Athena Exploring to Investigate its Habitability

Athena Exploring Enceladus to Investigate its Habitability

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September 2016 MSc Space Exploration Systems University of Leicester Department of Physics and Astronomy

Athena Exploring Enceladus to Investigate its Habitability

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Athena Exploring Enceladus to Investigate its Habitability

Change Log

Date Issue Revision Section Reason for Change 12/09/2016 1 1 Final Report N/A

Study Participants, Acknowledgement

Rawaa Algareb Samuel Bryan Narmin Manme Ashley Surridge

This research was carried out at the University of Leicester as part of the project component of the MSc Space Exploration Systems course.

The team would like to thank the following individuals for their valuable contributions:

Dr Nigel Bannister Dr Richard Ambrosi Dr Ian Hutchinson Hannah Lerman

Structure of the Enceladus Phase-A Study Final Report

This document provides a scientific, technical and management summary of the Athena mission phase-A study that was performed from March 2016 to September 2016. This report is structured in such a way to demonstrate the flow down of the proposed mission from scientific themes and objectives to concrete measurements and areas of further study. Section 1 gives an overview of the mission background and establishes the top level science goals, investigations, measurements and instruments. Section 2 highlights the requirements of the mission. Section 3 outlines the model payloads of the lander and orbiter systems. Section 4 details the mission design which has been proposed to tackle the science goal and its six sub-goals. Section 5 gives an overview of the future management required to develop the study further.

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Athena Exploring Enceladus to Investigate its Habitability

Table of Content

1 Science Goals, Measurements and Instruments ...... 1 1.1 A review of Voyager and Cassini Observations ...... 1 1.1.1 Location and Orbital Properties ...... 1 1.1.2 Composition and Physical Properties ...... 1 1.1.3 The Sub-Surface Reservoir ...... 4 1.1.4 Life in Enceladus ...... 4 1.2 Science Goals, Objectives, Investigations and Measurements ...... 5 1.2.1 Science Goals – Explore Enceladus to Investigate its Habitability...... 5 1.2.2 Objectives and Investigations ...... 5 1.3 Measurements and Instruments ...... 5 1.3.1 Instruments...... 8 1.3.2 Payload A – Sicily Lander ...... 8 1.3.3 Payload B – Athena Orbiter ...... 8 1.3.4 Considered Systems ...... 8 2 Science Requirements ...... 9 2.1 Athena ...... 9 2.2 Sicily ...... 9 3 Model Payloads ...... 10 3.1 Model Payload Definition ...... 10 3.2 Athena Model Payload ...... 10 3.2.1 Narrow Angle Camera (NAC) ...... 10 3.2.2 Wide Angle Camera (WAC) ...... 10 3.2.3 A Mid-Infrared Spectro-Imager -THERMAP (TM-MIRS) ...... 11 3.2.4 Altimeter (LA)...... 11 3.2.5 The Ice Penetrating Radar (IPR) ...... 12 3.2.6 Athena Model Payload Summary ...... 12 3.3 Sicily Model Payload ...... 13 3.3.1 Gas-Chromatograph – Ion Trap – Mass Spectrometer (GC-IT-MS) ...... 13 3.3.2 Gas-Chromatograph – Time of Flight – Mass Spectrometer (GC-TOF-MS) ...... 14 3.3.3 Signs Of Life Detector (SOLID) ...... 14 3.3.4 Raman Spectrometer (RS) ...... 15 3.3.5 Alpha Particle X-ray Source (APXS) ...... 15 3.3.6 Enceladus Surface Drill (ESD)...... 16 3.3.7 Enceladus Sample Collector (ESC) ...... 16 3.3.8 Pancam (PC) ...... 17 3.3.9 Sicily Model Payload Summary ...... 17 3.4 Conclusions and Recommendation ...... 19 4 Mission Design ...... 19 4.1.1 Mission Profile ...... 19 4.1.2 Earth to Saturn (Cruise) ...... 20 4.1.3 Launch ...... 20 iv

Athena Exploring Enceladus to Investigate its Habitability

4.1.4 Saturn to Enceladus (Moon’s Tour) ...... 21 4.1.5 Enceladus Science Orbits ...... 21 4.1.6 Regions of Enceladus ...... 22 4.1.7 The Sicily Landing Site ...... 22 4.1.8 Athena Operation Life...... 22 4.1.9 Sicily Descent ...... 24 4.1.10 Sicily Operation Life ...... 25 4.1.11 End of Mission ...... 26 4.2 Radiation Environment and Shielding ...... 26 4.3 Thermal Environment ...... 28 4.3.1 Earth to Saturn (Cruise) ...... 28 4.3.2 Saturn to Enceladus (Moon’s tour)...... 28 4.3.3 Sicily ...... 29 4.4 Propulsion ...... 29 4.4.1 Athena ...... 29 4.4.2 Sicily ...... 29 4.5 Power ...... 30 4.5.1 Athena ...... 30 4.5.2 Sicily ...... 30 4.6 Telecommunication...... 31 4.6.1 Athena ...... 31 4.6.2 Sicily ...... 32 4.7 Command and Data Handling ...... 33 4.8 Guidance, Navigation and Control ...... 33 4.8.1 Athena ...... 33 4.8.2 Sicily ...... 33 4.9 Mechanical Design...... 34 4.9.1 Athena Structure ...... 34 4.9.2 Sicily Structure...... 35 4.10 Mass Budget ...... 35 4.11 and Contamination Control ...... 36 4.11.1 Planetary Protection ...... 36 4.11.2 Contamination Control ...... 37 4.12 Mitigation of Technical Risk...... 38 4.13 Conclusions and Recommendations ...... 39 5 Management ...... 40 5.1 Schedule...... 40 5.2 Public Outreach and Science Communication ...... 40 5.3 Conclusions and Recommendations ...... 40 6 References ...... 41

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Athena Exploring Enceladus to Investigate its Habitability

List of Figures

Figure 1-1. Enceladus with its icy surface and distinctive ‘Tiger Stripe’ pattern (bottom right). Also shown are surface trenches caused by tidal forces and Saturn’s gravity (NASA/JPL/Space Institute)...... 1 Figure 1-2. Enhanced-colour maps of the northern and southern hemispheres of Enceladus. The yellowish and magenta tones are believed to be due to the differences in surface deposits. The fractures on the South Pole have a stronger UV signature and appear blueish (Lunar and Planetary Institute & NASA/JPL/Space Institute)...... 2 Figure 1-3. A perspective view of Damascus Sulcus with 12 to 30-meter resolution. The 'ropy' plains are a few tens of meters high with the relief being exaggerated by a factor of ~10 to enhance clarity ( (Bland, 2014) & NASA/JPL/Space Institute)...... 2 Figure 1-4. Enceladus showing its plume to the ISS on Cassini. Situated at the Tiger Stripes are ~100 jets which feed the plume that extends to a distance of ~435 km and consists of a solid water ice and a gas water vapour component (NASA/JPL/Space Institute)...... 2 Figure 1-5. The mass deposition rate of plume particles between 0.5 μm and 5 μm on the surface of Enceladus. The y axis is latitude and the x axis is longitude. Labelled are the strongest Enceladus dust jets deduced from the simulations; A: Alexandria, B: Baghdad, C: Cairo and D: Damascus. The rate is between 0.5 mm/yr at the location of the jets and 10-5 mm/yr at the equator. A thick band of deposition lies between -45° and - 90° (Kempf, 2010)...... 3 Figure 1-6. Abundance vs mass for chemicals sampled by Cassini’s Ion and Neutral Mass Spectrometer (INMS) during a flyby in 2008 (NASA/JPL/Space Institute) ...... 3 Figure 1-7. Model temperature distribution at an idealized tiger stripe, showing temperature vs height and distance from the warm fracture ( (O.Abramov, 2015) & NASA/JPL/Space Institute) ...... 3 Figure 1-8. Possible structure of Enceladus if a global ocean is present. (a) The amplitude of the physical forced liberation as a function of the icy shell's thickness. (b) Schematic cross section of the upper 60 km of Enceladus from the South Pole to the equator. Vertical exaggeration is 4.5 (Thomas P. , 2016)...... 4 Figure 1-9. Possible methane cycle on Enceladus. This cycle is present on Earth in hydrothermal vents and 'Biology' could be microbes like Bacteria or Archea (Mckay, 2008) ...... 4 Figure 1-10. Athena Science Traceability Matrix (STM)...... 7 Figure 4-1. The designed trajectory for the Athena mission. The combined system (Athena + Sicily) will launch from Earth, then perform flybys of and Earth and finally arrive at Saturn at ’s orbital radius. The total ΔV for 2 2 this cruise is 44 m/s and the required C3 is a maximum of 14 km /s (MacKenzie et al., 2016)...... 20 Figure 4-2. Characteristic energy (Atlas V series) (Elvperf.ksc.nasa.gov, 2016)...... 20 Figure 4-3. An illustration of the proposed moon’s tour in the Saturn system. The combined system would complete 2.5 years orbiting these four moons of Saturn (MacKenzie et al., 2016)...... 21 Figure 4-4. Shows the four regions of Enceladus. The northernmost region (green) is the North Polar Terrain (NPT) and is between 60 and 90°N. The southernmost region (orange) is the South Polar Terrain (SPT) and is between 70 and 90°S. The region above the SPT (black) is the resurfacing region (RR) and is between 45 and 70°S. The final region is the rest of Enceladus (REST) and is between 45°S and 60°N...... 22 Figure 4-5. Steps of descent Sicily Lander...... 24 Figure 4-6. Jovian and Saturnian radiation environments for a 1-week exposure. Rs is Saturn radii and Rj is Jupiter radii. Dose in rad (Si) for a 1 m radius aluminium spherical shell (H.Garett, 2005). The radiation environment at Saturn is ~ 1000 weaker than the radiation environment at Jupiter. For orbital distance greater than 9.47 Rs the 9.47 Rs plot has been used. For 4 Rs (Enceladus’ orbital distance) the values that are half way between 2.55 Rs and 5.95 Rs has been used...... 27 Figure 4-7. Shows the lander duty cycle ...... 31 Figure 4-8. Shows three separate perspectives for the combined orbiter and lander system. On the left is the structure sub-system and consists of the propellant tanks, CFRP propulsion tube, CFRP structure shell and aluminium shell. Any holes, tubing or extra detail is not included for clarity. At the top shows the placement of all the sub-systems and the Sicily lander. The bottom-right perspective shows the orientation of Athena during the initial imaging stages. The arrow indicates the direction of all the instruments FOV...... 34 Figure 4-9. Shows the main components of lander ...... 35 Figure 4-10. Shows the three legs of lander in opposite directions...... 35 Figure 4-11. The steps of the risk management process (ECSS-M-ST-80C, RD01)...... 38 Figure 4-12. Yellow and green retired risks of Athena mission with their mitigations ...... 38 Figure 4-13. Red, yellow and green retired risks of Athena mission with their mitigations ...... 39

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Athena Exploring Enceladus to Investigate its Habitability

List of Tables

Table 1-1. Orbital properties for Enceladus and Saturn. Enceladus’ orbit is almost circular and it always has the same face pointing at Saturn (NASA/JPL/Space Institute)...... 1 Table 1-2. The four measurements, including discussion, which will meet the investigation L1i (shown in the STM, Figure 1-10) in the goal Life ...... 5 Table 1-3. The measurements and discussion, which are designed to meet the Hotspot, Ocean and Secondary Investigation investigations (shown in the STM, Figure 1-10)...... 6 Table 3-1. Orbiter module instruments and the associated science goals ...... 10 Table 3-2. Lander module instruments and the associated science goals ...... 10 Table 3-3. Baseline NAC parameters ...... 10 Table 3-4. Baseline WAC parameters ...... 11 Table 3-5. Baseline THERMAP parameters ...... 11 Table 3-6. Some key parameters for the Laser altimeter of Athena. All parameters are the same as its heritage instrument LOLA (Neumann et al, 2011) ...... 11 Table 3-7. Basic operational requirements for the laser diode arrays of Enceladus (Neumann et al, 2011)...... 11 Table 3-8. Some key parameters for the Ice penetrating radar of Athena. All parameters are the same as its heritage instrument Rime radar of Juice mission, except the altitude which is 200-500km for the RIME (Bruzzone et al., 2013), while the altitude of Athena radar 50 km...... 12 Table 3-9. Athena model instruments summary ...... 13 Table 3-10. Some key properties for the GC-IT-MS instruments. All parameters are the same as its heritage instrument Ptolemy, except (*) the number of ovens has increased (Wright, 2006)...... 14 Table 3-11. Some key parameters for the GC-TOF-MS instrument. All parameters are the same as its heritage instrument COSAC, except (*) the number of ovens has increased (Goesmann, 2006) ...... 14 Table 3-12. Baseline Raman spectrometer parameters ...... 15 Table 3-13. Baseline APXS parameters ...... 16 Table 3-14. Some key parameters for ESD. All parameters are the same as its heritage instrument SD2, except (*) the operating temperature, which has decreased and (**) sampling distribution, which includes distribution to the ovens and instruments, instead of only the ovens...... 16 Table 3-15. Some suggestive parameters for the ESC...... 17 Table 3-16. Some key parameters for the PC. All parameters are the same as its heritage instrument CIVA-P (Bibring, 2007)...... 17 Table 3-17. Sicily model instruments summary. *The SOLID3 dimensions have been estimated from earlier versions of this instrument. **This is the power potentially needed to move the samples from the collector to the main handling unit...... 17 Table 4-1. The complete operation period for the Athena orbiter. All orbits are with an eccentricity of zero and stated inclinations are the maximum values. Number of orbits correspond to the instruments coverage requirements. (*) Inclination will increase from 65 – 90 ° in increments of 5 degrees (see section 4.2.5) (**) Depends on selected landing site latitude...... 23 Table 4-2. The total velocity phase of descent Sicily...... 24 Table 4-3. Sicily operation life consists of seven phases and is ~ 80 hours. After the batteries have been recharged in the hibernation phase the lander could have an extended mission. In bold are the sample collection systems. Further information on the instruments can be found in Section 3...... 25 Table 4-4. Key parameters of end of mission options ...... 26 Table 4-5. A top level assessment of the amount of aluminium shielding required on the list of components. The Max Assigned Dose is based on an assessment of the components susceptibility to the effects of radiation. With 200 rads being assigned to components with electronics, silicon detectors or other radiation sensitive components. The Predicted Aluminium Shielding is calculated from the component’s surface area, its duration in each of the four sections and its Max Assigned Dose. The NAC and WAC are located together and therefore shielded by one set of aluminium. The ESC contains no sensitive radiation components so a shielding thickness of zero is predicted. Total shielding mass for Sicily ~107 kg and Athena ~190 kg...... 27 Table 4-6. Operating Temperatures of Athena mission ...... 28 Table 4-7. The properties of radiator, the instruments, and battery, electronics, and propulsion systems will be maintained the thermal system the temperature of within operating limits. For example, a number of radioisotope heater units (RHUs) will be used in order ...... 28 Table 4-8. RHU distribution for lander ...... 29 Table 4-9. RHU distribution for orbiter ...... 29 Table 4-10. Summary of engines selected for orbiter module ...... 29 Table 4-11. Baseline thruster parameters ...... 30 vii

Athena Exploring Enceladus to Investigate its Habitability

Table 4-12. Baseline hydrazine propellant tank parameters ...... 30 Table 4-13. Baseline 241-AmRTGs parameters ...... 30 Table 4-14. Baseline Li-Ion parameters ...... 31 Table 4-15. Comparison of uplink and downlink antennae for Orbiter communication. *This value is for the smallest of the DSN dishes which was used to calculate up and downlink rates in the worst case environment ...... 31 Table 4-16. Some example parameters for an omni-directional antenna on Athena and Sicily (Technology, 2016). .. 32 Table 4-17. Some example parameters for the S-band transmitter on Athena and Sicily (Technology, 2016)...... 32 Table 4-18. Some example parameters for the S-band receiver on Athena and Sicily (Technology, 2016)...... 32 Table 4-19. Summary of key launch mass parameters ...... 35 Table 4-20. Breakdown of orbiter mass budget ...... 35 Table 4-21. Breakdown of lander mass budget ...... 35 Table 4-22. List of mission types/ target body of planetary protection that categorized by COSPAR (Nicholson, Schuerger and Race, 2009)...... 36 Table 4-23. Some possible active and passive methods for contamination control (Y.Bar-Choen, 2009) (JPL/NASA, 2011)...... 37

Work Breakdown

The table below relates the contributions of each team member to the associated chapter

1 Sam Bryan 4.1.3 Narmin Manme 1.1 Sam Bryan 4.1.4 Sam Bryan 1.1.1 Sam Bryan 4.1.5 Sam Bryan 1.1.2 Sam Bryan 4.1.6 Sam Bryan 1.1.3 Sam Bryan 4.1.7 Sam Bryan 1.1.4 Sam Bryan 4.1.8 Sam Bryan 1.2 Sam Bryan 4.1.9 Narmin Manme 1.2.1 Sam Bryan 4.1.10 Sam Bryan 1.2.2 Sam Bryan 4.1.11 Ashley Surridge 1.3 Sam Bryan 4.2 Sam Bryan 1.3.1 Sam Bryan 4.3 Narmin Manme 1.3.2 Sam Bryan 4.3.1 Narmin Manme 1.3.3 Sam Bryan 4.3.2 Narmin Manme 1.3.4 Ashley Surridge 4.3.3 Ashley Surridge 2.1 Ashley Surridge 4.4 Ashley Surridge 2.2 Sam Bryan 4.4.1 Ashley Surridge 3 Ashley Surridge 4.4.2 Rawaa Algareb 3.1 Ashley Surridge 4.5 Ashley Surridge 3.2 Ashley Surridge 4.5.1 Ashley Surridge 3.2.1 Rawaa Algareb 4.5.2 Rawaa Algareb 3.2.2 Rawaa Algareb 4.6 Sam Bryan 3.2.3 Rawaa Algareb 4.6.1 Sam Bryan 3.2.4 Narmin Manme 4.6.2 Ashley Surridge 3.2.5 Narmin Manme 4.7 Narmin Manme & Rawaa Algareb 3.2.6 Narmin Manme 4.8.1 Rawaa Algareb 3.3 Sam Bryan 4.8.2 Narmin Manme 3.3.1 Sam Bryan 4.9 Sam Bryan 3.3.2 Sam Bryan 4.9.1 Sam Bryan 3.3.3 Ashley Surridge 4.9.2 Rawaa Algareb 3.3.4 Ashley Surridge 4.10 Ashley Surridge 3.3.5 Ashley Surridge 4.11.1 Narmin Manme 3.3.6 Sam Bryan 4.11.2 Sam Bryan 3.3.7 Sam Bryan 4.12 Narmin Manme 3.3.8 Sam Bryan 4.13 Ashley Surridge 3.3.9 Ashley Surridge 5 Ashley Surridge 3.4 Ashley Surridge 5.1 Ashley Surridge 4.1.1 Ashley Surridge 5.2 Ashley Surridge 4.1.2 Sam Bryan 5.3 Ashley Surridge

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Athena Exploring Enceladus to Investigate its Habitability

1 Science Goals, Measurements and 1.1.1 Location and Orbital Properties Instruments Enceladus is the one of the inner moons that 1.1 A review of Voyager and Cassini orbit Saturn. It is located in the densest part of the E Observations ring, at a distance of ~238,000 km between Mimas and Tethys, Table 1-1. Enceladus rotates once in Saturn’s Enceladus Figure 1-1, named after one of the Gigantes equatorial plane every 1.37 days and its axial inclination in Greek Mythology was first discovered in 1789 by Sir is zero (NASA, Saturnian Satellite Fact Sheet, 2015). It William Herschel. The first close up observations came has an almost near circular orbit because its eccentricity in 1980 from and then in 1981 by Voyager is forced and kept finite by its sister moon Dione 2. Together they identified a highly reflective surface (Spohn, 2009). Therefore, both moons are locked in a including a heavily cratered north pole and a sparsely 2:1 mean motion resonance where Enceladus’ orbital cratered south pole (A.F.Cook, 1981). The next period period is half that of Dione’s. of observation occurred in 2005 by the Cassini-Huygens space probe. Cassini successfully entered an orbit Orbital Property Value Unit around Saturn in July 2004 and started its first Semi-Major Axis 3.94 Saturn Radii observation period of Enceladus at the start of 2005. During this first period Cassini’s Imaging Science Rotational Period 32.9 Hours Subsystem (ISS), (MAG) and Visual, Orbital Period 32.9 Hours Infrared and mapping spectrometer (VIMS) identified Eccentricity 0.0045 an unexpected active region at the south pole of Enceladus Axial Tilt 0.0 Degrees Enceladus (Porco, 2006). This activity has sparked great interest in Enceladus from the scientific community as Saturn Axial Tilt 26.9 Degrees a moon of this size should be an inactive world covered Enceladus Inclination 0.019 Degrees uniformly in craters. With its ‘Prime’ objective to (to Saturn’s equator) provide a thorough study of the Saturn system (ESA, Cassini-Huygens mission objectives, 2016) the probe Table 1-1. Orbital properties for Enceladus and Saturn. returned a vast amount of data on Enceladus. This data Enceladus’ orbit is almost circular and it always has the same has significantly improved our understanding of face pointing at Saturn (NASA/JPL/Space Institute). Enceladus and is therefore the main source for this report. 1.1.2 Composition and Physical Properties

Enceladus is the sixth largest moon of Saturn and it has an effective radius of 252.1 km, a mass of 1.08 x1020 kg and a surface gravity between 0.111 and 0.114 m/s2 (Thomas P. , 2007). The exact cross section of the moon is still greatly debated but it is believed to be differentiated with a 150 km radius core (C.Parkinson, 2007), a sub-surface reservoir and a crustal thickness between 10 to 40km (L.Iess, 2014) (Porco.C, 2014). Its surface is composed mostly of nearly pure water ice except at the SPT where there are light organics, CO2, and amorphous and crystalline water ice (Brown, 2006). The surface terrain varies on Enceladus with the obvious contrast being between the northern and southern hemispheres. Shown in Figure 1-2 the northern hemisphere has a distinctly greater number of craters than the southern hemisphere. Using crater frequency ratios (Porco, 2006) suggests that heavily cratered northern regions are around 4200 million years old and selected southern regions are less than 100 million years old. While the age of the northern region is comparable with the age of the solar system the southern hemisphere Figure 1-1. Enceladus with its icy surface and distinctive relatively young age is unusual and unexpected. ‘Tiger Stripe’ pattern (bottom right). Also shown are surface trenches caused by tidal forces and Saturn’s gravity (NASA/JPL/Space Institute).

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Athena Exploring Enceladus to Investigate its Habitability

Figure 1-2. Enhanced-colour maps of the northern and southern hemispheres of Enceladus. The yellowish and magenta tones are believed to be due to the differences in surface deposits. The fractures on the South Pole have a stronger UV signature and appear blueish (Lunar and Planetary Institute & NASA/JPL/Space Institute).

The South Pole more commonly known as the South At very high phase angles ISS imaging reveals ~100 Polar Terrain (SPT) is located between 70 °S and 90 °S. near-surface jets emanating from the Tiger Stripes, And it is distinguished from the rest of Enceladus by its which are seen, Figure 1-4, to feed a plume that extends lack of craters, unusual albedo signatures, unusual UV to a height of ~435 km. The jets, also called geysers, signatures, circular cracks and the four principal consist of two components: gas in the form of water trenches dubbed the ‘Tiger Stripes’ (Porco, 2006). The vapour and solid in the form of micron sized water ice four ‘Tiger Stripes’ identified as Alexandria, Cairo, particles. Globally the typical solid water-ice size is 50 Baghdad and Damascus are spaced by ~35 km plains to 150 μm but this increases to 100-300 μm at the SPT surrounded by a network of smaller trenches. From region (Brown, 2006). The mass of the gas and solid Cassini ISS measurements these principal trenches are components were measured using a combination of the each ~130 km long, 500 m deep and 2 km wide flanked Ultra Violet Imaging Spectrometer (UVIS) and the ISS on both sides by high 100 m ridges. The funiscular (Ingersoll, 2011). From the jets about 8 kg/s of gas and plains (named because of its ropy appearance) Figure 51 +/- 18 kg/s of solid are ejected into the plume. 1-3, are characterized by 50-100 m amplitude, tightly (Jürgen, 2008) (Porco.C, 2014). The source of the jets hinged ridges and troughs with a spacing of ~1 km is highly debated but the solid/gas mass ratio from the (Bland, 2014). jets (~6) strongly indicates the presence of a sub-surface reservoir.

Figure 1-3. A perspective view of Damascus Sulcus with 12 Figure 1-4. Enceladus showing its plume to the ISS on to 30-meter resolution. The 'ropy' plains are a few tens of Cassini. Situated at the Tiger Stripes are ~100 jets which feed meters high with the relief being exaggerated by a factor of the plume that extends to a distance of ~435 km and consists ~10 to enhance clarity ( (Bland, 2014) & NASA/JPL/Space of a solid water ice and a gas water vapour component Institute). (NASA/JPL/Space Institute).

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Athena Exploring Enceladus to Investigate its Habitability

Figure 1-5. The mass deposition rate of plume particles between 0.5 μm and 5 μm on the surface of Enceladus. The y axis is latitude and the x axis is longitude. Labelled are the strongest Enceladus dust jets deduced from the simulations; A: Alexandria, B: Baghdad, C: Cairo and D: Damascus. The rate is between 0.5 mm/yr at the location of the jets and 10-5 mm/yr at the equator. A thick band of deposition lies between -45° and - 90° (Kempf, 2010).

From study of the dust delivery to the E-ring the model The average surface temperature of Enceladus is ~75 K used by (Kempf, 2010) indicates that between 7.9% and (Nimmon, 2008), but located at the SPT and not at the 10.4% of particles exceeding 0.7 μm and between North Pole, is a thermal anomaly. Initially detected 95.3% and 96.9% of particles less than 0.7 μm escape using low resolution thermal scans 5.8 +/- 1.9 GW of to feed the E-ring. The majority of the mass ejected from thermal energy is emitted from the entire SPT region. the jets will not escape Enceladus’ gravity and therefore Using higher resolution imaging the heat sources fall to the surface as ‘snow’. The deposition rate against (otherwise called hot spots) are actually localised at the longitude and latitude is shown in Figure 1-5 and this Tiger Stripes locations. The hot spots are typically rate decreases with distance from the four Tiger Stripe hundreds of meters wide with temperatures between locations. The largest particles of about 5 μm are much 100 and 200 K, Figure 1-7 (Spencer J. , 2006) (Porco.C, smaller than the suggested 100 -300 μm by others 2014) (O.Abramov, 2015). (Brown, 2006) and this is thought to be attributed to other physically processes taking place on the surface, which can decrease the solid size (Kempf, 2010).

Figure 1-2. Model temperature distribution at an idealized tiger stripe, showing temperature vs height and distance from the warm fracture ( (O.Abramov, 2015) & NASA/JPL/Space Institute)

The mechanism responsible for the thermal emission is still highly debated but the leading theory involves tidal deformation and energy dissipation from the 2:1 orbit Figure 1-1. Abundance vs mass for chemicals sampled by resonance between Enceladus and Dione. As Cassini’s Ion and Neutral Mass Spectrometer (INMS) during Enceladus’ eccentricity is maintained by this resonance a flyby in 2008 (NASA/JPL/Space Institute) over one complete orbit there are measured periodic stresses from Saturn’s gravity. These stresses cause the The gas component sampled by the Ion and Neutron opening & closing of vents, and shear heating in the Mass Spectrometer (INMS) during a flyby through the Tiger Stripes trenches. These effects are thought to be plume includes water with trace amounts of ammonia, enhanced by the presence of a sub-surface reservoir and CO2, methane and light hydrocarbons see Figure 1-6. the thermal heat because they increase the flexibility in The solid component primarily (~99%) is salt rich with the ice shell overlaying the water body (Porco.C, 2014) a salinity between 0.5% and 2%, which is comparable (Bland, 2014). There are complications however with with Earth’s oceans. This salinity and the before research conducted by (J.Meyer, 2007) suggesting that mentioned high solid/gas mass ratio strongly suggests a tidal heating, and radiogenic heating at ~ 0.32 GW, global ocean or a regional sea sitting directly above a cannot account for the thermal emission of ~6 GW. rocky core (Porco.C, 2014) (Spencer J. , 2013). 3

Athena Exploring Enceladus to Investigate its Habitability

1.1.3 The Sub-Surface Reservoir 1.1.4 Life in Enceladus

There is growing evidence on the presence of a sub- Life on Earth has evolved in a manner suggestive of a surface reservoir on Enceladus. Some key evidence minimum set of prerequisites, which comprises includes the large solid/gas mass ratio, high salinity, the biologically useable energy, liquid water and essential flexibility in the ice shell, the presence of a negative biogenic chemicals. The hotspots or core could source mass anomaly in the south-polar region (L.Iess, 2014) the biologically useable energy; the global or local and that (Thomas P. , 2007) concludes that Enceladus is reservoirs could source the liquid water, and the differentiated, with a large low-density core, 30-40 km essential biogenic chemicals could be present in the crust (L.Iess, 2014) and a ~0.5 km topographic form of simple and complex organics, Figure 1-6. The depression at the SPT (Porco, 2006). The extent of the biosphere on Enceladus would therefore be a sub- reservoir is still highly debated with the two main surface hydrothermal reservoir deprived of solar energy theories being a global ocean or a ‘local’ sea (located at and likely to be in contact with the core and SPT crust. the SPT). For both cases a model made by (C.Glein, From research conducted by (Mckay, 2008) it is 2015) suggests that Enceladus’ reservoir is a Na-Cl-CO3 difficult to determine the exact mechanism that led to solution with an alkaline pH of ~11-12 and a salinity the origin of life on Earth, but some theories suggest that similar to the oceans on Earth. This reservoir would separated biological systems could evolve functionally erupt to the surface leading to the formation of the water equivalent adaptation by means of a convergent vapour plume and ice above the SPT. evolution. Therefore, a biosphere on Earth could mimic According to (L.Iess, 2014) “A regional sea is the conditions found on Enceladus. Some examples of consistent with the gravity, topography and high local these are hydrothermal vents and seafloor sediments. In heat fluxes and does not suffer from the thermal these biospheres, microbial life would have begun problems that a global ocean encounters”. It therefore where chemical energy was available from a mix of H, would be constrained to a region no larger than the S and Fe compounds, and would resemble southern hemisphere with a layer ~10 km that decreases methanogenic chemoautotrophs. These microbes feed in thickness as it approaches the equator (Porco.C, off molecular hydrogen and CO2, thereby releasing CH4 2014) (L.Iess, 2014). Evidence for the presence of a as a metabolic by-product, Figure 1-9. Molecular global ocean is based on Enceladus’ physical liberation hydrogen is likely to be present from water, while CO2 of 0.120° +/- 0.014°, Figure 1-8a (Thomas P. , 2016). and CH4 has already been detected by Cassini’s INMS The team reports that this value is too large to be instrument, Figure 1-6. consistent with Enceladus’s core being rigidly connect to its surface, and thus implies the presence of a global ocean rather than a localized sea. While evidence is present for this extent of the ocean their model is based on some simple approximations and therefore the exact characteristics of this ocean are not defined. Figure 1-8b points to a crustal thickness that is constant at ~25 km from latitudes between 0° and -50°, but then decreases to ~15 km at -90°, all with an ocean layer thickness of at least 30 km (Thomas P. , 2016). Figure 1-9. Possible methane cycle on Enceladus. This cycle is present on Earth in hydrothermal vents and 'Biology' could be microbes like Bacteria or Archea (Mckay, 2008)

Figure 1-3. Possible structure of Enceladus if a global ocean is present. (a) The amplitude of the physical forced liberation as a function of the icy shell's thickness. (b) Schematic cross section of the upper 60 km of Enceladus from the South Pole to the equator. Vertical exaggeration is 4.5 (Thomas P. , 2016).

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Athena Exploring Enceladus to Investigate its Habitability

1.2 Science Goals, Objectives, reservoir. Our investigation will concentrate on the Investigations and Measurements Tiger Stripes physical properties, southern hemisphere 1.2.1 Science Goals – Explore Enceladus to sub-surface imaging and North Pole & South Pole rotation features. Investigate its Habitability The main three sub goals for Athena will focus

on the characterisation of Enceladus from data collected Enceladus with its thermal presence, spewing geysers once during the 10 months after it enters an orbit around and chemical composition could hold a potential habitat Enceladus in July 2039. The three secondary sub goals for microbial life under its crustal layer. Athena will investigate long term (25-35 years) temporal therefore proposes the goal to explore Enceladus to variability in the hotspot, jet activity and plume activity investigate its habitability. Habitability relies on the by comparing the Cassini data to the Athena data. presence and interaction of three main elements: (1) Cassini completed its final (22nd) close flyby in Dec Essential biogenic chemicals (2) biologically useable 2015 after a 10-year period. Even though Cassini energy and (3) liquid water. For Athena we have split performed a few flybys it still achieved high resolution our main goal into three primary sub goals (1) Life (2) optical, chemical and thermal data sets of specific Hotspot and (3) Ocean and three selected secondary sub regions on the SPT. Upon completion of the Athena goals. The primary sub goals shall investigate a data set observation life an extended mission could see repeat collected once during Athena’s operation life (~10 observations over designated areas of Enceladus, such months long) and the secondary sub goals investigate as the SPT. This will allow the comparison of even temporal variability by comparing the data set collected higher resolution data sets with a period of ~ 10 months by Cassini to the data set collected by Athena, see separating them. below.

1.2.2 Objectives and Investigations 1.3 Measurements and Instruments

Athena three primary and one secondary sub goals are The objective of the sub goal ‘Life’ is based on a split into three primary and three secondary convergent evolution, which is discussed in Section investigations respectively. 1.1.4 Life in Enceladus, and determining if there are The first investigation is to provide evidence for extant methanogenic chemoautotrophs in Enceladus’ the presence of essential life-creating chemicals and sub-surface reservoir. Our investigation will identify if they are biogenic in origin. To answer this concentrate on finding evidence for essential life- investigation, we will perform four measurements creating chemicals and whether they are biogenic in shown in Table 1-2. The identifier column refers to the origin. identifiers used in the Science Traceability Matrix The objective of sub goal ‘Hotspot’ is based on (Figure 1-10). L1i(a-c) are molecular based improving the understanding of the hotspot located at measurements designed on the detection and the SPT. Our investigation will concentrate on characterisation of microbes (for example bacteria and characterising the hotspot’s physical properties. archaea) while L1id will provide complimentary The objective of sub goal ‘Ocean’ confirms the presence elemental analysis. and constrains the extent and location of a sub-surface Investigation Measurement Discussion L1ia Analyse water sample for presence of Positive identification of biopolymers such as DNA and RNA replicatory polynucleotides. which are present in all microbes will provide strong evidence for the presence of life. L1ib Analyse water sample for presence of Identification and then analysis of the microbe cell structure, cellular structures, membrane material membrane material and methane based products, which are and expected metabolic products. Shall present in most microbes. Chirality is included as many also analyse chirality, carbon pattern biologically active molecules are chiral. Carbon pattern and and C-C bond saturation ratio. C-C bond saturation ratio assist in the analysis of which are the buildings blocks of life. L1ic Shall identify and provide isotopic Isotopic ratios of molecular species and elemental species (in ratios of individual molecular species particular methane) will help distinguish between biogenic (Radiogenic and non-Radiogenic and environmental sources. H2, H2S and FeS are believed to including, N, H, C, O). Shall also provide the chemical energy in hydrothermal and deep sea search for the presence of H2, H2S and vents, see 1.1.4 Life in Enceladus. FeS. L1id Analyse water sample for presence of Complimentary elemental measurements to the molecular the common 'life elements' – CHNOPS. measurements performed in L1i(a-c). Table 1-2. The four measurements, including discussion, which will meet the investigation L1i (shown in the STM, Figure 1-10) in the goal Life 5

Athena Exploring Enceladus to Investigate its Habitability

The second and third investigation are to understand the will perform three measurements, which focuses on the hotspot and ocean on Enceladus. These investigations physical properties of the Tiger Stripes, the sub surface and any secondary investigations are shown in Table structure and rotation features at the North and South 1-3. The second investigation is about improving the Pole, the final three investigations comprise the understanding of the hotspot located at the south Polar secondary goals for Athena. The investigations are Terrain (SPT). And to meet this we will perform a shown in the STM, Figure 1-10 with the identifiers S1i, single measurement, which focuses on the thermal S1ii and S1iii. The purpose of the secondary physical properties of the hotspot, the Tiger Stripes and investigations is to help understand Enceladus’ the potential cause, tidal heating. The third investigation temporal properties and how features such as the Tiger is to confirm the presence and constrain the extent and Stripes, hotspot, jets and plume vary on timescales of location of a sub-surface reservoir. And to meet this we 25-30 years, or longer.

Investigation Measurement Discussion H1ia Characterise the sub-surface hotspot, Full analysis of the hot spots thermal properties. the 'Tiger Stripes' thermo-physical Measurements will involve spectroscopy and imaging scans properties and tidal heating. of the hotspot at a high resolution to characterise the source and mechanisms involved.

O1ia Monitor motion and rotation features at Monitoring tides and rotation features (vis measurements of the North and South Pole of Enceladus altimetry and surface tracking) can provide information on the to confirm the presence of a sub- ocean’s location, extent and density constraining its surface ocean and then determine the composition (Tobie, 2014). ocean’s extent, location and density. O1ib Characterise the crust’s sub-surface Research suggests the crustal thickness for a global ocean to structure to a maximum depth of 40 km be greater than 15 km, Figure 1-8, or for a local ocean to be between 30 – 40 km. Measurements on the sub-surface structure to depths of a few km will also be valuable for understanding the physical properties of the jets.

O1ic Characterise the physical (except To reach a reservoir through a thickness of at least 15 km is thermal) properties of the ‘Tiger impossible for current technology with the greatest depths on Stripes’ and jets. Earth are ~ 4km (About Ice Cores, n.d.). Therefore, measurements shall be made on the physical properties of the ‘Tiger Stripes’ and jets, as they are easier accessible and provide a direct connection to any possible reservoir.

S1ia Plume structure, ejection rates, particle Characterisation of the jets and plume dynamics will help size, particle velocity, time variability understand the physical processes involved including the and location of jets in the South Polar structures, mass loss rates and mechanisms in the sub-surface. Terrain (SPT)

S1iia Surface optical and sub-surface High coverage optical and sub-surface mapping can provide mapping details on Enceladus’ terrain and indicate the evolution history of Enceladus.

S1iiia Variance in the hotspot’s thermal The timeframe initially is the 25-35-year period between output over a timeframe observations by Cassini and observations by Athena, but it might also be the period between measurements in the Athena operation period and measurements in the Athena extended mission. The focus of this measurement is to determine if there is any variability in the thermal output of the hotspot over a period of time.

Table 1-3. The measurements and discussion, which are designed to meet the Hotspot, Ocean and Secondary Investigation investigations (shown in the STM, Figure 1-10).

.

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Athena Exploring Enceladus to Investigate its Habitability

Figure 1-4. Athena Science Traceability Matrix (STM).

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Athena Exploring Enceladus to Investigate its Habitability

1.3.1 Instruments 1.3.3 Payload B – Athena Orbiter

To answer the measurements set out in section 1.3.1 two The scientific payload located on the Athena orbiter payloads will be used. Payload A will be on a soft lander contains three groups of instruments designed to meet (called Sicily) and payload B will be on an orbiter the hotspot, ocean and all secondary investigations. (called Athena). Sicily is a soft lander designed to Listed below are the three groups. collect samples from the plume and the surface and Thermal remote sensing will be performed by three Athena is an orbiter that will provide remote sensing instruments: observations and provide a telecommunications relay 1) Thermal Mapper (TM) for Sicily. 2) Mid – Infrared Spectrometer (MIRS) Both payloads are composed of three sub systems: 3) Laser Altimeter (LA) Scientific Payload – the instruments that perform Optical remote sensing will be performed by two measurements, Non Scientific Payload – the elements instruments: used for outreach observations or used for providing the 1) Wide Angle Camera (WAC) instruments with samples, Communications Payload – 2) Narrow Angle Camera (NAC) the antennae, transmitters, receivers and any other Radio remote sensing will be performed by one elements that are required for telecommunication. instrument the Ice Penetrating Radar (IPR) This section looks to summarise the instruments used in There are no non-scientific instruments due to the Sicily and Athena sub-systems. Further instrument the remote sensing function of all the Athena information including their mass, power, heritage, TRL instruments thus there are no sample acquisition sample, environment and other system properties can be requirements. All outreach requirements will be met found in section 3 Model Payloads. Details on the through a combination of the Athena thermal, optical communication payloads are in Section 4.7. and radio remote sensing instruments.

1.3.2 Payload A – Sicily Lander 1.3.4 Considered Systems

The scientific payload located on the Sicily lander Ambitious missions, such as deep space missions are contains three groups of instruments designed to meet often highly competitive from a scientific deliverable the life investigation L1i, and the secondary goals perspective. Many mission proposals will see their investigation S1i. All the instruments are included to model payloads and their science objectives reduce in meet either one or more measurements, except the scope over the course of development, and in rare cases Raman Spectrometer which was only included to missions will be added to. This mission is no different compliment the GC-MS systems and SOLID and the potential for altering the mission aims and goals instrument. is ever-present. Often missions will also feature Life detection and molecular analysis will be performed multiple separate configurations in order to demonstrate by three instruments: multiple ways of achieving the science goals. 1) Signs of Life Detector (SOLID) The scope of this project only allowed for the 2) Gas Chromatograph – Ion Trap consideration of one configuration but there were a – Mass Spectrometer (GC-IT-MS) couple of alternative designs which were considered but 3) Gas Chromatograph – Time of Flight not developed with regards to project size. The two – Mass Spectrometer (GC-TOF-MS) main configurations which were considered were the Complimentary molecular analysis will be performed possibility of a sample return mission and an alternative by one instrument, the Raman Spectrometer (RS) orbiter/lander configuration where they would be a Elemental analysis will be performed by one combined system where the entire system would land instrument, the Alpha Particle X-ray Source (APXS) on the surface of Enceladus after completing all the orbital science. After a brief look at the available The non-scientific payload on Sicily contains two technologies, the feasibility of a sample return, without sample collection systems and one instrument included substantial technological development, was to meet outreach requirements. questionable and the more achievable lander and orbiter Sample collection will be performed by two systems: combination mission was selected. The technology for 1) Enceladus Surface Drill (ESD) an orbiter/lander combined system is more mature than 2) Enceladus Sample Collector (ESC) a sample return however is still a technological Outreach will be achieved through one instrument, the challenge with a greatly reduced amount of heritage Pancam (PC) available for a combination system. There are potential scientific benefits to using this approach however this is an area for further study. The final consideration is that the model payload may be subject to some alterations over the course of development - this will be covered briefly in section 4.18. 8

Athena Exploring Enceladus to Investigate its Habitability

2 Science Requirements 2.2 Sicily

There are three types of requirements that are relevant The science requirements are: to Athena’s operation and the Sicily operation life. For 1) Every instrument shall perform at least two Athena these are: altitude, resolution and coverage. For measurements on each type of sample; and Sicily these are: science, telecommunication and there are three types of samples: A sample from instrument. a depth of 0.5 cm, a sample from a depth of 5 cm and a sample from the plume 2.1 Athena 2) Any sample collection system shall not operate at the same time as any scientific instrument 3) The measurements for L1ia shall only be The altitude requirements are: performed on matter that has not fallen on the 1) Every instrument shall obtain the desired surface measurements from an orbital altitude of 50 km 4) The measurements for L1ib and L1ic shall be 2) The instruments shall be capable of imaging performed on all three types of samples from an altitude of 43 km for the purpose of 5) The measurements for L1id shall only be selecting a landing site performed on matter that has fallen on the 3) The communication system shall connect with surface the lander from an altitude of 43 km The telecommunication requirement is: The resolution requirements are: 1) To provide a communication every 32.9 hours 1) Every instrument shall capture images with a by powering its transmitter and receiver resolution of up to 25 m/pixel when observing the non-active area of Enceladus The instrument requirements are: 2) Every instrument shall capture images with a 1) All instruments (except SOLID) require a resolution of up to 1 m/pixel when observing sample size of 0.03 cc per measurement. And the active area of Enceladus SOLID requires a sample size of 0.06 cc (two 3) The resolution of the images obtained for the samples of 0.03 cc) purpose of landing site selection shall have a 2) To analyse one sample, it takes (for each higher resolution that 1 m/pixel instrument): 0.5 hrs for the GC-IT-MS, 0.5 hrs for the GC-TOF-MS, 1 hr for the RS, 3 hrs for The coverage requirements are: the APXS, and 0.52 hrs for the SOLID 1) Every instrument shall achieve a coverage of 3) It is required that the ESD shall take 1 hr per >60% during observations of the non-active sample delivery to the lander and the ESD shall region of Enceladus take 0.25 hr per sample collection and delivery 2) The IPR shall achieve a coverage of >80% during observations of the non-active region of Enceladus 3) Every instrument shall achieve a coverage of >98% during observations of the active region of Enceladus

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Athena Exploring Enceladus to Investigate its Habitability

3 Model Payloads 3.1 Model Payload Definition

The model payload consists of 14 instruments split and their associated science goals for the orbiter and between an orbiter and a lander: 6 instruments on the lander respectively. orbiter module with a total mass of ~209kg with a ~25% The instruments and their connections to each goal can margin calculated from various TRL levels & 8 be found in the science traceability matrix (|Figure instruments on the lander module with a total mass of 1-10). ~58kg with a ~40% margin calculated from various TRL levels. Shielding, power and other systems which impact mass have been estimated separately and will be covered in their relevant sub-sections. Table 3-1 and Table 3-2 summarise the model payload instruments

Model Instrument – Orbiter Science Contribution Narrow Angle Camera Physical characterization of ‘tiger stripes’ Rotation features at north and south poles Visual analysis of plumes and surrounding area Enceladus topography Wide Angle Camera Physical characterization of ‘tiger stripes’ Rotation features at north and south poles Visual analysis of plumes and surrounding area Enceladus topography Mid-Infrared Spectrometer Characterization of the hotspot Identification of heat sources Thermal analysis of plumes Temporal thermal variations Thermal Mapper Characterization of the hotspot Identification of heat sources Thermal analysis of plumes Temporal thermal variations Laser Altimeter Identification of heat sources Rotation features at north and south poles Location of the geysers Enceladus topography Ice Penetrating Radar Sub-surface structure Enceladus topography Table 3-1. Orbiter module instruments and the associated science goals

Model Instrument – Lander Science Contribution Gas-Chromatograph – Ion Trap – Mass Spectrometer Life-creating chemicals and isotopic ratios of any CHNOPS elements Plume composition Gas-Chromatograph – Time Of Flight – Mass Life-creating chemicals and any structural indicators Spectrometer Plume composition Signs Of Life Detector Life creating chemicals and specifically replicatory polynucleotides Raman Spectrometer Life creating chemicals Complementary analysis for life detecting instruments Alpha Particle X-ray Spectrometer Life creating chemicals and full elemental analysis Plume composition Enceladus Surface Drill Sample collection Enceladus Sample Collector Sample collection Pancam Complementary pictures for all lander instruments Table 3-2. Lander module instruments and the associated science goals

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Athena Exploring Enceladus to Investigate its Habitability

3.2 Athena Model Payload 3.2.1 Narrow Angle Camera (NAC)

The model payload for the Athena mission has been Narrow angle camera has the ability to image selected selected from a range of various different missions places with high resolution (1 m/pixel) at the chosen which will be examined in more detail within each altitude on Enceladus (50 km). Imaging south polar individual instrument sub-section. The instrument terrain of Enceladus from the high orbit then using NAC selections made in the model payload are derived from for imaging from the low altitude (few meters per pixel) the science objectives established in section 1.2. The could give clear information about the physical model payload serves as an indication that the outlined properties of tiger stripes, size and velocity of particles, goals are achievable using current, or near development location of jet source, structure of plume, rates of technologies. It also forms a placeholder instrument ejection particles and geyser locations which are selection which can be used to perform standard mission emitted in the South Polar Terrain (SPT). This camera analysis and define any mission parameters to examine uses an electromagnetic wave detector called a charge the viability of a mission of this type. coupled device (CCD). Therefore, NAC is really a very The instruments presented herein have been important instrument since it provides context imaging selected from various levels of heritage and all require that could be used by other instruments in synergy with some modification in order to be fully suitable for the the Wide Angle Camera (WAC) and the Laser Altimeter proposed mission and its goals. The impact of these (LA). The narrow angle camera can image black-and- changes has been accounted for in margin calculations white at high resolution through its 2.347-degree field but the parameters defined here should not be taken as of view. The heritage for the narrow angle camera of the final instrument suite. Due to the varying technical Athena is selected based on Mercury Dual Imaging readiness level (TRL) of the chosen model payload System (MDIS) of MErcury Surface, Space instruments, the accuracy of these instrument ENvironment, GEochemistry, and Ranging parameters is directly connected to their TRL. (MESENGER) mission (Solomon et al. 2007). The The environment of the Athena mission is quite main properties of wide angle camera are outlined in the diverse with many different temperature and radiation Table 3-3. factors that need to be taken into consideration. Radiation shielding has been analysed to a basic level NAC - Parameters Value for the survival of the instruments within the model payload. This change in shielding will affect the overall Focal length, m 0.65 payload but applied margins have accounted for these FOV Width on ground, m 2048 potential changes. Thermal analysis has also been performed at a basic level to ensure that the operation of Number of pixels 2048 the instruments is possible. Any alterations which are Width of each pixel 1.30E-05 required for the survival of the instruments will be Mass, kg 2 discussed in the relevant sub-sections below. Power, W 11 After the selection of the suite forming the Dimensions, m 0.27 × 0.26 × 0.12 model payload from the established science goals, these Table 3-3. Baseline NAC parameters instruments have formed the basis of all the decisions made in regards to all other systems. This model 3.2.2 Wide Angle Camera (WAC) payload is the mechanism to achieve the science goals so this study used this suite as the base line to examine the feasibility of this mission. Unlike Narrow Angle Camera, Wide Angle Camera has The conclusion reached using this model ability to image selected places with higher resolution payload is that with the appropriate shielding (both (5 m/pixel) at the chosen altitude on Enceladus (50 km). thermally and radiation sensitive) and sufficient It could provide wide spectral imaging for large development of the instruments that this payload will be different regions of Enceladus. This camera also uses an sufficient to fulfil the science requirements. electromagnetic wave detector called a charge coupled device (CCD). It will be explained with more detail in individual appendix (SES-RA-00001). The heritage for the Wide Angle Camera of Athena mission is also selected based on Mercury Dual Imaging System (MDIS) of MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESENGER) mission (Solomon et al. 2007). The main properties of wide angle camera are explained in the Table 3-4.

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Athena Exploring Enceladus to Investigate its Habitability

WAC - Parameters Value 3.2.4 Laser Altimeter (LA)

Focal length, m 0.13 Primary objectives of Enceladus Laser Altimeter will be FOV Width on ground, m 10240 to generate a global topographic model of high- Number of pixels 2048 resolution and geodetic outline which will support a safe landing, precise targeting, and surface mobility for Width of each pixel, m 1.30E-05 future exploration activities and future science. Mass, kg 2.5 Scientific goals and measurements. Power, W 11 The distance between the orbiter and the Dimensions, m 0.27 × 0.26 × 0.12 surface Table 3-4. Baseline WAC parameters The returned laser pulse spreading The energies of transmitted and returned laser The heritage for the Laser Altimeter of Athena mission 3.2.3 A Mid-Infrared Spectro-Imager - is Lunar Orbiter Laser Altimeter of Lunar THERMAP (TM-MIRS) Reconnaissance Orbiter (LOLA), which is an altimeter pulse detection time-of-flight that combines a five-spot pattern. The precise distance to the Enceladus surface A mid-infrared spectro‐imager (TM-MIRS) is based on simultaneously is measured at 5 spots. Therefore, the THERMAP instrument could be used for space providing 5 profiles across of the Enceladus surface, missions to small bodies in the inner solar system. The where dashes are receiver -of-view, and the solid heritage for THERMAP of Athena mission depends on circles are the transmitted laser footprints. the ESA Marco Polo-R mission (Groussin et al, 2016). Performance requirements. The LA of Athena mission THERMAP is used for imaging the surface and has some requirements as shown in Table 3-6below. It subsurface areas of Enceladus. It could provide has a high power of about 35W also has specific rang coverage of 60% of Resurfacing Region (RR) on operating temperatures. Enceladus and 100% of South Polar Terrain (SPT) with

same Width on Ground (2000 m) on both regions. It can LA - Parameter Value do two functions; thermal mapper and spectroscopy with excellent resolution imaging. It works within Orbit height, km 50 temperature range (5 – 15 Cᵒ). It could also explain the F.O.V. 1/e2, µrad 400 characteristics of hot spot and detect the sources of Frequency 28-Hz rate thermal heating and chemical properties of tiger strips. 2 (1 cold spare) In addition, Thermal mapper can determine location of Mass, kg 13 geysers, ejection rates, particle velocity, particle size Power, W 35 and the plume structure emitted in the South Polar Dimensions, m 0.37 x 0.25 x 0.13 Terrain (SPT). The main properties of THERMAP are Table 3-6. Some key parameters for the Laser altimeter of explained in the Table 3-5. Athena. All parameters are the same as its heritage instrument LOLA (Neumann et al, 2011) This instrument must operate in vacuum for one billion pulses. It has two of G2 laser diode arrays, and it THERMAP - Optical Imaging Spectroscopic required 50 km circular orbit (Neumann et al, 2011), as Properties Channel Value channel Value it shows in Table 3-7 the requirements operation of laser Focal length, 0.075 0.05 diode. The Athena Laser altimeter has different required m coverage for each region and it will use to map all FOV Width 8875 13850 regions of Enceladus except the rest of Enceladus and it on Ground, has width on ground of 200km. For example, the North m Pole terrain (NPT) requires 60% coverage, while the Wave length 8-29 8-29 South Pole terrain and resurfacing region need 100% range, μm coverage. Resolution, 8.67 25 m/pixel Package G2 (tow-bar array) Power, W 20.4 20.4 Mass, kg 7.5 7.5 Operational wavelength, nm 808 Dimensions, 0.28x0.255x0.14 0.287x0.255x0.14 Repetition frequency, Hz 28 m Pulse current, A 70 (max 90) Table 3-5. Baseline THERMAP parameters Current pulse width, ms 170 (max 200) Table 3-7. Basic operational requirements for the laser diode arrays of Enceladus (Neumann et al, 2011). 11

Athena Exploring Enceladus to Investigate its Habitability

3.2.5 The Ice Penetrating Radar (IPR)

Scientific goals and measurements. The study of the Other mission requirements. The radar sounder will subsurface and subsurface of Enceladus with a radar look nadir with accuracy of ±5°, and required altitude of sounder instrument will bring new data on the icy crust circular orbit 50 Km. According to a nadir view, the of icy moon (JUICE Science Study Team, 2011). antenna should illuminate the surface. Moreover, the • Physical global identification and physical local antenna should be parallel to the ground during the characterization and dielectric subsurface operations. Concerning yaw, a deviation should be less horizons than 1° from parallel to the ground. Furthermore, • Thermally controlled subsurface horizons another requirement of yaw, is having a small roll identify inside the ice shell. (<10°). Subsurface horizons of Enceladus will be identified and The antenna should be in the cross-track locally characterized at relatively high vertical direction (perpendicular to the flight direction). A resolution by obtaining subsurface sounding profiles of Doppler processing can be applied. The final down to a few kilometres. This instrument with the laser requirement is a high pulse repetition frequency (PRF) altimeter will achieved several goals in a (JUICE Science Study Team, 2011). complementary fashion. Athena radar has width on ground of 300 m, and Performance requirements. The Ice Penetrating Radar it will use to map all the region but it will coverage each (IPR) of Athena is a heritage of Rime radar sounder region with different percentage such as the North pole system at low frequency, which can penetrate the terrain the rest of Enceladus required 80% coverage the surface with a vertical resolution of several meters and south pole terrain need 100% coverage. can perform a subsurface analysis (JUICE Science Study Team, 2011). IPR Parameters Value Operating Temperature, K -73C -43C The central frequency selecting will rely on: Altitude, km 50 1) The radiation noise is higher at frequencies Antenna length, m 16 below 20 MHz Transmitted central frequency, 9 2) The operational mode of circular orbit around MHz Enceladus. Chirp length, µs 50 - 100 Mass, kg 10 In addition, this type of radar could be used as altimeter Power. W 25 with a moderate resolution. Decreasing the lower Dimensions, m 0.37 x 0.25 x 0.13 frequency to 5 MHz could lead to penetrate deeper into Table 3-8. Some key parameters for the Ice penetrating radar the Enceladus ice. of Athena. All parameters are the same as its heritage Performance requirements and concepts. The instrument Rime radar of Juice mission, except the altitude architecture of the radar of Enceladus which is based on which is 200-500km for the RIME (Bruzzone et al., 2013), RIME radar which is similar to the radar sounders of while the altitude of Athena radar 50 km. MARSIS and SHARAD. The Rime Radar has not been flown yet and has low TRL. This instrument is made up of a transmitter, a receiver, an antenna, and a digital 3.2.6 Athena Model Payload Summary system. The antenna is a dipole of 16 meters (two arms of 8 m), assuming a central frequency of 9 MHz (Bruzzone et al, 2013). The radar sounder can explore Table 3-9 gives an overview of the main parameters of different depths by changing the choice of the central the instruments on the model payload of the orbiter frequency. The less depth which is about 3 km at 50 module (Athena). The masses and average power values MHz could be accepted. The most important listed here have all had various margins applied based measurement is characterizing the sub-surface structure on their varying TRL levels. This is highlighted more in crust to a depth of 40 km. In addition, it can answer the section 4.11. The dimensions stated in this table are the secondary scientific questions, which will help in approximate total instrument sizes and these are the understanding the topography of Enceladus and values which have been used in any thermal, radiation indicators for Enceladus’ long-term geological or structural analysis. evolution

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Athena Exploring Enceladus to Investigate its Habitability

Instrument Acronym Mass, kg Size, cm Average TRL Heritage power, W Narrow NAC 5 27X26X12 10 5 Messenger Angle Camera Wide-Angle WAC 5 27X26X12 10 5 Messenger Camera Thermal 7.5 29X26X14 20 4 Bepi-Columbo (MERTIS) Mapper and TM Mid-IR Spectrometer Ice IPR 10 37x25x13 25 5 RIME radar /Juice mission Penetrating Radar Laser LA 13 35X35X29 35 4 Luna Orbiter Laser Altimeter Altimeter Table 3-9. Athena model instruments summary 3.3 Sicily Model Payload 3.3.1 Gas-Chromatograph – Ion Trap – Mass Spectrometer (GC-IT-MS) Payload A on the Sicily lander is comprised of five scientific instruments, three non-scientific instruments Science goals and measurements. The GC-IT-MS and the telecommunication package. The scientific and (complimented by the RS) will provide high resolution non-scientific payloads are primarily based on the analysis of the molecular structure. It will be designed instruments, which are included on the Philae lander as to use a GC capable of separating various compounds part of the mission to 67P/Churyumov- and isotopes chemically for subsequent detection with Gerasimenko (67P). The exceptions being the Raman an Ion trap MS and it will measure by chemistry and Spectrometer (RS), which is based on the RS found in mass analysis the H, C, N and O stable isotopes. The the , and the Enceladus Sample instrument will use step-wise heating from the two Collector (ESC) which has no heritage and will require available ovens (max temperature is shown in table full design and development. The Philae instruments below) to produce volatiles to allow analysis of the were primarily chosen because of our science objectives moon’s atmosphere (J. Biele, 2008). and the strict power (limited solar power) and mass Performance requirements. The GC-IT-MS instrument budgets (launch restrictions) required for the Philae is based on Philae’s Ptolemy instrument. The Ptolemy instruments by ESA. The Sicily lander will be required instrument was capable of measuring the overall to stay within a temperature range of 0 and 45 °C abundances of volatiles to a few % precision and throughout the duration of the mission and therefore this isotopic ratios with a precision up to 5%. And a is the required operating temperature range for all comparable performance will allow us to meet our Athena and Sicily elements. measurement requirements set out in section 2.1. From Listed below are the scientific and non-scientific the duty cycle in section 4.2.8 the GC-IT-MS will instruments that are included in payload A. The GC-IT- perform nine measurements and therefore a total of 18 MS and GC-TOF-MS complement each other and both ovens are required. From L1ic the heaviest nuclei that will be complimented by the RS. The RS will also need to be measured is FeS with a mass of ~88 amu. The compliment the SOLID instrument. The Enceladus Ptolemy instrument had a mass range up to and Surface Drill (ESD) and Enceladus Sample Collector including 150 amu therefore there is no reason to (ESC) will collect samples from the surface and plume modify the detectors performance (Wright, 2006). respectively. Each listed instrument includes a table with some key parameters. And a full parameter list can be found in the individual appendices Sam, Ash, Narmin and Rawaa.

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Athena Exploring Enceladus to Investigate its Habitability

GC-IT-MS Parameter Value 3.3.3 Signs Of Life Detector (SOLID) Operating Temperature, K 273 - 318 Number of ovens * 9 and 9 Science goals and measurements. The SOLID3 Oven Temperature, °C 400 and 800 instrument will perform molecular analysis using a Mass Range 12 - 150 microarray-based optical sensor detector. SOLID3 Target Ratios C, N, O, H makes use of -like molecules in order to detect Precision, % Few key (Parro et al., 2011) and will be a ground- Mass, kg 4.5 breaking discovery. If any complex identifiers of life are Power W 10 identified using this relatively untested technique, it will Dimensions, m 0.25 x 0.33 x 0.11 revolutionize life detecting instruments as well as the Table 3-10. Some key properties for the GC-IT-MS search for life within our solar system. This system instruments. All parameters are the same as its heritage forms part of the instrument suite upon the lander instrument Ptolemy, except (*) the number of ovens has module of the mission. Once on Enceladus SOLID will increased (Wright, 2006). begin its analysis towards the end of the lander lifecycle in science phase #3 (Table 4-3), this is due to allow fresh sample collection to be completed as these samples are 3.3.2 Gas-Chromatograph – Time of Flight – more likely to contain active biomarkers than ‘ground- Mass Spectrometer (GC-TOF-MS) deposited’ samples. SOLID is a highly focused instrument and as such will Science goals and measurements. The GC-TOF-MS only serve to detect biomarkers and signs of life with will provide high resolution analysis of the molecular particular focus on the detection of amino acids, structure. It will be designed to use multi-column gas and whole cells. The outcomes of SOLID will chromatography and high-resolution time-of-flight be very useful for directing any future missions to (TOF) to identify the molecular Enceladus by revealing more about the unknowns of the composition including complex organics, and chirality. surface. It can also address questions related to the prebiotic relevance, and the elemental, isotopic, chemical and Performance requirements. The SOLID instrument is mineralogical composition of the material. yet to be flown and as such has a relatively low TRL of Performance requirements. The GC-TOF-MS between 3 and 4. As this is a largely unproven instrument is based on the Cometary Sampling and technology (at least within space missions), the Composition experiment (COSAC) on the Philae requirements placed upon the SOLID3 are simply lander. The instrument The MS will be identical to the survival based. Thermal and radiation concerns are the one found on COSAC as there is no resolution biggest threat to any instrument at Enceladus which can requirements that are different. The number of columns be mitigated through shielding and heating solutions required is undecided but a similar number to that used which will be discussed in their relevant sub-sections. by COSAC (10 including two dedicated to chirality) is With its low TRL, some of the requirements have been foreseen. The instrument will use step-wise heating difficult to obtain in regards to SOLID3 so large (similar to Ptolemy) to heat the samples to 800 degrees margins have been applied where necessary to to produce volatiles to allow analysis of the moon’s accommodate the large changes as the instrument is atmosphere. From the duty cycle in section 4.2.8 the fully developed. The science traceability matrix (Figure GC-TOF-MS will perform nine measurements and 1-10) supports the inclusion of this instrument in the therefore a total of nine 800°C ovens are required model payload despite its low TRL. (Goesmann, 2006). Possible instrument concept. As discussed above, the GC-TOF-MS Parameter Value design of this instrument is outlined by (Parro et al., Operating Temperature, K 273 - 318 2011) and the only changes which will be made to the Number of ovens * 9 design revolve around the protection of the instrument Mass Range 12 - 1500 in a previously untested environment to ensure Oven Temperature, °C 800 measurements can be completed. With regards to Precision, % Few sample collection and handling there may need to be Mass, kg 4.5 some modifications to the original design which does Power W 10 not account for a roof mounted collection mechanism Dimensions, m 0.40 x 0.5 x 0.10 however this is not something which can accurately be Table 3-11. Some key parameters for the GC-TOF-MS examined before SOLID is developed further. instrument. All parameters are the same as its heritage instrument COSAC, except (*) the number of ovens has increased (Goesmann, 2006)

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Athena Exploring Enceladus to Investigate its Habitability

Other mission requirements. As this is a lander based Possible instrument concept. With the Raman instrument, the handling of any data outputs is handled spectrometer being placed within the lander, the thermal through communication windows with the orbiter. It is and radiation protection methods for the main body of important to keep the data volume produced by SOLID the lander will ensure that the instrument will survive in check as the storage solution for the lander, and by the harsh surface conditions of Enceladus. The heritage extension the orbiter, will need to accommodate all the this instrument is based upon is the Raman spectrometer data produced alongside any other instrument outputs. from the ExoMars Pasteur payload. This Raman The scheduling of SOLID3 places it towards the end of spectrometer used a laser based Raman technique. Table the mission lifetime of the lander to both allow the less 3-3 outlines some of the key parameters of the selected specific instruments to perform analysis and ensure the Raman spectrometer in order for it to perform at a level maximum outcomes of the model payload as well as to which satisfies the goals of the mission. allow for sample collection. Other mission requirements. As this is another instrument on the lander module of the mission, it will only be operating for a period of the mission on the 3.3.4 Raman Spectrometer (RS) surface of Enceladus. The preliminary research suggests

that the volume sample requirements and sample Science goals and measurements. The Raman analysis time are lower than SOLID and as such it can Spectrometer based on the lander module of the mission be operated at an earlier time during the lander life will compliment many of the instruments which are cycle. This is also due to the fact that the Raman is going present in the model payload with a focus on the search to perform analysis on the samples extracted by the drill for biomarkers, performing structural and molecular as well as the samples collected by the roof collector analysis and examining the composition of any which enables earlier usage. collected samples. The main measurements expected from the Raman spectrometer are: Stable elemental analysis on both freshly 3.3.5 Alpha Particle X-ray Source (APXS) deposited material collected by the roof collector and older deposited material Science goals and measurements. The APXS will be collected by the lander drill. used to obtain a full compositional analysis of any Complimentary analysis for samples which collected samples both from the surface of Enceladus have been analysed by and its plumes. The full elemental analysis will o SOLID3 compliment many of the other instruments and analysis o TOF GCMS performed by the lander however it’s primary function o IT GCMS will be searching for key signs of life. The elements the o APXS APXS will be searching for are the 6 most abundant The analysis performed by the Raman spectrometer will elements associated with life as we know it on Earth: be vital to contextualising any measurements obtained Carbon by the suite of instruments present on the lander. Hydrogen

Nitrogen Performance requirements. The operation of the Raman spectrometer, like most of the lander based instruments, Oxygen has thermal and radiation challenges which need to be Phosphorus overcome. The design of the Raman spectrometer in the Sulphur model payload is based on the ExoMars Pasteur payload These are known as the CHNOPS elements. and shielding and heating solutions are will be applied in order for the instrument to survive the vastly different Performance requirements. Performing a full elemental environment which Enceladus’ surface offers. The CCD analysis using an APXS requires low operation included in the model payload will be a 2048x512 pixel temperature, when compared to many other lander- detector which will allow for high resolution based instruments, to detect some of the ‘lighter’ spectroscopy. CHNOPS elements. The environment present on the surface of Enceladus helps facilitate this requirement RS Parameter Value and will be outlined further in the thermal sub-section Spectral resolution, cm-1 7 4.4. The instrument will still require some heating but Raman shift range, cm-1 150 – 3800 this will be less stringent compared to other instruments. As the instrument features a sensitive detector head and Laser wavelength, nm 532 using x-rays and alpha particles to perform its analysis, Instrument size, m 0.10 x 0.065 x 0.034 it is vital that the APXS has sufficient radiation CCD size, pixels 2048 x 512 shielding to survive for the duration of its operation. Table 3-12. Baseline Raman spectrometer parameters

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Athena Exploring Enceladus to Investigate its Habitability

Possible instrument concept. The heritage of this Therefore, verification that the drill can operate on instrument is the APXS from the Rosetta lander. It Enceladus is required and thus a potential redesign. The makes use of a curium-244 source to generate 5.5MeV SD2 system, and thus the ESD system consists of the alpha particles and 14keV x-rays (Klingelhöfer et al., following components (modified for operation on 2007). Alpha particle backscattering will be used to Enceladus) (Y.Bar-Choen, 2009): look for lighter elements (such as hydrogen and helium) 1) A drill/sampler that drilled the soil at depths of 0.5 while the x-ray mode will be used to complete the cm and 5 cm, collected the sample and discharged elemental analysis. The basic design is unaltered from them into the ovens and instruments the Rosetta APXS except that some basic shielding has 2) A carousel that accommodates the ovens been proposed to protect the instrument from radiation 3) A volume checker for estimating the collected and the thermal considerations on the surface will have sample volume an impact on the instrument design. Table 3-4 outlines 4) Electronics unit for managing the system with a some of the key parameters of the selected APXS high degree of autonomy required for it to achieve the established science goals. The SD2 system had an autonomy of three, however we are uncertain at this stage whether this is appropriate for APXS Parameter Value the Athena mission. Sensor head dimensions, 0.08 x 0.12 x 0.12 ESD Parameter Value m Operating Temperature, K * ~ 50 Detector area, mm2 10 Storage Temperature, K 50 Source Curium-244 Sampling Depth, m 0.05 Source activity, mCi 30 Material Strength, Pa 50 – 50 x106 Detector-sample distance, <20 Sampling Distribution ** Ovens + Instruments mm Sample Type Core/Powder Sample diameter, mm 17 Mass, kg 5 Resolution, eV 155 Power, W 6 – 14.5 Low energy threshold, 0.7 Dimensions, m 0.20 x 0.16 x 0.76 keV Table 3-14. Some key parameters for ESD. All parameters Table 3-13. Baseline APXS parameters are the same as its heritage instrument SD2, except (*) the operating temperature, which has decreased and (**) sampling distribution, which includes distribution to the 3.3.6 Enceladus Surface Drill (ESD) ovens and instruments, instead of only the ovens.

Performance requirements. The ESD is one of the 3.3.7 Enceladus Sample Collector (ESC) sample collection systems that Sicily will carry. Through a trade-off the drill shall resemble Philae’s Performance requirements. The ESC is a roof mounted Sample Drill and Distribution (SD2) system. Operation sample collected designed to funnel the plume’s matter of the ESD will follow very closely to the operation of into the ovens or instruments. The deposition rate the SD2 system, with a few minor changes. The ESD depends on the landing site location but our initial drill shall obtain a surface sample at a depth of 0.5 cm, design assumes a deposition rate of 0.2 mm/yr. and an extra sample extraction at a depth of 5 cm. Both Unfortunately, a deposition rate of this magnitude are samples will be used to compare to the sample collected only in the immediate vicinity of the Tiger Stripes and by the ESC, section 3.3.7. As the deposition rate on the thus a more likely deposition rate will be a smaller by a surface varies between ~10-1 to ~ 10-4 mm/year, section factor of 10 or even 100, see 4.2.7. 4.2.7, the drill will use both sample depths to meet the secondary sub-goals objective regarding temporal Possible instrument concept. The ESC has no heritage variability in the jets and plume. The ESD drill will then and no initial designs but for a deposition rate of 0.2 distribute samples to both the ovens and selected mm/yr we foresee a collection area between 0.56 and 2 instruments, see section 4.2.10. 1m . With a collection area between these two sizes the The drill will operate in a much colder required minimum sample volume of 0.93cc shall be environment (surface temperature ~75 K reached within 42 hours. With 0.93cc and 42 hours (ESA, Comet_Vital_Statistics_Article, 2015)) and it provided from our duty cycle in section 4.2.8. For a 2 will have to penetrate ice (compressive strength ~100 sample collector area of 0.56 m and the parameters, MPa at 100K). We envision the top few centimetres of Table 3-4, a preliminary mass of 13 kg is calculated. A the surface to have a compressive strength much less power budget of 5 W is for the mechanisms that collect, than 100 MPa, and the pressure will increase as depth compact and distribute the matter to the ovens or increases to a maximum of ~ 100 MPa. The SD2 drill instruments. For the material Aluminium was chosen meets some criteria for operation in a lunar environment because it is commonly used for a structure component. (Y.Bar-Choen, 2009) which will be similar to the environment on Enceladus.

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Athena Exploring Enceladus to Investigate its Habitability

3.3.9 Sicily Model Payload Summary ESC Parameter Value Material Aluminium Table 3-17 gives an overview of the main parameters of Mass, kg 13 the instruments on the model payload of the lander Power, W 5 module (Sicily). The masses and average power values Dimensions, m 0.75 x 0.75 x 0.1 listed here have all had various margins applied based Thickness, m 0.02 on their varying TRL levels. This is highlighted more in Table 3-15. Some suggestive parameters for the ESC. section 4.11. The dimensions stated in this table are the approximate total instrument sizes and these are the 3.3.8 Pancam (PC) values which have been used in any thermal, radiation or structural analysis. The max power values are also

listed in this table as these are the values which were The PC is the only instrument included for purely used to form the power duty cycle of the mission as a outreach purposes in either payload A or payload B. worst-case scenario. Sicily does not include an instrument that is capable of obtaining panoramic views of the lander’s surroundings. The camera will operate when power demands are low and it shall obtain one set of images with the capability of obtaining more. The camera will be an identical instrument the CIVA-P camera on Philae as its mass and power parameters and resolution requirements meet the ones set in section 2.1 (Bibring,

2007).

Parameter Value Number of Heads 5 (single) + 2 (stereo) Detector Type 1024 x 1024 CCD FOV, degrees 60 Spectral Sampling, nm Max @ 700 Spectral Range, nm 400 - 1000 Spatial Sampling 1 mm/pixel @ 1 m Mass, kg 1.19 Power, W 10.5 Dimensions, m 0.2 x 0.2 x 0.05 Table 3-16. Some key parameters for the PC. All parameters are the same as its heritage instrument CIVA-P (Bibring, 2007).

Instrument Acronym Mass, kg Size, m Average Max Heritage Power, W Power, W Gas-Chromatograph GC-IT-MS 6 0.25 x 0.33 x 0.11 11 12 Philae – – Ion Trap – Mass COSAC Spectrometer Gas-Chromatograph GC-TOF-MS 6 0.40 x 0.50 x 0.10 11 12 Philae – – Time Of Flight – PTOLEMY Mass Spectrometer Signs Of Life SOLID 10 0.40 x 0.40 x 0.40* 26 26 N/A Detector Raman Spectrometer RS 2.5 0.11 x 0.07 x 0.03 6 22 ExoMars Alpha Particle X-ray APXS 2 0.15 x 0.17 x 0.20 2 2 Rosetta Spectrometer Enceladus Surface ESD 6 0.20 x 0.16 x 0.76 7 18 N/A Drill Enceladus Sample ESC 18 0.20 x 0.20 x 0.05 3** 6** N/A Collector Table 3-17. Sicily model instruments summary. *The SOLID3 dimensions have been estimated from earlier versions of this instrument. **This is the power potentially needed to move the samples from the collector to the main handling unit.

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Athena Exploring Enceladus to Investigate its Habitability

3.4 Conclusions and Recommendation 4 Mission Design 4.1.1 Mission Profile The model payloads selected for the orbiter and lander (comprising of 5 instruments on the orbiter and 7 The main driver for many aspects of the instruments on the lander) have been selected to achieve Athena/Sicily mission comes from the difficulties in a particular science goal by obtaining a specific set of reaching such a distant body. This has an impact on the measurements. These suites of instruments, if allowed ΔV of the mission (and hence fuel requirements), the appropriate development time, is capable of achieving large temperature ranges experienced during cruise and the science goals of Athena and the derived mission operation, power systems at such a great distance from properties are within achievable means. the sun and finally the challenges of maintaining Most of the instruments have good heritage and communications with such a distant spacecraft. The fuel the basic analysis indicates that with appropriate and ΔV requirements have been minimised where shielding and thermal considerations they would be able possible with the use of gravity assists with Venus and to survive the harsh environments during transit and Earth; at the cost of increasing the duration of the cruise mission operation and then operate at the required phase. The fuel requirements are also reduced by scientific level. making use of further gravity assists after arriving at the The key to the feasibility of this mission, as Saturnian system. demonstrated above, is ensuring a robust development Thermally the spacecraft has been designed in such schedule is allowed so that the less developed a way to keep all the instruments protected during instruments performed as expected to deliver the transit and operational after arrival. This includes the required scientific outputs. high temperatures experienced by the mission during the Venus fly-bys and the low temperatures experienced at Enceladus. RTG power has been selected to accommodate for the lack of solar flux at Enceladus (else the system would require a very large solar array to provide comparable power) and this obviously has a few knock-on effects to the mission’s design for both thermal and safety concerns. A large high gain antenna is included to compensate for the large distance signals have to travel (~109 m). The high gain antenna also serves as a sun shield to help manage the thermal concerns of the spacecraft. The spacecraft will perform a series of tours of the moons (Titan – Rhea – Dione – Tethys) before finally arriving at Enceladus. During these tours and the gravitational assists achieved at Venus and Earth there is the opportunity to conduct extended observations. This is a potentially highly valuable prospect however it has not been considered at this stage. A more detailed mission profile is outlined below.

The following phases of the mission can be identified: 1. Launch and Earth to Saturn (Cruise) (10 years) 2. Saturn to Enceladus (Moon’s tour) (2.5 years) 3. Enceladus science orbits (300 days) 4. Lander descent and operation life (70 hours) 5. End of mission Total mission duration is approximately 14 years where the spacecraft would be in the Saturn system for 4 years. The selected launch date in 2026 would give an end date to the mission at the end of 2040. Using these values, the required propellant mass would be 2126 kg which would provide a total ΔV of 1649 ms-1.

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Athena Exploring Enceladus to Investigate its Habitability

4.1.2 Earth to Saturn (Cruise) 4.1.3 Launch

The combined system (Athena and Sicily) shall use the Athena mission will be launched by one of the launchers trajectory chosen by another proposal mission, THEO from the Atlas V500 series, and is being operated by (MacKenzie et al., 2016). This was chosen after a trade- United Launch Alliance. It is found that the off with four other trajectories (for more information see characteristic energy (C3) for this mission is 14 km2/s2, SES-SB-00001). This trajectory (referred to as the therefore the Atlas V511 is the most suitable launcher Cruise) has a launch date in April of 2026 and a Saturn for the mass of Athena and Sicily. The Atlas V551 has orbital insertion date in April of 2036, Figure 4-1. This capacity to lift 4680 kg and the total budget mass for 10 yearlong cruise includes one flyby of Venus and two this mission is 4331 kg. The volume of the orbiter flybys of Earth. A flyby of Jupiter isn’t possible as Athena system is 5 m3 and the Lander Sicily system is Saturn and Jupiter are not in favourable positions in 1 m3. All these reasons make Atlas 551 appropriate their orbits. The total ΔV for this cruise is 0.044 km/s, launcher for this mission. The NASA launch program which is very low compared to similar trajectories was used to make the plot below, which is the available. The C3 required to launch the combined characteristic energy (C3) versus mass, as it shown in system on this cruise is a maximum of 14 km2/s2. This Figure 4-2 below. is a maximum because the THEO team didn’t state the C3 required for this trajectory, but after some analysis a max C3 of 14 km2/s2 is deemed appropriate. During the cruise the combined system will perform a flyby of Venus. While Venus observations weren’t considered when determining the science goals and investigations, it does provide an opportunity to test the remote sensing instruments on board, and obtain images for outreach objectives. Thermally Venus offers an issue too. As the combined system will spend multiple years at Saturn its thermal system shall have to be designed for an extreme range of temperatures. This has been considered and discussed in section, 4.4. Figure 4-2. Characteristic energy (Atlas V series) After the 10-year period the combined system shall (Elvperf.ksc.nasa.gov, 2016). arrive at Saturn and enter an orbit at the orbital radius of Titan. This is desirable because it allows the opportunity to reduce the ΔV required to enter an orbit around In addition, it seems that the Atlas V C3 information has Enceladus, see Section 4.2.3. been removed from the user guide to plot it vs deliverable payload. Atlas V 551 has many specifications which are the launcher Atlas V 551 can carry payloads of up to 4680 kg and can contain a 5-meter payload Fairing to Low Earth Orbit. In addition, 8,700 kg is Geostationary Transfer Orbit Capability and it can be directly injected in Geostationary Earth Orbit (GEO) of Payloads up to 3,960 kg (Spaceflight101.com, 2016). The volume of Athena and lander is 6 m3 and the volume of Atlas 551 short is about 100 m3 fairing, it looks very big to payloads of mission. This launcher has five Solid Rocket Boosters, on top is installed on a Common Core Booster with the Payload Fairing 5-m and a common single engine Centaur that can do multiple burns in order to carry payload to a different types of orbits such as Low Earth Orbit, Earth Escape Trajectories, Geostationary Transfer Orbit and Geostationary Orbit. Many missions have been used this Atlas 551 such as NASA’s New Horizons (2006) and the US Navy’s MUOS Satellite Fleet as well as Figure 4-1. The designed trajectory for the Athena mission. The combined system (Athena + Sicily) will launch from Spacecraft (2011) (Spaceflight101.com, 2016). This Earth, then perform flybys of Venus and Earth and finally mission will launch from SLC-41 at CCAFS (Cape st arrive at Saturn at Titan’s orbital radius. The total ΔV for this Canaveral Air Force Station) on the 1 of April 2026 cruise is 44 m/s and the required C3 is a maximum of 14 (Spaceflight101.com, 2016). 2 2 km /s (MacKenzie et al., 2016). 20

Athena Exploring Enceladus to Investigate its Habitability

4.1.4 Saturn to Enceladus (Moon’s Tour) 4.1.5 Enceladus Science Orbits

After the combined system has launched from CCAFS The design of Athena’s science orbits will be based on and completed the cruise phase it will perform an orbital the work done by (R. Russell, 2009) on the design of an insertion manoeuvre at Saturn at the orbital radius of Enceladus science orbit. Science orbits are required to Titan. The following phase of the mission is called the have low altitude, low eccentricity and a high moon’s tour and it is based on the THEO concept inclination. A low altitude will improve the resolution mission (MacKenzie et al., 2016) and the suggest values allowing for a greater science return, a low eccentricity and properties are based on the work presented in their will help keep the achieved resolution constant at report. After the orbital insertion the combined system different points in an orbit, and a high inclination will will, over a period of 2.5 years, perform multiple tours allow both tidal bulge observations and coverage of the of four moons of Saturn. In total the combined system entire body without the need for inclination changes, will spend 8 months touring Titan, 7 months touring therefore saving propellant mass. Unfortunately, at Rhea, 11 months touring Dione and 4 months touring Enceladus, Saturn’s third body effects drives an Tethys, Figure 4-3. As time progresses the semi major instability leading to a possible impact on the order of 7 axis of the combined system’s orbit will decrease until – 14 days. Athena’s orbits therefore will be split into it is comparable to the semi major axis of Enceladus. At two types: low inclination and high inclination. this point an Enceladus orbital insertion manoeuvre Athena’s low inclination orbit is 65° inclined, shall occur. After this the Athena operation life shall because 65° is the near limit for long term stable orbits start. Although this tour shall take 2.5 years it is more suggested by (R. Russell, 2009). The combined system feasible than a direct cruise to Enceladus orbital (Athena and Sicily) will use the Laser Altimeter (LA) insertion because this requires a ΔV of several km/s. throughout multiple orbits to monitor any altitude Instead the total ΔV is 956 m/s and this is because the changes or signs of instability. Low inclination orbits cruise requires a ΔV of 44 m/s, the orbital insertion are required for the majority of the operation and by requires a ΔV of 712 m/s and the moon’s tour requires using the LA we aim to better understand Enceladus’s a ΔV of ~200 m/s giving. gravity field. If an orbit at 65° encounters instability, it is possible to decrease its latitude until long term stability is found. Athena’s high inclination orbit is 90° inclined. Between 65° and 90° there is a great risk of an impact with Enceladus (R. Russell, 2009). Unfortunately, in order to meet the spatial resolution requirements (through nadir imaging) of any one Tiger Stripe and the surrounding SPT, it is necessary to spend long periods of time between these two latitudes. Athena will therefore carry a substantial weight of propellant (to counter any instability in the high Figure 4-3. An illustration of the proposed moon’s tour in the inclination orbit (see section 4.5.1). Constant Saturn system. The combined system would complete 2.5 monitoring will be provided with the LA instrument by years orbiting these four moons of Saturn (MacKenzie et al., using active bursts at specific times throughout an orbit. 2016). To mitigate any further risk when the system is required to be in a high inclination orbit it will increase its latitude in increments of 5° to reduce any severe perturbation or gravitational issues. For the majority of Athena’s operation period the high and low inclination orbits will be at an altitude of 50 km, an eccentricity of zero and therefore with an orbital period of 3.41 hours. A 50 km altitude was selected because it meets the altitude requirement for all Athena’s instruments, and an eccentricity of zero was selected because there are no respective mission requirements and science orbits require a low eccentricity. For certain phases of the mission Athena will drop its altitude to 43 km which has two benefits: (1) higher NAC spatial resolution of 0.86 m/pixel, required for detecting obstacles that could risk the landers descent (2) the orbital period is then equal to 1/10th the moon’s rotational period, so every 10 orbits Athena will pass over the same position on the moon’s surface. 21

Athena Exploring Enceladus to Investigate its Habitability

4.1.6 Regions of Enceladus 4.1.7 The Sicily Landing Site

Enceladus has a surface area of 7.99 x105 m2 and for the The average Cassini optical spatial resolution of SPT mission this area has been split into four regions of region is several hundred metres/pixel (Th.Roatsch, study, see Figure 4-1. The four regions are designed 2008). As Sicily has a landing area of a couple of meters around the science investigations, which involves the and a ground clearance of ~1 m we require a resolution plumes’ mass deposition rate, Figure 1-5. The mass less than 1 m/pixel. Therefore, a landing site will be deposition rate is vital for the collection of samples selected after a period of 1 and 0.86 m/pixel images has through the ESC on the roof of Sicily, and optical, been acquired by Athena, see section 4.2.8. The SPT thermal and radar remote sensing measurements on region is the primary choice for the Sicily landing site Athena. Figure 1-5 is therefore the basis used for the but some of the terrain here is likely to pose a great risk duty cycle and thus operation period of Athena, see because of the 50-100 m amplitude, tightly hinged section 4.2.7. ridges and troughs that surround the Tiger Stripes. The South Polar Terrain (SPT) radius reaches Landing in a Tiger Stripe trench or in the surrounding the edge of the Tiger Stripes so all Tiger Stripe and SPT troughs is possible as there is a characteristic spacing of measurements will occur in this region. The a ~1 km, but it will require a guidance system on board Resurfacing Region (RR) spans from the SPT edge to a Sicily that is capable of achieving a landing site with a latitude of 45°S, which is the edge of a ‘thick’ band of radius of 0.5 km (or less). deposition. The North Polar Terrain (NPT) covers the To meet the life sub goal, it is crucial to land region of Enceladus where there is no observed Sicily where the deposition rate is high but the risks of deposition at any longitude and the rest of the surface landing is low. Sicily will carry on its roof the ESC (REST) covers the region between the NPT and RR system which will funnel matter into a volume measurer region and consists of a tooth like deposition pattern. and compacter inside the lander. The dimensions for These regions are based on the requirements of the this ESC is fixed prior to launch, the landing site is mission and might vary in size or definition to those selected after launch and will therefore depend on the labelled in the science community. predicted deposition rate provided by (Kempf, 2010). The NPT region is also useful for rotation Study into the ISS data suggests that the lower two- feature observations, comparison measurements and thirds of the main trunk of Damascus sees the greatest end of mission impact. The SPT region is also useful for jetting activity; with the lower part of Baghdad also rotation feature observations and landing site selection. seeing strong jetting activity (Porco.C, 2014). And from The RR region is also useful for comparison the model a high rate of ~ 0.5mm/year is found near any measurements and landing site selection. Tiger Stripe (Figure 1-5). A landing site around the Damascus and Baghdad Tiger Stripes is preferred, but a site in the immediate vicinity of any Tiger Stripe is likely to be acceptable. If a landing site near a Tiger Stripe is unavailable, the deposition rate will decrease from ~10-1 to ~ 10-4 mm/year as the distance from the Tiger Stripes increases (Figure 1-5). The landing site might be in the RR region as the deposition rate is less than~ 10-4 mm/year. The ESC initial design is based on an average deposition rate of 0.2 mm/year so a likely increase in its dimensions, and thus mass (to account for a more probable deposition rate) is foreseen.

4.1.8 Athena Operation Life

The Athena operation life is the period between its Enceladus orbital insertion and end of mission trajectory, Table 4-1. It contains ~2130 orbits spanning 300 days between 07/2039 and 05/2040. This period is split between science acquisition and non-science Figure 4-4. Shows the four regions of Enceladus. The acquisition. During the science acquisition phases northernmost region (green) is the North Polar Terrain (NPT) Athena is collecting science using two types of orbits, a and is between 60 and 90°N. The southernmost region low inclination orbit and a high inclination orbit. This is (orange) is the South Polar Terrain (SPT) and is between 70 necessary because only using low inclination orbits the and 90°S. The region above the SPT (black) is the resurfacing mission won’t achieve the resolution requirements of region (RR) and is between 45 and 70°S. The final region is the Tiger Stripes and the SPT region (north of ~80° the rest of Enceladus (REST) and is between 45°S and 60°N. latitude) for its NAC and WAC instruments. 22

Athena Exploring Enceladus to Investigate its Habitability

The measurements are prioritized with the NAC and LA and radar science and then enter its end of mission in the SPT and RR region being the top priority. This trajectory. There are two possible end of mission will allow the mission to select a landing site, reduce the trajectories for Athena (1) a Hohmann impact orbit on radiation dose on Sicily and the propulsion demands for the North Pole or (2) an escape trajectory with a 90° Athena. inclination. Option (1) is preferred but it might not be The combined system (Athena and Sicily) will possible due to a planetary protection issue, either way, enter an equatorial orbit around Enceladus and remain Athena will carry the propellant for both manoeuvres. there for completing system checks and assessing the During all the mission phases Athena will gravitational field. It will then start taking its landing remain in telecommunications contact with Earth for site imaging phase to locate as many potential landing 50% of an orbit. For the other 50% of the orbit Athena sites in the SPT and RR regions. The system will then will spend it acquiring science data. For the moments decrease its altitude to 43 km and start localised higher when Athena is not acquiring science it will be available resolution images of specific landing sites. Once the for either science acquisition or any other mission localised imaging is completed and the landing site is activity. In addition to communication with Earth, confirmed the Athena orbiter will adjust its orbital during the Lander Descent Orbit and Sicily Operation parameters to allow the Sicily lander to separate. The Life phases Athena will be in communication contact lander will then detach using its 1N and 20N thrusters with the lander (see section 4.2.10 for more and enter a Hohmann transfer orbit with the apoapsis at information). the selected site. Athena will then adjust its orbit to The length of the Athena operation life assumes place Sicily at its nadir where it will remain for ~ 20 that each orbit is split 50/50 with no interruptions orbits at an altitude of 43 km. It can remain in this orbit because of eclipse periods, obstructions to line of sight for longer if required. During the descent and the by Saturn/other objects, over lapping of multiple swaths following ~ 21 orbits Sicily, Athena and Earth will (more than the predicted 10%), system checks (beyond communicate and transmit the lander science data. the initial 10 orbits at EOI), attitude or inclination Upon completion of all the lander science and manoeuvres, and any other reason that could cause lander communication Athena will complete its thermal delay. An increase in the mission length is predicted. Instruments used (Required Coverage) Phase Altitude Inc (°) # of SPT RR REST NPT (km) Orbits Enceladus 50 0 10 Orbital Insertion (EOI) Landing Site 50 65 452 NAC (100%) NAC (100%) WAC (90%) WAC (90%) Imaging #1 LA (100%) LA (100%) LA (60%) LA (60%) Landing Site 50 * 301 NAC (100%) NAC (100%) WAC (90%) WAC (90%) Imaging #2 LA (100%) LA (100%) LA (60%) LA (60%) Localised 43 ** 100 NAC NAC Landing Site (Specific) (Specific) Imaging LA (Specific) LA (Specific) Lander 43 ** 1 Descent Orbit Sicily 43 ** 20 Operation Life Thermal 50 65 330 TM (100%) TM (60%) Imaging #1 MIRS (100%) MIRS (60%) Thermal 50 90 223 TM (100%) TM (60%) Imaging #2 MIRS (100%) MIRS (60%) Radar Imaging 50 65 475 IPR (100%) IPR (90%) IPR (80%) Radar (80%) #1 Radar Imaging 50 90 222 IPR (100%) IPR (90%) IPR (80%) Radar (80%) #2 End of 50 90 1 Mission Trajectory Table 4-1. The complete operation period for the Athena orbiter. All orbits are with an eccentricity of zero and stated inclinations are the maximum values. Number of orbits correspond to the instruments coverage requirements. (*) Inclination will increase from 65 – 90 ° in increments of 5 degrees (see section 4.2.5) (**) Depends on selected landing site latitude. 23

Athena Exploring Enceladus to Investigate its Habitability

4.1.9 Sicily Descent Orbit Value Unit Transfer/Landing Sicily Lander design is based on the Surveyor 1 Lander Parameter to performs soft landings for the lunar mission will land Altitude at 43 km on the surface of Enceladus and the gravity turn will be Injection used as simple control strategy, the descent has many Altitude at end of 0 km stages as shown in Figure 4-5. transfer Velocity at 156.31 m/s Mapping: this stage starts after arrival Enceladus orbit Injection (Vi) at 50 km. Velocity at end 169.12 m/s Choosing landing site: this stage begins after mapping of transfer (Vf) the South Pole terrain and resurfacing region by Time of transfer 1.47 hr different types of instruments such as Laser Altimeter Velocity initial 150.04 m/s and NAC 100% at 50km. transfer (Vtrans i) Releasing: Table 4-2, Sicily releases from 43km to enter Velocity final 175.65 m/s Hohmann transfer orbit with orbit inclination dependent transfer (Vtrans f) on Lander Latitude (max 90 degrees), transfer ∆v to ∆ va 6.27 m/s next phase is 12.79 (m/s), Perturbation ∆v (km/s) = ∆ vb 6.52 m/s 0.155 3.02E-06 for total Time 1.47 (hrs) and number of Total ∆V (for 12.79 m/s orbit is one. At altitude 162 km, the six 20 N thrusters transfer orbit) will be activated. Total ∆V (for 175.65 m/s Free falling: at 5 m height the thrusters will be turned landing orbit) off and the Lander starts to free fall and settles on Table 4-2. The total velocity phase of descent Sicily. Enceladus surface. It will be supported by three hinged landing gear legs, Level requirements and contained thruster engine with one monopropellant The need to land the largest scientific payload from the tank. The primary objective of Sicily is to do highest altitude with the greatest precision of mapping measurement for elements that can support life on before descent drives the Athena descent and landing Enceladus ice moon by using different types of architecture. Mission requirements dictate that the instrument that will be on Sicily. system of the descent and landing shall deliver 617 kg Sicily from an altitude of 43km within 1km of the desired landing site (Way et al, 2007). The incentive of landing requirements and driving descent is to let the scientific community to choose the Sicily landing site of safe landing sites from the largest probable set to place the Lander in a location with the highest possibility of achieving the science goals (Way et al, 2007).

Figure 4-5. Steps of descent Sicily Lander.

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Athena Exploring Enceladus to Investigate its Habitability

4.1.10 Sicily Operation Life After hibernation, the first period will be an array of system checks and monitoring to test its on-board Sicily and Athena are a combined system until the systems to find any issues. In parallel with this Sicily lander descent orbit phase (see Table 4-1) where the will send the data to Earth through Athena. This phase lander detaches from orbiter and descends to the is where controllers on Earth will know about any issues surface. Until its successful landing Sicily will be in that exist and subsequently provide some further hibernation using power from its two RTG and commands. The majority of Sicily’s commands will be potentially Athena’s RTG. During this period of stored on board prior to launch because a hibernation only Sicily’s key thermal and electrical communication signal takes ~ 66 minutes to travel systems will use power. After landing Sicily operation between Earth and Enceladus, thus automation is a key life will consist of at least four different periods; part of Sicily’s operation life. (1) System Checks Science #1 and Science #2 use the same (2) Science (#1, #2, #3 and #4) instruments in the same sequence. The only difference (3) Communication is the ESD system will achieve a 0.5 cm depth in (4) Hibernation Science #1 and a 5 cm depth in Science #2. The total The operation life (Table 4-3) has been designed length of Science #1 and Science #2 is 44 hours (22 assuming that every component is working as intended. hours each). These two phases are devoted to sampling Further designing is required to mitigate the risk of the sub-surface content to meet the all their sub-goals components developing issues and to reorganise the except L1ia. operation period if these risks occur. Science #3 will see the first measurements on After awaking from hibernation the first phase allows the plume’s matter by the RS and SOLID instruments. for the initial system checks and its length has not been This phase is ~6 hours long. The SOLID instrument finalised however a few hours is likely sufficient. After won’t operate in Science #1 or #2 because matter that the system checks there is the science period, which is has been deposited on the ground prior to sampling will comprised of the four separate science phases and is ~67 have received greater radiation and surface processes hours long. The communications phase is ~ 10 hours compared to matter that is in the plume, potentially long and its length has been determined by the impacting our measurements for the life sub-goal L1ia. telecommunications payload on Sicily and Athena. Science #4 will also use the ESC system, but it Following the communications phase is the hibernation will feed the matter into the two GC-MS systems phase, where Sicily will enter its hibernation state again. instead of the RS and SOLID instruments. This phase is The complete Sicily operation life is ~ 80 hours. ~ 5 hours long. The distinction between Science #3 and #4 is that the priority is to perform measurements using Operation Life Instruments Communication the SOLID and RS instruments to meet sub-goal L1ia. Phases used with Athena The GC-MS systems will only provide complimentary System Checks Yes molecular analysis and will have already analysed Science #1 GC-IT-MS Yes samples in Science #1 and #2. GC-TOF-MS In the communications phase only the RS communication sub-system and the critical sub-systems APXS will be using power. The length of this period is a ESD (0.5cm) conservative ~ 10 hours and is determined by a Science #2 GC-IT-MS Yes combination of the lander-orbiter bit rate, margins and GC-TOF-MS the volume of collected science data (see section 4.7.2). RS The final phase is when Sicily will enter hibernation and APXS power only its key sub-systems. During this hibernation ESD (5cm) period it is possible to recharge the two batteries using Science #3 RS Yes the two RTGs. This can allow for an extended lander SOLID mission if desired, however this has not been ESC determined and will require further study. Science #4 GC-IT-MS Yes At the end of the four science phases a total of GC-TOF-MS 757 Mbits of data would have been collected. ESC Throughout each phase communication with Athena is Communication Yes possible. Athena will have one communication window Hibernation No with the lander per orbit, and with an upload bitrate of Table 4-3. Sicily operation life consists of seven phases and ~ 0.5 Mbits/s communication will take ~ 25 minutes or is ~ 80 hours. After the batteries have been recharged in the 2 to 19 windows (for more information see Section 4.6). hibernation phase the lander could have an extended mission. In bold are the sample collection systems. Further information on the instruments can be found in Section 3.

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Athena Exploring Enceladus to Investigate its Habitability

4.1.11 End of Mission 4.2 Radiation Environment and Shielding

The uncertainty of the gravitational environment Enceladus is located ~ 4 Rs (Saturn Radii) from Saturn present at Enceladus poses a significant concern to the in the densest part of the E ring. During the Cassini spacecraft at the end of its operational lifetime. This flybys its magnetometer instrument detected a bending leads to ensuring a robust plan is in place for disposing of Saturn’s magnetic field lines caused by the ionized of the spacecraft at the end of its lifecycle. To comply particles within the jets and plumes of Enceladus. The with the planetary protection requirements, which are radiation environment outside of the combined system highlighted in section 4.12.1, two end of mission will therefore be a combination of solar energetic proposals have been selected. The primary end of particles (SEPS), cosmic rays, trapped electrons and mission solution is a controlled impact into the north protons in Saturn’s magnetosphere, bremsstrahlung and pole of Enceladus. This eliminates the danger of leaving ionised particles from the jets at Enceladus. Gamma- Athena in an unstable orbit with an uncertain final rays and neutrons will also be produced by the RTG trajectory. This is the most desirable outcome with a ΔV system, but their impact on the total ionising dose (TID) of 7 ms-1. However, with a mission which is searching has not been verified. For the radiation analysis of the for signs of life, there is the potential for reclassification Athena mission, the mission has been split into four of the moon during the mission from a planetary sections (below) and for the sections (2), (3) and (4) the protection standpoint; this leads to the need for a second radiation analysis will be from a graph produced by the alternative end of mission plan. Saturn Radiation Model (SATRAD) (H.Garett, 2005). The backup end of mission proposal involves For section (1) a 10 krads TID is predicted from putting Athena into an escape trajectory away from radiation models (NASA, Enceladus Flagship Mission Enceladus into deep space where it will no longer pose Concept Study, 2007). a threat to Enceladus. The ΔV for this option is a slightly The assessment of the radiation environment is higher 64 ms-1. As this is a backup proposal which may split into four sections: have to be selected during the mission, the propellant 1. Earth to Saturn (Cruise) mass for both of these options. The key parameters of 2. Saturn to Enceladus (Moons Tour) each option are summaries in Table 44. 3. Athena at Enceladus 4. Sicily at Enceladus Option ΔV, ms-1 Propellant mass, kg North pole impact 7 4 Earth to Saturn (Cruise) describes the radiation Escape trajectory 64 32 environment that the combined system (Athena + Table 4-4. Key parameters of end of mission options Sicily) will experience between its launch on Earth and its arrival at Titan in the Saturnian system. The Unfortunately, with Sicily being a stationary combined system will be in this environment for 10- lander, there is little that can be done with it after the years. 10 krads TID has been predicted. end of the mission. Therefore, ensuring the lander is Saturn to Enceladus (Moon’s Tour) describes the designed to minimise risks to Enceladus (in particular radiation environment that the combined system from the radioisotope thermoelectric generators aboard (Athena + Sicily) will experience between its arrival at the system). Titan in the Saturnian system and its arrival at Enceladus. The combined system will be in this environment for 2.5-years. Throughout the 2.5 years the combined system will spend 8 months touring Titan (20 Rs), 7 months touring Rhea (9 Rs), 11 months touring Dione (6.3 Rs) and 4 months touring Tethys (4 Rs). For the radiation environment at Titan, as a top level assessment, it will be similar to the radiation environment at Rhea. Athena at Enceladus describes the radiation environment that Athena will experience between its arrival at Enceladus (4Rs) and its end of mission. Athena shall be in this environment for ~ 10 months. Sicily at Enceladus describes the radiation environment that Sicily will experience between its arrival at Enceladus (4Rs) and the end of its operation life. Sicily shall be in this environment for ~ 4 months, therefore the end date for Sicily’s operation life is ~ 6 months prior to Athena’s.

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Athena Exploring Enceladus to Investigate its Habitability

Component Max Aluminium Mass of For sections (1), (2) and (3) the radiation analysis is Assigned Shielding Shielding, based on the SATRAD model. Both systems (Athena + Dose, Thickness, kg Sicily) are comprised of multiple components (systems, krads mm sub-systems and elements) some of which are more SOLID 200 50 82 susceptible to the effects of radiation. Table 4-3 shows RS 200 1 0.04 a top-level approximation on the amount of shielding GC-TOF-MS 200 10 8 required for Payload A, Payload B and the electronics GC-IT-MS 200 10 4.3 compartments. Figure 4-6 is based on the radiation APXS 500 1 0.1 absorbed on a 1 m radius probe with a surface area of ESD 200 10 4 2 4π m , in selected orbits around Saturn. In section (2) ESC N/A 0 0 the combined system will spend certain amounts of time PC 200 10 2 touring four different moons, therefore four different Lander 200 10 5 Rs. In section (3) and (4) Athena will spend 10 months Electronics and Sicily will spend 4 months at the Rs of Enceladus. MIRS + TM 500 1 0.2 Using the components surface area, the time spent at LA 200 50 29 each distance and Figure 4-6, the shielding requirements NAC 200 10 36 for each component were calculated. The TID was kept WAC 200 10 below the Max Assigned Dose (krads) Table 4-5. The major shielding mass is used on the SOLID and orbiter IPR 500 5 1 electronics box; therefore, further analysis is required to Orbiter 200 50 124 reduce these values. The lander electronics box is a Electronics factor ~6 less than the orbiter electronics box because Table 4-5. A top level assessment of the amount of aluminium shielding required on the list of components. The the electronics box surface area is four times greater. Max Assigned Dose is based on an assessment of the components susceptibility to the effects of radiation. With 200 rads being assigned to components with electronics, silicon detectors or other radiation sensitive components. The Predicted Aluminium Shielding is calculated from the component’s surface area, its duration in each of the four sections and its Max Assigned Dose. The NAC and WAC are located together and therefore shielded by one set of aluminium. The ESC contains no sensitive radiation components so a shielding thickness of zero is predicted. Total shielding mass for Sicily ~107 kg and Athena ~190 kg.

Figure 4-6. Jovian and Saturnian radiation environments for a 1-week exposure. Rs is Saturn radii and Rj is Jupiter radii. Dose in rad (Si) for a 1 m radius aluminium spherical shell (H.Garett, 2005). The radiation environment at Saturn is ~ 1000 weaker than the radiation environment at Jupiter. For orbital distance greater than 9.47 Rs the 9.47 Rs plot has been used. For 4 Rs (Enceladus’ orbital distance) the values that are half way between 2.55 Rs and 5.95 Rs has been used. 27

Athena Exploring Enceladus to Investigate its Habitability

4.3 Thermal Environment spacecraft at approximately 300K which is within the acceptable range for each instrument. The value used There are many reasons for most equipment and for the total internal power comes from the heating electronic functions on Earth optimally at around same requirement during operation at Enceladus (covered in temperature of human. Each instrument has required section 4.4.4). rang temperatures which allows obtaining the desire measurements. In addition, instruments that are sent to 4.3.2 Saturn to Enceladus (Moon’s tour) deep space will operate in different environment which could have temperature higher or lower than what Earth It is clear that for Athena mission during Cruise Saturn has and could vary between day and light. Therefore, it Enceladus and orbiting the spacecraft will experience should be control the temperature of equipment by using low temperature that could affect the operations on different ways such as isolating from new environment instruments therefore, series action has been taken. The (Ball et al, 2007). Table 4-6 below shows that each thermal requirement is maintained the range operation instrument has range of operating temperature, and temperature of each instrument for the lander and controlling the temperature for each cruise of the orbiter are showing in Table 4-6. The effect of solar mission many techniques have to be used. ionizing radiation and illumination on the spacecraft is much higher when it flies to an inner Solar System body Instrument Module Operating such as Venus, because it is nearest to the Sun as Name Temperatures, °C compared to Mars. In fact, the solar flux of the MIRS + TM Orbiter 5 to +15 spacecraft heating is greater four times at Venus than at Orbiter 10 to +4 Mars (European Space Agency, 2016). Radiator LA surfaces ultimately reject spacecraft waste heat to space Orbiter −30 to +25 at Venus cruise there are many forms of radiators such NAC as panels deployed after the spacecraft is on, flat plate Orbiter −30 to +25 radiators mounted to the side of the spacecraft and WAC spacecraft structural panels (David et al, 2002, p207). Orbiter -73C to -43 For Athena mission we will use the flat plate radiators IPR at Venus orbit, therefore the requirements for radiator RS Lander -10 to 20 of Athena are all infrared (IR), waste heat and any GC-TOF-MS Lander 0 to 45 radiant heat loads from other spacecraft surface that are absorbed or the environment have to be rejected by GC-IT-MS Lander 0 to 45 radiator from its surface. Thus, area of Athena has been APXS Lander -20 to 5 calculated and the required area is 0.2 m2 the Table 4-7 PC Lander -103 below show the properties of our radiator. Table 4-6. Operating Temperatures of Athena mission Parameters Value Emissivity 0.8 4.3.1 Earth to Saturn (Cruise) Q, W 346 Temperature, K 443.9 The period between leaving Earth and arriving at Saturn Area, m2 0.20 is a portion of the mission where it is important to keep Density, kg/m3 1980 the instruments cool as the thermal environment Thickness, m 0.05 experienced during this phase is often higher than the Volume, m3 0.010 rated temperatures of each instrument. Mass, kg 20 To achieve this, the instrument payloads will be Table 4-7. The properties of radiator, the instruments, and protected from the Sun by making use of the high gain battery, electronics, and propulsion systems will be antenna as a Sun shield. The design for cooling the maintained the thermal system the temperature of within spacecraft is based on the temperature environment at operating limits. For example, a number of radioisotope Venus as this is the highest temperature the spacecraft heater units (RHUs) will be used in order will be subjected to. At this distance, the angular diameter of the Sun is ~0.75m while the diameter of the dish is 4m so the payloads will be protected. The other method of ensuring a suitable thermal environment is radiating any internally generated heat. The selected radiator design is a carbon-carbon composite material, to ensure a low mass with a suitable area and high emissivity (0.8), with an area of ~0.2 m2 radiating a total of 346 W and a mass of ~20 kg. This design maintains the equilibrium temperature of the 28

Athena Exploring Enceladus to Investigate its Habitability

4.3.3 Sicily 4.4 Propulsion

The thermal considerations which need to be made for The propulsion solutions required for the two separate Sicily are identical to those which need to be made for orbiter and lander modules are different as Athena Athena as the difference in thermal environment requires longer burns to successfully arrive at Enceladus between the orbiter and the lander is negligible. while Sicily requires relatively short burns to arrive at During operation at Enceladus, Sicily the surface of the moon from a stable orbit. The experiences a much cooler (~70 K) environment. To propellant mass accounts for the decrease in mass after ensure the instruments are maintained within their the lander detaches from the payload. More information operating temperatures and are able to deliver their key on these calculations can be found in [SES-SB-00001]. measurements, most instruments need to be heated during operation at Enceladus. This will be achieved through the use of radioisotope heating units (RHUs) to 4.4.1 Athena maintain a set temperature for each sub-system. 126 RHUs will be required (29 on the lander and 97 on the The propellant system for the main thruster of the orbiter) each delivering 1 W of thermal power to the mission is based on MMH/N2O4 bi-propellant with a desired location. Table 4-8 and Table 4-9 offer a total propellant mass of 2111 kg. The main engine is a detailed breakdown of the distribution of the RHUs 400 N apogee motor with a specific impulse (Isp) of 318 within the lander and orbiter respectively. s and a mass of 3.60 kg. The required propellant mass has been minimised by utilising gravity assists where possible as outline in section 4.2. Instrument Number of RHUs required Six 4 N thrusters are employed in the orbiter RS 1 design for attitude control and will also be using the APXS 1 MMH/N2O4 bi-propellant. The mass of propellant Pancam 1 required for any attitude control manoeuvres has been GC-TOF-MS 18 encompassed in the total mass of propellant listed above GC-IT-MS 8 as both engine types use the same propellant. The 4 N Total 29 thrusters have a Isp of 290 s and a mass of 0.65 kg each. Table 4-8. RHU distribution for lander Parameter Main thruster Attitude thruster Instrument Number of RHUs required Thrust, N 400 4 TM/MIRS 8 Isp, s 318 290 WAC 37 Fuel MMH MMH NAC 37 Oxidiser N2O4 N2O4 LA 13 Engine mass, kg 3.60 0.65 IPR 2 Table 4-10. Summary of engines selected for orbiter module Total 97 Table 4-9. RHU distribution for orbiter The propellant and oxidiser used by the two different engine types of the orbiter will be stored in respective Some multi-layer insulation will likely be employed in tanks each with a dry mass of 45kg. order to prevent latent heat loss from the spacecraft structure itself during operation. Should the power generated from the RHUs and the systems themselves 4.4.2 Sicily become too great, the radiator design for the Venus thermal environment is suitable to provide cooling as it The propulsion system used by Sicily is chemical has been designed around the most demanding thermal system which uses monopropellant hydrazine. The total environment of the mission. Any heat transport will be wet mass of lander propellant is 661 kg include 20% managed within both the orbiter and the lander by heat margin. One hydrazine propellant tank with total mass pipes which will provide heating to instruments from 6.4 kg (including 10% margin) and its characteristics the RHUs and transport excess heat to the radiators to are highlighted in Table 4-12. In order to control the be dissipated. spacecraft attitude, velocity and decent, six 20N thrusters are used at the bottom of the lander while four 1N thrusters would be distributed on the sides of the lander. Table 4-11 shows the main parameters of two different types of thrusters. The operational steps of a hydrazine thruster start when the attitude control system triggers the thruster operation. The impulse produced by monopropellant hydrazine thrusters is typically between 220 to 240 seconds. 29

Athena Exploring Enceladus to Investigate its Habitability

There are other propellant systems that could be used 4.5.2 Sicily but Hydrazine was selected through a trade-off which will explain with more details in (SES-RA-00001). This Radioisotope Thermoelectric Generators (RTGs) can system could provide good performance for the supply electrical power for space craft in space missions functions of orbit maintenance and attitude control. by decaying the radioactive materials. The primary However, it could not provide high ΔV manoeuvres that power source for Sicily is 241-Americium Radioisotope are required for orbital insertion. However, bipropellant Thermoelectric Generators (241-AmRTGs) which was systems have ability to provide all these functions but chosen as it is cheaper, lighter and meets the scientific they are more complex (Sutton & Biblarz, 2010). requirements of Athena mission that need 70 hours of surface operation. Plutonium (Pu-238) has shown Parameters 1N Thruster 20N thruster excellent performance therefore it is the ideal fuel but it Volume, m3 7.68E-06 1.67E-04 cannot provide low electrical power in small missions Mass, kg 0.319 0.715 like Athena mission. The best alternative could be Specific Impulse, s 220 222 Americium-241 since it has a longer half-life more than Propellant Hydrazine Hydrazine 150 years and it is produced as a by-product of regular (N2H4) (N2H4) power reactors therefore, it exists in large quantities. Table 4-11. Baseline thruster parameters Two 241-AmRTGs will be enough to power Sicily. However, its power is around ¼ power density Parameters compared to Pu-238 and it could be emitting neutrons Volume, litre 104 to 177 more than Pu-238, therefore, it is a greater radiation Propellant volume, litre 78 hazard for people who handle it. It is not ready for using Propellant Hydrazine (N2H4) on the flight and a high level of development is required Mass, kg 6.4 (Perkinson et al., 2013). The main properties of 241- Table 4-12. Baseline hydrazine propellant tank parameters AmRTGs are shown in Table 4-13. 4.5 Power Another source of Sicily power is secondary Lithium-ion batteries (Li-Ion) which can provide a As with the propulsion system selections for Athena and power bus 28 V to subsystem. They are ordered in two Sicily, different considerations had to be made for each blocks and each one contains eight cells on Enceladus system as dictated by the requirements each payload lander. These batteries can work in different conditions imposes on the orbiter or the lander module. such as very cold, very hot (-40 to +70 ˚C) and sun radiation (Fellner et al., 2003). Two Lithium-Ion batteries were selected with total required cell capacity 4.5.1 Athena nearly 166.14 Wh. Table 4-14 shows the main characteristics of lithium ion batteries. Figure (4-7) The major limiting factor on power generation at explains the relation between lander instruments and Enceladus is the low solar flux (~15 Wm-2). This communication max power and time ant it will be eliminates the possibility of using solar panels on a explained with more details in (SES-RA-00001). mission such as this without having excessively large panels (discussed more in [SES-AS-00001]). Type of reactor Americium 241 With the inability to make use of solar power, Half life 432.6 the selected power system be a plutonium based Multi- Rate of efficiency decay 1-0.5^(1/half-life) Mission Radioisotope Thermoelectric Generator (MMRTG). It is based on the Mars Science Laboratory Percentage decay per year 0.16 MMRTG system and will produce 110 W of electrical End of mission time (years) 14 power at the start of life, degrading to ~100 W end of life performance. This power output will be sufficient to Initial power output of system 10 supply the instruments on board the orbiter once the (watts) duty cycle has been accounted for as not all instruments will operate at the same time. The system will also have Power output of system at 7.76 approximately 2000 W of thermal energy which has to Saturn (watts) be dealt radiated from the system. This is achieved by Table 4-13. Baseline 241-AmRTGs parameters making us of the included radiators of the MMRTG which will radiate this energy into space. The inclusion of an RTG system also introduces a safety concern both during operation and after the end of the mission lifetime. These issues are addressed below in section 4.13.

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Athena Exploring Enceladus to Investigate its Habitability

4.6.1 Athena Parameters Lithium-Ion Batteries Battery mass, kg (10% 28.16 Athena will be available for communication with Earth margin) for 68% of each orbit which equates to ~134 minutes for 1 Battery dimensions, m 0.14 x 0.30 x 0.15 a 43 km orbit and ~139 minutes for a 50 km orbit. This (10% margin) percentage only accounts for the periods where the view of Earth will be obscured by Enceladus and does not Number of Cells 8 take into account of any potential eclipsing of Enceladus Cell module NCP55-4 by Jupiter. The almost constant uplink between the Earth and Athena is possible thanks to the NASA Deep Required Cell Capacity, Ah 65.93 Space Network (DSN). With base stations placed around the Earth with 120° of separation the DSN is able to provide constant uplink which helps simplify the Table 4-14. Baseline Li-Ion parameters communication procedures between the orbiter and Earth. The HGA featured on Athena is based on the Cassini HGA and serves as the main relay between Athena and the DSN. Some key parameters of the Cassini HGA and the DSN are summarised in Table 4-15.

Parameter Cassini HGA DSN Operation 7.4 7.2 frequency, GHz Dish diameter, 4 34* m Transmitted 50 20000 power, W Bandwidth, Hz 400 100000 Transfer rate, 484 143000 bits/second Figure 4-7. Shows the lander duty cycle Table 4-15. Comparison of uplink and downlink antennae for Orbiter communication. *This value is for the smallest of the DSN dishes which was used to calculate up and downlink 4.6 Telecommunication rates in the worst case environment

To communicate with Earth Sicily will use Athena as a relay. Athena’s telecommunication system includes a The total estimated data which will be generated will be high-gain antenna with transmitter and receiver, and ~15450 Mbits. At the transfer rate of 484 bits/second two low gain antennae sharing one transmitter and this will take ~10368 hours of communication to receiver. Sicily’s telecommunication system includes broadcast all the data back to Earth which translates into one low-gain antenna with transmitter and receiver. The ~15247 hours of orbiting time. The transfer solution has uplink and downlink between Earth and Athena are not been finalised but an ideal option would see some through x-band, and the uplink and downlink between data transferred during Athena’s operation period rather Sicily and Athena are through s-band. The s-band than one single large communication window at the end transmitter, receiver and LGA antenna will be identical of the mission’s lifetime assuming the power system can on both the orbiter and lander. A transmission signal accommodate a solution of this nature. between Earth and Enceladus takes ~ 66 mins, so a return signal at a minimum will take ~132 mins. This poses restrictions on the type of commands that can be sent so a high level of autonomy is likely to exist in both systems. One orbit around Enceladus takes ~ 197 mins

(43 km altitude) or ~204 mins (50 km altitude).

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Athena Exploring Enceladus to Investigate its Habitability

4.6.2 Sicily With these parameters and the maximum separation between both systems (78 km) the predicted uplink bitrate is ~ 500 kbits/s and the predicted downlink Communication with Sicily will occur once every orbit bitrate is ~ 270 kbits/s. when Sicily is ‘near’ the nadir of Athena. Throughout To determine the number of windows required the descent and operation life of Sicily Athena will orbit to transmit 757 Mbits of data we used our own model at an altitude of 43 km, which allows its orbital period with the following assumptions: Every 10 orbits Sicily to be exactly 1/10th of the rotation period of Enceladus. will be directly below Athena (nadir), antenna gain is 3 This means that every 10 orbits Athena will have Sicily in all degrees of elevation, and once Sicily enters the at its nadir, and the communication window will be at FOV of Athena’s antenna, uplink ~500 kbits/s and its longest. The window varies in length depending on downlink ~270 kbits/s is achieved. With an antenna the orbiter’s antenna field of view (FOV), the lander’s FOV in Table 4-6, it is possible to achieve a max antenna FOV and the distance between Sicily’s landing communication window of ~15 minutes. Depending on site and Athena’s nadir point on the surface. It is the landing site latitude the number of communication therefore desirable for Athena to use attitude control to windows per day (10 orbits @ 3.29 hours each) varies. keep the antenna’s peak gain always pointing at Sicily, The ideal position would be at 90 °S where Athena however this hasn’t been included in the calculations. would be in constant contact with Sicily and Because we are using batteries on the lander, communication of 757Mbits could be achieved in two power is a driver for the telecommunications payload. orbits. In contrast to this the worst latitude would be -45 Therefore, the lander shall communicate all of its °S (RR-REST region edge) and 757 Mbits could be science and systems data within 76 hours (see section transferred in 19 orbits. Even for the worst case scenario 4.2.10). To achieve a bit rate great enough to transmit the current power configuration (see section 4.6.2) will the predicted 757 Mbits of science data an omni- provide power for all the communication windows. directional antenna with the parameters in Table 4-16, a Sicily operation starts with a descent to the transmitter with the parameters in Table 4-17 and a surface, secondly its operation life and finally a receiver with the parameters in Table 4-18. The antenna, hibernation mode (see section 4.2.10). Throughout the transmitter and receiver are based on products provided lander’s ~ 100 min descent there will communication by (Technology, 2016) and are not designed and tested contact with Earth. During this phase communication for missions outside of Earth’s sphere of influence. will take the form of data that has been acquired from Therefore, the values are approximate and require the attitude control sensors, Pancam (PC) and other on- further analysis. The products chosen from SSTL board systems. Once the lander has settled on the (antenna, transmitter and receiver) are examples of the surface it will initiate the system checks phase and products and designs that meet the telecommunications transmit this data to Earth. This is the first opportunity requirements set for this mission. where controllers on Earth, if necessary, can input new

commands for the lander. Following this phase are the Antenna Parameter Assumed Value four science phases where data will be up-linked to Frequency, GHz 2.3 – 2.8 Athena during the communication windows. It is likely Gain, dB 3 that all the data won’t be transmitted to Athena during Beam width, ° 360 (Az) x 100 (El) the four science phases, so an additional 10 hours has Mass, kg 0.021 been included (communication phase). Following this Table 4-16. Some example parameters for an omni- phase is the hibernation phase where no communication directional antenna on Athena and Sicily (Technology, 2016). contact is predicted. However, after the hibernation Transmitter Parameter Assumed Value phase (if commanded) it is possible to initiate another operation life for Sicily. If this happens Athena will be RF Frequency Range, GHz 2.2 – 2.9 put into a 43 km orbit with the relevant inclination (few RF Output Power, mW 250 m/s transfer ΔV required) to allow communication with Mass, kg 0.6 the lander. Power Consumption, W 6 Table 4-17. Some example parameters for the S-band transmitter on Athena and Sicily (Technology, 2016).

Receiver Parameter Assumed Value RF Frequency Range, GHz 2.2 – 2.9 Mass, kg 1.3 Power Consumption, W 1.5 Table 4-18. Some example parameters for the S-band receiver on Athena and Sicily (Technology, 2016).

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Athena Exploring Enceladus to Investigate its Habitability

4.7 Command and Data Handling to be prevent vibrations. This subsystem can do all these functions by using a number of sensors and control devices. The roles of sensors are checking whether the The key parts of the command and data handling spacecraft is pointed correctly or not, turning and the system are: changing of speed. While the roles of control devices Space Flight Computer are changing the spacecraft direction rate and speed of Flight Software turning (Johnston et al. 2005) more explanation in Solid State Recorder (JPL, 2016) [SES-NM-00001]. On-board data and Telemetry tracking and command (TTC) are often associated with each other. Availability 4.8.2 Sicily is an important requirement of Telemetry tracking and command (TTC) (Maral and Bousquet, 2009). Sicily attitude control system includes redundant laser On the Lander, instruments will produce data after rangefinders, IMUs, and the imaging instruments. The measurements which will then be transported to storage, IMU will be used by the lander for navigation when it and then the low gain antenna of the Lander will separates from the orbiter Athena until Sicily reaches transmit to the orbiter. After that, the low gain antenna about 20 km from the surface, and prior to release the on the orbiter will receive the data which transmitted by IMU will be calibrated from Athena. The navigation high gain antenna to the ground station on the Earth. transitions, at that point make use of the laser The on-board data handling (OBDH) is the rangefinder and imager to complete the descent. The subsystem which stores and manages data between the four redundant attitude control (1N) thrusters’ strings different electronics units and the ground station, by mounted on the bottom decks and top provide for using the tracking, telemetry and command (TTC) degree-of freedom control (ENCELADUS Saturn’s subsystem. Each Sicily instrument has the maximum Active Ice Moon, 2007). raw data output. The Sicily Lander instrument suite will Initial measurement unit (IMUs) includes a gyro an generate approximately 0.09 GBytes of data, therefore accelerometer. The gyros, is the device that provide a the required storage must exceed this. proportional output to the angle which they had been There is a possibility to use OBC750 LEO with (about their input axes) rotated. The gyros of Sicily memory 512 MBytes (Sstl.co.uk, 2016) as the space lander will be used as the sensing elements in null flight computer. This is a processor with a rich heritage looking for servos, with each gyro output connected to and has been proven to perform in deep space. a servo motor driving the appropriate gimbal, thus they However, it is important to consider the radiation in inertial space will keep the gimbal in a constant environment of Enceladus for the storage solution. In orientation (King.1998). addition, the other requirement that has to be considered The Laser rangefinders: is an instrument that generates is data compression that is recommended by the data which easy to obtain range (figure or ground) and Consultative Committee for Space Data Systems its data has some pre-processing will be used for (CCSDS) (CCSDS 121.0-B-2). removing mixed range pixels of the laser beam when a Further definition study will be required to measurement of the time is being made, the inherent examine the full automation structure necessary to ambiguity interval issues of data (Brady, 1989). achieve the science goals and ensure successful Imaging instrument has a miniaturised CCD camera. operation. This is especially important for this mission The primary purpose is function as an imaging device as a 66-minute one-way signal travel time to when Sicily will be descent to obtain high resolution communicate with the orbiter makes manual control of images of the landing site. It based on Rosetta Lander the orbiter and lander systems difficult. Imaging System (ROLIS) which take pictures of the Enceladus surface below the lander. It will emit a series 4.8 Guidance, Navigation and Control of light diodes by using several wavelengths 4.8.1 Athena (Open.ac.uk, 2016).

GNCS can control the spacecraft’s position in space and direct it into its target place with maintaining knowledge of the places of celestial bodies such as the sun and Earth. This information is very important in order to get the correct way to Enceladus or keeping communication with Earth through pointing it toward the Earth by doing right manoeuvres. The system can maintain information about location of Athena while it is in its orbit and accurately point the cameras. Therefore, it is this systems role to take accurate image for the desire location. The guidance, navigation, and control subsystem and structures subsystem should be designed 33

Athena Exploring Enceladus to Investigate its Habitability

4.9 Mechanical Design aluminium. This shell will provide radiation shielding to the internal components of Athena. All three shells Regarding the mechanical design of Athena and Sicily are designed to provide a platform for the other sub- a top level consideration has been made. Both system systems and provide resistance to shocks, vibration, include a 3d computer aided design (CAD) indicating axial g-loads and lateral g-loads during the launch the volume and placements of the system’s sub-systems phase. These shocks, vibrations and loads have not been and elements. The majority of dimensions include considered yet and further analysis is required. structure, thermal and shielding considerations. The Composite CFRP was chosen for the structure Athena orbiter dimensions are approximately 2.3 x 1.5 because of its low density (1570 kg/m3), high stiffness x 1.4 m and the Sicily lander dimensions are (160 GPa) and very high specific strength (2520 MPa approximately 1.0 x 1.40 x 0.80 m. @ 0°) (Williams, 2016). The design reduces the number of possible bending loads by keeping the RTG, 4.9.1 Athena Structure pressurant tank and Sicily lander attached as close to the central CFRP shell structure. The main loads are likely The structure for Athena is designed (8) around its two to be tension and compression of the central CFRP shell 1m diameter bi-propellant tanks. These are located caused by the ~660 kg Sicily lander, both propellant vertically in the central column of the orbiter to reduce tanks and the 100 kg 4 m diameter high gain antenna. any variation of the spacecraft’s centre of mass, thus The aluminium shell will provide radiation protection to reducing attitude control demands. Below the two the internal components but the thickness is propellant tanks is the 400 N apogee thruster and the 4 approximate and needs verification. The thickness of x 4N attitude control thrusters. These 4N thrusters are the composite CFRP shell and propulsion tubes are inserted inside an accommodating, but currently approximate. Also shown in Figure 4-3 is the unspecified material. Surrounding the propellant tanks orientation of the combined system during observation is a 2 cm thick cylinder made of composite CFRP. To of Enceladus. The IPR is stowed until its necessary add support and provide a platform for the aluminium deployment in the radar imaging phases, at which point shell another structure surrounds the propulsion tube. the lander is not attached to the orbiter. All instruments This is also made of composite CFRP and is 1 cm in are placed so there are no obstructions to their FOV. thickness. The final layer is a 5mm shell made of

Figure 4-8. Shows three separate perspectives for the combined orbiter and lander system. On the left is the structure sub-system and consists of the propellant tanks, CFRP propulsion tube, CFRP structure shell and aluminium shell. Any holes, tubing or extra detail is not included for clarity. At the top shows the placement of all the sub-systems and the Sicily lander. The bottom-right perspective shows the orientation of Athena during the initial imaging stages. The arrow indicates the direction of all the instruments FOV. 34

Athena Exploring Enceladus to Investigate its Habitability

4.9.2 Sicily Structure 4.10 Mass Budget

The lander structure is based on a hexagon design which The overall mass composition of the mission is shown consists of a shell that encloses two batteries, one in Table 4-19. Summary of key launch mass parameters monopropellant hydrazine tank would be installed at the with orbiter and lander specific breakdowns featured in centre of the spacecraft which made of composite CFRP Table 4-20 and Table 4-21 respectively. The margins and instruments box. It also contains three legs which applied are covered more clearly in the specific Athena are specially designed to move in different direction to and Sicily sub-sections but the proposed configuration help the lander during the descent and they can provide for this mission is compatible with the selected launcher stability on an icy surface, two RTGs in both sides of with 357 kg of spare mass. the lander, four 1N and six 20N thrusters and the Enceladus Sample Collector (ESC) which is connected Parameter Mass, kg to the shell from outside. The dimension of the lander Orbiter dry mass 1293 would be 1.0 × 1.4 × 0.8 m and the total mass is Lander dry mass 484 approximately 659 kg. To take samples from plumes in Propellant mass 2111 the SPT, the spacecraft should be protected as it and its Mission wet mass 4323 instruments could be at a high risk from the plume Max launch mass 4680 particles. This problem can be solved by shielding Launch margin 357 lander structures with a high tolerance shell which could Table 4-19. Summary of key launch mass parameters be made of resistance materials such as composite CFRP and aluminium alloy as shown in Figure 4-9. The For the payload subsystems, mass margins have been lander will separate from the orbiter and there legs could applied in accordance with the respective instruments be orientated in opposite side Enceladus surafce. TRL status (a 20% margin for TRL 4 or 5 instrument and a 40% margin for a TRL 2 or 3 instrument). A 20 % system margin has then been applied to obtain these total budget values. The values for the attitude and orbital control, telemetry, tracking and command and wiring subsystems have been estimated from other similar sized missions and then had margins applied to them.

Parameter Mass, kg Total Dry 1292.88

Propulsion (dry) 106.50 Mechanical 608.37 Figure 4-9. Shows the main components of lander Payload 209.84 The three legs have the ability to move in opposite Attitude and orbital control 180.00 directions as explain in Figure 4-9 and carry the lander Telemetry, tracking and command 7.20 structures and position it in the right place with more Power 47.30 stability Thermal 8.94 Wiring 7.20 System margin 117.53 Table 4-20. Breakdown of orbiter mass budget

Parameter Mass, kg Total Dry 484.36 Propulsion (dry) 13.16 Mechanical 273.10 Payload 57.76 Attitude and orbital control 6.00 Telemetry, tracking and command 6.00 Power 43.16 . Thermal 35.16 Figure 4-10. Shows the three legs of lander in opposite Wiring 6.00 directions. System margin 44.03 Table 4-21. Breakdown of lander mass budget

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Athena Exploring Enceladus to Investigate its Habitability

4.11 Planetary Protection and Contamination measures for this mission, it is recommended that Control following a binary hierarchical judgment to make a 4.11.1 Planetary Protection framework which is used a familiar Coleman–Sagan equation due to the unknowable and/or uncertain factors that is included in the latter (Konstantinidis et al., 2015). Planetary protection is a practice term of NASA’s term Due to Enceladus being considered as a potentially to protect solar system bodies such as (planets, habitable body, therefore it only remains to determine asteroids, comets, and moons) from Earth life if a the contamination that is posed by each mission module mission is one way and protect Earth from solar system (lander, orbiter), and the necessary measures to mitigate body’s life if mission is a sample return (Rummel and that these types contamination risks of within the Billings, 2004). Forward and backward are two types of passable possibility of 10-4 (Konstantinidis et al., 2015). interplanetary contamination (Beauchamp and Belz, The lander will come with direct contact with the 2012). Enceladus mission has to consider forward type surface of the icy moon, therefore applying Viking-level contamination because this mission is not returned to PP stringent is required (Konstantinidis et al., 2015). In Earth. The Committee on Space Research (COSPAR) addition, the sensitive Enceladus regions for the orbiter classified five categories groups of planetary protection are during non-nominal contingencies, and the orbiter missions (Beauchamp and Belz, 2012). The will have direct contact with North Pole of Enceladus. recommendations of depend on the celestial body The orbiter contamination could be reduced explored and the type of space mission. This mission is sufficiently, by clean room assembly which is required, classified as category four for orbiter (Athena) and for but further measures is not required for bio load lander (Sicily), which has specific requirements as reduction (Konstantinidis et al., 2015) as below: shown in Table 4-22 below (Beauchamp and Belz, 2012). Techniques for Sterilizing and Cleaning (Ball, Many reasons lead to consider the South-Polar 2007) Terrain Enceladus as a restricted area from a perspective Filtration and intrinsically clean assembly of planetary protection (PP). For example, remaining Thermal stress cracks of the surface open to a few kilometres depth and Radiation exposure active fissures in the ice that is venting could lead Sterilizing chemicals directly to the liquid subsurface water (Konstantinidis et al., 2015). Determining the planetary protection

Table 4-22. List of mission types/ target body of planetary protection that categorized by COSPAR (Nicholson, Schuerger and Race, 2009).

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Athena Exploring Enceladus to Investigate its Habitability

4.11.2 Contamination Control Space hardware sterilization for both Sicily and Athena will be a combination of different technologies or Unlike planetary protection, organic contamination methods, Table 4-23. This list does not represent the control requirements flow down from the science only technologies that can be used for the mission, but requirements (JPL/NASA, 2011). Typically, the instead a small selection indicating the key methods that assessment is based on the impact that contamination are available to contamination control. will make to the science investigations. For this report a top-level assessment of the contamination control has Technology Used on been made. Sicily will receive very stringent Dry Heat (e.g. All but some components such as requirements because (1) it will be drilling into the sub- 125°C for 5 some polymers, batteries and surface of Enceladus with its ESD and (2) it will be hours) magnets. It is a proven technology. collecting matter from the plume with its ESC. It is Solvent Used on all components except critical to minimize the forward contamination because Cleaning some surfaces. It is easy and cheap, it shall ruin the science data obtained by the high (routine) but a manual process. resolution chemical and molecular instruments in the Active Gamma Has not been brought to maturity but scientific payload. The samples are extracted from and electron could be used on radiation hard Enceladus using the ESD and collected using the ESC. beam components. Every element in contact with these systems shall Sterilization receive the same contamination control requirements. Passive Sterilization through the expected This will include the complete science payload, Sterilization radiation types at Saturn. It is an structure sub-system, some thermal tubing and some accepted practice for and electrical wires. It will possibly include the other likely possible for Enceladus. Likely internal sub-systems; the two batteries, electronics box, damage by high energy protons and remaining thermal sub-system, radiation shielding, and electrons so careful designing and the propulsion sub-system. considerations are required. Not Sources of contamination will exist on Earth, suitable for radiation sensitive with the rings of Saturn and the plume and jets at components. Enceladus. Reduction of contamination will be near Table 4-23. Some possible active and passive methods for impossible after the combined system has launched contamination control (Y.Bar-Choen, 2009) (JPL/NASA, from Earth due to the limited in-situ contamination 2011). procedures, but it is possible to minimise contamination perhaps through the use of mechanisms that are removed when necessary (louvres). Such mechanisms are foreseen for the camera lens and other surfaces that require a surface cleared of particulates. Reduction of contamination on Earth is common practice with a significant amount of planetary protection technologies, such as microbial reduction and validation, have matured significantly over the last 10 years, lacking only formal space agency approval. If these technologies are unsuitable there are further tests and technologies that are unapproved which could be requested for this mission (JPL/NASA, 2011). Effective organic contamination control calls for materials compatibility studies early in their design, to reduce and minimise outgassing. Sensitivity to contamination can be reduced through extensive periods of cleaning and minimising sample and equipment exposure times post- cleaning. We envision Athena and especially Sicily to be subjected to stringent protocols similar to that on the Viking lander, for example the bio load reduction protocols, cleanroom use, bio shield, organics inventory and a microbial reduction plan, (Y.Bar-Choen, 2009). What is promising is the technology available for missions nowadays has improved since the Viking mission with materials and processes used to fabricate space probes has advanced considerably.

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Athena Exploring Enceladus to Investigate its Habitability

4.12 Mitigation of Technical Risk Standard of the European Cooperation for Space Standardization (ECSS). Furthermore, another Risks are dangers that threat any project since they element to mitigate risks is communication between negatively effect on the project cost, technical team members, which is done for this mission. There are performance, and schedule. However, applicably many other ways to identify causes and provide controlling risks practices can give new positive impact probability data risks that are FMECA, Fault-tree, Event opportunities. In this work packages several risks have tree (Williams, 2015). There are four steps for risk been identified and mitigated during different stages of management process, Figure 4-11 and the live and the project. In addition, we have retired some of risks retired risk are explained in Figure 4-12 and Figure 4-13 and others risks are still live. All risks have medium or which will be discussed in detail in [SES-NM-00001]. low-level (yellow or green) risks but the only high (red) risk remaining is development time of Am RTG. Athena mission risk management has many aims, which are identify, assess, reduce, accept, and control. In addition, these aims of space project risks should be a systematic, comprehensive, proactive, and cost effective manner considering the technical and programmatic project constraints. Risk is taken in accounted tradable against the traditional identified project resources within technical such as mass, power, dependability and safety domains the programmatic, management such as cost and schedule. Therefore, in a project the overall risk management is a repeated process during the project life cycle. The project progress start to determine iterations through the Figure 4-11. The steps of the risk management process different project phases, and changes in order to a given (ECSS-M-ST-80C, RD01). project baseline that influence project resources (ECSS-

M-ST-80C, RD01). During this Athena project, many types of risks continually were identified, assessed, and managed. Action to reduce all types of risks from high level to medium or low level were taken by using and applying

Figure 4-12. Yellow and green retired risks of Athena mission with their mitigations 38

Athena Exploring Enceladus to Investigate its Habitability

Figure 4-13. Red, yellow and green retired risks of Athena mission with their mitigations 4.13 Conclusions and Recommendations

The trajectory and manoeuvre analysis was completed barriers to completing this mission and achieving the to minimise the required ∆V to reduce fuel mass and science goals. In the one potential planetary protection allow the maximum scientific payload to be delivered concern, an alternative solution has been proposed to with a minimal cost. The source of this trajectory shows overcome it. it to be feasible and the minor adaptions to the final orbit Ultimately, verifying the suitability for the design highlight that it is suitable to achieve all the science proposed here is pertinent to the missions’ success. goals. Highlighting where developments are needed, back-up Radiation and thermal considerations require solutions offered and more in-depth review completed. some further study to ensure the survival of the payload under a more robust thermal and radiation model. However, the initially proposed shielding and heating solutions demonstrate feasible design solutions based on readily available materials. Giving more definition to some of the thermal and radiation solutions described above is key to developing this mission further. The power system calculations display a potential power solution with the development concern being the primary risk within the mission. The failure for the AM-RTG to reach development deadlines is a potential show stopper and should this project be continued a potential alternative is vital to ensure mission success. A viable communications system has been demonstrated but the length of time for data transmission back to Earth is a concern as the proposed configuration would take longer to transmit the data than the scientific operation period. This is an issue which can be addressed by altering the configuration of the communications systems, however this could have a knock-on effect to many other systems (e.g. if the power requirement of the antenna is increased). Command and data handling and guidance, navigation and control systems have yet to be examined in any detail. As such, further research into the potential solutions for automation and control are essential in analysing the feasibility of this mission. The mechanical outlines show that the sizing of the lander and orbiter are acceptable from a launch perspective and also serve to give a rough idea of the configuration of these two modules. Material selection needs to be refined but these preliminary designs are suitable as proof of concept.

The planetary protection and contamination control analysis demonstrates that there are no immovable 39

Athena Exploring Enceladus to Investigate its Habitability

5 Management 5.3 Conclusions and Recommendations 5.1 Schedule A technical development plan needs to be established in For all of the instruments within the model payloads, a order to ensure the schedule requirements and varying level of technical development needs to be development targets are met. This is achieved by completed before the mission launches in 2026. This is assigning dates to each development and each goal. This also true for the Americium based RTG for use on the plan is also to provide definitive cut-off points for sub- lander if it is to be a viable power source. Instrument system development which will shape the final mission. finalisation would be completed by 2018 giving 5 years The plan must cover all the sub-systems as there is for the completion of any technological developments development needed in every area of the mission but as required. focus needs to be directed to the low TRL instruments As the launch opportunity approaches, the and the lander power system. instruments and other developing subsystems would be The outreach plan also needs to be fully realised set to stringent TRL goals or risk ejection from the in order to provide targets for public outreach and payload. This would be conducted in such a way that highlighting dates where such releases will be made. instruments would be ready to fly by 2023 which gives Finally, it is important to examine the possibility 3 years for assembly and implementation. The mission of an alternative launch date. With some of the lower only has one planned for one suitable launch window TRL instruments and the lander power system, the and missing this window would result in a total overhaul original launch window in 2026 may not be achievable of any orbital configurations as well as substantial and making arrangements for missing this launch propulsion changes. window are critical in mission design. Up until 2018, there is plenty of room for changes to be made to the model payloads and the mission can be altered in scope where necessary depending on TRL advances. The more immovable development requirement is that of the Am-RTG which would result in a total redesign of the mission should it fail to reach an acceptable TRL by the deadline.

5.2 Public Outreach and Science Communication

Landing for the first time on any celestial body is likely to generate interest from the public. With this in mind, Athena will run alongside a robust outreach program to establish interest within the scientific community and the wider public. ESA’s channels would be used for communication in order to generate maximum exposure and ESA would be responsible for any inter-agency communications. These communications will include any key milestones as the mission is planned and developed, awareness and broadcast of launch, milestone updates during the transit period (making use of Athena’s cameras where conditions allow) and finally the publication of the top level data outputs.

As the mission develops, the specifics of these outreach goals will alter and keeping communication with leading agencies, alongside ESA, and the public is key. Such communication events will take the form of mission outlines on purpose built web pages, media releases (such as images and video of development progress), traditional press releases to highlight milestones and social media presence which has proved successful for generating interest in recent missions.

The main difficulty will be ensuring public engagement does not wane too greatly over the long cruise phase; careful planning will help combat this issue.

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