1. Introduction and state-of-the-art

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1. Introduction and state-of-the-art

In this introductory chapter a brief explanation of the motivations of the work and the state-of-the-art in this field are shown, together with an insert about IDS Company, which has hosted the project. In the end, the global structure of the thesis is presented.

1.1 The importance of knowing temperature in space

Mathematically speaking, after having introduced an inertial reference system (e.g.

a Cartesian one) to describe the spatial domain of a generic system that can evolve in time, its temperature is a scalar field defined as:

𝑇 = 𝑇(𝒙, 𝑡)

For spacecraft, the most important scope of obtaining a temperature field as precise as possible is to ensure:

 Thermal control: any component of the system must be kept in a certain range of T that guarantees its functionality throughout the mission;

 Thermal stability: any thermal perturbation must not produce a variation of T in time that overcomes a fixed value;

 Thermal uniformity: the gradient of T at a given time must be less than a defined (small) value, [1];

A detailed and accurate thermal analysis is critical to the success of any spacecraft mission and is thus an integral part of the development cycle. Without an adequate

1 On one hand, Kelvin scale is suitable for qualitative analyses, where ΔT can be better appreciated if a scale with all positive values is employed; on the other hand, Celsius scale is preferable for grasping more intuitively quantitative data.

14 Thermal Control Subsystem (TCS), the extreme temperature ranges and gradients encountered by spacecraft in flight may cause temperatures of components to exceed operating or survivability limits, leading to decreased performance or even permanent damage, [2].

Table 1.1 - Operating temperature requirements for common spacecraft hardware

According to NASA, the main tasks of thermal analysis during the whole design of a spacecraft can be summarized as follows, [3]:

 provide an optimum thermal design within the constraints of the overall system design;

 provide temperature distributions and temperature histories to the level of detail required;

 provide early identification of design problems;

 provide the basis for predicting and evaluating thermal performance in test and flight.

The result of this iterative procedure is the final TCS, which can vary, for example, from simply choosing the correct coatings of each component up to designing a rather complex subsystem based on louvers and heat pipes. Therefore, thermal control can be achieved in an active or passive way, according to the need to supply

15 or not power to the TCS components. The table below highlights the main elements employed in designing a TCS.

Table 1.2 - Common TCS components, [4]

An important concept linked to thermal analysis is the building of a “thermal model”;

intuitively, the basis would be the geometric model of the spacecraft, but it would be quite proud to think of analyzing in detail each single element without being mistaken. Thermal modelling is a continuous challenge between making the geometrical model as simple as possible and maintaining acceptable levels of solution accuracy. The ability of the thermal analyst plays an important role in selecting the best thermal model capable to produce the best solution compatible with the objectives of the analysis. For large spacecraft, this can be a problem and what is usually done is to divide the geometric model at subsystem level or component level that are then analyzed in greater detail by means of expensive and time-consuming routines. An alternative is to reduce the detailed models of components, which are eventually combined at system level. These reduced models have to appropriately approximate the detailed models, in order to analyze the critical points with an acceptable degree of accuracy on a system level and solution runtime.

Usually, a maximum temperature difference in the solutions of detailed and reduced model should be below 3 – 5 K. Model reduction is currently mostly done manually, which afford considerable effort, time and cost. Thus, there is potential for automation to increase time efficiency with the help of model reduction tools, [5].

16 Figure 1.1 - Spacecraft model reduced at subsystem/component level

Being this work conceived with the development of a thermal code for general analysis of spacecraft, and taking into consideration the didactic scope of the topic, already simplified geometric models have been used, based on simple shapes like parallelepipeds, cubes and spheres. As a first approximation, these shapes can represent most of the common components in a satellite, [6]. Each component can be individually analyzed in detail and then return the solutions of interest that will set the constraints for next analyses at system level.

Originally, most analysts developed thermal models by hand with the time- consuming use of punch-cards. In the 1980s, with the development of minicomputers and workstations, the time required to build a thermal model could be greatly reduced through the interactive use of software codes that aid the analyst in model construction, debugging, and execution. In the 1990s, several integrated thermal- analysis programs were developed: they allowed the analyst to generate both thermal and geometrical models, execute them and display the results in a user- friendly environment on a workstation or PC, [7]. The path for today’s current

17 software packages was set and the most commercial of them that also include a shape-based model builder are:

 Thermal Synthesizer System (TSS) by SPACEDESIGN under license to NASA/JSC,

 THERMICA by Network Analysis Inc. under license of ASTRIUM,

 ITAS by Analytix Corporation, [7].

Sometimes there is interest in the knowledge of temperature only in a subdomain of V, like its boundary S, the external surface that confines V. Although the mechanisms of heat transfer do not change, the process of analysis and the instruments employed must be adapted according to the output. This concept has been crucial throughout the work and has played a key-role in the development of the thermal code, as underlined in the next section.

1.2 Requirements and constraints of the project

IDS S.p.A. Company has a Space Lab that has been working closely with other IDS divisions, external customers and the European Space Agency (ESA) for over 10 years, taking part in successful space programs such as ARTEMIS, COSMO- SKYMED, GALILEO and LISA PATHFINDER. The definitive consolidation in the space field has been obtained with the release of Antenna Design Framework (ADF), an electromagnetic simulation framework suitable to model both the interactions between antennas and the plasma emitted by ion thrusters and the RF links between a re-entry vehicle and ground or satellite relay stations.

Besides, IDS has developed products in both naval and aeronautical fields. One of these is VIRtual Aircraft Framework (VIRAF), a suite of tools for prediction of Radar and InfraRed (IR) signatures of all types of aerial platforms such as civil and military aircraft, unmanned vehicles and missiles, [8].

18 Figure 1.2 - Mesh model detail, [8]

The picture above shows a detail of one of the tools IDS has made available for this work: the Mesher. As explained in detail in Chapter 5, the Mesher takes as input a geometric model and allows discretizing it by means of a Finite-Element Model (FEM) mesh. The result is what is called in literature a Geometric Mathematical Model (GMM), [5-7] and, consequently, the first requirement of the work is that the thermal code be able to handle such GMMs as input.

Secondly, it is common knowledge that vacuum in space environment makes radiation the most important mechanism for heat transfer. IDS frameworks have their own tools to compute radiative exchanges among surfaces and the View-Factor Solver (VFS) is an executable file that provides the computation of the view factors, even in presence of obstructions, based on Monte Carlo ray-tracing method.

Choosing VFS as a source for view-factor determination has made possible to simplify the whole work and save computational time.

The “drivers” above have their own drawbacks, which have represented constraints in the project:

 using a more intuitive volumetric mesh as a starting point for the construction of the Thermal Mathematical Model (TMM), [5-7], has not been possible since the Mesher accepts only geometric models based on surface skins;

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 the geometric modelling with skins and the results provided by VIRAF suggest that a prediction of T should be as accurate as possible along the external surfaces of the spacecraft, while some lack of information can be tolerated inside;

 VFS handles 2D triangular mesh elements only; therefore, its use has influenced the definition of both materials and conductive/radiative couplings, as discussed in Chapters 5-6.

The prediction of IR signature of a generic system, obtained by means of radiation analysis codes, will have an accuracy proportional to the one of the temperature field of the system, which acts as an input for these kinds of codes. Should something similar to what VIRAF performs for atmospheric vehicles be required for space applications, then one can understand the need of a specific thermal-analysis code that considers the space environment, so different from the atmospheric one.

Chapter 3 discusses the typical heat loads a spacecraft must withstand during its mission.

1.3 Overview on spacecraft architectures

Before starting with technical details, a brief period has been devoted to collecting general information about what kinds of space activities are currently “in fashion”, the most interesting projects in the forthcoming future and their relation to this work.

The importance of TCS has already been highlighted; but one can identify a class of space missions or activities (and relative spacecraft architectures) whose success is in a certain sense “proportional” to the knowledge of T:

1. Atmospheric re-entry: with the end of the Space Shuttle Program, an architecture that seems to be coming back in fashion is the capsule, which has been proved to be cheaper and more reliable thanks to its simplicity if compared with Space Shuttle. When re-entry begins, the complex aerothermodynamics and the heaviest heat loads must be overcome in order for the mission to succeed. In doing so, the most precise temperature field of the whole capsule

20 is to be known at the end of the orbital phase, so that an appropriate Thermal Protection System (TPS) be designed.

Any re-entry vehicle has its own TPS for re-entry phase that does not interact with TCS. As an example, SpaceX’s Dragon V2 Capsule has a TPS composed of a heat shield of ablative phenolic impregnated carbon, capable to sustain up to 1600°C.

Figure 1.3 - SpaceX's Dragon V2 Capsule

A more and more sensitive topic atmospheric re-entry is involved in (but with a reversed goal) is the disposal of space structures after the end of their mission (the so-called Phase F, [9]). With the exponential increase of orbital debris, it is necessary that any spacecraft be able to self-destroy in 25 years from its End- Of-Life (EOL) time. A controlled burning in atmosphere seems to be the best solution, at least for small structures, to succeed.

2. Interplanetary exploration: the capability of interplanetary probes to maintain communication with Earth is central for a successful mission. For this purpose, antennas must be correctly pointing in order to guarantee data transmission.

This is directly related to the thermal status of the telecommunication subsystem, since thermal distortions on one hand and excessive white noise on the other hand restrict the operating range of T, [10]. Voyager 1 probe, which

21 was launched on 5 September 1977 and is currently flying in the interstellar space, represents an excellence in this kind of missions. Its ability to send even now signals to Earth astonishes more and more the scientific community.

Figure 1.4 - Voyager 1 and its Golden Record

3. National security: saying that the “Space Race” was an aspect of the Cold War should not upset. Both USA and USSR had to protect national borders and the development of space technology offered new solutions in this field. The concept of space-based surveillance was born at that time and the first documented spy-satellite, officially called “reconnaissance satellite”, is the Corona satellite in 1956 as a part of the Weapon System 117L program. It was composed of a camera with a film to record images, which would be returned to Earth in a re-entry capsule rather than sending them back electronically, [11].

Currently, the Space-Based InfraRed System (SBIRS) mission is considered one of the USA’s highest priority space programs to provide global IR surveillance capabilities to meet increasing demands in national security mission areas like missile warning and defense and battlespace awareness.

The SBIRS architecture consists of hosted sensor payloads in Highly Elliptical Orbit (HEO), dedicated Geosynchronous Earth Orbiting (GEO) satellites, and the associated ground infrastructure to receive, process, and deliver IR information to key decision makers, [12].

22 Figure 1.5 - Main components of SBIRS GEO satellite

To successfully complete their missions, this class of satellites rely on staying undetected for as long as possible and to detect their targets before they themselves are detected. If detected, it is important to increase the difficulty to be identified and this can be achieved by reducing satellite’s IR signature. A satellite or, more generally, a vehicle designed to satisfy these criteria is called

“stealth”: VIRAF pursues this scope in the aeronautical field.

It is out of the scopes of this work to treat at the same level of depth all the aspects related to a thermal code, but at least an attempt has been done and is discussed in Chapter 3, where a comparison between different models of eclipse and albedo is shown.

1.4 Structure and objectives of the project

The figure below shows the rationale followed in this work. A reference architecture that allows describing all the choices and the features of the thermal code has been iteratively updated, in accordance with the improvements supplied during the project.

As outputs, the temperature Ti at any point i of interest and, if required, the global heat path inside the system have been obtained.

23 Figure 1.6 - Work flow-chart

Large attention has been devoted in setting the appropriate boundary conditions for the thermal problem by evaluating the external heat loads for closed Keplerian orbits and for some pointing laws. Then, the GMM is extracted from the Mesher after uploading a skin modeled with CATIA, while a simplified thermo-optical model of materials has been used with the main physical quantities the TMM needs in order to enter the Heat Transfer Code. The chart above is inspired to the one in [3], Fig. 1, but a substantial difference exists: the former aims at developing an algorithm for thermal analysis of a spacecraft architecture as generic as possible; the latter is inserted in the bigger context of thermal design of spacecraft hardware, for which the geometric model is given as input. In this second case, the effort is concentrated on meeting all temperature limits for the “worst-case combinations” (hot and cold) of possible orbital environments, spacecraft or instrument operating modes and tolerances on all major thermal properties. The activity of thermal design could be one possible evolution of the present project.