Poseidon-Trident: Mission to the Neptunian System

Blue Team 2nd August 2012 -Alpbach Summer School-

Abstract 2 Scientific Background

Poseidon-Trident is a light weight mission to explore the The composition ’s atmosphere is mainly com- Neptunian system on a flyby trajectory. It shall pro- posed of H2 and He (82 % and 15 % respectively) as well vide knowledge on Neptune and after a 13.4 year as CH4 (3 %) (de Pater and Lissauer, 2010). The tem- cruise. A probe is released into Neptune’s atmosphere perature and pressure profile has been modelled down to gather compositional data up to 80 bars, deep enough to 10 bars, with radio occultation data from . to observe below all expected cloud layers. Imaging in- The tropopause has been detected near 100 mbar with struments from UV to IR will provide context and gen- a temperature of 52 ±2 K, and the stratosphere has a eral information. The closest approach on both bodies temperature of 130±12 K at 0.3 mbar (Lindal, 1991). will help improve our knowledge of the gravity fields. The probe measured the wind velocity in A general observation package will measure the parti- atmosphere and unexpectedly showed that the cle/plasma/field environment during the majority of the wind speed increases at lower altitudes. This has not mission. been done in Neptune’s atmosphere. The rotational As a giant planet, Neptune is crucial to our under- period of Neptune has been measured by radio waves. standing of formation. It is also believed Also, zonal wind speeds have been measured on Nep- to be representative of icy giants and can therefore shed tune by tracking cloud features relative to Neptune’s light on the way we understand exoplanets. rotation period. In contrast to the other giant planets A comparison of Triton and Pluto will provide there is a large dispersion in the wind speeds, which is about Objects (KBO) and shed light on the not well understood (Martin et al. 2012). origin of Triton. On Neptune, wind speeds reach approximately 350m/s. At higher latitudes (>50◦), the winds are pro- grade, and reach velocities of approximately 200m/s 1 Introduction (Sromovsky et al. 2001). Additionally, Voyager 2 de- tected the (GDS), a smaller dark spot The Neptunian system was visited for the first time by (DS2), and a small bright cloud (Scooter). When Nep- the Voyager 2 spacecraft on August 25, 1989. Neptune tune was imaged several years later, all three features is one of the giant icy planets of our Solar System, lo- had vanished. cated at a distance of 30 AU from the Sun. So far, 13 Models of interior structure of Neptune suggest the satellites have been observed in the Neptunian system, existence of a rocky core subtending over 20 % of the including Triton, its largest moon. A ring structure pos- planetary radius. The mantle comprises roughly 80% of sibly maintained by small moons, has also been observed the planet’s mass and consists of fluid rather than solid (Salo and H¨anninen, 1998). ices. Although the pressure and temperature in the in- Triton has been partly imaged by Voyager 2, leaving terior are rather high, we do not expect the pressure half of the northern hemisphere unexplored. Its retro- to be high enough to have metallic hydrogen. Therefore grade orbit and relative lack of other large another source for the magnetic field is required. It is as- moons in the Neptunian system suggests Triton was sumed that between 0.3 - 0.7 R (Neptunian radii), an previously a Kuiper belt object (Agnor and Hamilton, N ionic ocean exists, which is most likely the source of the 2006) and was captured by the Neptunian system. Tri- magnetic field. Comparing the equilibrium and effective ton’s average density is relatively large (2 g cm−3), sug- temperatures - 46 K and 59.3 K respectively - shows gesting a combination of ice and rock. Triton also has a that Neptune must have a large internal heat source, tenuous atmosphere consisting mainly of nitrogen, with presumably partially caused by primordial heat from a trace of methane (de Pater & Lissauer, 2010). Ad- the planets accretion (de Pater and Lissauer, 2010). ditionally, geyser-like eruptions have been observed by The Voyager 2 flyby of Neptune revealed a very com- Voyager 2 (Soderblom et al., 1990), implying that Triton plicated magnetospheric configuration. Within half a is geologically active. planetary rotation (8˜ h) a complete reconfiguration of A flyby to Neptune will comply with ESA’s Cosmic the magnetosphere can take place (Bagenal, 1992) due Vision long-term goals of exploration: from the Sun to to the large dipole tilt. This constant reconfiguration the edge of the Solar System, and the giant planets and of the magnetosphere will cause large loss processes of their environment. highly energetic particle which are usually trapped in a

1 planetary magnetosphere. of the unique magnetic field configuration. Radio emis- All magnetised planets of our solar system are known sions assumed to originate at the auroral regions will to be sources of radio emission, generated by high en- help to explain radio spectra from exoplanets. ergetic electrons. The Neptunian magnetosphere is also the source of radio emission, the so-called Neptunian 3.4 Secondary Objectives kilometric radiation (NKR). Voyager 2 measured radio bursts in the frequency range of 100 kHz to 1300 kHz. As a secondary goal, we will investigate the dust parti- From these measurements the Neptunian rotation pe- cles in the satellites and ring system of Neptune to un- riod has been obtained providing the first evidence that derstand formation and evolution of the Neptunian sys- Neptune has a magnetic field (Zarka, 1998 and 1992). tem. Also, before the main mission a possible synergy with JUICE at will give measurements of the magnetosphere and plasma morphology at two widely 3 Mission Objectives seperated points in space at the same time for good spatial resolution. After the main mission the space- The main goal of our mission is to explore the Neptu- craft will investigate Kuiper belt objects. During the nian system as an archetype system for the icy giant whole mission, targets of oppurtunity such as centaurs, planets and to investigate the nature of Triton. comets and interplanetary dust will be investigated if possible. Primary objectives: • Characterise the atmosphere and the interior of 4 Scientific Requirements Neptune In order to fulfil the missions objectives, certain scien- • Characterise the atmosphere, the interior and sur- tific requirements must be met. These will be described. face features of Triton • Improve our current understanding of the magne- 4.1 To characterise the atmospheres tosphere in the Neptunian system 1. For Neptune we shall measure the elemental abun- Secondary objectives: dances of C, O, N, S and the noble gases, down to a depth of at least 50 bar over a mass range 100 • Investigate the dust particles in the ring system and amu and with a mass resolution of m/dm 3000 at the less massive satellites of Neptune a peak height of 1 %, with an absolute accuracy of Targets of opportunity: 10 %.

• Investigate the Kuiper Belt objects 2. For Neptune we shall measure the absorption lines 13 15 of isotopes (DH, CH3D, C2H6, NH3), hydrocar- • Investigate targets of oppurtunity both before and bons (C2H2,C2H6, CH3, CH4,C2H4,C3H4,C6H6, after the flyby at Neptune (e.g Jupiter, , aster- C4H2,C3H8,C4H10) and PH3, GeH4, NH3,H2, oids, centaurs) CO, and HCN with a spectral resolution of 0.5–20 cm−1. 3.1 Neptune 3. For Neptune we shall measure the atmospheric By investigating the composition, chemistry, structure pressure with an accuracy of 1 % and a resolution of and dynamics of Neptune’s atmosphere and also map- 0.01 mbar, and the atmospheric temperature with ping interior structure, we will be able to constrain plan- an accuracy of 0.5 K and a resolution of 0.02 K etary formation and evolution models. down to preferably 100 bar but at least 50 bar. 4. For Neptune we shall measure wind velocities in the 3.2 Triton troposphere and the stratosphere with an accuracy By observing the atmosphere and surface features of of 1 m/s and a resolution of 0.1 m/s down to at least Triton, the nature of its nitrogen plumes and seemingly 10 bar, and we shall measure zonal wind speeds young surface will be more understood. The interior of with an accuracy of 0.5 m/s and a resolution of Triton will be mapped to see if Triton is differentiated, 0.02 m/s. and also to discover a possible subsurface ocean. 5. For Triton we shall investigate the atmospheric composition by observing a solar occultation in UV 3.3 Magnetosphere with 0.01 ◦/s stability within 1σ accuracy, and a stellar occultation with 0.1 ◦/s stability within 2σ Understanding the magnetospheric interactions is a key accuracy. to understand how Neptune and Triton interact to- gether in a certainly unique magnetic field configura- tion of our Solar system. The plasma morphology of 4.2 To characterise the interiors the magnetosphere will be measured to understand Tri- 1. For Neptune we shall measure the gravity field up ton’s role as a possible plasma source, and the dynamics to spherical harmonics (SH) degree 4 to estimate

2 the interior structure, using range - rate accuracy 5.2 Wide Angle Camera (WAC) in the range of 0.015mm/s and 0.1mm/s at 60s in- The Wide Angle Camera (WAC) will provide multi- tegration times. spectral imaging (in spectral range 350 − 1050 nm) of 2. For Triton we shall measure the gravity field up to Neptune, its moons and rings, to address the goals in SH degree 2 to estimate the moment of inertia, us- geology, geophysics, topography and meteorology. By ing range - rate accuracy in the range of 0.015mm/s this it will be possible to study cloud morphology, par- and 0.1mm/s at 60s integration times. ticle properties and dynamics. Similar to NAC, WAC requires that it maintains nadir-pointing. In addition, 3. For Neptune we shall measure ELF and TLF imaging from close distances requires that the scan line Schuhmann resonances as well as DC field strengths is perpendicular to the flight direction. generated by lightning to determine water content and depth of the lower layer of the resonant cavity. 5.3 Ultraviolet Imaging Spectrometer 4. For Triton we shall image the surface features to (UVIS) search for past and present activity at a global spa- The Ultraviolet Imaging Spectrometer (UVIS) will give tial resolution of 100 m, and for local interests, a occultation measurements of Triton’s and Neptune’s at- 1–5 m/px resolution shall be used. mospheres which will give high resolution information on the stratospheric temperatures and the atmospheric 4.3 To characterise the magnetosphere composition (in spectral range 110−320 nm). Investiga- 1. We shall measure the 3-dimensional magnetic field tion of the interaction between Triton’s and Neptune’s vectors with a fluxgate magnetometer with a range magnetospheres will give us a better understanding of of ±16384 nT per magnetic component, with a res- their ionospheres and exospheres. UVIS requires that ◦ olution of 0.2 nT. the Sun should be at least 30 away from the field of view of the instrument, and that the maximum angular 2. We shall measure the energy of particles using a speed of the spacecraft during operation is 0.1 deg/s. plasma and an energetic particle analyzer that mea- sures the electrons and ions (1 eV – 5 MeV) and 5.4 Magnetometer (MAG) also their electric charge using an ion neutral mass spectrometer in a mass range of 1–100 amu. The Magnetometer (MAG) will characterize Neptune’s and Triton’s three-axis intrinsic and induced magnetic 4.4 For the secondary objectives field with implications for the deep interior, e.g. a sub- surface ocean. In addition, it will investigate plasma 1. We shall measure the radio emissions in frequency sources, the magnetospheric response to solar wind vari- ranges from 1 Hz to 1300 kHz for the electric fields ability and planetary rotation effects. MAG will operate and from 1Hz to 100 kHz for the magnetic fields. at all times, and it will be calibrated on the ground prior to launch. In-flight calibration will determine the space- 2. We shall measure dust particles and their electrical craft induced magnetic field, verify the extent to which charge with a velocity in the range 4,000–300,000 the ground calibration remains valid and also quantify km/h, a size of 0.05–100 m, and a mass resolution changes in calibration parameters. m/dm = 200–500 amu. 5.5 Radio and Plasma Wave Experi- 5 Instruments ment

The instrumentation that we shall need to fulfil our sci- The Radio and Plasma Wave Experiment (RPW) will ence requirements will be described below [4]. consist of three 5 meter long antenna for measuring the electric field and three magnetic coils for magnetic field 5.1 Narrow Angle Camera (NAC) measurements. The antenna will be directly placed on the spacecraft and the coils on the magnetometer boom. The Narrow Angle Camera (NAC) will provide high res- It will measure on a frequency range of a fraction of a olution images (in spectral range 350−1050 nm) with a Hz up to 20−40 MHz. With this device we will be able resolution of less than 10m/px at 500km and less than to study local plasma wave phenomena, identify charac- 5 km/px at 1 Mkm. It will be used to image the sur- teristic frequencies of the plasma, detect lightning and face and by this determine zonal wind speeds. In ad- measure NKR (Neptunian kilometric radiation) spec- dition, monitoring of lightning flashes on the night side tra. The measurements shall be performed during all of Neptune, Triton’s geysers and Neptune’s rings can science phases of the mission. The RPW is heritage of be studied by the camera. NAC requires that it main- the Cassini spacecraft (Cassini Heritage). tains nadir-pointing. The pointing prediction shall be sufficiently accurate to point at selected targets and to 5.6 Particle Package guarantee sufficient image overlap to cover as much of the surface as possible. The Particle Package (PP) will observe the plasma dy- namics in the Neptunian system. It will investigate

3 whether Triton has an induced magnetic field, as well 5.12 Reflectron time-of-flight mass as observe Triton’s interaction with Neptunes magneto- spectrometer sphere and the plasma therein. In addition, the struc- ture and composition of the exosphere of Triton and the The reflectron time-of-flight mass spectrometer (RTOF) ionosphere of Neptune can be investigated. PP requires shall measure the composition of the atmosphere and continuous operations. shall measure isotopic ratios e.g. D/H. It has a mass resolution 3000 at full width half max and a mass range of 1 to 500 amu. It’s located in the probe. 5.7 Radio Science Instrument The Radio Science Instrument (RSI) consists of a Radio 5.13 ELFR Science Transponder (RST) at Ka-band and an Ultra- stable Oscillator (USO). It will determine the gravity 5.14 Other Instruments field of Neptune and Triton, and characterize the inter- • Very low frequency radio. It measures AC fields, nal structure, inclusive of a possible subsurface ocean at detects natural wave phenomena and lightning. Triton. The best performance of the RST is obtained It’s located in the probe. The sensitivity is 10 when simultaneous transmission and reception both at mV/m/Hz, the dynamic range 80 dB and the fre- X-band and Ka-band are carried out. The spacecraft quency range 0 − 500 Hz. antenna shall be constantly pointed toward the . • Electric field sensor It detects atmospheric effects 5.8 Composite Infrared Spectrometer and measures the Schumann resonances. ([?]). The sensitivity is 1 − 30 mV/m and the dynamic range The Infrared Spectrometer (IRS) (Cassin heritage) con- is 50 dB. sists of two spectrometers (a far-infrared and a mid- infrared). It can measure infrared emission from Nep- • Temperature sensor is located in the probe. tune and Triton’s atmospheres, rings, and surfaces (over • Pressure sensor is located in the probe. wavelengths from 7 to 1000 µm) to determine their com- position and temperatures. By this it can give infor- • Accelerometer is located in the probe. mation on energetic processes, hazes and clouds within Neptune’s atmosphere. 6 Mission design 5.9 Doppler Wind Experiment In order to bring the instruments to the Neptunian sys- The Doppler Wind Experiment (DWE) consists of two tem, several top-level architectures of the mission design Ultra-Stable Oscillators (USO), one on the probe and have been considered, analysed and traded-off. one on the orbiter. It is a high-precision tracking in- vestigation to determine wind velocities in Neptune’s 6.1 Transfer trajectory, Analysis and atmosphere. In addition, it can be used to measure Trade-off Doppler fluctuations to determine turbulence. The overall mission design is driven to a large degree by 5.10 Dust Analyser the following factors: The Dust Analyser (DA) is a reflectron-type, time-of- • Neptunian system mission architecture (flyby flight impact mass spectrometer, which has heritage vs. orbiter). from the Cassini CDA and the CIDA instru- • Transfer trajectory time-of-flight. ments. It could be used to detect the dust particles pa- rameters in the satellites and ring system of Neptune. • Launcher selection. Furthermore, it could be used to investigate the inter- planetary and Jupiter system dust particles. The launcher selection is limited to European oper- ated launchers, i.e. Vega, Soyuz or Ariane 5. However, 5.11 Visible InfraREd Hyperspectral the Vega launcher has been discarded as an option due Imaging Spectrometer to its limited size. Three transfer options were identified as the most The Visible InfraREd Hyperspectral Imaging Spectrom- competitive candidates. Two with an EVEEJ flyby se- eter (VIRHIS) consists of a three mirror anastigmatic quence with 22 and 13 years time-of-flight, generated telescope joined to a entrance slit of an Offner spectrom- using the PaGMO trajectory optimisation tool, and one eter. The spectral range is 0.4−5.2 µm and has a FOV of with an EEJ flyby sequence with 11 years time-of-fligt 3.4◦ and the IFOV is 0.125 − 0.25 mrad. It shall deliver (from Marley and Dudyinski 2010). For each of these information about composition and temperature struc- trajectories, both the mass that can be inserted into a ture of Neptune’s atmosphere and Triton’s exosphere. Neptune orbit and sent on a Neptune flyby were inves- As well it shall characterise surfaces properties of Tri- tigated for both an Ariane 5 and a Soyuz launcher. ton. For the orbit insertion, a low periapsis distance of 10,000 km was assumed, with an apoapsis distance of

4 Flyby Orbiter Source Delta V [m/s] NAC DSM (including 10 % margin) 523 WAC Navigation correction (for 5 125 MAG (x2) + boom legs) CIRS Altitude control (including 100 UVIS 100 % margin) PP Total 748 RPW SUDA Table 4: Delta V budget for selected Neptune transfer ELFR trajectory (excluding altitude control budget). 132 kg 80 kg 140 kg 80 kg %, the engineering matters (including cost, competitive- ness and mass). The result of the trade-off can be seen Table 3: Selected payload on Neptune orbiter/flyby to be the flyby spacecraft, launched by an Ariane 5, with mission. Red: Absent. Orange: Downsized. Green: a transfer time of 13.4 years. The main arguments for Present. selecting the flyby option are the following:

70 NR. This was chosen to minimise the required DV, while retaining a safe margin. The results of this anal- ysis, assuming a 10 % margin on the total DV require- ment, 25 m/s per transfer trajectory leg for navigation manoeuvres and an Isp of 300 s, are shown in Table 1(see last page).

The propellant system dry mass is Mp,dry taken as 20 % of the propellant mass. The probe mass Mprobe is assumed to be 300 kg. The payload system mass on the spacecraft is calculated as 10 % of the remaining spacecraft mass. From this the total instrument mass in Neptune orbit/flyby, with an additional 20 % margin on the payload mass, are given in Table2(see last page).

Figure 1: Tradeoff table between flyby and an orbiter. The negative numbers are discarded immediately, as they represent solutions with unfeasibly high propellant system masses. Of the remaining solutions, the 22 year • The main science goals can be successfully Ariane 5 trajectory to orbit and the 13 year Ariane 5 tra- achieved. jectory in flyby were investigated further. An important advantage of the flyby option is the fast transfer time, • A relatively more cost-efficient and faster mission combined with its ability of being capable of bringing can be achieved with an orbiter. more mass to (but not in orbit of) Neptune. The esti- • Flyby mitigates orbiter data rate bottleneck. mated available payload masses were compared with the desired instrument package described in Section 5. For • The chance to have a full, state-of-the-art instru- the flyby option, the full payload package can be flown. ment package on board, as opposed to orbiter. For the orbiter option, a selection of instruments to not The main downside of the flyby mission is that only fly on the spacecraft was made as shown in Table 3. a single flyby of both Triton and Neptune can be per- The capability of the two options to fulfil the science formed. This is a compromise on the temporal distri- objectives was traded-off, as shown in Figure1. It was bution of the measurements. The DV budget of the taken into account that the total data volume of the two selected trajectory is shown in Table 4. concepts would be roughly similar. This is due to the extremely low data rate at Neptune (12 kbit/s) and very high amount of data that can be collected by the flyby 6.2 Flyby trajectory design and trade- mission (200 Gbit). A flyby mission could spend a simi- off lar time sending data back to Earth as an orbiter, which The in-Neptunian system trajectory has been designed would cause a bottleneck for the orbiter. However, the to maximise the total amount of data, as well as the va- temporal and spatial distribution of the obtained data riety of the measurements. The design of the in-system will be less with a flyby mission, decreasing the total trajectory is constrained and driven by the following: science return that can be retrieved from the data. The fulfilment of the science objectives was weighted at 60 • The incoming hyperbolic excess velocity vector.

5 • The requirement for a Triton flyby. Neptune for gathering data will be mounted such that they point in the same direction as the Yagi antenna, al- • The requirement for observing the probe commu- lowing concurrent probe communications and Neptune nications signal during its atmospheric entry. observation. • The requirement for a close Neptune flyby for both Due to the high curvature of the trajectory, during observations and gravity science. the close flyby phase, the fulfilment of the second of the above requirement seems to be not feasible. In order • The desire for a high Neptune flyby inclination for to accomodate it, the angle between the trajectory and measuring zonal gravity harmonics. the required orientation of the fixed instruments and the HGA should vary instant by instant. The incoming excess velocity vector, combined with Several options were investigated and traded-off for the small required periapsis distance, precludes the pos- the solution of this problem, namely combinations of sibility of combining the Triton flyby requirement and single-direction fixed pointing, multiple-direction fixed high inclination Neptune flyby. However, a flyby trajec- pointing and variable pointing instruments (by use of a ◦ tory with an inclination of approximately 30 is possible, turntable). To accommodate these requirements, the for which the observable amplitude of, for instance, the instruments requiring Neptune pointing are mounted J2 gravity field signal will still be ≈50 % that of a polar obliquely on a turntable on the side of the satellite, al- orbit. lowing this combined pointing requirement and maxi- mizing the science return, at the cost of a system mass 7 Spacecraft Design and complexity increase. The two required degrees of freedom are attained by, firstly, a rotational degree of The design of the spacecraft subsystems was performed freedom provided by the rotation about the axis of the using a concurrent design environment, allowing the HGA and secondly, a degree of freedom provided by work of each subsystem engineer to be immediately the turntable. Similarly, this will allow the combined flown up and down to related calculations. In this man- instrument and HGA usage during the Triton flyby. ner, it is ensured that all analyses are performed in a The Yagi antenna is mounted at the same oblique consistent and up-to-date manner. Margins are applied angle (fixed to the spacecraft) as the instruments on consistently on a subsystem and system level, with rela- the turntable, to accommodate the concurrent usage of tively high values due to the weak heritage of spacecraft the instruments and the antenna. components in the outer solar system. A preliminary configuration of the whole spacecraft The main drivers of the spacecraft design were the is shown in Figure2. following: • Concurrent tracking and Neptune observation • Concurrent probe data collection and Neptune ob- servation • Flyby tracking accuracy • Flyby power requirement • Long mission duration • Neptunian environment

7.1 Spacecraft layout The mission requirements drive much of the placement of the instruments on the spacecraft, namely due to the following: • The requirement to observe the probe during its entry and descent, in order to prevent the loss of observation time during this period. • The requirement for both gravity science (pointing of HGA towards Earth) and Neptune observation Figure 2: A cartoon of the spacecraft. during closest approach (pointing of instruments towards Neptune) For probe communications, which have to be per- 7.2 Mass budget formed in UHF, a dedicated Yagi antenna is used, al- lowing more freedom for the pointing of the HGA. In addition, the instruments which need to be pointed at

6 Mass estimates have been obtained in a bottom-up is used mostly for TT&C and the transmission of the manner, i.e from mass of constituent components, where scientific data. The Ka-transponder enables the precise this was deemed feasible for the current study and as a Doppler tracking of the spacecraft for gravity science representative fraction of totaldry/propellantmass oth- measurements. erwise (in case of structural, propellant system dry The Cebreros station hosts a 35m antenna and is able mass)(see Figure 5 at the last page). Since it is ex- to receive and transmit data in the X-band, and by the pected that an extensive testing will be required for time of the mission this will also be true for the Ka- flight qualifying (existing) components for a Neptunian band. mission, relatively high margins were put on most sub- A diameter of 4m for the HGA (High Gain Antenna) system mass estimates, in addition to the 20% system should suffice to achieve a SNR (Signal to Noise Ratio), mass margin. The relatively low payload mass fraction both in uplink and downlink, suitable for the communi- that is achieved (< 10%) is largely due to the addition cations with Earth and the transmission of the scientific of the probe on the spacecraft, which takes up a consid- data. erable percentage of the spacecraft mass. It should be The Ka-transponder transmits with a sample time noted that the estimated payload mass obtained here is of 60s. The ground segment can estimate the range very similar to the payload mass estimated during the and the range-rate of the spacecraft by measuring the mission architecture trade-off phase, putting extra con- time of flight, and the Doppler-shift of the signal re- fidence on the robustness of the numerical basis of the spectively. This operational mode imposes the pointing trade-off. requirement accuracy of 0.1◦ from this subsystem. The main part of the on-board data is transmitted in 7.3 Power system and budget the X-band. At a distance of 31 AU, the telecommu- nications system should ideally be able to have a data For the primary spacecraft power source, it was chosen rate of 16 Kbps but is assumed to be 12 Kbps with a to use ASRGs (Advanced Stirling Radioisotope Gener- margin. ators), currently under development by ESA. A single The telecommunications between the spacecraft and unit will weigh 34 kg, providing 160W at the beginning the probe will be kept using a Yagi-Uda antenna on- of life, with an annual degradation of 0.8The power re- board the spacecraft and an omni-directional antenna quirements of the separate subsystems have been ana- on-board the probe, transmitting in UHF. The dimen- lyzed from their required constituent parts and opera- sions of the devices have been estimated considering the tional phases in order to size the power subsystem. The severe attenuation due to Neptune’s atmosphere. close flyby mission phase strongly drives the system de- sign, since the full instruments package, HGA and atti- 7.5 Command and Data Handling sub- tude control system will need to be active at the same time. The required power at each in-system phase was system analyzed to obtain the power and energy requirement In order to achieve a successful commands handling and during this phase and a trade-off was made between a reliable data storage, a fully redundant processing sys- the number of batteries and ASRGs. It was chosen to tem has been implemented. It consists of two radiation- use three ASRGs, providing 375 W of power after 16.4 hardened microprocessors which are capable of manag- years,taking into account three years of pre-launch stor- ing the data reception for all the instruments, reaching age and batteries with a capacity of 3580 Wh after 13.4 speeds up to 132 MHz. As the total amount of data years. This option was chosen over smaller batteries expected during the flyby shall reach 200 Gb, a 2x240 and four ASRGs to avoid the inclusion of an additional, Gb data storage system has been included, allowing the expensive, ASRG, as well as from a spacecraft configu- microprocessors saving all the information obtained dur- rational point-of-view. ing the mission as well as having a secondary back-up The design of the thermal subsystem is driven by the memory. combination of a flyby, where the environment is For the probe a data handling system able to collect hot and the Neptune operational phase, where the en- and send all the data from the instruments to the space- vironment is cold. At the Venus flyby, the HGA will be craft has been designed. This system is also fully redun- used as a heatshield to deflect the majority of the incom- dant and it has been designed to protect the data pro- ing radiation, with louvered radiators, sized at 1.8 m2 cessing and transmission due to the critical importance using the heat balance equation. These radiators will be of a synchronized communication between the space- open at Venus and closed in the outer solar system cold craft and the probe. The microprocessor selected is environment. The heat path from the hot ASRGs to the a radiation-hardened device capable of operating at 15 spacecraf is regulated using heat pipes. Active thermal MHz. control power requirements have been determined to be 60 W. 7.6 Additional subsystems 7.4 Telecommunication sub-system The additional spacecraft subsystems have been sized according to their requirements, which have been de- The spacecraft allocates two transponders for communi- rived from higher level requirements. cations. The Deep Space Transponder provides a link in The AOCS system has been sized from the pointing X-band between the Earth and the main spacecraft and accuracy and stability requirements imposed by the in-

7 Failures Likelihood Consequence Peak Power [W ] 382.46 (1-5) (1-5) Data rate [kb/s] 25 ASRG 3 3 FPA [deg] 35 Reaction Wheels 3 3 TPS Mass fraction[%] 28.21 Navigation Cam- 2 4 Relative velocity [km/s] 22.6 era Descent time[h] 1.5 Turntable 3 3 Total data collection[s] 5400

Table 6: Risk assessment table. Table 7: A summary of the main probe.

Probe 200 M e strumets and communications systems, as well as in- Bus 300 M e flight stabilization conditions. Sizing has been per- Power 150 M e formed using (Werty and Larson, 1999) and (Hubbard, 2010). A set of four momentum wheels is used for fine Safety 20 M e pointing and control, with coarse pointing and maneu- Flight operations 100 M e vers performed by a set of 16 attitude control thrusters. Science operations 50 M e Taking data from the aforementioned references, the Management 100 M e system pointing requirements given in Section Instru- Launcher 175 M e ments are met by a good margin. Payload mass (probe) 100 M e The propulsion system has been sized based on the Payload mass (spacecraft) 250 M e amount of required ∆V and an Isp of 320s. This results Earth ground stations (2 sta- 120 M e in a propellant mass of 517.6 kg. From this, a 20% mass tions, 12h/day) fraction was taken to obtain the propulsion system dry mass. Total 1,565 M e For the structure and harness mass, an empirical mass estimate based on previous missions, such as JUICE Table 8: A table showing the cost of the mission (JUICE yellow book) was used, taking into account the required additional secondary sructure for the probe at- A mass budget for the probe was made, based on the tachment. required instrument package and ... barring significant advances in TPS (Thermal protection system) technol- 7.7 Risk Assessment ogy there are a few materials able to withstand the heat fluxes that will be encountered. The fully dense carbon A risk assessment was performed to identify and mit- phenolic has been selected for this purpose. igate the main risks associated with the system. The The probe design is based on ([11]) and it has been main risks that were identified are shown in Figure7.7. adapted to our instrumentation requirements. The Only the main risks are shown for the sake brevity. In main values are reported in the following tables. order to mitigate the reaction control wheel failure risk, The Square Kilometer Array (SKA) will possibly be the momentum wheel system is made redundant. Simi- designed for the ground based observation campaign, if larly, either a single battery assembly or a single ASRG its ready at that time. With a frequency of 500 MHz can fail without compromising the capability of the sys- and an extension radius of 3000 km we will be able to tem to provide sufficient power. The turntable will be get an array resolution of 0.025 and a spatial resolution launched in a configuration that will maximiye the pos- of about 600 km2. sible science return in case the system fails and can no longer rotate. Similarly, it would be possible to include a safe mode that the system reverts to in case of un- 8 Cost foreseen circumstances. In case of failure of the navi- gation camera, the optical wavelength segment of the A breakdown of the cost of the mission can be seen in instrument package can take over the function of this Table 8. instrument. In case of a development delay, a 2 year later launch option has been identified, taking roughly 2 years longer than the nominal mission shown here. 9 Descoping Options

In order to account for the possibility of cost or schedul- ing overruns, as well as a decrease in available budget, 7.8 Probe Design it is important to have a number of possibilities for de- scoping the mission. These options would decrease the In order to fulfill the science objectives, the probe is de- science returns of the mission, but allow the mission to signed to reach 100 bar within the required range. The continue in a downscaled fashion. A number of these probe uses its own Power system. The batteries selected options have been identified for the current mission : are Li-Ion. Considering the length of the interplanetary transfer, a battery loss of 3% per year is assumed for • Increasing the flight time and launch date of the our cells. mission. A launch opportunity, two years after the

8 projected date and taking approximately two years [2] Bagenal, F., Giant planet magnetospheres, An- longer, has been generated using the PaGMO soft- nual review of earth and planetary sciences, 20, ware to validate this option. 1992. • Removing instruments related (largely) to achiev- [3] Cecconi, B. and Zarka, P., Direction finding and ing the secondary science objectives. Due to the antenna calibration through analytical inversion mass margin available in the current design itera- of radio measurements performed using a sys- tion, several instruments deemed to be non-critical tem of two or three electric dipole antennas on for accomplishing the primary science requirements a three-axis stabilized spacecraft., Radio Science, would be removed from the payload. 40, RS3003, 2005.

• Eliminating or downscaling the probe. Although [4] The JUICE Science Study Team, JUICE Explor- the probe is crucial to attaining the primary sci- ing the emergence of habitable worlds around gas ence objectives, a downsized version of the mission giants, ESA, 2011. with less emphasis on Neptune atmospheric sci- ence could be flown without a probe, or one with a [5] Lindal, F. G., The Atmosphere of Neptune: An smaller payload package or a less deep design depth. Analysis of Radio Occultation Data Acquired with Voyager 2, The Astronomical Journal, 103, 1991. 9.1 [6] Marley, M. and Dudzinski, L., Planetary Science Decadal Survey JPL Rapid Mission Architecture Jupiter and Venus are a Planetary Protection Category Neptune-Triton-KBO Study Final Report, NASA I targets. This means that they are not of direct interest JPL, 2010. for understanding the process of chemical evolution or the origin of life. [7] Martin, S. C., Neptune’s zonal winds from near- Triton is of significant interest relative to the process IR Keck adaptive optics imaging in August 2001. chemical evolution and the origin of life, but there is Astrophys and Space Sci. 337, 65-78, 2012. only a remote chance that contamination carried by a spacecraft could compromise future investigations of the [8] Meeus, J., Equinoxes and solstices on Uranus and satellite. Therefore it represents a Planetary Protection Neptune. J. Br. Astron. Assoc 107, 6, 1997. Category II target. [9] De Pater, I. and Lissauer, J. J., Planetary Sci- , whose flyby is performed at a safe distance, ences, Cambridge University Press, 2nd ed., is a Planetary Protection Category III target (planet of 2012. chemical evolution and/or origin of life interest or for which scientific opinion provides a significant chance of [10] De Pater, I., et al., Possible Microwave Absorp- contamination which could compromise future investi- tion by H2S Gas in Uranus’ and Neptune’s At- gations.) mospheres, Icarus, 91, 1991. Consequently the mission either needs to demonstrate that the likelihood of collision with Europa/Triton is [11] Planetary Entry Probes., PEP-Assessment study, < 10−4, or undergo active bioburden reduction to Prepared by PEP/CDF Team, June 2010. meet the requirement that the probability of inadver- −4 [12] Salo, H. and Hanninen, J., Neptune’s Partial tent contamination is < 10 (dry heat or plasma Rings: Action of on Self-Gravitating Arc sterilisation)([11]). Particles, Science, 282, 1998. [13] Soderblom, L. A., et al., Triton’s geyser-like 10 Conclusion plumes - Discovery and basic characterization, Science, 250, 1990. The current mission proposal is the outcome of an ex- tensive trade-off process, resulting in the conclusion that [14] Sromovsky, L. A., et al., Neptune’s Atmospheric the flyby mission architecture is a competitive option for Circulation and Cloud Morphology: Changes Re- a cost-effective and fast option for achieving the primary vealed by 1998 HST Imaging, Icarus, 150, 2001. science objectives that were identified. The current pro- posal allows the full instrument package that was pro- [15] Wertz, W. J. and Larson, J. R. (Larson), Space posed to be flown. The mission and system design was Mission Analysis and Design., 3rd edition, Micro- performed using extensive and consistent margins and cosm press, 1999. is largely driven by the required large data collection [16] Zarka, P., The auroral radio emissions from plan- during the flyby. etary magnetospheres - What do we know, what don’t we know, what do we learn from them?, Ad- References vances in Space Research, 12, 1992. [17] Zarka, P., Auroral radio emissions at the outer [1] Agnor, C. B. and Hamilton, D. P., Neptune’s cap- planets: Observations and theories, Journal of ture of its moon Triton in a binary-planet gravi- Geophysical Research, 103, 1998. tational encounter, Nature, 441, 2006.

9 Flyby trajectory Orbiter trajectory Time-of-flight [years] Neptune dry Neptune dry Neptune dry Neptune dry mass (Soyuz) mass (Ariane mass (Soyuz) mass (Ariane [Kg] 5)[Kg] [Kg] 5)[Kg] 11 483 1788 148 554 22 1028 3798 543 2020 13 679 2481 182 678

Table 1: Estimated dry mass in Neptune orbit/flyby trajectory using Soyuz/Ariane5 as launcher for three different trajectories.

Flyby trajectory Orbiter trajectory Time-of-flight [years] Neptune dry Neptune dry Neptune dry Neptune dry mass (Soyuz) mass (Ariane mass (Soyuz) mass (Ariane [Kg] 5)[Kg] [Kg] 5)[Kg] 11 -16.09 85.69 -44.99 -47.72 22 21.22 257.32 -9.15 79.26 13 -3.11 143.25 -42.93 -40.06

Table 2: Estimated payload mass in Neptune orbit/flyby trajectory using Soyuz/Ariane5 as launcher for the same three trajectories.

Dry mass contribution Without Margin (%) Total Fraction of margin (kg) total (with- out probe) (%) Structure (including sec- 378.80 20.00 454.56 26.69 ondary structure) Harness 106.33 20.00 127.60 7.49 Thermal control 72.23 20.00 86.67 5.09 Communication 96.48 20.00 115.78 6.80 AOCS 38.70 10.00 42.57 2.50 Propulsions 103.52 20.00 124.22 7.29 Power/ ASRG 102.00 10.00 112.20 6.59 Power/ battery 98.00 5.00 102.90 6.04 Instruments (excluding 110.50 20.00 132.60 7.78 probe) Electronics/command data 38.50 10.00 42.35 2.49 handling (including power control and distribution) Mechanism (TurnTable) 16.80 20.00 20.16 1.18 Probe 289.04 20.00 346.85 Total Dry 1,708.46 System Margin % 20.00 341.69 Total dry with margin (excl 2,050.15 launch adapter)

Table 5: Mass budget

10