Titanium matrix composites (TMC's) are commonly made up of a titanium
alloy matrix reinforced by silicon carbide fibers that are oriented
parallel to the loading axis. These composites can provide high
strength at lower densities than monolithic titanium alloys and
superalloys in selected gas turbine engine applications. The use
of TMC rings with unidirectional SiC fibers as reinforcing rings
within compressor rotors could significantly reduce the weight
of these components (ref. 1). In service, these TMC reinforcing
rings would be subjected to complex service mission loading cycles,
including fatigue and dwell excursions. Orthorhombic titanium
aluminide alloys are of particular interest for such TMC applications
because their tensile and creep strengths are high in comparison
to those of other titanium alloys (ref. 2). The objective of this
investigation was to assess, in simulated mission tests at the
NASA Lewis Research Center, the durability of a SiC(SCS-6)/Ti-22Al-23Nb
(at.%) TMC for compressor ring applications, in cooperation with
the Allison Engine Company.
The composite consisted of Ti-22Al-23Nb (at.%) alloy reinforced
by 41 vol % of unidirectional SCS-6 SiC fibers and consolidated
by hot isostatic pressing. A typical specimen cross-section is
shown in the photomicrograph. Specimens having a uniform reduced
midsection with the fibers oriented parallel to the loading axis
were machined and tested on a computer-controlled servohydraulic
fatigue test system heated by a quartz lamp.
Isothermal fatigue load-controlled tests were first performed
at a frequency of 0.33 Hz, a temperature of 538 °C,
and a maximum applied stress (
s
max)
of 1035 MPa. The effects of a more realistic simulated mission
cycle were then assessed. The Allison baseline mission cycle was
designed to simulate aircraft engine operation in a simplified
manner. The mission, illustrated in the left graph, is made up
of a "Type I" major cycle and "Type III" subcycles.
The Type I cycle represents starting the engine, accelerating
and stabilizing at maximum engine power at the beginning of an
aircraft mission, and later shutting down the engine at the end
of an aircraft mission. This cycle is simulated in the mechanical
test specimen by an excursion from minimum temperature and zero
stress through smax
and maximum temperature (
T
max), with a cyclic
stress ratio (
R
s)
of zero. Type III subcycles represent going from engine idle to
maximum power, stabilizing at maximum power, then returning to
idle at different times during a mission. This cycle is simulated
in the mechanical test specimen by an excursion from intermediate
stress and temperature through
T
max and smax,
with
R
s = 0.5. One
total mission cycle is composed of one Type I and six Type III
subcycles. Baseline conditions of smax
= 1035 MPa and
T
max = 538 °C
were chosen for detailed evaluations. The right graph shows a
typical stabilized stress-strain hysteresis loop with segment
descriptions.
The average mission life was 1235 cycles, significantly lower
than the average isothermal life of 8149 cycles in duplicate tests
with the same maximum temperature and stress. The mission test
specimens had fatigue cracks initiating from damaged fibers along
the machined specimen edges, as in isothermal specimens. However,
the percentage of fatigue-cracked area in mission tests was significantly
lower than in isothermal tests. This appeared to be associated
with a process of enhanced cyclic stress relaxation of the matrix
during the mission. The process encouraged load transfer from
the matrix to the fibers, which suppressed fatigue cracking and
induced fiber overload. In future work, mission tests will be
performed on orthorhombic TMC's that contain fibers with greater
inherent strength.
Previous articleLast updated April 30, 1997
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