Advancement In Thermal Protection System Engineering Essay

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An improved TPS not only perform its primary function but also have so many improved characteristic such as durable, cost effectiveness, operable and low mass. Improved durability provides more resistance to damage and various impacts such as rain impact, highvelocity impact, etc. The mass of the TPS will gradually reduce the overall vehicle weight which in turn increases performance on the vehicle [81].

The TPS of the Space Shuttle Orbiter has several variations on the concept of an insulated structure. Reusable surface insulation (RSI) was represented by three rigid ceramic insulation materials [74]:

1. High-temperature reusable surface insulation (HRSI) consisting of black-coated LI-900 and LI-2200 tiles,

2. Low-temperature reusable surface insulation (LRSI) consisting of white-coated LI-900 tiles, and

3. Fibrous refractory composite insulation (FRCI-12) with black coating.

Kelly and Blosser [72] reported the results from experiments in which convective cooling was used to protect test panels from a severe aerothermodynamic environment. It was concluded that: ''Material properties, which are sometimes considered secondary (such as: conductivity, thermal expansion, ductility, fracture toughness, etc.) assume primary importance because of the impact of heat transfer and the wide range of temperatures encountered.

Rasky reported [75] that researchers at the Ames Research Center (NASA) have developed very high-temperature zirconium-based and hafnium-based CMCs. Tests have revealed that diborides of zirconium and of hafnium were the most oxidation-resistant of the high-temperature materials studied.The wing leading edges would be made of ceramic matrix composites such as zirconium diboride or coated carbon/carbon. Kelly and Blosser [72] reported the results from experiments in which convective cooling was used to protect test panels from a severe aerothermodynamics environment.The presently used coating on the Space Shuttle reuseable surface insulation is a reaction cured glaze (RCG). Low resistance to physical impact is the major disadvantage of reaction cured glaze [92].

A new Adaptable, Robust, Metallic, Operable, Reusable (ARMOR) thermal protection system (TPS) concept has been designed, analyzed and fabricated by group of Research scholar. In addition to the inherent, tailorable robustness of metallic TPS, the ARMOR TPS offers improved features based on lessons learned from previous metallic TPS development efforts. ARMOR TPS panels may be attached directly to a smooth substructure or supported on TPS support structure ARMOR TPS provides an attractive solution for the next generation of reusable launch vehicles that are striving for economic viability. The robust ARMOR TPS panels offer the potential to greatly reduce maintenance costs and increase the range of weather conditions acceptable for flight compared to competing TPS alternatives [83].

The Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions was proposed and to create a high strain-to-failure TPS with dramatic reduction in costand complexity. In this work, Carbon Phenolic Matrix and Carbon Silicon Matrix based ablative material were used, which yield good result as compare to pervious ablative TPS [84].

A new approach was introduce for the development of a single, multifunctional protective shield, employing nanotechnology-based materials, to serve simultaneously as a TPS, an impact shield and as the first line of defense against radiation. The approach is first to choose low molecular weight ablative TPS materials, (existing one) and add functionalized carbon nanotubes. Together they provide both thermal and radiation (TR) shielding [90]

Another attempt was made to develop a 100 mission reusable thermal protection system for the Space Shuttle Orbiter has led to the investigation of lightweight, refractory ceramic insulation materials. A new closed-cell glass-ceramic foam insulation material (closed pore insulation) was developed from sintering low cost hollow alumino-silicate glass microspheres (Cenosphere) which was obtained from fly ash residues with high emittance binders. With the natural Cenosphere Cobalt Oxide (CoO) was introduced. The outstanding characteristics of this material are superior mechanical properties when compared to rigidized fibrous insulations, rigidity, water repellancy, high emittance, machinability and good thermal shock resistance. Thermal expansion increased with cobalt additions due to the increasing concentration of cobalt aluminate in the glass. Total normal emissivity increased with cobalt additions over the range of temperatures investigated (RT-1700K). Increasing cobalt additions appeared to reduce the thermal conductivity of the foam at elevated temperatures. This may be due to the increased radiation absorption contribution of the cobalt ion in the glass. Differential thermal analysis has shown no physical or chemical changes occurring over the range of temperatures investigated (RT-1700K). Increasing cobalt oxide levels caused a reduction in the softening point of the glass, resulting in corresponding large decreases in elastic modulus and strength above 1135K with large increases in ductility. The Creep deflections depended strongly on temperature, time and cobalt level; the lower cobalt level being the most creep resistant [85-88].

An attempt was made to use Cured Thermal Protection Material Blocks in Heat Shield for re-entry vehicles. These Cured blocks can be applied directly over the aircraft's external surface because these materials are bonded into a compatible structure with large-cell honeycomb matrix. Also the use of uniform-sized thermal protection material blocks and the use of a faceted subsurface under the honeycomb will increase the effectiveness of thermal protection system and also reduce protection cost to certain extern [89].

A ceramic composite tile structure was developed specially for thermal protection system (TPS) in such manner that these materials can directly applied over the surface of an aerospace vehicle. This composite structure comprising of a ceramic lower (base) tile, a ceramic coated/ ceramic fibre upper shell and one or more layers of flexible rigid [92].

Two year research program was conducted to develop and evaluate the high temperature insulations and packaging materials for use under the radiative metal heat shield for the reusable shuttle vehicle. Here the insulation concept used was typically low density refractory fibre felts, enclosed in thin metal foil packages. In this research, simulated shuttle mission cycle tests were conducted on a variety of insulations and metal foil packaging with maximum service temperatures of 1000°F, 22000F, and 2500°F.The Packaging Material and insulation materials used here are as follows, for service up to 1000F, titanium foil and glass fibre was satisfactory; for service to 2000F, Inconel 702 was satisfactory. For slightly lower temperatures (1800F), Hastelloy X; Inconel and silica fibre such as Astroquartz and Refrasil A-100 601 were satisfactory. Of the superalloys, only TD-NiCr and Fiberfrax SKX (Ceramic Fibre) provided good service at 22000F. For 25000F service, coated columbium showed good result [93-94].

Hafnium Diboride 9 HfBr2) and Zirconium Diboride (ZrBr2) are relatively new ceramic materials which has good potential to replace traditional thermal protection systems (TPS) in re-entry vehicles and hypersonic aircraft. This composite material shows good oxidation resistance at temperatures up to 2000 ° C and possess high melting point of around 3200 ° C. Hence, these materials are put into Ultra High Temperature Ceramic Composite (UHTC) Category [95-96].

Carbon/Silicon Carbide ceramic matrix composite (C/SiC) consists of carbon fibre reinforced with silicon carbide which is manufactured by liquid polymer infiltration techniques. This composite retains a relatively constant strength to weight ratio up to 1600 ° C and is lightweight due to the use of low density carbon fibre [97].

A metal matrix composite is produced by using TAZ-8A super alloy as reinforced material with high temperature ceramic material such as SiCs matrix. TAZ-8A superalloy is a combination of exotic alloying elements such as Columbium, Tantalum and Molybdenum in certain ratio.TAZ-8A was selected as it possessed the least amount of oxidation amongst the family of nickel superalloys and also have high melting point [98].

Titanium Metal Matrix Composite (TMC) has capable of withstanding thousands of hours of cyclic loading in an oxidative and high temperature environment (Stephens 1987). This makes TMC excellent materials for its use in hypersonic propulsion and structural components since temperatures can reach as high as 1600 ° C [99].

In Reinforced Carbon-Carbon (RCC) Thermal Protection System, several applications of overcoat sealers were required to be applied on the RCC surface to prevent severe oxidation damage while re-entry. This feature made the RCC Thermal Protection System as a failure one []. To overcome this difficulty and make it suitable for hypersonic flights, modified coating is applied over it to prevent oxidation at temperatures exceeding 2000°C. The coating consists of a SiC substrate, then chemical vapor deposited SiC layer followed by layer of Hafnium Carbide [100-101].

Another attempt was made to use reusable phase change material (PCM) as a thermal protection system. This system comprises of a thermally conductive casting, PCM, thermally conductive open cell foam and heat pipes. Here heat is radiated away by changing of phase. The PCM change its phase twice, from solid to liquid to radiate the heat [102].

Materials used in TPS

The TPS covers essentially the entire aircraft surface, and based on the amount of required heat protection, seven different materials are used in different locations in typically system. In general, only oxide materials are more suitable for TPS because of their high degree of chemical stability in the maximum operating temperatures, 2400 to 3050°F [82].These materials are as follows [71&78]:

Reinforced carbon-carbon is used on the nose cap, the wing leading edges, and around the external tank and its structural attachment. RCC protects areas where temperatures exceed 1,260 °C (2,300 °F).

Black high-temperature reusable surface insulation (HTRSI) 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 aircraft maneuvering system and reaction control system pods, 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. This HTRS insulation made of coated LI-900 Silica ceramics. The HRSI tiles protect areas where temperatures are below 1260 °C (2,300 F).

Black tiles also called fibrous refractory composite insulation (FRCI) were developed later in the thermal protection system program. FRCI tiles replace some of the HRSI tiles in selected areas of the aircraft.

Low-temperature reusable surface insulation (LTRSI) white tiles are used in selected areas of the fuselage, upper wing and vertical tail. These tiles protect only upto temperature 650°C (1,200 F). In order to have better thermal characteristics on vehicle, these tiles are coated with white surface coating.

An advanced flexible reusable surface insulation (AFRSI) is a type of composite structure which consists of composite quilted fabric insulation batting between two sewn layers of white fabric.  The AFRSI blankets provide reduced fabrication and installation time improved producibility and durability, and costs, and a weight reduction over that of the LRSI tiles. The AFRSI blankets protect areas where temperatures are below 650 °C (1,200 F).

White blankets made of coated Nomex reusable surface insulation are used on the portions of the upper wing surface, the upper payload bay doors, a portion of the OMS/RCS pods and portions of the mid fuselage and aft fuselage sides. The FRSI blankets protect from the temperatures only below 400 ° C (700 F).

Some other additional materials are used in special areas. These materials are metal for the control system fairings and elevon seal panels on the upper wing to elevon interface, thermal panes for the windows, silica cloth for thermal barriers, egress and ingress flight crew side hatch, mid fuselage vent doors and gap fillers around operable penetrations, such as main and nose landing gear doors, umbilical doors, forward RCS, RCS thrusters, , payload bay doors, rudder/speed brake, and OMS/RCS pods.


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 HRSI tiles are made of a low-density, high-purity silica 99.8-percent amorphous fiber (fibers derived from common sand, 1 to 2 micron 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 flight. Generally, the HRSI tiles are thicker at the forward areas of the aircraft 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 aircraft. 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.This tile cannot withstand airframe load deformation; therefore, stress isolation is necessary between the tiles and the aircraft structure. This isolation is provided by a strain isolation pad. SIPs isolate the tiles from the airframe'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 aircraft structure by an RTV process.The RTV silicon adhesive is applied to the aircraft's 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 entryand temperatures below minus 170 F on orbit. The tile/SIP bond is cured at room temperature under pressure applied by vacuum bags.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/aircraft umbilical doors, vent doors and vertical stabilizer leading edge. The remaining areas use tiles that weigh 9 pounds per cubic foot.


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 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 aircraft. 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 aircraft 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 andHRSI 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. TheAFRSI 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 aircraft by RTV silicon adhesive 0.20 inch thick. The very thin blue 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 aircraft 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 aircraft 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 aircraft's upper surfaces.


Thermal barriers are used in the closeout areas between various components of the aircraft 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 borosilicafibers 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 Macormachinable ceramic.