AAWSAP DIRD, Materials for Advanced Aerospace Platforms, January 2010

UNCLASSIFIED/JFOR OFFICIAL USE 014[ I 13-Ti alloys would be an attractive option for a launch vehicle with a very heavy payload. These alloys can be processed and heat-treated to ultimate tensile strengths greater than 1,300 mega-Pascals (MPa), making them very efficient structural alloys. However, at these strength levels, the fracture toughness is decreased to about 40 MPa-m0 •5 , making damage tolerance marginal. For example, using a design stress that is twothirds that of ultimate tensile strength, the critical crack size for an alloy with these properties is about 4 millimeters. Such a small critical crack size poses a challenge to any required field inspections associated with reusability requirements. p-Ti alloys' stiffness also is as much as 10 percent lower than that of a+J3 alloys. Ten percent lower stiffness generally is not an issue in a tension-loaded structure, but it can be an issue for a compression-loaded structure because of the potential for buckling. Designs with a higher section modulus can eliminate this concern, but the additional shape complexity will almost certainly add cost. In extreme circumstances, Ti alloys can be reinforced with ceramic fibers (typically silicon carbide) to increase their intrinsic stiffness, as will be discussed later. Ti alloys are reactive when exposed to air at temperatures of 550 °Celsius or above. Consequently, any manufacturing operations that exceed this temperature limit must be performed in a protective atmosphere of argon (Ar) or helium (He) gas. An exception occurs during forging if enough excess material is left on the raw forging to contain the oxygen-contaminated layer so it can be machined away during the final machining of the finished component. How to deal with this reactivity issue is well understood, and it poses no concern other than the additional costs associated with the excess material and additional machining. In other operations, such as welding, use of specially designed fixtures incorporating inert shielding gas also effectively eliminates concerns about oxygen contamination. In hot-forming applications, the as-formed part is typically chemically milled to remove the oxygen-rich surface layer because this layer typically has lower ductility and can cause fatigue cracking in service. The practice of eliminating all oxygen-contaminated material has served the aerospace industry well over the years but is quite conservative and restrictive. This matter is discussed in greater detail in the Reusable Reentry Vehicle section of this document. In summary, although the choices of materials for launch vehicles are in principle numerous, in practice these choices are reduced by a variety of application-specific considerations that include manufacturing capability for large components and manufacturing cost. The foregoing discussion has attempted to examine the prospects for advanced Al alloys, PMCs, and Ti alloys in light of these perceived practical constraints. On a case-by-case basis, a variety of requirements are imposed by design constraints, which some materials meet more readily than others. In all cases, discussion of available material options was constrained by the assumption that the maximum service temperatures would be relatively low. Consequently, the material classes discussed here are all intended for relatively low-temperature use. The separate case of a reusable single-stage-to-orbit vehicle, where operating temperature requirements are considerably higher, is discussed in a later section of this document. Ultimately, materials are selected to optimize structural performance, and a coordinated approach of materials selection and geometric design is essential to this. Going forward, a design using a synthesis process that treats form, fit, function, and materials capability as equal constraints is needed to achieve true optimum structural efficiency. 8 UNCLASSIFIED/ /570A OFFl&IAk W&E 8HL\«

UNCLASSIFIED//FOR OFFICIAL USE ONLY REUSABLE CREW MODULES The concept of manned orbital crew modules has evolved from the Mercury capsules to the Gemini and Apollo programs to the space shuttle, the first fully reusable crew module. The shuttle also has a combined payload bay used for transporting satellites into orbit and hardware for developing the International Space Station and for repairing and refurbishing the Hubble telescope, among other uses. The reusable nature of the shuttle crew module introduced a number of design and materials selection challenges. Perhaps foremost among these is the requirement for a thermal protection system (TPS) that would protect the crew during reentry and also minimize the intermission refurbishment requirements of the spacecraft itself. During the early days of the shuttle development program, there was much initial interest in a metallic TPS because it appeared to better meet the program's needs. Ultimately, however, ceramic tiles were used on the underbody and carbon-carbon composites (C-CCs) were used on the leading edges of the wings. The shuttle design itself can therefore be characterized as a "cold structure" with an insulating TPS. For example, much of the shuttle load-bearing structure is made of the Al alloy 2219, in part because it is fusion weldable and in part because it retains its strength at moderately elevated temperatures better than other high-strength Al alloys can. The refurbishment needs of the ceramic shuttle tiles after each flight reputedly are considerable and increase with vehicle age. The C-CC wing leading edges are basically not repairable but require scrutiny. Hindsight shows that C-CCs "age" and lose much of their fracture toughness during repeated thermal exposure. (This loss of toughness was a prime factor in the Columbia disaster. Had a metallic heat shield that included the wing leading edges been used, this disaster arguably could have been avoided.) If the discussion of a metallic TPS were held today, the outcome might not be much different. Certainly any serious consideration of a reusable single-stage-to-orbit vehicle today would need to reopen the discussion of a metallic TPS. The challenges and opportunities associated with a metallic TPS are discussed later. The design efficiency of an integrated TPS and load-bearing structure is extremely attractive. Such a design requires availability of high-temperature alloys that also have good strength at the moderate temperatures to withstand the aerodynamic and vibrational loads encountered during launch and orbital insertion. Any attractive alloy also must have reasonable intrinsic resistance to oxidation at the reentry temperatures and should be capable of being fabricated into sheet gauges at reasonable cost. Meeting these various requirements in combination becomes quite daunting. Earlier programs such as the DynaSoar reusable reentry glider devoted considerable time and resources to examining the use of refractory alloys such as Mo-0.5%Ti (Moly half Ti) and several Columbia-based alloys for the TPS and some hot structure. All the refractory metal alloys are solid solution strengthened and, consequently, have relatively low ambient temperature strengths. They also are quite dense, making the density-corrected strength even less attractive. Refractory metals and their alloys react extensively when exposed to air at elevated temperatures. Therefore, even if the mechanical property limitations could be overcome, any hot structure would require an oxidation-resistant protective coating. In the case of the DynaSoar program, scientists extensively investigated a surface conversion coating of MoSi2 formed by reacting the Moly half-Ti alloy with Si powder in a high-temperature fluidized bed. With the benefit of hindsight, it is now unclear how a large structure could have been successfully coated in this manner. This is particularly true when the brittle nature of MoSi2 is 9 UNCLASSIFIED//FOR OFFICIAL USE ONLY

UNCLASSIFIED//FOR OFFICIAL USE ONLY considered in light of the thermal stresses that were certain to develop in a large structure placed in the fluidized bed. Reusable Single-Stage-to-Orbit Vehicles A reusable single-stage-to-orbit (SSO) vehicle will require metallic materials for the TPS and for much of the other hot structure. This will be challenging from the standpoint of an empty vehicle weight. With the exception of military applications, which are outside the scope of this document, empty vehicle weight is a critical metric because every additional pound of empty weight reduces the payload by the same amount (assuming a fixed propulsion capability). As mentioned earlier, the key to a lightweight vehicle is the use of design methods that integrate the TPS and the load-bearing structure to minimize structural redundancy and single-function structure (for example, a TPS that is not load bearing). Achieving this goal will require new design paradigms that incorporate true synthesis of new structural concepts. In reality, such designs can be completed only if they are based on a detailed set of mission requirements, including the number of missions and expectations for turnaround time between missions. Furthermore, operational parameters such as the value of inserting a pound of payload into orbit are needed to bound the cost of the initial vehicle and the maintenance cost per mission (translated into cost per pound of payload). Absent such specific data, the following discusses possibilities for materials systems that can enable a reusable SSO vehicle. It is perhaps more efficient to discuss these materials according to their principal capability and the anticipated temperature regime in which they can be used most productively. This categorization method is illustrated in Table 1. Table 1. Potential Materials by Use Temperature Regime and Property Temperature Regime | Specific Strength | Specific Stiffness | Fatigue Resistance | Fracture Toughness | Creep Resistance Ambient up to 250 °C | PMCs; Advanced Al alloys; AMCs*; Ti alloys; TMCs* | PMCs; AMCs; TMCs | PMCs; AMCs; Ti Alloys; TMCs | PMCs; Ti alloys | Ti alloys; TMCs 250 °C up to 550 °C | Ti alloys; TMCs | TMCs | Ti alloys; TMCs | Ti alloys | Ti alloys; TMCs Above 550 °C | Ni-base alloys; Ti aluminides; Refractory metal alloys; CMCs*; C-C Cs* | Ti aluminides; CMCs; C-CCs | Ni-base alloys; Refractory metal alloys | Ni-base alloys; CMCs; Refractory metal alloys | Ni-base alloys; Ti aluminides; CMCs; Refractory metal alloys; C-C Cs AMCs = Al matrix composites; TMCs = titanium matrix composites; CMCs = ceramic matrix composites; C-CCs = carbon-carbon composites 10 UNCLASSIFIED//FOR OFFICIAL USE ONLY

UNCLASSIFIED/ /FOR OFFICIAL USE ONLY Table 1 shows there are essentially 10 distinct classes of candidate materials for use in a reusable SSO vehicle. Their usefulness for specific applications and components depends on the operating temperature regime and the design-limiting material property. Some background and characteristics of each of these materials are described in this section. Applications of some of these materials have already been addressed, so the discussion here is limited to applications for a reusable SSO vehicle. Advanced Al Alloys This class of materials is limited by its temperature capability because, even if used in an embedded structure, the thermal soakback from the hot external structure can lead to softening in real time or overaging during extended exposure (that is, after several missions). Scientists have worked to develop high-temperature powder-metallurgy Al alloys, mainly alloys containing Fe, cobalt (Co) and cesium (Ce) or Mn and Si. These alloys are not routinely produced today owing to a lack of demand stemming in part from their cost and their relative lack of maturity as judged by the high variability in mechanical properties between lots of material. Consequently, the main application for Al alloys is in the crew compartment itself, where temperatures must be maintained at levels that are tolerable for the human occupants. Al alloys have marginal stiffness, even on a density-corrected basis. However, using fabricated panels with Al face sheets and a lightweight core can minimize this limitation. Such panels have a high section modulus, which increases the structural stiffness without adding much weight. Earlier uses of phenolic honeycomb cores experienced only limited success because of the tendency of the core to absorb water from the environment if a breach of the face sheet or the face sheet core bond occurred. Today, Al phenolic honeycomb sandwich construction is unpopular among aircraft and spacecraft designers because of this earlier, unfavorable experience. However, a new possibility for lightweight stiff structures Al face sheets with a foamed Al alloy core is worthy of consideration. Considerable progress in making uniform-density Al and other metallic foams has been realized in the past 10 years. This new class of sandwich materials is sufficiently different from the earlier versions to merit a careful evaluation. Polymer Matrix Composites As described in the Launch Vehicle section, polymer matrix composites have excellent strength, stiffness, fatigue resistance, and fracture toughness. All polymeric materials are to some degree susceptible to degradation when exposed to ultraviolet (UV) radiation. For an orbiting vehicle, a major limitation is the effects of UV radiation on the polymer matrix. Given their limited time at altitude, this vulnerability is not an issue for launch vehicles; however, it is a concern for structures that remain in orbit for an extended time at altitudes where the UV intensity is much greater. Although there are coatings that protect substrates (PMCs in this case) from UV radiation, the risk of these coatings being breached and the uncertainty about their effectiveness in orbit make UV exposure an ongoing concern. Any consideration of using PMCs for exterior applications would need to include an evaluation of their effectiveness. As discussed for Al alloys, PMCs also could be considered for interior applications where temperatures are within the material capabilities. As discussed earlier, some polyimide resins are usable up to about 300 °Celsius; therefore, the low density and excellent specific stiffness of PMCs make them viable candidates for some components. As also discussed, some polyimides, such as PMR-15, contain methylenedianiline, and care must be taken to 11 UNCLASSIFIED/ /FOR OFFICIAL USE ONLY

ensure that any outgassing that occurs in space would not liberate any of this toxic material—at least not where it could be ingested by the crew. Al Matrix Composites Al matrix composites (AMCs) are typically made by mixing short fibers or even particulate silicon carbide (SiC) with Al alloy powder and hot-pressing or mechanically consolidating the mixture by extrusion or forging. AMCs have higher modulus and strength but suffer ductility losses owing to the hard, nondeforming SiC second phase. They also are much better in fatigue because the SiC particles or short fibers mitigate the effects of planar slip in precipitation-hardening alloy matrices. Machining, fusion-welding, or cold-forming AMCs is difficult. Mechanical fastening is possible, but drilling fastener holes is difficult and expensive because the wear caused by the hard, reinforcing phases shortens tool life. Most of the AMCs produced to date have used relatively simple Al matrix alloys such as 6061 and 5083. This presumably is because of the conventional wisdom that AMCs are not heat treatable and the perception that no benefit is gained by using more complex alloy matrices. This is probably the case for a conventional solution treat-and-age approach. Consequently, there has been little effort to optimize AMC systems. For the right application where AMCs could provide a significant benefit, this could present an opportunity. For example, using the high-temperature Al-Fe-Co-Ce alloy powder as the matrix could prove interesting and might permit use of AMCs at up to 200 °Celsius—a temperature at which Ti alloys do not provide any significant advantage, but one that is too high for conventional Al alloys to be suitable. In sum, while AMCs are not really a commercial materials system today, sufficient research has been performed to establish proof of concept. If an adequate market for AMCs were to emerge, the time and cost to make them commercially available could prove acceptable. Ti Alloys Ti alloys also have been discussed earlier, but mainly in the context of heavy-section, large-load-bearing applications for launch vehicles. Here, the potential of Ti alloys for lighter gauge applications in the warm structure and the TPS is considered. Table 1 shows that at intermediate temperatures, Ti alloys are attractive in all aspects except for specific stiffness. What this table does not capture is Ti alloys' propensity to react with oxygen in the air to form an oxygen-stabilized α phase layer on the surface known as α case. An example of α case is shown in Figure 3. 12 UNCLASSIFIED//FOR OFFICIAL USE ONLY

UNCLASSIFIED/;'FOR OFFl&IAl W&IE 9fslllf Figure 3. Micrograph Showing a Phase Formation at the Surface of a Ti Alloy That has Been Exposed to Air at Elevated Temperature The a case is harder than the matrix because oxygen is a potent a-phase solid-solution strengthener. As with almost all other strengthening reactions, the increased strength is accompanied by reduced ductility. Again, drawing on conventional wisdom, the presence of a case in sheet structures has been forbidden by specification, design practice, or whatever means a company uses to manage its hardware. Less clear is how truly detrimental a case is to properties. Essentially no effort has been made to determine whether it can be tolerated if the affected hardware is allowed to operate at a modestly reduced stress. The industry standard for a case has essentially been one of zero tolerance. Given the significant potential weight advantage associated with use of Ti alloys in portions of the TPS and warm structure, this conservative approach needs to be revisited. Several key questions related to this are: • Is a case truly detrimental to the load-bearing capability of Ti alloy sheet structures? • If so, is there a limiting amount that can be tolerated without significantly degrading the structural capability? • Which properties are the most severely degraded? • Are a case formation and property degradation alloy dependent? • If so, which alloys are the most tolerant of a case formation? 13 UNCLASSIFIED/ /FOR OFFI&IAl Wliliii QfslL¥

UNCLASSIFIED/ /FOR OFFICIAL USE ONLY • Once α case is present, can the structure be repaired (for example, by fusion or friction stir welding)? The Air Force Materials and Manufacturing Directorate is starting a new project intended to address these questions. The motivation for this program is hypersonic flight vehicles. The results from this U.S. Air Force program should prove highly useful to the design of future reusable SSO vehicles. Earlier uses of Ti alloys at high temperatures included the skin and much of the load-bearing structure of the SR-71 Blackbird. This airplane flew successfully at peak speeds in excess of mach 3.2 for 34 years (1964-1998). While the maximum skin temperatures are not readily available, they were in excess of 300 °Celsius. There were no known issues involving α case during the SR-71's service. Notably, the primary alloy used for the SR-71 was one of the original β-Ti alloys, B-120 VCA, the composition of which is Ti-13V-11Cr-3Al. The primary reason for choosing this alloy was that it is much easier than any of the α+β Ti alloys are to roll into sheet gauges. Today there are newer β-Ti sheet alloys featuring a better balance of properties that could be used in the same way as B-120 VCA. The most common of these is Ti-15V-3Cr-3Sn-3Al. However, the successful use of B-120 VCA raises the question of whether β-Ti alloys are more resistant than α+β alloys such as Ti-6-4 are to α case formation. The attraction to using Ti alloys, in addition to their structural efficiency, is the extensive industrial base for making the material in a variety of product forms and the extensive knowledge base resulting from the many successful applications of Ti alloys in high-performance products. For example, the ability to superplastically form Ti alloys such as Ti-6-4 creates the opportunity for design of a structure that functions both as load bearing and as thermal protection. Ti Matrix Composites As Table 1 showed, Ti alloys are not especially attractive for their specific stiffness. One way to overcome this limitation is to reinforce a Ti alloy matrix with SiC fibers. In this case, the fibers are "long" fibers-they have sufficient length for the matrix to transfer the maximum possible fraction of the external load to the fiber. The fibers are essentially monofilaments and must be carefully placed so adjacent fibers do not touch one another. Areas of contact between fibers essentially are incipient cracks that degrade the mechanical strength. As Table 2 shows, TMCs have excellent properties. Table 2. Example of Properties of Ti Matrix Composites Property | Property Value (English / Metric Units) Ultimate tensile strength | 276 ksi / 1902 MPa Young's modulus | 32.8 msi / 226 GPa Strain to fracture | 0.95% Density | 0.16 lb/in3 / 4.43 g/cm3 Fiber Volume fraction | 0.39 14 UNCLASSIFIED/ /FOR OFFICIAL USE ONLY

UNCLASSIFIED/ FOR OFFICIAL USE ONLY Figure 4 shows an example of a TI matrix composite (TMC) cross section. [IMAGE: Cross-section micrograph of a Ti matrix composite showing a regular array of small dark fiber centers] Figure 4. Cross-Section Micrograph of a Ti Matrix Composite. Small, dark centers of fibers are C monofilament substrates for deposition of SiC (light micrograph). Material system: Matrix alloy - Ti-6242; Fiber - SiC about 5.6-mil-diameter SCS-6 Table 2 shows that the strength and stiffness properties of TMCs can exceed those of the Ti matrix (or other a.+β Ti alloys) by more than a factor of two with no increase in density. So why are TMCs not in widespread use? The foremost reason is cost. Another reason is the availability of SiC fiber for TMCs. • During the 1990s, when the U.S. government (mainly DoD) was interested in and provided development money for TMC R&D, the most attractive reinforcing fiber was SCS-6™, which was produced exclusively by Textron Specialty Materials (TSM) in Lowell, MA. • Unfortunately (at least in hindsight), TSM made a business decision not to sell fiber and to instead forward-integrate and sell TMC components or finished TMC mill products. This decision stemmed in part from TSM's negative business experience with development of B-based fiber (Boro-SiC™) for first-generation TMCs. • Prior to making this business decision, TSM did not develop sufficient TMC-manufacturing expertise to properly position itself as a producer of TMC products with consistent properties. 15 UNCLASSIFIED/ FOR OFFICIAL USE ONLY

• A couple of costly and highly visible TMC component test failures using TSM materials called into question the viability of TMCs. In truth, the real issue was one of quality and not the fundamental viability of the TMC material concept. • These failures led to extreme caution by the government, which promptly imposed stringent quality requirements on TSM. Consequently, TSM became a serious bottleneck for availability of TMCs with acceptable quality. As a result, numerous large, government-funded R&D programs fell behind schedule to the point that some were canceled and others were abandoned. • Other, less attractive SiC fibers were being developed during this period, including Nicalon in Japan and Sigma in the United Kingdom. Because SiC fiber and TMCs were deemed strategic materials and neither of these alternate fibers was produced in the United States, there was reluctance to experiment with them, and the U.S. government occasionally prohibited their use in federally funded R&D programs. • The issue of fiber availability became a major distraction from the real business at hand—that is, improving the consistency and reducing the cost of the TMC product. • In the end, TSM partially relented, but by then such limited interest in TMCs existed that the original opportunity was lost. Furthermore, there still was no concrete evidence that TMCs could be produced with sufficiently consistent properties that they could be considered an engineering material (at any cost). • In the mid-1990s, a cost study that assumed fiber availability showed that TMCs that meet specification properties could be manufactured for about $500/lb (in 1995 $) if the use volume was about 10,000 lb/year. Finding enough low-risk, high-value applications to consume this quantity of TMCs was not deemed possible. • Consequently, work on TMCs halted after an investment of about $500 million of U.S. government funds and a (presumably) comparable, but less well known private sector (mostly independent R&D $) investment. • One lesson from this is that a credible market and cost study should be undertaken before embarking on a major R&D program to develop revolutionary materials such as TMCs. The foregoing discussion exemplifies the challenges associated with developing and commercializing a revolutionary new materials system. New material concepts originating from nonproduction sources such as national laboratories or research universities should be approached with caution and never be put on the critical path of product design. Even if a new material is vetted through an established production source, the timing of full commercialization should be carefully examined. Ni-base Alloys Ni-base alloys have been considered real engineering materials for at least 40 years. Because of their relatively low rate of strength loss with increasing temperature, Ni-base alloys are also called "superalloys." They are commonly used at temperatures well in excess of half of their melting point (Tm), which is the accepted useful limit for most structural metallic materials. There is no other class of structural alloys for which this is true. Ni-base alloys have been used in turbine engines of all types almost since their 16

UNCLASSIFIED/ /FOR OFFICIAL USE ONLY inception. In fact, Ni-base alloys are a prime enabler of the modern gas turbine engine. As a result, this class of materials has a well-established industrial base for the production of literally all product forms. There is a second class of Ni-base alloys that has excellent resistance to attack in aggressive environments, such as those encountered in petroleum recovery, but these alloys are not of interest for high-temperature applications. To avoid any confusion, the term Ni-base alloys is used in this document to mean high-temperature Ni-base alloys. Ni-base alloys have a high elevated temperature strength because they typically contain Al and Ti, which enable them to be precipitation strengthened. The alloys with the highest temperature capability are strengthened by precipitates of Ni3Al (called γ'), which have a positive temperature coefficient of strength up to about 1,000 °Celsius. This unusual characteristic accounts for Ni-base alloys' elevated temperature capability. In combination with the ordinary temperature dependence of the matrix strength, the overall temperature dependence is quite low compared with that of any other class of structural alloys. Ni-base alloys also contain alloying additions of refractory elements such as tungsten (W), tantalum (Ta), and niobium (Nb). These additions also improve the elevated temperature strength because they diffuse slowly and strengthen the matrix. Because Ni-base alloys are used at elevated temperatures, they contain Cr additions in significant concentrations to improve their oxidation resistance. The tailoring of the properties of Ni-base alloys has been studied extensively and is quite well understood empirically, albeit less so in a fundamental sense. Current-generation Ni-base alloys contain as many as 10 alloying additions, and interelement interactions make gaining a fundamental understanding of alloying effects a major challenge. Production of Ni-base alloys has evolved over several decades. Both ingot metallurgy and powder metallurgy methods are used to produce a full range of mill products. New alloys have periodically been introduced in response to specialized needs. Sometimes these alloy development efforts have been so focused on the "specialized need" that retention of other generally accepted Ni-base alloy properties has been overlooked. An example of this is the development of early "low coefficient of thermal expansion" (α) alloys (the so-called low-α alloys). An example of an application for which low-α alloys are attractive is structures in which there are two or more concentric rings connected by radial struts. In the presence of a radial temperature gradient, these alloys tend to crack during thermal cycling situations owing to "thermal fatigue." In principle, the availability of an "Invar type" alloy with structural characteristics reduces the thermal strains and improves the thermal fatigue life. Invar is so named because it uses the magnetic characteristics of Ni to offset normal thermal expansion, resulting in much lower α values. It was recognized that this concept could also be adapted to Ni-base structural alloys. The first-generation low-α alloy, Inco 903, had about half the coefficient of thermal expansion compared with ordinary Ni-base alloys up to the Curie temperature. This remarkable alteration of a physical property was achieved by removing the Cr from the alloy to intensify the magnetic effects of magnetostriction. An unfortunate, unintended consequence of this composition alteration was that the oxidation resistance was seriously degraded. In particular, Inco 903 had a propensity for intergranular cracking in air under stress, rendering it of limited utility as a structural alloy. In fact, the cracking was so severe it was given a name: stress-assisted grain boundary oxidation (SAGBO). A second-generation low-α alloy, Inco 909, contained Si to help improve the oxidation resistance, but this alloy has not gained widespread acceptance, possibly because Inco 903 problems were so severe. Low-α 17 UNCLASSIFIED/ /FOR OFFICIAL USE ONLY

alloys' superior thermal fatigue resistance gives them great potential for use in TPS; thus, it would be appropriate to revisit this class of alloy after addressing its past problems. Many of the Ni-base alloys are fusion weldable, but the higher strength grades are prone to cracking without special precautions such as preheating the work piece. Most alloys capable of being processed into sheet also can be processed to create a fine-grained structure. The fine-grained material can be superplastically formed, which is potentially of very important benefit to making light-gauge, load-bearing structures and TPS components. Finally, although almost all Ni-base alloys contain Cr and Al, both of which improve oxidation resistance by forming a stable, protective scale, this scale begins to lose its effectiveness above about 1,000 °Celsius. A number of effective environmental coatings have been developed by the gas turbine industry for Ni-base alloys. The only concern or unknown for application in a reusable SSO vehicle is the durability of these coatings under very high mass flow conditions, such as during reentry. Refractory Metal Alloys Refractory metal alloys were discussed briefly earlier in connection with the DynaSoar project. Subsequent to this, there has been limited systematic interest in designing aerospace structures that incorporate refractory metal alloys. There has been interest in using these alloys for high-temperature gas-cooled nuclear reactors, where the operating environment is benign and well controlled, but this has limited relevance to the current discussion of space vehicles required to withstand reentry into the earth's atmosphere. The attraction of refractory metals as a class is the high melting temperature of Mo, Nb, and Ta. Although alloys of these metallic elements exhibit "normal" temperature dependence of mechanical strength, 0.5Tm of Nb or Mo is still a higher temperature than Ni-base alloys can withstand. Of the three metals, Nb is by far the most attractive because of its lower density, lower elastic modulus, and better (but not good) oxidation resistance. The clear barrier to use of Nb alloys is their reactivity in air at high temperatures. There are coatings for these alloys, but at 1,250 °Celsius, a breach in the coating will cause immediate, catastrophic failure. A reasonable design practice for critical structure, such as the TPS, is to disallow use of any coated material if the material fails catastrophically when the coating is breached. If this practice were the norm, then refractory metal alloys would not be usable in the TPS or other critical hot structure in a reusable SSO vehicle. Ceramic Matrix Composites Ceramic matrix composites (CMCs) are attractive because they are much tougher than, but retain the high-temperature capability of, monolithic ceramics. CMCs' toughness is derived from the fibers used to reinforce the ceramic matrix. When a CMC is loaded in tension and the stress in the ceramic matrix reaches a critical value, microcracks develop. This stress level is known as the matrix microcrack stress. In a monolithic ceramic such as silicon carbide (SiC) or silicon nitride (Si3N4), microcracks would propagate rapidly, causing immediate failure. For this reason, monolithic ceramics are not suitable for tension-loaded applications. This characteristic is mitigated in CMCs by reinforcing the ceramic matrix with high-strength ceramic fibers, typically Al2O3 or SiC, whose role is to bridge and arrest the cracks. Under continued loading, the fibers also 18

will eventually fracture at their weakest points, which, statistically, is often somewhere in the matrix rather that at the mouth of the crack . These fractured fibers then must be pulled out of the matrix during crack extension under continued loading, an action that dissipates additional energy. Therefore, the total energy expended during fracture of a CMC is much greater than it would be in the case of a monolithic specimen of the ceramic matrix. The crack bridging and pullout mechanisms of toughening, with the attendant increase in energy dissipation during fracture, are the conceptual foundation for CMCs. In SiC fiber-reinforced SiC matrix (called SiC-SiC) CMCs, the toughness can be as high as 10 times that of the unreinforced matrix. Clearly, the nature of the fibermatrix interface determines the resistance to fiber pullout after fracture and controls the toughness. In high-temperature CMCs, the fibers are coated to achieve an intermediate fiber-matrix interfacial bond strength that optimizes pullout toughening. If the interface is too weak, the fibers pull out too easily. If it is too strong, the fibers break without any pullout. Neither of these situations maximizes the toughness. Therefore, the challenge in creating a tough CMC is not only creating this interface but also finding a coating that will remain stable over time during exposure to elevated temperatures. As service temperature increases, this latter requirement becomes more challenging. For extended service, SiC-SiC CMCs currently are limited to about 1,400 °Celsius, but this is higher than the capability of metallic materials. The matrix microcrack stress is also significant because once the matrix develops microcracks, it allows the environment to gain access to the fiber-matrix interfaces. The role of environment can be to alter the nature of the interface and reduce the extent of pullout toughening. The obvious difficulty with this is that the properties are better initially than after a period of exposure in service. This leads to a nonconservative design that is dangerous. Absent an arbitrary knockdown, there currently exists no means of estimating the reduced toughness as a function of service life. Nevertheless, CMCs are the material class that holds the greatest promise of defeating the temperature limits of current metals. An important issue at present is the limited industrial base for producing CMCs and the even more restrictive range of suitable ceramic fibers available for use as the reinforcement in CMCs. The cost of CMCs is currently very high but should come down with increased demand, as happened with TMCs. Even so, CMCs will always be expensive, making development of efficient designs that make optimal use of this class of material important. Carbon-Carbon Composites Carbon-carbon composites (C-CCs) consist of carbon fibers in a typically amorphous carbon matrix. In principle, this class of materials is very attractive because it has very low density and high strength at elevated temperatures. C-CCs also have good oxidation resistance at very high temperatures (1,200 °Celsius) because they use an Si-rich coating that forms a stable, protective Si02 film on the surface. This coating system is "self healing" if breached because, among other things, it is glassy (viscous) and reforms over cracks. The principal vulnerability occurs at intermediate temperatures where the viscosity of the glass is high enough that it may not flow rapidly enough to heal a breach, exposing the reactive C substrate to the damaging oxidizing environment. As with CMCs, there also are some questions concerning longterm changes in the fiber-matrix interfaces that are at least partially responsible for the loss of impact toughness over time, as experienced in the Columbia accident. 19 UNCLASSIFIED//FOUO