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Alpbach Summer School 2012

GREEN TEAM

URANUS SYSTEM EXPLORER

USE2/08/2012

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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.

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Hubble Space Trelescope / NASA

We will achieve this with an orbiter and an atmospheric probe.

www.planeten.ch

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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?

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[1] Scientific Rationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Outline

Voyager 2 / NASA

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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

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Hubble Space Trelescope / NASA

Voyager 2

URANUS

Interior

Atmosphere

Magnetosphere

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Atmosphere of Uranus

6

Composition ?

Drivers of atmospheric chemistry ?

Dynamics (transport of heat)

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Magnetosphere of Uranus

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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?

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Magnetosphere of Uranus

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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

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Molecular H2

Inhomogeneous

Metallic H

Ices + RocksCore?

Molecular H2

Helium + Ices

Ices mixed with Rocks?

Rocks?

Interior of UranusRel. low heat flux

Uranus Jupiter

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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?

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[1] Scientific Rationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Outline

Voyager 2 / NASA

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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

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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

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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

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Voyager detections

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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

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We resolve the upper hybrid frequency < 1 MHz

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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

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Traceability Matrix - Interior

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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 ?  

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40 Ru

1.5 Ru

15 RuPeriod ~11 days

Mission Requirements - Science Phase I

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~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

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20 Ru

1.5-1.05 Ru

Period 4.3 days

10 Ru

Mission Requirements

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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

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Mission requirements

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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

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[1] Scientific Rationale

[2] Baseline design[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Outline

Voyager 2 / NASA

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[1] Scientific Rationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Outline

Voyager 2 / NASA

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Mission Baseline

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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

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Launch and Cruise phase

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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

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Launch and Cruise phase

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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).

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Orbit insertion in Uranus

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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.

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• 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

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2029 2030 2031 2033 2036 2049

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Probe insertion and descent

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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

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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

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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

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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

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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.

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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.

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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 )

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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

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Science operations6 kpbs / Downlink time 25% / Dedicated & normal modes

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[1] Scientific Rationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Outline

Voyager 2 / NASA

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Payload Configuration Payload panel 1: Remote Sensing Boom: Magnetometers Payload panel 2 and 3 (opposite sides): Plasma package

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Subsystems Configuration ASRGs:

3 ASRGs 90° apart. Back panel:

Probe Sides panels:

Radiators Low gain antennas

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Launcher Ariane 5 ECA launcher

Total launch = 4185 kg Fairing

Maximum diameter = 4570 m Maximum height = 15589 mm

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Adapted from Ariane V user manualAdapted from Ariane V user manual

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Propulsion Main engine: Leros-1b by AMPAC™ (JUNO Heritage)

Bipropellant engine: NTO-Hydrazins Specific Impulse = 318 s Nominal Thrust = 645 N Status: Flight Proven

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Adapted from AMPAC™ website

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Probe layout

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Probe configuration during cruise phase

Elements of the probe:

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Attitude Control

The AACS provides accurate dynamic control of the satellite in both rotation and translation.

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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

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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

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Q & A – Inertia Calculations

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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

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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

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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

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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)

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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

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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)

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Low Gain Antenna

Low gain antenna for communications during NEOP.

Communications over S-band with Kourou.

Low mass and power patch antenna.

LGA

LGA

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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)

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Probe UHF link budget

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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

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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

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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.

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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

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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.)

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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

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Venus

Solar backscattering from Venus (but high altitude)

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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)

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Power budget: science orbit

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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)

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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

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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

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USE IT!

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Thank you!

&

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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

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Backup: Probe Mass Budget

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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.

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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)

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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)