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===Sphere or spherical section===
[[File:Apollo cm.jpg|thumb|[[Apollo command module]] flying with the blunt end of the [[heat shield]] at a non-zero [[angle of attack]] in order to establish a lifting entry and control the landing site (artistic rendition)]]
The simplest axisymmetric shape is the sphere or spherical section.<ref>{{cite web|last1=Przadka|first1=W.|last2=Miedzik|first2=J.|last3=Goujon-Durand|first3=S.|last4=Wesfreid|first4=J.E.|title=The wake behind the sphere; analysis of vortices during transition from steadiness to unsteadiness.|url=http://sphere.meil.pw.edu.pl/publi/AoM_60_2008_6.pdf|website=Polish french cooperation in fluid research.|publisher=Archive of Mechanics., 60, 6, pp. 467–474, Warszawa 2008. Received May 29, 2008; revised version November 13, 2008.|access-date=3 April 2015|archive-date=December 21, 2016|archive-url=https://web.archive.org/web/20161221044217/http://sphere.meil.pw.edu.pl/publi/AoM_60_2008_6.pdf|url-status=live}}</ref> This can either be a complete sphere or a spherical section forebody with a converging conical afterbody. The aerodynamics of a sphere or spherical section are easy to model analytically using Newtonian impact theory. Likewise, the spherical section's heat flux can be accurately modeled with the [[Fay-Riddell equation|Fay–Riddell equation]].<ref name="Fay-Riddell">{{cite journal|last1=Fay |first1=J. A. |last2=Riddell |first2=F. R. |title=Theory of Stagnation Point Heat Transfer in Dissociated Air |journal=Journal of the Aeronautical Sciences |volume=25 |pages=73–85 |date=February 1958 |url=http://pdf.aiaa.org/downloads/TOCPDFs/36_373-386.pdf |format=PDF Reprint |access-date=2009-06-29 |issue=2 |doi=10.2514/8.7517 |url-status=dead |archive-url=https://web.archive.org/web/20050107202757/http://pdf.aiaa.org/downloads/TOCPDFs/36_373-386.pdf |archive-date=2005-01-07 }}</ref> The static stability of a spherical section is assured if the vehicle's center of mass is upstream from the center of curvature (dynamic stability is more problematic). Pure spheres have no lift. However, by flying at an [[angle of attack]], a spherical section has modest aerodynamic lift thus providing some cross-range capability and widening its entry corridor. In the late 1950s and early 1960s, high-speed computers were not yet available and [[computational fluid dynamics]] was still embryonic. Because the spherical section was amenable to closed-form analysis, that geometry became the default for conservative design. Consequently, crewed capsules of that era were based upon the spherical section.
Pure spherical entry vehicles were used in the early Soviet [[Vostok programme|Vostok]] and [[Voskhod programme|Voskhod]] [[space capsule|capsule]]s and in Soviet Mars and [[Venera]] descent vehicles. The [[Apollo command module]] used a spherical section forebody heat shield with a converging conical afterbody. It flew a lifting entry with a hypersonic trim angle of attack of −27° (0° is blunt-end first) to yield an average L/D (lift-to-drag ratio) of 0.368.<ref>{{Cite web |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |title=Hillje, Ernest R., "Entry Aerodynamics at Lunar Return Conditions Obtained from the Flight of Apollo 4 (AS-501)," NASA TN D-5399, (1969). |access-date=July 7, 2017 |archive-date=September 16, 2020 |archive-url=https://web.archive.org/web/20200916020329/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |url-status=live }}</ref> The resultant lift achieved a measure of cross-range control by offsetting the vehicle's center of mass from its axis of symmetry, allowing the lift force to be directed left or right by rolling the capsule on its [[Flight control surfaces#Longitudinal axis|longitudinal axis]]. Other examples of the spherical section geometry in crewed capsules are [[Soyuz programme|Soyuz]]/[[Zond program|Zond]], [[Project Gemini|Gemini]], and [[Project Mercury|Mercury]]. Even these small amounts of lift allow trajectories that have very significant effects on peak [[g-force]], reducing it from 8–9 g for a purely ballistic (slowed only by drag) trajectory to 4–5 g, as well as greatly reducing the peak reentry heat.<ref>{{cite report|last1=Whittington|first1=Kurt Thomas|title=A Tool to Extrapolate Thermal Reentry Atmosphere Parameters Along a Body in Trajectory Space|url=http://repository.lib.ncsu.edu/ir/bitstream/1840.16/6817/1/etd.pdf|website=NCSU Libraries Technical Reports Repository|date=April 11, 2011|publisher=A thesis submitted to the Graduate Faculty of North Carolina State University in partial fulfillment of the requirements for the degree of Master of Science Aerospace Engineering Raleigh, North Carolina 2011, pp.5|access-date=5 April 2015|archive-date=April 11, 2015|archive-url=https://web.archive.org/web/20150411064311/http://repository.lib.ncsu.edu/ir/bitstream/1840.16/6817/1/etd.pdf|url-status=live}}</ref>
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====Real (non-equilibrium) gas model====
A non-equilibrium real gas model is the most accurate model of a shock layer's gas physics, but is more difficult to solve than an equilibrium model. The simplest non-equilibrium model is the ''Lighthill-Freeman model'' developed in 1958.<ref>{{cite journal |last=Lighthill |first=M.J. |title=Dynamics of a Dissociating Gas. Part I. Equilibrium Flow |journal=Journal of Fluid Mechanics |volume=2 |pages=1–32 |date=Jan 1957 |doi=10.1017/S0022112057000713 |issue=1|bibcode = 1957JFM.....2....1L |s2cid=120442951 }}</ref><ref>{{cite journal |last=Freeman |first=N.C. |title=Non-equilibrium Flow of an Ideal Dissociating Gas |journal=Journal of Fluid Mechanics |volume=4 |pages=407–425 |date=Aug 1958 |doi=10.1017/S0022112058000549 |issue=04|doi-broken-date=November 1, 2024 |bibcode = 1958JFM.....4..407F |s2cid=122671767 }}</ref> The Lighthill-Freeman model initially assumes a gas made up of a single diatomic species susceptible to only one chemical formula and its reverse; e.g., N<sub>2</sub> = N + N and N + N = N<sub>2</sub> (dissociation and recombination). Because of its simplicity, the Lighthill-Freeman model is a useful pedagogical tool, but is too simple for modelling non-equilibrium air. Air is typically assumed to have a mole fraction composition of 0.7812 molecular nitrogen, 0.2095 molecular oxygen and 0.0093 argon. The simplest real gas model for air is the ''five species model'', which is based upon N<sub>2</sub>, O<sub>2</sub>, NO, N, and O. The five species model assumes no ionization and ignores trace species like carbon dioxide.
When running a Gibbs free energy equilibrium program,{{clarify|date=August 2018}} the iterative process from the originally specified molecular composition to the final calculated equilibrium composition is essentially random and not time accurate. With a non-equilibrium program, the computation process is time accurate and follows a solution path dictated by chemical and reaction rate formulas. The five species model has 17 chemical formulas (34 when counting reverse formulas). The Lighthill-Freeman model is based upon a single ordinary differential equation and one algebraic equation. The five species model is based upon 5 ordinary differential equations and 17 algebraic equations.{{Citation needed|date=December 2017}} Because the 5 ordinary differential equations are tightly coupled, the system is numerically "stiff" and difficult to solve. The five species model is only usable for entry from [[low Earth orbit]] where entry velocity is approximately {{cvt|7.8|km/s|km/h mph}}. For lunar return entry of 11 km/s<!-- 36545 ft/s in NASA 1960s units -->,<ref>[https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435.pdf Entry Aerodynamics at Lunar Return Conditions Obtained from the Fliigh of Apollo 4] {{Webarchive|url=https://web.archive.org/web/20190411091352/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435.pdf |date=April 11, 2019 }}, Ernest R. Hillje, NASA, TN: D-5399, accessed 29 December 2018.</ref> the shock layer contains a significant amount of ionized nitrogen and oxygen. The five-species model is no longer accurate and a twelve-species model must be used instead.
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===== HARLEM =====
PICA and most other ablative TPS materials are either proprietary or classified, with formulations and manufacturing processes not disclosed in the open literature. This limits the ability of researchers to study these materials and hinders the development of thermal protection systems. Thus, the High Enthalpy Flow Diagnostics Group (HEFDiG) at the [[University of Stuttgart]] has developed an open carbon-phenolic ablative material, called the HEFDiG Ablation-Research Laboratory Experiment Material (HARLEM), from commercially available materials. HARLEM is prepared by impregnating a preform of a carbon fiber porous monolith (such as Calcarb rigid carbon insulation) with a solution of resole phenolic resin and [[polyvinylpyrrolidone]] in [[ethylene glycol]], heating to polymerize the resin and then removing the solvent under vacuum. The resulting material is [[Curing (chemistry)|cured]] and machined to the desired shape.<ref>{{Cite journal |last1=Poloni |first1=E. |last2=Grigat |first2=F. |last3=Eberhart |first3=M. |last4=Leiser |first4=David |last5=Sautière |first5=Quentin |last6=Ravichandran |first6=Ranjith |last7=Delahaie |first7=Sara |last8=Duernhofer |first8=Christian |last9=Hoerner |first9=Igor |last10=Hufgard |first10=Fabian |last11=Loehle |first11=Stefan |display-authors=3|date=12 August 2023 |title=An open carbon–phenolic ablator for scientific exploration |journal=Scientific Reports |volume=13 |issue=1 |page=13135 |article-number=13135|doi=10.1038/s41598-023-40351-x |doi-access=free|issn=2045-2322 |pmc=10423272 |pmid=37573464|bibcode=2023NatSR..1313135P }}</ref><ref>{{Cite journal |last1=Poloni |first1=E. |last2=Bouville |first2=Florian |last3=Schmid |first3=Alexander L. |last4=Pelissari |first4=Pedro I.B.G.B. |last5=Pandolfelli |first5=Victor C. |last6=Sousa |first6=Marcelo L.C. |last7=Tervoort |first7=Elena |last8=Christidis |first8=George |last9=Shklover |first9=Valery |last10=Leuthold |first10=Juerg |last11=Studart |first11=André R. |display-authors=1 |date=2022 |title=Carbon ablators with porosity tailored for aerospace thermal protection during atmospheric re-entry |journal=Carbon |volume=195 |pages=80–91 |doi=10.1016/j.carbon.2022.03.062 |doi-access=free|issn=0008-6223|arxiv=2110.04244 |bibcode=2022Carbo.195...80P }}</ref>
====SIRCA====
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| archive-url = https://web.archive.org/web/20210905190358/https://ntrs.nasa.gov/citations/20160001151
| url-status = live
}}</ref> The temperature of the surface rises to incandescent levels, so the material must have a very high melting point, and the material must also exhibit very low thermal conductivity. Materials with these properties tend to be brittle, delicate, and difficult to fabricate in large sizes, so they are generally fabricated as relatively small tiles that are then attached to the structural skin of the spacecraft. There is a tradeoff between toughness and thermal conductivity: less conductive materials are generally more brittle. The space shuttle used multiple types of tiles. Tiles are also used on the [[Boeing X-37]], [[Dream Chaser]], and [[SpaceX Starship (spacecraft)|
Because insulation cannot be perfect, some heat energy is stored in the insulation and in the underlying material ("thermal soaking") and must be dissipated after the spacecraft exits the high-temperature flight regime. Some of this heat will re-radiate through the surface or will be carried off the surface by convection, but some will heat the spacecraft structure and interior, which may require active cooling after landing.<ref name=TPSpaper/>
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In some early ballistic missile RVs (e.g., the Mk-2 and the [[sub-orbital spaceflight|sub-orbital]] [[Project Mercury|Mercury spacecraft]]), ''radiatively cooled TPS'' were used to initially absorb heat flux during the heat pulse, and, then, after the heat pulse, radiate and convect the stored heat back into the atmosphere. However, the earlier version of this technique required a considerable quantity of metal TPS (e.g., [[titanium]], [[beryllium]], [[copper]], etc.). Modern designers prefer to avoid this added mass by using ablative and thermal-soak TPS instead.
Thermal protection systems relying on [[emissivity]] use high emissivity coatings (HECs) to facilitate [[radiative cooling]], while an underlying porous ceramic layer serves to protect the structure from high surface temperatures. High thermally stable emissivity values coupled with low thermal conductivity are key to the functionality of such systems.<ref name="rtps">{{cite journal | last1=Shao| first1=Gaofeng|display-authors=et al| title= Improved oxidation resistance of high emissivity coatings on fibrous ceramic for reusable space systems | journal= Corrosion Science | year=2019 | volume=146| pages= 233–246 | doi= 10.1016/j.corsci.2018.11.006 | arxiv=1902.03943 | bibcode=2019Corro.146..233S| s2cid=118927116}}</ref>
Radiatively cooled TPS can be found on modern entry vehicles, but [[reinforced carbon–carbon]] (RCC) (also called ''carbon–carbon'') is normally used instead of metal. RCC was the TPS material on the Space Shuttle's nose cone and wing leading edges, and was also proposed as the leading-edge material for the [[X-33]]. [[Carbon]] is the most refractory material known, with a one-atmosphere sublimation temperature of {{Convert|3825|C}} for graphite. This high temperature made carbon an obvious choice as a radiatively cooled TPS material. Disadvantages of RCC are that it is currently expensive to manufacture, is heavy, and lacks robust impact resistance.<ref>{{Cite web|url=https://history.nasa.gov/columbia/CAIB_reportindex.html|title=Columbia Accident Investigation Board|website=history.nasa.gov|access-date=July 12, 2017|archive-date=December 25, 2017|archive-url=https://web.archive.org/web/20171225231135/https://history.nasa.gov/columbia/CAIB_reportindex.html|url-status=live}}</ref>
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In early 2019, [[SpaceX]] was developing an actively cooled heat shield for its [[SpaceX Starship|Starship]] spacecraft where a part of the thermal protection system will be a [[transpiration cooling|transpirationally cooled]] outer-skin design for the reentering spaceship.<ref name=sdc20190123>[https://www.space.com/43101-elon-musk-explains-stainless-steel-starship.html Why Elon Musk Turned to Stainless Steel for SpaceX's Starship Mars Rocket] {{Webarchive|url=https://web.archive.org/web/20190203064031/https://www.space.com/43101-elon-musk-explains-stainless-steel-starship.html |date=February 3, 2019 }}, Mike Wall, space.com, 23 January 2019, accessed 23 March 2019.</ref><ref name=trati20190123>[https://www.teslarati.com/spacex-ceo-elon-musk-starship-transpiring-steel-heat-shield-interview/ SpaceX CEO Elon Musk explains Starship's "transpiring" steel heat shield in Q&A] {{Webarchive|url=https://web.archive.org/web/20190124041422/https://www.teslarati.com/spacex-ceo-elon-musk-starship-transpiring-steel-heat-shield-interview/ |date=January 24, 2019 }}, Eric Ralph, ''Teslarati News'', 23 January 2019, accessed 23 March 2019</ref> However, SpaceX abandoned this approach in favor of a modern version of heat shield tiles later in 2019.<ref name="musk20190924">{{cite tweet |last=Musk |first=Elon |author-link=Elon Musk |user=elonmusk |number=1176561209971101696 |date=24 September 2019 |title=@OranMaliphant @Erdayastronaut Could do it, but we developed low cost reusable tiles that are much lighter than transpiration cooling & quite robust|access-date=9 May 2021 |archive-url=https://web.archive.org/web/20210427153543/https://twitter.com/elonmusk/status/1176561209971101696 |archive-date=27 April 2021 |url-status=live}}</ref><ref name="musk20190724">{{cite tweet |last=Musk |first=Elon |author-link=Elon Musk |user=elonmusk |number=1154229558989561857 |date=24 July 2019 |title=@Erdayastronaut @goathobbit Thin tiles on windward side of ship & nothing on leeward or anywhere on booster looks like lightest option|access-date=9 May 2021 |archive-url=https://web.archive.org/web/20210427154113/https://twitter.com/elonmusk/status/1154229558989561857 |archive-date=27 April 2021 |url-status=live}}</ref>
The [[Stoke Space Nova]] second stage, announced in October 2023 and not yet flying, uses a regeneratively cooled (by liquid hydrogen) heat shield.<ref>{{Cite web |last1=Volosín |first1=Trevor Sesnic |last2=Morales |first2=Juan I. |date=2023-02-04 |title=Full Reusability By Stoke Space |url=https://everydayastronaut.com/stoke-space/ |access-date=2023-02-05 |website=Everyday Astronaut |language=en-US}}</ref>
In the early 1960s various TPS systems were proposed to use water or other cooling liquid sprayed into the shock layer, or passed through channels in the heat shield. Advantages included the possibility of more all-metal designs which would be cheaper to develop, be more rugged, and eliminate the need for classified and unknown technology. The disadvantages are increased weight and complexity, and lower reliability. The concept has never been flown, but a similar technology (the plug nozzle<ref name="auto"/>) did undergo extensive ground testing.
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=== LOFTID ===
[[File:LOFTID inflates in space while attached to Centaur upper stage.gif|thumb|LOFTID inflating in orbit]]
A {{convert|6|m|ft|adj=on|sp=us}} inflatable reentry vehicle, ''Low-Earth Orbit Flight Test of an Inflatable Decelerator'' ([[LOFTID]]),<ref>{{Cite news|url=https://spacenews.com/noaa-finalizes-secondary-payload-for-jpss-2-launch/|title=NOAA finalizes secondary payload for JPSS-2 launch|date=March 10, 2020|website=SpaceNews|access-date=March 14, 2020|archive-date=October 1, 2021|archive-url=https://web.archive.org/web/20211001035259/https://spacenews.com/noaa-finalizes-secondary-payload-for-jpss-2-launch/|url-status=live|last1=Foust |first1=Jeff }}</ref> was launched in November 2022, inflated in orbit, reentered faster than Mach 25, and was successfully recovered on November 10.
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# Peak dynamic pressure
Peak heat flux and [[dynamic pressure]] selects the TPS material. Heat load selects the thickness of the TPS material stack. Peak deceleration is of major importance for crewed missions. The upper limit for crewed return to Earth from low Earth orbit (LEO) or lunar return is 10''g''.<ref name=autogenerated1>{{Cite web |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740007423_1974007423.pdf |title=Pavlosky, James E., St. Leger, Leslie G., "Apollo Experience Report - Thermal Protection Subsystem," NASA TN D-7564, (1974). |access-date=July 7, 2017 |archive-date=October 1, 2020 |archive-url=https://web.archive.org/web/20201001133219/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740007423_1974007423.pdf |url-status=live }}</ref> For Martian atmospheric entry after long exposure to zero gravity, the upper limit is 4''g''.<ref name=autogenerated1/> Peak dynamic pressure can also influence the selection of the outermost TPS material if [[spallation]] is an issue. The reentry vehicle's design parameters may be assessed through numerical simulation, including simplifications of the vehicle's dynamics, such as the [[planar reentry equations]] and heat flux correlations.<ref>{{Cite journal |last1=Sutton |first1=Kenneth |last2=Graves, Jr. |first2=Randolph A. |date=1971 |title=A general stagnation-point convective heating equation for arbitrary gas mixtures |url=https://ntrs.nasa.gov/api/citations/19720003329/downloads/19720003329.pdf |journal=NASA Tr R-376}}</ref>
Starting from the principle of ''conservative design'', the engineer typically considers two [[Best, worst and average case|worst-case]] trajectories, the undershoot and overshoot trajectories. The overshoot trajectory is typically defined as the shallowest-allowable entry velocity angle prior to atmospheric [[Boost-glide|skip-off]]. The overshoot trajectory has the highest heat load and sets the TPS thickness. The undershoot trajectory is defined by the steepest allowable trajectory. For crewed missions the steepest entry angle is limited by the peak deceleration. The undershoot trajectory also has the highest peak heat flux and dynamic pressure. Consequently, the undershoot trajectory is the basis for selecting the TPS material. There is no "one size fits all" TPS material. A TPS material that is ideal for high heat flux may be too conductive (too dense) for a long duration heat load. A low-density TPS material might lack the tensile strength to resist spallation if the dynamic pressure is too high. A TPS material can perform well for a specific peak heat flux, but fail catastrophically for the same peak heat flux if the wall pressure is significantly increased (this happened with NASA's R-4 test spacecraft).<ref name=autogenerated1/> Older TPS materials tend to be more labor-intensive and expensive to manufacture compared to modern materials. However, modern TPS materials often lack the flight history of the older materials (an important consideration for a risk-averse designer).
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*[[Genesis (spacecraft)|''Genesis'']] – The parachute failed to deploy due to a G-switch having been installed backwards (a similar error delayed parachute deployment for the [[Galileo Probe|''Galileo'' Probe]]). Consequently, the Genesis entry vehicle crashed into the desert floor. The payload was damaged, but most scientific data were recoverable.
*[[Soyuz TMA-11]] – The Soyuz propulsion module failed to separate properly; fallback ballistic reentry was executed that subjected the crew to accelerations of about {{convert|8|standard gravity|m/s2}}.<ref>{{cite web|url=http://www.spaceflightnow.com/station/exp16/080502peggywhitson.html|title=Whitson describes rough Soyuz entry and landing|access-date=July 12, 2008|publisher=Spaceflight Now|year=2008|author=William Harwood|archive-date=December 19, 2008|archive-url=https://web.archive.org/web/20081219073602/http://spaceflightnow.com/station/exp16/080502peggywhitson.html|url-status=live}}</ref> The crew survived.
*[[SpaceX Starship integrated flight test 3|Starship IFT-3]]: The SpaceX Starship's third integrated test flight was supposed to end with a hard splashdown in the
Some reentries have resulted in significant disasters:
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File:Cross section of Gemini 2 heatshield.jpg|Cross section of Gemini 2 heat shield
</gallery>
==Environmental impact==
[[File:ISS-46 Soyuz TMA-17M reentry.jpg|thumb|A plume in Earth's upper atmosphere left behind by a Soyuz spacecraft having reentered]]
Atmospheric entry has a measurable impact on [[Earth's atmosphere]], particularly the [[stratosphere]].
Atmospheric entry by spacecrafts have reached 3 % of all atmospheric entries by 2021, but in a scenario in which the number of satellites from 2019 are doubled artificial entries would make 40 % of all,<ref name="h473">{{citation | title=Burned-up satellites are polluting the atmosphere | publisher=American Association for the Advancement of Science (AAAS) | date=23 July 2024 | doi=10.1126/science.zub5l4y | page=}}</ref> which would cause atmospheric [[aerosols]] to be 94 % artificial.<ref name="p330">{{cite journal | last1=Schulz | first1=Leonard | last2=Glassmeier | first2=Karl-Heinz | title=On the anthropogenic and natural injection of matter into Earth's atmosphere | journal=Advances in Space Research | publisher=Elsevier BV | volume=67 | issue=3 | year=2021 | issn=0273-1177 | doi=10.1016/j.asr.2020.10.036 | doi-access=free | pages=1002–1025| arxiv=2008.13032 | bibcode=2021AdSpR..67.1002S }}</ref> The impact of spacecrafts burning up in the atmosphere during artificial atmospheric entry is different to meteors due to the spacecrafts' generally larger size and different composition. The atmospheric pollutants produced by artificial atmospheric burning-up have been traced in the atmosphere and identified as reacting and possibly negatively impacting the composition of the atmosphere and particularly the [[ozone layer]].<ref name="h473"/>
Considering [[space sustainability]] in regard to atmospheric impact of re-entry is by 2022 just developing<ref name="b448">{{cite journal | last1=Miraux | first1=Loïs | last2=Wilson | first2=Andrew Ross | last3=Dominguez Calabuig | first3=Guillermo J. | title=Environmental sustainability of future proposed space activities | journal=Acta Astronautica | publisher=Elsevier BV | volume=200 | year=2022 | issn=0094-5765 | doi=10.1016/j.actaastro.2022.07.034 | doi-access=free | pages=329–346| bibcode=2022AcAau.200..329M }}</ref> and has been identified in 2024 as suffering from "atmosphere-blindness", causing global [[environmental injustice]].<ref name="p583">{{cite journal | last1=Flamm | first1=Patrick | last2=Lambach | first2=Daniel | last3=Schaefer-Rolffs | first3=Urs | last4=Stolle | first4=Claudia | last5=Braun | first5=Vitali | title=Space sustainability through atmosphere pollution? De-orbiting, atmosphere-blindness and planetary environmental injustice | journal=The Anthropocene Review | publisher=SAGE Publications | date=6 June 2024 | issn=2053-0196 | doi=10.1177/20530196241255088 | doi-access=free | page=}}</ref> This is identified as a result of the current end-of life spacecraft management, which favors the [[Orbital station-keeping|station keeping]] practice of controlled re-entry.<ref name="p583"/> This is mainly done to prevent the dangers from uncontrolled atmospheric entries and [[space debris]].<ref name="p583"/>
Suggested alternatives are the use of less polluting materials and by in-orbit servicing and potentially in-space recycling.<ref name="b448"/><ref name="p583"/>
==Gallery==
<gallery widths="200px" heights="150px">
Soyuz TMA-05M spacecraft reentry.jpg|Close up of reentry trail (Soyuz)
Soyuz TMA-05M capsule reentry.jpg|Early reentry [[Plasma (physics)|plasma]] trail (Soyuz)
File:Re-entry of Progress Spacecraft 42P - NASA Earth Observatory.jpg|[[Progress (spacecraft)|Progress]] during atmospheric entry over Earth
==See also==
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