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URANUS SYSTEM EXPLORER . GREEN TEAM. USE. Alpbach Summer School 2012. 2/08/2012. Mission Summary. www.planeten.ch. We will achieve this with an orbiter and an atmospheric probe . Hubble Space Trelescope / NASA. 2. - PowerPoint PPT Presentation
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1
Alpbach Summer School 2012
GREEN TEAM
URANUS SYSTEM EXPLORER
USE2/08/2012
2
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.
2
Hubble Space Trelescope / NASA
We will achieve this with an orbiter and an atmospheric probe.
www.planeten.ch
3
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
4
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
[7] Conclusion
Outline
Voyager 2 / NASA
5
Basic facts of Uranus
One of the 4 giant planets Distance: 19 AU Rotation Period: 17h Orbit Period: 84 years Only visited by Voyager 2 in 1986
5
Hubble Space Trelescope / NASA
Voyager 2
URANUS
Interior
Atmosphere
Magnetosphere
6
Atmosphere of Uranus
6
Composition ?
Drivers of atmospheric chemistry ?
Dynamics (transport of heat)
7
Magnetosphere of Uranus
7
Rotation axis tilt 98° Dipole axis tilt by 59° Large quadrupole
momentVoyager 2
Source: Nicholas et al., AGU, 2011
Field Intensity @ 1.4 Ru
How and where is the intrinsic field generated? A new class of dynamo?
8
Magnetosphere of Uranus
8
Rotation axis tilt 98° Dipole axis tilt by 59° Large quadrupole
momentVoyager 2
LASP, University of Colorado, Boulder
Is there a significant internal plasma source on Uranus?
How is plasma transported in the Uranian magnetosphere?
How does the magnetosphere interact with solar wind?
Insight into Earth’s magnetosphere during magnetic reversals
9
Molecular H2
Inhomogeneous
Metallic H
Ices + RocksCore?
Molecular H2
Helium + Ices
Ices mixed with Rocks?
Rocks?
Interior of UranusRel. low heat flux
Uranus Jupiter
10
Molecular H2
Inhomogeneous
Metallic H
Ices + RocksCore?
Molecular H2
Helium + Ices
Ices mixed with Rocks?
Rocks?
Interior of UranusRel. low heat flux
Uranus Jupiter
Why does Uranus have such a strong intrinsic magnetic field? How do its characteristics constrain the interior?
Why is the heat flux lower than expected? Implications for the interior and thermal evolution of the planet?
Is there a rocky silicate core? Implications for solar system formation?
11
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
[7] Conclusion
Outline
Voyager 2 / NASA
12 12
Orbiter
Atmospheric Probe
Imaging Camera (CAM)Visible and Infrared Spectrometer (VIR-V & VIR-I)Thermal IR Spectrometer (TIR)UV-Specrtometer (UVS)Microwave Radiometer (MR)Electron and ion spectrometer (EIS)Scalar and Vector Magnetometer (SCM & MAG)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)
Mission Payload
In situ
Remote
13
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
14
Orbiter Payload
Electron and ion spectrometer – Rosetta/EIS Measures electrons and ions Range: 1-22 keV
Ion composition instrument – Rosetta / ICA 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
Voyager detections
15
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
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
We resolve the upper hybrid frequency < 1 MHz
16
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
17
Traceability Matrix - Interior
17
Instrument MAG TIR RAD
Intrinsic magnetic field / dynamo
What is the origin of the intrinsic magnetic field?
Is there secular variation in the Uranian magnetic field?
Mass distribution in the interior
Extent of mass of Si core
Are there different layers with different composition, states of matter?
Heat flux Is it uniform / Are there hotspots ? Heat transport mechanisms ?
18
40 Ru
1.5 Ru
15 RuPeriod ~11 days
Mission Requirements - Science Phase I
25
~20 Ru
Magnetosphere Globally probe
magnetosphere Cross
magnetopause
Atmosphere Global coverage on
day- and nightside Occultation
Interior (Gravity) HGA visible from
Earth Low altitude
Sun
19
20 Ru
1.5-1.05 Ru
Period 4.3 days
10 Ru
Mission Requirements
26
Atmosphere Global coverage on
day- and nightside
Interior (Gravity) HGA visible from
Earth Low altitude
Interior (Magnetic Field) Global coverage with
low altitude
Science Phase II and III
20
Mission requirements
20
Signal decays exponentially with altitude
Higher orders decay more efficiently
Higher orders can only be resolved at lower altitudes
Here: 2.5 Ru for degree 11
Gravity and magnetic field
21
[1] Scientific Rationale
[2] Baseline design[3] Mission analysis
[5] Spacecraft and ground segment design
[7] Conclusion
Outline
Voyager 2 / NASA
22
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
[7] Conclusion
Outline
Voyager 2 / NASA
23
Mission Baseline
23
08 Oct 2029 Launch
Feb 2031Earth GA
Jun 2033Earth GA
Mar 2036Jupiter GA
Nov 2049UOI
May 2051-May 2052Science phase 2
Sep 2049Probe release
Mar 2030Venus GA
26 Nov 2052End of nominal mission
2029-Oct 2031-Feb 2036-Mar 2052
CRUISE PHASE SCIENCE PHASE
Nov 2049- Sep 2050Science phase 1
May-Nov 2052Science phase 3
24
Launch and Cruise phase
24
2029 2030 2031 2033 2036 2049
Launch 8 Oct 2029 02:18:41
Ariane 5 launch. 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
25
Launch and Cruise phase
25
Total time cruise phase: 20.139 years
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).
26
Orbit insertion in Uranus
26
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.
27
• 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
Probe insertion and descent
27
2029 2030 2031 2033 2036 2049
28
Probe insertion and descent
28
100
Probe Entry, t = 0 min
Drogue Parachute
Drogue Parachute Release
Top Cover Removed
Heat Shield Drops Off
Probe Mission Terminates t = 90 min
0.1
0
Pres
sure
(bar
)
Δt ≈ 5 min
Δt ≈ 2 min
29 29
Science Phase Profile
Science Phase 1
10 months
Total science phase duration: 34 months
Insertion
Nov 2049 Sep 2050
6 months 12 months
Science Phase 2
May 2051 May 2052
6 months
Nov 2052
Science Phase 3
End of nominal mission
30 30
Science Phase 1 Orbits
[3] Mission analysis
Highly elliptical polar orbit.
10 months
Nov 2049 Sep 2050125 orbits
Large apoapsis to sample magnetosphere and cross magnetopause.
Low periapsis for gravity field measurements.
Dayside/Nightside global coverage.
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
31 31
Science Phase 2 Orbits6 months
Sep 2050 May 2051 May 2052
12 months
Detailed magnetosphere sampling at different Ru.
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
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
84 orbits30 orbits
32 32
Science Phase 3 Orbits
Internal gravity field sampling. Enhanced magnetic field sampling.
END OF MISSION: deorbiting maneuver at apoapsis of ΔV = 0.04 km/s to deliberately crash the orbiter to Uranus (avoiding satellite contamination).
Untargeted Uranian satellites fly-bys.
May 2052
6 monthsNov 2052
42 orbits
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
Highly elliptical polar orbit with low periapsis. Argument of perigee gain of 10 deg. Avoiding dust hazards from the rings.
33 33
Extended mission orbits
Enhanced magnetic field sampling.
• END OF MISSION?: Remaining ΔV or aerobraking
Untargeted Uranian satellites fly-bys.
Nov 2052
? months
???????orbits
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
Highly elliptical polar orbit with low periapsis. Argument of perigee gain (20 deg per year).
Aerobraking.
34 34
ΔV and fuel budget - CruiseID Maneuver ΔV [km/s] Total Mass
[kg]Used fuel [kg]
Remaining Fuel [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 Jupiter-Uranus DSM 0.00 3946.95 0.00 1856.956 Probe release maneuver 0.00 3607.42 307.99 1517.42
2029 2030 2031 2033 2036 2049
12 3 4 5 6
Total ΔV = 0.21 km/s (includes 5% margin and 25m/s maintenance )
35 35
ID Maneuver ΔV [km/s] Used fuel [kg]
Remaining Fuel [kg]
Total Mass [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
ΔV and fuel budget – Science Phase
10 monthsInsertion
Nov 2049 Sep 2050
6 months 12 months
May 2051 May 2052
6 months
Nov 2052
End of nominal mission
12 3 4 5 6 7
• Total ΔV = 1.23 km/s Mission total ΔV = 1.44 km/s
Remaining ΔV = 0.47 km/s
36 36
Science operations6 kpbs / Downlink time 25% / Dedicated & normal modes
37
[1] Scientific Rationale
[2] Baseline design
[3] Mission analysis
[5] Spacecraft and ground segment design
[7] Conclusion
Outline
Voyager 2 / NASA
38
Payload Configuration Payload panel 1: Remote Sensing Boom: Magnetometers Payload panel 2 and 3 (opposite sides): Plasma package
38
39 39
Subsystems Configuration ASRGs:
3 ASRGs 90° apart. Back panel:
Probe Sides panels:
Radiators Low gain antennas
40
Launcher Ariane 5 ECA launcher
Total launch = 4185 kg Fairing
Maximum diameter = 4570 m Maximum height = 15589 mm
40
Adapted from Ariane V user manualAdapted from Ariane V user manual
41
Propulsion Main engine: Leros-1b by AMPAC™ (JUNO Heritage)
Bipropellant engine: NTO-Hydrazins Specific Impulse = 318 s Nominal Thrust = 645 N Status: Flight Proven
41
Adapted from AMPAC™ website
42
Probe layout
42
Probe configuration during cruise phase
Elements of the probe:
43
Attitude Control
The AACS provides accurate dynamic control of the satellite in both rotation and translation.
43
Payload• 4 x Reaction Wheels• 4 xThrusters Clusters• 2 x Star Trackers• 2 x Sun Sensors• 3 x MIMURequired 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
44 44
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.
Achievable Pointing Accuracies - angle (degrees)
Typical AccuracyAttitude Maneuverability
Thrusters 0,2 Fast, least accurate
RWA 0,01 Slow, very accurate
Spinning 0,027from new horizons heritage
Sensors' Resolution
IRU - uncalibrated < 0.5 deg/h
IRU - calibrated < 0.05 deg/h
Sun Sensors < 0.01 deg
Star Trackers < 0.001 deg
45
Q & A – Inertia Calculations
45
Inertia Matrix w/ Probe
Ix 5458,0 Kg.m2
Iy 4828,7 Kg.m2
Iz 1590,4 Kg.m2 Inertia Matrix w/o Probe
Ix 3137,8 Kg.m2
Iy 2203,6 Kg.m2
Iz 1285,5 Kg.m2Good maneuverability !
-> 1 N thrusters-> 0.13 Nm
Change in the CM
46
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
47
High Gain Antenna 4m Cassini-derived HGA
for Earth comms to ESTRACK 35m network. Ka Band downlink
(35GHz) X-band uplink (7.2Ghz)
Ultrastable oscillator (HGA used for radio science)
HGA
48
High gain antenna link budgetHGA Downlink UplinkWaveband Ka (35 GHz) X (7.2 GHz)Transmitter power 20 dBW (100 W) -Transmitter line Loss -0.458 dBAntenna pointing Loss -1.33 dBTransmission losses -315 dB -301 dBEIRP 80.3 dBW 92 dBWReceiver G/T 62.8 dBi/K 23.1 dBi/KCarrier-to-noise C/N0 55.2 dB Hz 42.3 dB Hz
Coding BPSK/RS Viterbi DPSKRequired Eb/N0 2.7 dB 12 dB
Data rate 6 kbps 70 bpsImplementation loss -2 dB -2 dBRain attenuation -3 dB -10 dBLink margin 3.17 dB (>3 dB) 6.04 dB (>6 dB)
49
Medium Gain Antenna
Medium gain antenna for communications with orbiter near Venus when HGA used as sun shield.
Communications over X-band with Kourou.
0.8m diameter steerable antenna.
Rosetta heritage.ESA
MGA
50
Medium gain antenna link budgetMGA Downlink UplinkWaveband 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 dBEIRP 50.6 dBW 92 dBWReceiver G/T 41.0 dBi/K 9.14 dBi/KCarrier-to-noise C/N0 39.3 dB Hz 38.6 dB Hz
Coding BPSK/RS Viterbi DPSKRequired Eb/N0 2.7 dB 12 dB
Data rate 500 bps 30 bpsImplementation loss -2 dB -2 dBRain attenuation -3 dB -3 dBLink margin 3.81 dB (>3 dB) 6.03 dB (>6 dB)
51
Low Gain Antenna
Low gain antenna for communications during NEOP.
Communications over S-band with Kourou.
Low mass and power patch antenna.
LGA
LGA
52
Low gain antenna link budgetLGA Downlink UplinkWaveband S (2.2 GHz) S (2.1 GHz)Transmitter power 10 W -Transmitter Antenna Losses -0.458 dBAntenna pointing Loss -19.8 dBTransmission losses -197.3 dB -197 dBEIRP 14.4 dBW 74.7 dBWReceiver G/T 29.1 dBi/K -19.6 dBi/KCarrier-to-noise C/N0 55.0 dB Hz 86.6 dB Hz
Coding BPSK/RS Viterbi DPSKRequired Eb/N0 2.7 dB 12 dB
Data rate 500 bps 500 bpsImplementation loss -2 dB -2 dBRain attenuation -3 dB -3 dBLink margin 19.5 dB (>3 dB) 41.8 dB (>6 dB)
53
Probe UHF link budget
53
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
54
Communications system mass
Unit mass [kg]
Qty Mass (CBE) [kg]
DMM Mass (CBE+DMM) [kg]
HGA 100 1 100 5% 105X/Ka band Rx/Tx
5.05 2 10.1 20% (average)
11.8
MGA 10.4 1 10.4 30% 13.6X band transponder/filters
5.3 2 10.6 10% 11.7
LGA 0.08 2 0.16 20% 0.19S-band transponder
2.6 2 5.2 10% 5.72
UHF receiver 1.852 2 3.704 10% 4.07
Total 152 kg
55
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 22Probe 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.
56
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.
Command & Data Handling
57 57
Ground SegmentMOC
Mission Operation Center
SOCScience Operation
Center
PGSProbe Ground Segment
ESTRACK35-m antennas
15-m antenna for LEO
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
58 58
Radiometric TrackingMeas. Type Precisionrange-rate ~0.1mm/srange ~1mang. pos. (VLBI) ~0.03mas1 (s/c)
~2.43mas2 (probe)
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
Tracking schedule
cruise once per week
some months before UOIuntil 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.)
59
Thermal control – Hot case
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
Venus
Solar backscattering from Venus (but high altitude)
60
Thermal control – Cold case
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
Uranus Eclipse158W electrical power for payload (assume 10% dissipation)
0.55*(Rv/Rorbit)²
61
Thermal control
61
Payload MLI+Conductive insulation
HGAα/ε <<Teflon aluminized Teflon silveredOSR
TanksMLI
-Radiatorα/ε <<Teflon aluminized Teflon silveredOSR-Cryo-radiator for IR payload
-Louvers
BatteriesMLI
ASRGEff=28%, EOL electrical power=130WÞ Heat load dissipated~334W
IR payload
Outer S/C coverMLI betacloth outer layer
62
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.
62
w/o margin w margin
BUS (W) 129.5 142.9
PAYLOAD (W) 152.9 166.4
TOTAL (W) 282.4 309.30
3xASGs EOL (W) 390 312
Assumption: peak load (without COMM subsystem)
63
Power budget: science orbit
63
COMM
Occultation science ~158W
64
Mass budget
64
(kg) CBE+DMMTotal dry mass (excl. adaptator)With 20% system margin
2,115
Propellant needed until after orbit insertion
935
Propellant needed for attitude control
158
Propellant needed for science orbit
791
Total wet mass (excl. adaptator) 3,999Adaptator (incl. separation mechanism)
186
Total 4185Launch capability = 4185kg
62%
8%
13%
S/C BUS S/C ORBITER PAYLOADS/C PROBE
31%
4%7%
22%
4%
19%S/C BUSS/C ORBITER PAYLOADS/C PROBES/C PROPELLANT (until insertion)S/C PROPELLANT (AOCS)MASS MARGIN
65
Risk management
65[5] Spacecraft design
66
Critical items
66[5] Spacecraft design
Critical itemsUranus rings plane hazards
Thermal design (heat load variation)
Probe (TPS, trajectory)
Structure (FEA)
67
Development Timeline
68
Cost estimation
Description M€ Sub-Total M€ %
Launcher 175 175 9
Main S/CPF 1.250
1.450 74PL 200
ProbePF 200
230 12PL 30
Operations 100 100 5
Total 1.955 100
69
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
70
USE IT!
70
Thank you!
&
71
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.
Backup: Command & Data Handling
72
Backup: Probe Mass Budget
72
73
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.
74
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
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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.
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All mission is of Category II :
Type of Mission Planet Category
Flyby Venus II
Jupiter* II
Orbiter Uranus II
Probe Uranus II
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.
* Case of Europa
Backup: Planetary protection
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Backup: Planetary protection
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Backup: Thermal control (hot case)
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
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Venus
Solar backscattering from Venus (but high altitude)
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2649.7 W/m² 178.4*(Rv/Rorbit)²
payload panel
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)
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Venus
Solar backscattering from Venus (but high altitude)
1 node S/C
Backup: Thermal control (hot case)
80 80
Uranus Eclipse158W electrical power for payload (assume 10% dissipation)
0.55*(Rv/Rorbit)²
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)
Backup: Thermal control (cold case)
81 81
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
Cryo-radiator
T° IR instrument
Backup: Thermal control (cold case)