Objects entering an atmosphere experience [[Drag (physics)|atmospheric drag]], which puts mechanical stress on the object, and [[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 ([[ablation]]) or even complete disintegration of smaller objects, and objects with lower [[compressive strength]] can explode.

 

 

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 [[Space diving|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.

 

 

==History==

==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 [[wikt:Appendix:Glossary of atmospheric reentry|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).

===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)]]

 

Pure spherical entry vehicles were used in the early Soviet [[Vostok programme|Vostok]] and [[Voskhod programme|Voskhod]] [[space capsule|capsule]]s and in Soviet Mars and [[Venera]] descent vehicles. The [[Apollo command module]] used a spherical section forebody heat shield with a converging conical afterbody. It flew a lifting entry with a hypersonic trim angle of attack of −27° (0° is blunt-end first) to yield an average L/D (lift-to-drag ratio) of 0.368.<ref>{{Cite web |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |title=Hillje, Ernest R., "Entry Aerodynamics at Lunar Return Conditions Obtained from the Flight of Apollo 4 (AS-501)," NASA TN D-5399, (1969). |access-date=July 7, 2017 |archive-date=September 16, 2020 |archive-url=https://web.archive.org/web/20200916020329/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |url-status=live }}</ref> The resultant lift achieved a measure of cross-range control by offsetting the vehicle's center of mass from its axis of symmetry, allowing the lift force to be directed left or right by rolling the capsule on its [[Flight control surfaces#Longitudinal axis|longitudinal axis]]. Other examples of the spherical section geometry in crewed capsules are [[Soyuz programme|Soyuz]]/[[Zond program|Zond]], [[Project Gemini|Gemini]], and [[Project Mercury|Mercury]]. Even these small amounts of lift allow trajectories that have very significant effects on peak [[g-force]], reducing it from 8–9&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>

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>

 

 

===Shock layer gas physics ===

Almost all aeronautical engineers are taught the [[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.

 

 

====Real (equilibrium) gas model====

 

====Real (non-equilibrium) gas model====

 

When running a Gibbs free energy equilibrium program,{{clarify|date=August 2018}} the iterative process from the originally specified molecular composition to the final calculated equilibrium composition is essentially random and not time accurate. With a non-equilibrium program, the computation process is time accurate and follows a solution path dictated by chemical and reaction rate formulas. The five species model has 17 chemical formulas (34 when counting reverse formulas). The Lighthill-Freeman model is based upon a single ordinary differential equation and one algebraic equation. The five species model is based upon 5 ordinary differential equations and 17 algebraic equations.{{Citation needed|date=December 2017}} Because the 5 ordinary differential equations are tightly coupled, the system is numerically "stiff" and difficult to solve. The five species model is only usable for entry from [[low Earth orbit]] where entry velocity is approximately {{cvt|7.8|km/s|km/h mph}}. For lunar return entry of 11&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.

==Thermal protection systems==

{{Main|Heat shield#Spacecraft}}

 

===Ablative===

=====PICA-3=====

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

 

====SIRCA====

| archive-url = https://web.archive.org/web/20210905190358/https://ntrs.nasa.gov/citations/20160001151

| url-status = live

 

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/>

In some early ballistic missile RVs (e.g., the Mk-2 and the [[sub-orbital spaceflight|sub-orbital]] [[Project Mercury|Mercury spacecraft]]), ''radiatively cooled TPS'' were used to initially absorb heat flux during the heat pulse, and, then, after the heat pulse, radiate and convect the stored heat back into the atmosphere. However, the earlier version of this technique required a considerable quantity of metal TPS (e.g., [[titanium]], [[beryllium]], [[copper]], etc.). Modern designers prefer to avoid this added mass by using ablative and thermal-soak TPS instead.

 

 

Radiatively cooled TPS can be found on modern entry vehicles, but [[reinforced carbon–carbon]] (RCC) (also called ''carbon–carbon'') is normally used instead of metal. RCC was the TPS material on the Space Shuttle's nose cone and wing leading edges, and was also proposed as the leading-edge material for the [[X-33]]. [[Carbon]] is the most refractory material known, with a one-atmosphere sublimation temperature of {{Convert|3825|C}} for graphite. This high temperature made carbon an obvious choice as a radiatively cooled TPS material. Disadvantages of RCC are that it is currently expensive to manufacture, is heavy, and lacks robust impact resistance.<ref>{{Cite web|url=https://history.nasa.gov/columbia/CAIB_reportindex.html|title=Columbia Accident Investigation Board|website=history.nasa.gov|access-date=July 12, 2017|archive-date=December 25, 2017|archive-url=https://web.archive.org/web/20171225231135/https://history.nasa.gov/columbia/CAIB_reportindex.html|url-status=live}}</ref>

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

 

 

In 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.

===Russia===

 

 

[[File:Inflatable Re-entry Vehicle Experiment.jpg|thumb|right|NASA engineers check 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 [[vacuum packing|vacuum-packed]] into a {{convert|15|in|adj=mid|cm|-diameter}} payload shroud and launched on a [[Black Brant (rocket)|Black Brant 9]] [[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 [[silicone]]-coated <nowiki>[</nowiki>[[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 }}, [[NASA]] Mission News, 2009-08-17, accessed 2011-01-02.</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}}

 

 

=== LOFTID ===

 

==Entry vehicle design considerations==

# Peak dynamic pressure

 

 

Starting from the principle of ''conservative design'', the engineer typically considers two [[Best, worst and average case|worst-case]] trajectories, the undershoot and overshoot trajectories. The overshoot trajectory is typically defined as the shallowest-allowable entry velocity angle prior to atmospheric [[Boost-glide|skip-off]]. The overshoot trajectory has the highest heat load and sets the TPS thickness. The undershoot trajectory is defined by the steepest allowable trajectory. For crewed missions the steepest entry angle is limited by the peak deceleration. The undershoot trajectory also has the highest peak heat flux and dynamic pressure. Consequently, the undershoot trajectory is the basis for selecting the TPS material. There is no "one size fits all" TPS material. A TPS material that is ideal for high heat flux may be too conductive (too dense) for a long duration heat load. A low-density TPS material might lack the tensile strength to resist spallation if the dynamic pressure is too high. A TPS material can perform well for a specific peak heat flux, but fail catastrophically for the same peak heat flux if the wall pressure is significantly increased (this happened with NASA's R-4 test spacecraft).<ref name=autogenerated1/> Older TPS materials tend to be more labor-intensive and expensive to manufacture compared to modern materials. However, modern TPS materials often lack the flight history of the older materials (an important consideration for a risk-averse designer).

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 [[Lander (spacecraft)|impactor]] and [[Pioneer Venus project|Pioneer Venus]] probes.

 

[[File:Ingreso reentrada.svg|center|thumb|800px|Reentry window

{{ordered list

*[[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.

*[[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]].

*[[Space Shuttle Atlantis|Space Shuttle ''Atlantis'']] [[STS-27#Tile damage|STS-27]]&nbsp;– Insulation from the starboard [[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]]

*[[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 [[Galileo Probe|''Galileo'' Probe]]). Consequently, the Genesis entry vehicle crashed into the desert floor. The payload was damaged, but most scientific data were recoverable.

*[[Soyuz TMA-11]]&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:

==Uncontrolled and unprotected entries==

{{See also|List of reentering space debris}}

 

 

On January 24, 1978, the [[Soviet Union|Soviet]] [[Kosmos 954]] ({{convert|3800|kg|disp=sqbr}}) reentered and crashed near [[Great Slave Lake]] in the [[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 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]] 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>

 

 

 

===Deorbit disposal===

File:Cross section of Gemini 2 heatshield.jpg|Cross section of Gemini 2 heat shield

</gallery>

 

==Gallery==

Soyuz TMA-05M capsule reentry.jpg|Early reentry [[Plasma (physics)|plasma]] trail (Soyuz)

File:Re-entry of Progress Spacecraft 42P - NASA Earth Observatory.jpg|[[Progress (spacecraft)|Progress]] during atmospheric entry over Earth

 

==See also==