Sentinel 3 – Spacecraft Thermal Control: design, analysis

45th International Conference on Environmental Systems ICES-2015-19 12-16 July 2015, Bellevue, Washington

Sentinel 3 – Spacecraft Thermal Control: design, analysis and verification approach

M. De Stefanis1 Thales Alenia Space Italy, Rome, Italy, 00131

I. Melendo 2 Airbus Defence & Space, Madrid, Spain, 28022

G. Cluzet3 Thales Alenia Space France, Cannes, France, 06156

and

S. Dolce4 ESA-Estec P.O. Box 299, 2200 AG Noordwijk, The Netherlands

ESA is developing a family of spacecrafts called Sentinels for the operational needs of Copernicus. Copernicus is a programme of the European Commission which aims at achieving an autonomous, multi-level operational Earth observation. Each Sentinel mission is based on a constellation of two spacecrafts to fulfill revisit and coverage requirements, providing robust datasets for Copernicus Services. The Sentinel-3 mission's main objective is to measure sea-surface topography, sea- and land-surface temperature and ocean- and land-surface color with high-end accuracy and reliability in support of ocean forecasting systems, and for environmental and climate monitoring. The Sentinel-3 spacecraft prime is TASF. The Platform main contractor is TASI and the Thermal Control Subcontractor is ADS Spain. The paper summarise the Spacecraft thermal control design and relevant analysis. The Spacecraft key thermal design approach has been to decouple the payload from the platform. Each payload is responsible for its own thermal control within the given thermal interface conditions defined at system level. The Platform selected thermal control is passive supplemented by heaters. The radiators coating is 5 mil ITO Silver Teflon. The radiating area has been defined taking into account the estimated thermo-optical properties degradation with time and the need to minimize heater power demand. Interface filler is used to increase the thermal coupling between the electronic units and the Aluminum honeycomb panels. The internal part of the panels is black painted to maximize the radiative exchange. Aluminum doublers and heat pipes are used for high dissipative units. The heat pipes layout is compatible with the ground testing configuration. Heaters commanded by temperature data from thermistors are used during operational and non-operational phases to maintain the units in the specified temperature range. Multilayer insulation blankets are used on all non-radiating surfaces. The external layer is 1 mil Black Kapton. The main reason for this selection is to provide electrical conductivity to avoid surface charging under UV light in space. This design leads to a ‘black’ spacecraft. The verification of the thermal control performances is by analysis and by TB test. Due to the S/C conventional thermal control, the TBT is performed on the Flight Model. The main objective of the test is to validate the Platform TMM and the thermal interfaces to the P/L. The paper describes the test configuration, test set-up and main phases. The S/C TBT is foreseen in April 2015.

1 System Thermal Engineer, Thermo-mechanical department, [email protected] 2 Thermal Engineer, Thermal Control Department, [email protected] 3 Satellite Thermal Architect, Thermal engineering department, [email protected] 4 Thermal Engineer, ESTEC TEC-MTT, [email protected]

International Conference on Environmental Systems

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Nomenclature α = absorptivity ADS = Airbus Defence & Space AOCS = Attitude and Orbit Control System BOL = Beginning Of Life CRS = Coarse Rate Sensor DORIS = Doppler Orbitography & Radiopositioning Integrated by Satellite ε = emissivity EOL = End Of Life EPS = Electrical Power System ESA = European Space Agency GFRP = Glass fiber reinforced plastic GMM = Geometrical Mathematical Model GNSS = Global Navigation Satellite System I/F = InterFace ITO = Indium Tin Oxide LEO = Low Earth Orbit LRR = Laser Retro-Reflector MHSTR = Multi-Head Star Tracker MLI = Multi Layer Insulation MOD = MODulator MWR = MicroWave Radiometer OLCI = Ocean and Land Colour Instrument PCDU = Power Conditioning & Distribution Unit PDHT = Payload Data Handling and Transmission PDHU = Payload Data Handling Unit PIM = Payload Integrated Module P/L = PayLoad S/C = SpaceCraft SBT = S-Band Transponder SLSTR = Sea and Land Surface Temperature Radiometer SMU = Satellite Management Unit SRAL = Sar Radar ALtimeter SSM = Second Surface Mirrors SVM = SerVice Module TBT = Thermal Balance Test TCS = Thermal Control Subsystem TMM = Thermal Mathematical Model TRP = Temperature Reference Point TWT = X-Band Travelling Wave Tube TASF = Thales Alenia Space France TASI = Thales Alenia Space Italy

I. Introduction The main objective of the Sentinel 3 mission is to measure sea surface topography, sea and land surface

temperature, and ocean and land surface colour with high accuracy and reliability to support ocean forecasting systems, environmental monitoring and climate monitoring. The mission is driven by the need for continuity in provision of ERS, ENVISAT and SPOT vegetation data, with improvements in instrument performance and coverage.

The Sentinel-3 mission is characterized by a multi-payload implementation, see also Figure 1. The spacecraft embarks the following instruments:

• SLSTR: the instrument, based on Envisat's Advanced Along Track Scanning Radiometer (AATSR), is designed to determine global sea-surface temperatures to an accuracy of better than 0.3 K. The SLSTR has a 1420 km swath at nadir and a 750 km inclined view and measures in nine spectral channels


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including two additional bands optimized for fire monitoring. The SLSTR has a spatial resolution in the visible and shortwave infrared channels of 500 m and 1 km in the thermal infrared channels.

• OLCI: the instrument is based on heritage from Envisat's Medium Resolution Imaging Spectrometer (MERIS). The 1270 km swath of OLCI fully overlaps with that of SLSTR and measures surface radiance in 21 bands, compared to the 15 on MERIS. The OLCI design is optimized to minimize sun-glint and provides measurements at a resolution of 300 m over all surfaces. OLCI marks a new generation of measurements over the ocean and land.

• Dual-frequency (Ku and C band) advanced SRAL. The instrument is based on CryoSat heritage that provides measurements at a resolution of ~300m in SAR mode along track. SRAL is supported by a microwave radiometer for atmospheric correction

and a DORIS receiver for orbit positioning. The combined topography package will provide exact measurements of sea -surface height, which are essential for ocean forecasting systems and climate monitoring. SRAL will also provide accurate topography measurements over sea ice, ice sheets, rivers and lakes.

• MWR to support the SRAL to achieve the overall altimeter mission performance by providing the wet atmosphere correction.

• GNSS Assembly, suitable for the Precise Orbit Determination (POD) processed on ground to achieve the overall altimeter mission performance. Real time navigation and dating information from this equipment will provide spacecraft navigation and dating functions as well as the control of the Radar Altimeter open-loop tracking function.

• DORIS Assembly as a Customer Furnished Item, which constitutes a complementary POD data provider for the Ground Segment as well as a potential backup to the GNSS Assembly for the specific commanding of the SRAL Open Loop tracking mode.

• LRR will be used by laser tracking stations to measure the distance to Sentinel-3 to within 2 cm

The Sentinel 3 spacecraft is 3-axis stabilized. It has a launch mass of about 1200 kg compatible with small launchers like Rockot and VEGA. The overall power consumption is 1100 W. The design life is 7 years, with ~100 kg of hydrazine propellant for 12 years of operations, including deorbiting at the end. The orbit is sun synchronous, 800 [Km] and 10:00 DN. The spacecraft layout has been driven by the need to provide a large face viewing cold space for thermal control and a modular design for payload accommodation, simplified management of all on-board interfaces. In addition to Structure and Thermal control the main Spacecraft S/S are:

• EPS comprising one solar array wing providing power to the spacecraft and payload • AOCS composed of a Coarse Sun Sensor, Magnetometers, Coarse Rate Sensors, Star Trackers, a GNSS

receiver and control actuators including thrusters, magneto-torquers and reaction wheels. • SMU for spacecraft commanding and monitoring • PDHU for payload data handling • Mass Memory Units for the spacecraft and its payload • Telecommunication subsystems including an S-Band subsystem for both telecommand (TC) uplink and

telemetry (TM) downlink, and a dedicated high volume (2×280 Mbps) X-Band subsystem for mission data downlink.

Figure 1. Sentinel 3 configuration


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All S-3 mission data are stored in dedicated mass memory Packet Stores (PS) within the on-board PDHU. Three memory modules are implemented providing a storage capacity of 576 Gbit (72 Gb).

II. Spacecraft Configuration The S/C general layout is based on a body with parallelepiped shape, compatible in size, including antennas,

instruments, solar array, with the VEGA and Rockot launchers. The S/C will assume two main configurations: the stowed configuration during Launch phase (see Figure 2) and

the fully deployed configuration during flight operative phase in orbit (see Figure 3).

Figure 2. Spacecraft in stowed configuration

Figure 3. Spacecraft in deployed configuration


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The spacecraft main features are summarized in table 1.

The spacecraft Thermal Control main requirements are:

Table 1. Sentinel-3 main characteristics


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• maintain the temperature of all units within the specified ranges in all the mission phases; • provide an interface reference temperature for individually controlled items (e.g. antennas, instruments ) • provide the sensors (thermistors) for the spacecraft temperature monitoring and control • provide the heaters for the spacecraft temperature control • provide the required thermal environment to the structural parts in order to maintain alignment and

ensure its stability; • minimize any thermal misalignment and gradient • minimize mass and heater power demand

The spacecraft thermal control design is based on a passive design supplemented by heaters. The main features are:

• Main heat rejection achieved through dedicated radiators (SSM tape) • Heat flows to and from the external environment reduced by MLI blankets • High emissivity (black paint) on internal units and panels to reduce gradients and minimize heater

power • Heater system • Batteries enclosed in a MLI tent • Antenna and instrument , externally mounted , to ensure their own thermal control • External surfaces (MLI, radiators) electrically conductive

Instrument thermal control A.All instruments are thermally autonomous and are as much as possible thermally decoupled from the platform

with MLI blankets and low conductive thermal washers at I/F levels. This decoupling allows an easier thermal regulation of the instrument. Exceptions are the electronics of SLSTR and SRAL located in the bus and controlled together with platform units. Each instrument does the management of its thermal regulation for all the spacecraft modes, only power lines being provided by the platform. Figure 4 provides the general spacecraft thermal architecture for the different instruments.

Table 2 provides the detailed responsibility sharing for the thermal control of the different instruments.

Spacecraft general heating architecture B.Figure 5 describes the general spacecraft heating system including the platform and the instruments.

Table 2. Responsibility sharing for instruments thermal control

Figure 4. General spacecraft thermal architecture - instruments


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III. Platform thermal design and analysis The Sentinel-3 TCS is designed to control the thermal environment of the Platform as well as to maintain the

internal equipment within allowable temperature ranges during all the mission phases. The design approach is to use primarily passive solutions as MLI blankets, SSM, thermal interface fillers (i.e. Sigraflex) and coatings (i.e. black paint). The use of active components is limited to constant conductance heat pipes, heaters and thermistors.

Sentinel-3 thermal control uses lateral panels to reject the internal heat dissipation. The radiative areas are covered with SSM used in space where a surface finish with a low value of solar absorptivity and a high infrared emissivity is needed. For Sentinel-3, it has been designed a 5mil ITO Silver Teflon tape for external radiators. Typical thermo-optical characteristics are α (BOL) = 0.10, α (EOL) ≤ 0.24 (i.e. 0.22 for +Y lateral panels and 0.19 for remaining lateral panels) and ε =0.78. The sizing of the radiative areas are defined by taking into account the worst conditions of maximum heat dissipation, maximum solar illumination (solstices), end-of life thermo-

optical properties. Sentinel-3 radiative areas are summarized in Figure 6. The platform is isolated from the space environment by MLI blankets minimizing temperature variations. MLI is also used on internal components, e.g. battery, tank and propulsion components to emphasize heating efficiency by reducing heat leakage and it is composed of various Mylar or Kapton aluminized layers (with low infrared emissivity) separated by Polyester spacer. The number of layers and type of material depends on which temperature level MLI has to experience. The external main layer of blankets mounted on the external side of the spacecraft is made of Black Kapton while the thrusters’ zones are protected with high temperature MLI with a titanium foil as external side. Dedicated lay-up has been chosen for different components of the platform. Lay-up definition is shown in Table 3.

Figure 5. General spacecraft heating system

Figure 6. Platform radiator areas


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High infrared emissivity coating (i.e. Aeroglaze Z307) is designed for internal side of lateral panels, equipment’s

and doublers in order to increase thermal radiation exchange inside the S/C and to minimise temperature gradients.

The most dissipative unit (i.e. PCDU) is mounted on constant conductance heat pipes (CCHP) that are installed on the inner surface of +Y panel. There are 4 equal linear heat pipes and 1 C-shape heat pipe made from Aluminum Alloy extrusion and filled with high purity ammonia as working fluid. Heat pipe layout has been designed such a way that it prevents the failure of any heat pipe without consequent degradation of satellite performance. PCDU panel layout is presented in Figure 7. Heaters are bonded on the panel due to lake of surface on heat pipes induced by PCDU large surface contact area.

Aluminum doubler with 3 mm of thickness is foreseen for SBTs, TWTs and MODs in order to spread the heat. Figure 8 shows the thermal hardware mounted on PDHT panel, particularly the black painted doubler for TWTs (on the right with four unpainted areas for units mounting) and the black painted doubler for modulators (on the left with four unpainted areas for units mounting) while Figure 9 shows the designed doubler for SBTs (see two unpainted areas for units mounting).

Table 3. MLI lay-up definition

Figure 7. HPs layout on +Y panel


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The Sigraflex interfiller (Type F03510TH, 0.35mm thick) is chosen below units, doublers and heat pipes. It is

also mounted between units and heat pipes and between unit and doubler. GFRP washers are mounted below external appendages and payloads in order to thermally decouple them from platform structure. Same washers are also mounted between thruster’s support and structure.

Heaters and thermistors are the active components of thermal control subsystem. Electrical heaters with automatic control are used for the following purposes:

• Provide temperature control of battery, propellant tank and external sensors during mission nominal operations

• Substitute/compensate the power dissipated by the electronic equipment when they are not operating and the reduction of orbital flux when the satellite is in eclipse

• Maintain equipment operating/not operating minimum temperatures, to minimize seasonal temperature variation and to assure safe hold conditions in every emergency phase.

The above mentioned heater lines are main + redundant heater lines (M+R HTRs). Double layered foil heaters, with the nominal resistance in the lower face and the redundant in the upper face are designed for most of items, apart the piping lines where wire heaters (i.e. Clayborn type) are installed.

Each main and redundant heater line will be controlled via the same dedicated set of 3 thermistors taking into account the middle temperature reading (i.e. triple redundancy concept). Thermistors are monitoring and command ones (M+C THRs), because they both command heater circuits and monitor the actual on-board temperatures. Thermistors used on Sentinel-3 are Betatherm G15K4D489 (i.e. 15kΩ).

Sentinel-3 platform is provided of 35 + 35

Figure 8. Thermal HW mounted on PDHT panel

Figure 9. Thermal HW mounted on +Y panel

Table 4. Heater lines characteristics


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HTR lines (Main + Redundant), controlled in automatic by SMU with maximum power demand of 20 W at 28 V for 20 + 20 lines and 40 W at 28 V for 15 + 15 lines. Table 4 shows main characteristics and location of platform heater lines.

The main functions of the heating system are summarized here below: • Acquisition of the 3 thermistors value (done by SMU) • Determination of the median temperature (tolerance to one thermistor failure) (done by SMU) • Calculation of the required heating power with a Proportional Integrative (PI) algorithm taking into account

the difference between the median sensor temperature value and the temperature target and the previous heater power (done by SMU)

• Calculation of the heating duration (duty cycle) corresponding to the required heating power (done by SMU)

• Injection of the calculated power with a Pulse Width Modulation (PWM) for heater operation (done by PCDU)

The required heating power is calculated each 32 s. The power injection cycle duration is 32 s and the power injection step is 1s. The intrinsic stability performance of such a regulated heating system is roughly ±1.0°C. For each heating line i and for each computation cycle k, the regulation algorithm has the following form:

With, for each line i:

Pi (k-1): Heating power at the previous cycle k-1 Ti (k): Measured temperature at the current cycle k Ti (k-1): Measured temperature at the previous cycle k-1 Tci: Temperature target C1i , C2i: PI regulation coefficients And with following initial values: k=1, Pi (0)=0, Ti (0)=Ti

Thermal analysis description A.

A TMM representative of the spacecraft is built to complete the Sentinel 3 thermal design sizing and provide temperature and heat fluxes profiles. The model includes all the major assembly components represented, either as individual nodes or as group of nodes. A GMM is built to compute view factors between the different elements of the satellite and external fluxes. The GMM includes spacecraft platform (MLIs, radiators, doublers, lateral panels…), SVM internal units (modeled as one node when reduced TMM not provided), tank, propulsion components and I/F ring.

THERMICA software was used to build the geometrical model and to compute the view factors, radiative exchanges and external heat inputs. The overall GMM is composed by 5400 thermal nodes; satellite model overview is shown by Figure 10.

The TMM also includes heat pipes (considered isothermal and modeled as one node per each), PIM panels and struts internal nodes, piping, TMMs of each instrument and heaters modeled as heat applied in the geometrical node where located. ESATAN software was used to compute the temperature results.

1: 1 1 ∗ 2 ∗ 1 (1)


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1. Environment definition – Nominal orbit

The paragraph describes all mission relevant parameters used in the analysis to compute external fluxes. Spacecraft operates in a sun-synchronous orbit

14+7/27 with an average altitude of 814.15 km and a local time descending node at 10:00 am. Detailed description for the orbit is provided in Table 5.

The satellite pointing is such as +Z axis points towards the Earth. The flight direction is along –X axis such that the payload, located on the +X axis, is not exposed directly to the ATOX flux. Therefore, the main features of the Sentinel 3 mission guidance are:

• +Z axis pointing toward Nadir direction (yaw axis) • -X axis pointing toward spacecraft velocity

direction (roll axis) • +Y axis pointed in the opposite direction of the

orbit normal (pitch axis) • Solar array rotates around -Y axis in order to point

the Sun Figure 11 illustrates the satellite attitude in operational

mode.

Concerning the satellite steady state safe mode, it can be reached after an initial heliocentric pointing acquisition phase during which the spacecraft may have any attitude (2.5 hours maximum after an in-flight reconfiguration). In this mode, the spacecraft will point its –Z axis to the Sun within an error cone of ± 15 degrees during daylight part of the orbit. When leaving eclipse, the error cone angle may be as high as 30 degrees due to the drift during eclipse. During the safe mode steady state, the rate around Z axis is controlled at around 0.06º/s.

2. Environment definition - LEOP PHASE Launch and early operations phase is constituted by three different stages: Launch, Rate Dumping Phase and Sun

Pointing Phase. During launch solar array stays in stowed configuration, S/C may have any attitude and it is exposed

Figure 10. External model breakdown of Sentinel-3

Table 5. S3 nominal orbit description

Figure 11. Satellite attitude in operational mode


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to an aero-thermal flux pointing +X S/C axis as shown in Figure 12. This phase has a maximum duration of 90 minutes.

In the course of rate dumping phase solar array becomes deployed, S/C is placed in the nominal orbit and it may have any attitude. This phase takes a maximum of 150 minutes. Sun pointing phase is extended during 90 minutes. Solar array is maintained in deployed position and S/C may have any attitude in the nominal orbit. 3. Thermal cases definition

An extensive analysis campaign has been performed in order to evaluate all the mission phases and satellite performances. Table 6 presents a summary of the analysis cases definition.

4. Analysis results

Design temperatures limits are reported in Table 7 (i.e. minimum and maximum values). It was necessary to consider, for the evaluated temperatures, the effect of the uncertainty of calculation (Calculation Uncertainty Margin, U.M.). It was assumed an UM of + 8ºC for maximum temperature; then, - 3ºC for minimum temperature for units with a dedicated heating line whereas an UM of - 5ºC has been taken into account for units without a dedicated heating line (controlled by mean of the environment).

Unit’s dissipation for different satellite modes is shown in Table 8. There are values for nominal mode (maximum and nominal figures), safe mode, launch, rate dumping phase and sun pointing phase.

Table 7 shows a summary of maximum and minimum guaranteed temperatures (i.e. including UM) of SVM units for the whole mission.

The power budget calculated for TCS during nominal mode is 34.1W while during safe mode is 171.2W as worst case, obtained taking into account BOL conditions. These values are compliant with respect to requirements figures (i.e. 90W in nominal mode and 220 W in safe mode).

Also the calculated mass budget is compliant with the respect to requirement. Indeed, TCS mass

Table 6. Thermal cases definition summary

Figure 12. Aero-thermal flux

Table 7. Summary of analysis results for SVM units


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is 21.7 Kg that is less than specified value of 22 Kg.

IV. Platform thermal verification Platform thermal control design and analysis main vehicle

of verification is through a Thermal Balance Test (TBT). The main objectives of the Platform TCS TBT are listed below:

• Verification of the Platform Thermal Control performances, including:

o Verification of the TMM, as part of the TCS qualification

o Demonstration of the suitability of the thermal control design

o Verification of the performances of the thermal control hardware

o Verification of the active thermal control in flight configuration (temperature targets and control constants)

• Verification that thermal interfaces with the instruments with a dedicated thermal control remain within the specified range as used in the instrument thermal balance test. It concerns OLCI, SLOSU, MWR and SRAL antenna/feed.

TBT Success Criteria & Test Sequence A.TBT success criteria are as defined below:

• No major damage occurs to the test specimen as a result of testing

• The stated test conditions have been achieved and stable conditions have been reached as defined below

• All required measurements are recorded completely and correctly The stabilization criteria for the steady state thermal cases shall be fulfilled with the three following

conditions: • The temperature variation of all PF units remain under ±0,2°C during 3 successive hours

• For each PF flight heater line, the one-hour calculated average value of the heating duration X parameter of duty cycle does not exceed 10% of variation out of 3 hours (as a reminder this parameter indicates the time during which the installed power will be injected out of every cycle of 32s)

• For each test heater implemented on PF, the consumption variation should not exceed 5%

The TBT will start after an outgassing phase (duration from 2.5 to 4 days). Outgassing phase stabilized temperatures have been predicted, according to inputs provided (test set-up, platform units dissipation, etc.), as a starting point for the TBT predictions. Thermal Balance Test sequence, after the outgassing phase, is recalled here below and shown in Figure 13:

• TB1 Hot Nominal Phase. It is composed by a steady state phase (TB1A) simulating hottest in-flight conditions for most of the Platform units, and a transient phase (TB1B) simulating the PDHT downlink (maximum dissipation for PDHT units)

• TB2 Cold Nominal Phase. It is a steady state phase simulating coldest in-flight conditions for nominal

Figure 13. S3 TBT phases

Table 8. Unit dissipation


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operation • TB3 Cold Safe Mode Phase. It is a steady state phase simulating coldest in-flight conditions for safe

mode

TBT Facilities and Set-up B.For TBT, the spacecraft will be installed inside the TAS-F thermal vacuum chamber Espace 500, located in

Cannes. The thermal chamber will be able to reach a secondary pressure inferior to 10-5 mbar (1.3 10-5hPa) for all the duration of the vacuum test. Figure 14 shows the satellite set-up that will be used for thermal test in Cannes.

The ground launch vehicle adapter (GLVA) of S3 will be attached to the thermal vacuum test adaptor (TVTA, see Figure 15) conductively decoupled from S/C. The interface between the TVTA and the spacecraft will be such that the conductive heat flux from/to satellite is kept within 2 Watt. To respect this requirement the temperature of the thermal test adapter will be controlled with dedicated test heaters and thermocouples. The thermal test adaptor will also be insulated by MLI in order to minimize the radiative thermal impact on S/C.

During thermal cycling, the thermal control of the thermal test adapter will allow to control the interface temperature within the range [-40°C; +60°C]. 300 W heater power will be installed on the TVTA. During thermal balance test the temperature gradient between the GLVA and the TVTA shall be lower than 5°C.

Platform configuration during TBT will be the complete PFM, with the following exceptions:

• Solar Array not installed • HRMs not installed; test MLI patches will cover the HRM areas in the platform panels • 2 S-band antennas present but disconnected • X-band antenna removed and replaced by a RF load on +Z panel • Tank filled with 2 bar gaseous Nitrogen

Infra-red lamps are installed in front of several areas of the spacecraft with the aim to maintain the satellite in safe conditions in case of total loss of power at satellite level and also to simulate the external fluxes for MWR and SLSTR radiators. Figure 16 depicts the IR lamps location chosen for Sentinel 3 thermal test.

Figure 15. Thermal test adapter configuration

Figure 14. S3 satellite configuration in TV chamber


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Only IR lamps located

in areas A5, B4, C4 and C5 will be used during TBT. Their maximum simulated heat flux will be 420 W/m2. Heated plates are installed in front of several areas of the spacecraft in order to maintain external components above their minimum temperature.

Figure 17 details the different areas where heated plates are located, so as the heater power installed for each one.

In addition, platform test heaters have been installed on four radiators to simulate the external absorbed heat fluxes where the flight heater is insufficient (i.e. PCDU and CRS units) or not present (i.e. SMU and MHSTR units). Then, other four test heaters have been mounted on fixations of external elements (i.e. LRR, Doris antenna, GNSS antennas). Test heaters are Kanthal A1 2 wire heaters (N+R) with 11.3 Ohm/m resistance, FEP/polymide insulated. Main parameters definition is presented in Table 9.

Figure 18 shows an example of test

heaters installed around platform radiators, in particular test heater on SMU panel (inner side) is shown on the left and heater for MHSTR units is depicted on the right.

Flight heaters have also a specific ground use during TBT different from flight. For heaters located around PF units on radiator areas, they are used in the TBT hot case to simulate orbital fluxes on PF radiators. TBT cold cases (operational and safe mode), are simulated without orbital fluxes,

therefore, the function of flight heaters is to maintain the PF units at their minimum temperatures.

Figure 16. Infra-red lamps configuration

Figure 17. Heated plates location and related heater power

Table 9. Platform test heater definition


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For TBT hot case, simulation of orbital fluxes on radiators can be done on nearly all radiators, as they have individual flight heating lines. A specific TCS mode “Test boost mode” will be used for this purpose. For all heating lines installed on radiators:

• The “Overloading_flag” will be used • A fixed heating duration will be set with “Overloading_Xk” • As flight voltage is lower than TB test voltage (27 V vs 37 V), the range in the “Overloading_Xk” flag

will be [0 to 17] instead of [0 to 32] applicable for flight configuration heating pulses, in order not to stress the heaters more than in flight configuration

- If needed, the power limitation value (already foreseen in the software) will be increased for some areas with limited installed power

Platform units’ dissipation will be as per Table 10. Platform units dissipations have been measured on air prior to TBT and these values have been used for TBT predictions.

A total amount of 294 thermocouples (TCs) have been installed, in order to measure the temperature of different areas during TBT. TCs measurements will be used after TBT for TMM correlation purposes. Temperatures have been predicted for each TC.

A total amount of 126 thermistors (THs) have been installed, in order to control the different heating lines (3x33 THs) and to monitor the Platform in flight (27 items). THs measurements will also be used after TBT for TMM correlation purposes. Temperatures have been predicted for each TH, special attention will be paid during TBT and TMM correlation to the gradients between TCs installed in Platform units TRPs and THs associated to heating lines used to control the unit temperature, installed on panels (close to TRP).

TBT PREDICTION C.TBT prediction has been performed for all the

test phases. Temperature has been calculated for all the THs and TCs on the Platform. Transient analysis has been performed and stabilization criteria have been calculated for each TC and TH in order to determine the duration of each phase. Test heaters and flight heaters power have also been calculated.

Software used for radiative computation is Systema 4.5.3 and software used for thermal analysis network solution is Esatan 10.10

A dedicated TBT thermal mathematical model has been built. The basis is the TMM built for TCS CDR analysis (using BOL thermal properties), with the following modifications:

• Deletion of S/C flight elements not present during TBT (Solar array, HRMs, X-band antenna)

Figure 18. Test heaters installed on SMU and MHSTR panels

Table 10. Measured unit dissipation [W]

Status Dissipation Status Dissipation Status Dissipation Status DissipationSADM Hold/SAS 11 Hold/SAS 11 Hold/SAS 11 Hold/SAS 11PCDU / 65.3 / 73.6 / 58.3 / 64.9BTA1 / 0 / 0 / 0 / 0BTA2 / 0 / 0 / 0 / 0SMU / 53 / 53 / 53 / 45SBT1 Rx 5.15 Rx 5.15 Rx + Tx 25.6 Rx + Tx 25.6SBT2 Rx + Tx 25.6 Rx + Tx 25.6 Rx 5.15 Rx 5.15

GNSS1 On 11 On 11 Off 0 Off 0GNSS2 Off 0 Off 0 On 11 Off 0CRS1 On 6 On 6 Off 0 Off 0CRS2 Off 0 Off 0 On 6 On 6

MHSTR1 On 7.8 On 7.8 Off 0 Off 0MHSTR2 Off 0 Off 0 On 7.8 Off 0STROH1 On 0.8 On 0.8 On 0.8 Off 0STROH2 On 0.8 On 0.8 On 0.8 Off 0STROH3 On 0.8 On 0.8 On 0.8 Off 0

MTB1 50 A.m² 0.1 50 A.m² 0.1 50 A.m² 0.1 190 A.m² 2MTB2 50 A.m² 0.1 50 A.m² 0.1 50 A.m² 0.1 190 A.m² 2MTB3 50 A.m² 0.1 50 A.m² 0.1 50 A.m² 0.1 190 A.m² 2MAG1 Off 0 Off 0 Off 0 Off 1MAG2 Off 0 Off 0 Off 0 Off 0RW1 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6RW2 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6RW3 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6 Off 0RW4 ~25 rad/s 6 ~25 rad/s 6 ~25 rad/s 6 Off 0

PRES. TRANS / 0.2 / 0.2 / 0.2 / 0.2MOD1 0 0 15 0EPC1 0 0 9.4 0TWT1 0 0 2.1 0HPI1 0 0 0 0

MOD2 15 15 0 0EPC2 10 13.2 0 0TWT2 2.2 33.7 0 0HPI2 0 5.3 0 0

MOD3 0 0 15 0EPC3 0 0 10.7 0TWT3 0 0 2.3 0HPI3 0 0 0 0

MOD4 15 15 0 0EPC4 10 13.2 0 0TWT4 2.2 33.7 0 0HPI4 0 5.3 0 0

OMUX / 0 / 10.6 / 0 / 0WGSW1 / 0 / 1.2 / 0 / 0WGSW2 / 0 / 1.2 / 0 / 0PDHU On 30 On 35 On 30 Off 0DPU1 31.5 31.5 0 0RFU1 46.9 46.9 0 0DPU2 0 0 29.3 0RFU2 0 0 41.2 0

SLCPE Htr M+s/s 28 Htr M+s/s 28 Htr M+s/s 28 Off 0DORIS BDR / 20.5 / 20.5 / 20.5 Off 0

423.1 529.4 408.3 176.9

Off

TB3

Off

Off

Off

Off

On SAR

Off

Off

AOCS & Propu

PDHT

Inst

On SAR

On LRM

Total dissip inside PF

Off

Off

Pow+TTC

Off

TX-On

Off

TX-On

Off

TX-On +Tx-Conf

Off

TX-On +Tx-Conf

TB1A TB2

TX-On

Off

TX-On

Off

TB1B


17

• Inclusion of the following elements present during TBT: Vacuum chamber, Test adaptor, IR lamps and heated plates, with supporting structure, RF load, HRMs MLI patches

Figure 19 shows the geometrical model used for TBT. Units’ temperatures have been calculated for each TBT case.

However, only results obtained for phase TBT1A are shown in Table 11. Hot nominal phase (TB1A) predicted duration is 30 hours including transition from outgassing phase to nominal hot phase. Stabilized temperatures, unit’s dissipations during this phase, test heaters power, and flight heater power and number of pulses @37V are included in Table 11.

Table 11. TBT prediction for hot case TB1A

S3TB1A01

UNIT

DESIGN LIMITS

(ºC) STABILIZED

T (ºC)

UNIT

DISSIPATION

(W)

TEST

HEATERPI FLIGHT HEATER

MIN MAXPOWER

(W)Nk@37V

POWER

(W)

DORIS -10 40 23.3 20.5 7 14.0

MHSTR 1 -10 40 22.3 7.80.0

MHSTR 2 -20 50 16.6 0.0

DPU 1 -10 40 28.8 31.5 14 28.0

DPU 2 -20 50 11.2 0.0 5 10.0

RFU 1 0 40 28.3 46.9 10 40.0

RFU 2 -20 50 -1.2 0.0 6 24.0

PCDU -10 45 34.1 65.3 100.0 17 34.0

SADM -10 55 33.5 11.0

SBT 1 -10 50 20.9 25.617 34.0

SBT 2 -10 50 15.6 5.2

BTA 1 0 30 9.6 0.0 17 34.0

BTA 2 0 30 10.2 0.0 17 34.0

JUNCTION

BOX-10 40 14.0 0.0

SLCPE -10 40 15.7 28.0 17 34.0

GNSS 1 -10 40 16.5 11.017 18.2

GNSS 2 -20 45 10.5 0.0

SMU -10 45 24.0 53.0 20.0

OMUX -20 65 20.8 0.0

PDHU -10 40 18.2 30.0 17 18.2

Switch 1 -40 90 19.5 0.0

Switch 2 -40 90 19.4 0.0

HPI 1 -40 85 20.8 0.0

HPI 2 -40 85 22.1 0.0

HPI 3 -40 85 21.4 0.0

HPI 4 -40 85 22.5 0.0

MOD 1 -20 50 15.5 0.0

17 12.2MOD 2 -10 40 28.5 15.0

MOD 3 -20 50 16.1 0.0

MOD 4 -10 40 29.4 15.0

EPC 1 -20 50 16.6 0.0

17 18.2EPC 2 -10 40 27.5 10.0

EPC 3 -20 50 21.0 0.0

EPC 4 -10 40 23.9 10.0

TWT 1 -20 65 22.8 0.0

17 18.2TWT 2 -10 65 21.9 2.2

TWT 3 -20 65 22.0 0.0

TWT 4 -10 65 21.0 2.2

GYRO 1 -10 40 30.5 6.00.0 15 30.0

GYRO 2 -20 50 28.9 0.0

-Y+X_RWS -10 55 22.3 6.00 0.0

-Y-X_RWS -10 55 22.9 6.0

+Y+X_RWS -10 55 25.3 6.00 0.0

+Y-X_RWS -10 55 23.6 6.0

MAG 1 -20 50 19.0 0.0

MAG 2 -20 50 19.4 0.0

MTB X -20 55 14.6 0.1

MTB Z -20 55 15.4 0.1

MTB Y -20 55 20.4 0.1

Figure 19. GMM overview for TBT prediction


18

V. Conclusion The Sentinel 3 TCS thermal design and relevant analysis has been presented. The chosen approach is the result

of a complex activity, with which all the main criticalities have been successfully solved. The challenge for the thermal design has been to decouple the payloads from the platform being each instrument responsible for its own thermal control. The thermal decoupling allowed the simplification of the interfaces and reduction of the development risks. The satellite layout as well the TCS thermal hardware has been optimized in order to meet this requirement. The Platform selected thermal control is passive supplemented by heaters controlled by SMU using a PI control. The analyzed thermal control performances will be verified by thermal balance test. Indeed, the test will validate the Platform TMM and the thermal interfaces to the P/L. The test configuration, test set-up and main phases of TBT have been described in the paper. Test prediction has been performed for each test case and the foreseen temperatures of the SVM units during hot case TB1A have been presented. The S/C TBT will be performed in April 2015.

Acknowledgments The authors would like to thank all the members of the Sentinel 3 team for their constant and constructive

support.

References 1GMES Sentinel-3 Team, “GMES Sentinel-3 System Requirement Document” S3-RS-ESA-SY-0010, Issue 3, 17 October

2007 2ESA Communications, “Sentinel 3: ESA’s Global Land and Ocean Mission for GMES Operational Services” ESA SP-

1322/3, October 2012 3Donlon, C. et al, “The Global Monitoring for Environment and Security (GMES) Sentinel-3 mission” Remote Sensing of

Environment 120, 37–57, (2012)

Documents

Sentinel 3 – Spacecraft Thermal Control: design, analysis