Editing Atmospheric entry (section)

==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 [[dynamic pressure]] selects the TPS material. Heat load selects the thickness of the TPS material stack. Peak deceleration is of major importance for crewed missions. The upper limit for crewed return to Earth from low Earth orbit (LEO) or lunar return is 10''g''.<ref name=autogenerated1>{{Cite web |url=https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740007423_1974007423.pdf |title=Pavlosky, James E., St. Leger, Leslie G., "Apollo Experience Report - Thermal Protection Subsystem," NASA TN D-7564, (1974). |access-date=July 7, 2017 |archive-date=October 1, 2020 |archive-url=https://web.archive.org/web/20201001133219/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740007423_1974007423.pdf |url-status=live }}</ref> For Martian atmospheric entry after long exposure to zero gravity, the upper limit is 4''g''.<ref name=autogenerated1/> Peak dynamic pressure can also influence the selection of the outermost TPS material if [[spallation]] is an issue. The reentry vehicle's design parameters may be assessed through numerical simulation, including simplifications of the vehicle's dynamics, such as the [[planar reentry equations]] and heat flux correlations.<ref>{{Cite journal |last1=Sutton |first1=Kenneth |last2=Graves, Jr. |first2=Randolph A. |date=1971 |title=A general stagnation-point convective heating equation for arbitrary gas mixtures |url=https://ntrs.nasa.gov/api/citations/19720003329/downloads/19720003329.pdf |journal=NASA Tr R-376}}</ref>

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

Based upon Allen and Eggers discovery, maximum aeroshell bluntness (maximum drag) yields minimum TPS mass. Maximum bluntness (minimum ballistic coefficient) also yields a minimal [[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 ([[Huygens (spacecraft)|''Huygens'']] probe in a nitrogen atmosphere) for Mars entry (''[[Beagle 2]]'' in a carbon dioxide atmosphere).{{citation needed|date=September 2016}}{{original research inline|date=September 2016}} Prior to being abandoned, the [[Mars program|Soviet Mars lander program]] achieved one successful landing ([[Mars 3]]), on the second of three entry attempts (the others were [[Mars 2]] and [[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 [[Lander (spacecraft)|impactor]] and [[Pioneer Venus project|Pioneer Venus]] probes.