Alpbach Summer School 2012
URANUS SYSTEM EXPLORER
USE
2/08/2012
GREEN
TEAM
1
Mission Summary
Study the Uranian system with a
focus on the interior, atmosphere
and magnetosphere in order to
better constrain the solar formation
model and to understand how the
icy giants formed and evolved.
www.planeten.ch
We will achieve this with an
orbiter and an atmospheric
probe.
Hubble Space Trelescope / NASA
2
2
ESA Cosmic Vision 2015-2025
What are the conditions for Planet Formation and the
Emergence of Life?


Observations of Uranus will help to improve existing
models of planetary system formation
Understand icy giant planets (exoplanets)
How does the Solar System Work?


What is the structure and dynamics of the icy giants?
How do they interact with their space environment?
3
3
Outline
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
Voyager 2 / NASA
[7] Conclusion
4
Basic facts of Uranus
URANUS
Interior
Voyager 2





Magnetosphere
Atmosphere
One of the 4 giant planets
Distance: 19 AU
Rotation Period: 17h
Orbit Period: 84 years
Only visited by Voyager 2 in 1986
Hubble Space Trelescope / NASA
5
5
Atmosphere of Uranus
Composition ?
Drivers of
atmospheric
chemistry ?
Dynamics
(transport of
heat)
6
6
Magnetosphere of Uranus
Field Intensity @ 1.4 Ru
 Rotation axis tilt 98°
 Dipole axis tilt by 59°
 Large quadrupole
moment
Voyager 2
Source: Nicholas et al., AGU, 2011
How and where is the intrinsic field
generated? A new class of dynamo?
7
7
Magnetosphere of Uranus
How does the magnetosphere
interact with solar wind?
Voyager 2
How is plasma transported in the
Uranian magnetosphere?
 Rotation axis tilt 98°
 Dipole axis tilt by 59°
 Large quadrupole
moment
Is there a significant internal plasma
source on Uranus?
Insight into Earth’s magnetosphere
during magnetic reversals
LASP, University of Colorado, Boulder
8
8
Interior of Uranus
Rel. low heat flux
Molecular H2
Molecular H2
Helium + Ices
Inhomogeneous
Metallic H
Ices mixed with
Rocks?
Rocks?
Uranus
Ices + Rocks
Core?
Jupiter
9
Interior of Uranus
Why is the heat flux lower than
expected? Implications for the interior
Molecular H2
Molecular H2
and thermal evolution of the planet?
Rel. low heat flux
Helium + Ices
Why does Uranus have such a strong
intrinsic magnetic field? How do its
Ices mixed with
characteristics constrainRocks?
the interior?
Inhomogeneous
Metallic H
Is there a rocky
silicate core? Implications
Rocks?
for solar system formation?
Uranus
Ices + Rocks
Core?
Jupiter
10
Outline
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
Voyager 2 / NASA
[7] Conclusion
11
Mission Payload
Orbiter
Atmospheric Probe
Imaging Camera (CAM)
Visible and Infrared Spectrometer (VIR-V & VIR-I)
Remote
Thermal IR Spectrometer (TIR)
UV-Specrtometer (UVS)
Microwave Radiometer (MR)
Electron and ion spectrometer (EIS)
Scalar and Vector Magnetometer (SCM & MAG)
In situ
Energetic Particle Detector (EPD)
Radio and Plasma Wave Instrument (RPWI)
Ion composition instrument (ICI)
Mass Spectrometer (ASS & GCMS)
Nephelometer (NEP)
Doppler wind instrument (DWI)
Atmosphere Physical Properties Package (AP3)
12
12
Orbiter Payload
 Imaging camera - New Horizons / Lorri


Study the cloud motion and winds of Uranus
Range: 0.35 – 0.85 μm ; FOV: 0,29 x 0,29 deg
 Visible and Infrared Spectrometer - Dawn / VIR



Study chemical composition of the atmosphere
Range: 0.25 – 1.05 μm ; FOV: 3,67 x 3,67 deg
Range: 1.0 – 5.0 μm ; FOV: 0,22 x 0,22 arcmin
 Thermal IR Spectrometer - Cassini / CIRS


Heat flux at different points to constrain models of the interior and
thermal evolution
Range: 7.67 – 1000 µm ; Spectral Resolution 0.5 – 20/cm
 UV Spectrometer - New Horizons


Morphology and source of Uranus auroral emission
Range: 52 – 187 nm ; Spectral Resolution < 3nm ; spatial res < 500 km
13
13
Orbiter Payload
 Electron and ion spectrometer – Rosetta/EIS


Measures electrons and ions
Range: 1-22 keV
 Ion composition instrument – Rosetta / ICA


Voyager detections
Measure magnetospheric plasma particles in order to study plasma
composition and distribution
Range: 1eV/e to 22 keV/e ; Resolution: dE/E = 0.04
 Energetic Particle Detector - New Horizons / PEPPSI


Energetic charged particles that can be used to characterize and locate
radiation belts
Range: 15 keV – 30 MeV ; energy resolution: 8 keV
14
14
Orbiter Payload
 Magnetometer - Juno


globally measure the magnetic field from low altitude to constrain the
dynamics of the field generation layer
resolution < 1nT in range of 0.1 – 120000 nT
 Radio Wave and Plasma Instrument - Cassini


Measure plasma waves
range: kHz – MHz
We resolve the upper hybrid
frequency < 1 MHz
 Microwave Radiometer - Juno / MWR


atmospheric and terrestrial radiation, air temperature, total amount of
water vapor and total amount of liquid water
range: 1.3 – 50 cm
 High gain antenna

Space craft tracking to make gravity field measurements
15
15
Probe Payload
 Aerosol sampling system / Gas Chromatograph & Mass Spectrometer Galileo
sample aerosols during descent and a gas chromatograph and measure heavy
elements, noble gas abundances, key isotope ratios
 range: 1 – 150 amu/e

 Nephelometer - Galileo

studies dust particles in the clouds of Uranus' upper atmosphere
 Doppler Wind Instrument - Huygens / DWE
height profile of Uranus zonal wind velocity
 resolution: 1 m/s

 Atmosphere Physical Properties Package - Huygens / HASI

measure the physical characteristics of the atmosphere
 temperature sensor
 pressure sensor
 3 axis accelerometer
 electric field sensor
16
16
Traceability Matrix - Interior
Instrument
Intrinsic magnetic field /
dynamo
Mass distribution in the
interior
Heat flux
MAG TIR RAD
What is the origin of the intrinsic magnetic
field?
Is there secular variation in the Uranian
magnetic field?
Extent of mass of Si core
Are there different layers with different
composition, states of matter?
Is it uniform / Are there hotspots ? Heat
transport mechanisms ?
17
17
Mission Requirements Science Phase I
Interior (Gravity)
 HGA visible from
Earth
 Low altitude
Period ~11 days
15 Ru
40 Ru
Sun
Magnetosphere
 Globally probe
magnetosphere
 Cross
magnetopause
1.5 Ru
Atmosphere
 Global coverage on
day- and nightside
 Occultation
18
25
Mission Requirements
 Science Phase II and III
Interior (Magnetic Field)
 Global coverage with
low altitude
Period 4.3 days
10 Ru
20 Ru
1.5-1.05 Ru
Interior (Gravity)
 HGA visible from
Earth
 Low altitude
Atmosphere
 Global coverage on
day- and nightside
19
26
Mission requirements
 Gravity and magnetic field
 Higher orders can only be
resolved at lower altitudes
 Here: 2.5 Ru for degree 11
 Signal decays
exponentially with altitude
 Higher orders decay more
efficiently
20
20
Outline
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
Voyager 2 / NASA
[7] Conclusion
21
Outline
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
Voyager 2 / NASA
[7] Conclusion
22
Mission Baseline
SCIENCE PHASE
CRUISE PHASE
Nov 2049
UOI
Jun 2033
Earth GA
Mar 2030
Venus GA
Feb 2031
Earth GA
2029-Oct
2031-Feb
08 Oct 2029
Launch
May-Nov 2052
Science phase 3
Mar 2036
Jupiter GA
Sep 2049
Probe release
Nov 2049- Sep 2050
Science phase 1
2036-Mar
May 2051-May 2052
Science phase 2
2052
26 Nov 2052
End of nominal mission
23
23
Launch and Cruise phase
2029
2030 2031
2033
2036
2049
Launch 8 Oct 2029 02:18:41
 Ariane 5 launch.
 𝑉𝑒𝑠𝑐 = 3.56 km/s (C3=12.67)
 Total Mass available: 4185.1 kg -> Launch driven by mass maximization.
Total time cruise phase: 20.139 years
24
24
Launch and Cruise phase
2029
2030 2031
2033
2036
2049
Launch 8 Oct 2029 02:18:41




Gravity Assist sequence: Venus-Earth-Earth-Jupiter.
Total ΔV = 0.21 km/s
5% Margin and 25m/s maintenance for the 5 legs applied.
The mission is classified category II (COSPAR Planetary Protection).
Total time cruise phase: 20.139 years
25
25
Orbit insertion in Uranus
2029
2030 2031
2033
2036
2049
Orbit insertion in Uranus: 19 Nov 2049 13:33:00
 Uranus Orbit Insertion: 19 Nov 2049 with ΔV = 0.60 km/s burn.
 Velocity at Uranus arrival: 3.36 km/s
 Final orbit Inclination set to 90° at arrival.
26
26
Probe insertion and descent
2029
2030 2031
2033
2036
2049
• Probe release:
 Probe released 19 Sep 2049, 2 months before orbit insertion.
 Release maneuver ΔV = 0.001 km/s burn.
• Probe insertion
 Entry at the atmosphere at 23 km/s.
 Arrival at latitude of 20 deg.
 Dayside arrival.
• Probe descent
27
27
Probe insertion and descent
0
Probe Entry, t = 0 min
Δt ≈ 5 min
Pressure (bar)
Drogue Parachute
Drogue Parachute Release
Δt ≈ 2 min
Top Cover Removed
0.1
Heat Shield Drops Off
100
Probe Mission Terminates
t = 90 min
28
28
Insertion
Science Phase Profile
10 months
6 months
Science Phase 1
Nov 2049
12 months
Science Phase 2
Sep 2050
6 months
Science Phase 3
May 2051
May 2052
End of
nominal
mission
Nov 2052
Total science phase duration: 34 months
29
29
Science Phase 1 Orbits
10 months
Nov 2049
Sep 2050
125 orbits
 Highly elliptical polar orbit.
Eccentricity
0.93
-
Inclination
90
°
Arg. of perigee
280.39
°
Apogee radius
996803.63
km
Perigee radius
38338.58 (1.5 Ru) km
Period
11.67
days
 Large apoapsis to sample magnetosphere
and cross magnetopause.
 Low periapsis for gravity field
measurements.
 Dayside/Nightside global coverage.
[3] Mission analysis
30
30
Science Phase 2 Orbits
6 months
Sep 2050
12 months
May 2051 84 orbits May 2052
30 orbits
 Orbit circularization lowering the apoapsis
in 4 steps: 1.40-1.35-1.30-1.25-1.20 Ru
 10 orbits at each step, 84 at last orbit.
 Total ΔV = 0.55 km/s
Eccentricity
0.86
-
Inclination
90
°
Arg of perigee
°
Apogee radius
511215.10 (20 Ru)
km
Perigee radius
37331.12 (1.05 Ru)
km
Period
4.34
days
 Detailed magnetosphere sampling at
different Ru.
31
31
6 months
Science Phase 3 Orbits
May 2052 42 orbits Nov 2052



Highly elliptical polar orbit with low periapsis.
Argument of perigee gain of 10 deg.
Avoiding dust hazards from the rings.
Eccentricity
0.90
-
Inclination
90
°
Arg. of perigee
297.176-307.16
°
Apogee radius
510334 (20 Ru)
km
Perigee radius
26837 (1.05 Ru)
km
Period
4.51
days


Internal gravity field sampling.
Enhanced magnetic field sampling.

Untargeted Uranian satellites fly-bys.
 END OF MISSION: deorbiting maneuver at apoapsis of ΔV = 0.04 km/s to
deliberately crash the orbiter to Uranus (avoiding satellite contamination).
32
32
Extended mission orbits
? months
Nov 2052


?orbits
??????
Highly elliptical polar orbit with low periapsis.
Argument of perigee gain (20 deg per year).
Eccentricity
0.90
-
Inclination
90
°
Arg. of perigee
297.176-307.16
°
Apogee radius
510334 (20 Ru)
km
Perigee radius
26837 (1.05 Ru)
km
Period
4.51
days
 Enhanced magnetic field sampling.
 Untargeted Uranian satellites fly-bys.
 Aerobraking.
• END OF MISSION?: Remaining ΔV or aerobraking
33
33
ΔV and fuel budget - Cruise
ID
Total Mass
ΔV [km/s]
[kg]
Maneuver
Used
Remaining
fuel
Fuel [kg]
[kg]
1
Initial State
3.56
4185.1
2095.1
2
Venus-Earth DSM
0.04
4093.86
57.80
2003.86
3
Earth-Earth DSM
0.04
4010.78
50.37
1920.78
4
Earth-Jupiter DSM
0.00
3978.74
0.00
1888.74
5
6
Jupiter-Uranus DSM
Probe release maneuver
0.00
0.00
3946.95
3607.42
0.00 1856.95
307.99 1517.42
 Total ΔV = 0.21 km/s (includes 5% margin and 25m/s maintenance )
1
2029
2
3
4
2030 2031
2033
5
6
2036
34
2049
34
ΔV and fuel budget – Science Phase
ID
Maneuver
Used fuel
[kg]
ΔV [km/s]
Remaining Total Mass
Fuel [kg]
[kg]
1
Orbit Insertion
0.6343
664.37
853.05
2943.05
2
Apo 40Ru-35Ru
0.0994
92.37
760.67
2850.67
3
Apo 35Ru-30Ru
0.0603
54.61
706.07
2796.07
4
Apo 30Ru-25Ru
0.0833
73.75
632.31
2722.31
5
Apo 25Ru-20Ru
0.1228
105.15
527.16
2617.16
6
Per 1.5Ru-1.05Ru
0.1875
152.79
374.37
2464.37
7
De-orbit
0.0443
34.79
339.59
2429.59
• Total ΔV = 1.23 km/s
Mission total ΔV = 1.44 km/s
Insertion
Remaining
10 months
6 months
ΔV = 0.47 km/s
12 months
6 months
1
2 3
Nov 2049
Sep 2050
4
5
May 2051
6
May 2052
35
7
End of
nominal
mission
Nov 2052
35
Science operations
6 kpbs / Downlink time 25% / Dedicated & normal modes
36
36
Outline
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground
segment design
Voyager 2 / NASA
[7] Conclusion
37
Payload Configuration



Payload panel 1: Remote Sensing
Boom: Magnetometers
Payload panel 2 and 3 (opposite sides): Plasma package
38
38
Subsystems Configuration

ASRGs:
 3 ASRGs 90° apart.

Back panel:
 Probe

Sides panels:
 Radiators
 Low gain antennas
39
39
Launcher

Ariane 5 ECA launcher
 Total launch = 4185 kg

Fairing
 Maximum diameter = 4570 m
 Maximum height = 15589 mm
Adapted from Ariane V user manual
Adapted from Ariane V user manual
40
40
Propulsion

Main engine: Leros-1b by AMPAC™ (JUNO Heritage)
 Bipropellant engine: NTO-Hydrazins
 Specific Impulse = 318 s
 Nominal Thrust = 645 N
 Status: Flight Proven
Adapted from AMPAC™ website
41
41
Probe layout
 Probe configuration during cruise phase
 Elements of the probe:
42
42
Attitude Control

The AACS provides accurate
dynamic control of the satellite
in both rotation and
translation.
Required Pointing Accuracies
angle (degrees)
Comms.
0,107
Probe Relay
0,107
Remote Sensing
0,061086524
Pointing Stability (1sec)
0,001396263
Pointing Stability (1h)
0,048869219
 Payload
•
•
•
•
•
4 x Reaction Wheels
4 xThrusters Clusters
2 x Star Trackers
2 x Sun Sensors
3 x MIMU
43
43

Attitude phases:
Possible + Z spinning during cruise. It is required to protect sensors, pointing HGA antenna to
the Sun. AACS is automated with coarse Sun sensors.
3-axis stability when approaching with RWA, compensation the realease of the proabe with
thrusters;
During nominal phase, 3-axis attitude control is done with reaction wheels. The largest
reaction torque is 0.13 Nm. Angular momenta less than 34 Nms (approx.: 2000 rpm);
fast maneuvers or accelerations must be achieved with less precise but faster thrusters
(RCS);
Inertia Tensors calculated before and after probe releasing. In both cases the
values are inferior to those in Cassini which uses the same actuators.
Sensors' Resolution
IRU - uncalibrated
< 0.5 deg/h
IRU - calibrated
< 0.05 deg/h
Sun Sensors
< 0.01 deg
Star Trackers
< 0.001 deg
Achievable Pointing Accuracies - angle (degrees)
Attitude
Typical Accuracy Maneuverability
Thrusters
RWA
Spinning
0,2 Fast, least accurate
0,01 Slow, very accurate
from new horizons
0,027 heritage
44
44
Q & A – Inertia Calculations
-> 1 N thrusters
-> 0.13 Nm
Inertia Matrix w/ Probe
Ix
5458,0 Kg.m2
Iy
4828,7 Kg.m2
Iz
1590,4 Kg.m2
Good maneuverability !
Inertia Matrix w/o Probe
Ix
3137,8 Kg.m2
Iy
2203,6 Kg.m2
Iz
1285,5 Kg.m2
Change in the CM
45
45
Communication Overview
 HGA for Orbiter-Earth
communications


Ka-band downlink (35 GHz)
X-band uplink (7.2 GHz)
 MGA for Orbiter-Earth
communications near Venus


X-band downlink (8.1 GHz)
X-band uplink (7.2 GHz)
 LGA for LEOPS


S-band downlink (2.2 GHz)
S-band uplink (2.1 GHz)
 UHF for Probe-Orbiter
communications

UHF (400/420 MHz) dual uplink
46
High Gain Antenna
 4m Cassini-derived
HGA
HGA for Earth comms
to ESTRACK 35m
network.


Ka Band downlink
(35GHz)
X-band uplink (7.2Ghz)
 Ultrastable oscillator
(HGA used for radio
science)
47
High gain antenna link budget
HGA
Downlink
Uplink
Waveband
Ka (35 GHz)
X (7.2 GHz)
Transmitter power
20 dBW (100 W)
-
Transmitter line Loss
-0.458 dB
Antenna pointing Loss
-1.33 dB
Transmission losses
-315 dB
-301 dB
EIRP
80.3 dBW
92 dBW
Receiver G/T
62.8 dBi/K
23.1 dBi/K
Carrier-to-noise C/N0
55.2 dB Hz
42.3 dB Hz
Coding
BPSK/RS Viterbi
DPSK
Required Eb/N0
2.7 dB
12 dB
Data rate
6 kbps
70 bps
Implementation loss
-2 dB
-2 dB
Rain attenuation
-3 dB
-10 dB
Link margin
3.17 dB (>3 dB)
6.04 dB (>6 dB)
48
Medium Gain Antenna
 Medium gain antenna for
communications with
orbiter near Venus when
HGA used as sun shield.
 Communications over Xband with Kourou.
 0.8m diameter steerable
antenna.
 Rosetta heritage.
MGA
49
ESA
Medium gain antenna link budget
MGA
Downlink
Uplink
Waveband
X (8.1 GHz)
X (7.2 GHz)
Transmitter power
20 dBW (50 W)
-
Transmitter line Loss
-0.458 dB
-
Antenna pointing Loss
-0.286 dB
-
Transmission losses
-281 dB
-301 dB
EIRP
50.6 dBW
92 dBW
Receiver G/T
41.0 dBi/K
9.14 dBi/K
Carrier-to-noise C/N0
39.3 dB Hz
38.6 dB Hz
Coding
BPSK/RS Viterbi
DPSK
Required Eb/N0
2.7 dB
12 dB
Data rate
500 bps
30 bps
Implementation loss
-2 dB
-2 dB
Rain attenuation
-3 dB
-3 dB
Link margin
3.81 dB (>3 dB)
6.03 dB (>6 dB)
50
Low Gain Antenna
 Low gain antenna for
communications
during NEOP.
LGA
 Communications over
LGA
S-band with Kourou.
 Low mass and power
patch antenna.
51
Low gain antenna link budget
LGA
Downlink
Uplink
Waveband
S (2.2 GHz)
S (2.1 GHz)
Transmitter power
10 W
-
Transmitter Antenna Losses
-0.458 dB
Antenna pointing Loss
-19.8 dB
Transmission losses
-197.3 dB
-197 dB
EIRP
14.4 dBW
74.7 dBW
Receiver G/T
29.1 dBi/K
-19.6 dBi/K
Carrier-to-noise C/N0
55.0 dB Hz
86.6 dB Hz
Coding
BPSK/RS Viterbi
DPSK
Required Eb/N0
2.7 dB
12 dB
Data rate
500 bps
500 bps
Implementation loss
-2 dB
-2 dB
Rain attenuation
-3 dB
-3 dB
Link margin
19.5 dB (>3 dB)
41.8 dB (>6 dB)
52
Probe UHF link budget
LGA
Uplink #1 (Coast)
Uplink #1 (100 bar)
Waveband
UHF (400 MHz)
UHF (400 MHz)
Antenna
Patch LGA on Aeroshell
Quad helix on Probe
Transmitter power
10 W
150W
Transmitter line loss
-0.45 dB
-0.45 dB
Antenna pointing Loss
-11.81dB
-9.33dB
Transmission losses
-182 dB
-197 dB
EIRP
24.18 dB
21.89 dB
Receiver G/T
-3.55 dB
-3.55 dB
Carrier-to-noise C/N0
40.57 dB
40.70 dB
Coding
BPSK/RS Viterbi
BPSK/RS Viterbi
Required Eb/N0
2.7 dB
2.7 dB
Data rate
2.25 kbps
2.32 kbps
Implementation loss
-3 dB
-3 dB
Atmospheric attenuation
0 dB
-15 dB
Link margin
2.7 dB
2.7 dB
53
53
Unit mass
[kg]
Qty
Mass (CBE)
[kg]
DMM
Mass
(CBE+DMM)
[kg]
HGA
100
1
100
5%
105
X/Ka band
Rx/Tx
5.05
2
10.1
20%
(average)
11.8
MGA
10.4
1
10.4
30%
13.6
X band
transponder
/filters
5.3
2
10.6
10%
11.7
LGA
0.08
2
0.16
20%
0.19
S-band
transponder
2.6
2
5.2
10%
5.72
2
3.704
10%
4.07
Communications system mass
UHF receiver 1.852
Total
152 kg
54
Communications system power budget
DMM
Uplink (CBE)
[W]
Uplink
(CBE+DMM)
[W]
Downlink
(CBE) [W]
Downlink
(CBE+DMM)
[W]
High gain
(averages)
20%
25
32.5
140
158
Medium
gain
10%
20
22
50
55
Low gain
10%
5
5.5
20
22
Probe UHF
10%
6
6.6
0
0
Uplink power consumption scaled from downlink using typical numbers from SMAD.
Probe UHF system values are from Mars Odessey.
Power consumption for TWTA (40W) comes from WFI CDF study report.
55
Command & Data Handling
 GNC and CDH Flight computers redundant
 Mass storage: Two High Speed Solid State
Data recorders 4 Gbyte (4x1 Gbyte DRAM)

Redundant storage – recorders operate in
parallel
 Primary data bus (MIL-1553)

Spacewire to high data rate instruments (ORS),
SSDRs and communications system.
56
Ground Segment
ESTRACK
MOC
35-m antennas
Mission Operation Center
15-m antenna for LEO
SOC
Science Operation
Center

PGS
Probe Ground Segment
MOC
 monitoring and control of the complete mission
 generation and provision of the complete raw-data sets

SOC
 scientific mission planning support
 creation of pre-processed scientific data

PGS
 supports operations of the Probe
 coordinates scientific mission planning
57
57
Radiometric Tracking
Meas. Type
Precision
Tracking schedule
range-rate
~0.1mm/s
cruise
once per week
range
~1m
ang. pos. (VLBI)
~0.03mas1 (s/c)
~2.43mas2 (probe)
some months before UOI
until end of science phase
once per day at
pericenter passage
probe descent
continuous
10.03mas
translates to 0.4 km at the mean distance of Uranus (35GHz freq.)
22.43mas translates to 34 km at the mean distance of Uranus (400MHz freq.)
POD during science phase: ranges, range-rates
 Position accuracy of the s/c in the
 cruise phase: 10-20km
 science phase: km range
 VLBI can be used to improve Uranus’ ephemeris
 Position accuracy of the probe: 34km

58
58
Thermal control – Hot case
Venus
Solar backscattering from Venus
(but high altitude)
2649.7 W/m²
178.4*(Rv/Rorbit)²
payload panel




Antenna towards sun for critical hot case
Avoid payload panel towards Venus
Radiators top, bottom or towards zenith
ARSGs shadowed by antenna
59
59
Thermal control – Cold case
0.55*(Rv/Rorbit)²
Uranus
Eclipse
158W electrical power for payload
(assume 10% dissipation)
Critical => heat load needed for cold case for balance
Possibility to use ASRG waste heat load in addition to decrease
need of heaters (15W for New Horizons)
Heat pipes for better transport to critical components (tanks,
batteries)
Classic solution: louvers
VCHP? Heat switch?
60
60
Thermal control
-Radiator
α/ε <<
Teflon aluminized
Teflon silvered
OSR
-Cryo-radiator for
IR payload
-Louvers
Batteries
MLI
HGA
α/ε <<
Teflon aluminized
Teflon silvered
OSR
IR payload
Payload
MLI+Conductive
insulation
Tanks
MLI
ASRG
Eff=28%, EOL electrical power=130W
Heat load dissipated~334W
Outer S/C cover
MLI betacloth
outer layer
61
61
Power budget
We plan to use ASRGs (Am241, 27.8kg, 140W
BOL, 130W EOL).
 Scaled from the Nasa plutonium ARSRGs, taking
into account the lower activity level of Am241
(requires 5x more radioactive material). 20%
margin applied.

w/o margin
w margin
BUS (W)
129.5
142.9
PAYLOAD (W)
152.9
166.4
TOTAL (W)
282.4
309.30
390
312
3xASGs EOL (W)
Assumption:
peak load
(without COMM
subsystem)
62
62
Power budget: science orbit
Occultation
science
~158W
COMM
63
63
Mass budget
13%
8%
(kg)
Total dry mass (excl. adaptator)
With 20% system margin
Propellant needed until after
orbit insertion
Propellant needed for attitude
control
Propellant needed for science
orbit
Total wet mass (excl. adaptator)
Adaptator (incl. separation
mechanism)
Total
CBE+DMM
2,115
62%
935
S/C BUS
158
S/C ORBITER PAYLOAD
S/C BUS
S/C ORBITER
PAYLOAD
791
S/C PROBE
19%
31%
S/C PROBE
4%
S/C PROPELLANT
(until insertion)
3,999
186
4185
S/C PROPELLANT
(AOCS)
22%
4%
7%
MASS MARGIN
Launch capability
= 4185kg
64
64
Risk management
[5] Spacecraft design
65
65
Critical items
Critical items
Uranus rings plane hazards
Thermal design (heat load variation)
Probe (TPS, trajectory)
Structure (FEA)
[5] Spacecraft design
66
66
Development Timeline
67
Cost estimation
Description
M€
Sub-Total M€
%
Launcher
175
175
9
1.450
74
230
12
100
100
5
Total
1.955
100
Main S/C
Probe
Operations
PF
1.250
PL
200
PF
200
PL
30
68
Summary
 USE is equipped with sufficient instruments to carry out
sufficient measurements to answer the scientific questions
 mass, power and cost budget allow the mission to be
feasible
 technologies proposed use heritage from previous space
missions and can be easily implemented for future space
missions
69
Thank you!
&
USE IT!
70
70
Backup: Command & Data Handling
 GNC and CDH Flight computers: Leon3FT, 89 MIPS
 Secondary GNC/CDH computers


Hardware watchdog timer based redundancy
If Primary computer does not reset the timer, backups are
brought online.
 Mass storage: Two High Speed Solid State Data
recorders (derived from LRO-SSDR)


4 Gbyte (4x1 Gbyte DRAM)
Redundant storage – recorders operate in parallel
 Primary data bus (MIL-1553)

Spacewire to high data rate instruments (ORS), SSDRs and
communications system.
71
Backup: Probe Mass Budget
72
72
Backup: Cruise phase science
 Take measurements of Venus and Jupiter during successive fly

bys.
Calibration of instruments during Earth fly-bys.
During the Venus fly-by use the ‘Energetic Particle Detector’
and the ‘Radio and Plasma wave instrument’, to measure the
interaction of Venus with the solar wind.
 During Jupiter flyby, we can use the newer instrumentation to


obtain more, accurate, results then previous flybys.
During the two flyby’s of Earth, the obiter's systems can be
calibrated.
Instruments shall be calibrated every year and engines tested.
73
Backing: Calibrating instruments
 Using the Earth to calibrate the obiters instruments
means that we can rely on ground based observations
as well as satellite. This would lower error margins, and
help to signify any problems the instruments may be
having
 Orbiting the Earth twice will allow ground operations
to check twice the working order of the instruments.
 Calibrating the magnetometer: as satellite passes
through Earths Magnetic field, the reading it samples
can be compared to the known value for the Earths
magnetic field and the instruments can be calculated
accordingly
74
Backup: Calibrating instruments
 Other instruments that rely on the interaction of the Earth’s


magnetic field can be calibrated using the orbiting satellites and
taking measurements around the earth. For example the ‘Energetic
Particle Detector’ and the ‘Electron and Ion Spectrometer’ can be
calculated using the current orbiting satellite’s data.
Other instruments such as the imaging camera and visible and
infrared Spectrometer need to take images of certain sections of
the Earth of which the wavelengths are known. Using those
previously obtained values and comparing our results, to see if they
fall within the acceptable range, we can determine if and by how
much the instruments need calibrating.
Since Earth sends out a large number of radio waves, we can use
these known radio waves to calibrate our radio instruments.
75
Backup: Planetary protection
All mission is of Category II :
Type of Mission
Planet
Category
Flyby
Venus
II
Jupiter*
II
Orbiter
Uranus
II
Probe
Uranus
II
* Case of Europa
Category II: All types of missions to target bodies where there is significant interest
relative to the process chemical evolution and the origin of life, but where there is only a
remote chance that contamination carried by a spacecraft could compromise future
investigations.
76
Backup: Planetary protection
77
Backup: Thermal control (hot case)
Venus
Solar backscattering from Venus
(but high altitude)
2649.7 W/m²
178.4*(Rv/Rorbit)²
payload panel
 HGA: α=0.1; ε=0.8; A=14,3 m²
(teflon aluminized)
 Payload panel: α=0.5; ε=0.5; A=3.7*1.7 m²
 Side panels+Back panel: α=0.4; ε=0.9; A=3.7*1.7 m²
(betacloth)
 Radiators not considered
 No critical power consumption for this case
78
78
Backup: Thermal control (hot case)
Venus
Solar backscattering from Venus
(but high altitude)
2649.7 W/m²
178.4*(Rv/Rorbit)²
payload panel
1 node S/C
α_antenna*Qsun*A_antenna + εside*Qir*Aside + Qdiss=
σ*ε_antenna*A_antenna*T^4
+σ*ε_payload_panel*A_payload_panel*T^4
+3*σ*ε_panel*A_panel*T^4
+σ*ε_bottom_panel*A_bottom_panel*T^4
(neglecting exchanges with area between antenna and top panel)
(neglecting VF of the other panels than probe panel)
79
79
Backup: Thermal control (cold case)
0.55*(Rv/Rorbit)²
Uranus
Eclipse
158W electrical power for payload
(assume 10% dissipation)
ε_payload_panel*A_payload_panel*Qir + Qdiss=
4*σ*ε_panel*A_panel*T^4
+σ*ε_bottom_panel*A_bottom_panel*T^4
+ σ*ε_antenna*A_antenna*T^4
(neglecting exchanges with area between antenna and top
panel)
(neglecting VF of the other panels than probe panel)
80
80
Backup: Thermal control (cold case)
T° IR instrument
Cryo-radiator
ε_payload*A_aperture*Qir + Qdiss=
σ*ε_payload*A_internal_surfaces*T^4
+GL*(T^4-T_cryo_radiator)
GL*(T^4-T_cryo_radiator) =
σ*ε_cryo_radiator*A_cryo_radiator*T_cryo_radiator^4
81
81