Spacecraft Materials Research Articles

Automated nested co-design framework for structural/control dynamics in flexible space systems

Modern space missions for Earth Observation (EO) purposes often rely on satellites equipped with very large flexible appendages, such as antennas and solar panels, which are demanded to perform agile slew manoeuvres. In most cases, the elasticity of such systems cannot be neglected in the design of the attitude controller, as excessive elastic displacements of the structural elements may compromise their stability and pointing performance. In this scenario, this paper aims at proposing an automated nested optimization framework to simultaneously optimize spacecraft structural and control dynamics, to be applicable to a wide range of flexible spacecraft. The objective of such a co-design architecture is to modify design parameters, at both structural and control levels, to minimize the mass of the spacecraft while maximizing its agility and satisfying imposed requirements. Moreover, as robust multivariable techniques have become more and more applied to ensure satisfactory robust performance margins, this paper's goal is to pose a multi-channel structured H∞ control architecture in the co-design problem. To guarantee the generality of implementation, a structural design tool (MSC Nastran) is interfaced with a coding environment (Matlab/Simulink) to set-up an autonomous exchange of information between structural and control domains. Starting from an initial definition of the spacecraft material, geometry and control requirements (in terms of loop-shaping transfer functions), relevant parameters are extracted from the structural tool and a linearized dynamic model assembled. Then, a controller is synthetized based on the provided requirements, followed by a verification on the nonlinear plant of the satellite. The procedure is repeated until the stop criteria (based on tolerance and max iterations) is satisfied. Finally, the output of the proposed architecture is obtained as an optimized structural model and robust controller tailored for the satellite dynamics.

Resurfacing: An Approach to Planetary Protection for Geologically Active Ocean Worlds

The putative and potential ocean worlds of our solar system span the asteroid belt to the Kuiper Belt, containing within their icy shells past or present global saltwater oceans. Among these worlds, those bearing signs of present-day geologic activity are key targets in the search for past or extant life in the solar system. As the icy surfaces of these bodies are modified by geologic processes, landforms are erased and replaced through what is called “resurfacing.” To avoid contaminating sites for robotic spacecraft exploration, planetary protection requirements obligate missions to these ocean worlds to demonstrate a less than 10−4 probability of introducing a viable terrestrial microorganism into a liquid water body. To constrain the probability of subsurface contamination, we investigate the interaction with geologic resurfacing on an active ocean world. Through the example of Europa, we show how the surface age can be used to constrain the resurfacing rate, a critical parameter to estimate the probability that nonsterile spacecraft material present on the surface is geologically incorporated into the subsurface, and extend this example to mission scenarios at Ganymede and Enceladus. This approach was critical to demonstrating compliance with planetary protection requirements for the Europa Clipper mission, reducing its probability of contamination by two to five orders of magnitude. We also show how a Europa lander mission might be brought close to complying with planetary protection requirements, that a Ganymede impactor could easily comply, and that the situation of Enceladus, while more complex, can greatly benefit from this approach.

Flame spread behaviour of Polydimethylsiloxane (PDMS) membranes in 1 g and µg environments

Diffusion flame behaviour and silica ash (SiO2) production were experimentally studied for various Polydimethylsiloxane (PDMS) membrane thicknesses (0.125 mm to 1.0 mm) in normal gravity and during microgravity flight experiments. The flames were established on vertical samples (300 mm in length) and subjected to either opposed or concurrent forced flows (both laminar and turbulent), assimilating the NASA Test 1 that is in use for spacecraft material selection. The opposed flame spread rate was observed to be steady and could be estimated using classical theory. Under concurrent flow, the flame spread rate was only steady for very high forced flows. The opposed flame-spread rate ranged from 0.5 to 1.5 mm/s, while the concurrent case ranged between 0.1 and 12 mm/s. The transport of silica ash (SiO2) was found to affect the heat balance of the concurrent flame spread in a manner that resulted in unsteady flame spread. For opposed flame spread, on the other hand, the transport of silica ash showed to be irrelevant. The extinction behaviour for the concurrent flame spread was heavily dominated by the transport of silica-ash, while for opposed flames, extinction was due to kinetics (at high forced flows). In microgravity environments, the transport and deposition of silica ash is anticipated to dominate flame spread and near-limit as well. These results suggest that silica-based products might be less flammable in microgravity than other similar materials such as common thermoplastics (PP or PE) used as wire jackets.

Asynchronous Synergistic Damage Effect of Atomic Oxygen and Space Micro Debris on Kapton Film

In the low earth orbit environment, many environmental factors lead to the degradation of material properties. The synergistic effect of long-term atomic oxygen (AO) irradiation and instantaneous impact of micro debris (MD) on long-term and transient space environmental factors has attracted more and more attention. In this paper, the performance evolution of Kapton films under the conditions of MD, AO single factor load spectrum and MD + AO, AO + MD asynchronous synergistic load spectrum were studied by laser driven flyer and microwave atomic oxygen technology. The macro morphology, optical properties and quality changes of Kapton films before and after each load spectrum were compared, and the mechanism of micro morphology and structure changes was explored. The results show that compared with MD + AO loading spectrum, the surface holes of Kapton films are larger under AO + MD load spectrum condition, the residual aluminum particles formed by reverse sputtering of Al particles during impact are less, the average transmittance of the film decreases slightly, and the weight loss of Kapton film is slightly more under the same atomic oxygen exposure time. Under the condition of MD + AO load spectrum, plastic tearing cracks, craters and holes are formed on the surface of Kapton film; the edge of the hole formed under the condition of AO + MD load spectrum is straight, without obvious depression and tear characteristics. Under the condition of MD + AO load spectrum, due to the adhesion of Al after the impact of micro debris, the subsequent atomic oxygen erosion of the film is reduced, so the C-C bond is not seriously damaged, and a considerable part of the residual aluminum flyer is oxidized to alumina by atomic oxygen; The AO + MD loading spectrum test makes the film first eroded by atomic oxygen, resulting in the reduction in C–O bond and C–C bond. The fracture of C–N bond is caused by the hypervelocity impact of micro debris. Hypervelocity impact leads to the thermal decomposition of the material, destroys the C–N bond in the imide ring and generates an N–H bond. This study will provide a method reference and a reference for the multi-factor ground collaborative simulation of space environment of spacecraft materials.

Solvay to provide spacecraft materials

Solvay to provide spacecraft materials

Solvay to provide spacecraft materials to Avio

Solvay to provide spacecraft materials to Avio

Clouds of Spacecraft Debris Liberated by Hypervelocity Dust Impacts on Parker Solar Probe

Abstract Hypervelocity impacts on spacecraft surfaces produce a wide range of effects including transient plasma clouds, surface material ablation, and for some impacts, the liberation of spacecraft material as debris clouds. This study examines debris-producing impacts on the Parker Solar Probe spacecraft as it traverses the densest part of the zodiacal cloud: the inner heliosphere. Hypervelocity impacts by interplanetary dust grains on the spacecraft that produce debris clouds are identified and examined. Impact-generated plasma and debris strongly perturb the near-spacecraft environment, producing distinct signals on electric, magnetic, and imaging sensors, as well as anomolous behavior of the star tracker cameras used for attitude determination. From these data, the spatial distribution, mass, and velocity of impactors that produce debris clouds are estimated. Debris-cloud expansion velocity and debris fragment sizes are constrained by the observational data, and long-duration electric potential perturbations caused by debris clouds are reported, along with a hypothesis for their creation. Impact-generated plasma-cloud expansion velocities, as well as pickup acceleration by the solar wind and driven plasma waves are also measured. Together, these observations produce a comprehensive picture of near-spacecraft environmental perturbations in the aftermath of a hypervelocity impact.

Space debris generation in GEO: Space materials testing and evaluation

The aim of this work is to evaluate what happens to the spacecraft materials beyond the spacecraft End of Life. A review of spacecraft external materials and effects of space environment is presented. This paper results from a continued study on spacecraft material degradation, and space debris formation in geostationary orbit (GEO). In this paper a 20-year GEO dose profile that combines simultaneous UV, particles irradiation and thermal cycling was applied to a set of external spacecraft materials. These materials comprised MLI assemblies, Velcros fixation and spacecraft painting. The evaluation of these external spacecraft materials, exposed to simulated space environment have confirmed the criticality of degradation of MLI, Velcros fixation and painting, with delamination mechanisms and particulate contamination. The synergy of space radiation (particles, UV) and thermal cycling ages the material and induces mechanical stress, causing creation of brittle surfaces, cracks and delamination. These phenomena cause serious damage to exposed surfaces, changing the surfaces thermo-optical properties, and may induce the generation of space debris. In particular, experimental results show the delamination of internal MLI layers and the severe degradation of the Velcros.

Cosmic-ray flux predictions and observations for and with Metis on board Solar Orbiter

Context.The Metis coronagraph is one of the remote sensing instruments hosted on board the ESA/NASA Solar Orbiter mission. Metis is devoted to carry out the first simultaneous imaging of the solar corona in both visible light (VL) and ultraviolet (UV). High-energy particles can penetrate spacecraft materials and may limit the performance of the on-board instruments. A study of the galactic cosmic-ray (GCR) tracks observed in the first VL images gathered by Metis during the commissioning phase is presented here. A similar analysis is planned for the UV channel.Aims.We aim to formulate a prediction of the GCR flux up to hundreds of GeV for the first part of the Solar Orbiter mission to study the performance of the Metis coronagraph.Methods.The GCR model predictions are compared to observations gathered on board Solar Orbiter by the High-Energy Telescope in the range between 10 MeV and 100 MeV in the summer of 2020 as well as with the previous measurements. Estimated cosmic-ray fluxes above 70 MeV n−1have been also parameterized and used for Monte Carlo simulations aimed at reproducing the cosmic-ray track observations in the Metis coronagraph VL images. The same parameterizations can also be used to study the performance of other detectors.Results.By comparing observations of cosmic-ray tracks in the Metis VL images with FLUKA Monte Carlo simulations of cosmic-ray interactions in the VL detector, we find that cosmic rays fire only a fraction, on the order of 10−4, of the whole image pixel sample. We also find that the overall efficiency for cosmic-ray identification in the Metis VL images is approximately equal to the contribution ofZ ≥ 2 GCR particles. A similar study will be carried out during the whole of the Solar Orbiter’s mission duration for the purposes of instrument diagnostics and to verify whether the Metis data and Monte Carlo simulations would allow for a long-term monitoring of the GCR proton flux.

Carbon fiber reinforced elastomeric thermal interface materials for spacecraft

Due to the high-heat flux issue caused by the increasing power consumption of electronic devices in spacecraft, there is an urgent need but still a significant challenge to develop high thermal conductivity elastomeric thermal interface materials (TIMs) for spacecraft. In this study, mesophase pitch-based carbon fibers (CFs) are added as fillers on the basis of conventional alumina/silicone rubber TIMs, and the CFs are oriented in the matrix by a strong magnetic field during the preparation process. Attributed to the phonon-transport highways composed of vertically-oriented CFs and the synergistic improvement based on CFs and alumina, the through-plane thermal conductivity of the final elastomeric composite with 20 vol% CFs is increased by nearly 14 times compared with the alumina/silicone rubber composite, and it also confirms to the vacuum outgassing performance standard for spacecraft. Moreover, with the help of a verification system, the actual heat transfer capability of the composite is evaluated through the ground test and on-orbit test, and the results suggest that it has a perfect interface heat transfer capability both on the ground and in space, which reveals that this composite can potentially be applied as elastomeric TIM for spacecraft thermal control.

Optical Characterization of DebriSat Fragments in Support of Orbital Debris Environmental Models

The NASA Orbital Debris Program Office (ODPO) develops, maintains, and updates orbital debris environmental models, such as the NASA Orbital Debris Engineering Model (ORDEM), to support satellite designers and operators by estimating the risk from orbital debris impacts on their vehicles in orbit. Updates to ORDEM utilize the most recent validated datasets from radar, optical, and in situ sources to provide estimates of the debris flux as a function of size, material density, impact speed, and direction along a mission orbit. On-going efforts within the NASA ODPO to update the next version of ORDEM include a new parameter that highly affects the damage risk – shape. Shape can be binned by material density and size to better understand the damage assessments on spacecraft. The in situ and laboratory research activities at the NASA ODPO are focused on cataloging and characterizing fragments from a laboratory hypervelocity-impact test using a high-fidelity, mock-up satellite, DebriSat, in controlled and instrumented laboratory conditions. DebriSat is representative of present-day, low Earth orbit satellites, having been constructed with modern spacecraft materials and techniques. The DebriSat fragment ensemble provides a variety of shapes, bulk densities, and dimensions. Fragments down to 2 mm in size are being characterized by their physical and derived properties. A subset of fragments is being analyzed further in NASA’s Optical Measurement Center (OMC) using broadband, bidirectional reflectance measurements to provide insight into the optical-based NASA Size Estimation Model. Additionally, pre-impact spectral measurements on a subset of DebriSat materials were acquired for baseline material characterization. This paper provides an overview of DebriSat, the status of the project, and ongoing fragment characterization efforts within the OMC.

Методичний підхід до розрахунку вартості дослідно-конструкторських ро-біт зі створення космічних апаратів

The aim of this work is to develop a methodological approach to the development of a unified, common for all developers, computer-aided branch procedure for spacecraft development cost estimation based on a spacecraft cost model and practical recommendations on computer-aided branch procedure development. The cost parameters of the development, manufacturing, and operation of new spacecraft and their technical level (perfection) are the determining factors in competitiveness assessment. To decide on the advisability of starting or resuming the development of a new spacecraft, one has to correctly estimate the development cost. Using the standardized calculation method in estimating the development cost for new space hardware is unacceptable for lack of bug-free design and production documentation, which is the end product of any development activity. Parametric methods (the basic methods used to estimate the spacecraft development cost in the USA and Europe) cannot be used in the development of a high-quality procedure for spacecraft development cost estimation for lack of a branch statistical database on spacecraft development labor intensiveness and materials consumption at the State Space Agency of Ukraine. This calls for a nonstandard cost model of spacecraft development. The authors’ cost model is based on a method of componentwise analogy for simple spacecraft components, moving (up and down) along the edges of a weighted oriented tree graph that models the spacecraft technical structure, and fuzzy analysis methods. The tree graph Gi(V(С),D) models the spacecraft technical structure (V, C, and D are the sets of graph vertices and edges and spacecraft components, respectively; to each graph vertex there corresponds a spacecraft component). The paper presents a nonstandard cost model of spacecraft development, which in its essence is close to a nonlinear parametric cost model, and a scientific methodology for the development of an advanced branch procedure for spacecraft development cost calculation with component and stage detailing.

Resistance investigation of diamond-like carbon coatings to cyclic temperature changes

Optical elements used in outer space must be designed considering the effects of such factors as space vacuum, atomic oxygen in low Earth orbit, solar and space radiation, large temperature drops, gas release of spacecraft materials and structural elements, space dust and debris. In order to harden and protect mirror surfaces of optical elements from external factors, it has been promisingly applied diamond-like carbon coatings on their surface. These coatings are characterized by high strength and wear-resistant properties, in particular, high hardness, low friction coefficient, high wear resistance and chemical inertness. This leads to their widespread use in various fields of science and technology, including optical instrumentation. This paper presents the results of testing an aluminum mirror with a diamondlike carbon coating under the effect of cyclic temperature changes for determining their ability to withstand a rapid cyclic ambient temperature change, and specifically, to maintain optical and mechanical properties.

High-Performance Planar Thermal Diode With Wickless Components

Abstract The characterization “thermal diode” (TD) has been used to portray systems that spread heat very efficiently in a specific direction but obstruct it from flowing in the opposite direction. In this study, a planar vapor chamber (VC) with a wickless, wettability-patterned side and an opposing wick-lined side is fabricated and tested as a thermal diode. When the chamber operates in the forward mode, heat is naturally driven away from the heat source; in the reverse mode, the system blocks heat flow, thus acting as a thermal diode. The low-profile assembly takes advantage of the phase-changing properties of water inside a sealed chamber. The wettability-patterned plate—when on the cooled side, e.g., forward operation mode—enables spatially controlled dropwise condensation (high heat transfer rate) and filmwise condensation (high drainage rate), thus facilitating an efficient transport mechanism of the condensed medium on superhydrophilic wedge tracks by way of Laplace pressure-driven capillary forces. The same chamber acts as a thermal blocker when the wick-covered plate is on the cooled side (reverse operating mode), trapping the condensate in the wick pores and blocking heat flow to the opposite side. The system's thermal behavior is similar to the theoretical electrical diode. This work explores the effect of the condenser's wettability pattern design and the chamber's fluid charging ratio (CR). With this system, thermal diodicities exceeding 20 have been achieved, and are tunable by altering the wettability pattern. The thermal rectification concept and its proper quantification in terms of possible definitions are discussed. The present vapor chamber—thermal diode design could be well-suited for an extensive range of thermal-management applications, ranging from aerospace, spacecraft, and smart-building construction materials, to electronics protection, electronics packaging, refrigeration, thermal control during energy harvesting, thermal isolation, etc.

Spectroscopic behavior of various materials in a GEO simulated environment

Artificial objects currently populating Earth orbital regimes can be distinguished by comparing remote observational data to that of optical measurements within the visible (VIS) and near infrared (NIR) spectral regions of materials obtained in the laboratory. Careful comparison of spectrally resolved observations with laboratory-based material irradiation experiments can provide a tool for remote diagnosis and anomaly resolution. However, the spectral signatures of many materials change as a function of exposure to radiation in geosynchronous Earth orbit (GEO). In order for remote characterization to be a viable characterization technique, the evolution of material optical properties (VIS-NIR) upon exposure to the GEO environment must be understood. To this end, we have studied the reflectance of several commonly used spacecraft materials including organic polymers, solar cell coverglasses, black thermal control paint, and carbon-carbon (C–C) matrix composite as a function of exposure to high energy electron irradiation in order to mimic the harsh environment of GEO. It was found that different classes of materials exhibit large variation in radiation stability resulting in the development of unique spectral characteristics.

Cassini‐Plasma Interaction Simulations Revealing the Cassini Ion Wake Characteristics: Implications for In‐Situ Data Analyses and Ion Temperature Estimates

AbstractWe have used Spacecraft Plasma Interaction Software (SPIS) simulations to study the characteristics (i.e., dimensions, ion depletion, and evolution with the changing spacecraft attitude) of the Cassini ion wake. We focus on two regions, the plasma disk at 4.5–4.7 RS, where the most prominent wake structure will be formed, and at 7.6 RS, close to the maximum distance at which a wake structure can be detected in the Cassini Langmuir probe (LP) data. This study also reveals how the ion wake and the spacecraft plasma interaction have impacted the Cassini LP measurements in the studied environments, for example, with a strong decrease in the measured ion density but with minor interference from the photoelectrons and secondary electrons originating from the spacecraft. The simulated ion densities and spacecraft potentials are in very good agreement with the LP measurements. This shows that SPIS is an excellent tool to use for analyses of LP data, when spacecraft material properties and environmental parameters are known and used correctly. The simulation results are also used to put constraints on the ion temperature estimates in the inner magnetosphere of Saturn. The best agreement between the simulated and measured ion density is obtained using an ion temperature of 8 eV at ∼4.6 RS. This study also shows that SPIS simulations can be used in order to better constrain plasma parameters in regions where accurate measurements are not available.

Analysis of the Effects of Dynamic Alloying on the Structure of Aluminium and its Alloys

The dynamic alloying of aluminum and its alloy with a high-speed stream of silicon carbide (SiC) particles simulates the effect of a stream of cosmic dust on spacecraft materials. The study showed a structure change in the volume of aluminum and its alloy and the formation of new structural elements. The transformation of the structure during dynamic alloying leads to a change of the composition and mechanical properties of the matrix material.

A review on characteristics of composite and advanced materials used for aerospace applications

Space is a stupendous and elephantine field that needs immense exploration and the advancements in the field of spacecraft materials are the top-notch field of today's research. The objective of the article is to examine the compositions of advance and composite materials used in the designing of space shuttle. The present work will discuss the applications and mechanical properties of reinforced materials used in the aerospace industry. Moreover, the paper is focused to highlight the usage and endurance limits of aluminum alloys in the air frame structure. An attempt has been made to provide the distribution of metallic materials in the structure of space shuttle through graphical representation. The thermodynamic characteristics of composite materials and possible advancements in the composite materials are discussed. The merits and demerits of composites such as advance carbon, reinforced carbon, flexible insulation blanket etc. in air shuttle fabrication are discussed to provide further directions for researchers. The developmental stages and growth of space materials are provided to select the suitable compositions of space shuttle materials.

Property changes in materials due to atomic oxygen in the low Earth orbit

We expect satellites at altitude below 300 km, very low Earth orbit (VLEO), making observations of the Earth at optical wavelength with increasingly higher resolution. The density of atomic oxygen (AO) at VLEO is significantly higher than that at LEO; severe degradation of spacecraft materials (polymers) due to the high-flux AO is a serious concern. To clarify VLEO environmental effects on spacecraft materials, we designed the Material Degradation Monitor (MDM) and MDM2 missions. The MDM is a material exposure experiment onboard the Super Low-Altitude Test Satellite (SLATS). It aims to understand reactions and degradation of polymeric materials depending on AO fluence in VLEO. In the MDM, samples of spacecraft material were exposed at altitude of 160–560 km; their degradation behaviors were observed optically by a CCD camera for 1.8 years. The MDM2 is a material exposure experiment onboard the International Space Station (ISS) and aims to correctly understand surface reactions and degradation of the same samples used in the MDM at a given AO fluence. In the MDM2, the samples were exposed at altitude of 400 km for 1 year and then returned to Earth for analysis. Based on the results from both missions, we will help in the molecular design of more-durable materials, and establish design standards for future VLEO satellites. This study aims to quantitatively understand the surface reactions and degradation of the 11 types of thermal control materials exposed on the ISS in the MDM2. Five types of multilayer insulation (MLI) films (three types of Si-containing AO protective materials (a silsesquioxane-(SQ-) containing coated polyimide film, two types of polysiloxane-block polyimide (BSF-30) films), an ITO-coated polyimide film, and a Beta Cloth), and flexible optical solar reflectors (flexible OSRs) were found to have a high durability against erosion by AO. This was determined by measuring their loss of mass and thermo-optical properties. The Ag/Inconel layer’s discoloration and peeling were observed for three types of FEP/Ag films as determined by the Ag layer’s oxidation by AO. Also, X-ray photoelectron spectroscopy (XPS) showed that reactions of the Si-containing materials, the SQ-coated polyimide film and the BSF-30 film, form a layer of silica that protects against AO. Even though the concentration of Si in the SQ-coating is the same or greater than in the BSF-30 film, the amount of the SQ-coating that reacted was larger than that of the BSF-30 film under the same AO fluence. Moreover, the effective ability of the UV-shielding coating, composed of ITO and CeO2 coated onto one of the BSF-30 films, was demonstrated by UV–Vis spectrometry. Its sufficient AO protection was confirmed by mass measurements, XPS analyses, and FE-SEM observations.

Atomic oxygen effect of Zr-Al-C coatings on ZrNb alloys used in space environment

The preparation of suitable coating materials is crucial for enhancing the atomic oxygen erosion of zirconium alloys. Zr-Al-C ceramics composed of MAX phase have excellent inherited properties of metal and ceramic, making them useful as spacecraft materials for atomic oxygen protection. The MAX phases are a group of layered ternary compounds with the general formula Mn+1AXn (M: early transition metal; A: group A element; X: C and/or N; n = 1–3), which combine some properties of metals, such as machinability, low hardness and damage tolerance, with those of ceramics, such as high elastic moduli, high temperature strength, oxidation and corrosion resistance. In this paper, Zr-Al-C coatings were successfully prepared by the magnetron sputtering method. The macroscopy morphology, microstructure, adhesion and the effect of atomic oxygen of Zr-Al-C coatings were all investigated. The results showed the successful deposition of Zr-Al-C coatings consisting of Zr2Al3C4, Zr3Al3C5, and ZrAl3 phases on zirconium alloy surface by adding an intermediate layer using magnetron sputtering route. The adhesion between the coating and the substrate was estimated to 18 N by scratch testing. After 30 cycles of cold and hot impact tests, the Zr-Al-C coating remained undamaged without crack or peel off. The oxidation kinetics of ZIRLO with Zr-Al-C coating followed approximatively a linear variation law. The oxidation rate of ZIRLO containing Zr-Al-C coating was lower than that of uncoated Zr alloy within 30 h of atomic oxygen exposure. In other words, Zr-Al-C coating could protect Zr alloy under certain exposure dosage of atomic oxygen. Meanwhile, further increase in exposure time of atomic oxygen led to the formation of ZrO, ZrO2, Al2O3 on Zr-Al-C coating. Also, more pores appeared on Zr-Al-C coating surface due to the oxidation of carbon elements in the coatings to form volatile oxides, such as CO and CO2. As atomic oxygen erosion continued, the existing holes became connected to form complex channel-like labyrinthine two-dimensional network structures. The latter penetrated further the atomic oxygen into the interior of Zr-Al-C coating to continue the corrosion effect of atomic oxygen. By considering the protection effect of Zr-Al-C coating on Zr alloy under a certain exposure dosage of atomic oxygen, the protection time can be calculated and experiments can be designed according to the linear law obtained as a function of the thickness of Zr-Al-C coating. In sum, these results look promising for future application of MAX or MAX-like phase in spacecraft environment.