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Figure 7. Interlocking approaches of continuous filament substrates: (a) tape wrapped, shingle; (b) filament wound, helix; and (c) multidimensional.

reduced pressures (6 kPa (50 torr)) with the flow rates primarily determined by the substrate surface area. This technique produces a crust on the outer surfaces of the substrate, thus requiring machining and multiple infiltration cycles.

In the thermal gradient technique (fig. 9), the part to be infiltrated is supported by a mandrel that is inductively heated. Therefore, the hottest portion of the substrate is the inside surface, which is in direct contact with the mandrel. The outer surface of the low-density substrate is exposed to a cooler environment and results in a temperature gradient through the substrate thickness. Surface crusting is eliminated because the deposition rate is greater on the heated fibers near the mandrel, whereas the cooler outer fibers receive little or no deposit. Under proper infiltration conditions, the carbon is first deposited on the inside surface and, in a continuous process, progresses radially through the substrate as the densified substrate itself becomes inductively heated. Infiltration is normally accomplished at atmospheric pressure with a mandrel heated to «1100°C (ref. 14).

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Figure 8. Isothermal chemical vapor deposition to infiltrate fibrous carbon substrate.

Carbonized Organic Composites

Carbonized organic composites have fabrication procedures that are similar to those of conventional fiber-reinforced, resin-laminating techniques. The starting material is usually a prepregged# fabric or yam. These precursor materials are staged nominally at «100°C to achieve the desired degree of tack and flow of the resin. A laminate is then constructed and cured under pressure to compact the stack-impregnated fabric. Curing temperatures normally range from 125°C to 175°C with curing pressures on the order of 2.76 MPa (400 psi). The reinforced resin laminate is then postcured at 200°C to 275°C. As pyrolysis is initiated, shrinking occurs as the organic phase decomposes. Simultaneously, the release of vapors from pyrolysis expands the composite material. A slow release of these volatile by-products is required to minimize structural damage to the char. Finally, as higher temperatures are reached, thermal expansion of the carbon char itself occurs after pyrolysis is complete. After the initial carbonization, the material is then subjected to a series of reimpregnation and carbonization cycles until the desired density or the maximum density is achieved. The reimpregnation process is usually conducted under vacuum and pressure to aid in maximizing the pore filling.

* A fabric impregnated with a matrix material in a tacky state.

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Figure 9. Thermal gradient chemical vapor deposition.

If graphitization is desired, the high-temperature heat treatment may be used after each carbonization step or at the end of the reimpregnation and recarbonization cycles.

To summarize, a typical manufacturing cycle of a 2-D CC part is shown in figure 10. First, a woven graphite fabric that is preimpregnated with phenolic resin is laid-up as a phenolic-graphite laminate in a mold and is autoclave-cured. Once cured, the part is pyrolyzed to form a carbon matrix surrounding the graphite fibers. The part is then densified by multiple furfural alcohol reimpregnations and pyrolyzations. The resulting CC part then is ready for use in inert or vacuum environments. This process is very time-consuming. A single pyrolysis may take >70 hr in a low-temperature, inert-atmosphere furnace.

Although CC materials can withstand temperatures >3000°C in a vacuum or in an inert atmosphere, they oxidize and sublime when in an oxygen atmosphere at 600°C. To allow for use of CC parts in an oxidizing atmosphere, they must be compounded with materials that produce oxidation-protective coatings through thermochemical reaction with oxygen at >2000°C (ref. 15) or they must be coated and sealed to protect them (ref. 16). For applications such as the Space Shuttle

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Figure 10. Fabrication steps involved in manufacture of2-D carbon-carbon part tetraethylorthosilicate (TEOS).

CC leading edges and nose caps, surfaces are converted to silicon carbide in a high-temperature diffusion-coating process (fig. 10). Because of differences in thermal expansion between the silicon carbide and the CC part, the coating develops microcracks when the part is cooled from the coating temperature. To maintain oxidation protection on space vehicles such as the Space Shuttle, cracks are impregnated with tetraethylorthosilicate (TEOS). The TEOS process leaves silica in all of the microcracks, greatly enhancing the oxidation protection of the CC substrate. Current improvements being developed for oxidation protection of the CC Space Shuttle components are additions of low-temperature glass formers that enhance the sealing capability of the existing coating-TEOS system.

Mechanical Properties

The extreme thermomechanical requirements of the Space Shuttle have been the impetus for evaluating properties of low-density CC. The use of CC on the nose cap and leading edges of the Space Shuttle makes it imperative to know as much as possible about all the characteristics of this material. The effect of temperature on the ratio of tensile strength to density for several classes of high-temperature materials is shown in figure 11. The major advantage of CC materials for hightemperature applications is that they do not lose strength as the use temperature is increased. This property is in contrast to other materials such as superalloys and ceramics. Figure 11 shows three levels of CC strength efficiency. The first, labeled Space Shuttle material, is the strength level of the reusable carbon-carbon (RCC) material used in the Space Shuttle thermal protection system. Even though this material is made with low-strength carbon fibers, its strength efficiency is superior to both superalloys and ceramics at >1000°C. Development of advanced carbon-carbon (ACC) composites has produced a material that is twice as strong as the CC composite first put on the Space Shuttle. The ACC material is made using woven carbon cloth. When unidirectional carbon fiber tapes are interplied with woven cloth to create a hybrid ACC, its strength in at least one direction can be increased to >345 MPa (>50 000 psi). Current data on thermomechanical and thermochemical properties of some of the advanced CC systems show that material composition, oxidation resistance, processing, joining, and fiber architecture are producing noticeable improvements in CC materials and structures (refs. 17 to 26).

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Figure II. Strength-to-density ratio for several classes of high-temperature materials.


An example of the stale of the art in CC composite applications is shown in figure 12. This figure shows a prototype CC component (ref. 27). The component is a one-piece, bladed turbine rotor that, in service, would be coated to prevent oxidation. At the lower left in figure 12 is a representative coated CC microstructure showing the reinforcing filaments aligned in several directions. At the top of this inset in figure 12. the silicon carbide oxidation coating can be seen. Other gas

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