Accepted Manuscript

Europa Planetary Protection for Juno Jupiter Orbiter

Douglas E. Bernard, Robert D. Abelson, Jennie R. Johannesen, Try Lam, William J. McAlpine, Laura E. Newlin

PII: S0273-1177(13)00151-8 DOI: http://dx.doi.org/10.1016/j.asr.2013.03.015 Reference: JASR 11293

To appear in: Advances in Space Research

Received Date: 9 August 2012 Revised Date: 21 February 2013 Accepted Date: 3 March 2013

Please cite this article as: Bernard, D.E., Abelson, R.D., Johannesen, J.R., Lam, T., McAlpine, W.J., Newlin, L.E., Europa Planetary Protection for Juno Jupiter Orbiter, Advances in Space Research (2013), doi: http://dx.doi.org/ 10.1016/j.asr.2013.03.015

This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting proof before it is published in its final form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain. 1 2 3 4 5 Europa Planetary Protection for Juno Jupiter Orbiter 6 7 Douglas E. Bernard, Robert D. Abelson, Jennie R. Johannesen, Try Lam, William J. 8 McAlpine, and Laura E. Newlin 9 10 11 Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, 12 Pasadena, CA 91109, USA, douglas.e.bernard@jpl.nasa.gov, 13 [email protected], [email protected], 14 [email protected], [email protected], [email protected] 15 16 17 © 2011 California Institute of Technology. Government sponsorship acknowledged. 18 19 Abstract 20 21 22 NASA's Juno mission launched in 2011 and will explore Jupiter and its near 23 environment starting in 2016. Planetary protection requirements for avoiding the 24 contamination of Europa have been taken into account in the Juno mission 25 26 design. In particular Juno's polar orbit, which enables scientific investigations of 27 parts of Jupiter's environment never before visited, also greatly assist avoiding 28 close flybys of Europa and the other Galilean satellites. 29 30 The science mission is designed to conclude with a deorbit burn that disposes 31 32 of the spacecraft in Jupiter’s atmosphere. Compliance with planetary protection 33 requirements is verified through a set of analyses including analysis of initial 34 bioburden, analysis of the effect of bioburden reduction due to the space and 35 Jovian radiation environments, probabilistic risk assessment of successful 36 37 deorbit, Monte-Carlo orbit propagation, and bioburden reduction in the event of 38 impact with an icy body. 39 40 41 1. Introduction 42 43 44 NASA’s next mission to the outer planets will send a spacecraft to Jupiter for 45 a detailed scientific investigation of the planet itself. The Juno mission—named 46 for the mythological wife of the Roman god Jupiter—was launched in 2011 and 47 will arrive at Jupiter in 2016. Like the eponymous Roman Juno, the Juno mission 48 49 is designed to figuratively brush away the clouds to learn about Jupiter’s true 50 behavior. NASA’s New Frontiers Program selects among competing high- 51 science-return Principal Investigator-led planetary science investigations. In 52 2005, NASA selected Juno as the second mission of this Program with Scott 53 54 Bolton of Southwest Research Institute as the Principal Investigator. Juno is 55 managed by the California Institute of Technology’s Jet Propulsion Laboratory 56 (JPL). 57 58 Juno’s highly elliptical polar orbit and low perijove (above the cloud tops by 59 about 7% of Jupiter’s radius) allow it to explore portions of the Jovian 60 61 62 63 64 65 1 2 3 4 environment never before visited. Its suite of instruments is designed to 5 6 investigate the atmosphere, gravitational fields, magnetic fields, and auroral 7 regions of Jupiter from this vantage point. This orbit also helps avoid or limit 8 exposure to the most intense portion of Jupiter’s radiation belts. This limited 9 radiation exposure, along with the low-power spacecraft and instrument design, 10 make it possible for Juno to perform its observation using solar power – making it 11 12 the first solar-powered mission to the outer planets. 13 14 The Juno mission is not orbiting or flying close to Europa or any other 15 Galilean satellite – its polar orbit does not even bring it close to them. 16 Nevertheless, planetary protection requirements for avoiding the contamination 17 18 of Europa have been taken into account in the Juno mission design. 19 20 The science mission is designed to conclude with a deorbit burn that disposes 21 of the spacecraft in Jupiter’s atmosphere. Compliance with planetary protection 22 (PP) requirements is verified through a set of analyses including analysis of initial 23 bioburden, analysis of the effect of bioburden reduction due to the space and 24 25 Jovian radiation environments, probabilistic risk assessment of successful 26 deorbit, Monte Carlo orbit propagation, and bioburden reduction due to impact 27 with an icy body. 28 29 This paper discusses the Juno mission’s approach to meeting planetary 30 31 protection requirements for missions in orbit at Jupiter. The approach that meets 32 Europa planetary protection requirements easily meets planetary protection 33 requirements for the other Galilean satellites, so this paper focuses on Europa. 34 35 2. Juno Overview 36 37 38 Juno is a science mission, so this overview starts with the science objectives. 39 The mission design and spacecraft design are then outlined. 40 41 2.1. Science 42 43 44 The overall goal of the Juno mission is to improve our understanding of the 45 solar system by understanding the origin and evolution of Jupiter. The science 46 objectives for Juno are laid out in the New Frontiers Program Plan: 47 48 49  Atmospheric Composition: Juno investigates the formation and origin 50 of Jupiter’s atmosphere and the potential migration of planets through 51 the measurement of Jupiter’s global abundance of oxygen (water) and 52 nitrogen (ammonia). 53 54  Atmospheric Structure: Juno investigates variations in Jupiter’s deep 55 atmosphere related to meteorology, composition, temperature profiles, 56 cloud opacity, and atmospheric dynamics. 57 58  Magnetic Field: Juno investigates the fine structure of Jupiter’s 59 magnetic field, providing information on its internal structure and the 60 nature of the dynamo. 61 62 63 2 64 65 1 2 3 4  Gravity Field: Gravity sounding by Juno explores the distribution of 5 6 mass inside the planet. 7  Polar Magnetosphere: Juno explores Jupiter’s three-dimensional polar 8 magnetosphere and aurorae. 9 10 11 The science objectives will be accomplished by a payload suite supporting 9 12 investigations: a magnetometer (MAG), a microwave radiometer (MWR), a 13 gravity science experiment using the spacecraft’s X-band communications 14 15 system plus a Ka-band translator, a plasma instrument with four sensors to 16 measure electron and ion particle distributions (JADE), an energetic particle 17 detector with three sensors (JEDI), a electromagnetic field sensor (Waves), an 18 ultraviolet imaging spectrometer (UVS), an infrared imaging spectrometer 19 (JIRAM), and an optical camera (JunoCam) (Matousek, 2005). 20 21 22 2.2. Interplanetary Trajectory 23 24 Juno was launched from Cape Canaveral Air Force Station in August 2011 on 25 an Atlas V 551 launch vehicle. The Juno spacecraft interplanetary trajectory to 26 27 Jupiter is shown in Figure 1. After separation of the spacecraft from the launch 28 vehicle, the flight system (spacecraft) began an almost 5-year cruise phase to 29 Jupiter with two Deep Space Maneuvers (DSM) roughly at aphelion (~ 1 year 30 after launch), and an Earth flyby (EFB) at 500 km altitude, 26 months after 31 32 launch. The spacecraft will arrive at Jupiter in July 2016. Interplanetary cruise 33 concludes with a Jupiter Orbit Insertion (JOI) burn that places Juno into a polar 34 107-day capture orbit around Jupiter. 35 36 Only the spacecraft and the upper stage escape Earth orbit. Earth and Jupiter 37 are the only solar system bodies encountered—at no time does the spacecraft 38 39 come closer than 1.6 AU to Mars. 40 41 2.3. Orbital Mission 42 43 At Jupiter, Juno’s initial 107-day capture orbit and its ~11-day science orbits 44 45 are all polar orbits, allowing unprecedented views of the North and South polar 46 regions and in-situ measurements at all latitudes. See Figure 2. The red circles in 47 Figure 2 represent the orbits of Io, Europa, Ganymede, and Callisto (in 48 increasing distance from Jupiter). The Period Reduction Maneuver (PRM) at the 49 50 end of the capture orbit sets up the 10.9725-day orbit that keeps each 51 successive perijove (PJ) visible at the Goldstone Deep Space Network (DSN) 52 complex in California. This timing is necessary so that mapping of the gravity 53 field at Jupiter may be done using X-band and the Ka-band uplink/downlink 54 capability at that complex. The precise timing of the orbit period is achieved by 55 56 an Orbit Trim Maneuver (OTM) scheduled 4 hours after each perijove; this 57 enables the equator crossings to be equally spaced with respect to each other. 58 Following the full set of science orbits plus one spare orbit, the spacecraft will 59 perform a deorbit maneuver near apojove of the last orbit to impact Jupiter, 60 61 62 63 3 64 65 1 2 3 4 thereby ending the mission and providing continued avoidance of the Galilean 5 6 satellites as required for planetary protection. 7 8 2.4. Juno Flight System 9 10 The Juno spacecraft is shown in Figure 3. The most prominent features are 11 the three solar array (SA) wings. Two wings have four solar panels each and the 12 13 third has three solar panels plus a magnetometer boom. The magnetometer 14 boom is the home for MAG investigation sensors that include Fluxgate 15 Magnetometers (FGM) and the Advanced Stellar Compass (ASC) optical heads. 16 17 Figure 4 provides close-up views of the central core of the spacecraft. Of 18 19 particular interest from a planetary protection viewpoint are the Stellar Reference 20 Unit (SRU) Optical Heads and the electronics vault. In Figure 4, the electronics 21 vault is situated beneath the High Gain Antenna (HGA). In addition to sensors for 22 Instruments mentioned above, this view also shows the Medium Gain Antenna 23 24 (MGA), the Fore Low Gain Antenna (FLGA), and the Reaction Control System 25 Rocket Engine Modules (RCS REM). 26 27 The vault walls are solid titanium and the electronics boxes are packed fairly 28 closely together to allow each box to help shield its neighbors from radiation. The 29 Inertial Measurement Units (IMUs) and SRU electronics are the most heavily 30 31 shielded units inside the vault. 32 33 3. Europa Planetary Protection requirements 34 35 The Planetary Protection requirements and goals for Juno address Mars, Io, 36 37 Europa, Ganymede, and Callisto. This paper addresses the most difficult 38 requirements: Europa planetary protection. NASA has defined three Europa- 39 related planetary protection requirements for the Juno mission (Newlin 2008). 40 They are (paraphrased): 41 42 1. Keep the probability of Europa impacts during the prime Juno mission 43 –4 44 less than 1x10 ; 45 2. Provide an end-of-mission plan that addresses disposition of the 46 spacecraft and ensures continued avoidance of Galilean satellite 47 48 impact after the mission has completed its observations; 49 3. Demonstrate that the probability of contamination of an Europan ocean 50 is less than 1x10–4, taking into account the possibility of an 51 unsuccessful deorbit burn. 52 53 54 The first requirement is met by a mission design that avoids close flybys of 55 Europa. The second requirement is met by the inclusion of a deorbit burn at the 56 57 end of the science mission. 58 59 60 61 62 63 4 64 65 1 2 3 4 The third requirement has required much deeper analysis than the other two 5 6 and is the subject of this paper. For contamination of the Europan ocean to 7 occur, the following low probability events would all need to occur: 8 9 1. A spacecraft failure prevents it from being able to successfully complete 10 the deorbit maneuver; 11 12 13 2. A failed spacecraft impacts the Europan surface prior to radiation 14 sterilization (7 Mrad); 15 16 3. Microbes survive the impact long enough to contaminate the ocean. 17

18 Each of these will be considered in turn below. 19 20 21 4. Juno PP Approach 22 23 Juno’s meets its planetary protection requirements through design of the 24 trajectory, sterilization in the Jupiter environment, and sterilization by impact. 25 26 27 4.1. Trajectory Design 28 29 Note that Juno’s science objectives are all related to the planet Jupiter. There 30 are no science objectives related to the moons of Jupiter and Juno’s trajectory is 31 designed to avoid—rather than encounter—the Galilean satellites. In addition, 32 33 the Jupiter mission design includes an end-of-mission deorbit burn that targets 34 disposal of Juno in Jupiter’s atmosphere. 35 36 Based on the above, the Juno mission, operating as designed, will not impact 37 any Galilean satellites. Therefore, in order to completely assess the Juno 38 39 mission’s compliance with planetary protection (PP) requirements, an analysis of 40 Juno failure modes and implications is also required. 41 42 Juno’s interplanetary trajectory has been analyzed to determine the 43 probability of impact in the event that a failure in cruise prevents later re- 44 45 targeting. The Europa impact probabilities are negligible—and the Juno 46 spacecraft would fly by the Jupiter system without going into orbit. 47 48 In addition, Juno’s Jupiter orbit trajectory has been analyzed to determine the 49 probability of impact in the event that a failure in orbit prevents later re-targeting. 50 The Europa impact probabilities during the Juno science mission are also 51 52 negligible. 53 54 The more interesting analysis assesses the probability of spacecraft-Europa 55 impact following the planned deorbit maneuver in the event that the spacecraft is 56 in Jupiter orbit but is unable to perform the planned deorbit maneuver. 57 58 59 60 61 62 63 5 64 65 1 2 3 4 For the discussion of spacecraft reliability and the probability of spacecraft 5 6 failure, see Section 5. For a discussion of the probability of Europa impact given 7 a spacecraft failure, see Section 6. 8 9 4.2. Space Environmental effects 10 11 Several aspects of the space environment are lethal to some organisms and 12 13 prevent others from repairing radiation damage. 14 15 Jupiter’s radiation environment is much more intense than the radiation 16 environment at the Earth (or any other planet in our solar system). Although 17 Juno’s orbit is designed to avoid the worst of the radiation during the mission, the 18 19 orbit naturally evolves in such a manner that there would be a much higher 20 radiation dose per orbit for times after the planned deorbit burn. For a discussion 21 of the orbit evolution if the deorbit burn is not successful, see Section 6. For a 22 discussion of the radiation environment for Juno components in that eventuality, 23 24 see Section 7. 25 26 About half of the solar power harvested by the solar panels is used for 27 heaters at Jupiter. This means that after a complete spacecraft failure, the 28 heaters would cease functioning and spacecraft components would experience 29 much colder temperatures than they do on a working spacecraft. A failure that 30 31 left Juno uncommandable would terminate the periodic commanded precessions 32 that keep the solar arrays pointed at the sun. Within a few years as Jupiter 33 continues to orbit the sun, the solar arrays would be expected to be pointed so 34 far from the sun that the electrical power to the vehicle would be lost and never 35 36 recovered. And that is generously assuming that the spacecraft would survive for 37 years without ground commanding. Thermal analysis of the Juno spacecraft with 38 no internal power dissipation predicts steady state temperatures in the -110 C to 39 -170 C range. 40 41 The estimated initial bioburden and the combined effects of vacuum, 42 43 radiation, and cold on micro-organisms are discussed in Section 8. Section 8 44 also shows a high likelihood that complete sterilization is achieved after 150 45 years in Jupiter orbit. However, Juno is using a 400 year time to sterility 46 (corresponding to 7 Mrad) in order to demonstrate compliance to NASA PP 47 48 requirements. 49 50 4.3. Sterilization due to impact 51 52 Impact of Juno into Europa does not imply contamination of the Europan 53 54 ocean. Because Juno would be impacting from Jupiter orbit, in almost all cases 55 the impact would be at a speed greater than 20 km/s. In addition, an impact 56 would be a surface impact and would not penetrate the thick Europan ice. 57 Contamination of the Europan ocean would require surface contamination 58 followed by the transfer of viable micro-organisms to the deeply buried ocean. 59 60 61 62 63 6 64 65 1 2 3 4 Section 9 analyzes the physics of such an impact and discusses the 5 6 likelihood of microbe survival. 7 8 5. Spacecraft Reliability 9 10 Juno has performed a limited-scope probabilistic risk assessment (PRA) to 11 quantify the reliability of the Juno spacecraft and its ability to successfully deorbit 12 13 at the end of the mission. PRA is a systematic and comprehensive methodology 14 used to numerically evaluate the risks associated with complex systems, and is 15 widely used in the nuclear and aircraft industries. The use of PRA for NASA 16 robotic spacecraft missions has previously been limited, and thus a combination 17 18 of standard tools and techniques, along with new approaches and reliability 19 models, was used for Juno. 20 21 The scope of the Juno PRA was limited to those spacecraft components 22 required for deorbit whose failure could result in the spacecraft being unable to 23 24 perform its deorbit maneuver. The PRA modeled the entire mission duration, 25 from launch to end-of-mission (EOM), in order to account for any partial 26 spacecraft failures that might occur during Cruise and JOI (potentially resulting in 27 a less than fully functional spacecraft reaching Jupiter). However, the interval of 28 interest for assessing compliance with Europa Planetary Protection requirements 29 30 is the ~15 month period between JOI and End-of-Mission (EOM). Critical 31 spacecraft failures occurring earlier in the mission (e.g., launch, cruise, or JOI) 32 would prevent Juno from achieving JOI, thus mitigating the Europa impact threat. 33 34 The spacecraft subsystems analyzed included: 35 36 37  Electrical Power System (EPS) 38  Guidance, Navigation, and Control (GNC) 39 40  Command and Data Handling (C&DH) 41  Telecommunications (Telecom) 42 43  Thermal Control System (TCS) 44  Propulsion (Prop) 45 46 47 For each of the subsystems above, the critical components required for 48 spacecraft operations and control between JOI and deorbit were identified, along 49 with the profile of how they were to be used (i.e., number of power on-off cycles, 50 exposure durations, etc.). Additionally, spacecraft component redundancy and 51 52 cross strapping were also included in the PRA. 53 54 In addition to spacecraft hardware components, the reliability of system-level 55 interactions was modeled, including the contributions of software and mission 56 operations. Analysis of previous mission failures had identified these as potential 57 risk drivers for robotic missions in general. As there was no existing approach to 58 59 estimating this risk for JPL robotic missions, Juno developed a novel approach 60 based on heritage mission data and Juno’s specific operating profile. 61 62 63 7 64 65 1 2 3 4 Lastly, micrometeoroids were explicitly accounted for in the spacecraft 5 6 reliability model. 7 8 The PRA methodology is typically used to compare relative risks between 9 options. The numerical results are usually skewed to the conservative side, 10 particularly when there is limited flight data available for evaluation. Results of 11 the PRA indicate a 95% probability of successfully performing the deorbit 12 13 maneuver at end of mission. The key risk drivers are the spacecraft C&DH/power 14 subsystems (~2.8% failure probability per the PRA methodology), and systems- 15 level interactions (~2% failure probability, extrapolating from historical in-flight 16 anomaly data). The risk contribution from micrometeoroids and other spacecraft 17 18 subsystems is small relative to the above risk drivers. 19 20 There is additional conservatism in the above PRA results in that many 21 mitigations options were identified but not included in the reliability model. These 22 mitigations include operating the spacecraft in a degraded mode (the model 23 currently assumes a part either works or doesn’t work) and attempting recovery 24 25 of a ―failed‖ component. Had these mitigations been included in the reliability 26 model, they would have further increased the probability of successfully 27 deorbiting the spacecraft. 28 29 Given this PRA-based estimate of a 5% chance of failing to perform the 30 31 deorbit burn, we turn our attention to the mission design and the possibility of 32 Europa impact in that eventuality. 33 34 6. Mission Design 35 36 37 Throughout this section, the term RJ represents one Jupiter Radius. So, for 38 example, 5 RJ is a non-dimensional distance equal to 5 Jupiter radii. 39 40 6.1. Jupiter approach 41 42 43 There is negligible probability of impacting Europa during the Jupiter 44 approach phase. Five maneuvers are used to set up the Jupiter arrival 45 conditions. The first of these, at Earth flyby + 10 days, has the largest 46 uncertainty in position and timing (5 R and 2 hours), but even this 3-sigma 47 J 48 uncertainty is much less than the orbital range of Europa at 9.4 RJ. If no 49 additional maneuvers were possible, the spacecraft would not impact Europa. 50 Additional maneuvers at Earth + 6 months and JOI - 6 months reduce the 51 uncertainty to the order of 0.1 RJ and 2 minutes, while maneuvers at JOI - 35 52 53 days and JOI - 9 days refine the targeting to further reduce uncertainties by 54 another factor of 10, to 0.01 RJ and 14 seconds. Additionally, if JOI did not occur 55 for some reason, the spacecraft would pass by Jupiter on a polar trajectory with a 56 perijove altitude of about 0.1 RJ and would not be captured with respect to 57 Jupiter. 58 59 60 6.2. Polar orbit 61 62 63 8 64 65 1 2 3 4 5 6 One key feature of the orbital design that greatly reduces the probability of 7 impact with the Galilean satellites is that the Juno orbit is polar (90º inclination), 8 while the satellites orbit in the plane of Jupiter’s equator. A spacecraft in an 9 equatorial orbit with comparable perijove and apojove ranges (< 0.1 RJ x 38.8 RJ) 10 would cross the orbits of the Galilean satellites twice on each orbit, since the orbit 11 12 ranges for the Galilean satellites (Io: 5.9 RJ, Europa: 9.3 RJ, Ganymede: 15.0 13 RJ, Callisto: 26.3 RJ) are less than the apojove range. As a consequence, a 14 spacecraft in an equatorial orbit could have its orbit greatly perturbed if it crossed 15 a satellite orbit when the satellite was nearby, or it might even impact the 16 17 satellite. Because Juno is in a polar orbit, it cannot encounter the Galilean 18 satellites except when crossing the equatorial plane, and therefore, its trajectory 19 maintains an orbital shape throughout the mission. In addition the crossing of the 20 equatorial plane just after perijove occurs at a range < 0.1 RJ (well away from 21 Galilean satellites), so there is only one possibility per orbit of impacting a 22 23 satellite, when crossing the equatorial plane on approach to perijove (i.e., at the 24 ascending node). 25 26 6.3. Apsidal rotation 27 28 29 At the beginning of Juno’s mission, the perijove latitude is about 3 deg (north) 30 while the apojove position is slightly below the equatorial plane. Juno’s 31 ascending-node equatorial crossing (traversing from apojove to the next perijove 32 and ascending from below the equatorial plane to above the plane) is initially at a 33 very large distance (160 R ) on the capture orbit and near 37 R early in the 11- 34 J J 35 day orbits. Because Jupiter’s shape is markedly oblate, the trajectory undergoes 36 a rotation of the line of apsides (the imaginary line through Jupiter between 37 perijove and apojove locations). The apsidal rotation is about –0.95 deg per 38 orbit, causing more northerly latitudes at perijove and causing the ascending 39 40 nodal range to move closer to Jupiter. This crossing of the equator on approach 41 to the next perijove provides the only opportunity for Juno to impact a satellite. 42 The apsidal rotation also causes the spacecraft to enter more severe radiation 43 areas as the mission evolves. 44 45 Near the end of mission, the perijove latitude reaches about 34 deg north, 46 47 while the ascending node range decreases to just under Europa’s orbital range. 48 Figure 5 shows that Callisto’s orbit is crossed near orbit 12, Ganymede’s orbit is 49 crossed between orbits 23 and 24, and Europa’s orbit is crossed on orbit 34 (the 50 final ascending node between deorbit and Jupiter impact). The selection of PRM 51 52 date (when 11-day orbits are established) greatly influences the proximity of the 53 satellites at the ascending node crossings. The choice of October 19, 2016 as 54 the PRM date provides good separation between Juno near equatorial crossings 55 and the individual Galilean satellite positions throughout the orbital mission. This 56 is true even at the final equatorial crossing, where the separation of Juno and 57 58 Europa is about 10 hours. 59 60 61 62 63 9 64 65 1 2 3 4 Fairly distant satellite encounters cause some perturbations to the Juno 5 6 trajectory. These perturbations behave as a small gravity assist delta-V, much 7 as Juno’s flyby of the Earth produced a (much larger) gravity assist delta-V for 8 the trajectory. The perturbations affect the orbit inclination and the orbit period 9 (related to perijove altitude). Ganymede and Callisto, because of their larger 10 sizes, have more influence on the trajectory and cause larger altitude variations 11 12 than do Europa and Io. Callisto, at 240,000 km range (orbit 4) and 170,000 km 13 range (orbit 7), causes drops in the perijove altitude of about 400 km. The figure 14 also shows that the close approaches with Ganymede and Callisto (and Europa) 15 do not occur near the equatorial crossings at the ascending node. This is also a 16 consequence of the choice of the PRM date. 17 18 19 The orbit trim maneuvers (OTMs) at PJ+4hrs target the spacecraft to the 20 proper equator-crossing longitude at the descending node of the following orbit 21 (for the magnetic field investigation), and they help dampen the effect of the 22 satellite flybys by maintaining the orbital period. These OTMs range in size from 23 24 0.1 m/s to about 3 m/s, with a total of about 40 m/s needed for OTMs for the 25 entire orbital mission. The orbit inclination is not controlled and changes by only 26 a few degrees over the course of the mission. 27 28 The nominal mission plan calls for Juno to deorbit into Jupiter on PJ34 29 30 (October 16, 2017), after 33 11-day orbits. This is accomplished by making a 31 burn on the RCS thrusters at apojove to reduce the perijove altitude to a 32 subsurface value. The current design uses –700 km as the target perijove 33 altitude. 34 35 36 If deorbit is not possible because of spacecraft failure, an analysis of the 37 propagated trajectory shows that the apsidal rotation continues with the orbit 38 remaining basically polar and with similar orbit period and shape. Figure 6 39 shows this apsidal rotation behavior. The location of the equator crossing (where 40 satellite impact could occur) switches twice between inbound and outbound legs 41 42 with respect to perijove before 360 deg of apsidal rotation causes the pattern to 43 repeat. The full cycle of rotation takes about 12 years, which is about the same 44 as the period of Jupiter around the sun. During that period of time, the perijove 45 altitude rises and falls, with the possibility of impact with Jupiter for low perijove 46 altitudes. 47 48 49 A Monte Carlo analysis was undertaken to determine the probability of impact 50 with Europa (and with Jupiter and other Galilean satellites) if deorbit were not 51 possible. This analysis (Lam, et al., 2008) included sampling uncertainties in 52 spacecraft state, satellite ephemeris, and Jupiter’s spherical harmonics, and was 53 54 performed on supercomputers using JPL’s Mission analysis and Operational 55 Navigation Toolkit Environment (MONTE) software. An atmospheric model 56 adapted from one provided by Sushil Atreya of the University of Michigan was 57 used to provide the basis for drag modeling. Solar radiation pressure modeling 58 was also included. Figure 7 shows the results of a total of 4000 cases, each 59 60 propagated for 300 years. The probability of impact with Europa over 150 years 61 62 63 10 64 65 1 2 3 4 for a failed deorbit is 0.81% ± 0.28% (2-sigma uncertainty). During this period, 5 6 3/5 of the trajectories impacted Jupiter and about another 1/3 had no impacts 7 with Europa or any other Galilean satellite. 8 9 In addition to looking at impact probabilities for the failed deorbit case, a more 10 complete impact probability analysis was performed for failure anywhere along 11 the nominal mission while in orbit around Jupiter. Here, the apojove state at 12 13 each orbit was sampled and included with uncertainties in satellite ephemeris 14 and Jupiter spherical harmonics. Atmospheric drag and solar radiation pressure 15 modeling were also included. The results of propagation of 4000 trajectories per 16 apojove for 400 years are shown in Figure 8 on an individual orbit basis and in 17 18 Table 1, where the impact probabilities for all orbits were averaged. For 150 19 years the average percentage of Europa impacts was 0.51% ± 0.04% (2-sigma 20 uncertainty). For 400 years, the average percentage of Europa impacts grew to 21 1.06% when weighting each orbit equally. In will be shown in Section 10 that the 22 average percentage of Europa impacts over 400 years is 0.93% when weighting 23 24 each orbit by the probability of failure in that orbit. 25 26 7. Radiation 27 28 Bioburden within the Juno spacecraft will decrease over time once in Jupiter 29 30 orbit due to increasing radiation exposure from magnetically trapped high-energy 31 electrons and protons around Jupiter. The Juno spacecraft will be exposed to 32 these high-energy particles once per orbit, from Jupiter Orbit Insertion (JOI) to 33 the nominal end of mission at Jupiter orbit 33, as shown in Figure 9 34 35 36 In the very unlikely scenario that the spacecraft cannot perform the deorbit 37 burn, analyses were performed to calculate the long-term radiation exposure to 38 the Juno spacecraft as an input into the bioburden reduction analyses. The 39 computation of the radiation exposure to the Juno spacecraft over long time 40 durations were determined by first propagating the spacecraft trajectory over one 41 42 complete period of apsidal rotation around Jupiter (approximately 12 years). The 43 spacecraft orbit trajectory over 12 years was integrated with the Jovian trapped 44 radiation models to yield the external charged particle environment expected 45 over one apsidal rotation period. Charged particles will deposit energy, primarily 46 47 through ionization, as they interact with the Juno spacecraft materials. The 48 deposited energy (henceforth called total ionizing dose (TID)) is dependent on 49 energy spectra and species of the external charged particle environment, the 50 thickness and composition of the material that is traversed between the external 51 environment (henceforth called shielding), and the composition of the absorbing 52 53 material. A dose depth curve, Figure 10, is often used to show how the energy 54 deposited by the external environment is attenuated by shielding material. The 55 TID in water is expressed in terms of rad (water) as a function of aluminum shield 56 thickness. Radiation exposure in rad (water) is used in the bioburden 57 calculations. 58 59 60 61 62 63 11 64 65 1 2 3 4 A dose depth curve, while useful for generating rough estimates of TID, is of 5 6 limited usefulness when complex shielding configurations and numerous different 7 materials are used in spacecraft construction. For these types of complex 8 spacecraft shielding configurations and multiple construction materials, radiation 9 transport codes are used to determine the total ionizing dose at specific locations 10 within the spacecraft. For the Juno project, the NOVICE radiation transport code 11 12 was used to calculate radiation exposures in materials, electronic parts, and 13 sensors throughout the spacecraft for the 33-orbit mission. The mechanical 14 configuration of the spacecraft, instruments, and electronics units used in the 15 radiation exposure calculations were generated using CAD models, and contain 16 mechanical information down to the electronic part and circuit card level, as 17 18 shown in Figure 11. The flat planar structures in Figure 11 are circuit boards 19 populated with electronic devices. The composition and density of materials 20 used in unit construction and surrounding spacecraft structure are used in dose 21 calculations at the electronic components on the circuit boards. Similar models 22 23 were used to generate the dose input to bioburden reduction calculations for all 24 electronics boxes and spacecraft structure. 25 26 The same mechanical models and radiation transport code that were used to 27 calculate doses for electronic devices, materials, and sensors were used to 28 29 calculate doses in water at several points within each electronics unit and within 30 structural elements on the spacecraft for one complete apsidal rotation. For each 31 electronics unit, the best-protected point in the box was used to determine the 32 amount of shielding surrounding the entire box volume. Assuming that the 33 maximum shielding is present over the entire volume of each electronics unit 34 35 provides a worst case bound for bioburden reduction calculations. 36 37 All spacecraft electronics units and spacecraft structure elements were then 38 grouped by shielding thickness into different zones. The average annual dose 39 for each of the units was derived by dividing the dose accumulated over one 12- 40 41 year apsidal rotation by 12 years. The most heavily shielded electronics within 42 the Juno spacecraft are listed in Table 2. Assuming sterilization at 7 Mrad, note 43 that the last item sterilized will be the SRU optical head – at about 400 years. 44 These electronics units have significantly more shielding than the rest of the 45 spacecraft to protect sensitive electronics parts from the charged particle 46 47 environment. These are also the units that would have the lowest annual doses 48 available for bioburden reduction. 49 50 8. Bioburden 51 52 53 8.1. Approach 54 To determine the time to spacecraft sterility, Juno is using 7 Mrad as the 55 threshold dose value as clarified by the NASA Planetary Protection Office 56 (Conley, 2011, personal communication). This value is independent of initial 57 bioburden and spacecraft temperature. 58 59 60 61 62 63 12 64 65 1 2 3 4 In addition to using a 7 Mrad threshold value, Juno performed a detailed 5 6 bioburden assessment to determine the time for the bioburden population to be 7 reduced to non-reproducible levels (0.1 organisms). The remainder of this 8 section details this approach, which yielded <150 years to sterility were the 9 spacecraft to fail in any of Juno’s 11-day science orbits, and <1500 years were 10 the spacecraft to fail in the capture orbit. 11 12 13 The estimation of spacecraft bioburden and time to sterility in the Jovian 14 radiation environment must consider the number of organisms on the spacecraft 15 at launch as a function of location; the radiation sensitivities of the various 16 organisms likely to be present on the spacecraft at launch; the sensitivities of the 17 18 various organisms to the deep space environment experienced during cruise ; 19 and the total ionizing dose of radiation expected over time for various spacecraft 20 locations (discussed in Section 7). 21 22 The spacecraft bioburden or total number of microorganisms at launch, was 23 estimated by considering the spore bioburden specifications in NPR 8020.12C 24 25 (NASA, 2005) and historical spore bioburden information from past Mars 26 missions; reviewing Appendix A of the SSB Report (Space Studies Board, 2000) 27 for their non-spore-forming organism population assumptions; reviewing the work 28 done by the JIMO project (Kohlhase, 2004) for their microorganism population 29 30 assumptions; and by using Mars Reconnaissance Orbiter (MRO) historical 31 bioburden information, where applicable, for Juno’s MRO heritage hardware. 32 Focus was placed on the hardware with the best protection from the Jovian 33 radiation environment. As shown in Table 2, the best-protected hardware 34 components are the SRU Optical Heads, followed by the IMUs and the SRU 35 36 electronics modules. As will be discussed below, the SRU optical heads would 37 be the last items sterilized if Juno were to remain in Jupiter orbit. 38 39 To determine the radiation sensitivities of organisms, the Juno project 40 performed a literature search for radiation sensitivity data; reviewed Appendix A 41 42 of the Space Studies Board (SSB) Report, Preventing the Forward 43 Contamination of Europa (Space Studies Board, 2000) for the SSB’s radiation 44 sensitivity assumptions; and reviewed work done by the JIMO project (Kohlhase, 45 2004) for its radiation sensitivity assumptions. 46 47 48 The microbial responses to low water activity (Grant, 2004) and to low 49 temperatures (Price, 2004) were investigated. The spacecraft bioburden was 50 then reduced based upon those sensitivities and on the survivability in deep 51 space assumptions stated in Appendix A of the SSB Report (Space Studies 52 Board, 2000). 53 54 55 8.2. Parameter Specifications and Assumptions 56 57 8.2.1. Planetary Protection Parameter Specifications 58 The following PP Parameter Specifications of NPR 8020.12C (NASA, 2005) 59 60 have been used in the bioburden estimation: 61 62 63 13 64 65 1 2 3 4 Time-Temperature for Sterility: A surface, the temperature of which 5 6 exceeds 500°C for more than 0.5 seconds, may be considered sterile (bioburden 7 equal to zero). 8 9 D Value, Z Value: Dry heat microbial reduction at a relative humidity less 10 than 25% (referenced to 0°C and 1 atmosphere) for spores has a D value at 11 125°C (i.e., the time for a reduction by 10 at that temperature) of 0.5 hour for free 12 13 surfaces, 1 hour for mated surfaces, and 5 hours for encapsulated non-metallic 14 material, and a Z value of 21°C (i.e., the change in temperature for a factor of 10 15 change in the D value), on the temperature interval 104°C to 125°C. 16 17 Hardy Organisms: The maximum reduction factor that may be taken for dry 18 heat microbial reduction is 10–4. (This parameter specification is based on an 19 –3 20 assumed fraction of ―hardies‖ of 10 and a reduction of only one order of 21 magnitude in a nominal sterilization cycle.) 22 23 Surface Microbial Density: Various parameter specifications permitted for 24 surface burden estimates: 1 x 105 spores/m2 for uncontrolled manufacturing; 1 x 25 4 2 3 26 10 spores/m for class 100,000 clean room with normal controls; 1 x 10 2 27 spores/m for class 100,000 clean room with stringent controls. Surface density 28 for vegetative microorganisms is 10 times the surface density for spores. 29 30 Encapsulated Microbial Density: Various parameter specifications 31 32 permitted for encapsulated burden estimates (in non-metallic materials): average -3 -3 33 130 cm , specific to electronic parts, 3-150 cm ; and specific to other non- 34 metallic materials, 1-30 cm-3. 35 36 8.2.2. Planetary Protection Parameter Assumptions 37 38 The following parameter assumptions were used in the bioburden estimation: 39 40 Microbial Population: The SSB Europa Report (Space Studies Board, 41 2000) suggests that the microbial population of a spacecraft be described by four 42 types of organisms, as shown in Table 3. The SSB also made an estimate of the 43 44 fraction of each type of organism within the total population. Later, the JIMO 45 project proposed alternative fractions of each type of organism (Kohlhase, 2004) 46 and then further defined four disjoint population sets based on the four types of 47 organisms of the SSB report, as shown in Table 4. The Juno project has adopted 48 the use of these disjoint groups and associated population fractions. Table 4 also 49 50 cites the source of the radiation survival curves to be used by the Juno project. 51 52 Microbial Abundance: It was conservatively assumed that laboratory 53 cultivation by the NASA Standard Assay (NASA, pending), underestimates actual 54 microbial abundance by a factor of 1,000 for each type of microbial 55 56 subpopulation (Space Studies Board, 2000). 57 58 59 60 61 62 63 14 64 65 1 2 3 4 Encapsulated Microbial Density: It was assumed that the encapsulated 5 6 density for vegetative microorganisms is 5 times the encapsulated density for 7 spores (Newlin, 2008). 8 9 Exposure to ultrahigh vacuum of deep space: It was assumed that 10 exposure to the ultrahigh vacuum of deep space has the following effect on the 11 microbial population (Space Studies Board, 2000): (i) the bacterial spores of 12 13 Types B and C are known to be generally resistant to high vacuum, so no 14 reduction credit should be taken; (ii) the vegetative cells of Type A are often 15 susceptible to inactivation by extreme vacuum, so a survival fraction of 0.1 will be 16 assumed; (iii) some of the radiation-resistant vegetative cells of Type D are 17 18 highly resistant to desiccation and others are not, so a survival fraction of 0.5 will 19 be assumed. These are conservative assumptions, as illustrated by work done 20 by Gerda Horneck of DLR and others (e.g., (Hornek, 1993), (Mileikowsky, 2000), 21 and (Saffary, 2002)). 22 23 Exposure to water activity below 0.2: Water activity, a , is defined as the 24 w 25 ratio of vapor pressure of water in the substance to the vapor pressure of pure 26 water at the same temperature. For the Juno spacecraft hardware in deep space, –17 –15 27 aw values will vary from 10 to 10 depending on the temperature. It was 28 conservatively assumed that the exposure to water activity below 0.2 will cause 29 30 the radiation-resistant non-spore-formers to be unable to repair radiation damage 31 (Grant, 2004), so Group 4 will be treated as Group 1. 32 33 Exposure to extreme cold of deep space: It will be conservatively 34 assumed that the exposure to the extreme cold of deep space, temperatures 35 36 below –80 C (193 K) will cause the radiation resistant non-spore-formers to be 37 unable to repair radiation damage (Price, 2004), so Group 4 organisms at this 38 temperature (i.e., in/on a non-operational spacecraft at Jupiter) will be treated as 39 Group 1. 40 41 8.2.3. Other Assumptions and Considerations 42 43 The following additional assumptions were made, and considerations used, 44 during the bioburden estimation: 45 46  Spacecraft components are to be developed and assembled in a class 47 100K clean room or better; 48 49  Hardware exterior surfaces are alcohol wiped to class 100K typical 50 cleanliness; 51  MRO heritage bioburden densities (based on actual bioassays and 52 manufacturing conditions/processes), surface areas, and volumes were 53 54 used where appropriate; 55  Radiation levels used in calculations are the minimums for each module; 56  SRU Electronics Module assumptions: 57 o Surface area of internal surfaces of module is 10 times the external 58 59 surface area of the module; 60 o Board volume 50% of module volume; 61 62 63 15 64 65 1 2 3 4 o Three orders of magnitude reduction in spores due to cure of 5 6 component boards and no vegetative organisms; 7 o Juno is not taking any credit for bioburden reductions due to 8 manufacture or burn-in of electronic components. 9 10 8.3. Initial Microorganism Populations 11 12 Initial microorganism population densities were developed based on the 13 parameters, assumptions, and considerations listed above. Table 5 summarizes 14 the population densities by Type. Table 5 starts with organisms culturable by the 15 NASA Standard Assay (NASA, pending) present at launch. Then the densities for 16 17 the culturable organisms are scaled (up) to include all organisms that are present 18 at launch. Finally the densities are adjusted for space effects to summarize the 19 total number of organisms at JOI. The initial microorganism populations for 20 Jupiter orbit are established by combining the population densities at JOI with the 21 surface areas and non-metallic material volumes for the components of interest. 22 23 This data by type is the starting point for developing population densities by 24 group using the set relations in Table 4. 25 26 8.4. Time to Sterilization 27 The time to sterilization is estimated by applying the radiation dose for each 28 29 component and the radiation sensitivities for each Group (see Section 8.2.2), to 30 the number of organisms for each Group in each component. Table 6 provides 31 the microorganism survival fraction for each component by Group as a function 32 of years after launch. Table 7 provides the number of surviving microorganisms 33 for each component by Group, also as a function of years after launch. The time 34 35 to sterility is predicted to be 126 years, and is dependent on the microorganisms 36 of Groups 2 and 3 in the SRU optical heads. 37 38 As previously mentioned, Juno is using a 400 year time to sterility 39 (corresponding to 7 Mrad) in order to demonstrate compliance to NASA PP 40 41 requirements for the 11-day orbits (4000 years is used for the capture orbit). 42 43 44 45 9. Impact Analysis 46 47 48 Previous Sections addressed the reliability of the spacecraft, the low 49 probability of Europa impact even if the deorbit maneuver is unsuccessful, and 50 the rate of microbial sterilization in the low-temperature, high radiation 51 environment that Juno would be in after a failure. 52 53 54 This Section explores the question of the likelihood of microbe survival if Juno 55 were to impact Europa prior to sterilization. 56 57 9.1. Impact Geometry 58 59 60 61 62 63 16 64 65 1 2 3 4 Section 6 described the Monte Carlo analysis that was used to determine the 5 6 probability of Europa impact. This same analysis can be used to understand the 7 statistics related to the dynamics of the impact; both impact velocity and impact 8 angle are important. In the development that follows, some situations are not 9 analyzed, and since there is adequate margin against the requirement, Juno 10 takes the very conservative position of assuming surface contamination in all un- 11 12 analyzed situations. 13 14 Over a collection of Monte Carlo runs, all 220 instances of Europa impact 15 were analyzed to understand the impact velocity and geometry. Figure 12 shows 16 the cumulative distribution function of impact velocity. Note that the impact 17 18 velocity is typically around 22 km/s and that only 1% of the cases have an impact 19 velocity of less than 20.8 km/s. Higher impact velocities give greater likelihood of 20 sterilization, so to simplify the analysis, the Juno Project adopts 20.8 km/s as the 21 impact velocity and assumes that surface contamination occurs for the 1% of 22 cases where the impact velocity is less that 20.8 km/s. 23 24 25 These same 220 cases were also analyzed to understand the distribution of 26 impact angle. The angle of obliquity is defined to be 0° for a normal impact and 27 90° for a grazing impact. The distribution of angle of obliquity for this sample is 28 shown in Figure 13. Note that 7% of the cases have obliquity greater than 75° 29 30 and only 2% have obliquity higher than 80°. Since the analysis becomes more 31 difficult for higher obliquity cases, Juno does not analyze cases with obliquity 32 higher than 80°. Instead, the conservative assumption is made that surface 33 contamination is possible for obliquity higher than 80°. 34 35 36 37 9.2. Assumptions on Sterilization 38 39 Before analyzing the impact, the Juno Project established criteria to 40 determine if microbes should be assumed to have been killed or sterilized by the 41 42 conditions of impact. The criteria in Table 8 are used: 43 44 Note that if the substrate on which the microbe is located is merely 45 fragmented by the hypervelocity shock of impact (but not vaporized or liquefied), 46 the microbe is conservatively assumed to have survived and the analysis 47 48 continues. 49 50 51 9.3. Impact Analysis 52 53 54 The analysis of high energy/ high velocity impacts is a specialized field of 55 study. To gain access to state-of-the-art analytical tools, the Juno Project 56 contracted with Steve Hancock of Foils Engineering. At Juno’s direction, Mr. 57 Hancock analyzed: 58 59 60 61 62 63 17 64 65 1 2 3 4  Spacecraft impact into Europa at 20.8 km/s at varying angles of 5 6 obliquity; 7  Resulting temperatures of spacecraft materials for intact multiple- 8 substrate impact, and for isolated single material substrate impact 9 (giving more conservative results); 10 11  Physical effect on spacecraft materials: fragmentation? Liquefaction? 12 Vaporization? 13 14  Particle cooling analysis and maximum size of sufficiently cooled 15 particle; 16  Fragment size analysis as a function of obliquity ; 17 18  Burial depth analysis as a function of fragment size. 19 20 The Hancock analysis used the CTH tool with a library of SESAME materials 21 22 models. CTH is a Sandia Shock Physics Code for hypervelocity impact modeling. 23 SESAME, developed by Los Alamos National Laboratory, captures the equations 24 of state in tabular form. Most models include melting, vaporization, ionization, 25 and molecular dissociation for molecules. 26 27 28 The shock strength is shown in Figure 14. In this figure, intersections of the 29 curves give shock pressure and particle velocity for each spacecraft material/ice 30 combination. The specific internal energy of a shocked material is ½ Du2 , where 31 Du is its velocity change. Note that the ice density has a significant effect on 32 shock strength. Since the lower shock strength (and lower resulting 33 34 temperatures) occurs for the 60% density ice, the analysis that follows assumes 35 60% ice, since it is less effective in sterilizing the impacting material. The plate 36 material is heated nearly uniformly through its thickness by the shock; the 37 thickness influences the duration, but not the strength of the shock. 38 39 40 An early conclusion of the study is that at this impact velocity, there is a 41 potential for very high temperatures even if only a small fraction of the kinetic 42 energy is converted to heat, and one needs to look very closely to find a way for 43 a spore or other microbe to survive. For scale, the specific total energy is 216 44 MJ/Kg or about 50 times the energy density of TNT. 45 46 47 Looking next at the expansion after a shock, Figure 15 shows the variation in 48 temperature and volume for a variety of engineering materials following the 49 shock. Solid and liquid materials expand and will fragment when tension occurs, 50 then cool over a much longer time scale. Vaporized material will continue to 51 52 expand and cool. 53 54 Figure 16 shows the results of applying the CTH tool to the case of a solid 55 aluminum plate impacting 60% density ice at 0° obliquity. 56 57 In this simulation, the vapor expands and envelops the plate, leading to 58 59 temperatures of around 6000 K on the back of the plate, with much higher 60 temperatures on the front. The entire plate exceeds 3500 K for 2.5 ms. 61 62 63 18 64 65 1 2 3 4 The above is just a single dense plate. Temperatures rise higher if the 5 6 material is less dense or there are multiple separated plates. 7 8 Hancock explains the higher temperatures for the multiple plate situation 9 associated with a spacecraft as follows: 10 11 12 The impact of the corresponding stack of materials will be much more complex 13 than for a single slab. An impact involving multiple layers will result in residual 14 temperatures which are much higher than for a single layer. Among the reasons 15 are: 16 17 18 • There are multiple shocks and consequently more irreversible shock 19 heating. 20 • The plates may impact rebounding material, thus producing higher 21 22 effective impact velocities and shock strengths than for the incident impact 23 velocity alone. 24 • Plates impacting and recompressing vaporized material will greatly 25 increase the heating of the vaporized material. 26 27 28 Even the last layer to impact will have a higher amount of heating than it 29 would have if it impacted alone. This seems somewhat counter-intuitive if we 30 think about a low speed impact, such as a car crash, in which the first 31 structural elements to impact can have a cushioning effect on the following 32 33 regions. However, at an impact velocity well above the speed of sound in 34 metal, there is no time for the first impacting layers to send stress waves back 35 to slow the trailing layers, so there is no time for any buckling to cushion the 36 later impacts. The final layer probably impacts into rebounding material, 37 increasing its effective impact velocity, and it is probably subjected to many 38 39 shocks from the impacts ahead of it, each one doing irreversible heating. 40 41 A very simple but more realistic model for the entire spacecraft impacting ice 42 is 90% porous aluminum (which approximates Juno’s average density). Figure 43 17 is a simulation of this case. Note that in this simulation, the entire spacecraft 44 45 vaporizes at temperatures above 20,000 K. 46 47 A key conclusion is that bringing a spacecraft fragment to a stop from 20.8 48 km/s with temperature below 773 K is not credible. This would require more than 49 99% of the kinetic energy to be absorbed by the environment and that is not 50 51 considered reasonable for a fragment that is being destroyed. So the 52 temperature is expected to exceed 773 K, but the time above 773 K may be less 53 than the required 0.5 s if the fragment is small enough to cool rapidly. This leaves 54 rapid cooling on a very small particle as one chance for survival. 55 56 Rapid cooling needs a cold environment, away from the (hot) impact site, so 57 58 such survival requires: 59 60  A particle small enough to cool rapidly is created by the initial impact; 61 62 63 19 64 65 1 2 3 4  A microbe survives the destruction of the substrate and is carried by 5 6 the particle; 7  The small particle is ejected at high speed up and away from the 8 impact site so that rapid cooling occurs; 9 10  The particle later impacts at the velocity that gives it the highest 11 possible penetration into the ice. 12 13 14 A small particle cannot penetrate deeply (Possible Survival Case 1), but it might 15 fall into a crack with some finite probability (Possible Survival Case 2). 16 17 For high obliquity, the initial impact does not release most of the kinetic energy, 18 and multiple ricochets might be envisioned that slowly bleed off energy (although 19 20 it is still not credible to expect almost all of this energy to be absorbed by the 21 environment). Since analysis of such multiple ricochets to prove this is extremely 22 complex, an alternate ―fragmentation followed by radiation sterilization‖ analysis 23 is used to address these cases (Possible Survival Case 3). 24 25 26 9.3.1. Possible Survival Case 1: Small Particle Blasted Clear, Cooled, and Buried in Ice 27 28 This case assumes that a particle is heated above 800 K. Temperatures 29 30 above 773 K for > 0.5 s will sterilize the particle, and large particles cannot 31 radiate heat away fast enough to avoid this sterilization. However, if the particle 32 is small enough, it can cool in less than half a second. For these particles, 33 sterilization relies on radiation. The Hancock report assesses the maximum burial 34 depth in 50% density ice for particles of differing materials and mass. The SSB 35 36 report on Forward Contamination of Europa (Space Studies Board, 2000) 37 estimates the radiation dose accumulated under varying depths of ice. 38 Combining the Hancock depths with the NRC radiation dose, and an assumption 39 of sterility after 7 Mrad (Space Studies Board, 2000), gives the results in Table 9. 40 41 42 Note that for such small particles, rapid sterilization is expected following such 43 a shallow burial in the ice. Survival probability: 0%. 44 45 9.3.2. Possible Survival Case 2: Small Particle Blasted Clear, Cools and Falls in deep 46 crevasse 47 48 49 This is the same as the case just discussed, but relies on a small particle 50 following a trajectory to the bottom of an existing crevasse that is so deep that 51 radiation will not sterilize the particle in times sufficient for geological processes 52 to transport surviving organisms to liquid water. Since this possibility seems 53 54 remote, but cannot be excluded, this is included in our analysis with a 55 conservative 1% probability. 56 57 For time scale context, the time to sterilization of a particle buried 1.3 m deep 58 is estimated to be 30,00 years as discussed in the next section and the average 59 60 61 62 63 20 64 65 1 2 3 4 resurfacing time for Europa is estimated to be greater than 20 million years 5 6 (National Research Council, 2012). 7 8 9.3.3. Possible Survival Case 3: Oblique Impact and Fragmentation 9 10 For high obliquity impacts, the magnitude of residual velocity increases with 11 increasing obliquity because less energy is absorbed in the impact. The debris 12 13 center-of-mass velocity greatly exceeds Europa’s escape velocity of 2 km/s in all 14 cases, meaning debris will either return to space or impact other terrain at very 15 high velocity. See Figure 18. 16 17 Higher obliquity impacts provide less heating and lower temperatures from 18 19 the initial impact since the debris can outrun the vapor, and most of the kinetic 20 energy is still with the material. The material may either return to space or later, 21 less oblique impacts may result in the high temperatures required for sterilization. 22 23 Hancock cites Cordelli’s work (Cordelli et al, 1998) on fragmentation of space 24 25 debris to assess the debris that results from even highly oblique impacts at such 26 high velocities. He concludes that fragment size increases with angle of obliquity. 27 This provides an analytic alternative. It is not necessary to analyze the possible 28 multiple collisions that might allow a fragment of a spacecraft to come to rest 29 after an oblique impact. It is sufficient to understand the maximum fragment size 30 31 resulting from the first oblique impact and determine how deep fragments of that 32 size can be buried and how long it will take for them to be sterilized by radiation. 33 Starting from an assumption that any time to radiation sterilization of < 100,000 34 years is acceptably small compared to the time required to transport living 35 36 microbes to liquid water, we find that we can demonstrate sterilization for all 37 particles from oblique impacts up to 80°. The vault wall ends up being the driving 38 case. Table 10 shows several of the most significant fragments after an 80° 39 oblique impact and the time required for radiation sterilization – taken as 7 Mrad 40 if no credit is taken for clean-room assembly (Space Studies Board, 2000). 41 42 43 Conclusion: all fragments from impacts at angles < 80° are sterilized. 44 Although fragments from the 2% of impacts at angles > 80° are most likely also 45 sterilized, the numbers are so few that planetary protection requirements are met 46 even if the project assumes that all such impacts lead to contamination, so they 47 48 are not analyzed further. Note: although one could imagine a nearly horizontal 49 cave that lines up with the debris path following an oblique impact, it is assumed 50 that the likelihood the debris finding its way into that cave and then decelerating 51 through a special combination of multiple bounces—none of which heats the 52 debris above 773 K—is considered to be negligible compared to the 2% already 53 54 allocated to the case of oblique impacts. 55 56 57 9.4. Impact Summary 58 59 60 61 62 63 21 64 65 1 2 3 4 Summing together all instances of possible contamination due to impact 5 6 gives: 7 8  1% possible contamination from impacts at less than the analyzed 9 impact velocity of 20.8 km/s (Section 9.1, 9.3.1); 10 11  1% possible contamination from small particles landing in deep 12 crevasses (Section 9.3.2); 13  2% possible contamination from impacts at angles of obliquity greater 14 than the 80° case analyzed (Section 9.1, 9.3.3). 15 16 17 Note that none of these cases are cases of likely contamination. Rather, they 18 are all cases where the limitations of the analytical tools plus the low likelihood of 19 20 Europa impact allow the project to accept conservatively large estimates and still 21 meet planetary protection requirements. Another project with higher likelihood of 22 Europa impact could reasonably expect to ―sharpen the pencil‖ in this analysis 23 and demonstrate even lower probability of contamination given a non-sterile 24 25 impact. 26 27 10. Capability vs. Requirement 28 29 Juno’s project-level allocation for Europa impact (not contamination) 30 -3 31 probability is <1.5x10 . The approach used to demonstrate compliance with this 32 requirement was to combine the results from the spacecraft reliability model and 33 the Monte Carlo-based trajectory analysis into a single, integrated result. 34 35 Spacecraft reliability was computed over a series of 11-day intervals 36 (corresponding to the period of each MWR and Gravity Science orbit), Figure 19. 37 38 As seen from the figure, spacecraft reliability over each interval varies depending 39 on the activities performed during that interval, driven primarily by the number of 40 ground commands executed by the spacecraft. The relatively low failure 41 probability during the capture orbit (between JOI and PRM) is due to the low level 42 43 of commanding during that period. The higher failure probabilities just prior to the 44 PRM maneuver and during the MWR and Gravity science orbits is due to 45 heightened commanding during those periods. 46 47 The spacecraft reliability results were then combined with the probability of 48 Europa Impact within 400 years assuming the spacecraft failed during a specified 49 50 orbit number (i.e., the spacecraft could no longer be controlled from that point 51 forward), Figure 20. 52 53 The spacecraft reliability and Europa impact results were combined on an 54 orbit-by-orbit basis such that the temporal variations of each could be properly 55 56 accounted for in the composite result, shown in Figure 21. Note that the 57 spacecraft reliability results are integrated over time to show how the overall 58 Europa impact probability varies with mission duration at Jupiter. As shown in the 59 figure, the composite Europa impact probability is 4.6x10–4 at end of mission. 60 61 62 63 22 64 65 1 2 3 4 To demonstrate compliance with NASA’s <1x10–4 requirement of not 5 –4 6 contaminating the Europa ocean, Juno’s Europa impact probability (4.6x10 ) is 7 combined with its ocean contamination probability (<0.04, Section 9), to yield an –5 8 overall probability of <1.8x10 . Thus, Juno is shown to meet NASA’s Europa 9 planetary protection requirement with a large degree of margin (~5x), Table 11. 10 11 12 13 11. Summary 14 15 The Juno science mission does not pass close to Europa or the other 16 Galilean satellites. The mission is designed to conclude with a deorbit burn that 17 18 disposes of the spacecraft in Jupiter’s atmosphere. If that deorbit burn is 19 successful as planned, planetary protection compliance of the Juno mission with 20 respect to Europa is assured and the story is complete. 21 22 23 This paper has included analysis of the finite possibility that the deorbit 24 maneuver does not occur as planned, examining the initial bioburden, the 25 bioburden reduction due to the space and Jovian radiation environments, the 26 likelihood of successful deorbit, the expected orbit propagation with and without a 27 deorbit burn, and the bioburden reduction in the event of impact into an icy body. 28 29 These combined analyses demonstrate that Juno satisfies NASA’s planetary 30 protection requirements. 31 32 Acknowledgements 33 34 35 The authors were members of the development team for the Juno Project. Doug 36 Bernard and Rob Abelson were Project System Engineer and Deputy Project 37 System Engineer, respectively. Jennie Johannesen was the Mission Design 38 Manager and Try Lam was a member of her team. Bill McAlpine was the 39 40 Radiation Manager and Laura Newlin was the Planetary Protection engineer for 41 the Juno Project. 42 43 Stuart Stephens of JPL contributed to the mission and trajectory design portions 44 of this paper including creating Figures 1, 2, and 9 using SOAP software. 45 46 Amanda Briden and Chet Everline of JPL contributed technical content to 47 mission design and probabilistic risk analysis portions of this paper, respectively. 48 Steve Hancock of Foils Engineering helped the Juno team appreciate the 49 physical implications of high-speed impact into an icy body and created Figures 50 14-18. Lockheed Martin is the Juno spacecraft developer and produced Figures 51 52 3 and 4. The research was carried out at the Jet Propulsion Laboratory, 53 California Institute of Technology, under a contract with the National Aeronautics 54 and Space Administration. 55 56 57 58 59 References 60 61 62 63 23 64 65 1 2 3 4 Cordelli, A., P. Farinella, and A. 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E., et al., ―Meeting the Planetary Protection Requirement for 22 JIMO/Europa‖, presentation to JIMO SDT, May 13, 2004, California. 23 24 25 La Duc, M., Dekas, A., Osman, S., Moissl, C., Newcombe, D., and 26 Venkateswaran, K., ―Isolation and Characterization of Bacteria Capable of 27 Tolerating the Extreme Conditions of Clean Room Environments,‖ Applied and 28 Environmental Microbiology, Vol. 73, No. 8, Apr. 2007, p. 2600–2611. 29 30 Lam, T., Johannesen. J. R., and Kowalkowski, T. D., "Planetary Protection 31 32 Trajectory Analysis for the Juno Mission", AIAA/AAS Astrodynamics Specialist 33 Conference, AIAA-2008-7368, Aug. 2008. 34 35 Li, S. W., et al., ―Isolation and Characterization of Ionizing Radiation-Resistant 36 Soil Bacteria from the Hanford Site, Washington,‖ 107th General Meeting of the 37 38 American Society for Microbiology, ASM Press, Toronto, Canada, May 21-25, 39 2007. 40 41 Matousek, S.W., ―The Juno New Frontiers Mission,‖ International Astronautical 42 Congress, Fukuoka, Japan, IAC-05-A3.2.A.04, October 17-21, 2005 43 44 45 Mileikowsky, Curt, Cucinotta, Francis A. Wilson, John W., Gladman, Brett, 46 Horneck, Gerda, et al., "Risks threatening viable transfer of microbes between 47 bodies of our solar system," Planetary and Space Science, Vol. 48, Issue 11, 48 September, 2000. 49 50 51 National Aeronautics and Space Administration, Science Mission Directorate, 52 Biological Contamination Control for Outbound and Inbound Planetary 53 Spacecraft, NPD 8020.7F, Washington, D.C., February 19, 1999 (Revalidated 54 10/23/03). 55 56 National Aeronautics and Space Administration, Science Mission Directorate, 57 58 Planetary Protection Provisions for Robotic Extraterrestrial Missions, NPR 59 8020.12C, Washington, D.C., April 27, 2005. 60 61 62 63 24 64 65 1 2 3 4 National Aeronautics and Space Administration, Science Mission Directorate, 5 6 NASA Handbook for the Microbiological Examination of Space Hardware, NASA- 7 HDBK-6022, Washington, D.C. (release pending). 8 9 National Research Council, Assessment of Planetary Protection Requirements 10 for Spacecraft Missions to Icy Solar System Bodies, The National Academies 11 Press, 2012. 12 13 14 Newlin, L. E., Juno Project Planetary Protection Plan, Rev. A, JPL D-34003, May, 15 2008. 16 17 Price, P. Buford, and Sowers, Todd, ―Temperature Dependence of Metabolic 18 Rates for Microbial Growth, Maintenance, and Survival,‖ Proceedings of the 19 20 National Academy of Sciences of the United States of America, Vol. 101, No. 13, 21 pp. 4631-4636, March 30, 2004. 22 23 Saffary, Roya, et al., "Microbial Survival of Space Vacuum and Extreme 24 Ultraviolet Irradiation: Strain Isolation and Analysis During a Rocket Flight," 25 26 FEMS Microbiology Letters, Vol. 215, Issue 1, pp. 163-168, September, 2002. 27 28 Space Studies Board, National Research Council, Preventing the Forward 29 Contamination of Europa, National Academy Press, Washington, D.C., 2000. 30 31 Urgiles, E., Wilcox, J., Montes, O., Osman, S., Venkateswaran, K., et al., 32 33 "Electron Beam Irradiation for Microbial Reduction on Spacecraft Components," 34 2007 IEEE Aerospace Conference, Big Sky, Montana, March, 2007. 35 36 37 38 Figure Captions 39 40 Figure 1. Interplanetary Trajectory 41 Figure 2. Juno Orbits of Jupiter Shown in Earth-to-Jupiter and Jupiter North Pole 42 43 Views 44 Figure 3. Juno Spacecraft Post Solar Array Deployment 45 46 Figure 4. Close-up View of Juno Flight System Core 47 48 Figure 5. Satellite Range at Closest Approaches 49 50 Figure 6. Apsidal Rotation 51 Figure 7. Impact Probability Evolution Over Time Following Failed Deorbit 52 53 Attempt 54 Figure 8. Europa Impact Probability Assuming S/C Failure After Each Orbit 55 56 Figure 9. Juno Orbit Trajectory and the Jovian Magnetically Trapped Radiation 57 Belts. 58 59 Figure 10. The Total Ionizing Dose Absorbed in Water per 12-Year Period. 60 61 62 63 25 64 65 1 2 3 4 Figure 11. NOVICE Model of a Juno Electronics Unit. 5 6 7 8 Figure 12. Impact Velocity 9 10 Figure 13. Obliquity 11 Figure 14. Shock Strength for Ice Impact 12 13 Figure 15. Shock, Expansion, and Temperature 14 15 Figure 16. Normal Impact of a Solid Aluminum Plate 16 17 Figure 17. Normal Impact of a Porous Aluminum Plate

18 Figure 18. Oblique Impacts 19 20 Figure 19. Juno Spacecraft Failure Probability per Orbit Number 21 22 Figure 20. Europa Impact Probability as a Function of Orbit Number 23 24 Figure 21. Composite Europa Impact Probability Versus Mission Duration 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 26 64 65 Table

Table 1: Average Impact Probability Assuming S/C Failure before Deorbit

% Occurrence over % Occurrence over Impact type 150 years 400 years No Impacts 24.1% +/– 1.24% 13.2% Jupiter impacts 72.9% +/– 1.29% 80.8% Io impacts 1.75% +/– 0.40% 3.52% Europa impacts 0.52% +/– 0.22% 1.06% Ganymede impacts 0.55% +/– 0.22% 1.10% Callisto impacts 0.18% +/– 0.12% 0.35%

Table 2. Average Annual Dose in Water for most Heavily Shielded Juno Equipment.

Shielding Average Annual Unit Name (mils equiv. Dose (krad Al) water/yr)

Stellar Reference Unit Opt. Head 3300 17.7

Inertial Measurement Unit A 2400 35.6

Inertial Measurement Unit B 2280 39.9

Stellar Reference Unit Elect. A 2200 43.3

Stellar Reference Unit Elect. B 2190 43.7

Small Deep Space Transponder B 2170 44.6

Command and Data Handling Unit A 2140 46.1

Command and Data Handling Unit B 2030 51.9

Small Deep Space Transponder A 2020 52.5

Power Distribution and Drive Unit 2020 52.5

Table 3: Summary of Micro-organism Types Proposed by SSB Europa Report Population Fractions Type Definition SSB JIMO / Juno Culturable using the NASA Standard A Assay TSA plating technique (NASA, – – pending) B Spore-forming bacteria 2% of A ~ 10% of A

Radiation resistant spores, with ≥ 10% C ~ 0.1% of B ~ 10% of B survival above 0.8 Mrad Radiation resistant non-spore-forming D bacteria, with ≥ 10% survival above 4.0 ~ 0.1% of A ~ 2% of A Mrad

Table 4: Summary of Micro-organism Groups Proposed by JIMO and Juno Relation to Group Definition Curve Source Types Radiation sensitive non- 1 A–B–C Kohlhase, 2004 spore-forming bacteria* Radiation sensitive spore- 2 B–C Kohlhase, 2004 forming bacteria* Radiation resistant spore- B. megaterium 3 C forming bacteria* Urgiles, 2007 Radiation resistant non- Isolate 3B1 4 D spore-forming bacteria* Li, 2007 * culturable using the NASA Standard Assay TSA plating technique

Table 5: Initial Microorganism Population Densities by Type Structure Interior & module exterior Area Electronics Structure Mixed ( / m2) populated Non- Electronics Non- Class board metal Non-metal metal Fluids Organism 100K Uncontrolled Area Volume Volume Volume Volume Type Cleanroom Manufacture ( / m2) ( / cm3) ( / cm3) ( / cm3) ( / cm3)

Culturable A 1E+05 1E+06 3E+04 150 750 650 50 B 1E+04 1E+05 3000 30 150 130 10 C 1000 10000 300 3 15 13 1 D 2000 20000 600 3 15 13 1

Total at Launch = Culturable + Non-Culturable A 1E+08 1E+09 3E+07 1.5E+05 7.5E+05 6.5E+05 5.0E+04 B 1E+07 1E+08 3E+06 3E+04 1.5E+05 1.3E+05 1.0E+04 C 1E+06 1E+07 3E+05 3000 1.5E+04 1.3E+04 1000 D 2E+06 2E+07 6E+05 3000 1.5E+04 1.3E+04 1000

Total at JOI (Cruise reduction) A 1E+07 1E+08 3E+06 1.5E+04 7.5E+04 6.5E+04 5E+03 B 1E+07 1E+08 3E+06 3E+04 1.5E+05 1.3E+05 1E+04 C 1E+06 1E+07 3E+05 3000 1.5E+04 1.3E+04 1000 D 1E+06 1E+07 3E+05 1500 7.5E+03 6.5E+03 500

Table 6: Microorganism Survival Fraction as a Function of Years After Launch Years After Launch N / No 0 25 50 75 100 125 150 Group 1 1 6.1E-03 3.8E-05 2.3E-07 1.4E-09 8.7E-12 5.3E-14 SRU Group 2 1 3.4E-02 1.1E-03 3.8E-05 1.3E-06 4.2E-08 1.4E-09 Optical Group 3 1 5.3E-02 2.8E-03 1.5E-04 7.7E-06 4.0E-07 2.1E-08 Heads Group 4 1 6.2E-01 3.8E-01 2.3E-01 1.4E-01 8.8E-02 5.4E-02 Group 1 1 3.6E-05 1.3E-09 4.5E-14 1.6E-18 5.6E-23 2.0E-27 Group 2 1 1.1E-03 1.2E-06 1.3E-09 1.4E-12 1.5E-15 1.6E-18 IMUs Group 3 1 2.7E-03 7.2E-06 1.9E-08 5.1E-11 1.4E-13 3.7E-16 Group 4 1 3.8E-01 1.4E-01 5.4E-02 2.0E-02 7.6E-03 2.9E-03

Group 1 1 3.9E-06 1.5E-11 5.8E-17 2.2E-22 8.7E-28 3.4E-33 SRU Group 2 1 2.5E-04 6.1E-08 1.5E-11 3.7E-15 9.1E-19 2.2E-22 Electronics Group 3 1 7.4E-04 5.5E-07 4.1E-10 3.1E-13 2.3E-16 1.7E-19 Modules Group 4 1 3.1E-01 9.3E-02 2.8E-02 8.7E-03 2.7E-03 8.1E-04 Group 1 1 4.9E-06 2.5E-08 1.6E-10 9.5E-13 5.8E-15 3.6E-17 Combined Group 2 1 4.4E-04 5.0E-06 1.6E-07 5.5E-09 1.8E-10 6.1E-12 Total Group 3 1 1.1E-03 1.3E-05 6.3E-07 3.3E-08 1.8E-09 9.2E-11 Group 4 1 2.3E-05 1.2E-07 7.3E-10 4.5E-12 2.7E-14 1.7E-16

Table 7: Surviving Microorganisms as a Function of Years After Launch Years After Launch N 0 25 50 75 100 125 150 Group 1 1.1E+07 6.8E+03 4.2E+01 2.6E-01 1.6E-03 <0.001 <0.001 SRU Group 2 1.4E+06 4.6E+04 1.5E+03 5.1E+01 1.7E+00 5.7E-02 1.9E-03 Optical Group 3 1.5E+05 7.9E+03 4.2E+02 2.2E+01 1.2E+00 6.1E-02 3.2E-03 Heads Group 4 2.6E+05 7.9E+02 4.8E+00 3.0E-02 <0.001 <0.001 6.8E-09 Total 1.3E+07 6.1E+04 2.0E+03 7.3E+01 2.9E+00 1.2E-01 5.1E-03 Group 1 3.3E+08 7.0E+02 2.1E-02 7.0E-07 <0.001 <0.001 <0.001 Group 2 3.7E+07 2.7E+04 2.4E+01 2.3E-02 <0.001 <0.001 <0.001 IMUs Group 3 4.1E+06 7.6E+03 1.7E+01 4.1E-02 <0.001 <0.001 <0.001

Group 4 7.6E+06 8.0E+01 2.4E-03 8.0E-08 <0.001 <0.001 <0.001 Total 3.8E+08 3.5E+04 4.1E+01 6.4E-02 <0.001 <0.001 <0.001 Group 1 1.3E+09 4.7E+02 1.7E-03 6.4E-09 <0.001 <0.001 <0.001 SRU Group 2 2.8E+08 6.5E+04 1.5E+01 3.7E-03 <0.001 <0.001 <0.001 Elec- Group 3 3.1E+07 2.2E+04 1.6E+01 1.1E-02 <0.001 <0.001 <0.001 tronics Group 4 3.3E+07 6.0E+01 2.2E-04 8.0E-10 <0.001 <0.001 <0.001 Modules Total 1.6E+09 8.7E+04 3.1E+01 1.5E-02 <0.001 <0.001 <0.001 Group 1 1.7E+09 8.0E+03 4.2E+01 2.6E-01 1.6E-03 <0.001 <0.001 Com- Group 2 3.1E+08 1.4E+05 1.6E+03 5.1E+01 1.7E+00 5.7E-02 1.9E-03 bined Group 3 3.5E+07 3.7E+04 4.5E+02 2.2E+01 1.2E+00 6.1E-02 3.2E-03 Total Group 4 4.1E+07 9.3E+02 4.8E+00 3.0E-02 <0.001 <0.001 <0.001 Total 2.0E+09 1.8E+05 2.1E+03 7.3E+01 2.9E+00 1.2E-01 5.1E-03

Table 8: Assumed Sterilized Conditions

 Temperature exceeds 500 C (773 K) for at least 0.5 s  Substrate vaporized or liquefied by hypervelocity shock  Radiation dose exceeds 7 Mrad  Fragmented and escapes Europa orbit (negligible protection from radiation in Jupiter orbit)

Table 9: Burial of a Small Particle Maximum Time to 7 Max non- Mrad in Material burial sterilized Europa depth size ice Tungsten 0.4 mm 5 cm 30 years Titanium 0.2 mm 1 cm 2 years

Table 10: Burial of a Spacecraft Fragment

Largest solid Diameter of Max Time to 7 Material Dose rate at piece prior largest fragment burial Mrad in (density) depth to impact after impact depth Europa ice

Titanium 34 kg 4 cm 130 cm 20 rad/mo 30,000 years (4.5 g/cc) (vault wall) Tungsten 1.0 kg (JEDI 0.8 cm 110 cm 50 rad/mo 12,000 years (19 g/cc) shield wall)

Aluminum 48 kg 5.5 cm 110 cm 50 rad/mo 12,000 years (2.7 g/cc) (top deck)

Table 11. Juno Project-Level Planetary Protection Requirements

Current Sub- Probability suballocations Best Allocation Estimate  Spacecraft failure analysis: Probability that a spacecraft failure prevents it from being able to 0.05 <0.10 deorbit by EOM  Mission Design: Probability of Europa impact of a 0.0093 <0.015 non-sterile spacecraft (assumes no deorbit burn)  Spacecraft/Europa Impact Analysis: Probability of <0.04 <0.06 contamination in the event of a non-sterile impact

Combined Probability of Contamination of the Europan <1.8x10–5 <9x10–5 Ocean NASA Requirement <1x10–4

Figure

Figure 1. Interplanetary Trajectory

Figure 2. Juno Orbits of Jupiter Shown in Earth-to-Jupiter and Jupiter North Pole Views

Figure 3. Juno Spacecraft Post Solar Array Deployment

Figure 4. Close-up View of Juno Flight System Core

Figure 5. Satellite Range at Closest Approaches

Figure 6. Apsidal Rotation

Figure 7. Impact Probability Evolution Over Time Following Failed Deorbit Attempt

Figure 8. Europa Impact Probability Assuming S/C Failure After Each Orbit

Figure 9. Juno Orbit Trajectory and the Jovian Magnetically Trapped Radiation Belts.

Figure 10. The Total Ionizing Dose Absorbed in Water per 12-Year Period.

Figure 11. NOVICE Model of a Juno Electronics Unit.

Figure 12. Impact Velocity

Figure 13. Obliquity

Figure 14. Shock Strength for Ice Impact

Figure 15. Shock, Expansion, and Temperature

Figure 16. Normal Impact of a Solid Aluminum Plate

Figure 17. Normal Impact of a Porous Aluminum Plate

Figure 18. Oblique Impacts

Probability of S/C Failure within the Indicated Orbit Number

0.17% Capture Orbit PRM MWR Gravity Orbits Deorbit 0.16%

0.15%

0.14%

0.13%

0.12%

0.11%

Probability of S/C Failure(%) 0.10%

0.09%

0.08%

Orbit No. Figure 19. Juno Spacecraft Failure Probability per Orbit Number

Probability of Europa Impact Assuming the S/C Fails During the Indicated Orbit Number 2.5% Propagated for 4000 yrs Propagated for 400 yrs

2.0%

1.5%

(%) 1.0%

0.5%

Probability Probability of Europa Impact Given Failed a Spacecraft, 0.0% 0 0b 0d 0f 0h 1 3 5 7 9 11 13 15 17 19 21 23 25 27 29 31 33 Orbit Number (Measured from PJn to PJn+1)

Figure 20. Europa Impact Probability as a Function of Orbit Number

Integrated Composite Europa Impact Probability Accounting for Spacecraft Reliability 0.050% Composite Europa Impact Probability at EOM = 4.6E-4 0.045%

0.040%

0.035%

0.030%

0.025%

0.020%

0.015%

0.010% Integrated Probability of Europa Impact, (%) Impact, Europaof Probability Integrated 0.005%

0.000% 0 0b 0d 0f 0h 1 3 5 7 9 11 13 15 17 19 21 23 25 27 29 31 33 Orbit Number (Measured from PJn to PJn+1) Figure 21. Composite Europa Impact Probability Versus Mission Duration Highlights

 The Juno Mission to Jupiter meets planetary protection requirements by design and analysis

 The nominal mission is designed to be concluded with a deorbit maneuver leading to disposal of the spacecraft in Jupiter’s atmosphere

 Juno’s orbit is never on an impact trajectory with Europa and the most likely end-state even without a deorbit maneuver is disposal in Jupiter’s atmosphere

 Analyses include initial bioburden at launch and bioburden reduction due to space and Jovian environments including temperature, vacuum, and (most importantly) radiation

 Analysis of hypothetical impacts between the Juno spacecraft and Europa shows a high likelihood of sterilization by the impact event