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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Solar Orbiter EUV Spectrometer
Thermal Design Considerations
Bryan Shaughnessy
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
The Thermal Challenge
Phase
Cold Non-Operational
Hot Non-Operational
Cold Operational
Hot Operational
Sun Distance
(AU)
1.2
0.8
0.45
0.2
Heat Flux
(kW/m2)
1.0
2.2
6.8
34.4
Note
Cruise phase
Aphelion
Start/end 30 day solar encounter
Perihelion
• Reject heat input to instrument of order 100 W at 0.2 AU
• Maintain sensible temperatures through the solar
encounter
• Reduce heat loss when instrument is further from the Sun
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Spacecraft Thermal Interface
• Preliminary interfaces (SCI-A/2005-307/SO/AJ Issue 1):
– Instrument contained within spacecraft
– Cold finger interface provided for detector cooling
– Interfaces to fluid loops/heat pipes for hot elements.
• Spacecraft rejects heat using louvered radiators (ESA
CDF study)
• Radiators likely to needed embedded heat-pipes or loop
heat pipes to distribute heat.
• Modelling assumptions
– 50 W/K thermal link from interfaces to radiator
– Radiator efficiency 90%
– Louvers : Fully open at 40C; effective emissivity 0.7
Fully closed at 20C; effective emissivity 0.1
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Instrument Configuration
• Normal Incidence (baseline)
– Uncoated SiC or Au coated SiC primary mirror (for
medium/long wavelengths)
• Solar absorptivity ~ 0.8
– Au coated SiC primary mirror (for medium/long wavelengths)
• Solar absorptivty ~ 0.1
– Multilayer coated SiC primary mirror (for short wavelengths)
• Solar absorptivity ~ 0.4 - 0.6
• Grazing Incidence (backup)
– Coated SiC optics (short, medium and long wavelengths)
• Solar absorptivity ~ 0.5 - 0.6
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Normal Incidence Thermal Concept
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Normal Incidence Thermal Concept
PRIMARY MIRROR
ENTRANCE BAFFLE
HEAT STOP / HEAT REJECTION MIRROR
HEAT REJECTION I/F (HOT)
HEAT REJECTION I/F (COLD)
DETECTOR THERMALLY
ISOLATED ENCLOSURE
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Grazing Incidence Thermal Concept
Plane mirror for
rastering
Entrance slit
Hyperbolic mirror
From the Sun

Parabolic mirror

TVLS grating
Detector
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Grazing Incidence Thermal Concept
Plane mirror for
rastering BAFFLE
Entrance slit
Hyperbolic mirror
From the Sun
HEAT STOP

Parabolic mirror
ENTRANCE BAFFLE

TVLS grating
Detector
DETECTOR THERMALLY
ISOLATED ENCLOSURE
HEAT REJECTION I/F (HOT)
HEAT REJECTION I/F (COLD)
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Heat Load Summary (at 0.2 AU)
GI
Through Aperture
Entrance Baffle Absorbed
PM Absorbed
SM Absorbed
SM Baffle Absorbed
RM Absorbed
Slit Incident
86
3
54
17
9
1
1
NI (various PM finishes)
SiC
Multilayer
Au
132
132
132
28
28
28
82
52
10
21
52
93
Note
Partially reflect out of instrument?
Partially reflect out of instrument?
Heat stop / heat reflection mirror
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Basic Thermal Requirements
• Detector temperature: < -60 C (target -80 C)
• Optics: < 100 C assumed
– Coatings (if used) are limiting factor
• Hot heat rejection interface < +50 C
– Assuming NH3 heat-pipes
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Primary Mirror flexible thermal link
•
•
•
•
High conductance flexible thermal link required:
Alignment with spacecraft interfaces
Allow PM scanning
Assume PM can operate hot (~ 100 C) but spacecraft interface
limited to 50 C :
– Conductance required: ~ 1.6 W/K (NI with absorbing PM)
– Approximately 180 x 0.1 mm Al foils (25 mm wide, 50 mm length)
and bolted clamps
• Careful design required
– Thermally induced deformation of mirror surface
– Need to ensure that spacecraft interface is not heated above is
maximum temperature requirement
• Similar link could be used for all heat rejection interfaces
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Primary Mirror flexible thermal link
PM interface plate
Bolted clamp between foil bundle
and PM interface plate
Foils
PM
Strap interface to
spacecraft heat rejection
point
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Thermal Predictions
• ESATAN/ESARAD thermal models have been developed for the
NI and GI configurations
• Predictions presented for NI (absorbing PM) and GI
• Further assumptions:
– No MLI around instrument
– Spacecraft conductive/radiative interfaces temperatures 40 C (Hot)
and 0 C (Cold)
– Detector dissipation 1.6 W
– NI configuration assumes mirror at heat stop reflects unwanted
radiation out of instrument
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Thermal Predictions
TEMPERATURE PREDICTIONS
NI
GI
0.2 AU
0.8 AU
0.2 AU
0.8 AU
BAFFLE
52
0
45
-2
BAFFLE RADIATOR
51
0
PRIMARY MIRROR
95
0
80
-2
PM RAD I/F
47
-4
49
-4
PRIMARY RADIATOR
45
-4
48
-4
SECONDARY MIRROR
60
-1
SM RAD I/F
50
-2
SECONDARY RADIATOR
49
-2
RASTER MIRROR
69
2
SLIT/HEAT STOP
47
1
54
1
GRATING
41
0
45
0
DETECTOR
-85
-93
-84
-93
DETECTOR RADIATOR
-116
-121
-116
-121
Detector
Primary Mirror
Entrance Baffle
Secondary Mirror
RADIATOR AREAS (m2)
0.09
0.21
0.07
-
0.09
0.13
0.04
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Orbital Solar Load Variation
90
80
Note variation over the 30 day
(+/- 15 day) observation period
PM absorbed load, W
70
60
50
40
30
20
10
0
0
15
30
45
60
75
90
105
120
Time, days
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Orbital Solar Load Variation – impact on NI
Primary Mirror
50
0.7
45
35
0.5
30
0.4
25
0.3
20
15
0.2
10
Radiator effective emissivity
40
Temperature, deg C
0.6
T400 - Primary Mirror
T405 - S/C thermal i/f
T410 - S/C louvered radiator
Radiator effective emissivity
0.1
5
0
0
0
15
30
45
60
75
90
105
120
Time, days
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
Conclusion
• Thermal design concepts outlined for EUS
• Thermal design is highly dependent on spacecraft thermal
interfaces
– Heat sink temperature requirements
– Variation in heat rejection, especially over solar encounter period
• Critical areas:
– Design of high conductance flexible straps, particularly interfaces to
optical surfaces (i.e., thermal distortion)
– Feasibility of ‘heat rejection mirrors’
– Qualification of coatings (if used)
– Intensity of solar beam at heat-stop if reflective primary mirror used
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Solar Orbiter EUS: Thermal Design Considerations
Bryan Shaughnessy, Rutherford Appleton Laboratory
THE END
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