REINFORCED CARBON-CARBON (RCC)

¶The basic RCC composite is a laminate of graphite-impregnated rayon fabric, further impregnated with phenolic resin and layered, one ply at a time, in a unique mold for each part, then cured, rough-trimmed, drilled, and inspected. The part is then packed in calcined coke and fired in a furnace to convert it to carbon and is made more dense by three cycles of furfuryl alcohol vacuum impregnation and firing.

¶To prevent oxidation, the outer layers of the carbon substrate are converted into a 0.02-to-0.04-inch-thick layer of silicon carbide in a chamber filled with argon at temperatures up to 3,000 degrees Fahrenheit. As the silicon carbide cools,

¶"craze cracks" form because the thermal expansion rates of the silicon carbide and the carbon substrate differ. The part is then repeatedly vacuum-impregnated with tetraethyl orthosilicate to fill the pores in the substrate, and the craze cracks are filled with a sealant.

¶wing leading edge would not need to withstand impact from debris or ice, since these objects would not pose a threat during the launch phase.6

¶Reinforced Carbon-Carbon

¶The development of Reinforced Carbon-Carbon (RCC) as part of the Thermal Protection System was key to meeting the wing leading edge design requirements. Developed by Ling-Temco-Vought (now Lockheed Martin Missiles and Fire Control), RCC is used for the Orbiter nose cap, chin panel, forward External Tank attachment point, and wing leading edge panels and T-seals. RCC is a hard structural material, with reasonable strength across its operational temperature range (minus 250 degrees Fahrenheit to 3,000 degrees). Its low thermal expansion coefficient minimizes thermal shock and thermoelastic stress.

¶Each wing leading edge consists of 22 RCC panels (see Figure 3.3-1), numbered from 1 to 22 moving outward on each wing (the nomenclature is "5-left" or "5-right" to differentiate, for example, the two number 5 panels). Because the shape of the wing changes from inboard to outboard, each panel is unique.¹¹ program. NASA actually spent $9.9 billion in real year dollars to George Mueller, Associate Administrator for Manned Space Flight, take the Shuttle through design, development and initial testing. This NASA, "Honorary Fellowship Acceptance," address delivered to the sum, when converted to fixed year 1971 dollars using the aerospace British Interplanetary Society, University College, London, England, price deflator, equals $5.9 billion, or a 15 percent cost overrun on August 10, 1968, contained in John M. Logsdon, Ray A. Williamson, the original estimate for phase one. Compared to other complex Roger D. Launius, Russell J. Acker, Stephen J. Garber, and Jonathan L. development programs, this was not a large cost overrun." See Howard Friedman, editors, Exploring the Unknown: Selected Documents in the McCurdy, "The Cost of Space Flight," Space Policy 10 (4) p. 280. For History of the U.S. Civil Space Program Volume IV: Accessing Space, a program budget summary, see Jenkins, Space Shuttle, p. 256. NASA SP-4407 (Washington: Government Printing Office, 1999), pp. 11 202-205. STS stands for Space Transportation System. Although in the years just Report on the Commission on the Future of the United States Aerospace Industry, November 2002, p. 3-3. NASA Agency Contingency Action Plan for Space Flight Operations, January 2003, p. D-2.Show full note 2 3

¶4 5

¶21 22

¶Figure 3.3-1. There are 22 panels of Reinforced Carbon-Carbon on each wing, numbered as shown above.

¶Wing Leading Edge Damage Leading Edge Maintenance

¶The risk of micrometeoroid or debris damage to the RCC panels has been evaluated several times. Hypervelocity impact testing, using nylon, glass, and aluminum projectiles, as well as low-velocity impact testing with ice, aluminum, steel, and lead projectiles, resulted in the addition of a 0.03- to 0.06-inch-thick layer of Nextel-440 fabric between the Inconel foil and Cerachrome insulation. Analysis of the design change predicts that the Orbiter could survive re-entry with a quarter-inch diameter hole in the lower surfaces of RCC panels 8 through 10 or with a one-inch hole in the rest of the RCC panels.

¶RCC components have been struck by objects throughout their operational life, but none of these components has been completely penetrated. A sampling of 21 post-flight reports noted 43 hypervelocity impacts, the largest being 0.2 inch. The most significant low-velocity impact was to Atlantisʼ panel 10-right during STS-45 in March and April 1992. The damaged area was 1.9 inches by 1.6 inches on the exterior surface and 0.5 inches by 0.1 inches in the interior surface. The substrate was exposed and oxidized, and the panel was scrapped. Analysis concluded that the damage was caused by a strike by a man-made object, possibly during ascent. Figures 3.3-2 and 3.3-3 show the damage to the outer and inner surfaces, respectively.

¶Figure 3.3-2. Damage on the outer surface of RCC panel 10-right from Atlantis after STS-45.

¶Figure 3.3-3. Damage on the inner surface of RCC panel 10-right

¶Post-flight RCC component inspections for cracks, chips, scratches, pinholes, and abnormal discoloration are primarily visual, with tactile evaluations (pushing with a finger) of some regions. Boeing personnel at the Kennedy Space Center make minor repairs to the silicon carbide coating and surface defects.

¶With the goal of a long service life, panels 6 through 17 are refurbished every 18 missions, and panels 18 and 19 every 36 missions. The remaining panels have no specific refurbishment requirement.

¶At the time of STS-107, most of the RCC panels on Columbiaʼs left wing were original equipment, but panel 10-left, T-seal 10-left, panel 11-left, and T-seal 11-left had been replaced (along with panel 12 on the right wing). Panel 10-left was tested to destruction after 19 flights. Minor surface repairs had been made to panels 5, 7, 10, 11, 12, 13, and 19 and T-seals 3, 11, 12, 13, 14, and 19. Panels and T-seals 6 through 9 and 11 through 17 of the left wing had been refurbished.

¶Reinforced Carbon-Carbon Mission Life

¶The rate of oxidation is the most important variable in determining the mission life of RCC components. Oxidation of the carbon substrate results when oxygen penetrates the microscopic pores or fissures of the silicon carbide protective coating. The subsequent loss of mass due to oxidation reduces the load the structure can carry and is the basis for establishing a mission life limit. The oxidation rate is a function of temperature, pressure, time, and the type of heating. Repeated exposure to the Orbiterʼs normal flight environment degrades the protective coating system and accelerates the loss of mass, which weakens components and reduces mission life capability.

¶Currently, mass loss of flown RCC components cannot be directly measured. Instead, mass loss and mission life reduction are predicted analytically using a methodology based on mass loss rates experimentally derived in simulated re-entry environments. This approach then uses derived re-entry temperature-time profiles of various portions of RCC components to estimate the actual re-entry mass loss.

¶For the first five missions of Columbia, the RCC components were not coated with Type A sealant, and had shorter mission service lives than the RCC components on the other Orbiters. (Columbiaʼs panel 9 has the shortest mission service life of 50 flights as shown in Figure 3.3-4.) The predicted life for panel/T-seals 7 through 16 range from 54 to 97 flights.7

¶Localized penetration of the protective coating on RCC components (pinholes) were first discovered on Columbia in 1992, after STS-50, Columbiaʼs 12th flight. Pinholes were later found in all Orbiters, and their quantity and size have increased as flights continue. Tests showed that pinholes were caused by zinc oxide contamination from a primer from Atlantis after STS-45. used on the launch pad.

¶Columbia Wing Leading Edge

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 — 18

¶Panel/T-Seal Assembly

¶Figure 3.3-4. The expected mission life for each of the wing leading edge RCC panels on Columbia. Note that panel 9 has the shortest life expectancy.

The Orbiter wing leading edge structural subsystem consists of the RCC panels, the upper and lower access panels (also called carrier panels), and the associated attachment hardware for each of these components.

On Columbia, two upper and lower A-286 stainless steel spar attachment fittings connected each RCC panel to the aluminum wing leading edge spar. On later Orbiters, each upper and lower spar attachment fitting is a one-piece assembly.

The space between each RCC panel is covered by a gap seal, also known as a T-seal. Each T-seal, also manufactured from RCC, is attached to its associated RCC panel by two Inconel 718 attachment clevises. The upper and lower carrier panels, which allow access behind each RCC panel, are attached to the spar attachment fittings after the RCC panels and T-seals are installed. The lower carrier panel prevents superheated air from entering

¶The Wing Leading Edge Structural System on Columbia.

¶In October 1993, panel 12-right was removed from Columbia after its 15th flight for destructive evaluation. Optical and scanning electron microscope examinations of 15 pinholes revealed that a majority occurred along craze cracks in the thick regions of the silicon carbide layer. Pinhole glass chemistry revealed the presence of zinc, silicon, oxygen, and aluminum. There is no zinc in the leading edge support system, but the launch pad corrosion protection system uses an inorganic zinc primer under a coat of paint, and this coat of paint is not always refurbished after a launch. Rain samples from the Rotating Support Structure at Launch

¶Complex 39-A in July 1994 confirmed that rain washed the unprotected primer off the service structure and deposited it on RCC panels while the Orbiter sat on the launch pad. At the request of the Columbia Accident Investigation Board, rain samples were again collected in May 2003. The zinc the RCC panel cavity. A small space between the upper carrier panel and the RCC panel allows air pressure to equalize behind the RCC panels during ascent and re-entry.

¶The mid-wing area on the left wing, behind where the breach occurred, is supported by a series of trusses, as shown in red in the figure below. The mid-wing area is bounded in the front and back by the Xo1040 and Xo1191 cross spars, respectively. The numerical designation of each spar comes from its location along the Orbiterʼs X-axis; for example, the Xo1040 spar is 1,040 inches from the zero point on the X-axis. The cross spars provide the wingʼs structural integrity. Three major cross spars behind the Xo1191 spar provide the primary structural strength for the aft portion of the wing. The inboard portion of the mid-wing is the outer wall of the left wheel-well, and the outboard portion of the mid-wing is the wing leading edge spar, where the RCC panels attach.

¶Xo1365

¶Xo1191

¶Xo1040

¶The major internal support structures in the mid-wing are constructed from aluminum alloy. Since aluminum melts at 1,200 were destroyed and wing structural integrity was lost.

¶degrees Fahrenheit, it is likely these truss tubes in the mid-wing fallout rate was generally less than previously recorded except for one location, which had the highest rate of zinc fallout of all the samples from both evaluations. Chemical analysis of the most recent rainwater samples determined the percentage of zinc to be consistently around nine percent, with that one exception.

¶Specimens with pinholes were fabricated from RCC panel 12-right and arc-jet-tested, but the arc-jet testing did not substantially change the pinhole dimensions or substrate oxidation. (Arc jet testing is done in a wind tunnel with an electrical arc that provides an airflow of up to 2,800 degrees Fahrenheit.) As a result of the pinhole investigation, the sealant refurbishment process was revised to include clean- ing the part in a vacuum at 2,000 degrees Fahrenheit to bake out contaminants like zinc oxide and salt, and forcing sealant into pinholes.

¶Post-flight analysis of RCC components confirms that sealant is ablated during each mission, which increases subsurface oxidation and reduces component strength and mission life. Based on the destructive evaluation of Columbiaʼs panel 12-right and various arc-jet tests, refurbishment intervals were established to achieve the desired service life.

¶In November 2001, white residue was discovered on about half the RCC panels on Columbia, Atlantis, and Endeavour. Investigations revealed that the deposits were sodium car- bonate that resulted from the exposure of sealant to rainwater, with three possible outcomes: (1) the deposits are washed off, which decreases sealant effectiveness; (2) the deposits remain on the partʼs surface, melt on re-entry, and combine with the glass, restoring the sealant composition; or (3) the deposits remain on the partʼs surface, melt on reentry, and flow onto metal parts.

¶The root cause of the white deposits on the surface of RCC parts was the breakdown of the sealant. This does not damage RCC material.

¶Non-Destructive Evaluations of Reinforced Carbon- F3.3-1 Carbon Components

¶Over the 20 years of Space Shuttle operations, RCC has performed extremely well in the harsh environment it is exposed to during a mission. Within the last several years, a few instances of damage to RCC material have resulted in a re-examination of the current visual inspection process. Concerns about potential oxidation between the silicon carbide layer and the substrate and within the substrate has resulted in further efforts to develop improved Non-Destructive Evaluation methods and a better understanding of subsurface oxidation.

¶Since 1997, inspections have revealed five instances of RCC silicon carbide layer loss with exposed substrate. In November 1997, Columbia returned from STS-87 with three damaged RCC parts with carbon substrate exposed. Panel 19-right had a 0.04 inch-diameter by 0.035 inch-deep circu- lar dimple, panel 17-right had a 0.1 inch-wide by 0.2 inch- long by 0.025-inch-deep dimple, and the Orbiter forward External Tank attachment point had a 0.2-inch by 0.15-inch

¶by 0.026-inch-deep dimple. In January 2000, after STS-103, Discoveryʼs panel 8-left was scrapped because of similar damage (see Figure 3.3-5).

¶In April 2001, after STS-102, Columbiaʼs panel 10-left had a 0.2-inch by 0.3-inch wide by 0.018-inch-deep dimple in the panel corner next to the T-seal. The dimple was repaired and the panel flew one more mission, then was scrapped because of damage found in the repair.

Panel 8L (Discovery)

¶igure 3.3-5. RCC panel 8-left from Discovery had to be scrapped fter STS-103 because of the damage shown here.

¶Findings:

¶F3.3-1 The original design specifications required the

RCC components to have essentially no impact resistance.

¶F3.3-2 Current inspection techniques are not adequate

to assess structural integrity of the RCC components.

¶F3.3-3 After manufacturerʼs acceptance non-destructive

evaluation, only periodic visual and touch tests are conducted.

¶F3.3-4 RCC components are weakened by mass loss

caused by oxidation within the substrate, which accumulates with age. The extent of oxidation is not directly measurable, and the resulting mission life reduction is developed analytically.

¶F3.3-5 To date, only two flown RCC panels, having

achieved 15 and 19 missions, have been destruc- tively tested to determine actual loss of strength due to oxidation.

¶F3.3-6 Contamination from zinc leaching from a primer

under the paint topcoat on the launch pad structure increases the opportunities for localized oxidation.

¶Recommendations:

¶R3.3-1 Develop and implement a comprehensive in-

spection plan to determine the structural integrity of all Reinforced Carbon-Carbon system components. This inspection plan should take advantage of advanced non-destructive inspection technology.

¶R3.3-2 Initiate a program designed to increase the

Orbiterʼs ability to sustain minor debris damage by measures such as improved impact-resistant Reinforced Carbon-Carbon and acreage tiles. This program should determine the actual impact resistance of current materials and the effect of likely debris strikes.

¶R3.3-3 To the extent possible, increase the Orbiterʼs abil-

ity to successfully re-enter the Earthʼs atmosphere with minor leading edge structural sub-system damage.

¶R3.3-4 In order to understand the true material character-

istics of Reinforced Carbon-Carbon components, develop a comprehensive database of flown Reinforced Carbon-Carbon material characteristics by destructive testing and evaluation.

¶R3.3-5 Improve the maintenance of launch pad struc-

tures to minimize the leaching of zinc primer onto Reinforced Carbon-Carbon components.