X-ray Calorimeter

1. X-ray Calorimeter Thermal Kimberly Brown 2 – 6 April, 2012

2. Thermal System Overview for S/C Bus • S/C bus is thermally isolated from FMA • Insulated MLI on exterior s/c bus, metering structure, backside of Solar Array, non-cell area on front side of Solar array • 15 layer make-up • Silver conductive composite coating ITO/SiOx/Al203/Ag • Low absorptance (0.08 at BOL) and high emittance 0.6 • Radiator Panel for Bus (located on anti-sun side) • Coating NS43G yellow paint • High emittance 0.9 • Heat Pipes (CCHPs) • Transfer heat from boxes to radiator panel • Heat pipes embedded in radiator panel to spread heat • Heater Control for Propulsion, Gimbals, Battery, Star Trackers • Mechanical thermostats (operational and survival) • Primary and redundant heater circuits • Two thermostats in series per circuit • Kapton film heaters attached to components

3. S/C Bus Thermal Control Subsystem Functional Block Diagram Thrusters RW RW RW RW RWE RWE RWE RWE Battery (Li-Ion) Radiators Reject Waste Heat to Space FMA PSE Comm System Avionics Gyro Propellant Tanks, Lines and Fill-and-Drain Valves Radiators MLI Thermostatically controlled heaters

4. Spacecraft Temperature Limits • Propulsion System • +10°C to 40˚C • S/C Components Electronics • -10°C to +40°C operational and -20°C to +50°C survival • Avionics, comm system • Solar Array Temperature - Operational -100°C to +106°C (10°C above predict) • CommSystem: • HGA, two axis Gimbal motors 0°C to 40°C • Antenna Dish -40°C to 65°C • Xband: Operational +10°C to 40°C • Li Ion Battery Temperatures - Operational 0°C to +30°C

5. S/C MEL to Include • MLI for Metering Structure • FMA isothermalized from spacecraft – heat pipes and MLI internal

6. Instrument Block Diagram Jettison cover

7. S/C Bus Thermal Control For FMA MLI on interior to radiatively isolate from FMA CCHP isothermalizes mounting interfaces for FMA (redundancy not shown) CCHP embedded in honeycomb radiator panel for spacecraft load

8. XMS Instrument Radiator Sizes Blue is updated configuration Sun MLI on Backside 1.0005 m2Total for 3 Panels (Slightly Canted) NS43G Paint on Front Side 0.7888 m2 NS43G Paint on Front Side 0.6324 m2 M. Choi, 4/4/2012

9. Radiators Sized for Mission S/C Radiator

10. Summary of X-Ray Cal Radiator Sizes s/c radiator based on 492 Power Watt load Updated in study this week Customer provided

11. Spacecraft Total Heater Power * Operational heater circuits

12. Heater Circuits Summary

13. Propulsion Subsystem Schematic X-Ray Calorimeter MLI tanks Mechanical thermostats Heater control for all Propulsion components P P N2 N2H4 N2 N2H4 Diaphragm Tanks F 4 N Thrusters FD Valve Latch Valve Pressure Transducer Filter F P

14. Propulsion System Heater Power

15. Propulsion System Thermal Design • Tank (2) • Conductively isolated tanks from Bus structure with isolation like aluminum support struts • Radiatively isolated by blanketing the tanks (6 layer) inside the S/C bus with a low emittance coating • Bottom deck covered with MLI blanket (15 layer) • Heaters for tanks thermostatically controlled, prime and redundantheaters with mechanical thermostats • Fuel Line Design • Fuel Lines assumed to be internal from tank to thrusters • All lines wrapped with heater elements spirally wrapped • Heaters are thermostatically controlled • All lines spirally wrapped with 5 mil aluminum tape with 50 % overwrap • Lines to be low ε taped then wrapped with MLI sleeve blanketing (15 layer) external • Zonal heaters • Thrusters (4) 1 lbf (4N) • Heaters thermostatically controlled • Prime heater per thruster with two mechanical thermostats per circuit • Thruster has MLI boot blanket cap with over-temperature outer layer for soak back • Current Heater Power Estimate for Propulsion 20.4 Watts

16. Instrument TB/TV test • Instruments test 8 hot/cold TV cycles. • Each Instrument conduct TV qualification and TB testing during instrument level testing of FMA, Calorimeter Dewar and Instrument Electronics Radiator. • Instrument’s Radiators and CCHPs tested fully at Instrument level test.

17. Verification of Thermal System • Perform Thermal Vacuum Thermal Balance Testing Per GEVS at System level. • Perform 4 Hot/Cold Thermal Vacuum Cycles • Perform Thermal Balance Tests Subjecting X-Ray Gratings to Worst Hot and Cold Case Conditions. • Verify Thermal Models, Perform Model Correlation to Test Data. • Verify Proper Operation and Design of Heater Circuits • Verify Proper Thermistor Calibration, Operation and Placement • Verify FMA Interface to S/C Interface • CCHP for Spacecraft Bus • Conduct detail TB tests for Spacecraft Bus during Observatory level testing. • Component Level Testing Shall be Performed by the Vendor Prior to Shipment to S/C Vendor: • Heaters, Thermostats, Thermistors • Electronic boxes test 8 hot/cold TV cycling • S/A deployment.

18. Issues / Potential Risks / Future work Issues – None Risks: • Risk minimized from previous studies for thermal interface definition • Instrument’s radiators are better defined (no dependency on spacecraft) so they can be tested for design at the Instrument level test Future work: • STOP analysis to determine gradients of the metering structure and the effects of any structural distortion • Layout of CCHPs to test in TV

19. Acronym List • MLI: Multi-layer insulation • CCHP: Constance conductance heat pipe • FMA: Flight mirror assembly • TCS: Thermal control system • BOL: Beginning of Life • HGA: High Gain Antenna • MEB: Main Electronics Box • CFEE: Control Front End Electronics • CE: Control Electronics • CDEEP: Calorimeter Data Processing Unit