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The purpose of this paper is to describe the design and construction of a custom‐built low‐cost thermal vacuum chamber (TVC) for spacecraft environmental testing and verification. The paper provides detailed analysis and an insight into the design and development of the chamber. The chamber was specifically constructed for carrying out the thermal and vacuum environmental tests in a 16″ dia × 16″ long horizontal thermal vacuum chamber. The chamber is constructed using a combination of mechanical (roughing) pump and turbo‐molecular pump, used to pump the chamber down to 10⁻⁵ Torr and a combination of radiation heaters and nitrogen gas is used to vary the temperature within the chamber from +80 to −50°C.

The TVC equipment is built as part of the picosatellite and nanosatellite program at Space Systems Research Laboratory of Saint Louis University. The equipment is built at a low cost and is suited for testing an entire picosatellite and several components and subsystems of nanosatellite simulating thermal and vacuum conditions similar to space environment. The different main parts of the equipment are described in a way which explains the choice of construction and partly makes it possible to replicate similar equipment.

The TVC equipment is successfully used to simulate the thermal and vacuum conditions of space similar to the conditions experienced by a picosatellite or nanosatellite in low earth orbit.

The design and construction of TVC in this paper have broader implications and can be a platform for future research on low‐cost TVC. This equipment can be utilized in the research areas of electronics and communications, biology and medicine to name a few. Thermal and vacuum experiments on several astro‐biological experiments can be performed.

The paper is intended to be a source of inspiration for industrial or academic space research laboratories which would like to design and construct a similar test‐equipment, instead of investing expensive commercially available alternatives.

The paper discusses in detail, the simplified cost‐effective approach of constructing TVC and also outlines the various issues to be considered. The TVC equipment is custom‐built and is described in an easily understandable way, which makes this a helpful paper for those who wish to produce similar equipment. This will be the only known manuscript in the literature to detail the design and construction of low‐cost, economical TVC.

The purpose of this paper is to present a new approach in the concept and implementation of autonomous micro‐spacecraft. The one true “artificial agent” approach to autonomy requires the micro‐spacecraft to interact in a direct manner with the environment through the use of sensors and actuators. As such, little computational effort is required to implement such an approach, which is clearly of great benefit for limited micro‐satellites. Rather than using complex world models, which have to be updated, the agent is allowed to exploit the dynamics of its environment for cues as to appropriate actions to achieve mission goals. The particular artificial agent implementation used here has been borrowed from studies of biological systems, where it has been used successfully to provide models of motivation and opportunistic behaviour. The so‐called “cue‐deficit” action selection algorithm considers the micro‐spacecraft to be a non‐linear dynamical system with a number of observable states. Using optimal control theory rules are derived which determine which of a finite repertoire of behaviours the satellite should select and perform. The principal benefits of this approach is that the micro‐spacecraft is endowed with self‐sufficiency, defined here to be the ability to achieve mission goals, while never placing itself in an irrecoverable position.

The filamentary growths of single crystals on material surfaces are termed whiskers. They are seen to nucleate and grow on certain electronic materials either from vapour and liquid phases or by a process induced by residual stresses in electroplated surfaces. Whisker growth does not depend on the existence of an electric field and surfaces prone to their growth may nucleate and form whiskers as a result of exposure to a space environment. This paper includes a detailed examination of tin whiskers which were found to have 1 to 4 micron diameters and lengths exceeding 2 mm. Some were found to carry currents between 22 and 32 mA before burning out. Conductive whiskers can cause extensive short circuit damage to spacecraft electronics particularly as miniature devices progressively employ closer spacings between conductors. Several modes of whisker growth on spacecraft electronic materials (molybdenum, tungsten, Kovar, tin) have been observed and are described. Tin, cadmium and zinc surfaces can support stress‐induced whisker growth and it is recommended that these metal finishes are excluded from spacecraft design and possibly replaced by a tin‐lead alloy.

Alternative energy sources and power generation techniques for long‐term space missions are gaining importance in recent years for future bases and colonies on the Moon or Mars. Current technologies used for manned or unmanned missions to the Moon or Mars use either solar panels (bulky, expensive/kilogram to space, and inefficient) or nuclear energy (extremely dangerous and unpopular). Enzyme based bio fuel cells can be used as alternative energy sources, but its survival depends on maintaining appropriate temperature and pressure in space. The purpose of this paper is to detail the concept design and development of a payload tank to house bio fuel cells for operations in space environment.

Details about the development of the design methodology for such housing are discussed. A full‐scale payload tank is designed to house a small biological fuel cell using space grade materials. Requirements analysis, design, validation, and manufacturing process are covered.

The outcome is a dimensionally optimized housing structure for housing biological fuel cells and maintaining the temperature and pressure for survival of the fuel cell.

The designed payload housing satisfies all the constraints and requirements. Furthermore, its advantages include scalability and modularity by virtue of using optimized design approach. The final product provides a planned procedure for designing larger housing for other missions.

The purpose of this study is to optimize multilayer vacuum thermal insulation (MLI) of modern high-weight spacecrafts. An adequate mathematical simulation of heat transfer in the MLI is impossible if there is no available information on the main insulation properties.

The results of experiments in thermo-vacuum facilities are used to re-estimate some radiative properties of metallic foil/metalized polymer foil and spacer on the basis of the inverse problem solution. The experiments were carried out for the sample of real MLI used for the BP-Colombo satellite (ESA). The recently developed theoretical model based on neglecting possible near-field effects in radiative heat transfer between closely spaced aluminum foils was used in theoretical predictions of heat transfer through the MLI.

A comparison of the computational results and the experimental data confirms that there are no significant near-field effects between the neighboring MLI layers. It means that there is no considerable contradiction between the far-field model of radiative transfer in MLI and the experimental estimates.

An identification procedure for mathematical model of the multilayer thermal insulation showed that a modified theoretical model developed recently can be used to estimate thermal properties of the insulation at conditions of space vacuum.

This paper seeks to provide an insight into the design and development of the thermal test model (TTM) of X‐Sat, a 120 kg class micro‐satellite, being developed at the Centre. This model was specifically constructed for carrying out a thermal balance test (TBT) in a 4 m diameter vertical thermal vacuum chamber.

The construction of the thermal model followed a structural mock‐up model which was modified thermally to suit the purpose. Specific and careful consideration was given to the geometry and, more importantly, thermal characteristics such as thermal mass, surface properties, etc. to mimic the actual satellite configuration as closely as possible. Test plans were devised to qualify the fabricated components to meet the out‐gassing and other thermal requirements for the model. Design and qualification of supporting frame and linkages for TBT are also covered.

It is possible to simulate the thermal characteristics of a micro‐satellite in orbit under a different mission scenario through proper scaling and using alternative material options while developing TTM.

The paper discusses in detail the simplified cost‐effective approach of constructing TTM and also outlines the various issues to be considered for a TBT. It provides valuable information needed for micro‐satellite designers.

The purpose of this paper is to describe experimental and numerical investigations focussed on the shock wave modification induced by a dc glow discharge. The model is a flat plate in a rarefied Mach 2 air flow, equipped with a plasma actuator composed of two electrodes. The natural flow without actuation exhibits a shock wave with a hyperbolic shape. When the discharge is on, the shock wave shape remains hyperbolic but the shock wave is pushed forward, leading to an increase in the shock wave angle. In order to discriminate thermal effects from purely plasma ones, the plasma actuator is then replaced by an heating element.

The experimental study is carried out with the super/hypersonic wind tunnel MARHy located at the ICARE Laboratory in Orléans. The experimental configuration with the heating element is simulated with a code using the 2D full compressible Navier-Stokes equations adapted for the rarefied conditions.

For heating element temperatures equal to the flat plate wall surface ones with the discharge on, experimental and numerical investigations showed that the shock wave angle was lower with the heating element, only 50 percent of the values got with the plasma actuator, meaning that purely plasma effects must also be considered to fully explain the flow modifications observed. The results obtained with the numerical simulations are then used to calculate the aerodynamic forces, i.e. the drag and the lift. These numerical results are then extrapolated to the plasma actuator case and it was found that the drag coefficient rises up to 13 percent when the plasma actuator is used, compared to only 5 percent with the heating element.

This paper matters in the topic of atmospheric entries where flow control, heat management and aerodynamic forces are of huge importance.

The purpose of this study is to explore the applications of 3D printing in space sectors. The authors have highlighted the potential research gap that can be explored in the current field of study. Three-dimensional (3D) printing is an additive manufacturing technique that uses metallic powder, ceramic or polymers to build simple/complex parts. The parts produced possess good strength, low weight and excellent mechanical properties and are cost-effective. Therefore, efforts have been made to make the adoption of 3D printing successful in space so that complex parts can be manufactured in space. This saves a considerable amount of both time and carrying cost. Thereof the challenges and opportunities that the space sector holds for additive manufacturing is worth reviewing to provide a better insight into further developments and prospects for this technology.

The potentiality of 3D printing for the manufacturing of various components under space conditions has been explained. Here, the authors have reviewed the details of manufactured parts used for zero-gravity missions, subjected to onboard international space station conditions and with those manufactured on earth. Followed by the major opportunities in 3D printing in space which include component repair, material characterization, process improvement and process development along with the new designs. The challenges like space conditions, availability of power in space, the infrastructure requirements and the quality control or testing of the items that are being built in space are explained along with their possible mitigation strategies.

These components are well comparable with those prepared on earth which enables a massive cost saving. Other than the onboard manufacturing process, numerous other components as well as a complete robot/satellite for outer space applications were manufactured by additive manufacturing. Moreover, these components can be recycled onboard to produce feedstock for the next materials. The parts produced in space are bought back and compared with those built on earth. There is a difference in their nature, i.e. the flight specimen showed a brittle nature, and the ground specimen showed a denser nature.

This review discusses the advancements of 3D printing in space and provides numerous examples of the applications of 3D printing in space and space applications. This paper is solely dedicated to 3D printing in space. It provides a breakthrough in the literature as a limited amount of literature is available on this topic. This paper aims at highlighting all the challenges that additive manufacturing faces in the space sector and also the future opportunities that await development.