Optimization (MALTO) tool.1 The electric propulsion margins and assumptions have significant ... The broad search included launch opportunities ... high specific impulse engines was identified as the best option for all launch years considered. ..... dynamics Specialist Conference and Exhibit, Mackinac Island, MI, Aug 2007.
AAS 11-188
CONCEPTUAL MISSION DESIGN OF A POLAR URANUS ORBITER AND SATELLITE TOUR James McAdams,* Christopher Scott,† Yanping Guo,§ John Dankanich2 and Ryan Russell‡ In response to NASA’s planetary science decadal survey, this paper outlines the conceptual mission design of a Uranus orbiter. In the baseline design the spacecraft launches during a 21-day launch period in 2020, followed by a 13-year cruise with solar electric propulsion and a single Earth flyby. Repeatable launch opportunities are available from 2021-2023. An atmospheric probe is released 29 days prior to Uranus orbit insertion. After completion of the probe descent phase the spacecraft inserts into a highly inclined elliptical orbit for 431 days, followed by the satellite tour with targeted flybys of five satellites.
INTRODUCTION This conceptual mission design was part of a broad system-level effort to determine whether a scientifically compelling mission to Uranus could be flown within a New Frontiers or “subFlagship” classification launching between 2020 and 2023. Primary objectives include determining Uranus’ bulk composition, internal structure, and source of the magnetic field, and identifying the mechanisms responsible for internal heat transport to the surface. General investigations of the atmosphere, satellites, and rings are secondary objectives. This study was managed by The Johns Hopkins University Applied Physics Laboratory, with key direction from a science steering committee composed of outer planet experts from NASA centers and universities. These objectives translate into mission design guidelines and constraints. The spacecraft must perform at least 20 complete Uranus orbits with a large inclination and eccentricity to sample a broad range of distances, latitudes, and longitudes. The radius of periapse must be approximately 1.3 Uranus radii while avoiding the ring system. The satellite tour targets Miranda, Ariel, Umbriel, Titania, and Oberon while maintaining flyby speeds conducive to measurement objectives. Study guidelines prohibit the use of a Jupiter gravity assist, which eliminates the feasibility of high-thrust chemical propulsion with multiple planetary flybys, set the maximum cruise duration to 13 years, and requires any Earth flyby to be at or above 1,000 km altitude.
*
Mission Design Lead Engineer, The Johns Hopkins University Applied Physics Laboratory, 11100 Johns Hopkins Rd, Laurel MD, 20723. † Mission Design Analyst, The Johns Hopkins University Applied Physics Laboratory, 11100 Johns Hopkins Rd, Laurel MD, 20723. § Mission Design Lead Engineer, The Johns Hopkins University Applied Physics Laboratory, 11100 Johns Hopkins Rd, Laurel MD, 20723. Ľ Mission Analyst, Gray Research, Inc, 21000 Brookpark Rd. M/S 77-4, Cleveland, OH, 44135. ‡ Assistant Professor, Georgia Institute of Technology, Georgia Institute of Technology, Atlanta, Georgia, 30332.
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Trades conducted for each Uranus orbiter mission phase yielded recommendations for launch vehicle, launch period, cruise duration and propulsive mode, atmospheric probe deployment and targeting, Uranus orbit insertion (UOI), primary science orbit definition, and the secondary satellite tour. Primary considerations in conducting these trades include achievement of science objectives, delivered payload mass, launch opportunity repeatability, total mission duration, spacecraft safety (e.g., Uranus ring avoidance), and satellite encounter variety. This paper provides concise overviews of selected trades, as well as summarizes key elements of the recommended Uranus orbiter baseline mission. JOURNEY TO URANUS Definition of an optimal interplanetary cruise trajectory to Uranus requires careful selection of a configuration that includes a capable launch vehicle, propulsion system, planetary flyby number and altitude, and trip duration. The New Frontier or sub-Flagship mission class designation led to consideration of the Atlas V series expendable launch vehicle. Propulsion system selection combines the every-launch-year flexibility of solar electric propulsion (SEP) with the high-thrust capability needed from a chemical propulsion system for UOI and subsequent maneuvers. One Earth flyby at the minimum allowable 1,000-km altitude supplements the onboard high-'V available via SEP. Uranus arrival velocity and therefore orbit insertion 'V are minimized by using a 13-year duration launch to Uranus arrival. Propulsion Trades The electric propulsion mission trades were conducted using the Mission Analysis Low-Thrust Optimization (MALTO) tool.1 The electric propulsion margins and assumptions have significant effects on the results. The baseline assumptions are provided in Table 1 unless otherwise stated. The SEP thrusters assume a thruster model that can throttle from minimum to maximum power while changing thrust and specific impulse to coincide with the demonstrated performance of the thruster. For radioisotope powered electric propulsion trades, the thruster is allowed to operate at the optimal specific impulse for the specified thruster efficiency. Table 1. Thruster and optimization assumptions. Solar Electric Propulsion
Radioisotope Electric Propulsion
Power, kW* 20 0.7 Housekeeping Power, kW 0.0 0.0 Thruster Efficiency, % NEXT 55% Specific Impulse, s NEXT Optimized Duty Cycle, % 90% 90% Solar Array Model Ultraflex NA Number of Thrusters 2 1 Launch Vehicle Atlas 551 Atlas 551 w/ Star 48 * Solar power specified at 1 AU, radioisotope power is constant throughout the mission
As was shown by Landau, Lam and Strange2, the combination of solar electric propulsion and gravity assists enable missions with larger payloads than those with chemical propulsion over a broad range of flight times and power levels. The broad search included launch opportunities between 2018 and 2030 for 10-year transfer times using the NASA's Evolutionary Xenon Thruster (NEXT) 2+1 SEP stage with 15 kW of power as proposed for the Titan Saturn System Mission.3 This JPL study indicates that significant benefits in delivered mass to Uranus come from utilizing a Jupiter flyby for launch dates from 2018 to 2020. Since the launch years for the NASA
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Ice Giants Decadal Survey include 2020 to 2023, reliance on a Jupiter gravity-assist flyby had to be removed. For launch dates from 2021 to 2023, when Jupiter gravity assists are no longer helpful, Saturn flybys will provide greater delivered mass capabilities. However, for a maximum 13year duration trip to Uranus and launches from 2020 to 2023, a Saturn flyby effectively increases Uranus arrival velocity too much. Inclusion of only one Earth flyby with a pair of NEXT P10 high specific impulse engines was identified as the best option for all launch years considered. Trades on the number of NEXT thrusters and array power level for 13-year fixed cruise duration showed that performance (measured by mass after UOI) increases significantly with more power and increases minimally with an extra thruster. Table 2 demonstrates this observation for a hybrid SEP\chemical propulsion system chemical propellant for UOI and the satellite tour. Table 2. Thruster number and array power effects on mass to Uranus orbit. System
LaunchWet Mass,kg
SEPStage Mass,kg
EPPropellant, kg
2NEXT,15kW 2NEXT,17kW 2NEXT,19kW 2NEXT,21kW 2NEXT,23kW 2NEXT,25kW 3NEXT,15kW 3NEXT,17kW 3NEXT,19kW 3NEXT,21kW 3NEXT,23kW 3NEXT,25kW
4133 4491 4801 5008 5169 5275 4281 4832 5206 5428 5523 5600
900 940 980 1020 1060 1100 984 1024 1064 1104 1144 1184
749 808 890 953 1026 890 760 866 977 1060 1075 1096
ArrivalVь, km/s 7.20 7.15 7.15 7.17 7.19 7.21 7.16 7.16 7.20 7.21 7.22 7.22
UOIȴV,km/s
Chemical Propellant,kg
Estimated Mass after UOI, kg
1.54 1.53 1.53 1.53 1.54 1.55 1.53 1.53 1.54 1.55 1.55 1.55
973 1065 1139 1184 1209 1286 987 1144 1241 1283 1302 1308
1511 1678 1792 1850 1873 1999 1550 1798 1924 1980 2002 2011
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Launch Opportunities Limiting planetary gravity assist options to a single Earth encounter provides 21-day period launch opportunities each year from 2020 to 2023. Figure 1 provides an example Uranus orbiter trajectory profile at the start of the 2020 21-day launch period, where launch energy is highest.
Figure 1. Ecliptic plane view of 13-year duration Uranus orbiter heliocentric trajectory.
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A timeline showing cruise-phase thrust and coast periods (see Figure 2) reveals that SEP thrust periods occur only when the spacecraft is within a few AU from the Sun. After about 2000 days after launch, when array power is limited due to increasing solar distance, the spacecraft coasts with only infrequent small course-correction maneuvers using the chemical propulsion system. The timeline in Figure 2 has seven coast periods, including 30 days for spacecraft check out after launch, 42 days prior to the Earth gravity-assist flyby, and five other coast periods that last 20 to 30 days each.
Figure 2. Cruise Phase timeline showing Uranus orbiter SEP thrust and coast times. An opportunity analysis for consecutive year launch opportunities shows a gradual trend of increased delivered spacecraft mass to Uranus orbit (see Table 3 for 20-kW, 3 NEXT thruster, 13year cruise results). This improvement comes primarily from a decrease of about 0.06 AU/year in the Uranus-Sun distance at Uranus arrival. Table 3 also shows the significant performance improvement is available if the spacecraft arrives with a higher arrival velocity and relies on aerocapture at Uranus orbit insertion. However, aerocapture presents additional technical challenges including ring avoidance constraints for the Uranus arrival geometry in the 2033 to 2036 time frame. Table 3. Thruster number and array power effects on mass to Uranus orbit. Comment
LaunchWet Mass,kg
SEPStage Mass,kg
EPPropellant, kg
13yrͲAero 13yrͲ2020 13yrͲ2021 13yrͲ2022 13yrͲ2023
6269 5260 5313 5340 5322
800 800 625 625 625
430 853 909 926 910
ArrivalVь, km/s 12.0 7.3 7.2 7.2 7.2
UOIȴV,km/s
Chemical Propellant,kg
Estimated Mass after UOI, kg
Aerocapture 1.57 1.56 1.55 1.54
NA 1393 1449 1450 1439
3360 2214 2330 2339 2349
Sensitivity to Earth Flyby Altitude and Cruise Phase Duration Trades were performed to assess the effect of Earth flyby altitude and cruise-phase trip time on spacecraft mass delivered to Uranus orbit. While Earth flyby minimum altitudes of less than 1,000 km offer significant performance improvement, 1,000 km is the established minimum for this study. Analysis revealed that changes in mass performance become more gradual as Earth flyby altitude increases. For example, the change in mass after UOI is about the same between 1,000 km and 2,000 km as it is between 2,000 km and 4,200 km. Increasing the cruise-phase transfer time from 10 years to the maximum allowable 13 years provides continual improvement in spacecraft mass delivered to Uranus orbit. This near linear relationship is a result of the decreasing Uranus arrival velocity, and therefore UOI 'V, with increasing heliocentric transfer time.
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ARRIVAL AT URANUS Mission activities from atmospheric probe release through Uranus orbit insertion mark the transition from cruise phase to orbit phase. Uranus arrival requirements include: 1) both the probe and orbiter trajectories must avoid crossing the Uranian ring plane at an altitude below that of the most distant prominent “epsilon” ring (2.017 Uranus radii) 2) orbiter trajectories must avoid crossing the ring plane within the “nu” and “G” rings (2.582 to 2.739 Uranus radii) 3) maintain a line-of-sight probe-to-orbiter relay communications link during the probe’s atmospheric descent 4) probe deceleration must never exceed 400 times g's 5) preference for gravity field measurements to be able to observe periapsis from Earth 6) periapsis radius as close as possible to 1.1 Uranus radii 7) 20 primary science orbits providing repeated coverage of the same region at ~2-hour intervals for cloud tracking For a mid-2033 Uranus arrival date, Figure 3 reveals that Uranus’ north pole is pointed near the direction of Earth and the incoming spacecraft trajectory direction is such that part of the trajectory near periapsis will not be visible from Earth (an unavoidable partial violation of requirement #5). The approach trajectory will also support probe entry into Uranus’ atmosphere before encountering the ring plane.
Figure 3. Uranus orbiter arrival geometry in August 2033. Probe Release and Atmospheric Entry In order to maximize compliance with Uranus arrival constraints, several trades were conducted to characterize operational and system requirements for the atmospheric entry probe. After the late-cruise checkout of the probe system functionality, a small TCM will target the orbiter/probe spacecraft toward the desired probe entry point. Prior to releasing the probe, the spacecraft will spin up to a determined spin rate needed to attain the desired probe attitude at atmospheric entry. The orbiter and entry probe separate 29 days prior to the probe’s atmosphere entry.
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One day after orbiter-probe separation the orbiter completes an orbit deflection maneuver (ODM) of about 30 m/s to target the spacecraft to the UOI B-plane aim point. As seen in Figure 4, the ODM places the orbiter behind the probe, such that the orbiter arrives at Uranus slightly later to be in position to relay probe data to Earth-based tracking antennas.
Figure 4. Uranus orbiter and probe arrival trajectories on June 28, 2033. Remaining atmospheric probe analyses are centered on the journey from atmosphere entry through end of data transmission. Probe mechanical assumptions included a 127 kg mass, a 1.05 drag coefficient, and a 0.45 m2 area in the velocity direction. The Uranus atmosphere model4 starts at 550 km altitude and rotates with the planet with zero relative wind speed. A reference point of 5 bars atmospheric pressure at approximately -61 km altitude marks the point beyond which reliable data transmission may not be possible. A drogue parachute opens less than one minute after the probe descends past 550 km altitude. The parachute area selection trade (see Figure 5) sought a 60-minute total descent5, with 50 minutes of this descent after parachute deployment at Mach 0.9 to Mach 1.0. For a parachute coefficient of drag of 0.55 and 51 minutes from parachute deployment to 5 bar pressure, and an 8.13-m2 area, the diameter of the parachute was chosen to be 3.25 m. Two computations of peak deceleration revealed 372 and 389 g’s, both below the 400-g limit. These acceleration results compare well with a previous study with a Uranus atmosphere probe6.
Figure 5. Uranus atmospheric probe parachute area size trade (B = ballistic coefficient).
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Orbit Insertion While the requirement to avoid Uranus’ rings and the need to place Uranus orbit insertion near periapsis prevent all study requirements from being met, the June 28, 2033 UOI design chosen lowers risk by placing the start and end of UOI within the direct line of sight of Earth. As shown earlier in Figure 4, the UOI maneuver starts about 62 minutes after completion of the probe communications link, with most of UOI not visible from Earth. This hour-long delay allows sufficient time for the orbiter to slew to the UOI burn attitude and complete preparation for the 67-minute, 1661-m/s UOI maneuver. The UOI maneuver assumes a 150-lb thrust engine with 332-s specific impulse. The ultimate purpose of UOI is accomplished by placing the spacecraft into a 1.3 RU (33,425-km altitude) periapsis by 21.0-day Uranus orbit with inclination of 97.7°. Initial periapsis longitude and latitude are 12.5° and -72.6°, respectively. ORBIT PHASE AND SATELLITE TOUR Most of the mission’s primary science objectives will be achieved during the 20.5-orbit, 431day baseline orbit phase that begins after UOI and ends with the start of the optional satellite tour. During the baseline orbit phase, data collected and transmitted to Earth will greatly increase understanding of the atmosphere, magnetic field, and the gravity field of Uranus.
Figure 6. Uranus orbiter primary science and satellite tour trajectories relative to selected satellite orbits. Primary Science Orbit Analysis of the first 20 integrated orbits reveals a slowly changing orbit. Two small, but notable changes affecting gravity determination include a 1,150 km increase in periapsis altitude and a 2.7° reduction in argument of periapse (due primarily to the large Uranus oblateness). Figure 6
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shows in magenta the 20.5 orbits after UOI and a representative satellite tour in dark blue. Orbits of the major satellites of Uranus appear in red, with the orbiter’s heliocentric approach trajectory in light green. At the end of the science orbit phase, and before the satellite encounters, lowering periapsis radius from 1.345 RU to either 1.025 RU or 1.041 RU (25,600 km or 26,000 km) enables significant improvement in determining the Uranus gravity field J6 term. Reducing periapse radius by 25,600 km requires an apoapse 'V of 56 m/s. Analyses of spacecraft heating and drag, assuming a 1500 kg spacecraft with 4.7 m2 area in the velocity direction, reveals unacceptably high heating rates – exceeding 0.007 W/cm2. Satellite Tour Options Lasting 424 days, the baseline tour has ten targeted flybys passing by Miranda (body ID 705), Ariel (ID 701), Umbriel (ID 702), Titania (ID 703), and Oberon (ID 704) twice each and four additional close untargeted flybys with Umbriel. In most cases, nearby satellite:spacecraft resonances were targeted for orbits between repeat flybys of the same satellite (16:1 for Miranda, 10:1 for Ariel, 6:1 for Umbriel, 3:1 for Titania, and 2:1 for Oberon). Many small maneuvers not shown in the satellite tour summary (see Table 4) are required in the design because precise resonances are absent in a full ephemeris model. Consecutive maneuvers indicate a two-impulse sequence that guarantees phasing in about two spacecraft orbits. The first maneuver is targets the ring plane crossing, while the second maneuver is at the ring plane and adjusts the period to ensure proper phasing. The two-impulse solutions are less favorable for Titania and Oberon because their periods are much larger than the other moons, providing for fewer targeting opportunities. Therefore, a single-impulse solution is used at Titania and Oberon. Note the italicized labels in Table 4 indicate a very long flight time after the maneuver that achieves the Titania transfer. This extended time was considered acceptable compared to the ~50 m/s extra required for the minimum time solution. Maneuvers at flybys occur at the sphere of influence. Large maneuvers target the next moon. The preliminary tour design is based on the zero-radius sphere of influence patched-conic model with satellite ephemeris locations provided by ‘ura083.bsp’. The total 'V for the tour is 619 m/s. The tour was designed in part using the graphical methods based on V globe maps that reduce feasible options for post flyby orbits onto maps with contours of desired quantities. Figure 7 shows an example globe map of the final flyby of the tour (at Oberon), showing contours of ring plane crossing distances associated with post flyby orbits. The tip of the pre-flyby V vector location denoted by an 'x' and the circle represents the possible V vector locations after a ballistic 50 km flyby. Locations outside the circle require flybys with altitudes less than 50 km and are therefore unreachable. In order to return to Oberon on a 2:1 resonance, there are two options as denoted by the intersections of the circle with the 2:1 resonance band. Successive encounters with a single body are achieved by targeting the resonant bands while encounters with new moons are enabled by targeting intersections of the 50 km flyby circle with the ring plane crossing corresponding to the targeted moons' orbital radius. (Note that a later maneuver and multiple revolutions then lead to correct phasing.) Successive resonant free-returns can be used to reduce delta-v requirements to reach the next moon, although the efficiency is low due to the low mass of all the moons. Instead, in this tour we favored short flight times and used only one resonant return for each moon in order to slightly reduce delta-v but more importantly achieve a second flyby to enhance the science return.
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Table 4. Satellite flyby and maneuver dates for the baseline satellite tour. JulianDate 2464214.50 2464228.63 2464239.46 2464261.51 2464271.33 2464284.13 2464295.36 2464307.06 2464330.76 2464355.96 2464368.42 2464407.76 2464419.13 2464432.63 2464547.64 2464557.96 2464569.19 2464584.08 2464598.08 2464611.57 2464622.64
BodyID Ͳ Ͳ Ͳ 705 Ͳ
Event ref_orbit maneuver maneuver flyby(16:1) maneuver
ȴv(m/s) 0 88.4 8.239 4.979 1.401
Ͳ Ͳ 701 Ͳ 701 Ͳ Ͳ 702 Ͳ 702 Ͳ 703 Ͳ 703 Ͳ 704 Ͳ
maneuver maneuver flyby(10:1) maneuver flyby(10:1) maneuver maneuver flyby(6:1) maneuver flyby(6:1) maneuver flyby(3:1) maneuver flyby(3:1) maneuver flyby(2:1) maneuver
79.107 8.932 0 0.984 0 99.205 2.687 0 0.956 0 181.205 0 2.094 0 139.335 0 1.548
Various views of the baseline satellite tour enhance understanding of the relative locations of maneuvers and satellite flybys. Figure 8 shows a Uranus inertial frame view of the satellite tour trajectory with red markers at maneuver locations and satellite flybys. An ecliptic plane projection view in Figure 9 shows the orbit’s orientation relative to the Sun and Earth directions (approximately in the –y direction). With such highly elliptical orbits, spacecraft-satellite encounter velocities vary from 10.9 km/s at Miranda to 8.8 km/s at Ariel to 7.3 km/s at Umbriel to 5.6 km/s at Titania to 4.7 km/s at Oberon. For Oberon and Titania, these encounter speeds allow for only ~1° of turning due to a hyperbolic flyby while the other moons provide ~0.25° or less turning.
Figure 7. Representative V map that aids the tour design process.
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Figure 8. Uranus orbiter baseline satellite tour trajectory in IAU Uranus J2000 frame.
Figure 9. Uranus orbiter baseline satellite tour trajectory in Ecliptic J2000 frame.
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Integrating the Uranus centric orbit using the patched-conic satellite ephemeris model and checking the distances to all five major Uranus moons at each time step provides a representative list of non-targeted satellite encounters with spacecraft minimum range less than 100,000 km. Out of 13 such non-targeted encounters with Miranda, Ariel, and Umbriel, the four closest encounters are with Umbriel and occur between 1,400 and 3,700 km. Spacecraft ground tracks were plotted on equidistant cylindrical surface image maps of the five major Uranus satellites, with imagery obtained from the January 1986 Voyager 2 encounter. Figures 10 to 14 show ground tracks at ranges less than 25,000 km for encounters with Miranda, Ariel, Umbriel, Titania, and Oberon, respectively. On each ground track plot red numbers indicate the flyby number in the satellite tour, a circle marks approach, an “x” marks periapse, a “¸” marks departure, and a yellow “*” indicates the sub-solar point.
Figure 10. Uranus orbiter ground tracks on the surface of Miranda.
Figure 11. Uranus orbiter ground tracks on the surface of Ariel.
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Figure 12. Uranus orbiter ground tracks on the surface of Umbriel.
Figure 13. Uranus orbiter ground tracks on the surface of Titania.
Figure 14. Uranus orbiter ground tracks on the surface of Oberon.
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Some observations regarding satellite tour design provide insight about limitations to achieving all requirements as well as recommendations for future work. The tour is not optimized end to end. Each leg is a global minimum for its associated single-impulse maneuver, or an approximate global minimum for the two-impulse maneuver. It is anticipated that extra maneuvers alongside an end-to-end optimization may provide some overall 'V savings. The near-polar inclined tour is highly constrained because 1) very high excess velocity limits flyby capabilities, and 2) the only potential for moon encounters is at the node crossings of the spacecraft. These limitations heavily constrain the design space and therefore the presented solutions are expected to be close to the global minimum for an end to end optimization. While non-trivial, J2 effects are not expected to qualitatively change the tour results. Moving to an integrated trajectory instead of patched conics could have 'V penalties (or reductions) on the same order of magnitude as the lack of J2 consideration. An additional 5 m/s/flyby is allocated (see Table 5) to accommodate for navigation errors and model fidelity errors (based on numbers from Cassini tour design). A future study should look at the coupling effect of optimizing the Uranus arrival conditions from the interplanetary trajectory to benefit the moon tour. The moon tour of the current study was constrained due to the primary science requiring a near polar orbit with very low Uranus attitudes. Results of an alternative satellite tour with 5 Miranda, 1 Ariel, 7 Umbriel, 4 Titania, and 5 Oberon flybys are not presented here because the tour flight time was more than 50% longer than the 424-day duration for the baseline satellite tour. An alternative, Galileo-style Uranus satellite tour was offered by Heaton and Longuski9. Table 5. Uranus Orbiter mission delta-V budget.
Phase Cruise
Probe Release Orbit Insertion
Science Orbit Satellite Tour
Total
Event LaunchInjection Cleanup EarthFlybyTargeting Interplanetary Statistical
DeltaͲV PropulsionType (m/s) (1=Hybrid2=biprop) Comment 0.0 0.0
0 0
ProvidedbySEPStage ProvidedbySEPStage
30.0
1
CruisepostͲSEPstageseparation
OrbitDeflection 30.0 UOIBͲplane targeting 10.0 OrbitInsertion 1661.0 Cleanup 25.0 OrbitMaintenance 20.0
2
afterproberelease
PeriapsisReduction MirandaTargeting MirandaStatistical ArielTargeting ArielStatistical UmbrielTargeting UmbrielStatistical TitaniaTargeting TitaniaStatistical OberonTargeting OberonStatistical
2 2 1 2 1 2 1 2 1 2 1
1 2 1 1
56.0 103.0 10.0 89.0 10.0 102.8 10.0 183.3 10.0 140.9 10.0 2501.1
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Finiteburn ~1m/sperorbit Lowerperiapsisnearendof scienceorbitphase 5m/sperflyby(2x) 5m/sperflyby(2x) 5m/sperflyby(2x) 5m/sperflyby(2x) 5m/sperflyby(2x)
CONCLUSION The Uranus orbiter mission concept study resulted in a viable trajectory design for a lowthrust 13-year transfer to Uranus with deployment of an atmospheric probe, a 20-orbit primary science phase, and a 14-month satellite tour. To meet decadal study cost guidelines for a New Frontiers or “sub-Flagship” class mission classification, an Atlas V-551 expendable launch vehicle was chosen in conjunction with 20-kW of power with two NEXT solar-electric propulsion thrusters for interplanetary cruise combined with bi-propellant chemical propulsion system for maneuvers from Uranus arrival through orbit insertion and the secondary satellite tour. Additional trades were conducted for the Uranus arrival and orbit mission phases. Atmospheric probe separation occurs 29 days prior to Uranus arrival and is followed a day later by an orbit deflection maneuver that targets the orbiter for Uranus orbit insertion. Parachute sizing trades yielded an 8.13 m2 primary parachute with 3.25-m diameter. The 51 minutes from parachute release to reaching 5 bar atmospheric pressure will be accompanied by data transfer to the lagging Uranus orbiter. Uranus orbit insertion can be no lower than 1.3 RU (vs. the 1.1 RU objective for gravity field measurements) to keep the Uranus ring plane crossing beyond the most distant ring. After a 1661-m/s UOI maneuver the orbiter enters a 21-day, 97.7°-inclination orbit. After a 431day primary science orbit the spacecraft periapse radius, providing ring avoidance can be assured, can be lowered to 1.1 RU. Upon completion of the primary science phase, a baseline 424-day, 619-m/s satellite tour design provides two targeted flybys each of the five largest moons of Uranus: Miranda, Ariel, Umbiel, Titania, and Oberon. ACKNOWLEDGMENTS The authors acknowledge NASA sponsorship for the Ice Giant Orbiter/Probe Mission Decadal study under contract NNN06AA01C, task NNN08AA03T with The Johns Hopkins University Applied Physics Laboratory (JHU/APL), where Helmut Seifert provided management oversight. REFERENCES 1
Sims, J. A., Finlayson, P. A., Rinderle, E. A., Vavrina, M. A., and Kawalkowski, T. D., “Implementation of a LowThrust Trajectory Optimization Algorithm for Preliminary Design,” AIAA 2006-6746, AIAA/AAS Astrodynamics Specialist Conference, Keystone, CO, August 21-24, 2006.
2
Landau, D., Lam, T., and Strange, N., “Broad Search and Optimization of Solar Electric Propulsion Trajectories to Uranus and Neptune,” AAS 09-428.
3
“Titan Saturn System Mission Final Report (on the NASA Contribution to a Joint Mission with ESA,” January 30, 2009.
4
Lindal, G. F., “The Atmosphere of Neptune: An Analysis of Radio Occultation Data Acquired with Voyager 2”, The Astronomical Journal, Vol. 103, No. 3, pp. 967-982, 1992.
5
Kazeminejad, B., Pérez-Ayúcar, M., Lebreton, J.P., Sanchez-Nogales, M., Belló-Mora, M., Strange, N., Roth, D., Popken, L., Clausen, K., and Couzin, P., “Simulation and analysis of the revised Huygens probe entry and descent trajectory and radio link modeling”, Planetary and Space Science, pp. 799-814, 2004.
6
Tauber, M., Wercinski, P., Henline, W., and Paterson, J., “Uranus and Neptune Atmospheric-Entry Probe Study”, Journal of Spacecraft and Rockets, Vol. 31, No. 5, pp. 799-805, 1994.
7
Strange N.J., Russell, R. P., Buffington, B., “Mapping the V-infinity Globe,” Paper AAS 07-277, AAS/AIAA Astrodynamics Specialist Conference and Exhibit, Mackinac Island, MI, Aug 2007.
8 Russell, R. P., Ocampo, C. A., “Geometric Analysis of Free-Return Trajectories Following a Gravity-Assisted Flyby,” Journal of Spacecraft and Rockets, Vol. 42, No. 1, pp. 138-151, 2005. 9
A. Heaton and J. Longuski, “Feasibility of a Galileo-Style Tour of the Uranian Satellites,” Journal of Spacecraft and Rockets, Vol. 40, No. 4, July-August 2003, pp. 591-596.
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