACE extreme ultraviolet observations in particular illustrate the
existence of fine-scale structures in the coronal loops, and reveal
continuous dynamic activity in particular at the smallest scales. There
is strong evidence that the size of the actual brightness structures
lies well below the best current spatial resolution. This points to the
need for still higher spatial resolution, which, as in the visible, can
be obtained by modest means with the Solar Orbiter.

The wide coverage of coronal temperatures by the telescopes on the Solar
Orbiter will enable complete images to be obtained in fast cadence. From
these the density, temperature and flow distributions can be derived
reliably for the full range of magnetic structures, allowing detailed
studies such that coronal heating processes in current sheets, shock
fronts, or acceleration in small explosive events and rapid plasma jets
become clearly visible and can be resolved in time and space for the
first time.

![](media/image4.png){width="5.6in" height="3.7625in"}

Figure 2.2: Images of plasma in magnetic loops observed on the Sun. In
addition to hydrogen, the Sun's atmosphere contains ions of common
elements like helium, oxygen and magnesium. The images are produced by
detecting the emission lines from a number of ions, and their
characteristic temperatures are given. One of the surprises that the
SOHO/CDS data shows is that loops at widely differing temperatures can
exist side-by-side in the same regions of the Sun's atmosphere. The disc
plotted on the oxygen image shows the Earth to the same scale.

## LINKING THE PHOTOSPHERE AND CORONA TO THE HELIOSPHERE: CO-ROTATION OBSERVATIONS

Studies of the evolution of solar features such as active regions,
loops, prominences or sunspots are greatly complicated by the fact that
their evolution time scales are comparable to the solar rotation period.
Thus the evolution is entangled with other effects such as
centre-to-limb variation, foreshortening and projection effects. In
order to disentangle these effects it is necessary to co-rotate with the
Sun. The Solar Orbiter will for the first time provide such an
opportunity and will thus help resolve old and otherwise intractable
problems related to the solar dynamo and the diffusion of the magnetic
field across the solar surface. It will also allow for the first time to
follow the evolution of sunspots and active regions as well as their
influence on the corona above.

### Global solar corona and solar wind 

There are two characteristic types of solar corona and wind, prevailing
at different heliographic latitude regions of the Sun and heliosphere.
The high-speed flow is the basic equilibrium state of the solar wind. It
is most conspicuous near solar minimum, when it emanates steadily from
the magnetically open coronal holes around the poles, whereas the slow
wind originates in a more gusty fashion from the equatorial streamers.
These are for most of the time magnetically closed, but appear to open
intermittently to release discrete mass ejections (CMEs) (see the
subsequent Figure 2.4) and more continuously the slow wind embedding the
heliospheric current sheet. The transient nature of the solar wind is
also evident in the high variability of its abundances in helium and
other heavier elements found in association with large-scale magnetic
activity on the Sun. The solar magnetic field in the equatorial regions
reveals a rich morphology and many fine-scale structures, such as the
low-latitude rays resembling the polar plumes, which are clearly evident
in visible light coronagraph pictures and in the SOHO/EIT images.
Remnants of these features are even found in the meso-scale stream
variations of the solar wind, as observed *in-situ* by Helios and
Ulysses and through interplanetary scintillations.

We can only resolve between the slow and transient solar wind, and
understand them properly, by direct sampling close to the solar wind
sources at a fixed heliographic longitude. It is a key advantage of the
Solar Orbiter that from the co-rotational vantage point the temporal and
spatial variations of the solar wind can be disentangled unambiguously.
The observations to be made by the Solar Orbiter will concentrate on the
slow streams and on solar disturbances associated with magnetic solar
activity, flares, loops, erupting prominences, and CMEs, when the Solar
Orbiter is close to the ecliptic, and on the fast streams, when the
Solar Orbiter is close to maximum inclination. The temporal evolution
and spatial structure of such phenomena at the coronal base will for the
first time be measured at very high spatial (\< 100 km) and temporal
resolution.

It is well known that the solar wind carries angular momentum from the
Sun through kinetic and magnetic stresses. This will lead ultimately to
a spin-down of the Sun\'s rotation within its remaining life time, a
conclusion derived from the Helios measurements of the angular momentum
as carried away by the wind near the ecliptic plane. The determination
of the overall angular momentum loss associated with the solar wind
outflow at higher latitudes is an important goal for the Solar Orbiter.
Also, the analysis of the partial co-rotation of the outer corona and
the detachment of the wind from solar rotation near the Alfvén surface
may be possible by exploiting the co-rotational passes. Such
measurements will produce novel results concerning the global angular
momentum loss of the Sun as a star, which is an important issue in
astrophysics.

The Solar Orbiter mission provides unique possibilities for
complementary remote sensing and *in-situ* observations of the Sun from
close distances (perihelion of 45 solar radii). Unprecedented spatial
resolution at scales below 100 km will be achieved in the images
obtained in various wavelength bands. Solar rotation will have a
negligible influence on the observations during the heliosynchronous
orbital phases of the mission. This will allow us to separate spatial
from temporal variations. Therefore, time variability of the magnetic
field and its optical and interplanetary-particles manifestations can be
studied extensively for many days at a given heliographic longitude. The
favourable vantage points of the Solar Orbiter along a heliosynchronous
trajectory are unattainable by any other means.

The scientific success of the Solar Orbiter mission in this area is
ensured by a comprehensive and well-focused suite of *in-situ* and
remote-sensing instruments, which allow the local plasma and field
environment to be connected with its plasma and magnetic sources on the
Sun, and thus to establish spatial and temporal links with the location
and evolution of these source regions in the highly structured and
widely varying solar atmosphere. The Solar Orbiter will fly through
field lines with foot points in the fields of view of its imagers, and
therefore the characterisation of the properties of the plasma and its
sources can be made, essentially by virtue of the co-rotation of the
spacecraft with the Sun during the perihelion passes.

### Global coronal sources of the solar wind 

Helios and Ulysses have clarified the global origin of the fast and slow
solar wind in the large-scale corona. More recently, SUMER on SOHO has
provided intriguing insight into the small-scale sources of the fast
solar wind (see Figure 2.3). However, the detailed origin of the slow
wind remains unclear. There are two main alternative ideas about the
origin of the slow solar wind. One is that it arises from the outer
edges of coronal holes, in magnetic field regions continuously open to
interplanetary space. The other is that the slow wind arises in the
middle and lower corona from the tops of helmet streamers, which are
primarily closed magnetic field regions but do open intermittently to
release the flow at heights beyond about two solar radii in the form of
plasma blobs or plasmoids. Perhaps the slow wind consists of these two
types of flow that intermingle at a few solar radii. The remote sensing
observations of the corona combined with *in-situ* solar wind
observations made by the Solar Orbiter during the co-rotational passes
will allow us to distinguish between these two hypotheses, by providing
accurate connections of in-situ solar wind structures with
remotely-sensed coronal structures.

### Magnetic network 

The supergranulation network, which dominates the chromospheric plasma
dynamics, is apparent in the EUV emission pattern as seen by the SUMER
instrument on SOHO. Magnetograms from SOHO have revealed the ubiquitous
appearance of small magnetic bipoles at the solar surface. After
emergence, the polarities separate and are carried to the network
boundaries by the supergranular flow, where they merge with the
pre-existing network flux. This leads to flux cancellation, submergence
and reconnection events. The magnetograms also show that the magnetic
field exists in the network in two components side-by-side, i.e. in
uncanceled unipolar fields together with a carpet of closed loops and
flux tubes. The small loops will either emerge or contract downwards and
collide, and thus constitute a permanent source of energy, which can be
tapped by the particles through magnetic field dissipation. Numerical
simulations suggest that many of the bipolar structures have scales
smaller than 100 km.

![](media/image5.jpeg){width="4.081944444444445in"
height="3.8854166666666665in"}

Figure 2.3: SOHO/SUMER observations of the solar-wind source regions and
magnetic structure of the chromospheric network. The insert shows the
measured Doppler shifts of Neon ions, indicating blue-shifts, i.e.
outflow, at the network cell boundaries and lane junctions below the
polar coronal hole, and red-shifts (downflow) in the network regions
underlying the globally closed corona.

### Boundaries and fine structures in the corona and solar wind 

The boundaries and gradients between the fast and slow solar wind have
been well characterised by the Helios and Ulysses spacecraft at 0.3 AU
and beyond, but little is known about their microphysics and dynamics.
For example, instabilities can be caused by velocity shear across a
boundary and become a source of turbulence and friction, and thus alter
the nature of the wind in the vicinity of stream boundaries and
interaction regions. There are plenty of other fine-scale structures in
the solar wind and corona, one example of which are the conspicuous
plumes over the poles, which have been studied by SOHO in considerable
detail. Plumes have been observed to extend beyond many solar radii, yet
their *in-situ* signatures in the wind plasma are elusive. Material in
plumes has been found to flow at much lower speeds than in the ambient
darker lanes throughout the altitudes observed in coronal holes. By
combining the Solar Orbiter data from the *in-situ* and remote-sensing
instruments, we will be able to identify the genuine contributions of
plumes to the fast wind.

Little is known currently about the projection of coronal substructures
into the outer corona and near-Sun solar wind. There are many other
meso-scale structures in the corona and particularly the slow solar
wind, which is intrinsically variable in space and time and convects
structures such as tangential discontinuities, the coronal origin of
which remains a mystery. Many structures exist in the corona at a wide
variety of length scales. Structures very similar to the polar plumes
have also been observed in inter-streamer regions. At even smaller
spatial scales, coronal spikes and EUV spicules with widths of order
10^3^ km have been observed. Ground-based white light eclipse images
show elongated and thin, thread-like structures that extend outward in
the innermost corona and seem to be ubiquitous outside coronal holes.
Further evidence of fine structure well below the spatial resolution of
current optical instruments comes from interplanetary radio
observations. The Solar Orbiter, by virtue of its proximity to the
corona will be able to resolve such features with sufficient brightness
contrast. It will trace their extensions from the solar atmosphere into
space by measuring the remnant signatures in the more distant solar wind
during the perihelion passes. Thus, the Solar Orbiter will examine
fine-scale structures in the corona at much higher spatial resolution
than ever employed before and associate them with their interplanetary
manifestations.

### Connections between the internal plasma states of the solar corona and the solar wind

There are many closely-connected heliospheric and solar science
objectives of the Solar Orbiter, which can be addressed in particular
during the co-rotation passes enabling steady observational conditions
from a fixed vantage point. The principal goals are:

-   to determine the relationship between coronal and solar wind
    structure on all scales that can be resolved;

-   to identify the coronal energy source for the solar wind and trace
    by remote-sensing the flow of energy through the different layers of
    the atmosphere.

Remnants of the processes that heat and accelerate the ions will still
be detectable in the *in-situ* measured microscopic features of the
solar wind particle velocity distributions (such as double ion beams,
minor ion differential streaming, pronounced temperature anisotropies
indicating cyclotron or Landau resonance between plasma waves and
particles, or heat flux-carrying suprathermal tails, as already observed
by Helios and Ulysses) and will allow clear inferences to be made on the
coronal plasma processes. A related goal is to determine the role of
turbulence and waves in heating the corona, accelerating the solar wind,
and energising particles, and to determine how the particle populations,
plasma waves, and magnetohydrodynamic turbulence evolve together with
heliocentric distance in the inner-most heliosphere.

### Coronal and solar wind abundances and fractionation effects

Understanding the abundances of elements in the Sun and solar wind is an
important issue in solar physics. The solar wind carries the material
delivered from the outer convective zone of the Sun into the
heliosphere. With the exception of the H and He isotopes, the elemental
composition of the outer convective zone is close to the proto-solar
nebula composition. The element helium, which is not visible in the
photosphere, is highly variable and rather under-abundant in the solar
wind and solar energetic particles. The two types of solar wind, fast
and slow, differ considerably in their charge-state composition (a
coronal signature) and elemental abundance (a chromospheric signature).
Apparently, the elemental fractionation processes are fundamentally
different in the fast and slow solar wind, although fractionation on the
basis of the First Ionization Potential (FIP) is present to a different
degree in both types of wind. The FIP effect seems to be active also in
the chromosphere beneath the polar coronal holes, in the sources of the
fast wind, but at a reduced strength. Charge-state spectra are also
differing in the two types of streams: thermal equilibrium spectra
prevail in fast streams, indicating a simple freezing-in process. The
slow wind charge-state distribution, on the other hand, shows an excess
of high charge states and indicates that different coronal temperature
regimes may coexist in the source region.

The Solar Orbiter, while taking high-resolution images and making
spectroscopic measurements of solar wind source regions that are
magnetically linked to the spacecraft location and making simultaneous,
*in-situ* measurements over the long co-rotation intervals, is ideally
suited to address the critical issues of the chromospheric fractionation
process. The coronagraph on the Solar Orbiter will determine for the
first time the helium abundance (high FIP element) in those atmospheric
layers where the acceleration of the slow and fast streams actually
occurs, and where the charge states freeze in. The results expected will
provide keys for the understanding of the processes at the origin of the
solar wind and of the elemental composition in the heliosphere.

For the Solar Orbiter a new class of UV and visible-light coronagraph
observations is envisioned, aiming at measuring directly and
characterising in detail the properties of the two most abundant
elements, hydrogen and helium. In particular, this innovation will
provide new and better views of the solar corona than possible with SOHO
presently and STEREO in the future. In particular, the Solar Orbiter
will be able to provide

-   the first UV images of the full corona for the two most abundant
    elements,

-   the first global maps of the solar wind outflow,

-   the first images of the He II coronal emission.

The Solar Orbiter will allow us to do these new measurements of the
corona also from out of the ecliptic, and will thus provide the first
view of the entire equatorial corona and global observations of the
coronal expansion near the equator.

### Coronal transients

A subject that has recently attained much attention is forecasting of
"space weather". It is concerned with transient events such as flares,
coronal mass ejections (CMEs), eruptive prominences, and shock waves and
their impacts upon the Earth\'s magnetosphere and atmosphere.
Fundamental questions in this area have not been answered, e.g.,

-   What are the indicators for imminent violent eruptions? Can they be
    predicted?

-   Why are there disparate types of solar transients?

-   What determines their propagation properties?

-   How far around the Sun do the resulting shock waves reach?

The Solar Orbiter will be ideally located, being closer to the sources
of transients in the solar atmosphere, to measure the input into the
heliosphere and to determine the boundary conditions near the Sun. These
dramatic events can literally shatter the whole heliosphere, and their
effects can be felt at all planets. The Solar Orbiter will be a key link
in a chain of solar terrestrial observatories to be stationed in
Earth\'s orbit and at the libration points in that it provides
near-real-time event alerts from its unique orbit close to the Sun.

![](media/image6.png){width="3.9479166666666665in"
height="2.7708333333333335in"}

Figure 2.4: Giant coronal mass ejection as seen by LASCO on SOHO in
1998. The filament material of the ejected prominence exhibits twisted
helix-shaped structures. The Solar Orbiter at its perihelion of 45 R~s~
would be able to measure the ejecta *in-situ* and out of the ecliptic in
much more detail than possible from the Earth\'s orbit.

The Solar Orbiter will provide high resolution images of the solar
atmosphere as well as allow plasma diagnostics using spectroscopy.
These, combined with a capability to determine the photospheric fields
at the footpoints of coronal loops, provide the tools for a thorough,
detailed analysis of the processes leading to eruptive events. The
influence of the eruptions on the outer corona can be determined with
coronagraph instrumentation, and the associated effects on the
interplanetary medium can be studied with in-situ instrumentation
without significant delays or transport-related changes of the particles
and fields. The multi-wavelength, multi-disciplinary approach of the
Solar Orbiter, combined with the novel location, produces a powerful
tool for studies of the influence of eruptive events such as CMEs on
interplanetary space.

## PARTICLES AND FIELDS: *IN-SITU* MEASUREMENTS IN THE INNER HELIOSPHERE

According to recent SOHO findings, one must conclude that coronal
expansion arises because of the high temperatures of the coronal ions,
with the minor species reaching even 10^8^ K at a few solar radii. In
contrast, electrons are comparatively cool, in fact they are found to
hardly reach the canonical coronal temperature of 10^6^ K, and
consequently the electric field (related to the electron partial
pressure gradient) has a minor role in accelerating the ions. The high
pressure of the coronal ions and the low pressure of the local
interstellar medium lead to a supersonic solar wind extending to long
distances from the Sun to some 100 AU. Yet, even after the SOHO mission
the detailed physical mechanisms that heat the corona and accelerate the
plasma to supersonic speed remain poorly understood, because the
resolution of the SOHO imagers and spectrometers were still not
sufficient, and because the solar wind plasma has never been directly
sampled closer to the Sun than the Helios perihelion (0.3 AU).

The Solar Orbiter will provide the first opportunity of going closer to
the Sun and into the inner heliosphere to 0.21 AU, but unlike Helios and
Ulysses, the Solar Orbiter will also carry powerful, high-resolution
optical instruments together with the *in-situ* instruments. In
particular, the plasma and field instruments will have high temporal
resolutions, ranging between 0.01 s and 1 s, and offer unique
possibilities for resolving physical processes at their intrinsic
scales. Therefore, this mission will reveal new insights in the plasma
kinetic processes that structure the Sun\'s atmosphere, heat the
extended corona and accelerate the solar wind as well as energetic
particles.

### Microstate of the interplanetary solar wind

The ultimate causes of interplanetary kinetic phenomena are to be found
in the dynamic corona itself. The closer a spacecraft comes to the Sun
the more likely it is to detect remnants of coronal heating and related
plasma processes, occurring on a broad range of scales in space and
time. The radial evolution of the internal state of the expanding wind
resembles a complicated relaxation process, in which free energy stored
in the form of stream structure and shear as well as in non-Maxwellian
particle velocity distributions (see Figure 2.5) is converted into wave
and turbulence energy.

![](media/image7.png){width="4.204861111111111in"
height="3.0944444444444446in"}Plasma waves in the collisionless solar
corona and wind play a role analogous to collisions in ordinary fluids.
These wave modes can theoretically be excited by a variety of free
energy sources, including drifts, currents, temperature anisotropies and
beams, which must be resolved in detail. All the wave modes of primary
importance together with the ions and electrons will be measured by the
Solar Orbiter at high time resolution, in order to provide the
comprehensive wave and particles diagnostics necessary to study the
wave-particle interactions and kinetic processes.

The Solar Orbiter, while approaching the Sun to about 0.2 AU with its
plasma and wave analysers, will enable high-resolution measurements of
kinetic processes, and this has great potential for unexpected
discoveries. It will address fundamental solar wind science and key
plasma-physics questions such as

-   How do particle distributions develop velocity-space gradients and
    deviations from Maxwellians?

-   What processes drive plasma instabilities and cause wave growth and
    damping?

-   How are new-born ions, e.g., sputtered from dust grains,
    incorporated into the solar wind flow in the inner heliosphere?

-   What regulates transport and ensures the observed fluid behaviour in
    the collisionless solar wind?

-   What are the radial, latitudinal and longitudinal gradients of
    plasma parameters in the inner heliosphere?

-   What is the microstructure of, e.g. stream interfaces and boundary
    layers near the Sun?

-   How does the chemical and charge-state composition of the plasma
    vary spatially?

### Solar wind ions as tracers of coronal structures

The solar wind carries information on its coronal source regions, e.g.
through its elemental and isotope composition, and the ionisation states
of the various atoms. At the coronal base, a compositional bias is
introduced according to the first ionisation potential (FIP) of the
elements. This FIP effect appears to be significantly different for slow
and fast solar wind flows. Further, the ionisation state of various
species indicates that the slow wind must have undergone substantially
more heating than the fast wind. These signatures will be used for the
first time to resolve fine structures and boundaries in the solar wind.
This diagnostic technique works most reliably close to the Sun where the
flow has not yet been processed (compressed, deflected, etc.) by
interactions between streams of different speed. With Helios it was
found that at least in the high-speed wind a basic flow-tube structure
was still recognisable in-situ at 0.3 AU. The scale-size observed
matches that of supergranules fairly well, if an appropriate flux-tube
expansion is taken into account. The Solar Orbiter, using modern
ion-composition instruments and being closer to the Sun, will

-   reveal, through precise determination of compositional variations,
    the fine structures in all types of solar wind;

-   link the flow tubes directly with the underlying chromospheric
    network observed remotely;

-   identify pick-up ions stemming from dust and interplanetary sources.

### Magnetohydrodynamic turbulence

The solar wind plasma is in a highly turbulent state composed of various
components. The energetically dominant component consists of largely
incompressible Alfvénic fluctuations (see Figure 2.6 for the power
spectrum). The minor component has a much lower amplitude than the
Alfvénic fluctuations, is compressible and clearly enhanced in the mixed
low-speed flows. The Solar Orbiter will scan a belt ranging roughly from
-40° to 40° in heliographic latitude, while being within 0.3 AU of the
Sun. From these vantage points the Solar Orbiter will be able to answer
important questions such as:

-   How does MHD turbulence evolve spatially at higher latitudes near
    the Sun?

-   What are the crucial conditions for in-situ turbulence generation?

-   How does the turbulence pattern vary with stream structure closer to
    the Sun?

The overall radial trends as seen by Ulysses and Helios suggest strong
variations of the local production rate of the Alfvénic fluctuations in
the region just inside 0.3 AU. The reduction of the perihelion distance
from 0.3 AU (Helios) to 0.2 AU (Solar Orbiter) will offer unique
opportunities to study the local generation, non-linear coupling and
spatial evolution of MHD waves near the Sun. Measuring Alfvénic
fluctuations and MHD turbulence *in-situ* represents also a means of
diagnosing their coronal sources. The Solar Orbiter will address these
issues from its unique vantage point and help to answer basic questions
such as:

-   How and where are Alfvénic fluctuations generated in the solar
    corona?

-   How does MHD turbulence evolve radially and dissipate in the inner
    heliosphere?

-   Do the spectra contain indications or relics of high-frequency wave
    heating of the corona?

The solar wind is the only available plasma "laboratory" where detailed
studies of MHD turbulence can be carried out free from interference with
spatial boundaries, and in the important domain of very large magnetic
Reynolds numbers. Detailed comparison between experimental *in-situ*
data and theoretical concepts will allow us to put MHD turbulence theory
on more solid physical ground, which will be of critical importance for
understanding the solar (stellar) coronal heating mechanism and the role
or turbulence in the solar (a stellar) wind.

### Acceleration and transport of solar energetic particles 

A permanent source of difficulty has been our inability to predict the
intensity of solar energetic particles at the Earth from observed
transient activity on the Sun. An important part of the problem is that
we do not know the suprathermal population that feeds the acceleration
processes near the Sun. The efficiency for transferring energy from
flares to energetic particles cannot be inferred from remote
observations, because an unknown fraction of the accelerated ions remain
trapped by strong magnetic fields near the Sun for a significant time
after acceleration. Subsequent γ-ray and neutron emissions resulting
from their eventual loss to the atmosphere are often too weak to be
observable. Present observations indicate that small transients occur
sufficiently often to allow a determination of the efficiency, e.g. by
neutrons as proxies for the magnetically bound component. Our ability to
solve this issue will be greatly enhanced by the Solar Orbiter, because
during its multiple perihelion passages we will

-   gain a better knowledge of the source spectrum;

-   obtain new observations on particle motion in the hypothetical
    storage region;

-   measure changes in the spectrum as the ions and electrons propagate
    from the Sun to the spacecraft after escape from the trapping
    region.

![](media/image8.png){width="3.8291666666666666in"
height="3.2777777777777777in"}

The Solar Orbiter will for the first time investigate the particle
environment in close proximity to the different source regions on the
Sun, such as coronal holes, streamers, coronal mass ejections (CMEs) and
associated shocks, active regions, and flare locations. With regard to
CMEs in particular, the Solar Orbiter will

-   determine the solar source conditions for different particle species
    (e.g. e, p, ^3^He, heavy ions, p/He ratios) from composition
    measurements, energy spectra and time evolution;

-   distinguish clearly between gradual (shock-associated) CME events
    and impulsive flare-type events related to magnetic reconnection;

-   study the effects of particle acceleration and turbulence-moderated
    propagation at different locations with respect to the CME centre;

-   find the differences between the particle signatures associated with
    parallel and perpendicular shocks at the east and west flanks of
    CMEs;

-   probe the effects of magnetic reconfigurations in the aftermath of
    CME launches, at times when the acceleration processes still occur
    in the corona;

-   detect perhaps for the first time energetic particle populations
    from microflares, a measurement which is not possible further away
    from Sun due to background problems.

In addition, we can study with the Solar Orbiter important global
aspects of the Sun and heliosphere (see Figure 2.7) by

-   utilising the energetic particles as probes for the coronal and
    heliospheric magnetic field;

-   analysing the propagation of solar particles and modulation of
    galactic cosmic rays.

![](media/image9.png){width="4.588888888888889in" height="5.8in"}

Figure 2.7: A series of solar energetic particle events observed in July
1996 with the COSTEP/EPHIN experiment onboard SOHO. The upper panel
shows images of the Sun taken on July 9, 12, 14 with SOHO/EIT in Fe XV
(28.4 nm). The bottom panel shows intensities of 0.25-0.7 MeV electron,
4.3-7.8 MeV proton and 4.3-7.8 MeV/n helium nuclei measured with the
COSTEP/EPHIN. The intensity increases of energetic particles at SOHO
were caused by flares and coronal mass ejections in an active region in
the Sun\'s western hemisphere.

Generally, the processes which accelerate particles to very high
energies are of great interest in astrophysics. Observations of
energetic particles close to their sources on the Sun will allow us to
study our nearest star**,** the Sun, as a particle accelerator.

### Neutral particles from the Sun

Neutral atoms are widespread in the heliosphere and their energies are
in the range from a few eV to more than 100 keV. Contrary to charged
particles, they can travel large distances through space, undisturbed by
the interplanetary magnetic field. In the solar corona, the neutral
hydrogen atoms are closely coupled to the emerging solar wind plasma and
give rise to the prominent Lα corona. The ratio of the densities of
hydrogen and protons is very low, 10^-6^ to 10^-7^, and the neutral
particles therefore have a negligible impact on the plasma. The neutral
atoms are a trace particle population originating from the solar wind
plasma by charge exchange.

The in-situ observation of neutral hydrogen by a space-borne instrument
in the vicinity of the Sun would be a \"first\" and an essential
contribution to our knowledge of the solar corona. The Solar Orbiter
provides the opportunity of observing neutral particles emitted from the
Sun. By measuring the "neutral solar wind", we obtain an independent
means of determining the solar wind velocity profile in the outer
corona.

In addition, Energetic Neutral Atoms (ENAs) produced by charge-exchange
processes in interplanetary space can be detected by the Solar Orbiter
for the first time near the Sun. ENAs may be used for imaging the outer
solar corona. The ENAs originating from a seed-ion population or
background neutral gas will allow us to derive properties of the
background gas as well as to remotely sense the energy distribution,
spatial distribution and temporal evolution of the seed populations. The
neutral gas could be either of interstellar origin or from a neutral
component stemming from the solar wind itself, e.g. being produced by
neutralisation of solar wind ions on interplanetary dust particles close
to the Sun. Possible seed ion populations are the solar wind protons,
energetic solar particles and anomalous cosmic rays. Measuring the
neutrals means probing and inferring the physical properties of the
circumsolar dust. Such measurements have great potential for unforeseen
discoveries.

Direct observation of the neutral atoms will help to refine the physical
models of the Lα corona, by providing information complementing the
remote-sensing observations of the coronagraph on the Solar Orbiter, and
thus will enhance our understanding of the coronal plasma processes and
wave-particle interactions out to a few solar radii. Further out, within
a few solar radii more and more of the neutral atoms become de-coupled
from the plasma.

Coronal mass ejections (CMEs) are known to include a wide range of ion
charge distributions, originating from very cold to very hot coronal
plasma. Consequently, one expects to find also neutral particles in
CME-related solar wind. However, up to now, no direct observations of
such neutral particles exist. Since these particles are not magnetically
trapped, they could give a clue to the understanding of the trigger
process of the CMEs. This would be another \"first\" and important
observation of the Solar Orbiter.

### Circumsolar and interplanetary dust 

The Solar Orbiter will also measure the size and flux distribution of
dust particles surrounding the Sun. These particles are known to produce
the zodiacal light and the F-corona and may stem from Sun-grazing
comets. The large number of Sun-grazers discovered by the SOHO
coronagraphs indicates that these comets must be a prolific dust source.

The observed distribution of interplanetary dust particles (IDPs), i.e.
micrometeorites of submicron to millimetre size, forms a flat spheroid
centred on the ecliptic plane. Its spatial density decreases with
increasing heliocentric distance and increasing latitude. Different
forces and effects are acting on IDPs: gravitational attraction,
radiation pressure, corpuscular pressure, magnetic forces, planetary
perturbations, erosion processes (sputtering, sublimation), and mutual
catastrophic collisions. IDPs originate from comets and asteroids or
even from interstellar space. They approach the Sun on time scales of
10000 to 100000 years due to the deceleration by the Poynting-Robertson
effect. In the vicinity of the Sun, IDPs go through a complicated
evolution process, which depends on the particle properties and chemical
composition, caused by their progressive sublimation, as their
temperature becomes higher.

On the other hand, these processes are the source of an additional ion
population in the solar wind: outgassing, sublimation and sputtering
produce neutrals which are then ionised and picked up by the solar wind.
The study of these particles will contribute significantly to the
understanding of the evolution of interplanetary dust in the vicinity of
the Sun. The interaction between the plasma and the solar magnetic field
with IDPs makes this picture even more complex: far-reaching, dense
coronal structures as well as the changing pattern of the magnetic field
can perturb the IDPs dynamics.

The present understanding of IDPs results from Zodiacal Light
observations, *in-situ* impact detectors and laboratory analysis of
collected particles. The processes taking place in the circumsolar
region remain poorly understood. Their description is, to a large
extent, based on theoretical studies. The possibilities of remote
sensing experiments are limited due to instrumental problems. Especially
the chemical composition and the dynamics of submicron grains can only
be investigated by *in-situ* methods. The dust observations on the Solar
Orbiter will

-   render possible the first *in-situ* detection of grains which have
    suffered extreme radiation damage and corpuscular impacts, processes
    relevant in accretion discs and important for stellar winds from
    late-type stars and for star- and stellar-system formation,

-   help in determining the extent of the dust-free zone near the Sun;

-   provide the potential for discovery of a dust disc fed partly by
    Sun-grazing comets;

-   deliver data relevant for understanding the physics of
    protoplanetary discs.

### Solar neutrons 

A large fraction of the particles accelerated by flares (and
microflares) interact in closed magnetic field lines of the solar
atmosphere (at heights much less than 3 R~s~) and do not escape into
interplanetary space. They are detected through X-ray, γ-ray, microwave
and decimetric-metric radio emissions, and through secondary particles,
such as neutrons produced in interactions with the dense layers of the
solar atmosphere. The neutron flux varies dramatically with distance
from the Sun because of beta decay. Note that the maximum distance
neutrons can reach is given by the product of their speed and their mean
lifetime (which is only eight minutes for non-relativistic neutrons!).
In fact, at the Earth's orbit neutrons with energies less than 100 MeV
can not be detected at all. Very few solar neutron events have been
reported as compared with several hundreds of γ-ray events. During the
perihelion passages of the Solar Orbiter, it will be possible to measure
directly also the less energetic neutrons for the first time.

Neutrons (and γ-rays) are produced in interactions between the primarily
accelerated ions and the constituents of the solar atmosphere.
Information about the composition and charge states of energetic
particles can provide important insights into the existence of distinct
phases of proton acceleration. This process is either very prompt and
occurs within seconds of the release of flare energy, or it is delayed
and associated with the shock wave possibly generated by the flare
ejecta. The Solar Orbiter will provide new information about the first
phase of this process and about nuclear reactions in the solar
atmosphere. Thus, the Solar Orbiter will open up a new window: solar
neutron astronomy.

### Solar radio emission

The Sun is an intense source of thermal as well as nonthermal radio
emission. The nonthermal radiation is generated by suprathermal and/or
highly energetic electrons produced by sudden releases of magnetic
energy in the active corona. Nonthermal radiation at frequencies below 1
GHz is generally assumed to be plasma radiation emitted near the local
electron plasma frequency which is a unique function of the local
electron density. Consequently, the higher and lower frequencies
correspond to the inner and outer corona, respectively, and remote
sensing of the radial electron density profile is possible.

In the corona the non-thermal radio radiation is generated by energetic
electrons. The measurement of this radiation will allow us to study the
plasma processes associated with such electron populations and to
address questions as:

-   What is the elementary process of magnetic energy release?

-   How are ions and electrons accelerated up to high energies within a
    few seconds?

-   What are the relationships between flares, coronal mass ejections,
    radio bursts, interplanetary shocks, and energetic particle events?

-   How does the quiet Sun produce suprathermal particles?

The radio spectrometer data are particularly valuable since they will be
obtained in connection with complementary data from the optical
telescopes imaging the activity site, and data of the particles and
fields instruments onboard. Detailed correlative studies between the
coronal/interplanetary electrons and the radio waves they generate will
be possible for the first time from a near-Sun and out-of-ecliptic
vantage point on the Solar Orbiter.

## THE SUN\'S POLAR REGIONS AND EQUATORIAL CORONA: EXCURSION OUT OF THE ECLIPTIC

Despite the great achievements of Ulysses and SOHO, significant
scientific questions remain unanswered. They will be addressed by the
Solar Orbiter with its combination of high latitude vantage points and
remote-sensing instruments. For example, progress in understanding the
solar dynamo will depend on how well we understand differential rotation
and the circumpolar and meridional flows near the poles of the Sun.
Torsional oscillations associated with the magnetically-active latitudes
seem to originate close to the polar regions and then propagate toward
the equator. Are the meridional flows observed at lower latitudes and
these torsional oscillations associated? The poles appear to rotate
comparatively slowly, but the polar vortex is still not well
characterised due to the serious limitations of in-ecliptic
observations. The Solar Orbiter will provide the first opportunity to
measure directly the magnetic field at the poles, as well as the surface
and subsurface flows there. We therefore anticipate that some of the
outstanding problems related to the origin of solar magnetism will be
solved.

The large-scale unipolar character of the magnetic field in coronal
holes leads to the fast solar wind which super-radially expands into the
heliosphere. The Solar Orbiter offers the opportunity of sampling the
fast wind at distances where the plasma has retained some memory of the
acceleration processes.

With regard to the sites of acceleration, SOHO spectroscopic
measurements have identified the special role of the network, but only
through very difficult Doppler measurements affected by a foreshortened
line of sight. High latitude observations allow a normal line of sight
which will provide a definite answer to the role of the network.

Even with new dedicated missions such as STEREO, it is impossible to
determine the mass distribution of large-scale structures such as
streamers and the true longitudinal extent of Coronal Mass Ejections
(CMEs). The Solar Orbiter will provide the first observations of the
complete equatorial corona and its expansion in the equatorial plane.
Furthermore, it will provide the third dimension of CMEs.

### The Sun\'s polar magnetic field and the dynamo

The internal structure and dynamics of the near polar regions of the Sun
is of paramount importance, and perhaps *the* key to our understanding
of the solar activity cycle. There are strong empirical indications that
the magnetic properties of the polar regions at solar minimum determine
the strength of the forthcoming solar maximum. It has been found that
the value of the geomagnetic aa index at its minimum is related to the
sunspot number during the ensuing maximum. At solar minimum this
geomagnetic index mostly depends on the solar wind from the polar
regions. This means that a new sunspot cycle starts in the polar
regions. This idea was used successfully to make a correct prediction of
the current solar cycle from direct measurements of the polar magnetic
field. This has been called the "Solar Dynamo Amplitude index method".

However, our understanding of how the solar dynamo actually operates is
still very poor. Apparently, the polar field is directly related to the
dynamo process, presumably as a source of a poloidal field to be wound
up by the differential rotation at the shear layer at the base of the
convection zone. Theoretical models indicate that the meridional
circulation pattern found by SOHO could be crucial for the solar dynamo.
Such a circulation pattern transports surface magnetic flux toward the
poles and downward to the shear layer. The models predict a substantial
magnetic flux concentration near the poles. The Solar Orbiter will allow
us

-   to measure reliably the strength of the polar field for the first
    time, and thus

-   to provide a new and crucial observational constraint for dynamo
    models.

Studying the internal properties of the polar regions will help to solve
the puzzle of the solar cycle, and it will provide a better
understanding of how the solar dynamo works and how other stars generate
magnetic fields.

### Polar rotation and internal flows

The polar regions are extremely dynamic. There is evidence that the
differential rotation near the poles may vary on a short time scale.
Evidence has been found for a significant variation of the subsurface
high-latitude rotation with the solar cycle. While there are large
uncertainties in these measurements because of our limited view of the
polar regions, it is evident that mapping the internal flows near the
poles will improve our understanding of the mechanism of the
differential rotation and its role in the global circulation of the Sun.
Various dynamic phenomena in the polar regions of the solar atmosphere
and corona also show close connection to the solar cycle, e.g. the
\"rush\" of Hα filaments to the poles near the periods of magnetic
polarity reversals and the sudden increase of the coronal temperature
during these periods.

However, there is no basic understanding of how these variations are
related to each other and to the internal dynamo processes. Therefore,
studying the polar regions of the Sun is fundamental to one of the key
open questions of solar physics and astrophysics:

-   How and why does the Sun vary?

With the advent of helioseismology from space, our knowledge of the
solar interior has made considerable progress. SOHO has provided
helioseismic data of extreme precision. Large-scale differential
rotation has been mapped in most of the convection zone, and temporal
variations in the zonal flows beneath the Sun\'s surface have been
detected as the present activity cycle unfolds. There have been
unexpected and still controversial discoveries, such as a buried
jet-like flow and a surprisingly slow rotation rate at high latitudes
(see Figure 2.8a).

To complement the techniques of global seismology, which has no
resolution in longitude and cannot distinguish the northern from the
southern hemisphere, new techniques of local helioseismology have
recently been developed. In particular, time-distance helioseismology
gives the opportunity to make 3-D tomographic maps of the
sub-photospheric temperature inhomogeneities, magnetic field and flows.
An important result concerns the observation at depth of the poleward
meridional circulation (see Figure 2.8b).

The Solar Orbiter offers a unique opportunity to learn about the polar
regions of the Sun, which remain largely unexplored. The main objective
will be to map rotation and meridional flows near the poles using
time-distance helioseismology. One week of data is sufficient to derive
the rotation rate in the upper convection zone, and confirm or refute
the existence of the high-speed jet at 75°. We will also be able to
study the convergence of the meridional flow at the poles and observe
how and where the solar plasma dives back into the Sun. Convective
motions over the poles, in particular supergranulation, will also be
studied.

![](media/image10.jpeg){width="2.8201388888888888in"
height="2.917361111111111in"}![](media/image11.jpeg){width="2.8125in"
height="2.8125in"}

Figure 2.8: (a) Solar internal rotation. Lines are contours of constant
angular velocity (in units of nHz). The uncertainty is large in the
shaded region. Note the slow near-surface rotation at high latitudes,
and the jet stream at latitude 75^0^. (b) Meridional circulation
measured by time-distance helioseismology as a function of latitude. The
scale on the left vertical axis gives the travel time difference in
seconds between waves travelling in the southward and northward
directions. On the right axis is the corresponding velocity, which is an
average in the 34 Mm below the surface.

### Polar observations of CMEs

The high inclinations to be reached by the Solar Orbiter will allow
unique observations of CMEs and their distributions. This will allow us:

-   to determine for the first time the longitudinal extent of CMEs,

-   to provide in conjunction with a spacecraft in Earth orbit and
    ground-based observations a 360°-longitude observation of the entire
    Sun,

-   to monitor global effects and to determine the extension of coronal
    waves associated with CMEs.

The high-latitude vantage point will allow us to determine for the first
time the longitudinal extent of CMEs, a parameter which has never been
measured yet is a basic measure of CME topology and is crucial for the
identification of CME source regions. The high-latitude observations
also allow a complete view of the equatorial corona and thus a global
view of CME activity. This has never been done and is extremely
important for the determination of CME longitudinal distribution, the
occurrence of sympathetic CME activity and the identification of large,
or even global (360°), CMEs. Global and large-area effects which may be
related to CMEs can also be determined for the first time, such as the
extent and timing of coronal waves and dimming in coronal emission
lines.

### Solar luminosity variations

The \"solar constant\", or total solar irradiance, is a measure of the
total radiative output of the Sun from the X-rays over the visible to
the infrared. As we have learned from satellite measurements, it is
changing on two different time-scales. The solar output and activity
exhibit variations that are anti-correlated over days, when large dark
sunspots cross the disc, but they are correlated over the solar cycle.
This apparent contradiction can be solved by understanding fully the
role that the excess energy in faculae and the deficit energy in
sunspots are playing on the different time scales. The variability of
solar irradiance over a solar cycle, approximately 0.1 %, is too small
to affect significantly the Earth\'s climate, however changes over
longer periods might be responsible for larger global climatic
variations.

The influence of faculae and sunspots on the value of the solar constant
value can be eliminated only by observing the Sun above the poles, where
these features are not expected to occur. Therefore, the Solar Orbiter,
at its highest inclination with respect to the ecliptic, will allow a
measurement of the \"true solar irradiance\" (TSI), as well as a
quantification of the effects of solar activity through a comparison of
these high-latitude measurements with the traditional irradiance
measurements simultaneously made in the ecliptic.

By carrying out irradiance studies from out of the ecliptic and on the
side away from the Earth, the Solar Orbiter will provide a unique
opportunity to answer questions such as:

-   How does the solar luminosity (i.e. the radiation escaping in all
    directions) vary? Does it change globally? Or is a brightening at
    the equator compensated for by a darkening over the poles?

-   Why is the irradiance variability of the Sun a factor of three
    smaller than that of Sun-like stars?

One major difference between the Sun and Sun-like stars is that whereas
we observe the Sun from almost directly above the equator, stars are on
average seen from a latitude of approximately 30^o^. The Solar Orbiter
will reach this latitude and test whether this inclination effect really
is responsible for the Sun\'s apparently anomalous behaviour.

The TSI record is not only a highly sensitive diagnostic tool for the
study of solar cycle related changes of the Sun, but also an important
contribution to the issues of natural versus anthropogenic global
change. For the latter, information over time scales of centuries is
needed, which obviously cannot be assessed from measurements during a
few 11-year solar activity cycles. Thus, an understanding of how the
solar irradiance and luminosity are modulated is of paramount
importance.

### Polar observation of the fast solar wind

The Solar Orbiter will for the first time image directly the polar
regions, where the fast solar wind is formed, from an out-of-ecliptic
position. There, the expansion of the fast solar wind has not yet been
determined directly, since the expansion is essentially perpendicular to
the line of sight of spectroscopic instruments observing from the
ecliptic, and therefore the Doppler effect is negligible. With
increasing inclination angle of its orbit, the Solar Orbiter will
perform out-of-ecliptic measurements and

-   render possible velocity measurements due to a sizeable flow
    component along the line of sight,

-   image the border of polar coronal holes and their deformation in
    response to the coronal waves that develop during CMEs,

-   observe globally the entire streamer belt in the extended corona
    (beyond 2.6 solar radii at the end of the nominal mission and beyond
    1.3 solar radii at the end of the extended mission),

-   find the magnetic signal associated with coronal plumes and clarify
    their nature.

### Secular variation of the heliospheric magnetic flux

The solar wind drags magnetic flux out of the Sun into the heliosphere
in the form of the weak interplanetary magnetic field. Magnetic flux is
also carried into the heliosphere by the sporadic and violent coronal
mass ejections, that are frequently observed also during the minimum of
solar activity. Some of the solar magnetic energy is then transferred to
the near-Earth environment through magnetic reconnection of the
interplanetary and the magnetospheric magnetic field. Recently, it has
been shown that the total magnetic flux leaving the Sun has increased by
a factor 1.4 since 1964. Indirect interplanetary magnetic field
measurements seem to indicate that the increase has been a factor of 2.3
over a century. A doubling of the Sun\'s coronal magnetic field may thus
have occurred during the past century. The change in magnetic flux may
be related to a long-term change in the Sun\'s dynamo action. The
possible influence of this secular magnetic flux variation on the global
climate on Earth has not yet been quantified.

The Solar Orbiter will determine with unprecedented accuracy the
relative contributions of the solar wind and coronal mass ejections to
the interplanetary magnetic field flux. From the out-of-ecliptic vantage
point the Solar Orbiter will allow us for the first time to derive the
longitudinal extent of coronal mass ejections and therefore give a much
more accurate assessment of the magnetic flux dragged into the
heliosphere during such huge events. Present space-borne and
ground-based coronagraphs can only measure the latitudinal extent of
coronal mass ejections. Furthermore, in the periods of co-rotation with
the Sun, the ability of the Solar Orbiter to observe the long-term
evolution of the magnetic configuration of a coronal streamer will allow
us to study the processes that lead to magnetic reconnection. The
contribution of ejected plasmoids to the heliospheric magnetic flux when
detaching from the streamers as a consequence of reconnection events
will also be determined.

# SCIENTIFIC PAYLOAD

## MEASUREMENT REQUIREMENTS

The Helios, Ulysses, Yohkoh, SOHO and TRACE missions may be used as a
baseline for designing the payload of the Solar Orbiter. However, this
mission will take a scientific space probe much closer to the Sun than
has ever been reached before by a man-made object. In particular, the
thermal load of almost 25 solar constants presents an unprecedented
technical challenge. The instruments have to be designed accordingly.
They may dwell on developments implemented in many space missions in the
past, but the extreme thermal environment requires substantial redesign.
In order to make sure that the ambitious mission goals of the Solar
Orbiter can really be achieved, all individual instruments have been
given extraordinary attention even at this stage. Indeed, some of the
instruments have been designed to a level of detail not usually achieved
at this phase of a mission. In addition, the study team has ensured that
the technology required for these instruments is available within
Europe.

The strawman payload discussed in this chapter contains sufficient
design details to demonstrate feasibility. However, it is important to
note that no pre-selection of instruments or instrument principles is
implied.

The payload must be state-of-the-art, withstand the considerable thermal
load at 45 solar radii, comply with the general requirements of a
low-mass, compact and integrated design, make use of on-board data
compression/storage and require a modest data transmission rate. The
selected payload, which meets the solar and heliospheric science
objectives of the mission, encompasses two instrument packages:

-   *In-situ* heliospheric instruments (see Table 3.13 below): Solar
    wind plasma analyser measuring electrons and ions, magnetometer for
    DC and AC fields, particle detectors for solar energetic ions and
    electrons, interplanetary dust detector, neutral particle detector,
    solar neutron detector, radio instruments. The heritage for these
    instruments is from Helios, SOHO, Ulysses, Wind and ACE.

-   Remote-sensing solar instruments (see Table 3.26 below): EUV
    full-Sun and high-resolution imager; high-resolution EUV
    spectrometer covering selected emission lines from the chromosphere
    to corona; high-resolution visible light telescope and magnetograph;
    EUV and visible-light coronagraph; and a radiometer. The heritage
    for these instruments is from Yohkoh, SOHO, TRACE.

## *In-situ* HELIOSPHERIC INSTRUMENTS 

### Solar Wind Plasma Analyser (SWA)

Scientific drivers

Three principal science goals require the inclusion of a solar wind
plasma analyser in the payload:

-   to provide observational constraints on kinetic plasma properties
    for a fundamental and detailed theoretical treatment of all aspects
    of coronal heating;

-   to investigate charge- and mass-dependent fractionation processes of
    the solar wind acceleration process in the inner corona;

-   to correlate comprehensive in-situ plasma analysis and compositional
    tracer diagnostics with space-based and ground-based optical
    observations of individual stream elements.

Furthermore, the SWA will

-   investigate in detail ^3^He and "unusual" charge states in
    CME-related flows and to correlate the observations with other
    in-situ particle and field data and with optical observations;

-   investigate the "recycling" of solar wind ions on dust grains in the
    distance range which has been located as the "inner source". Freshly
    produced pick-up ions from this inner source are specially suited as
    test particles for studying the dynamics of incorporation of these
    particles into the solar wind.

The SWA will measure separately the three-dimensional distribution
functions of the major solar wind constituents: protons, α-particles and
electrons. The basic moments of the distributions, such as density,
velocity, temperature tensor, and heat flux vector will be obtained
under all solar wind conditions and be sampled sufficiently rapidly to
characterise fully the fluid and kinetic state of the wind. In this way
we will be able to determine possible non-gyrotropic features of the
distributions, ion beams, temperature anisotropies, and particle
signatures of wave excitation and dissipation.

In addition, measurements of representative high-FIP elements (the C, O,
N group) and of low-FIP elements (such as Fe, Si or Mg) will be carried
out in order

-   to obtain their abundances, velocities, temperature anisotropies and
    charge states;

-   to probe the wave-particle couplings (heavy-ion wave surfing);

-   determine the freeze-in temperatures (as a proxy for the coronal
    electron temperature).

Instrument concept

In view of the limited resources of mass, volume and telemetry allocated
to the SWA, a compromise between sensitivity, mass/charge- and mass- and
time resolution has to be found. The SWA has to cover a large dynamic
range in ion fluxes. Since there is an enormous difference between the
proton fluxes at perihelion (typically 10^14^ m^-2^ s^-1^) and the
fluxes of relevant minor ion tracers at 1 AU (e.g. Fe^10+^ at typically
10^8^ m^-2^ s^-1^ etc.) it is suggested to implement three different
sensors:

1.  A Proton/α-particle Sensor (PAS) with the principal aim to
    investigate the velocity distribution of the major ionic species at
    a time resolution equivalent to the ambient proton cyclotron
    frequency.

2.  An Electron Analysers System (EAS) consisting of two or three
    combined sensor heads covering a solid angle of almost 4π and
    allowing the determination of the principal moments of the electron
    velocity distribution at high time resolution.

3.  A Minor Ion Sensor (MIS) which allows the independent determination
    of the major charge states of oxygen and iron and a coarse mapping
    of the three-dimensional velocity distribution of some prominent
    minor species. Also, pick-up ions of various origins, such as
    weakly-ionised species (C^+^, N^+^, Ne^+^, Mg^+^, Si^+^, etc.), will
    be measured.

The PAS is designed to resolve the relevant ion-kinetic scales. It will
achieve a time resolution which is compatible with the characteristic
ion cyclotron periods near perihelion. The characteristic time scale is
about 10 ms. For the minor species (heavier than He) a time resolution
of 1 s is needed.

To fulfil the requirements of compositional diagnostics, at least the
charge state distributions of oxygen and iron should be measured
routinely and with a time resolution of better than 10 s.

The instrument specifications are outlined in Table 3.1:

  ------------ ------------------ ---------- -------------- -------------- --------------
  **Sensor**   **Particles**      **Field of **Energy       **Angular      **Time
                                  View**     resolution**   resolution**   resolution**

  PAS          Protons and        Conical    E/∆E = 10      Azimuth 30°\   10 ms
               α-particles        ±45°                      Zonal ±5°      

  EAS          Electrons          4π         E/∆E = 10      Azimuth 30°\   10 ms
                                                            Zonal ±5°      

  MIS          O, (3He, C, Ne,    Conical                                  1 s
               Mg, Si), Fe        ±45°                                     
  ------------ ------------------ ---------- -------------- -------------- --------------

Table 3.1: Specifications of the SWA sensors.

The instrument package proposed for SWA is a further development of
instruments successfully flown on several missions in the past, e.g.,
SWICS on Ulysses, CELIAS/CTOF on SOHO, TRIPLET on INTERBALL. A possible
scheme for PAS based upon the design of TRIPLET is shown in Figure 3.1.

Resources and requirements

Telemetry

The transmission of velocity distributions of the major species
(protons, electrons and α−particles) drives the telemetry requirements
of SWA. In the normal mode of operation this information should be
compressed to transmit the principal moments of the distributions.
Periodically however, the transmission of the complete velocity
distributions of protons and α−particles with a (velocity) resolution of
10 x 10 x 10 pixels at a depth of 5 bits is needed. This requires 5 kb
per spectrum. Correspondingly, at a time resolution of 10 ms, the
required telemetry rate is 500 kb/s. This rate should be maintained
during 1s, i.e. for some 100 ion cyclotron periods to cover a typical
growth time of plasma instabilities. With an average telemetry rate of 5
kb/s a duty cycle of the order of 1% can be achieved.

![](media/image12.png)

Figure 3.1: Schematics of PAS (following the design of TRIPLET on
INTERBALL). A hyperbolic mirror maps the angular velocity distribution
of solar wind ions onto a coded aperture. With this measure the
information on the position at the entrance of the hemispherical energy
analyser is transmitted (at the cost of \~ 50% reduction in beam
intensity) to the exit aperture of the analyser. The additional
information on energy per charge is contained in the exit location and
convoluted with the entrance location. Deconvolution of the entrance
angle and the energy per charge by the onboard DPU allows the
reconstruction of the complete three-dimensional velocity distribution
of solar wind protons and α particles with a time resolution of 10 ms.

Allocations of mass, volume, and power.

-   A mass allocation for MIS of 4 kg, 1 kg for PAS and 1 kg for EAS
    (excluding the electronics) is foreseen.

-   The volume allocation for MIS is 20 x 20 x 20 cm^3^, and 10 x 10 x
    10 cm^3^ for PAS and for each of the two EAS sensors.

-   The total power allocated to SWA is 5 W.

Thermal and radiation requirements

The instrument should be able to operate under the specified spacecraft
conditions. Special attention has to be devoted to the extreme solar UV
radiation flux. The design of the particle instruments should reject
this radiation to a sufficiently low level in order not to compromise
the detection of rare particles.

Electrostatic cleanliness

Special care has to be taken to ensure near-equipotential conditions in
the vicinity of the EAS entrance apertures.

Instrument summary

  ---------------- ------------ ------------- ---------------- ----------------
  **Instrument**   **Mass       **Power (W)** **Volume         **Data rate
                   (kg)**                     (cm^3^)**        (b/s)**

  PAS              1            1.2           1000             2000

  EAS              1            0.8           2000             2000

  MIS              4            3             7000             1000

  Total SWA        6            5             10000            5000
  ---------------- ------------ ------------- ---------------- ----------------

Table 3.2: SWA resource requirements.

### Radio and Plasma Waves Analyser (RPW)

ScienTIFIC drivers

*In-situ* wave investigations on many spacecraft (e.g., Ulysses, Wind)
have amply demonstrated the crucial role of wave observations in the
study of a broad range of solar terrestrial phenomena. Wave observations
have been particularly useful in correlation with SOHO optical and EUV
observations. Typical waves instrumentation provides measurements of
both the electric field and magnetic field in a broad frequency band,
usually from a fraction of a Hertz up to several tens of MHz, covering
characteristic frequencies in the solar corona and interplanetary
medium. Both electrostatic waves and electromagnetic waves can be
measured, leading to different diagnostics: electrostatic waves mostly
provide *in-situ* information in the vicinity of the spacecraft, while
electromagnetic waves can provide an extensive remote sensing of
energetic phenomena in the solar corona and interplanetary medium.

Solar radio astronomy in a broad range of frequencies provides a unique
means to monitor, track and analyse energetic phenomena taking place in
the solar corona and interplanetary medium, in association with solar
activity. Radio emission is essentially generated through the
interaction of energetic electrons (produced in solar flares and shock
waves) with the ambient solar wind plasma. Most radiation mechanisms
produce radiation at a frequency close to the local plasma frequency
which is directly proportional to the square root of the electron
density N~e~ , leading to the strong frequency-distance dependence shown
in Table 3.3.

  ------------------ ----------------------------- ----------------------
                     **electron density Ne         **radio frequency**
                     (cm^-3^)**                    

  Low corona         \~ 10^8^ -- 10^10^            100 MHz -- 1 GHz

  Medium corona      \~ 10^6^ -- 10^8^             10 MHz -- 100 MHz

  \~ 10 Rs           \~ 10^4^                      \~1 MHz

  \~ 40 Rs           \~ 10^3^                      \~ 300 kHz

  \~ 1 AU            \~ 10                         \~ 30 kHz
  ------------------ ----------------------------- ----------------------

Table 3.3: Radio frequency as function of electron density.

Using proven techniques, the Solar Orbiter mission profile will allow a
radio and plasma wave package to face quite different and new
situations: access to waves and turbulence that occur much closer to the
Sun; access to viewing angles from well out of the ecliptic plane,
permitting for the first time studies of the north-south symmetry of the
radio radiation in the solar corona.

Instrument concept

The Radio and Plasma Waves Analyser (RPW) comprises two main
sub-systems: the Plasma Waves Analyser (PWA) covering *in-situ*
measurements and the Radio Spectrometer (RAS) for remote sensing. The
two sub-systems share some of the sensors (essentially the low frequency
boom antenna) and have a common digital signal processing unit and
interface with the spacecraft.

The same receivers can be used to analyse the different types of waves
detected by different sensors, for instance electric antennae and
magnetic coils. There is a lot of experience in the design of the
sensors and receivers, as well as in the analysis techniques. Also all
aspects of electromagnetic cleanliness, are well understood and are no
longer a problem for modern spacecraft designs (e.g. Ulysses,
Wind/Polar, Cassini, IMAGE, Cluster).

Concept of the PWA

The PWA will identify the various plasma waves and kinetic modes
comprising the high-frequency part of the fluctuation and turbulence
spectra. We provide below a brief description of the characteristics of
the plasma waves expected to be present at 0.2 AU.

The expected field strength may range between a few μV/m and mV/m, and
up to about 1 V/m for the convection electric field. The magnetic field
strength is expected to vary between a few nT and μT, with large
differences between the longitudinal and transverse components with
respect to the mean magnetic field and the solar wind flow direction
(substantial Doppler shifts are to be expected).

From Helios observations one can estimate that a sensitivity equal to
10^‑6^nT/(Hz)^1/2^ will be required at 0.2 AU to identify without
ambiguity whether the observed waves are electromagnetic or
electrostatic.

At and above the electron plasma frequency intense electron plasma
oscillations and solar radio waves are expected. Helios observations
show that wave intensities associated with type III bursts increase very
rapidly with decreasing radial distance from the Sun. Intense emission
could extend up to 10^-4^ nT/Hz^1/2^.

The measurement requirements necessary for the identification and
analysis of the various electromagnetic phenomena expected to be
measured at 0.2 AU are given in Table 3.4.

  ----------------------- ------------- ------------------- ----------------
  **Phenomenon**          **Frequency   **Minimum amplitude **Maximum
                          band**        (nT/Hz^1/2^)**      amplitude
                                                            (nT/Hz^1/2^)**

  Doppler-shifted ELF/VLF 1 Hz -- 10    10^--5^ at 10 Hz    10^--1^ at 10 Hz
  waves and turbulence    kHz                               

  Ion acoustic waves and  10 kHz -- 10  A few 10^--7^ at 1  10^--4^ at 1MHz
  Radio waves             MHz           MHz                 
  ----------------------- ------------- ------------------- ----------------

Table 3.4: Measurement requirements for the fluctuating magnetic field.

The PWA will cover a broad band in frequencies, extending from about 1
Hz into the MHz range. Resolving the vector components of the electric
and magnetic fields is scientifically highly desirable to determine wave
modes unambiguously. However, due to different constraints of the
mission, we propose to measure one electric field component only, with a
boom antenna mounted in the anti-Sun direction, i.e. in the shadow of
the spacecraft. The three components of the fluctuating magnetic field
can be easily measured with a 3-axial search coil magnetometer arranged
in a compact configuration and mounted on a short boom which should also
point into the anti-Sun direction. This boom is required for magnetic
cleanliness reasons and could be shared with the MAG (see 3.2.4) as long
as a minimum distance between the two sensors is respected. There is a
strong heritage for those sensors mounted on three-axis stabilised
spacecraft, e.g. on Galileo, Cassini, STEREO. The PWA will also benefit
from studies and developments under progress for the Solar Probe.

The PWA will perform on-board processing of the data and deliver
selectable wave forms (Time Domain Sampler, TDS), to identify non-linear
coherent fluctuations and correlate them with measurements from the EPD.

Concept of the RAS

The RAS will measure the solar and interplanetary radio waves in the
frequency range from 100 kHz to 1 GHz, with a sweep period between 0.1 s
and 10 s and a high spectral resolution (∆f/f ≈ 0.07). The RAS will
observe plasma processes associated with energetic electrons from the
low corona up to about 0.5 AU. It will probe the plasma at distances
ranging from the solar surface to the spacecraft location, thereby
connecting the low-altitude coronal regions observed by the optical
instruments with the near-Sun heliospheric conditions specified by the
in-situ measurements. Since radio radiation is generally beamed (beam
widths sometimes down to a few tens of degrees) more or less along a
radial direction from the Sun, this technique is particularly relevant
for different vantage points, for instance when the Solar Orbiter
observes the far side of the Sun. The time history provided by the
regular acquisition of the radio dynamic spectrum will help to trace the
development of an active region in a synthetic manner. Activity indices
could thus be produced.

The time resolution required to detect the rapidly varying solar bursts
varies with the radio frequency. Typically, the duration of a type III
burst (stream of energetic electrons) is ∆t(s) = 220 / f(MHz). This
points to time resolutions of the order of 0.1 s or better for the high
frequencies and of 10 s for the low frequencies.

A single boom antenna can be used for the low frequency part of the band
(below, e.g. 20 MHz), where the sensitivity is essentially determined by
the sky background radiation. Above 20 MHz and particularly above
200 MHz, the background is determined by the receiver temperature and
gain is needed in order to detect solar radio bursts: a simple boom
cannot be used anymore. A broad-band antenna with significant gain is
thus required. Each one of the receivers will be fed by different
antenna or search coil systems.

The RAS instrument could be divided into three sub-spectrometers, for
instance:

-   RAD1: 100 kHz - 16 MHz

-   RAD2: 16 MHz - 200 MHz

-   RAD3: 200 MHz - 1 GHz

Different designs are used on many space missions and the heritage goes
right back to the birth of space radio astronomy at long wavelengths.

RPW instrument description

An overall block diagram of the Solar Orbiter RPW instrument is shown on
Figure 3.2.

![](media/image13.png){width="6.522222222222222in"
height="4.049305555555556in"}

Figure 3.2: Block diagram of the Solar Orbiter RPW instrument.

Plasma waves and low frequency activity will be monitored by a receiver
covering the band from about 1 kHz up to 100 kHz. In order to allow for
a high time resolution, whilst minimising resources, it can be of the
type flown on Wind and planned for STEREO, using a digital filtering
technique.

The wave form analyser or Time Domain Sampler (TDS) is a broad band
waveform sampler with 4 channels (1 electric field component and 3
magnetic field components). The signal from the boom antenna and search
coil are sampled at a rate up to 10^6^ samples per second. Events are
detected on board and transmitted according to the available bit-rate.
This is done using "intelligent" algorithms which can return a quality
value for the events. A number of random events can be sent to the
ground without regard of their quality in order to provide unbiased
statistics. Real time triggers can be sent to the EPD for special
studies of wave-particle phenomena.

The radio receivers can be of the classical super-heterodyne type.
Frequency synthesisers will allow for a maximum flexibility in the
choice of the observing frequencies (for instance to avoid "polluted"
frequencies on the spacecraft). Each sub-receiver can consist of up to
256 selectable channels. Only a selection of these channels could be
transmitted to the telemetry stream.

The DPU will command the different subsystems and interface with the
spacecraft. The variety of instruments call for a packetised telemetry
system, that can be managed by the DPU itself, as is done for similar
packages on the Wind spacecraft (fixed format telemetry system) and
STEREO. The DPU will interface with other instruments on board the
spacecraft, particularly the EPD. The DPU could be shared with some
other instruments from the payload, e.g. with MAG.

Tables 3.5a, 3.5b, 3.5c provide the main characteristics of the sensors
and analysers comprised in the RPW investigation. Numbers are indicative
and can vary according to design options

  --------- -------------- ------------- ---------------- ------------ -------- -----------
            **Magnetic     **frequency   **Sensitivity    **Size**     **mass   **power**
            sensors**      band**        (nT/Hz^1/2^)**                (g)**    

  B1        Search coils   1 Hz -- 10    10^--5^ at 10 Hz 3 rods, 18   500      200 mW
                           kHz                            cm each               

  B2        Magnetic loop  10 kHz -- 10  10^--6^ at 1MHz  20 cm        600      200 mW
            (optional)     MHz                            diameter              

  *Total*                                                              *1100*   
  --------- -------------- ------------- ---------------- ------------ -------- -----------

Table 3.5a: Main characteristics of the RPW magnetic sensors.

  --------- ------------------ -------------- ----------- ----------- -----------
            **Electric         **frequency    **Size**    **mass      **power**
            sensors**          band**                     (g)**       

  E1        Radio boom         1 Hz -- 20 MHz 5 m boom    1300        N/A

  E2        Active dipole      20 MHz -- 200  \~ 1m x 2cm 800 (TBC)   300 mW
                               MHz            x 2 cm                  

  E3        Log-periodic       200 MHz -- 1   \~ 1m x 1m  1500 (TBC)  N/A
            antenna            GHz            x 1cm                   

  *Total*                                                 *2600*      
  --------- ------------------ -------------- ----------- ----------- -----------

Table 3.5b: Main characteristics of the RPW electric sensors.

  -------------- -------------- --------------- ------------------ -------- ----------- ---------
  **Analyser**   **frequency    **Sensor(s)**   **Sensitivity**    **mass   **Bitrate   **Power
                 band**                                            (g)**    (b/s)**     (W)**

  TDS            wave form      E1, B1          30mV/m, 0.2 pT     800      1700        1.0
                 1,000,000                                                              
                 samples/s                                                              

  PWR            2 kHz -- 128   E1, B1          10 nV/Hz^1/2^,     800      1250        0.7
                 kHz                            10^--5^ nT/Hz                           

  RAD1           100 kHz -- 16  E1, (B2)        10 nV/Hz^1/2^,     800      800         0.8
                 MHz                            10^--6^ nT/Hz^1/2^                      

  RAD2           16 MHz --200   E2              \~ 10^-20^ W m^2^  700      600         0.7
                 MHz                            Hz^-1^                                  

  RAD3           200 MHz -- 1   E3              \~ 10^-20^ W m^2^  700      600         0.7
                 GHz                            Hz^-1^                                  

  DSP/DPU                                                          1200                 1.5

  Converter                                                        800                  1.2

  Harness etc.                                                     500      50 (H/K)    

  Total                                                            6300     5000        6.8
  -------------- -------------- --------------- ------------------ -------- ----------- ---------

Table 3.5c: Main characteristics of the RPW analysers and subsystems.

Resources and requirements

Antenna systems

Antenna systems for the range from 20 MHz to 1 GHz are readily available
off the shelf for ground based systems, including "portable" ones.
Different types and techniques are available, and the theory of the
measurement is well known. They would have to be developed specifically
for the Solar Orbiter mission. The RAD2 band (16 MHz - 200 MHz) could be
covered by an active dipole antenna and the RAD3 band (200 MHz - 1 GHz)
by a log-periodic antenna. These two antenna systems can be grouped
together and mounted on the same boom in order to have a common
mechanical (deployment mechanism) and electrical interface with the
spacecraft. They would have to be pointed to the Sun (accuracy of a few
degrees) and thus protected from the direct sunlight by a shield not
affecting the radio wave propagation. The whole system can be kept to a
reasonable size (longest element of the order of 1 m) and provide the
required gain. The beam shape of the antennae mounted on the spacecraft
will have to be modelled and measured using existing techniques.

Electro-Magnetic Cleanliness (EMC)

Many of the RPW scientific objectives are based on the measurement of
low-level signals for which maximum sensitivity is required. Several
inexpensive measures can be taken at the spacecraft and project level to
ensure that the Solar Orbiter spacecraft is clean from the point of view
of both conducted and radiated electromagnetic interference. The
sensitivity of the RPW instrument will be approximately 10^-8^
V/m/Hz^1/2^ in the frequency range of 1 kHz to 20 MHz and 3×10^-7^
V/m/Hz^1/2^ below 100 Hz. In addition, the wave form analyser is
sensitive to impulsive interference of duration as short as a fraction
of a microsecond. Although these sensitivities may appear to require an
excessively \"clean\" spacecraft, they are not difficult to achieve if
good EMC practices are incorporated in the spacecraft design.

Instrument summary

RPW Sensors:

-   triaxial search coil mounted on the magnetometer boom,

-   magnetic loop antenna,

-   boom antenna (5 to 6 m on side opposed to Sun),

-   VHF/UHF system comprising an active dipole and a log-periodic
    antenna.

Role of main subsystems:

-   RPW/PWA: identify plasma waves and kinetic modes (high frequency
    part of the fluctuation and turbulence spectra)

-   RPW/RAS: measure and track solar and interplanetary radio bursts;
    connect low-altitude coronal regions observed by the optical
    instruments with the *in-situ*, local solar wind conditions at the
    spacecraft, study directivity of radio emissions

  ---------------- ------------ ------------- ---------------- ----------------
  **Instrument**   **Mass       **Power (W)** **Volume         **Data rate
                   (kg)**                     (cm^3^)**        (b/s)**

  **Total RPW**    10           7.5           7500             5000
  ---------------- ------------ ------------- ---------------- ----------------

Table 3.6: RPW resource requirements.

### Coronal Radio Sounding (CRS)

A passive radio science experiment can be carried out using the
available radio links. The spacecraft will pass behind the Sun (superior
solar conjunction) on many occasions, making it possible to investigate
the solar corona by radio sounding at solar distances much less than the
perihelion distance and down to at least 2 solar radii. Integrated
line-of-sight parameters such as electron content (densities), Faraday
rotation, scintillations and angular broadening can be recorded. The
Solar Orbiter orbit and the proximity of the spacecraft to the Sun will
create unique situations for this kind of analysis. The CRS
investigation with a two-way radio link via the spacecraft high-gain
antenna (ranging capability) would require a radio subsystem with
dual-frequency phase coherent downlinks at X-band and Ka-band. Linear
polarisation of the downlink signals would enable Faraday rotation
measurements as an option. The RF power for both radio links (X and Ka)
planned for this mission is sufficient for this investigation. A two-way
dual-frequency coherent radio link (X-band uplink; X-band and Ka-band
simultaneous coherent downlinks) is considered as the optimal
configuration for a sufficiently stable link and for detecting
signatures of CME events traversing the radio ray path.

It is presently unclear whether or not an ultra-stable oscillator would
be necessary to guarantee the frequency stability of the downlinks. If
only a one-way radio link (spacecraft to ground) is feasible due to
operational constraints during solar conjunction, the transmitted radio
signal must be stabilised by an onboard ultra-stable oscillator
(frequency stability 10^-11^ to 10^-12^ at 3 s integration time, 200g, 3
W). The practicality of using linear polarised signals for Faraday
rotation measurements is another trade-off to be explored. The final
decision depends critically on the geometry and mission operation plan.
Considerable interaction between the radio scientists and the
spacecraft's radio subsystem contractor is a prerequisite for conducting
a successful coronal sounding experiment. The occultation geometry
during the solar conjunctions is another aspect that needs to be studied
in detail.

### Magnetometer (MAG) 

Scientific drivers

Knowledge of the magnetic field is of crucial importance for
characterising the state of the local interplanetary medium and is
indispensable for the interpretation of measurements of energetic
particles and thermal plasma particles.

The magnetometer will measure the three orthogonal components of the
interplanetary magnetic field vector over a broad dynamic range from a
few nT up to the μT range, and for frequencies well below the local
heavy-ion or proton gyrofrequencies.

Instrument concept

The magnetometer may consist of a triaxial identical sensor system. A
compromise has to be found between the requirements of the spacecraft
and payload systems and the demand of the magnetometer for magnetically
clean conditions.

The heritage of MAG includes dozens of similar instruments flown on
various magnetospheric, planetary and deep space missions. The MAG
proposed for the STEREO mission might serve as a guideline. It is in
turn based on the design used for the NEAR and ACE missions. The MAG
sensor and electronics for the Solar Orbiter require no new development.
State of the art instruments can be built for less than 1 kg. Further
mass savings may be possible.

MAG might consist of a 3-axis fluxgate magnetometer with the sensor
mounted at the end of an extendable boom (length \<1m) which is
positioned in the shadow of the Orbiter. Other mounting locations such
as on the high-gain antenna structure appear possible, provided MAG
stays in permanent shadow. It measures fields in 8 ranges (automatically
selected) from 4 nT to 65536 nT with a resolution of 30 ppm, from DC up
to AC frequencies of 500 Hz. The parameters can be adjusted according to
the particular needs of the Solar Orbiter mission. The DPU of MAG could
be shared with either the plasma instruments or the wave instruments.

Resources and requirements

The mass of the sensor package will be less than 1 kg, the average power
consumption is about 1 W, and the required telemetry is of the order of
200 b/s.

Magnetic cleanliness of a spacecraft as required by a magnetometer on
board has been a cost driver for some missions in the past. However,
more recent experience has shown that there are in fact acceptable
low-cost engineering solutions possible. A reasonable magnetic
cleanliness program can be performed, if the MAG experimenters work in
close contact with the all other teams in order to ensure proper
engineering at an early phase and to make other precautions without
impacting the mission schedule. Similar schemes were applied in the past
for several successful missions.

Instrument summary:

  ------------------- ---------------------------------------------------
  Instrument          triaxial fluxgate magnetometer

  Telemetry           \< 200 b/s

  Mass                \< 1 kg

  Power               \< 1 W

  Special             mounting on short boom (\<1 m), in shadow of
  requirements        spacecraft,\
                      basic magnetic cleanliness of spacecraft.
  ------------------- ---------------------------------------------------

Table 3.7: MAG resource requirements.

### Energetic Particle Detector (EPD) 

Scientific drivers

The Solar Orbiter will allow us to study the sources, acceleration and
propagation of solar energetic particles in association with coronal and
interplanetary shocks. Energetic pick-up particles originating from
outgassing or sputtering of near-Sun dust should be measured as well in
conjunction with the plasma measurements. It is essential that all
measurements are done fast, typically at 1 s, with complete angular
coverage to resolve the pitch-angle distributions. The EPD will

-   determine *in-situ* the generation, storage, release and propagation
    of different species of solar energetic particles in the inner
    heliosphere;

-   identify the links between magnetic activity and acceleration on the
    Sun of energetic particles, by virtue of combined remote-sensing of
    their source regions and *in-situ* measurements of their properties,
    while staying magnetically connected with the acceleration sites
    during solar co-rotation passes;

-   characterise gradual (CMEs) and impulsive (flares) particle events
    and trace their spatial and temporal evolution near the Sun.

Instrument concept

The EPD will determine chemical and charge composition and energy
spectra of ions in a wide energy range, from about the typical solar
wind energies of a few keV to several 100 MeV/nucleon for protons and
heavy ions. Electrons should be measured from 10 keV to 10 MeV. The
combination of electrostatic E/Q-analysis with time-of-flight
E/M-determination and subsequent direct energy measurement in a solid
state detector has been employed in many EPDs in the past and is also a
possible design option for the Solar Orbiter.

The specific design suggested here resembles closely the one recently
selected for the IMPACT instrument on the STEREO mission. It is a
multi-head sensor system using solid state technology. Ongoing research
in this field is promising and will certainly lead to further
improvements concerning energy thresholds and noise level as well as to
size reduction.

Resources and requirements

The measurement requirements can be met readily by a low-mass and
compact sensor assembly with different lines of sight and modest
aperture sizes. We propose to place several detectors at different
positions onboard the spacecraft, in order to obtain a sufficient field
of view and angle-of-incidence coverage without using a scanning
platform. The whole EPD package will have a total mass not exceeding 4
kg and consume an average power of 3 W. The required telemetry rate is
1.8 kb/s.

The characteristics of the EPD could, e.g. resemble the 4 sensors of the
IMPACT instrument on STEREO, as given in Table 3.8.

  ------------ ----------------- -------------- -------- ------------------------
  **Sensor**   **Energy          **GF (cm^2^sr  **FOV (° **Name**
               (MeV/nucleon)**   )**            x °)**   

  EPT          e: 0.02 - 0.40    4 x 0.30, 4 x           Electron Proton
               p:\               0.24,\                  Telescope
               0.02 - 7.00       4 view cones            

  SIT          0.03 - 2          0.3, 1 view    17 x 44  Supra-thermal Ion
                                 cone                    Telescope (He -- Fe)

  LET          1.5 - 40          4.5, 10 view   130 x 30 Low Energy Telescope (He
                                 cones                   -- Ni)

  HET          13 - 100          2.4, 2 view    48 x 48  High Energy Telescope
                                 cones                   (He -- Fe)
  ------------ ----------------- -------------- -------- ------------------------

Table 3.8: EPD instrument characteristics.

Instrument summary

  ----------------- ----------- ------------ --------------- ---------------
  **Instrument**    **Mass      **Power      **Volume        **Data rate
                    (kg)**      (W)**        (cm^3^)**       (b/s)**

  **EPT**           0.94        0.60         300             300

  **SIT**           0.78        0.55         620             500

  **LET**           0.44        0.12         430             700

  **HET**           0.27        0.04         210             300

  **DPU/H-L-VPS**   1.37        1.17         3000            

  **Total EPD**     3.80        2.48         4560            1800
  ----------------- ----------- ------------ --------------- ---------------

Table 3.9: EPD resource requirements.

### Dust Detector (DUD) 

Science drivers

The DUD will analyse interplanetary dust particles with respect to their
composition and orbital characteristics and determine *in-situ* the
distribution, composition and dynamics of dust particles in the near-Sun
heliosphere in and out of the ecliptic. The spatial distribution of
particles with masses ranging between 10^-16^ g and 10^-6^ g will be
determined. DUD will

-   help outlining the extent of the dust-free zone around the Sun;

-   discover the sources of dust, e.g. from Sun-grazing comets;

-   unravel the role played by near-Sun dust for pick-up ions;

-   deliver data relevant for an understanding of proto-planetary discs.

Instrument concept

The science objectives can be met with an instrument of, e.g. Giotto
heritage. The dust experiment may consist of a multiple sensor assembly,
with each sensor looking in a different direction to resolve incidence
angles. A low-mass dust-detecting element could look as sketched in
Figure 3.3. The dust particles enter the detector through a thin
aluminium foil and evaporate during impact on a gold target. A set of
wires in front of the target plate collects the ions produced. The
signal is analysed to deduce the mass and possibly speed of the dust
particle.

![](media/image14.png)

Figure 3.3: Instrument design of DUD.

Full chemical analysis and mass resolution of the important elements H,
C, N, O of carbonaceous material and the metals like Na, Mg, Al, Si, Ca
and Fe contained in chondritic silicates is desirable but may require a
more sophisticated design.

Resources and requirements

Two identical detector units should be located on the sides of the
spacecraft body, one looking 90° off the spacecraft-Sun line in the
orbital plane to measure the near-ecliptic dust particles, the other one
in the direction perpendicular to the orbital plane in order to measure
dust particles in orbits highly inclined to the ecliptic. The field of
view of each sensor is ± 80°.

The mass of each sensor is 0.5 kg. Each one consumes a power of 0.3 W
and has a volume of 1200 cm^3^. The average bit-rate will not exceed 50
b/s. The DPU should be shared with other particle instruments.

Instrument summary

  ------------ ------------ ------------- ---------------- ----------------
  Instrument   Mass (kg)    Power (W)     Volume (cm^3^)   Data rate (b/s)

  DUD          1            0.6           2400             50
  ------------ ------------ ------------- ---------------- ----------------

Table 3.10: DUD resource requirements.

### Neutral Particle Detector (NPD)

Scientific drivers

The Lα line of neutral hydrogen is used in spectroscopic observations on
SOHO to determine the solar wind velocity profile in the corona up to
several solar radii. The NPD on the Solar Orbiter will measure primarily
neutral hydrogen atoms and thus render possible the detection of the
interplanetary \"neutral solar wind\". The instrument should detect
neutral atoms in the typical solar-wind and supra-thermal velocity range
between about 350 and 5000 km/s.

The neutral atom flux is expected to be about 100-1000 atoms cm^-2^
s^-1^ at 0.21 AU, but it could be up to 10^6^ atoms cm^-2^ s^-1^ in a
CME. The NPD will also measure energetic neutral atoms emitted from
various coronal sources, and thus enable images of these coronal
emission regions to be constructed from rays of neutral atoms with
different velocities.

Instrument concept

Neutral atom detectors have been flown on SOHO, Cassini and IMAGE.
Time-of-flight instruments are for example flown on Ulysses, SOHO and
ACE.

As a baseline, the instrument parameters of the NPD on the Solar Orbiter
could be: Energy/mass range: 0.6 to 100 keV/nucleon, moderate velocity
resolution; mass resolution is a secondary objective. The field of view
(for the imaging capability) of the NPD should be centred at about 17°
off the spacecraft-Sun-centre line.

The instrument entrance system must block radiation emerging from the
solar disc. The UV scattered in the solar corona, i.e. Lα, must be
suppressed by a factor of more than 10^-10^. The entrance system must
reduce ions and electrons of the solar wind plasma by a factor of about
10^-8^ whilst transmitting neutral atoms.

A feasible neutral atom detector (see Figure 3.4) may consist of two
sections: the first compartment accommodates the baffle plates keeping
out solar UV and the deflection unit for keeping off unwanted charged
particles. The second section contains a time-of-flight (TOF)
double-coincidence detector. Between the two sections is the
pin-hole/slit that defines the optical properties of the neutral atom
detector.

The direct solar radiation is blocked by a collimator. For the
suppression of scattered light within the collimator, the plates are
covered with sawtooth-like structures. The collimator consists of,
alternatively positive and negative, high-voltage biased plates and
therefore acts as a charged-particle deflector. The collimator plate
length, distance and potential determine the cut-off of the
charged-particle transmission (about 100 keV/q). Lα light is transmitted
through the collimator. The grid in front of the pin-hole reflects the
UV. The grid consists of 2 gratings with a grating constant matching the
Lα peak. The grids are more transparent for particles than for UV
photons (ratio up to 105 per grating, in dependence on the UV
wavelength).

The neutral atoms pass through the pin-hole (or slit) and a 0.5 to 1.1
μg/cm2 carbon foil. As they pass the carbon foil, they release secondary
electrons which are accelerated towards a START MCP (multi-channel
plate). The atoms travel through a 3-cm-wide TOF drift region and hit
the rear STOP MCP. The time of flight for 1 keV/n-100 keV/n atoms is
about 8 to 80 ns. The position resolution is 1-D for the START MCP and
2-D for the STOP MCP. The flight pass of the neutral atom can be
calculated from these measurements by an inversion of the instrument
function. The detector determines the velocity (energy/nucleon) and the
flight pass (imaging) of the neutral atoms.

Resources and requirements

The NPD can be built for a mass of 1 kg. The power consumption is 2 W.
The required telemetry rate is less than 0.3 kb/s. It is envisaged to
use a common DPU and to share the mechanical structure with the SWA. The
FOV has to be 17° off the spacecraft-Sun line because of the typical 10°
aberration of a 1 keV hydrogen atom travelling at solar wind speed.

![](media/image15.png){width="4.825in" height="3.1006944444444446in"}

Figure 3.4: Instrument schematics of the Neutral Particle Detector.

Instrument summary

  ------------- ------------ ------------- ---------------- ----------------
  Instrument    Mass (kg)    Power (W)     Volume (cm^3^)   Data rate (b/s)

  NPD           1.00         2.00          2000             300
  ------------- ------------ ------------- ---------------- ----------------

Table 3.11: NPD resource requirements.

### Neutron Detector (NED)

Scientific drivers

Solar neutrons (and γ rays) are a by-product of solar flares due to
spallation reactions between energetic protons and α-particles. The
overwhelming majority of these neutrons are thermalised within minutes
and later absorbed by hydrogen to form deuterium and the well known 2.2
MeV γ−line. Some neutrons, however, escape into space and can eventually
be detected (before they decay with a half-life of 619.9 s) by a sensor
located in close proximity to the Sun. The temporal correlation of
neutron fluxes with solar energetic particle fluxes and with optical
observations in the EUV provides valuable information on the solar
particle acceleration process, which is inaccessible by other means.

The advantage of neutrons over charged particles not being delayed by
travel along complicated paths through magnetic fields will allow an
unambiguous temporal correlation with optical observations of features
of flares in different energy ranges.

Including the NED on the Solar Orbiter offers routine monitoring of
neutron fluxes and, optionally, γ- rays and will be a "first" in solar
science. A major contribution to the solar γ-ray flux stems from the 2.2
MeV line, which is due to the reaction of thermal neutrons with protons
to form deuterium. Hence, the γ-ray detector registering "extinct" solar
neutrons would ideally complement the neutron detector, which is to
record "live" neutrons in-situ.

Instrument concept

Following the design of the instrument proposed for the Solar Probe the
heart of the suggested NED instrument consists of a cell containing thin
(\<10 μm) thorium foils. With a cross section of typically 100 mbarn (at
energies above 1 MeV) an efficiency of 10^-5^ for producing fission
fragments - to be detected in a pulse ionisation chamber - can be
achieved. The anti-coincidence chamber surrounding the sensitive volume
contains a counting gas, e.g. methane or argon, and energetic particles
which might produce false counts in the ionisation chamber will be
detected with an efficiency of close to 100%. The angular resolution of
the NED will not be sufficient to map individual features on the solar
surface.

Optionally, incorporation of a facility to monitor solar γ-rays into the
experiment would greatly enhance the scientific return.

Planetary flybys could be used to block the solar neutron flux and to
determine the intrinsic background rate of the spacecraft. Further
information, also on the average energy of solar neutrons will be gained
by comparison of long time averages of the neutron flux at different
solar distances.

Resources and requirements

The main concern with NED will be to minimise the energetic particle
background in the sensor cell. On one hand the anti-coincidence cell
should be as efficient as possible to prevent false coincidences, on the
other hand, the location of the sensor within the payload has to be
chosen such that as to keep the particle radiation level to a minimum.

![](media/image16.png){width="4.708333333333333in" height="3.0625in"}

Figure 3.5: Schematics of the NED sensor.

Instrument summary

  ------------------------------ -------- --------- ----------- ---------- --
  Instrument                     Mass     Power     Volume      Data rate  
                                 \[kg\]   \[W\]     \[cm^3^\]   \[b/s\]    

  NED                            1        1         1000        100        

  NED+ \*)                       2        1.5       2000        150        

  \*) Version including γ-ray                                              
  detection facility                                                       
  ------------------------------ -------- --------- ----------- ---------- --

Table 3.12: NED resource requirements.

### Heliospheric instruments summary

  -------------- ------------- ----------------- -------------------- -------- -------------- --------- -------------
  **Name**       **Acronym**   **Measurement**   **Specifications**   **Mass   **Size (cm x   **Power   **Telemetry
                                                                      (kg)**   cm x cm)**     (W)**     (kb/s)**

  Solar Wind     SWA           Thermal ions and  0-30 keV/Q;\         6        20 x 20 x 20   5         5
  Plasma                       electrons         0-10 keV                                               
  Analyser                                                                                              

  Radio and      RPW           AC electric and   μV/m - V/m 0.1nT -   10       15 x 20 x 30   7.5       5
  Plasma Wave                  magnetic fields   μT                            (electronics             
  Analyser                                                                     box)                     

  Radio Sounding CRS           Wind density and  X-band\              (USO     5 x 5 x 5      3         0
                               velocity          Ka-band              200g)                             

  Magnetometer   MAG           DC magnetic field to 500 Hz            1        10 x 10 x 10   1         0.2

  Energetic      EPD           Solar and         Ions and Electrons   4        10 x 20 x 20   3         1.8
  Particle                     cosmic-ray        .01-10 MeV                                             
  Detector                     particles                                                                

  Dust Detector  DUD           Interplanetary    Mass (g):            1        10 x 10 x 10   1         0.05
                               dust particles    10^-16^-10^-6^                                         

  Neutral        NPD           Neutral hydrogen  0.6 -- 100 keV       1        10 x 10 x 20   2         0.3
  Particle                     and atoms                                                                
  Detector                                                                                              

  Neutron        NED           Solar neutrons    e\>1MeV              2        10 x 10 x 10   1         0.15
  detector                                                                                              
  -------------- ------------- ----------------- -------------------- -------- -------------- --------- -------------

Table 3.13: *In-situ* heliospheric instrumentation.

## REMOTE SENSING SOLAR INSTRUMENTS 

### General considerations

An integrated ensemble of high resolution remote sensing instruments is
proposed for the Solar Orbiter. The instruments are described in this
section with emphasis on the scientific requirements and the feasibility
of the instrument concepts. All the disc-observing instruments will
resolve the small-scale dynamical processes in the solar atmosphere. The
combination of instruments proposed here will provide a complete set of
multi-wavelength measurements required to understand the Sun's
magnetised atmosphere -- from below the photosphere up into the extended
corona.

Critical issues like heat input, mechanical structure, mass and power
requirements have been investigated for each instrument.

### Visible-Light Imager and Magnetograph (VIM)

Scientific drivers

The purpose of the VIM is to determine the magnetic field boundary
conditions for the MHD processes observed by other remote sensing
instruments. It will observe the morphology, dynamics, and strength of
the magnetic elements and flux tubes at the photospheric level with a
resolution that is consistent with the resolution of the EUV telescopes.
It will also provide images, Dopplergrams and magnetograms from the side
of the Sun, which is invisible from Earth.

VIM will be a vector magnetograph. Having vector capabilities is of
fundamental importance to understand the nature of photospheric fields.
Are they vertical unipolar fields? Do they harbour complex neutral lines
with horizontal sheared fields? Having vector capabilities is the only
way in which quantitative inferences of the magnetic field in the
transition region and corona can be made (from force-free or full 3D MHD
extrapolations). The importance of such data for the understanding of
coronal processes (like, e.g., the initiation of CMEs) has been amply
demonstrated by SOHO.

VIM will also produce line-of-sight velocity maps by observing two
points on either side of a spectral line. These maps can be used,
through local helioseismology techniques, to investigate subsurface
flows. The internal structure and dynamics of the near-polar regions of
the Sun is of paramount importance and perhaps *the* key to our
understanding of the solar cycle.

The key scientific goals of the VIM are

-   to resolve solar magnetism down to its fundamental length scale
    (\<100 km).

-   to provide measurements of the "magnetic carpet" which drives
    chromospheric and coronal activity.

-   to observe and accurately quantify for the first time the surface
    polar magnetic field of the Sun.

-   to unveil the small-scale photospheric dynamo,

-   to provide the first magnetograms and Dopplergrams of the far side
    of the Sun.

-   to measure rotation and flows near the Sun's poles using techniques
    of local area helio-seismology and thereby provide crucial
    constraints on solar dynamo theories.

Instrument concept

The optics of the VIM include

-   25 cm diameter High Resolution Telescope (HRT),

-   5 cm diameter Full Disc Telescope (FDT), and

-   Filtergraph Optics (FO).

The two telescopes, HRT and FDT, deliver an image of the Sun at a common
telescope focal plane. The FO produces a magnified image of the
telescope focus on the instrument\'s only detector. While the HRT
provides an image with a high resolution at its diffraction limit, the
FDT adjusts its focal length such as to fill the detector with the solar
disc. The filtergraph optics include a double Fabry-Perot interferometer
equipped with a narrow-band interference filter as a prefilter, tuned to
the Fe I 630.2 nm line. This filtergraph is easily adjusted to the
lineshifts which occur due to Sun-spacecraft relative velocities up to
40 km s^-1^. The three components are described in the following
sections.

High Resolution Telescope (HRT)

*- General description*

The HRT is a 250 mm diameter Gregorian reflector. It provides an image
with a spatial resolution as high as 75 km on the solar surface during
the near-Sun orbital phase. A polarisation modulator is foreseen in the
optical path behind the main mirror in the f/17 beam and before the
first oblique reflection.

The secondary mirror is actively mounted with three translation degrees
of freedom for telescope alignment control and pointing.

Figure 3.6 shows an annotated perspective view of the HRT and FO.

*- Thermal control*

The HRT will have to withstand the heat load during the near-Sun orbital
phases. The primary mirror will be made from SiC and passively cooled to
dispose off the approximately 170W of radiation which is absorbed in its
reflective layer. SiC enjoys better static and dynamic thermal stability
than other mirror blank materials.

The Gregorian design permits rejecting 96% of the incident radiation at
its primary focus with a field stop. The field stop is an inclined
mirror with a 2 mm central hole which transmits the observed field. The
reflected light is directed out of the spacecraft. The field stop will
also absorb about 170 W which will require passive cooling.

A broadband interference filter will be included after M2 to reduce the
amount of light in the filtergraph optics.

Full Disc Telescope (FDT)

It is impractical to combine the functions of full-disc and high
resolution viewing of the Sun into a single telescope. Therefore, a
separate telescope will provide a full disc image at the same position
as the secondary image of the HRT. The filtergraph is therefore shared
by both telescopes. A folding mirror which delivers the FDT image is
movable and can be inserted into the optical path of the HRT. The
mechanism will be such that the mirror will be removed from the optical
beam in the case of a mechanism failure, the high resolution function
will therefore have precedence.

![](media/image17.wmf){width="6.271527777777778in"
height="3.9194444444444443in"}

Figure 3.6: Perspective view of the optical layout of the high
resolution telescope and filtergraph.

![](media/image18.wmf){width="6.271527777777778in"
height="3.9194444444444443in"}

Figure 3.7: Perspective view of the optical layout of the full disc
telescope.

The FDT has its own pointing alignment control system, which consists of
a simple limb sensor and the same combination of tip-tilt mirror and
dichroic as the HRT. It shares the control electronics with the HRT.

The FDT consists of a refractor with a focal length which varies between
233 mm and 888 mm. The telescope provides a full-disc image of the Sun
of constant diameter on the detector during all orbital phases between
0.22 and 0.9 AU. The zoom lens consists of a classical afocal PNP
triplet followed by a fixed aperture stop and a camera lens which
provides a focus at a fixed position. All lenses will be singlets as
chromatic aberration is not a concern with the narrow-band filtergraph.
A mechanism which drives the middle lens of the triplet and the camera
lens provides the zoom function. The diameter of the entrance pupil
varies between 20 mm and 50 mm as the focal length increases, resulting
in a diffraction limited sampling of the detector at all distances from
the Sun. The optical design is diffraction limited for all settings of
the focal length.

An interference filter at the entrance aperture reflects most of the
incident sunlight so that the radiation load is not a problem for the
FDT.

Figure 3.7 shows the annotated optical layout of the FDT from a similar
point of view as the HRT in Figure. 3.6. Note that the layout is
\"horizontal\" (in the XZ plane) compared to the vertical (YZ plane)
layout of the HRT proper. A separate modulator is necessary for the FDT
in order to perform polarisation modulation within the centred optical
path before the first oblique reflection.

Filtergraph Optics (FO)

The FO consists of two singlet lenses which combine the functions of
collimator and camera with a 1:2 magnification. A system of two off-axis
ellipsoid and paraboloid mirrors provide a 1:1 pupil relay function and
produce an f/66 telecentric intermediate image plane.

Two 50 mm diameter Fabry-Perot etalons at the telecentric image plane
provide the required spectral resolution of typically 5 pm. One etalon
will provide the spectral resolution while the other blocks the
secondary transmission maxima of the first. A separate interference
filter with a 1 nm band blocks the secondary transmission maxima of the
combined etalons.

We consider using two solid state etalons with fixed resonator widths
which are mounted on a common rotating stage for fine tuning. About ±1°
of rotation angle is required for a shift of the passband of ±100 pm,
which is sufficient to cover both line width and Sun-spacecraft velocity
shifts.

The FO is mounted together with the FDT on a common optical bench.

Figure 3.8 shows a plane view of the filtergraph optics. The etalons are
not included but their position indicated.

![](media/image19.wmf){width="6.271527777777778in"
height="3.9194444444444443in"}

Figure 3.8: Arrangement of filtergraph optics.

Overall configuration

The combination of all optical subsystems is shown in Figure 3.9. The
HRT is the large cylinder on top of the Y and Z views. The FDT is to the
lower left of the Z view and near the bottom of the X view. The FO are
to the lower left of the Z view and near the top of the X view.

Polarisation Modulation Package (PMP)

![](media/image20.png){width="3.8354166666666667in"
height="1.9465277777777779in"}![](media/image21.png){width="1.6458333333333333in"
height="2.0284722222222222in"}

![](media/image22.png){width="3.9340277777777777in"
height="1.6291666666666667in"}

Figure 3.9: Plain views (top left: Y, top right: Z, bottom: X) of the
combined VIM. The HRT is arranged above the FDT and the FO, both of
which are mounted on a common optical bench. The optics fit a 300 mm x
400 mm x 1100 mm envelope.

The PMP will allow VIM to provide vector magnetograms of the region
being observed. The PMP will produce the modulation of the intensity at
the CCD as a function of the input polarisation state. These intensity
changes of the CCD measurements will be used to recover the Stokes
vector of the solar light.

VIM will use two PMPs, one for the HRT and one for the FDT. Each PMP
will be composed of a rotating retarder and followed by a fixed linear
polariser. The retarder follows the design of the Solar-B instrument,
which is based on the Advanced Stokes Polarimeter (ASP) used at
Sacramento Peak observatory. The retarder will be positioned at 8
different positions per half rotation, each separated by 22.5°. At each
of these positions the CCD will be read and the 8 readouts combined in
different ways to determine the final Stokes parameters (I,Q,U,V).

Optical alignment and image stabilisation

The optical alignment of the VIM is an important and critical issue. Due
to large temperature gradients during the mission, it is important to
have reliable and accurate in-flight alignment capabilities. To this
end, we propose a wavefront control system which is capable to detect
low order modes of wavefront deformations in the telescope. A wavefront
sensor measures the actual state of the optical alignment and generates
an appropriate error signal. A control system converts this error signal
into actuation signals which are used to drive the position of the
secondary mirror M2 and the tip-tilt mirror (M3).

Figure 3.10 shows a sketch of the wavefront sensor which is located in
the vicinity of the secondary focus. The sensor consists of a
collimating lens, a hexagonal lenslet array with seven - one central and
six peripheral - elements which cover the pupil of the telescope, and
lenses which reimage the lenslet array focal plane onto a matrix array
detector (CCD camera). The image scale of the pupil on the array is
chosen such that the illuminated areas of the lenslets are equal, taking
also into account the central obstruction. All lenslets are partially
illuminated.

CCD cameras

A variety of CCD detectors exist on the market that are fully adequate
for the VIM. The HRT and FDT telescopes have been designed for a 2k x 2k
CCD with 10 μm pixels. Such CCDs are already common on the market today.

Figure 3.10: Optical configuration of wavefront sensor.

Resources and requirements

Telemetry needs -- data compression.

VIM will have various high cadence data acquisition modes, each
optimised for the specific scientific requirements of local-area
helioseismology measurements and magnetic field dynamics observations.
For example, one can read only 1024 x 1024 pixels (smaller FOV, but
still large enough for magnetic field studies) or readout the CCD with 2
x 2 on-chip summing (preserving FOV with reduced spatial resolution, but
still adequate for local helioseismology). In this case, 1 full frame
(1024 x 1024, 12 b/pixel) at a cadence of 1 per minute produces a data
rate of 200 kb/s. By using efficient lossy compression algorithms with a
reduction factor of 10, the average date rate will be around 20 kb/s. In
this mode only one physical quantity would be stored (e.g. the LOS
magnetic field or the LOS velocity). Another option will be to read all
2048 x 2048 pixels, and send 4 physical magnitudes (for example all 4
Stokes parameters) at a cadence of 1 frame each 15 minutes using a
similar data rate. In the low rate observing mode (3 kb/s), one can send
2 physical magnitudes at full resolution at a cadence of 1 per hour.

Mass breakdown and power requirements

  ---------------------------------------------------------- ------------
                                                             **Mass (g)**

  **1. High Res. Telescope (incl. ISS, re-imaging, filter,   **18000**
  CCD)**                                                     

  \- 25 cm telescope (mirrors + structure)                   3000

  \- prefilter (@ primary hole)                              400

  \- rotating quarter wave plate                             800

  \- linear polariser                                        200

  \- folding mirror                                          400

  \- tip-tilt mirror + PZT (incl. electronics)               1700

  \- cube beam-splitter                                      200

  \- ISS lens                                                200

  \- ISS detector                                            800

  \- re-imaging system (4x200)                               800

  \- Fabry-Perot and Oven                                    3000

  \- shutter                                                 300

  \- CCD camera + electronics                                2700

  \- harness (25% of total above)                            3500

  **2. Full Disc Telescope**                                 **3000**

  \- prefilter                                               200

  \- focal lenses (4\*200)                                   800

  \- rotating quarter wave plate                             800

  \- linear polariser                                        200

  \- folding mirror                                          400

  \- harness (25% of total above)                            600

  **3. VIM enclosure**                                       **5000**

  **4. VIM TOTAL**                                           **26000**
  ---------------------------------------------------------- ------------

Table 3.14: VIM mass break-down.

  ----------------------------------------------- -----------------------
                                                  **Power (W)**

  CCD                                             5

  Image Stabilisation System                      5

  Fabry Perot & Oven                              10

  PMP (x2)                                        5

  **TOTAL**                                       **25**
  ----------------------------------------------- -----------------------

Table 3.15: VIM power requirements.

*INSTRUMENT SUMMARY*

+-------------------+--------------------------------------------------+
| **Telescope**     | HRT: Gregorian reflector\                        |
|                   | - 250 mm diameter\                               |
|                   | - effective focal length 8500 mm                 |
|                   |                                                  |
|                   | FDT: zoom lens\                                  |
|                   | - aperture: 20 -- 50 mm\                         |
|                   | - effective focal length: 233 -- 888 mm          |
+-------------------+--------------------------------------------------+
| **Spatial         | 0.28 arcsec pixel size (40 km on Sun at 0.2 AU)  |
| resolving         |                                                  |
| element**         |                                                  |
+-------------------+--------------------------------------------------+
| **Spectrometer**  | -   Fabry-Perot double etalon\                   |
|                   |     - tuning by rotation                         |
+-------------------+--------------------------------------------------+
| **Spectral        | FWHM: \<50 mÅ                                    |
| resolution**      |                                                  |
+-------------------+--------------------------------------------------+
| **Field of        | 80 x 80 Mm^2^; pointing to anywhere on Sun by    |
| View/Pointing**   | tilting and decentring secondary mirror.         |
+-------------------+--------------------------------------------------+
| **Detector**      | CCD:\                                            |
|                   | - 10 μm pixels\                                  |
|                   | - 2048 x 2048 array\                             |
|                   | - back thinned                                   |
+-------------------+--------------------------------------------------+
| **Stability**     | Active image stabilisation system and low order  |
|                   | wavefront correction.                            |
+-------------------+--------------------------------------------------+
| **Telemetry**     | 20 kb/s                                          |
+-------------------+--------------------------------------------------+
| **Mass**          | 26 kg                                            |
+-------------------+--------------------------------------------------+
| **Power**         | 25 W                                             |
+-------------------+--------------------------------------------------+
| **Size**          | Optics box: 110 cm x 40 cm x 30 cm               |
+-------------------+--------------------------------------------------+

Table 3.16: VIM instrument summary.

### EUV spectrometer (EUS)

Scientific drivers

Observations of the UV/EUV spectral range are critical for the
determination of plasma diagnostics in the solar atmosphere across the
broad temperature range from tens of thousands to several million K.
Analysis of the emission lines, mainly from trace elements in the Sun's
atmosphere, provides information on plasma density, temperature,
element/ion abundances, flow speeds and the structure and evolution of
atmospheric phenomena. Such information provides a foundation for
understanding the physics behind a large range of solar phenomena.

Current spacecraft instrumentation (SOHO) provides EUV spatial and
spectral resolving elements of order 2-3 arcsec and 0.01 nm,
respectively, and UV resolutions of 1 arcsec and 0.002 nm. There are no
plans for EUV or UV spectroscopic measurements from the NASA STEREO
mission (2004 launch), the NASA Solar Dynamics Observatory (\~2005
launch) and the Solar Probe (\~2007 launch). The only planned EUV
spectrometer for a future mission is the EIS instrument on Solar-B with
1 arcsec (750 km on the Sun) and 0.001 nm resolving elements. However,
this instrument is tuned to active region observations with almost no
transition region capability.

The European solar physics community has a well established expertise in
solar EUV/UV spectroscopy as illustrated by the successful CDS and SUMER
instruments on SOHO.

Instrument concept

Whilst recognising that an EUV spectrometer is an essential component of
the Solar Orbiter, we must be aware that it must be compact and
light-weight, must not be too telemetry 'thirsty' and must be able to
cope with the thermal and particle environment of such an orbit.

A Ritchey-Chretien telescope design, feeding a spectrometer, has been
chosen to minimise the size of the EUS instrument. It has the
disadvantage of one extra reflection over, for example, a single-mirror
off axis paraboloid system, but the wavelength range selected is geared
to bright solar EUV lines.

The instrument structure would be made of carbon fibre, with silicon
carbide (SiC) optical components. Multilayers will be considered if the
final wavelength selection requires it.

Basic optical layout

In the current design, the primary and secondary mirrors are separated
by 900 mm, with the focal plane 200 mm beyond the primary, making a
total telescope length of 1.1 m. The telescope diameter is 120 mm,
making an overall diameter of the telescope "tube" of about 150 mm. The
telescope effective focal length is 3.7 m.

The slit assembly lies at the focal plane and beyond this is the
spectrometer, with a toroidal grating forming a focus at a 2-D detector.
The grating ruling spacing is yet to be decided but we take 4800 l/mm as
a guide for the current calculations. For the wavelength bands selected
(see Table), and assuming a spectrometer magnification of 1.5, the
spectrometer will project 1.2 m beyond the focal plane with a 1.76 m
distance to the detector. The detector will lie 0.55 m off-axis,
adjacent to the primary mirror. Thus, without a spectrometer, the
instrument cross-section would be of order 15 x 15 cm^2^ but the
off-axis path of the spectrometer will ensure that the back-end of the
instrument is wider, making an overall boot-shaped instrument of maximum
dimensions of order 230 cm x 15 cm x 55 cm. The 55 cm dimension is
dependent on the choice of grating and the wavelength selection.

The secondary mirror presents a portion of the Sun at the slit and can
be rotated to allow rastered images (i.e. exposures interlaced with
mechanism movements to build up images simultaneously in selected
wavelengths). Only some 20% of the Sun will be visible to the secondary.
The instrument can be re-pointed to acquire particular targets.

Calculations of the spot-size at the focus from ray-tracing codes have
been performed for various angles of the secondary. This confirms that
the concept of rastering with the secondary is workable for the
scientific requirements over the planned range. We plan a 9 μ detector
pixel size and a field of view of 34 x 34 arcmin.

The heritage of this instrument comes from the SOHO/CDS, SOHO/SUMER and
Solar-B/EIS projects.

Resolution/detector

We size the instrument to an observing distance of 0.2 AU (perihelion).
A spatial resolving element of 0.1 arcsec represents 75 km on the Sun
from the Earth; the same resolution can be achieved at 0.5 arcsec from
0.2 AU. With regard to spectral resolution, our basic aim is to return
sufficient numbers of well separated spectral lines to allow thorough
spectroscopic studies.

A detector array of 4k x 4k 9 μ pixels is baselined. Thus, the EUS has a
spectral range of 4 nm at 0.001 nm/pixel. The same array will give a
spatial extent (vertical distance on the detector = slit length) of 0.5
arcsec x 4096 = 2048 arcsec = 34 arcmin. The solar diameter at 0.2 AU is
170 arcmin, i.e. we have a slit length of 0.2 of the solar diameter.
Thus, a pointing mechanism is required. For a given pointing location,
rastered imaging will be made up from movement of the secondary mirror
in only one direction.

We baseline an APS (Active Pixel Sensor) detector system. The SOHO CCD
detectors receive many "cosmic ray" particle hits, occasionally have
"bleeding" of charge for the brightest intensities (e.g. EIT flares and
LASCO planets) and there must be some degradation due to protons causing
damage to the silicon lattice (this creates "traps" that can steal
charge which can be transferred to other parts of the image).

In an APS device, the signal charge is sensed by an amplifier within the
pixel (each pixel has its own). Therefore, it avoids the charge transfer
problem. Pixels will still be hit, but the damage will not propagate
because of the transfer process used for CCDs.

![](media/image23.jpeg){width="5.823611111111111in"
height="4.358333333333333in"}

Figure 3.11: Optical scheme of the EUS.

This charge "trap" problem has been experienced by Chandra, mainly
because the CCDs have a direct line of sight to space for protons and
the damage was done mainly by relatively low energy protons (\~100 keV)
of which there are many. In SOHO, there is a lot more shielding.
However, this effectively means that the degradation will be at a slower
rate, not that there will be no degradation. For a long duration mission
(7 years) heading into a more intense particle environment, we have to
consider this seriously. We expect a more intense particle environment
because the Solar Orbiter is more likely to encounter accelerated
protons from solar events.

Intense sources display bleeding on CCD devices. This is not the case
for the APS detector system. This alone could be a tremendous advantage.

The APS detectors are still silicon chip devices, it is just that they
incorporate individual pixel amplification and charge extraction. The
EUV sensitivity will be provided in the same way as with CCDs, in this
case with back-thinned devices. There is no mass overhead in using such
a system, relative to the CCD approach.

Resources and requirements

Telemetry

For EUS we baseline 17 kb/s telemetry. The instrument detector has 4096
x 4096 pixels. It is clear that data from individual exposures will have
to be selected carefully, as given by pre-defined observing sequences --
i.e. only selected fractions of the full range would be returned in any
exposure. In addition, a number of compression schemes should be
available with a common ("lossless") compression of order 3-5 but other
schemes up to a factor of 10 would be available.

In order to cope with the limited telemetry resources, a combination of
novel compression schemes, large dedicated instrument memory and careful
emission line selection will be applied. In the light of the successful
operation of the CDS instrument on SOHO -- where a similar selection and
compression consideration was applied - and the new opportunities given
by the Orbiter's location, it is anticipated that this instrument can
make significant advances with a telemetry rate of order 17 kb/s.

Stability/pointing

Given our plan to achieve 0.5 arcsec resolution elements, we must choose
one of the following options:

-   do not include an image stabilisation system, assuming that the
    variations of the spacecraft stability occur on timescales much less
    than the exposure time of the spectrometer and thus any corrections
    could be done on the ground;

-   include an image stabilisation system, possibly making use of VIM
    limb sensor error signals to drive adjustments to the secondary
    mirror.

The EUS instrument would require an independent pointing system to
enable independent observation of any solar location. This would be done
using actuators at the legs of the instrument.

Thermal environment

At 1 AU the average solar intensity is 1.371 W/m^2^. During the nominal
phase, in each 149 day period, the spacecraft will encounter a range
from 2142 W/m^2^ (0.8 AU) to 34275 W/m^2^ (0.2 AU -- 25 times the value
at 1 AU). This presents a severe thermal challenge which we tackle in a
number of ways.

The SiC primary mirror "sees" the full Sun. SiC optical components can
run hotter than traditional components, but the primary receives 310.6 W
at 0.2 AU. Gold-coating is used to reduce the absorption (to 0.2). We
run the primary mirror at 70^o^C, and require a dedicated radiator of
0.11 m^2^ (33 cm x 33 cm).

Some 248.5 W are reflected from the primary. The secondary is sized such
that it catches only 20% of this, i.e. it only sees 20% of the solar
image, and much of the image is reflected back out of the aperture.
Thus, 49.7 W is received by the secondary mirror which will also be gold
coated. Thus it absorbs 9.9 W which will require dissipation via a small
radiator of area about 11 cm x 11 cm. We run the secondary at 83^o^C.

The reflected component, now 39.8 W continues toward the spectrometer
aperture in the primary. This beam can be reduced considerably by stops;
only about 1/100 of this incoming beam is required for each exposure.

Finally, the APS detector system should run cold and this would require
a cold finger to a further radiator. For the current design we assume
passive radiators to space. With regard to the remaining structure, the
design would include thermal blankets and the front panel could include
a solar shield. Heaters will be required to maintain the correct
temperature gradients throughout the orbit.

Instrument pointing and the spacecraft shield

The EUS can view a 34 x 34 arcmin area of the Sun, which has an apparent
diameter of about 170 arcmin at 0.2 AU. Thus an independent pointing
mechanism is required which needs to cover the range of about **±**3°.
The instrument has an aperture of 120 mm. This must match a viewing port
in the spacecraft +X-plate shield. To minimise the viewing port size,
the instrument will be mounted on six legs (like the SOHO/CDS
instrument) with the actuators at the back and the pivot near the front
of the instrument. Thus, we envisage a port size of approximately 130
mm. The front shield of the EUS instrument will act as a secondary to
the spacecraft shield when the aperture edge and viewing port edge are
not aligned.

Mass breakdown

A rough breakdown of the estimated mass is given in Table 3.17. We take
the view that this is a new generation instrument made of light-weight
materials which really ought to be significantly lighter than, for
example, the equivalent (100 kg) instruments on SOHO.

  -------------------------------------------------- --------------------
  **Component**                                      **Mass (kg)**

  Primary mirror                                     0.5

  Mirror support                                     0.3

  Secondary mirror                                   0.1

  Mirror scan mechanism                              0.6

  Slit assembly                                      0.3

  Grating assembly                                   0.6

  OPS                                                1.5

  Detector                                           1

  Detector electronics.                              1.5

  Baffles                                            0.5

  Structure                                          5.4

  Thermal subsystem                                  3.5

  Harness                                            1.2

  Main electronics incl. PSU                         5

  GRAND TOTAL                                        22
  -------------------------------------------------- --------------------

Table 3.17: Mass breakdown of the EUS.

*INSTRUMENT SUMMARY*

+---------------------+------------------------------------------------+
| **Telescope**       | -   Ritchey-Chretien type:\                    |
|                     |     - 120 mm diameter\                         |
|                     |     - 900 mm primary to secondary\             |
|                     |     - 1100 mm secondary to focus\              |
|                     |     - effective focal length 3.7 m\            |
|                     |     - gold-coated SiC mirrors                  |
+---------------------+------------------------------------------------+
| **Spectrometer**    | Normal Incidence:\                             |
|                     | - 4800 l/mm toroidal grating\                  |
|                     | - slit/grating distance 1200 mm\               |
|                     | - grating/detector distance 1760mm             |
+---------------------+------------------------------------------------+
| **Spatial resolving | 0.5 arcsec (75 km on Sun at 0.2 AU)            |
| element**           |                                                |
+---------------------+------------------------------------------------+
| **Spectral          | 0.001 nm/pixel (5 km/s)                        |
| resolving element** |                                                |
+---------------------+------------------------------------------------+
| **Raster            | Through motion of the secondary                |
| Mechanism**         |                                                |
+---------------------+------------------------------------------------+
| **Detector**        | Active Pixel Sensor (APS):\                    |
|                     | - 9 μ pixels\                                  |
|                     | - 4096 x 4096 array\                           |
|                     | - back thinned                                 |
+---------------------+------------------------------------------------+
| **Preliminary       | \- First Order: 58 --62 nm\                    |
| Wavelength          | - Second Order: 29 --31 nm\                    |
| Selection**         | (e.g. Fe XII 291, Si X 293, He I 584, Si IX    |
|                     | 296, O III 599, He II 304, Mg X 610, Fe XIX    |
|                     | 592).                                          |
+---------------------+------------------------------------------------+
| **Slit              | 34 arcmin length; width \<9 μ                  |
| Length/Width**      |                                                |
+---------------------+------------------------------------------------+
| **Field of          | Raster over 34 arcmin; pointing to anywhere on |
| View/Pointing**     | Sun with independent pointing mechanism.       |
+---------------------+------------------------------------------------+
| **Telemetry**       | 17 kb/s                                        |
+---------------------+------------------------------------------------+
| **Mass**            | 22 kg                                          |
+---------------------+------------------------------------------------+
| **Power**           | 25 W                                           |
+---------------------+------------------------------------------------+
| **Size**            | Instrument : 230 cm x 15 cm x 55 cm            |
|                     |                                                |
|                     | Electronics box: 20 cm x 20 cm x 20 cm         |
+---------------------+------------------------------------------------+
| **Thermal**         | Mirror operating temperatures 70 and 83°C,     |
|                     | both with radiators. APS detector cooled with  |
|                     | cold finger and radiator. Possible shield at   |
|                     | front. Standard thermal blankets with heaters  |
|                     | for thermal control.                           |
+---------------------+------------------------------------------------+
| **Operation**       | On board deferred command store in which can   |
|                     | be loaded time-tagged observation sequences    |
|                     | designed on the ground.                        |
+---------------------+------------------------------------------------+
| **Stability**       | Active image stabilisation system at secondary |
|                     | may be considered if jitter is significant     |
|                     | over common exposure times.                    |
+---------------------+------------------------------------------------+

Table 3.18: EUS instrument summary.

### EUV imager (EXI)

Scientific drivers

Observations from Yohkoh, SOHO and TRACE in extreme ultraviolet and
X-ray wavelengths have revealed a truly complex, highly dynamic solar
atmosphere with magnetic loops confining plasmas at widely varying
temperatures. The TRACE EUV observations in particular illustrate the
existence of fine-scale structures in coronal loops, and reveal
continuous dynamic activity in particular at the smallest scales. In the
quiet Sun, various \"events\" of different sizes (e.g., bright points,
explosive events, jets, blinkers) all provide evidence for small-scale
heating, probably related to magnetic reconnection. The observed
distribution functions have self-similarity properties which point at
sub-resolution processes. The results from Yohkoh, SOHO, and TRACE led
to new questions concerning the real basic dimensions of coronal
structures, the role played by nanoflares in the heating of the quiet
solar corona, and the structuring of the corona above the poles.

Dramatic new results can be obtained by increasing the spatial
resolution by one order of magnitude on one hand and from connecting
chromospheric/coronal phenomena with close *in-situ* measurements on the
other hand.

This is where an EUV Imager (EUI) on the Solar Orbiter can fully exploit
the unique capabilities of the mission. EUI will provide EUV images with
an angular resolution increased by one order of magnitude in order to
reveal the fine-scale structure of coronal features. The closeness of
the Sun makes this task easier. EUI will also provide full disc EUV
images of the Sun in order to reveal the global structure of
inaccessible regions such as the \"far side\" of the Sun and the polar
regions.

The European solar physics community has a well established expertise in
solar EUV/UV imaging as illustrated by the successful EIT instrument on
SOHO. Some studies concerning the thermal (radiative) load on the optics
have been shared with European teams involved in the remote sensing
package of the Solar Probe mission.

Instrument concept

A single telescope, providing both high spatial resolution and a full
disc field-of-view, faces two difficulties:

1.  limited number of pixels on current detectors and telemetry
    constraints;

2.  thermal (radiative) load at perihelion (25 solar constants on the
    first optical element if the field-of-view is not reduced).

These considerations led to a separation of the EUI into two
instruments: a High Resolution Imager (HRI) and a Full Sun Imager (FSI).
In this way, one can find specific solutions for the thermal load and
have realistic detectors.

High Resolution Imager (HRI)

The design driver is the need to get rid of the unwanted solar radiation
which goes along the direct light path. A first step consists in
reducing the entrance aperture of the instrument to about 2 cm and to
fully use the available length of the payload module. In doing so, one
can cluster a few telescopes together, each one dedicated to one
wavelength, which eliminates any moving part, takes full use of the
aperture and allows for *simultaneous* observations in different
wavelengths.

Essentially one reduces the FOV through the use of an entrance baffle
and one sets the FOV to 70 000 km (or 500 arcsec when at perihelion) for
a 2 x 2 K detector. This sets the equivalent pixel size to 35 km on the
Sun (or 0.25 arcsec, equivalent to 0.05 arcsec from Earth). The pixel
size of the detector is baselined at 9 μ, so the equivalent focal length
is 7.2 m.

We baseline a set of three telescopes working at 13.3, 17.4, and 30.4
nm. These three wavelengths cover a very wide range of temperatures
(from 5 x 10^4^ K up to 1.6 x 10^7^ K) and targets (from quiet Sun to
flares). Note that the 13.3 nm band includes a very hot line (Fe XXIII),
visible only during flares and that the 17.4 nm band only selects the
17.4 nm Fe X line, excluding the cooler Fe IX line at 17.1 nm.

The basic feature of each telescope relies upon the use of a long baffle
to reduce the FOV and the stray light, an entrance metallic filter and
an off-axis Gregory telescope with multilayers fitted to each
wavelength. The Gregory concept allows to put stops at the focal plane
and at the image of the entrance aperture, which leads to a significant
reduction of straylight. Figure 3.12 shows the optical scheme of the
HRI.

A carbon-carbon composite structure is planned.

We baseline 2k x 2k pixel APS detectors (see EUS section in 3.3.3).
Diamond or GaN/AlGaN new generation detectors are interesting
alternatives. They are robust, solar-blind (allowing a better
instrumental sensitivity) and hard-rad.

For the 17.4 nm and 30.4 nm channels, the small aperture will give a few
counts per pixel per second. One can cope with this low number by
increasing the exposure time and/or on-board binning. The count number
for the 13.3 nm channel can be much higher but only on very active
regions and flares.

The electronics (including the DPU) is common to the three detectors and
to the FSI telescope. Its volume should fit within the overall volume
given in the Table 3.22.

Full Sun Imager (FSI)

The Full Sun Imager needs a wide field (about 5°) when at perihelion in
order to cover the full lower solar corona with about 50 % margin in
each direction. In order to solve the issue of thermal (radiative) load
on the entrance filter of the FSI, different schemes have been
investigated. One possibility is to move the entrance filter as far back
as possible (1800 mm) and as close as possible to the M1 mirror. Then
the mirror combination (all off-axis) allows for an image quality of the
order of the pixel size (9 μ). The entrance filter is crossed twice by
the light. This system can work e.g. in the 17.4 nm line with an
appropriate multilayer (Mo/Si) coating. Another solution exists if one
works above 120 nm (e.g. in the Lα line at 121.6 nm). The dimensions of
such an instrument are of the order 2000 mm x 250 mm x 400 mm, i.e.
shorter but wider than the HRI.

Figure 3.13 shows the optical scheme of the FSI.

![](media/image24.jpeg){width="6.0256944444444445in"
height="2.2944444444444443in"}

Figure 3.12: HRI optical scheme with its basic components (bottom) : B:
baffle, M1: primary mirror, M2: secondary mirror, F: entrance filter, D:
detector. The figure also shows two possible implementations of the
three telescopes (top and bottom; boxes on the right-hand side).

![](media/image25.jpeg){width="6.019444444444445in"
height="2.420138888888889in"}

Figure 3.13: FSI optical scheme with the same components as in Fig.
3.12.

RESOURCES AND REQUIREMENTS

HRI telemetry

With all three telescopes working together at a 10 seconds cadence, we
arrive at a *peak* telemetry rate of 720 kb/s after data compression. An
average rate of less than 20 kb/s can be obtained by a combination of
further compression/selection schemes (e.g. on-board 2x2 binning, or
cascade time differentiation) and by adopting appropriate observing
strategies (e.g. interleaved high/low time resolution sequences,
observations in only one channel, etc.). We assume that we can compress
the data by a factor 20. Actually, with a Rosetta-type compression
scheme, tests performed on solar images indicate that compression
factors as high as 50 can be achieved.

HRI mass budget

The overall HRI mass budget is given in the Table 3.19.

  -------------------------------------------------------- --------------
  **Component**                                            **Mass (kg)**

  Structure (including the baffle) :                       6.0

  Optics (including mounts) :                              0.2

  Detector (including electronics and shielding) :         0.5

  Miscellaneous (harness, ..)                              0.1

  Margin                                                   0.7

                                                           

  Total per telescope                                      7.5

  Total for 3 telescopes                                   22.5

  Pointing mechanism                                       2.0

  Electronics shared between the 4 telescopes              1.0

                                                           

  Grand total for the HRI                                  25.5
  -------------------------------------------------------- --------------

> Table 3.19: Mass breakdown of the HRI of EUI

HRI power

The power requirements have been considered for the detector,
electronics and the pointing actuator (common to the three telescopes).
For a given telescope, we anticipate a constant power requirement of
about 4 W for the cycling (power on, integration, dump to memory). With
reasonable margins and the required power for pointing, the total is 5
W. If the three telescopes are working simultaneously, they will consume
about 15 W for basic functions (above) to which one has to add about 2 W
for data compression and dumping by the common DPU. Non-simultaneous
(sequential) observations of course lead to a smaller power consumption.

HRI stability/pointing

Given our plan to achieve 0.25 arcsec resolution elements and the fact
that exposure times are of the order of 10 seconds, an image
stabilisation system should be included, possibly making use of the VIM
limb sensor error signals to drive adjustments to the secondary mirror.

The three telescope cluster has its own pointing mechanism since the
limitation of the FOV prevents any independent internal pointing. The
pointing is performed by a joint displacement of the three telescopes.
The spacecraft has an offset capability of about 10°, much larger than
that required (2.5° at perihelion), and thus in principle it would be
possible to point to a solar target using the whole spacecraft.

FSI requirements

An average rate of 0.5 kb/s is sufficient to transmit one compressed
full Sun image every 4800 s (about every hour and 20 minutes).

The FSI mass is about 10.5 kg (structure: 6.5 kg; optics and mountings:
3 kg; detector and associated electronics: 1 kg).

Power is only required for the detector (3 W) since the electronics is
common with HRI.

The FSI also relies on the pointing stability of the payload module but
with constraints much less stringent than for the HRI.

The HRI and FSI characteristics are summarised in the Tables 3.20 and
3.21, respectively..

+----------------+-----------------------------------------------------+
| 3 Telescopes   | Off-axis Gregory; 20 mm diam.                       |
+----------------+-----------------------------------------------------+
| Spatial        | 0.25 arcsec pixel size (35 km on Sun at perihelion) |
| resolving      |                                                     |
| element        |                                                     |
+----------------+-----------------------------------------------------+
| Pointing       | Common pointing mechanism for the three components  |
+----------------+-----------------------------------------------------+
| Detector       | Baseline 9μ, 2000 x 2000 pixels                     |
+----------------+-----------------------------------------------------+
| Field of View  | 70x70 Mm^2^ at perihelion                           |
+----------------+-----------------------------------------------------+
| Peak Telemetry | 720 kb/s (with compression 20)                      |
|                |                                                     |
| Average        | 19.5 kb/s                                           |
| Telemetry      |                                                     |
+----------------+-----------------------------------------------------+
| Mass           | 25.5 kg                                             |
+----------------+-----------------------------------------------------+
| Power          | 17 W (essentially detectors; shared electronics and |
|                | DPU)                                                |
+----------------+-----------------------------------------------------+
| Size           | Instrument = 2500 mm x 150 mm x 240 mm              |
+----------------+-----------------------------------------------------+
| Thermal        | Operating temperature 20°C with passive cooling;    |
|                | 140 cm long baffle at the entrance of each          |
|                | telescope                                           |
+----------------+-----------------------------------------------------+
| Stability      | Active image stabilisation at secondary.            |
+----------------+-----------------------------------------------------+

**Table 3.20:** Characteristics of the HRI of EUI

+----------------+-----------------------------------------------------+
| Telescope      | Off-axis Gregory; 20 mm diam.                       |
+----------------+-----------------------------------------------------+
| Spatial        | 9 arcsec (1300 km on Sun at perihelion)             |
| resolving      |                                                     |
| element        |                                                     |
+----------------+-----------------------------------------------------+
| Pointing       | Pointed at Sun centre                               |
+----------------+-----------------------------------------------------+
| Detector       | Baseline - 9 μ, 2000 x 2000 detector                |
+----------------+-----------------------------------------------------+
| Field of View  | 4 x 4 solar radii at perihelion                     |
+----------------+-----------------------------------------------------+
| Peak Telemetry | 240 kb/s (with compression 20)                      |
|                |                                                     |
| Average        | 0.5 kb/s                                            |
| Telemetry      |                                                     |
+----------------+-----------------------------------------------------+
| Mass           | 10.5 kg                                             |
+----------------+-----------------------------------------------------+
| Power          | 3 W                                                 |
+----------------+-----------------------------------------------------+
| Size           | Instrument = 2000 mm x 250 mm x 400 mm              |
+----------------+-----------------------------------------------------+
| Thermal        | Operating temperature 20°C with passive cooling;    |
|                | 600 cm long baffle at the entrance of each          |
|                | telescope                                           |
+----------------+-----------------------------------------------------+

Table 3.21: Characteristics of the FSI of EUI

*INSTRUMENT SUMMARY*

+----------------+-----------------------------------------------------+
| HRI and FSI    | Off-axis Gregory Telescopes; 20 mm diam.            |
+----------------+-----------------------------------------------------+
| Spatial        | HRI : 0.25 arcsec pixel size (35 km on Sun at       |
| resolving      | perihelion)                                         |
| element        |                                                     |
|                | FSI : 9 arcsec (1300 km on Sun at perihelion)       |
+----------------+-----------------------------------------------------+
| Pointing       | HRI : pointing mechanism to point anywhere on the   |
|                | Sun                                                 |
|                |                                                     |
|                | FSI : pointed at Sun centre                         |
+----------------+-----------------------------------------------------+
| Detector       | Baseline - 9 μ, 2000 x 2000 pixels                  |
+----------------+-----------------------------------------------------+
| Field of View  | HRI : 70x70 Mm^2^ at perihelion                     |
|                |                                                     |
|                | FSI : 4 x 4 solar radii at perihelion               |
+----------------+-----------------------------------------------------+
| Average        | 20 kb/s                                             |
| Telemetry      |                                                     |
+----------------+-----------------------------------------------------+
| Mass           | 36 kg                                               |
+----------------+-----------------------------------------------------+
| Power          | 20 W                                                |
+----------------+-----------------------------------------------------+
| Size           | Instrument = within 2500 mm x 400 mm x 400 mm       |
+----------------+-----------------------------------------------------+
| Thermal        | Operating temperature 20°C with passive cooling     |
+----------------+-----------------------------------------------------+

Table 3.22: EUI instrument summary.

### Ultraviolet and Visible-light Coronagraph (UVC) 

Scientific drivers

Co-rotation, during the helio-synchronous phases of the orbit, will
freeze for many days coronal structures in the plane of the sky. This
condition will be most favourable to investigate the evolution of the
magnetic configuration of streamers in order to test the hypothesis of
magnetic reconnection as one of the main processes leading to the
formation of the slow solar wind.

The out-of-ecliptic vantage point will allow for the first time a unique
view of the plasma distribution and solar wind expansion in the coronal
low-latitude/equatorial belt. Therefore it will be possible to measure
the longitudinal extent of coronal streamers and coronal mass ejections.
These parameters, that at present are unknown, are essential to
determining the magnetic flux carried by plasmoids and coronal mass
ejections in the heliosphere.

UVC will allow the first determination of the absolute abundance (i.e.
relative to hydrogen) of helium in corona. Helium, the second largest
contributor to the density of coronal plasma, is important for the
dynamics of solar wind, and it may act as a regulator to maintain a
nearly constant solar wind mass flux.

UVC will also determine the differential outflow speed of the major
components (H, He) of the solar wind and discriminate the mechanisms of
solar wind acceleration.

The European solar community has a well established expertise in visible
light and UV coronagraphy as illustrated by the successful LASCO and
UVCS instruments on SOHO.

Instrument concept

UVC is an externally o**c**culted telescope designed for narrow-band
imaging of the EUV corona in the He II 30.4 nm and H I 121.6 nm lines,
and for broad-band polarisation imaging of the visible K-corona, in an
annular field of view between 1.2 and 3.5 solar radii, when the Solar
Orbiter perihelion is 0.21 AU. When it moves to 0.3 AU, at the end of
the nominal mission, the UVC field-of-view is within 1.8 and 5.3 solar
radii.

The telescope optical configuration is an off-axis Gregorian. The EUV Lα
lines are separated with multi-layer mirror coatings and EUV
transmission filters. These mirrors with coating optimised for 30.4 nm
still have a good reflectivity at 1216 nm and in the visible. The
visible light channel includes an achromatic polarimeter, based on
electro-optically modulated liquid crystals.

Optical design

*-Telescope*

The external occulter ensures both thermal protection for the optics and
better stray-light rejection. A Gregorian telescope design has been
chosen because it gives real images of the external occulter and the
edges of the telescope primary mirror. Field and Lyot stops will be used
to reduce stray light and diffracted radiation from the telescope
primary mirror. In addition, use of an annular secondary mirror and
light trap captures and removes diffracted light from the external
occulting disc and its supports. A light trap behind the primary mirror
ensures that only the light reflected by the mirror enters the filters
and detector assembly structure. The direct light from the solar disc is
mostly dumped into space, though a portion of it may be used for
radiometry and / or coarse imaging of the solar disc in the three
wavelength bandpasses.

*-UV and Visible-Light Channels*

The instrument structure is made of carbon fibre, with Zerodur optical
components. Zerodur is used for the mirrors\' substrate because of its
extremely small thermal expansion coefficient. Optics made with this
material is also easy to polish into non-spherical figures with
sub-angstrom rms surface roughness.

The basic specifications and drivers of the instrument are summarised in
Table 3.24.

Structure

The instrument structure is of carbon fibre/cyanate or carbon-carbon
composite, designed for zero coefficient of thermal expansion in the
axial direction. It is in the form of a rigid rectangular baseplate to
which are bonded appropriately placed brackets (also of composite) for
the support of the optical and detector elements. The focus and
orientation of each element is set and maintained by a precision spacer
located between the element and its support bracket. The optical system
is surrounded by a lightweight rectangular composite enclosure and
baffle structure that attaches to and further strengthens the baseplate.
Removable covers on the top of the enclosure allow for installation and
servicing of the optical and detector elements. An external door
mechanism shields the telescope aperture when the instrument is not
observing.

![](media/image26.jpeg){width="5.923611111111111in"
height="3.036111111111111in"}

Fig.3.14: Two views of the optical layout of UVC.

Thermal

The thermal control approach is based on shielding the instrument from
most direct solar radiation, rejecting light from the solar disc that
enters the annular entrance aperture, and an opto-mechanical design that
is insensitive to bulk temperature changes. The use of a carbon fibre
composite structural system allows the optical system to be almost
completely athermalised. Moreover, the use of a low secondary
magnification factor in the telescope (i.e. 3:1) makes the optical
system less sensitive to dimensional changes in comparison with more
traditional space instruments such as TRACE. Preliminary estimates
suggest that the allowable temperature range of the opto-mechanical
system may be as much as **±**50^o^ C. Local temperature control will be
needed for the liquid crystal polarisation module, which has a smaller
allowable temperature range. Thus we consider it prudent to include a
small power allocation for thermal control heaters (e.g. 1 W) pending
further analysis.

The occulter is the element of the coronagraph, which reaches the
highest temperature since it is directly exposed to the Sun,
particularly when the Solar Orbiter is near perihelion. The occulter has
an exiguous conductive link (represented by the thin rods of the
support) with the rest of the structure. The occulter and supporting
links are made of titanium and are arranged in a conical geometry such
that thermally induced strain is absorbed without de-centring the
occulter disc.

The temperature achieved by the occulter has been estimated by assuming
a purely radiative heat transfer and an external coating of the same
type as that utilised for the sunshield of the payload module. The
thermal flux incident on the occulter is 31 kW/m^2^ at 0.21 AU and 5.5
kW/m^2^ at 0.5 AU. The temperature difference experienced by the
occulter along the observation phase is at maximum 266.5 C. A small
reduction of the temperature will be caused by the transport of part of
the heat through the occulter supports.

Detectors

There are two detectors, optimised for the Visible region (450 -- 600
nm) and for the Extreme UV (304-121.6 nm). Both detectors have 15 μ
pixels and an array size of 4096 x 4096. The baseline visible detector
is an Active Pixel Sensor (APS). This detector system is based on CMOS
technology, which is attractive for high radiation dose environment. In
addition, the APS architecture has the potential to allow on-chip
differencing of the polarimetric signals, yielding substantial
improvements in signal-to-noise performance in comparison with
conventional CCD detectors.

The baseline EUV detector is a Charge Injection Device (CID) array. This
detection system is highly resistant to space radiation. It can easily
withstand more than 100 times the lethal dose of a typical CCD and
radiation-hardened versions can tolerate doses in excess of 1 MegaRad
(Si total dosage). CID detectors have randomly addressable pixels,
permitting the dynamically programmable read-out of individual pixels
and sub-arrays. This feature will be used in UVC for limiting the
readout area to the solar corona (excluding the occulted portion field)
and for defining \"Regions of Interest\" to be observed with higher time
resolution. EUV sensitivity will be provided by a Micro-Channel Plate
image intensifier coupled to the CID array with fibre-optics. This
approach results in a detector that is blind in the visible,
substantially improving the stray light rejection performance of the EUV
channel.

We note that both detector approaches are based to some extent on
anticipated developments in rapidly evolving technologies (e.g. array
sizes, APS noise performance). Heavily shielded CCD detectors are
suitable alternatives in either case.

Resources and requirements

Data rates and volume

The primary driver of the data rate is UVC efficiency in the three
channels. In the two EUV lines, the coronal signal is weaker and the
instrument efficiency lower than in the visible band. Therefore, longer
exposure times will be required for the EUV coronal observations. On the
basis of the count-rate for H I and He II Lyman-α emission estimated
from a coronal hole (i.e. worst case), we assume an average exposure
time of 600 sec. A single image (4096 x 4096 pixel) with 16 bit per
pixel takes about 270 Mb of memory. Assuming that only 50% of the image
is used (the rest is occulted disc, extreme corners of the square
matrix, etc.), and that a factor of 3 lossless compression can be
achieved, we expect a data rate of 0.07 Mb/s per detector. Since the
visible-light detector works simultaneously with the H I Lyman-α
(therefore, about 50% of the time), the data volume accumulated per
orbit is about 260 Gb. A factor 20 image compression is needed to reach
a data rate of about 5 kb/s. It is expected that this level of
compression can be achieved with an acceptable loss of information with
schemes such as the "Adaptive Discrete Cosine Transform\'\' (ADCT). The
ADCT technique is being used in the Large-Angle Spectroscopic
Coronagraph (LASCO) on SOHO.

Mass

The UVC mass breakdown is given in Table 3.23.

  --------------------------------------------- -------------------------
  **Component**                                 **Mass (kg)**

  Structure                                     5

  Optics                                        2

  Mechanisms                                    1

  Detectors (2)                                 2

  CPU & Interface                               2

  DC/DC converter                               1

  ADC                                           1

  Data compressor                               1

  Motor drive                                   2

  TOTAL instrument                              17
  --------------------------------------------- -------------------------

Table 3.23: Mass breakdown of the UVC.

+-----------------+----------------------------------------------------+
| **Telescope**   | **Off-axis Gregorian**                             |
+-----------------+----------------------------------------------------+
| Occultation     | External occulter: 60 mm diam.                     |
+-----------------+----------------------------------------------------+
| Optics          | Primary mirror: 33 mm diam.                        |
|                 |                                                    |
|                 | Secondary mirror: 94 mm diam.                      |
|                 |                                                    |
|                 | (multilayer coated mirrors)                        |
+-----------------+----------------------------------------------------+
| Eff. focal      | 490 mm                                             |
| length          |                                                    |
+-----------------+----------------------------------------------------+
| Field of View   | Annular, Sun-centred. Coverage:                    |
|                 |                                                    |
|                 | 1.2 -- 3.5 R at 0.21 AU,(perihelion at mission     |
|                 | start)                                             |
|                 |                                                    |
|                 | 1.7 - 5.0 R at 0.30 AU, (perihelion at mission     |
|                 | end)                                               |
+-----------------+----------------------------------------------------+
| Spatial         | UV (30.4 nm, 121.6 nm): 8 arcsec at 1.2 R (1200    |
| resolution      | km)                                                |
|                 |                                                    |
|                 | 40 arcsec at 3.5 R (6000 km)                       |
|                 |                                                    |
|                 | Visible (500 nm): 40 arcsec entire FOV (6000 km)   |
+-----------------+----------------------------------------------------+
| Stray-light     | \< 10^-8^ (visible light); \< 10-7 (30.4 nm, 121.6 |
| levels          | nm)                                                |
+-----------------+----------------------------------------------------+
| Wavelength.     | 1\) He II (30.4 ± 2) nm                            |
|                 |                                                    |
| bandpasses      | 2\) H I (121.6 ± 10) nm                            |
|                 |                                                    |
|                 | 3\) Visible (450 - 600) nm                         |
+-----------------+----------------------------------------------------+
| Detectors       | Visible-light: Active Pixel Sensor (APS)           |
|                 |                                                    |
|                 | EUV (30.4 nm., 121.6 nm): Charge Injection Device  |
|                 | (CID)                                              |
|                 |                                                    |
|                 | (EUV intensified); 4096 x 4096 array, 15 μm pixels |
+-----------------+----------------------------------------------------+
| Data rate       | 5 kb/s                                             |
+-----------------+----------------------------------------------------+
| Dimensions      | 80 cm x 30 cm x 30 cm                              |
+-----------------+----------------------------------------------------+
| Mass (kg)       | 17                                                 |
+-----------------+----------------------------------------------------+
| Power (W)       | 25 (operational mode, peak); 20 (stand-by mode)    |
+-----------------+----------------------------------------------------+

Table 3.24: UVC instrument summary.

### Radiometer (RAD)

Scientific Drivers

The mission profile of the Solar Orbiter is ideally suited for studies
to improve our understanding of the angular distribution of the radiance
from active regions of the photosphere. This can be done by observing
e.g. the influence of faculae on the True Solar Irradiance (TSI) from
different angles: from near the equator (from Earth-bound missions) and
from higher latitudes (from the Solar Orbiter). Similar studies can also
be done and have been done from one platform (e.g. SOHO) using the fact
that the Sun rotates, which allows us to view the same feature from
different angles at different times; the uncertainty, however, is
limited by the evolution of the active regions themselves which are
known to change on time scales of days at the level of up to 0.1%. From
our present understanding, an uncertainty of about 1 part in 10000 is
needed for such comparative studies. Thus, we have to ensure that the
RAD measurements can be projected back to other radiometers at 1 AU for
all operational modes and at all times with sufficient precision. This
appears to be feasible, as is shown below, but has to be studied in
detail and confirmed during a Phase A study.

Instrument concept

RAD is a new generation, room temperature radiometer with four channels,
which can be operated independently or in groups of two for internal
comparisons. Each channel has its shutter and an independent cover for
protection. RAD has electrically calibrated cavity receivers. Two
cavities are operated together with one closed and the other with cyclic
shutter phases of 50s open and 50s closed duration. Phase sensitive
detection at the fundamental shutter frequency is used to analyse the
signal. This operational mode has many advantages over the normal
"active cavity" operation in e.g. SOHO/VIRGO, which relies on reaching a
steady state before measurements can be taken.

The operation will be controlled by software in the instrument\'s
computer. There are essentially only two analogue circuits:

-   the amplifier of the thermistor bridge measuring the temperature
    difference along the thermal impedance and between two cavities (one
    measuring irradiance the other as thermal reference),

-   the highly stable voltage source for the electrical power fed to the
    cavity.

The bridge is powered by a sine wave at about 100Hz produced by software
and its output is A-D converted and synchronously demodulated by
software. The demodulated signal is used to control the heater power of
the shuttered cavity to compensate for the extra radiant power during
the open periods. The heater power is deduced from the highly stable
voltage source which is pulse-width modulated by the output of the
controller. This power signal is frequency analysed and the amplitudes
and phases of the shutter fundamental frequency and the first few
harmonics of it (up to about the 8^th^) are transmitted to ground. Only
the in-phase signal at the shutter fundamental is needed to calculate
the irradiance, the harmonics are for the in-flight characterisation of
the temporal behaviour of the detectors. Some of the correction factors,
needed for the final evaluation of the irradiance can be determined
and/or checked in flight.

Resources and requirements

The RAD will have a mass of 4 kg. The power consumption is less than 6.5
W.

The outer dimensions of the RAD instrument will be about 110 mm x 110 mm
x 220 mm. The 110 x 110mm face has to be pointed towards the sun; in
front there will be some (probably 2) radiation shields which will
extend the length of the box by 40 mm and the area of the front to
140x140mm.

For normal operation, i.e. when measuring with one radiometer, a
telemetry rate of 500 b/s is required. For backup operations (about once
a week) 280 b/s have to added. During check-out intervals a total of
3500 b/s will be needed.

  ---------------- ----------- ------------ --------------- ---------------
  **Instrument**   **Mass      **Power      **Volume        **Data rate
                   (kg)**      (W)**        (cm3)**         (b/s)**

  RAD              4           6.5          3000            500
  ---------------- ----------- ------------ --------------- ---------------

Table 3.25: RAD resource requirements.

### Other potential solar instruments 

A valid option to be considered in future studies of the Solar Orbiter
is the addition of two corona-related instruments:

-   A heliospheric imager might extend the view of the heliosphere to
    almost a full hemisphere, from about 3 solar radii (at 0.2 AU) to as
    much as the anti-solar direction. Thus, it would have almost the
    full Earth-Sun line within its field of view for major parts of the
    observation time, depending on the spacecraft attitude. This
    instrument would allow us for the first time to observe the
    propagation of transient disturbances towards the Earth from various
    distances and aspect angles, in particular from outside the ecliptic
    plane. Preliminary design studies have demonstrated that such an
    instrument can probably be built for less than 2 kg. A crucial
    requirement would be that the hemisphere to be watched must be kept
    free of Sun-illuminated spacecraft appendages.

-   An X-ray instrument dedicated to the high-temperature portion of the
    coronal emission would enhance and complement the Solar Orbiter
    payload. A novel configuration for an imaging spectroscopic
    instrument would operate in the wavelength range from 4 nm to 6 nm
    and detect emissions from plasma at temperatures between a few 10^6^
    K and a few 10^7^ K. Such an instrument working in extreme-grazing
    incidence will obtain very high spectral and spatial resolution. It
    would also be able to measure densities equal to or greater than
    10^9^ cm^-3^ (through density-sensitive lines in this
    high-temperature range).

### Solar remote-sensing instruments summary

  --------------- --------- ------------- ---------------- ------ --------- -------- ------------
  Name            Acronym   Measurement   Specifications   Mass   Size\     Power    Tele-metry
                                                           (kg)   (cm x cm  (W)      (kb/s)
                                                                  x cm)              

  Visible-light   VIM       High-res.     Fe 630 line      26     30 x 40 x 25       20
  Imager and                disc imaging                          120                
  Magneto-graph             and                                                      
                            polarimetry                                              

  EUV Imager and  EUS       Imaging and   EUV emission     22     30 x 15 x 25       17
  Spectrometer              diagnostics   lines                   140                
                            of TR and                                                
                            corona                                                   

  EUV Imager      EXI       Corona        He and Fe Ion    36     40 x 40 x 20       20
                            imaging       lines                   250                

  Ultraviolet and UVC       Imaging and   Coated mirror    17     20 x 20 x 25       5
  Visible Light             diagnostics   coronagraph             50                 
  Corona-graph              of the corona                                            

  Radiometer      RAD       Solar         Visible light    4      11 x 11 x 6.5      0.5
                            constant                              22                 
  --------------- --------- ------------- ---------------- ------ --------- -------- ------------

Table 3.26: Remote-sensing solar instruments.

# Mission profile

The Solar Orbiter scientific requirements provide the outline for a
mission profile, which defines the basic outline of the required
mission, in terms of orbital parameters, launch windows, payload mass
etc, and this profile becomes the driver for the spacecraft design.

The energy requirements for a mission such as the Solar Orbiter can be
derived by studying the simple problem of a Hohmann transfer orbit to
the required final orbit (perihelion as close as possible to the Sun and
high orbit inclination with respect to solar equator). To achieve such a
trajectory, however, requires a high Delta V and/or a long transfer
period even if electric propulsion is used. To overcome this, a strategy
based on low-thrust electrical propulsion and planetary gravity assists
is baselined.

The orbit is based on the use of a standard Soyouz-Fregat as launcher
from Baikonur. The mission profile calls for the interleaving of
Earth/Venus gravity assists and SEP (Solar Electric Propulsion) firings.
Stationary Plasma Thrusters (SPT) are baselined for costing purposes.
The ground operations will be provided by ESOC and will involve using
the 35-m ground station near Perth, in Australia. The mission profile is
composed of three phases:

The **cruise phase**, starting at spacecraft separation from the
launcher, and ending at the start of scientific operations (some science
may be performed during the cruise phase):

1.  the **nominal mission phase**, during which the main scientific
    mission is performed,

2.  the **extended mission phase**[,]{.underline} when further gravity
    assist manoeuvres will allow higher inclination observations.

During the **cruise phase** (0 - 1.86 years) there are SEP thrust
phases, with durations ranging from 6 to 105 days, and Venus swing-bys
driving the semi-major axis changes and the inclination increase.

During the **nominal mission** (1.86 - 4.74 years, i.e. duration: 2.88
years), the orbit is typically of order 150 days, taking the spacecraft
from about 0.2 to 0.9 AU. There are 2 Venus swing-by's in this 7-orbit
phase taking the orbital inclination to over 30°.

During the **extended mission** (4.74 - 7.02 years, i.e. duration: 2.28
years), there are two further Venus swing-by's over 6 orbits for a
continued increase of the orbital inclination.

The selected orbit assures the observation of the Sun with a
satisfactory scientific heliographic latitude, while the design of the
spacecraft is still thermally feasible. The celestial constellation of
the Sun, Venus and Earth leads to a launch window of 3 weeks in every 19
months.

The ecliptic projection of the Solar Orbiter trajectory as well as the
perihelion/aphelion radius and solar latitude with respect to mission
time are shown in Figures 4.1 to 4.4.

![](media/image27.wmf){width="5.7625in" height="5.582638888888889in"}

![](media/image28.png){width="5.820138888888889in"
height="3.392361111111111in"}Figure 4.1: Ecliptic projection of the
Solar Orbiter trajectory.

Figure 4.2: Perihelion distance as function of time.

![](media/image29.png){width="5.8in" height="3.615972222222222in"}Figure
4.3: Aphelion distance as function of time.

Figure 4.4: Spacecraft latitude with respect to the Sun\'s equator as
function of the flight time in days.

# System design

## 

## 

## 

## 

## Design requirements and drivers

The system design of the Solar Orbiter is driven by the mission profile
as well as by general requirements and cost objectives. The spacecraft
is designed to get as close to the Sun as the materials and engineering
will allow (0.21 AU). Solar radiation increases with decreasing distance
from the Sun, and is, of course, the most challenging aspect of the
present mission. The solar constant at a distance r (in Astronomical
Units \[AU\]) is given by C~s~(r) = C~so~ /r^2^, where C~so~ is the
solar constant measured at 1 AU, i.e. 1367 W/m^2^. Figure 5.1
illustrates the maximum heat load, which the Solar Orbiter will
encounter. Two other spacecraft, Cassini and the Mercury Orbiter, are
shown for comparison.

The orbit is as highly inclined to the Sun\'s equator as the launcher
and propulsion capability will allow. The power demand is higher during
the cruise phase, the greatest demand being for the electric propulsion.
The solar array sizing is set to provide adequate power at the furthest
distance from the Sun. Two sets of solar arrays are used. The larger
arrays are for the SEP system and these will be jettisoned after the
cruise phase, i.e. when the firings are completed. The remaining array,
for spacecraft/instrument power is a thermal "burden" to the system. It
is protected by angling the array away from the Sun during the
perihelion passages. The basic spacecraft size and shape is directly
linked to those of the instruments and the service module performance.

Figure 5.1: Solar Constant versus distance from Sun.

The major design driving features for the spacecraft were:

-   instruments requirements, mainly field of view, mass, pointing
    requirements, stability, operations and size;

-   launcher mass envelope and interfaces;

-   Earth communication, mainly accommodation of the high gain antenna
    (HGA), and HGA pointing for data transmission; and

-   use of existing hardware to minimise risk and cost.

## Main system design features

The basic spacecraft configuration is box-like - about 3.0 m long, 1.6 m
wide and 1.2 m deep. It has internally mounted instruments, a 2-axis
steerable HGA, two sets of one degree of freedom steerable solar arrays
and the electrical propulsion system.

The spacecraft must provide accommodation for all of the subsystems and
ensure compatibility between them throughout the mission. This demands a
detailed consideration for all phases of the mission on each subsystem.
The technical difficulties of this mission encountered during the study
were many and arose in many areas: the major ones are identified below.

### Requirements and constraints

The major drivers for the overall configuration of the Solar Orbiter can
be summarised as follows:

-   **Payload:** Instrument sensors Sun-pointing (except for the
    magnetometer and dust detector) with no obstacles in the fields of
    view. Detector radiators, assumed to be mounted on the sensor body,
    open to cold space. Instruments provided with stable mounting but
    accessible.

-   **Thermal:** Cater for solar radiation input with shielding,
    radiating surfaces and active orientation of components.

-   **Propulsion:** SEP thruster line of force through the spacecraft
    centre of gravity during the whole cruise phase.

-   **Launcher:** Accommodation of the spacecraft in a stowed position
    under the SOYUZ type S fairing and mechanical interface with FREGAT.

-   **Power:** Accommodation of solar array area in two sets, i.e. the
    cruise phase solar array (for the SEP) and the "standard" solar
    array.

-   **Communications:** Accommodation of the large HGA.

-   **Others:** Equipment mounting area and particular equipment
    accommodations.

### Spacecraft baseline design

Figures 5.2 to 5.4 show the Solar Orbiter spacecraft overall
configuration. The spacecraft body has a cross-section of 1600 mm (Y) x
1200 mm (Z) and its length is 3000 mm (X). The +X side, which faces the
sun (±30°), is covered by a thermal shield shadowing the spacecraft
body. In order to minimise the energy input from the Sun the
cross-section of the spacecraft has been minimised (in the Y, Z plane),
i.e. the +X panel area is minimised.

![](media/image32.png){width="3.1118055555555557in"
height="3.316666666666667in"}

![](media/image33.png){width="3.279166666666667in"
height="3.157638888888889in"}

Figure 5.2a: Launch. Figure 5.2b: Cruise.

![](media/image34.png){width="3.0854166666666667in"
height="3.0569444444444445in"}

Figure 5.2c: Observation. Figure 5.2d: Downlink.

Figure 5.2a-d: Solar Orbiter configurations during the mission phases.

![](media/image36.wmf){width="4.790972222222222in"
height="6.1097222222222225in"}

Figure 5.3: Solar Orbiter main dimensions. Launch configuration.

![](media/image37.wmf){width="5.8284722222222225in"
height="5.247222222222222in"}

Figure 5.4: Solar Orbiter main dimensions. In-orbit configuration.

The spacecraft is of modular design, with a Service Module (SVM) and
Payload Module (PLM). The +/-Y sides of the PLM accommodate the cruise
solar array (2 wings, 3 panels per wing) and the top shield radiators.
The optical instruments are right beneath the top shield, pointed in the
+X direction. They are isostatically attached to the central cylinder
and may be cooled by radiators to cold space in the +\\- Z sides. The
instrument electronic boxes are mainly mounted on the bottom panel of
the PLM.

The SVM accommodates the orbit solar panels and the SEP/equipment
radiators on the +/-Y panels. The HGA and the thrusters are attached on
the +Z panel. The propellant tank is at the Centre Of Gravity (COG) of
the S/C, inside the central SVM cylinder mounted on a dedicated ring.
This design can account for possible changes on the COG height as the
spacecraft design matures. The spacecraft equipment boxes are mounted
internally on the SVM panels.

In order to cope with different attitudes and distances from the Sun,
both solar arrays incorporate a 1 Degree of Freedom (DOF) driving
mechanism. The cruise solar array can be jettisoned. The HGA mast (2 m
long) is mounted on a 2 DOF mechanism and ensures coverage throughout
the mission. Four configurations are considered for the 'active'
externally mounted components of the Solar Orbiter (see Figs. 5.2):

-   **Launch:** The solar arrays, HGA and magnetometer are stowed.

-   **Cruise Phase:** The cruise solar array is deployed (19 m tip to
    tip). The orbit solar array is deployed (at 75° with respect to the
    Y-axis). The HGA is protected from the Sun behind the orbit solar
    array. When required, the SEP thrusters fire in the +Z direction.

-   **Nominal/Extended Missions - Observation:** The cruise solar array
    will be jettisoned at the start of the nominal mission. The HGA will
    be protected behind the orbit solar array and the --X face of the
    spacecraft (i.e. in spacecraft shadow).

-   **Nominal/Extended Missions - Downlink:** The orbit solar array
    angle will be varied between 0 and 75° with respect to the Y-axis,
    depending on the spacecraft-Sun distance. The HGA will be in the
    operational position.

### Thermal Design

The proximity to the Sun is particularly demanding for the spacecraft
appendages (e.g. the arrays, booms, HGA), which cannot be protected by
thermal shielding. Figure 5.5 shows the temperature profile
(perpendicular surface to the Sun is considered) with distance from the
Sun using different technologies for external coatings. Special
attention has been given to the selection of suitable coatings, with
absorption/emissivity (α/ε) values below 0.25 even after ageing effects.
In addition, the requirements for ElectroStatic Discharge (ESD) and
spall-resistant external coatings have to be satisfied. This is a
technological challenge.

![](media/image38.wmf)

Figure 5.5: Temperature of a generic spacecraft surface vs the α/ε-ratio
(abs/emi) and as a function of solar constant (Θ= 0°).

The spacecraft is 3-axis stabilised and always Sun pointed (X-axis),
except during SEP firing. The Sun pointing face of the Solar Orbiter is
as small as is possible in order to minimise the thermal input from the
Sun. The remaining spacecraft surfaces will carry radiators. With this
approach, only one thermal shield, made of 3 titanium foils plus 15
layers Kapton/Mylar/Dracon (MLI) shall be used to protect the spacecraft
bus and spacecraft +/-Z faces during the most demanding mission phases.
The external surface will be painted to maintain a very low α/ε (see
above). Furthermore, behind the structure another 15 layers of MLI will
be used for further insulation of the spacecraft from the residual flux
leaking through the MLI, as well as to avoid heat leak during the
coldest phases of the mission.

The thermal shield has been tailored to extend beyond the +X face of the
spacecraft. It will shadow the spacecraft walls during thruster firing
(Sun de-pointing) at the minimum Sun-spacecraft distance. Furthermore,
it is also been designed to keep the HGA and solar array mechanisms in
shadow so that their expected temperatures always remain within standard
limits (i.e. temperatures experienced at 1 AU) to increase reliability.
In order to cool the sunshield and to avoid heat leaks inside the
spacecraft, dedicated radiators are accommodated on the spacecraft +/-Y
walls. Heat pipes are used to increase thermal conductance between the
sunshield structure and its radiators.

The thermal shield design was driven by the following:

**Cruise Phase:** SEP firing at 0.33 AU with ±10° off X-axis pointing
(cf. Figure 6.2)

**Nominal Mission - Observation:** Minimum Sun/spacecraft distance 0.21
AU.

Particular attention had to be paid to the spacecraft appendages, such
as the solar arrays and HGA.

### Power Supply

Photovoltaic materials convert light to electricity but are more
efficient when kept to lower temperatures. Since the spacecraft working
environment is not readily compatible with this, the following strategy
has been conceived.

In order to be compliant with mission requirements and spacecraft needs
it is necessary to carry two sets of solar arrays on the Solar Orbiter.
The first one will be used during the cruise phase for electrical
propulsion. The maximal allowed temperature of this array is 130°C and a
complex operation or design would be required to prevent extreme
temperatures on the solar cells. Thus, since it is not possible or
desirable to keep the array during the nominal mission it will be
jettisoned after the last firing. One other reason why the array cannot
be used during the nominal mission, over and above the thermal concerns,
is the complexity imposed on the attitude and orbit control system
(AOCS) design in keeping the required instrument pointing stability.

The cruise solar array will take advantage of a standard design (for GEO
Missions) to power the electrical propulsion system and other functions
during the cruise phase. It is made up of 2 symmetrical wings of 3
panels each, able to provide 6.2 kW at 1 AU.

The second solar array will be used during the nominal and extended
mission phases. It has a maximal allowed temperature of 150°C. Its
design is different from the cruise array in order to increase the upper
temperature limit. Thus, panels with OSR\'s mounted on them, are to be
made of a honeycomb core with aluminium face sheets. This reduces the
thermal gradient within the panels when illuminated. The solar array
size (and power output) is kept to a minimum.

Extensive work has been done to define the strategy for the variation,
over the orbits of solar radiation on the solar cells, to prevent the
maximum temperature being exceeded during both the cruise and
nominal/extended mission phases. Figure 5.6 shows how the solar
incidence angle can reduce the temperature of the solar arrays.

![](media/image39.png){width="4.5in" height="2.0569444444444445in"}

![](media/image40.png){width="4.500694444444444in"
height="2.0652777777777778in"}

Figure 5.6: Cruise/orbit (top/bottom) solar array temperature versus
distance from the Sun and solar angle incidence.

## Data collection and transmission

The Solar Orbiter mission will have variable transmission durations and
distances from Earth within each orbit, so a clever strategy for data
transmission to Earth has to be considered and some trade-offs made.

During the cruise phase the downlink data consist of housekeeping
information (up to 1 kb/s). Communication via wide beam-width Low Gain
Antennae (LGA) in X-Band is baselined for both uplink and downlink. If
required, the High Gain Antennea (HGA) could be pointed toward the Earth
and in this case a careful co-ordination with the propulsion activities
and attitude manoeuvres is required.

During the nominal/extended mission phases most orbits have perihelion
and aphelion figures of about 0.2 AU and 0.9 AU. The spacecraft-Earth
distance will change with every orbit, from 0.3 to 1.8 AU, and
consequently the downlink data rate that can be supported will vary
considerably. Each orbit takes about 150 days.

The severe thermal environment close to the Sun does not allow the use
of the HGA to downlink data to Earth during the perihelion passages in
real-time. The data are therefore stored in a large on-board mass memory
(baseline 240 Gb) and downlinked later when the Sun-spacecraft distance
is larger then 0.5 AU.

Three sets of scientific observation periods (ten days each) per orbit
are considered as baseline, with the observation strategy for each orbit
tailored such that these periods are centred on the passages through
(see Figure 5.7):

-   maximum southern latitude,

-   maximum northern latitude,

-   perihelion.

We anticipate a high-data-rate acquisition of 75 kb/s in the nominal
observation mode.

![](media/image41.png){width="5.9in" height="3.75in"}Figure 5.7: Modes
of operation during the nominal and extended mission.

In order to avoid antenna pattern distorsion as well as technological
problems, a minimum Sun-spacecraft distance of 0.5 AU for downloading
has been fixed. Only for larger distances the high-rate link via the HGA
is feasible (approximately 110 days/orbit). Out of the 30 days of
observations at 75 kb/s, scientific observations at a lower rate (11.5
kb/s) will take place. When possible these low rate data will be
downlinked in real time, otherwise stored on board and dumped later as
all the other stored data, according to a data dump strategy that takes
into account the slant range spacecraft/Earth and the radio link
capability (750 kb/s maximum at 0.6 AU spacecraft/Earth slant range).

![](media/image42.png){width="5.5in"
height="3.1166666666666667in"}Figure 5.8 shows the data downlink
capabilities for the various orbits. It should be noted that for orbits
2, 3, 5, 7, 8, 10, and 12 much higher data rates than the baselined 75
kb/s can be supported. For these orbits, the limiting factor is the size
of the on-board mass memory. Using 4 Gb chips (which should be readily
available in a few years time, considering the fast pace of progress in
memory technology) instead of the baselined 1 Gb chips, the data
acquisition rate can be doubled for orbits 3 and 8, tripled for orbit 2,
and even quadrupled for orbits 5, 7, 10, and 12. The instruments will
allow for a flexible data acquisition rate. A second ground-station
provided e.g. by an international partner would allow even higher data
rates.

Figure 5.8: Downlink capabilities during the nominal and extended
mission phases.

The communication link calculations show that the required telemetry and
telecommand links are feasible. Figure 5.9 shows the HGA maximum
telecommand data rate against the spacecraft-Earth distance and Figure
5.10 the LGA maximum telemetry data rate against the spacecraft-Earth
distance, respectively.

  ------------------------ ---------- ------------- ----------------------
  Cruise Phase             Uplink     X - Band      34m Perth ⇒ LGA

                           Downlink   X - Band      34m Perth ⇒ LGA

  Nominal + Extended       Uplink     X - Band      34m Perth ⇒ LGA
  phases                                            

                           Downlink   X/KA - Band   34m Perth ⇒ LGA/HGA
  ------------------------ ---------- ------------- ----------------------

![](media/image43.png){width="5.5in" height="2.8847222222222224in"}Table
5.1: Communication link summary.

Figure 5.9: TM (Ka-Band) data rate against S/C-Earth distance via HGA.

Figure 5.10: TM (X-Band) data rate against S/C-Earth distance via LGA
(under clear sky conditions).

## 

## Radiation

An evaluation of the environment radiation dose over the three mission
phases has been made. Results indicated that a nominally shielded (4mm
Al) silicon component is expected to receive a dose of 48 krad over the
entire mission. The total ionising dose is within current engineering
standards.

## Mass Budgets

The mass identified in the system budget is based on the specified
values of the individual units and subsystems. Depending on the maturity
status of the items, contingency is applied on unit/item level. For each
piece of equipment a mass margin was applied in relation to its level of
development, specifically:

-   5% for off-the-shelf items,

-   10% for items to be modified,

-   20% for items to be developed.

A system level margin of 10% was placed on the spacecraft dry mass (dry
mass including sub-system margins).

The Soyuz-ST Fregat launcher from Baikonur allows for a 1310 kg
spacecraft (including adapter mass). The design mass of the Solar
Orbiter with margin is 1308 kg, i.e. within the launch capabilities of a
Soyuz-Fregat launch from Baikonur.

  ---------------------------------------------------------- ------------
  Structure                                                  152 kg

  Thermal control                                            41 kg

  Mechanisms                                                 74 kg

  Pyrotechnics & pyro. harness                               10 kg

  Communications                                             30 kg

  Data handling                                              46 kg

  AOCS (including RCS)                                       37 kg

  Propulsion (SEP)                                           95 kg

  Power (including solar arrays)                             241 kg

  Harness                                                    41 kg

  Payload                                                    130 kg

  \[Spacecraft dry mass                                      897 kg\]

  System margin                                              90 kg

  **Total dry mass (incl. system margin & excl. adapter)**   **987 kg**

  Propellant mass (incl. margins)                            271 kg

  Adapter                                                    50 kg

  Total launch mass without system margin                    1218 kg

  []{#__RefHeading___Toc489691215 .anchor}TOTAL LAUNCH MASS  **1308 kg**
  ---------------------------------------------------------- ------------

Table 5.1: Solar Orbiter mass budget.

[]{#__RefHeading___Toc489691216 .anchor}Subsystem design

## 

## 

## 

## Attitude and orbit control

The basic functions of the attitude and orbit control subsystem (AOCS)
are to support the pre-operational transfer phase, the acquisition and
control of attitude to meet scientific requirements for target
observation, the acquisition and control of safe attitude in case of
failures, the maintenance of the operational orbit and to maintain the
spacecraft in a safe Sun-pointing attitude using a minimum of on-board
resources while ensuring power generation and ground communication.

The spacecraft attitude is three-axis controlled from release in orbit
and throughout the mission. After initial rate reduction and Sun
acquisition using gyroscopes, coarse sun sensors, and thrusters for
actuation, an initial attitude (nominally Sun-pointed) is acquired using
the autonomous star pattern recognition of the star tracker. The routine
operations of the Solar Orbiter mission will all be performed in the
Normal Mode. This mode includes all the functionality necessary to
perform the communications and the nominal science observations (Sun
pointing), and the slew manoeuvres necessary between all these
operations. The attitude measurement is performed with the autonomous
star tracker, while a set of four 4 Nms reaction wheels in a pyramid
configuration is used as primary actuator. This wheel selection has been
made after considering the mass constraints, and frequent wheel
off-loading is required during nominal science observations. The ∆V and
insertion manoeuvres are covered by the SPT-firing mode.

The AOCS hardware architecture, Figure 6.1, is inherited from on-going
ESA scientific missions. The Control and Data Management System (CDMS)
includes a dedicated processor for the AOCS software and the high level
monitoring functions required during the ∆V and insertion manoeuvres for
the autonomy and fail-operational capabilities. Considering the
programmatic design constraints, the AOCS uses many units already
developed and validated for other programmes. In the baseline, the
Inertial Measurement Unit (IMU), the Sun Acquisition Sensors (SAS), the
Star Tracker (STR), and the Reaction Wheels Assembly (RWA) are copies of
units used for Mars Express and Smart-1, providing a good level of reuse
of the AOCS interface unit (AIU). The total mass and maximum power of
the AOCS baseline are 25.8 kg (without margin, which is applied at
system level) and 68.4 W respectively.

A single-axis linear covariance analysis has been performed to assess
the minimum AOCS error contribution, mainly star tracker noise and wheel
quantisation, to the instrument pointing stability budget. For a star
tracker noise of 4 arcsec (1σ), resp. 1 arcsec (1σ), the contribution of
the AOCS errors amounts to 1.26 (resp. 0.42) arcsec which is above the
pointing stability requirement (Table 6.1). Therefore, either a better
noise performance star tracker or a relaxation of the system pointing
stability requirement is needed.

![](media/image45.wmf){width="0.47834645669291337in"
height="1.2562992125984251in"}![](media/image46.wmf){width="0.9in"
height="0.38818897637795274in"}![](media/image47.wmf){width="1.0192913385826772in"
height="0.4216535433070866in"}![](media/image48.png){width="0.2283464566929134in"
height="0.18346456692913385in"}![](media/image49.wmf){width="0.40433070866141735in"
height="0.4779527559055118in"}![](media/image50.png){width="0.45551181102362204in"
height="0.42362204724409447in"}![](media/image51.wmf){width="0.6720472440944882in"
height="0.6318897637795275in"}![](media/image52.png){width="0.16377952755905512in"
height="0.23976377952755906in"}![](media/image53.png){width="0.1673228346456693in"
height="0.2389763779527559in"}![](media/image54.wmf){width="0.5007874015748032in"
height="0.8480314960629921in"}

Figure 6.1: AOCS hardware architecture.

+----------------------+--------+-------------+-----------------------+
|                      | Cruise | Observation | Comments              |
|                      | Mode   | Mode        |                       |
+----------------------+--------+-------------+-----------------------+
| Attitude Control     |        |             | Covariance analysis   |
+----------------------+--------+-------------+-----------------------+
| -   STR noise + RW   | 0.     | 1.17/0.42   | STR: 4 arcsec/ 1      |
|     quantisation     | 83/0.3 |             | arcsec                |
|                      |        | 0.10        |                       |
| ```{=html}           | 0.10   |             | Allocation            |
| <!-- -->             |        | 0.10        |                       |
| ```                  | 0.10   |             | Allocation            |
| -   Delay            |        |             |                       |
|                      |        |             |                       |
| ```{=html}           |        |             |                       |
| <!-- -->             |        |             |                       |
| ```                  |        |             |                       |
| -   Environmental    |        |             |                       |
|     disturbances     |        |             |                       |
+----------------------+--------+-------------+-----------------------+
| Dynamic interaction  |        |             |                       |
+----------------------+--------+-------------+-----------------------+
| -   High-frequency   | 0.10   | 0.10        | Allocation            |
|     disturbance (RW) |        |             |                       |
|                      | 0.40   | 0.40        | Allocation            |
| ```{=html}           |        |             |                       |
| <!-- -->             | 0.20   | 0.20        | Allocation            |
| ```                  |        |             |                       |
| -   Thermo-elastic   |        |             |                       |
|                      |        |             |                       |
| ```{=html}           |        |             |                       |
| <!-- -->             |        |             |                       |
| ```                  |        |             |                       |
| -   Others           |        |             |                       |
+----------------------+--------+-------------+-----------------------+
| Pointing Stability   | 0.9    | 1.26/0.67   | RSS - single axis, 1σ |
| (15 min)             | 6/0.56 |             |                       |
+----------------------+--------+-------------+-----------------------+
| Requirement (single  | 0.29   | 0.29        | Assumptions: 1 arcsec |
| axis)                |        |             | over 15 min specified |
|                      |        |             | half-cone angle at 3  |
|                      |        |             | σ                     |
+----------------------+--------+-------------+-----------------------+

Table 6.1: Pointing stability error budget with STR noise of 4 arcsec /
1 arcsec (1σ) in the S/C body reference frame.

## Reaction Control System 

A monopropellant (hydrazine) system has been selected as the Reaction
Control System (RCS) for spacecraft attitude control after separation
and safe mode, and reaction wheel off-loading. It consists of 2x6 5 N
thrusters, one hydrazine tank, valves, piping, heaters, and insulation.
The thruster configuration has been finalised on the basis of minimised
cosine losses, minimised plume impingement, and best moment arm around
the spacecraft Y-axis.

## Propulsion

The main requirement for the propulsion subsystem is to enable the
completion of Earth to solar orbit manoeuvres, as designed in the
mission analysis. This translates into the total delta-V figure of 4770
m/s, with an additional 140 m/s for the gravity assist auxiliary
manoeuvres. Such a high delta-V drives the selection of the thrusters
towards a high specific impulse system, in order not to exceed the
launch mass limit of 1310 kg imposed by the selected launcher.

A second set of requirements is originated by the attitude control
subsystem. The total angular momentum needed for the off-loading of the
reaction wheels is 15932 Nms, and a minimum thrust level of a few N is
desired for the rate reduction and safe modes.

THRUSTER POSITIONING

To maintain thermal control of the spacecraft, a Sun-pointing attitude
is required throughout the mission. The thruster placement on one side
of the spacecraft (the +Z panel) is a consequence of this requirement.
In order to be able to reach the nominal mission orbits, it is necessary
to have the right momentum variation. Therefore, a firing strategy needs
to be defined. For the conceived mission during firing the spacecraft is
positioned up to a maximum of 30° off Sun pointing at 1 AU. The maximum
Sun-spacecraft pointing angle is a direct consequence of the fact that
momentum variation, through the use of electrical propulsion, is
achieved by firing for a time longer than typical times used for
chemical propulsion since the thrust provided by that propulsion device
is low.

In accordance with the SEP firing strategy illustrated in Figure 6.2,
the thrusters would be canted by the maximum offset angle mentioned
above, and thrust modulation would be used in order to rotate the thrust
vector and keep it tangential to the orbit while the spacecraft would
remain Sun-pointed. The thrusters would have to be placed in a line, at
the intersection of the orbit plane with the +Z panel; only two
thrusters would fire at the same time (one „left" and one „right"),
while the other two would be the redundant pair. Each thruster must be
capable of providing the maximum (0.3 N) thrust when aligned with the
thrust direction.

Figure 6.2: SEP firing philosophy.

## Structure

The structure is split into two modules. This provides several benefits,
including flexibility in the AIV programme, the possibility to integrate
and verify concurrently the two modules, and it allows easy access to
the propulsion system. A high thermo-elastic stability is provided by
use of a high stability Carbon Fibre Reinforced Plastic (CFRP)
structure.

The payload module includes a central cylinder that provides the
required stiffness. Two shear walls link the central cylinder to the +Y
and --Y panels. This results in a lightweight H-shaped beam that
provides a very good combination of stiffness and accessibility to the
instruments and equipment. The central cylinder is also connected to the
service module with an aluminium interface ring that supports also the
--X platform. At its other end is attached a heat-screen made of a CFRP
panel.

The service module does not have the same stringent thermo-elastic
stability requirements as the payload module and is made of a honeycomb
sandwich with aluminium face-sheets. This is a cost-effective
lightweight design.

The cruise solar generator is made of two wings with three panels each.
The wings are attached to two opposite walls of the satellite body when
stowed. In addition, another solar generator is required for the orbital
phase.

## Thermal Control

The extreme environments encountered by the spacecraft throughout the
whole mission drive the thermal design of the Solar Orbiter. At one
extreme, the spacecraft is orbiting the Sun at distances as close as
0.21 AU. At the other extreme, the spacecraft is as far as 1.21 AU from
the Sun. Another important factor is the electrical propulsion that
generates intermittently, i.e. during firing, a considerable amount of
heat inside the spacecraft. Therefore the thermal design has to
accommodate a wide range of heat load levels and locations.

The thermal design is of conventional type (selected surface finishes,
MLI, non-operational heaters) and uses proven hardware supplemented by
the use of heat pipes on the sunshield to cope with the extreme
environmental conditions. The thermal design of critical items such as
the cruise and the orbiter solar arrays, the HGA, the payload detector
(low temperature) radiators, the sunshield and the spacecraft (ambient
temperature) radiators have been investigated. As a result, the
associated heater power and mass budgets have been established.

The spacecraft configuration is represented in the sub-figures of Figure
5.2. One of the major design challenges has been to cope with the
extremely high solar fluxes especially during the spacecraft depointing
required when firing the SPT thrusters. Several combinations of SPT
thrusters location/firing strategy were studied. The SPT thrusters have
been located on the +Z spacecraft panel and the firing strategy has been
tailored to the spacecraft thermal control needs.

The basic design concept consists of shading as much as possible any
part of the spacecraft from the Sun at the closest approach to the Sun
and during the SPT firings in the cruise phase. To this aim, a sunshield
has been accommodated at the +X side of the spacecraft. It extends on
the +/-Z sides to make sure that the spacecraft is shadowed even when
the +X axis is depointing as much as 10º from the Sun. The +/-Z
depointing of 10º occurs during the SPT firing at the closest distance
to the Sun (0.33 AU) during the cruise phase. However, other SPT firings
occur at higher angles: as high as 28º at the closest Sun-spacecraft
distance of 0.67 AU. For that particular case, it has been verified that
the solar intensity is harmless to the spacecraft protruding parts
(detector radiators, SPT thrusters, HGA, mechanisms...) provided they
are insulated with typical MLI blankets.

# Ground systems and operations

## Ground segment facilities and services

All facilities established for the Solar Orbiter ground operations will
be based on an extension of existing ground segment infrastructure,
tailored to meet the requirements of the Solar Orbiter mission.

The ground operations facilities consist of:

-   The ground stations and the communications network;

-   The mission control centre (infrastructure, computer hardware);

-   The flight control system (data processing and flight dynamics
    software).

An overview of the ground system is shown in Figure 7.1 below.

![](media/image55.png){width="5.5159722222222225in"
height="5.210416666666666in"}

Figure 7.1: Ground systems overview.

## Ground stations and communications network 

It is assumed that the 35 m ESA ground station at Perth will be used for
contact with the spacecraft during all mission phases. Despite heavy
demands for support from Perth for other ESA missions, for the cruise
phase, the Perth station can be assumed to be sufficiently available for
operations. For the nominal/extended mission phase, we assume that the
Perth station is available for at least 8 hours per day, for costing
purposes, and recognise that the actual stations used have to be
clarified at a later time. A possibility to be investigated is the use
of the 70 m station in Sardinia for those times when Perth is not free.

During the first 10 days of the mission and during critical mission
phases, the ESA 15 m station at Kourou will also be used. The 15 m
station at Villafranca will be available as back up.

All ESA stations will interface with the control centre in ESOC
(Darmstadt) via the OPSNET communications network. OPSNET is a closed
Wide Area Network for data (telecommand, telemetry, tracking data,
station monitoring and control data) and voice.

## Mission control centre

The Solar Orbiter will be operated by ESOC, and it will be controlled
from the Mission Control Centre (MCC) which basically consists of the
Main Control Room (MCR) augmented by the Flight Dynamics Room (FDR) and
dedicated control rooms and project support rooms. During periods around
major mission events mainly during launch and critical periods of the
transfer phase the MCR will be used for Solar Orbiter mission control.
During the science operations phase (nominal and extended mission) and
also during low activity periods, the mission control will be conducted
from a dedicated control room.

The control centre is equipped with workstations giving access to the
computer systems used for different tasks of operational data
processing. It will be staffed by dedicated Solar Orbiter spacecraft
operations staff, and experts in spacecraft control, flight dynamics and
network control, who will be shared between ESA programmes but will be
available for the full mission duration.

## Computer facilities and network

The computer configuration used in the Mission Control Centre for the
Solar Orbiter mission will be derived from the existing infrastructure.
The computer system consists of

-   a mission dedicated computer system (workstation hosting SCOS 2000)
    used for real time telemetry processing and for command preparation
    and execution, telemetry and command log archiving, and also for non
    real-time mission planning and mission evaluation;

-   a workstation hosting the flight dynamics system (ORATOS);

-   a workstation hosting the science telemetry data distribution and
    instrument command reception system (data servers);

-   the simulation computer, providing an image of the spacecraft system
    during ground segment verification, for staff training and during
    operations.

All computer systems in the control centre will be redundant with common
access to data storage facilities and peripherals. All computing systems
will be connected by a Local Area Network (LAN) to allow for rapid
transfer of, and for joint access to, data.

The external connections to the science operations team, science data
processing centres and PI's will use commercial/public networks (Wide
Area Network).

## flight control software system

A flight control system based on planned infrastructure development
(SCOS 2000), using a distributed hardware and software architecture for
all spacecraft monitoring and control activities will be established
including

-   telemetry reception facilities for acquisition, quality checking,
    filing and distribution;

-   telemetry analysis facilities for status/limit checking, trend
    evaluation;

-   telecommand processing facilities for the generation of command for
    control, master schedule updates, and on-board software maintenance,
    including their uplink and verification;

-   monitoring of instrument housekeeping telemetry for certain
    parameters which affect spacecraft safety and command acceptance and
    execution verification;

-   separation and forwarding of payload telemetry to science data
    processing centres;

-   checking, reformatting, scheduling command request for payload.

Within the SCOS 2000 system, mission specific software will be developed
wherever necessary.

## Mission operations concept

The Solar Orbiter mission operation will comprise

-   spacecraft operations, consisting of mission planning, spacecraft
    monitoring and control, and all orbit and attitude determination and
    control;

-   science instrument operations, consisting of the implementation of
    the observation schedules and collection and data quality control of
    the science telemetry.

Mission operations proper will commence at separation of the Solar
Orbiter spacecraft from the launcher and will continue until the end of
the mission, when ground contact to the spacecraft will be aborted.
Mission operations will comprise the following tasks:

-   mission planning, long term planning and short term planning (24
    hours to one week time frame);

-   spacecraft status monitoring;

-   spacecraft control, based on monitoring and following the flight
    operations plan and the short-term plan;

-   orbit determination and control using tracking data and implementing
    orbit manoeuvres;

-   attitude determination and control based on the processed attitude
    sensors data in the spacecraft telemetry and by commanded updates of
    control parameters in the on-board attitude control system;

-   on-board software maintenance;

-   operations support for the experimenters in terms of telemetry
    packet routing and command checking with respects to spacecraft
    safety, and telecommand uplink;

-   maintenance of ESA ground facilities and network.

## Science instrument operations and team structure

The Solar Orbiter is a Principal Investigator (PI) mission. The
instruments will be provided and run by PI teams who will take on
responsibility for the health and scientific operation of their
instruments.

Each team will interface with the MCC to monitor all instrument health
and operational aspects. This may be done through dedicated workstation
facilities at the home institutes, which are linked to ESOC for the
monitoring of instrument parameters and the receipt of data.

The Solar Orbiter is not run by day to day real-time operation, as is
SOHO. We will have a 150 day orbit with approximately 30 days of
observation for the remote sensing instruments centred on the perihelion
and maximum latitude passages. The observation period around perihelion
is called the Nominal Observing period and the rest of the orbit is the
Time-Share Observation period. During the latter, the high-resolution
remote sensing instruments will usually be off. The full disc imagers
and the particle and field instruments will be in a synoptic mode and
stored data from the last solar passage will be relayed to Earth.

The scientific planning for the solar passes (Nominal Observation
periods) will be determined through regular PI-team workshops held once
per orbit, i.e. once in every 150 days (5 months), about 2 months prior
to the onset of the start of the Nominal Observation period. These
workshops may be held at any PI institute, ESA institute or a convenient
meeting location. The 30 days of the next pass will be mapped out and
instrument teams will return home to prepare the observation sequences
ready to install prior to the onset of the pass. Observations for the
pass must be loaded into deferred command stores (time-tagged) due to
the lack of contact during the pass, i.e. commands for the 30 days must
be well defined prior to the onset of the pass. Some features of the
plan may be deferred to the last few days before the 30 day period, in
order to make target selection at the last minute, but the basic
sequences should be arranged well ahead of time.

It is expected that most operations will involve a co-ordination of the
instruments, i.e. the bulk of the observations will be through Joint
Observing Programmes (JOPs) which may be defined in detail and posted on
the Web for inspection.

Further details will be arranged through a suite of Web pages with a
planning and scheduling tool facility. The Web pages will also include
public understanding facilities, publication and background information.

The PI teams will be responsible for the calibration of data and the
archiving of their instrument data. Full archives will be available,
possibly located at the sites of current solar archives such as the SOHO
archives at IAS and RAL, with Web interfaces to the world community.

The Solar Orbiter mission exploitation must not be restricted to the
instrument PI teams. The data should be made public and accessible to
the general scientific community, a procedure which has been shown by
the SOHO community to be very successful and allowing the best possible
use of the data.

# Communication and outreach

Previous solar missions have received great public and media interest.
This is partly due to a raised public awareness of the Sun and its
influence on Earth, and due to the highly appealing and startling
imagery -- in particular from the SOHO and TRACE missions. The public
can relate to the Sun and is in awe of its complexity and power.

SOHO results have been the subject of hundreds of media interviews over
the last 4.5 years, and countless talks have been given by the SOHO
community to public audiences. There have been 14 SOHO-related stories
on CNN interactive and 8 on BBC interactive since January 1, 2000 and
over 5,000,000 requests per month to the SOHO Web servers. The SOHO
real-time image screensaver has been requested by over 150,000 distinct
hosts since its launch in November 1999. SOHO is a major feature of the
IMAX movie "SolarMax" which was released internationally in July 2000.
People see the Sun as an exciting astronomical body which has a direct
influence on them.

There is no doubt that the ambitious mission profile of the Solar
Orbiter, passing close to the Sun and climbing out of the ecliptic using
flybys at Venus will also be of great interest to the public. High
resolution images taken close to the Sun and pictures from over the
polar regions will capture the attention of the media. Exploring the
inner regions of our solar system is something that the person in the
street can understand and be excited by. Exploring the sources of "space
weather" is seen by the public as a topic which agencies such as ESA
should address. The Solar Orbiter is a mission which has enormous
potential for public interest.

The mission will provide excellent material to fuel the public interest
with close-up images of the Sun. A Web site will be maintained with the
most recent images displayed and with good background information on the
mission and its objectives, including a picture gallery. The mission
Project Scientist and the instrument PIs are responsible for this. The
Web site includes contact information for all of the groups; the media
are always looking out for contacts in their own countries. Major new
results will be highlighted as "hotshots" with good images and captions
for public use and press releases should be written whenever appropriate
and certainly at the time of major mission milestones.

Basic, colourful brochure information will be made available, as well as
CD-ROMs, which can be distributed widely (see e.g. the SOHO CD-ROM).

A close liaison will be maintained between the Solar Orbiter project and
the ESA public relations offices to ensure the best possible utilisation
of Orbiter results.

Special attention will be paid to the provision of information and
materials for schools and students. Astronomy is a popular topic and the
Solar Orbiter is able to provide excellent material for use by
educational establishments.

# Management

## F2/F3 procurement approach

Here, we describe a possible management approach for the Solar Orbiter
mission implementation. The exact approach to be taken may differ from
this and will be known when the mission is approved. Furthermore, the
procurement schedule and model philosophy as shown in the following
sections is provided for illustrative purposes only, as this was the
basis for the assessment study. The approach is based on the Mars
Express Project approach. This approach gives more responsibility to
industry than in previous ESA scientific projects, with a fixed price
contract instead of a cost reimbursement type of contract.

## Procurement philosophy

The proposed procurement scheme for the Solar Orbiter is based on the
concept that the payload, consisting of a number of instruments and the
associated control and data handling/processing electronics, will be
provided by Principal Investigators (PI\'s) with funding from ESA member
states. For the purpose of the study, it has been assumed that all booms
relevant to instruments will be provided by the PI\'s.

ESA would be responsible for

-   the overall spacecraft and mission design,

-   provision of the spacecraft bus and payload module (PLM)

-   integration of the PLM onto the spacecraft bus,

-   system testing,

-   spacecraft launch and operations,

-   acquisition and distribution of data to Science Data Centre(s)
    (SDC).

## Scientific management and payload selection

After the approval of the mission, the agency will issue an open call
for tender (ITT) for the selection of two (TBC) competitive industrial
contractors for a definition phase. This is equivalent to a Phase A
study, as performed in the past, but includes studies of the payload
design and requirements. The payload will be selected by means of an
Announcement of Opportunity (AO) which is released in parallel with the
definition phase ITT. In this way, a large part of the definition phase
will be performed taking into account the design of the selected
instruments and not on the basis of the model payload.

The AO will outline the spacecraft mission and resources, technical
interfaces, schedule, deliverable items, the responsibilities of all
parties, etc. The AO contains a description of the model payload and
requests proposals for the complete instrumentation (and their
integration) and for the Solar Orbiter scientific operations concept
(see Section 7.). These proposals must describe in detail the scientific
objectives and the design and development of the envisaged
instrumentation and associated software. The proposals must also
describe the management structure for each instrument and show how the
responsibilities for the scientific, technical, operational and analysis
aspects are discharged.

Selection of the Solar Orbiter experiments will take place through
normal procedures which include a technical evaluation by the ESA
executive, supported by the selected industrial contractor for the
definition phase, and a scientific evaluation by ESA\'s scientific
advisory bodies and approval by the Science Programme Committee (SPC).

The selected PI\'s will be responsible for obtaining the necessary
funding from the appropriate national authorities. Following selection,
refined resource allocations and instrument interfaces will be
negotiated with the PI\'s prior to the start of the Phase B study. These
resources and interfaces will be frozen at the end of Phase B. Special
items of mechanical ground support equipment and electrical/optical
check‑out equipment will also have to be supplied by the Solar Orbiter
collaboration. This applies in particular for payload tests required at
satellite level.

The payload module (PLM) and associated electronics are considered to be
qualified and accepted at payload level, with a very well defined set of
interfaces at satellite level. The satellite prime contractor will
integrate the instruments with the service module (SVM), forming the
protoflight model (PFM) satellite, to be tested for acceptance.

During the instrument development phase, the project team conducts, with
the Solar Orbiter collaboration, a preliminary design review, a critical
design review and a flight model review. A Solar Orbiter science team
comprising the scientists from the collaboration, discipline scientists
and the ESA project scientist will be established to support the
project.

## Industrial management

It is proposed that an industrial prime contractor, having
responsibility for the integration and testing of the Solar Orbiter
payload assembly and the spacecraft, should carry out the Solar Orbiter
spacecraft procurement. Furthermore, the prime contractor will assist
ESA in the procurement of the launcher, and related services, and for
the acquisition of the PLM. The industrial contracts will be funded and
placed by ESA. The responsibility for control and monitoring the
contracts and provision for liaison between partners, contractors and PI
groups are with the ESA project team.

## Development philosophy and schedule

### General

The proposed Solar Orbiter development scheme has the aim of minimising
the spacecraft development cost at reasonable risk. It is based on a
minimum number of modules and spacecraft models and the maximum
utilisation of each model. It is assumed that all parts, units and
assemblies would be available off-the-shelf or taken from the
BepiColombo project. No technology developments have been considered,
only adaptation of design when necessary, like for the orbit solar
array.

The first system level test at spacecraft level will be conducted on the
Structural Model (SM), to establish the structural characteristics and
compatibility at the system level. Other tests, like mass properties and
alignment, will be also performed on the spacecraft SM module. Then, the
PFM models of the PLM and of the SVM will be assembled at spacecraft
level, and the final testing is performed to finally accept the PFM
spacecraft. The facilities to test the extreme thermal aspects of the
mission are assumed to be available also from BepiColombo, and only the
adaptation to the Solar Orbiter requirements have been considered.

### Payload model philosophy

The model philosophy chosen is based on the requirement for a single
flight instrument supported by adequate spares, e.g. a set of kits of
parts to cope with random failure and allowing a unit repair within 3-4
weeks, when spare units do not exist.

To meet the early system level test a Structural Thermal Model (STM)
will be built, fitted with at least one active unit of the instrument
detectors, the rest being representative dummies. This would allow a
full verification of the thermo‑mechanical design of the payload. For
electrical and functional testing an Engineering Model (EM) is required.
All payload elements will undergo qualification testing, if necessary,
prior to delivery for AIV (Assembly, Integration, and Verification) and
test activities.

### Service module model philosophy

A protoflight approach is proposed where only one model, the PFM, is
built to flight standard level. However, to perform the early system
level tests with the payload, a SM model of the bus would be utilised,
using mechanically representative dummies instead of flight parts. The
SM spacecraft structure will be at flight standard. The SM will be
subjected to a reduced test programme (sinus vibration, modal survey,
mass properties, alignment) which allows verification of the design
concept and interfaces. It is expected that most of the units will have
been qualified on previous programmes. However, when deemed necessary to
qualify a unit for the project, this unit, duly refurbished, will become
a flight spare. It is assumed that spare units/components kits will be
available to cope with random failures and to allow repair of any unit
within 3 to 4 weeks.

### Schedule

The schedule of the Solar Orbiter programme is shown in Figure 9.1.
While the assumed programmatic launch date for the Solar Orbiter is the
year 2008, the schedule shows that the earliest possible launch date is
less than 6 years after the start of the programme. It should be noted
that the schedule doesn't account for funding availability problems. It
is only a technical assessment of the time required to implement the
Solar Orbiter Project. There are other constraints such as the need to
have satellite parts designs available from the BepiColombo cornerstone,
for cost and availability purposes, or that the cruise module of
BepiColombo be launched six months ahead of the Solar Orbiter, for
operations resources availability and cost. There are launch windows in
May 2007, January 2009, August 2010, and in principle every 18 months.

Figure 9.1: Solar Orbiter programme schedule.

# List of Acronyms

  ------------ ----------------------------------------------------------
  ACE          Advanced Composition Explorer

  ADCT         Adaptive Discrete Cosine Transform

  AIU          AOCS Interface Unit

  AIV          Assembly, Integration, and Verification

  AO           Announcement of Opportunity

  AOCS         Attitude and Orbit Control Subsystem

  APS          Active Pixel Sensor

  ASP          Advanced Stokes Polarimeter

  AU           Astronomical Unit, 150 Mio km

  BOL          Beginning of Life

  CCD          Charge Coupled Device

  CDMS         Control and Data Management System

  CDS          Coronal Diagnostic Spectrometer (on SOHO)

  CELIAS       Charge and Element and Isotope Analysis System (on SOHO)

  CFRP         Carbon Fibre Reinforced Plastic

  CID          Charge Injection Device

  CME          Coronal Mass Ejection

  COG          Center of Gravity

  CRS          Coronal Radio Sounding

  CTOF         Charge determining Time-of-Flight sensor (of CELIAS on
               SOHO)

  DOF          Degree of Freedom

  DPU          Data Processing Unit

  DUD          Dust Detector

  EAS          Electron Analyser System (of SWA)

  EIT          EUV Imaging Telescope (on SOHO)

  ELF          Extreme Low Frequency Waves

  EM           Engineering Model

  EMC          Electro-Magnetic Cleanliness

  ENA          Energetic Neutral Atoms

  EOL          End of Life

  EPD          Energetic Particle Detector

  ESA          European Space Agency

  ESD          ElectroStatic Discharge

  ESOC         European Space Operations Centre

  EUI          EUV Imager

  EUS          EUV Spectrometer

  EUV          Extreme UltraViolet

  FDR          Flight Dynamics Room

  FDT          Full Disc Telescope (of VIM)

  FIP          First Ionisation Potential

  FO           Filtergraph Optics (of VIM)

  FOV          Field-of-View

  FSI          Full Sun Imager (of EUI)

  HGA          High Gain Antenna

  HRI          High Resolution Imager (of EUI)

  HRT          High Resolution Telescope (of VIM)

  IDP          Interplanetary Dust Particles

  IMPACT       In-situ Measurements of PArticles and CME Transients
               (STEREO)

  IMU          Inertial Measurement Unit

  ISS          Image Stabilisation System

  ITT          Invitation To Tender

  JOP          Joint Observing Programme

  KIS          Kiepenheuer-Institut für Sonnenphysik (in Freiburg,
               Germany)

  LAN          Local Area Network

  LASCO        Large Angle and Spectrometric Coronagraph (on SOHO)

  LGA          Low Gain Antenna

  LOS          Line-of-Sight

  MAG          Magnetometer

  MCP          Micro Channel Plate

  MCR          Mission Control Room

  MHD          Magneto-Hydrodynamics

  MIS          Minor Ion Sensor (of SWA)

  MLI          Multi Layer Insulation

  NASA         National Aeronautics and Space Administration

  NEAR         Near Earth Asteroid Rendezvous

  NED          Neutron Detector

  NPD          Neutral Particle Detector

  OSR          Optical Surface Reflector

  PAS          Proton and a-particle Sensor (of SWA)

  PFM          ProtoFlight Model

  PLM          Payload Module

  PMP          Polarisation Modulation Package (of VIM)

  PWA          Plasma Wave Analyser of RPW

  RAD          Radiometer

  RAS          Radio Spectrometer (of RPW)

  RCS          Reaction Control System

  RPW          Radio and Plasma Wave Analyser

  RW           Reaction Wheel

  RWA          Reaction Wheel Assembly

  SAS          Sun Acquisition Sensor

  SEP          Solar Electric Propulsion

  SM           Structural Model

  SPC          Science Programme Committee

  SPT          Stationary Plasma Thruster

  STM          Structural Thermal Model

  S/C          Spacecraft

  SiC          Silicon Carbide

  SOHO         Solar and Heliospheric Observatory

  STEREO       Solar-TErrestrial Relations Observatory

  STR          Star Tracker

  SUMER        Solar Ultraviolet Measurement of Emitted Radiation (on
               SOHO)

  SVM          Service Module

  SWA          Solar Wind Analyser

  SWICS        Solar Wind Ion Composition Sensor (on Ulysses)

  TBC          To Be Confirmed

  TDS          Time Domain Sampler

  TM           TeleMetry

  TOF          Time-of-Flight system

  TRACE        Transition Region And Coronal Explorer

  TSI          Total Solar Irradiance

  UVC          Ultraviolet an Visible-Light Coronagraph

  UVCS         UltraViolet Coronagraph Spectrometer (on SOHO)

  VIM          Visible-light Imager and Magnetograph

  VIRGO        Variability of solar Irradiance and Gravity Oscillations
               (on SOHO)

  VLF          Very Low Frequency waves
  ------------ ----------------------------------------------------------