Editing Atmospheric entry (section)

==Thermal protection systems==
{{Main|Heat shield#Spacecraft}}
A '''thermal protection system''', or TPS, is the barrier that protects a [[spacecraft]] during the searing heat of atmospheric reentry. Multiple approaches for the thermal protection of spacecraft are in use, among them ablative heat shields, passive cooling, and active cooling of spacecraft surfaces. In general they can be divided into two categories: ablative TPS and reusable TPS. 

Ablative TPS are required when space craft reach a relatively low altitude before slowing down.{{dubious|date=November 2024}} Spacecraft like the space shuttle are designed to slow down at high altitude so that they can use reuseable TPS. (see: [[Space Shuttle thermal protection system]]). 

Thermal protection systems are tested in high enthalpy ground testing or plasma wind tunnels that reproduce the combination of high enthalpy and high stagnation pressure using [[Induction plasma]] or DC plasma.

===Ablative===
[[File:Apollo 12 heat shield.JPG|thumb|left|Ablative heat shield (after use) on [[Apollo 12]] capsule]]
The [[Ablation|ablative]] heat shield functions by lifting the hot shock layer gas away from the heat shield's outer wall (creating a cooler [[boundary layer]]). The boundary layer comes from ''blowing'' of gaseous reaction products from the heat shield material and provides protection against all forms of heat flux. The overall process of reducing the heat flux experienced by the heat shield's outer wall by way of a boundary layer is called ''blockage''. Ablation occurs at two levels in an ablative TPS: the outer surface of the TPS material chars, melts, and [[Sublimation (physics)|sublimes]], while the bulk of the TPS material undergoes [[pyrolysis]] and expels product gases. The gas produced by pyrolysis is what drives blowing and causes blockage of convective and catalytic heat flux. [[Pyrolysis]] can be measured in real time using [[thermogravimetric analysis]], so that the ablative performance can be evaluated.<ref>Parker, John and C. Michael Hogan, "Techniques for Wind Tunnel assessment of Ablative Materials", NASA Ames Research Center, Technical Publication, August, 1965.</ref> Ablation can also provide blockage against radiative heat flux by introducing carbon into the shock layer thus making it optically opaque. Radiative heat flux blockage was the primary thermal protection mechanism of the [[Galileo Probe]] TPS material (carbon phenolic). 

Early research on ablation technology in the USA was centered at [[NASA]]'s [[Ames Research Center]] located at [[Moffett Field]], California. [[Ames Research Center]] was ideal, since it had numerous [[wind tunnels]] capable of generating varying wind velocities. Initial experiments typically mounted a mock-up of the ablative material to be analyzed within a [[hypersonic]] wind tunnel.<ref>Hogan, C. Michael, Parker, John and Winkler, Ernest, of [[NASA]] Ames Research Center,  "An Analytical Method for Obtaining the Thermogravimetric Kinetics of Char-forming Ablative Materials from Thermogravimetric Measurements", AIAA/ASME Seventh Structures and Materials Conference, April, 1966</ref> Testing of ablative materials occurs at the Ames Arc Jet Complex. Many spacecraft thermal protection systems have been tested in this facility, including the Apollo, space shuttle, and Orion heat shield materials.<ref>{{Cite web|title = Arc Jet Complex|url = http://www.nasa.gov/centers/ames/research/technology-onepagers/arcjetcomplex.html|publisher = NASA|website = www.nasa.gov|access-date = 2015-09-05|archive-date = October 5, 2015|archive-url = https://web.archive.org/web/20151005011405/http://www.nasa.gov/centers/ames/research/technology-onepagers/arcjetcomplex.html|url-status = live}}</ref>

[[File:Mars Pathfinder.jpg|thumb|upright|''[[Mars Pathfinder]]'' during final assembly showing the aeroshell, cruise ring and solid rocket motor]]

====Carbon phenolic====
Carbon phenolic was originally developed as a rocket nozzle throat material (used in the [[Space Shuttle Solid Rocket Booster]]) and for reentry-vehicle nose tips.
The [[thermal conductivity]] of a particular TPS material is usually proportional to the material's density.<ref name="Di Benedetto">{{cite book|last1=Di Benedetto|first1=A.T.|last2=Nicolais|first2=L.|last3=Watanabe|first3=R.|title=Composite materials : proceedings of Symposium A4 on Composite Materials of the International Conference on Advanced Materials – ICAM 91, Strasbourg, France, 27–29 May 1991|date=1992|publisher=North-Holland|location=Amsterdam|isbn=978-0444893567|page=111}}</ref> Carbon phenolic is a very effective ablative material, but also has high density which is undesirable.

The NASA [[Galileo Probe]] used carbon phenolic for its TPS material.<ref name=HSAE>{{cite journal |last=Milos |first=Frank S. |journal=Journal of Spacecraft and Rockets |issn=1533-6794 |year=1997 |doi=10.2514/2.3293 |title=Galileo Probe Heat Shield Ablation Experiment |volume=34 |issue=6 |pages=705–713 |bibcode=1997JSpRo..34..705M |url=https://zenodo.org/record/1235941 }}
</ref>

If the heat flux experienced by an entry vehicle is insufficient to cause pyrolysis then the TPS material's conductivity could allow heat flux conduction into the TPS bondline material thus leading to TPS failure. Consequently, for entry trajectories causing lower heat flux, carbon phenolic is sometimes inappropriate and lower-density TPS materials such as the following examples can be better design choices:

====Super light-weight ablator====
''SLA'' in ''SLA-561V'' stands for ''super light-weight ablator''. SLA-561V is a proprietary ablative made by [[Lockheed Martin]] that has been used as the primary TPS material on all of the 70° sphere-cone entry vehicles sent by NASA to Mars other than the [[Mars Science Laboratory]] (MSL). SLA-561V begins significant ablation at a heat flux of approximately 110&nbsp;W/cm<sup>2</sup>, but will fail for heat fluxes greater than 300&nbsp;W/cm<sup>2</sup>. The MSL aeroshell TPS is currently designed to withstand a peak heat flux of 234&nbsp;W/cm<sup>2</sup>. The peak heat flux experienced by the ''[[Viking 1]]'' aeroshell which landed on Mars was 21&nbsp;W/cm<sup>2</sup>. For ''Viking 1'', the TPS acted as a charred thermal insulator and never experienced significant ablation. ''Viking 1'' was the first Mars lander and based upon a very conservative design. The Viking aeroshell had a base diameter of 3.54&nbsp;meters (the largest used on Mars until Mars Science Laboratory). SLA-561V is applied by packing the ablative material into a honeycomb core that is pre-bonded to the aeroshell's structure thus enabling construction of a large heat shield.<ref>{{cite tech report
| first=Huy
| last=Tran
| author2=Michael Tauber
| author3=William Henline
| author4=Duoc Tran
| author5=Alan Cartledge
| author6=Frank Hui
| author7=Norm Zimmerman
| title=Ames Research Center Shear Tests of SLA-561V Heat Shield Material for Mars-Pathfinder
| number=NASA Technical Memorandum 110402
| institution=NASA Ames Research Center
| year=1996
| url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19960049758_1996080506.pdf
| access-date=July 7, 2017
| archive-date=September 25, 2020
| archive-url=https://web.archive.org/web/20200925143246/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19960049758_1996080506.pdf
| url-status=live
}}</ref>

====Phenolic-impregnated carbon ablator====
[[File:OSIRIS-REx Sample Return (NHQ202309240002).jpg|thumb|OSIRIS-REx Sample Return Capsule at USAF Utah Range.|left]]
''Phenolic-impregnated carbon ablator'' (PICA), a [[carbon fiber]] preform impregnated in [[phenolic resin]],<ref>{{cite conference |url=http://www.jeanlachaud.com/research/Lachaud2010_AbstractOFWS5.pdf |first1=Jean |last1=Lachaud |first2=Nagi |last2=N. Mansour |title=A pyrolysis and ablation toolbox based on OpenFOAM |conference=5th OpenFOAM Workshop |place=Gothenburg, Sweden |date=June 2010 |page=1 |access-date=August 9, 2012 |archive-date=September 12, 2012 |archive-url=https://web.archive.org/web/20120912052713/http://www.jeanlachaud.com/research/Lachaud2010_AbstractOFWS5.pdf |url-status=live }}</ref> is a modern TPS material and has the advantages of low density (much lighter than carbon phenolic) coupled with efficient ablative ability at high heat flux. It is a good choice for ablative applications such as high-peak-heating conditions found on sample-return missions or lunar-return missions. PICA's thermal conductivity is lower than other high-heat-flux-ablative materials, such as conventional carbon phenolics.{{Citation needed|date=February 2009}}

PICA was patented by [[NASA Ames Research Center]] in the 1990s and was the primary TPS material for the [[Stardust (spacecraft)|Stardust]] aeroshell.<ref>Tran, Huy K, et al., "Qualification of the forebody heat shield of the Stardust's Sample Return Capsule", AIAA, Thermophysics Conference, 32nd, Atlanta, GA; 23–25 June 1997.</ref> The Stardust sample-return capsule was the fastest man-made object ever to reenter Earth's atmosphere, at 28,000&nbsp;mph (ca. 12.5&nbsp;km/s) at 135&nbsp;km altitude. This was faster than the Apollo mission capsules and 70% faster than the Shuttle.<ref name=stardust>{{cite web|url=http://stardust.jpl.nasa.gov/cool.html|title=Stardust – Cool Facts|website=stardust.jpl.nasa.gov|access-date=January 9, 2010|archive-date=January 12, 2010|archive-url=https://web.archive.org/web/20100112063823/http://stardust.jpl.nasa.gov/cool.html|url-status=live}}</ref> PICA was critical for the viability of the Stardust mission, which returned to Earth in 2006. Stardust's heat shield (0.81&nbsp;m base diameter) was made of one monolithic piece sized to withstand a nominal peak heating rate of 1.2&nbsp;kW/cm<sup>2</sup>. A PICA heat shield was also used for the [[Mars Science Laboratory]] entry into the [[Martian atmosphere]].<ref name="N+SX_picaX"/>

=====PICA-X=====
An improved and easier to produce version called PICA-X was developed by [[SpaceX]] in 2006–2010<ref name="N+SX_picaX"/> for the [[SpaceX Dragon 1|Dragon]] [[space capsule]].<ref name=srdc20090223>{{cite web|url=http://www.spaceref.com/news/viewpr.html?pid=27612|title=SpaceX Manufactured Heat Shield Material Passes High Temperature Tests Simulating Reentry Heating Conditions of Dragon Spacecraft|website=www.spaceref.com|date=February 23, 2009 }}</ref> The first reentry test of a PICA-X heat shield was on the [[Dragon C1]] mission on 8 December 2010.<ref name=clog20101208>[https://web.archive.org/web/20101211200945/http://cosmiclog.msnbc.msn.com/_news/2010/12/08/5614525-dragon-could-visit-space-station-next Dragon could visit space station next], ''[[msnbc.com]]'', 2010-12-08, accessed 2010-12-09.</ref> The PICA-X heat shield was designed, developed and fully qualified by a small team of a dozen engineers and technicians in less than four years.<ref name="N+SX_picaX">
{{cite web
 |last=Chambers 
 |first=Andrew 
 |title=NASA + SpaceX Work Together 
 |url=http://www.nasa.gov/offices/oce/appel/ask/issues/40/40s_space-x_prt.htm
 |publisher=NASA 
 |access-date=2011-02-16 
 |author2=Dan Rasky 
 |date=2010-11-14 
 |quote=''SpaceX undertook the design and manufacture of the reentry heat shield; it brought speed and efficiency that allowed the heat shield to be designed, developed, and qualified in less than four years.''' 
 |url-status=dead 
 |archive-url=https://web.archive.org/web/20110416170908/http://www.nasa.gov/offices/oce/appel/ask/issues/40/40s_space-x_prt.htm 
 |archive-date=2011-04-16 
}}</ref>
PICA-X is ten times less expensive to manufacture than the NASA PICA heat shield material.<ref name="a&s201201">{{cite news
| last=Chaikin
| first=Andrew
| title=1 visionary + 3 launchers + 1,500 employees = ? : Is SpaceX changing the rocket equation?
| url=http://www.airspacemag.com/space/is-spacex-changing-the-rocket-equation-132285884/?page=2
| access-date=2016-06-03
| newspaper=Air & Space Smithsonian
| date=January 2012
| quote=''SpaceX's material, called PICA-X, is one-tenth as expensive than the original [NASA PICA material and is better], ... a single PICA-X heat shield could withstand hundreds of returns from low Earth orbit; it can also handle the much higher energy reentries from the Moon or Mars.''
| archive-date=September 7, 2018
| archive-url=https://web.archive.org/web/20180907221220/https://www.airspacemag.com/space/is-spacex-changing-the-rocket-equation-132285884/?page=2
| url-status=live
}}</ref>

=====PICA-3=====
A second enhanced version of PICA—called PICA-3—was developed by SpaceX during the mid-2010s.  It was first flight tested on the [[Crew Dragon]] spacecraft in 2019 during the [[Crew Dragon Demo-1|flight demonstration mission]], in April 2019, and put into regular service on that spacecraft in 2020.<ref>[https://www.nasa.gov/multimedia/nasatv/#public NASA TV broadcast for the Crew Dragon Demo-2 mission departure from the ISS] {{Webarchive|url=https://web.archive.org/web/20200802031316/https://www.nasa.gov/multimedia/nasatv/#public |date=August 2, 2020 }}, NASA, 1 August 2020.</ref>

===== 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====
[[File:Ds 2.jpg|thumb|[[Deep Space 2]] [[Lander (spacecraft)|impactor]] aeroshell, a classic 45° sphere-cone with spherical section afterbody, enabling aerodynamic stability from atmospheric entry to surface impact]]
Silicone-impregnated reusable ceramic ablator (SIRCA) was also developed at NASA Ames Research Center and was used on the Backshell Interface Plate (BIP) of the ''[[Mars Pathfinder]]'' and [[Mars Exploration Rover]] (MER) aeroshells. The BIP was at the attachment points between the aeroshell's backshell (also called the afterbody or aft cover) and the cruise ring (also called the cruise stage). SIRCA was also the primary TPS material for the unsuccessful [[Deep Space 2]] (DS/2) Mars [[Lander (spacecraft)|impactor]] probes with their {{Convert|0.35|m|ft|-base-diameter|adj=mid|sp=us}} aeroshells. SIRCA is a monolithic, insulating material that can provide thermal protection through ablation. It is the only TPS material that can be machined to custom shapes and then applied directly to the spacecraft. There is no post-processing, heat treating, or additional coatings required (unlike Space Shuttle tiles). Since SIRCA can be machined to precise shapes, it can be applied as tiles, leading edge sections, full nose caps, or in any number of custom shapes or sizes. {{as of|1996}}, SIRCA had been demonstrated in backshell interface applications, but not yet as a forebody TPS material.<ref>Tran, Huy K., et al., "Silicone impregnated reusable ceramic ablators for Mars follow-on missions," AIAA-1996-1819, Thermophysics Conference, 31st, New Orleans, June 17–20, 1996.</ref>

====AVCOAT====
[[AVCOAT 5026-39|AVCOAT]] is a [[NASA]]-specified ablative heat shield, a glass-filled [[epoxy]]–[[novolac]] system.<ref name=nasa196808>[https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19680021275_1968021275.pdf Flight-Test Analysis Of Apollo Heat-Shield Material Using The Pacemaker Vehicle System] {{Webarchive|url=https://web.archive.org/web/20200925144304/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19680021275_1968021275.pdf |date=September 25, 2020 }} [[NASA]] Technical Note D-4713, pp. 8, 1968–08, accessed 2010-12-26. ''"Avcoat 5026-39/HC-G is an epoxy novolac resin with special additives in a fiberglass honeycomb matrix. In fabrication, the empty honeycomb is bonded to the primary structure and the resin is gunned into each cell individually. ... The overall density of the material is 32 lb/ft3 (512 kg/m3). The char of the material is composed mainly of silica and carbon. It is necessary to know the amounts of each in the char because in the ablation analysis the silica is considered to be inert, but the carbon is considered to enter into exothermic reactions with oxygen. ... At 2160O R (12000 K), 54 percent by weight of the virgin material has volatilized and 46 percent has remained as char. ... In the virgin material, 25 percent by weight is silica, and since the silica is considered to be inert the char-layer composition becomes 6.7 lb/ft3 (107.4 kg/m3) of carbon and 8 lb/ft3 (128.1 kg/m3) of silica."''</ref><!-- the AVCOAT article claims this material was used on the Apollo command module, but no citation is provided so the claim is not [[WP:V|verified]]. -->

NASA originally used it for the [[Apollo command and service module#Command Module (CM)|Apollo command module]] in the 1960s, and then utilized the material for its next-generation beyond low Earth orbit [[Orion (spacecraft)|Orion]] crew module, which first flew in a December 2014 test and then operationally in November 2022.<ref name=nasa20090407>[http://www.nasa.gov/home/hqnews/2009/apr/HQ_09-080_Orion_Heat_Shield.html NASA.gov NASA Selects Material for Orion Spacecraft Heat Shield] {{Webarchive|url=https://web.archive.org/web/20101124220318/http://www.nasa.gov/home/hqnews/2009/apr/HQ_09-080_Orion_Heat_Shield.html |date=November 24, 2010 }}, 2009-04-07, accessed 2011-01-02.</ref> The Avcoat to be used on Orion has been reformulated to meet environmental legislation that has been passed since the end of Apollo.<ref name=fg20090310>{{Cite web|url=http://www.flightglobal.com/articles/2009/03/10/323585/nasas-orion-heat-shield-decision-expected-this-month.html|title=Flightglobal.com NASA's Orion heat shield decision expected this month 2009-10-03, accessed 2011-01-02|access-date=January 2, 2011|archive-date=March 24, 2009|archive-url=https://web.archive.org/web/20090324160130/http://www.flightglobal.com/articles/2009/03/10/323585/nasas-orion-heat-shield-decision-expected-this-month.html|url-status=live}}</ref><ref>{{cite web|url=http://www.thefreelibrary.com/Company+Watch+-+NASA-a0198584187|title=Company Watch – NASA. – Free Online Library|website=www.thefreelibrary.com|access-date=January 2, 2011|archive-date=October 22, 2012|archive-url=https://web.archive.org/web/20121022003103/http://www.thefreelibrary.com/Company+Watch+-+NASA-a0198584187|url-status=live}}</ref>

===Thermal soak===
Thermal soak is a part of almost all TPS schemes. For example, an ablative heat shield loses most of its thermal protection effectiveness when the outer wall temperature drops below the minimum necessary for pyrolysis. From that time to the end of the heat pulse, heat from the shock layer convects into the heat shield's outer wall and would eventually conduct to the payload.{{Citation needed|date=February 2009}} This outcome can be prevented by ejecting the heat shield (with its heat soak) prior to the heat conducting to the inner wall.

===Refractory insulation===
[[File:Shuttle heat shield.jpg|thumb|left|Astronaut [[Andy Thomas|Andrew S. W. Thomas]] takes a close look at TPS tiles underneath [[Space Shuttle Atlantis|Space Shuttle ''Atlantis'']].]]
[[File:Silica Space Shuttle thermal protection (TPS) tile, c 1980. (9663807484).jpg|thumb|right|Rigid black [[LI-900]] tiles were used on the [[Space Shuttle]].]]

Refractory insulation keeps the heat in the outermost layer of the spacecraft surface, where it is conducted away by the air.<ref name=TPSpaper>{{cite conference
| url = https://ntrs.nasa.gov/citations/20160001151
| title = Thermal Protection Systems: Past, Present and Future
| last1 = Johnson
| first1 = Sylvia M.
| date = January 25, 2015
| publisher = 
| book-title = 
| pages = 
| location = International Conference and Exposition on Advanced Ceramics and Composites (Daytona Beach, FL)
| id = ARC-E-DAA-TN29151
| conference = 
| access-date = September 5, 2021
| archive-date = September 5, 2021
| 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)|Starship's upper stage]].

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/>

Typical [[Space Shuttle thermal protection system|Space Shuttle TPS]] tiles ([[LI-900]]) have remarkable thermal protection properties. An LI-900 tile exposed to a temperature of 1,000 K on one side will remain merely warm to the touch on the other side. However, they are relatively brittle and break easily, and cannot survive in-flight rain.<ref>https://ntrs.nasa.gov/api/citations/19880011857/downloads/19880011857.pdf Damage testing on effect of Rain by Robert R. Meyer and Jack Barneburg</ref>

===Passively cooled===
[[File:Mercury capsule HD.jpg|thumb|upright|left|The Mercury capsule design (shown here with its [[Launch escape system|escape tower]]) originally used a radiatively cooled TPS, but was later converted to an ablative TPS.]]
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>

Some high-velocity [[aircraft]], such as the [[SR-71 Blackbird]] and [[Concorde]], deal with heating similar to that experienced by spacecraft, but at much lower intensity, and for hours at a time. Studies of the SR-71's titanium skin revealed that the metal structure was restored to its original strength through [[annealing (metallurgy)|annealing]] due to aerodynamic heating. In the case of the Concorde, the [[aluminium]] nose was permitted to reach a maximum [[operating temperature]] of {{convert|127|°C|°F}} (approximately {{convert|180|C-change|F-change|0}} warmer than the normally sub-zero, ambient air); the metallurgical implications (loss of [[tempering (metallurgy)|temper]]) that would be associated with a higher peak temperature were the most significant factors determining the top speed of the aircraft.

A radiatively cooled TPS for an entry vehicle is often called a ''hot-metal TPS''. Early TPS designs for the Space Shuttle called for a hot-metal TPS based upon a nickel [[superalloy]] (dubbed [[René 41]]) and titanium shingles.<ref name="auto">{{Cite web|url=http://www.astronautix.com/s/spaceshuttle.html|title=Space Shuttle|website=www.astronautix.com|access-date=April 22, 2022|archive-date=March 18, 2022|archive-url=https://web.archive.org/web/20220318082348/http://www.astronautix.com/s/spaceshuttle.html|url-status=dead}}</ref> This Shuttle TPS concept was rejected, because it was believed a silica tile-based TPS would involve lower development and manufacturing costs.{{Citation needed|date=July 2010}} A nickel [[superalloy]]-shingle TPS was again proposed for the unsuccessful [[X-33]] [[single-stage-to-orbit]] (SSTO) prototype.<ref>{{Cite web |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040095922_2004100223.pdf |title=X-33 Heat Shield Development report |access-date=July 7, 2017 |archive-date=January 26, 2021 |archive-url=https://web.archive.org/web/20210126222704/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040095922_2004100223.pdf |url-status=live }}</ref>

Recently, newer radiatively cooled TPS materials have been developed that could be superior to RCC. Known as [[Ultra high temperature ceramic matrix composite|Ultra-High Temperature Ceramics]], they were developed for the prototype vehicle Slender Hypervelocity Aerothermodynamic Research Probe (SHARP). These TPS materials are based on [[zirconium diboride]] and [[hafnium diboride]]. SHARP TPS have suggested performance improvements allowing for sustained [[Mach number|Mach]] 7 flight at sea level, Mach 11 flight at {{convert|100000|ft|adj=on}} altitudes, and significant improvements for vehicles designed for continuous hypersonic flight. SHARP TPS materials enable sharp leading edges and nose cones to greatly reduce drag for airbreathing combined-cycle-propelled spaceplanes and lifting bodies. SHARP materials have exhibited effective TPS characteristics from zero to more than {{Convert|2000|C}}, with melting points over {{Convert|3500|C}}. They are structurally stronger than RCC, and, thus, do not require structural reinforcement with materials such as Inconel. SHARP materials are extremely efficient at reradiating absorbed heat, thus eliminating the need for additional TPS behind and between the SHARP materials and conventional vehicle structure. NASA initially funded (and discontinued) a multi-phase R&D program through the [[University of Montana]] in 2001 to test SHARP materials on test vehicles.<ref>{{cite web|url=http://hubbard.engr.scu.edu/docs/thesis/2003/SHARP_Thesis.pdf |title=SHARP Reentry Vehicle Prototype |access-date=2006-04-09 |url-status=dead |archive-url=https://web.archive.org/web/20051215231157/http://hubbard.engr.scu.edu/docs/thesis/2003/SHARP_Thesis.pdf |archive-date=2005-12-15 }}</ref><ref>{{Cite web|url=http://www.coe.montana.edu/me/faculty/cairns/sharp/sharp.htm|archive-url=https://web.archive.org/web/20151016071845/http://www.coe.montana.edu/me/faculty/cairns/sharp/sharp.htm |url-status=dead |title=sharp structure homepage w left<!-- Bot generated title -->|archive-date=October 16, 2015}}</ref>

===Actively cooled===
Various advanced reusable spacecraft and hypersonic aircraft designs have been proposed to employ heat shields made from temperature-resistant metal [[alloy]]s that incorporate a refrigerant or cryogenic fuel circulating through them.

Such a TPS concept was proposed for the [[Rockwell X-30|X-30 National Aerospace Plane]] (NASP) in the mid-80s.{{cn|date=January 2024}} The NASP was supposed to have been a [[scramjet]] powered hypersonic aircraft, but failed in development.{{cn|date=January 2024}}

In 2005 and 2012, two unmanned [[lifting body]] craft with actively cooled hulls were launched as a part of the German [[Sharp Edge Flight Experiment]] (SHEFEX).{{cn|date=January 2024}}

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.

===Propulsive entry===

Fuel permitting, nothing prevents a vehicle from entering the atmosphere with a retrograde engine burn, which has the double effect of slowing the vehicle down much faster than atmospheric drag alone would, and forcing the compressed hot air away from the vehicle's body.  During reentry, the first stage of the SpaceX [[Falcon 9]] performs an entry burn to rapidly decelerate from its initial hypersonic speed.{{Citation needed|date=June 2021}}