European Space Thermal Engineering Workshop 2026

6 Oct 2026, 09:00 → 8 Oct 2026, 18:00 Europe/Amsterdam

Einstein, Newton (ESA/ESTEC)

Tuesday 6 October

09:30 → 10:30 Opening

ESA_Tech_Thermal_Overview_ESTEW_2026_REV1.pptx

10:30 → 11:00 Break

11:00 → 13:00 Small Satellites & Cubesats

Convener: Andreas Mussger (ESA)

11:00 Design Guidelines for Dynamic Radiative Thermal Environment Simulation in CubeSat TVAC Testing

Simulating the orbital thermal environment remains challenging even for leading space agencies and industry, as heat transfer in orbit occurs solely through radiation and is strongly affected by rapidly varying orbital position, flux, and temperature. During testing of larger spacecraft, thermal radiation is typically simulated using existing or dedicated structures composed of black radiative panels and controllable heaters, designed to fit within thermal vacuum chambers. As CubeSats become increasingly prominent thanks to their compact electronics, lower cost, and short development cycles, the need for accurate and accessible thermal test methods is growing. Yet, within the CubeSat community, the development of representative radiative thermal test setups remains limited. This paper identifies the different existing solutions to simulate a radiative thermal environment and proposes a methodology for dynamically adjusting radiative flux in resource-constrained settings. Drawing on experience and lessons learned from several large-satellite thermal vacuum tests, the paper evaluates how these heritage methods can be adapted for CubeSat applications. The paper presents a comparative analysis of current approaches used across European and North American CubeSat teams, supported by selected case studies. Test strategies are examined in the context of Low Earth Orbit, Medium Earth Orbit, and Geostationary missions to provide a broader applicability. Ultimately, this paper aims to propose practical guidelines and adaptable test setups that enable CubeSat teams worldwide to more accurately reproduce orbital radiative conditions using the resources available to them.

Speaker: Daphne Papadatos

Presentation_ICES2026_PuertoRico_ESTEW2026_DaphnePapadatos.pdf

11:30 Thermal Analysis of MOVE-IIIa and PEANUT Student Projects

Thermal simulation and analysis impacts mission success and is becoming increasingly critical as thermal system complexity grows in the space industry. Despite this, it remains largely underrepresented in student satellite and low-gravity experiment projects. Beyond a limited number of papers covering general concepts, there is almost no documentation on concrete step-by-step analysis approaches. To address this gap, we conducted thermal analyses of two projects from the Scientific Association for Rocketry and Spaceflight at the Technical University of Munich.
Both simulations were conducted in COMSOL Multiphysics. For the MOVE-IIIa CubeSat, we performed orbital hot- and cold-case simulations and correlated the model against results from two thermal-vacuum test campaigns, which yielded improvements to the simulation and validated the design. For PEANUT — a student experiment on piezo actuated reaction wheels within the German-Swedish REXUS/BEXUS programme, organized by the German Aerospace Center (DLR) and the Swedish National Space Agency (SNSA) — we conducted a one way coupled 2D CFD and 3D heat-load simulation. The results directly influenced the experiment's design, demonstrating the need for a battery heater and confirming PCB stack survival under high Mach number convective heating during ascent.
With MOVE-IIIa's launch scheduled for late 2026, we will correlate our thermal model against real in-orbit flight data. The presented workflow and results are intended to serve as a reproducible reference for future student teams working in thermal engineering, lowering the barrier to hands-on experience in this field.

Presenting this work at ESTEW would offer the opportunity to gather industry feedback and further validate our approach, supporting its use as a reproducible reference for future student teams.

Speaker: David Hangen

Thermal Analysis of MOVE-IIIa and PEANUT Student Projects.pptx

12:00 Thermal Balance and Cycle Testing of the ALEASAT 1U In-Orbit Demonstration CubeSat

ALEASAT is a 1U In-Orbit Demonstration (IOD) CubeSat intended for Earth Observation. Like other such projects, ALEASAT is generally characterized by higher risk of acceptance profile, low level of complexity (relative to other ESA space projects), low cost and short schedule, short operational lifetime (typically < 5 years in LEO), acceptance of single point failures, and robust safe mode (thermal and power safe in any attitude). In order to assess the thermal environment within ALEASAT, a thermal model (TM) was constructed in which the structure was decomposed into a number of discrete regions with specified thermal characteristics. Earlier this year, an extensive series of thermal tests were conducted at ESA’s CubeSat Support Facility in Belgium to verify this model and inform the spacecraft system engineering process. In the case of ALEASAT, this testing was conducted earlier in the development phase and in more detail than is usual for similar projects. Here, we present the results of such testing including thermal balance testing to provide data for model correlation and thermal cycle testing in which functional testing of the payload module and power amplifier at hot and cold temperatures was used to characterize performance of mission critical components. The results will inform the thermal modelling of other CubeSat projects as well as provide useful insights concerning the impact of thermal issues on both ALEASAT’s camera and RF power amplifier.

Speaker: Shreya Shah (University of British Columbia)

ESTEW 2026 Conference Presentation.pptx

11:00 → 13:00 Thermal Design

Convener: Heiko Ritter (ESA)

11:00 PCM-Based Heat Capacitor and Heat Spreader Development and Design Optimization

The Copper Alloy for Additive Manufacturing (CufAM) project was envisioned as a technology demonstration activity. In particular, a green (515 nm) laser was utilized which offers high energy absorption of up to 60% in copper thus enabling improved manufacturing capabilities for sophisticated geometries. The aim of the project was to advance the end-to-end AM process chain, and deliver reliable properties and part characteristics using oxygen-free, high-conductivity (OFHC) copper powder as a feedstock.

In order to showcase the advantages of this technology, phase change material (PCM) based devices were selected. Given the low thermal conductivity of the PCM material itself, such devices require an internal lattice in order to aid spreading heat across the volume. To this end, lattices made out of high thermal conductivity copper have been assessed, designed, compared and finally manufactured to be integrated into the devices or directly included at the fabrication phase.

The presentation is going to focus on the design and development process of two different types of devices: a heat capacitor and a heat spreader (doubler). The presentation will cover design considerations, the development process and manufacturing of the devices. Post-processing, additional machining and other steps to prepare the devices for testing are going to be discussed as well. An overview of the testing approach, characterization results and challenges encountered during testing will also be provided.

The CufAM project was a collaboration between the Manufacturing Technology Centre (MTC), Airborne Engineering (AEL), Fraunhofer-Institut für Lasertechnik (ILT) and Azimut Space GmbH, funded through the ESA General Support Technology Programme (GSTP).

Speaker: Matvei Andreev (Azimut Space GmbH)

PCM-Based Heat Capacitor and Heat Spreader Development and Design Optimization.pptx

11:30 Passive Cooling System Design of the Ariel Telescope

The Ariel mission is part of ESA’s Cosmic Vision program and is dedicated to studying the composition and physical properties of exoplanets. Cooling of the ARIEL payload is achieved by a combination of active and passive systems which lower the telescope's temperature down to an isothermal environment below 60K. The passive cooling to cryogenic temperature ranges is achieved by high-efficiency thermal shielding in the shape of V-Grooves, which utilize the favorable thermal conditions of the L2 orbit. This presentation outlines the development and resulting thermal design of the V-Grooves, including their thermal control surfaces, thermal interfaces, temperature monitoring instrumentation and the thermal properties of their mechanical mounting structure. The design is strongly based on heritage of the Planck mission and key developments resulting from critical differences between the two missions are presented.

Speaker: Ms Antonia Grethen-Bussmann (Beyond Gravity Schweiz AG)

ESTEW_2026_Passive_Cooling_System_Design_of_the_Ariel_Telescope_v3.pdf

12:00 HIROS Atmospheric Sounding Payload: Achieving Milli-Kelvin Stability in a Miniaturised Instrument

SOLSTICE is a UK-led programme developing novel miniaturised scientific payloads for the high-resolution profiling of atmospheric composition. One of these payloads is the High-resolution InfraRed Occultation Sounder (HIROS). This presentation describes the thermal design and verification of the instrument, with particular focus on its demanding thermal control requirements.

HIROS is a thermal infrared laser heterodyne spectro-radiometer using a quantum cascade laser (QCL) as local oscillator. To ensure it operates within its selected spectral micro-window, HIROS must maintain its QCL within ±3 K of a set temperature, while achieving a thermal stability of <3 mK/s for periods of up to 120 s. In addition, the detector must be cooled below −66 °C using a TEC-based cooling system. Such requirements present a significant challenge within a compact 7.9 dm³ payload architecture, further constrained by a limited power budget, restricted radiator area, and the highly localised dissipation of 2.5 W within the mm²-scale QCL package.

Initial payload-level thermal-vacuum testing identified difficulties in meeting these requirements. A dedicated thermal test campaign was subsequently conducted to characterise the instrument’s thermal behaviour, investigate the underlying causes, and assess compliance with operational requirements. The campaign identified the thermal mechanisms preventing compliance with the ±3 K temperature requirement, demonstrated compliance with the QCL thermal stability requirement (<3 mK/s) and highlighted conductive thermal limitations affecting detector cooling. The presentation describes the test methodology, key findings, and lessons learned applicable to the thermal control and verification of miniaturised scientific instruments.

Speaker: Theo Dantichan (RAL Space)

ESTEW 26 - HIROS Thermal Design and Testing.pptx

12:30 Design and optimization of open cell radiator for cryogenic application

Cryogenic radiators are often required to cool down sensitive, high precision instruments to their optimal operation temperature. However, IR emissivity of thermo-optical coatings below 140K tends to drop rapidly, making it difficult to comply with scientific requirements, when radiative area is limited by design constraints. A potential solution is to exploit cavity effect, that can increase effective emissivity of a given area pending on cavity geometry and IR reflectivity of cavity surfaces. In this presentation it will be demonstrated how the radiative performance can be maximized at cryogenic temperature using an open cell honeycomb. The presentation will include aspects of coating selection and geometrical design of ARIEL Instrument Radiator considering the key design drivers.

Speaker: Mark Alexa (Admatis Ltd.)

ARIEL-ADM-PL-PRS-189_DRAFT.pptx

ARIEL-ADM-PL-PRS-189_FINAL.pptx

13:00 → 14:00 Lunch

14:00 → 16:00 Thermal Analysis

Convener: Duncan Gibson (Telespazio BE for ESA)

14:00 ESATAN Thermal Modelling Suite - Product Development

ESATAN-TMS provides an advanced thermal modelling environment for the thermal analysis of spacecraft and launch vehicles. The suite is continually being enhanced to meet current and future requirements of space projects, and to support the specific needs of thermal engineers. This presentation will focus on the latest development to be included in the upcoming ESATAN-TMS 2027 release.

Speaker: Samuel Ekregbesi (ITP Aero)

ESATAN-TMS 2027 Product Overview.pptx

14:30 SYSTEMA-THERMICA recent updates, model interoperability and AI usage

The SYSTEMA-THERMICA software suite continues to evolve to support space thermal engineers facing increasingly complex simulation and data exchange challenges. This presentation provides an overview of the latest developments focusing on upcoming software features, model interoperability and new workflow methods.

First, we will introduce the upcoming Systema 4.9.4P3 release scheduled for the end of 2026.
A major highlight is the introduction of a new interoperability facilitator enabling bidirectional TMM (Thermal Mathematical Model) conversion between Systema and Thermal Desktop. Additionally, a new "H5 converter" utility, available on demand, will be presented to easily convert standard text and Excel spreadsheets into the native Systema H5 format.

Following these software updates, a brief status update on the ongoing Systema V5 development will be provided to outline the next steps of its roadmap.

Finally, addressing the emergence of modern digital assistants, the presentation will conclude with practical insights on the synergy between AI and Systema. We will share concrete tips and best practices on how to leverage AI tools to accelerate the creation of Python scripts or to directly pilot the Systema environment through its API.

Speaker: Léa Galeron

ESTEW_2026_Systema_v1.pptx

15:30 Radian Developments in 2026: Cloud-Based Thermal Analysis Software

Radian is a thermal analysis software conceived to provide agility to engineers, both at modelling and computing processes. It is accessible through a regular web browser and relies on a scalable network of computing resources in the cloud. This paper reports the advancements introduced over the course of 2026, in the user interface as well as in the simulation engine.
The most prominent new capability is the support of orbital mechanisms. Geometries can now move along the mission, either following a time series of rotations and translations provided by the user or tracking a celestial body such as the Sun. View factors, radiative exchange factors and environmental heat fluxes are handled as epoch-dependent series, and the Monte Carlo ray tracer only recomputes the configurations that actually differ, keeping the computational cost bounded. Environmental fluxes are now also stored per face, which improves the post-processing of shells exposed on both sides.
The parametric analysis feature has been widened. Beyond material and optical properties, variations can now target geometry thickness, mass, boundary conditions and scenario parameters, and solar and planetary flux computations are distributed as parallel tasks over the cluster. Mission definition has been enriched with operational modes, which group heater and dissipation states and bind them to mission events, and with a generalised pointing attitude with configurable primary and secondary vectors and additional target bodies. Heliocentric orbits are now supported as well.
On the computing side, periodic solutions can be paused, inspected iteration by iteration and resumed from the stored temperature field, and the memory footprint of the ray tracer has been reduced substantially for large models. Interoperability has progressed with an importer for ESATAN-TMS geometry, linear and convective conductors, and with the extension of the STEP-TAS importer to carry the thermal mathematical model alongside the geometry.

Speakers: Bernat Frangi Mahiques, David Criado

ESTEW Thermal Workshop 2026 Keynote-0.1.pdf

14:00 → 16:00 Thermal Control

Convener: Paolo Ruzza (ESA)

14:00 A Composite Architecture for Radiation Shielding and Thermal Management

Modern space missions increasingly rely on high-performance commercial-off-the-shelf (COTS) electronics to maximize processing capabilities. However, traditional aluminum shielding forces design trade-offs: thick metallic walls add significant mass without addressing the substantial heat generated by modern processors. To overcome these size, weight, and power (SWaP) limitations, we developed a lightweight, non-metallic dual-layer polymer composite that combines radiation protection and thermal dissipation within a single conformal architecture. This system pairs a high-attenuation metal-oxide polymer composite (MOPC) layer for radiation shielding with a fiber-reinforced polymer composite (FRPC) layer engineered for exceptional lateral thermal conductivity (3.0 - 3.8 W/m·K). Ground evaluation demonstrated that the MOPC layer delivers up to seven times higher mass attenuation than aluminum in low-photon-energy regimes typical of electron-dominated space environments. The integrated material architecture successfully withstood 600 krad(Si) total ionizing dose and 500 thermal-vacuum cycles (-20 °C to 125 °C) without cracking or delamination, while meeting stringent NASA ASTM E595 outgassing standards. When applied as localized spot shielding, the composite architecture reduced total shielding mass by 70% to 85% compared to conventional aluminum enclosures. Finally, the environmental durability of this dual-layer technology was validated under operational flight conditions via integration onto camera control PCBs aboard a suborbital sounding rocket. The recovered flight hardware demonstrated pristine structural cohesion and mechanical strength, successfully surviving launch and atmospheric reentry forces. Together, these findings establish a scalable, multifunctional material platform that eliminates traditional structural-thermal boundaries, enabling the safe and efficient deployment of next-generation electronics in compact spacecraft.

Speakers: Dr Srujan Rokkam (Advanced Cooling Technologies, Inc.), Dr Nathan Van Velson (Advanced Cooling Technologies, Inc.)

ESTEW_26, (VanVelson) Composite Architecture for RadShield, Thermal Management.pdf

15:00 Lightweight insulation for high-temperature ISRU systems: Material characterization test results

Future lunar and Martian surface missions will increasingly rely on high-temperature processes for energy generation and in-situ resource utilization. Solid Oxide Electrolysis (SOXE) stacks are particularly relevant in this context, but their operation at temperatures typically between 600 and 1000 °C creates a major thermal control challenge. Heat losses must be minimized to reduce power demand, while surrounding spacecraft elements must be protected from excessive temperatures. These requirements are especially demanding in low-pressure CO$_2$ environments and in vacuum.

The ESA-funded LIGHTS activity addresses this need by developing lightweight high-temperature insulation solutions applicable to both Martian and lunar surface missions. We present the current status of the activity, focusing on the material characterization campaign carried out to support the integration of insulation concepts into a SOXE breadboard demonstrator. The campaign investigated selected microporous insulation materials by measuring their steady-state thermal conductivity in air, low-pressure CO$_2$, low-pressure N$_2$, and vacuum. Their thermal cycling stability was also assessed up to 850 °C under the same environmental conditions. These tests were complemented by post-cycling thermal conductivity measurements, enabling the degradation of the microporous materials to be evaluated. In addition, the campaign assessed outgassing and contamination behavior, mechanical response under compression, manufacturability aspects such as shaping, encapsulation, and dust mitigation, and electrical insulation properties at high temperature.

The results provide an experimental basis for comparing candidate materials under representative operating environments and for consolidating the insulation architecture to be implemented in the next project phase. In particular, the campaign demonstrated the relevance of microporous insulation for lightweight high-temperature thermal performance.

Future work will focus on the manufacturing, integration, and testing of a thermally representative breadboard based on a dummy SOXE stack. The breadboard campaign will evaluate the insulation performance at system level, including heat-loss behavior, external-surface temperature, mechanical integration, and performance stability over time under representative environmental conditions.

Speaker: Francisco J. Guerrero-Gonzalez (Maana Electric)

2026_Guerrero-Gonzalez_ESTEW.pdf

2026_Guerrero-Gonzalez_ESTEW.pptx

15:30 Design Update and Breadboard Validation of an Electrohydrodynamic Thermal Switch with Improved Volume Compensation

APR Technologies, a Swedish SME, has developed an updated electrohydrodynamic (EHD) thermal switch for spacecraft thermal control and validated the design through a breadboard test campaign. The EHD thermal switch transfers up to 150 W of heat in the ON state while consuming less than 3 W of electrical power, and requires no power in the OFF state. Since the device contains no moving parts, it is free from mechanical wear and does not generate vibrations.

Although previous generations of the thermal switch demonstrated good performance, sporadic reductions in ON-state thermal conductance were observed during testing. Root-cause analysis identified the free-floating gas bubble used for volume compensation as the cause of the observed performance instability. To address this issue, a revised design incorporating an external accumulator for volume compensation has been developed.

Breadboard testing in thermal vacuum evaluated the thermal switch over operating temperatures ranging from −30 °C to +70 °C and heat loads between 30 W and 150 W. Repeated measurements at identical operating points were conducted to assess long-term performance stability. ON- and OFF-state thermal conductance were characterized for combinations of thermal switch and accumulator temperatures to establish the required accumulator temperature relative to the thermal switch cold-side temperature. Thermal vacuum cycling was also performed, with performance measurements repeated at all operating points following the test.

The test campaign demonstrated stable performance of the revised thermal switch, with no degradation observed over the 25-day thermal vacuum campaign. An ON-state thermal conductance of 330–700 W/(m^2K) and an OFF-state conductance of 41–51 W/(m^2K) was measured across the tested temperature and power range.

Speaker: Dr Magnus Heldin (APR Technologies)

APR_THSW_ESTEW2026_DRAFT.pdf

APR_THSW_ESTEW2026_FINAL.pdf

16:00 → 16:30 Break

16:30 → 17:30 Heat Transport Technology

Convener: Emmanuel Caplanne (ESA)

16:30 Towards Multifunctional Satellite Structures: Embedding Pulsating Heat Pipes in Lightweight CFRP Laminates

Efficient thermal management in carbon fibre reinforced polymers (CFRPs) is crucial for many space applications. When used in the vicinity of heat-generating components, the relatively poor thermal conductivity of CFRPs requires additional measures to transport and remove heat, in order to keep the structure within its required operating temperature range. Prior work has explored the integration of thermally conductive materials, such as metallic inclusions, into the composite. However, this can introduce thermal expansion mismatches, degrade mechanical performance, and increase mass. This work presents a novel approach to thermal management in CFRP structures that embeds pulsating heat pipes (PHPs) directly within composite laminates. PHPs are passive, two-phase heat transfer devices that generate self-sustained pulsating fluid flow in a network of partially filled, meandering channels, driven by temperature-induced pressure oscillations. To fabricate CFRP PHPs, a sacrificial material vaporisation technique is employed to form continuous, sealed microchannel networks within the cured composite. This is achieved without introducing foreign materials or compromising structural homogeneity. Operation of the embedded PHPs demonstrated an over 400% improvement in thermal conductivity compared to solid CFRP, with performance governed by orientation, channel geometry, internal surface roughness, and anisotropic through-thickness thermal conductivity of the material. The pulsating fluid flow was captured in-situ using high-frame-rate 2D radiography during PHP operation. The utilisation of embedded PHPs could transform a CFRP structure into a multifunctional composite, combining load-bearing capabilities with integrated thermal management.

Speaker: Robert Oxford Pope (University of Bristol)

TowardsMultifunctionalSatelliteStructures_RobertOxfordPope.pptx

17:00 Thermal and microvibration characterization of a pulsating heat pipe for space application

Pulsating Heat Pipes (PHPs) are promising passive thermal control devices for space applications due to their high heat transport capability, design flexibility, and relatively simple wickless construction. However, the oscillatory motion of the working fluid may generate microvibrations that could affect sensitive spacecraft systems.
This work presents the thermal and microvibration characterization of two PHP breadboards developed within the MITIGATE project carried out for the European Space Agency (ESA). A large breadboard was developed for phased array antenna cooling with a target transported power of 200 W, and a small breadboard for detector plane cooling with a target of 30 W. Under nominal operating conditions, the large breadboard achieved a thermal resistance of 0.019 K/W at 200 W and 65 °C, while the small breadboard achieved 0.192 K/W at 30 W and 45 °C. Thermal performance and operational stability were strongly influenced by condenser temperature, orientation, and heat load, particularly near the operational limits. The presentation will discuss these effects and highlight the key operational trends observed during the test campaign.
Microvibration measurements demonstrated benign dynamic behavior, with RMS force levels well below the 0.2 N requirement. Occasional impulsive force events exceeding the specified peak-force limit were observed; however, additional tests without working fluid demonstrated that these events originated from the ground support equipment rather than from the internal two-phase flow. Overall, the results confirm the feasibility of PHP technology for space thermal control applications requiring efficient heat transport with low exported microvibrations.

Speakers: Artur Jurkowski, Marcin Wójcik

16:30 → 17:30 Thermo Elastic

Convener: Matthew Vaughan (ESA)

16:30 Thermoelastic Distortion Analysis and Optimization of a CubeSat Reflectarray Aperture

Reflectarray antennas (RAs) offer the gain of a parabolic reflector at a fraction of the stowed thickness, making them well suited to high-rate downlink from CubeSats. Once deployed, the large, thin aperture is directly exposed to the orbital thermal environment, and the resulting thermoelastic distortion (TED) degrades surface accuracy and beam pointing. Published work on deployable RAs focuses on electromagnetic performance; the chain from orbital thermal loading through structural deformation to antenna performance metrics remains under addressed in open literature.

This work presents a coupled thermal-structural analysis and optimization framework applied to a flight-representative X-band RA mounted on a 6U CubeSat. The study originates from ESA activity GT17-110EF and was selected for an In-Orbit Demonstration by the European Commission. Mission analysis over a quasi-polar SSO
defines the thermal cases across a beta angle range of 58° to near 90° throughout the mission duration, yielding panel temperatures ranging from approximately +40°C to −70°C. At high beta angle, the spacecraft body is continuously illuminated while the panel faces receive only grazing incidence, so the spacecraft hot case constitutes the antenna cold case. The thermal cases are evaluated through orbital thermal analyses and mapped to the structural model via a mesh export and a custom mapper, verified by comparing the through-thickness gradient across the dissimilar meshes.

A novel thermal control approach based on partial alteration of the thermo-optical properties of selected locations of the antenna is proposed, and its effectiveness is evaluated directly against the structural stability of the antenna, by accounting both BOL and EOL conditions. In addition to the thermal control strategies, the paper presents the structural choices taken to improve the antenna pointing and the results obtained are assessed for the envisioned employment in higher frequencies.

Speaker: Mark Gonzi (Robin Space GmbH, TUM)

ESTEW_2026_TED_of_Reflectarrays_Gonzi.pptx

17:00 Considerations for the Solar Coronagraph for OPErations (SCOPE)

The Solar Coronagraph for OPErations (SCOPE) is a coronagraph that has been developed at RAL Space to support Europe’s space weather monitoring capabilities. Recent development of the design has focussed on an adaptable approach that ensures the instrument can be accommodated on both an L5 and LEO mission. This presentation will cover the considerations and thermal analysis involved in the latest phase of development to achieve performance in these two environments.

The success of the coronograph depends on its optical performance, which the structural alignment of the instrument plays a role in. The structural alignment is influenced by thermo-elastic deformations which are determined by the outcome of the thermal analysis. Therefore there are no explicit thermal requirements on the coronograph, but an optical requirement that depends on collaboration between structural and thermal design to achieve. The STOP (Structural, Thermal Optical) analysis performed for both L5 and LEO options will be described, and the interdependencies of different engineering disciplines will be discussed.

Speaker: Dr Nicole Melzack (STFC RAL Space)

Melzack_ESTEW_SCOPE_Presentation (3).pptx

17:30 → 20:30 Cocktails

Wednesday 7 October

09:00 → 11:00 Heat Transport Technology

Convener: Paula Prado Montes (ESA)

09:00 Argon Loop Heat Pipe with Large Area of Heat Collection

Loop heat pipes (LHPs) are passive, two-phase thermal management devices capable of transporting large heat loads over long distances without the use of mechanical pumps. They are widely employed in spacecraft thermal control systems to transfer waste heat from payloads to radiators. This work presents the development and experimental evaluation of an argon loop heat pipe designed to operate at cryogenic temperatures near 120 K, investigated in two distinct configurations. The first configuration corresponds to a conventional LHP architecture commonly used in space applications, while the second employs a modified architecture intended to reject heat loads distributed over a large surface area. Thermal performance testing was conducted over input power levels ranging from 50 W to 100 W at sink temperatures near 120 K. Experimental results demonstrate stable operation of the LHP across the tested conditions, with the conventional configuration successfully transporting heat loads of up to 100 W. In the modified configuration, the LHP provided effective cooling of a large-area heat source up to 40 W applied on the large area when 15 W was applied to the capillary pump. This work was conducted under the DOE SBIR Phase II program (DE-SC0022896).

Speaker: Dr Calin Tarau (Advanced Cooling Technologies)

09:30 Passive Variable Heat Rejection System for Lunar and Planetary Habitats

NASA seeks new technologies that will facilitate low-mass and highly reliable thermal control systems for the solar system. There is interest in developing thermal control technologies that will enable crewed habitats to survive and operate through the lunar night for extended periods of time on the lunar surface where temperatures range from -193 °C or lower in shadowed regions (including night) to 120 °C at the equator. In addition, these technologies would allow habitats or pressurized rovers to operate in all these environments, transit from Earth to the Moon as well as Deep Space while meeting requirements like reduced mass, volume, and power usage, avoiding thermal control heaters, micrometeoroid and orbital debris (MMOD) robustness and protection potential, high heat rejection turndown ratio, freeze tolerance and low toxicity.
In response to the above stated requirements, Advanced Cooling Technologies, Inc (ACT) develops a high turndown ratio, freeze-tolerant and passive Variable Heat Rejection Radiator for Planetary Surface and Space Habitats based on multiple Non-Integrated Hot Reservoir Variable Conductance Heat Pipes (HR-VCHPs). Figure 1 shows a CAD design of the proof-of-concept prototype that is currently in development. Habitat waste heat is acquired by a single-phase pumped loop (SPL) using a benign fluid and further transferred to a series of these VCHPs charged with ammonia (or propylene depending on freezing and toxicity requirements) and argon (or neon, again, depending on the freezing requirements) that in turn transfer the heat (with self-adjusting resistance) to the radiators for ultimate rejection. In addition, the HR-VCHPs, which mainly consist of aluminum extrusions, will be flexible/deployable because of stainless steel bellows inserted as adiabatic sections that would also act as thermal barriers during survival. Therefore, the passive high turndown ratio of the system is provided by the following three components:
1. The hot biased reservoir that, by its nature, provides high sensitivity to the NCG front motion to modulate the conductance.
2. The fact that ammonia freezes relatively early during sink temperature decay (especially during the Lunar night)
3. The bimetallic section/bellows that, in addition to flexibility, acts as thermal barrier significantly reducing the parasitic dry conductance of the HR-VCHPs during low sink temperatures/survival.
An additional feature that would significantly contribute to control and survival power savings is the use of the Non-Integrated HR-VCHPs. It means that the VCHP reservoirs are installed inside the habitat on the SPL upstream of each corresponding evaporator. During survival conditions, to minimize parasitic conduction, the reservoir may need to be heated with a low control power to further push the NCG front all the way into the evaporator. This control power/energy will remain inside the habitat, and, at most, will count as habitat survival power to compensate for the already minimized conduction based parasitic leaks.

Speaker: Calin Tarau (Advanced Cooling Technologies)

ESTEW2026_ACT_Variable Conductance Radiator System for Space and Planetary Habitats.docx

10:00 Experimental Characterisaton of Two-Phase Thermal Spreader for Space Electronic Cooling

Thermal control is a key challenge in spacecraft electronic components, particularly for compact high power density units. Given the limitation of conventional aluminium ammonia grooved heat pipes regarding the maximum heat flux density, a two-phase spreader can be utilised as a primary thermal control technology to reduce the heat flux density to acceptable levels, allowing secondary heat transport equipment to transfer the dissipation to dedicated radiators. In this context Airbus with the help of CEA, developed an experimental test bench to characterise three types of thermal spreaders; a solid aluminium spreader (benchmark), a copper / water spreader and a nickel / water two phase spreader.
The two-phase spreaders were tested over a range of saturation temperatures (30-90°C), orientations and input powers (up to 250W) that correspond to high heat flux densities up to 50 W/cm². The thermal resistance and spreading efficiency were evaluated using direct temperature measurements and numerical modelling. The aluminium reference showed the expected conductive behaviour with limited spreading performance (R_total ≈ 0.10 K/W). The copper water spreader provided significantly lower thermal resistance (down to 0.04 K/W), with better temperature homogeneity (<2°C), and a more effective redistribution of heat at the condenser. The nickel/water spreader was also characterised under the same conditions with its performance found to be inferior compared to the copper/water spreader. This paper presents the activities that allowed the characterisation of the spreading technologies including the design of the experimental apparatus, the experimental results and the efficiency calculation.

Speaker: Dr Georgios Gkounis

Experimental Characterisaton of Two-Phase Thermal Spreader for Space Electronic Cooling.pdf

10:30 Recent Development of High-Temperature Alkali Metal Heat Pipes at ACT

Alkali metal heat pipes are passive two-phase heat transfer devices that utilize alkali metals such as potassium and sodium to transfer heat at high temperatures (>300°C). Advanced Cooling Technologies (ACT) has a long history of developing unique high-temperature alkali metal heat pipes for wide variety of applications such as nuclear power, solar power, and thermal energy storage. This presentation will highlight several recent RD programs at ACT to develop high-temperature and high-power alkali metal heat pipe systems. Highlighted programs include:
• The development of a unique high-temperature multi-material heat pipe system for transporting heat from a nuclear reactor to a Stirling engine as part of the NASA/DOE FSP program. The system used sodium-Kovar thermosyphons and a sodium-Haynes 230 loop thermosyphon.
• Active development of high-power wicks for alkali metal heat pipes for cooling nuclear reactors. Wicks have shown >10x improvement over conventional wick design.
• Large high-power heat pipes and thermosyphons, power > 10kW.
• A high-temperature (>600°C) thermal storage system using alkali metal heat pipe technology.
• An alkali metal heat-pipe heat spreader and pulsating heat pipe for a particle receiver board in concentrated solar power that reduced peak temperature by 130°C.
For each program, the application, design, prototype, and experimental results will be highlighted.

Speaker: Calin Tarau (Advanced Cooling Technologies)

09:00 → 11:00 Thermal Analysis

Convener: Matthew Vaughan (ESA)

09:00 Automated Thermal Control Design: A Coupled NLP-MILP Framework for Spacecraft Heater Sizing and Placement

Ensuring spacecraft components remain within strict temperature limits depends heavily on the strategic placement and sizing of active thermal control heaters. The industry standard for this design process is manual and iterative, which is not only time-intensive but often results in suboptimal power consumption. Furthermore, while metaheuristics have been proposed for automation, they fall victim to combinatorial explosion when applied to large-scale, flight-representative models. To overcome these scalability issues, this work presents a novel optimization framework that integrates Non-Linear Programming (NLP) and Mixed-Integer Linear Programming (MILP). By leveraging the direct sensitivity method to compute exact analytical gradients, the physical properties of the Thermal Mathematical Model are directly embedded into the optimization loop to solve discrete placement and continuous power allocation under steady-state conditions. We validate this approach using the flight-representative thermal model of the ARIEL Telescope Assembly, focusing on the highly coupled decontamination heater lines of the Primary Mirror and Telescope Optical Bench. Despite evaluating a massive search space of approximately $10^{19}$ configurations, the proposed algorithm successfully identifies optimal layouts in under 10 minutes. The optimized configurations satisfy all minimum temperature constraints and significantly outperform traditional manual designs in both power efficiency and thermal gradient reduction.

Speaker: Carlos Arroyo Ruiz (Universidad Politécnica de Madrid)

NLP_MILP_ESTEW2026_v4.pptx

09:30 Physics-Guided Neural Surrogate Modelling for Rapid Spacecraft Thermal Analysis

Spacecraft thermal analysis relies on high-fidelity numerical models, but repeated transient simulations across different orbital and operational conditions can be computationally expensive. This work presents a physics-guided modular neural surrogate developed using ESATAN-TMS to accelerate system-level transient thermal predictions.

The reference spacecraft model comprises 204 lumped thermal nodes organised into 16 physical subsystems. Environmental heat loads are estimated by two dedicated neural models. The first predicts direct solar radiation and planetary albedo outside eclipse, while a deterministic physics-based gate sets both contributions to zero during eclipse. The eclipse detector achieved complete agreement with ESATAN-TMS across all evaluated cases. The second model predicts planetary infrared radiation throughout the entire orbit, as this contribution remains present during eclipse. Temperature evolution is then predicted by a Stable Block Graph GRU combining local subsystem encoders, inter-block attention and a stable update formulation designed to limit autoregressive error accumulation.

The dataset comprises 10,000 ESATAN-TMS cases, each containing 1,001 time samples with a 10 s timestep and covering different orbital, attitude and internal-dissipation conditions. The cases were divided into non-overlapping training, validation and test sets using a 70%/15%/15% split. On the test set, the solar plus albedo and planetary infrared models achieved mean absolute errors of 0.0521 W and 0.0141 W, corresponding to 1.06% and 1.18% of their respective mean heat loads. The thermal surrogate achieved a global MAE of 0.0991 K over 997 consecutive free-run prediction steps. The mean node-wise MAE corresponded to only 0.225% of the nodal temperature excursion. The maximum absolute temperature error was 1.607 K, while the maximum sustained relative error was 2.94%. At the final prediction step, the MAE remained limited to 0.1184 K, confirming that error accumulation remained contained throughout the transient sequence.

Once trained, the surrogate evaluates a complete 10,000 s transient scenario in approximately 4 s, compared with 53–59 s for a single ESATAN-TMS simulation in the tested configuration. The proposed framework therefore enables rapid exploration of operating conditions and may also support future transfer-learning-assisted updating and correlation of high-fidelity thermal models.

Speaker: Alessandro Palomba (Politecnico di Milano)

Physics-Guided Neural Surrogate Modelling for Rapid Spacecraft Thermal Analysis.pptx

10:00 AI-Based Thermal Correlation with In-Flight Telemetry

Satellite thermal models are traditionally developed during the design phase and correlated using ground-based thermal vacuum tests. Their primary purpose is to predict temperature extremes under representative worst-case conditions and to demonstrate compliance with the spacecraft thermal requirements. While this approach is essential for spacecraft design and qualification, it provides only limited information on the dynamic thermal behaviour of the spacecraft during nominal operations. In particular, the actual in-flight temperature response results from the combined effects of the orbital environment, spacecraft operating modes, equipment dissipations and thermal control system behaviour, which are difficult to reproduce comprehensively during ground testing.
The use of in-flight telemetry provides an opportunity to further characterize this behaviour and to improve the representativeness of the thermal model under operational conditions. However, correlating a thermal model against flight data introduces additional challenges, as several physical phenomena and model parameters can produce similar temperature signatures. The identification of equipment dissipations is therefore an inverse problem involving a potentially very large number of parameters and requiring a large number of thermal simulations.
This work is in the frame of the ESA contract ESA/CNES study Rationalization of Thermal Simulators for Operations, led by DOREA with help of Thales Alenia Space.. This study focuses on the correlation of the Sentinel-3B thermal model developed by Thales Alenia Space. The starting model had previously been validated against thermal vacuum test measurements, providing a physically consistent basis for the in-flight correlation. The objective is to determine whether the model can reproduce the observed thermal behaviour in orbit while identifying parameters representative of the actual spacecraft operating conditions.
The proposed approach combines several levels of analysis. The orbital environment is reconstructed from the spacecraft position and orbital configuration, including eclipse entry and exit, variations in the Earth albedo resulting from the changing Earth projection, and other relevant environmental conditions. The scope was first to reach the convergence of the thermal model. The behaviour of the thermal control system is also analysed, including PI-controlled and thermostatically controlled equipment. Spacecraft operational modes are identified from telemetry to distinguish changes in thermal behaviour associated with different configurations and operating states.
The main identification task concerns the estimation of equipment thermal dissipations. Depending on the level of detail considered in the thermal model, several hundred dissipation parameters may potentially need to be adjusted within the ranges defined during the design phase. A systematic exploration of all possible combinations is therefore computationally impractical. Several optimization strategies are investigated, including gradient-based approaches and evolutionary algorithms, with the objective of efficiently minimizing the discrepancy between simulated and measured temperatures while maintaining physically meaningful parameter values.
The computational cost of the optimization process is addressed through the use of high-performance computing. minimizing the residuals (cost function) between simulated outputs and measured telemetry temperatures over the last orbit. A new convergence is needed by input dissipations. These steps resources provided by the CRIANN cluster. CRIANN is a French Normandy’s regional high-performance computing and digital research center. Its provides advanced computing infrastructure and expertise to support scientific research and innovation. The thermal simulations corresponding to successive optimization iterations are parallelized, allowing a large number of candidate solutions to be evaluated simultaneously and significantly reducing the overall computation time. Particular attention is given to the balance between optimization performance, computational resources and the physical interpretability of the resulting parameters.
The study also investigates the respective contributions of expert knowledge and automated optimization. While environmental conditions and thermal control behaviour can often be identified through expert analysis of telemetry profiles, the simultaneous estimation of a large number of equipment dissipations is better suited to automated optimization techniques. This combination aims to reduce the complexity of the identification process while preserving the physical understanding required to interpret the results.
The presentation will describe the methodology developed for the in-flight correlation, the main technological and computational challenges encountered, and the solutions implemented to overcome them. The resulting correlation performance will be presented together with an analysis of the identified thermal parameters and their consistency with the expected spacecraft behaviour. The respective advantages and limitations of the investigated optimization strategies will also be discussed.
Beyond the Sentinel-3B case study, the objective is to assess the applicability of this methodology to other spacecraft thermal models and operational scenarios. Improved knowledge of the actual in-flight thermal behaviour could provide a more accurate assessment of thermal margins and reduce uncertainties associated with thermal design assumptions. In the longer term, such an approach could contribute to more efficient thermal design processes and to the development of thermal simulators better suited to spacecraft operations.

Speaker: Mr Wayne Croisetu

ESA_Presentation_TCS_Correlation.pdf

10:30 Accelerating Spacecraft Thermal Analysis via Physics-Informed Neural Networks

Spacecraft thermal analysis traditionally relies on full scale physical simulations performed with industry standard tools such as ESATAN-TMS and Thermica. While essential for mission verification, these simulations can be computationally intensive, creating a major bottleneck for large-scale sensitivity analyses, design space exploration, and statistical Monte Carlo uncertainty quantification.
To address this limitation, OBO Space developed PIN-TAN (Physics-Informed Neural – Thermal Analysis Network), an end-to-end surrogate modelling framework, enabling real-time transient temperature predictions by combining deep neural networks with governing thermal physics equations.
The framework directly interfaces with ESATAN-TMS through a dedicated parser that ingests geometry files (.erg) and input decks (.d) to extract nodal properties (diffusive, arithmetic, and boundary), linear and radiative conductors, absorbed environmental fluxes and parametric design variables defined in GMM and TMM such as thermo-optical properties, material conductivities, contact conductances etc. A Latin Hypercube Sampling (LHS) engine explores the user-defined parameter space, orchestrates batch executions of the GMM and TMM solvers, and compiles the resulting temperature results into structured training datasets. The core surrogate relies on a Multi-Layer Perceptron (MLP) trained to predict multi-node temperatures as a function of time and operational parameters. The optimization is governed by a composite loss function that combines data regression with a physics-informed residual enforcing the nodal energy balance differential equation across conductive paths, non-linear radiative couplings, environmental fluxes, power dissipations and nodal thermal capacities. The complete workflow is packaged into an interactive dashboard, allowing engineers to monitor dataset creation, model training, inspect node-level validation, and perform real-time transient predictions and Monte Carlo sensitivity analyses.
Preliminary evaluations demonstrate that PIN-TAN achieves high physical consistency and thermal accuracy compatible with space engineering requirements for reduced models, while cutting evaluation times from traditional solver timescales down to sub-second responses. PIN-TAN provides space thermal engineers with a practical tool for rapid trade studies, thermal digital twins, and uncertainty quantification directly compatible with standard industrial workflows.

Speaker: Lorenzo Rabagliati (OBO Space)

ESTEW 2026 Rabagliati - Accelerating Spacecraft Thermal Analysis via Physics-Informed Neural Networks.pptx

11:00 → 11:30 Break

11:30 → 13:00 Thermal for surface missions

Convener: Giulio Tonellotto (ESA)

11:30 Surviving the Lunar Night: Passive MLI Shelter Design and Restart-Criterion Assessment for a 5 kg-Class Lunar Micro-Rover

Lunar night survival is a defining constraint on commercial micro-rover missions. Non-polar surface temperatures can fall to approximately 92 K, well below the storage limits of typical rover electronics and batteries, motivating approaches that do not rely on radioisotope heating.

This work evaluates the technical feasibility of a passive multilayer-insulation (MLI) shelter for a 5 kg-class micro-rover under non-polar lunar conditions. Following a tailored Phase-A systems-engineering approach, it combines a one-dimensional regolith model, which screens shelter reflectivity, deployment time and removal time to identify candidate operating windows, with a two-dimensional planar cross-sectional finite-element model that evaluates selected deployment timings and predicts internal temperatures for a two-compartment rover representation. Results are evaluated against published electronics and battery temperature limits, and contextualised by a functional decomposition, a traceable requirement set, and a weighted multi-criteria trade-off accounting for imported mass, stowed volume and the operational deployment window.

Shelter configuration and nominal MLI layer count strongly influence the modeled compartment minimum. At 15 MLI layers, the best sampled hemi-spheroidal case is approximately 24 K warmer than the best sampled box case, although these occur at different deployment times. The configurations differ in external radiating area and support-pole exposure as well as enclosure shape. For the 15-layer hemi-spheroid, the restart-critical compartment minimum varies by less than 1 K across the sampled 12:00–16:00 lunar local solar time (LLST) interval, with fixed component initial temperatures across deployment cases. Here, 24 LLST hours span one lunar solar day. At the next sampled deployment time, 18:00 LLST, the minimum falls below −40 °C. The selected baseline—hemi-spheroid, 15 MLI layers, 0.850 kg accounted concept-level mass excluding additional deployment and integration hardware—reaches a best sampled minimum of −38.05 °C. This meets a fallback battery storage-temperature criterion of −40 °C, but not the primary criterion of −20 °C for the selected commercial off-the-shelf (COTS) cell; it does not establish post-night restart capability. Two options warrant further assessment: substitution of a cold-rated cell, reducing specific energy from 248 to 207 Wh kg⁻¹, or active heating of the COTS cell. A preliminary ideal heating-energy estimate is approximately 38.6 kJ over the remainder of the night, equivalent to 7.3% of nominal battery energy, before accounting for system losses and reduced usable capacity at low temperature.

Key assumptions include nominal MLI effective emissivity, an idealised skirt-to-regolith interface and prescribed initial temperatures. Numerical verification, shelter-level experimental validation and battery restart qualification remain future work.

Speaker: Marc Alexander Wegener

02_10_Thermal_Workshop_2026.pdf

12:00 Active Thermal Control Mechanisms for Extreme Environments

With increasing interest and plans evolving for a range of lunar surface missions, there are considerable opportunities to deploy lunar payloads along with emerging plans for development of significant infrastructure on the surface of the moon. ESA, having anticipated such needs, have supported the development of new active thermal technologies which are generic in nature. These have included a dust-resilient lunar shutter and more recently a versatile retractable MLI device.
The shutter is a linear device referred to as LTS (formerly LDRLR) which began in 2020, is expected to reach TRL 5 in Q1 2027, while the retractable MLI device (REM), which started in 2024, is expected to achieve a similar status in Q4 of 2027.
Both devices have been developed with a view to be readily scaled, to support a wide range of applications. Both are suitable for radiator or solar panel applications, with a view to providing dust protection, while 3D nature of the REM is very adaptable to a range of surface applications, as well as for interplanetary missions. Specific surface applications would include surface payloads, docking ports, infrastructure service covers, radiators and rover storage facilities. Two versions of REM are being developed with a modular 90° actuation as the baseline and a full 180° version as an option.
Both developments have involved ESR Technology and Almatech, while Admatis joined the consortium as system and thermal lead for the Retractable MLI development. ESR have been responsible for both mechanism developments while Almatech have supported thermal and structural analysis of the shutter, as well as critical flex-pivot applications in both devices.
The presentation will provide an overview of both developments to explain the design approaches taken and the status of each. The results of early REM prototyping will be highlighted, as well as the evolution of the EM design, as established for the Detailed Design Review.

Speakers: János Szőke (Admatis), Matthew Oldfield (ESR Technology)

ESA-ESRT-HO-0180 01- Active Thermal Control Mechanisms.ppsx

12:30 Adaptive Heat Rejection for Lunar Surface Thermal Control Applications

Future lunar surface missions require thermal control systems capable of operating across the prolonged temperature extremes associated with the lunar day-night cycle. Unlike orbiters characterized by frequent thermal transitions, lunar surface systems must survive up to fourteen days of constant sunlight followed by 14 days of constant darkness while maintaining critical subsystems within operational and survival temperature limits, Conventional thermal control architectures often struggle to provide adequate control for both these extremes using a single system.
As part of the LUX-Thermal program, Lunar Outpost EU is developing a thermal control mechanism called “Thermal Cover” for providing adaptive heat rejection for lunar surface applications. The Thermal Cover employs a scalable, gravity-independent, radiator- and solar panel-covering mechanism made of rollable Multi-Layer-Insulation (MLI), capable of autonomously modulating heat rejection in response to changing environmental and operational conditions. By regulating surface exposure through an integrated feedback system, the technology enables dynamic adjustment of thermal performance while maintaining a mechanically simple and robust design.
By providing autonomous and variable thermal rejection tailored to the slow thermal cycling characteristics of the lunar environment, the Thermal Cover represents a promising pathway toward resilient and energy-efficient thermal control for future long-duration lunar missions.
Representative breadboards have been successfully tested in laboratory and vacuum environments, including dedicated performance characterization of the MLI and the mechanism functionality. The initial results confirm the feasibility of adaptive thermal rejection using the proposed architecture. A full-scale verification campaign is planned for the coming months to assess system performance, scalability, and operational robustness under mission-relevant conditions.
Moreover, a related mechanism called the “Thermal Curtain” is included in LUX-Thermal and focuses on opening and closing an aperture to a solar receiver cavity, which is used to heat high-temperature phase change material (PCM) that store heat for the system to utilize during the lunar night . Unlike conventional fixed architectures and existing variable thermal control solutions primarily developed for orbital applications, the proposed concept combines adaptive heat rejection, gravity-independent operation, scalability, and autonomous control within a single thermal control element tailored to the unique challenges of the lunar environment.
In addition, LUX-Thermal includes also several Active Thermal Switches (ATS), a compact device which actively changes the thermal conductance on-demand between a heat source and sink from a low to a high value and vice-versa.
All these technologies that are being developed at Lunar Outpost EU reduces excessive heat loss during lunar night while preserving efficient heat rejection during lunar day conditions. This capability has the potential to decrease heater power demand, improve system mass efficiency, and reduce overall thermal subsystem complexity compared with conventional fixed solutions. Owing to its scalability, this technology is applicable also to larger lunar infrastructures, including surface habitats, where parasitic heat losses significantly impact energy budgets during extended periods of darkness.

Speaker: Kaja Dabrowska (Lunar Outpost EU)

ESTEW2026_LOEU_Slides_v0.pdf

11:30 → 13:00 Thermal Testing

Convener: Arturo González-Llana (ESA)

11:30 K-CORE® RADIATOR PANELS TVAC TEST AND TMM CORRELATION: AN INNOVATIVE THERMAL CONTROL SOLUTION FOR HIGH POWER DENSITY CUBESATS

The HEliospheric pioNeer for sOlar and interplanetary threats defeNce (HENON) satellite is a 12U cubesat aimed at monitoring the Sun’s weather while orbiting around the Lagrangian point L2. In order to transmit almost real-time data back to Earth, the spacecraft requires a powerful yet compact communication suite, which comprises the satellite’s highest dissipating component. As a result, the spacecraft presents a very high power density, which poses a significant challenge from the heat dissipation point of view. Thermal management of the satellite has thus become a critical factor within the design process. To overcome this issue, two K-Core® radiator panels, provided by Boyd Thermal by Eaton, were chosen to be used on the highest-dissipating side of the satellite. These panels are composed of an annealed pyrolytic graphite (APG) core encapsulated in an aluminum shell. Such panels are intended to strongly increase the in-plane heat spreading capabilities of the spacecraft panels with respect to typical aluminum panels, in order to maximize the radiative capabilities of the satellite to dissipate the excessive heat towards deep space.

As the HENON satellite is set to launch early next year, a rigorous thermal test campaign on the two K-Core® radiator panels was required in order to obtain a correlated thermal mathematical model (TMM) of the panels to be coupled with the correlated TMM of the HENON satellite, resulted from the HENON STM test campaign.
The thermal test campaign was performed on the engineering model (EM) of the K-Core® panels at the Argotec facility in San Mauro Torinese.

This work presents an overview of the two custom-made K-Core® radiator panels, along with the rationale that led to their choice. The TVAC test campaign is then presented, comprising the rationale behind the test and the test set-up itself. Finally, TMM definition and the subsequent correlation process on the ESATAN-TMS software are presented, highlighting two different approaches that were considered: firstly, a layered model was adopted, where each layer of material comprising the K-Core® panels was represented by different geometries; then an equivalent model was analyzed, where the whole panel was modeled by a single geometry, to which equivalent properties were assigned. This work also presents the difference between the two modeling approaches and their respective results, along with the rationale that led to the ultimate choice.
This work is intended to provide guidelines for thermal design, test and correlation of such peculiar thermal management solutions, which could one day become common solutions for high power density deep space cubesats.

Speaker: Mattia Benericetti

Abstract Thermal Workshop 2026 - KCore Testing and Correlation.docx

K-Core Radiator Panels TVAC Test and TMM Correlation.pptx

12:00 Thermoplastic structures - modelling and testing

Thermoplastic composite (TC) structures offer a promising alternative to conventional
aluminium spacecraft architectures by combining low mass, high specific stiffness, and
increased manufacturing flexibility. This paper presents the thermal characterization
and thermal vacuum (TVAC) validation of the TeSat to assess the applicability of LM￾PAEK thermoplastic composites for small satellite primary structures.
The test article consisted of a ~100 kg class satellite platform featuring a primary
structure manufactured from LM-PAEK reinforced with carbon fibres. A comprehensive
TVAC campaign was performed, including bakeout at +80°C, thermal balance tests at
−40°C and +40°C, and thermal cycling between −40°C and +80°C under vacuum levels
below 1×10⁻⁵ mbar. More than 240 hours of thermal data were collected from
distributed temperature reference points located on structural panels, payload
simulators, avionics units, and external antennas.
The spacecraft successfully completed all test phases without structural degradation,
sensor detachment, or violation of operational temperature limits. During thermal
balance testing, the highest measured temperature reached approximately 76-80°C,
while the coldest structural regions reached −18°C, resulting in local thermal gradients
approaching 100°C across the platform. The thermal response was primarily driven by
internal power dissipation rather than environmental boundary conditions. In the cold￾active case, top patch antennas dissipating 14.5 W reached temperatures of
approximately 65°C despite the chamber environment being maintained at −42°C. A 15
kg steel mass simulating the SVPX assembly dominated transient behaviour, extending
stabilization times to up to 21 hours for the most demanding thermal balance cases.
Correlation of the thermal mathematical model against TVAC measurements revealed
significant discrepancies in the initial analysis. The non-correlated model
underestimated temperatures of composite walls and interface regions by up to 28°C,
mainly due to overestimated surface emissivity and uncertainties in composite-to-metal
thermal contact conductance. Following correlation, including adjustment of thermo￾optical properties and interface conductances, the maximum error was reduced to
14°C, while prediction accuracy for most equipment temperatures improved to within 1-
7°C of measured values. The average error for critical wall sensors was reduced by
approximately 50-70%.
The results demonstrate that LM-PAEK composite structures exhibit substantially
different thermal behaviour than conventional aluminium satellites. While the assumed
bulk thermal conductivity values (10 W/mK in-plane and 3 W/mK through-thickness)
were largely validated, metallic inserts and fasteners were found to create significant
conductive pathways that strongly influence overall thermal performance. The study
highlights the need for dedicated characterization of emissivity and composite-to-metal
interface conductance and provides valuable experimental data for future thermal
design, modelling, and qualification of thermoplastic composite spacecraft structures.

Speakers: Mr Marcel Daniluk (Creotech Instruments), Ms Sylwia Ciorga (Creotech Instruments / Warsaw University of Technology)

Prezka_ESTEW_2026_ opy.pptx

12:30 Orbital thermal cycling as an excitation source for on-orbit non-destructive inspection: passive shearography for large space structures

On-orbit inspection of large space structures mostly relies on visual and video checks, which reveal only major visible damage. Subsurface damage such as micrometeoroid and orbital debris impacts, disbonds in thermal protection and solar-array stacks, and thermal-fatigue cracking stays hidden until it affects structural integrity. The ground-based NDT methods that could find this damage largely depend on an active excitation source: flash lamps, heaters, shakers. In orbit, that excitation is limited by the available power, mass and safety budgets. As an alternative, the NASA EVA infrared camera flown on the ISS in 2005 demonstrated that solar heating and shadowing can be used to build a through-thickness temperature gradient suitable for infrared thermography inspection.
This presentation examines whether the orbital thermal environment can act as the excitation for a full-field, non-contact, strain-sensitive inspection. The technique is shearography (speckle-pattern shearing interferometry), which measures surface displacement gradients at the microstrain level. When compared to thermography, shearography inspection results support the assessment of structural integrity and residual life.
Within the ESA OSIP ShearScope project, an orbital sunrise was replicated in the TU Delft Optical Metrology for Aerospace (OMA) laboratory using power-modulated halogen lamps, reaching a temperature change of up to 100 °C. Under this replicated passive excitation, a disbond in an aluminium-cork thermal protection panel and a delamination in a carbon-carbon plate were detected, together with bonding issues in solar-array panels. In addition, FEM analysis was made to support and guide the detection limits of the technique for these materials.
The project presentation is supported with two contributions by MSc researchers. The first addresses the orbital thermal environment and the resulting thermo-elastic response, comparing various orbits and the foreseen temperature gradients to feed a thermomechanical model. The second addresses the inspection of solar arrays, where an orbital heat-flux profile is applied as excitation to detect delaminations, trapped-air bubbles in the adhesive layers and voids in the composite facesheet of a solar-panel. Both MSc projects conclude with simulated shearography phase maps.
We would like to discuss with the thermal engineering community the realism of laboratory-replicated orbital heating, the flux profiles and possible orbital cases. Representative (space-grade) specimens and in-flight thermal data are very welcome to start new collaborations.
This research is performed as part of the ESA OSIP ShearScope project (No. 4000148089).

Speakers: Mr Nan Tao (TU Delft), Mr Matthijs van Heck (TU Delft), Mafalda Caramelo Assis (TU Delft)

ShearScope ESTEW_3.pptx

13:00 → 14:00 Lunch

14:00 → 16:00 Heat Transport Technology

Convener: Paula Prado Montes (ESA)

14:00 Impact of Environmental and Life Qualification Campaigns on the Thermal Performance of Space-Grade Copper-Water Heat Pipes

Impact of Environmental and Life Qualification Campaigns on the Thermal Performance of Space-Grade Copper-Water Heat Pipes
Georgios Gkounis1, Kevin Lynn
(1) Airbus Defence and Space Ltd, Gunnels Wood Rd, Stevenage, SG1 2AS, UK, Email: georgios.gkounis@airbus.com

Thermal management is critical for high-power density spacecraft electronics, requiring robust heat transport technologies embedded within electronic structures. Copper-water heat pipes are a robust and cost-effective technology where their relatively small size delivers substantial performance compared to traditional AGHP. Their internal wick structure allows for operating capabilities in adverse tilt, thus minimising the orientation restrictions at system level testing. In this context, ADS and Boyd carried out a qualification campaign based on the ECSS-E-ST-31-02C, tailored to the specific needs of the space application. Flight-representative heat pipes were subjected to a series of tests including thermal endurance, mechanical/pressure as well as gamma radiation exposure.
Functional performance tests were performed across a range of saturation temperatures, input powers, and operating angles relative to gravity to fully characterize performance and capture end-of-life (EOL) performance. Following the environmental exposure to the test conditions, including gamma radiation, the heat pipes demonstrated a minor reduction in their thermal transport limit. Targeted low temperature diagnostic tests identified that the reduction in the performance was predominantly attributed to an increased temperature gradient localised within the condenser region, indicating a potential generation of Non-Condensable Gas (NCG). Under nominal operating temperatures, this NCG is compressed into the end of the condenser and imposes no operational constraints while the performance was within specification range. An equivalent (evaporating and condensing) heat transfer coefficient in the range of 15,000 W/m²K was derived and the test data was subsequently used for flight prediction. This paper outlines the qualification methodology and qunatifies the performance degradation at the end of a mission’s life.

Speaker: Dr Georgios Gkounis

Impact of Environmental and Life Qualification Campaigns on the Thermal Performance of Space-Grade Copper-Water Heat Pipes.pdf

14:30 Development of an Additively Manufactured Titanium – Water Loop Heat Pipe

Next-generation space missions require efficient, lightweight, and highly reliable passive thermal management systems to support extended operational lifetime. Loop Heat Pipes (LHPs) are a key thermal management technology for satellites, planetary orbiters, and surface exploration missions due to their ability to passively transport heat over long distance with minimum temperature drop. In this work, Advanced Cooling Technologies (ACT) and its collaborators have developed additively manufactured titanium-water loop heat pipes. LHP evaporators have been fabricated using Laser Power Bed Fusion (LPBF) technique. ACT carried out titanium LHP fabrication using additively manufactured titanium evaporator. The titanium evaporator consists of an integrated fine-pore primary wick fabricated directly through additive manufacturing, whereas a secondary wick was fabricated using a coarse-pore screen of titanium. Additionally, ACT conducted theoretical analysis to determine the LHP power carrying capability for a range of operating temperatures from 100 °C to 125 °C. Following the fabrication, ACT will conduct thermal testing to demonstrate the working of a one-of-a-kind additively manufactured titanium-water LHP. Initial testing will focus on transporting nominal heat loads on the order of several hundred watts to demonstrate stable LHP operation. Additional testing will investigate startup behavior at low powers, thermal transport characteristics, and effective thermal conductivity, while experimental results will be compared against analytical predictions. The results will provide insight into the fabrication feasibility and performance of additively manufactured titanium-water LHPs for next-generation space applications.

Speaker: Calin Tarau (Advanced Cooling Technologies)

15:00 Additive Manufactured Topology Optimized Cooling Structures for Electronic Equipment

Data processing capacity and power demand are rapidly increasing in telecom and scientific payloads. Together with system miniaturization, this trend often leads to performance limitations caused by insufficient cooling, which can limit the revenue obtainable from flight equipment. Optimized satellite structures and cooling systems are therefore key enablers for next-generation high-throughput small satellites and future space missions. This work presents the application of topology optimization to the design of an electronics frame for a high-power satellite application. The objective is to obtain a minimum-mass design that withstands demanding mechanical and thermal loads. The concept incorporates embedded heat pipes to improve thermal management through phase-change heat transport, and it is intended for additive manufacturing, allowing agile production of optimized geometries tailored to specific use cases. The Solid Isotropic Material with Penalization (SIMP) method was applied using HyperMesh and OptiStruct. The resulting electronics frame supports a printed circuit board inside a satellite, withstands launch mechanical loads, and dissipates heat generated by the electronic components. An in-house tool was developed to design the heat-pipes, predicting heat transfer and subsequently fed with empirical results (cross sections, filling ratios, heat loads). Current designs can extract up to 75 W locally at component level, translating into a power density of 34 W/cm2. Compared with a conventional frame, the optimized design achieves about 15% mass reduction while the heat-pipe concept reduces critical electronic temperatures by tens of degrees Celsius, increasing the available thermal margin for higher power throughput. Additionally, AIT activities are considerably simplified as externally mounted hardware such as thermal braids or traditional heat pipes are not required.

Speaker: Jaime Ruiz

Madrid Space - ESTEW 2026 - TopologyOptimizationElectronicFrame.pdf

Madrid Space - ESTEW 2026 - TopologyOptimizationElectronicFrame_v2.pdf

15:30 Titanium-Water Heat Pipe Radiators for Fission Surface Power – Structural Assessment

Advanced Cooling Technologies (ACT) conducted structural and thermal assessment of titanium-water (Ti-H2O) heat pipe and Graphite Fiber Reinforced Composite (GFRC) face sheet-based radiators for Fission Surface Power (FSP) applications. ACT previously developed and fabricated these radiator panels by directly bonding GFRC face sheets to titanium-water heat pipes. The heat pipes incorporated a proprietary wick structure within the evaporator. In addition, non-condensable gas (NCG) was used to provide startup assistance and freeze-thaw tolerance. The GFRC face sheets featured an unique alignment of carbon fibers to cancel the Coefficient of Thermal Expansion (CTE) mismatch with titanium in the axial direction. Recently, ACT performed numerical analysis of structural integrity of the radiators under launch vehicle shock and vibration environments. The analysis identified optimized structural reinforcements configurations that exhibited successful structural response to random vibration loads corresponding to launch conditions. Following the fabrication and assembly of structural reinforcements, radiators will undergo experimental shock and vibration testing. Thermal testing will be conducted before and after the shock and vibration testing and will be used to evaluate the response of radiators to launch vehicle loads. In addition, face sheet material and its bonding with titanium heat pipe will be characterized for any structural damage. Success of shock and vibration testing will be evaluated from the structural integrity of the radiators, and their power rejection capability. In addition, ACT conducted detailed freeze-thaw tolerance performance testing of the heat pipe using two different freezing methodologies. First, in the case of freezing only the condenser; whereas the evaporator remained warm, the heat pipe showed 12 successful freeze-thaw tolerant cycles of testing without failure. In the second case of freezing the entire heat pipe, a total of 10 successful freeze-thaw tolerance tests were conducted without any failure. Following each freeze-thaw cycle, the heat pipe regained its performance, and the radiator module rejected same amount of power at similar operating conditions. Finally, ACT performed a theoretical assessment of the probability of heat pipe failure and face sheet area loss resulting from micrometeoroid impacts over a 10-year operational lifetime in the lunar environment.

Speaker: Dr Calin Tarau (Advanced Cooling Technologies)

14:00 → 16:00 Thermal Analysis

Convener: Arthur Dunlop (ESA (ESTEC))

14:00 METIS CORONAGRAPH: FLIGHT DATA THERMAL CORRELATION

Metis, one of the ten instruments of the Solar Orbiter mission, is the first solar coronagraph capable of
performing simultaneous imaging of the off-limb solar corona in both visible and UV light. Due to the
uniqueness of the Solar Orbiter mission profile, Metis has already observed the Sun corona from a
minimum distance of ~0.28AU and is approaching higher latitude views with the increase of the orbit
inclination above the ecliptic plane.
Scope of this paper is to describe the correlation of the Metis temperature measured in flight during the
Solar Orbiter mission at different solar distances (from the first aphelion at 0.952 AU to the first perihelion
at 0.52 AU and to the one of the closest perihelion at 0.292 AU) with those predicted from the Metis
thermal model correlated with the ground tests.
Objectives of this correlation were to validate the prediction of the instrument thermal behaviour in all
flight conditions, to provide recommendations about the temperature ranges to be applied at the
interfaces between the Metis detectors (VLDA, UVDA) cold fingers and the spacecraft interfaces during
the scientific acquisitions and the VLDA annealing, and to compute the temperatures reached by the
internal optical elements in off-pointing from the sun-centred condition with the Metis external door open.
The comparison with the flight data demonstrated the validity of the Metis thermal model resulting from
the correlation with the data of the thermal-balance test campaigns performed first on the STM with Sun
Simulator in ESA/ETS facilities and finally on the PFM in TAS-I Rome facility. In fact, it has been verified
that the thermal model correctly predicts within the uncertainty values the temperatures measured inflight from the aphelion to the minimum perihelion.
Keywords: Coronagraph, Thermal Correlation, Sun simulator, ESATAN-TMS, flight data

Speaker: TIZIANO SCHILLACI (THALES ALENIA SPACE ITALIA)

ESTEW06_Metis_final_.pdf

ESTEW06_Metis_final_.PPTX

14:30 Improved post-processing workflow for space thermal design insight

The renewed global interest in large satellite constellations for high-bandwidth communication has accelerated the demand for accurate, efficient, and traceable thermal engineering tools for space programs. This paper presents recent developments in the TMG solver for thermal analysis workflows supporting the ongoing satellite communication space race, with emphasis on constellations operating in Very-Low Earth Orbit (VLEO). The study focuses on thermal design assessment during critical mission phases, including launch and early orbit, where rapid environmental transitions and coupled conductive-radiative effects can drive design margins and hardware qualification decisions.
A key contribution of this work is the development of a built-in reporting tool for allowable temperature margins within Simcenter 3D Space Systems Thermal. The tool enables thermal analysts to rapidly evaluate component-level temperature compliance across steady-state and transient load cases, identify hot and cold margin drivers, and communicate results in a structured and repeatable format. By consolidating margin calculations, allowable temperature limits, and case-by-case thermal results, the reporting capability reduces manual post-processing effort and improves consistency across design reviews, subsystem assessments, and program-level reporting.
The paper also presents a new graphical heat-flow representation capability that supports interpretation of conductive and radiative energy exchanges within the thermal model. Used together, the temperature-margin reporting tool and heat-flow visualization provide analysts with a more efficient way to diagnose critical thermal paths, understand the impact of boundary conditions, and prioritize design changes early in the development cycle. These capabilities are particularly valuable for spacecraft programs requiring frequent design iterations, large numbers of analysis cases, and clear traceability between simulation results and thermal requirements.
Representative steady-state and transient analyses are presented for launch and early-orbit scenarios relevant to VLEO communication spacecraft. The results demonstrate how automated margin reporting and heat-flow visualization can support rapid assessment of thermal compliance, improve communication between thermal analysts and system stakeholders, and strengthen confidence in early thermal design decisions. These developments enhance the efficiency, reliability, and scalability of spacecraft thermal control system analysis, contributing to high-throughput workflows needed for future satellite constellation programs.

Speaker: Jean-Frederic Ruel (Maya HTT)

ESA Thermal Workshop 2026 Improved post-processing workflow.pptx

15:00 UPMpyTMD: An Open-Source Tool for Efficient Post-Processing of ESATAN-TMS Thermal Models

Post-processing thermal mathematical models is often a bottleneck in the thermal engineering workflow, requiring repetitive manual checks and case-by-case analysis, typically relying on ad-hoc in-house scripts. UPMpyTMD is an open-source tool that reads TMD result files directly and provides fast, interactive post-processing.

The tool enables rapid model checking (node counts, conductor status, group hierarchy) in seconds, generates steady-state summary tables, and visualises the thermal balance of any group in an intuitive, appealing graphical format. For transient cases, it produces fully interactive temperature and flux plots. A dedicated module evaluates temperature gradients across mechanical interfaces (including multi-group conductive paths) to automatically identify the worst-case instant for thermoelastic structural analysis.

UPMpyTMD has already supported the thermal analysis of SUNRISE III, ARIEL, Vigil PMI E-Unit, and UPMSat-3, validating it as a reliable tool for space thermal engineering workflows.

Speaker: Dr Alejandro Fernández-Soler (IDR/UPM)

UPMpyTMD_ESTEW2026.pptx

15:30 Pycanha: current status of an open-source thermal analysis tool

Pycanha is an open-source tool for thermal analysis based on the lumped parameter method. It is written in C++ and Python. Data structures, solvers and any other computationally intensive operations are implemented in C++ while the user interface is implemented in Python and installed as a regular Python package, currently available for the latest versions of Windows, Linux and macOS.

Pycanha offers simplicity and usability while still providing advanced users with complete control over the thermal model data and solver processes, and great introspection and debugging capabilities through standard Python tools. Pycanha uses state-of-the-art linear algebra libraries for its solvers and GPU accelerated raytracing for the calculation of REFs and view factors.

Since the tool was first presented, the internal geometrical model has been remade and connected to the thermal mathematical model and to a 3D interactive viewer. Parametric capabilities have been expanded, and interoperability with other tools and standards, mainly ESATAN-TMS and STEP-TAS, through import and export operations, has been introduced. The tool is still evolving quickly, and more features are expected soon.

In its current state, Pycanha has the full set of tools that allow a complete stand-alone thermal analysis: geometrical modelling, automatic conductive and radiative couplings generation, steady-state and transient solvers and interactive post-processing.

Speaker: Javier Piqueras Carreño

PycanhaESTEW2026.pptx

PycanhaESTEW2026_v4.pptx

16:00 → 16:30 Break

16:30 → 18:00 Thermal Control

Convener: Andreas Mussger (ESA)

16:30 STARTEC, thermo-electric coolers for star trackers

Under the “European Thermo-Electric Cooler for Star Trackers” ITT Activity in the esa-star Tendering system, Azimut Space GmbH, in collaboration with the Fraunhofer Institute for Physical Measurement Techniques (IPM), embarked on this project in order to develop a European supply chain of thermo-electric coolers (TECs, but also known as Peltier coolers) for European star trackers. Furthermore, this activity encompasses the design, manufacture, and testing of TECs suitable for star tracker applications, which in turn, are specified by European producers of star trackers.

The TECs were designed based on the performance requirements from the star tracker manufacturers and consequently manufactured and screened by Fraunhofer IPM. In turn, environmental and reliability test campaigns were carried out by Azimut Space.

The project was split into a preliminary phase, a de-risking phase and a functional verification phase. During the preliminary phase, a single TEC design was developed and used for assessing and overcoming manufacturing and testing challenges. The de-risking phase was introduced as manufacturing challenges were indeed encountered, so it was necessary to identify failure causes before moving on to the next phase. Finally, during the functional verification phase, two TEC designs were developed using the previously established manufacturing process, as well as being able to survive realistic (and expected) environmental conditions (mechanical shock and vibration, thermal cycling, operational measurements) and to keep an acceptable BOL/EOL performance decay. Upon the completion of the project, the TECs will achieve TRL6 and set the path for a full qualification campaign for space applications, aiming to reach TRL 8.

Speaker: Mr Sebastian Ospina (Azimut Space GmbH)

ASG_Startec ESTEW 2026.03.pptx

ESTEW2026_Abstract_STARTEC.docx

17:00 Advanced Manufacturing Phase Change Material Heat Accumulators for New Space

Phase Change Materials (PCM) offer the possibility to store thermal energy directly as latent heat of fusion. This old principle is now coming on the scene due to its numerous advantages: stability of temperature control, absence of moving parts, reduced radiator size, reduced housekeeping power.

Advanced manufacturing techniques allow now to propose low-cost solutions. Adaptation to tricky volumes is also an advantage. The main benefits and challenges of series production are presented together with some practical applications.

Speaker: JP Collette (Walopt)

abstract Thermal Workshop ESTEC 2026.docx

PCM_CRM_ESTEW_2026_NN final.pptx

17:30 Magnetocaloric Thermal Control for Space Applications

The current research will explore the feasibility of employing a heat pump loop based on
active magnetic regeneration (AMR) of magnetocaloric materials for enhancing standard
heater/radiator-based thermal regulation on small (3U–6U) LEO spacecraft, which would
involve moving heat from heated regions to cold regions without rejection or independent
resistive regeneration. A rotary Halbach array type of AMR device using gadolinium is
analyzed according to a representative orbit and compared to radiators, resistive heaters,
heat pipes/variable conductance heat pipes, and PCM energy storage units. The estimated
active-mode electrical power (0.9–3.3 W) compares favourably to reported eclipse heater
loads (3–6 W), but this is achieved at materially higher mass (280–810 g), control com
plexity, and reliability risk, and at a substantial technology-readiness gap (TRL 2–3 versus
TRL 9 for every conventional alternative). The radiator is not eliminated under any op
erating mode. Drawing on published AMR regenerator performance data [11, 12], this
assessment further highlights regenerator geometry (aspect ratio, particle size) and operat
ing frequency as primary design parameters not accounted for by the current mass/power
sizing analysis, which should be the focus of future breadboard tests. This work is clearly
meant to be an initial order of magnitude feasibility assessment, not a detailed design.

Speaker: Patryk Cwik

ESA_Submission (4).pdf

ESA_Workshop_Presentation v3.pptx

16:30 → 18:00 Thermal Design

Convener: Vito Laneve (ESA)

16:30 In flight thermal correlations of ECLAIRs instrument Thermal Control System

SVOM (Space-based multi-band astronomical Variable Objects Monitor) mission objective is a thorough monitoring of Gamma Ray Burst (GRB) phenomena. Based on a collaboration between France (CNES: French Space Agency) and China (CNSA: China National Space Administration and CAS: Chinese Academy of Science), the SVOM payload includes four scientific instruments (ECLAIRs and MXT provided by CNES, GRM and VT provided by CNSA/CAS) installed on SVOM spacecraft. On Earth, several telescopes and a data center contribute to GRBs observations in addition to SVOM payload.

This paper covers the ECLAIRs instrument, designed to detect GRBs autonomously in near real time in the X-Gamma ray energy range, and then to quickly transmit to the ground telescopes their direction in the universe. After a GRB detection, a change of the spacecraft orientation points the other instruments in the GRB direction. Therefore, the ECLAIRs TCS (Thermal Control System) main constraint is to allow ECLAIRs instrument to detect GRBs whatever the attitude and the position of the spacecraft on its orbit. Indeed, the ECLAIRs TCS copes with a large variation of external environment during the mission with low heating power allocations. Thus, the ECLAIRs detection plane thermal bus uses Variable Conductance Heat Pipes (VCHP) with a regulation loop, in order to minimize heating power consumptions and to allow temperature control on the detectors. The presentation of the ECLAIRs instrument TCS highlights its development and qualification phases. Moreover, it presents also the in-flight thermal correlations that were performed, as well as a discussion about the lessons learnt - including real in-flight examples.

Speaker: Yann Cervantes

ESTEW_CNES_2026_VF2.pdf

17:00 Thermal design challenges in the preliminary architecture of ASLSTR, successor to the Sea and Land Surface Temperature Radiometer

The Advanced Sea and Land Surface Temperature Radiometer (ASLSTR) represents the next-generation successor to the SLSTR instrument flying on Sentinel-3 missions, with enhanced performance requirements in terms of radiometric accuracy and lower detector operating temperatures. Achieving these objectives requires a robust thermal design supported by thermal modelling and iterative system-level analysis. The preliminary design phase has highlighted a complex balance between demanding detector temperature requirements, calibration stability needs, and constraints imposed by the evolving instrument configuration. Compared with the previous generation, the new instrument accommodation leads to a different exposure to external environmental heat fluxes, placing additional emphasis on instrument-level radiative shielding. Particular attention is given to the optical bench assembly, composed of two optical benches, with emphasis on the section accommodating the TIR detectors, which represents one of the main thermal design drivers at instrument level. The contribution will present the thermal modelling activity, focusing on the thermal challenges emerging from the ongoing architecture definition.

Speaker: Edoardo Maria Benivegna

Thermal design challenges in the preliminary architecture of ASLSTR, successor to the Sea and Land Surface Temperature Radiometer.pptx

17:30 Thermal Control and Marangoni Effects in a Two-Phase Flow of the METRO ISS Experiment

The METRO experiment, being developed for the International Space Station, uses a gas–liquid slug-flow loop to investigate mass transfer in microgravity. This contribution presents the thermal control concept and CFD analysis of the METRO fluid loop, focusing on the influence of thermal effects on multiphase flow behaviour.

Temperature gradients across the gas–liquid interface generate surface-tension gradients (Marangoni stresses) that can significantly affect interface shape, slug motion and flow stability. The simulations investigate these effects under representative operating conditions and assess their importance relative to other forces governing the flow.

Particular attention is given to the temperature dependence of surface tension as a potential means of controlling and scaling two-phase flow for ground-based drop-tower microgravity experiments. The study explores whether thermocapillary effects can be deliberately exploited to reproduce relevant microgravity flow behaviour at different geometrical scales.

The contribution provides insight into the coupling between thermal control and two-phase flow dynamics and supports the design, verification and ground testing of the METRO experiment.

Speaker: Jiri Teichman

estew2026_teichman_Thermal_Control_and_Marangoni_Effects_in_a_Two-Phase_Flow_of_the_METRO_ISS_Experiment_final.pptx

18:00 → 20:30 Networking buffet

Thursday 8 October

09:00 → 11:00 Thermal Analysis

Convener: Duncan Gibson (Telespazio BE for ESA)

09:00 Assessment of the thermal model uncertainty from flight data comparison with thermal model predictions: Sentinel 1C case study

Thermal uncertainty margin philosophy has been one of the critical factors in ensuring space mission success since the dawn of space era. Building on a review of its historical evolution, this work aims to quantitatively assess the current state of thermal margins, by comparing in-flight temperatures with thermal model predictions for the Sentinel 1C satellite.
Specifically, the platform’s detailed thermal mathematical model (TMM) has been updated with registered telemetry data – including unit dissipations, attitude profile and heater line statuses – collected during Summer Solstice period. The refined model’s predictions were benchmarked against flight thermistor measurements from the same interval. The same process was then repeated on the non – correlated thermal mathematical model of S1C, isolating the impact of the correlation exercise.
Statistical post-processing of the results – focusing on bias, standard deviation, Uncertainty Flight Prediction (UFP) and its associated confidence level – revealed discrepancies between predicted and in-flight temperatures on both the mathematical models, but highlighted the critical role of the correlation exercise in reducing satellite UFP. In particular, considering a confidence level of 98% the calculated uncertainty decreases of 46% (from 13°C to 7 °C).
These findings highlighted persistent gaps in predictive accuracy during on-orbit operations, primarily driven by the inability of accurately replicate the unit dissipations (due to missing dedicated telemetry) and missing harness thermal effect. Conversely, the results outlined the vital role of TMM correlation in narrowing the uncertainty on temperature predictions.

Speaker: Francesco Latella (Thales Alenia Space)

Assessment of the thermal model uncertainty from flight data comparison with thermal model predictions Sentinel 1C case study_V2.pdf

09:30 Integrated Thermal Analysis and Battery Energy Balance Assessment of a Spacecraft Orbital Transfer and Return Vehicle Using TAITherm

This paper presents a transient thermal analysis with integrated electrical power assessment of the Spacecraft Orbital Transfer and Return Vehicle (OTRV) using TAITherm, with particular emphasis on Spacecraft thermal control and battery thermal management during orbital operations. Spacecraft incorporates both active and passive thermal control systems, including a fluid-loop heat transport system, thermostatically controlled heaters, multilayer insulation (MLI), and radiator systems including a variable-emissivity radiator. The thermal mathematical model incorporates Spacecraft geometry, material thermo-optical properties, conductive interfaces, internal heat dissipation, and the relevant thermal-control components. The orbital thermal environment accounts for direct solar radiation, Earth infrared emission, albedo, eclipse transitions, Spacecraft attitude, and mission-dependent orbital conditions.

Multiple orbital conditions and operational modes, including safe mode, detumbling, and payload operations, are analysed to determine Spacecraft and subsystem temperature excursions and identify limiting hot and cold cases. Particular attention is given to the battery subsystem, where transient temperatures are assessed against operational limits. Battery heater-control strategies are evaluated in terms of required heater power, duty cycle, and energy consumption. The contribution of the fluid-loop system and radiator configuration to Spacecraft heat transport and heat rejection is considered as part of the overall thermal-management strategy.

The simulations are performed using the finite difference-based solver TAITherm for three-dimensional transient thermal analysis, including multi-bounce ray-tracing view-factor calculations. The thermal analysis is coupled with the Spacecraft's flexible photovoltaic arrays to account for the interaction between solar-array temperature, photovoltaic power generation, Spacecraft power consumption, and battery energy balance. Transient photovoltaic output is evaluated considering solar incidence angle, Spacecraft attitude, self-shadowing, eclipse periods, and temperature-dependent cell efficiency. The generated electrical power is used to supply Spacecraft loads and charge the battery, enabling the thermal and electrical energy behaviour to be evaluated over a common orbital timeline.

The integrated analysis enables the identification of thermally critical orbital and operational conditions, assessment of component and battery thermal margins, and evaluation of the influence of eclipse periods and operational modes on battery temperature and heater demand. The model supports assessment of the combined active and passive thermal-control architecture, including heater operation, fluid-loop heat transport, insulation, and radiator performance, together with battery heater sizing and heater-energy requirements. The resulting methodology provides a Spacecraft-level approach for evaluating the coupled thermal management, battery thermal performance, and electrical energy balance of Spacecraft during orbital operations.

Speakers: NAGAPP PRADHANI (Atmos Space Cargo gmbh), Mr Sacha Jelic (ThermoAnalytics, GmbH)

TAI_ATMOS_ESAthermalworkshop2026_v04.pdf

TAI_ATMOS_ESAthermalworkshop2026_v04.pptx

TAI_ATMOS_ESAthermalworkshop2026_v05.pdf

TAI_ATMOS_ESAthermalworkshop2026_v05.pptx

10:00 Thermal Vacuum Testing of the MIRMIS Instrument for ESA's Comet Interceptor Mission and Automated Thermal Model Correlation

The Modular Infrared Molecules and Ices Sensor (MIRMIS), developed for ESA's
Comet Interceptor mission, underwent a dedicated thermal vacuum (TVAC) test
campaign at ESA/ESTEC. The campaign was performed in the Medium Vacuum
Facility (MEVAF) to characterize the instrument's thermal behavior and support the
correlation of its Thermal Mathematical Model (TMM). The MIRMIS STM was
subjected to cold and hot survival conditions, including different levels of simulated
solar illumination. Five representative steady-state cases were selected, using
temperature and survival-heater power measurements as test references.
A detailed thermal model of the MIRMIS instrument and the MEVAF test environment
was developed in ESATAN-TMS. The five TVAC cases were implemented as
independent thermal analyses and compared with the corresponding test measurements.
To automate the correlation process, a dedicated Python-based framework was
developed and coupled with TMM. The framework updates the correlation factors,
executes the thermal analyses, extracts the results, evaluates the correlation, and stores
the solutions throughout the iterative process.
A key feature is the parallel execution of the five TMM TVAC cases. This allows all
cases to be evaluated simultaneously, with the computational time of each iteration
mainly governed by the slowest case rather than the sum of their execution times.
MIRMIS was used as a representative test case to demonstrate the integration of TVAC
testing, detailed Thermal Mathematical modelling, and automated parallel correlation.
The developed approach reduces manual intervention while improving the efficiency,
repeatability, and traceability of thermal model correlation.

Speaker: João Castanheira (Synopsis Planet)

ESTEW_2026_Synopsis_Planet_submited.pdf

ESTEW_2026_Synopsis_Planet_submited.pdf

ESTEW_2026_Synopsis_Planet_improved ESTEW_2026_Synopsis_Planet_improved_submited.pdf (Protected)

10:30 Comet Interceptor Scan Mirror Assembly : Thermal Model Correlation

Comet Interceptor is a fast-class mission in ESA’s Cosmic Vision Program comprising three spacecrafts. With the challenge of a short development process, it will be the first mission to visit a Long Period Comet that is starting its journey in the inner Solar System. These objects are difficult to target because while detected, a small amount of time remains before leaving again the inner Solar System. The Comet Interceptor mission aims to be deployed as soon as such objects are detected to get valuable observations on material untouched since the early millennials of the Solar System.

The Rotating Mirror Assembly (RMA) is a mechanism that is mounted on the Comet Interceptor spacecraft A. Its main function is to rotate the entrance mirror facing the Comet Camera (CoCa) instrument. The rotation allows to perform continuous observations during the flyby, without moving the spacecraft. It is composed of 2 elements: the Scan Mirror Assembly (SMA) which is the contribution of the Centre Spatial de Liège (CSL), and the Scan Mirror Electronics (SME) which is a Thales Alenia Space Switzerland (TASCH) contribution, driving the mechanism and electrically interfacing with the spacecraft.

The SMA is constituted of two major assemblies: the Optical Assembly, including a flat mirror, optical baffles and a dust shield for protection against dust particles from the comet tail during fly-by; and the Drive Mechanism at its root, interfacing with the spacecraft and the SME. The main function of the SMA is to reflect the comet object to the CoCa telescope during the whole flyby, without observation interruption.

The end of the qualification campaign of the SMA Qualification Model (QM) took place with the TVAC tests, with a duration of 2 weeks during the end of May and the beginning of June 2026. The performances of the mechanism as well as its behavior under representative hot and cold conditions have been assessed successfully. With the temperature monitoring of the SMA during the campaign, the TVAC test enables access to experimental data that will help to build an accurate and efficient thermal model of the SMA.

In this presentation, the thermal correlation of the TVAC model is discussed with the associated thermal flight model and updated predictions. The differences between both models are discussed. The methodology and the correlation process are presented, and the main difficulties encountered are discussed. A conclusion summarizes the lesson learnt until now as a young graduate thermal engineer.

Speaker: Célestin Libert (Centre Spatial de Liège (CSL))

261001_ESTEW26_presentation_celestin_libert_v1.1.pdf

261001_ESTEW26_presentation_celestin_libert_v1.1.pptx

09:00 → 11:00 Thermal Testing

Convener: Andreas Mussger (ESA)

09:00 LISA Optical Metrology Subsystem Thermal Testing at CNES – Setup and Thermal Environment Characterization PART 1

LISA (Laser Interferometer Space Antenna) mission objective is to detect low frequency gravitational waves from space, opening a new era in astronomical science observations. Driven by ESA, the mission relies on strong partnerships with international agencies and scientific laboratories. The French instrument contribution to the LISA mission is managed by CNES with the support of French institutes such as APC (Astroparticule et Cosmologie), ARTEMIS, CEA-IRFU, LT2, L2IT, CPPM, LPC Caen and LMA. The French instrument contribution oversees the definition and realization of the ground performance testing of the optical metrology system (OMS), which is the metrological heart of the MOSA instrument aboard LISA. In this frame, end-to-end performance tests of the OMS will be performed in CNES facilities.
This paper presents the OMS test setup definition and its main thermal engineering constraints. The LISA mission requires a drastic thermal stability and a unique requirement on thermal noise filtering over the frequency range 1mHz – 1Hz. For this performance test, the optical assembly and electronic units are separated into two vacuum chambers to minimize perturbations. In the optical chamber, an optical stack composed of the Optical Bench (OB), the Beam Simulator (BSIM) and the Test Mass Simulator (TMSIM) is integrated. The OB is provided by ESA and is the heart of the system where interferometric measurements are performed. The BSIM is provided by APC and simulates the laser source coming from a distant satellite as well as the presence of a second MOSA. The TMSIM is provided by CEA and simulates the gravitational reference sensor. The electronic chamber welcomes four electronic boxes, including the LASER provided by NASA.
The modelling of the frequential domain temperature levels now needs to be confronted with real measurements. A first thermal characterization test of the setup, without any critical hardware, was conducted in 2026. The goal of this test is to measure the noise levels inside the optical chamber and demonstrate the ability to predict correctly the thermal noise propagation. This paper presents the setup and the results of this first thermal characterization test.

Speaker: Corentin Buti (Centre National d'Etudes Spatiales - CNES)

Abstract_ESTEW_2026_LISA_CBU.pdf

ESTEW_2026-CNES-LISA_Thermal_Carac_PART_1.pptx

09:30 Frequency-Domain Identification of Thermal Contact Conductance under Ambient and Vacuum Conditions

Thermal contact conductance is one of the main sources of uncertainty in spacecraft thermal models. Its experimental characterization is generally performed under vacuum because the presence of convection and gas conduction at ambient pressure can significantly affect the measured thermal response. This requirement increases testing complexity and restricts the characterisation of thermal interfaces during early development phases.  

This works presents a methodology to evaluate the contact conductance with tests in the frequency domain. The proposed method applies periodic heating at different excitation frequencies and evaluates the resulting phase lag across the mechanical interface. Ideally, the phase lag should be independent of the ambient pressure influence and still contains enough information to estimate the contact conductance value. Thus, this method could in principle be used without a vacuum chamber.  

An experimental campaign was conducted to develop and assess the methodology, including the selection of excitation frequencies, the treatment of transient drift, and the sensitivity to heater and temperature-sensor positioning.  

A first validation experiment looked into just the thermal conductance to assess the test set-up. The principal validation experiment consists of two mechanically joined aluminium plates tested under ambient and vacuum conditions.  

The conductance values identified in both environments are compared to determine whether suitable excitation frequencies can reduce the influence of atmospheric heat-transfer mechanisms and preserve sensitivity to the solid-contact conductance. 

Preliminary observations indicate that the method is promising, although the experimental validation is still ongoing. If successful, the method could provide a faster and more accessible approach for preliminary characterization of spacecraft thermal interfaces, while retaining vacuum testing for final qualification and verification.

Speaker: Alejandro Barbosa Lizarbe

ESTEW 2026 Alejandro Frequency Domain FINAL.pptx

10:00 Frequency-Domain Correlation of a LISA Payload Thermal Model

The Laser Interferometer Space Antenna (LISA) mission will detect gravitational waves through picometre-level interferometric measurements between three spacecraft separated by 2.5 million kilometres. Achieving the required sensitivity demands accurate prediction of the thermo-optic and thermo-elastic behaviour of the Payload Module. Within the Optical Bench Assembly (OBA), the Quadrant Photoreceiver (QPR) is one of the dominant thermal disturbance sources, and thermal transfer functions describing the propagation of temperature fluctuations through the instrument are directly used in performance analyses. Consequently, confidence in both the magnitude and dynamic behaviour of the thermal model is essential.

To support these activities, a dedicated thermal test campaign was performed on three pre-Structural Model (pSM) QPR units incorporating different thermo-optical surface treatments. In addition to conventional steady-state thermal balance correlation, the campaign investigated methods for dynamic model validation. Several excitation approaches were considered, including discrete single-frequency sinusoidal inputs and broadband thermal excitation techniques. A pseudo-random binary heater input was ultimately selected as it enabled efficient excitation across a wide frequency range within a practical test duration while maintaining representative thermal operating conditions. The resulting temperature measurements were processed in the frequency domain to derive transfer functions and power spectral densities relating thermal inputs and responses.

Measured frequency responses were compared directly with transient thermal model predictions. The correlation exercise identified deficiencies associated with thermal capacitances, contact conductances and radiative coupling paths, leading to an improved model representation and significantly enhanced agreement between prediction and test. The correlated model demonstrated good reproduction of both steady-state temperatures and dynamic thermal behaviour, providing confidence in the transfer functions subsequently used within the LISA thermo-optic performance analyses. Furthermore, the work demonstrates that frequency-domain thermal testing can provide substantially greater insight into model fidelity than traditional thermal balance testing alone and offers a practical methodology for future spacecraft thermal verification programmes.

Speaker: Arthur Dunlop (ESA (ESTEC))

Frequency-Domain_Correlation_of_a_LISA_Payload_Thermal_Model.pptx

10:30 TVAC experimental verification of a temperature-induced excess noise subtraction methodology for sub-millihertz ultra-low-noise space magnetometers.

In recent years, nanosatellites have enabled the testing and in-orbit validation of advanced instrumentation intended for larger space science missions. MELISA is a magnetic measurement payload designed to demonstrate the in-orbit performance of Anisotropic Magnetoresistance (AMR) sensors featuring dedicated noise reduction techniques at sub-milliHertz frequencies (0.1 mHz to 1 Hz). This bandwidth is critical for future space-borne gravitational wave (GW) observatories, where the local magnetic environment yields a significant contribution to the overall noise budget, requiring demanding noise floors down to $10\,\rm nT\,Hz^{-1/2}$.

The in-orbit characterisation in Low Earth Orbit (LEO) presents a major challenge for evaluating sensor performance at ultra-low frequencies, as environmental thermal and magnetic variations become the dominant contribution to the total noise budget, severely impeding the characterization of the intrinsic performance of the sensor. Nevertheless, the thermal noise contribution to the magnetic field measurements can be disentangled and subtracted through on-ground cross-correlation with onboard housekeeping data and post-processing.

This work presents the initial steps toward demonstrating the capabilities of a methodology for subtracting temperature-induced excess noise applied to MELISA data. The unit is integrated into a Thermal Vacuum (TVAC) chamber to first excite sine-wave thermal profiles at distinct frequencies and, finally, recreate the expected full in-orbit temperature profiles. While environmental magnetic stability is strictly maintained, any observed fluctuations in the magnetic field are expected to be solely provoked by temperature variations. The magnetic measurements are correlated with onboard temperature sensors, and the resulting thermal-magnetic effects are fitted to physically based models to quantify the exact amount of thermal noise that can be subtracted.

Furthermore, this TVAC-validated decoupling strategy is proposed as an in-flight operational tool to mitigate thermal noise in magnetically demanding space science missions with long integration times.

Speaker: Pablo Cárdenas

ESTEW 2026 Cardenas Pablo.pdf

ESTEW 2026 TVAC validation [WIP].pdf

11:00 → 11:30 Break

11:30 → 13:00 Heat Transport Technology

Convener: Emily Tipper

11:30 Development, Qualification and Upcoming In-Orbit Demonstration of a Heat Pipe Module for Small Satellite Thermal Control

Furukawa Electric has a long heritage in the development and manufacturing of heat pipe modules for terrestrial thermal management applications. Based on these proven technologies, the company has been developing a heat pipe module optimized for spacecraft thermal control applications.

The developed module integrates heat transport and heat spreading functions to efficiently transfer heat from electronic equipment to radiator surfaces. By optimizing the working fluid and internal structure, stable operation has been achieved over a temperature range from approximately −40°C to +100°C. In addition, the module exhibits anti-gravity performance, enabling flexible thermal design independent of installation orientation. Thermal performance testing of a module incorporating a heat pipe equivalent to 6 mm outer diameter demonstrated a thermal resistance below 0.1°C/W at a heat load of 15 W.

To verify its applicability to space missions, the heat pipe module will be installed on FUNADE, a demonstration satellite independently developed by Furukawa Electric and scheduled for launch in 2027. Prior to launch, a qualification campaign was conducted to evaluate the module under spacecraft environmental conditions. Environmental tests, including vibration and thermal-vacuum tests, were successfully completed, confirming the module's suitability for space applications without performance degradation.

This presentation introduces the development background, design concept, and qualification activities of the heat pipe module for spacecraft use. Thermal performance data obtained after environmental testing will be presented, together with the objectives and evaluation plan of the in-orbit demonstration mission on FUNADE. The results are expected to contribute to the future application of lightweight and efficient heat pipe modules for spacecraft thermal control systems.

Speaker: AKIRA TAKAGI

ESTEW2026_Furukawa_Heat_Pipe_1.1.pptx

12:00 Advanced composite materials using high quality diamond particles to achieve superior thermal conductivity

Conduction-cooled embedded computing applications for space platforms face escalating thermal management challenges as processor power densities continue to increase.

Traditional two-phase thermal systems (heat pipes, loop heat pipes, vapor chambers) exhibit gravity-dependent performance and are vulnerable to malfunction under high G-forces during launch, requiring extensive qualification testing and inspection protocols.

This presentation introduces a gravity-agnostic solid-state thermal solution combining PHENOMENA™ II Copper-Diamond (Cu-Di) composite alloy heat spreaders with PHENOMENA™ I diamond-enhanced thermal interface material (TIM) paste.

The PHENOMENA II Cu-Di composite alloy achieves exceptional thermal conductivity ranging from 700-800 W/m·K through a proprietary sintering process that integrates high-quality mono-crystalline diamonds (Tc > 2000 W/m·K) with pure copper powder (Tc = 394 W/m·K). Unlike two-phase systems, solid conduction materials maintain consistent thermal performance regardless of orientation, acceleration, or microgravity conditions, significantly improving reliability factors for space missions while reducing inspection requirements and qualification complexity.
The material demonstrates thermal conductivity attenuation of less than 5% after 1000 thermal cycles (-55°C to +175°C), making it ideal for long-duration space missions.

The synergistic performance is realized through PHENOMENA™ I TIM paste, which employs ultra-high concentration nano-particle formulation to bind high-quality diamonds in a monolayer configuration between processors and Cu-Di heat spreaders. This creates direct diamond-to-diamond thermal pathways that minimize interface resistance while maintaining electrical isolation. The paste demonstrates vacuum compatibility to 6.0 × 10⁻⁷ kPa with minimal outgassing (TML 0.06% at +200°C), meeting stringent space-flight requirements.

An additional advantage is the integration of HEATBIND solid retainers, which provide structural stability under launch vibration and thermal cycling while enhancing the thermal pathway between modules and their chassis.

Compared to conventional copper or aluminum heat spreaders with standard TIMs, the PHENOMENA™ diamond combination offers up to 50% reduction in junction-to-case thermal resistance while providing mass savings of 15-20% due to the composite’s lower density (5-6 g/cm³) relative to copper (8.96 g/cm³). The low coefficient of thermal expansion (5.0-7.2 × 10⁻⁶/K) closely matches semiconductor packages and aluminum chassis, reducing thermomechanical stress.

This presentation will demonstrate measured thermal performance data from ADHA plug-in modules operating FPGA devices, discuss design considerations for space-qualified applications, and present lessons learned from prototype implementations.

Speaker: Gaby Cristian Mindreci (Performance Interconnect SAS)

ESTEW-2026-Mindreci-PHENOMENA-HEATBIND.pptx

PHENOMENA_example.png

12:30 Analysis and Test of a Heat Pipe System under a Centrifugal Acceleration Field

The Copernicus Imaging Microwave Radiometer (CIMR) is part of ESA’s Copernicus Sentinel Expansion missions and is designed to monitor Earth’s surfaces through radiometric measurements of parameters such as sea surface temperature, sea-ice concentration, and salinity. CIMR operates in a quasi-polar, near-circular, Sun-synchronous orbit, ensuring global coverage including polar regions. The spacecraft includes a rotating microwave radiometer crate and a deployable mesh reflector. The instrument crate thermal control relies on heat pipes (HPs), which spread the heat dissipated by onboard electronic units over a radiating surface. This paper presents the HP architecture developed to withstand the centrifugal forces generated by the instrument rotation. The HPs are mounted along a flat panel at the crate perimeter, parallel to the rotation axis. Differential centrifugal accelerations affect HP performance and, at high force levels, may drive the working fluid towards both ends of the pipe, depleting the central section. The initial design was based on HP performance under a constant acceleration field because data for the actual flight configuration were unavailable. This imposed limitations on pipe length and reduced thermal performance. To overcome these limitations, a dedicated test campaign was conducted in a centrifuge facility to evaluate HP performance under a realistic rotation-induced acceleration field. Different pipe designs were tested under various spin rates and adverse conditions, while residual performance margins were assessed. This paper presents and discusses the test setup and results. Based on the experimental data, system-level analyses were conducted. The thermal conductance of the HPs was degraded with rotational speed, especially for the longest HPs. In spite of that, the results showed better performance than tilted configuration, leading to an optimized heat pipe layout that maximizes overall efficiency while minimizing losses, system complexity, and cost.

Speaker: Paula Prado Montes (ESA)

ESTEW-2026_ROTATING_PIPES.pdf

11:30 → 13:00 Thermal Design

Convener: Paolo Ruzza (ESA)

11:30 Thermal Control System Design and Verification for Comet Interceptor

Comet Interceptor is an ESA-led mission, in collaboration with JAXA, designed to perform the first flyby of a dynamically new comet—a pristine object making its first passage through the inner Solar System. Rather than targeting a known body, the spacecraft, comprising a main module (A) and two probes (B1, B2), will be placed in a halo orbit around the Sun–Earth L2 point and remain there for potentially several years, until a suitable target is discovered and a transfer trajectory designed.
This mission profile presents thermal-control challenges with few precedents in previous scientific missions. During the extended waiting phase, the thermal control system (TCS) must maintain all units within survival limits in a stable, cold environment, relying primarily on passive insulation supplemented by heater lines designed for minimal power consumption while most of the spacecraft remains powered down.
The transfer phase can require solar illumination from any direction within the XZ plane to maintain communication with ground, potentially exposing most of the payload, propulsion system and optical instruments directly to the Sun. A lunar gravity-assist manoeuvre may further introduce a short-duration cold excursion through lunar eclipse, requiring the TCS to accommodate a rapid transient outside nominal environmental conditions.
During the comet flyby, trajectory correction manoeuvres and instrument pointing may result in direct solar illumination of the two main platform radiators. The TCS must therefore accommodate a wide range of external heat loads while remaining within stringent mass and power budgets.
This presentation describes the thermal-control approach adopted for Comet Interceptor and the principal challenges encountered during its design and analysis. Particular attention is given to worst-case definition under uncertain mission conditions, variable spacecraft interface temperatures, passive thermal-control design, heater-power limitations and thermal verification. Key design trades and the resulting approach to achieving robust thermal performance within mass and power constraints are presented.

Speaker: Mr Alessandro Spalla (OHB Italia S.p.A.)

Thermal_Control_System_Design_and_Verification_for_Comet_Interceptor_presentation.pdf

Thermal_Control_System_Design_and_Verification_for_Comet_Interceptor_presentation.pptx

12:00 Envision: mission status at PDR

The presentation will provide an overview of the Envision mission status at PDR that has been recently completed.
The S/C thermal challenges will be presented as well as design options selected to overcome them.
Thermal justification will be also shown, as well as procurement status.

Speakers: Mr Sergio Zanola (TAS), Mr Stephan Ptacek (OHB), Mr emmanuel caplanne (ESA)

12:30 Thermal and thermo-mechanical design of the ESA Genesis mission

The Genesis mission, part of ESA’s FutureNAV programme, aims at significantly improving the accuracy and stability of the International Terrestrial Reference Frame (ITRF), the foundation of satellite navigation and Earth science. This extreme accuracy will be achieved by co-locating the four main geodetic techniques: GNSS, very-long-baseline interferometry (VLBI), satellite laser ranging and DORIS, onboard a single well-calibrated satellite, all synchronized by an ultra-stable oscillator (USO). These ambitious geodetic objectives impose demanding requirements on the spacecraft's thermal and thermo-mechanical design, making thermal control a key driver of the overall system architecture.
To achieve these objectives, a very particular mission and satellite thermo-mechanical configuration had to be developed. Firstly, the Genesis satellite will be placed on a 6000 km MEO orbit, one of the harshest radiation environments in Earth Orbit, significantly driving the satellite design. To mitigate radiation effects, the platform is built around thick aluminium honeycomb panels that provide radiation shielding. Where additional protection is required, dedicated radiation vaults enclose critical electronic units, significantly reducing their ability to dissipate heat. The selected orbit, combined with continuous nadir pointing and the absence of deployable or rotating solar arrays, results in large seasonal variations of solar illumination. Consequently, nearly every spacecraft face is exposed to direct solar flux at some point during the year, creating a particularly challenging thermal environment, critical to maintaining the ultra-stable oscillators and other thermally sensitive equipment within their stability requirements, going as low as 0.1 °C/min
The second major challenge of the Genesis mission is achieving the thermo-mechanical stability required to maintain the relative geometry of the scientific payloads throughout the mission lifetime. This requirement drove the development of an integrated thermal and structural architecture, in which the radiation-shielded electronics are housed within a dedicated platform module, while the antennas, instruments and star trackers are mounted on a payload module constructed entirely from CFRP sandwich panels. This configuration minimizes thermo-elastic effects and enables the dimensional stability required to meet the mission's geodetic performance. The design was supported by an extensive thermo-elastic distortion analysis campaign, comprising of hundreds of thermal load cases representative of all operational and environmental conditions throughout the mission lifetime, with the use of PySINAS as the main mapping tool.
The presentation will describe how stringent geodetic performance requirements drove the satellite’s thermal and thermo-mechanical architecture, with emphasis on the trade-offs between radiation shielding, thermal stability, heat rejection, materials and thermo-elastic performance.

Speaker: Gianfranco Salih (OHB-I)

Thermal and thermo-mechanical design of the ESA Genesis mission.pdf

Thermal and thermo-mechanical design of the ESA Genesis mission.pptx

13:00 → 14:00 Lunch

14:00 → 15:00 Thermal for surface missions

Convener: Matthew Vaughan (ESA)

14:00 Progress on the capabilities of the thermal simulation environment for moving objects on the lunar surface

This contribution presents enhancements and features of the lunar simulation environment for transient thermal analyses of moving ground systems. The basic version of this lunar simulation environment is revisited briefly, whose verification has been presented previously.
Especially for autonomous lunar rovers, an accurate prediction of heat fluxes and temperature evolution along a planned trajectory is essential, as it enables the incorporation of thermal considerations into path planning and operational decision-making.
The proposed framework integrates the local surrounding lunar surface into lunar topography data along a predefined trajectory. A thermal model of the environment is generated, including local surface features such as boulders. Transient thermal analyses are performed, considering the dynamically changing thermal conditions along the traverse. This facilitates a detailed prediction of heat fluxes and temperature distributions for both the moving ground system and the lunar surrounding.
Enhancements with respect to the basic version of the lunar simulation environment are e.g. identification of the local horizon for surface meshing, thermo-optical properties depending on the current region, statistical boulder placement, and the introduction of local surrounding topography.
The presented approach provides a foundation for thermally informed lunar mission planning and advanced thermal assessment of autonomous lunar rovers.

Speaker: Joel Guetzlaff (FH Aachen University of Applied Sciences)

Presentation_ESTEW_26_Guetzlaff.pptx

Presentation_ESTEW_26_Guetzlaff_public.pdf

14:30 Thermal Modelling of lunar surface temperature using Systema (Airbus Defence and Space software).

In the current context, lunar exploration, particularly with NASA's Artemis program, has become a crucial issue for the space industry. Several agencies are investing in systems able to withstand this extremely hostile environment. This presentation focuses on modelling lunar surface temperature using Systema (Airbus Defence and Space software). The approach involves comparing numerical simulations with reference cases from the bibliography, such as the ESA article “ICES-2024-154 - Application Cases of the ESATAN-TMS Python API”.
First, a simplified geometrical model of the lunar surface has been created, with thermo-optical properties of regolith, using a flat meshing and including the Moon's orbitography. These results were compared with measured data (NASA – Apollo 17). Then, a sub-surface non-geometrical thermal model has been introduced, with physical properties of regolith, to evaluate its impact on Moon surface temperature. For this example, sensitivity cases have been performed (landing sites, number of sub–surface nodes etc..) to find the good compromise between calculation time and model precision. Thanks to Systema Pyton API, ADS developed the Python « Lunar Reconnaissance Orbiter Data Importer », to import topography in the thermal model and to create sub-surfaces non-geometrical nodes. Using these tools, a more complex ground thermal model has been created, with geometrical nodes on lunar surface (with reliefs and craters) and a non-geometrical model for the sub-surface. This example allowed to study the impact of shading effects on a simplified geometry on lunar surface. To conclude, the thermal modeling process of lunar surface temperature implemented using Systema has been validated with these reference cases.

Speaker: Simona Calarco (CNES)

CALARCO_S_Thermal_Modelling_Lunar_Surface_Temperature_Systema_V2.pdf

CALARCO_S_Thermal_Modelling_Lunar_Surface_Temperature_Systema_V2.pptx

14:00 → 16:00 Small Satellites & Cubesats

Convener: Nektarios Chari (ESA)

14:00 Parametric Sensitivity Analysis on Thermal Management Solutions for High-Power 3U CubeSats

3U CubeSats typically achieve maximum power levels of around 30 W, having the on-board power generation capacity constrained by the surface area available for mounting solar panels. However, through advancements in the area of high-capacity batteries, some satellite platform developers provide solutions of up to one order of magnitude higher in payload power. But as the external surface area of the satellite remains constant, its thermal regulation capacity may not accommodate the increased heat dissipation required for maintaining it within nominal temperature limits. Our study explores the influence high-power payloads and their duty cycle have on a 3U CubeSat payload’s temperature profile. The results demonstrate that even short durations of payload activity can cause temperature spikes beyond operational values, calling for the implementation of thermal control methods. We then assess the satellite’s thermal response to common passive thermal control strategies by carrying out individual sensitivity analyses on their parameters, namely: (i) heat storages’ mass; (ii) external faces optical surface properties’ values; and (iii) thermal couplings’ values across components. The results demonstrate how the satellite’s temperature outputs are influenced by each distinct change brought to its thermal design, helping identify the most influential parameters, and facilitating a more informed decision-making regarding the parameters’ values in the design process.

Speaker: Ioana Teodora Oprea (Instituto Superior Técnico)

ESTEW26_Parametric Sensitivity Analysis on Thermal Management Solutions for High-Power 3U CubeSats.pdf

15:00 3D-Printed Loop Heat Pipes and Deployable Radiator Systems for CubeSat, SmallSat, and Lunar Habitat Applications

ACT has increased the technology readiness level (TRL) of 3DP LHP systems to TRL 6 and has developed three flight-ready 3DP LHPs to meet a variety of mission needs: a 50 W LHP for a 12U CubeSat, a 250 W LHP for an ESPA-class SmallSat, and a 1kW LHP featuring a passive thermal control valve (TCV) for lunar night survival. Two of the 3DP LHPs (12U CubeSat and 250W ESPA-class SmallSat) are configured with flexible transport lines and deployable radiator panels (DRPs). This presentation shows the design choices, build procedure, testing, and qualification of the three configurations. The covered topics include radiator panel design, hinge line design, DRP deployment modeling and testing, ambient thermal testing, and space-simulating thermal-vacuum testing.

Speaker: Calin Tarau (Advanced Cooling Technologies)

ESTEW_26 3D LHP Presentation_ACT.pdf

15:30 Thermal modelling of ELECTRA: lessons learned for a student team

The thermal design of a CubeSat can present unique challenges that stem from the small size and desired low complexity of the spacecraft. It is often impractical to include proper thermal control components. As such, thermal modelling becomes the main tool for ensuring that the satellite can operate using passive systems. These challenges are more demanding when the project is led by students.

ELECTRA (Electron Layer Exploration using CubeSat for TEC Research and Analysis) is a 3U CubeSat being developed by the CubeSat PoliTO Team for ionospheric measurements and technological demonstration of an innovative electric solid-state thruster. During the last two years, the Team, with the support of ESA’s experts within the Fly Your Satellite! Design Booster 2 programme, designed a mainly passive thermal control system based on surface coatings and a heater for the batteries. To assess the functionality of the system, we modeled the satellite using Thermal Desktop. The modelling showed that the satellite can maintain an acceptable temperature range. However, at this stage of development there are a lot of details that have to be refined. Finally, a TVAC test will be conducted, and the thermal model will be refined.

The work carried out provided valuable experience both in thermal modelling and project management. This paper details the thermal modelling process and the lessons that might be valuable for other university teams.

Speaker: Riccardo Dipietro (Politecnico di Torino)

PresentationPolitoCubesatTeam.pdf

PresentationPolitoCubesatTeam.pptx

15:00 → 16:00 Thermal Testing

Convener: Miguel Copano (ESA)

15:00 Thermal test and correlation of SeRANIS Athene-1 small satellite

Athene-1 is the first small satellite of the University of the Bundeswehr Munich and is part of SeRANIS (Seamless Radio Access Networks for Internet of Space), a national space program funded to develop innovative space technologies and enhance satellite communication capabilities in Germany. The program establishes a new benchmark for national space capabilities by embracing New Space principles, including reduced development costs, agile engineering processes, and accelerated access to orbit. This satellite is the result of a successful co-engineering effort between an institutional customer and OHB LuxSpace, acting as platform owner, satellite integrator, and unit designer.

SeRANIS Athene-1 hosts one of the most diverse and challenging payload suites flown on a small satellite (<300 kg) to date. The spacecraft accommodates 15 experiments and more than 30 individual payload units, encompassing a broad range of technologies, including RF antennas operating across multiple frequency bands (L, Ka, Ku, S, UHF/VHF, and X), infrared and visible imaging systems, propulsion technologies, laser communication systems, and software-based and artificial intelligence experiments. This unprecedented payload diversity introduces significant challenges for spacecraft thermal design, driven by the wide range of operational modes, power dissipation profiles and accommodation requirements.

Athene-1 is embarked on the Triton-X Heavy platform and incorporates two IAUs (Integrated Avionic Unit) acting as onboard computers. Both the platform and avionics architectures were designed and qualified by OHB LuxSpace under ESA funding and will be demonstrated in orbit for the first time. The mission will provide valuable in-orbit validation of these technologies and support the development of future satellite constellations. Electronic units designed by OHB Czech, a fellow company within the OHB group, have also been installed on the satellite to support payload operations.

This presentation shows the thermal design approach adopted for SeRANIS Athene-1, beginning with a high-level overview of the mission architecture and the associated thermal challenges. The thermal vacuum (TVAC) test campaign is then described, including the test philosophy and verification objectives. Finally, the results of the campaign are presented, together with the thermal model correlation process and the lessons learned for future small satellite missions.

Speaker: Mr Miguel Cruzat (OHB LuxSpace)

ESTEW-Presentation - Thermal test and correlation of SeRANIS Athene-1 Small Satellite.pptx

Thermal test and correlation of SeRANIS Athene-1 small satellite.pdf

15:30 FORUM FEI: Thermo-Mechanical architecture and TVAC verification

FORUM (Far-infrared Outgoing Radiation Understanding and Monitoring) is the ninth Earth Explorer mission selected by ESA. This mission aims to measure infrared radiation emitted from Earth across the entire far-infrared part of the electromagnetic spectrum, improving the understanding of the greenhouse effect and enhancing the accuracy of climate change assessments.
The FEI instrument consists of two main assemblies: the Optical Head (OH) and the Front-End Electronics (FEE), designed by Leonardo S.p.A. in Campi Bisenzio, Florence. The OH is composed of the Optical Barrel (OB), which houses the optical components and their associated mechanical interfaces, and the Focal Plane Assembly (FPA), which includes the detector, cold-shield aperture, and the electronics required for detector operation. The FEE provides the acquisition signals required by the FPA and performs the main data acquisition and processing functions.
A key innovative feature of the FEI thermo-mechanical architecture is the implementation of a passive thermal coupling between the OH and the FEE, also electrically connected (through a FLEX), using a high-conductivity pyrolytic-graphite thermal strap. The strap provides a conductive thermal path from the OH to the upper panel of the FEE: therefore, the electronic box radiator is the only cold sink for both FEE and OH. In this configuration, the heat dissipated by the FEE is directly rejected to space through its radiator, while the heat generated within the OH – which is located in a separate closure – is transferred via the thermal strap and rejected employing the same radiating surface. This shared-radiator concept enables the thermal control of two physically separated assemblies through a single passive heat-rejection interface (no need for a dedicated radiator on the OH). This innovative architecture seems unusual but allows for a reduction in terms of thermal HW, making the system more compact, while the heat rejection of the inner items is guaranteed even though there is no space available for the allocation of a second radiator.
The proposed thermal architecture has been experimentally verified during the Thermal Vacuum Test (TVAC) campaign performed on the Proto-Flight Model (PFM) in Leonardo S.p.A. facilities. The thermal strap exhibited the expected behaviour throughout the test, enabling the correct thermal control of both the electronic unit and Optical Head and allowing all planned test phases to be successfully completed. The achieved thermal performances ensure compliance with the applicable test requirements and contributed to the successful qualification of the PFM.
This work demonstrates the successful verification of the thermal architecture and its effectiveness as a passive thermal control solution for the FEI instrument, compatible with the stringent accommodation and thermal stability requirements of optical payloads. The versatility of this thermal architecture and the reduced costs (with respect, for example, to the implementation of a heat pipes) paves the way for the application of this approach to other electro-optical payloads.

Speakers: Giammarco Claroni (Leonardo S.p.A.), Monica Mozzato (ALTEN Italia)

FORUM_FEI_Thermo-mechanical_architecture_and_TVAC_verification.pptx

FORUM_FEI_Thermo-Mechanical_architecture_and_TVAC_verification_draft.pdf

FORUM_FEI_Thermo-Mechanical_architecture_and_TVAC_verification_draft.pptx

16:00 → 16:15 Break

16:15 → 17:45 Thermal Analysis

Convener: Alexandre Darrau (ESA)

16:15 Early Stage Satellite Thermal Analysis: Validating IDM-Thermal with Systema-Thermica

Early stage thermal analyses are essential for mission feasibility studies, yet the commercial packages that dominate the industry (NX Thermal, ANSYS Thermal Desktop, Systema-Thermica, etc.) are costly, feature rich, and often complex to implement for conceptual design work. To bridge this gap, the Concurrent Engineering Centre at CNES has released the Integrated Design Model (IDM), a free, modular, and user friendly framework that lets thermal engineers build, size, and simulate a satellite’s thermal behaviour within minutes. IDM’s core, the IDM Editor, allows users to define spacecraft geometry, mass inertia data, material properties, and mission scenario, including orbit, attitude, power mode, articulation sequence, and ephemeris. Complementary modules (IDM CAD, IDM Scenario, and IDM View) provide rapid CAD integration, scenario management, and orbit visualisation, while a built in thermal add-on creates the Thermal Mathematical Model, solves the discretised heat balance equations, and delivers steady state and transient nodal temperatures together with automatic pre-sizing of radiators and heaters. To verify IDM Thermal’s results fidelity, CNES launched a second generation validation campaign on a complex model under similar constraints to the MicroCarb mission. Using similar geometry, material, and dissipation data, parallel simulations were performed in IDM and in Airbus’s proven tool Systema -Thermica. The comparison focused on incoming fluxes (solar, albedo, and Earth IR) computed with and without the same orbital ephemerides and nodal temperatures in the spacecraft structure . These results demonstrate that IDM can deliver accurate thermal predictions at a fraction of the cost and complexity, making early phase thermal feasibility studies accessible to the space community, especially start-ups and academia, such as university CubeSat teams, and other resource constrained programmes. Its STEP-TAS import export capability also ensures downstream integration with commercial solvers to enable detailed design refinement.

Speaker: Daphne Papadatos

ESTEW 2026 - Daphne Papadatos official pres_final - JLLG.pdf

ESTEW 2026 - Daphne Papadatos official pres_final - JLLG.pptx

16:45 A Simplified Methodology for LEOP Thermal Analyses with Evolving Radiative Enclosures: the HENON Case Study

Abstract
This work focuses on the thermal analysis methodology developed for the Launch and Early Orbit Phase (LEOP) of the HEliospheric pioNeer for sOlar and interplanetary threats defeNce (HENON) cubesat. Although conceived for a specific phase of a mission, the proposed approach is applicable to any condition that faces similar constraints in terms of available information, development schedule and configuration variability.

Launch and Early Orbit Phase (LEOP) thermal analyses pose a series of unique challenges. These are a consequence of the transient nature of the phenomena in question and the necessity to represent multiple mission phases, from launch to spacecraft commissioning.
Specifically, the launch stage is characterised by significant changes in the radiative environment that the spacecraft experiences, including, but not limited to, lift-off, fairing jettisoning, and deployment. This variability calls for the utilisation of distinct radiative cases, wherein the activity of certain surfaces is subject to variation. These analyses are often required when the availability of mission information is limited, and where rapid delivery of results is critical. In such circumstances, modelling approaches must balance realism, time-effort, simplicity, and ease of maintenance while minimising the number of models and analysis cases to be managed.

The proposed strategy, therefore, focuses on the efficient management of evolving radiative surroundings within a single Thermal Mathematical Model (TMM), achieved through the manual activation and deactivation of surface radiative activity as the spacecraft configuration evolves across mission phases. This enables sequential transient analyses to be performed in a controlled, repeatable and traceable manner while reducing the workload associated with modelling.

The presentation discusses the rationale behind the adopted modelling assumptions, the implementation within the ESATAN-TMS Workbench environment, and the workflow developed to manage changing spacecraft configurations and associated radiative cases for the HENON case study, where the methodology has been successfully developed and applied. The resulting methodology provides a practical and robust framework for performing preliminary transient thermal analyses when simplicity, reliability and rapid development are primary design drivers. Future work will explore automating the case-switching process to further reduce turnaround time and the risk of human error.

Speaker: Marco Guzzon

Abstract Thermal Workshop 2026 - LEOP Thermal Analysis Approach - Guzzon Marco.docx

ESTEC Thermal Workshop 2026 - Guzzon Marco.pptx

17:15 A Custom Two-Phase Nitrogen Property Library and Fluid-Solver-Free Reduced-Order Model for LN2 TVAC Cooling Prediction in ESATAN-TMS

Thermal Vacuum Chamber (TVAC) qualification of large-scale hardware such as solar arrays increasingly relies on two-phase liquid nitrogen (LN2) cooling systems to achieve cryogenic transients from ambient to below -180°C. High-fidelity ESATAN-TMS thermo-hydraulic simulation of these systems is computationally expensive, with a single two-phase fluid-network transient exceeding 13 hours of runtime, limiting iterative test-plan development and control-strategy evaluation.

As ESATAN-TMS's default fluid library does not include nitrogen for two-phase modelling, a custom nitrogen thermophysical property library was developed from NIST real-gas reference data (4-0.1 bar, -190°C to +180°C), using a positive-definite enthalpy reference to eliminate sign-convention errors in energy-balance calculations. The library was validated through staged checks: known-power-input response, correct isothermal saturation behaviour across the two-phase regime, and round-trip consistency against the source pressure-enthalpy data.

Building on this library, a reduced-order model (ROM) was developed that eliminates the fluid solver entirely while preserving full closed-loop control for shroud heating and LN2 mass flow regulation. The fluid domain is represented as an imaginary node co-located within the existing thermal network, updated via energy-balance/enthalpy tracking against the custom library rather than a solved momentum/continuity field. A film-temperature-dependent heat transfer coefficient is applied within the node power calculation, mapped via a custom subroutine built from three correlations derived from initial two-phase simulation results across the loop's liquid, two-phase, and vapour regimes. The power calculation itself takes the minimum of a convective estimate ($Q = hA\Delta T$) and a mass-flow enthalpy estimate ($Q = \dot{m}\Delta H$), preserving the dominant physics without re-solving the flow field. Valve-induced pressure drop is represented as a calibrated saturation-temperature offset (-185°C at full line pressure, 3.15 bar, dropping to -190°C under throttled flow control), preserving the dominant thermodynamic effect of valve operation without solving for the pressure field directly. Mass flow boundary conditions were derived from full two-phase simulation results (lower bound) and empirical LN2 tank depletion during chamber operation (upper bound). The solar array payload was represented using effective front/rear thermal properties from manufacturer-supplied material test data.

The cooling-phase ROM described in this abstract was correlated and integrated with automated heating-phase control to form the predictive model. This predictive model was independently validated against a test not used in model development. The predictive model correctly predicted an anomalous heating behaviour during hot-dwell — one panel sensor failed to reach its target set-point — matching the observed deviation to within 0.262°C. A 40-45-minute cold-dwell timing discrepancy was attributed to overcooling-protection heater sensitivity combined with the simplified payload representation. The approach reduces simulation runtime from over 13 hours to under 1 hour — a >13x speedup — while retaining predictive accuracy sufficient for test-plan optimization and control-strategy trade studies ahead of future payload campaigns.

Keywords: ESATAN-TMS, Reduced-Order Modelling, Two-Phase Cryogenic Simulation, TVAC, Nitrogen Property Library, Heat Transfer Coefficient Mapping, PID Control

Speakers: Mr Thilan Chandika Ranasinghe Ranasinghe Arachchillage (Technical University of Munich), Mr Thilan Ranasinghe (Technical University of Munchen)

Custom N2 Library and Fluid-Solver-Free Reduced-Order Model for LN2 TVAC Cooling Prediction ESATAN (1).pptx

Custom N2 Library and Fluid-Solver-Free Reduced-Order Model for LN2 TVAC Cooling Prediction ESATAN.pdf

Supporting Diagrams.pdf

16:15 → 17:45 Thermal Control

Convener: Miguel Copano (ESA)

16:15 Graphene electrochromic variable emissivity radiator

Variable emissivity radiators (VERs) can offer significant advantages in satellite thermal control systems by reducing power consumption. SmartIR has developed a graphene-based VER that uses electrochromism to modulate infrared emissivity by 0.3 while maintaining high rejection in the solar spectrum. The SmartIR VER has been incorporated into a subsystem integrated on one of ATMOS Space Cargo's satellites, which is due to be launched into low Earth orbit and recovered after re-entry. Application of low voltage to the VER enables the transition from a high infrared emissivity state of 0.8 to the low emissivity state of 0.5, where the former is the property it defaults to when unpowered. This low voltage requirement and minor current draw enable the subsystem to be very low power, enabling significant power budget savings in the thermal control system as a whole by reducing the reliance on heating during cold phases of the orbit.

Speakers: Dr Ciaran Mullan (SmartIR), NAGAPP PRADHANI (Atmos Space Cargo gmbh)

SmartIR-Atmos ESTEW 2026

16:45 Opportunities for Solid-State Elastocaloric Technology in Active Thermal Management for Space Applications

Thermal management is essential for maintaining spacecraft systems and payloads within their required operating temperature ranges, supporting reliable operation and mission performance. Elastocalorics is an emerging technology for active thermal management based on solid-state refrigerants, such as shape-memory alloys, which undergo reversible, stress-induced phase transformations. Combined with solid–solid contact heat transfer, it enables heat pumping without a fluid heat-transfer circuit within the active core. Active regeneration or cascaded configurations can extend device-level temperature spans beyond the adiabatic temperature change of the elastocaloric material.
This contribution introduces the operating principle of elastocaloric technology and presents findings from the ESA-supported SPACECOOL study. The study explores space applications in which elastocaloric technology could offer benefits relative to established thermal-control solutions, reviews candidate materials and their operating temperature ranges, and uses preliminary numerical simulations to assess achievable temperature spans and coefficient of performance.
A technology de-risking roadmap is also presented, addressing key challenges in material durability, thermal interfaces and actuation, and outlining priorities for experimental validation. These activities aim to establish a basis for developing a solid-state elastocaloric breadboard demonstrator and assessing its suitability for selected space applications.

Speaker: Andrej Žerovnik

Opportunities for Solid-State Elastocaloric Technology in Active Thermal Management for Space Applications - final2.pptx

17:45 → 18:00 Opening: Closure

ESA_Tech_Thermal_Overview_ESTEW_2026_REV1.pptx