AAWSAP DIRD Materials for Advanced Aerospace Platforms January 12 2010
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Defense
Intelligence
Reference
Document
Acquisition Threat Support
12 January 2010
! COD: 1 December 2009
DIA-08-0912-008
Materials for Advanced
Aerospace Platforms
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Materials for Advanced Aerospace Platforms
Prepared by:
Acquisiti on Support Di vision (DW0-3)
Defense warning Office
Directorate for Analysis
Defense Intelligence Agency
Author:
AAP Person 66
Administrative Note
COPYRIGHT WARNING: Further dissemination of the photographs in this publication is not authorized.
This product is one in a series of advanced technology reports produced in FY 2009
under the Defense Intelligence Agency, Defense Warning Office's Advanced Aerospace
Weapon System Applications (AAWSA Program. Comments or questions pertaining to
this document should be addressed toiMP Person 1 ! AAWSA Program
Manager, Defense Intelligence Agency, ATTN: CLAR/DW0-3, Bldg 6000, Washington,
DC 20340-5100.
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Contents
Introduction ...........................................................................................................iv
Launch Vehicles ................................................................................................. 1
Reusable Crew Modules...................................................................................... 9
Reusable Single-Stage-to-Orbit Vehicles ...................................................... 10
Advanced Al Alloys ....................................................................................... 11
Polymer Matrix Composites .......................................................................... 11
Al Matrix Composites .................................................................................... 12
Ti Alloys ....................................................................................................... 12
Ti Matrix Composites .................................................................................... 14
Ni-base Alloys .............................................................................................. 16
Refractory Metal Alloys ................................................................................ 18
Ceramic Matrix Composites .......................................................................... 18
Carbon-Carbon Composites .......................................................................... 19
Titanium Aluminides .................................................................................... 20
Propulsion Systems .......................................................................................... 21
Summary and Recommendations ..................................................................... 22
Figures
1. Schematic Diagram of Friction Stir Welding ....................................................... 2
2. Specially Modified 747 Transporter Unloading a Boeing 787 Composite
Fuselage Barrel Section ..................................................................................... 4
3. Micrograph Showing Phase Formation at the Surface of a Ti Alloy That has
Been Exposed to Air at Elevated Temperature ................................................. 13
4. Cross-Section Micrograph of a Ti Matrix Composite .......................................... 15
Tables
1. Potential Materials by Use Temperature Regime and Property ......................... 10
2. Example of Properties of Ti Matrix Composites ................................................ 14
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Materials for Advanced Aerospace Platforms
Introduction
"Advanced aerospace platforms" is a broad topic that can be divided into
several narrower subtopics to enable a more concise discussion of materials
advances, challenges, and opportunities. Consequently, this document
discusses the areas of launch vehicles, space vehicles, and space propulsion
systems separately because their key requirements are often application
specific, which affects materials selection decisions. In addition, single-use
and reusable boosters have different durability requirements that directly
impinge on design and materials selection. Furthermore, current engineering
practice has evolved to the point that design synthesis must integrate the
structure and construction materials to achieve optimum product performance.
For example, the space shuttle was designed to meet customer-imposed
mission requirements (range, payload, empty weight, landing capability, and
so forth) without significant real-time consideration of materials capability.
This approach led to significant compromises at later stages in the shuttle's
development and maturation. (Arguably, the shuttle could be designed as a
more efficient vehicle today.) In the extreme, a spectacular engineering
failure was the National Aerospace Plane (also dubbed the Orient Express),
which was launched as a military project and was intended to be a mach 12
reusable strike vehicle. This project rapidly became materials limited and was
canceled in 1993, after about $750 million in federal R&D expenditures and a
substantial private sector investment. The point is that any "clean sheet of
paper design" must start with an assessment of the requirements for
construction materials and be accompanied by a realistic assessment of the
capability of currently available materials to meet these needs. If these two
assessments indicate a gap between requirements and existing materials
capability, a risk assessment and a risk-mitigation plan must be developed
before expending engineering hours and funds.
Since the inception of manned space flight, the approach to design has
changed to include the concept of damage tolerance. This shift in design
philosophy was prompted by the (eventual) recognition that complex
structures cannot be designed and produced with zero defects. With the
maturation of fracture mechanics and means of reducing these concepts to
practice, the transition from zero defects to defect tolerance became the norm.
This new approach in turn led to recognition that high-performance materials
required not only high specific strength and stiffness but pacing increases in
strength with simultaneous improvements in fracture toughness and fatigue
crack growth resistance. The introduction of damage tolerance was
accompanied by a renewed emphasis on nondestructive inspection capabilities.
This latter thrust was driven by the need to demonstrate the capability to
reproducibly locate small flaws that could become failure initiation sites,
either because of static or because of cyclic loading conditions. In the case of
atmospheric flight, the U.S. Air Force has introduced standards for airframes
(the Aircraft Structural Integrity Program) and propulsion systems (the
Engine Structural Integrity Program) that tie structural life and reliability to
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this demonstrated inspections capability. Implementation of these standards,
starting with the B-1 bomber and the F-100 and F-110 engines, has
dramatically reduced (but not eliminated) the incidence of catastrophic
failures of critical components that endanger crews, vehicles, or both.
Taking these changes into account, this document outlines the current
situation regarding the design and production of high-performance structures
for aerospace platforms, including launch vehicles, space vehicles, and
propulsion systems for transporting space vehicles (and payloads) into orbit.
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Materials for Advanced Aerospace Platforms
LAUNCH VEHICLES
For the purposes of this document, launch vehicles are defined as the structure that
supports and/or encloses the propulsion system, the fuel supply, and the crew or
payload module. Launch vehicles today are either single use or multiple use after
recovery and extensive refurbishment. This approach adds considerably to the cost of
transporting a pound of payload into earth orbit, regardless of whether an unmanned
satellite or a manned orbiting crew module that must withstand the temperatures and
loads associated with safe reentry to earth. Furthermore, the larger the payloads are,
the greater are the reaction forces the launch veh icle must withstand during launch.
With the total weight of the payload, the empty weight of the launch vehicle, and fuel
all needing to be lifted initially, fuel-efficient propulsion and lightweight launch vehicles
are essential to maximizing the payload. Except in the area around the propulsion
system exhaust, the temperatures experienced by launch vehicles during launch are not
demanding. Therefore, advanced, high-strength aluminum (Al} alloys and polymer
matrix carbon fiber composites (PMCs) are prime candidates for the parts of the
structure that experience aerodynamic loads and where aerodynamic heating does not
exceed about 125 °Celsius. One class of advanced Al alloys is the lithium (Li)-bearing
alloys, such as Al alloy 2090. This alloy contains enough Li to reduce its density by 8
percent while increasing the elastic modulus (E) by 10 percent. Other, newer advanced
Al alloys, such as 7050 and 2050, have been developed to have improved damage
tolerance. These alloys have excellent specific strength at or near room temperature
and experience no major loss of ductility at cryogenic temperatures. The newer variants
of the 2000 and 7000 Al alloys also have substantially improved resistance to most
types of corrosion, including exfoliation and stress corrosion cracking. This can be
important in a reusable vehicle.
Perhaps the most important aspect of the improved Al alloys is their higher fracture
toughness, accomplished through a combination of alloy composition control and
improved processing. In alloy composition control, the concentrations of the residual
elements iron (Fe), chromium (Cr), manganese (Mn), and silicon (Si) are reduced at
the ingot stage. These elements combine with Al to form hard, brittle intermetallic
compounds known as constituent phases. The advanced alloys contain fewer, smaller
constituent phases, leading to improved fracture resista nce and higher fracture
toughness values. In applications such as body skins for commercial aircraft, this
improved toughness has enabled an increase in t he spacing of the circumferential
fuselage frames, or "hat sections," that serve both as stiffeners and as crack stoppers
to prevent a catastrophic failure during pressurization. For any given operating stress
in this case the pressurization stress-the spacing of the frames is directly related to
the critical crack size of the body skin . Higher toughness alloys have larger critical crack
sizes, and the frames can be spaced further apart without increasing t he risk of
catastrophic fa ilure. The increased spacing ultimately allows a fuselage design that
requires fewer frames. Consequently, the airplane benefits from a commensurate
reduction both in weight and in manufacturing cost. Similar possibilities exist for the
design of a fail-safe launch vehicle that has a lower empty weight. Clearly, the
advanced Al alloys offer intrinsic improvements over the alloys used in the Saturn
launch vehicle and introduce the prospect of new, more efficient launch vehicle designs.
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The newer Al alloys also can be specially processed to render them superplastically
formable . This capability opens a realm of possibilities to replace structures that, in the
absence of this capability, are machined from thick plate. Very large structures are
produced in sections that must be joined. Conventional fusion welding techniques do
not work for high-strength Al alloys such as 7075, 7050, 2024, or 2050 because either
the welds lead to cracks or the welds made under conditions that avoid cracking have
greatly reduced tensile properties. Because these alloys are not amenable to welding,
heavier, fatigue-prone mechanically fastened joints must be used. Recently, scientists
developed a joining process that permits joining of Al alloys such as 7050. This process,
called friction stir welding (FSW), allows joint designs in a variety of configurations that
were not considered possible when fusion welding was the only alternative.
In FSW, a rotating steel tool is inserted into the seam between the two Al alloy pieces
to be joined. As the rotating tool is driven forward, the friction between the tool and the
work piece generates enough heat to soften the Al alloy without melting it. A schematic
of this process is shown in Figure 1.
Tool
retreating
/
side
Figure 1. Schematic Diagram of Friction Stir Welding
The combined action of the rotation and the traversing of the tool essentially kneads
the two pieces together, leaving a mechanically sound joint. Although the weld
properties may be somewhat inferior to those of the base metal, they are good enough
that a relatively small increase in thickness at the joint position can compensate .
Although substantial development of the FSW process is ongoing, FSW already has
been put into practice. For example, the current external propellant tank on the space
shuttle is made from an Al-Li alloy fabricated through FSW. The weight advantage of
using welded as opposed to bolted joints in a large structure such as a launch vehicle is
considerable. With earlier high-strength alloys such as 7075, concerns about fracture
toughness in conjunction with monolithic structures would have caused a welded
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construction to be considered too risky. Today, the combination of higher toughness
alloys and FSW opens up the possibility of greater design flexibility resulting in lighter
large structures with equal or greater reliability than ear lier ones.
In sum, metallic, nonreusable (at least nominally so) launch vehicles made from
advanced Al alloys and fabricated through FSW constitute an incremental but significant
improvement over earlier versions .
In recent years, PM Cs have matured significantly. For many components that are not
exposed to elevat ed temperatures, PMCs provide a degree of design flexibility not
read ily available in meta ls. Consequently, PMC materials have begun to supplant Al
alloys in the construction of commercial subson ic aircraft. The use of PMCs in the
empennage of the Boeing 777 was one of the first examp les of Al alloys being
displaced. Subsequently, the new Boeing 787 has more structure made from
composites than from metallic materials. Once PMCs are introduced into a structure in
sign ificant quantit ies, a constraint related t o galvanic incompatibility between the PMC
st ructure and any adjoining Al all oys also is introduced. When a PMC structure is in
direct contact with an Al alloy structure, catastrophic corrosion of the Al alloy
components can occur. In the Boeing 787, the remedy for this concern is the use of
titan ium (Ti) all oys in areas where there is direct contact between t he metallic and the
PMC structures. This is directly analogous to the plastic bushing a plumber puts in the
j oint between copper and iron piping. Notwithstanding this constra int, the specific
strength and stiffness of PMC structures make a compell ing argument for their
appl ication in high-performance structures, such as launch veh icles .
Composite structures can be manufactured using one of th ree methods: hand layup of
pre-preg, automated tow placement, and resin transfer molding.
• The most rudimentary of these, but also the most flexible, is hand layup of pre -preg.
This method uses sheets of material that contain both the fiber and the polymeric
matr ix (called pre-preg). The polymeric matrix can be either a thermoset (for
example, epoxy) or a thermoplastic. Individual plies are cut from the pre-preg
typica lly using a numerically controlled laser or mechanical cutting device and are
la id up to form the desired shape. Areas that have heavier loads contain more plies
locally, and the plies are cut in an or ientation with respect to the fiber direction in
the pre-preg to ach ieve the desired strength relative to the principa l load path.
These plies are carefully placed according to a drawing (blueprint), making hand
layup a labor-intensive process and, therefore, making parts made using th is
method expensive . During ply placement, it is crit ical that no ply wrinkles are
introduced, as these create severe reductions in the local load-bearing capability of
the fina l component. Once all the plies are in their proper places , the article is
placed in a vacuum-tight bag that is evacuated and placed in an autoclave for curing
of the epoxy matrix or fusing of the thermoplastic. A disadvantage of a pre-preg
whose matrix is a thermoset is limited shelf life. In practice, this is managed to a
degree by storing the pre-preg in a freezer to slow the rate of chemica l reaction t hat
sets the epoxy. However, th is does not completely halt the reaction, causing these
materials to have a shelf life beyond wh ich they are not easily manipulated during
layup and do not develop full strength after curing in the autoclave. An additional
issue is out t ime-the time the pre-preg can be out of the freezer during layup
before the reaction proceeds at an accelerated rate and reaches a point at which the
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material is not suitable for the reasons stated earlier. Clearly, the time required for
layup places practical limitations on component size.
• In automated tow placement, thin ribbons of a pre-preg are fed off a drum or rolled
into a computer numerically controlled machine that places them in the desired
position. In principle, this process trades recurring labor cost for up-front capital
investment (the tow placement machine) and programming time . If the anticipated
volume of identical parts is high enough to amortize the capital investment and,
particularly, the programming cost, this can be an attractive means of reducing
manufacturing costs. For axisymmetric shapes, such as cylinders, this essentially
becomes a winding process and is quite efficient. An example of a finished
composite fuselage barrel section for the Boeing 787 is shown in Figure 2. For more
irregular three -dimensional shapes, such as a spar or a strut, placing the tows
becomes much more difficult and presents a fundamental limitation. Consequently,
PMC structures with complex shapes are still for the most part made using the hand
layup process. A variant of automated t ow placement is compression, whereby a
preform, made by automated tow placement, is forced by a press into a preshaped
die. This process allows fabrication of more complex shapes, but the rigidity of the
fiber and the extreme anisotropy of the tows can lead to wrinkles, which are not
acceptable because of the reductions in properties these cause.
Figure 2. Specially Modified 747 Transporter Unloading a Boeing 787 Composite Fuselage Barrel
Section
• The third main composite fabrication method, resin transfer molding (RTM), begins
with a woven fiber mat or preform. The polymeric matrix is injected into this mat to
create a fully dense composite. The major benefit of RTM is that it permits use of a
three-dimensional weave that minimizes the risk of delamination between plies.
Note that, with the other two methods, the material is reinforced in only two
dimensions (the plane of the pre-preg or tows). RTM's limitations include the
viscosity of the resins used. If the resin is too viscous, injecting it will either distort
the fiber architecture of the woven preform or not fully penetrate the preform,
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leaving voids. This limitation makes it difficult to use many thermoplastic resins that
are otherwise attractive because they are recyclable and have much longer shelf
lives than thermosets (epoxies). Another, more obvious limitation of RTM is
component size. This is in part because weaving of very large fiber preforms is
challenging and requires a very large weaving machine. It also requires a large
injection mach ine capable of multiple injection sites to ensure the complete
infiltration necessary to avoid formation of voids.
In practice, there are several significant challenges associated with manufacturing large
PMC structures today. Among these, perhaps the most significant are manufacturing
cost and the difficulty associated with making large, nonaxisymmetric components.
Large axisymmetric components that can be produced by winding tows of pre-preg are
clearly possible today, as demonstrated by the Boeing 787 fuselage whose section is
shown in Figure 2. For other shapes, because the load-bearing capability of PMCs
depends on the transfer of externally imposed loads to the strong, stiff fibers, joints
that intersect the principal load path become problematic. This is because the fibers
and, as a direct consequence, the load path are discontinuous, requiring local section
size adjustments to offset this local structural inefficiency. Thus, achieving the most
structurally efficient use of PMCs requires monolithic structures with continuous fibers.
In large structures, this is at best a challenge. Even in the Boeing 787's PMC fuselage,
mechanically fastened joints are used to connect adjacent barrel sections. In the
fuselage, as in any cylindrical pressure vessel, the principal stresses are hoop stresses
stemming from pressurization during flight. However, the compromise in structural
efficiency is minimized by the circumferential orientation of the joints. Although these
bolted joints add weight, the overall structural efficiency of the PMC structure is still
better, albeit considerably more expensive, than a longitudinally and circumferentially
stiffened Al structure. The circumferential joints also create discontinuous longitudinal
crack paths that improve the structure's damage tolerance. Some of the added expense
of the PMC fuselage stems from the use of Ti alloy fasteners because of the galvanic
coupling issues that would accompany Al or steel fasteners.
For other applications, which are limited by different material properties, the PMC
system can be tailored to optimize structural performance. This is possible because
both the matrix and the fiber can be independently selected. Moreover, the fiber
"architecture" (fiber orientation, weave geometry, and fiber volume fraction) can be
varied spatially to optimize load-bearing capability under complex stress states. For
example, again drawing on recent applications in commercial aircraft, the fan blades of
the large, high-bypass-ratio turbofan engine {GE90) produced by General Electric for
the Boeing 777 are made from PMCs. The limiting design consideration for these fan
blades is resistance to bird strikes. To optimize the PMCs' impact resistance, a medium
modulus, high-tensile-strength carbon fiber was selected in combination with a
thermoplastic toughened epoxy matrix. Furthermore, the fiber architecture was set to
optimize the bending strength under the impact of a bird. The GE90 fan blades are
produced by hand layup and are quite costly to produce. In the 10-plus years that
these fan blades have been in service, not a single unscheduled engine removal related
to the PMC fan blades has occurred. The competitive fan blade technology for B-777-
class engines is hollow Ti, which is used by both Pratt & Whitney and Rolls Royce. By all
informal accounts, these blades are cost intensive. This example supports the unwritten
rule that the pathways leading to high-performance, high-value structures typically are
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not technology dependent but are cost intensive, no matter what technology is
employed to meet the requirements.
An additional characteristic of PMC structures-one related to the laminated
construction of components made by hand layup or automated tow placement
methods-is their susceptibility to formation of delaminations when impacted
perpendicular to the plane of the plies. This is due to the mismatch in bending stiffness
between adjacent plies that have different unidirectional fiber orientations. This
mismatch causes shear stresses to develop that can exceed the shear strength of the
interlaminate bonds, causing small, embedded cracks to form. Under subsequent
inplane compression loading, the laminates bow because of the Poisson stresses and
separate because of the lack of an interlaminar bond to hold them together. In
significant compression loads, the laminates buckle, and this leads to structural failure.
This phenomenon, called compression after impact, is an insidious failure mode because
the delaminations are not externally detectable unless ultrasonic inspection methods
are used. Sources of such an impact include dropped tools, foreign objects (for
example, meteorites), and, perhaps most commonly, hail storms. Ultrasonic inspection,
if required, is expensive and time consuming. The latter concern in turn affects vehicle
availability and turnaround time. PMC parts made using RTM typically have reinforcing
fibers in the through-thickness direction, so concerns about compression after impact
are minimal.
The maximum temperature at which PMCs can be used is limited by PMCs' susceptibility
to oxygen degradation of the polymeric matrix. The maximum-use temperature for
prolonged exposure is determined by the thermal oxidative stability (TOS) of a
particular resin. The TOS, like any chemical reaction, is determined by both time and
temperature . The glass transition temperature (T9 ) of the polymeric matrix also
imposes strength and dimensional stability limitations independent of the TOS limits.
This is particularly true for thermoplastics. However, the TOS limits usually impose
lower temperature limits than T9 if prolonged thermal exposure is contemplated. Three
distinct classes or groups of resins exist, each with a successively higher temperature
capability. These are conventional epoxies and most thermal plastics, bismaleimides
(BMis), and linear polyimides. The first class is limited by TOS to about 125° Celsius.
BMis can be used to about 175° Celsius. Linear polyimides, such as the in situ
polymerization of monomer reactants (PMR) group of thermosetting formulations, can
be used to about 300° Celsius. Many of the PMR resins contain the hazardous
compound methylenedianiline, which requires special care during use, including
protective clothing to limit personnel exposure (for example, during ply cutting and
hand layup). This requirement reduces productivity, adds cost, and creates a degree of
liability concern for the manufacturer of the PMC components. For RTM, the BMis and
PMR resins typically have higher viscosity and require commensurately higher injection
temperatures to reduce the viscosity to manageable levels for reasons discussed
earlier. New resins are constantly being developed, including some with attractive
properties. However, many of the sources of these resins are startup companies that
exist on R&D fund in g, often in the form of U.S. government SBIR (Small Business
Innovative Research) projects. Such companies are good at innovation but often have
limited experience transitioning new products from the laboratory to large -scale
production. Furthermore, in the current economic climate, access to sufficient capital to
set up production-scale capacity can be a form idable problem for a small company.
Larger companies (for example, BASF, DuPont, GE Plastics) typically are not interested
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in materials with small annual sales volume, so even licensing the new material to one
of these companies may not be commercially feasible .
This foregoing discussion highlights a commonly encountered inconsistency between
technical innovation and commercial progress. Until a production-scale source exists,
pricing of new materials is at best highly uncertain and potentially unstable over time.
This is in part due to the uncertainty surrounding demand and the associated volume of
material that will be required. Taken together, these factors act as a clear deterrent to
the adoption of new materials of literally all classes (polymers, metals, and ceramics).
For heavily loaded structures or structures that will experience temperatures higher
than about 200 °Celsius, Ti alloys are the preferred material class. Ti alloys are about
half as dense as steel or Ni-base alloys and possess a density-corrected strength and
stiffness competitive with that of other metallic materials. Numerous grades of Ti alloys
are in use today, and a strong domestic industrial base of suppliers exists for nearly all
these grades . Ti alloys for structural applications can be divided into three groups based
on their metallurgy: near a alloys, a+p alloys, and metastable p alloys (commonly called
p alloys for short) . Both the aircraft and propulsion original equipment manufacturers
and several private engineering firms have extensive design experience with Ti alloys.
Although most of this experience is with Ti-6Al-4V (Ti-6-4 ), other alloys also are widely
used, particularly in jet engines and liquid-fueled rocket engines. The alloy most likely
to be used in a launch vehicle, Ti-6-4, has been available and in use for more than 40
years but is still highly competitive with newer grades in large part because of its
versatility. Potential applications of other alloys that have particularly attractive
characteristics, such as a higher temperature capability, are discussed in the Reusable
Reentry Vehicle and Propulsion Systems sections of this document.
The most commonly used structural grade of Ti alloy today, Ti-6-4, can be readily
fusion welded, formed both by forging and as a sheet product, and conventionally
machined, although each of these operations requires special precautions. In addition,
the feasibility of friction stir weld ing has been demonstrated. Ti-6-4 also can be
superplastically formed and diffusion bonded, enabling synthesis of innovatively shaped
components. For launch vehicles, the most likely application for Ti alloys is in the
structure that carries the reactions from the propulsion system to the vehicle itself.
These applications typically involve heavy sections to accommodate the large loads,
and the main property requirements are high strength, fatigue resistance, and fracture
toughness. Ti alloys have been used in both military and commercial aircralt. Heavily
loaded components in service today include the wing box of the B-1B bomber (Ti-6-4),
the landing gear beam in the B-747 (Ti-6-4), and the landing gear truck beam in the B-
777 (higher strength Ti-10V-2Fe-3AI [Ti-10-2-3]). The choice of Ti-10-2-3 for the
landing gear truck beam reflects the time-based maturation of the newer p-Ti alloys,
such as Ti-10-2-3. These alloys have the advantage of being "deep hardenable"
compared with Ti-6-4, they can develop full strength in thicker sections during heat
treatment . For example, that the B-777's truck beam is up to 6 inches th ick in some
locations factored significantly in the choice of Ti-10-2-3. Because Ti alloys are about
half as dense as steel, they are very competitive on a density-corrected basis. However,
mass is not the only driver for some applications; the volume of a component also must
be compatible with the space available for it. This factor also is a consideration in the
choice ofTi -1 0- 2-3 for the B-777 truck beam. And it also applies to an application such
as land ing gear, because it is retracted into the fuselage for aerodynamic reasons.
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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-m 0 •5 ,
making damage tolerance marginal. For example, using a design stress that is two
thirds 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 requ irements. 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
conta in 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 th is 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 capab ility 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 veh icle, 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.
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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 t he 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 ree ntry glider devoted considerable time and
resources to examining the use of refractory alloys such as Mo-0.5%Ti (Moly half Ti)
and several Columb ia-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 MoSii is
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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 mater ials 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 in itial 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 Specific Specific Fatigue Fracture Creep
Regime Strength Stiffness Resistance Toughness Resistance
PMCs;
Advanced Al PMCs;
PMCs;
Ambient up to alloys; AMCs; PMCs; Ti alloys;
AMCs;
250 °C AMCs * ; Ti Alloys; Ti alloys TMCs
TMCs
Ti alloys; TMCs
TMCs *
250 °Cup to Ti alloys; Ti alloys; Ti alloys;
TMCs Ti alloys
550 °C TMCs TMCs TMCs
Ni-base Ni-base
alloys; alloys;
Ti Ni-base Ti
Ti Ni-base
aluminides; alloys; aluminides;
aluminides; alloys;
Above 550 °C Refractory CMCs; CMCs;
CMCs; Refractory
metal Refractory Refractory
c-ccs metal alloys
alloys; metal alloys metal
CMCs * ; alloys;
C-C Cs* C-C Cs
AMCs = Al matrix composites; TMCs = titanium matrix composites; CMCs = ceramic matrix composites; C-CCs =
carbon-carbon composites
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Table 1 shows there are essentially 10 distinct classes of candidate materials for use in
a reusable 550 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 S5O veh icle.
Advanced Al Alloys
Th is 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 averaging 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 lim itation. 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 pheno lic 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 carefu l 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
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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 a phase layer on the surface known
as a case. An example of a case is shown in Figure 3.
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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?
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• Once a 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 th is 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 a case during the SR-71's service. Notably, the primary alloy used for
the SR-71 was one of the original ~-Ti alloys, B-120 VCA, the composition of which is
Ti-13V-11Cr-3AI. The primary reason for choosing this alloy was that it is much easier
than any of the a +~ Ti alloys are to roll into sheet gauges. Today there are newer ~- 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-3AI. However, the
successful use of B- 120 VCA raises the question of whether ~- Ti alloys are more
resistant than a+ ~ alloys such as Ti-6-4 are to a 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 strenqth 276 ksi / 1902 MPa
Young's modulus 32.8 msi / 226 GPa
Strain to fracture 0.95%
Density 0.16 lb/in 3 / 4.43 q/cm 3
Fiber Volume fraction 0.39
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Figure 4 shows an example of a TI matrix composite (TMC) cross section.
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).
Materia l 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.+13 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, t he 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.
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• 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
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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 Ni 3AI (called y'),
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 ord inary 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" (a)
alloys (the so-called low-a alloys) . An example of an application for which low-a 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 therma l expansion, resulting in much
lower a values. It was recognized that this concept could also be adapted to Ni-base
structural alloys. The first-generation low-a 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-a alloy, Inco 909,
contained Si to help improve the oxidation resistance, but t his alloy has not gained
widespread acceptance, possibly because Inco 903 problems were so severe. Low-a
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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 550 vehicle is the
durability of these coatings under very high mass flow conditions, such as during
reentry.
Refractory Meta I 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.ST m 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 550 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 Al 2O3 or SiC,
whose role is to bridge and arrest the cracks. Under continued loading, the fibers also
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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 add itional 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 fiber
matrix 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 su itable
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 long
term 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.
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Therefore, instabilities that cause time-dependent reductions in toughness must be
understood before these materials can be considered candidates for future TPS
applications.
Titanium Aluminides
Titanium aluminides are intermetallic compounds that form between Ti and Al. There
are three such compounds-TbAI, TiAI and AbTi-but the one of principal interest is
TiAI, often referred to as gamma titanium aluminide or y-Ti AI. This compound contains
about 35 weight% Al, but ternary and quaternary alloys based on TiAI typically contain
somewhat less than th is amount. y-TiAI is interesting for intermediate-temperature
applications for several reasons. These include a lower density, at least a 125 °Celsius
higher temperature capability, and better surface stability (oxidation resistance)
compared with conventional Ti alloys. The surface scale that forms during oxidation of
y-TiAI is Al rich and is more protective than the TiO2 scale that forms on conventiona l Ti
alloys. Like most other intermetallic compounds, y-TiAI has very limited ductility at
temperatures up to about 600 °Celsius but is not brittle in the classical sense; that is, it
has the capabi lity to deform plastically before fracturing. The ductility increases as
temperatures increase and is quite good at normal service temperatures.
y-TiAI alloys' limited ductility poses manufacturing problems for any wrought products.
These issues are manageable but reduce product yields, which in turn impacts cost.
Recent work in Austria has demonstrated that some y-TiAI alloys can be made into
sheet. This is particularly relevant to applications such as metallic TPS. Other work has
demonstrated that some y-TiAI alloys can be conventionally forged, albeit with care.
Should one or more of the efforts currently under way to produce low-cost, prealloyed
Ti powder prove successful, the availability of affordable y-TiAI alloy powder could be a
major cost breakthrough.
The development and maturation of y-TiAI alloys has taken more than 30 years. This is
in part due to the lack of actual production applications because of designer's concerns
regarding the limited low-temperature ductility of all the y-TiAI alloys. Today, finally,
there are several applications of alloys based on y-TiAI, the most significant of which are
two stages of low-pressure turbine blades in the engine General Electric is providing for
the Boeing 787 and the growth 747 (called the 747-8). These blades are cast to near
net shape and will operate at temperatures up to about 750 °Celsius. Another
application is a cast y-TiAI turbocharger rotor Mitsubishi uses in a production auto
because it has half the mass of an equivalent Ni-base alloy component.
A potentially ideal, high-volume application of y-TiAI alloys is exhaust valves for autos.
If the low-cost powder mentioned earlier becomes a reality, th is factor, coupled with
the current pressure to improve fuel consumption in American cars, could make this
prospect a reality. The point of this short discussion of non-space-related applications of
y-TiAI alloys is that rea lization of any or all of these appl ications will create a stronger
production base for this class of alloy, which in turn will make their introduction into
future space platforms easier and more cost effective.
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PROPULSION SYSTEMS
Among propulsion systems, only reusable rocket engines are considered here because
they hold potential for significant progress that could reduce the cost of placing
payloads into orbit.
The basic concept of the space shuttle main engine is still viable, but the durability of
the materials used to make the hardware has been a major expense and source of
concern for NASA. Much of the concern is related to the effects of hydrogen on the Ni
base alloys used in the turbo pumps. The turbo pumps are the heart of a liquid-fueled
hydrogen-oxygen rocket engine because they deliver the fuel and oxidizer to the thrust
chamber. The turbo pumps rotate at a very high speed-up to 35,000 rpm. Such
speeds create enormous centrifugal stresses in the rotating components, particularly
the disks that hold the air foils used to extract work from the hot gas stream. One side
of the turbo pump rotor operates in the hot gas stream created by the combustion of
the oxygen-hydrogen mixture; the other end is a cryopump that operates at cryogenic
temperatures, either in liquid oxygen (90 K) or in liquid hydrogen (20 K). The hot side,
a turbine, resembles the turbine rotor in a gas turbine engine and is made of many of
the same Ni-base alloys. The cryopump is made of Ti alloys.
The turbine essentially operates in a hydrogen-rich supercritical steam environment at
a maximum temperature of about 1,050 °Celsius. The turbine disk is a forged Ni-base
alloy, and the air foils are single-crystal investment castings. Neither material is well
suited to operate in a hydrogen-rich environment; however, no other material class can
withstand the operating temperatures and has better hydrogen tolerance-a situation
that still prevails today. The mechanism of hydrogen-induced cracking is much better
understood today as the result of extensive research over the past 25 years. Therefore,
it would be very useful to use this improved understanding to design a Ni-base alloy
that has improved hydrogen tolerance. If successful, a turbine that has improved
resistance to hydrogen cracking would greatly reduce the intermission refurbishment
time and cost.
Traditionally, the cryopump rotating parts have been made from forgings of the near-o,
phase Ti alloy Ti-SAl-2.SSn (Ti-5-2.5). This alloy has been chosen for its superior
notched tensile strength compared with Ti-6-4 when tested at cryogenic temperatures .
Hindsight suggests it is unclear that notched tensile strength is the best criterion for
selecting a cryogenic rotor material. The use of notched tensile strength originated with
steels. Here, the hydrostatic stress state at t he notch root could trigger the onset of
brittle fracture of the type seen in smooth tensile tests below the ductile brittle
transition temperature (DBTT). Ti alloys do not exhibit a DBTT; therefore, true fracture
toughness measured at the re levant temperature is a more accurate indication of the
fracture resistance of the rotor. An examination of the limited available fracture
toughness data for the two alloys shows no clear advantage in using Ti -5-2.5. Because
Ti-5-2.5 is more difficult to produce than Ti-6-4, using Ti-5-2.5 adds cost to the
cryopump. Other, newer, higher strength Ti alloys-for example, several of the newer
~-Ti alloys-may be even better suited than Ti-6-4 is for this application. The critical
question of how much toughness is real ly required cannot be answered at present
because t he notched tensile ratio has been the deciding criterion. Therefore, a new
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design of a cryopump potentially can benefit from these new materials and from
abandoning the use of notched strength as a design paradigm .
The combustion chamber of a rocket engine is made from a copper alloy and is cooled
by passing liquid hydrogen through channels in the outer wall. The space shuttle main
engine combustion chamber is made of NARloy-Z, a Cu-3%Ag-0.5%Zr alloy that has
excellent thermal conductivity and better strength than the commercially available Cu-
0.15%Zr alloy known as AMZIRC. While a lighter combustion chamber would be
welcome, developing hydrogen-resistant turbine materials would be a much better use
of available rocket engine alloy development resources.
SUMMARY AND RECOMMENDATIONS
This document has attempted to review a range of materials that may have promise for
all aspects of aerospace platforms. Because of the breadth of this topic, discussion has
focused on opportunities and associated risks, but with little technical detail. Practical
issues such as manufacturing capabilities and costs and the availability of materials also
have been addressed where appropriate. Another recurring issue is the ability to
achieve better design efficiency through a design synthesis process that concurrently
treats form, fit and function, manufacturing capability, and materials capability as
equally important design constraints. Although this approach has yet to be successfully
used for a major high-performance structure project, engineering is nearing a state of
development where it may now be feasible.
If the space shuttle is used as a benchmark, it is clear that numerous opportunities will
exist to improve structural efficiency. This possibility reflects the progress made since
the shuttle and its engine were designed. Improved durability and reduced operating
cost also are possible. The issue of long lead time holding back these improved design
methods is most often one of timely availability of new, attractive materials that are
mature enough to be used in a system without increasing the risk of failure. The
materials community is actively developing and using computational models and
simulation methods to address this concern. Despite some progress, further advances
are required to bring computational materials engineering to the desired level of
maturity.
1. Hatch, J. ed . Aluminum, Properties and Physical Metallurgy. American Society for Metals, 1984.
2. Chawla K. Composite Materials, Science and Engineering 2 nd edition. Springer, 1998.
3. Baker, A., Dutton, S. and Kelly, D. Composite Materials for Aircrah Structures 2 nd edition. AIAA, 2002.
4. Lutjering, G and Williams, J. Titanium, 2 nd edition . Springer, 2007.
5. Sims, C. and Hagel, W. The Superalloys. Wiley, 1972.
6. Mishra, R. and Ma, Z. Friction Stir Welding and Processing, Materials Science and Engineering. v .50, pp. 1-78,
2005.
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