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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.

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green team

URANUS SYSTEM EXPLORER

GREEN TEAM

USE

AlpbachSummerSchool 2012

2/08/2012

mission summary
Mission Summary

www.planeten.ch

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

Hubble Space Trelescope / NASA

2

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.

esa cosmic vision 2015 2025
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?
slide4

Outline

[1] ScientificRationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Voyager 2 / NASA

slide5

Basic facts of Uranus

URANUS

Interior

Magnetosphere

Atmosphere

Voyager 2

Hubble Space Trelescope / NASA

One of the 4 giant planets

Distance: 19 AU

RotationPeriod: 17h

Orbit Period: 84 years

OnlyvisitedbyVoyager 2 in 1986

5

atmosphere of uranus
Atmosphere of Uranus

Composition ?

Drivers of atmospheric chemistry ?

Dynamics (transport of heat)

slide7

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

slide8

Magnetosphere of Uranus

How does the magnetosphere interact with solar wind?

  • Rotation axis tilt 98°
  • Dipole axis tilt by 59°
  • Large quadrupole moment

How is plasma transported in the Uranian magnetosphere?

Voyager 2

Is there a significant internal plasma source on Uranus?

Insight into Earth’s magnetosphere during magnetic reversals

LASP, University of Colorado, Boulder

8

slide9

Interior of Uranus

Molecular H2

Inhomogeneous

Metallic H

Ices + Rocks

Core?

Rel. low heat flux

Molecular H2

Helium + Ices

Ices mixed with Rocks?

Rocks?

Uranus

Jupiter

slide10

Interior of Uranus

Molecular H2

Inhomogeneous

Metallic H

Ices + Rocks

Core?

Why is the heat flux lower than expected? Implications for the interior and thermal evolution of the planet?

Rel. low heat flux

Molecular H2

Helium + Ices

Why does Uranus have such a strong intrinsic magnetic field? How do its characteristics constrain the interior?

Ices mixed with Rocks?

Is there a rocky silicate core? Implications for solar system formation?

Rocks?

Uranus

Jupiter

slide11

Outline

[1] ScientificRationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Voyager 2 / NASA

slide12

Mission Payload

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)

Remote

Orbiter

In situ

Mass Spectrometer (ASS & GCMS)

Nephelometer (NEP)

Doppler wind instrument (DWI)

Atmosphere Physical Properties Package (AP3)

Atmospheric Probe

12

orbiter payload
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
orbiter payload1
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

Voyager detections

orbiter payload2
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

We resolve the upper hybrid frequency < 1 MHz

probe payload
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
slide18

Mission Requirements -

Science Phase I

  • Interior (Gravity)
  • HGA visible from Earth
  • Low altitude

15 Ru

Period ~11 days

  • Magnetosphere
  • Globally probe magnetosphere
  • Cross magnetopause

40 Ru

1.5 Ru

~20 Ru

Sun

  • Atmosphere
  • Global coverage on day- and nightside
  • Occultation

25

slide19

Mission Requirements

  • Science Phase II and III
  • Interior (Magnetic Field)
  • Global coverage with low altitude

Period 4.3 days

  • Interior (Gravity)
  • HGA visible from Earth
  • Low altitude

10 Ru

20 Ru

1.5-1.05 Ru

  • Atmosphere
  • Global coverage on day- and nightside

26

mission requirements
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
slide21

Outline

[1] ScientificRationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Voyager 2 / NASA

slide22

Outline

[1] ScientificRationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Voyager 2 / NASA

mission baseline
Mission Baseline

SCIENCE PHASE

CRUISE PHASE

Nov 2049

UOI

May-Nov 2052

Science phase 3

Mar 2036

Jupiter GA

Jun 2033

Earth GA

Mar 2030

Venus GA

Sep 2049

Probe release

Feb 2031

Earth GA

Nov 2049- Sep 2050

Science phase 1

May 2051-May 2052

Science phase 2

2031-Feb

2036-Mar

2029-Oct

2052

08 Oct 2029

Launch

26 Nov 2052

End of nominal mission

launch and cruise phase
Launch and Cruise phase

2036

2049

2033

2029

2030

2031

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

launch and cruise phase1
Launch and Cruise phase

2036

2049

2033

2029

2030

2031

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

o rbit insertion in uranus
Orbit insertion in Uranus

2036

2049

2033

2029

2030

2031

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.
probe insertion and descent
Probe insertion and descent

2036

2049

2033

2029

2030

2031

  • 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 descent1
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

Heat Shield Drops Off

0.1

Probe Mission Terminates

t = 90 min

100

science phase profile
Science Phase Profile

Insertion

10 months

6 months

12 months

6 months

End of nominal mission

Science Phase 1

Science Phase 2

Science Phase 3

Sep 2050

May 2051

May 2052

Nov 2052

Nov 2049

Total science phase duration: 34 months

science phase 1 orbits
Science Phase 1 Orbits

10 months

Sep 2050

Nov 2049

125 orbits

  • Highly elliptical polar orbit.
  • Large apoapsis to sample magnetosphere and cross magnetopause.
  • Low periapsis for gravity field measurements.
  • Dayside/Nightside global coverage.

[3] Mission analysis

science phase 2 o rbits
Science Phase 2 Orbits

6 months

12 months

84 orbits

Sep 2050

May 2051

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
  • Detailed magnetosphere sampling at different Ru.
science phase 3 o rbits
Science Phase 3 Orbits

6 months

Nov 2052

May 2052

42 orbits

  • Highly elliptical polar orbit with low periapsis.
  • Argument of perigee gain of 10 deg.
  • Avoiding dust hazards from the rings.
  • 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).
extended mission orbits
Extended mission orbits

? months

Nov 2052

?orbits

??????

  • Highly elliptical polar orbit with low periapsis.
  • Argument of perigee gain (20 deg per year).
  • Enhanced magnetic field sampling.
  • Untargeted Uranian satellites fly-bys.
  • Aerobraking.
  • END OF MISSION?: Remaining ΔV or aerobraking
v and fuel budget cruise
ΔV and fuel budget - Cruise
  • Total ΔV = 0.21 km/s (includes 5% margin and 25m/s maintenance )

1

2

3

5

4

6

2036

2049

2033

2029

2030

2031

slide35

ΔV and fuel budget – Science Phase

Mission total ΔV = 1.44 km/s

  • Total ΔV = 1.23 km/s

Insertion

Remaining ΔV = 0.47 km/s

10 months

6 months

12 months

6 months

End of nominal mission

1

2

4

3

5

6

7

Sep 2050

May 2051

May 2052

Nov 2052

Nov 2049

science operations
Science operations

6 kpbs / Downlink time 25% / Dedicated & normal modes

slide37

Outline

[1] ScientificRationale

[2] Baseline design

[3] Mission analysis

[5] Spacecraft and ground segment design

[7] Conclusion

Voyager 2 / NASA

payload configuration
Payload Configuration
  • Payload panel 1: Remote Sensing
  • Boom: Magnetometers
  • Payload panel 2 and 3 (opposite sides): Plasma package
subsystems configuration
Subsystems Configuration
  • ASRGs:
    • 3 ASRGs 90° apart.
  • Back panel:
    • Probe
  • Sides panels:
    • Radiators
    • Low gain antennas
launcher
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

propulsion
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

probe layout
Probe layout
  • Probe configuration during cruise phase
  • Elements of the probe:
attitude control
Attitude Control
  • The AACS provides accurate dynamic control of the satellite in both rotation and translation.
    • Payload
  • 4 x Reaction Wheels
  • 4 xThrusters Clusters
  • 2 x Star Trackers
  • 2 x Sun Sensors
  • 3 x MIMU
slide44

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.
q a inertia calculations
Q & A – Inertia Calculations

-> 1 N thrusters

-> 0.13 Nm

Good maneuverability !

Change in the CM

communication overview
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
high gain antenna
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

medium gain antenna
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.

MGA

ESA

low gain antenna
Low Gain Antenna
  • Low gain antenna for communications during NEOP.
  • Communications over S-band with Kourou.
  • Low mass and power patch antenna.

LGA

LGA

communications system power budget
Communications system power budget

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.

slide56

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.
ground segment
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
radiometric tracking
RadiometricTracking
  • POD during science phase: ranges, range-rates
  • Position accuracy of the s/c in the
    • cruise phase: 10-20km
    • science phase: km range
  • VLBI canbeusedto improve Uranus’ ephemeris
  • Position accuracy of the probe: 34km
  • 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.)
thermal control hot case
Thermal control – Hot case

Venus

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

Solar backscattering from Venus (but high altitude)

thermal control cold case
Thermal control – Cold case
  • 0.55*(Rv/Rorbit)²

Uranus

  • 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?

Eclipse

158W electrical power for payload (assume 10% dissipation)

thermal control
Thermal control

-Radiator

α/ε <<

Teflon aluminized Teflon silvered

OSR

-Cryo-radiator for IR payload

-Louvers

HGA

α/ε <<

Teflon aluminized Teflon silvered

OSR

IR payload

Payload

MLI+Conductive insulation

Batteries

MLI

Tanks

MLI

  • ASRG
  • Eff=28%, EOL electrical power=130W
  • Heat load dissipated~334W

Outer S/C cover

MLI betacloth outer layer

power budget
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.

Assumption: peak load (without COMM subsystem)

power budget science orbit
Power budget: science orbit

Occultation science ~158W

COMM

mass budget
Mass budget

Launch capability = 4185kg

risk management
Risk management

[5] Spacecraft design

critical items
Critical items

[5] Spacecraft design

slide69

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
use it
USE IT!

Thank you!

&

slide71

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)
    • 4Gbyte (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 cruise phase science
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.
backing calibrating instruments
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
backup calibrating instruments
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.
slide76

Backup: Planetary protection

All mission is of Category 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.

backup thermal control hot case
Backup: Thermal control (hot case)

Venus

2649.7 W/m² 178.4*(Rv/Rorbit)²

payload panel

  • HGA: α=0.1; ε=0.8; A=14,3 m² (teflonaluminized)
  • Payload panel: α=0.5; ε=0.5; A=3.7*1.7 m²
  • Sidepanels+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

Solar backscattering from Venus (but high altitude)

backup thermal control hot case1
Backup: Thermal control (hot case)

Venus

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 betweenantenna and top panel)

(neglecting VF of the other panels than probe panel)

Solar backscattering from Venus (but high altitude)

1 node S/C

backup thermal control cold case
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)
backup thermal control cold case1
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