Atmospheric reentry: Encyclopedia II - Atmospheric reentry - Thermal Protection Systems

Atmospheric reentry - Thermal Protection Systems

Atmospheric reentry - Ablative

The type of heat shield that best protects against high heat flux is the ablative heat shield. The ablative heat shield functions by lifting the hot shock layer gas away from the heat shield's outer wall (creating a cooler boundary layer) through blowing. The overall process of reducing the heat flux experienced by the heat shield's outer wall is called blockage. Ablation causes the TPS layer to char, melt, and sublimate through the process of pyrolysis. The gas produced by pyrolysis is what drives blowing and causes blockage of convective and catalytic heat flux. 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). Thermal protection can also be enhanced in some TPS materials through coking. Coking is the process of forming solid carbon on the outer char layer of the TPS. TPS coking was discovered accidentally during development of the Apollo-CM TPS material (Avcoat 5026-39).

Early research on ablation technology in the USA was centered at NASA's Ames Research Center, also known as Moffett Field, with ancillary work at other NASA facilities. Ames Research Center was ideal, since it had numerous wind tunnels capable of gererating varying wind velocities. Initial experiments typically mounted a mock-up of the ablative material to be analyzed within a hypersonic wind tunnel. The pyrolysis was measured in real time using thermogravimetric analysis, so that the ablative performance could be carefully evaluated.^[7]

The thermal conductivity of a TPS material is proportional to the material's density. 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 inappropriate and lower density TPS materials like SLA-561V, SIRCA or PICA can be better design choices.

PICA (Phenolic Impregnated Carbon Ablator) was the primary TPS material for the Stardust aeroshell. It was with Stardust that PICA first flew in space.

SIRCA (Silicone Impregnated Reuseable Ceramic Ablator) was used on the Backshell Interface Plate (BIP) of the Mars Pathfinder and Mars Exploration Rover (MER) aeroshells. The BIP was at the attachment points between the aeroshell's backshell (also called the "afterbody" or "aft cover") and the cruise ring. SIRCA was also the primary TPS material for the unsuccessful Deep Space 2 (DS/2) Mars probes.

"SLA" from SLA-561V stands for "Super Light weight Ablator". All of the 70 degree sphere-cone entry vehicles sent by NASA to Mars used SLA-561V as their primary TPS material. SLA-561V begins significant ablation at a heat flux of approximately 80 W/cm² but will fail for heat fluxes greater than 300 W/cm². SLA-561V would be unusable as an Apollo-CM TPS material for lunar return where the peak heat flux is around 497 W/cm². The peak heat flux experienced by the Viking-1 aeroshell which landed on Mars was 21 W/cm². For Viking-1, the TPS acted as a pure thermal insulator and never experienced significant ablation (an inappropriate design choice). However for the Mars Pathfinder aeroshell, the peak heat flux was 106 W/cm². SLA-561V was an appropriate design choice for Mars Pathfinder.

Atmospheric reentry - 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 soaks into the heat shield's outer wall and would eventually convect to the payload. This outcome is prevented by ejecting the heat shield (with its heat soak) prior to the heat convecting to the inner wall.

Thermal soak TPS is intended to shield mainly against heat load and not against a high peak heat flux (a long duration heat pulse of low intensity is assumed for the TPS design). The Space Shuttle orbit vehicle was designed with a reusable heat shield based upon a thermal soak TPS. A Space Shuttle's underside is coated with thousands of tiles made of silica foam that are intended to survive multiple reentries with only minor repairs between missions. When a Space Shuttle lands, there is a significant amount of heat stored in the TPS. Shortly after landing, a ground support cooling unit connects to the Space Shuttle's internal freon coolant loop to remove heat soaked in the TPS and orbiter structure.

Typical Space Shuttle's TPS tiles (LI-900) have remarkable thermal protection properties but are relatively brittle and break easily. An LI-900 tile could be exposed to a temperature of a 1000 K on one side, but merely warm to the touch on the other side. An impressive stunt that can be performed with a cube of LI-900 is to remove it glowing white hot from a furnace and then hold it with one's bare fingers without discomfort along the cube's edges (the author has done this).

Atmospheric reentry - Passively cooled

In some early ballistic missile RVs, e.g. the Mk-2 and the sub-orbital Mercury spacecraft, radiatively cooled TPS were used to initially store heat flux during the heat pulse and then latter radiate and convect the heat away from the vehicle. However the technique required a considerable quantity of metal TPS (e.g. titanium, beryllium, copper, etc.), adding greatly to the vehicle's mass. Consequently ablative and thermal soak TPS have been more common.

Some high-velocity aircraft, such as the SR-71 Blackbird and Concorde, had to deal with heating similar to that experienced by spacecraft but at much lower intensity. Studies of the SR-71's titanium skin revealed the metal structure was restored to its original strength through annealing due to aerodynamic heating. In the case of Concorde the nose was permitted to reach a maximum operating temperature of 127 °C (typically 180 °C warmer than the sub-zero ambient air).

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 titanium shingles. Unfortunately the earlier Shuttle TPS concept was rejected because it was incorrectly believed a silica tile based TPS offered less expensive development and manufacturing costs. A titanium shingle TPS was again proposed for the unsuccessful X-33 Single-Stage to Orbit (SSTO) prototype.

Atmospheric reentry - Actively cooled

Various advanced reusable spacecraft and hypersonic aircraft designs have been proposed to employ heat shields made from temperature-resistant metal alloys that incorporated a refrigerant or cryogenic fuel circulating through them. Such a TPS concept was proposed for the X-30 National Aerospace Plane (NASP). The NASP was supposed to have been a scramjet powered hypersonic aircraft but failed in development.

In the early 1960s various TPS systems were proposed to use water or other cooling liquid sprayed into the shock layer. Such concepts never got past the proposal phase since ordinary ablative TPS is much more reliable and mass efficient.

Adapted from the Wikipedia article "Thermal Protection Systems", under the G.N U Free Docmentation License. Please also see http://en.wikipedia.org/wiki