1. Solar Orbiter EUV Spectrometer
Thermal Design Considerations
Bryan Shaughnessy

2. The Thermal Challenge
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

3. Spacecraft Thermal Interface
Preliminary interfaces (SCI-A/ /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 Fully closed at 20C; effective emissivity 0.1

4. 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 ~
Grazing Incidence (backup)
Coated SiC optics (short, medium and long wavelengths)
Solar absorptivity ~

5. Normal Incidence Thermal Concept

6. Normal Incidence Thermal Concept
ENTRANCE BAFFLE
HEAT STOP / HEAT REJECTION MIRROR
PRIMARY MIRROR
DETECTOR THERMALLY
ISOLATED ENCLOSURE
HEAT REJECTION I/F (HOT)
HEAT REJECTION I/F (COLD)

7. Grazing Incidence Thermal Concept

8. Grazing Incidence Thermal Concept
HEAT REJECTION I/F (HOT)
HEAT REJECTION I/F (COLD)
ENTRANCE BAFFLE
DETECTOR THERMALLY
ISOLATED ENCLOSURE
HEAT STOP
BAFFLE

9. Heat Load Summary (at 0.2 AU)

10. 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

11. 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

12. Primary Mirror flexible thermal link
Bolted clamp between foil bundle
and PM interface plate
PM interface plate
Strap interface to spacecraft heat rejection point
Foils PM

13. 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

14. Thermal Predictions

15. Orbital Solar Load Variation
Note variation over the 30 day (+/- 15 day) observation period

16. Orbital Solar Load Variation – impact on NI Primary Mirror

17. 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

18. THE END