On possibilities for development of the common-sense concept of habitats beyond the Earth

Acta Astronautica 170 (2020) 487–498

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Research paper

On possibilities for development of the common-sense concept of habitats beyond the Earth

T

A.V. Degtyareva, L.M. Lobanovb, A.P. Kushnar’ova, Ie.Yu. Baranova, V.S. Volkovb,∗, A.O. Perepichayb, V.V. Korotenkob, O.A. Volkovac, G.G. Osinovyya, Yu.A. Lysenkoa, M.D. Kaliapina a

Yuzhnoye State Design Office, 3 Kryvorizka str., Dnipro, Ukraine, 49008a E.O. Paton Electric Welding Institute of the NASU, 11, Kazimir Malevich str., 03680, Kyiv, Ukraineb c National Technical University of Ukraine “Igor Sikorsky Kyiv Polytechnic Institute”, 37, Prosp. Peremohy, 03056, Kyiv, Ukrainec b

A R T I C LE I N FO

A B S T R A C T

Keywords: Space habitats Lunar base Shielding technologies Deployable structures Load-carrying shells

The article reviews the contemporary concepts of space habitats, focusing on the habitats’ mass, dimensions, and resistance to the space environment. The authors discuss a concept of constructing a multifunctional settlement on the Moon's surface that suggests sequential solutions for main problems of assuring habitability and operational versatility. Different aspects of the efficient algorithm for lunar colony development were analyzed, including the efficiency of payload delivery and a possibility for assuring proper protection of payload. The article describes the advantages of using the modular design and standard components that are delivered ready-to-use and enable the shortest time needed to provide primary habitability of the on-planet colony. It was suggested to use load-carrying deployable structures as standard pressure shells of the generic modules because of the ability of these structures to acquire sufficient protection properties and three-dimensional stiffness once deployed. The article presents a scenario that gives an understanding of the way of using the proposed space transportation systems with a step-by-step expansion of the colony's functionality. In conclusion, the authors assume that the compact folding of standard habitat modules’ pressure shells during delivery can result in fewer transportation missions.

1. Introduction For decades, the largest state space agencies and corporations have been studying the possibility of creating permanent habitats for deep space exploration. Building the temporary or permanent settlements for a long-term stay on Mars or the Moon is part of the paradigm of further space exploration. The closest celestial body to the Earth, the Moon is considered a part of the Earth's space infrastructure and naturally is number one among other celestial bodies in the plans of satisfying humanity's growing demands for resources and expansion of boundaries for space exploration. The active international project Lunar Orbital Platform-Gateway is also planned to serve as a platform for manned lunar missions that are expected to have been fulfilled by the

late 2020s. The goal of operating a permanent settlement for using Moon's resources, conducting research activities, and trying out technologies for further human expansion in space seems more feasible with such an intermediate platform in lunar orbit. Constructing a lunar settlement will make it possible to perform medical and biological research in the near future, try out systems of human life support beyond the Earth, and run unique physical experiments that are not possible in the Earth's conditions. The design of lunar modules can be eventually adapted for the conditions of Mars, for which some of the habitability support problems are not critical; for instance, the presence of atmosphere eliminates the problem of micrometeoroid protection. After the first manned mission to the Moon roughly half a century ago, there have been no other similar projects, first of all, because of the

∗

Corresponding author. E-mail address: [email protected] (V.S. Volkov). a [email protected] b offi[email protected] c fl[email protected] https://doi.org/10.1016/j.actaastro.2020.02.014 Received 16 December 2019; Received in revised form 23 January 2020; Accepted 6 February 2020 Available online 11 February 2020 0094-5765/ © 2020 IAA. Published by Elsevier Ltd. All rights reserved.

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financial outlay. In the last decade, the concepts of lunar colonies have been focused on some major problems, with prime importance on the effectiveness of transportation vehicles and payload delivery, followed by a sensible step-by-step algorithm of settlement construction and the effectiveness of settlement habitability support. In this article, the authors analyze the ways of solving these problems and suggest their concept of a multifunctional, self-sufficient research and production colony on the lunar surface.

contained in regolith, mix the product with a molten glass matrix, and use it as a material for habitat walls. Paper [20] suggests using flexible fiberglass and dehydrated regolith concrete produced from sulfur for lunar habitat structures. In this case, a cylindrical concrete structure is reinforced by round tension cables which compensate the tension force induced by inner pressure. The mentioned concepts, though, imply the necessity of delivering additional equipment and trying out ISRU technologies. The authors of [21] draw a conclusion that it is necessary to study additive processes carefully, with a proper account for the properties of regolith at a suggested area of moon landing; furthermore, an adequate and effective process of solidification shall be tried out for the case of in-situ production of construction materials. The same conclusion can be drawn for some other promising materials such as magnesium, the main properties of which are summarized in Ref. [22] in the context of using it as a material for in-situ construction of habitats on the Moon. In the analysis of a structure made of magnesium, applied on sintered regolith, and shielded additionally with regolith sand, paper [23] arrives at a conclusion that shielding with regolith 3 m in thickness can provide proper insulation of a structure regardless of its location on the Moon. The seismic structural analysis of lunar base, see paper [24], proves sintered regolith to be a preferable material for the base's foundation; however, it is noted that the technologies of regolith sintering are not suitable for the suggested technique of construction so far. Obviously, the drawn conclusion about damping properties of regolith in conditions of seismic activity requires an additional analysis as well. Temperature deformations in lunar base structures’ walls are also a question of great importance for the case of using the materials produced in situ. The Moon surface temperature can vary between −171 о С and 111 оС [25] while the design temperature inside a protected module can be near 20 оС. The maximum values of stresses σMAX, caused by temperature deformations in a structure under the specified conditions, can be evaluated by using the following equation [26]:

2. Design concept The concepts of lunar colonies have varied for the last decade due to the emergence of new materials and technologies, but the general design principles have remained unchanged. A higher level of standardization of components is obviously the top priority, as well as a step-bystep extension of lunar base's capabilities during construction. The rich experience gained during the development of the ISS has been applied as a basis for the European Baseline Roadmap, which incorporated feasible conventional mission scenarios, including Moon Next among others [1]. According to the concept, it was planned to lift two lunar modules to a low earth orbit by a 130-ton space launch system and then deliver them to a lunar transfer orbit by using a cryogenic propulsion system. The Global Exploration Roadmap [2] gives a clearer understanding of the future habitats' design, and its Moon Village concept implies gradual progress in multifunctional capabilities. Paper [4] provides the analysis of the ways suggested by ESA Director General Jan Wӧrner for the implementation of the Moon Village concept [3]. The paper explains that design constraints for lunar modules arise from the available transportation vehicles such as Orion, Dragon, and CST100 type transportation capsules. It is also necessary to mention the concept of modular assembly in low Earth orbit (MALEO) of the initial operational capability lunar habitation base [5]. Delivering the structural components of the base to the lunar surface in a finished form provides several important advantages, including the absence of contact with the lunar soil, which has undesirable abrasive and cohesive properties, and a safer radiation environment during the extra-vehicular activity in low Earth orbit (LEO). The critical technology in this concept, apparently, is the lunar landing system of the large structure, which should ensure the absence of shock and asymmetry when touching the lunar surface. Without a doubt, the questions of payload effectiveness and transportation capabilities remain of paramount importance, yet the problems of assuring functionality of payload draw the attention of authors of multiple papers. Paper [6] describes two approaches which should be implemented in future sustainable habitats: recycling technologies and primary production of material resources by applying the principles of in-situ resource utilization (ISRU). It also suggests using local resources for habitat protection. Paper [7] reasonably proposes to build lunar structures in successive phases and assure natural protection by embedding them in lava tubes. Paper [8] proposes a similar concept, stressing its important advantages such as protection not only from the meteoroid hazard but also from lunar dust which causes risks for equipment and crew's health [9,10]. There are also proposals for the use of space station modules as the first lunar base habitats and laboratories, covered with a 2 m thick lunar regolith protective layer [11]. Earlier, paper [12] described the ideas of fabricating load-bearing structures from regolith sintered by the focused sunlight [13] or microwave [14], and paper [15] summarized the known concepts of lunar bases, including those made of pneumatic, prefabricated, framed, combined, and other structures, focusing on using lunar regolith. According to some concepts, radiation and micrometeoroid protection of a deployable membrane structure with pre-assembled rigid elements can be provided by using a regolith shield [17] or a shell made of sintered regolith and 3D printed in situ [16]. Paper [18] applies a similar approach to deployable mobile lunar modules that were described in Ref. [19]. One of the ideas is to produce fiberglass from silicon oxide

σMAX =

αT ΔTE , 2(1 − ν )

(1)

where αT is the coefficient of thermal expansion, ΔT is the temperature variation, E is the modulus of elasticity, ν is the structural material's Poisson ratio. Taking design values E = 8358 ± 748 MPa and ν = 0.246 … 0.25 for sintered regolith [27], the maximum stresses in the conditions of the shadow will be σMAX ≈ 6.5 MPa, which poses a risk to the wall's structural integrity. For comparison, the strength limit of the lunar concrete produced from samples returned by Apollo 16 is just about 8.3 MPa [28]. And still, for the case of delivering ready-touse protected modules, it is possible to provide the smallest daily temperature fluctuations by selecting a proper location of the lunar base, and the availability of in-situ resources for the construction of protective shells will not be a decisive factor. In the authors’ opinion, a combined approach which predominates in the known conceptual projects is most effective, provided that some of the main associated problems are solved. The lunar base described in Ref. [29] combines traditional permanent-pressure ISS-type modules that can have preinstalled equipment with vertical and horizontal, inflatable and telescopic expandable modules that enable larger net volume when deployed. Delivery with preinstalled equipment can also work out for the deployable modules that have a specific configuration and feature sufficient rigidity during transportation, and analysis of this possibility is one of the most relevant matters. Fig. 1 shows the general configuration of the suggested concept of Industrial Research Lunar Base (IRLB) in successive phases of construction [30–32]. Fig. 1a shows the phase of IRLB minimal configuration, during which primary habitability is provided and the first power plant (PP), mechanized auxiliaries for assembly, and a vehicle for module transportation (MT) are delivered. This stage includes a manned expedition, and crew members have at their disposal a longterm shelter from the very beginning of the mission. The construction of 488

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Fig. 1. IRLB construction phases.

enable obtaining resources in situ (Si, Fe, Ca, Al, Mg, Ti, and other elements) as by-products of extraction of chemically bound oxygen that forms about 45% of mass of lunar soils and rocks, according to Ref. [36]. The process of electrolysis can be applied to water obtained from ice to produce propellant components such as liquid hydrogen and oxygen [37] while temperature conditions of the polar region enable proper storage of cryogenic components. Among other advantages, this location allows the longest time to obtain solar energy during lunar day, possibility for the team to evacuate and return to the Earth in case of an emergency, convenient transport connection with lunar orbit, and the least effect of the Moon's gravitational field on a manned spacecraft which travels on stand-by in lunar orbit. An important criterion for the selection of a suitable lunar landing site is a certain degree of natural protection, which is provided in the selected place [38]. Lastly, terrain features of the Moon and the ability of lunar module mounting supports to compensate for level differences of up to 400 mm enable easy connection of separate lunar modules. The tried-and-true technologies of fabricating lightweight, leaktight load-bearing shells from grid stiffened structures were taken into account in designing standard structures of vertical and horizontal lunar modules. A free volume required for one mission team member was determined by using known contemporary methods (such as that described in Ref. [39]) and taking into consideration the problem of monotony in long space missions [40]. The Celentano curve [41] was also applied to find the smallest habitable volume reduced to the duration of the mission (for one man). The accommodation capability for a 180-day lunar mission is four people for an IRLB horizontal module with net volume of 39.5 m3 and eight people for a vertical module with net volume of 99 m3. For comparison, the authors of [42] suggest that at least 19.4 m3 of habitable volume will be required for each of the four team members in a 501-day mission to Mars. The strategy for determining a relation between the level of air pressure in a habitable module and the content of oxygen in the air proceeded from the analysis of known factors such as the necessity of prebreathable airlocks. Accordingly, the design value of the module's inner pressure (101 kPa at 21% of O2) was one of the initial parameters for determining a required structural strength of membrane-framework modules. It may be reasonable to decrease the pressure since this makes a mission more economical and causes no effect on people's safety and physiology, though results in higher O2 content and, consequently,

the initial lunar settlement using teleoperated from Earth robotic equipment was not considered a priority for some reasons, including roundtrip time delay, which reduces the effectiveness of this approach in aberrant situations or anomalies [33]. A gateway module (GM) which provides access from base's inner premises to the lunar surface and back, as well as a habitation module (HM) designed to support crews' life during normal operation of the colony, are built around the standard configuration of a horizontal lunar module [30–32]. Fig. 1b and c shows the phases of IRLB expansion through introducing one by one a command module (CM) for base operation control and communication with the Earth and a production and repair module (PRM) for equipment maintenance. The CM and PRM are built around a typical vertical lunar module with four mating flanges that are attached to the corresponding interfaces of horizontal modules [30–32]. A space experiment module (SEM) for basic and applied research, a vivarium module (WM) for gradient transition of lunar base's life support systems to using own resources, a storage module (SM), and a research rover (R) are introduced during this phase as well. The phase is considered accomplished upon the construction of two framework-membrane hangars with opening shells (H1 and H2) for keeping the equipment (MR, R, and other units). The final phase (Fig. 1d) includes the construction of a fully selfsufficient life support system comprising two vivarium modules (WM1 and WM2) and a production complex with facilities for production of propellants, constructional and structural materials, rare-earth minerals, and crew life support resources (water, oxygen). The goal of the phase is to assure the possibility of permanent human presence and activity on the Moon. The phase also includes the construction of a space tourism complex and a lunar observatory at least 1 km away from the IRLB for studying the Universe [30–32]. In addition, the sufficient distance of the landing site from the assembly site of IRLB infrastructure elements should provide protection against high-energy ejecta missiles, which in the absence of atmosphere and with very low gravity on the moon can have a devastating effect on the crew and deployed surface assets [34]. According to the conceptual design, the IRLB will be located on the rim of the Shackleton crater that lies at the geographical south pole of the Moon. Locating the base in the polar region will enable obtaining water ice which is available in average concentrations of about 1,5% in perpetually shadowed craters near the lunar poles [35]. It will also

489

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Fig. 2. Standard horizontal (a) and vertical (b) configuration of stationary lunar modules.

poses a risk of flame spread in case of an emergency. Fig. 2 shows standard configurations of stationary lunar modules, which are the most suitable for permanent accommodation of the crew throughout a long-term mission. The outer diameter of horizontal module's pressure shell (1) is 3 m and the length along the mating device is 6 m (Fig. 2a); for the vertical module (Fig. 2b), these parameters are 5 m and 6 m, respectively. The latches which connect the modules with airtight door (2) and mating device (4) have the passage diameter of 1.5 m and assure integrity and airtightness of joints in case of seismic activity. The modules are equipped with a pressure relief device (5). Foldable mounting supports (3) are designed with due account of the physical and mechanical properties of lunar soil and results of lunar soil simulant tests (as described in Ref. [43], for instance). The outer 10 mm thick aluminum surface of the pressure shell has a multilayer, up to 100 mm thick cover to assure radiation, meteoroid, and thermal protection of the modules. Horizontal modules have an additional radiation barrier which is mounted on the pressure shell's inner surface. Fig. 3 shows the arrangement of vertical and horizontal configurations of the habitable module (a), vivarium module (b), and production and repair module (c). As it was mentioned above, almost each of the popular lunar base concepts is based on using deployable technologies in the structures of lunar base modules, and this obviously will not change for future missions to the Moon. For instance, the authors of [44] express the opinion that the inability to obtain structural materials from the resources in situ is one of the main weaknesses of inflatable on-surface structures and suggest that this problem can be solved by using lunar fiberglass [45]. The possibility of using the deployables in standard structures of the

modules for permanent crew accommodation is described below in this article. Using the deployables is a more obvious option for the hangartype semispheric modules (H1, H2, WM1, and WM2) because of their supporting role in the IRLB infrastructure. Paper [46] describes the details of the Lunar Greenhouse project (LGH), in the design of which pneumatic structures are used [47]. The suggested structure of LGH comprises four independent cylindrical growth chambers, each of which has about 19 m3 of net volume. For hangar modules of larger net volume, the known options of making large-sized developable or deployable shells in the form of inflatable membranes with metal framework can appear to be more appropriate owing to guaranteed structural strength. Pneumatic modules designed for the Martian Base [48] may be more space-saving with regard to delivery despite having the base of more than 20 m in diameter. This is explained by little thickness of modules’ shells (about 7 mm), which is like that because these shells should not assure protection from micrometeoroids that can burn down in the Martian atmosphere. It is possible to enlarge habitable space of the lunar base by means of the multispherical arrangement [49], in which the basic load due to the action of inner pressure is carried by cables which separate spherical membrane structures. In this concept, the thickness of Kevlar and Spectra membranes of the shell is 0.005 m, the maximum diameter of the structure is 22 m, and the pressure drop is 0.1 MPa. Obviously, an additional analysis should be done to prove the sufficiency of protective properties and the possibility of folding a membrane of the suggested shape in a space-saving manner for delivery. The dimensions of hangars H1 and H2 are defined by minimal combined dimensions of lunar equipment being in service; the dimensions of vivariums WM1 and WM2, by the values of gross life support

Fig. 3. The arrangement of IRLB modules of horizontal (a), (b) and vertical configuration. 490

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Fig. 4. Overall view of IRLB membrane-framework structures and the procedure of compact folding of the outer shell's membrane.

possibility of implementing the concept, strategies of shielding and possibilities of applying them in transformable modules shall be studied.

requirements. According to Ref. [50], gross requirements for a 180-day mission to the Moon are estimated at about 22.2 tons, and about 30–40 m2 of cultivated area is needed to provide enough food for one man [51]. The valuable experience of the Arthur Clarke Mars Greenhouse [52] in greenhouse engineering and training for using biologically-based life support systems beyond the Earth was applied to design the IRLB vivariums. The resources available in situ can include buffer gas, a basic consumable substance which is used for desaturation of oxygen for breathing. As shown in the example of a mission to Mars described in Ref. [53], a required amount of oxygen can be counted in metric tons. In order to build the IRLB membrane-framework hangars and vivariums with the base of 176 m2, it is suggested to optimize the sequence of load-bearing structure assembling and the method of membrane's compact folding for the outer shell's membranes (Fig. 4). The shell comprises several multilayer membranes of approximately equal thicknesses, with specific functions of a radiation, ballistic, and thermal barrier. A flexible, 7 mm thick protective membrane (a) with a base of 15 m in diameter can be transformed into a 1 m high disk (b) by making concentric folds; the disk can be next compacted and coiled to a diameter of 5 m (d), with height remaining unchanged (c). The points of structural ruptures that affect the reliability of the shell made of soft materials coincide with elements of the load-bearing framework and require no additional strengthening such as, for instance, reinforcement with flat ribs as described in Ref. [54]. Fig. 4e shows a pattern of a shell folding between stages (b) and (c) and equivalent stress distribution fields σe (Von Mises) in the material of shell's layers when the shell is folded up to maximum compaction. The maximum values of stresses are equal to σe = 37.5 MPa and do not exceed the value of tensile strength σU = 43 MPa of the used material (HDPE). At the same time, the value of total deformations does not exceed 45%. Discussing the methods of connecting the lunar modules, S. Timoshenko and S. Woinowsky-Krieger [26] mentioned that the welding of large-sized components in space could be the best solution for assembly. To assemble load-bearing structures of IRLB hangars, electron-beam welding with the help of small-sized hand tools can be an acceptable solution for framework (F) sections that are made of cast porous aluminum [55]. Therefore, the authors incline to the concept of a lunar base with structural elements that are either ready-to-use or ready-for-assembly and do not need any technologies applied in situ for shielding, stressing the advantages of weight and dimensions of those elements. A configuration of the base is proposed which can assure the required functionality, has a modular structure that can be expanded successively, and introduces deployable technologies in the modules wherever this does not affect module's habitability. Accordingly, to assess the

2.1. Shielding technologies The known concepts of using lunar terrain features for colony protection are quite reasonable ([7], for instance), and yet they cannot be a single option because of their feasibility problems. Some approaches to shielding suggest the use of lunar resources for the in-situ production of even traditional materials such as Nextel™, the ceramic cloth which is used effectively for meteoroid protection [56]. However, in addition to being not properly ready, in-situ technologies require the delivery of the additional payload. Using a water jacket for radiation protection of a four-people shelter involves additional equipment which weighs about 200 kg [57]. For any of the base expansion scenarios, the shielding strategies introduced at the phase of implementing future technologies of lunar resource utilization should be built on ready solutions. For this reason, the authors considered systems that are either fast-deployable or delivered ready-to-use as the best option for IRLB module protection. In the analysis of protection from factors of the space environment ([58–61]), including thermal, ballistic, and radiation barriers, the questions of building the latter are discussed most of all. A number of papers studied the materials involved in a spacecraft radiation analysis [62], assessed the possibilities of crew habitation in conditions of radiation exposure in a long mission [63] and analyzed the efficiency of methods of passive and active protection from ionizing radiation for astronauts [64]. However, most of the authors incline to the conclusion that radiation protection cannot rely on spacecraft's mechanical structures only and, therefore, a temporary shelter is necessary. It is proposed in Ref. [65] to use superconductive magnetic systems for a shelter, although they have some disadvantages such as superconductor's great mass and low operating temperature (10–15 K) and accordingly, require thermal protection and consume power for cooling. The authors of [57] suggest building a shelter for protection from solar particle events (SPE) by using temporary expandable or deployable shelters made of hydrogen-rich materials such as polyethylene which, as compared to aluminum, is about a 30% more effective absorber of radiated high charge and energy (HZE) particles [66]. Risks from secondary radiation appear to be another weakness of aluminum. Paper [67] determines the values of whole body effective doses, including secondary radiation, for various materials, aerial densities, and types of particles. Shielding with aluminum has a hazardous side effect of a large number of released neutrons, while highdensity polyethylene is one of the most effective and cheapest shielding materials. It is said in Ref. [42] that shielding with transformable HDPE 491

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Fig. 5. Cylindrical pressure shell of an IRLB horizontal module (a) and configuration of a multilayer ballistic barrier (b).

structural integrity and long-term protection to modules of the future lunar base. The authors focus mainly on the problem of providing habitability of module's pressure shells which can be folded compactly for delivery.

panels shows promise for reducing a passive mass required for protection from both solar events and galactic cosmic rays which consist of protons, electrons, and ionized light elements and the biological effect of which has not been studied fully so far [68]. In the authors' opinion, such known and effective methods as shielding an aluminum shell with a graphite-epoxy resin [69] or using hydrogenated graphite nanofibers, or HGNF [57], have no obvious advantages as compared to available and easy-to-shape polyethylene. Standard structures of IRLB horizontal modules have additional protection by 20 mm thick HDPE panels attached to the grid-stiffened structure on the inner side of the cylindrical shell (Fig. 5a). To determine an optimum design of meteoroid barrier, a critical thickness of a Whipple shield which comprises an aluminum wall and multilayer insulation was analyzed by using standard ballistic limit equation (BLE) [70,71] for the case of a bombarding particle impacting along the normal. The suggested design of multilayer barrier (Fig. 5b) includes the configuration features of stuffed Whipple shields which have been designed by NASA [72] and ESA [73] and have similar protection parameters. The structure made of Nextel™ and Kevlar© [72] may appear more preferable due to easy assembly and a possibility of replacement, as against a Nextel™ structure with Kevlar and epoxy composite plates, suggested in Ref. [73]. Paper [74] describes an example of a meteoroid barrier made of a 0.1 cm thick aluminum plate attached to a 0.276 cm thick Kevlar panel. The barrier demonstrated the effective performance in conditions of being bombarded by aluminum spheres 0.32 cm in diameter at a speed of 7 km/s. Paper [75] makes a comparative evaluation of the effectiveness of different types of barriers, according to which a more complex configuration of a stuffed Whipple shield can be less effective than a regular Whipple shield's configuration, while a foam-core sandwich panel can be twice as light as a honeycomb sandwich panel with the same ballistic characteristics. The suggested configuration of the barrier (Fig. 5b) can use open-cell foam as filler (for instance, polyamide foam AC 550 (0.018 g/cm2) or AC 530 (0.014 g/cm2)). External thermal insulation of standard modules was made of standard 10 mm thick MLI and disrupter layers of Beta-cloth and fiberglass cloth, both of which have demonstrated good performance in the Integrated Thermal Micrometeoroid Garment. It was taken into account for calculations of insulation that the temperature of the module's surfaces can be higher than that of the lunar surface due to the albedo of surrounding regolith fluff's radiation [76]. Thermal insulation of IRLB modules can include Dacron scrim separators and layers of standard metalized Kapton® that demonstrates stable mechanical properties and little thermos-optical degradation under long-term radiation. Obviously, a better situation with MLI mass owing to the use of new materials (for example, polymer films [77]) is not a decisive factor; however, coating the module's pressure shell with heat-insulating polymers with micro- and nano-channels can show promise [78]. In general, the analysis of lunar surface's conditions proves that there are no problems that could be critical in relation to assuring both

2.2. Deployable technologies The importance of the advantages of deployable structures, such as high compactness and excellent mass-dimensional characteristics, is growing with the increasing time of space missions. For this reason, structures of future lunar habitats include various types of deployables in most of the known concepts. It is necessary to find drastic solutions to decrease the tremendous costs of lunar missions which are going to be more complex and last longer than the Apollo missions ([79]). This can be done in two ways: by upgrading the transportation vehicles or by optimizing the mass and dimensions of payload. Other things being equal, compact folding of large-sized shells of lunar habitats enables fewer launches, but the problem of arranging cargo and preinstalled equipment in the shells still needs a solution. Perhaps, the main weakness of available deployable and inflatable structures lies in their low dimensional stiffness as compared with traditional ISS-type modules. Inflexible combinations of shielding materials provide the best protection of habitable modules, and this raises a problem of assuring habitability of inflatable habitats with flexible synthetic shells. According to Ref. [80], for instance, the analysis of radiation effect on the crew for the TransHab transformable module (see Ref. [81] for the detailed description of TransHab's shell) and the Columbus ISS module shows the best shielding capabilities of the latter. As compared to metal and synthetic shields used in ISS-type modules, the TransHab's MMOD shield made of alternating Kevlar and polyurethane foam layers demonstrates poor performance, although being about 0.4 m thick [82]. On the other hand, the use of soft materials decreases the destructive effects of secondary ionizing particles being typical of metal structural materials [83]. The authors sought to find solutions for the aforesaid contradictions through a brief analysis of most representative concepts of habitable deployable structures. There are lots of designs of deployable and inflatable structures for space applications, by using which it is possible to solve more or less the problem of providing sufficient stiffness and required protective properties to the shell structure when it is deployed. For higher stiffness of the structure, it was suggested to combine inflatable shells with structural ribs such as tape springs [84,85] or carbon and Vectran enhanced bars [86]. In some designs, particular inflatable elements are used as ribs while an external rotary solar shield provides protection from radiation [79]. In this case, a temporary shelter's shell with the inner volume of 12 m3 has proper compactness owing to shelter's thickness (near 5 mm) being close to 4.8 mm, the thickness of the Extravehicular Mobility Unit (EMU). It is also possible to apply chemical polarization, heat supply, and UV radiation to increase the stiffness of a shell made of reinforced, thermo-curing, prepreg fibrous materials. In Ref. [87], the 492

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for TransHab, in particular, with the help of a rigid cylindrical core. In the development of IRLB deployable modules, the authors studied the possibility of combining the aforesaid advantages of both the deployables and the stationary ISS-type modules in a single protected pneumatic structure. If it is possible to fold a thin pressure shell compactly and unfold it next with keeping the structural integrity, then the structure's pressure shell can be made of the materials that are traditional for rigid load-bearing modules, such as aluminum. Combining stiffness and transformability, this shell can also meet high protection criteria as compared with soft pneumatic structures, which is assured by alternating metal, synthetic, and composite barriers like in ISS modules. Lastly, the compactness of this module can be greater than that of pneumatic structures reinforced with ribs and a foldable, transformable, or solidifying framework. The module's inside can be made habitable and ergonomic by applying common solutions such as a floor [48] made of composite materials [87]. The authors of [101] provide a comparative analysis of a deployable cone or cylinder to show that metal shells with surfaces of zero Gaussian curvature can be folded compactly. The use of a conical shell for IRLB transformable modules is explained by its considerably higher compactness, the absence of touching and overlapping of compact folds. Due to this, a module's folded metal surface can be coated with standard flexible protective barriers which can be easily unfolded ready-fitted by generating overpressure inside the shell's structure. Fig. 6 depicts the suggested solution for a multilayer folded wall of a horizontal module. A module's bilayer load-bearing shell made of aluminum alloy 2219 is, in fact, an analogue to a standard stuffed Whipple shield with filler which acts as a radiation barrier and partially increases the stiffness of a deployed shell. Fig. 6a shows a folded horizontal module of the IRLB. The module's pressure shell comprises eight deployable conical sections joined by frame rings. When deployed, the module keeps its load-bearing capability owing to isometric folding of the smooth conical shell, which enables folding with almost no tension or compression of structural material. Accordingly, for proper folding and deploying of a module's multilayer wall afterwards, all its layers should be isometric to each other as much as possible and pleated one by one into annular folds that are similar to each other. In case of a small thickness of protective layers which form the conical shell, folding is possible if middle surfaces of the folds are geometrically equivalent. The outer surface of the shell comprises a ballistic and a thermal barrier and is made of flexible materials, and its ability for folding with no harm to structural integrity by analogy with [48,81] or [102] is beyond question. A clearance between the frames enables forming of random folds of the soft outer protective surface, and a damping layer with polyurethane open-cell foam (which is resistant to radiation degradation, see Ref. [103]) compensates for displacements of surface elements of the coating with a total thickness of about 50 mm. The loadbearing frame rings can be used during delivery as attachment points for preinstalled equipment, which gives the deployable structure the advantages of rigid or combined structures. Besides, intermediate layers of meteoroid screens can implement various non-standard solutions such as metal grids, the endurance of which in conditions of micrometeoroid activity is described in Ref. [104]. A multilayer radiation barrier comprising polyethylene and Teflon layers for mutual friction reduction is placed in the clearance between two load-bearing aluminum walls. For instance, paper [105] gives an analysis of how the mechanical properties of Teflon insulation degrade under exposure to the factors of the space environment. It is assumed that no stable diffusion bonding is produced between the layers' surfaces in the restricted airtight space between the walls and, therefore, nothing impedes mutual sliding of the surfaces. Small gaps between the layers at the folds’ peaks are removed upon full deployment, and a multilayer shell of the module acquires the features of an integrated load-bearing structure. Spatial eight-node volumetric final elements were applied to

authors suggest applying the method of polymerization of a composite material with a liquid polymer matrix for glass cloth shells of largesized lunar modules. Obviously, these designs should take into account the whole complex of space environment factors (for example, [88]), including negative surface temperatures, as well as the complex of mechanical loads that may act on the structure's shell, for instance, when connecting lunar base modules. Widely represented concepts of modules built around a foldable load-bearing framework with a multilayer shell demonstrate a reliable way of achieving sufficient stiffness of the shell and allow delivery of ready-to-use modules, although they still remain just ideas. For example, a shell with a foldable carbon fiber frame [89] can be extended twice as long or more. If a module with a radius R has a flexible cylindrical shell with a length L, then a random fragment of the shell is a circle with radii R and (R-L) when folded, that is, its square is less by πL2 as compared with the initial position. Obviously, an additional analysis is required for the possibility of a fold formation on the protective blanket and an effect of fold's geometry on the structure's compactness. Papers [90,91] describe the designs of deployable lunar modules which are built on the principles of architectural bionics, suggest a design of the deployable shell made of multilayer panels, and describe how to attach these panels by using swivel joints. In this case, an analysis shall be done for the compactness of the structure as well as for the ability of flexible swivel joints of multilayer panels to compensate for the displacements upon deploying with no loss of tightness, particularly in the nodes where folds cross. A number of papers suggest using bilayer foam-filled membranes for lunar base shells. In addition to a complicated process of construction, the load-bearing capacity of such structures depends on the pressure in the dome. Paper [92] describes a lunar module made of a bilayer composite cloth membrane filled with foam (polyurethane) which is injected inside with the help of a compressed gas and solidifies. Paper [93] makes an assessment of how long it would take to provide auxiliary infrastructure on the Moon for this purpose. For a shell with a diameter of 18.3 m and a net volume of about 566.3 m3, the thickness of a bilayer wall filled with foam is 30.5 cm. At the same time, for 6.1 × 6.1 × 3.0 m regolith-shielded modules with the load-bearing framework, the thickness of an inflatable Kevlar49 dome's membrane is only 0.3 mm [94]. In the project Konpneu [95], on the contrary, the load-bearing and protective capabilities of the lunar module are provided by a shell which is made of regolith concrete applied onto an inflatable formwork which is removed afterwards. Each of the mentioned concepts of shielding with regolith may be implemented if the appropriate industrial infrastructure is provided at the stationary lunar base. Another problem that draws attention concerns the stresses due to temperature deformations in the shell (see Eq. (1)). Taking into account technical complexities of their implementation, such projects can find an alternative through applying technologies in situ, with no necessity in delivering and assembling a pneumatic membrane, see Refs. [96,97], or [98]. Lastly, projects that focus on the optimization of operational flexibility and compatibility of individual modules represent a combined approach to the construction of lunar habitat by using deployable technologies. Paper [99] describes the concepts, according to which both rigid and pneumatic modules, as well as load-bearing and flexible segments, can share the same structure. In these concepts, rigid elements with appropriate attachment points enable transportation of built-in and preinstalled equipment and assure air tightness and protection of inner systems during liftoff and landing, while a transformable shell ensures lighter mass and smaller volume for delivery. The flexibility of the shell's structural layers and, in addition, a minimal number of these layers give high compactness of deployable shells. The authors of [100], for instance, suggest that protective layers, functions of which do not include radiation protection specifically, can be grouped in standard combinations and tailored to different sizes and purposes of modules. As it was mentioned above, this problem is solved 493

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Fig. 6. IRLB horizontal transformable module: folded (a) and during unfolding the pressure shell (1:10) (b), and a result of numerical simulation of bilayer pressure shell's segment unfolding (c). For (b) and (c), the arrows indicate the heading of shell unfolding.

The deployable technologies are used in two types of IRLB modules: storage module (SM) and gateway module (GM), both of which do not require permanent habitability and, accordingly, need no additional radiation protection on the inner surface of the pressure shell. Besides, the designs of these modules include a minimum of target hardware for an entire structure (SM) or its segment which takes about a third of net volume (GM's airlock chamber). The process of unfolding an SM's scale model gives the understanding of a pressure shell transformation nature (Fig. 6b). Fig. 7 shows an overall view of the IRLB gateway module and some elements of its design [109]. The stationary shell (1) which is built around a grid-stiffened structure is attached to the load-bearing deployable shell (2) with an inner partition wall (3). Accordingly, the module's airlock chamber is delivered folded together with target hardware and consumable resources. The target hardware is secured on the frame rings inside the folded structural shell (Fig. 6a). The internal space of the folded airlock chamber is enough to accommodate space suit maintenance station (4), bottles with oxygen/nitrogen mixture (5) for airlock chamber pressurization, and magnetic arch trap (6). Emergency reserves (7), reservoirs with water for sublimation heat exchangers of spacesuit cooling system (8), spacesuits on stands (9), and other things are held in the GM stationary section. The GM sections have symmetrical points such as supports and fasteners for rigging operations (10), and module's three-dimensional stiffness during tilting operations is assured by a lower conduit which is fastened to the stationary section and used as a support and a guide for the unfolding of the deployable shell. The shell (2) makes it possible to reduce the initial length of the GM to about 4.5 m and, together with a length of about 2 m of the folded SM, to improve significantly the procedure of IRLB payload delivery discussed below.

simulate the process of shell layer deployment in a numerical threedimensional final element model. The interaction between load-bearing layers and filler's layers was simulated by using contact pairs. No hardening was considered for the filler material. For the material of load-bearing layers, a bilinear isotropic hardening model based on the Mises yield criterion

F (σe, σy ) + σe σy = 0,

(2)

was applied, where σe is the equivalent stresses, σу is the yield stress of the particular material. The gas pressure in the inner volume of the airtight shell is taken as a basic load. Taking into account the known nonlinear relation between stresses and deformations, the material is considered elastic and plastic. The design method was a Euler-Lagrange method for nonlinear dynamic analysis, which is applicable for the simulation of flows of materials with large deformations. The results of the three-dimensional numerical simulation (Fig. 6c) illustrate the values of equivalent stresses in structural layers of the deployable module's pressure shell. At the end of the transformation, the values of equivalent stresses in filler are σе = 13.4 MPa, which is about 54% of the yield strength of high-density polyethylene (HDPE) (RP0.2 = 22 … 25 MPa). The maximum values of equivalent stresses in multilayer shell's load-bearing layers are σеmax = 363 MPa (see Fig. 6c), which is slightly more than the yield strength of the used aluminum alloy. The maximum equivalent stresses occur in the fold peaks where the shell's material experiences plastic deformations. The values of maximum displacements in a numerical model after pressure relief make it possible to estimate the transformation coefficient of the whole structure, that is, the ratio of its lengths in the folded and deployed states. When the module is fully deployed, it can be evident from a constant value of reversible elastic deformations of the shell upon release of inner pressure, the maximum value of which was about 60 kPa in a transformation experiment run for a 1/10 scaled dummy (Fig. 6b). When the structure is fully unfolded, the annular ribs help maintain structural integrity along the meridian and circumferential directions, similar to the tendons of a soft “pumpkin” balloon [106] or a toroidal inflatable element with two curvatures in two basic directions [107,108]. With this configuration of protective layers, the design parameters of folds allow the unfolding of module's conical sections, thereby tripling its linear dimension from 2000 mm to 6000 mm.

3. Results and discussion Highly unified lunar base components and ready solutions available for all project phases and adopted from experience (like [110], for instance) are the main advantages of the suggested concept of IRLB. Reusable transportation vehicles are suggested for an optimum mission scenario, as well as a smaller number of their flights because of using the deployable technologies [111]. A space transportation system (STS) plays the primary role in the 494

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Fig. 7. Overall view of the IRLB gateway module.

descend, and land on the Moon. The moonship with the AC, PS1, and PS2 stays in orbit (Phase II). When operations on the Moon are completed, the MC with the crew returns to the CO in the TM and docks with the moonship; then the TM and PS1 are separated (Phase III). Next, PS2 brings the MS from CO to LEO. To return the crew to the Earth, the moonship docks with an orbiter which is delivered to LEO by a medium-lift launch vehicle with a capacity of 13–14 tons (Phase IV). The orbiter with the crew separates from the moonship, leaves the LEO, descends, and lands on the Earth like a plane. The moonship with the AC and MC stays in LEO. For next flights with new crews, the medium-lift launch vehicle delivers the orbiter with a new crew and moonship life support system consumables to LEO again, the orbiter docks with the moonship, and the crews change places. Then the super-heavy RC delivers new pairs PS1+PS2 and LP2+TM to LEO (Phase V) where they are positioned towards the direction of travel and put on standby. The MS docks with the LP2+TM pair, the previous PS2 separates, and the new PS1 and PS2 dock instead. Phases II-IV of the flight procedure are repeated. This is how the core of the moonship (AC and MC) is reused to send people to the Moon's surface and back to LEO with no descent and landing. With a lunar orbital station deployed in CO, it will be possible to use the moonship to send people to the station and back. While building the IRLB minimal configuration, a special emergency ship with a take-off module can be provided at the station to ensure an urgent evacuation of people from the Moon's surface to LEO or the Earth if the standard manned ship is unavailable in CO at that moment or in case of its sudden failure. To return cargoes from the Moon's surface to the Earth by using the STS, the MS is substituted with a transport ship with a cargo bay instead of the MC. The IRLB construction schedule, see Fig. 9, shows the sequence and the number of RC launches for the case of 3-month intervals between

delivery of cargo and crew. The STS is lifted into a low Earth orbit (200 × 200 km) by a super-heavy rocket carrier (RC) with a payload capacity of 91.5 tons, which is based on the Mayak-C3.9 launch vehicle and operates engines RD815 and RD835 in the first and second stages, respectively [112]. To deliver spacecraft to a circumlunar orbit (CO) at an altitude about 100 km, an unmanned version of the STS comprises a moon upper-stage rocket (MUSR) with payload capacity of 30.3 tons and a circumlunar space tug (CST) with payload capacity of 20.5 tons that are used, respectively, to boost the transportation system up to a speed of achieving the Moon and for circumlunar maneuvers and corrections to achieve the CO. To deliver IRLB components to the Moon's surface, the STS has a landing platform (LP1) which is used to leave the CO above the area of expected landing, descend, and land on the Moon's surface [32]. The first IRLB modules will be delivered during the phase of IRLB minimal configuration (see Fig. 1a), together with a four-man crew to assemble it. For this purpose, the STS has a manned spacecraft (MS) with a moon cabin (MC), accessory compartment (AC), and propulsion systems PS1 and PS2 that are used to sustain flight from LEO to CO and from CO to LEO, respectively. This STS configuration includes another landing platform (LP2) and a takeoff module (TM) used to deliver the moon cabin from the Moon's surface to CO. It is planned to send four manned missions with four men in each during the IRLB expansion phase (Fig. 1b and c); altogether the team will consist of 16–24 people on completion of building the IRLB infrastructure during the final phase of base deployment [30–32]. Fig. 8 shows an approximate procedure of using the STS of unmanned (a) and manned (b) configurations. The flight procedure for sending and returning the first IRLB team starts with delivering the MS, LP2, and TM to LEO by using the RC (Fig. 8b, Phase I). The spacecraft flies to CO with the help of PS1; then, the TM and the MC with crew separate from the MS, leave the CO,

Fig. 8. Using the STS of unmanned (a) and manned (b) configurations. 495

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Fig. 9. Schedule of IRLB construction.

configuration and purposes of particular modules can vary in accordance with the expansion of general infrastructure's functionality. Owing to the delivery of the base's protected modules ready-to-use, there is no necessity for using additional equipment for ISRU technologies, which contributes to a shorter time needed to assure primary habitability. The suggested mission scenario proposes the solutions for optimal employment of transportation vehicles due to multiple use of their particular components. The method of fast unfolding is suggested for protected deployable modules with rigid load-bearing shells that are designed for various purposes and are habitable once deployed, with no necessity of additional means of protection. The article points out that the use of deployable technologies can increase the mass-dimensional payload effectiveness.

them. All necessary explorations of the Moon by using unmanned spacecraft (remote sensing spacecraft, RSS) shall be completed during the preparatory phase: detailed mapping, 3D modeling of the surface, investigation, and analysis of lunar soils, selection of most suitable areas to locate the base. The operations of the preparatory phase include also the delivery of mobile research rover (R) to the location of the future colony and conducting reconnaissance. If the area is found suitable for the colony, the landing beacons are installed and a 3D model of the colony is built (updated). The delivery of IRLB modules starts when the minimal configuration of the base is completed and should be finished by the beginning of the production phase. The Figure demonstrates that using the deployable technologies during the phase of base expansion enables fewer flights of the super-heavy RC (13 against 18) to deliver the standard horizontal modules. The deployable modules acquire all necessary protective properties as soon as unfolded by inner pressure and do not need any additional measures for providing stiffness to the load-bearing shell. One of the arguments for IRLB construction by using components delivered ready-to-use is the inevitable long period of time taken to provide full self-efficiency of the base, which is expected to happen in the mid lunar base expansion phase, according to the suggested concept. Most of the mentioned ISRU technologies of construction and protection of the base's structural components require the availability of primary infrastructure to assure self-sufficiency of these technologies. On the other hand, fully completed habitable modules help a mission team get quickly involved in the research and production activity, while mass and dimensions being appropriate in the view of delivery means fewer launches of RC to deliver the auxiliary equipment.

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4. Conclusions The suggested concept of the Industrial Research Lunar Base shows the ways of solving the urgent problems concerning the construction of a permanent inhabited colony on the Moon's surface. Being versatile and compatible, standard structural elements of the IRLB enable an optimal sequence of constructing the base, according to which the 496

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