41th International Conference on Environmental Systems, July 17-21, 2011, Portland, Oregon

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GMES Sentinel 1: Thermal Design and Verification Approach of the Thermal Control Subsystem

C. Bruno, M. L’Abbate, A. Panetti and D. Selci Thales Alenia Space Italia, Rome, Italy, 00131

and

S. Dolce European Space Agency (ESA/ESTEC), Nordwijk – The Netherlands

This paper describes the main features of the selected thermal design of Sentinel 1 satellite with a summary of the thermal analyses performed to optimize the passive Thermal Control Subsystem versus the satellite design drivers. Finally, an overview of the satellite verification approach is also described.

Nomenclature AO = Atomic Oxygen ASS = Antenna Support Structure BOL = Beginning Of Life CAPS = C-Band Antenna Power Supply CFRP = Carbon Fibre Reinforced Plastic DSHA = Data Storage and Handling Assembly ECE = Eads Casa Espacio EOL = End Of Life EPC = X-Band Electronic Power Conditioner ESA = European Space Agency ESD = ElectroStatic Discharge FSS = Fine Sun Sensor GMES = Global Monitoring for Environment and Security GMM = Geometrical Mathematical Model GPSRE = GPS REceiver box GPSA = GPS Antenna ICE = Integrated Central Electronics ITO = Indium Tin Oxide LEO = Low Earth Orbit LN2 = Liquid Nytrogen LVA = Launch Vehicle Adapter MLI = Multi Layer Insulation MOD = MODulator OCP = Optical Communications Payload PCDU = Power Conditioning & Distribution Unit PDHT = Payload Data Handling and Transmission PFM = ProtoFlight Model PLM = PayLoad Module PPM = ProPulsion Module PRIMA = Piattaforma Riconfigurabile Italiana Multi Applicazione RCT = Reaction Control Thruster RW = Reaction Wheel SBA = S-Band Antenna S/C = SpaceCraft

41st International Conference on Environmental Systems17 - 21 July 2011, Portland, Oregon

AIAA 2011-5165

Copyright © 2011 by Thales Alenia Space Italia S.p.A. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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SAR = Synthetic Aperture Radar SAS = SAR Antenna Subsystem SAW = Solar Array Wing SES = SAR Electronics Subsystem SBT = S-Band Transponder SMU = Spacecraft Management Unit SSM = Second Surface Mirrors SRM = SAW Reel Mechanism STT = STar Tracker SVM = SerVice Module TBT = Thermal Balance Test TCS = Thermal Control Subsystem TMM = Thermal Mathematical Model TWT = X-Band Travelling Wave Tube TASI = Thales Alenia Space Italia XBAA = X-Band Antenna Assembly wrt = with respect to

I. Introduction In the frame of the Global Monitoring for Environment and Security programme (GMES), the Sentinel 1 satellite

is an European polar orbit satellite for the continuation of SAR operational applications in C-Band, providing user services (e.g. mapping, monitoring and assessing, surveillance, etc) with continuous images, day and night, in all weather conditions. In particular, the system is capable to image all global land masses, coastal zones, ice areas and maritime transport zones (e.g. shipping routes) at high and medium resolution with C-band SAR operating in STRIPMAP and TOPSAR (Interferometric Wide Swath and Extra Wide Swath) modes. Oceans are acquired in WAVE mode.

The Sentinel 1 satellite is commissioned by European Space Agency (ESA) with Thales Alenia Space Italia (TASI) as Prime Contractor, Astrium Germany as subcontractor for the SAR instrument and Eads Casa Espacio (ECE) as subcontractor for the TCS S/S.

The selected orbit for Sentinel 1 satellite is a Near-Polar Sun-Synchronous orbit at an altitude of about 693 km. The foreseen nominal mission lifetime is 7.25 years.

Sentinel 1 satellite design as well as the test verification approach of the Sentinel 1 TCS S/S thermal design has been outlined in line with the TASI heritage, i.e. Radarsat-2 bus and COSMO-SkyMed constellation (Figure 1).

Radarsat-2 COSMO-SkyMed

Figure 1. TASI heritage

However, due to the presence of several design drivers (e.g. presence of fixed SAR Antenna Support Structure

(ASS) on ±X satellite sides which highly reduces the view factor of the ±X radiators, equipment allowable temperature ranges, etc), the optimized passive Thermal Control Subsystem (TCS) is particular challenging even if

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based on proven passive thermal control elements, i.e. coatings (paints, MLI blankets, SSM), heat pipes, doublers, heaters and thermistors.

Detailed thermal analyses of the sizing cases were performed in order to determine the equipment temperatures along the nominal mission lifetime. The selected tools are THERMICA and ESATAN softwares, using a Finite Difference method. The TCS guarantees that all the equipment predicted temperatures, which are calculated considering the uncertainty of TCS design, are within their allowable temperature ranges.

Finally, an overview of the test verification approach of the Sentinel 1 TCS S/S thermal design is presented.

II. Satellite Layout and Mission: Main Project Drivers The Sentinel 1 satellite is based on the Prime heritage of PRIMA multipurpose platform concept, which foresees

the S/C body divided in three main modules structurally and functionally decoupled to allow the parallel modules integration and testing activities up to the satellite final integration. The modules are:

• SerVice Module (SVM), which carries most of bus equipment apart from the propulsion ones; • ProPulsion Module (PPM), which carries all the propulsion components connected by tubing and

connectors; • PayLoad Module (PLM), which carries mainly all the payload equipment including the instrument SAR

antenna. Most of the PPM is enclosed in the SVM, being mounted into the cone section ending with the external Launch

Vehicle Adapter (LVA) on the bottom side of satellite, while the PLM is mounted onto the SVM allowing mainly the SAR Antenna accommodation.

The Sentinel 1 external satellite is shown in Figure 2.

Where:

Stowed Configuration Deployed Configuration

Figure 2. Sentinel 1 external satellite

Z

Flight direction

SAR Antenna

S/C body

Flight direction

Earth direction

Sun side

X

Y NOMINAL ATTITUDE

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The internal equipment allocation is driven by the satellite attitude in the overall mission phases during the mission lifetime.

The Sentinel 1 orbit is a Near-Polar Sun-Synchronous orbit at an altitude of about 693 km (LEO) and RAAN 18:00.

When the satellite is stowed, it is rotating around its Y axis with 2 revolutions per orbit (wrt an orbital reference frame, barbeque mode) while the XZ plane is in the orbital plane and -Y axis is pointing towards the Sun (Figure 3).

When the satellite is deployed in nominal operations (routine phase), it flies with the +X axis parallel to the velocity axis with a roll angle (steering law) of about -30° (SAR antenna pointing the dark side of the Earth) while the -Y axis is pointing to the Sun (Figure 3).

When the satellite is deployed in emergency phases (e.g. Safe Hold Mode (SHM) and Ultimate Safe Mode (USM)), it is stabilized in a gravity-gradient attitude while the -Y axis is pointing to the Sun (Figure 3).

Stowed Configuration

Deployed Configuration – Nominal Operations

Deployed Configuration – Emergency Phases

Figure 3. Sentinel 1 main mission phases wrt mission lifetime

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Summarizing, the Sentinel 1 satellite attitudes are characterized by having only one side of the satellite mainly exposed to the Sun (-Y side), as for some TASI past programs (i.e. Radarsat-2 bus, COSMO-SkyMed constellation).

Therefore, taking into account the above similarity and advantage of the previous satellite layouts, most internal equipment are accommodated on the other three sides of the satellite, i.e. the velocity and anti-velocity panels (±X sides) and on the anti-sun panels (+Y side), as shown in Figure 4.

Internal Equipment on the satellite lateral panels

Internal Equipment on the satellite central structure

Figure 4. Sentinel 1 internal equipment layout

SVM+X Battery panel

PLM-X PDHT panel

PLM+Y SES panel

DSHA

PCDU

ICEs

TWTs

Z

X

Y

Z

X Y

Z

X

Y

Flight direction

MODs

REMARK: +X panels are unfolded

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The Sentinel 1 SAR antenna is a large planar phased array antenna which is composed of a fixed central panel and two deployable wings. Due to the extended envelope of the antenna on the satellite in stowed configuration and in order to fulfill the mechanical satellite requirements (i.e. global stiffness and interface forces), the following two antenna design concepts are needed:

• a great number of antenna attachment points on the top panel (+Z side) and on most of the velocity and anti-velocity panels (±X sides);

• the presence of a complex fixed Antenna Support Structure (ASS) under the two deployable antenna wings (±X sides).

Even if the ASS temperatures are not relevant anymore after antenna deployment from the antenna point of view, the ASS±X (Figure 5) with several antenna attachment points and a complex frame of beams is the major constraint for the thermal design of the equipment on the velocity and anti-velocity panels, giving also an impact on the preliminary internal equipment layout. In particular, the shift of the modulators (MODs) in the PDHT panel was needed causing a general review of all the PDHT internal equipment positions; the final PDHT layout is shown in Figure 4.

Figure 5. Sentinel 1 SAR ASS±X

III. TCS S/S Thermal Design 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 S/C 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 (Figure 4). Other lower dissipation equipment have to be installed on internal panels (Figure 4) and their temperatures are radiatively controlled by the radiators.

The thermal design of all the equipment mounted on the S/C body and the following external mechanisms and optical sensors, i.e. SAW reel mechanisms (SRMs), star trackers (STTs) and fine sun sensors (FSSs), is under the TCS S/S responsibility, In particular the first group has a thermal design using the panel on which they are mounted while the second group has dedicated local thermal designs using the appendage on which they are mounted.

X

Y

Z

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The thermal design of the remaining external appendages (e.g. antennas, instruments, solar arrays, thrusters, etc) is under the relevant Equipment Supplier responsibility.

The TCS S/S thermal design is achieved by passive thermal control elements, ensuring the fulfilment of the mission requirements, thermal control performance requirements, thermal interface requirements and conforming to the environmental factors encountered during all mission phases. In particular the passive elements are thermal filler, paints, SSM, MLI blankets and thermal washers. The active elements are limited to heat pipes, doublers, heaters and thermistors.

For sizing the TCS S/S thermal hardware at hot, the worst measurement modes of the SAR Antenna and PDHT S/S have been selected: a SAR operational timeline comprising 25 minutes of Image mode (highest dissipation) and continuous Wave mode (lowest dissipation) for the rest of the orbit and a PDHT timeline of 30 minutes have to be taken into account. This results in a cyclic variation of the overall satellite thermal dissipation with an orbit average of about 1.0 kW.

The internal equipment layout with relevant thermal hardware (i.e thermal filler, paints, heat pipes and doublers) is shown in Figure 6.

-X +Y +X

Figure 6. Sentinel 1 internal thermal design for lateral panels

All internal equipment, which are mounted on satellite 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

= Doubler (GPRSE doubler and SBT doubler are not modelled in the GMM)

= Heat pipes (ICE heat pipes are not modelled in the GMM)

DSHA

GPSRE SBT SMU

GPSRE SBT

PCDU

MODs

EPCs

TWTs

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area. The selected thermal filler type for equipment except the battery modules is the SIGRAFLEX . For the battery modules, the CHOTHERM is selected for their electrical insulation.

The internal equipment, which are mounted on satellite central structure except the reaction wheels (RWs), usually do not need the thermal filler. For the RWs, the silicon grease NUSIL is used.

The internal side of the lateral panels, except the battery panel (SVM+X) since it is covered by MLI, is with black paint in order to obtain an isothermal internal satellite temperature.

Several heat pipes 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 thermal dissipation of PCDU, DSHA and ICE in nominal operations during the routine phase is respectively 176.3 W, 134.47 W and 173 W. Heat pipes will be mounted with thermal filler in between equipment and heat pipe, and between heat pipe and mounting panel. The selected thermal filler type is again the SIGRAFLEX.

Several Aluminium doublers 3 mm thick 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 will be mounted with thermal filler in between equipment and doubler, and between doubler and mounting panel. The selected thermal filler type is again the SIGRAFLEX.

The external satellite body with relevant thermal hardware (i.e. SSM and MLI) is shown in Figure 7.

+X +Y -X

Figure 7. Sentinel 1 external thermal design

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 dissipation towards the deep space. In particular, the SSM thermo-optical properties at BOL/EOL are shown in Table 1.

= SSM

= MLI

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Absortivity Emissivity

ααααBOL ααααEOL εεεε SSM (Radiators) 0.11 0.18 0.8

Table 1. SSM Thermo-optical properties

The presence of the ASS±X on satellite radiators made the sizing of the SSM areas on the velocity and the anti-

velocity panels extremely complex with several not foreseen iterations also in the ASS±X and SAW yoke thermal design. In particular, in order to preserve the satellite radiator functionality, the temperature of the ASS surfaces in the field of view of the radiators has to be maintained as low as possible. Initially the proposed ASS thermal design foresees the structure completely wrapped in MLI to control its temperature and its heat flux at S/C I/F. This proposal had to be discarded because the ASS beams on the sun side reach high temperatures at EOL (above +70°C) and the radiative heat exchange between the ASS MLI and the satellite radiators also causes high temperatures of the radiators of the high power dissipation equipment. The most affected equipment is the DSHA temperature at EOL (above +50°C, which is the maximum design operating temperature) due to its high thermal dissipation and its accommodation inside the S/C body. Then, the ASS thermal design was updated considering the structure with white painted beams and S/C I/F brackets covered with SSM foil. The presence of the white paint on the beams improved the temperature results for both beams and equipment but it was not sufficient to guarantee a successful radiator sizing due to the low radiator view factor towards the deep space (average value of 0.4, see Figure 8). In fact, even with the updating in the PDHT internal equipment layout, which was performed in order to improve equipment thermal dissipation distribution on the panel, the satellite was still too hot. Several equipment, i.e. DSHA, MODs, battery, LF and propellant tank, had negative temperature margins at hot and the DSHA had the most significant temperature negative margin (value of -5°C).

+X -X

Figure 8. Sentinel 1 satellite body view factor towards deep space

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Therefore, further revisions of the thermal design of the ASS beams were needed: the final ASS thermal design foresees the structure with beams covered with SSM and S/C I/F brackets covered with SSM foil.

During the above iteration loops on the SAR ASS, also the thermal design of the SAW yoke was revised being it a full CFRP structure close to the satellite radiators. As for the ASS, the radiative heat exchange between the SAW yoke and the satellite radiators influences the temperatures of the internal equipment mounted on the velocity and the anti-velocity panels. The most affected equipment is the battery panel due to its average temperature and temperature gradient requirements. However, due to its limited effect and after the decreasing of the battery thermal dissipations, the option to have the sun side of the SAW yoke white painted has been finally discarded.

The MLI blankets are used to insulate, minimising temperature variations, all the other external surfaces out of the radiators (unless specific trimming requirements on SSM to balance heat rejection), including the panel areas around the thrusters (RCTs) which may be subjected to high temperature (Figure 7). In addition, the MLI blankets are also used to insulate some internal equipment (i.e. battery, propellant tank and propulsion piping).

A MLI blanket is usually composed of various Mylar or Kapton aluminised layers (with low infrared emissivity) separated by Dacron net. The number of layers and type of material depends on which temperature level MLI has to experience. The external layer of external MLI blankets is made of Kapton coated with electrically conductive ITO in order to prevent ElectroStatic Discharge (ESD) and Atomic Oxygen (AO) degradation. Therefore, several types of MLIs are used for Sentinel-1 satellite.

The SRM, STT and FSS thermal design, in which the relevant bracket is partially covered by MLI blankets with the exception of radiator, are shown in Figure 9.

SRM (*) STT FSS (*)

(*) Only the thermal hardware of the relevant appendage is reported. The colors of the rest of the satellite is not coherent with the above colors.

Figure 9. Sentinel 1 SRM, STT and FSS thermal design

In addition, all other appendages are completely covered by MLI blankets or have a dedicated independent thermal design (i.e. SAR, SAW, XBAA, GPSA, SBA, OCP).

The thermal washers are used to decouple them from the satellite body. The required thermal washers material type is Vetronit or Titanium based case by case.

Finally, the temperature monitoring and control will be performed through the use of thermistors and heater lines. The thermistors triplets (according to the selected triple redundancy approach) from SMU and CAPS command the heater circuits respectively from PCDU and CAPS. In particular, both PCDU and CAPS feed and rule their automatically controlled heater circuits using a “flat”/hot redundancy policy in which all heater lines are active: in both PCDU and CAPS, half of them are the main heaters lines and the other half are the redundant ones, which have different activation/de-activation thresholds from the previous group in order to avoid simultaneous activations but still having the redundant heater lines ready in case of the main heater line failure.

= SSM

= MLI

= SSM or white paint

= MLI

= SSM

= MLI

= SRM MLI

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In general, the heater lines will be controlled via a dedicated set of thermistors, which are S/W-controlled from SMU, with temperature set-points re-programmable from ground. The temperature set-points will be set according the predictions and, when necessary, up-linked via TC on the base of the on-orbit telemetries.

The above thermistors triplets and the additional thermistors, which are available from SMU, also monitor the actual on-board temperatures.

For sizing the TCS S/S thermal hardware at cold, the worst less dissipative satellite modes have been selected: the USM at Beginning Of Life, the P/L in stand-by mode and the LEOP. Therefore, the position of the heaters and thermistors as well as the temperature set-point has been properly selected taking into account the analysis results. In particular, the definition of most of the heater lines occurs in USM at Beginning Of Life, which results in an orbit average overall satellite thermal dissipation of about 271 W.

The above Sentinel 1 TCS S/S thermal design ensures that all equipments are maintained within the allowed temperature limits during all mission lifetime, in compliance with the mass requirement (< 50 kg).

The summary for the TCS mass budget is shown in Table 2. Taking advantage of the consolidated thermal design of the above described previous programs and being the total TCS mass well inside the mass limit, any mass optimization had been performed but only local solution refinements. In fact the doublers were finally selected for battery modules instead of heat pipes being them adequate to meet the battery requirements and very simple thermal hardware. As consequence the doubler mass is the highest contribution in the budget.

Table 2. Sentinel 1 TCS mass budget

IV. Verification Approach The Sentinel 1 TCS S/S thermal design is verifiable by analysis, using an overall Thermal Mathematical Model

(TMM) of the satellite, and by ground test, performed during the Thermal Balance Test (TBT) in the frame of spacecraft environmental test campaign of the ProtoFlight Model (PFM). In addition, the PFM TBT has the target to validate the overall TMM of satellite by means of correlation of test results.

The overall Thermal Mathematical Model (TMM) of the satellite has been established for the prediction of the equipment temperatures during the mission phase (i.e. LEOP, worst hot and cold cases of routine and emergency phases). The selected tools are THERMICA software for the overall Geometrical Mathematical Model (GMM) and ESATAN software for the overall TMM.

The overall TMM is composed of S/C body detailed TMM and simplified thermal models of the equipment not under TCS responsibility (e.g. some external appendages and instruments) in order to take into account heat fluxes to/from the spacecraft structure as well as interface temperatures. It comprises about 8000 thermal nodes; among them, only 287 thermal nodes are dedicated to the internal equipment while the great amount of the thermal nodes (i.e. about 4000) is used to model the structure. The degradation of the external surfaces thermo-optical properties has been also accounted for in the End Of Life (EOL) analysis.

The overall GMM of the satellite is mainly shown in Figure 6, Figure 7, Figure 9 and Figure 10.

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Stowed Configuration Deployed Configuration

Figure 10. Sentinel 1 external GMM

Once the model was completed, consolidated and checked, the sensitivity analysis was performed to quantify the uncertainties affecting the design and they have been opportunely reduced. For the hot cases, the uncertainty nominal value of +10ºC has been reduced to +7ºC/+8ºC for internal equipment and +7ºC for propulsion while +10°C is maintained for external appendages. For the cold cases, an increment of +10ºC in the heating lines thresholds involves a similar increment in the minimum calculated temperatures, excluding some local equipment. Therefore, the uncertainty nominal value is reduced to -5ºC in most of the equipment since, sufficient heating power is available for maintaining the minimum temperatures inside margins.

The overall TMM of the satellite will be also validated through model correlation using the PFM TBT temperature results and then it will be used for final flight predictions. Therefore, it will be modified accordingly to the S/C configuration during test and then it will be used to predict the equipment temperatures during the TBT in the frame of the ground test phase.

The S/C will be tested without deployable antennas and solar arrays. In particular, the SAR central panel and SAR ASS will be replaced by dummies mechanically/thermally representative. In addition, the propellant tank will be empty, the pyrotechnics devices will be not installed and the battery will be integrated but not armed. Finally, some flight MLIs will be replaced by test MLIs, which will be fully flight representative with the exception of the external coating of the final layer.

The presence of the mechanically/thermally representative SAR ASS±X dummies is mandatory for the PFM TB purposes, being them fixed structures in front of panel radiators.

The PFM TBT will be performed with a simulation of external heat fluxes using an infrared heating system. The use of infrared heating system is the preferred one to simulate Sentinel 1 thermal environment during TBT for the following considerations:

• the orbital heat fluxes on satellite radiator panels (+X, -X, +Y panels) are mainly due to infrared radiation from Earth;

• sun flux impinges on satellite radiator panels at a variable small angle along the orbit and the equivalent average absorbed heat flux can be calculated and added to heat flux provided by infrared heating system;

• different thermal conditions expected in orbit can be simulated in the test phases individually changing the heat flux absorbed by the radiator panels without need to move the satellite;

• Sentinel 1 TCS S/S thermal control design is flight proven on similar satellites (e.g. Radarsat-2, Cosmo PFM, FM2, FM3, FM4) in terms of size, mass, power and mission;

• the main scope of TBT is the correlation of satellite TMM with test results, that can be achieved only using a well known thermal environment;

• the same mechanical configuration can be used for TB and TV execution without need to reopen vacuum chamber.

A preliminary study of the S/C, which is represented without SAR antenna central panel, fixed appendages and support structure dummies, inside the baseline selected TVAC chamber for TV/TB test is shown in Figure 11.

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Test Configuration

Figure 11. Sentinel 1 satellite in TVAC chamber

The selected IR heating system approach is further facilitated by the special features of the selected TVAC chamber (Figure 11). In particular, its thermal regulation is based on a Cooling S/S and a Heating S/S. The Cooling S/S is composed by LN2 lines reaching the cold sandwich black painted shrouds which emissivity is 0.896. The shroud has a temperature uniformity of 0/+5°C. Two types of cooling techniques are available: LN2 partial and total flooding. The Heating S/S is composed by several ceramic Infrared Tiles groups and each group has a maximum power of 2.4 KW. The IR Tiles are placed along cylinder, door and bottom sides.

Due to Heating and Cooling S/Ss, the lowest temperature inside the chamber is around 113 K (-160°C) and the highest temperature 413 K (+140°C). In fact the reduction of shrouds field of view for the infrared tiles presence is 10% max, maintaining a good cooling capability (up to -160°C on test article external surfaces).

In addition, several radiometers will be installed in selected locations (radiators in general) in view of S/C panels in order to monitor/record the radiative heat flux incident on external surfaces of the S/C during the test. The radiometers shall have an accuracy of ±10% in the range 0-1400 W/m2.

Therefore the TVAC chamber features, in terms of Heating S/S and radiometers, allow to produce controllable heat fluxes on selected areas, which will be used as known boundary conditions and input parameters for the TMM correlation.

S/C interfaces to GSE and test cables will be also thermally controlled by TC and test heaters, to avoid unwanted heat exchange through non-flight interfaces.

V. Conclusion The Sentinel 1 TCS S/S thermal design has been presented. Its maturity is the result of a complex activity, with

which all the main criticalities have been successfully solved. The presence of the SAR ASS±X has posed a challenge to the thermal design, requiring the optimization of the satellite layout, the ASS thermal design and the TCS thermal hardware.

The presence of the SAR ASS±X is also fundamental for the PFM TBT. The definition of the detailed test set-up during the PFM TBT is in progress.

Y

Z

X

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Acknowledgments The authors would like to thank all the members of the Sentinel 1 team for their constant and constructive

support.

References

1E. Attema, “Mission Requirements Document for the European Radar Observatory Sentinel-1” ES-RS-ESA-SY-0007, Issue 1.4, 11 July 2005.

2E. Attema, P. Snoeij, R. Torres, A. Pietropaolo, D. Scaranari, V. Mastroddi, S. Occhigrosssi, “ANALYSIS OF SENTINEL-1 MISSION CAPABILITIES”, EUSAR 2010, 7 June 2010

3R. Torres, S. Lokas, C. Bruno, R. Croci, M. L'Abbate, M. Marcozzi, A. Panetti, A. Pietropaolo, P. Venditti, “GMES SENTINEL-1: MISSION CONCEPT AND SYSTEM DESIGN” 16th Ka Band Conference, 20-22 October 2010.