Vane subelements were fabricated from a silicon carbide fiber-reinforced silicon carbide matrix (SiC/SiC) composite. A cross-sectional slice of an aircraft engine metal vane was the basis of the vane subelement geometry. To fabricate the small radius of the vane's trailing edge using stiff Sylramic SiC fibers, a unique SiC fiber architecture was developed. A test configuration for the vanes in a high pressure gas turbine environment was designed and fabricated. Testing was conducted using a pressure of 6 atm and combustion flow rate of 0.5 kg/sec, and consisted of fifty hours of steady state operation followed by 102 2-minute thermal cycles. A surface temperature of 1320 C was obtained for the EBC-coated SiC/SiC vane subelement. This paper will briefly discuss the vane fabrication, test configuration, and results of the vane testing. The emphasis of the paper is on characterization of the post-test condition of the vanes

Calomino, Anthony

Robinson, R. Craig

Thomas, David J.

Verrilli, Michael

NASA Technical Reports Server

CHARACTERIZATION OF CERAMIC MATRIX COMPOSITE VANE SUBELEMENTS 
SUBJECTED TO RIG TESTING IN A GAS TURBINE ENVIRONMENT 
Michael Verrilli and Anthony Calomino 
NASA Glenn Research Center 
2 1000 Brookpark Road 
Cleveland, OH 44135 
R. Craig Robinson 
QSS Group, Inc., NASA Glenn Research Center 
2 1000 Brookpark Road 
Cleveland, OH 44135 
David J. Thomas 
Ohio Aerospace Institute 
22800 Cedar Point Road 
Brook Park, OH 44 142 
ABSTRACT J 
Vane subelements were fabricated from a silicon carbide fiber-reinforced silicon carbide matrix 
(SiC/SiC) composite. A cross-sectional slice of an aircraft engine metal vane was the basis of the vane 
subelement geometry. To fabricate the small radius of the vane's trailing edge using stiff Sylramic Sic 
fibers, a unique Sic fiber architecture was developed. 
A test configuration for the vanes in a high pressure gas turbine environment was designed and 
fabricated. Testing was conducted using a pressure of 6 atm and comb $ion flow rate of 0.5 kg/sec, and 
consisted of fifty hours of steady state operation followed by 102 -minute thermal cycles. A surface 
briefly discuss the vane fabrication, test configuration, and results of the vane testing. The emphasis of 
the paper is on characterization of the post-test condition of the vanes. 
temperature of 1320°C was obtained for the EBC-coated SiC/Si ,r" vane subelement. This paper will 
INTRODUCTION 
Ceram'c matrix co p ites CMCs re being evaluated for hot sections of advanced gas turbine 
engines &&tg E$vh$r&ance, such as lower emissions and higher cycle efficiency, 
relative to today's engineiwith superalloy hot section components. The capability of one ceramic matrix 
composite for use as combustor liners, silicon carbide fiber-reinforced silicon carbide matrix composite 
(SiC/SiC), has been demonstrate$ through rig',2 and engine tests3. 
In order to gain full eysfem performance benefits from these advanced materials in combustor liner 
applications, a high temperature first stage vane assembly must be successfully introduced to the engine. 
A recent study demonstrated fabrication of an exhaust guide vane from a SiC/SiC composite having 3D 
fiber architecture4. The resulting vane had a fiber architecture which accounted for the inherent low 
through-thickness strength of CMCs. However, it did not have the aerodynamic shape (including the 
sharp trailing edge) required to maintain engine performance. 
The overall objective of this research was to demonstrate the scale up of the SiC/SiC material system 
from the panels and simple component architectures previously p r~duced~ ,~  to the more complex 
architecture of a turbine vane. This activity included the development of a unique Sic fiber architecture 
for the vane's sharp trailing edge, fabrication of two SiC/SiC prototypical vane designs, development of 
an advanced environmental barrier coating (EBC) for airfoils, and rig testing in a turbine environment. 
This paper will summarize the vane fabrication, test configuration, and results of the vane testing. These 
results are discussed in detail el~ewhere~,~. The emphasis of the paper is on characterization of the post- 
test condition of the vanes. 
L, "0, 
b I n  
VANE GEOMETRY AND MATERIAL 
Turbine vanes possess a rather complex geometry to optimize engine efficiency. In today's engines, 
a vane design typically incorporates trailing edges with radii less than 0.50 mm, and a cross section that is 
often tapered in the radial direction, and cantered in the circumferential direction of the engine. Our 
This is a preprint or reprint of a paper intended for presentation at a 
conference. Because changes may be made before formal 
publication, this is made available with the understanding that it will 
not be cited or reproduced without the permission of the author. 
https://ntrs.nasa.gov/search.jsp?R=20050198903 2019-08-29T20:30:25+00:00Z
Figure 1- Geometry of vane 
sube€ement. Dimensions are in mm. 
approach to fabricate the turbine vane was to first demonstrate that a 
prototypical shape could be manufactured utilizing the material 
processing established for multiple ply lay-ups of two-dimensional 
cloth. In order to have a vane geometry considered relevant to 
actual engine hardware, a cross-sectional slice of an aircraft engine 
metal vane served as the basis for the vane dimensions used in this 
fabrication demonstration. This representative slice was then 
extruded to produce the vane subelement geometry with a constant 
cross-section as shown in Figure 1. The SiC/SiC material used in 
this activity was devel d under NASA's Enabling Propulsion 
Materials and Ultra E f i  Engine Technologies programssd. The 
vane subelement consisted of BN-coated Sylramicm Sic fiber cloth 
reinforced with a CWslurry-castlmelt infiltrated SiC/Si matrix. 
VANE FABRICATION 
a web between the suction and the 
have six plies in the outside walls and the same outside dimensions. The web consists of four plies. 
Two configurations of the vane subelement were manufactured. One had internal cavities formed by 
sides of the vane and the other had no web. Both vane types 
5HS Split Plane 
b 
d e 
Vane trailing e d g y  - 
Joined plain weave cloth v 
Figure 2: Y-cloth fiber architecture used to 
form vane trailing edge. 
It is difficult to fabricate components with small radii, as 
exists in the trailing edge of this vane subelement, using the 
stiff stoichiometric Sic fibers, such as Sylramic, that are 
cumntly preferred for high temperature-capable SiC/SiC 
composites. The challenges posed by the aggressive 
geometric and structural requirements associated with the 
vane configuration were addressed using a new Y-cloth fiber 
architecture shown in Figure 2. The Y-cloth provided a fiber 
architecture in the remaining regions of the vane that had been 
well characterized and suc ly demonstrated in 
combustor As igure 2, the Y-cloth 
was used as the solution to forming a sharp 
splitting the cloth into two planes during the 
. The fiber tows forming the trailing edge 
erlocked thereby enhancing the through 
six plies of Y- 
s CM so that the 
thickness strength capability of the composite material. 
cloth, the length of each Y-cl 
For vanes constructed fiom 
two legs would wrap around the perimeter of the vane &eform ID graphite tool with a 10 mm overlap. 
The additional length was used to join the two seetions with an overlapping h g e  splice. To make the 
external sixth ply, a standard woven cloth was cut to the required fmal length and a fringe splice joined 
the two sections of the cloth at the trailing edge. The cloth was then . The entire assembly was 
then placed into an aluminum compaction tool designed to form the of the vane. After the 
prepreg material was allowed to dry, the preform was removed from the aluminum tooliig and placed 
into an external graphite tool be€ore being shipped to a vendor for matrix infiltration. 
To make the preform for the vane with the internal cavities and connecting web wall, a slightly 
different initial approach was used. Each section forming the internal cavities (and ultimately the web) 
was created by first dipping two concentric layers of a two-dimensional, 2-by-2, M5" braided tube 
around net-shape graphite mandrels. Four 
additional plies were wrapped around the assembled graphite mandrels in the same fashion used for the 
six ply shell vane configuration. 
Two venders were selected to manufkcture both vane . Both General Electric Power Systems 
Composites (GEPSC) and Goodrich used their own c ite processing techniques to infiltrate 
Both mandrels were prepregged, and allowed to dry. 
Sylramic vane preforms to make completed SiC/SiC vanes. Examples of both completed vane types are 
shown in Figure 3.  
Figure 3: Completed SiC/SiC vanes using a 
layup of Sylramic Y-cloth, a) six ply shell 
configuration, b) webbed eonfiguration. 
mr 
Figure 4 -Test section configuration used for 
testing SiC/SiC vanes. 
To protect the SiCISiC substrate fiom recession in the 
environment, an env barrier coating 
applied to the vanes. coating. system 
a silicon bond coat, a mullite intermediate 
coat, and a proprietary rare earth silicate topcoat. Cooling 
holes were machined in the trailing edge via laser drilling 
prior to application of the EBC. 
VANE TESTING 
The test conf3guration was comprised of three parallel 
vanes contained in a led pressure section (Figure 
4). The center v EBC-coated SiC/SiC test 
with superalloy vanes on either side to help 
flow around the SiC/SiC vane. The 
hadthe same external geometry and wereh 
using superalloy platforins as shown in Figure 5. Vane 
cooling air flowed through these superalloy plates into 
ends of the vanes and out through holes in the trailing 
for all three components. The test section design 
led gas temperature adjacent to the 
vanes and downstxeam couples. Optical 
used to measure external vane material 
igure 4). Two six ply shell vanes 
tested. 
tested using NASA‘s High 
. The rig burns jet fuel and 
enables material 
to pressures of 4 to 15 atm and 
dsec ,  under either lean-burn or rich-burn combustion 
mixtures’. Flow &om th or was directed &om the 
150 mm inner diameter 
cool the internal cavities of all vanes in the test section. Gas 
(and ultimately material temperature) was 
ough variation of the combustor’s fueVair ratio 
under lean m i x t u ~ s .  
Prior to testing co 
element analyses were 
ss conditions present in the vane during 
results are described in ref. 8. To 
predicted maximum temperature (1190 “C) was within the material eapabilities (-1300 “C). 
in-plane tensile stresses were low (maximum axial stress of 27 MPa and maximum transverse or “hoop” 
stress of 105 MPa), but the interlaminar tensile (ET) stresses predicted were found to be high enough to 
be of concern. ILT values in the leading edge (14 MPa) approach the flat panel measured strength value 
of 17 MPa for this material. In the vane's the trailing edge, the ILT stress of 54 MPa exceeded the 
measured strength. However, this 54 MPa ILT stress is artificially inflated in this analysis8. Since the 
vane's Y-cloth architecture provides through thickness fiber reinforcement at the trailing edge, these 
stresses were considered to be less of a concern than those at the leading edge. 
RESULTS 
Test Results 
After completion of 20 hours of rig testing using gas temperatures of about 1200 "C, the SiC/SiC 
vane subelement was inspected and found to have no visual degradation, whereas the superalloy vanes 
and other superalloy rig hardware had visually identifiable cracks. One superalloy vane was removed to 
enable measurement of the SiC/SiC vane temperature using optical pyrometry. Operation continued for 8 
hours under the same rig conditions and a peak EBC surface temperature of about 1 170 "C was measured. 
The third superalloy vane was reintroduced and 50 hours of steady state operation were completed. 
The rig was disassembled to inspect the test section after 
completion of 50 hours of testing. Water was found in the 
test section. One of the water-cooled thermocouples was 
found to have leaked prior to the beginning of the last 3.5 
hours of testing. This lead to the accumulation of water in 
the test section between rig runs. During this period of time, 
the SiC/SiC vane absorbed some of this water. Upon 
restarting the test, the absorbed water vaporized, producing 
spallation damage (Figure 6). About 26% of the total coating 
surface area had spalled. 
Because of the damage to the first test vane (number 
909), a second EBC-mated vane subelement (number 910) 
~ 
consisted of holding the gas temperature at a constant 
Figure 6 - SiC/SiC vane (number 909) after minimum value for 45 seconds, then heating to a maximum 
~mlvertmt Water elrposm fO&med by rig over 15 seconds, holding for 45 seconds, and finally cooling 
testing for 3.5 hours. back to the minimum temperature in 15 seconds. Initial peak 
temperatures were based on ence obtained during the 
steady state testing. However, higher gas temperatures of up to 1420 "C were used for latter cycles. 
Under the higher gas s, a peak material temperature of 1320 "C was obtained. After 102 
cvcle~ were completed 
Inspection of the test section after 102 cycles revealed 
considerable damage of the test section metal hardware. 
Meking of the superalloy vanes occurred, as well as other test 
section metal hardware. The SiC/SiC vane subelement was 
intact but had deposits on the EBC 
diffraction revealed evidence of the Co- 
vane on the suction side of the SiC/SiC vanes, (which can be 
seen in Fig. 7) and deposits of Ni-based Waspaloy on the 
pressure side. Note that both these alloys melt above 1400 
"C and were used to fabricate test section hardware. 
Examination of Vanes 
m. - *Trailing Edge 
I ^ .^.^.^ 
NDE inspection via computed tomography (CT) was 
Figure 7 - SiC/SiC (number 910) and 
spmuoy vane subelements after the 
testing. 
used to inspect both SiC/SiC vanes after completion 
testing. Images of vane 909 both in the as-fabricated s 
0 hours of steady state rig operation are compared 
1 
I 
in Figure 8. Spallation damage can be seen in the post-test image, particularly on the pressure (is. 
images of vane 909, a) as-fabricated, b) after rig testing for 50 hours 
Figure 9 - Computed tomography. of vane 910, a) as-fabricated, b) after rig testing for 102 2-minute cycles. 
Both SiC/SiC vanes were sliced at the midline to obtain sections for mierostmctural examination. A 
coating was removed fiom the vane tested under steady state conditions (vane 
recoat it for further testing. However, since and NDE inspection 
damage, a second vane (number 910) was 
sliee from each vane was mounted in florescent epoxy and polished. 
microstructure of the first vane was not 
s mounted in the as-tested condition. 
nominal proffie are superimposed over the mi 
The cross sections of both vanes are shown in Figure 10. Engineering draw 
after exposure to severe test conditions. C 
matrix is evid 
holes can be seen near the trailing 
on the inside surface. 
Figure 10 - Cross sections of SiC/SiC vanes after rig testing, a) vane 910 tested under eyclic conditions, b) vane 909 
tested under steady state conditions. 
Since the finite element analyses predicted the highest stresses to exist in the leading and trailing 
edges, it is important to examine the microstructure in these regions. In Figure 1 I, details of the leading 
edges are shown. A coating of superalloy can be seen on vane 910 on both the pressure and suction sides, 
but not at leading edge. Spallation of outer composite plies exists in vane 909 at locations away from the 
leading edge. However, no cracks exist. In both vanes, the gap on the ID between the actual vane profile 
and the drawing near the leading edge are the result of the manufacturing process. 
Details of the trailing edge 
(Figures 12 
similar to those seen in the 
leading edge. Superalloy 
deposits of about 0.4 mm 
fhickness extend toward 
trailing edge radius in van 
aled fe 
Figure 11 - Microstructure ofthe leading edge of S 
after rig testing, a) vane 910 tested under cyclic c 
under steady sfate conditions. 
most damaged reaons. Figure 13@) shows 
with the original six plies. It is likely 
in higher stresses in the trailing edge of van 
in Figs. 12 and 13. In spite of this damage, the trailing edge of vane 909 remained 
Figure 12 -Microstructure ofthe trailing edge of SiC/SiC vane sub 
910 tested under cyclic conditions, b) vane 909 tested under 
L 
13- Mierostmctur 
under eyclic c 
side of the vanes, 20 mm away ftom the tip of the tr 
e 909 tested under steady state conditions. 
e, a) vane 
SUMMARY AND CONCLUSIONS 
Vane subelements were fabricated fiorn a SiC/SiC composite. The Y-cloth fiber architecture was 
developed at NASA Glenn Research Center to provide a straightforward means to form the sharp trailing 
edge using stoichiometric Sic (Sylramic) fibers. This approach also utilizes mature composite processing 
methods used for standard ply lay-ups of two-dimensional cloth, such as panels. Vane subelements with 
and without an internal web connecting the pressure and suction walls were fabricated. 
Two vanes without internal webs were tested in a simulated turbine environment. Testing consisted 
of fifty hours of steady state operation and 102 2-minute thermal cycles. A surface temperature of 
1320°C was obtained for the EBC-coated SiC/SiC vane subelement. The vanes were inadvertently 
ly severe test conditions (such as exposure to standing water prior to combustion 
exposure and deposition of molten superalloy parts) which lead to damage in the vanes. NDE and 
microstructural examinations revealed the extent of spallation in one vane due to vaporization of absorbed 
water, and deposition of molten metal on both the pressure and suction sides of the second vane. Both 
vanes retained their shape even after exposure to these severe test conditions and they maintained full 
functionality. These encouraging results show the potential of SiC/SiC composites for turbine airfoil 
applications. 
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Characterization of Ceramic Matrix Composite Vane Subelements Subjected to Rig Testing in a Gas Turbine Environment

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