Thermal The Space Shuttle design presented many thermal insulation challenges. The system not only had to perform well, it had to integrate Protection with other subsystems. The Orbite r’s surfaces were exposed to Systems exceedingly high temperatures and needed reusable, lightweight, low-cost thermal protection. The vehicle also required low vulnerability to orbital debris and minimal thermal conductivity. NASA decided to Introduction bond the Orbite r’s thermal protection directly to its aluminum skin, Gail Chapline which presented an additional challenge. Orbiter Thermal Protection System Alvaro Rodriguez The External Tank required insulation to maintain the cryogenic fuels, Cooper Snapp Geminesse Dorsey liquid hydrogen, and liquid oxygen as well as to provide additional Michael Fowler structural integrity through launch and after release from the Orbiter. Ben Greene William Schneider The challenge and solutions that NASA discovered through tests and Carl Scott flight experience represent innovations that will carry into the next External Tank Thermal Protection System generation of space programs. Myron Pessin Jim Butler J. Scott Sparks Solid Rocket Motor Joint—An Innovative Solution Paul Bauer Bruce Steinetz Ice Detection Prevents Catastrophic Problems Charles Stevenson Aerogel-based Insulation System Charles Stevenson

182 Engineering Innovations Orbiter Thermal most difficult problems one can materials would have to be developed, imagine. It is certainly a problem that as the technology from the previous Protection System constitutes a challenge to the best brains Mercury, Gemini, and Apollo flights working in these domains of modern were only single-mission capable. Throughout the design and development aerophysics.” He was referring to Engineers embraced this challenge by of the Space Shuttle Orbiter Thermal protecting the intercontinental ballistic developing rigid silica/alumina fibrous Protection System, NASA overcame missile nose cones. Fifteen years later, materials that could meet the majority many technical challenges to attain a the shuttle offered considerably greater of heating environments on windward reusable system that could withstand the difficulties. It was vastly larger. Its surfaces of the Orbiter. On the nose high-temperature environments of thermal protection had to be reusable, cap and wing leading edge, however, re-entry into Earth’s atmosphere. and this thermal shield demanded both the heating was even more extreme. Theodore von Karman, the dean of light weight and low cost. The In response, a coated carbon-carbon American aerodynamicists, wrote in requirement for a fully reusable system composite material was developed to 1956, “Re-entry is perhaps one of the meant that new thermal protection

Thermal Protection System Could Take the Heat Orbiter remained protected during catalytic heating.

While the re-entry surface heating of the an exothermic reaction. Since the surface Orbiter to reject a majority of the chemical Orbiter was predominantly convective, acted as a catalyst, it was important that the energy. Engineers performed precise arc sufficient energy in the shock layer interfacing material/coating have a low jet measurements to quantify this effect dissociated air molecules and provided the propensity to augment the reaction. Atomic over a range of surface temperatures for potential for additional heating. As the air recombination influenced NASA’s selection both oxygen and nitrogen recombination. molecules broke apart and collided with the of glass-type materials, which have low This resulted in improved confidence in the surface of the vehicle, they recombined in catalycity and allowed the surface of the Thermal Protection System.

Flow

NitrogenNitrogen and oxygen molecules areare dissociated in the shock layer.layerr..

Atomsoms may recombinereecombine and form molecules on the vehicle surface.

Boundary Layer Shock Layer

Oxygen molecules Recombination of atoms on in the shock layer the surface of the vehicle separate into adds heat of dissociation to O+ and O- atoms. the Thermal PrProtectionrotection System.

Engineering Innovations 183 form the contours of these structural externally to the outer structural skin System to prevent a transition to components. NASA made an of the Orbiter to passively maintain the turbulent flow early in the flight when exhaustive effort to ensure these skin within acceptable temperatures, heating was at its highest. materials would operate over a large primarily during the re-entry phase Requirements for the Thermal spectrum of environments during of the mission. During this phase, the Protection System extended beyond launch, ascent, on-orbit operations, Thermal Protection System materials the nominal trajectories. For abort re-entry, and landing. protected the Orbiter’s outer skin from scenarios, the systems had to continue to exceeding temperatures of 176 °C perform in drastically different (350 °F). In addition, they were reusable Environments environments. These scenarios included: for 100 missions with refurbishment and Return-to-Launch Site; Abort Once During re-entry, the Orbiter’s external maintenance. These materials performed Around; Transatlantic Abort Landing; surface reached extreme temperatures— in temperatures that ranged from and others. Many of these abort up to 1,648°C (3,000°F). The Thermal -156°C (-250°F) in the cold soak of space scenarios increased heat load to the Protection System was designed to to re-entry temperatures that reached vehicle and pushed the capabilities of provide a smooth, aerodynamic surface nearly 1,648 °C (3, 000° F). The Thermal the materials to their limits. while protecting the underlying metal Protection System also withstood structure from excessive temperature. the forces induced by deflections The loads endured by the system of the Orbiter airframe as it responded Thermal Protection included launch acoustics, aerodynamic to various external environments. System Materials loading and associated structural At the vehicle surface, a boundary deflections, and on-orbit temperature Several types of Thermal Protection layer developed and was designed variations as well as natural System materials were used on the to be laminar—smooth, nonturbulent environments such as salt fog, wind, Orbiter. These materials included tiles, fluid flow. However, small gaps and and rain. In addition, the Thermal advanced flexible reusable surface discontinuities on the vehicle surface Protection System had to resist insulation, reinforced carbon-carbon, could cause the flow to transition from pyrotechnic shock loads as the Orbiter and flexible reusable surface laminar to turbulent, thus increasing separated from the External Tank (ET). insulation. All of these materials used the overall heating. Therefore, tight high-emissivity coatings to ensure The Thermal Protection System fabrication and assembly tolerances the maximum rejection of incoming consisted of various materials applied were required of the Thermal Protection convective heat through radiative heat

Orbiter Tile Placement System Configuration

Reinforced Carbon-Carbon Coating High-temperature Reusable Surface Insulation Tile Low-temperature Reusable Surface Insulation Tile Advanced Flexible Reusable Surface Insulation Blanket Flexible Reusable Surface Insulation Blanket

184 Engineering Innovations H Orbiter Tile Attachment System High-temperature Reusable Surface Insulation

Reaction-cured Tile-to-Tile Gap Glass Coating Tile Densi ed Layer

Filler Bar Koropon®-primed Structure Room-temperature Vulcanizing Adhesive Strain Isolation Pad transfer. Selection was based on the reusable surface insulation. Surface Fibrous Refractory Composite temperature on the vehicle. In areas coating constituted the primary Insulation tiles helped reduce the in which temperatures fell below difference between these two categories. overall weight and later replaced the approximately 1,260°C (2,300°F), High-temperature reusable surface LI-2200 tiles used around door NASA used rigid silica tiles or fibrous insulation tiles used a black borosilicate penetrations. Alumnia Enhanced insulation. At temperatures above glass coating that had an emittance Thermal Barrier was used in areas in that point, the agency used reinforced value greater than 0.8 and covered areas which small particles would carbon-carbon. of the vehicle in which temperatures damage fragile tiles. As part of the reached up to 1,260°C (2,300°F). post-Columbia Return to Flight Tiles Low-temperature reusable surface effort, engineers developed Boeing insulation tiles contained a white Rigidized Insulation. Overall, the The background to the shuttle’s tiles coating with the proper optical major improvements included lay in work dating to the early 1960s properties needed to maintain the reduced weight, decreased at Lockheed Missiles & Space appropriate on-orbit temperatures for vulnerability to orbital debris, and Company. A Lockheed patent vehicle thermal control purposes. minimal thermal conductivity. disclosure provided the first description The low-temperature reusable surface of a reusable insulation made of Orbiter tiles were bonded using strain insulation tiles covered areas of the ceramic fibers for use as a re-entry isolation pads and room-temperature vehicle in which temperatures reached vehicle heat shield. In other phased vulcanizing silicone adhesives. The up to 649°C ( 1,200°F). shuttle Thermal Protection System inner mold line of the tile was densified development efforts, ablatives and hot The Orbiter used several different prior to the strain isolation pad bond, structures were the early competitors. types of tiles, depending on thermal which aided in the uniform distribution However, tight cost constraints and a requirements. Over the years of of the stress concentration loads at the strong desire to build the Orbiter with the program, the tile composition tile-to-strain isolation pad interface. an aluminum airframe pointed toward changed with NASA’s improved The structure beneath the tile-to-tile the innovative, lightweight, and understanding of thermal conditions. gaps was protected by filler bar that reusable insulation material that could The majority of these tiles, prevented gas flow from penetrating be bonded directly to the airframe skin. manufactured by Lockheed Missiles into the tile bond line. NASA used gap & Space Company, were LI-900 fillers (prevented hot air intrusion and NASA used two categories of Thermal (bulk density of 144 kg/m 3 tile-to-tile contact) in areas of high Protection System tiles on the [9 pounds/ft 3]) and LI-2200 (bulk differential pressures, extreme Orbiter—low- and high-temperature density of 352 kg/m 3 [22 pounds/ft 3]).

Engineering Innovations 185 aero-acoustic excitations and to allowable strength for the interface a solution of colloidal silica particles passivate over-tolerance step and gap was approximately 50% less than to the non-coated tile underside and conditions. The structure used for the the LI-900 tile material used on the baked in an oven at 1,926°C (3,500°F) bonding surface was, for the most part, Orbiter. This reduction was caused by for 3 hours. The densified layer aluminum; however, several other stress concentrations in the reusable produced measured about 0.3 cm substrates used included graphite epoxy, surface insulation because of the (0 .1 in.) in thickness and increased beryllium, and titanium. formation of “stiff spots” in the strain the weight of a typical 15-by -1 5-cm isolation pad by the needling felting (6-by-6-in.) tile by only 2 7 grams Design Challenges process. Accommodating these stiff (0.06 pounds). For load distribution, Determining the strength properties spots for the more highly loaded tiles the densified layer served as a of the tile-to-strain isolation pad was met by locally densifying the structural plate that distributed the interface was no small feat. The underside of the tile. NASA applied concentrated strain isolation pad loads evenly into the weaker, unmodified reusable surface insulation tiles. NASA faced a greater structural Other Thermal Protection System design challenge in the creation of numerous unique tiles. It was Materials? NASA had it Covered. necessary to design thousands of these tiles that had compound curves, Flexible Reusable Surface Insulation interfaced with thermal barriers and White blankets made of coated Nomex ® Felt Reusable Surface Insulation protected hatches, and had penetrations for areas where surface temperatures fell below 371°C (700°F). The blankets were used on instrumentation and structural access. the upper payload bay doors, portions of the mid-fuselage, and on the aft fuselage sides. The overriding challenge was to ensure the strength integrity of the tiles had a Advanced Flexible Reusable Surface Insulation probability of tile failure of no greater After initial delivery of Columbia to the assembly facility, NASA developed an advanced than 1/10 8. To accomplish this flexible reusable surface insulation consisting of composite quilted fabric insulation magnitude of system reliability and still minimize the weight, it was batting sewn between two layers of white fabric. The insulation blankets provided necessary to define the detailed loads improved producibility and durability, reduced fabrication and installation time and and environments on each tile. To costs, and reduced weight. This insulation replaced the majority of low-temperature verify the integrity of the Thermal reusable surface insulation tiles on two of the shuttles: Discovery and Atlantis. Protection System tile design, each Following Columbia’s seventh flight, the shuttle was modified to replace most of the tile experienced stresses induced by low-temperature reusable surface insulation tiles on portions of the upper wing. the following combined sources: For Endeavour, the advanced flexible reusable surface insulation was directly built n Substrate or structure out-of-plane into the shuttle. displacement n Aerodynamic loads on the tile Additional Materials n Tile accelerations due to vibration NASA used additional materials in other areas of the Orbiter, such as in thermal glass and acoustics for the windows, Inconel ® for the forward Reaction Control System fairings, and elevon n Mismatch between tile and structure seal panels on the upper wing. Engineers employed a combination of white and black at installation pigmented silica cloth for thermal barriers and gap fillers around operable penetrations n Thermal gradients in the tile such as main and nose landing gear doors, egress and ingress flight crew side hatch, n Residual stress due to tile umbilical doors, elevon cove, forward Reaction Control System, Reaction Control manufacture System thrusters, mid-fuselage vent doors, payload bay doors, rudder/speed brake, n Substrate in-plane displacement and gaps between Thermal Protection System tiles in high differential pressure areas.

186 Engineering Innovations Reinforced Carbon-Carbon

The temperature extremes on the nose Orbiter Wing Panel Assembly cap and wing leading edge of the S Orbiter required a more sophisticated Access Panel material that would operate over a large Insulator spectrum of environments during launch, ascent, on-orbit operations, Reinforced Carbon-Carbon re-entry, and landing. Developed by Panel the Vought Corporation, Dallas, Texas, in collaboration with NASA, reinforced carbon-carbon formed the contours of the nose cap and wing leading edge structural components. Reinforced carbon-carbon is a composite made by curing graphite Attach fabric that has been pre-impregnated Fitting with phenolic resin laid up in complex shaped molds. After the parts are rough trimmed, the resin polymer is Reinforced converted to carbon by pyrolysis— Carbon-Carbon a chemical change brought about by T-seal the action of heat. The part is then Spanner Beam impregnated with furfuryl alcohol and Access Panel pyrolyzed multiple times to increase its density with a resultant improvement and coating have a different coefficient onto the surface; however, since there in its mechanical properties. of thermal expansion. Impregnating is sufficient viscosity, the sealant the carbon part with tetraethyl remains on the part. When the Since carbon oxidizes at elevated orthosilicate and applying a brush-on reinforced carbon-carbon cools down, temperatures, a silicon carbide coating sealant provides additional protection the glass fills back into the craze crack. is used to protect the carbon substrate. against oxygen paths to the carbon Any oxidation of the substrate directly from the craze cracks. Why Reinforced Carbon-Carbon? affects the strength of the material and, therefore—in the case of the Orbiter— The tetraethyl orthosilicate is applied The functionality of the reinforced had to be limited as much as possible to via a vacuum impregnation with the carbon-carbon is largely due to its ensure high performance over multiple intent of filling any remaining porosity ability to reject heat by external missions. Silicon carbide is formed within the part. Once the tetraethyl radiation (i.e., giving off heat from by converting the outer two plies of the orthosilicate has cured, a silicon surface to the surroundings) and carbon-carbon material through a dioxide residue coats the pore walls cross-radiation, which is the internal diffusion coating process, resulting in throughout the part, thus inhibiting reinforced carbon-carbon heat a stronger coating-to-substrate oxidation. After the tetraethyl transfer between the lower and upper interlaminar strength. orthosilicate process is complete, structures. Reinforced carbon-carbon a sodium silicate sealant is brushed has an excellent surface emissivity As a result of the silicon carbide onto the surface of the reinforced and can reject heat by radiating to formation, which occurs at carbon-carbon. The sealant fills in the space similar to the other Thermal temperatures of 1,648°C (3,000°F), craze cracks and, once cured, forms a Protection Systems. It is designed as craze cracks develop in the coating glass. The craze cracks close at high a shell section with an open interior on cool-down as the carbon substrate temperatures and the sealant will flow cavity that promotes cross-radiation.

Engineering Innovations 187 Since the highest heating is biased ensured that the metallic mechanisms traced to a change in maintenance of toward the lower surface, heat can be worked in concert with the hot structure the launch pad structure. Engineers cross-radiated to the cooler upper as a complete system in addition to altered the silica/cement topcoat over surfaces, thus reducing temperatures meeting the multi-mission requirements. the zinc primer such that zinc particles of the lower windward surface. were able to come into contact with Another benefit is that the thermal Reinforced Carbon-Carbon the wing leading edge and react with gradients across the part are minimized. Flight Experience Lessons Learned the silicon carbide coating during re-entry, thereby forming pinholes. While reinforced carbon-carbon is While NASA confirmed the NASA developed criteria for the designed to withstand high fundamental concepts and design pinholes as well as vacuum heat clean temperatures and maintain its structural sufficiency through the wing leading and repair methods. shape, the material has a relatively edge subsystem certification work and high thermal conductivity so it did early flight test phase of the Space not significantly inhibit the heat flow Shuttle Program, the agency also Improved Damage Assessment to reach the internal Orbiter wing identified design deficiencies. In most and Repair With Return to structure. The metallic attachments that cases, modifications rectified those Flight After Columbia Accident mated the reinforced carbon-carbon to deficiencies. These modifications the wing structure were crucial for included addressing the gap heating NASA performed rigorous testing and accommodating the thermal expansion between the reinforced carbon-carbon analysis on the Thermal Protection of reinforced carbon-carbon and and reusable surface insulation to System materials to adequately identify maintaining a smooth outer mold line inhibit hot gas flow-through and risks and to mitigate failure as much as of the vehicle. Protecting these retrofitting hardware to the wing practical. Engineers developed impact attachments and the spar structure itself leading edge subsystem design to testing, damage-tolerance assessments, required internal insulation. Incoflex ®, account for a substantial increase in and inspection and repair capabilities as an insulative batting encased by a thin the predicted airloads. With increasing part of the Return to Flight effort. Inconel ® foil, protected the metal design environment maturity, structural components from the internal temperature predictions on the attach Impact Testing cavity radiation environment. fittings were significantly lowered, The greatest lesson learned was that which allowed a design change from failure of the reinforced carbon-carbon Certification steel to titanium and a weight reduction and the catastrophic loss of the vehicle of 136 kg (300 pounds). Prior to the Orbiter’s first flight, NASA was caused by a large piece of foam performed extensive test and analysis to Over the 30 years of flight, the shuttle debris that was liberated from the ET. satisfy all requirements related to the encountered many anomalies that While modifications to the thermal natural and induced environments. The required investigative testing and protection foam on the tank reduced space agency accomplished certification analysis. Inspections revealed several the risk of shedding large debris of the wing leading edge subsystem cracks in the T-seals—i.e., components during launch, NASA still expected for flight by analyses verified with made of reinforced carbon-carbon that smaller-sized debris shedding. It was development and qualification tests fit between reinforced carbon-carbon critical that engineers understand the conducted on full-scale hardware. panels that allowed for thermal impact of foam shedding on the Engineers performed subscale testing expansion of those components while Orbiter’s wing leading edge and tiles. to establish thermal and mechanical keeping a smooth outer mold line. The Southwest Research Institute, properties, while full-scale testing The cracks were later found to be San Antonio, Texas, conducted many ensured the system performance and caused by convoluted plies from the of these impact tests to understand the provided the necessary data to correlate original layup of the T-seals. NASA important parameters that governed analytical models. This included a corrected the cracking by modifying structural failure of reinforced full-scale nose cap test article and twin the manufacturing techniques and carbon-carbon and tile materials. wing leading edge panel configuration implementing additional inspections. Additionally, NASA developed finite tested through multiple environments In 1993, the agency identified small element modeling capabilities to (i.e., acoustic/vibration, static loads, pinholes that went down to the carbon derive critical-damage thresholds. and radiant testing). Full-scale testing substrate and were subsequently

188 Engineering Innovations Tile Repair—A Critical Capability Was Developed

Prior to the first shuttle launch, NASA name of the material to Shuttle Tile Ablator, recognized the need for a capability to 865 kg/m 3 (54 pounds/ft 3) (STA-54). repair tiles on orbit. The loss of a tile during This material decreased the amount of launch due to an improper bond posed the swell during re-entry while maintaining greatest threat. In response, NASA a low enough viscosity to dispense prioritized the development of an ablative with the extravehicular activity hardware. material, MA-25S, for repairs of missing or The material did not harden and would damaged tiles. The biggest obstacle, remain workable for approximately 1 hour however, was finding a stable work but still cured within 24 hours in the platform. Thus, NASA cancelled the early on-orbit environments. repair effort in 1979. Simulating a damaged shuttle tile After the Columbia accident in 2003, NASA created dust that prevented the STA-54 prioritized tile repair capability. Prior to the from penetrating the surface of the tiles. Ground test of Orbiter tile repair. Columbia accident, the inspections after This led to the development of additional Finally, NASA performed an on-orbit every flight revealed damage greater than materials: a gel cleaning brush that was experiment during STS -1 23 (2008). Crew 2.5 cm ( 1 in.) in approximately 50 to 100 coated with a sticky silicone substance member Michael Foreman dispensed locations. The original ablative material used to clean tile dust from the repair STA-54 into several damaged tile formed the basis for the repair material cavity prior to filling; and primer material specimens. The on-orbit experiment was developed in the Return to Flight effort. that provided a contact surface to which a success, showing that the material the STA-54 could adhere. Once the primer Some reformulation of MA-25S began behaved exactly as it had during vacuum was cured, the bond strength was stronger in 2003. At that time, NASA changed the dispenses on the ground. than the shuttle tile.

Damage Tolerance Criteria different environments. Testing in an an imagery sensor package attached Arc Jet facility provided the closest to the Shuttle Robotic Arm was used to To make use of the inspection data, ground simulation for the temperature perform the inspection. The sensor NASA developed criteria for critical and chemical constituents of re-entry. package contained two laser imaging damage. Damage on reinforced Engineers performed numerous tests for systems and a high-resolution digital carbon-carbon ranged from spallation both reinforced carbon-carbon and tile camera. Additionally, astronauts residing (i.e., breaking up or reducing) of the to establish damage criteria and verify on the International Space Station (ISS) silicon carbide coating to complete newly developed thermal math models photographed the entire Orbiter as it penetration of the substrate. Tiles used for real-time mission support. executed an aerial maneuver, similar could be gouged by ascent debris to to a backflip, 182 m (600 ft) from the varying depths with a wide variety of ISS. The crew transmitted photographs cavity shapes. The seriousness of any Inspection Capability to Houston, Texas, where engineers given damage was highly dependent NASA developed an inspection on the ground evaluated the images for on local temperature and pressure capability to survey the reinforced any potential damage. environments. NASA initiated an carbon-carbon and tile surfaces. This extensive Arc Jet test program during NASA employed an additional capability provided images to assess Return to Flight activities to detection system to gauge threats from any potential impact damages from characterize the survivability of ascent and on-orbit impacts to the wing ascent and orbital debris. A boom with multiple damage configurations in leading edge. As part of preparing the

Engineering Innovations 189 Reinforced Carbon- Carbon Repair— Damage Control in the Vacuum of Space

Following the Space Shuttle Columbia accident in 2003, a group of engineers and scientists gathered at Johnson Space Center to discuss concepts for the repair of damaged reinforced carbon-carbon in the weightless vacuum environment of space. Few potential repair materials could Astronaut Andrew Thomas (left) watches as Charles Camarda tests the reinforced withstand the temperatures and pressures carbon-carbon plug repair (STS-114 [2005]). on the surface. Of those materials, few were developed this system and the NASA For larger damages, a plug repair system compatible with the space environment and repair team slightly modified it to optimize protected the reinforced carbon-carbon none had been tested in this type of its material properties for use in space. using a series of thin, flexible composite application. Thus, the team developed two Technicians used a modified commercial discs designed to fit securely against repair systems that were made available for caulk gun to apply the material to the the curvature of the surface. Engineers contingency use on the next flight. damaged wing. The material was spread developed 19 geometric shapes, which The first system—Non-Oxide Adhesive out over the damage using spatulas similar were flown to provide contingency Experimental—was designed to repair to commercial trowels. Once dried and repair capability. An attach mechanism coating damage or small cracks in cured by the sun, Non-Oxide Adhesive held the plugs in place. The anchor was reinforced carbon-carbon panels. This Experimental used the heat of re-entry to made up of a refractory alloy called pre-ceramic polymer had the consistency convert the material into a ceramic, which titanium zirconium molybdenum that was of a thick paste. COI Ceramics, Inc., protected exposed damage from extreme capable of withstanding the 1,648°C headquartered in San Diego, California, temperatures and pressures. (3,000°F) re-entry temperature.

Orbiter for launch, technicians placed Conclusion modify and upgrade both design and accelerometers on the spar aluminum materials, thus increasing the robustness The Orbiter Thermal Protection Systems structure behind the reinforced and safety of these critical systems on the shuttle proved to be effective, carbon-carbon panels at the attachment during the life of the program. Through with the exception of STS -107 (2003). locations. Forty-four sensors across the tragedy of the Columbia accident, On that flight, the catastrophic loss both wings detected accelerations NASA developed new inspection and was caused by a large piece of foam from potential impacts and relayed the repair techniques as protective measures debris that was liberated from the ET. data to on-board laptops, which could to ensure the success and safety of Advanced materials and coatings be transmitted to ground engineers. subsequent shuttle missions. were key in enabling the success of Using test-correlated dynamic the shuttle in high-temperature models, engineers assessed suspected environments. Experience gathered impacts for their level of risk based over many shuttle missions led the on accelerometer output. Thermal Protection Systems team to

190 Engineering Innovations External Tank Thermal tanks as well as the intertank—also retardant; a surfactant (which controls referred to as the tank “sidewalls.” surface tension and bubble or cell Protection System The other major component was a formation); and a catalyst (to enhance composite ablator material (a heat the efficiency and speed of the The amount of Thermal Protection shield material designed to burn away) polymeric reaction). The blowing System material on the shuttle’s made of silicone resins and cork. agent—originally chlorofluorocarbon External Tank (ET) could cover an (CF C)-11, then hydrochlorofluorocarbon NASA oversaw the development of acre. NASA faced major challenges (HCFC)-141b—created the foam’s the closed-cell foam to keep propellants in developing and improving cellular structure, making millions of at optimum temperature—liquid tank-insulating materials and processes tiny bubble-like foam cells. hydrogen fuel at -253°C (-423°F) and for this critical feature. Yet, the space liquid oxygen oxidizer at -182°C NASA altered the Thermal Protection agency’s solutions were varied and (-296°F)—while preventing a buildup System configuration over the course innovative. These solutions represented of ice on the outside of the tank, even of the Space Shuttle Program; however, a significant advance in understanding as the tank remained on the launch pad by 1995, ET performance requirements the use of Thermal Protection System under the hot Florida sun. led the program to baseline four materials as well as the structures, specially engineered closed-cell foams. aerodynamics, and manufacturing The foam insulation had to be durable The larger sections were covered in processes involved. enough to endure a 180-day stay at polyisocyanurate (an improved version the launch pad, withstand temperatures The tanks played two major roles of polyurethane) foam (NCFI 24 -1 24) up to 46°C (11 5°F) and humidity as during launch: containing and provided by North Carolina Foam high as 100%, and resist sand, salt fog, delivering cryogenic propellants to Industries. NCFI 24 -1 24 accounted for rain, solar radiation, and even fungus. the Space Shuttle Main Engines, and 77% of the total foam used on the tank During launch, the foam had to serving as the structural backbone and was sprayed robotically. A similar tolerate temperatures as high as 649°C for the attachment of the Orbiter and foam, NCFI 24-57, was sprayed (1 ,200°F) generated by aerodynamic Solid Rocket Boosters. The Thermal robotically on the aft dome of the liquid friction and rocket exhaust. As the Protection System, composed of hydrogen tank. Stepanfoam ® BX-265 tank reentered the atmosphere spray-on foam and hand-applied was sprayed manually on closeout approximately 30 minutes after insulation and ablator, was applied areas, exterior tank feedlines, and launch, the foam helped hold the tank primarily to the outer surfaces of the internal tank domes. The tank’s ablator, together as temperatures and internal tank. It was designed to maintain the Super-Lightweight Ablator (SLA)-5 61 , pressurization worked to break it up, quality of the cryogenic propellants, was sprayed onto areas subjected to allowing the tank to disintegrate safely protect the tank structure from ascent extreme heat, such as brackets and over a remote ocean location. heating, prevent the formation of ice other protuberances, and the exposed, (a potential impact debris source), Though the foam insulation on the exterior lines that fed the liquid and stabilize tank internal temperature majority of the tank was only about oxygen and liquid hydrogen to the during re-entry into Earth’s atmosphere, 2.5 cm (1 in.) thick, it added shuttle’s main engines. NASA used thus helping to maintain tank structural approximately 1,700 kg (3,800 pounds) Product Development Laboratory -1 034, integrity prior to its breakup within to the tank’s weight. Insulation on the a hand-poured foam, for filling a predicted landing zone. liquid hydrogen tank was somewhat odd-shaped cavities. thicker—between 3.8 and 5 cm Basic Configuration (1 .5 to 2 in.). The foam’s density varied with the type, but an average density Application Requirements NASA applied two basic types of was 38.4 kg/m 3 (2.4 pounds/ft 3). Application of the foam, whether Thermal Protection System materials to The tank’s spray-on foam was a automated or hand-sprayed, was the ET. One type was a low-density, polyurethane material composed of five designed to meet NASA’s requirements rigid, closed-cell foam. This foam was primary ingredients: an isocyanate for finish, thickness, roughness, sprayed on the majority of the tank’s and a polyol (both components of density, strength, adhesion, and size “acreage”—larger areas such as the the polymeric backbone); a flame and frequency of voids within the liquid hydrogen and liquid oxygen foam. The foam was applied in

Engineering Innovations 191

External Tank Thermal Protection Systems Materials

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The External Tank’s Thermal Protection System consisted of a number of different foam formulations displayed here. NASA selected materials for their insulating properties, and for their ability to withstand ascent aerodynamic forces.

specially designed, environmentally applied at Lockheed Martin’s Michoud As the ET was the only expendable controlled spray cells and sprayed Assembly Facility in New Orleans, part of the shuttle, NASA placed in several phases, often over a period Louisiana, where the tank was particular emphasis on keeping tank of several weeks. Prior to spraying, manufactured. Some closeout Thermal manufacturing costs at a minimum. engineers tested the foam’s raw Protection System was applied either by To achieve this objective, the agency material and mechanical properties hand or manual spraying at the Kennedy based its original design and to ensure the materials met NASA Space Center (KSC) in Florida. manufacturing plans on the use specifications. After the spraying was of existing, well-proven materials complete, NASA performed multiple Design and Testing and processes with a planned visual inspections of all foam evolution to newer products as they surfaces as well as tests of “witness” In the early 1970s, NASA developed became available. specimens in some cases. a spiral “barber pole” Thermal Protection System application The original baseline Thermal More than 90% of the foam was technique that was used through Protection System configuration called ® sprayed onto the tank robotically, the end of the program. This was an for the sprayable Stepanfoam BX-250 leaving 10% to be applied by manual early success for the ET Program, foam (used on the Saturn S-II stage) on spraying or by hand. Most foam was but many challenges soon followed. the liquid hygrogen sidewalls (acreage)

192 Engineering Innovations Solid Rocket Motor Joint —An Innovative Solution

Alliant Techsystems (ATK) Aerospace Systems, in partnership with NASA Glenn Research Center, developed a solution for protecting the temperature-sensitive O-rings used to seal the shuttle reusable solid rocket motor nozzle segments. The use of a Reusable Solid Rocket Motor carbon fiber material promoted safety and enabled joint assembly Aft Segment

in a fraction of the time required by previous processes, with Carbon Fiber Rope Thermal Barrier enhanced reproducibility.

The reusable solid rocket motors were fabricated in segments and pinned together incorporating O-ring seals. Similarly, nozzles consisted of multiple components joined and sealed at six joint locations using O-rings. A layer of rubber insulation, referred to as “joint fill” compound, kept the 3,038°C (5,500°F) combustion gases a safe distance away from these seals. In a few instances,

however, hot gases breached the compound, leaving soot within Nozzle-to- .

Case Joint d e

the joint. NASA modified the compound installation process and v r e s e r instituted reviews of postflight conditions. Although the s t h g i r

modifications proved effective, damage was still possible in the l l A

. K

unlikely event that gases breached the compound. T A

© ATK chose an innovative approach through emerging technologies. Using carbon fiber rope instead of rubber insulation in solid rocket motor Rather than attempt to prevent gas intrusion with manually nozzle joints simplified the joint assembly process and improved shuttle safety margins. applied rubber fill compound, the heat energy from internal gases would be extracted with a special joint filler and the O-ring seals structure allowed it to conform to tolerance assembly conditions. would be pressurized with the cooled gas. The thermal barrier provided flexibility and resiliency to ATK’s solution was based on a pliable, braided form of high- accommodate joint opening or closing during operation. Upon performance carbon material able to withstand harsh temperature pressurization, the thermal barrier seated itself in the groove to environments. The braided design removed most of the thermal obstruct hot gas flow from bypassing the barrier. energy from the gas and inhibited flow induced by pressure The carbon fiber solution increased Space Shuttle safety margins. fluctuations. The carbon fiber thermal barrier was easier to install Carbon fibers are suited to a nonoxidizing environment, and significantly reduced motor assembly time. withstanding high temperatures without experiencing degradation. In a rocket environment, carbon fibers withstood temperatures up The barrier provided a temperature drop across a single diameter, to 3,816°C (6,900°F). The braided structure and high surface reducing gas temperature to O-rings well below acceptable levels. area-to-mass ratio made the barrier an excellent heat exchanger The thermal barrier also kept molten alumina slag—generated while allowing a restricted yet uniform gas flow. The weave during solid fuel burn—from contacting and affecting O-rings.

and forward dome, and SLA-5 61 (used In the late 1970s, however, design of detaching from the ET. This caused a on the Viking Mars Lander) on the aft the Orbiter tiles advanced to the point reassessment of the Thermal Protection dome, intertank, and liquid oxygen where it became apparent that they System design to prevent the formation tank in the areas of high heating. were susceptible to damage from ice of ice anywhere on the tank forward

Engineering Innovations 193 of the liquid hydrogen tank aft-end NASA accepted ice formation on these plasma arc facility at NASA’s Ames structural ring frame. The Orbiter/ice brackets as unavoidable. Research Center in California, issue drove the requirement to cover which could deliver the required high While attempting to prevent ice the entire tank with Stepanfoam ® heating rates. Better understanding of buildup on the tank, NASA also BX-250, except for the high-heating ablation rates and the flow fields worked to characterize both the ablator aft dome, which remained SLA-5 61 . around ET protuberances permitted material and the foams for expected Ice was to be prevented on tank refinement of the Thermal Protection heating rates. NASA worked with pressurization lines through the use System configuration. Arnold Engineering Development of a heated purge. Certain liquid Center in Tennessee to modify its wind Another unique project was the testing oxygen feedline brackets, subject to tunnel to provide the capability to test of spray-on foam insulation on a extensive thermal contraction, could foam materials under realistic flight subscale tank, measuring 3 m (10 ft) in not be fully insulated without motion conditions. SLA-5 61 was tested in the diameter, in the environmental hanger breaking the insulation. Therefore, at Eglin Air Force Base, Florida. The insulated tank was filled with liquid nitrogen and subjected to various rain, wind, humidity, and temperature conditions to determine the rate of ice growth. These data were then converted to a computer program known as Surfice, which was used at KSC to predict whether unacceptable ice would form prior to launch. To provide information on application techniques, the agency ran cryogenic flexure tests that verified substrate adhesion and strength as well as crush tests on the Thermal Protection System materials.

A secondary function of the Thermal Protection System was to stabilize tank internal temperature In a continuous search for optimum during re-entry into Earth’s atmosphere, thus helping to maintain tank structural integrity prior to its Thermal Protection System breakup over a remote ocean location. performance, NASA—still in the Thermal Protection System design and testing phase—decided to use Chemical Products Research (CPR)-4 21 , a commercial foam insulation with good high-heating capability. Lockheed Martin developed a sprayable Thermal Protection System to apply to tank sidewalls and aft dome. Application needed a relative humidity of less than 30%, which resulted in the addition of a chemical dryer at Michoud. Also, the tank wall had to be heated to 60°C (1 40°F). This required passing hot gas through the tank while it was being rotated for the “barber pole” foam application mode. The key to the External Tank’s foam Thermal Protection System insulating properties was its cellular structure, creating millions of tiny bubble-like foam cells. The sprayed foam (NCFI 24-124) can be seen here after application to an area of the tank’s aluminum “acreage,” consisting of the liquid oxygen tank, liquid hydrogen tank, and intertank.

194 Engineering Innovations Ice Detection Prevents Catastrophic Problems

NASA had a potentially catastrophic pad. NASA and its partners initiated a problem with ice that formed on the program to develop a system capable of cryogenic-filled Space Shuttle External detecting ice on the External Tank spray-on Tank. Falling ice could have struck and foam insulation surfaces. This system was damaged the crew compartment windows, calibrated for those surfaces and used an reinforced carbon-carbon panels on the infrared strobe, a focal plane sensor array, wing leading edge of the Orbiter, or its and a filter wheel to collect successive thermal protection tiles, thus placing the images over a number of sub-bands. crew and vehicle at risk. The camera processed the images to determine whether ice was present, and it Kennedy Space Center and the US Army also computed ice thickness. The system Tank Automotive and Armaments Research, was housed in nitrogen-purged enclosures Robert Speece, NASA engineer, is shown Development and Engineering Center that were mounted on a two-wheeled operating the ice detection system at the pad, confirmed that a proof-of-concept system, prior to shuttle launch. portable cart. It was successfully applied tested by MacDonald, Dettwiler and to the inspection of the External Tank Associates Ltd. of Canada, offered potential The system can be used to detect ice on on STS-116 (2006), where the camera to support cryogenic tanking tests and any surface. It can also be used to detect detected thin ice/frost layers on two ice debris team inspections on the launch the presence of water. umbilical connections.

First Flight Approaches coefficient of thermal expansion of testing of foam samples on the roof of the ablator binder, as compared to the the Michoud Assembly Facility, As the Space Shuttle Program moved aluminum, would cause the ablator however, showed the damage to be so toward the first shuttle flight in 1981 , to shrink. This would introduce shallow that it was insignificant. NASA NASA faced another challenge. biaxial tension in the ablator and decided not to paint the tanks, resulting 2 2 Approximately 37 m (400 ft ) of corresponding shear forces at the bond in a weight savings of about 260 kg ablator became debonded from the line near any edges, discontinuities, (580 pounds), lowered labor costs, and tank’s aluminum surface the first time or cracks. Then, when the tank was the introduction of the “orange” tank. a tank was loaded with liquid hydrogen. pressurized, tank expansion from While the failure analysis was pressure would compound this shear Environmental Challenges inconclusive, it appeared that the force, possibly causing the bond line production team had tried to bond too to fail. NASA decided to pre-pressurize Knowledge of toxic properties and large an area and did not get the ablator the liquid hydrogen tank with helium environmental contaminations panels under the required vacuum gas prior to filling the tank for increased over the 30 years of the before the adhesive pot life ran out. launch—and to pressures higher than Space Shuttle Program. Federal laws Technicians at Michoud Assembly flight pressures—to stretch the ablator reflected these changes. For instance, Facility reworked the application when it was warm and elastic. ozone-depleting substances, including process for the ET at their facility and some Freon ® compounds, reduced the Because early test data showed the tank the first tank at KSC. protecting atmospheric ozone layer. insulation could be adversely affected NASA worked with its contractors to Following the ablator bonding by ultraviolet light, NASA painted the reduce both toxicity and environmental problem, NASA intensified its analysis first several tanks white, using a consequences for the cooling agents of the ablator/aluminum bond line. fire-retardant latex paint. Exposure This analysis showed that the higher and the foam compounds.

Engineering Innovations 195 During the 1990s, the University of Utah published data showing that Liquid CPR-4 21 was potentially toxic. Based Oxygen Tank on this analysis, Chemical Products Liquid Oxygen Research withdrew CPR-4 21 from the Feedline Feedline market. NASA’s ET office had Bracket Chemical Products Research reformulate this foam, with the new product identified as CPR-488. New challenges arose related to emerging environmental policies that Intertank necessitated changes to Thermal Protection System foam formulations. Intertank Flange In 1987, the United States adopted the Montreal Protocol on Substances Liquid Hydrogen that Deplete the Ozone Layer, Tank which provided for the eventual The foam’s approximately 2.5-cm (1-in.) thickness borders the circumferential flange that joins the international elimination of intertank with the liquid hydrogen tank. The ribbed area is the intertank, that, like the liquid oxygen tank ozone-depleting substances. The in the background and the liquid hydrogen tank in the foreground, was robotically sprayed with NCFI United States implemented the protocol 24-124 foam. The flange would later be hand-sprayed with Stepanfoam ® BX-265. The liquid oxygen by regulations under the Clean Air feedline at the top of the tank and a feedline bracket have been hand-sprayed with BX-265 foam. Act. Ozone-depleting substances, including CFC -11 —the Freo n® blowing agent used in the production of the Thermal Protection System sprayable foams for the tanks—were scheduled to be phased out of production. After the phaseout, CFC -11 would only be available for such uses through a rigorous exemption process. To prepare for the upcoming obsolescence of the foam blowing agent, Marshall Space Flight Center (MSFC) along with Lockheed Martin tracked and mitigated the effect of emerging environmental regulations. After extensive research and testing of potential substitutes, NASA proposed A technician at NASA’s Michoud Assembly Facility sprays the flange that connects the intertank and that HCFC-141b replace the CFC -11 liquid hydrogen tank. Stepanfoam ® BX-265 was sprayed manually on closeout areas, exterior tank blowing agent. NASA continued to use feedlines, internal tank domes, closeout areas of mating External Tank subcomponent surfaces, and stockpiled supplies of CFC -11 -blown small subcomponents. foam until the HCFC-141b foam was sidewall foam, CPR-488. North and MSFC developed an environmental certified for tank use and phased in Carolina Foam Industries reformulated test. This test used a flat aluminum beginning in 1996. CPR-488 and developed a new product. plate machined to match aft dome NASA undertook the development stress levels. The plate was attached As part of qualifying this new product , and qualification of a foam to be to a cryostat filled with liquid helium Lockheed Martin, Wyle Laboratories, phased in as a replacement for the tank and then strained with hydraulic jacks

196 Engineering Innovations to the flight biaxial stress levels. To save the weight of this ablator and was sprayed manually—with a Radiant heat lamps were installed to its associated cost, NASA had North CFC -11 blowing agent—on the tank’s match the radiant heating from the solid Carolina Foam Industries develop “closeout” areas. During STS -1 08 rocket motor plumes, and an acoustic a foam adequate for the aft dome (20 01 ), Stepanfoam ® BX-265—with horn blasted the test. This simulated environment without ablator. The foam HCFC -1 41b as its blowing agent— the aft dome ascent environment as was phased in on the aft dome, flying first flew as a replacement for BX-250. well as possible. The test results first on Space Transportation System BX-250 continued to be flown in indicated the need to spray ablator on (STS)-79 in 1996. The first usage of certain applications as BX-265 was the aft dome. To provide the capability the new foam on the tank sidewalls phased into the manufacturing process. to spray the ablator, personnel at was phased in over three tanks starting The use of HCFC-141b as a foam Michoud Assembly Facility built two with STS-85 in August 1997. blowing agent, however, was also spray cells, with an additional cell to Environmental Protection Agency problematic. It was classified as clean and prime the liquid hydrogen regulations also required NASA to a Class II ozone-depleting substance tank before ablator application. replace Stepanfoam ® BX-250, which and was subject to phaseout under the

Aerogel-based Insulation System Precluded Hazardous Ice Formation

During the ST S-1 14 (2005) tanking test, the External Tank Gaseous Hydrogen Vent Arm Umbilical Quick Disconnect formed ice and produced liquid nitrogen/air. Testing of gaseous hydrogen vent arm umbilical disconnect equipment at Kennedy Space Center. The phenomenon was repeated during subsequent testing and launch. For the accomplished with two changes to the shroud was maintained above freezing shuttle, ice presented a debris hazard umbilical purge shroud. First, the space with no ice formation and that no nitrogen to the Orbiter Thermal Protection System agency improved the shroud purge gas penetrated into the shroud purge cavity. and was unacceptable at this umbilical flow to obtain the desired purge cavity NASA used the modified design on location. The production of uncontrolled gas concentrations. Second, technicians STS-121 (2006) and all subsequent flights. liquid nitrogen/air presented a hazard to wrapped multiple layers of aerogel Aerogel insulation is a viable alternative to the shuttle, launch pad, and ground blanket material directly onto the quick the current technology for quick support equipment. disconnect metal surfaces within the disconnect shrouds purged with helium or purged shroud cavity. NASA incorporated a fix into the existing nitrogen to preclude the formation of ice design to preclude ice formation and NASA tested the design modifications at and liquid nitrogen/air. In most cases, the uncontrolled production of liquid the Kennedy Space Center Cryo Test Lab. aerogel insulation eliminates the need for nitrogen/air. The resolution was Tests showed that the outer surface of the active purge systems.

Engineering Innovations 197 Clean Air Act effective January 2003. environment. Wind tunnel tests the tank. This required extensive NASA was granted exemptions demonstrated Thermal Protection engineering. NASA created enhanced permitting the use of HCFC-141b in System closeout capability to structural dynamics math models to foams for specific shuttle applications. withstand maximum aerodynamic better define the characteristics of These exemptions applied until the end loads without generating debris. this area of the tank and performed of the program. numerous wind tunnels tests. The ET protuberance air load ramps were manually sprayed wedge-shaped The ET fuel tank Main Propulsion Post-Columbia Accident layers of insulating foam insulation System pressurization lines and cable Advances in Thermal Protection along the pressurization lines and trays were attached along the length cable tray on the side of the tank. They of the tank at multiple locations by Following the loss of Space Shuttle were designed as a safety precaution metal support brackets. These were Columbia in 2003, NASA undertook to protect the tank’s cable trays and protected from forming ice and frost the redesign of some tank components pressurization lines from airflow that during tanking operations by foam to reduce the risk of ice and foam could potentially cause instability in protuberances called ice frost ramps. debris coming off the tank. These these attached components. Foam loss The feedline bracket configuration hardware changes drove the need to from the ramps during ascent, however, had the potential for foam and ice improve the application of Thermal drove NASA to remove them from debris loss. Redesign changes were Protection System foam that served as an integral part of the components’ function. The major hardware addressed included the ET/Orbiter attach bipod External Components Redesign closeout, protuberance air load ramps, ice frost ramps, and the liquid hydrogen tank-to-intertank flange area. Original Configuration Orbiter Belly The ET bipod attached the Orbiter to the tank. The redesign removed the Bipod foam ramps that had covered the bipod Attach Fitting attach fittings, and which had been designed to prevent the formation of ice when the ET was filled with cold liquid hydrogen and liquid oxygen on the launch pad. This left the majority of each fitting exposed. NASA Foam Ramp installed heaters as part of the bipod (removed) configuration to prevent ice formation on the exposed fittings. Redesigned Attach NASA developed a multistep process Fitting Foam Thermal to improve the manual bipod Thermal Protection System Protection System spray technique. Closeout Validation of this process was Redesigned Configuration accomplished on a combination of high-fidelity mock-ups and a full-scale After the Columbia accident, NASA implemented a number of improvements to External ET test article in a production Tank components and related Thermal Protection System elements. One such measure was the redesign of the Orbiter/External Tank attach bipod fitting mechanism, which included a meticulous reworking of the attach fitting Thermal Protection System configuration.

198 Engineering Innovations In all post-Columbia Thermal Protection System enhancement efforts, NASA modified process controls to ensure that defects were more tightly kept within the design envelope. The space agency simplified application techniques and spelled out instructions in more detail, and technicians had the opportunity to practice their application skills on high-fidelity component models. MSFC and Lockheed Martin also developed an electronic database to store information for each spray. New application certification requirements were added. Improvements included the forward bellows heater, the liquid In what used to be a one-person operation, a team of technicians at NASA’s Michoud Assembly Facility prepares to hand-spray BX-250 foam on the bipod attach fittings. The videographer (standing) records oxygen feedlines, and titanium the process for later review and verification. A quality control specialist (left) witnesses the operation, brackets. Improved imagery analysis while two spray technicians make preparations. and probabilistic risk assessments also allowed NASA to better track incorporated into the 17 ice frost The NASA/Lockheed Martin team and predict foam loss. Thermal ramps on the liquid hydrogen tank to also developed an enhanced three-part protection debris could never be reduce foam loss. BX-265 manual procedure to improve the Thermal completely eliminated, but NASA spray foam replaced foam in the ramps’ Protection System closeout process on had addressed a complex and closeout areas to reduce debonding the liquid hydrogen tank-to-intertank unprecedented set of problems with and cracking. flange area. determination and innovation.

With Ramps Without Ramps

Protuberance Air Protuberance Air Load Ramps Load Ramps Have Been Removed

Liquid Liquid Hydrogen Hydrogen Tank Ice Tank-to-Intertank Frost Ramps Flange Foam Thermal Protection System Closeout

NASA decided to delete the tank’s protuberance air load ramps and implement design changes to the 17 ice frost ramps on the liquid hydrogen tank. Both these measures required adjustments in the components’ Thermal Protection System configuration and application processes. Materials and techniques were also altered to improve the Thermal Protection System closeout of the flange joining the liquid hydrogen tank with the intertank.

Engineering Innovations 199