View source for Astronomy:Atmospheric entry

{{Short description|Passage of an object through the gases of an atmosphere from outer space}}
[[File:Entry.jpg|thumb|Mars Exploration Rover (MER) aeroshell, artistic rendition]]
'''Atmospheric entry''' (sometimes listed as V<sub>impact</sub> or V<sub>entry</sub>) is the movement of an object from [[Astronomy:Outer space|outer space]] into and through the gases of an [[Astronomy:Atmosphere|atmosphere]] of a [[Astronomy:Planet|planet]], [[Astronomy:Dwarf planet|dwarf planet]], or [[Astronomy:Natural satellite|natural satellite]]. There are two main types of atmospheric entry: ''uncontrolled entry'', such as the entry of [[Astronomy:Astronomical object|astronomical object]]s, [[Astronomy:Space debris|space debris]], or [[Astronomy:Bolide|bolide]]s; and ''controlled entry'' (or ''reentry'') of a spacecraft capable of being navigated or following a predetermined course. Technologies and procedures allowing the controlled atmospheric ''entry, descent, and landing'' of spacecraft are collectively termed as ''EDL''.

[[File:Meteoroid meteor meteorite.gif|Animated illustration of different phases as a [[Astronomy:Meteoroid|meteoroid]] enters the Earth's atmosphere to become visible as a meteor and land as a meteorite|thumb|right]]
[[File:Reentry video as Orion returns from Artemis I.webm|thumb|Video of Orion's reentry on Artemis 1, showing the entire reentry process unedited from space to splashdown]]
Objects entering an atmosphere experience [[Physics:Drag|atmospheric drag]], which puts mechanical stress on the object, and [[Astronomy:Aerodynamic heating|aerodynamic heating]]—caused mostly by compression of the air in front of the object, but also by drag. These forces can cause loss of mass ([[Physics:Ablation|ablation]]) or even complete disintegration of smaller objects, and objects with lower [[Engineering:Compressive strength|compressive strength]] can explode.

Reentry has been achieved with speeds ranging from 7.8&nbsp;km/s for [[Astronomy:Low Earth orbit|low Earth orbit]] to around 12.5&nbsp;km/s for the [[Engineering:Stardust (spacecraft)|Stardust]] probe.<ref name=stardust/> Crewed space vehicles must be slowed to subsonic speeds before parachutes or air brakes may be deployed. Such vehicles have high kinetic energies and atmospheric dissipation is the only way of expending this, as it is highly impractical to use [[Engineering:Retrorocket|retrorocket]]s for the entire reentry procedure.

Ballistic warheads and expendable vehicles do not require slowing at reentry, and in fact, are made streamlined so as to maintain their speed. Furthermore, slow-speed returns to Earth from near-space such as high-altitude parachute jumps from balloons do not require heat shielding because the gravitational acceleration of an object starting at relative rest from within the atmosphere itself (or not far above it) cannot create enough velocity to cause significant atmospheric heating.

For Earth, atmospheric entry occurs by convention at the [[Earth:Kármán line|Kármán line]] at an altitude of {{Convert|100|km|mi nmi|abbr=in}} above the surface, while at [[Astronomy:Venus#Atmospheric entry|Venus atmospheric entry]] occurs at {{Convert|250|km|mi nmi|abbr=on}} and at [[Astronomy:Mars atmospheric entry|Mars atmospheric entry]] at about {{Convert|80|km|mi nmi|abbr=on}}. Uncontrolled objects reach high velocities while accelerating through space toward the Earth under the influence of Earth's [[Physics:Gravity|gravity]], and are slowed by friction upon encountering Earth's atmosphere. Meteors are also often travelling quite fast relative to the Earth simply because their own orbital path is different from that of the Earth before they encounter Earth's [[Physics:Gravity well|gravity well]]. Most objects enter at [[Physics:Hypersonic speed|hypersonic speeds]] due to their [[Engineering:Sub-orbital spaceflight|sub-orbital]] (e.g., [[Engineering:Intercontinental ballistic missile|intercontinental ballistic missile]] reentry vehicles), [[Physics:Orbit|orbital]] (e.g., the [[Engineering:Soyuz (spacecraft)|Soyuz]]), or [[Astronomy:Hyperbolic trajectory|unbounded]] (e.g., meteors) trajectories. Various advanced technologies have been developed to enable atmospheric reentry and flight at extreme velocities. An alternative method of controlled atmospheric entry is [[Physics:Buoyancy|buoyancy]]<ref>{{cite web |url=http://www.jpaerospace.com/atohandout.pdf |title=ATO: Airship To Orbit |publisher=JP Aerospace |access-date=December 14, 2013 |archive-date=October 13, 2013 |archive-url=https://web.archive.org/web/20131013190704/http://jpaerospace.com/atohandout.pdf |url-status=live }}</ref> which is suitable for planetary entry where thick atmospheres, strong gravity, or both factors complicate high-velocity hyperbolic entry, such as the atmospheres of [[Astronomy:Venus|Venus]], [[Astronomy:Titan (moon)|Titan]] and the [[Astronomy:Gas giant|gas giant]]s.<ref>{{cite book|title=Unmanned Spacecraft Meeting 1965|first=F.|last=Gross|publisher=American Institute of Aeronautics and Astronautics|doi=10.2514/6.1965-1407|chapter=Buoyant Probes into the Venus Atmosphere|year=1965}}</ref>

==History==
[[File:Blunt body reentry shapes.png|thumb|Early reentry-vehicle concepts visualized in [[Schlieren photography|shadowgraph]]s of high speed wind tunnel tests]]
The concept of the ablative [[Astronomy:Heat shield|heat shield]] was described as early as 1920 by Robert Goddard: "In the case of meteors, which enter the atmosphere with speeds as high as {{convert|30|mi}} per second, the interior of the meteors remains cold, and the erosion is due, to a large extent, to chipping or cracking of the suddenly heated surface. For this reason, if the outer surface of the apparatus were to consist of layers of a very infusible hard substance with layers of a poor heat conductor between, the surface would not be eroded to any considerable extent, especially as the velocity of the apparatus would not be nearly so great as that of the average meteor."<ref>{{cite web|last=Goddard|first=Robert H.|title=Report Concerning Further Developments|date=Mar 1920|url= http://siarchives.si.edu/history/exhibits/documents/goddardmarch1920.htm|publisher=The Smithsonian Institution Archives|access-date= 2009-06-29| archive-url= https://web.archive.org/web/20090626145252/http://siarchives.si.edu/history/exhibits/documents/goddardmarch1920.htm| archive-date= 26 June 2009 | url-status= live}}</ref>

Practical development of reentry systems began as the range, and reentry velocity of ballistic missiles increased. For early short-range missiles, like the [[Engineering:V-2 rocket|V-2]], stabilization and aerodynamic stress were important issues (many V-2s broke apart during reentry), but heating was not a serious problem. Medium-range missiles like the Soviet R-5, with a {{convert|1200|km|nmi|sp=us|adj=on|abbr=off}} range, required ceramic composite heat shielding on separable reentry vehicles (it was no longer possible for the entire rocket structure to survive reentry). The first ICBMs, with ranges of {{convert|8000|to|12000|km|nmi|sp=us|abbr=on}}, were only possible with the development of modern ablative heat shields and blunt-shaped vehicles.

In the United States, this technology was pioneered by H. Julian Allen and A. J. Eggers Jr. of the [[Organization:National Advisory Committee for Aeronautics|National Advisory Committee for Aeronautics]] (NACA) at [[Organization:Ames Research Center|Ames Research Center]].<ref>Boris Chertok, "Rockets and People", NASA History Series, 2006</ref> In 1951, they made the counterintuitive discovery that a blunt shape (high drag) made the most effective heat shield.<ref>{{cite book |url=https://history.nasa.gov/SP-4305/sp4305.htm |last=Hansen |first=James R. |date=Jun 1987 |title=Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917–1958 |series=The NASA History Series |volume=sp-4305 |chapter=Chapter 12: Hypersonics and the Transition to Space |chapter-url=https://history.nasa.gov/SP-4305/ch12.htm |publisher=United States Government Printing |isbn=978-0-318-23455-7 |access-date=July 12, 2017 |archive-date=July 14, 2019 |archive-url=https://web.archive.org/web/20190714121746/https://history.nasa.gov/SP-4305/sp4305.htm |url-status=live }}</ref> From simple engineering principles, Allen and Eggers showed that the heat load experienced by an entry vehicle was inversely proportional to the [[Physics:Drag coefficient|drag coefficient]]; i.e., the greater the drag, the less the heat load. If the reentry vehicle is made blunt, air cannot "get out of the way" quickly enough, and acts as an air cushion to push the shock wave and heated shock layer forward (away from the vehicle). Since most of the hot gases are no longer in direct contact with the vehicle, the heat energy would stay in the shocked gas and simply move around the vehicle to later dissipate into the atmosphere.

The Allen and Eggers discovery, though initially treated as a military secret, was eventually published in 1958.<ref>
{{cite journal
 | url=http://naca.central.cranfield.ac.uk/reports/1958/naca-report-1381.pdf
 | last1=Allen | first1=H. Julian
 | last2=Eggers | first2=A. J. Jr.
 | title=A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth's Atmosphere at High Supersonic Speeds
 | journal=NACA Annual Report | issue=NACA-TR-1381
 | volume=44.2 | pages=1125–1140 | year=1958
 | publisher=NASA Technical Reports
 | url-status=dead
 | archive-url=https://web.archive.org/web/20151013182336/http://naca.central.cranfield.ac.uk/reports/1958/naca-report-1381.pdf
 | archive-date=October 13, 2015
}}</ref>

==Terminology, definitions and jargon==
{{Hatnote|Over the decades since the 1950s, a rich technical jargon has grown around the engineering of vehicles designed to enter planetary atmospheres. It is recommended that the reader review the jargon glossary before continuing with this article on atmospheric reentry.}}

When atmospheric entry is part of a spacecraft landing or recovery, particularly on a planetary body other than Earth, entry is part of a phase referred to as ''entry, descent, and landing'', or EDL.<ref>{{Cite web|url=http://www.nasa.gov/pdf/501326main_TA09-EDL-DRAFT-Nov2010-A.pdf|title=NASA.gov|access-date=April 9, 2015|archive-date=February 20, 2017|archive-url=https://web.archive.org/web/20170220165240/https://www.nasa.gov/pdf/501326main_TA09-EDL-DRAFT-Nov2010-A.pdf|url-status=live}}</ref> When the atmospheric entry returns to the same body that the vehicle had launched from, the event is referred to as '''reentry''' (almost always referring to Earth entry).

The fundamental design objective in atmospheric entry of a spacecraft is to [[Dissipation|dissipate]] the <!-- kinetic plus potential -->[[Physics:Mechanical energy|energy]] of a spacecraft that is traveling at [[Physics:Hypersonic speed|hypersonic speed]] as it enters an atmosphere such that equipment, cargo, and any passengers<!-- using "passengers" rather than "crew" as, increasingly after 2020, we are seeing a transition in spacecraft designs for vastly more passengers than crew --> are slowed and land near a specific destination on the surface at zero velocity while keeping stresses on the spacecraft and any passengers within acceptable limits.<ref name=nasa197203>{{cite conference |last1=Graves |first1=Claude A. |last2=Harpold |first2=Jon C. |title=Apollo Experience Report – Mission Planning for Apollo Entry |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19720013191.pdf |conference=NASA Technical Note (TN) D-6725 |date=March 1972 |quote=The purpose of the Apollo entry maneuver is to dissipate the energy of a spacecraft traveling at high speed through the atmosphere of the earth so that the flight crew, their equipment, and their cargo are returned safely to a preselected location on the surface of the earth. This purpose must be accomplished while stresses on both the spacecraft and the flight crew are maintained within acceptable limits.|access-date=October 25, 2019 |archive-date=October 25, 2019 |archive-url=https://web.archive.org/web/20191025121155/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19720013191.pdf |url-status=live }}</ref> This may be accomplished by [[Engineering:Propulsion|propulsive]] or [[Physics:Drag|aerodynamic]] (vehicle characteristics or [[Engineering:Parachute|parachute]]) means, or by some combination.

==Entry vehicle shapes==
{{Main|Engineering:Aeroshell}}
There are several basic shapes used in designing entry vehicles:

===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.<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 [[Physics:Angle of attack|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|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 [[Astronomy:Vostok programme|Vostok]] and [[Astronomy:Voskhod programme|Voskhod]] [[Engineering:Space capsule|capsule]]s and in Soviet Mars and [[Astronomy:Venera|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 longitudinal axis. Other examples of the spherical section geometry in crewed capsules are [[Engineering:Soyuz programme|Soyuz]]/[[Astronomy:Zond program|Zond]], [[Astronomy:Project Gemini|Gemini]], and [[Astronomy:Project Mercury|Mercury]]. Even these small amounts of lift allow trajectories that have very significant effects on peak [[Physics:G-force|g-force]], reducing it from 8–9&nbsp;g for a purely ballistic (slowed only by drag) trajectory to 4–5&nbsp;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>

===Sphere-cone===
The sphere-cone is a spherical section with a [[Frustum|frustum]] or blunted cone attached. The sphere-cone's dynamic stability is typically better than that of a spherical section. The vehicle enters sphere-first. With a sufficiently small half-angle and properly placed center of mass, a sphere-cone can provide aerodynamic stability from Keplerian entry to surface impact. (The ''half-angle'' is the angle between the cone's axis of rotational symmetry and its outer surface, and thus half the angle made by the cone's surface edges.)

[[File:Mk 2.jpg|thumb|left|Prototype of the Mk-2 Reentry Vehicle (RV), based on blunt body theory]]
The original American sphere-cone aeroshell was the Mk-2 RV (reentry vehicle), which was developed in 1955 by the [[Company:General Electric|General Electric]] Corp. The Mk-2's design was derived from blunt-body theory and used a radiatively cooled thermal protection system (TPS) based upon a metallic heat shield (the different TPS types are later described in this article). The Mk-2 had significant defects as a weapon delivery system, i.e., it loitered too long in the upper atmosphere due to its lower [[Physics:Ballistic coefficient|ballistic coefficient]] and also trailed a stream of vaporized metal making it very visible to [[Engineering:Radar|radar]]. These defects made the Mk-2 overly susceptible to anti-ballistic missile (ABM) systems. Consequently, an alternative sphere-cone RV to the Mk-2 was developed by General Electric.{{citation needed|date=November 2013}}

[[File:Mk 6 reentry vehicle on display at National Atomic Museum.jpg|thumb|upright|right|Mk-6 RV, Cold War weapon and ancestor to most of the U.S. missile entry vehicles]]
This new RV was the Mk-6 which used a non-metallic ablative TPS, a nylon phenolic. This new TPS was so effective as a reentry heat shield that significantly reduced bluntness was possible.{{citation needed|date=November 2013}} However, the Mk-6 was a huge RV with an entry mass of 3,360&nbsp;kg, a length of 3.1 m and a half-angle of 12.5°. Subsequent advances in nuclear weapon and ablative TPS design allowed RVs to become significantly smaller with a further reduced bluntness ratio compared to the Mk-6. Since the 1960s, the sphere-cone has become the preferred geometry for modern ICBM RVs with typical half-angles being between 10° and 11°.{{citation needed|date=November 2013}}

[[File:Rv film pod.jpg|thumb|upright|left|"Discoverer" type reconnaissance satellite film Recovery Vehicle (RV)]]
[[File:Galileo probe.jpg|thumb|right|Galileo Probe during final assembly]]
[[Astronomy:Reconnaissance satellite|Reconnaissance satellite]] RVs (recovery vehicles) also used a sphere-cone shape and were the first American example of a non-munition entry vehicle ([[Engineering:Corona (satellite)|Discoverer-I]], launched on 28 February 1959). The sphere-cone was later used for space exploration missions to other celestial bodies or for return from open space; e.g., [[Engineering:Stardust (spacecraft)|''Stardust'']] probe. Unlike with military RVs, the advantage of the blunt body's lower TPS mass remained with space exploration entry vehicles like the [[Astronomy:Galileo Probe|Galileo Probe]] with a half-angle of 45° or the [[Engineering:Viking program|Viking aeroshell]] with a half-angle of 70°. Space exploration sphere-cone entry vehicles have landed on the surface or entered the atmospheres of [[Astronomy:Mars|Mars]], [[Astronomy:Venus|Venus]], [[Astronomy:Jupiter|Jupiter]], and [[Astronomy:Titan (moon)|Titan]].
{{clear}}

===Biconic===
[[File:Delta Clipper DC-X first flight.jpg|thumb|right|upright|The [[Engineering:McDonnell Douglas DC-X|DC-X]], shown during its first flight, was a prototype single-stage-to-orbit vehicle, and used a biconic shape similar to AMaRV.]]
The biconic is a sphere-cone with an additional frustum attached. The biconic offers a significantly improved L/D ratio. A biconic designed for Mars aerocapture typically has an L/D of approximately 1.0 compared to an L/D of 0.368 for the Apollo-CM. The higher L/D makes a biconic shape better suited for transporting people to Mars due to the lower peak deceleration. Arguably, the most significant biconic ever flown was the ''Advanced [[Engineering:Maneuverable reentry vehicle|Maneuverable Reentry Vehicle]]'' (AMaRV). Four AMaRVs were made by the [[Company:McDonnell Douglas|McDonnell Douglas]] Corp. and represented a significant leap in RV sophistication. Three AMaRVs were launched by [[LGM-30 Minuteman|Minuteman-1 ICBMs]] on 20 December 1979, 8 October 1980 and 4 October 1981. AMaRV had an entry mass of approximately 470&nbsp;kg, a nose radius of 2.34&nbsp;cm, a forward-frustum half-angle of 10.4°, an inter-frustum radius of 14.6&nbsp;cm, aft-frustum half-angle of 6°, and an axial length of 2.079 meters. No accurate diagram or picture of AMaRV has ever appeared in the open literature. However, a schematic sketch of an AMaRV-like vehicle along with trajectory plots showing hairpin turns has been published.<ref>Regan, Frank J. and Anadakrishnan, Satya M., "Dynamics of Atmospheric Re-Entry", AIAA Education Series, American Institute of Aeronautics and Astronautics, Inc., New York, {{ISBN|1-56347-048-9}}, (1993).<!-- I think this refers to diagrams of simulations on pg. 268 & 269. Is there a way to note that? ([https://books.google.com/books?as_isbn=1563470489 Google Books]) --></ref>

AMaRV's attitude was controlled through a split body flap (also called a ''split-windward flap'') along with two yaw flaps mounted on the vehicle's sides. [[Physics:Hydraulic machinery|Hydraulic actuation]] was used for controlling the flaps. AMaRV was guided by a fully autonomous navigation system designed for evading [[Engineering:Anti-ballistic missile|anti-ballistic missile]] (ABM) interception. The [[Engineering:McDonnell Douglas DC-X|McDonnell Douglas DC-X]] (also a biconic) was essentially a scaled-up version of AMaRV. AMaRV and the DC-X also served as the basis for an unsuccessful proposal for what eventually became the [[Engineering:Lockheed Martin X-33|Lockheed Martin X-33]].

===Non-axisymmetric shapes===
Non-axisymmetric shapes have been used for crewed entry vehicles. One example is the winged orbit vehicle that uses a [[Engineering:Delta wing|delta wing]] for maneuvering during descent much like a conventional glider. This approach has been used by the American [[Engineering:Space Shuttle|Space Shuttle]] and the Soviet [[Engineering:Buran (spacecraft)|Buran]]. The [[Engineering:Lifting body|lifting body]] is another entry vehicle geometry and was used with the X-23 PRIME (Precision Recovery Including Maneuvering Entry) vehicle.{{citation needed|date=November 2013}}

==Entry heating==
[[File:Gemini2reentry.jpg|thumb|View of plasma trail of Gemini 2 reentry]]
Objects entering an atmosphere from [[Astronomy:Outer space|space]] at high velocities relative to the atmosphere will cause very high levels of [[Physics:Heat flux|heating]]. Atmospheric entry heating comes principally from two sources:
  
* [[Physics:Convection (heat transfer)|convection]] of hot gas flow past the surface of the body and catalytic chemical recombination reactions between the surface and atmospheric gases; and 
*[[Physics:Thermal radiation|radiation]] from the energetic [[Physics:Shock wave|shock layer]] that forms in the front and sides of the body<ref name=nasa201401/>

As velocity increases, both convective and radiative heating increase, but at different rates. At very high speeds, radiative heating will dominate the convective heat fluxes, as radiative heating is proportional to the eighth power of velocity, while convective heating is proportional to the third power of velocity. Radiative heating thus predominates early in atmospheric entry, while convection predominates in the later phases.<ref name=nasa201401>{{cite conference |title=Biologically-Derived Photonic Materials for Thermal Protection Systems |date=30 January 2014 |last1=Johnson |first1=Sylvia M. |last2=Squire |first2=Thomas H. |last3=Lawson |first3=John W. |last4=Gusman |first4=Michael |last5=Lau |first5=K-H |last6=Sanjuro |first6=Angel |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140012475.pdf |conference=38th Annual Conference on Composites, Materials, and Structures January 27–30, 2014 |access-date=December 29, 2018 |archive-date=December 1, 2019 |archive-url=https://web.archive.org/web/20191201012540/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140012475.pdf |url-status=live }}</ref>

During certain intensity of ionization, a ''radio-blackout'' with the spacecraft is produced.<ref name="Poddar Sharma 2015 pp. 5899–5902">{{cite journal | last1=Poddar | first1=Shashi | last2=Sharma | first2=Deewakar | title=Blackout mitigation during space vehicle re-entry | journal=Optik | publisher=Elsevier BV | volume=126 | issue=24 | year=2015 | issn=0030-4026 | doi=10.1016/j.ijleo.2015.09.141 | pages=5899–5902| bibcode=2015Optik.126.5899P }}</ref>

While Earth entry interface is considered to take place at the [[Earth:Kármán line|Kármán line]] {{convert|100|km|ft}}, the main heating during controlled entry takes place at altitudes of {{convert|65|to|35|km|ft}}, peaking at {{convert|58|km|ft}}.<ref name="Di Fiore Maggiore Mainini pp. 3017–3035">{{cite journal | last1=Di Fiore | first1=Francesco | last2=Maggiore | first2=Paolo | last3=Mainini | first3=Laura | title=Multifidelity domain-aware learning for the design of re-entry vehicles | journal=Structural and Multidisciplinary Optimization | publisher=Springer Science and Business Media LLC | volume=64 | issue=5 | date=2021-10-04 | issn=1615-147X | doi=10.1007/s00158-021-03037-4 | pages=3017–3035| s2cid=244179568 | doi-access=free }}</ref>

===Shock layer gas physics ===
At typical reentry temperatures, the air in the shock layer is both ionized and [[Chemistry:Dissociation|dissociated]].{{citation needed |date=May 2021}}<ref>{{Cite web|url=https://www.icas.org/ICAS_ARCHIVE/ICAS2018/data/papers/ICAS2018_0837_paper.pdf|title=Ionization And Dissociation Effects On Hypersonic Boundary-Layer Stability|access-date=May 13, 2021|archive-date=October 1, 2021|archive-url=https://web.archive.org/web/20211001035258/https://www.icas.org/ICAS_ARCHIVE/ICAS2018/data/papers/ICAS2018_0837_paper.pdf|url-status=live}}</ref> This chemical dissociation necessitates various physical models to describe the shock layer's thermal and chemical properties. There are four basic physical models of a gas that are important to aeronautical engineers who design heat shields:

====Perfect gas model====
Almost all aeronautical engineers are taught the [[Physics:Ideal gas|perfect (ideal) gas model]] during their undergraduate education. Most of the important perfect gas equations along with their corresponding tables and graphs are shown in NACA Report 1135.<ref>{{cite journal |title=Equations, tables, and charts for compressible flow |publisher=NASA Technical Reports |issue=NACA-TR-1135 |year=1953 |url=http://www.nasa.gov/sites/default/files/734673main_Equations-Tables-Charts-CompressibleFlow-Report-1135.pdf |journal=NACA Annual Report |volume=39 |pages=613–681 |access-date=June 17, 2015 |archive-date=September 4, 2015 |archive-url=https://web.archive.org/web/20150904043857/http://www.nasa.gov/sites/default/files/734673main_Equations-Tables-Charts-CompressibleFlow-Report-1135.pdf |url-status=live }}</ref> Excerpts from NACA Report 1135 often appear in the appendices of thermodynamics textbooks and are familiar to most aeronautical engineers who design supersonic aircraft.

The perfect gas theory is elegant and extremely useful for designing aircraft but assumes that the gas is chemically inert. From the standpoint of aircraft design, air can be assumed to be inert for temperatures less than 550 K at one atmosphere pressure. The perfect gas theory begins to break down at 550 K and is not usable at temperatures greater than 2,000 K. For temperatures greater than 2,000 K, a heat shield designer must use a ''real gas model''.

====Real (equilibrium) gas model====
An entry vehicle's pitching moment can be significantly influenced by real-gas effects. Both the Apollo command module and the Space Shuttle were designed using incorrect pitching moments determined through inaccurate real-gas modelling. The Apollo-CM's trim-angle angle of attack was higher than originally estimated, resulting in a narrower lunar return entry corridor. The actual aerodynamic center of the [[Engineering:Space Shuttle Columbia|''Columbia'']] was upstream from the calculated value due to real-gas effects. On ''Columbia''{{'}}s maiden flight ([[Engineering:STS-1|STS-1]]), astronauts John Young and Robert Crippen had some anxious moments during reentry when there was concern about losing control of the vehicle.<ref>Kenneth Iliff and Mary Shafer, ''Space Shuttle Hypersonic Aerodynamic and Aerothermodynamic Flight Research and the Comparison to Ground Test Results'', pp. 5–6 {{ISBN?}}</ref>

An equilibrium real-gas model assumes that a gas is chemically reactive, but also assumes all chemical reactions have had time to complete and all components of the gas have the same temperature (this is called ''[[Physics:Thermodynamic equilibrium|thermodynamic equilibrium]]''). When air is processed by a shock wave, it is superheated by compression and chemically dissociates through many different reactions. Direct friction upon the reentry object is not the main cause of shock-layer heating. It is caused mainly from [[Physics:Isentropic process|isentropic]] heating of the air molecules within the compression wave. Friction based entropy increases of the molecules within the wave also account for some heating.{{original research inline|date=November 2013}} The distance from the shock wave to the [[Physics:Stagnation point|stagnation point]] on the entry vehicle's leading edge is called ''shock wave stand off''. An approximate rule of thumb for shock wave standoff distance is 0.14 times the nose radius. One can estimate the time of travel for a gas molecule from the shock wave to the stagnation point by assuming a free stream velocity of 7.8&nbsp;km/s and a nose radius of 1 meter, i.e., time of travel is about 18 microseconds. This is roughly the time required for shock-wave-initiated chemical dissociation to approach [[Chemistry:Chemical equilibrium|chemical equilibrium]] in a shock layer for a 7.8&nbsp;km/s entry into air during peak heat flux. Consequently, as air approaches the entry vehicle's stagnation point, the air effectively reaches chemical equilibrium thus enabling an equilibrium model to be usable. For this case, most of the shock layer between the shock wave and leading edge of an entry vehicle is chemically reacting and ''not'' in a state of equilibrium. The Fay–Riddell equation,<ref name="Fay-Riddell"/> which is of extreme importance towards modeling heat flux, owes its validity to the stagnation point being in chemical equilibrium. The time required for the shock layer gas to reach equilibrium is strongly dependent upon the shock layer's pressure. For example, in the case of the [[Engineering:Galileo (spacecraft)|''Galileo'']] probe's entry into Jupiter's atmosphere, the shock layer was mostly in equilibrium during peak heat flux due to the very high pressures experienced (this is counterintuitive given the free stream velocity was 39&nbsp;km/s during peak heat flux).

Determining the thermodynamic state of the stagnation point is more difficult under an equilibrium gas model than a perfect gas model. Under a perfect gas model, the ''ratio of specific heats'' (also called ''isentropic exponent'', adiabatic index, ''gamma'', or ''kappa'') is assumed to be constant along with the [[Physics:Gas constant|gas constant]]. For a real gas, the ratio of specific heats can wildly oscillate as a function of temperature. Under a perfect gas model there is an elegant set of equations for determining thermodynamic state along a constant entropy stream line called the ''isentropic chain''. For a real gas, the isentropic chain is unusable and a ''Mollier diagram'' would be used instead for manual calculation. However, graphical solution with a Mollier diagram is now considered obsolete with modern heat shield designers using computer programs based upon a digital lookup table (another form of Mollier diagram) or a chemistry based thermodynamics program. The chemical composition of a gas in equilibrium with fixed pressure and temperature can be determined through the ''Gibbs free energy method''. [[Physics:Gibbs free energy|Gibbs free energy]] is simply the total [[Physics:Enthalpy|enthalpy]] of the gas minus its total [[Physics:Entropy|entropy]] times temperature. A chemical equilibrium program normally does not require chemical formulas or reaction-rate equations. The program works by preserving the original elemental abundances specified for the gas and varying the different molecular combinations of the elements through numerical iteration until the lowest possible Gibbs free energy is calculated (a Newton–Raphson method is the usual numerical scheme). The data base for a Gibbs free energy program comes from spectroscopic data used in defining [[Physics:Partition function (statistical mechanics)|partition functions]]. Among the best equilibrium codes in existence is the program ''Chemical Equilibrium with Applications'' (CEA) which was written by Bonnie J. McBride and Sanford Gordon at NASA Lewis (now renamed "NASA Glenn Research Center"). Other names for CEA are the "Gordon and McBride Code" and the "Lewis Code". CEA is quite accurate up to 10,000 K for planetary atmospheric gases, but unusable beyond 20,000 K ([[Physics:Double ionization|double ionization]] is not modelled). [https://web.archive.org/web/20200205195610/https://www.grc.nasa.gov/WWW/CEAWeb/ CEA can be downloaded from the Internet] along with full documentation and will compile on Linux under the [[Software:GNU Fortran|G77 Fortran]] compiler.

====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|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 [[Astronomy:Low Earth orbit|low Earth orbit]] where entry velocity is approximately {{cvt|7.8|km/s|km/h mph}}. For lunar return entry of 11&nbsp;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.
Atmospheric entry interface velocities on a Mars–Earth trajectory are on the order of {{Cvt|12|km/s|km/h mph}}.<ref>[https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080023907.pdf Overview of the Mars Sample Return Earth Entry Vehicle] {{Webarchive|url=https://web.archive.org/web/20191201012539/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080023907.pdf |date=December 1, 2019 }}, NASA, accessed 29 December 2018.</ref>
Modeling high-speed Mars atmospheric entry—which involves a carbon dioxide, nitrogen and argon atmosphere—is even more complex requiring a 19-species model.{{citation needed|date=August 2018}}

An important aspect of modelling non-equilibrium real gas effects is radiative heat flux. If a vehicle is entering an atmosphere at very high speed (hyperbolic trajectory, lunar return) and has a large nose radius then radiative heat flux can dominate TPS heating. Radiative heat flux during entry into an air or carbon dioxide atmosphere typically comes from asymmetric diatomic molecules; e.g., [[Chemistry:Cyanogen|cyanogen]] (CN), [[Chemistry:Carbon monoxide|carbon monoxide]], [[Chemistry:Nitric oxide|nitric oxide]] (NO), single ionized molecular nitrogen etc. These molecules are formed by the shock wave dissociating ambient atmospheric gas followed by recombination within the shock layer into new molecular species. The newly formed diatomic molecules initially have a very high vibrational temperature that efficiently transforms the [[Physics:Quantum harmonic oscillator|vibrational energy]] into [[Physics:Radiant energy|radiant energy]]; i.e., radiative heat flux. The whole process takes place in less than a millisecond which makes modelling a challenge. The experimental measurement of radiative heat flux (typically done with shock tubes) along with theoretical calculation through the unsteady [[Physics:Schrödinger equation|Schrödinger equation]] are among the more esoteric aspects of aerospace engineering. Most of the aerospace research work related to understanding radiative heat flux was done in the 1960s, but largely discontinued after conclusion of the Apollo Program. Radiative heat flux in air was just sufficiently understood to ensure Apollo's success. However, radiative heat flux in carbon dioxide (Mars entry) is still barely understood and will require major research.{{citation needed|date=August 2018}}

====Frozen gas model====
The frozen gas model describes a special case of a gas that is not in equilibrium. The name "frozen gas" can be misleading. A frozen gas is not "frozen" like ice is frozen water. Rather a frozen gas is "frozen" in time (all chemical reactions are assumed to have stopped). Chemical reactions are normally driven by collisions between molecules. If gas pressure is slowly reduced such that chemical reactions can continue then the gas can remain in equilibrium. However, it is possible for gas pressure to be so suddenly reduced that almost all chemical reactions stop. For that situation the gas is considered frozen.{{citation needed|date=August 2018}}

The distinction between equilibrium and frozen is important because it is possible for a gas such as air to have significantly different properties (speed-of-sound, [[Physics:Viscosity|viscosity]] etc.) for the same thermodynamic state; e.g., pressure and temperature. Frozen gas can be a significant issue in the wake behind an entry vehicle. During reentry, free stream air is compressed to high temperature and pressure by the entry vehicle's shock wave. Non-equilibrium air in the shock layer is then transported past the entry vehicle's leading side into a region of rapidly expanding flow that causes freezing. The frozen air can then be entrained into a trailing vortex behind the entry vehicle. Correctly modelling the flow in the wake of an entry vehicle is very difficult. [[Engineering:Space Shuttle thermal protection system|Thermal protection shield]] (TPS) heating in the vehicle's afterbody is usually not very high, but the geometry and unsteadiness of the vehicle's wake can significantly influence aerodynamics (pitching moment) and particularly dynamic stability.{{citation needed|date=August 2018}}

==Thermal protection systems==
A '''thermal protection system''', or TPS, is the barrier that protects a [[Engineering:Spacecraft|spacecraft]] during the searing heat of atmospheric reentry. A secondary goal may be to protect the spacecraft from the [[Astronomy:Solar constant|heat]] [[Physics:Thermal radiation|and]] [[Astronomy:Outer space#Environment|cold]] of space while in orbit. Multiple approaches for the thermal protection of spacecraft are in use, among them ablative heat shields, passive cooling, and active cooling of spacecraft surfaces.

===Ablative===
[[File:Apollo 12 heat shield.JPG|thumb|left|Ablative heat shield (after use) on [[Engineering:Apollo 12|Apollo 12]] capsule]]
The [[Physics:Ablation|ablative]] heat shield functions by lifting the hot shock layer gas away from the heat shield's outer wall (creating a cooler [[Physics:Boundary layer|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 sublimes, while the bulk of the TPS material undergoes [[Chemistry:Pyrolysis|pyrolysis]] and expels product gases. The gas produced by pyrolysis is what drives blowing and causes blockage of convective and catalytic heat flux. [[Chemistry:Pyrolysis|Pyrolysis]] can be measured in real time using [[Chemistry:Thermogravimetric analysis|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). Carbon phenolic was originally developed as a rocket nozzle throat material (used in the [[Engineering:Space Shuttle Solid Rocket Booster|Space Shuttle Solid Rocket Booster]]) and for reentry-vehicle nose tips.

Early research on ablation technology in the USA was centered at [[Organization:NASA|NASA]]'s [[Organization:Ames Research Center|Ames Research Center]] located at Moffett Field, California. [[Organization:Ames Research Center|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 [[Physics:Hypersonic|hypersonic]] wind tunnel.<ref>Hogan, C. Michael, Parker, John and Winkler, Ernest, of [[Organization:NASA|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]]
The [[Physics:Thermal conductivity|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. 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 [[Company:Lockheed Martin|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 [[Organization:Mars Science Laboratory|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 ''[[Astronomy:Viking 1|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 [[Engineering: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 [[Organization:Mars Science Laboratory|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 [[Company:SpaceX|SpaceX]] in 2006–2010<ref name="N+SX_picaX"/> for the [[Software:SpaceX Dragon 1|Dragon]] [[Engineering:Space capsule|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 [[Engineering: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>

====SIRCA====
[[File:Ds 2.jpg|thumb|Deep Space 2 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 ''[[Astronomy:Mars Pathfinder|Mars Pathfinder]]'' and [[Astronomy:Mars Exploration Rover|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 [[Astronomy:Deep Space 2|Deep Space 2]] (DS/2) Mars [[Engineering: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 is a [[Organization:NASA|NASA]]-specified ablative heat shield, a glass-filled [[Chemistry:Epoxy|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 }} [[Organization:NASA|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 verified. -->

NASA originally used it for the [[Engineering: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 [[Engineering: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 Andrew S. W. Thomas takes a close look at TPS tiles underneath 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 [[Astronomy:Boeing X-37|Boeing X-37]], [[Engineering:Dream Chaser|Dream Chaser]], and [[Engineering:SpaceX Starship (spacecraft)|Starships 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 [[Engineering:Space Shuttle thermal protection system|Space Shuttle TPS]] tiles ([[Physics:LI-900|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.

===Passively cooled===
[[File:Mercury capsule HD.jpg|thumb|upright|left|The Mercury capsule design (shown here with its [[Astronomy: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 [[Engineering:Sub-orbital spaceflight|sub-orbital]] [[Astronomy: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., [[Chemistry:Titanium|titanium]], [[Chemistry:Beryllium|beryllium]], [[Chemistry:Copper|copper]], etc.). Modern designers prefer to avoid this added mass by using ablative and thermal-soak TPS instead.

Thermal protection systems relying on [[Physics:Emissivity|emissivity]] use high emissivity coatings (HECs) to facilitate [[Astronomy:Radiative cooling|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 | s2cid=118927116}}</ref>

Radiatively cooled TPS can be found on modern entry vehicles, but [[Chemistry:Reinforced carbon–carbon|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. [[Chemistry:Carbon|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 [[Engineering:Aircraft|aircraft]], such as the SR-71 Blackbird and [[Engineering:Concorde|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 [[Engineering:Annealing (metallurgy)|annealing]] due to aerodynamic heating. In the case of the Concorde, the [[Chemistry:Aluminium|aluminium]] nose was permitted to reach a maximum [[Physics:Operating temperature|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 [[Engineering: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 [[Chemistry:Superalloy|superalloy]] (dubbed [[Chemistry:René 41|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 [[Chemistry:Superalloy|superalloy]]-shingle TPS was again proposed for the unsuccessful X-33 [[Astronomy:Single-stage-to-orbit|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 [[Chemistry: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 [[Chemistry:Zirconium diboride|zirconium diboride]] and [[Chemistry:Hafnium diboride|hafnium diboride]]. SHARP TPS have suggested performance improvements allowing for sustained [[Physics: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 [[Organization:University of Montana|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 [[Chemistry:Alloy|alloy]]s that incorporate a refrigerant or cryogenic fuel circulating through them.

Such a TPS concept was proposed for the [[Engineering: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 [[Astronomy:Scramjet|scramjet]] powered hypersonic aircraft, but failed in development.{{cn|date=January 2024}}

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

In early 2019, [[Company:SpaceX|SpaceX]] was developing an actively cooled heat shield for its [[Engineering:SpaceX Starship|Starship]] spacecraft where a part of the thermal protection system will be a [[Physics: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 |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 &amp; 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 &amp; 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>

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 [[Engineering:Falcon 9|Falcon 9]] performs an entry burn to rapidly decelerate from its initial hypersonic speed.{{Citation needed|date=June 2021}}

==Feathered entry==
In 2004, aircraft designer [[Biography:Burt Rutan|Burt Rutan]] demonstrated the feasibility of a shape-changing airfoil for reentry with the sub-orbital [[Engineering:SpaceShipOne|SpaceShipOne]]. The wings on this craft rotate upward into the ''feathered configuration'' that provides a [[Engineering:Shuttlecock|shuttlecock]] effect. Thus SpaceShipOne achieves much more aerodynamic drag on reentry while not experiencing significant thermal loads.

The configuration increases drag, as the craft is now less streamlined and results in more atmospheric gas particles hitting the spacecraft at higher altitudes than otherwise. The aircraft thus slows down more in higher atmospheric layers which is the key to efficient reentry. Secondly, the aircraft will ''automatically'' orient itself in this state to a high drag attitude.<ref>{{cite web|url=http://science.howstuffworks.com/spaceshipone4.htm|title=How SpaceShipOne Works|date=20 June 2004|access-date=April 23, 2011|archive-date=January 12, 2012|archive-url=https://web.archive.org/web/20120112003232/http://science.howstuffworks.com/spaceshipone4.htm|url-status=live}}</ref>

However, the velocity attained by SpaceShipOne prior to reentry is much lower than that of an orbital spacecraft, and engineers, including Rutan, recognize that a feathered reentry technique is not suitable for return from orbit.

On 4 May 2011, the first test on the [[Engineering:SpaceShipTwo|SpaceShipTwo]] of the feathering mechanism was made during a glideflight after release
from the White Knight Two. Premature deployment of the feathering system was responsible for the 2014 [[Engineering:VSS Enterprise crash|VSS ''Enterprise'' crash]], in which the aircraft disintegrated, killing the co-pilot.

The feathered reentry was first described by Dean Chapman of NACA in 1958.<ref name="NACA TN 4276">{{cite journal | url=http://naca.larc.nasa.gov/reports/1958/naca-tn-4276/naca-tn-4276.pdf | title=An approximate analytical method for studying reentry into planetary atmospheres | author=Chapman, Dean R. | journal=NACA Technical Note 4276 | date=May 1958 | page=38 | url-status=dead | archive-url=https://web.archive.org/web/20050127060632/http://naca.larc.nasa.gov/reports/1958/naca-tn-4276/naca-tn-4276.pdf | archive-date=January 27, 2005 | access-date=February 3, 2006 }}</ref> In the section of his report on ''Composite Entry'', Chapman described a solution to the problem using a high-drag device:

{{quote|
It may be desirable to combine lifting and nonlifting entry in order to achieve some advantages... For landing maneuverability it obviously is advantageous to employ a lifting vehicle. The total heat absorbed by a lifting vehicle, however, is much higher than for a nonlifting vehicle... Nonlifting vehicles can more easily be constructed... by employing, for example, a large, light drag device... The larger the device, the smaller is the heating rate.

Nonlifting vehicles with shuttlecock stability are advantageous also from the viewpoint of minimum control requirements during entry.

... an evident composite type of entry, which combines some of the desirable features of lifting and nonlifting trajectories, would be to enter first without lift but with a... drag device; then, when the velocity is reduced to a certain value... the device is jettisoned or retracted, leaving a lifting vehicle... for the remainder of the descent.
}}

=={{anchor|IRDT}} Inflatable heat shield entry==

Deceleration for atmospheric reentry, especially for higher-speed Mars-return missions, benefits from maximizing "the drag area of the entry system. The larger the diameter of the aeroshell, the bigger the payload can be."<ref name=nasa20090817/> An inflatable aeroshell provides one alternative for enlarging the drag area with a low-mass design.

===Russia===

Such an inflatable shield/aerobrake was designed for the penetrators of [[Astronomy:Mars 96|Mars 96]] mission. Since the mission failed due to the launcher malfunction, the NPO Lavochkin and DASA/ESA have designed a mission for Earth orbit. The Inflatable Reentry and Descent Technology (IRDT) demonstrator was launched on Soyuz-Fregat on 8 February 2000. The inflatable shield was designed as a cone with two stages of inflation. Although the second stage of the shield failed to inflate, the demonstrator survived the orbital reentry and was recovered.<ref name=IRDT2000>{{Cite web|url=http://www.esa.int/esapub/bulletin/bullet103/marraffa103.pdf|title=Inflatable Re-Entry Technologies: Flight Demonstration and Future Prospects|access-date=April 22, 2011|archive-date=January 29, 2012|archive-url=https://web.archive.org/web/20120129130903/http://www.esa.int/esapub/bulletin/bullet103/marraffa103.pdf|url-status=live}}</ref><ref>[http://www.spaceflight.esa.int/irdt/factsheet.pdf Inflatable Reentry and Descent Technology (IRDT)] {{webarchive|url=https://web.archive.org/web/20151231130516/http://www.spaceflight.esa.int/irdt/factsheet.pdf |date=2015-12-31 }} Factsheet, ESA, September, 2005</ref> The subsequent missions flown on the [[Engineering:Volna|Volna]] rocket failed due to launcher failure.<ref name=2R2Smissions>{{Cite web|url=https://www.2r2s.com/|archive-url=https://web.archive.org/web/20161207150236/http://www.2r2s.com/demo_missions.html |url-status=dead |title=502 Bad Gateway nginx openresty 208.80.154.81|archive-date=December 7, 2016|website=www.2r2s.com}}</ref>

[[File:Inflatable Re-entry Vehicle Experiment.jpg|thumb|right|NASA engineers check IRVE.]]

===NASA IRVE ===
NASA launched an inflatable heat shield experimental spacecraft on 17 August 2009 with the successful first test flight of the Inflatable Re-entry Vehicle Experiment (IRVE). The heat shield had been [[Physics:Vacuum packing|vacuum-packed]] into a {{convert|15|in|adj=mid|cm|-diameter}} payload shroud and launched on a [[Engineering:Black Brant (rocket)|Black Brant 9]] [[Engineering:Sounding rocket|sounding rocket]] from NASA's Wallops Flight Facility on Wallops Island, Virginia. "Nitrogen inflated the {{convert|10|ft|m|adj=mid|-diameter}} heat shield, made of several layers of [[Chemistry:Silicone|silicone]]-coated <nowiki>[</nowiki>[[Chemistry:Kevlar|Kevlar]]<nowiki>]</nowiki> fabric, to a mushroom shape in space several minutes after liftoff."<ref name=nasa20090817/> The rocket apogee was at an altitude of {{convert|131|mi}} where it began its descent to supersonic speed. Less than a minute later the shield was released from its cover to inflate at an altitude of {{convert|124|mi}}. The inflation of the shield took less than 90 seconds.<ref name=nasa20090817>[http://www.nasa.gov/topics/aeronautics/features/irve.html NASA Launches New Technology: An Inflatable Heat Shield] {{Webarchive|url=https://web.archive.org/web/20101219090403/http://www.nasa.gov/topics/aeronautics/features/irve.html |date=December 19, 2010 }}, [[Organization:NASA|NASA]] Mission News, 2009-08-17, accessed 2011-01-02.</ref>

=== NASA HIAD ===
Following the success of the initial IRVE experiments, NASA developed the concept into the more ambitious Hypersonic Inflatable Aerodynamic Decelerator (HIAD). The current design is shaped like a shallow cone, with the structure built up as a stack of circular inflated tubes of gradually increasing major diameter. The forward (convex) face of the cone is covered with a flexible thermal protection system robust enough to withstand the stresses of atmospheric entry (or reentry).<ref name = HughesOverview>{{cite web
 |url          = https://solarsystem.nasa.gov/docs/9.2%20Hughes%20Hypersonic%20Inflatable%20Aerodynamic%20Decelerator%20(HIAD)%20Technology%20Development%20Overview.pdf
 |title        = Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview
 |last         = Hughes
 |first        = Stephen J.
 |website      = www.nasa.gov
 |publisher    = NASA
 |access-date  = 28 March 2017
 |archive-url  = https://web.archive.org/web/20170126043753/http://solarsystem.nasa.gov/docs/9.2%20Hughes%20Hypersonic%20Inflatable%20Aerodynamic%20Decelerator%20(HIAD)%20Technology%20Development%20Overview.pdf
 |archive-date = 26 January 2017
 |url-status   = dead
}}</ref><ref name = Cheatwood2016>{{cite web
 |url          = https://www.nasa.gov/sites/default/files/atoms/files/gcd_industryday_hiad.pdf
 |title        = Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology
 |last         = Cheatwood
 |first        = Neil
 |date         = 29 June 2016
 |website      = www.nasa.gov
 |publisher    = NASA
 |access-date  = 28 March 2017
 |archive-date = February 24, 2017
 |archive-url  = https://web.archive.org/web/20170224080715/https://www.nasa.gov/sites/default/files/atoms/files/gcd_industryday_hiad.pdf
 |url-status   = live
}}</ref>

In 2012, a HIAD was tested as Inflatable Reentry Vehicle Experiment 3 (IRVE-3) using a sub-orbital sounding rocket, and worked.<ref>{{Cite web|url=http://www.ulalaunch.com/uploads/docs/Published_Papers/Supporting_Technologies/LV_Recovery_and_Reuse_AIAASpace_2015.pdf|title=''Launch Vehicle Recovery and Reuse ''|access-date=January 10, 2018|archive-date=July 6, 2016|archive-url=https://web.archive.org/web/20160706013800/http://www.ulalaunch.com/uploads/docs/Published_Papers/Supporting_Technologies/LV_Recovery_and_Reuse_AIAASpace_2015.pdf|url-status=live}}</ref>{{rp|8}}

See also [[Engineering:Low-Density Supersonic Decelerator|Low-Density Supersonic Decelerator]], a NASA project with tests in 2014 & 2015 of a 6m diameter SIAD-R.

=== LOFTID ===
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 Nov 2022, inflated in orbit, reentered faster than Mach 25, and was successfully recovered on November 10.

==Entry vehicle design considerations==

There are four critical parameters considered when designing a vehicle for atmospheric entry:{{Cn|date=September 2022}}

# Peak heat flux
# Heat load
# Peak deceleration
# Peak dynamic pressure

Peak heat flux and [[Physics:Dynamic pressure|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 [[Physics:Spallation|spallation]] is an issue.

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 [[Astronomy: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).

Based upon Allen and Eggers discovery, maximum aeroshell bluntness (maximum drag) yields minimum TPS mass. Maximum bluntness (minimum ballistic coefficient) also yields a minimal [[Physics:Terminal velocity|terminal velocity]] at maximum altitude (very important for Mars EDL, but detrimental for military RVs). However, there is an upper limit to bluntness imposed by aerodynamic stability considerations based upon ''shock wave detachment''. A shock wave will remain attached to the tip of a sharp cone if the cone's half-angle is below a critical value. This critical half-angle can be estimated using perfect gas theory (this specific aerodynamic instability occurs below hypersonic speeds). For a nitrogen atmosphere (Earth or Titan), the maximum allowed half-angle is approximately 60°. For a carbon dioxide atmosphere (Mars or Venus), the maximum-allowed half-angle is approximately 70°. After shock wave detachment, an entry vehicle must carry significantly more shocklayer gas around the leading edge stagnation point (the subsonic cap). Consequently, the aerodynamic center moves upstream thus causing aerodynamic instability. It is incorrect to reapply an aeroshell design intended for Titan entry ([[Engineering:Huygens (spacecraft)|''Huygens'']] probe in a nitrogen atmosphere) for Mars entry (''[[Astronomy:Beagle 2|Beagle 2]]'' in a carbon dioxide atmosphere).{{citation needed|date=September 2016}}{{original research inline|date=September 2016}} Prior to being abandoned, the [[Astronomy:Mars program|Soviet Mars lander program]] achieved one successful landing ([[Astronomy:Mars 3|Mars 3]]), on the second of three entry attempts (the others were [[Astronomy:Mars 2|Mars 2]] and [[Astronomy:Mars 6|Mars 6]]). The Soviet Mars landers were based upon a 60° half-angle aeroshell design.

A 45° half-angle sphere-cone is typically used for atmospheric probes (surface landing not intended) even though TPS mass is not minimized. The rationale for a 45° half-angle is to have either aerodynamic stability from entry-to-impact (the heat shield is not jettisoned) or a short-and-sharp heat pulse followed by prompt heat shield jettison. A 45° sphere-cone design was used with the DS/2 Mars [[Engineering:Lander (spacecraft)|impactor]] and [[Astronomy:Pioneer Venus project|Pioneer Venus]] probes.

==Notable atmospheric entry accidents==
[[File:Ingreso reentrada.svg|center|thumb|800px|Reentry window
{{ordered list
 | list-style-type = upper-alpha
 | Friction with air
 | In air flight
 | Expulsion lower angle
 | Perpendicular to the entry point
 | Excess friction 6.9° to 90°
 | Repulsion of 5.5° or less
 | Explosion friction
 | Plane tangential to the entry point
}}]]

Not all atmospheric reentries have been completely successful:

*[[Astronomy:Voskhod 2|Voskhod 2]]&nbsp;– The service module failed to detach for some time, but the crew survived.
*[[Engineering:Soyuz 5|Soyuz 5]]&nbsp;– The service module failed to detach, but the crew survived.
*[[Engineering:Apollo 15|Apollo 15]]&nbsp;- One of the three ringsail parachutes failed during the ocean landing, likely damaged as the spacecraft vented excess control fuel. The spacecraft was designed to land safely with only two parachutes, and the crew were uninjured.
*[[Astronomy:Mars Polar Lander|Mars Polar Lander]]&nbsp;– Failed during EDL. The failure was believed to be the consequence of a software error. The precise cause is unknown for lack of real-time [[Telemetry|telemetry]].
*[[Engineering:Space Shuttle Columbia|Space Shuttle ''Columbia'']] [[Engineering:STS-1#Mission anomalies|STS-1]]&nbsp;– a combination of launch damage, protruding gap filler, and tile installation error resulted in serious damage to the orbiter, only some of which the crew was aware. Had the crew known the true extent of the damage before attempting reentry, they would have flown the shuttle to a safe altitude and then bailed out. Nevertheless, reentry was successful, and the orbiter proceeded to a normal landing.
*[[Engineering:Space Shuttle Atlantis|Space Shuttle ''Atlantis'']] [[Engineering:STS-27#Tile damage|STS-27]]&nbsp;– Insulation from the starboard [[Engineering:Space Shuttle Solid Rocket Booster|solid rocket booster]] nose cap struck the orbiter during launch, causing significant tile damage. This dislodged one tile completely, over an aluminum mounting plate for a TACAN antenna. The antenna sustained extreme heat damage, but prevented the hot gas from penetrating the vehicle body.
[[File:Genesis wreck.jpg|thumb|Genesis entry vehicle after crash]]
*[[Engineering:Genesis (spacecraft)|''Genesis'']]&nbsp;– The parachute failed to deploy due to a G-switch having been installed backwards (a similar error delayed parachute deployment for the [[Astronomy:Galileo Probe|''Galileo'' Probe]]). Consequently, the Genesis entry vehicle crashed into the desert floor. The payload was damaged, but most scientific data were recoverable.
*[[Engineering:Soyuz TMA-11|Soyuz TMA-11]]&nbsp;– 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.

Some reentries have resulted in significant disasters:

*[[Engineering:Soyuz 1|Soyuz 1]]&nbsp;– The [[Astronomy:Spacecraft attitude control|attitude control]] system failed while still in orbit and later parachutes got entangled during the emergency landing sequence (entry, descent, and landing (EDL) failure). Lone cosmonaut Vladimir Mikhailovich Komarov died.
*[[Engineering:Soyuz 11|Soyuz 11]]&nbsp;– During tri-module separation, a valve seal was opened by the shock, depressurizing the descent module; the crew of three asphyxiated in space minutes before reentry.
*Space Shuttle ''Columbia'' [[Engineering:Space Shuttle Columbia disaster|STS-107]]&nbsp;– The failure of a [[Chemistry:Reinforced carbon–carbon|reinforced carbon–carbon]] panel on a wing leading edge caused by debris impact at launch led to breakup of the orbiter on reentry resulting in the deaths of all seven crew members.

==Uncontrolled and unprotected entries==
Of satellites that reenter, approximately 10–40% of the mass of the object is likely to reach the surface of the Earth.<ref>{{Cite web|url=http://www.aerospace.org/cords/spacecraft-reentry/how-much-material-from-a-satellite-will-survive-reentry/|archive-url=https://web.archive.org/web/20140302115307/http://www.aerospace.org/cords/spacecraft-reentry/how-much-material-from-a-satellite-will-survive-reentry/ |url-status=dead |title=Spacecraft Reentry FAQ: How much material from a satellite will survive reentry?|archive-date=March 2, 2014}}</ref> On average, about one catalogued object reenters per day.<ref>{{Cite web|url=https://www.nasa.gov/news/debris_faq.html|archive-url=https://web.archive.org/web/20140311033244/http://www.nasa.gov/news/debris_faq.html |url-status=dead |title=NASA - Frequently Asked Questions: Orbital Debris|archive-date=March 11, 2014|website=www.nasa.gov}}</ref>

Due to the Earth's surface being primarily water, most objects that survive reentry land in one of the world's oceans. The estimated chances that a given person will get hit and injured during their lifetime is around 1 in a trillion.<ref>{{cite web|url=http://www.aerospace.org/cords/spacecraft-reentry/|title=Animation52-desktop|website=www.aerospace.org|access-date=2013-03-04|archive-url=https://web.archive.org/web/20140302150713/http://www.aerospace.org/cords/spacecraft-reentry/|archive-date=2014-03-02|url-status=dead}}</ref>

On January 24, 1978, the [[Place:Soviet Union|Soviet]] [[Engineering:Kosmos 954|Kosmos 954]] ({{convert|3800|kg|disp=sqbr}}) reentered and crashed near Great Slave Lake in the [[Earth:Northwest Territories|Northwest Territories]] of Canada. The satellite was nuclear-powered and left radioactive debris near its impact site.<ref name="jaxa">{{cite web|url=http://www.jaxa.jp/library/space_law/chapter_3/3-2-2-1_e.html|title=3-2-2-1 Settlement of Claim between Canada and the Union of Soviet Socialist Republics for Damage Caused by "Cosmos 954" (Released on April 2, 1981)|website=www.jaxa.jp|access-date=December 28, 2010|archive-date=September 30, 2019|archive-url=https://web.archive.org/web/20190930070151/http://www.jaxa.jp/library/space_law/chapter_3/3-2-2-1_e.html|url-status=live}}</ref>

On July 11, 1979, the US [[Engineering:Skylab|Skylab]] space station ({{convert|77100|kg|disp=sqbr}}) reentered and spread debris across the Australian Outback.<ref>{{cite book|last=Hanslmeier|first=Arnold|title=The sun and space weather|date=2002|publisher=Kluwer Academic Publishers|location=Dordrecht; Boston|isbn=9781402056048|page=269}}</ref> The reentry was a major media event largely due to the Cosmos 954 incident, but not viewed as much as a potential disaster since it did not carry toxic nuclear or [[Chemistry:Hydrazine|hydrazine]] fuel. NASA had originally hoped to use a [[Engineering:Space Shuttle|Space Shuttle]] mission to either extend its life or enable a controlled reentry, but delays in the Shuttle program, plus unexpectedly high solar activity, made this impossible.<ref>{{cite book|last=Lamprecht|first=Jan|title=Hollow planets : a feasibility study of possible hollow worlds|date=1998|publisher=World Wide Pub.|location=Austin, Texas|isbn=9780620219631|page=326}}</ref><ref>{{cite book|last=Elkins-Tanton|first=Linda|title=The Sun, Mercury, and Venus|date=2006|publisher=Chelsea House|location=New York|isbn=9780816051939|page=56}}</ref>

On February 7, 1991, the Soviet [[Engineering:Salyut 7|Salyut 7]] space station ({{convert|19820|kg|disp=sqbr}}), with the [[Engineering:Kosmos 1686|Kosmos 1686]] module ({{convert|20000|kg|disp=sqbr}}) attached, reentered and scattered debris over the town of Capitán Bermúdez, Argentina.<ref>{{Cite web|url=http://www.aero.org/capabilities/cords/faq8.html|archive-url=https://web.archive.org/web/20120513052250/http://www.aero.org/capabilities/cords/faq8.html |url-status=dead |title=aero.org, ''Spacecraft Reentry FAQ:''|archive-date=May 13, 2012}}</ref><ref name="auto"/><ref>[https://www.nytimes.com/1991/02/07/world/salyut-7-soviet-station-in-space-falls-to-earth-after-9-year-orbit.html "Salyut 7, Soviet Station in Space, Falls to Earth After 9-Year Orbit"] {{Webarchive|url=https://web.archive.org/web/20161118131918/http://www.nytimes.com/1991/02/07/world/salyut-7-soviet-station-in-space-falls-to-earth-after-9-year-orbit.html |date=November 18, 2016 }} ''New York Times''</ref> The station had been boosted to a higher orbit in August 1986 in an attempt to keep it up until 1994, but in a scenario similar to Skylab, the planned [[Engineering:Buran programme|Buran shuttle]] was cancelled and high solar activity caused it to come down sooner than expected.

On September 7, 2011, NASA announced the impending uncontrolled reentry of the [[Engineering:Upper Atmosphere Research Satellite|Upper Atmosphere Research Satellite]] ({{convert|6540|kg|disp=sqbr}}) and noted that there was a small risk to the public.<ref>{{cite news | title=Huge Defunct Satellite to Plunge to Earth Soon, NASA Says | work=Space.com | url=http://www.space.com/12859-nasa-satellite-falling-space-debris-uars.html | author=David, Leonard | date=7 September 2011 | access-date=10 September 2011 | archive-date=May 6, 2021 | archive-url=https://web.archive.org/web/20210506183208/https://www.space.com/12859-nasa-satellite-falling-space-debris-uars.html | url-status=live }}</ref> The decommissioned satellite reentered the atmosphere on September 24, 2011, and some pieces are presumed to have crashed into the South [[Earth:Pacific Ocean|Pacific Ocean]] over a debris field {{convert|500|mi}} long.<ref name="finalupdate">{{cite web |url=http://www.nasa.gov/mission_pages/uars/index.html |title=Final Update: NASA's UARS Re-enters Earth's Atmosphere |access-date=2011-09-27 |archive-date=February 25, 2018 |archive-url=https://web.archive.org/web/20180225204106/https://www.nasa.gov/mission_pages/uars/index.html |url-status=live }}</ref>

On April 1, 2018, the Chinese Tiangong-1 space station ({{convert|8510|kg|disp=sqbr}}) reentered over the Pacific Ocean, halfway between Australia and South America.<ref>{{Cite web |url=http://www.aerospace.org/cords/reentry-predictions/tiangong-1-reentry/ |title=aerospace.org ''Tiangong-1 Reentry'' |access-date=2018-04-02 |archive-url=https://web.archive.org/web/20180404003614/http://www.aerospace.org/cords/reentry-predictions/tiangong-1-reentry/ |archive-date=2018-04-04 |url-status=dead }}</ref> The China Manned Space Engineering Office had intended to control the reentry, but lost [[Telemetry|telemetry]] and control in March 2017.<ref name="rogue">{{cite news|last=Jones|first=Morris|date=30 March 2016|title=Has Tiangong 1 gone rogue|url=http://www.spacedaily.com/reports/Has_Tiangong_1_gone_rogue_999.html|newspaper=Space Daily|access-date=22 September 2016|archive-date=September 13, 2017|archive-url=https://web.archive.org/web/20170913231010/http://www.spacedaily.com/reports/Has_Tiangong_1_gone_rogue_999.html|url-status=live}}</ref>

On May 11, 2020, the core stage of Chinese [[Engineering:Long March 5|Long March 5B]] (COSPAR ID 2020-027C) weighing roughly {{convert|20000|kg|disp=sqbr||abbr=}}) made an uncontrolled reentry over the Atlantic Ocean, near West African coast.<ref>{{cite tweet|user=18SPCS|number=1259891636189839360|date=11 May 2020|title=#18SPCS has confirmed the reentry of the CZ-5B R/B (#45601, 2020-027C) at 08:33 PDT on 11 May, over the Atlantic Ocean. The #CZ5B launched China’s test crew capsule on 5 May 2020. #spaceflightsafety|access-date=11 May 2020|language=en|author=18 Space Control Squadron}}</ref><ref>{{Cite web|title=China's massive Long March 5B's rocket falls out of orbit over Atlantic Ocean – Spaceflight Now|url=https://spaceflightnow.com/2020/05/11/chinas-massive-long-march-5bs-rocket-falls-out-of-orbit-over-ocean/|last=Clark|first=Stephen|language=en-US|access-date=2020-05-12|archive-date=May 14, 2020|archive-url=https://web.archive.org/web/20200514211903/https://spaceflightnow.com/2020/05/11/chinas-massive-long-march-5bs-rocket-falls-out-of-orbit-over-ocean/|url-status=live}}</ref> Few pieces of rocket debris reportedly survived reentry and fell over at least two villages in Ivory Coast.<ref>{{Cite web|title=Bridenstine Criticizes Uncontrolled Long March 5B Stage Reentry – Parabolic Arc|date=May 15, 2020 |url=http://www.parabolicarc.com/2020/05/15/bridenstine-criticizes-uncontrolled-long-march-5b-stage-reentry/|language=en-US|access-date=2020-05-16|archive-date=May 21, 2020|archive-url=https://web.archive.org/web/20200521234307/http://www.parabolicarc.com/2020/05/15/bridenstine-criticizes-uncontrolled-long-march-5b-stage-reentry/|url-status=live}}</ref><ref>{{Cite web|title=Chinese Rocket Debris May Have Fallen On Several African Villages After An Uncontrolled Re-Entry|url=https://www.forbes.com/sites/jonathanocallaghan/2020/05/12/parts-of-a-chinese-rocket-may-have-fallen-on-an-african-village/|last=O'Callaghan|first=Jonathan|website=Forbes|language=en|access-date=2020-05-13|archive-date=May 12, 2020|archive-url=https://web.archive.org/web/20200512235929/https://www.forbes.com/sites/jonathanocallaghan/2020/05/12/parts-of-a-chinese-rocket-may-have-fallen-on-an-african-village/|url-status=live}}</ref>

On May 8, 2021, the core stage of Chinese [[Engineering:Long March 5|Long March 5B]] (COSPAR ID 2021-0035B) weighing {{convert|23000|kg|disp=sqbr||abbr=}}) made an uncontrolled reentry, just west of the Maldives in the Indian Ocean (approximately 72.47°E longitude and 2.65°N latitude).<ref>{{Cite web|title=Huge Chinese rocket booster falls to Earth over Arabian Peninsula - Space.com|url=https://www.n2yo.com/satellite/?s=48275|language=en-US|access-date=2021-05-09|archive-date=July 23, 2021|archive-url=https://web.archive.org/web/20210723163817/https://www.n2yo.com/satellite/?s=48275|url-status=live}}</ref> Witnesses reported rocket debris as far away as the Arabian peninsula.<ref>{{Cite web|title=Huge Chinese rocket booster falls to Earth over Arabian Peninsula - Space.com|url=https://www.space.com/chinese-rocket-booster-long-march-5b-space-junk-crash|last=Wall|first=Mike|website=Space.com|date=May 9, 2021|language=en-US|access-date=2021-05-09|archive-date=July 23, 2021|archive-url=https://web.archive.org/web/20210723163815/https://www.space.com/chinese-rocket-booster-long-march-5b-space-junk-crash|url-status=live}}</ref>

===Deorbit disposal===
[[Engineering:Salyut 1|Salyut 1]], the world's first space station, was deliberately de-orbited into the Pacific Ocean in 1971 following the [[Engineering:Soyuz 11|Soyuz 11]] accident. Its successor, [[Engineering:Salyut 6|Salyut 6]], was de-orbited in a controlled manner as well.

On June 4, 2000, the [[Astronomy:Compton Gamma Ray Observatory|Compton Gamma Ray Observatory]] was deliberately de-orbited after one of its gyroscopes failed. The debris that did not burn up fell harmlessly into the Pacific Ocean. The observatory was still operational, but the failure of another gyroscope would have made de-orbiting much more difficult and dangerous. With some controversy, NASA decided in the interest of public safety that a controlled crash was preferable to letting the craft come down at random.

In 2001, the Russian ''[[Astronomy:Mir|Mir]]'' space station was deliberately de-orbited, and broke apart in the fashion expected by the command center during atmospheric reentry. Mir entered the Earth's atmosphere on March 23, 2001, near Nadi, Fiji, and fell into the South Pacific Ocean.

On February 21, 2008, a disabled U.S. spy satellite, [[Engineering:USA-193|USA-193]], was hit at an altitude of approximately {{convert|246|km|sp=us}} with an SM-3 missile fired from the U.S. Navy [[Engineering:Cruiser|cruiser]] {{USS|Lake Erie|CG-70|2}} off the coast of Hawaii. The satellite was inoperative, having failed to reach its intended orbit when it was launched in 2006. Due to its rapidly deteriorating orbit it was destined for uncontrolled reentry within a month. U.S. Department of Defense expressed concern that the {{convert|1000|lb|adj=on}} fuel tank containing highly toxic [[Chemistry:Hydrazine|hydrazine]] might survive reentry to reach the Earth's surface intact. Several governments including those of Russia, China, and {{wipe|Belarus}} protested the action as a thinly-veiled demonstration of US anti-satellite capabilities.<ref>{{cite news |first=Andrew |last=Gray |title= U.S. has high confidence it hit satellite fuel tank |url= https://www.reuters.com/article/scienceNews/idUSN1930844420080222 |work=Reuters |date=2008-02-21 |access-date=2008-02-23| archive-url= https://web.archive.org/web/20080225204300/https://www.reuters.com/article/scienceNews/idUSN1930844420080222| archive-date= 25 February 2008 | url-status= live}}</ref> China had previously caused an international incident when it tested an anti-satellite missile in 2007.

<gallery>
File:Closeup of Gemini 2 heatshield.jpg|Closeup of Gemini 2 heat shield
File:Cross section of Gemini 2 heatshield.jpg|Cross section of Gemini 2 heat shield
</gallery>

==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 [[Physics:Plasma|plasma]] trail (Soyuz)
File:Re-entry of Progress Spacecraft 42P - NASA Earth Observatory.jpg|[[Engineering:Progress (spacecraft)|Progress]] during atmospheric entry over Earth
STS-135 Space Shuttle Atlantis reentry seen from the ISS.jpg|Space Shuttle reentry
ISS-46 Soyuz TMA-17M reentry.jpg|Plume remains after reentry of a Soyuz</gallery>

==See also==
{{div col}}
*{{annotated link|Astronomy:Van Allen radiation belt}}
*{{annotated link|Engineering:Aerocapture}}
*{{annotated link|Engineering:Intercontinental ballistic missile}}
*{{annotated link|Engineering:Lander (spacecraft)}}
*{{annotated link|Engineering:Landing footprint}}
*{{annotated link|Astronomy:List of reentering space debris}}
*{{annotated link|Engineering:NASA reentry prototypes}}
*{{annotated link|Engineering:Space capsule}}
*{{annotated link|Engineering:Space Shuttle thermal protection system}}
*{{annotated link|Engineering:Paper plane launched from space}}
{{div col end}}

== References ==
{{Reflist}}
<!-- Dead note "congress_88": 88th Congress, First Session, Part 3, House Appropriations Committee (Independent Appropriations for 1964). -->
<!-- Dead note "rainey": Rainey, Larry B., "Space Modelling and Simulation Roles and Applications Throughout the System Lifecycle," [[Organization:American Institute of Aeronautics and Astronautics|American Institute of Aeronautics and Astronautics]] (2004) ISBN 18849891512{{Please check ISBN|reason=Invalid length.}}. -->

==Further reading==
*{{cite book |last1=Launius |first1=Roger D. |last2=Jenkins |first2=Dennis R. |title=Coming Home: Reentry and Recovery from Space |date=October 10, 2012 |publisher=NASA |isbn=9780160910647 |oclc=802182873 |url=http://www.nasa.gov/connect/ebooks/coming_home_detail.html |access-date=August 21, 2014}}
*{{cite book|  last=Martin| first= John J.| title=Atmospheric Entry – An Introduction to Its Science and Engineering| publisher=Prentice-Hall| location=Old Tappan, New Jersey| year=1966}}
*{{cite book| last=Regan| first=Frank J.| title=Re-Entry Vehicle Dynamics (AIAA Education Series)| publisher=American Institute of Aeronautics and Astronautics, Inc.| location=New York| isbn= 978-0-915928-78-1| year=1984}}
*{{cite book|  last=Etkin| first=Bernard| title=Dynamics of Atmospheric Flight| publisher=John Wiley & Sons, Inc.| location=New York| isbn=978-0-471-24620-6| year=1972}}
*{{cite book| last=Vincenti| first=Walter G. |author2=Kruger Jr, Charles H. | title=Introduction to Physical Gas Dynamics| publisher=Robert E. Krieger Publishing Co.| location=Malabar, Florida| isbn=978-0-88275-309-6| year=1986}}
*{{cite book| last=Hansen| first=C. Frederick| title=Molecular Physics of Equilibrium Gases, A Handbook for Engineers| id=NASA SP-3096| publisher=NASA| year=1976| bibcode=1976mpeg.book.....H}}
*{{cite book| last=Hayes| first=Wallace D.|author2=Probstein, Ronald F. | title=Hypersonic Flow Theory| publisher=Academic Press| location=New York and London| year=1959}} A revised version of this classic text has been reissued as an inexpensive paperback: {{cite book|last=Hayes| first=Wallace D.| title=Hypersonic Inviscid Flow| publisher=Dover Publications| location=Mineola, New York| isbn=978-0-486-43281-6| date=1966}} reissued in 2004
*{{cite book| last=Anderson| first=John D. Jr.| title=Hypersonic and High Temperature Gas Dynamics| publisher=McGraw-Hill, Inc.| location=New York| isbn=978-0-07-001671-2| year=1989}}

==External links==
* [https://amat.readthedocs.io Aerocapture Mission Analysis Tool (AMAT)] provides preliminary mission analysis and simulation capability for atmospheric entry vehicles at various Solar System destinations.
* [http://www.aerospace.org/cords/ Center for Orbital and Reentry Debris Studies (The Aerospace Corporation)]
* [https://www.youtube.com/watch?v=aW5ozq4Tqew Apollo Atmospheric Entry Phase], 1968, NASA Mission Planning and Analysis Division, Project Apollo. video (25:14).
* [http://www.buran-energia.com/bourane/bourane-consti-bouclier.php Buran's heat shield]
* [https://web.archive.org/web/20110805043431/http://www.astronautix.com/fam/rescue.htm Encyclopedia Astronautica article] on the history of space rescue crafts, including some reentry craft designs.


{{DEFAULTSORT:Atmospheric entry}}
[[Category:Atmospheric entry| ]]
[[Category:Aerospace engineering]]

{{Sourceattribution|Atmospheric entry}}