<table><caption>**Table 1** Science objectives of Comet Interceptor</caption><thead><tr><th>Science Theme</th><th>Level 0 - Science Objective</th></tr></thead><tbody><tr><td rowspan="3">Comet Nucleus Science - What is the surface composition, shape, morphology, and structure of the target object?</td><td>Comet Interceptor shall characterise the target shape, size and rotation state.</td></tr><tr><td>Comet Interceptor shall characterise the target morphology.</td></tr><tr><td>Comet Interceptor shall assess the bulk composition of the target surface.</td></tr><tr><td rowspan="7">Comet Environment Science - What is the composition of the coma, its connection to the nucleus (activity) and the nature of its interaction with the solar wind?</td><td>Comet Interceptor shall characterise and map the bulk neutral composition of the coma and determine any local structure and connection to the nucleus.</td></tr><tr><td>Comet Interceptor shall determine the isotopic composition of the coma.</td></tr><tr><td>Comet Interceptor shall characterise the structure of the dust environment of the coma and determine any connection to the nucleus.</td></tr><tr><td>Comet Interceptor shall characterise the coma dust properties, including reflectance and polarimetric properties and determine dust fluxes.</td></tr><tr><td>Comet Interceptor shall determine motion and evolution of ion rays and other coma and tail features including dust and gas.</td></tr><tr><td>Comet Interceptor shall characterise the plasma environment around the target, determine any resulting boundaries and assess energy, mass and momentum transfer.</td></tr></tbody></table>

<table><caption>**Table 2** Top-level science requirements corresponding to the science objectives related to the cometary nucleus</caption><thead><tr><th>Level 0 - Science Objective</th><th>Level 1 - Science Requirement</th></tr></thead><tbody><tr><td rowspan="4">Comet Interceptor shall characterise the target shape, size and rotation state.</td><td>Comet Interceptor shall establish the size and shape of the directly illuminated portion of the target to an accuracy of &lt; 60 m.</td></tr><tr><td>Comet Interceptor shall establish the size and shape of the unilluminated (potentially indirectly illuminated by backscattered sunlight) portion of the target to an accuracy of &lt; 600 m.</td></tr><tr><td>Comet Interceptor shall determine existence and location of any non-surface bound boulders with sizes &gt; decametres in the vicinity of the nucleus.</td></tr><tr><td>Comet Interceptor shall assess the rotation rate and angular momentum orientation with an accuracy of 1%.</td></tr><tr><td rowspan="2">Comet Interceptor shall characterise the target morphology</td><td>Comet Interceptor shall establish the existence and characterise surface features on the directly illuminated surface to an accuracy of &lt; 20 m.</td></tr><tr><td>Comet Interceptor shall establish the existence and characterise surface features in unilluminated areas (potentially indirectly illuminated by backscattered sunlight) to an accuracy of &lt; 300 m.</td></tr><tr><td rowspan="3">Comet Interceptor shall assess the bulk composition of the target surface</td><td>Comet Interceptor shall determine the target phase curve by measuring the geometric albedo of the surface &gt; 5 different phase angles (φ) where max phase angle should be least 90° and include measurement at φ &lt; 5°.</td></tr><tr><td>Comet Interceptor shall assess the bulk surface roughness via polarimetric observations.</td></tr><tr><td>Comet Interceptor shall determine the colour of the object and its colour diversity at visible wavelengths.</td></tr><tr><td></td><td>Comet Interceptor shall determine the thermal and spectral characteristics of the surface.</td></tr></tbody></table>

The two science themes are first broken down into top level Science Objectives (Level 0) that describe the properties of the target to be investigated. Each Science Objective is further split in a series of Science Requirements (Level 1) that quantitatively describe the features, characteristics and processes that need to be measured to achieve the Science Objective. Finally, Level 2 Requirements, not included here for brevity, describe the measurements by each instrument that contribute to meeting the Level 1 requirements. Tables 1, 2, and 3 list the Science Objectives and the Level 1 Science Requirements.
<table><caption>**Table 1** Science objectives of Comet Interceptor</caption><thead><tr><th>Science Theme</th><th>Level 0 - Science Objective</th></tr></thead><tbody><tr><td rowspan="3">Comet Nucleus Science - What is the surface composition, shape, morphology, and structure of the target object?</td><td>Comet Interceptor shall characterise the target shape, size and rotation state.</td></tr><tr><td>Comet Interceptor shall characterise the target morphology.</td></tr><tr><td>Comet Interceptor shall assess the bulk composition of the target surface.</td></tr><tr><td rowspan="7">Comet Environment Science - What is the composition of the coma, its connection to the nucleus (activity) and the nature of its interaction with the solar wind?</td><td>Comet Interceptor shall characterise and map the bulk neutral composition of the coma and determine any local structure and connection to the nucleus.</td></tr><tr><td>Comet Interceptor shall determine the isotopic composition of the coma.</td></tr><tr><td>Comet Interceptor shall characterise the structure of the dust environment of the coma and determine any connection to the nucleus.</td></tr><tr><td>Comet Interceptor shall characterise the coma dust properties, including reflectance and polarimetric properties and determine dust fluxes.</td></tr><tr><td>Comet Interceptor shall determine motion and evolution of ion rays and other coma and tail features including dust and gas.</td></tr><tr><td>Comet Interceptor shall characterise the plasma environment around the target, determine any resulting boundaries and assess energy, mass and momentum transfer.</td></tr></tbody></table>

<table><caption>**Table 2** Top-level science requirements corresponding to the science objectives related to the cometary nucleus</caption><thead><tr><th>Level 0 - Science Objective</th><th>Level 1 - Science Requirement</th></tr></thead><tbody><tr><td rowspan="4">Comet Interceptor shall characterise the target shape, size and rotation state.</td><td>Comet Interceptor shall establish the size and shape of the directly illuminated portion of the target to an accuracy of &lt; 60 m.</td></tr><tr><td>Comet Interceptor shall establish the size and shape of the unilluminated (potentially indirectly illuminated by backscattered sunlight) portion of the target to an accuracy of &lt; 600 m.</td></tr><tr><td>Comet Interceptor shall determine existence and location of any non-surface bound boulders with sizes &gt; decametres in the vicinity of the nucleus.</td></tr><tr><td>Comet Interceptor shall assess the rotation rate and angular momentum orientation with an accuracy of 1%.</td></tr><tr><td rowspan="2">Comet Interceptor shall characterise the target morphology</td><td>Comet Interceptor shall establish the existence and characterise surface features on the directly illuminated surface to an accuracy of &lt; 20 m.</td></tr><tr><td>Comet Interceptor shall establish the existence and characterise surface features in unilluminated areas (potentially indirectly illuminated by backscattered sunlight) to an accuracy of &lt; 300 m.</td></tr><tr><td rowspan="3">Comet Interceptor shall assess the bulk composition of the target surface</td><td>Comet Interceptor shall determine the target phase curve by measuring the geometric albedo of the surface &gt; 5 different phase angles (φ) where max phase angle should be least 90° and include measurement at φ &lt; 5°.</td></tr><tr><td>Comet Interceptor shall assess the bulk surface roughness via polarimetric observations.</td></tr><tr><td>Comet Interceptor shall determine the colour of the object and its colour diversity at visible wavelengths.</td></tr><tr><td></td><td>Comet Interceptor shall determine the thermal and spectral characteristics of the surface.</td></tr></tbody></table>

The two science themes are first broken down into top level Science Objectives (Level 0) that describe the properties of the target to be investigated. Each Science Objective is further split in a series of Science Requirements (Level 1) that quantitatively describe the features, characteristics and processes that need to be measured to achieve the Science Objective. Finally, Level 2 Requirements, not included here for brevity, describe the measurements by each instrument that contribute to meeting the Level 1 requirements. Tables 1, 2, and 3 list the Science Objectives and the Level 1 Science Requirements.
<table><caption>**Table 3** Top-level science requirements related to the cometary coma environment</caption><thead><tr><th>Level 0 – Science Objective</th><th>Level 1 – Science Requirement</th></tr></thead><tbody><tr><td>Comet Interceptor shall characterise and map the bulk neutral composition of the coma and determine any local structure and connection to the nucleus.</td><td>Comet Interceptor shall map coma volatiles including H<sub>2</sub>O, CO and CO<sub>2</sub> with an accuracy of 300 m<br>Comet Interceptor shall determine absolute and relative densities of volatiles (e.g., H<sub>2</sub>O, CO, and CO<sub>2</sub>) along the fly-by trajectory.<br>Comet Interceptor shall detect minor species (e.g., O<sub>2</sub>) and potential distributed sources.<br>Comet Interceptor shall make Hydrogen Lyman-alpha observations of the coma<br>Comet Interceptor shall determine the large-scale gas coma activity, morphology and any variability therein.</td></tr><tr><td>Comet Interceptor shall determine the isotopic composition of the coma.</td><td>Comet Interceptor shall determine D/H and <sup>18</sup>O/<sup>16</sup>O ratios in H<sub>2</sub>O.<br>Comet Interceptor shall investigate isotopes in other species including <sup>34</sup>S/<sup>32</sup>S and <sup>13</sup>C/<sup>12</sup>C.</td></tr><tr><td>Comet Interceptor shall characterise the structure of the dust environment of the coma and determine any connection to the nucleus.</td><td>Comet Interceptor shall determine the large-scale dust coma activity, morphology and any variability therein.<br>Comet Interceptor shall determine the inner dust coma brightness distribution to potentially capture evidence of non-1/r<sup>2</sup> particle distributions resulting from processes such as acceleration, sublimation, and/or fragmentation.<br>Comet Interceptor shall establish evidence for (and any variation of) dust emission from the surface by measuring reflected light from dust in the vicinity of the nucleus.<br>Comet Interceptor shall determine the existence (or not) of local dust structures in the coma.</td></tr><tr><td>Comet Interceptor shall characterise the coma dust properties, including reflectance and polarimetric properties and determine dust fluxes.</td><td>Comet Interceptor shall measure the mass and fluence of dust in the range 1-200 µm, where possible with multipoint measurements.<br>Comet Interceptor shall search for the existence of, and if found characterise, dust particles in the range centimetre – decimetre (chunks and snowballs).<br>Comet Interceptor shall determine the compositional diversity of the dust within the comet coma.<br>Comet Interceptor shall determine coma dust reflectance and polarimetric properties at visible wavelengths.</td></tr><tr><td>Comet Interceptor shall determine motion and evolution of ion rays and other coma and tail features including dust and gas.</td><td>Comet Interceptor shall produce maps of ion rays via wide field imaging of the coma<br>Comet Interceptor shall determine the existence (or not) of in-situ plasma signatures correlating to features observed through wide field imaging of coma.</td></tr><tr><td>Comet Interceptor shall characterise the plasma environment around the target, determine any resulting boundaries and assess energy, mass and momentum transfer.</td><td>Comet Interceptor shall determine the electron, ion, ENA, E and B field environment around the comet.<br>Comet Interceptor shall assess and quantify coupling between neutral, dust and plasma, including potential assessment of dust charging.<br>Comet Interceptor shall identify plasma structures and boundaries and characterise, where possible, with multipoint measurements.<br>Comet Interceptor shall characterise the plasma wave environment (λ≥2 km) to constrain energy and momentum transfer across the coma.</td></tr></tbody></table>

## 4 Planned Measurements and Payload Overview

The Comet Interceptor payload (Table 4) was selected to fulfil the mission Science Objectives.

The Comet Camera (CoCa) on Spacecraft A will contribute to both mission themes by determining the physical properties of the nucleus and coma with high resolution images. CoCa will determine the physical properties of the nucleus with high resolution images, tracking the nucleus through the flyby. The current CoCa instrument design will

- Determine the size, shape, and albedo of the nucleus

- Identify features on the nucleus at <20 metres resolution (assuming a nominal 1000 km flyby) and allow comparison with similar observations of the surfaces of short-period comets

- Establish the reflectivity spectral gradient of the nucleus and of sub-regions on the nucleus (Thomas and Keller 1989; Fornasier et al. 2015)
<table><caption>**Table 3** Top-level science requirements related to the cometary coma environment</caption><thead><tr><th>Level 0 – Science Objective</th><th>Level 1 – Science Requirement</th></tr></thead><tbody><tr><td>Comet Interceptor shall characterise and map the bulk neutral composition of the coma and determine any local structure and connection to the nucleus.</td><td>Comet Interceptor shall map coma volatiles including H<sub>2</sub>O, CO and CO<sub>2</sub> with an accuracy of 300 m<br>Comet Interceptor shall determine absolute and relative densities of volatiles (e.g., H<sub>2</sub>O, CO, and CO<sub>2</sub>) along the fly-by trajectory.<br>Comet Interceptor shall detect minor species (e.g., O<sub>2</sub>) and potential distributed sources.<br>Comet Interceptor shall make Hydrogen Lyman-alpha observations of the coma<br>Comet Interceptor shall determine the large-scale gas coma activity, morphology and any variability therein.</td></tr><tr><td>Comet Interceptor shall determine the isotopic composition of the coma.</td><td>Comet Interceptor shall determine D/H and <sup>18</sup>O/<sup>16</sup>O ratios in H<sub>2</sub>O.<br>Comet Interceptor shall investigate isotopes in other species including <sup>34</sup>S/<sup>32</sup>S and <sup>13</sup>C/<sup>12</sup>C.</td></tr><tr><td>Comet Interceptor shall characterise the structure of the dust environment of the coma and determine any connection to the nucleus.</td><td>Comet Interceptor shall determine the large-scale dust coma activity, morphology and any variability therein.<br>Comet Interceptor shall determine the inner dust coma brightness distribution to potentially capture evidence of non-1/r<sup>2</sup> particle distributions resulting from processes such as acceleration, sublimation, and/or fragmentation.<br>Comet Interceptor shall establish evidence for (and any variation of) dust emission from the surface by measuring reflected light from dust in the vicinity of the nucleus.<br>Comet Interceptor shall determine the existence (or not) of local dust structures in the coma.</td></tr><tr><td>Comet Interceptor shall characterise the coma dust properties, including reflectance and polarimetric properties and determine dust fluxes.</td><td>Comet Interceptor shall measure the mass and fluence of dust in the range 1-200 µm, where possible with multipoint measurements.<br>Comet Interceptor shall search for the existence of, and if found characterise, dust particles in the range centimetre – decimetre (chunks and snowballs).<br>Comet Interceptor shall determine the compositional diversity of the dust within the comet coma.<br>Comet Interceptor shall determine coma dust reflectance and polarimetric properties at visible wavelengths.</td></tr><tr><td>Comet Interceptor shall determine motion and evolution of ion rays and other coma and tail features including dust and gas.</td><td>Comet Interceptor shall produce maps of ion rays via wide field imaging of the coma<br>Comet Interceptor shall determine the existence (or not) of in-situ plasma signatures correlating to features observed through wide field imaging of coma.</td></tr><tr><td>Comet Interceptor shall characterise the plasma environment around the target, determine any resulting boundaries and assess energy, mass and momentum transfer.</td><td>Comet Interceptor shall determine the electron, ion, ENA, E and B field environment around the comet.<br>Comet Interceptor shall assess and quantify coupling between neutral, dust and plasma, including potential assessment of dust charging.<br>Comet Interceptor shall identify plasma structures and boundaries and characterise, where possible, with multipoint measurements.<br>Comet Interceptor shall characterise the plasma wave environment (λ≥2 km) to constrain energy and momentum transfer across the coma.</td></tr></tbody></table>

## 4 Planned Measurements and Payload Overview

The Comet Interceptor payload (Table 4) was selected to fulfil the mission Science Objectives.

The Comet Camera (CoCa) on Spacecraft A will contribute to both mission themes by determining the physical properties of the nucleus and coma with high resolution images. CoCa will determine the physical properties of the nucleus with high resolution images, tracking the nucleus through the flyby. The current CoCa instrument design will

- Determine the size, shape, and albedo of the nucleus

- Identify features on the nucleus at <20 metres resolution (assuming a nominal 1000 km flyby) and allow comparison with similar observations of the surfaces of short-period comets

- Establish the reflectivity spectral gradient of the nucleus and of sub-regions on the nucleus (Thomas and Keller 1989; Fornasier et al. 2015)
<table><caption>**Table 4** Summary of scientific instruments</caption><thead><tr><th>Instrument</th><th>Sensor</th><th>Application</th></tr></thead><tbody><tr><td colspan="3">Spacecraft A</td></tr><tr><td>CoCa</td><td></td><td>Visible-NIR hi-res imager</td></tr><tr><td rowspan="3">MIRMIS</td><td>NIR</td><td>NIR Hyperspectral imager</td></tr><tr><td>MIR</td><td>MIR Point spectrometer</td></tr><tr><td>TIRI</td><td>Multispectral thermal imager</td></tr><tr><td>MANiaC</td><td></td><td>Mass spectrometer</td></tr><tr><td rowspan="5">DFP</td><td>DISC</td><td>Dust detector</td></tr><tr><td>FGM-A</td><td>Magnetometer</td></tr><tr><td>COMPLIMENT</td><td>Electric field</td></tr><tr><td>SCIENA</td><td>Ion and energetic neutral atom detector</td></tr><tr><td>LEES</td><td>Electron spectrometer</td></tr><tr><td colspan="3">Probe B1</td></tr><tr><td>HI</td><td></td><td>Hydrogen imager</td></tr><tr><td>PS</td><td></td><td>Magnetometer</td></tr><tr><td></td><td></td><td>Ion mass spectrometer</td></tr><tr><td rowspan="2">NAC/WAC</td><td>NAC</td><td>Narrow-angle imager</td></tr><tr><td>WAC</td><td>Wide-angle imager</td></tr><tr><td colspan="3">Probe B2</td></tr><tr><td>EnVisS</td><td></td><td>All-sky imager</td></tr><tr><td>OPIC</td><td></td><td>Forward imager</td></tr><tr><td rowspan="2">DFP</td><td>FGM-B2</td><td>Magnetometer</td></tr><tr><td>DISC</td><td>Dust detector</td></tr></tbody></table>

- Search for evidence of surface ices for comparison with Rosetta and Deep Impact observations of SPCs (Sunshine et al. 2006; Pommerol et al. 2015)

- Constrain the rotation of the nucleus through, as a minimum, observation of dust coma structures

- Determine the spatial distribution of dust emission from the nucleus (including the day-side/nightside emission ratio, which may have implications for the presence of super-volatiles such as CO and CO₂, Gerig et al. 2020)

- Investigate and constrain the properties of the acceleration region of the inner coma by studying dust column density profiles (Gerig et al. 2018; Zakharov et al. 2018)

- Establish the direction and relative magnitudes of jet-like structures for correlation with dust impact events on the spacecraft

- Provide pointing information including determination of the flyby geometry (Curdt et al. 1988)

- Identify impact events on the spacecraft by providing evidence of uncontrolled Spacecraft Attitude changes (Curdt and Keller 1990).

The infrared spectra taken by the Modular InfraRed Molecules and Ices Sensor (MIRMIS) will provide information about the composition and thermal properties. They will be supported by the Narrow Angle Camera (NAC) on probe B1 and the Optical Periscope Imager for Comet (OPIC) on probe B2, which will image the nucleus from perspectives
<table><caption>**Table 4** Summary of scientific instruments</caption><thead><tr><th>Instrument</th><th>Sensor</th><th>Application</th></tr></thead><tbody><tr><td colspan="3">Spacecraft A</td></tr><tr><td>CoCa</td><td></td><td>Visible-NIR hi-res imager</td></tr><tr><td rowspan="3">MIRMIS</td><td>NIR</td><td>NIR Hyperspectral imager</td></tr><tr><td>MIR</td><td>MIR Point spectrometer</td></tr><tr><td>TIRI</td><td>Multispectral thermal imager</td></tr><tr><td>MANiaC</td><td></td><td>Mass spectrometer</td></tr><tr><td rowspan="5">DFP</td><td>DISC</td><td>Dust detector</td></tr><tr><td>FGM-A</td><td>Magnetometer</td></tr><tr><td>COMPLIMENT</td><td>Electric field</td></tr><tr><td>SCIENA</td><td>Ion and energetic neutral atom detector</td></tr><tr><td>LEES</td><td>Electron spectrometer</td></tr><tr><td colspan="3">Probe B1</td></tr><tr><td>HI</td><td></td><td>Hydrogen imager</td></tr><tr><td>PS</td><td></td><td>Magnetometer</td></tr><tr><td></td><td></td><td>Ion mass spectrometer</td></tr><tr><td rowspan="2">NAC/WAC</td><td>NAC</td><td>Narrow-angle imager</td></tr><tr><td>WAC</td><td>Wide-angle imager</td></tr><tr><td colspan="3">Probe B2</td></tr><tr><td>EnVisS</td><td></td><td>All-sky imager</td></tr><tr><td>OPIC</td><td></td><td>Forward imager</td></tr><tr><td rowspan="2">DFP</td><td>FGM-B2</td><td>Magnetometer</td></tr><tr><td>DISC</td><td>Dust detector</td></tr></tbody></table>

- Search for evidence of surface ices for comparison with Rosetta and Deep Impact observations of SPCs (Sunshine et al. 2006; Pommerol et al. 2015)

- Constrain the rotation of the nucleus through, as a minimum, observation of dust coma structures

- Determine the spatial distribution of dust emission from the nucleus (including the day-side/nightside emission ratio, which may have implications for the presence of super-volatiles such as CO and CO₂, Gerig et al. 2020)

- Investigate and constrain the properties of the acceleration region of the inner coma by studying dust column density profiles (Gerig et al. 2018; Zakharov et al. 2018)

- Establish the direction and relative magnitudes of jet-like structures for correlation with dust impact events on the spacecraft

- Provide pointing information including determination of the flyby geometry (Curdt et al. 1988)

- Identify impact events on the spacecraft by providing evidence of uncontrolled Spacecraft Attitude changes (Curdt and Keller 1990).

The infrared spectra taken by the Modular InfraRed Molecules and Ices Sensor (MIRMIS) will provide information about the composition and thermal properties. They will be supported by the Narrow Angle Camera (NAC) on probe B1 and the Optical Periscope Imager for Comet (OPIC) on probe B2, which will image the nucleus from perspectives
different to those of Spacecraft A, providing stereo views and increasing the fraction of the surface that can be investigated. Probe B1 will also have a wide-angle visible camera (WAC) to image the nucleus and its surroundings from a fixed orientation during closest approach, and the OPIC instrument, on probe B2, will return resolved images of the nucleus and the inner coma from a different angle shortly before closest approach. Composition information will come from multi-colour imaging (four broadband filters in the visible range in CoCa; a tuneable hyperspectral imager in the 0.9 to 1.7 µm near-infrared region and fixed narrowband filters between 8.9 and 21.6 µm in MIRMIS) and point spectroscopy in the 2.5 to 5 µm range (MIRMIS/MIR). Unique insights into the physical properties of the surface layers (particle sizes, thermal inertia) will come from visible wavelength polarimetry measurements by the Entire Visible Sky instrument, EnVisS, on B2, and thermal-infrared temperature measurements (MIRMIS).

For the Comet Environment Theme, the coma composition will be measured *in situ* by the Mass Analyzer for Neutrals in a Coma (MANiaC), including measurement of isotopic ratios. Provided that the outgassing activity of the comet is sufficiently high, elemental, molecular, and isotope abundances of the gas will be derived by MANiaC. This includes monitoring the major coma volatiles H₂O, CO, and CO₂, key molecules such as O₂, highly volatile species, and organic compounds, as well as hydrogen and oxygen isotopes in water. Even (icy) grains entering the instrument may be measured. MANiaC provides key measurements to study the comet's activity and the material from which its nucleus formed and how it evolved. Gas composition measurements through remote sensing will be performed by MIRMIS. In addition, the Hydrogen Imager (HI) on probe B1 will monitor the cometary water production rate from months before the flyby until the end of operations. The spatial distribution of dust will be investigated from three different viewpoints by CoCa on Spacecraft A, the WAC on B1, and OPIC and the all-sky imager EnVisS on probe B2. EnVisS will additionally constrain dust properties through polarimetric measurements, and the Dust Impact Sensor and Counter (DISC), part of the Dust, Fields, and Plasma (DFP) instrument, will measure the mass distribution of dust particles colliding with both Spacecraft A and probe B2. DFP's COMPLIMENT and LEES sensors will detect nanograin impacts.

The Dust Field and Plasma (DFP) package is a combined experiment dedicated to the multi-point *in situ* study of the multi-phased ionized and dusty environment in the comet's coma, and its interaction with the surrounding space environment. DFP will measure the magnetic field, electric field, plasma parameters (density, temperature, and speed), the distribution functions of electrons, ions, and energetic neutrals, spacecraft potential, and cometary dust, in order to:

- Identify boundaries and regions in the cometary environment of a comet and its interaction with the solar wind (e.g., bow shock, diamagnetic cavity) and to assess their structure.

- Map the dust and plasma phases around the target.

- Assess mass, momentum, and energy transfer in the cometary environment.

- Provide simultaneous magnetic field, plasma, and dust measurements to identify the interplay between the ionized and dusty phases around a comet and characterize dusty plasma properties.

- Map the solar wind – coma interaction.

- Describe and map the (i) electron, (ii) negative and positive ion, and (iii) energetic neutral atom distribution functions in the vicinity of the comet and in the interaction region with the solar wind.

- Identify electron and ion kinetic processes that mediate the solar wind-comet interactions from ion kinetic scales down to electron scales.
different to those of Spacecraft A, providing stereo views and increasing the fraction of the surface that can be investigated. Probe B1 will also have a wide-angle visible camera (WAC) to image the nucleus and its surroundings from a fixed orientation during closest approach, and the OPIC instrument, on probe B2, will return resolved images of the nucleus and the inner coma from a different angle shortly before closest approach. Composition information will come from multi-colour imaging (four broadband filters in the visible range in CoCa; a tuneable hyperspectral imager in the 0.9 to 1.7 µm near-infrared region and fixed narrowband filters between 8.9 and 21.6 µm in MIRMIS) and point spectroscopy in the 2.5 to 5 µm range (MIRMIS/MIR). Unique insights into the physical properties of the surface layers (particle sizes, thermal inertia) will come from visible wavelength polarimetry measurements by the Entire Visible Sky instrument, EnVisS, on B2, and thermal-infrared temperature measurements (MIRMIS).

For the Comet Environment Theme, the coma composition will be measured *in situ* by the Mass Analyzer for Neutrals in a Coma (MANiaC), including measurement of isotopic ratios. Provided that the outgassing activity of the comet is sufficiently high, elemental, molecular, and isotope abundances of the gas will be derived by MANiaC. This includes monitoring the major coma volatiles H₂O, CO, and CO₂, key molecules such as O₂, highly volatile species, and organic compounds, as well as hydrogen and oxygen isotopes in water. Even (icy) grains entering the instrument may be measured. MANiaC provides key measurements to study the comet's activity and the material from which its nucleus formed and how it evolved. Gas composition measurements through remote sensing will be performed by MIRMIS. In addition, the Hydrogen Imager (HI) on probe B1 will monitor the cometary water production rate from months before the flyby until the end of operations. The spatial distribution of dust will be investigated from three different viewpoints by CoCa on Spacecraft A, the WAC on B1, and OPIC and the all-sky imager EnVisS on probe B2. EnVisS will additionally constrain dust properties through polarimetric measurements, and the Dust Impact Sensor and Counter (DISC), part of the Dust, Fields, and Plasma (DFP) instrument, will measure the mass distribution of dust particles colliding with both Spacecraft A and probe B2. DFP's COMPLIMENT and LEES sensors will detect nanograin impacts.

The Dust Field and Plasma (DFP) package is a combined experiment dedicated to the multi-point *in situ* study of the multi-phased ionized and dusty environment in the comet's coma, and its interaction with the surrounding space environment. DFP will measure the magnetic field, electric field, plasma parameters (density, temperature, and speed), the distribution functions of electrons, ions, and energetic neutrals, spacecraft potential, and cometary dust, in order to:

- Identify boundaries and regions in the cometary environment of a comet and its interaction with the solar wind (e.g., bow shock, diamagnetic cavity) and to assess their structure.

- Map the dust and plasma phases around the target.

- Assess mass, momentum, and energy transfer in the cometary environment.

- Provide simultaneous magnetic field, plasma, and dust measurements to identify the interplay between the ionized and dusty phases around a comet and characterize dusty plasma properties.

- Map the solar wind – coma interaction.

- Describe and map the (i) electron, (ii) negative and positive ion, and (iii) energetic neutral atom distribution functions in the vicinity of the comet and in the interaction region with the solar wind.

- Identify electron and ion kinetic processes that mediate the solar wind-comet interactions from ion kinetic scales down to electron scales.
DFP-SCIENA will measure energetic particles of solar wind and cometary origin, both with and without charge. The ion observational capabilities will allow the direct detection of solar wind and cometary ions, providing energy, direction, and estimated mass. These measurements are necessary to determine the 3D flow of plasma in the comet's induced magnetosphere in order to assess the mass, momentum, and energy flow and transfer between different plasma regions, identification of plasma boundaries such as the bow shock, solar wind void and similar. The energetic neutral atom (ENA) measurement capability allows the study of the direct interaction of the solar wind with the neutral atmosphere, providing continuous monitoring of the solar wind's dynamic pressure, an estimate of the position of the regions of strongest interaction between the solar wind and the coma as well as the coupling between the coma and the cometary ions.

Magnetometers are the only *in situ* instrument type present on all three spacecraft, though all three have different designs and instrument heritage. They will also detect various wave modes, such as mirror modes and ion cyclotron waves, to assess the energy transfer across these boundaries. Plasma properties will be derived from measurements of three spectrometers as part of DFP and an ion mass spectrometer as part of the B1 Plasma Suite (PS).

The following section provides more details of the mission's scientific instruments.

# 5 Scientific Instruments

## 5.1 Comet CCamera (CoCa)

The camera system on-board Spacecraft A, CoCa, is required to provide detailed imaging of the nucleus and the innermost coma of the target. The design uses previous heritage to establish a baseline performance (Table 5) that surpasses that of previous flyby missions to comets. The instrument is based upon two elements. Firstly, it uses the telescope of the Colour and Stereo Surface Imaging System (CaSSIS) that is successfully operating at Mars on the European Space Agency's ExoMars Trace Gas Orbiter (TGO) (Thomas et al. 2017). Secondly, the CoCa design uses the detector system of the JANUS instrument from ESA's JUICE mission (Witasse 2021). By integrating these two elements, CoCa can achieve an angular scale of 8 µrad px⁻¹, superior to cameras on all previous comet missions with the exception of Deep Impact's HRI (A'Hearn et al. 2011). The scale is nearly a factor of three superior to that of the Halley Multicolour Camera, HMC, on-board Giotto (Keller et al. 1987). The detector system uses a rolling shutter technique to allow rapid image read-out with a minimum possible exposure time of 220 µs to avoid motion smear at closest approach for even the highest velocity flybys. A major difference here is that, unlike Giotto which was a spinning spacecraft, Comet Interceptor is a three-axis stabilised system implying that the exposure times can be selected. The detector allows saturation of the nucleus without blooming of charge. This, in turn, implies that the exposure times of selected images can be programmed to provide high signal to noise observations of the dust coma while saturating on the nucleus. This capability will increase the flexibility of the mission if targets are eventually found that have only weak dust emission.

CoCa will be equipped with four selectable interference filters covering the sensitivity range of the detector (roughly 400 nm to 1000 nm). The filters will be around 150 nm in bandwidth and optimised to have approximately equal signal to noise ratio in all images of the nucleus. Low spectral resolution is not a disadvantage here as previous observations have shown that visible spectra are mostly featureless with a constant reflectivity gradient (e.g. Fornasier et al. 2015). Subtle broadband colour differences were seen by the Rosetta camera
DFP-SCIENA will measure energetic particles of solar wind and cometary origin, both with and without charge. The ion observational capabilities will allow the direct detection of solar wind and cometary ions, providing energy, direction, and estimated mass. These measurements are necessary to determine the 3D flow of plasma in the comet's induced magnetosphere in order to assess the mass, momentum, and energy flow and transfer between different plasma regions, identification of plasma boundaries such as the bow shock, solar wind void and similar. The energetic neutral atom (ENA) measurement capability allows the study of the direct interaction of the solar wind with the neutral atmosphere, providing continuous monitoring of the solar wind's dynamic pressure, an estimate of the position of the regions of strongest interaction between the solar wind and the coma as well as the coupling between the coma and the cometary ions.

Magnetometers are the only *in situ* instrument type present on all three spacecraft, though all three have different designs and instrument heritage. They will also detect various wave modes, such as mirror modes and ion cyclotron waves, to assess the energy transfer across these boundaries. Plasma properties will be derived from measurements of three spectrometers as part of DFP and an ion mass spectrometer as part of the B1 Plasma Suite (PS).

The following section provides more details of the mission's scientific instruments.

# 5 Scientific Instruments

## 5.1 Comet CCamera (CoCa)

The camera system on-board Spacecraft A, CoCa, is required to provide detailed imaging of the nucleus and the innermost coma of the target. The design uses previous heritage to establish a baseline performance (Table 5) that surpasses that of previous flyby missions to comets. The instrument is based upon two elements. Firstly, it uses the telescope of the Colour and Stereo Surface Imaging System (CaSSIS) that is successfully operating at Mars on the European Space Agency's ExoMars Trace Gas Orbiter (TGO) (Thomas et al. 2017). Secondly, the CoCa design uses the detector system of the JANUS instrument from ESA's JUICE mission (Witasse 2021). By integrating these two elements, CoCa can achieve an angular scale of 8 µrad px⁻¹, superior to cameras on all previous comet missions with the exception of Deep Impact's HRI (A'Hearn et al. 2011). The scale is nearly a factor of three superior to that of the Halley Multicolour Camera, HMC, on-board Giotto (Keller et al. 1987). The detector system uses a rolling shutter technique to allow rapid image read-out with a minimum possible exposure time of 220 µs to avoid motion smear at closest approach for even the highest velocity flybys. A major difference here is that, unlike Giotto which was a spinning spacecraft, Comet Interceptor is a three-axis stabilised system implying that the exposure times can be selected. The detector allows saturation of the nucleus without blooming of charge. This, in turn, implies that the exposure times of selected images can be programmed to provide high signal to noise observations of the dust coma while saturating on the nucleus. This capability will increase the flexibility of the mission if targets are eventually found that have only weak dust emission.

CoCa will be equipped with four selectable interference filters covering the sensitivity range of the detector (roughly 400 nm to 1000 nm). The filters will be around 150 nm in bandwidth and optimised to have approximately equal signal to noise ratio in all images of the nucleus. Low spectral resolution is not a disadvantage here as previous observations have shown that visible spectra are mostly featureless with a constant reflectivity gradient (e.g. Fornasier et al. 2015). Subtle broadband colour differences were seen by the Rosetta camera
<table><caption>**Table 5** The main characteristics of the CoCa instrument</caption><thead><tr><th>Parameter</th><th>Value</th></tr></thead><tbody><tr><td>Instrument pixel scale</td><td>8 µrad/px</td></tr><tr><td>Field of View</td><td>0.69° × 0.92°</td></tr><tr><td>Detector</td><td>CIS115 Back-side illuminated<br>CMOS image sensor</td></tr><tr><td>Pixels</td><td>1504 × 2000</td></tr><tr><td>Pixel size</td><td>7 × 7 µm</td></tr><tr><td>Exposure times</td><td>220 µs (flyby) to<br>15 min (identification),<br>rolling shutter</td></tr><tr><td>Imaging rate multi-colour</td><td>≥1 frame s<sup>-1</sup></td></tr><tr><td>Imaging rate single colour</td><td>≥2 frames s<sup>-1</sup></td></tr><tr><td>Filters</td><td>475 nm (Δλ = 150 nm) BLU<br>675 nm (Δλ = 100 nm) ORG<br>775 nm (Δλ = 100 nm) RED<br>900 nm (Δλ = 150 nm) NIR</td></tr><tr><td>Mass</td><td>13.5 kg (3 units)</td></tr><tr><td>Power</td><td>19 W average</td></tr><tr><td>Volume</td><td>CSU: 350 × 460 × 550 mm<sup>3</sup><br>PEU: 210 × 160 × 70 mm<sup>3</sup><br>ELU: 120 × 240 × 180 mm<sup>3</sup></td></tr><tr><td>Data I/F</td><td>Spacewire</td></tr><tr><td>Instrument memory<br>(holding science data)</td><td>2 × 128 Gbit</td></tr><tr><td>Max data volume</td><td>128 Gbit uncompressed</td></tr></tbody></table>

OSIRIS (Fornasier et al. 2016) requiring high signal to noise. A filter wheel mechanism has been designed to switch filters quickly to ensure minimum changes in phase angle between adjacent images. A goal of 1 second between images has been set and achieved in the prototype giving a change in phase angle for adjacent observations of <1°. The passage through closest approach and imaging at high cadence will also provide a data set for establishing the 3D spatial distribution of dust in the inner coma using post-processing tomography techniques. The instrument is capable of acquiring around 2500 images during the encounter. Data compression (using a JPEG algorithm used extensively on CaSSIS) and sub-framing are foreseen to ensure that the on-board storage produces no practical limit to the flexibility of the data acquisition. It is intended that the instrument will be set into a “mode” by command and will continue imaging in this mode until there is a mode change or until CoCa is told to stop. CoCa therefore operates autonomously with pre-defined modes that can be programmed in flight.

The CoCa design is shown in Fig. 16. The Camera Support Unit (CSU) has numerous elements. The open structure in Fig. 16 is a 13.5 cm diameter 4-mirror off-axis telescope with a focal length of 880 mm following the design for the CaSSIS telescope on TGO. The structure is carbon-fibre reinforced plastic (CFRP). Following the CaSSIS experience, a small change to the internal baffling at the intermediate stop (between mirror M2 and M3) has been made and improved front baffle has been designed. Otherwise, the design is unchanged. The mirrors are silver coated, providing a field of view that is larger than the active area of the detector in the focal plane. The open structure ensures low mass and
<table><caption>**Table 5** The main characteristics of the CoCa instrument</caption><thead><tr><th>Parameter</th><th>Value</th></tr></thead><tbody><tr><td>Instrument pixel scale</td><td>8 µrad/px</td></tr><tr><td>Field of View</td><td>0.69° × 0.92°</td></tr><tr><td>Detector</td><td>CIS115 Back-side illuminated<br>CMOS image sensor</td></tr><tr><td>Pixels</td><td>1504 × 2000</td></tr><tr><td>Pixel size</td><td>7 × 7 µm</td></tr><tr><td>Exposure times</td><td>220 µs (flyby) to<br>15 min (identification),<br>rolling shutter</td></tr><tr><td>Imaging rate multi-colour</td><td>≥1 frame s<sup>-1</sup></td></tr><tr><td>Imaging rate single colour</td><td>≥2 frames s<sup>-1</sup></td></tr><tr><td>Filters</td><td>475 nm (Δλ = 150 nm) BLU<br>675 nm (Δλ = 100 nm) ORG<br>775 nm (Δλ = 100 nm) RED<br>900 nm (Δλ = 150 nm) NIR</td></tr><tr><td>Mass</td><td>13.5 kg (3 units)</td></tr><tr><td>Power</td><td>19 W average</td></tr><tr><td>Volume</td><td>CSU: 350 × 460 × 550 mm<sup>3</sup><br>PEU: 210 × 160 × 70 mm<sup>3</sup><br>ELU: 120 × 240 × 180 mm<sup>3</sup></td></tr><tr><td>Data I/F</td><td>Spacewire</td></tr><tr><td>Instrument memory<br>(holding science data)</td><td>2 × 128 Gbit</td></tr><tr><td>Max data volume</td><td>128 Gbit uncompressed</td></tr></tbody></table>

OSIRIS (Fornasier et al. 2016) requiring high signal to noise. A filter wheel mechanism has been designed to switch filters quickly to ensure minimum changes in phase angle between adjacent images. A goal of 1 second between images has been set and achieved in the prototype giving a change in phase angle for adjacent observations of <1°. The passage through closest approach and imaging at high cadence will also provide a data set for establishing the 3D spatial distribution of dust in the inner coma using post-processing tomography techniques. The instrument is capable of acquiring around 2500 images during the encounter. Data compression (using a JPEG algorithm used extensively on CaSSIS) and sub-framing are foreseen to ensure that the on-board storage produces no practical limit to the flexibility of the data acquisition. It is intended that the instrument will be set into a “mode” by command and will continue imaging in this mode until there is a mode change or until CoCa is told to stop. CoCa therefore operates autonomously with pre-defined modes that can be programmed in flight.

The CoCa design is shown in Fig. 16. The Camera Support Unit (CSU) has numerous elements. The open structure in Fig. 16 is a 13.5 cm diameter 4-mirror off-axis telescope with a focal length of 880 mm following the design for the CaSSIS telescope on TGO. The structure is carbon-fibre reinforced plastic (CFRP). Following the CaSSIS experience, a small change to the internal baffling at the intermediate stop (between mirror M2 and M3) has been made and improved front baffle has been designed. Otherwise, the design is unchanged. The mirrors are silver coated, providing a field of view that is larger than the active area of the detector in the focal plane. The open structure ensures low mass and
<figure><img src="image_18.png" /><figcaption>**Fig. 17** The CoCa filter wheel assembly with four filters. The detector is shown in pink below one of the filters. The filter wheel includes a launch lock to prevent motion during launch.</figcaption></figure>

<figure><img src="image_19.png" /></figure>

will finally be wrapped in multi-layer insulation to produce a light-tight unit. The telescope is mounted on a baseplate that also supports the detector and filter wheel assembly. The detector is a spare of the development for JANUS, the imaging system on JUICE. The sensor is a back-side illuminated (BSI) CMOS device from e2v with 1504 × 2000 pixels and 7 µm pitch, a peak quantum efficiency exceeding 90% (Crews et al. 2020) and a full-well of 27,000 e⁻. A radiator will be used to reduce the sensor temperature, with -30 °C being the goal, although nominal operation can be achieved at 0 °C. Combined with the telescope system, the detector provides a field of view of 0.69° × 0.92°. The filters (Fig. 17) will use fused silica substrates with standard interference coatings, designed for high throughput with sharp cut-offs and high out-of-band rejection. The PEU houses the proximity electronics for the detector while the ELU provides power conversion, instrument control, and data management.

Considerable effort has been invested in protecting CoCa from hyper-velocity dust impacts during the flyby. It is to be recalled that HMC was damaged severely during the 1P/Halley encounter despite being mostly behind the Whipple shield of the spacecraft (Schwehm 1991). In the case of Comet Interceptor, a rotating mirror assembly (RMA) has been developed which will allow CoCa to be mounted behind the protection shield while still providing a continuous view of the nucleus. The RMA has two elements (Fig. 18): the SMA (Scan Mirror Assembly), and the SME (Scan Mirror Electronics). The SMA is a mechanism holding the folding mirror and that will rotate this mirror in order to orient the field of view of CoCa towards the comet during encounter. It is based on a brushless DC motor moving the mirror via a gear system and an optical position sensor in order to allow closed loop control. The mechanism will be driven by the SME that will take care of powering the motor to position the folding mirror based on encounter parameters provided by the spacecraft platform.
<figure><img src="image_19.png" /><figcaption>**Fig. 17** The CoCa filter wheel assembly with four filters. The detector is shown in pink below one of the filters. The filter wheel includes a launch lock to prevent motion during launch.</figcaption></figure>

<figure><img src="image_20.png" /></figure>

will finally be wrapped in multi-layer insulation to produce a light-tight unit. The telescope is mounted on a baseplate that also supports the detector and filter wheel assembly. The detector is a spare of the development for JANUS, the imaging system on JUICE. The sensor is a back-side illuminated (BSI) CMOS device from e2v with 1504 × 2000 pixels and 7 µm pitch, a peak quantum efficiency exceeding 90% (Crews et al. 2020) and a full-well of 27,000 e⁻. A radiator will be used to reduce the sensor temperature, with –30 °C being the goal, although nominal operation can be achieved at 0 °C. Combined with the telescope system, the detector provides a field of view of 0.69° × 0.92°. The filters (Fig. 17) will use fused silica substrates with standard interference coatings, designed for high throughput with sharp cut-offs and high out-of-band rejection. The PEU houses the proximity electronics for the detector while the ELU provides power conversion, instrument control, and data management.

Considerable effort has been invested in protecting CoCa from hyper-velocity dust impacts during the flyby. It is to be recalled that HMC was damaged severely during the 1P/Halley encounter despite being mostly behind the Whipple shield of the spacecraft (Schwehm 1991). In the case of Comet Interceptor, a rotating mirror assembly (RMA) has been developed which will allow CoCa to be mounted behind the protection shield while still providing a continuous view of the nucleus. The RMA has two elements (Fig. 18): the SMA (Scan Mirror Assembly), and the SME (Scan Mirror Electronics). The SMA is a mechanism holding the folding mirror and that will rotate this mirror in order to orient the field of view of CoCa towards the comet during encounter. It is based on a brushless DC motor moving the mirror via a gear system and an optical position sensor in order to allow closed loop control. The mechanism will be driven by the SME that will take care of powering the motor to position the folding mirror based on encounter parameters provided by the spacecraft platform.
<figure><img src="image_20.png" /><figcaption>**Fig. 18** CAD/CAM drawing of the RMA showing the opening (top) and the fold mirror mount (turquoise colour) which rotates and reflects light toward CoCa (left). The mounting feet (yellow colour) attach the RMA to the exterior of the spacecraft while the entire CoCa instrument (Fig. 16) is inside with their optical axes aligned</figcaption></figure>

combined with the read-out of the position sensor. The SMA includes a protection system that will hide the mirror from incoming dust particles during the most critical part of the encounter, when the spacecraft is closest to the nucleus.

CoCa is being built by an international consortium formed by the Instituto de Astrofísica de Andalucía (IAA Granada, Spain), the Research Center for Astronomy and Earth Science (CSFK, Budapest, Hungary), the Laboratoire d'Astrophysique de Marseille (LAM, France), and the DLR Institute for Planetary Research (Berlin, Germany) under the lead of the University of Bern (Switzerland). The University of Liège (Belgium) leads the development of the Rotating Mirror Assembly (RMA).

## 5.2 Modular InfraRed Molecules and Ices Sensor (MIRMIS)

MIRMIS is the hyper and multi spectral remote sensing instrument for Comet Interceptor (Fig. 19, Table 6). The instrument covers a wavelength range of 0.9 to 25 µm which samples spectral features of CO₂, H₂O, CO, mineral compositions and is rich in thermophysical data. The instrument is made up of three closely integrated modules: the Near-IR (NIR)/Mid-IR (MIR) spectrometer and a thermal infrared imager (TIRI), with one thermal, mechanical, and electrical interface to the spacecraft. NIR is a 0.9 to 1.7 µm hyperspectral imager, MIR is a 2.5 to 5 µm point spectrometer and TIRI is a 6 to 25 µm multispectral thermal imager.

The MIR and NIR modules are based on tunable Fabry-Pérot interferometers which are used as an adjustable bandpass filter (Näsilä and Kohout 2020). For the NIR channel the filter is combined with an InGaAs focal plane assembly to give a field of view (FoV) of ca. 6.7 × 5.4° from 640 × 512 pixels and spectral bandwidth of 20 nm. The NIR design is based on refractive optics with heritage from the ASPECT imager in the Milani CubeSat for ESA's
<figure><img src="image_21.png" /><figcaption>**Fig. 18** CAD/CAM drawing of the RMA showing the opening (top) and the fold mirror mount (turquoise colour) which rotates and reflects light toward CoCa (left). The mounting feet (yellow colour) attach the RMA to the exterior of the spacecraft while the entire CoCa instrument (Fig. 16) is inside with their optical axes aligned</figcaption></figure>

combined with the read-out of the position sensor. The SMA includes a protection system that will hide the mirror from incoming dust particles during the most critical part of the encounter, when the spacecraft is closest to the nucleus.

CoCa is being built by an international consortium formed by the Instituto de Astrofísica de Andalucía (IAA Granada, Spain), the Research Center for Astronomy and Earth Science (CSFK, Budapest, Hungary), the Laboratoire d'Astrophysique de Marseille (LAM, France), and the DLR Institute for Planetary Research (Berlin, Germany) under the lead of the University of Bern (Switzerland). The University of Liège (Belgium) leads the development of the Rotating Mirror Assembly (RMA).

## 5.2 Modular InfraRed Molecules and Ices Sensor (MIRMIS)

MIRMIS is the hyper and multi spectral remote sensing instrument for Comet Interceptor (Fig. 19, Table 6). The instrument covers a wavelength range of 0.9 to 25 µm which samples spectral features of CO₂, H₂O, CO, mineral compositions and is rich in thermophysical data. The instrument is made up of three closely integrated modules: the Near-IR (NIR)/Mid-IR (MIR) spectrometer and a thermal infrared imager (TIRI), with one thermal, mechanical, and electrical interface to the spacecraft. NIR is a 0.9 to 1.7 µm hyperspectral imager, MIR is a 2.5 to 5 µm point spectrometer and TIRI is a 6 to 25 µm multispectral thermal imager.

The MIR and NIR modules are based on tunable Fabry-Pérot interferometers which are used as an adjustable bandpass filter (Näsilä and Kohout 2020). For the NIR channel the filter is combined with an InGaAs focal plane assembly to give a field of view (FoV) of ca. 6.7 × 5.4° from 640 × 512 pixels and spectral bandwidth of 20 nm. The NIR design is based on refractive optics with heritage from the ASPECT imager in the Milani CubeSat for ESA's
<figure><img src="image_21.png" /><figcaption>**Table 6** MIRMIS instrument summary table</figcaption></figure>

<table><tr><td>Scientific performance summary</td></tr><tr><td>Spectral range 0.9–25 µm using four channels:<br>NIR, hyperspectral imager, 0.9–1.7 µm<br>@20 nm spectral bandwidth<br>MIR, single-point spectrometer, 2.5–5.0 µm<br>@30 nm spectral bandwidth<br>TIRI, multispectral thermal imager, 6–25 µm</td></tr><tr><td>Key instrument numbers</td></tr><tr><td>TIRI: FoV = 9 × 7° (7 µm diffraction limit), iFoV = 0.26 mrad<br>NIR: FoV = 6.7 × 5.4°, iFoV = 0.18 mrad<br>MIR: FoV = 2° circular<br>Total Mass (CBE): 8.8 kg with margin<br>Standby average: 8.3 W<br>Standby average with detector thermal control: 9.9 W<br>Average science operating mode (Nucleus pointing): 11.2 W<br>Average science operating mode (Coma monitoring): 9.7 W<br>Total module volume: 548.5 mm × 282.0 mm × 126.8 mm</td></tr></table>

Hera mission (Kohout et al. 2018; Michel et al. 2022). A prototype version has been flying in low Earth orbit since November 2018 on the Reaktor Hello World nanosatellite. The basic operating principle of the MIR channel is the same as NIR but has a single HgCdTe detector with a ca. 2° FoV to provide point spectroscopy and spectral bandwidth of 30 nm.

The TIRI module is based on the Lunar Thermal Mapper instrument (Bowles et al. 2020) due for launch on NASA's Lunar Trailblazer mission in 2023 with heritage from the Compact Modular Sounder (CMS) instrument flown in low Earth orbit between 2014–2017 on the UK TechDemoSat-1 spacecraft. TIRI uses gold-coated all-reflective optics with f/1.4 and a 50-mm aperture. A pointing mirror is used to direct the field of view onto the target object, a black body calibration target or a space view. Calibrations using the black body and space view will be performed immediately before and after each observation sequence. A two-mirror telescope directs the incoming infrared radiation onto a filter assembly used to define ten individual spectral channels (Shirley et al. 2022) for compositional and temperature mapping of the nucleus. This filter assembly is re-imaged onto a 640 × 480 uncooled microbolometer array using a three-mirror relay.
<figure><img src="image_22.png" /><figcaption>**Table 6** MIRMIS instrument summary table</figcaption></figure>

<table><tr><td>Scientific performance summary</td></tr><tr><td>Spectral range 0.9–25 µm using four channels:<br>NIR, hyperspectral imager, 0.9–1.7 µm<br>@20 nm spectral bandwidth<br>MIR, single-point spectrometer, 2.5–5.0 µm<br>@30 nm spectral bandwidth<br>TIRI, multispectral thermal imager, 6–25 µm</td></tr><tr><td>Key instrument numbers</td></tr><tr><td>TIRI: FoV = 9 × 7° (7 µm diffraction limit), iFoV = 0.26 mrad<br>NIR: FoV = 6.7 × 5.4°, iFoV = 0.18 mrad<br>MIR: FoV = 2° circular<br>Total Mass (CBE): 8.8 kg with margin<br>Standby average: 8.3 W<br>Standby average with detector thermal control: 9.9 W<br>Average science operating mode (Nucleus pointing): 11.2 W<br>Average science operating mode (Coma monitoring): 9.7 W<br>Total module volume: 548.5 mm × 282.0 mm × 126.8 mm</td></tr></table>

Hera mission (Kohout et al. 2018; Michel et al. 2022). A prototype version has been flying in low Earth orbit since November 2018 on the Reaktor Hello World nanosatellite. The basic operating principle of the MIR channel is the same as NIR but has a single HgCdTe detector with a ca. 2° FoV to provide point spectroscopy and spectral bandwidth of 30 nm.

The TIRI module is based on the Lunar Thermal Mapper instrument (Bowles et al. 2020) due for launch on NASA's Lunar Trailblazer mission in 2023 with heritage from the Compact Modular Sounder (CMS) instrument flown in low Earth orbit between 2014–2017 on the UK TechDemoSat-1 spacecraft. TIRI uses gold-coated all-reflective optics with f/1.4 and a 50-mm aperture. A pointing mirror is used to direct the field of view onto the target object, a black body calibration target or a space view. Calibrations using the black body and space view will be performed immediately before and after each observation sequence. A two-mirror telescope directs the incoming infrared radiation onto a filter assembly used to define ten individual spectral channels (Shirley et al. 2022) for compositional and temperature mapping of the nucleus. This filter assembly is re-imaged onto a 640 × 480 uncooled microbolometer array using a three-mirror relay.
**Fig. 20** MANiaC consisting of a time-of-flight mass spectrometer (SHU, Sensor Head Unit), the Neutral Density Gauge (NDG), and the ELectronic Unit (ELU). For reference the long axis of the SHU corresponds to ~470 mm. Only the antechamber spheres of both the NDG and the SHU (marked yellow) are exposed to the gas and dust flow of the coma and are covered by dedicated dust shields. The rest is enclosed and protected inside the spacecraft

<figure><img src="image_22.png" /></figure>

MIRMIS/NIR and MIRMIS/MIR are supplied by VTT Finland with the University of Oxford, UK providing MIRMIS/TIRI and overall integration of the instrument. The MIRMIS instrument team includes members from the UK, Finland, and the USA.

## 5.3 Mass Analyzer for Neutrals in a Coma (MANiaC)

MANiaC is dedicated to the *in situ* measurement of the neutral gas coma. MANiaC consists of two instruments, a time-of-flight mass spectrometer and a neutral density gauge (Fig. 20). The mass spectrometer obtains the relative abundances of the major and a subset of minor volatiles. The neutral density gauge measures the total gas density. Combining the measurements of both instruments yields the absolute densities of a suite of volatile species along the flyby trajectory.

Since the flyby velocity range of 10–70 km/s will be much larger than the neutral gas speed (~1 km/s) MANiaC will be mounted on the spacecraft such that the aperture is always pointing in the direction of relative motion of the spacecraft. To cope with the large range of possible flyby velocities, both the SHU and the NDG contain antechambers for the thermalization of the incoming gas. Afterwards, the neutral gas entering the ion source will be ionized by impacting 70 eV electrons emitted by a hot filament. In the NDG, the resulting ions are measured as a current by a sensitive electrometer and in proportion to the gas density inside the antechamber and hence the surrounding coma. In the SHU, the newly formed ions are accelerated by a sharp extraction voltage pulse into the drift section. After passing the reflectron, i.e., an opposing electric field, the ions cross the drift section again before impinging on the Micro Channel Plate detector. Since the voltage pulse provides the same energy to all extracted ions, their arrival time on the detector can be converted into a mass/charge ratio. Combining NDG and SHU measurements leads to *in situ* abundances of coma gases.

MANiaC is being built by an international consortium formed by the Instituto de Astrofísica de Andalucía (IAA Granada, Spain), the Institut für Weltraumforschung (IWF Graz, Austria), the Institut de Recherche en Astrophysique et Planétologie (IRAP Toulouse, France), and Creotech Instruments S. A. (Piaseczno, Poland) under the lead of the University of Bern (Switzerland).

## 5.4 Dust, Fields, and Plasma (DFP-A)

To enable multipoint *in situ* measurements of dust, plasma, and energetic atoms, five instrumental sensors, common data processing unit DAPU, and power supply system, PSU, will
**Fig. 20** MANiaC consisting of a time-of-flight mass spectrometer (SHU, Sensor Head Unit), the Neutral Density Gauge (NDG), and the ELectronic Unit (ELU). For reference the long axis of the SHU corresponds to ~470 mm. Only the antechamber spheres of both the NDG and the SHU (marked yellow) are exposed to the gas and dust flow of the coma and are covered by dedicated dust shields. The rest is enclosed and protected inside the spacecraft

<figure><img src="image_23.png" /></figure>

MIRMIS/NIR and MIRMIS/MIR are supplied by VTT Finland with the University of Oxford, UK providing MIRMIS/TIRI and overall integration of the instrument. The MIRMIS instrument team includes members from the UK, Finland, and the USA.

## 5.3 Mass Analyzer for Neutrals in a Coma (MANiaC)

MANiaC is dedicated to the *in situ* measurement of the neutral gas coma. MANiaC consists of two instruments, a time-of-flight mass spectrometer and a neutral density gauge (Fig. 20). The mass spectrometer obtains the relative abundances of the major and a subset of minor volatiles. The neutral density gauge measures the total gas density. Combining the measurements of both instruments yields the absolute densities of a suite of volatile species along the flyby trajectory.

Since the flyby velocity range of 10–70 km/s will be much larger than the neutral gas speed (~1 km/s) MANiaC will be mounted on the spacecraft such that the aperture is always pointing in the direction of relative motion of the spacecraft. To cope with the large range of possible flyby velocities, both the SHU and the NDG contain antechambers for the thermalization of the incoming gas. Afterwards, the neutral gas entering the ion source will be ionized by impacting 70 eV electrons emitted by a hot filament. In the NDG, the resulting ions are measured as a current by a sensitive electrometer and in proportion to the gas density inside the antechamber and hence the surrounding coma. In the SHU, the newly formed ions are accelerated by a sharp extraction voltage pulse into the drift section. After passing the reflectron, i.e., an opposing electric field, the ions cross the drift section again before impinging on the Micro Channel Plate detector. Since the voltage pulse provides the same energy to all extracted ions, their arrival time on the detector can be converted into a mass/charge ratio. Combining NDG and SHU measurements leads to *in situ* abundances of coma gases.

MANiaC is being built by an international consortium formed by the Instituto de Astrofísica de Andalucía (IAA Granada, Spain), the Institut für Weltraumforschung (IWF Graz, Austria), the Institut de Recherche en Astrophysique et Planétologie (IRAP Toulouse, France), and Creotech Instruments S. A. (Piaseczno, Poland) under the lead of the University of Bern (Switzerland).

## 5.4 Dust, Fields, and Plasma (DFP-A)

To enable multipoint *in situ* measurements of dust, plasma, and energetic atoms, five instrumental sensors, common data processing unit DAPU, and power supply system, PSU, will
<figure><img src="image_23.png" /><figcaption>**Fig. 21** DISC Unit. Left: assembled DISC breadboard. Right: Sensing plate with glued PZTs at 3 corners</figcaption></figure>

be constructed and placed on Spacecraft A, and two sensors with the respective DAPU and PSU will be present on probe B2 (see Sect. 5.10). Overall management of the DFP suite on both spacecraft A and B2 is by Centrum Badań Kosmicznych, CBK, Warsaw, Poland.

### 5.4.1 Dust Impact Sensor and Counter (DISC)

DFP-DISC will be provided in two twin units, mounted on spacecraft A and probe B2 (Fig. 21). DFP-DISC is devoted to the *in situ* characterization of cometary dust. In particular, DFP-DISC is designed to count the dust particles populating the coma and to determine the mass of each individual particle. The DFP-DISC design is a direct heritage of the Impact Sensor: one of the measurement sub-systems of the successful GIADA (Grain Impact Analyzer and Dust Accumulator) instrument. Full details of the instrument are provided in Della Corte et al. (2023).

DFP-DISC consists of:

- a square aluminum diaphragm, 0.5 mm thick with a sensitive area of 84 × 84 mm²;

- three lead zirconate titanate ceramic piezoelectrics (PZTs), with a resonant frequency of 200 kHz, placed at three corners of the aluminum diaphragm;

- one PZT, placed at the fourth corner of the aluminum diaphragm, is the internal calibrator.

Each dust grain impacting the aluminum diaphragm generates acoustic waves that propagate in the diaphragm reaching the PZTs. These waves' amplitudes are proportional to the momentum of the impacting dust particle. From the individual particle momentum measurement, knowing the relative speed between the spacecraft and the dust particles, DFP-DISC will determine the mass of individual impacting particles. DFP-DISC is a monitoring instrument with event driven acquisitions that will provide an *in situ* characterization of dust particles. It will count impacts and determine for individual grains their mass, impact duration, and density/structure. These measurements will allow the characterization of the dust coma structures along the spacecraft trajectory. The dust grain momentum measurement capabilities, combined with the foreseen flyby speeds, DFP-DISC will determine the dust mass distribution for grains in the range $10^{-15}-10^{-8}$ kg (for particles with mass $>10^{-8}$ kg, the dust particle count will be provided).

DISC is being built by an Italian consortium scientifically led by INAF Osservatorio Astronomico di Capodimonte (INAF-OACN, Napoli, Italy) and having Leonardo S.p.A (Campi Bisenzio (Fi), Italy) as the Prime industry and the Italian Space Agency (ASI) as the main supporting funder. The contributing consortium partners are the University of Napoli “Parthenope” (Napoli, Italy) and the Istituto di Astrofisica e Planetologia Spaziali (INAF-IAPS Roma, Italy). DISC benefits also of the support of ESA, Univ. Parthenope and MUR (Ministero Università e Ricerca).

### 5.4.2 FGM-A

The FGM-A fluxgate magnetometer is composed of two sensors (outboard and inboard) mounted on a deployable boom (with heritage from Venus Express and Kompsat-2A) and
<figure><img src="image_24.png" /><figcaption>**Fig. 21** DISC Unit. Left: assembled DISC breadboard. Right: Sensing plate with glued PZTs at 3 corners</figcaption></figure>

be constructed and placed on Spacecraft A, and two sensors with the respective DAPU and PSU will be present on probe B2 (see Sect. 5.10). Overall management of the DFP suite on both spacecraft A and B2 is by Centrum Badań Kosmicznych, CBK, Warsaw, Poland.

### 5.4.1 Dust Impact Sensor and Counter (DISC)

DFP-DISC will be provided in two twin units, mounted on spacecraft A and probe B2 (Fig. 21). DFP-DISC is devoted to the *in situ* characterization of cometary dust. In particular, DFP-DISC is designed to count the dust particles populating the coma and to determine the mass of each individual particle. The DFP-DISC design is a direct heritage of the Impact Sensor: one of the measurement sub-systems of the successful GIADA (Grain Impact Analyzer and Dust Accumulator) instrument. Full details of the instrument are provided in Della Corte et al. (2023).

DFP-DISC consists of:

- a square aluminum diaphragm, 0.5 mm thick with a sensitive area of 84 × 84 mm²;

- three lead zirconate titanate ceramic piezoelectrics (PZTs), with a resonant frequency of 200 kHz, placed at three corners of the aluminum diaphragm;

- one PZT, placed at the fourth corner of the aluminum diaphragm, is the internal calibrator.

Each dust grain impacting the aluminum diaphragm generates acoustic waves that propagate in the diaphragm reaching the PZTs. These waves' amplitudes are proportional to the momentum of the impacting dust particle. From the individual particle momentum measurement, knowing the relative speed between the spacecraft and the dust particles, DFP-DISC will determine the mass of individual impacting particles. DFP-DISC is a monitoring instrument with event driven acquisitions that will provide an *in situ* characterization of dust particles. It will count impacts and determine for individual grains their mass, impact duration, and density/structure. These measurements will allow the characterization of the dust coma structures along the spacecraft trajectory. The dust grain momentum measurement capabilities, combined with the foreseen flyby speeds, DFP-DISC will determine the dust mass distribution for grains in the range $10^{-15}$–$10^{-8}$ kg (for particles with mass $>10^{-8}$ kg, the dust particle count will be provided).

DISC is being built by an Italian consortium scientifically led by INAF Osservatorio Astronomico di Capodimonte (INAF-OACN, Napoli, Italy) and having Leonardo S.p.A (Campi Bisenzio (Fi), Italy) as the Prime industry and the Italian Space Agency (ASI) as the main supporting funder. The contributing consortium partners are the University of Napoli “Parthenope” (Napoli, Italy) and the Istituto di Astrofisica e Planetologia Spaziali (INAF-IAPS Roma, Italy). DISC benefits also of the support of ESA, Univ. Parthenope and MUR (Ministero Università e Ricerca).

### 5.4.2 FGM-A

The FGM-A fluxgate magnetometer is composed of two sensors (outboard and inboard) mounted on a deployable boom (with heritage from Venus Express and Kompsat-2A) and
<table><caption>**Table 7** FGM-A properties</caption><tr><td>Operation range</td><td>±16,000 nT (configurable)</td></tr><tr><td>Digital resolution</td><td>2 pT</td></tr><tr><td>Noise</td><td>&lt;10 pT/sqrt(Hz) at 1 Hz</td></tr><tr><td>Absolute accuracy (requirement)</td><td>±1 nT (goal), ±2 nT</td></tr><tr><td>Mass</td><td>1.8 kg boom with sensors, 0.5 kg electronics</td></tr><tr><td>Power</td><td>1.2 W total</td></tr><tr><td>Temperature range</td><td>[-80; +60] °C (both survival and operation). No heaters installed</td></tr></table>

<figure><img src="image_24.png" /><figcaption>**Fig. 22** Integration of the FGM-A sensor within the COMPLIMENT merged probe</figcaption></figure>

their electronic front-ends, which are hosted in the DFP central electronics box, CEBOX. Its main properties are shown in Table 7. The fluxgate outboard sensor has been merged with the COMPLIMENT Spherical Probe, shown in Fig. 22. The combined sensor consists of a hollow spherical Langmuir probe that harbours a fluxgate magnetometer at its centre. Special precautions have been taken to minimize the possible interference between both whilst also being very lightweight. FGM-A is being built by an international consortium formed by the Technische Universität Braunschweig (Germany), the Institut für Weltraumforschung (IWF Graz, Austria) and Imperial College London (UK).

### 5.4.3 COMPLIMENT

The COMetary Plasma Light InstruMENT unit (COMPLIMENT) is an electric field instrument that combines the capabilities of both a mutual impedance probe and a Langmuir probe. It is designed to probe the cometary plasma, the electric field and the (few nanometer-size) dust within the cometary environment. COMPLIMENT will provide the following measurements: electric field and waves (one component), high cadence independent ion and electron densities, electron temperature(s), spacecraft potential, integrated EUV flux, nanodust impacts (signal processed by DAPU) in order to address the structure and dynamics of the ionized and dusty cometary environment and its interactions with the escaping cometary atmosphere. COMPLIMENT (Table 8, Fig. 22) is composed of three sensors: two
<table><caption>**Table 7** FGM-A properties</caption><tr><td>Operation range</td><td>±16,000 nT (configurable)</td></tr><tr><td>Digital resolution</td><td>2 pT</td></tr><tr><td>Noise</td><td>&lt;10 pT/sqrt(Hz) at 1 Hz</td></tr><tr><td>Absolute accuracy (requirement)</td><td>±1 nT (goal), ±2 nT</td></tr><tr><td>Mass</td><td>1.8 kg boom with sensors, 0.5 kg electronics</td></tr><tr><td>Power</td><td>1.2 W total</td></tr><tr><td>Temperature range</td><td>[-80; +60] °C (both survival and operation). No heaters installed</td></tr></table>

<figure><img src="image_25.png" /><figcaption>**Fig. 22** Integration of the FGM-A sensor within the COMPLIMENT merged probe</figcaption></figure>

their electronic front-ends, which are hosted in the DFP central electronics box, CEBOX. Its main properties are shown in Table 7. The fluxgate outboard sensor has been merged with the COMPLIMENT Spherical Probe, shown in Fig. 22. The combined sensor consists of a hollow spherical Langmuir probe that harbours a fluxgate magnetometer at its centre. Special precautions have been taken to minimize the possible interference between both whilst also being very lightweight. FGM-A is being built by an international consortium formed by the Technische Universität Braunschweig (Germany), the Institut für Weltraumforschung (IWF Graz, Austria) and Imperial College London (UK).

### 5.4.3 COMPLIMENT

The COMetary Plasma Light InstruMENT unit (COMPLIMENT) is an electric field instrument that combines the capabilities of both a mutual impedance probe and a Langmuir probe. It is designed to probe the cometary plasma, the electric field and the (few nanometer-size) dust within the cometary environment. COMPLIMENT will provide the following measurements: electric field and waves (one component), high cadence independent ion and electron densities, electron temperature(s), spacecraft potential, integrated EUV flux, nanodust impacts (signal processed by DAPU) in order to address the structure and dynamics of the ionized and dusty cometary environment and its interactions with the escaping cometary atmosphere. COMPLIMENT (Table 8, Fig. 22) is composed of three sensors: two
<table><caption>**Table 8** COMPLIMENT properties</caption><tr><td>Electric field component δE(t)</td><td>1 Hz–1.4 MHz</td></tr><tr><td>Electron density</td><td>10<sup>2</sup>–10<sup>5</sup> cm<sup>-3</sup>, &lt;1 Hz</td></tr><tr><td>Density fluctuation δn/n</td><td>DC-1 kHz</td></tr><tr><td>Ion density N<sub>i</sub></td><td>10<sup>2</sup>–10<sup>5</sup> cm<sup>-3</sup>, &lt;1 Hz</td></tr><tr><td>Electron Temperature T<sub>e</sub></td><td>0.01–30 eV, &lt;1 Hz</td></tr><tr><td>spacecraft potential U<sub>sc</sub></td><td>&lt;100 Hz</td></tr><tr><td>Integrated solar flux</td><td>&lt;1 Hz</td></tr></table>

<figure><img src="image_25.png" /><figcaption>**Fig. 23** The SCIENA fully operational Technology Model, excluding some thermal hardware</figcaption></figure>

electric spherical probes (8 cm), one of which is a merged Electro-magnetic sensor (COMPLIMENT + FGM-A), mounted on booms and a transmitter, two electronic boards (LP and HMI) for signal generation, reception, and treatment, plus dedicated software hosted on the DAPU.

DFP-COMPLIMENT is being built by an international consortium formed by the Belgian Institute for Space Aeronomy (BIRA-IASB, Brussels, Belgium), the Swedish Institute of Space physics (IRF, Sweden) under the lead of the Laboratoire de Physique et de Chimie de l'Environnement et de l'Espace (LPC2E, CNRS, Université d'Orléans, CNES, Orléans, France).

### 5.4.4 Solar Wind and Cometary Ions and Energetic Neutral Atoms (SCIENA)

SCIENA is an instrument of the SWIM family (Wieser and Barabash 2016) with one ion and one ENA sensor head (Fig. 23). The ion sensor head achieves a near $2\pi$ steradian field of view through four directional electrodes, followed by a classical electrostatic analyser for energy determination in the range a few eV up to 15 keV (Table 9). Mass is determined through start and stop surfaces and a time-of-flight system. The ENA system is similar, but has only two direction electrodes and is performing a two-dimensional direction scan. In front of the direction electrodes are an ion rejector and a charge conversion surface, similar to what is used on the ASAN instrument operating on the far side of the moon (Wieser et al. 2020). The energy coverage is focussed on solar wind ENA at 300 eV to 3 keV. SCIENA is being built by the Swedish Institute of Space Physics (IRF, Sweden).
<table><caption>**Table 8** COMPLIMENT properties</caption><tr><td>Electric field component δE(t)</td><td>1 Hz–1.4 MHz</td></tr><tr><td>Electron density</td><td>10<sup>2</sup>–10<sup>5</sup> cm<sup>-3</sup>, &lt;1 Hz</td></tr><tr><td>Density fluctuation δn/n</td><td>DC-1 kHz</td></tr><tr><td>Ion density N<sub>i</sub></td><td>10<sup>2</sup>–10<sup>5</sup> cm<sup>-3</sup>, &lt;1 Hz</td></tr><tr><td>Electron Temperature T<sub>e</sub></td><td>0.01–30 eV, &lt;1 Hz</td></tr><tr><td>spacecraft potential U<sub>sc</sub></td><td>&lt;100 Hz</td></tr><tr><td>Integrated solar flux</td><td>&lt;1 Hz</td></tr></table>

<figure><img src="image_26.png" /><figcaption>**Fig. 23** The SCIENA fully operational Technology Model, excluding some thermal hardware</figcaption></figure>

electric spherical probes (8 cm), one of which is a merged Electro-magnetic sensor (COMPLIMENT + FGM-A), mounted on booms and a transmitter, two electronic boards (LP and HMI) for signal generation, reception, and treatment, plus dedicated software hosted on the DAPU.

DFP-COMPLIMENT is being built by an international consortium formed by the Belgian Institute for Space Aeronomy (BIRA-IASB, Brussels, Belgium), the Swedish Institute of Space physics (IRF, Sweden) under the lead of the Laboratoire de Physique et de Chimie de l'Environnement et de l'Espace (LPC2E, CNRS, Université d'Orléans, CNES, Orléans, France).

### 5.4.4 Solar Wind and Cometary Ions and Energetic Neutral Atoms (SCIENA)

SCIENA is an instrument of the SWIM family (Wieser and Barabash 2016) with one ion and one ENA sensor head (Fig. 23). The ion sensor head achieves a near $2\pi$ steradian field of view through four directional electrodes, followed by a classical electrostatic analyser for energy determination in the range a few eV up to 15 keV (Table 9). Mass is determined through start and stop surfaces and a time-of-flight system. The ENA system is similar, but has only two direction electrodes and is performing a two-dimensional direction scan. In front of the direction electrodes are an ion rejector and a charge conversion surface, similar to what is used on the ASAN instrument operating on the far side of the moon (Wieser et al. 2020). The energy coverage is focussed on solar wind ENA at 300 eV to 3 keV. SCIENA is being built by the Swedish Institute of Space Physics (IRF, Sweden).
<table><caption>**Table 9** SCIENA properties</caption><thead><tr><th></th><th>Ions</th><th>ENA</th></tr></thead><tbody><tr><td>Energy range</td><td>10 eV–15 keV</td><td>300 eV–3 keV</td></tr><tr><td>Angular coverage</td><td>Near 2 π</td><td>30° × 150°</td></tr><tr><td>Mass resolution</td><td>1, 2, 4, 8, 16, 32 amu</td><td>1, heavy</td></tr><tr><td>Time resolution</td><td>1 s / energy spectra<br>20–50 s / full distribution</td><td>1–10 s / energy spectra<br>5–50 s / full scan</td></tr></tbody></table>

<figure><img src="image_26.png" /><figcaption>**Table 10** LEES measurement parameters</figcaption></figure>

<table><thead><tr><th></th><th>ENA</th></tr></thead><tbody><tr><td>Energy range</td><td>10 eV–1 keV in 90 energy bins</td></tr><tr><td>Energy resolution</td><td>E/ΔE 0.07</td></tr><tr><td>Angular coverage</td><td>2.8 π sr</td></tr><tr><td>Time resolution</td><td>8 s / energy spectra</td></tr></tbody></table>

### 5.4.5 Low-Energy Electron Spectrometer (LEES)

LEES will determine the electron density, temperature, and the velocity distribution functions of the *in situ* plasma environment of the solar wind and comet (Fig. 24). LEES will detect the suprathermal photoelectrons created during neutral-plasma interactions in the coma and trace the magnetic connectivity between the spacecraft and the cometary environment. In addition LEES will measure the properties of negatively charged ions and dust of cometary origin. The energy range covered by LEES will be from a few eV up to 1 keV. Particles are measured in 360° azimuth angles with elevation angles ranging from –40° to 70° (Table 10).

LEES is a classical top-hat type electrostatic analyzer with a Field-of-View (FoV) deflector system to allow the coverage of the elevation range by electrostatic deflection of incoming electrons. Incident charged particles enter the sensor through the exterior electrically grounded toroidal aperture grid. The particles are steered from the arrival direction into the top-hat electrostatic analyzer (ESA) using voltages applied to the upper deflection plate. The ESA section permits only electrons of the selected energy to reach the detector subsystem with MCPs (microchannel plates) in chevron stack. The MCP anode is divided in
<table><caption>**Table 9** SCIENA properties</caption><thead><tr><th></th><th>Ions</th><th>ENA</th></tr></thead><tbody><tr><td>Energy range</td><td>10 eV–15 keV</td><td>300 eV–3 keV</td></tr><tr><td>Angular coverage</td><td>Near 2 π</td><td>30° × 150°</td></tr><tr><td>Mass resolution</td><td>1, 2, 4, 8, 16, 32 amu</td><td>1, heavy</td></tr><tr><td>Time resolution</td><td>1 s / energy spectra<br>20–50 s / full distribution</td><td>1–10 s / energy spectra<br>5–50 s / full scan</td></tr></tbody></table>

<figure><img src="image_27.png" /><figcaption>**Table 10** LEES measurement parameters</figcaption></figure>

<table><thead><tr><th></th><th>ENA</th></tr></thead><tbody><tr><td>Energy range</td><td>10 eV–1 keV in 90 energy bins</td></tr><tr><td>Energy resolution</td><td>E/ΔE 0.07</td></tr><tr><td>Angular coverage</td><td>2.8 π sr</td></tr><tr><td>Time resolution</td><td>8 s / energy spectra</td></tr></tbody></table>

### 5.4.5 Low-Energy Electron Spectrometer (LEES)

LEES will determine the electron density, temperature, and the velocity distribution functions of the *in situ* plasma environment of the solar wind and comet (Fig. 24). LEES will detect the suprathermal photoelectrons created during neutral-plasma interactions in the coma and trace the magnetic connectivity between the spacecraft and the cometary environment. In addition LEES will measure the properties of negatively charged ions and dust of cometary origin. The energy range covered by LEES will be from a few eV up to 1 keV. Particles are measured in 360° azimuth angles with elevation angles ranging from –40° to 70° (Table 10).

LEES is a classical top-hat type electrostatic analyzer with a Field-of-View (FoV) deflector system to allow the coverage of the elevation range by electrostatic deflection of incoming electrons. Incident charged particles enter the sensor through the exterior electrically grounded toroidal aperture grid. The particles are steered from the arrival direction into the top-hat electrostatic analyzer (ESA) using voltages applied to the upper deflection plate. The ESA section permits only electrons of the selected energy to reach the detector subsystem with MCPs (microchannel plates) in chevron stack. The MCP anode is divided in
<table><caption>**Table 11** The parameters of the PSU-A unit</caption><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Supply Voltage (unregulated) [V]</td><td>24–34</td></tr><tr><td>Power [W]</td><td>22</td></tr><tr><td>Redundancy</td><td>YES (NOM/RED)</td></tr><tr><td>Mass [kg]</td><td>0.5 (NOM+RED)</td></tr><tr><td>Secondary output voltages</td><td>3.7 V, ±5 V, ±12 V</td></tr></tbody></table>

16 sections that provide 16 azimuth bins with about 20° of angular resolution. Such a design allows fulfilling all scientific requirements for the electron and negative ions measurements.

LEES is being developed by the Institut de Recherche en Astrophysique et Planétologie in France, with significant contributions from Charles University in the Czech Republic and the Laboratoire d'Astrophysique de Bordeaux in France.

### 5.4.6 DAPU-A

The DAPU is a central data processing unit of the DFP instrument suite both on Spacecraft A and B2. It is a computer board serving as a common digital interface between the spacecraft and the DFP instruments and sensors. DAPU will:

- Perform last stage processing, compression and buffering of all science data,

- Store the data from the entire flyby in its large flash memory (as a backup copy in case),

- Manage common DFP suite modes, instrument commanding and configuration.

- DAPU-A will also count dust particle impacts in COMPLIMENT probe voltage data, by sensing the plasma clouds from the vapourised dust grains. This will allow the detection of small dust particles down to <100 nm across.

Development of the DAPU units for spacecraft A and probe B2 is led by the Institute of Atmospheric Physics (IAP) of the Czech Academy of Sciences, Prague, Czechia.

### 5.4.7 PSU-A

The Power Supply Unit (Table 11) shall generate, condition, control, monitor, and distribute electrical power to the DFP units from two unregulated 28 V buses, to fulfil the instrument power demands throughout all mission phases. The instrument power interfaces toward the spacecraft have been designed to prevent any single point failure, which could lead to a short circuit. The PSU modules include current and voltage monitoring, soft-start circuits, over-current protection (OCP), over-voltage protection (OVP), and under-voltage lock-out circuitry (UVLO) for the protection of the DFP units and subsystems. The current design of the PSU is based on previous and ongoing instrument power supply designs that have been implementing isolated DC/DC converters with embedded logic for switching control, protection, and HK monitoring. The design has heritage from instruments also built at Centrum Badań Kosmicznych, CBK, that were mainly used for radio and plasma diagnostics, i.e. Chronograph Control Block for PROBA-3 (ESA) and RELEC (Russia).

## 5.5 Hydrogen Imager (HI)

HI is an ultraviolet imager for hydrogen Lyman-α emission with parameters as listed in Table 12. The instrument aims to characterize the spatial distribution of the hydrogen coma
<table><caption>**Table 11** The parameters of the PSU-A unit</caption><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Supply Voltage (unregulated) [V]</td><td>24-34</td></tr><tr><td>Power [W]</td><td>22</td></tr><tr><td>Redundancy</td><td>YES (NOM/RED)</td></tr><tr><td>Mass [kg]</td><td>0.5 (NOM+RED)</td></tr><tr><td>Secondary output voltages</td><td>3.7 V, ±5 V, ±12 V</td></tr></tbody></table>

16 sections that provide 16 azimuth bins with about 20° of angular resolution. Such a design allows fulfilling all scientific requirements for the electron and negative ions measurements.

LEES is being developed by the Institut de Recherche en Astrophysique et Planétologie in France, with significant contributions from Charles University in the Czech Republic and the Laboratoire d'Astrophysique de Bordeaux in France.

### 5.4.6 DAPU-A

The DAPU is a central data processing unit of the DFP instrument suite both on Spacecraft A and B2. It is a computer board serving as a common digital interface between the spacecraft and the DFP instruments and sensors. DAPU will:

- Perform last stage processing, compression and buffering of all science data,

- Store the data from the entire flyby in its large flash memory (as a backup copy in case),

- Manage common DFP suite modes, instrument commanding and configuration.

- DAPU-A will also count dust particle impacts in COMPLIMENT probe voltage data, by sensing the plasma clouds from the vapourised dust grains. This will allow the detection of small dust particles down to <100 nm across.

Development of the DAPU units for spacecraft A and probe B2 is led by the Institute of Atmospheric Physics (IAP) of the Czech Academy of Sciences, Prague, Czechia.

### 5.4.7 PSU-A

The Power Supply Unit (Table 11) shall generate, condition, control, monitor, and distribute electrical power to the DFP units from two unregulated 28 V buses, to fulfil the instrument power demands throughout all mission phases. The instrument power interfaces toward the spacecraft have been designed to prevent any single point failure, which could lead to a short circuit. The PSU modules include current and voltage monitoring, soft-start circuits, over-current protection (OCP), over-voltage protection (OVP), and under-voltage lock-out circuitry (UVLO) for the protection of the DFP units and subsystems. The current design of the PSU is based on previous and ongoing instrument power supply designs that have been implementing isolated DC/DC converters with embedded logic for switching control, protection, and HK monitoring. The design has heritage from instruments also built at Centrum Badań Kosmicznych, CBK, that were mainly used for radio and plasma diagnostics, i.e. Chronograph Control Block for PROBA-3 (ESA) and RELEC (Russia).

## 5.5 Hydrogen Imager (HI)

HI is an ultraviolet imager for hydrogen Lyman-α emission with parameters as listed in Table 12. The instrument aims to characterize the spatial distribution of the hydrogen coma
<table><caption>**Table 12** HI parameters</caption><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Power</td><td>&lt;9 W (nominal), 12 W(max.)</td></tr><tr><td>Mass</td><td>&lt;1.6 kg</td></tr><tr><td>Field of View</td><td>±2.1°</td></tr><tr><td>Spatial resolution</td><td>0.02°</td></tr><tr><td>Bandpass filter</td><td>120 nm, FWHM 10 nm</td></tr><tr><td>Hydrogen filter</td><td>121.567 nm (FWHM ~3 pm, depending on filament temperature)</td></tr><tr><td>Deuterium filter</td><td>121.533 nm (FWHM ~3 pm, depending on filament temperature)</td></tr></tbody></table>

<figure><img src="image_27.png" /><figcaption>**Fig. 25** Design overview of HI</figcaption></figure>

through the imaging observation during the approach to the target. From the radial profile of hydrogen around the nucleus, the water production rate and its spatial variation will be deduced. HI is a Cassegrain-type telescope, with mirrors coated with Al/MgF₂. A bandpass filter is installed on the light axis (Fig. 25). The photon detection system consists of MCPs combined with a Resistive Anode Encoder (RAE), enabling two-dimensional imaging. The electron cloud generated by the MCPs is divided into the four corners of the RAE, and the arrival position of photons is determined based on the relative charge distribution. HI offers two observation modes: the “light curve mode” and the “imaging mode”. In the former, a detector counts the number of incident photons per frame, while in the latter, the main processor calculates the position of each photon and integrates them into a 256 × 256 matrix over a fixed time span. Two additional gas filters are installed along the light axis. The narrow bandpass filters contain atomic hydrogen and deuterium individually when activated by filaments. By using these filters, the brightness ratio of the isotopic components for the resonant absorption of Lyman-α radiation can be determined, which occurs at 121.567 nm for hydrogen and 121.534 nm for deuterium. However, it is important to note that the measurement accuracy achieved using the gas filters strongly depends on the geometric conditions of the target and the B1 probe, which can cause a Doppler shift in the line centre. The HI consortium is led by the University of Tokyo, with contributions from Rikkyo University and JAXA.

## 5.6 Plasma Suite (PS)

PS consists of an ion mass spectrometer and a 3-axis magnetometer (Fig. 26). It provides velocity distribution functions of individual ion species of the low-energy coma plasma, as well as the DC and low-frequency AC magnetic field data. The ion sensor consists of
<table><caption>**Table 12** HI parameters</caption><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Power</td><td>&lt;9 W (nominal), 12 W(max.)</td></tr><tr><td>Mass</td><td>&lt;1.6 kg</td></tr><tr><td>Field of View</td><td>±2.1°</td></tr><tr><td>Spatial resolution</td><td>0.02°</td></tr><tr><td>Bandpass filter</td><td>120 nm, FWHM 10 nm</td></tr><tr><td>Hydrogen filter</td><td>121.567 nm (FWHM ~3 pm, depending on filament temperature)</td></tr><tr><td>Deuterium filter</td><td>121.533 nm (FWHM ~3 pm, depending on filament temperature)</td></tr></tbody></table>

<figure><img src="image_28.png" /><figcaption>**Fig. 25** Design overview of HI</figcaption></figure>

through the imaging observation during the approach to the target. From the radial profile of hydrogen around the nucleus, the water production rate and its spatial variation will be deduced. HI is a Cassegrain-type telescope, with mirrors coated with Al/MgF₂. A bandpass filter is installed on the light axis (Fig. 25). The photon detection system consists of MCPs combined with a Resistive Anode Encoder (RAE), enabling two-dimensional imaging. The electron cloud generated by the MCPs is divided into the four corners of the RAE, and the arrival position of photons is determined based on the relative charge distribution. HI offers two observation modes: the “light curve mode” and the “imaging mode”. In the former, a detector counts the number of incident photons per frame, while in the latter, the main processor calculates the position of each photon and integrates them into a 256 × 256 matrix over a fixed time span. Two additional gas filters are installed along the light axis. The narrow bandpass filters contain atomic hydrogen and deuterium individually when activated by filaments. By using these filters, the brightness ratio of the isotopic components for the resonant absorption of Lyman-α radiation can be determined, which occurs at 121.567 nm for hydrogen and 121.534 nm for deuterium. However, it is important to note that the measurement accuracy achieved using the gas filters strongly depends on the geometric conditions of the target and the B1 probe, which can cause a Doppler shift in the line centre. The HI consortium is led by the University of Tokyo, with contributions from Rikkyo University and JAXA.

## 5.6 Plasma Suite (PS)

PS consists of an ion mass spectrometer and a 3-axis magnetometer (Fig. 26). It provides velocity distribution functions of individual ion species of the low-energy coma plasma, as well as the DC and low-frequency AC magnetic field data. The ion sensor consists of
<figure><img src="image_28.png" /><figcaption>**Table 13** Expected performance of PS</figcaption></figure>

<table><tr><td>Ion mass spectrometer:</td><td>Instantaneous 3-D ion distribution with mass discrimination</td></tr><tr><td>Energy:</td><td>10–20,000 eV/q, ΔE/E ~10%</td></tr><tr><td>Mass:</td><td>M/ΔM ~30</td></tr><tr><td>Field-of-view:</td><td>Hemispheric</td></tr><tr><td>Magnetometer:</td><td>DC and low-frequency AC magnetic field</td></tr><tr><td>Absolute accuracy:</td><td>1 nT @ ±512 nT Range</td></tr><tr><td>Directional accuracy:</td><td>&lt;5 degrees</td></tr><tr><td>Noise level:</td><td>~12 pT/Hz<sup>1/2</sup> @ 1 Hz<br/>~6 pT/Hz<sup>1/2</sup> @ 10 Hz</td></tr></table>

an electrostatic analyser and time-of-flight mass analyser (TOF). The incident energy and direction of each incoming ion are determined by the electrostatic analyser. Ions are then introduced to the TOF sector, where mass-per-charge is measured by the linearly increasing electric field. A large field-of-view (entire hemisphere if there is no exterior interference) is achieved by an entrance deflector in the electrostatic analyser unit. The PS magnetometer is based on the fundamental mode orthogonal fluxgate (FM-OFG) technique. FM-OFG adopts an amorphous wire sensor core driven with a unique excitation method where AC current is superposed on DC bias current. It enables low-noise detection of magnetic field with a compact and lightweight sensor design. The magnetometer is accommodated on the top of an extensible boom for magnetic cleanliness of the measurements. The boom is stowed during the launch and deployed during commissioning. The key performance aspects of PS are summarised in Table 13.

The PS consortium is led by the University of Tokyo, with contributions from Kyoto University, Osaka University, Kobe University, and JAXA.

## 5.7 Narrow Angle Camera (NAC) and Wide Angle Camera (WAC)

The Narrow Angle Camera (NAC) is an optical telescopic camera. NAC will obtain optical images of the target nucleus with a high solar phase angle, to address nucleus science. The field of view is 3.5 × 2.6°, which is anticipated to be wide enough to observe the entire nucleus. NAC is equipped with a CCD sensor of 3296 × 2472 pixels. The pixel resolution of the NAC is about 15.6 m/pix or better at a closest approach distance of 850 km. The Wide
<figure><img src="image_29.png" /><figcaption>**Table 13** Expected performance of PS</figcaption></figure>

<table><tr><td>Ion mass spectrometer:</td><td>Instantaneous 3-D ion distribution with mass discrimination</td></tr><tr><td>Energy:</td><td>10–20,000 eV/q, ΔE/E ~10%</td></tr><tr><td>Mass:</td><td>M/ΔM ~30</td></tr><tr><td>Field-of-view:</td><td>Hemispheric</td></tr><tr><td>Magnetometer:</td><td>DC and low-frequency AC magnetic field</td></tr><tr><td>Absolute accuracy:</td><td>1 nT @ ±512 nT Range</td></tr><tr><td>Directional accuracy:</td><td>&lt;5 degrees</td></tr><tr><td>Noise level:</td><td>~12 pT/Hz<sup>1/2</sup> @ 1 Hz<br/>~6 pT/Hz<sup>1/2</sup> @ 10 Hz</td></tr></table>

an electrostatic analyser and time-of-flight mass analyser (TOF). The incident energy and direction of each incoming ion are determined by the electrostatic analyser. Ions are then introduced to the TOF sector, where mass-per-charge is measured by the linearly increasing electric field. A large field-of-view (entire hemisphere if there is no exterior interference) is achieved by an entrance deflector in the electrostatic analyser unit. The PS magnetometer is based on the fundamental mode orthogonal fluxgate (FM-OFG) technique. FM-OFG adopts an amorphous wire sensor core driven with a unique excitation method where AC current is superposed on DC bias current. It enables low-noise detection of magnetic field with a compact and lightweight sensor design. The magnetometer is accommodated on the top of an extensible boom for magnetic cleanliness of the measurements. The boom is stowed during the launch and deployed during commissioning. The key performance aspects of PS are summarised in Table 13.

The PS consortium is led by the University of Tokyo, with contributions from Kyoto University, Osaka University, Kobe University, and JAXA.

## 5.7 Narrow Angle Camera (NAC) and Wide Angle Camera (WAC)

The Narrow Angle Camera (NAC) is an optical telescopic camera. NAC will obtain optical images of the target nucleus with a high solar phase angle, to address nucleus science. The field of view is 3.5 × 2.6°, which is anticipated to be wide enough to observe the entire nucleus. NAC is equipped with a CCD sensor of 3296 × 2472 pixels. The pixel resolution of the NAC is about 15.6 m/pix or better at a closest approach distance of 850 km. The Wide
<figure><img src="image_29.png" /><figcaption>**Fig. 27** Components of the NAC/WAC system</figcaption></figure>

Angle Camera (WAC) is also an optical camera with CMOS imaging sensor of 2048 × 2048 pixels, which observes the coma with wide-field of view of 90 × 90°. The instantaneous field of view is roughly 40 times wider than that of the NAC. Both cameras have panchromatic filters covering 0.4 to 0.75 µm in wavelength. They are utilized not only for the scientific objectives, but also for the optical navigation of the B1 probe. The NAC/WAC system (Fig. 27) consists of: (a) electronics box with control function of the both cameras and interface to the bus system (NAC-E), (b) Narrow angle camera sensor and optics (NAC-S), and (c) Wide angle camera sensor and optics (WAC). NAC-E and NAC-S will be based on the telescopic camera TENGOO on-board the Martian Moon Explorer – MMX – mission (Kameda et al. 2021).

The NAC and WAC consortium is led by the Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration Agency (JAXA), with contributions from Rikkyo University.

## 5.8 Entire Visible Sky (EnVisS)

The EnVisS camera (Fig. 28) has been conceived to study the comet's dust environment in the visible range, at wavelengths of 550-800 nm. The intensity, degree of linear polarisation, and polarisation angle orientation of the light scattered by the dust particles in the comet coma, with a full 180° phase angle coverage, will be studied. Such a measurement is unique, it has never been carried out in space. Giotto's Halley Opical Probe - HOPE - could only observe a very narrow angle FoV in the direction opposite to the motion of the spacecraft (Levasseur-Regourd et al. 1986).

The FoV of the instrument is designed to allow the entire sky to be acquired thanks to the rotation of the B2 spacecraft. EnVisS will feature a flexible push-broom/push-frame imaging technique, thus acquiring slices of the sky, while the probe rotates; the slices acquired will be stitched together, on-ground, to form a whole sky image. The coma investigation will be conducted throughout the full flyby of the comet from an advantageous point inside the coma itself. The probe spin-axis will be pointed at the comet nucleus for most of the time, except at closest approach when the comet nucleus will fall inside the camera FoV.

Depending on the target object activity, the map of the coma will be taken with different spatial resolution (i.e., smearing through adjustment of exposure length, and pixel binning) to achieve the desired SNR. The minimum angular resolution element corresponds to 0.2°. EnVisS features a fish-eye lens coupled to a commercial space-qualified detector and ad-hoc power and data handling units and software. The design solution adopted for the filters, i.e., fixed filter strips mounted as near as possible to the detector, allows for a compact, low mass and low complexity camera. Three broad band filters, all working in the same wavelength range, are foreseen for the camera: one non polarizing filter, centred on the detector, and, on
<figure><img src="image_30.png" /><figcaption>**Fig. 27** Components of the NAC/WAC system</figcaption></figure>

Angle Camera (WAC) is also an optical camera with CMOS imaging sensor of 2048 × 2048 pixels, which observes the coma with wide-field of view of 90 × 90°. The instantaneous field of view is roughly 40 times wider than that of the NAC. Both cameras have panchromatic filters covering 0.4 to 0.75 µm in wavelength. They are utilized not only for the scientific objectives, but also for the optical navigation of the B1 probe. The NAC/WAC system (Fig. 27) consists of: (a) electronics box with control function of the both cameras and interface to the bus system (NAC-E), (b) Narrow angle camera sensor and optics (NAC-S), and (c) Wide angle camera sensor and optics (WAC). NAC-E and NAC-S will be based on the telescopic camera TENGOO on-board the Martian Moon Explorer – MMX – mission (Kameda et al. 2021).

The NAC and WAC consortium is led by the Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration Agency (JAXA), with contributions from Rikkyo University.

## 5.8 Entire Visible Sky (EnVisS)

The EnVisS camera (Fig. 28) has been conceived to study the comet's dust environment in the visible range, at wavelengths of 550-800 nm. The intensity, degree of linear polarisation, and polarisation angle orientation of the light scattered by the dust particles in the comet coma, with a full 180° phase angle coverage, will be studied. Such a measurement is unique, it has never been carried out in space. Giotto's Halley Opical Probe - HOPE - could only observe a very narrow angle FoV in the direction opposite to the motion of the spacecraft (Levasseur-Regourd et al. 1986).

The FoV of the instrument is designed to allow the entire sky to be acquired thanks to the rotation of the B2 spacecraft. EnVisS will feature a flexible push-broom/push-frame imaging technique, thus acquiring slices of the sky, while the probe rotates; the slices acquired will be stitched together, on-ground, to form a whole sky image. The coma investigation will be conducted throughout the full flyby of the comet from an advantageous point inside the coma itself. The probe spin-axis will be pointed at the comet nucleus for most of the time, except at closest approach when the comet nucleus will fall inside the camera FoV.

Depending on the target object activity, the map of the coma will be taken with different spatial resolution (i.e., smearing through adjustment of exposure length, and pixel binning) to achieve the desired SNR. The minimum angular resolution element corresponds to 0.2°. EnVisS features a fish-eye lens coupled to a commercial space-qualified detector and ad-hoc power and data handling units and software. The design solution adopted for the filters, i.e., fixed filter strips mounted as near as possible to the detector, allows for a compact, low mass and low complexity camera. Three broad band filters, all working in the same wavelength range, are foreseen for the camera: one non polarizing filter, centred on the detector, and, on
<figure><img src="image_30.png" /><figcaption>**Table 14** Summary of EnVisS characteristics</figcaption></figure>

<table><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Wavelength coverage</td><td>550–800 nm<br>1 broad band filter<br>2 polarimetric filters</td></tr><tr><td>Instrument FoV</td><td>180° × 45° (fixed)<br>180° × 360° (dynamic)</td></tr><tr><td>Entrance aperture (F#)</td><td>1.23 mm (2.8)</td></tr><tr><td>Detector</td><td>CMOS 2k × 2k<br>5.5-μm px size</td></tr><tr><td>Scale factor</td><td>0.1°/px</td></tr><tr><td>MTF</td><td>>70% @ 45 lp/mm</td></tr><tr><td>Distortion & telecentricity</td><td>&lt;8% (f-theta distortion law)<br>and &lt;4° (at the FoV edges)</td></tr></tbody></table>

either side, two polarimetric filters with transmission axis angles oriented at 45° one to the other. The full EnVisS characteristics are summarized in Table 14.

A flexible approach has been devised to allow a SNR of 10 to be obtained in the case of broadband images and of 100 for the polarimetric images. The signal from the coma in the direction of the apparent motion of the scene is not expected to show big changes, so a high spatial resolution is not needed from a scientific point of view. The integration time for each filter strip can be tuned, allowing for some smearing in the along-track direction. The spatial resolution can be retained in the across-track direction to assure a sampling of the comet phase function every 0.2°. This strategy will also allow for an adjustment of the exposure time if the radiance of the coma is different to the expected. Should the signal be extremely low, further pixel binning on-board, or co-adding, on-ground, of the images over different rotations, could be considered.

EnVisS is being realized by an international consortium scientifically led by Istituto di Fotonica e Nanotecnologie – Consiglio Nazionale delle Ricerche (CNR-IFN, Padova, Italy) and having Leonardo S.p.A (Campi Bisenzio (FI), Italy) as Prime industry. Instituto de Astrofísica de Andalucía (IAA-CSIC, Granada, Spain) is co-leading and, with SENER
<figure><img src="image_31.png" /><figcaption>**Table 14** Summary of EnVisS characteristics</figcaption></figure>

<table><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Wavelength coverage</td><td>550–800 nm<br>1 broad band filter<br>2 polarimetric filters</td></tr><tr><td>Instrument FoV</td><td>180° × 45° (fixed)<br>180° × 360° (dynamic)</td></tr><tr><td>Entrance aperture (F#)</td><td>1.23 mm (2.8)</td></tr><tr><td>Detector</td><td>CMOS 2k × 2k<br>5.5-μm px size</td></tr><tr><td>Scale factor</td><td>0.1°/px</td></tr><tr><td>MTF</td><td>&gt;70% @ 45 lp/mm</td></tr><tr><td>Distortion & telecentricity</td><td>&lt;8% (f-theta distortion law)<br>and &lt;4° (at the FoV edges)</td></tr></tbody></table>

either side, two polarimetric filters with transmission axis angles oriented at 45° one to the other. The full EnVisS characteristics are summarized in Table 14.

A flexible approach has been devised to allow a SNR of 10 to be obtained in the case of broadband images and of 100 for the polarimetric images. The signal from the coma in the direction of the apparent motion of the scene is not expected to show big changes, so a high spatial resolution is not needed from a scientific point of view. The integration time for each filter strip can be tuned, allowing for some smearing in the along-track direction. The spatial resolution can be retained in the across-track direction to assure a sampling of the comet phase function every 0.2°. This strategy will also allow for an adjustment of the exposure time if the radiance of the coma is different to the expected. Should the signal be extremely low, further pixel binning on-board, or co-adding, on-ground, of the images over different rotations, could be considered.

EnVisS is being realized by an international consortium scientifically led by Istituto di Fotonica e Nanotecnologie – Consiglio Nazionale delle Ricerche (CNR-IFN, Padova, Italy) and having Leonardo S.p.A (Campi Bisenzio (FI), Italy) as Prime industry. Instituto de Astrofísica de Andalucía (IAA-CSIC, Granada, Spain) is co-leading and, with SENER
<figure><img src="image_31.png" /><figcaption>**Fig. 29** OPIC engineering model with internals exposed</figcaption></figure>

(Barcelona, Spain), is responsible for the DHU and PHU development. The contributing consortium partners are: Istituto Nazionale di Astrofisica (INAF-OAC Napoli, Italy), responsible for the instrument management, with the support of University of Napoli Parthenope (Napoli, Italy); and Huld (Espoo, Finland), with the support of Aalto University (Espoo, Finland), responsible for the ASW development. Preliminary instrument conceptual design was carried out by the University College London's Mullard Space Science Laboratory, MSSL, UK.

## 5.9 Optical Periscope Imager for Comets (OPIC)

OPIC, situated on probe B2 and looking over the edge of its dust shield, is an automated camera system for taking images of the target and its near environment (Fig. 29). It consists of an automated camera head (3D Plus 3DCM734-1 SS), imaging optics (lens assembly and baffled periscope) and interface electronics. It has a 2048 × 2048-pixel CMV4000 sensor and an integrated ProAsic3 FPGA. OPIC's field-of-view is ~18.2 × 18.2°. OPIC is connected to the EnVisS instrument for additional data handling and power.

When far from the nucleus, OPIC will take long exposure images of the area around the nucleus, when the nucleus is not resolvable. These images will show the amount and spatial distribution of gas and dust is in OPIC's viewing direction. The data can be also used to constrain the trajectory and rotation state of B2 after separation.

When the nucleus becomes resolvable, the captured images will be processed within OPIC before transmission to Spacecraft A. The images will show the low-resolution structure of the nucleus and of the gas and dust immediately around it, including potential satellites and fragments. This can be combined with imagery from the A and B1 platforms to generate a more detailed and less ambiguous 3D model of the target. OPIC is being built by Tartu Observatory, University of Tartu, Tõravere, Estonia.

## 5.10 Dust, Fields, and Plasma (DFP-B2)

The DFP sensors on probe B2 will measure the magnetic field and the cometary dust. Two instrumental sensors, DISC, and FGM-B2, a slightly modified common data processing unit DAPU, and power supply system PSU will be constructed and placed on probe B2. The Power Supply Unit on B2 is similar to PSU-A, but with the parameters listed in Table 15. The Dust Impact Sensor and Counter (DISC) on Probe B2 is identical to that on Spacecraft A (see Sect. 5.4).
<figure><img src="image_32.png" /><figcaption>**Fig. 29** OPIC engineering model with internals exposed</figcaption></figure>

(Barcelona, Spain), is responsible for the DHU and PHU development. The contributing consortium partners are: Istituto Nazionale di Astrofisica (INAF-OAC Napoli, Italy), responsible for the instrument management, with the support of University of Napoli Parthenope (Napoli, Italy); and Huld (Espoo, Finland), with the support of Aalto University (Espoo, Finland), responsible for the ASW development. Preliminary instrument conceptual design was carried out by the University College London's Mullard Space Science Laboratory, MSSL, UK.

## 5.9 Optical Periscope Imager for Comets (OPIC)

OPIC, situated on probe B2 and looking over the edge of its dust shield, is an automated camera system for taking images of the target and its near environment (Fig. 29). It consists of an automated camera head (3D Plus 3DCM734-1 SS), imaging optics (lens assembly and baffled periscope) and interface electronics. It has a 2048 × 2048-pixel CMV4000 sensor and an integrated ProAsic3 FPGA. OPIC's field-of-view is ~18.2 × 18.2°. OPIC is connected to the EnVisS instrument for additional data handling and power.

When far from the nucleus, OPIC will take long exposure images of the area around the nucleus, when the nucleus is not resolvable. These images will show the amount and spatial distribution of gas and dust is in OPIC's viewing direction. The data can be also used to constrain the trajectory and rotation state of B2 after separation.

When the nucleus becomes resolvable, the captured images will be processed within OPIC before transmission to Spacecraft A. The images will show the low-resolution structure of the nucleus and of the gas and dust immediately around it, including potential satellites and fragments. This can be combined with imagery from the A and B1 platforms to generate a more detailed and less ambiguous 3D model of the target. OPIC is being built by Tartu Observatory, University of Tartu, Tõravere, Estonia.

## 5.10 Dust, Fields, and Plasma (DFP-B2)

The DFP sensors on probe B2 will measure the magnetic field and the cometary dust. Two instrumental sensors, DISC, and FGM-B2, a slightly modified common data processing unit DAPU, and power supply system PSU will be constructed and placed on probe B2. The Power Supply Unit on B2 is similar to PSU-A, but with the parameters listed in Table 15. The Dust Impact Sensor and Counter (DISC) on Probe B2 is identical to that on Spacecraft A (see Sect. 5.4).
<table><caption>**Table 15** The parameters of the PSU-B2 unit</caption><thead><tr><th>Parameter</th><th></th></tr></thead><tbody><tr><td>Supply Voltage (unregulated) [V]</td><td>28 V</td></tr><tr><td>Power [W]</td><td>10</td></tr><tr><td>Redundancy</td><td>No</td></tr><tr><td>Mass [kg]</td><td>0.25</td></tr><tr><td>Secondary output voltages</td><td>3.4 V, ±5 V</td></tr></tbody></table>

<figure><img src="image_32.png" /><figcaption>**Fig. 30** Picture of its sensing elements (left) and CAD rendering (right) of one FGM-B2 sensor</figcaption></figure>

### 5.10.1 FGM-B2

FGM-B2, or BFG for short, (Fig. 30) is composed of two sensors (outboard and inboard) mounted on a rigid boom, and their electronic front-ends, which are hosted in the DFP CE-BOX. Whilst both sensors and electronics have a strong heritage (e.g., Rosetta/Philae, Venus Express, and THEMIS (sensors) and MMS, Geo-KOMPSAT-2A/SOSMAG (electronics)), their mass and power consumption have been optimised for the resource-constrained Probe B2. The instrument has a dynamic range of ±1000 nT (compensation range ±9000 nT), a digital resolution of 31 pT and an accuracy per component of at least ±2 nT. The noise level is <10 pT/sqrt(Hz) at 1 Hz.

FGM-B2 is being built by an international consortium formed by Imperial College London (UK), the Institut für Weltraumforschung (IWF Graz, Austria) and the Technische Universität Braunschweig (Germany).

## 6 Mission Design

### 6.1 Mission Analysis

To recap, the Comet Interceptor mission aims to intercept an LPC, ideally a DNC that is approaching the inner Solar System for the first time, or even an interstellar body. The encounter involves a close-approach flyby scenario using three elements: a mother spacecraft, spacecraft A, and probes B1 and B2 carried as payloads until the flyby and delivered to different flyby trajectories. This will allow the gathering of multi-point observations of the comet and its coma.

LPCs are typically discovered as they reach the inner Solar System, no more than a few years before perihelion passage. Their orbits are therefore not known in advance. An innovative and flexible mission concept is therefore required. By waiting in an orbit around Sun-Earth Lagrange Point 2, SEL2, typically for up to 3 years before being targeted, the probability of finding a suitable LPC that can be reached in time is considerably increased. The likelihood of a suitable LPC being found during the waiting period will be greatly
<table><caption>**Table 15** The parameters of the PSU-B2 unit</caption><thead><tr><td>Parameter</td><td></td></tr></thead><tbody><tr><td>Supply Voltage (unregulated) [V]</td><td>28 V</td></tr><tr><td>Power [W]</td><td>10</td></tr><tr><td>Redundancy</td><td>No</td></tr><tr><td>Mass [kg]</td><td>0.25</td></tr><tr><td>Secondary output voltages</td><td>3.4 V, ±5 V</td></tr></tbody></table>

<figure><img src="image_33.png" /><figcaption>**Fig. 30** Picture of its sensing elements (left) and CAD rendering (right) of one FGM-B2 sensor</figcaption></figure>

### 5.10.1 FGM-B2

FGM-B2, or BFG for short, (Fig. 30) is composed of two sensors (outboard and inboard) mounted on a rigid boom, and their electronic front-ends, which are hosted in the DFP CE-BOX. Whilst both sensors and electronics have a strong heritage (e.g., Rosetta/Philae, Venus Express, and THEMIS (sensors) and MMS, Geo-KOMPSAT-2A/SOSMAG (electronics)), their mass and power consumption have been optimised for the resource-constrained Probe B2. The instrument has a dynamic range of ±1000 nT (compensation range ±9000 nT), a digital resolution of 31 pT and an accuracy per component of at least ±2 nT. The noise level is <10 pT/sqrt(Hz) at 1 Hz.

FGM-B2 is being built by an international consortium formed by Imperial College London (UK), the Institut für Weltraumforschung (IWF Graz, Austria) and the Technische Universität Braunschweig (Germany).

## 6 Mission Design

### 6.1 Mission Analysis

To recap, the Comet Interceptor mission aims to intercept an LPC, ideally a DNC that is approaching the inner Solar System for the first time, or even an interstellar body. The encounter involves a close-approach flyby scenario using three elements: a mother spacecraft, spacecraft A, and probes B1 and B2 carried as payloads until the flyby and delivered to different flyby trajectories. This will allow the gathering of multi-point observations of the comet and its coma.

LPCs are typically discovered as they reach the inner Solar System, no more than a few years before perihelion passage. Their orbits are therefore not known in advance. An innovative and flexible mission concept is therefore required. By waiting in an orbit around Sun-Earth Lagrange Point 2, SEL2, typically for up to 3 years before being targeted, the probability of finding a suitable LPC that can be reached in time is considerably increased. The likelihood of a suitable LPC being found during the waiting period will be greatly
increased soon by the availability of the LSST, with which it is expected that during its routine operations, much earlier LPC discoveries, at heliocentric distances ~20 au will be achieved, giving warning times of >5 years before targeting. This may be early enough to know the target comet before launch, although not before the mission and spacecraft designs have to be frozen.

Comet Interceptor will transfer using its on-board propulsion system to a single flyby of the target during which all science measurements will occur (Sánchez et al. 2021). Science data from probes B1 and B2 will be transmitted to spacecraft A and stored on-board. The downlink to Earth of the science data will take place in the months immediately following the comet flyby. In the unlikely event that no suitable target is identified, Comet Interceptor will transfer to a backup target from a list of known short-period comets. A request for community observations to characterise the preliminary set of mission backup targets was published by Schwamb et al. (2020). A request for observations of the revised backup targets is to be published in the near future.

## 6.1.1 Target Comet Population

The characterization of the LPC population is based on numerical studies of the evolution of these objects' orbits and comparisons with observational data, in particular from the Pan-STARRS1 survey (Wiegert and Tremaine 1999; Boe et al. 2019). The Comet Interceptor Science Consortium provided a set of 1699 LPCs with perihelion occurring inside 2 au, which has been used to derive statistical distributions for the orbit parameters of the LPC population. Figure 31 shows the cumulative distribution functions (CDF) for the perihelion distance, eccentricity and inclination. LPC orbits are quasi-parabolic with a probability of eccentricity peaking steeply very close to 1 (the minimum eccentricity in the set is 0.9504), and more likely to be retrograde (64% probability of inclination >90°). The other orbital elements follow an essentially uniform distribution. Correlations between orbital elements are neglected for the simulations carried out in the context of Comet Interceptor.

## 6.1.2 Launch, Transfer to SEL2, and Waiting Phase

Comet Interceptor is scheduled to launch in late 2029 with the ESA mission ARIEL on an Ariane 62 launcher. The launch configuration envisages the use of the Ariane 6 short fairing (TBC) and the Dual Launch Structure, with Comet Interceptor as the upper passenger and ARIEL as the lower. The injection orbit for Comet Interceptor will be a 9° inclined high-apogee, nearly parabolic orbit with perigee at 180 km altitude and apogee at 1.5 million km. The Ariane 62 will use a direct ascent strategy, with a single boost of the upper stage's Vinci engine. An additional biasing boost between both spacecraft separations will be implemented to reach slightly different SEL2 transfer trajectories, unbiased for Comet Interceptor and biased for ARIEL. This launch scenario results in a total wet launch mass for Comet Interceptor limited to ~975 kg, excluding the mass of any required launch adapter.

A large amplitude quasi-halo orbit is selected on the basis of the following arguments:

- Compatibility with ARIEL also targeting a large amplitude SEL2 orbit.

- Minimization of the ΔV required for the transfer and no need for an insertion manoeuvre.

- Mitigation of eclipses during transfer and in the orbit around SEL2.

A sample direct transfer to a large SEL2 quasi-halo and waiting phase around SEL2 are depicted in Fig. 32, showing possible orbit features and geometry.
increased soon by the availability of the LSST, with which it is expected that during its
routine operations, much earlier LPC discoveries, at heliocentric distances ~20 au will be
achieved, giving warning times of >5 years before targeting. This may be early enough to
know the target comet before launch, although not before the mission and spacecraft designs
have to be frozen.

Comet Interceptor will transfer using its on-board propulsion system to a single flyby of the target during which all science measurements will occur (Sánchez et al. 2021). Science data from probes B1 and B2 will be transmitted to spacecraft A and stored on-board. The downlink to Earth of the science data will take place in the months immediately following the comet flyby. In the unlikely event that no suitable target is identified, Comet Interceptor will transfer to a backup target from a list of known short-period comets. A request for community observations to characterise the preliminary set of mission backup targets was published by Schwamb et al. (2020). A request for observations of the revised backup targets is to be published in the near future.

6.1.1 Target Comet Population

The characterization of the LPC population is based on numerical studies of the evolution
of these objects' orbits and comparisons with observational data, in particular from the Pan-
STARRS1 survey (Wiegert and Tremaine 1999; Boe et al. 2019). The Comet Interceptor
Science Consortium provided a set of 1699 LPCs with perihelion occurring inside 2 au,
which has been used to derive statistical distributions for the orbit parameters of the LPC
population. Figure 31 shows the cumulative distribution functions (CDF) for the perihelion
distance, eccentricity and inclination. LPC orbits are quasi-parabolic with a probability of
eccentricity peaking steeply very close to 1 (the minimum eccentricity in the set is 0.9504),
and more likely to be retrograde (64% probability of inclination >90°). The other orbital
elements follow an essentially uniform distribution. Correlations between orbital elements
are neglected for the simulations carried out in the context of Comet Interceptor.

6.1.2 Launch, Transfer to SEL2, and Waiting Phase

Comet Interceptor is scheduled to launch in late 2029 with the ESA mission ARIEL on an Ariane 62 launcher. The launch configuration envisages the use of the Ariane 6 short fairing (TBC) and the Dual Launch Structure, with Comet Interceptor as the upper passenger and ARIEL as the lower. The injection orbit for Comet Interceptor will be a 9° inclined high-apogee, nearly parabolic orbit with perigee at 180 km altitude and apogee at 1.5 million km. The Ariane 62 will use a direct ascent strategy, with a single boost of the upper stage's Vinci engine. An additional biasing boost between both spacecraft separations will be implemented to reach slightly different SEL2 transfer trajectories, unbiased for Comet Interceptor and biased for ARIEL. This launch scenario results in a total wet launch mass for Comet Interceptor limited to ~975 kg, excluding the mass of any required launch adapter.

A large amplitude quasi-halo orbit is selected on the basis of the following arguments:

- Compatibility with ARIEL also targeting a large amplitude SEL2 orbit.

- Minimization of the ΔV required for the transfer and no need for an insertion manoeuvre.

- Mitigation of eclipses during transfer and in the orbit around SEL2.

A sample direct transfer to a large SEL2 quasi-halo and waiting phase around SEL2 are
depicted in Fig. 32, showing possible orbit features and geometry.
<figure><img src="image_33.png" /><figcaption>**Fig. 31** Empirical CDFs of LPCs with perihelion <2 au</figcaption></figure>
<figure><img src="image_34.png" /><figcaption>**Fig. 31** Empirical CDFs of LPCs with perihelion <2 au</figcaption></figure>
<figure><img src="image_34.png" /><figcaption>**Fig. 32** Sample transfer to SEL2 quasi-Halo orbit and waiting phase in the Sun-Earth rotating frame</figcaption></figure>

The transfer geometry and amplitude of the achievable quasi-halo orbit will depend on the launch date and time. Assuming that a single Ariane 62 flight program is used, the natural variation of the perigee velocity over the launch window will have to be corrected with a manoeuvre 2 days into the mission. This manoeuvre will be combined with the correction of launcher injection errors. More trajectory correction manoeuvres are planned at days 5 and 20 to achieve an accurate manoeuvre-free transfer into the quasi-Halo orbit. Including the deterministic and the stochastic parts, 50 m/s are allocated overall, for the trajectory corrections during this part of the mission. The overall transfer duration is ~3–4 months, during which commissioning activities for the spacecraft and probes will be performed.

Comet Interceptor will orbit SEL2 for a currently-unknown duration, likely between a few months and 4 years. It is during this waiting time that Earth-based observatories are expected to discover one or more potential targets for the mission, if one was not found prior to launch. Following target selection, Comet Interceptor will remain near SEL2 waiting for the right conditions to start the transfer towards the target comet.

Quasi-Halo orbits around SEL2, up to an amplitude of ~1 million km, have a period of ~180 days. These orbits are inherently unstable; any small perturbation will lead to an exponential deviation from the reference orbit, hence periodic station keeping manoeuvres are planned. The frequency and size of these manoeuvres depends on the spacecraft's ability to reduce velocity perturbations and dynamic noise. The current plan considers station keeping manoeuvres every 28 days and allocates 2.3 m/s per year to stay at SEL2.

### 6.1.3 Transfer from SEL2 to Encounter

Newly detected potential targets will be monitored continuously from Earth, and the possible transfer trajectories and encounter/post-encounter profiles will be studied in detail for each candidate a minimum of 6 months in advance of initiating the transfer. Once the comet target is selected, or a decision is made to go to a backup, the spacecraft will wait until the optimum time to depart from its orbit around SEL2.
<figure><img src="image_35.png" /><figcaption>**Fig. 32** Sample transfer to SEL2 quasi-Halo orbit and waiting phase in the Sun-Earth rotating frame</figcaption></figure>

The transfer geometry and amplitude of the achievable quasi-halo orbit will depend on the launch date and time. Assuming that a single Ariane 62 flight program is used, the natural variation of the perigee velocity over the launch window will have to be corrected with a manoeuvre 2 days into the mission. This manoeuvre will be combined with the correction of launcher injection errors. More trajectory correction manoeuvres are planned at days 5 and 20 to achieve an accurate manoeuvre-free transfer into the quasi-Halo orbit. Including the deterministic and the stochastic parts, 50 m/s are allocated overall, for the trajectory corrections during this part of the mission. The overall transfer duration is ~3–4 months, during which commissioning activities for the spacecraft and probes will be performed.

Comet Interceptor will orbit SEL2 for a currently-unknown duration, likely between a few months and 4 years. It is during this waiting time that Earth-based observatories are expected to discover one or more potential targets for the mission, if one was not found prior to launch. Following target selection, Comet Interceptor will remain near SEL2 waiting for the right conditions to start the transfer towards the target comet.

Quasi-Halo orbits around SEL2, up to an amplitude of ~1 million km, have a period of ~180 days. These orbits are inherently unstable; any small perturbation will lead to an exponential deviation from the reference orbit, hence periodic station keeping manoeuvres are planned. The frequency and size of these manoeuvres depends on the spacecraft's ability to reduce velocity perturbations and dynamic noise. The current plan considers station keeping manoeuvres every 28 days and allocates 2.3 m/s per year to stay at SEL2.

### 6.1.3 Transfer from SEL2 to Encounter

Newly detected potential targets will be monitored continuously from Earth, and the possible transfer trajectories and encounter/post-encounter profiles will be studied in detail for each candidate a minimum of 6 months in advance of initiating the transfer. Once the comet target is selected, or a decision is made to go to a backup, the spacecraft will wait until the optimum time to depart from its orbit around SEL2.
An optimal transfer trajectory avoids costly out-of-plane $\Delta V$ manoeuvres and stays close to the ecliptic plane. As a result, the target comet has to be intercepted at one of its ecliptic nodal crossings, hence the encounter location can be defined by just 2 parameters: the heliocentric distance at encounter $R_c$ and the phase angle of Earth at encounter $\theta$, i.e. Comet-Sun-Earth angle. Analysed by 2-body dynamics under solar gravity, the transfer orbit needs to have the perihelion and/or the aphelion distance adjusted in such a way that $R_c$ can be reached, together with an orbital period such that the phase drift, ahead or behind Earth, that leads to the desired angle $\theta$ in a given transfer time. Typically, this requires $\le 2$ trajectory manoeuvres, though in a few cases a third manoeuvre can be beneficial. At the beginning of the transfer, when the spacecraft leaves SEL2, the gravity effect of Earth has a significant impact on the trajectory. Extensive analysis leads to two different strategies being envisaged:

1. *Direct Transfer.* The spacecraft performs a manoeuvre to leave the SEL2 parking orbit in order to exit the gravitational pull of the Earth-Moon system and is injected into a heliocentric orbit drifting towards encounter. A second deep-space manoeuvre, during the transfer orbit to the comet might be necessary to adjust the orbit or the phasing. These transfers can be Exterior, when the spacecraft leaves directly towards the outside of the Sun-Earth direction, or Interior, leaving towards Earth, performing a high-altitude Earth flyby before leaving in the direction of Sun-Earth Lagrange Point 1. The complex dynamics of the interior case can exploit multiple loops around the Earth and/or an Earth flyby to reduce the transfer $\Delta V$.

2. *Moon Gravity Assist.* The dynamics of the SEL2 manifold towards Earth allow a Moon flyby to be performed, after which the spacecraft can escape from Earth with a velocity at infinity of ~1-1.4 km/s and in a direction approximately opposite to the Earth's velocity vector. This is an efficient way to reach heliocentric orbits with perihelion <1 au and favours targets with negative phase angle at encounter ($\theta < 0°$, ahead of Earth). The Moon flyby allows $\Delta V$ savings, but introduces additional operational complexity.

The reachable domain of comet encounters for each strategy (Fig. 33), is driven by a trade between transfer time and $\Delta V$. Direct transfers favour $R_c > 1$ au and $\theta > 0°$ (behind Earth), while the opposite occurs for transfers with Moon gravity assist, whereas a region of overlap exists in which both strategies are feasible. Increasing the transfer time significantly impacts the reachable domain.

Figure 34 illustrates sample transfer trajectories to an identified backup target. Two transfer trajectories are shown: one optimised for minimum $\Delta V$, requiring 37 m/s and 847 days, and one optimised for minimum transfer time with a $\Delta V$ cap at 570 m/s that reduces the transfer duration to 529 days. In both cases the exterior direct transfer strategy is used, and the transfer requires 2 manoeuvres, one to depart from SEL2 and one deep space manoeuvre during the cruise towards the comet encounter.

It is likely that Comet Interceptor will be able to slightly adjust its transfer trajectory in order to also encounter a suitable inactive minor body. Such an additional flyby would be a good opportunity for an engineering test of the spacecraft systems and operational procedures required for the comet flyby in a similar scenario, i.e., the optical navigation cameras and autonomous tracking. This would provide valuable experience to increase the mission robustness and probability of success of the actual science flyby.

### 6.1.4 Probabilistic Reachability Analysis and Statistics of Key Mission Parameters

The results presented in this section have been obtained using a Monte Carlo tool that simulates possible Comet Interceptor missions, modelling from the target detection process to
An optimal transfer trajectory avoids costly out-of-plane $\Delta V$ manoeuvres and stays close to the ecliptic plane. As a result, the target comet has to be intercepted at one of its ecliptic nodal crossings, hence the encounter location can be defined by just 2 parameters: the heliocentric distance at encounter $R_c$ and the phase angle of Earth at encounter $\theta$, i.e. Comet-Sun-Earth angle. Analysed by 2-body dynamics under solar gravity, the transfer orbit needs to have the perihelion and/or the aphelion distance adjusted in such a way that $R_c$ can be reached, together with an orbital period such that the phase drift, ahead or behind Earth, that leads to the desired angle $\theta$ in a given transfer time. Typically, this requires $\le 2$ trajectory manoeuvres, though in a few cases a third manoeuvre can be beneficial. At the beginning of the transfer, when the spacecraft leaves SEL2, the gravity effect of Earth has a significant impact on the trajectory. Extensive analysis leads to two different strategies being envisaged:

1. *Direct Transfer.* The spacecraft performs a manoeuvre to leave the SEL2 parking orbit in order to exit the gravitational pull of the Earth-Moon system and is injected into a heliocentric orbit drifting towards encounter. A second deep-space manoeuvre, during the transfer orbit to the comet might be necessary to adjust the orbit or the phasing. These transfers can be Exterior, when the spacecraft leaves directly towards the outside of the Sun-Earth direction, or Interior, leaving towards Earth, performing a high-altitude Earth flyby before leaving in the direction of Sun-Earth Lagrange Point 1. The complex dynamics of the interior case can exploit multiple loops around the Earth and/or an Earth flyby to reduce the transfer $\Delta V$.

2. *Moon Gravity Assist.* The dynamics of the SEL2 manifold towards Earth allow a Moon flyby to be performed, after which the spacecraft can escape from Earth with a velocity at infinity of ~1-1.4 km/s and in a direction approximately opposite to the Earth's velocity vector. This is an efficient way to reach heliocentric orbits with perihelion <1 au and favours targets with negative phase angle at encounter ($\theta < 0°$, ahead of Earth). The Moon flyby allows $\Delta V$ savings, but introduces additional operational complexity.

The reachable domain of comet encounters for each strategy (Fig. 33), is driven by a trade between transfer time and $\Delta V$. Direct transfers favour $R_c > 1$ au and $\theta > 0°$ (behind Earth), while the opposite occurs for transfers with Moon gravity assist, whereas a region of overlap exists in which both strategies are feasible. Increasing the transfer time significantly impacts the reachable domain.

Figure 34 illustrates sample transfer trajectories to an identified backup target. Two transfer trajectories are shown: one optimised for minimum $\Delta V$, requiring 37 m/s and 847 days, and one optimised for minimum transfer time with a $\Delta V$ cap at 570 m/s that reduces the transfer duration to 529 days. In both cases the exterior direct transfer strategy is used, and the transfer requires 2 manoeuvres, one to depart from SEL2 and one deep space manoeuvre during the cruise towards the comet encounter.

It is likely that Comet Interceptor will be able to slightly adjust its transfer trajectory in order to also encounter a suitable inactive minor body. Such an additional flyby would be a good opportunity for an engineering test of the spacecraft systems and operational procedures required for the comet flyby in a similar scenario, i.e., the optical navigation cameras and autonomous tracking. This would provide valuable experience to increase the mission robustness and probability of success of the actual science flyby.

### 6.1.4 Probabilistic Reachability Analysis and Statistics of Key Mission Parameters

The results presented in this section have been obtained using a Monte Carlo tool that simulates possible Comet Interceptor missions, modelling from the target detection process to
<figure><img src="image_35.png" /><figcaption>**Fig. 33** Reachable $R_c-\theta$ regions with 750 m/s for sample transfer times of 1 year (top) and 3 years (bottom)</figcaption></figure>

the transfer and comet encounter and the following post-encounter phase. The simulations consider a rate of 14 LPCs per year with perihelia <2 au originating from the comet population of Sect. 6.1.1. This underlying assumption is consistent with the historical observations of 21 new LPCs in the 2010–2019 decade with nodal crossings in the accessible 0.9–1.25 au range.

The probability of finding at least one feasible LPC target within the given set of constraints and mission requirements is used as a figure of merit. Figure 35 shows how the mission duration and the transfer $\Delta V$ impact this probability. With the allocations at the time of writing of 600 m/s for the transfer and 6 years overall mission duration the probability is 80% (30% of single target plus 50% of multiple targets). Excluding the transfer option with the Moon gravity assist reduces the probability significantly to 63%.
<figure><img src="image_36.png" /><figcaption>**Fig. 33** Reachable $R_c-\theta$ regions with 750 m/s for sample transfer times of 1 year (top) and 3 years (bottom)</figcaption></figure>

the transfer and comet encounter and the following post-encounter phase. The simulations consider a rate of 14 LPCs per year with perihelia <2 au originating from the comet population of Sect. 6.1.1. This underlying assumption is consistent with the historical observations of 21 new LPCs in the 2010–2019 decade with nodal crossings in the accessible 0.9–1.25 au range.

The probability of finding at least one feasible LPC target within the given set of constraints and mission requirements is used as a figure of merit. Figure 35 shows how the mission duration and the transfer $\Delta V$ impact this probability. With the allocations at the time of writing of 600 m/s for the transfer and 6 years overall mission duration the probability is 80% (30% of single target plus 50% of multiple targets). Excluding the transfer option with the Moon gravity assist reduces the probability significantly to 63%.
**Fig. 34** Example geometry of sample transfers to comet 26P, which was an original backup target for the mission. a) Projection on Sun-Earth rotating frame. Solid blue line: minimum ΔV; dashed orange line: minimum time of flight. b) Zoom of SEL2 departure. c) Distances to Earth and Sun; dots every 15 days; 6 months post-encounter phase

<figure><img src="image_36.png" /></figure>

**Fig. 35** Influence of ΔV and mission duration on the probability of at least one LPC target. Time between launch and target detection <2 years, mission duration includes 6 months post-encounter phase

<figure><img src="image_37.png" /></figure>

The Monte Carlo mission simulator provided statistical information on durations relevant to the mission. We have to consider that missions finding multiple feasible targets might choose to favour a given parameter, thus two limiting cases are studied. The main results are summarised in Fig. 36 in which the case of the mission eventually intercepting a backup comet target has been disregarded. The median waiting time at SEL2 is 1–2 years, while waiting times >4 years rarely occur (<5% of cases).

For transfer from SEL2 to the comet, when aiming for the shortest transfer there is a preference for heliocentric transfers of durations close to an integer number of years. Aiming for
**Fig. 34** Example geometry of sample transfers to comet 26P, which was an original backup target for the mission. a) Projection on Sun-Earth rotating frame. Solid blue line: minimum ΔV; dashed orange line: minimum time of flight. b) Zoom of SEL2 departure. c) Distances to Earth and Sun; dots every 15 days; 6 months post-encounter phase

<figure><img src="image_37.png" /></figure>

**Fig. 35** Influence of ΔV and mission duration on the probability of at least one LPC target. Time between launch and target detection <2 years, mission duration includes 6 months post-encounter phase

<figure><img src="image_38.png" /></figure>

The Monte Carlo mission simulator provided statistical information on durations relevant to the mission. We have to consider that missions finding multiple feasible targets might choose to favour a given parameter, thus two limiting cases are studied. The main results are summarised in Fig. 36 in which the case of the mission eventually intercepting a backup comet target has been disregarded. The median waiting time at SEL2 is 1–2 years, while waiting times >4 years rarely occur (<5% of cases).

For transfer from SEL2 to the comet, when aiming for the shortest transfer there is a preference for heliocentric transfers of durations close to an integer number of years. Aiming for
**Fig. 36** Statistics of time parameters relevant for the mission. a) Shortest wait at SEL2; b) Longest wait at SEL2; c) Transfer time from SEL2 to comet encounter; d) Duration from launch to comet encounter

<figure><img src="image_38.png" /></figure>

**Fig. 37** Relative encounter speed (left) and flyby solar aspect angle (right) for reachable LPCs (0.7 au < Rc < 1.3 au)

<figure><img src="image_39.png" /></figure>

the longest transfer tends to smooth out the peaks and to result in more uniform distribution. The median of this parameter is 1.5–2.5 years.

For the mission duration from launch to comet encounter (limited to 5.5 years assuming 6 years maximum overall duration, including 6 months of post-encounter activities), the statistics show an increasing probability density followed by a flat region for durations >4 years. The median is observed ~3.5 years and the 90th percentile at ~5 years.

### 6.1.5 Comet Encounter

Figure 37 shows the distribution of the encounter relative speed, and the flyby solar aspect angle (the angle between the Comet-Sun vector and the relative velocity vector), for simulated feasible encounters obtained from the population of LPCs. The relative velocity is biased towards higher values and peaks around 60 km/s. The flyby solar aspect angle shows a symmetrical distribution around 90°. Constraining the encounter parameters has an impact on the availability of targets: having a 60 km/s maximum velocity would remove 33% of possible targets, while the baselined 70 km/s requirement removes only 8.5%. On the other hand, the requirement that constrains the flyby solar aspect angle to $90^\circ \pm 45^\circ$ removes ~7.5% of the targets.

In addition, it must be pointed out that the orbital mechanics of an LPC encounter at a given heliocentric distance, from 0.9 to 1.2 au, constrain the feasible combinations of relative speed and flyby solar aspect angle, as depicted in Fig. 38. The solar aspect angle provides information directly as to whether the encounter is on the inbound or the outbound leg of the comet's orbit, with the angle being >90° or <90°, respectively.
**Fig. 36** Statistics of time parameters relevant for the mission. a) Shortest wait at SEL2; b) Longest wait at SEL2; c) Transfer time from SEL2 to comet encounter; d) Duration from launch to comet encounter

<figure><img src="image_39.png" /></figure>

**Fig. 37** Relative encounter speed (left) and flyby solar aspect angle (right) for reachable LPCs (0.7 au < Rc < 1.3 au)

<figure><img src="image_40.png" /></figure>

the longest transfer tends to smooth out the peaks and to result in more uniform distribution. The median of this parameter is 1.5–2.5 years.

For the mission duration from launch to comet encounter (limited to 5.5 years assuming 6 years maximum overall duration, including 6 months of post-encounter activities), the statistics show an increasing probability density followed by a flat region for durations >4 years. The median is observed ~3.5 years and the 90th percentile at ~5 years.

### 6.1.5 Comet Encounter

Figure 37 shows the distribution of the encounter relative speed, and the flyby solar aspect angle (the angle between the Comet-Sun vector and the relative velocity vector), for simulated feasible encounters obtained from the population of LPCs. The relative velocity is biased towards higher values and peaks around 60 km/s. The flyby solar aspect angle shows a symmetrical distribution around 90°. Constraining the encounter parameters has an impact on the availability of targets: having a 60 km/s maximum velocity would remove 33% of possible targets, while the baselined 70 km/s requirement removes only 8.5%. On the other hand, the requirement that constrains the flyby solar aspect angle to $90^\circ \pm 45^\circ$ removes ~7.5% of the targets.

In addition, it must be pointed out that the orbital mechanics of an LPC encounter at a given heliocentric distance, from 0.9 to 1.2 au, constrain the feasible combinations of relative speed and flyby solar aspect angle, as depicted in Fig. 38. The solar aspect angle provides information directly as to whether the encounter is on the inbound or the outbound leg of the comet's orbit, with the angle being >90° or <90°, respectively.
<figure><img src="image_40.png" /><figcaption>**Fig. 38** Allowed regions of relative velocities and flyby solar aspect angles</figcaption></figure>

The approach to encounter phase is assumed to be defined as the last 60 days before the flyby. By trajectory design there will be no need for a deterministic manoeuvre during this phase, hence operations can focus on the navigation required to reach the comet. This navigation will rely on ground-based measurements of the comet's position, radio tracking of the spacecraft using the ESTRACK deep space antennas and, most importantly, on the optical data from the navigation cameras (see Sect. 6.1.7). The optical observations have the strength to directly relate the states of the spacecraft and target comet, improving the accuracy of the prediction of the flyby location and time, and allowing critical trajectory correction manoeuvres (TCMs) to be performed.

The flyby targets are defined in the B-plane, i.e. the plane perpendicular to the relative velocity $\vec{v}_{rel}$ and passes through the comet nucleus. Two perpendicular directions are defined: the T vector is the orthogonal projection of the Sun-to-target vector onto the B-plane. The R vector completes an orthogonal right-handed triad with $S = \vec{v}_{rel}$ and T. The B-plane targets for the spacecraft and probes are defined by the closest approach distance and the angle $\theta$, with the T-axis measured in the direction towards the R-axis.

The final approach operations will take place according to the timeline illustrated in Fig. 39, which is still subject of refinement and optimisation in future phases of the mission design.

- The navigation camera will be used continuously to improve the determination of the nucleus position and the flyby accuracy. A ground turn-around time of 12 h (seen as the data cut-off time before each TCM) is considered necessary to downlink the last image, perform the on-ground processing, orbit determination and next manoeuvre and/or separation planning, and to uplink telecommands. Therefore, the input data cut-off for each TCM is 12 hours beforehand.

- The spacecraft composite (A, B1 and B2) is assumed to be targeted at the B1 aim point, with a closest approach of 850 km and $\theta = 135^\circ$, thanks to navigation during the approach phase.

- A stochastic TCM at -44 hours to the flyby will target the composite precisely at the B1 aim point, making use of the most updated optical observations.

- Separation of Probe B1 will occur 2 hours later, following a post-TCM tranquilisation phase and a slew to the separation attitude.

- TCM @ 30 hours to the flyby will target the platforms A+B2 towards the aim point of Probe B2 at closest approach of 400 km and $\theta = 180^\circ$. This TCM will combine a deterministic part of ~6 m/s with a stochastic correction.
<figure><img src="image_41.png" /><figcaption>**Fig. 38** Allowed regions of relative velocities and flyby solar aspect angles</figcaption></figure>

The approach to encounter phase is assumed to be defined as the last 60 days before the flyby. By trajectory design there will be no need for a deterministic manoeuvre during this phase, hence operations can focus on the navigation required to reach the comet. This navigation will rely on ground-based measurements of the comet's position, radio tracking of the spacecraft using the ESTRACK deep space antennas and, most importantly, on the optical data from the navigation cameras (see Sect. 6.1.7). The optical observations have the strength to directly relate the states of the spacecraft and target comet, improving the accuracy of the prediction of the flyby location and time, and allowing critical trajectory correction manoeuvres (TCMs) to be performed.

The flyby targets are defined in the B-plane, i.e. the plane perpendicular to the relative velocity $\vec{v}_{rel}$ and passes through the comet nucleus. Two perpendicular directions are defined: the T vector is the orthogonal projection of the Sun-to-target vector onto the B-plane. The R vector completes an orthogonal right-handed triad with $S = \vec{v}_{rel}$ and T. The B-plane targets for the spacecraft and probes are defined by the closest approach distance and the angle $\theta$, with the T-axis measured in the direction towards the R-axis.

The final approach operations will take place according to the timeline illustrated in Fig. 39, which is still subject of refinement and optimisation in future phases of the mission design.

- The navigation camera will be used continuously to improve the determination of the nucleus position and the flyby accuracy. A ground turn-around time of 12 h (seen as the data cut-off time before each TCM) is considered necessary to downlink the last image, perform the on-ground processing, orbit determination and next manoeuvre and/or separation planning, and to uplink telecommands. Therefore, the input data cut-off for each TCM is 12 hours beforehand.

- The spacecraft composite (A, B1 and B2) is assumed to be targeted at the B1 aim point, with a closest approach of 850 km and $\theta = 135^\circ$, thanks to navigation during the approach phase.

- A stochastic TCM at -44 hours to the flyby will target the composite precisely at the B1 aim point, making use of the most updated optical observations.

- Separation of Probe B1 will occur 2 hours later, following a post-TCM tranquilisation phase and a slew to the separation attitude.

- TCM @ 30 hours to the flyby will target the platforms A+B2 towards the aim point of Probe B2 at closest approach of 400 km and $\theta = 180^\circ$. This TCM will combine a deterministic part of ~6 m/s with a stochastic correction.
<figure><img src="image_41.png" /><figcaption>**Fig. 39** Timeline of operations for the comet flyby</figcaption></figure>

- Separation of Probe B2 occurs 6 hours later.

- The diversion manoeuvre of spacecraft A occurs –20 hours to the flyby. It targets a greater closest approach distance of 1000 km and $\theta = 180^\circ$. The deterministic part of the diversion manoeuvre is 8.5 m/s.

### 6.1.6 Back-up Targets

It is critical for the mission's success that the availability of a backup target for any launch date is ensured. This has been investigated for a down-selected list of 11 comet candidates provided by the Science Consortium and for the 4-year launch timeframe 2029–2032. The analysis identified 3 backup targets compatible with a transfer $\Delta V$ allocation of 600 m/s and an overall mission duration of 6 years, chronologically as follows:

1. From 2029-03-24 to 2030-05-18: 15P/Finlay

2. From 2030-05-18 to 2030-12-14: 289P/Blanpain

3. From 2030-12-14 onwards: 300P/Catalina

Figure 40 shows a summary of the backup target selection. As an example, for launch on January 1st, 2030, the backup target is 15P/Finlay and the latest selection of a primary LPC target needs to occur before ~3 years after launch. Otherwise, a transfer to 15P will be used, which requires departure ~3.5 years after launch and arrives in September 2034; 6 months later, or 5.2 years after launch, the mission will be finished. We observe that more than 3 years are provided consistently from launch until the decision to go to the backup target. Only relatively short periods, in Q1–Q2 2030 and May 2032, provide shorter decision cut-off times of between 2.5 and 3 years.

### 6.1.7 Spacecraft Design Drivers and Schedule

The mission design is driven by the key objective of performing multi-point observations during a high relative velocity flyby with a target which will be identified after the finalisation of the spacecraft design. An Engineering Dust Coma Model, EDCM, was developed by members of the mission's proposing consortium, to assess the risk from dust particle impacts onto the spacecraft (Marschall et al. 2022), the results of which are available at https://doi.org/10.5281/zenodo.6906815. The main design drivers of spacecraft A and probe B2 are summarised in Table 16.

Following adoption of the mission by ESA, the project team carried out the satellite level Preliminary Design Review (PDR). For the mission's definition phase, two consortia were contracted to provide designs for spacecraft A and probe B2; these were led by OHB-IT
<figure><img src="image_42.png" /><figcaption>**Fig. 39** Timeline of operations for the comet flyby</figcaption></figure>

- Separation of Probe B2 occurs 6 hours later.

- The diversion manoeuvre of spacecraft A occurs –20 hours to the flyby. It targets a greater closest approach distance of 1000 km and $\theta = 180^\circ$. The deterministic part of the diversion manoeuvre is 8.5 m/s.

### 6.1.6 Back-up Targets

It is critical for the mission's success that the availability of a backup target for any launch date is ensured. This has been investigated for a down-selected list of 11 comet candidates provided by the Science Consortium and for the 4-year launch timeframe 2029–2032. The analysis identified 3 backup targets compatible with a transfer $\Delta V$ allocation of 600 m/s and an overall mission duration of 6 years, chronologically as follows:

1. From 2029-03-24 to 2030-05-18: 15P/Finlay

2. From 2030-05-18 to 2030-12-14: 289P/Blanpain

3. From 2030-12-14 onwards: 300P/Catalina

Figure 40 shows a summary of the backup target selection. As an example, for launch on January 1st, 2030, the backup target is 15P/Finlay and the latest selection of a primary LPC target needs to occur before ~3 years after launch. Otherwise, a transfer to 15P will be used, which requires departure ~3.5 years after launch and arrives in September 2034; 6 months later, or 5.2 years after launch, the mission will be finished. We observe that more than 3 years are provided consistently from launch until the decision to go to the backup target. Only relatively short periods, in Q1–Q2 2030 and May 2032, provide shorter decision cut-off times of between 2.5 and 3 years.

### 6.1.7 Spacecraft Design Drivers and Schedule

The mission design is driven by the key objective of performing multi-point observations during a high relative velocity flyby with a target which will be identified after the finalisation of the spacecraft design. An Engineering Dust Coma Model, EDCM, was developed by members of the mission's proposing consortium, to assess the risk from dust particle impacts onto the spacecraft (Marschall et al. 2022), the results of which are available at https://doi.org/10.5281/zenodo.6906815. The main design drivers of spacecraft A and probe B2 are summarised in Table 16.

Following adoption of the mission by ESA, the project team carried out the satellite level Preliminary Design Review (PDR). For the mission's definition phase, two consortia were contracted to provide designs for spacecraft A and probe B2; these were led by OHB-IT
<figure><img src="image_42.png" /><figcaption>**Fig. 40** Summary of backup target selected as a function of the launch date. As the launch date at the time of writing is in Q4 of 2029, 26P is no longer reachable</figcaption></figure>

and a TAS-UK. On completion of the PDR, the OHB-IT-led consortium was selected by ESA as the prime contractor. The design proposed by the selected OHB-led consortium is represented in Fig. 41. Given the fast development approach followed by Comet Interceptor, the design solutions had to rely on existing platform heritage, minimizing the need for qualification activities. The total spacecraft mass, including propellant and margins is limited to 975 kg by the presently estimated launcher performance (for a dual launch with Ariel). The overall dimensions of the stowed spacecraft are ~1600 x 1600 x 1500 mm. The configuration of spacecraft A includes:

- a cuboid shape, hosting all platform equipment and accommodating payload and probes.

- the Attitude, Orbit, Guidance and Navigation Control (AOGNC) subsystem, including different sensors (Coarse Sun Sensor, Gyro, Star Trackers, Navigation Cameras) and actuators (Reaction Wheels and Reaction Control System).

- a chemical propulsion system with a large ΔV capability.

- a communication system based on a fixed High Gain Antenna operating in X band and including an inter-satellite link operating in S band.

- two deployable solar arrays.

- a thermal control system, based on a classic passive design.

- a dust shield protecting spacecraft A and probes during the encounter phase on the ram face.

- two probes accommodated on the same spacecraft side.

The chemical propulsion system is designed to provide, within the maximum allowed launch mass, a minimum transfer ΔV capability of 600 $m s^{-1}$ (see Sect. 6.1.4); in addition,
<figure><img src="image_43.png" /><figcaption>**Fig. 40** Summary of backup target selected as a function of the launch date. As the launch date at the time of writing is in Q4 of 2029, 26P is no longer reachable</figcaption></figure>

and a TAS-UK. On completion of the PDR, the OHB-IT-led consortium was selected by ESA as the prime contractor. The design proposed by the selected OHB-led consortium is represented in Fig. 41. Given the fast development approach followed by Comet Interceptor, the design solutions had to rely on existing platform heritage, minimizing the need for qualification activities. The total spacecraft mass, including propellant and margins is limited to 975 kg by the presently estimated launcher performance (for a dual launch with Ariel). The overall dimensions of the stowed spacecraft are ~1600 x 1600 x 1500 mm. The configuration of spacecraft A includes:

- a cuboid shape, hosting all platform equipment and accommodating payload and probes.

- the Attitude, Orbit, Guidance and Navigation Control (AOGNC) subsystem, including different sensors (Coarse Sun Sensor, Gyro, Star Trackers, Navigation Cameras) and actuators (Reaction Wheels and Reaction Control System).

- a chemical propulsion system with a large ΔV capability.

- a communication system based on a fixed High Gain Antenna operating in X band and including an inter-satellite link operating in S band.

- two deployable solar arrays.

- a thermal control system, based on a classic passive design.

- a dust shield protecting spacecraft A and probes during the encounter phase on the ram face.

- two probes accommodated on the same spacecraft side.

The chemical propulsion system is designed to provide, within the maximum allowed launch mass, a minimum transfer ΔV capability of 600 $m s^{-1}$ (see Sect. 6.1.4); in addition,
<table><caption>**Table 16** Main design drivers for the Comet Interceptor spacecraft</caption><thead><tr><td>Design driver</td><td>Main implications</td></tr></thead><tbody><tr><td>Dual launch with ARIEL on A62</td><td>Max launch mass limited to 975 kg</td></tr><tr><td>Multi-point observation principle</td><td>Additional probes to be carried by the main spacecraft</td></tr><tr><td>Large payload complement</td><td>Accommodation of several <i>in situ</i> and remote sensing units on spacecraft A and Probe B2</td></tr><tr><td>Target defined at late stage</td><td>Spacecraft design compatible with range of possible targets, of encounter conditions and Sun-Earth-Target geometries</td></tr><tr><td>Maximise probability to reach a suitable target</td><td>Maximise ΔV capability. Navigation & Target tracking capabilities to remain compatible with multiple targets. High maximum flyby relative velocity (range 10 to 70 km/s)</td></tr><tr><td>Interplanetary mission</td><td>Spacecraft operating at ~1–2 au from Earth</td></tr><tr><td>Measurements performed during a high relative velocity flyby</td><td>“One shot” science. Data downlinked to Earth after the closest approach</td></tr><tr><td>Comet environment</td><td>Capability to survive the dust environment for a variety of possible targets</td></tr><tr><td>Programmatic constraints as from F-Mission call</td><td>Cost at completion boundaries. Fast development track. Incompatibility with dedicated technology developments and need to rely on existing, flight qualified solutions</td></tr></tbody></table>

<figure><img src="image_43.png" /><figcaption>**Fig. 41** Comet Interceptor spacecraft A and probe B2 design concepts (OHB)</figcaption></figure>

should the Ariane 6.2 performance improve by the time of Comet Interceptor's launch, the capability to load additional propellant, thus exceeding the minimum required ΔV performance, is requested. It is noted that during the flyby the target is maintained in the field of view of the high-resolution camera CoCa via the dedicated Rotating Mirror Assembly (Sect. 5.1), while the spacecraft maintains inertial attitude, maintaining the dust shield in the relative velocity direction.

Probes B1 and B2 are both smaller craft, without propulsion capability, and are deployed from spacecraft A via dedicated separation systems. Both B1 and B2 carry dust shields for protection during the encounter, and will transmit data from their scientific instruments to
<table><caption>**Table 16** Main design drivers for the Comet Interceptor spacecraft</caption><thead><tr><td>Design driver</td><td>Main implications</td></tr></thead><tbody><tr><td>Dual launch with ARIEL on A62</td><td>Max launch mass limited to 975 kg</td></tr><tr><td>Multi-point observation principle</td><td>Additional probes to be carried by the main spacecraft</td></tr><tr><td>Large payload complement</td><td>Accommodation of several <i>in situ</i> and remote sensing units on spacecraft A and Probe B2</td></tr><tr><td>Target defined at late stage</td><td>Spacecraft design compatible with range of possible targets, of encounter conditions and Sun-Earth-Target geometries</td></tr><tr><td>Maximise probability to reach a suitable target</td><td>Maximise ΔV capability. Navigation & Target tracking capabilities to remain compatible with multiple targets. High maximum flyby relative velocity (range 10 to 70 km/s)</td></tr><tr><td>Interplanetary mission</td><td>Spacecraft operating at ~1–2 au from Earth</td></tr><tr><td>Measurements performed during a high relative velocity flyby</td><td>“One shot” science. Data downlinked to Earth after the closest approach</td></tr><tr><td>Comet environment</td><td>Capability to survive the dust environment for a variety of possible targets</td></tr><tr><td>Programmatic constraints as from F-Mission call</td><td>Cost at completion boundaries. Fast development track. Incompatibility with dedicated technology developments and need to rely on existing, flight qualified solutions</td></tr></tbody></table>

<figure><img src="image_44.png" /><figcaption>**Fig. 41** Comet Interceptor spacecraft A and probe B2 design concepts (OHB)</figcaption></figure>

should the Ariane 6.2 performance improve by the time of Comet Interceptor's launch, the capability to load additional propellant, thus exceeding the minimum required ΔV performance, is requested. It is noted that during the flyby the target is maintained in the field of view of the high-resolution camera CoCa via the dedicated Rotating Mirror Assembly (Sect. 5.1), while the spacecraft maintains inertial attitude, maintaining the dust shield in the relative velocity direction.

Probes B1 and B2 are both smaller craft, without propulsion capability, and are deployed from spacecraft A via dedicated separation systems. Both B1 and B2 carry dust shields for protection during the encounter, and will transmit data from their scientific instruments to
<figure><img src="image_44.png" /><figcaption>**Table 17** The key dates for the Comet Interceptor high level schedule</figcaption></figure>

<table><thead><tr><td>Milestone</td><td>Date</td></tr></thead><tbody><tr><td>Selection as F1 mission</td><td>June 2019</td></tr><tr><td>PDR</td><td>Q2 2022</td></tr><tr><td>Mission Adoption</td><td>June 2022</td></tr><tr><td>Prime selection</td><td>Q4 2022</td></tr><tr><td>Start of phase C/D</td><td>Jan 2023</td></tr><tr><td>CDR</td><td>July 2023<br>(Instruments)<br>Q4 2024 (System)</td></tr><tr><td>Delivery of payload flight units</td><td>Q4 2025 (probe B2)<br>Q1 2026 (spacecraft A)<br>Q1 2026 (probe B1)</td></tr><tr><td>QAR</td><td>Dec. 2027</td></tr><tr><td>Launch ready</td><td>Mid-2028</td></tr><tr><td>Launch (L)</td><td>Dec. 2029<br>(shared with ARIEL)</td></tr><tr><td>Arrival at L2</td><td>L + 4 months</td></tr><tr><td>Waiting at L2</td><td>Maximum 5 years</td></tr><tr><td>Transfer to target Comet</td><td></td></tr><tr><td>Comet flyby</td><td>Latest L + 5.5 years</td></tr><tr><td>End of Operations</td><td>Latest L + 6 years</td></tr></tbody></table>

spacecraft A via dedicated S-band inter-satellite links. The masses of both probes are ~35 kg each, with a typical diameter/size of ~0.5 m.

B1 has a cuboid shape, and is 3-axis stabilised through a dedicated Attitude, Guidance and navigation sub-system (Fig. 42). Its electrical power system is based on deployable solar arrays and a secondary battery. B2 has an axisymmetric shape, is gyroscopically stabilised and has no attitude control capability. Its power system based on a primary battery, with an operational lifetime of ~30 hours.
<figure><img src="image_45.png" /><figcaption>**Table 17** The key dates for the Comet Interceptor high level schedule</figcaption></figure>

<table><thead><tr><td>Milestone</td><td>Date</td></tr></thead><tbody><tr><td>Selection as F1 mission</td><td>June 2019</td></tr><tr><td>PDR</td><td>Q2 2022</td></tr><tr><td>Mission Adoption</td><td>June 2022</td></tr><tr><td>Prime selection</td><td>Q4 2022</td></tr><tr><td>Start of phase C/D</td><td>Jan 2023</td></tr><tr><td>CDR</td><td>July 2023<br>(Instruments)<br>Q4 2024 (System)</td></tr><tr><td>Delivery of payload flight units</td><td>Q4 2025 (probe B2)<br>Q1 2026 (spacecraft A)<br>Q1 2026 (probe B1)</td></tr><tr><td>QAR</td><td>Dec. 2027</td></tr><tr><td>Launch ready</td><td>Mid-2028</td></tr><tr><td>Launch (L)</td><td>Dec. 2029<br>(shared with ARIEL)</td></tr><tr><td>Arrival at L2</td><td>L + 4 months</td></tr><tr><td>Waiting at L2</td><td>Maximum 5 years</td></tr><tr><td>Transfer to target Comet</td><td></td></tr><tr><td>Comet flyby</td><td>Latest L + 5.5 years</td></tr><tr><td>End of Operations</td><td>Latest L + 6 years</td></tr></tbody></table>

spacecraft A via dedicated S-band inter-satellite links. The masses of both probes are ~35 kg each, with a typical diameter/size of ~0.5 m.

B1 has a cuboid shape, and is 3-axis stabilised through a dedicated Attitude, Guidance and navigation sub-system (Fig. 42). Its electrical power system is based on deployable solar arrays and a secondary battery. B2 has an axisymmetric shape, is gyroscopically stabilised and has no attitude control capability. Its power system based on a primary battery, with an operational lifetime of ~30 hours.
The key dates for the Comet Interceptor high level schedule are given in Table 17. Compared to L- or M-class missions, the schedule is highly compressed, with 3 years of study phase (0, A, B) from ESA Science Programme Committee selection in June 2019 to formal adoption in June 2022, and 6 years of development phase (C, D) from adoption to launch readiness mid-2028.

## 7 Conclusion

Comet Interceptor, which will be the first dedicated comet mission launched by any space agency for over 25 years, promises to provide unique, multi-point measurements of a long-period comet, preferably dynamically-new. The mission also represents a new approach to a science mission by ESA, being the first in the agency's F-class of projects. The mission addresses ambitious science objectives in a modest-cost project, and is the first planetary mission selected by an agency where its primary science target has not been identified beforehand, and may not even be known at the time of launch.

**Acknowledgements** We dedicate this work to the memory of Prof. Anny-Chantal Levasseur-Regourd (1945–2022) and Dr. Maria Teresa Capria (1954–2022). Both were unfailing supporters of the Comet Interceptor mission.

**Funding** GHJ and CS acknowledge the support of the UK Space Agency; GHJ and AJC are grateful to STFC for partial support through consolidated grant ST/S000240/1. GHJ was supported for part of this work through the Visiting Scientist program of the International Space Science Institute (ISSI) in Bern. This work was supported by the Italian Space Agency (Contracts ASI-INAF n. 2020-4-HH.0, Addendum n. 2020-4-HH.1-2020 and 2020-4-HH.2-2021, and ASI-INAF agreement I/024/12/0) and by the European Space Agency (Contracts ESA-CNR n. 4000136673/21/NL/IB/ig and 3-17164/21/NL/GP/pbe COMET INTERCEPTOR-IT-3 – Comet Interceptor Dust Impact Effect Study). Work in Japan was supported by MEXT/JSPS KAKENHI Grants 21H04509, 22H01280, and 17KK0097. Co-authors in France acknowledge the support of CNES for the Comet Interceptor mission. Work in Germany was financially supported by the German Ministerium fuer Wirtschaft und Energie and the Deutsches Zentrum f¨ur Luft- und Raumfahrt under contract 50OW2101. Support by the Swedish National Space Agency through Grant 108/18 is acknowledged. Work at IRF-Uppsala is supported by the Swedish National Space Agency, DNr 2021-00047. Work at BIRA-IASB acknowledges BELSPO funding via PRODEX PEA 4000133230. Work at BIRA-IASB acknowledges BELSPO funding via PRODEX PEA 4000133230 and via B2/191/P1/SeVoCo. MRELM is partly supported by the internal grant 8474000336-KU-SPSC. The work at Imperial College London was supported by the UK Space Agency [grant numbers ST/W002604/1, ST/V002767/1 and ST/X002349/1] and by ESA [4000130837/20/NL/IB/ig]. ZL acknowledges support by STFC [ST/W507519/1]. Work at the Open University, UK was supported by ESA [4000134679/21/NL/GP/pbe]. CG acknowledges the Research Fellowship Programme at the European Space Agency (ESA) for support. AGL has received funding from the European Research Council (ERC) under the European Union’s Horizon 2020 research and innovation programme (Grant Agreement No 802699). The work of AK was supported by the PRODEX 4000132506 project of ESA and the KKM of Hungary. Work at IAA is supported by project PGC2018-099425-B-I00 (MCI/AEI/FEDER, UE). Work in Switzerland was supported by SSO/SNF (200020_207312)/ State of Bern/ESA. CSW and MV thank the Austrian Science Fund (FWF) P32035-N36 and P35954-N. This work was supported by NASA HQ, Washington DC USA, and JPL/NASA. The work of RH is funded by the Science and Engineering Research Board (SERB, grant no. SB/S2/RJN-080/2018), a statutory body of the Department of Science and Technology (DST), Government of India through the Ramanujan Fellowship. KM’s research supported by the Academy of Finland grants No. 1345115 and 1336546, and AP by their grant no. 325805. TK’s work was supported by ESA EXPRO funds, the Academy of Finland project No. 335595, the NASA Solar System Exploration Research Virtual Institute Center for Lunar and Asteroid Surface Science, and it was conducted within institutional support RVO 67985831 of the Institute of Geology of the Czech Academy of Sciences. MTB appreciates support by the Rutherford Discovery Fellowships from New Zealand Government funding, administered by the Royal Society Te Ap¯arangi. Centro de Qu´ımica Estrutural acknowledges the financial support of FCT-Fundac¸˜ao para a Ciˆencia e a Tecnologia (UIDB/00100/2020 and UIDP/00100/2020), and Institute of Molecular Sciences acknowledges the financial support of FCT-Fundac¸˜ao para a Ciˆencia e a Tecnologia (LA/P/0056/2020).
The key dates for the Comet Interceptor high level schedule are given in Table 17. Compared to L- or M-class missions, the schedule is highly compressed, with 3 years of study phase (0, A, B) from ESA Science Programme Committee selection in June 2019 to formal adoption in June 2022, and 6 years of development phase (C, D) from adoption to launch readiness mid-2028.

# 7 Conclusion

Comet Interceptor, which will be the first dedicated comet mission launched by any space agency for over 25 years, promises to provide unique, multi-point measurements of a long-period comet, preferably dynamically-new. The mission also represents a new approach to a science mission by ESA, being the first in the agency's F-class of projects. The mission addresses ambitious science objectives in a modest-cost project, and is the first planetary mission selected by an agency where its primary science target has not been identified beforehand, and may not even be known at the time of launch.

**Acknowledgements** We dedicate this work to the memory of Prof. Anny-Chantal Levasseur-Regourd (1945–2022) and Dr. Maria Teresa Capria (1954–2022). Both were unfailing supporters of the Comet Interceptor mission.

**Funding** GHJ and CS acknowledge the support of the UK Space Agency; GHJ and AJC are grateful to STFC for partial support through consolidated grant ST/S000240/1. GHJ was supported for part of this work through the Visiting Scientist program of the International Space Science Institute (ISSI) in Bern. This work was supported by the Italian Space Agency (Contracts ASI-INAF n. 2020-4-HH.0, Addendum n. 2020-4-HH.1-2020 and 2020-4-HH.2-2021, and ASI-INAF agreement I/024/12/0) and by the European Space Agency (Contracts ESA-CNR n. 4000136673/21/NL/IB/ig and 3-17164/21/NL/GP/pbe COMET INTERCEPTOR-IT-3 – Comet Interceptor Dust Impact Effect Study). Work in Japan was supported by MEXT/JSPS KAKENHI Grants 21H04509, 22H01280, and 17KK0097. Co-authors in France acknowledge the support of CNES for the Comet Interceptor mission. Work in Germany was financially supported by the German Ministerium fuer Wirtschaft und Energie and the Deutsches Zentrum f¨ur Luft- und Raumfahrt under contract 50OW2101. Support by the Swedish National Space Agency through Grant 108/18 is acknowledged. Work at IRF-Uppsala is supported by the Swedish National Space Agency, DNr 2021-00047. Work at BIRA-IASB acknowledges BELSPO funding via PRODEX PEA 4000133230. Work at BIRA-IASB acknowledges BELSPO funding via PRODEX PEA 4000133230 and via B2/191/P1/SeVoCo. MRELM is partly supported by the internal grant 8474000336-KU-SPSC. The work at Imperial College London was supported by the UK Space Agency [grant numbers ST/W002604/1, ST/V002767/1 and ST/X002349/1] and by ESA [4000130837/20/NL/IB/ig]. ZL acknowledges support by STFC [ST/W507519/1]. Work at the Open University, UK was supported by ESA [4000134679/21/NL/GP/pbe]. CG acknowledges the Research Fellowship Programme at the European Space Agency (ESA) for support. AGL has received funding from the European Research Council (ERC) under the European Union’s Horizon 2020 research and innovation programme (Grant Agreement No 802699). The work of AK was supported by the PRODEX 4000132506 project of ESA and the KKM of Hungary. Work at IAA is supported by project PGC2018-099425-B-I00 (MCI/AEI/FEDER, UE). Work in Switzerland was supported by SSO/SNF (200020_207312)/ State of Bern/ESA. CSW and MV thank the Austrian Science Fund (FWF) P32035-N36 and P35954-N. This work was supported by NASA HQ, Washington DC USA, and JPL/NASA. The work of RH is funded by the Science and Engineering Research Board (SERB, grant no. SB/S2/RJN-080/2018), a statutory body of the Department of Science and Technology (DST), Government of India through the Ramanujan Fellowship. KM’s research supported by the Academy of Finland grants No. 1345115 and 1336546, and AP by their grant no. 325805. TK’s work was supported by ESA EXPRO funds, the Academy of Finland project No. 335595, the NASA Solar System Exploration Research Virtual Institute Center for Lunar and Asteroid Surface Science, and it was conducted within institutional support RVO 67985831 of the Institute of Geology of the Czech Academy of Sciences. MTB appreciates support by the Rutherford Discovery Fellowships from New Zealand Government funding, administered by the Royal Society Te Ap¯arangi. Centro de Qu´ımica Estrutural acknowledges the financial support of FCT-Fundac¸˜ao para a Ciˆencia e a Tecnologia (UIDB/00100/2020 and UIDP/00100/2020), and Institute of Molecular Sciences acknowledges the financial support of FCT-Fundac¸˜ao para a Ciˆencia e a Tecnologia (LA/P/0056/2020).
MRELM acknowledges funding from the KU internal grant (8474000336-KU-SPSC). MK gratefully acknowledges funding from the European Union's Horizon Europe research and innovation programme under grant agreement No. 101079231 (EXOHOST), and from UK Research and Innovation (UKRI) under the UK government's Horizon Europe funding guarantee (grant number 10051045).

## Declarations

**Competing Interests** The authors declare no competing interests.

**Open Access** This article is licensed under a Creative Commons Attribution 4.0 International License, which permits use, sharing, adaptation, distribution and reproduction in any medium or format, as long as you give appropriate credit to the original author(s) and the source, provide a link to the Creative Commons licence, and indicate if changes were made. The images or other third party material in this article are included in the article's Creative Commons licence, unless indicated otherwise in a credit line to the material. If material is not included in the article's Creative Commons licence and your intended use is not permitted by statutory regulation or exceeds the permitted use, you will need to obtain permission directly from the copyright holder. To view a copy of this licence, visit http://creativecommons.org/licenses/by/4.0/.

## References

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Altwegg K, Balsiger H, Bar-Nun A, Berthelier JJ, Bieler A, Bochsler P, Briois C, Calmonte U, Combi MR, Cottin H, De Keyser J, Dhooghe F, Fiethe B, Fuselier SA, Gasc S, Gombosi TI, Hansen KC, Haessig M, Jäckel A, Kopp E, Korth A, Le Roy L, Mall U, Marty B, Mousis O, Owen T, Rème H, Rubin M, Sémon T, Tzou CY, Waite JH, Wurz P (2016) Prebiotic chemicals–amino acid and phosphorus–in the coma of comet 67P/Churyumov-Gerasimenko. Sci Adv 2(5):e1600285. https://doi.org/10.1126/sciadv.1600285

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MRELM acknowledges funding from the KU internal grant (8474000336-KU-SPSC). MK gratefully acknowledges funding from the European Union's Horizon Europe research and innovation programme under grant agreement No. 101079231 (EXOHOST), and from UK Research and Innovation (UKRI) under the UK government's Horizon Europe funding guarantee (grant number 10051045).

## Declarations

**Competing Interests** The authors declare no competing interests.

**Open Access** This article is licensed under a Creative Commons Attribution 4.0 International License, which permits use, sharing, adaptation, distribution and reproduction in any medium or format, as long as you give appropriate credit to the original author(s) and the source, provide a link to the Creative Commons licence, and indicate if changes were made. The images or other third party material in this article are included in the article's Creative Commons licence, unless indicated otherwise in a credit line to the material. If material is not included in the article's Creative Commons licence and your intended use is not permitted by statutory regulation or exceeds the permitted use, you will need to obtain permission directly from the copyright holder. To view a copy of this licence, visit http://creativecommons.org/licenses/by/4.0/.

## References

Adams FC (2010) The birth environment of the solar system. Annu Rev Astron Astrophys 48:47–85. https://doi.org/10.1146/annurev-astro-081309-130830. arXiv:1001.5444

A'Hearn MF, Millis RC, Schleicher DO, Osip DJ, Birch PV (1995) The ensemble properties of comets: results from narrowband photometry of 85 comets, 1976-1992. Icarus 118(2):223–270. https://doi.org/10.1006/icar.1995.1190

A'Hearn MF, Belton MJS, Delamere WA, Kissel J, Klaasen KP, McFadden LA, Meech KJ, Melosh HJ, Schultz PH, Sunshine JM, Thomas PC, Veverka J, Yeomans DK, Baca MW, Busko I, Crockett CJ, Collins SM, Desnoyer M, Eberhardy CA, Ernst CM, Farnham TL, Feaga L, Groussin O, Hampton D, Ipatov SI, Li JY, Lindler D. Lisse CM, Mastroemos N, Owen WM, Richardson JE, Wellnitz DD, White RL (2005) Deep impact: excavating comet Tempel 1. Science 310(5746):258–264. https://doi.org/10.1126/science.1118923

A'Hearn MF, Belton MJS, Delamere WA, Feaga LM, Hampton D, Kissel J, Klaasen KP, McFadden LA, Meech KJ, Melosh HJ, Schultz PH, Sunshine JM, Thomas PC, Veverka J, Wellnitz DD, Yeomans DK, Besse S, Bodewits D, Bowling TJ, Carcich BT, Collins SM, Farnham TL, Groussin O, Hermalyn B, Kelley MS, Kelley MS, Li JY, Lindler DJ, Lisse CM, McLaughlin SA, Merlin F, Protopapa S, Richardson JE, Williams JL (2011) EPOXI at comet Hartley 2. Science 332(6036):1396. https://doi.org/10.1126/science.1204054

Alfven H (1957) On the theory of comet tails. Tellus 9:92. https://doi.org/10.3402/tellusa.v9i1.9064

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