Sentinel 1 Spacecraft and SAR antenna thermal design, analysis, verification and flight performances

Transcription

1 45th July 2015, Bellevue, Washington ICES Sentinel 1 Spacecraft and SAR antenna thermal design, analysis, verification and flight performances M. Perellón 1 and R. Alvarez 2 Airbus Defense and Space, Madrid, Spain P. Petrini 3 Thales-Alenia Space, Rome, Italy Arne Sauer 4 Airbus Defense and Space, Friedrichshafen, Germany and S. Dolce 5 ESA-ESTEC, Noordwijk, The Netherlands Abstract Copernicus is the European Programme for monitoring the Earth to better understand how our planet and its climate are changing, the role played by human activities in these changes and how these will influence our daily lives. As part of the Copernicus space component, the Sentinel-1 mission is composed by two spacecrafts (A and B) each carrying an imaging C-band SAR instrument providing all weather, day-and-night images for land and ocean services and ensuring data continuity of ERS and ENVISAT SAR types of mission. Each Sentinel-1 (S1) spacecraft is designed for an operational lifetime of 7 years with consumables for 12 years. The Sentinel 1 spacecrafts are commissioned by ESA with Thales Alenia Space Italy as prime contractor, Airbus DS Germany as main contractor for the SAR and Airbus DS Spain as subcontractor for the Platform Thermal Control. The paper describes the main features of the S1 spacecraft, in particular the platform, SAR antenna thermal design, analysis and verification. The platform thermal design is passive supplemented by heaters. The SAR antenna thermal design has to cope with a cyclic variation of the dissipation and is based on heat rejection through the antenna front side. The rear side, exposed to seasonal sun inputs, is insulated by MLIs. Heaters driven by thermostats are used in the non-operational phases only The platform and the SAR thermal design performances have been verified by a combination of thermal analysis and separate, dedicated thermal balance tests. A spacecraft overall Thermal Balance test has not possible because - in stowed conditions the SAR wings completely block the Platform radiators - in deployed conditions the 12.3 x 1 [m] SAR antenna does not fit in any facility. The first Sentinel-1 spacecraft has been launched on a Soyuz rocket from Kourou on 3 April The second Sentinel-1 spacecraft (S1b) is foreseen to be launched ~ 18 months after the S1a. The in-flight performances of the S1a thermal control are nominal. All temperatures are well within the specified limits. 1 Thermal Engineer, Thermal Control Department, marc.perellon@eads-astrium.net 2 Senior Thermal Engineer, Thermal Control Department, raquel.alvarez@eads-astrium.net 3 Senior Thermal Engineer, Competence Center Platform & Integration / Thermo-mechanical Engineering Department, pierluigi.petrini@thalesaleniaspace.com 4 Thermal System Engineer, Thermal Engineering (TSOEM23), arne.sauer@airbus.com 5 Senior Thermal Engineer, TEC-MTT, silvio.dolce@esa.int

2 MLI a ε SVM PLM PPM SAR S/C RCT BOL EOL DSHA Nomenclature = Multilayer Insulation = Absorptivity = Emissivity = Service Module = Payload Module = Propulsion Module = Synthetic Aperture Radar = Spacecraft = Reaction Control Thrusters = Beginning of life = End of life = Data Storage and Handling Assembly I. Introduction S1 is a three-axis stabilised spacecraft. It operates in a sun-synchronous repeat-track low earth orbit at about 700 km of altitude; the repeat cycle is 12 days with 175 orbits. Figure 1: SC stowed configuration The spacecraft mass is 2300 kg including 130 kg of propellant. The S/S general layout, figure 1, is based on a body with parallelepiped shape, compatible in size, including the appendages, SAR antenna, solar array, with the SOYUZ launcher. The Platform configuration is based on the Thales Alenia Space Italia PRIMA multipurpose platform concept, also used for the 4 spacecrafts of the COSMO-SkyMed constellation (ASI) and in Radarsat-2 (CSA). The platform comprises three main modules, which are structurally and functionally decoupled to allow for a parallel module integration and testing up to the spacecraft final integration. 2

3 The modules are: Service Module (SVM), carrying all the bus units Propulsion Module (PPM), carrying all the propulsion items connected by tubing and connectors; Payload Module (PLM), carrying the SAR Instrument antenna The spacecraft attitude is measured by Sun, star, gyro and magnetic field sensors, and a set of four reaction wheels and three torque rods are used for attitude control. The spacecraft is equipped with two solar array wings capable of producing 5900 W (at end-of-life) to be stored in a modular battery. The orbit is accurately maintained by a dedicated propulsion system utilising hydrazine propellant and catalytic reaction thrusters. Figure 2: SC SAR deployed configuration Figure 3: Solar panels stowed configuration The S-1 payload, figure 2, is a multi-mode dual polarisation C-band SAR operating in Interferometric Wide Swath and Extra Wide Swath), STRIPMAP and W AVE (sampled STRIPMAP) modes. The main characteristic of the SAR antenna are : Active phased array antenna providing fast scanning in elevation and in azimuth Dimensions : 12.3 X 0.82 m 2 Mass : < 920 kg Transmitted peak power : 4400 W Center Frequency : 5405 MHz The main characteristic of the S1 platform are summarised in table 1. 3

4 Table 1. Platform Characteristics II. Platform thermal design and analysis The Sentinel 1 TCS S/S thermal design, as well as the equipment accommodation, has been established by taking advantage of experiences from Radarsat-2 program, which is a bus commissioned by MacDonald, Dettwiler and Associates (MDA) to TASI, and COSMO-SkyMed program, which is a 4-satellite constellation commissioned by Agenzia Spaziale Italiana (ASI) to TASI. As for the above previous programs, the heat rejection of the spacecraft body is achieved primarily through velocity, anti-velocity and anti-sun panels (radiator panels) on which, due to their minimum incident orbital flux, the high power dissipation equipment are mounted. Other lower dissipation equipment have to be installed on internal panels and their temperatures are radiatively controlled by the radiators. All internal equipment, which are mounted on spacecraft lateral panels, have the thermal filler under their baseplate/feet in order to maximise to conductive coupling with the mounting panel and therefore with the radiator. The selected thermal filler is the SIGRAFLEX Type F03510TH 0.35mm thick, figure 4. The base of this material is natural graphite. Principal properties are : anisotropic electrical and thermal conductivity, stability at high temperatures and reflectivity of thermal radiation. It can be used in air at temperatures to some 400ºC and in a reducing or inert atmosphere at up to more than 3000ºC. Figure 4: Sigraflex material 4

5 For the battery modules, the CHOTHERM 1671 is selected for electrical insulation. It is a silicone elastomer, precisely filled with a controlled dispersion of boron nitride particles to provide superior thermal and electrical performance characteristics. Reinforced with fiberglass cloth, it offer maximum resistance to tear, cut-through, and punctures due to burrs and other mating surface irregularities. For the Reaction Wheels, the silicon grease NUSIL is used. This material has an excellent resistance to moisture, ozone and oxidation over a wide temperature range, low outgassing and minimal volatile condensables under extreme operating conditions to avoid condensation in sensitive devices, high thermal conductivity, high and low temperature stability, and low bleed. The internal side of the lateral panels, has been painted with Chemglaze Z307, see figure 5. The units baseplate, as well as forbidden areas for harness supports, have not been painted. The paint is a black polyurethane conductive paint and it is used in order to obtain an isothermal internal spacecraft temperature and satisfy the superficial conductivity requirements. It exhibits low outgassing properties in high vacuum environments and provides excellent performance on rigid or flexible substrates. Figure 5: Internal panels view Several heat pipes, see figure 6 and 7, of various shapes are mounted under the high dissipating equipment (i.e. PCDU, DSHA and ICEs) in order to spread their heat load to larger areas of radiator panels. In particular, the average, sizing case, thermal dissipation specified for the PCDU, DSHA and ICE is respectively W, W and 173 W. Heat pipes are mounted with thermal filler in between equipment and heat pipe, and between heat pipe and mounting panel. The selected thermal filler type is SIGRAFLEX. Two different types of Axially Grooved Heat Pipes (AGHP) of constant conductance made from Aluminum Alloy extrusion and filled with high purity ammonia ( %) as working fluid are used. The type used for PCDU and DSHA has 30mm (total wings width). In the +Y PLM panel, heat pipes with width equal to 60mm must be used for better accommodation of the ICE s mounting feet. The outer diameter of the heat pipes is 12,5 mm. The flanges are drilled and machined according to the I/F necessities. Figure 6: Heat pipes For the electronic units located on heat pipes, spacers are mounted between heat pipes flanges in the units fixation bolts position to relieve the heat pipe from any mechanical loads. 5

6 Figure 7: Heat pipes on panels Several pure Aluminium doublers 3 mm thick, figure 8, are mounted to spread the heat of the equipment (i.e. SMU, SBTs, GPSREs and MODs) and to absorb the peaks dissipated by equipment (batteries, EPCs and TWTs). Doublers are mounted with thermal filler in between equipment and doubler, and between doubler and mounting panel. Figure 8: Aluminium doublers on panels The external spacecraft body with relevant thermal hardware (i.e. SSM and MLI) is shown in Figure 3. The external sized radiator area of the velocity, anti-velocity and anti-sun panels is covered with flexible SSM, whose optical features (e.g. low solar absorptivity, high infrared emissivity, limited properties degradation with respect to the mission profile) are exploited to ensure a good heat rejection capability of internal equipment thermal 6

7 dissipation towards the deep space. In particular, the SSM thermo-optical properties at BOL/EOL are shown in Table 2. Table 2 Thermo-optical Properties ε = 0.77 α BOL = 0.15 α EOL = 0.21 Absorptivity BOL and Emissivity values have been measured before and after the thermal test without differences. The Absorptivity EOL value had obtained from previous programs with available data. Figure 9: Radiators Figure 10: Grounding for radiators The SSM foils are attached to the panel by means of stripes of adhesive. For each SSM radiator, the grounding is achieved by two strips of electrically conductive adhesive (9703 PSA) as shown in figure 10. The SAR Antenna Support Structure (SAR-ASS) has an impact on the platform radiators (see also figure 11). This issue is discussed in detail in Ref.1 (AIAA ). MLI blankets are used to insulate, minimising temperature variations, all the external surfaces including the panel areas around the thrusters (RCTs). In addition, the MLI blankets are also used to insulate some internal equipment (i.e. battery, propellant tank and propulsion piping). Figure 11:: external MLI and SAR Antenna Support Structure 7

8 Figure 12: MLI types Six types of MLIs are used, see also figure 12. The main difference is in the number of layers and type of material,mylar or Kapton, separated by Dacron net, which depends on which temperature level the MLI has to experience. the External MLIs cover all the external structure surfaces beyond the radiator areas. This type of MLI is also be used to cover appendages The external layer of external MLI blankets, see also figure 13, is made of Kapton coated with electrically conductive ITO and SiO X in order to prevent ElectroStatic Discharge (ESD) and Atomic Oxygen (AO) degradation. Each cut in the MLI at the radiator area incorporate MLI flaps to allow the trimming of the radiator. In the thrusters, the Mylar is substituted by Kapton. The thermo-optical properties for the external face are : ε = 0.75 α BOL = 0.35 α EOL = 0.6 The BOL properties were measured previous and after the SC TB test without any degradation. Figure 13: External MLIs Panels and Thrusters Three types of internal MLI are used. One on the tank (aluminized both faces), another for the batteries panel ( aluminized externally and Kapton finish in the face looking to the batteries to avoid possible ElectroStatic Discharges) and another different type for the Propulsion piping. This last type of MLI is manufactured in stripes and is wrapped around the pipes. The SRM, STT and FSS thermal design, in which the relevant bracket is partially covered by MLI blankets with the exception of radiator. All other appendages are completely covered by MLI blankets or have a dedicated independent thermal design (i.e. SAR, SAW, XBAA, GPSA, SBA, OCP). Thermal washers are used to decouple them from the spacecraft body. The required thermal washers material type is Vetronit or Titanium. 8

9 For the Solar Array, the thermal control of the panels (wing) is obtained by means of natural balance between the bare carbon fiber on the rear side and the thermo-optical properties of solar array cells on the front side. SMU SBT GPSRE SBT The overall Thermal Mathematical Model (TMM) of the spacecraft has been established for the prediction of the equipment temperatures during the mission phases. Figure 14: External geometrical model The overall TMM, see also figures 14 and 15, is composed by the SPACECRAFT body detailed TMM and by simplified thermal models of the SAR antenna, appendages, Solar Array. It consists of 6863 thermal nodes; among them, 2588 thermal nodes are dedicated to the external part,including 1063 thermal nodes from the SAR antenna reduced model. 287 thermal nodes are dedicated to the internal equipment ( including units reduced models), 3922 to the structure and 66 to the LCT instrument. 9

10 Figure 15: Internal geometrical model Thermal analysis have been performed for the mission sizing cases and for specific mission phases i.e. Leop with deployment of Solar Array and SAR antenna. Table 3 presents a summary of the analyses cases. Table 3. Analysis cases Based on the analysis results, the radiator locations and size, heat pipes lay-out and power consumption of the heating system have been defined. In addition it has been requested to modify the Solar Array Yoke Support Structure coating to limit radiative interactions with the Platform radiators. Sensitivity analysis has been performed to quantify the uncertainties affecting the design and temperature calculation. This was important because, for some units, the temperature prediction was very close to the maximum allowed temperature. Table 4 shows the selected parameters used for the sensitivity analysis. Table 5 shows the changes in unit dissipations implemented for the sensitivity analysis. Table 4. Sensitivity parameters 10

11 VALUE MODEL FACTOR Radiator Alpha QS & QA *1.166 Antenna Support Structure Alpha QS & QA *1.125 Radiator Emissivity DR (99999)* Antenna Support Structure Emissivity DR (99999)* Radiator Area -5% DR (99999)*0.95 MLL Efficiency +50% Array MLL *1.5 MLL Efficiency -50% Array MLL *0.5 Lnternal MLL Efficiency -50% Array MLL *0.5 Panel thermal Conductivity in plane -30% DL*0.7 Panel thermal Conductivity out of plane -30% DL*0.7 Contact Conductance Unit to Panel -50% DL*0.5 Contact Conductance Unit to HP -20% DL*0.8 Contact Conductance Unit to Doubler -30% DL*0.7 Contact Conductance HP to Panel -20% DL*0.8 Contact conductance Doubler to Panel -30% DL*0.7 Unit Dissipation (>10W Nominal) +5% QL*1.05 Unit Dissipation (<10W Nominal) +10% QL*1.1 Note: Sun trapping, multireflection : flux + 10% for +/-X radiators affected by SAR ASS Table 5 Units dissipation variations UNIT BAT JUNC BOX CAPS SRM SRME PCDU SMU GEU CGYRO POWER MAX MAX NOM MAX MAX MAX MAX 0.00 FACTOR NA UNIT RWL's FSS MGM's MGT's GPRSE's STR ICU GEU PT POWER MAX MAX NOM MAX NOM MAX MAX FACTOR UNIT HYB SBT's S-BAND HYSOL WGSW TWT's EPC's MOD's XBAA POWER NOM MAX MAX MAX MAX NOM NOM MAX NOM FACTOR UNIT DSHA OMUX WG ICE's MDFE_Tx MDFE_Rx TGU DCU LIAU POWER MAX NOM NOM MAX NOM NOM NOM MAX MAX FACTOR III. SAR antenna thermal design The Sentinel-1 SAR antenna thermal design relies mainly on a passive design based on Multi-Layer Insulation and radiators, assisted by a software-controlled heater system close to the internal tile units. The challenge is to avoid high temperature limit exceedings during the high dissipative imaging mode, as well as low temperature limit exceedings during the cold Safe hold modes. The range of radiative heat discharge is thus limited in both directions. Based on TerraSAR-X experience, the heat rejection has been performed via the RF Waveguides, i.e. the tile front surface. The rear of the tiles is covered with Multi-Layer Insulation. In order to optimize the heat rejection and at the same time minimize the radiative absorption due to albedo reflection, the waveguides have been coated with white paint to achieve the necessary thermo-optical properties, see also figure

12 TPSU-1 Network Beam Cross Stiffener TPSU-2 EFE 01 TCU-B EFE 10 TA-A EMP TCU-A TA-B EPDN 1:4 Combiner Waveguide Radiator Panel Figure 16: Antenna Tile Configuration (left), White Waveguide Coating (right) The internal tile temperatures are significantly dependent on the thermo-optical properties of the radiator coating. For this reason the condition to trim the white paint thermo-optical properties has been preserved until the STM thermal test correlation, if the model correlation after thermal testing would show out-of-requirements. Nevertheless, the correlated values showed no necessity for trimming In case of the need for adjusting the thermo-optical properties, the blank copper sections around the long holes on the waveguides would have been adjusted in area, having influence on the total average infrared emissivity of the tile radiator surface. Thermal Waveguide Design and heat spreading doubler plate: The internal heat of the tile has to be discharged via the waveguide structure towards the external radiator surface, see figure 17. This is done via the waveguide walls, manufactured with carbon reinforced fiber material (HTA40). The thermal conductivity of this material is not very high (less than 1/10 of aluminium or high conductive carbon fibers) but had to be used due to mass and processibility reasons. Nevertheless for electrical and RF applications, the fiber had to be coated with copper. For this reason, the copper has been also used to increase the waveguide conductivity, thus the copper thickness has been applied with 20 µm during manufacturing, increasing the total material conductivity of the waveguides to a correlated value of 20 W/mK. In order to spread the heat dissipated by the EFE units, a thermal doubler called EMP (EFE Mounting Plate) has been glued to the internal waveguide surface. On this plate the EFE units and their foil heaters are applied. As material the high conductive carbon fiber K13C2U has been used. The originally assumed lay-up resulting conductivity of 200 W/mK could be confirmed by means of thermal model correlation after test. 12

13 Heat Transfer via Cross Stiffener Heat Transfer via EMP and Waveguides Heat Rejection to Space Figure 17: Concept of Heat Discharge via Waveguides Deployment Mechanism Thermal Design (figure 18) : The thermal design of the deployment mechanisms is also not trivial, for they are constantly exposed to sun and have a strict required temperature range, especially prior to deployment. The sun exposed drum shields have been beset with Second Surface Mirror foil (SSM), the rear part of the drum and the disc are black anodized in order to provide adequate heat rejection. SSM black anodized Betacloth MLI Figure 18: Thermal DEM Concept The harness has been protected with Betacloth sleeves, see figure 19, to prevent direct solar fluxes on the aluminium shielding which would cause very high temperatures. 13

14 Kapton/VDA Betacloth sleeve cable bundles with Al shielding fixation on harness clamps covered by MLI Figure 19: Betacloth Harness Shielding To maintain the necessary motor temperatures prior to deployment, software controlled heaters on the actuators have been used. After deployment they are deactivated. IV. Spacecraft verification The Spacecraft verification approach is based on separate testing for the platform and the SAR antenna. The Platform Thermal Balance (TB) test has been combined with the Thermal Vacuum (TV), Thermal Cycling test. The main objectives of the Thermal Balance test have been: Verification that the Thermal Control S/S design is adequate to maintain the SPACECRAFT equipment within the operating/non operating temperature requirements in the thermal environment simulated in TVAC chamber, representative of the expected flight thermal environment Provide SPACECRAFT equipment temperature measurements that satisfy the stabilization criteria in wellestablished electrical configurations and thermal environment to be used for the correlation of SPACECRAFT Thermal Mathematical Model The objectives of the Thermal Vacuum test have been: Verification of the SC performances during 3 thermal cycles (2 short cycle + 1 test cycle) Confirmation of the SC functionality when subjected to qualification temperature distribution derived in accordance to flight predictions The test configuration did not include the Solar Array and SAR antenna. It would have not been possible to include them in deployed configuration. Also, in stowed configuration, they would have completely blocked the heat rejection capability of the platform main radiators. The Battery modules were installed but not connected. In line with TAS-I heritage from all previous programs, the X-Band, S-Band, GPS Antenna were not installed and all RF tests were performed by connecting waveguides and coaxial cables to the relevant EGSEs. The SAR ASS was also not included due to schedule, programmatic reasons. As a consequence, the +/- X radiators heat rejection capability has been validated by analysis only. This, although far from being the preferred verification approach, has been eventually considered acceptable taking into account that, for flight, the maximum size of the radiators has been implemented and that the dissipations have been over-specified. The TVAC Chamber used for S1-A Thermal Balance and Thermal Vacuum Test is the HVT-60 of the Thales Italy Center for Small Satellites Integration located in Rome. In the test, the orbital fluxes have been simulated using the HVT-60 infrared heating system, that has the following main characteristics: 32 groups of Infrared (IR) ceramic heater plates: 14

15 24 groups on the chamber cylinder inner side 4 groups on the chamber door inner side 4 groups on the chamber rear inner side Electric power can be independently regulated from 0 to 2400 W IR heater plates temperature of each group can be independently regulated from 50 C to 500 C using pilot thermocouples (dedicated test campaign has been performed to correlate for each group temperature versus electrical power) Warm-up can be performed by setting temperature target of each group in manual mode or automatic (programmable) mode The layout of IR heating plates groups inside TVAC chamber is reported in figure 20: Figure 20: IR plates The S-1 overall TMM has been used to calculate the flight expected sink temperatures of radiator areas in the selected Cold and Hot On-orbit cases to be simulated in the TBT. For the TMM correlation a reduced TMM of the chamber has been used with the IR heating plates set as boundary nodes at the temperatures actually used during TBT and the chamber shrouds set as boundary nodes at LN2 temperature. The minimum infrared emissivity of IR heating plates has been evaluated equal to 0.8, the emissivity variation with temperature is reported in figure 21. Figure 21: Infrared Emissivity vs Temperature of various Materials (ceramic heater plates in red) 15

16 Figures 22 and 23 show the TMM of S1-A inside the HVT-60 chamber: Figure 22: IR plates on S/C model Figure 23: IR plates on S/C model inside thermal chamber 16

17 The S1-A PFM TB/TV test is divided into 20 main phases described in the following table: Table 6 Test Phases PHASE TITLE DESCRIPTION Phase 1 Initial Check Out P/L and S/S Equipment Power Consumption Reference Test is performed to verify readiness for TVAC test Phase 2 Pump Down & Shrouds Flooding Chamber depressurization is performed to required vacuum level with S/C in Lift Off electrical configuration; Chamber Shrouds are filled with liquid nitrogen (LN2) Phase 3 Hot Stress Chamber Heating System is switched-on to heat S/C equipment at maximum allowed temperatures to ease outgassing of S/C surfaces for cleanliness improve Phase 4 Cool Down Chamber Heating System is switched-off to cool down S/C equipment at Cold Start temperature level; test and flight heaters inside S/C are configured to stabilize the required temperature level Phase 5 Cold Turn On S/C equipment is switched-on when required Cold Start temperature level is reached Phase 6 Phase 7 Phase 8 Transient to Cold Thermal Balance phase-1 Cold TB phase-1 (Cold Calibration) Transient to Cold TB phase-2 S/C equipment is configured as required for Cold Thermal Balance phase-1; flight heaters inside S/C and guard heaters at Thermal Test Adapter I/F are configured as required to achieve stabilization S/C internal equipment is configured in Safe mode with P/L off to have minimum power dissipation; Chamber infrared heating system is off (worst not operative cold environment); stabilized temperatures are used to perform TMM cold calibration in not operative mode S/C equipment is configured as required for Cold TB phase-2; flight heaters inside S/C and guard heaters at Thermal Test Adapter I/F are configured as required to achieve stabilization Phase 9 Cold TB phase-2 S/C internal equipment is configured in On-orbit mode with P/L in Ready mode; Chamber infrared heating system is off (worst operative cold environment); stabilized temperatures are used to perform TMM cold calibration in operative mode Phase 10 Phase 11 Transient to Hot TB phase Hot Thermal Balance phase S/C equipment is configured as required for Hot TB phase; flight heaters inside S/C and guard heaters at Thermal Test Adapter I/F are configured as required to achieve stabilization S/C internal equipment is configured in On-orbit mode with P/L in Ready mode; Chamber infrared heating system is on to simulate orbital thermal fluxes absorbed on S/C radiators (operative hot environment); stabilized temperatures are used to perform TMM hot calibration in operative mode Phase 12 Short Cycle 1 S/C equipment is configured in On-orbit mode with P/L in Ready mode with equipment main chain active; no electrical verification test is performed Phase 13 Short Cycle 2 S/C equipment is configured in On-orbit mode with P/L in Ready mode with equipment main chain active; no electrical verification test is performed Phase 14 Transient to Cold Plateau S/C equipment is configured as required for Cold Plateau initial phase; test and flight heaters inside S/C and guard heaters at Thermal Test Adapter I/F are configured as required to achieve initial stabilization to start electrical tests Phase 15 Cold Plateau S/C equipment is configured in the various operative modes as necessary to perform the required TVAC electrical verification tests of P/L & S/S's; Chamber infrared heating system is off, test and flight heaters inside S/C are configured to achieve without exceeding minimum acceptance/qualification temperatures on equipment to be tested Phase 16 Transient to Hot Plateau S/C equipment is configured as required for Hot Plateau initial phase; test and flight heaters inside S/C and guard heaters at Thermal Test Adapter I/F are configured as required to achieve initial stabilization to start electrical tests Phase 17 Hot Plateau S/C equipment is configured in the various operative modes as necessary to perform the required TVAC electrical verification tests of P/L & S/S's; Chamber infrared heating system is on, test and flight heaters inside S/C are configured to achieve without exceeding maximum acceptance/qualification temperatures on equipment to be tested Phase 18 Warm-up LN2 is drained from Chamber Shrouds, Chamber internal environment is warmed to Ambient Temperature Phase 19 Pressure Recovery Chamber internal environment is repressurized to Ambient Pressure Phase 20 Final Check Out Final P/L and S/S Equipment Power Consumption Test is performed to compare measured values with initial Power Consumption Reference Test 17

18 The test profile is shown in figure 24: Figure 24: Thermal test profile The SC functional, electrical performances have been verified during the Cold and Hot plateau. The thermal cycling test has been completed by two short cycles during which no functional test have been executed. 18

19 The three thermal balance phases have been selected to be representative of in-flight spacecraft thermal environment: Phase 7 Cold TB-1 corresponds to SPACECRAFT electrical configuration in Safe Hold Mode and Cold environment Phase 9 Cold TB-2 corresponds to SPACECRAFT electrical configuration with P/L s in Operative Mode and Cold Environment Phase 11 Hot TB corresponds to SPACECRAFT electrical configuration with P/L s in Operative mode and Hot Environment The main findings in the TB test phases have been : A problem for two units of the Propulsion SS : the Pressure Transducer exceeded its lower temperature limit and the Liquid Filter its higher temperature limit. The root cause was due to a wrong thermal modelling that considered a conductive coupling of the Liquid Filter (LF) to the platform. On the contrary the Liquid Filter is conductively isolated and consequently its temperature sharply increases when the heater line is activated because of the flat heater mounted on its external surface. Since the same heater line warms also the Pressure Transducer (PT) but it is controlled by means thermistors triplet mounted on Liquid Filter, the results is excessive heating of Liquid Filter and insufficient heating of Pressure Transducer. The work around solution to allow the continuation of the test was the extensive use of test heaters mounted on SVM panels near the Propulsion Platform where the LF and PT are located. After the TB test, a rework has been implemented to solve the problem: Removal of the MLI on the Liquid Filter (remove the radiative isolation) Removal of foil heater from the Filter Relocation of thermistors on the Pressure Transducer Installation on the Filter of Clayborn heater ( same type used for all the pipes of this heating line) Foil heater equal to the one previously installed on the Filter, installed on the propulsion platform to maintain the same total resistance of the heating line. Analysis and assessment of thermal performances (test in clean room) Phase 7 Cold Thermal Balance 1 The SPACECRAFT is on average about 10 C warmer than the predicted. Phase 9 Cold Thermal Balance 2 The SPACECRAFT is on average about 10 C colder than predicted temperatures, which is the opposite of previous phase. Phase 11 Hot Thermal Balance The test temperatures are in agreement with predictions for this phase on PLM and Battery panel and colder than predicted of about 10 C on SVM. Measurements of Radiators size and thermo-optical properties have been performed before and after the TB, TV test to verify the values used in the TMM and any degradation effect due to test execution. No degradation has been observed. The correlation work was initiated with an audit of the GTMM used for the test predictions. This audit showed that in some cases, the test heaters associated to some nodes as QR where automatically deleted by the flight heaters QR associated to the same node. 19

20 In the correlation exercise, the main changes introduced in the model have been : a. introduction of a reduced model (RTMM) of the DSHA. This is a large unit with a significant dissipation. JUSTIFICATION: The RTMM of DSHA has been introduced in the model to better assess its temperature and gradients. b. Hemispherical Emissivity vs Normal emissivity : the emissivity of the SSM has been reduced from 0.8 to 0.77, with the exception of the BTAs panel radiators due to the fact that there is 3% less area of radiator than in the reality. JUSTIFICATION: Looking at the mean temperatures of the PLM and SVM a global deviation of 5 degrees has been observed in all the cases. This global deviation is estimated as a decrement of the heat rejection capacity of the spacecraft. The size of the radiator has been reviewed and the nominal dimensions were OK except the batteries with 3% difference. The emissivity of the radiators was measured with a value close to (normal emissivity). For this value the ratio between hemispherical and normal is in the range of In the mathematical model the emissivity was 0.8. The 5 degrees deviation was absorbed changing the radiator emissivity to 0.77 for all radiators except the batteries that it was maintained 0.8. c. Platform radiative model divided in two parts ( SVM internal, PLM internal) JUSTIFICATION: During the correlation exercise of the Propulsion system it was discovered that the radiative couplings, GRs, were incorrect due to overpassing of the maximum model memory. To avoid this problem the internal GMM model was divided in two models, one corresponding to the PLM enclosure and other one for the SVM enclosure. d. Thermal nodes of doublers checked and modified to have the real size. JUSTIFICATION: The doublers size was reviewed finding small changes in SMU and other units of the Communication panel. The updated size and corresponding conductive model was modified e. local conductive couplings ( HPs, doublers, on the panel) JUSTIFICATION: To match the test results, the conductive couplings of heat pipes, doublers and units were modified. For heat pipes the contact conductance was changed from 1000W/m 2 K to 700W/m 2 K (DSHA and ICES) and 500W/m 2 K for the PCDU due to the small HPs density. The effect on unit temperature is in the range of 1 degree. For doublers the general value of 1000W/m 2 K was maintained. f. Flight heaters, node heat distribution. JUSTIFICATION: Flight heaters node assignment has been checked/modified to take into account border effects (in case the heater is active on several nodes) g. Efficiency of MLIs. JUSTIFICATION: MLI efficiency for the SC - Y side (exposed to sun) has been decreased (disturbance factor from 4 to 2) during the correlation. It has been maintained with a factor 4 for flight due to the fact that this is conservative. h. STT internal model ( coupling between baffle and electronics) JUSTIFICATION: Full recalculation of conductive couplings of the supporting brackets and interfaces with lateral panel and shear panel i. MGT3 error in the position of the SPACECRAFT checked and modified on the internal model 20

21 JUSTIFICATION: The MGT3 position in the SVM was in an incorrect location due to a mistake of the axis between ICD and model j. Conductance coupling between panels ( brackets and inserts) JUSTIFICATION: Flight predictions and test predictions did not consider any inter-panels conductance. Due to the test results conductances were introduced. The modification improves the correlation of the batteries (major unit element responding to the change) and the units mounted close to the panel edges. The correlation exercise has been performed with minor changes in the thermal parameters of the Flight GTMM. The only significant change is on the external radiators emissivity and on the conductances between panels. V. SAR verification Thermal Test and Verification Campaign Test Set-up and challenges: In order to verify the thermal model and the thermal design, a detailed thermal test campaign has been performed. The thermal model was correlated on Tile STM and Centre Panel PFM basis. Tile STM Infrared Radiator for rear MLI thermal environment Figure 25: Tile STM in thermal chamber Due to the design status of the STM, black MLI had been used as it can be seen in Figure 25. Falsifications of results were avoided by controlling the rear MLI temperature by means of an infrared radiator, and not by application of a representative fix infrared flux as originally planned. Figure 26: Centre Panel PFM in chamber The Centre Panel PFM, see also figure 26, has been used only to correlate the effective heat transfer at the panel interfaces to the spacecraft. For hot orbit simulation with RF activity and the final READY mode verification (confirmation, that no heaters will switch during operational modes) with active units, the EQM test has been used. 21

22 RF Absorber with cold plate attached to the rear Figure 27: EQM in chamber with RF absorbers A significant challenge during the EQM test was the effects caused by the RF absorbers inside the chamber, figure 27. They caused a very long outgassing time due to the high absorber surface (about 3 times as high as during the STM test without absorber). Also the RF absorber surface was heated by the rear infrared rig which should affect the MLI only. This led to the necessity to drive the shroud to very low temperatures to generate the necessary sink temperature for the waveguide radiators. The further PFM and FM tests of the antenna parts have been performed on Wing basis (two attached antenna panels as shown in figure 28). Figure 28: Antenna Wing prior to testing The folded configuration of the Wing caused an insulation of the tile radiators towards the chamber shroud. As a consequence very long cool down times of the test sample were predicted (warm up could be performed by means of internal flight heaters). To speed up transitions, the cool down ramp was driven by gaseous Nitrogen, injected into the chamber prior to cool down, see also figure 29. The gaseous Nitrogen was evacuated prior to heater activation in order to avoid corona effects. Figure 29: Chamber pressure and temperature curve including GN2 flooding for cool down acceleration 22

23 Test and correlation outcomes: Following results and outcomes were gained in the course of the test campaign and the thermal model correlation afterwards: The Waveguide Infrared Emissivity was significantly higher than originally predicted, it had to be correlated from to 0.88 The EMP conductivity of K13C2U with originally assumed 200 W/mK was confirmed The cross stiffener conductivity was correlated to a higher value of 66 W/mK instead of originally assumed 44 W/mK The waveguide conductivity was correlated to a slightly lower value of originally assumed 24 W/mK to 20 W/mK The in-flight predictions with implemented correlation factors showed no out-of-requirements. VI. Spacecraft in flight performances The spacecraft in-flight thermal performances are well within expectation. All temperatures and heater power demand are well within the limits. Heaters when necessary have been activated during the launch, Leop phases. The heater lines activated are those with "high" set points, in particular RCS propulsion lines 22 to 24 [C] typical, Thruster brackets 32 to 34 [C], batteries 14 to 16 [C], large radiator. A successful health check of the remaining heater lines not activated has been performed in the first days of SC commissioning. A comparison between actual temperatures and flight predictions has been performed. This assessment has identified some minor discrepancies. In particular : a) the PCDU in flight temperature is much lower than predicted. The unit is accommodated on heat pipes and has a dedicated radiator.the discrepancy can be explained by : - unit dissipation used in the analysis is much higher than actual - poor temperature correlation achieved in the TB test. There is a residual discrepancy of ~ 8 [C] between the post-test correlated and measured temperatures ( TB test phase 9, Cold operating) b) the average temperature of the batteries is slightly higher than the required 15 [C]. There are 10 battery modules. For each module, the specified design temperature range is 10 to 40 [C] but, for performance reasons, the average temperature of all modules has to be maintained ~ 15 [C] throughout the lifetime. The modules are installed on doublers and have dedicated radiators. The radiator area is the maximum available on the panel. Heater power is required to maintain the minimum operating temperature. The slightly higher in flight temperature can be explained by the on-off logic selected for the heater control and the (relatively) high installed heater power ( 94 [W] at 65 [V]). c) the RW 1,2 temperatures are slightly colder than predicted while the RW 3,4 temperatures are in line with the predictions for the cold case. The Reaction wheels are accommodated on internal structural panels and are controlled by radiation to the SC internal environment. There is not yet a convincing explanation for the RW 1,2 lower temperatures. d) the Star Tracker temperatures are slightly warmer than predicted. They are accommodated on a dedicated bracket and have dedicated radiators. The discrepancy can be explained by the relatively complex design i.e. the radiator is a separate item installed between the unit and the bracket. Also large temperature deviations were present, for some test cases, after model correlation. SAR In-Orbit behavior and in-flight correlation: Comparison of flight data with analysis data: The first data received after launch showed that the heater systems were functional and able to control the temperature at the foreseen heater set points without exceeding the maximum allowed average heater power per orbit as required. 23

24 Data measured during the commissioning phase showed that the long term READY mode temperatures were well within the predicted range and well above the calculated temperatures. Heater activation during operational modes is thus not expected, as required. The model has been correlated by means of adjusting the radiative environment based on the season during measurements performed after a 24 hour swung-in 25% imaging mode to 75% WAVE mode condition. The infrared radiator emissivity has slightly been adapted to reproduce the measured temperatures. Figure 30: Measured in-flight TRP temperatures (coldest EFE) of all Tiles Figure 31: Correlated Tile thermal model EFE temperatures (TRP (coldest EFE) correlated to hottest measured TRP in Figure 30) Higher Imaging Mode Duty Cycle assessment: The correlated model allowed an assessment of maximum possible imaging mode time during Begin of Life conditions. The assessment has shown that the theoretically possible imaging mode time could go up to 42 minutes, which is well above the required 25 minutes. Nevertheless it has to be considered that this is valid for the Begin of Life condition only. For End of Life conditions, the 25 minute limitation is still valid based on the currently unknown degree of degradation expected at End of Life conditions. TPSU Dissipation: One important finding of the measured in-flight data evaluation was the fact that the TPSUs operate at more than 10 C below the analytically predicted value. This was investigated closely and it was traced back to the unit dissipation which apparently was lower than assumed. This is very likely justified by the higher operational voltage (66 V instead of 50 V) which, with constant 24

25 resistance and total consumption, leads to a lower unit current. The unit dissipation is very dependent on the current, for this reason it is assumed, that the lower operating current leads to this effect. Nevertheless, this is uncritical. It is even helpful, because in the course of the analysis, the upper TPSU operating temperature was the duty cycle limitation. There has therefore been risk mitigation due to this effect, as well as the opportunity in driving even higher imaging mode duty cycles. Figure 32: Correlated in-flight thermal model TPSU temperatures (between 40.2 C and 41.9 C predicted) Figure 33: Measured in-flight TPSU temperatures (between 28.9 C to 32.1 C measured) Generally it can be said, that the measured in-flight data fits well to the predicted values of the mathematical thermal model. The thermal design showed its positive functionality in every way: All heaters worked well so far and kept the designated units within their temperature ranges without exceeding the maximum allowed average heater power The heat discharge by means of the waveguide conduction and radiator concept works well and provides the predicted range of operating temperature The correlated model under Begin-of-Life conditions shows, that under current circumstances a significantly higher imaging duration than the required 25 minutes is possible The lower operating temperature of the TPSU raises questions, but reduces mission risk, and also increases possible durations of the imaging mode 25

26 VII. Conclusions The Sentinel 1 TCS S/S thermal design development, manufacturing and verification has been a challenging activity, that required to solve several criticalities mainly due to complex interface between Platform and SAR antenna and equipment thermal constraints. The thermal balance tests, performed separately at Spacecraft and SAR antenna levels, were successfully conducted and demonstrated the capability of thermal design to meet the temperature requirements. In-flight performances of the S1a thermal control are nominal confirming the validity of the selected design and verification. 26