The thermal protection system consists of various materials applied externally to the outer structural skin of the orbiter to maintain the skin within acceptable temperatures, primarily during the entry phase of the mission. The orbiter's outer structural skin is constructed primarily of aluminum and graphite epoxy.
During entry, the TPS materials protect the orbiter outer skin from temperatures above 350 F. In addition, they are reusable for 100 missions with refurbishment and maintenance. These materials perform in temperature ranges from minus 250 F in the cold soak of space to entry temperatures that reach nearly 3,000 F. The TPS also sustains the forces induced by deflections of the orbiter airframe as it responds to the various external environments. Because the thermal protection system is installed on the outside of the orbiter skin, it establishes the aerodynamics over the vehicle in addition to acting as the heat sink.
Orbiter interior temperatures also are controlled by internal insulation, heaters and purging techniques in the various phases of the mission.
The TPS is a passive system consisting of materials selected for stability at high temperatures and weight efficiency. These materials are as follows:
1. Reinforced carbon-carbon is used on the wing leading edges; the nose cap, including an area immediately aft of the nose cap on the lower surface (chine panel); and the immediate area around the forward orbiter/external tank structural attachment. RCC protects areas where temperatures exceed 2,300 F during entry.
2. Black high-temperature reusable surface insulation tiles are used in areas on the upper forward fuselage, including around the forward fuselage windows; the entire underside of the vehicle where RCC is not used; portions of the orbital maneuvering system and reaction control system pods; the leading and trailing edges of the vertical stabilizer; wing glove areas; elevon trailing edges; adjacent to the RCC on the upper wing surface; the base heat shield; the interface with wing leading edge RCC; and the upper body flap surface. The HRSI tiles protect areas where temperatures are below 2,300 F. These tiles have a black surface coating necessary for entry emittance.
3. Black tiles called fibrous refractory composite insulation were developed later in the thermal protection system program. FRCI tiles replace some of the HRSI tiles in selected areas of the orbiter.
4. Low-temperature reusable surface insulation white tiles are used in selected areas of the forward, mid-, and aft fuselages; vertical tail; upper wing; and OMS/RCS pods. These tiles protect areas where temperatures are below 1,200 F. These tiles have a white surface coating to provide better thermal characteristics on orbit.
5. After the initial delivery of Columbia from Rockwell International's Palmdale assembly facility, an advanced flexible reusable surface insulation was developed. This material consists of sewn composite quilted fabric insulation batting between two layers of white fabric that are sewn together to form a quilted blanket. AFRSI was used on Discovery and Atlantis to replace the vast majority of the LRSI tiles. Following its seventh flight, Columbia also was modified to replace most of the LRSI tiles with AFRSI. The AFRSI blankets provide improved producibility and durability, reduced fabrication and installation time and costs, and a weight reduction over that of the LRSI tiles. The AFRSI blankets protect areas where temperatures are below 1,200 F.
6. White blankets made of coated Nomex felt reusable surface insulation are used on the upper payload bay doors, portions of the midfuselage and aft fuselage sides, portions of the upper wing surface and a portion of the OMS/RCS pods. The FRSI blankets protect areas where temperatures are below 700 F.
7. Additional materials are used in other special areas. These materials are thermal panes for the windows; metal for the forward reaction control system fairings and elevon seal panels on the upper wing to elevon interface; a combination of white- and black-pigmented silica cloth for thermal barriers and gap fillers around operable penetrations, such as main and nose landing gear doors, egress and ingress flight crew side hatch, umbilical doors, elevon cove, forward RCS, RCS thrusters, midfuselage vent doors, payload bay doors, rudder/speed brake, OMS/RCS pods and gaps between TPS tiles in high differential pressure areas; and room-temperature vulcanizing material for the thick aluminum T-0 umbilicals on the sides of the orbiter aft fuselage.
RCC fabrication begins with a rayon cloth graphitized and impregnated with a phenolic resin. This impregnated cloth is layed up as a laminate and cured in an autoclave. After being cured, the laminate is pyrolized to convert the resin to carbon. This is then impregnated with furfural alcohol in a vacuum chamber, then cured and pyrolized again to convert the furfural alcohol to carbon. This process is repeated three times until the desired carbon-carbon properties are achieved.
To provide oxidation resistance for reuse capability, the outer layers of the RCC are converted to silicon carbide. The RCC is packed in a retort with a dry pack material made up of a mixture of alumina, silicon and silicon carbide. The retort is placed in a furnace, and the coating conversion process takes place in argon with a stepped-time-temperature cycle up to 3,200 F. A diffusion reaction occurs between the dry pack and carbon-carbon in which the outer layers of the carbon-carbon are converted to silicon carbide (whitish-gray color) with no thickness increase. It is this silicon-carbide coating that protects the carbon-carbon from oxidation. The silicon-carbide coating develops surface cracks caused by differential thermal expansion mismatch, requiring further oxidation resistance. That is provided by impregnation of a coated RCC part with tetraethyl orthosilicate. The part is then sealed with a glossy overcoat. The RCC laminate is superior to a sandwich design because it is light in weight and rugged; and it promotes internal cross-radiation from the hot stagnation region to cooler areas, thus reducing stagnation temperatures and thermal gradients around the leading edge. The operating range of RCC is from minus 250 F to about 3,000 F. The RCC is highly resistant to fatigue loading that is experienced during ascent and entry.
The RCC panels are mechanically attached to the wing with a series of floating joints to reduce loading on the panels caused by wing deflections. The seal between each wing leading edge panel is referred to as a T-seal. The T-seals allow for lateral motion and thermal expansion differences between the RCC and the orbiter wing. In addition, they prevent the direct flow of hot boundary layer gases into the wing leading edge cavity during entry. The T-seals are constructed of RCC.
Since carbon is a good thermal conductor, the adjacent aluminum and the metallic attachments must be protected from exceeding temperature limits by internal insulation. Inconel 718 and A-286 fittings are bolted to flanges on the RCC components and are attached to the aluminum wing spars and nose bulkhead. Inconel-covered cerachrome insulation protects the metallic attach fittings and spar from the heat radiated from the inside surface of the RCC wing panels.
The nose cap thermal insulation ues a blanket made from ceramic fibers and filled with silica fibers. HRSI or FRCI tiles are used to protect the forward fuselage from the heat radiated from the hot inside surface of the RCC.
During flight operations, damage has occurred in the area between the RCC nose cap and the nose landing gear doors from impact during ascent and excess heat during entry. The HRSI tiles in this area are to be replaced with RCC.
In the immediate area surrounding the forward orbiter/ET attach point, an AB312 ceramic cloth blanket is placed on the forward fuselage. RCC is placed over the blanket and is attached by metal standoffs for additional protection from the forward orbiter/ET attach point pyrotechnics.
The HRSI tiles are made of a low-density, high-purity silica 99.8-percent amorphous fiber (fibers derived from common sand, 1 to 2 mils thick) insulation that is made rigid by ceramic bonding. Because 90 percent of the tile is void and the remaining 10 percent is material, the tile weighs approximately 9 pounds per cubic foot. A slurry containing fibers mixed with water is frame-cast to form soft, porous blocks to which a collodial silica binder solution is added. When it is sintered, a rigid block is produced that is cut into quarters and then machined to the precise dimensions required for individual tiles.
HRSI tiles vary in thickness from 1 inch to 5 inches. The variable thickness is determined by the heat load encountered during entry. Generally, the HRSI tiles are thicker at the forward areas of the orbiter and thinner toward the aft end. Except for closeout areas, the HRSI tiles are nominally 6- by 6-inch squares. The HRSI tiles vary in sizes and shapes in the closeout areas on the orbiter. The HRSI tiles withstand on-orbit cold soak conditions, repeated heating and cooling thermal shock and extreme acoustic environments (165 decibels) at launch.
For example, an HRSI tile taken from a 2,300 F oven can be immersed in cold water without damage. Surface heat dissipates so quickly that an uncoated tile can be held by its edges with an ungloved hand seconds after removal from the oven while its interior still glows red.
The HRSI tiles are coated on the top and sides with a mixture of powdered tetrasilicide and borosilicate glass with a liquid carrier. This material is sprayed on the tile to coating thicknesses of 16 to 18 mils. The coated tiles then are placed in an oven and heated to a temperature of 2,300 F. This results in a black, waterproof glossy coating that has a surface emittance of 0.85 and a solar absorptance of about 0.85. After the ceramic coating heating process, the remaining silica fibers are treated with a silicon resin to provide bulk waterproofing.
Note that the tiles cannot withstand airframe load deformation; therefore, stress isolation is necessary between the tiles and the orbiter structure. This isolation is provided by a strain isolation pad. SIPs isolate the tiles from the orbiter's structural deflections, expansions and acoustic excitation, thereby preventing stress failure in the tiles. The SIPs are thermal isolators made of Nomex felt material supplied in thicknesses of 0.090, 0.115 or 0.160 inch. SIPs are bonded to the tiles, and the SIP and tile assembly is bonded to the orbiter structure by an RTV process.
Nomex felt is a basic aramid fiber. The fibers are 2 deniers in fineness, 3 inches long and crimped. They are loaded into a carding machine that untangles the clumps of fibers and combs them to make a tenuous mass of lengthwise-oriented, relatively parallel fibers called a web. The cross-lapped web is fed into a loom, where it is lightly needled into a batt. Generally, two such batts are placed face-to-face and needled together to form felt. The felt then is subjected to a multineedle pass process until the desired strength is reached. The needled felt is calendered to stabilize at a thickness of 0.16 inch to 0.40 inch by passing through heated rollers at selected pressures. The calendered material is heat-set at approximately 500 F to thermally stabilize the felt.
The RTV silicon adhesive is applied to the orbiter surface in a layer approximately 0.008 inch thick. The very thin bond line reduces weight and minimizes the thermal expansion at temperatures of 500 F during entry and temperatures below minus 170 F on orbit. The tile/SIP bond is cured at room temperature under pressure applied by vacuum bags.
Since the tiles thermally expand or contract very little compared to the orbiter structure, it is necessary to leave gaps of 25 to 65 mils between them to prevent tile-to-tile contact. Nomex felt material insulation is required in the bottom of the gap between tiles. It is referred to as a filler bar. The material, supplied in thicknesses corresponding to the SIPs', is cut into strips 0.75 inch wide and is bonded to the structure. The filler bar is waterproof and temperature-resistant up to approximately 800 F, topside exposure.
SIP introduces stress concentrations at the needled fiber bundles. This results in localized failure in the tile just above the RTV bond line. To solve this problem, the inner surface of the tile is densified to distribute the load more uniformly. The densification process was developed from a Ludox ammonia-stabilized binder. When mixed with silica slip particles, it becomes a cement. When mixed with water, it dries to a finished hard surface. A silica-tetraboride coloring agent is mixed with the compound for penetration identification. Several coats of the pigmented Ludox slip slurry are brush-painted on the SIP/tile bond interface and allowed to air-dry for 24 hours. A heat treatment and other processing are done before installation. The densification coating penetrates the tile to a depth of 0.125 inch, and the strength and stiffness of the tile and SIP system are increased by a factor of two.
There are two different densities of HRSI tiles. The first weighs 22 pounds per cubic foot and is used in all areas around the nose and main landing gears, nose cap interface, wing leading edge, RCC/HRSI interface, external tank/orbiter umbilical doors, vent doors and vertical stabilizer leading edge. The remaining areas use tiles that weigh 9 pounds per cubic foot.
The FRCI tiles were developed by NASA's Ames Research Center, Mountain View, Calif., and were manufactured by Lockheed Missiles and Space Division, Sunnyvale, Calif.
The FRCI-12 HRSI tiles are a higher strength tile derived by adding AB312 (alumina-borosilicate fiber), called Nextel, to the pure silica tile slurry. Developed by the 3M Company of St. Paul, Minn., Nextel activates boron fusion and, figuratively, welds the micron-sized fibers of pure silica into a rigid structure during sintering in a high-temperature furnace. The resulting composite fiber refractory material composed of 20-percent Nextel and 80-percent silica fiber has entirely different physical properties from the original 99.8-percent-pure silica tiles. Nextel, with an expansion coefficient 10 times that of the 99.8-percent-pure silica, acts like a preshrunk concrete reinforcing bar in the fiber matrix.
The reaction-cured glass (black) coating of the FRCI-12 tiles is compressed as it is cured to reduce the coating's sensitivity to cracking during handling and operations. In addition to the improved coating, the FRCI-12 tiles are about 10 percent lighter than the HRSI tiles. The FRCI-12 HRSI tiles also have demonstrated a tensile strength at least three times greater than that of the HRSI tiles and a use temperature approximately 100 F higher than that of HRSI tiles.
The FRCI-12 HRSI tile manufacturing process is essentially the same as that for the 99.8-percent-pure silica HRSI tiles, the only change being in the wet-end prebinding of the slurry before it is cast. It also requires a higher sintering temperature. When the material is dried, a rigid block is produced. These blocks are cut into quarters and then machined to the precise dimensions required for each tile. The FRCI-12 tiles are the same 6- by 6-inch size as HRSI tiles and vary in thickness from 1 inch to 5 inches. They vary also in size and shape at the closeout areas and are bonded to the orbiter in essentially the same way as the HRSI tiles.
The FRCI-12 tiles are used to replace the HRSI 22-pound-per-cubic-foot tiles. The FRCI-12 tiles have a density of 12 pounds per cubic foot and provide improved strength, durability, resistance to coating cracking and weight reduction.
The LRSI tiles are of the same construction and have the same basic functions as the 99.8-percent-pure silica HRSI tiles, but they are thinner (0.2 to 1.4 inches) than HRSI tiles. Thickness is determined by the heat load encountered during entry. The 99.8-percent-pure silica LRSI tiles are manufactured in the same manner as the 99.8-percent-pure silica HRSI tiles, except that the tiles are 8- by 8-inch squares and have a white optical and moisture-resistant coating applied 10 mils thick to the top and sides. In addition, the white coating provides on-orbit thermal control for the orbiter. The coating is made of silica compounds with shiny aluminum oxide to obtain optical properties. The coated 99.8-percent-pure silica LRSI tiles are treated with bulk waterproofing similar to the HRSI tiles. LRSI tiles are installed on the orbiter in the same manner as the HRSI tiles. The LRSI tile has a surface emittance of 0.8 and a solar absorptance of 0.32.
Because of evidence of plasma flow on the lower wing trailing edge and elevon leading edge tiles (wing/elevon cove) at the outboard elevon tip and inboard elevon, the LRSI tiles are replaced with FRCI-12 and HRSI 22 tiles along with gap fillers on Discovery (OV-103) and Atlantis (OV-104). On Columbia (OV-102), only gap fillers are being installed in this area.
AFRSI blankets replace the vast majority of the LRSI tiles. AFRSI consists of a low-density fibrous silica batting that is made up of high-purity silica and 99.8-percent amorphous silica fibers (1 to 2 mils thick). This batting is sandwiched between an outer woven silica high-temperature fabric and an inner woven glass lower temperature fabric. After the composite is sewn with silica thread, it has a quiltlike appearance. The AFRSI blankets are coated with a ceramic collodial silica and high-purity silica fibers (referred to as C-9) that provide endurance. The AFRSI composite density is approximately 8 to 9 pounds per cubic foot and varies in thickness from 0.45 to 0.95 inch. The thickness is determined by the heat load the blanket encounters during entry. The blankets are cut to the planform shape required and bonded directly to the orbiter by RTV silicon adhesive 0.20 inch thick. The very thin glue line reduces weight and minimizes the thermal expansion during temperature changes. The sewn quilted fabric blanket is manufactured by Rockwell in 3- by 3-foot squares of the proper thickness. The direct application of the blankets to the orbiter results in weight reduction, improved producibility and durability, reduced fabrication and installation cost, and reduced installation schedule time.
FRSI is the same Nomex material as SIP. The FRSI varies in thickness from 0.160 to 0.40 inch depending on the heat load encountered during entry. It consists of sheets 3 to 4 feet square, except for closeout areas, where it is cut to fit. The FRSI is bonded directly to the orbiter by RTV silicon adhesive applied at a thickness of 0.20 inch. A white-pigmented silicon elastomer coating is used to waterproof the felt and provide required thermal and optical properties. The FRSI has an emittance of 0.8 and solar absorptance of 0.32. FRSI covers nearly 50 percent of the orbiter's upper surfaces.
Thermal barriers are used in the closeout areas between various components of the orbiter and TPS, such as the forward and aft RCS, rudder/speed brake, nose and main landing gear doors, crew ingress and egress hatch, vent doors, external tank umbilical doors, vertical stabilizer/aft fuselage interface, payload bay doors, wing leading edge RCC/HRSI interface, and nose cap and HRSI interface. The various materials used are white AB312 ceramic alumina borosilica fibers or black-pigmented AB312 ceramic fiber cloth braided around an inner tubular spring made from Inconel 750 wire with silica fibers within the tube, alumina mat, quartz thread and Macor machinable ceramic.
Where surface pressure gradients would cause cross flow of boundary layer air within the intertile gaps, tile gap fillers are provided to minimize heating. The tile gap filler materials consist of white AB312 fibers or a black-pigmented AB312 cloth cover containing alumina fibers. These materials are used around the leading edge of the forward fuselage nose cap, windshields and side hatch, wing, trailing edge of elevons, vertical stabilizer, rudder/speed brake, body flap and heat shield of the shuttle's main engines.
Fused silica threaded inserts and plugs are used in the tiles to provide access for door removal or panel attachment.
Each TPS tile has an identification code that is painted in yellow. The paint does not burn off during entry and can be obtained commerically under the name Spearex.
Flags and letters are painted on the orbiter with a Dow Corning 3140 silicon-base material colored by adding pigments. It is basically the same paint used to paint automobile engines and will break down in temperature ranges between 800 to 1,000 F.
After each flight, the orbiter thermal protection system is rewaterproofed. Dimethylethoxysilane is injected into each tile through an existing hole in the surface coating with a needleless gun, and the AFRSI blankets are injected with DMES from a needle gun.
The contractors for the thermal protection system are Vought Corporation, Dallas, Texas (RCC); Lockheed Missiles and Space Co. Inc., Sunnyvale, Calif. (HRSI and LRSI tiles and HRSI FRCI-12 tiles); Albany International Research Co., Dedham, Mass. (Nomex felt); General Electric, Waterford, N.Y. (room-temperature vulcanizing adhesive); 3M Company, St. Paul, Minn. (AB312 fibers); Santa Fe Textiles, Santa Ana, Calif. (Inconel 750 wire spring and fabric sleeving); ICI United States Inc., Wilmington, Del. (alumina mat); J.P. Stevens Co., Los Angeles, Calif. (quartz thread); Corning Glass Works, Corning, N.Y. (Macor machinable glass ceramic); Velcro Corp., New York, N.Y. (Velcro hooks and loops); Prodesco, Perkasie, Pa. (fibrous pile-S glass); Johns Manville, Waterville, Ohio (high-purity silica glass); and Rockwell International, Downey, Calif. (AFRSI quilted fabric).
Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Wednesday October 11 17:43:09 EDT 1995
Jim Dumoulin (firstname.lastname@example.org)
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