CRaTER Thermal Analysis

Transcription

1 CRaTER Thermal Analysis Huade Tan

2 Contents System Overview Design & Requirements Inputs and Assumptions Power Dissipations Environment and Orbit Current Model Results Instrument temperatures Orbital temperature ranges Conclusions Uncertainties and Improvements Thermal Engineering 2

3 Design Approach & Requirements Design Approach Radiatively isolated with multi-layer thermal blanket over entire surface. Single layer blanket covering 10cm 2 telescope apertures nadir and zenith Tight conductive coupling to spacecraft optical bench Interface Requirements at Instrument Mounting Surface Survival 50 C -40 C Operational 35 C -30 C Thermal ICD para 6.1 Rate-of-Change n/a Thermal ICD para 6.2 Gradient n/a Thermal ICD para 6.3 Thermal Engineering 3

4 Model Requirements The CRaTER thermal model is required to represent, with as much detail as possible, the behavior of critical reference points in the CRaTER instrument in a computer simulated mission orbit environment in order to anticipate and correct for any possible hardware degradation or failure under similar circumstances. In order to ensure the survival of the CRaTER instrument, the thermal model should account for the worst case scenarios in both hot and cold temperature limits. The model must adhere to all RGMM and RTMM requirements given in the TICD. Thermal Engineering 4

5 Instrument Power Consumption Power dissipations in the instrument are modeled as heat loads. The relevant values of such heat loads are given in the following table. Hot case numbers are taken to be the maximum power consumption of each electrical component. Cold case numbers are assumed to be 80 % of the nominal power consumption of each electrical component. Hot Case (W) Nominal (W) Cold Case (W) digital board analog board V power supply V power supply telescope total power Thermal Engineering 5

6 Surface finish properties: Coating Location MLI and Optical Bench Cold Case Hot Case Absorptance? S Emittance? H Absorptance? S Emittance? H Kapton 3mil Black Kapton 3 mil Germanium Black Kapton Silver Teflon (5 mil) 3,4 MLI Blanket Silver Teflon (10 mil) 4 MLI Blanket Effective emittance: e* for MLI assumed to be.005 or.03 for best and worst cases. Modeled optical bench temperature are +35 hot case and 30 C cold case, which at least bounds the problem; we do not have a current best-estimate predict for the optical bench. Thermal Engineering 6

7 Orbital Heat Rate Factors: Environmental Parameters Hot Case Cold Case Solar Constant 1450 W/m W/m2 Albedo Factor Planetshine/Infrared Emission Lunar surface IR constants modeled after the characteristic Lambertian surface having a subsolar temperature of 1420 w/m 2 hot case and 1280 w/ m 2 cold case to a shadow IR emission of 5 w/m 2 for both cases.. Surface IR emissions across the bright side are described in the General Thermal Subsystem specification 431-SPEC Thermal Engineering 7

8 Orbit The current instrument model is assumed to be in a basic polar orbit at a cold case altitude of 70 km and a hot case altitude of 30 km. At a Beta angle of zero, the model simulates the operational worst case scenario where the instrument cycles from one temperature extreme to the other. The total heat absorbed (solar, albedo & IR) by the instrument through each orbit is computed using the Radcad Monte Carlo ray trace method. Thermal Engineering 8

9 Orbit The latest spacecraft geometric model received from GSFS (as seen to the left) corresponds to version E of the spacecraft layout and is subject to change. Subsequent changes to the optical bench layout directly affect the instrument s view factor to space and the heat load on the instrument through each orbit. Given the latest results of the model, minor changes in heat loads should not generate significant changes in temperature. Thermal Engineering 9

10 Current Instrument Model The coordinate system used in the CRaTER model corresponds with the reference coordinate system of the spacecraft as outlined in the TICD. The current instrument model consists of 69 nodes and 51 surfaces. Thermal Engineering 10

11 Current Instrument Schematic The CRaTER instrument is divided into three distinct radiatively coupled regions. Each housing consists of an isolated PCB or group of PCBs and a specific power dissipation as described in the model inputs. Thermal Engineering 11

12 Mounting Footprint CRaTER s current design mounts to the spacecraft at six points located at the base of the electronics box. Each modeled mounting plate is scaled to adjust for the true contact surface area. The model assumes a contact conductance between the mounting feet and the optical bench of 0.5 W/cm 2 C. The surface finish of the instrument panel directly facing the LRO is assumed to be anodized aluminum with an emissivity of 0.6. Thermal Engineering 12

13 Results: Instrument

14 Hot Case Heat Flows Watts time (s) ymin ymax interface zmin xmin xmax zmax film zmax film zmin xmax xmin ymax ymin zmax zmin Thermal Engineering 14

15 cold case heat map 100 Watts ymin ymax interface zmin xmin xmax zmax film zmax film zmin xmax xmin ymax ymin zmax zmin -100 time (s) Thermal Engineering 15

16 Orbital Temperatures (hot case) Temp (K) CR_FILM.T1 CR_FILM.T2 CR_BPCB.T4 CR_INT.T1 CR_MOUNT.T1 CR_XMAX.T1 CR_XMIN.T1 CR_YMAX.T1 CR_YMIN.T1 CR_ZMAX.T1 CR_ZMIN.T time (s) Thermal Engineering 16

17 Orbital Temperatures (cold case) Temp (K) CR_FILM.T1 CR_FILM.T2 CR_BPCB.T4 CR_INT.T1 CR_MOUNT.T1 CR_XMAX.T1 CR_XMIN.T1 CR_YMAX.T1 CR_YMIN.T1 CR_ZMAX.T1 CR_ZMIN.T time (s) Thermal Engineering 17

18 Results Summary CRaTER is driven by the temperature of the optical bench. Hot Case Max Operating Temperature [optical bench at 35C] Cold Case Min Operating Temperature [optical bench at -30 C} instrument interface 36 to 37C -29 to -30C pcb's 38 to 39C -27 to -29C nadir 36 to 38C -28 to -30C scope 36 to 38C -28 to -30C Instrument Internal temperatures vary <5 C from the optical bench temperature between extremes of hot and cold. Thermal Engineering 18

19 Summary and Conclusions Estimate of Internal Temperatures: Maximum internal temperatures are no more than 4 degrees C above the interface temperature. Need to run model with optical bench predicts to establish flight margins. Uncertainties and Modeling Improvements: Temperature dependence of material properties: variations in thermal conductivity can be neglected given an instrument temperature fluctuation of no more than a few degrees C through the beta 0 orbit. Incorporating TEPs into the thermal model Incorporating actual circuitry details on the PCBs Fine tuning MLI optical characteristics Thermal Engineering 19

20 Thermal Engineering 20

21 Backup Slides

22 Inputs Thermal and Physical properties: Material k (W/m/K) Cp (J/kg/K) rho (kg/m^3) e* Aluminum PCB mil Black Kapton Film MLI Optical Properties: Material a e Aluminum PCB mil Black Kapton Film Thermal Engineering 22

23 Material properties: Assumptions Thermophysical properties of Al-6061 were taken from Matweb databases Optical properties of Aluminum obtained from Cooling Techniques for Electronic Equipment: Second Edition MLI assumptions: Currently modeled using bulk properties PCB assumptions: 2 ground and power layers (80% fill) and 4 signal layers (20% fill), 1 mm total thickness PCB properties determined at TEP assumptions: Currently not modeled Thermal Engineering 23

24 Assumptions Conductive Resistances: Interface characteristics between PCB and Aluminum are assumed to be of copper to aluminum in vacuum at 30 C referred to in Heat Transfer. Holman, J.P Surfaces of the Ebox are assumed to behave under characteristic conduction of Al-6061 (assuming that the ebox is constructed out of a single block of aluminum) Conductive resistances are modeled between the top and bottom covers of the ebox, and the interface between the ebox and the telescope assembly. Internal Radiation: View factors between internal surfaces determined by Radcad using radk ray trace method Emissivity factors are calculated assuming either infinite parallel planes or general case for two surfaces from PCBs to the interior walls. Heat Flow to the Space Craft: Assuming interface temperatures of 30 and 35 degrees C Contact conductance of mounting feet to LRO assumed to be.5 W/cm 2 C Radiative heat transfer to the LRO through 15 layer MLI Thermal Engineering 24

25 Cold Case Orbit (bright to dark) Thermal Engineering 25