solar orbiter eus: thermal design considerations bryan shaughnessy, rutherford appleton laboratory 1...

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Solar Orbiter EUS: Thermal Design ConsiderationsBryan Shaughnessy, Rutherford Appleton Laboratory

Solar Orbiter EUV Spectrometer

Thermal Design Considerations

Bryan Shaughnessy

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The Thermal Challenge

PhaseSun Distance

(AU)Heat Flux (kW/m2) Note

Cold Non-Operational 1.2 1.0 Cruise phaseHot Non-Operational 0.8 2.2 AphelionCold Operational 0.45 6.8 Start/end 30 day solar encounterHot Operational 0.2 34.4 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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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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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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Normal Incidence Thermal Concept

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Normal Incidence Thermal Concept

ENTRANCE BAFFLE

HEAT STOP / HEAT REJECTION MIRROR

PRIMARY MIRROR

DETECTOR THERMALLYISOLATED ENCLOSURE

HEAT REJECTION I/F (HOT)

HEAT REJECTION I/F (COLD)

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Grazing Incidence Thermal Concept

Plane mirror forrastering

Parabolic mirror

Hyperbolic mirror

Entrance slit

Detector

TVLS grating

From the Sun

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Grazing Incidence Thermal Concept

Plane mirror forrastering

Parabolic mirror

Hyperbolic mirror

Entrance slit

Detector

TVLS grating

From the Sun

HEAT REJECTION I/F (HOT)

HEAT REJECTION I/F (COLD)

ENTRANCE BAFFLE

DETECTOR THERMALLYISOLATED ENCLOSURE

HEAT STOP

BAFFLE

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Heat Load Summary (at 0.2 AU)

GISiC Multilayer Au Note

Through Aperture 86 132 132 132Entrance Baffle Absorbed 3 28 28 28 Partially reflect out of instrument?PM Absorbed 54 82 52 10SM Absorbed 17 - - -SM Baffle Absorbed 9 - - - Partially reflect out of instrument?RM Absorbed 1 - - -Slit Incident 1 21 52 93 Heat stop / heat reflection mirror

NI (various PM finishes)

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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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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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Primary Mirror flexible thermal link

FoilsStrap interface to spacecraft heat rejection point

Bolted clamp between foil bundleand PM interface platePM interface plate

PM

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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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Thermal Predictions

0.2 AU 0.8 AU 0.2 AU 0.8 AUBAFFLE 52 0 45 -2BAFFLE RADIATOR 51 0 - -PRIMARY MIRROR 95 0 80 -2PM RAD I/F 47 -4 49 -4PRIMARY RADIATOR 45 -4 48 -4SECONDARY MIRROR - - 60 -1SM RAD I/F - - 50 -2SECONDARY RADIATOR - - 49 -2RASTER MIRROR - - 69 2SLIT/HEAT STOP 47 1 54 1GRATING 41 0 45 0DETECTOR -85 -93 -84 -93DETECTOR RADIATOR -116 -121 -116 -121

DetectorPrimary MirrorEntrance BaffleSecondary Mirror

NI GI

RADIATOR AREAS (m2)

TEMPERATURE PREDICTIONS

0.090.13

-0.04

0.090.210.07

-

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Orbital Solar Load Variation

0

10

20

30

40

50

60

70

80

90

0 15 30 45 60 75 90 105 120

Time, days

PM

abs

orbe

d lo

ad, W

Note variation over the 30 day (+/- 15 day) observation period

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Orbital Solar Load Variation – impact on NI Primary Mirror

0

5

10

15

20

25

30

35

40

45

50

0 15 30 45 60 75 90 105 120

Time, days

Te

mpe

ratu

re, d

eg C

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Rad

iato

r ef

fect

ive

emis

sivi

ty

T400 - Primary MirrorT405 - S/C thermal i/fT410 - S/C louvered radiatorRadiator effective emissivity

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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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THE END

solar orbiter eus: thermal design considerations bryan shaughnessy, rutherford appleton laboratory 1...

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