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.
1257
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
.
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.
1261
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).
1262
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.
1265
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.
1267
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.
1268
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
1269
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.
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2
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3
“Titan Saturn System Mission Final Report (on the NASA Contribution to a Joint Mission with ESA,” January 30,
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4
Lindal, G. F., “The Atmosphere of Neptune: An Analysis of Radio Occultation Data Acquired with Voyager 2”, The
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