1
advantages and limitations
General Electric Co is developing hot-section ceramic components for
future versions of its popular 7F gas turbines, reports Lee S Langston,
professor emeritus, UConn, who began his career as a research engineer
at Pratt & Whitney more than 50 years ago. Among the first will be a
ceramic matrix composite (CMC) shroud for the first stage on new 7F.04
and 7F.05 machines; it also is an option on full Advanced Gas Path
upgrades to the earlier 7F.01, .02, and .03 models.
CMCs, at about one-third the weight of conventionally used hot-section
nickel/cobalt alloys they replace, are more heat resistant and require less
cooling air. These properties promise enhanced engine durability, life,
fuel economy, and performance.
Ceramics have many favorable characteristics. Compared to metals now
used in gas turbines, they often can have superior corrosion resistance,
hardness, lower density, and high-temperature capability. Their main
drawbacks are comparatively low toughness and the possibility to
fracture in a catastrophic brittle mode.
Toughness is a measure of load or stress needed to drive a crack
through a material. A china dinner plate is not easy to break in half,
Langston points out, but if it has a slight crack, fragmentation is easy,
compared to, say, an identical ductile metal plate.
Ceramics subjected to compressive stresses, where crack defects are
made smaller, are very strong. Ceramics subject to tensile or bending
stresses (such as in rotating turbine blades), where crack defects are
made larger, can cause sudden failure. CMCs have been developed to
alleviate this characteristic ceramic brittle behavior.
Graceful failure. GE is introducing CMCs
both in its 7F gas turbines and latest jet engine for single-aisle
commercial aircraft—such as the Boeing 737 and the Airbus A320. Called
LEAP (Leading Edge Aviation Propulsion), it is scheduled to enter into
service this year. Since the LEAP CMC program is further along than that
for the 7F, let’s review it.
CMC first use in LEAP is as the shroud for the first-stage high-pressure
turbine (HPT). Note that the shroud is the segmented inner structure of
the engine casting and the closest stationary surface to the rotating firststage turbine blade tips. There are 18 segments in the LEAP engine
casting’s inner structure.
GE will expand its application of CMC use in the company’s 100,000-lbthrust GE9X engine, now under development for Boeing’s 777X airframe
and scheduled to enter service in 2020. It will feature CMC combustion
liners, HPT stators, and first-stage shrouds. Early in 2015, GE ran tests
on a turbine rotor with CMC blades—the ultimate structural test of this
new material. One can speculate that GE’s use of CFMs in its 7F gas
turbine will follow a similar path taken for the GE9X.
The LEAP CMC first-stage turbine shroud is a composite consisting of fine
intertwined ceramic silicon carbide (SiC) fiber, embedded in, and
reinforcing, a continuous silicon carbide-carbon (SiC-C) ceramic matrix.
The shroud also has an environmental barrier coating to protect the CMC
from chemical reactions with the turbine gases.
The CMC SiC fibers are continuous (greater than about 2 in. long), a
fraction of a human hair in diameter, and relatively free of oxygen
(which can degrade its high-temperature properties). The resulting
intertwined fiber reinforcers are covered with a multi-layer coating based
on boron nitride.
The fiber-reinforced CMC has a unique failure mechanism, which
Langston refers to as a “graceful failure” mode. As the SiC-C matrix
cracks develop under imposed thermal or mechanical stresses, the load
is transferred to the reinforcing SiC fibers. Their multi-layer boron nitride
coating then permits the fibers to slide in the matrix, allowing load
transfer and energy absorption. Thus multiple micro-cracks build up,
prior to actual fracture, resulting in increased toughness, and imitating
the ductile behavior of a metal.
This crack-mitigation tolerance, which resists the classic brittle failure of
a pure ceramic, should also yield gas-turbine parts that are not highly
sensitive to manufacturing flaws.
In sum, GE’s use of CMC gas-path parts looks very promising. CMC’s
graceful-failure mechanism will allow the use of this promising composite
ceramic, with its light-weight and high- temperature characteristics. GE
estimates an advantage of at least 180 to 360 deg F compared to metals
currently in use.
This means CMC parts could operate at about 2400F, well into and above
the softening/melting point of superalloys. (Pratt & Whitney estimates a
CMC operating temperature at 2700F.) Currently, CMCs are very
expensive—hundreds to thousands of dollars per kilogram. GE is
counting on cost reduction by process scale-up, automation, and
improved machining.
CMC future. All-in-all, GE has been working on CMCs for two decades,
has spent over $1 billion on the technology, and recently opened a $125million plant in Asheville, NC, to mass-produce CMC gas-turbine
components. Just last October, the company announced an investment
of over $200 million to create factories in Huntsville, Ala, to mass
produce silicon carbide materials and manufacture CMCs for both
aviation and land-based GTs.
The brief account given here to describe the management of tensile
cracking does not do justice to the research, analysis, and testing GE
and others have done to develop CMCs for gas turbines, says Langston.
Trying to manufacture a ceramic material structure which imitates what
nature provides in a ductile metal is a challenge.
However, success does not always favor the pioneer. For instance, in the
late 1960s, Rolls-Royce attempted to be the first to use a composite
ducted fan on its then new RB211 engines, for the Lockheed L-1011
airframe. The fan, using a carbon-fiber composite Hyfil, failed final
testing, contributing to the 1971 bankruptcy of the company. The jet
engine industry has since developed successful composite fans, but the
inaugurating company got off to a rocky start.
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Challenges
With ceramic materials turbine would then operate at higher temperatures, yielding higher
power with smaller engine sizes. Ceramic material are known for their capability to withstand
high temperatures. In addition they are quite tolerant to contaminants such as sodium or
vanadium which are present in lost cost fuels and highly corrosive e to the currently used nickle
base supperalloys. Ceramics are upto40% lighter than camparable high temperature alloys .
they are also cost much less . their cost is around 5%cost of supper alloys.
Brittleness is the main problem
The demand for ever higher performance, particularly in military aircraft, will motivate major changes in
the types of material used in future generations of gas turbine engine. The need to increase
thrust/weight ratios from the current level of 8/1 to around 20/1 will require design changes, many of
which will only be possible with new materials. These new materials must be stronger, stiffer, have
defect tolerance and be capable of operating at temperatures beyond the capability of current
materials. The degree of property improvement necessary may only be attainable in composite
materials, various types of which are likely to become predominant in major engine components. Resin
matrix composites Resin matrix composite materials have been in production use for many years. They
were first used in the RBI 62 lift jet engine more than 25 years ago<". The basic requirement of that
particular engine concept was maximum thrust/weight ratio and minimum weight was a key
contribution. Since glass reinforced plastic (GRP) was more than twice as strong as the aluminium alloys
available at that time and had a 25% lower density (Table 1) it was used for compressor blades and
casings. Subsequent application of the RBI 62 as a booster engine to improve take-off performance in
Trident lilB aircraft required an uprate in temperature capability. This was achieved by replacing the
epoxy resin matrix by a polymide resin. In contrast to the strength, the elastic modulus of GRP is no
higher than that of aluminium alloys. In stiffness-critical applications, such as engine nacelle
components, it was necessary to use carbon fibre as a reinforcement rather than glass fibre. A carbon
fibre reinforced plastic (CFRP) cowling door on an RB211 engine is shown in Rg 1. Replacement of
aluminium with CFRP typically results in a component weight saving of some 25%. This resin matrix
composite development sequence illustrates the basic objectives of future composite development for
gas turbine engines. These are to achieve the benefits of Tablet Properties of unidirectionally reinforced
composites compared with aluminium Material Aluminium L73 GRP CFRP Nominal fibre vol ume fraction
0.6. Density (Iv1g/m3) 2.63 2.03 1.59 Tensile Strength (GPa) 0.42 1.03 2.10 Tensile Modulus (GPa) 69 45
130 Fig 1 CFRP cowling door on RB211 engine. MATERIALS & DESIGN Vol. 10 No. 5
SEPTEMBER/OCTOBER 1989 231 increased strength and stiffness together with good defect tolerance
and minimum density at temperatures appropriate to the operation of the advanced compressors,
turbines and exhaust/reheat systems of future designs of militiary engine. Metal matrix composites
Composites based on aluminium with particulate, whisker and fibre reinforcement have been available
for some time'^' There are production applications including aerospace and reciprocating engines. The
relatively small improvement in temperature capability over resin matrix composites limits the potential
of aluminium matrix composites in gas turbine applications however. Titanium matrix materials do offer
the possibility of composite property benefits at higher temperature, possibly up to 900°C with
appropriate matrix compositions. Potential improvement of some 50% in strength and 100% in elastic
modulus over conventional titanium alloys are theoretically possible in uniaxial composites containing
around 0.4 volume fraction of reinforcement with silicon carbide monofilaments. Satisfactory stiffness
has been demonstrated but further improvement in tensile properties is required. One of the factors
that has limited properties to date is the reaction between fibre and matrix during composite fabrication
and subsequent simulated service exposure at elevated temperature. The fracture surface of a three
point bend test piece which has seen elevated temperature exposure is shown in Fig 2. This shows
separation of interface material from both monofilament and matrix. Further work on the factors
controlling the fibre/matrix interface is one of the elements required before titanium matrix composites
can be considered as reliable gas turbine materials. Glass matrix composites Glass has been considered
as a matrix material for composites with moderate temperature capability, and the improvement in
toughness of carbon fibre reinforced glass over unreinforced glass is shown in Rg 3'^'. The failure strain
of the matrix is lower than that of the fibres and the composite failure sequence is matrix cracking (A-B),
initial failure of reFig 2 Fracture surface of titanium matrix composite. aoo 600 'c i Ui oc *-I/I 200 »0 1 1 1
1 1 / // 1/ A / / carbon reinforce d glas s / y\ glas s l/^ 1 1 1 1 1 1 2 i S t 10 STRAIN |10-'l 1 1 12 - c \i. Fig 3
Stress strain relationship for glass and carbon fibre reinforced glass (Vf=0.4). inforcing fibre (B) and
stress relaxation in broken fibres together with fibre pullout (B-C). Composites of pyrex reinforced with
Nicalon fibres have also demonstrated good toughness and have a temperature capability of around 500
C. Glass ceramic matrices including the alumino-silicates have the potential to increase this to around
1100 C, the temperature limit for the Nicalon fibre. Because of the non-stoichiometry of Nicalon, its
elastic modulus is only some 50% that of stoichiometric silicon carbide and this provides a further
limitation in composite properties. High temperature non-metallic composites The need for materials
with higher temperature capability than nickel superalloys for components such as turbine aerofoils, in
combustors and reheat/exhaust units has been discussed'"'. Most of the possible successor materials
have major problem areas, for example the inadequate defect tolerance of monolithic ceramics in
stressed applications. Simplistically there are two basic approaches to developing tougher high
temperature non-metallic materials. The first involves finding ways 232 MATERIALS & DESIGN Vol. 10
No. 5 SEPTEMBER/OCTOBER 1989 MM^i-y Fig 4 Silicon carbide composite turbine disk. of improving the
oxidation resistance of carbon/carbon composite materia). The second is based on developing
composite versions of monolithic materials which have inherent oxidation resistance. Carbon-carbon
composites have attractive mechanical properties with significant strength at temperatures up to
2800°C and good toughness. Oxidation resistance is inadequate for anything other than applications
involving short time exposure at temperature in oxidising conditions. This problem is being addressed by
the use of silicon carbide coatings on carbon-carbon components and by the incorporation within the
matrix and oxide forming elements to inhibit oxidation of any material exposed by cracks in the external
silicon carbide coating. Ceramics such as silicon nitride, silicon carbide and many oxides have inherent
oxidation resistance at temperature of interest. Their inadequate defect tolerance is being addressed by
the development of ceramic composite materials, one example of which, a turbine disk in SIC reinforced
Table II Properties of some fibres for use at high temperatures Fibre C B Nicalon A 203 SiC E (GPa) 250
400 230 360 450 Tensile Strength (GPa) 4 3 1,5 1.8 3,5 Dia jum 8 150 15 20 150 App ox. Temperature
Capability (C) 400 700 1000 1300 1600 MATERIALS & DESIGN Vol. 10 No. 5 SEPTEMBER/OCTOBER 1989
233 60. SO 40 . weight 30. 20. 10. 1 ^ \ Steel Nickel/^ ^^'^^x ' ^ ^ y^ /^Titanium ~-~. y / CartXMi ~^;>'=~----__Mjminii ^ composites WO 1970 1980 1W0 yn« Spey RB211 RB199 534E4 V2400 EJ200 Metal matrix
composites Ceramic malrixN, S^ompoaitevx. // 2000 2010 Ytv Fig 5 Predicted trends of material usage in
Jet engines. with Nicalon fibres, is siiown in Fig 4''". Tiiere are many problems which still must be
resolved in the development of composite ceramics, but the biggest single problem is probably the lack
of suitable fibres for reinforcement. The major high temperature fibres which currently exist are listed in
Table 2. The highest temperature capability and elastic modulus are obtained in stoichiometric SiC
monofilaments. which are currently only produced at 100/150 /xm diameter. This large fibre diameter
results in significant design limitations. Smaller diameter hicalon fibre is available but as indicated earlier
elastic modulus and temperature capability are reduced by its non-stoichiometry. Alumina fibres have
attractive stiffness but inadequate strength and carlxjn fibre, despite its high temperature strength in
inert environments, can only be used up to about 400°C in presence of air. A further general problem is
that the currently available fibres are chemically compatible with few of the required matrix materials
and control and stability of the matrix/fibre interface is critical to the generation of strength combined
with toughness. Nevertheless, considerable progress is being made and the use of composite materials
in gas turbines will progressively increase (Fig 5)'^' as the current design, material and manufacturing
technology problems are solved. It is predicted that future generations of jet engine will use composite
materials in major compressor and turbine components. References 1. H E Qresham and C G Hannah, J
Roy Aero Soc71 677 1967. 2. E A Feest, Materials & Design VII 58 1986. 3. R W Davidge, I'roc World
Congress on High Tech Ceramics (6th CIMTEC) 1986 Bsevier. 4. G W Meetham, Materials & Design IX 213
1988. 5. G Bernhart et al, "Developments in Science and Technology of Composite Materials" ed Bunsell
1985 p 475 AEMC Bordeaux, France. 6. N A Payne et al, FVoc World Conf on Adv Materials for
Innovations in Energy, Transportation and Communication Tokyo, 1987
Advatages
high damage tolerance

high stiffness (because a lot of airline structures are sized for stability)

active controls (which, if sufficiently reliable, may reduce the need for
high stiffness)

low-cost raw materials and fabrication methods

low density (with high strength-to-weight ratios)

means to assure that material properties are satisfactory following
repairs and have not degraded unexpectedly over the life of aircraft in which
composites have been incorporated

modularity

resistance to lightning strikes (an area where Boeing, for one, is
investing millions of dollars)
Improvements in performa

Rotor case study
Rotor blades are subjected to extremely harsh conditions, both operational and environmental.
Rotational tip velocities of approximately 200 m/s (~480mph), and "flapping" during flight, are
coupled with extremes in both humidity and temperature. The latter can vary from -40�C to +90�C.
So, a number of specific material properties are required for efficient and effective rotor blades.
Composites can be made that fulfil these property requirements.
Figure 3 (click the picture to enlarge)
The manufacture of rotor blades begins with the ultrasonic profiling of partially cured fibre reinforced
plastics known as pre-pregs, which allows the production of advanced shaped and sectioned blades.
Such components are virtually impossible to fabricate economically from metal. The contoured prepregs are then positioned, using a specific 'lay-up' pattern, within a mould. This is then closed,
crushing the material into the desired shape and form, and an external hydraulic pressure is applied.
Curing is completed by means of a computer-controlled process, during which the pressure is
maintained and the temperature slowly increased to 125�C. Finally, the blade construction is
finished with the simple adhesion of the honeycomb core between the two constituent blade layers,
which are illustrated in figure 3.
Many other desirable properties and characteristics are achieved by the use of composites, including
good strength-to-density ratios, which are four to six times greater than those of steel or aluminium.
The specific modulus of certain composites is also far greater than those of steel and aluminium,
leading to composite blades that are up to 45% lighter than their metal equivalents. In addition,
complex blades are much easier to process and manufacture, are joined with adhesives, negating
the need for riveting and simplifying assembly and can be produced using much cheaper tooling
than for metals.
Developments in composite materials such as carbon fibre reinforced plastic have allowed the
creation of rotor blades that far surpass their predecessors in every way, and continued research
into new areas of Materials Science will no doubt improve on these blades in the future.
http://classroom.materials.ac.uk/caseRoto.php
1 Aeroengines of the future are expected to have a
thrust-weight ratio of 20: I and much higher
efficiency, which may require a turbine entry
temperature (TET) of 2500 °K..
2. Studies on tapered composite shafts6 have
shown that while meeting the torsional strength
requirements, it was possible to obtain
configurations which resulted in an increase of
20-30 percent in natural frequencies and reduction
of 50-60 per cent in maximum dynamical stresse
3 Fibre reinforced composites in epoxy or
glass matrix appear to be ideal materi.als for the
shaft on the compressor end because of their high
specifict strength and modulus up to 250 °c and
400 °C,
4 For the shaft on the turbine
end, glass ceramic, ceramic or metal matrix
composites will be required. The LP stage
compressor dis9s may be ofcal'bon fibre reinforced
plastics (FRPs), while others can be of glass matrix
or metal matrix composites. Turbine discs have to
be of carbon-carbon, ceramic or metal matrix
composite.' Compressor blades of LP stages can be
of carbon FRP, while compressor blades of HP
stages may be of glass matrix or metal matrix
composites. Turbine blades have to be necessarily
of ceramic.
5 The basic idea to be adopted in the design of
various components is to have, as far as possible,
uniform stress, which should automatically lead to
an optimum weight because of nonisotropic nature
of composites and possibility of matching their
directional properties/strength with the direction of
loads.
6 Because of a higher E/p value of composites
compared to metals, larger bearing spans without
actually lowering the rotor critical speeds are
possible.
7