A two-layer thermal barrier coating system with a bond coating of nickel-chromium-aluminum-yttrium and a ceramic coating of yttria-stabilized zirconia was tested for corrosion protection, thermal protection and durability. Full-scale gas-turbine engine tests demonstrated that this coating eliminated burning, melting, and warping of uncoated parts. During cyclic corrosion resistance tests made in marine diesel fuel products of combustion in a burner rig, the ceramic cracked on some specimens. Metallographic examination showed no base metal deterioration

Liebert, C. H.

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https://ntrs.nasa.gov/search.jsp?R=19790011947 2020-03-21T23:06:56+00:00Z
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e	 NASA Technical Memorandum 79115
(NASA —T!!-79116) 	 TESTS OF NASA CEFA4IC	 N79-20118
;HESMAL BARRTER COATING FOR CAS—TURSINE
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TESTS OF NASA CERAMIC
THERMAL BARRIER COATING
FOR GAS-TURBINE ENGINES
Curt H_ Liebert
Lewis Research Center
Cleveland, Ohio
`	 TECHNICAL PAPER to be presented at the
International Conference on Metallurgical Coatings
•	 San Diego, California, April 23-27, 1979
TESTS OF NASA CERAMIC THERMAL BARRIER COATING
FOR GAS-TURBINE ENGINES
by Curt H. Liebert
National Aeronautics and Space Administration
Lewis Research Center
Cleveland. Ohio
1. INTRODUCTION
A NASA ceramic thermal barrier coating (TBC) system wrs tested by in
dustrial and governmental organizations for a variety of aeronautical, marine,
and ground-based gas-turbine engine applications. This TBC is a two- layer
system with a bond coating of nickel-chromium,-aluminum-yttrium (Ni- 16Cr-6A1-
O.6Y, in wt. %) and a ceramic coating of yttria stabilized zirconia (ZrO2-12Y2O3,
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	in wt. %). Tests (ref. 1) have been conducted to determine corrosion resistance.
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thermal protection, durability, thermal conductivity, and fatigue characteristics.
The information presented herein covers some of the significant test results ob-
tained on the first three items. The information also includes photographs of coated
parts after tests, measurements of coating loss, amount of metal wall temperature
reduction when the TBC is used, and extent of base metal corrosion.
2. RESULTS AND DISCUSSION
2.1 Solid, High Temperature Alloy Pins
the corrosion protection and durability of the TBC and its capability for pro-
)n of base metal parts from corrosion in marine diesel fuel products of com-
on was tested with four coated, solid pin-type specimens made from Rene 80
1) . The fuel contained 1 weight percent of sulphur and 10 parts per million
a salt. The uncoated diameter of these pins was 0.318 centimeter. The tests
a burner rig simulated coating operation in mAriae gas-turbine engines.
4
r•	 2
One specimen withstood 1000 hours (40 cycles) of testing. A cycle consisted
of 23 hours at a gas (and metal) temperature of 978 K with a 2 hour cool-down
period to 300 K. Nominal, polished zirconia thickness before testing was
0.038 centimeter (bond coating thickness was about 0.01 cm) and metallo -
graphic examination after testing showed that about 0.008 centimeter of yttria
stabilized zirconia was lost during the test. The TBC on the second specimen
cracked and spalled after 322 hours or 14 of these cycles. The other two speci-
mens were heated with this rig from about 300 K to a higher temperature of
1170 K. and then cooled down to 300 K. One of these specimens ran for 450 hours
before cracks appeared in the ceramic: the other specimen was tested for 800
hours before cracking and spallation occurred. Metallographic examination
showed no base metal deterioration nn any of the four specimens. The results
of these corrosion protection and durability tests indicated to the government lab-
oratory That the TBC is as good as some other coating systems also tested at these
cyclic conditions.
This polished coating was also tested in this rig at more severe conditions:
in this case a cycle consisted of 48 hours at 978 K, 24 hours at 1061 K. 24 hours
at 1 170 K and cooling in still air to 300 K. The duration of a full cycle was about
100 hours. These burner rig tests simulated the same marine gas-turbine en-
vironment described above, but the uncoated pin diameter was 0.953 centimeter,
or three times larger. The TBC on the larger diameter pin did not crack after
10 cycles of operation for a test duration of 1000 hours. These tests suggest that
the TBC becomes more durable as pin diameter (or analgously, , as turbine vane
and blade leading edge diameter) is increased. Reasons for this better durability
(no cracking) are probably due to the reduced stress in the TBC on larger diam-
i
VJ^
3
eter pins and to the greater ease of plasma-spray application which can give
more consistent TBC deposition on larger diameters.
2.2 Turbine Vanes-Military Aircraft Engine
	
An uncoated vane located in a hot gas region (about 1700 K) melted and burned 	 .I
during accelerated cyclic endurance tests in a military aircraft engine. This
occurred during the first 60 hours of bimodal cyclic operation which included 240
cycles between an average gas temperature and pressure of 1550 K and 2.2 N1Pa ,
respectively, to shutdown, and 1920 cycles at this gas condition to idle. This
melted vane was replaced with a coated vane of identical design and subjected to
the same test conditions. After 60 hours of these cyclic tests, there was no
melting, burning or other deterioration of the base metal. But. loss of the TBC
was observed at the trailing-edge pressure surface aft of the cooling-air dump
holes and in small areas of the leading edge. Because the metal wall was very thin
at the trailing-edge region, it could only be lightly grit blasted (about half of the
normal blasting pressure) prior to application of the TBC. This could be partly
responsible for the premature coating deterioration in this region. This vane was
sectioned and areas were found primarily at the leading edge, where the bond coat-
ing thickness was 0.005 centimeter or less. For best ceramic coating adherence,
the bond layer should have been applied to thicknesses of 0.010 centimeter (ref. 2) .
The ceramic coating was applied at a nominal thickness, of 0.020 centimeter.
A second vane of identical design (fig. 2) was coated, vacuum heat treated for
4 hours at 1 .3x10 3 Pa and 1450 K (ref. 3) , and then tested at the same engine con-
ditions. Visual inspection of this vane after 100 hours of testing showed no de-
terioration of the TBC. This test time included 400 cycles and 3200 cycles of the
two types just described. Testing continued with this vane for a total of 300 hours;
	
this included 1200 cycles and 9600 cycles at each of the two types of cycles. Visual
	 4
4inspection after 300 hours showed minute coating loss at the leading- and trailing-
edge region, but again no deterioration of the cooled metal wall was noted. This
better durability is attributed to the vacuum heat treatment which can sinter and
densify the bond coating to improve its oxidation resistance and durability (ref. 3).
Iron oxide particles were deposited on the TBC during the test. Iron oxide
(Fe 2O 3 ) can react with yttria stabilized zirconia in the TBC (when monoclinic
ZrO 2
 is present) over times of 200 to 800 hours and temperatures of 1473 to 1673 K
(ref. 4) . Since the coating operated in this temperature range and is presumably
partially monoclinic, such a reaction may ba partly responsible for the minute
loss of the TBC after 300 hours of cyclic operation.
2.3 Combustor Dome and Scroll and Turbine Nozzle Shroud
Tests were conducted on the TBC in an air-cooled combustor dome and
scroll, and turbine nozzle shroud of a military ground-vehicle engine at gas
temperatures and pressures of 1467 K and 1.5 MPa . The uncoated parts de-
teriorated early at these gas conditions, but after the TBC was applied the de-
terioration stopped. Inspection of the coated parts after 500 hours of testing
showed that the metal parts remained coated and showed no visible deterioration.
Inspection of the engine, however, showed that some erosion of the TBC
occurred because mini-'e deposits of yttria stabilized zirconia meterial were
found lying the ducting recesses. Bond and ceramic coating thickness estimates
are 0.013 to 0.025 centimeter and 0.038 to 0.127 centimeter. respectively.
2.4 Turbine Plenum - Vehicular Engine
The hot-spot area of the turbine plenum was plasma-sprayed with the TBC.
F igure 3 shows the good condition of this patch of TBC after 50 hours of accel-
erated endurance testing in a burner rig. These test.- consisted of 120 cycles at
5 cycle variations (ref. 1). The highest gas temperature for these cycles was	 I
r1
r
5
1310 K and the lowest was 530 K. The cycle consisted of 5 minutes of heating
!	 and 20 minutes of cooling. Some uncoated ;,,etal areas were badly warped after
the test; however, no warping or coating deterioration occurred where the patch
of TBC was applied on the plenum metal wall. Because these cyclic burner tests
were so promising, two other plenums with fully coated internal surfaces are
undergoing tests in full-scale vehicular engines. One coated plenum has satis-
factorily accumulated 220 hours with no visible warping or coating deterioration;
the TBC has lowered hot-spot metal wall temperatures by 30 to 100 K (from a
temperature level of 1300 K) . Testing of the TBC on these plenums is continuing.
The bond and ceramic coating tLickness were about 0.015 and 0.050 centimeters,
respectively.
3. CONCLUDING REMARKS
The TBC. lowered metal temperatures, improved corrosion resistance of
metal parts, increased part life, and eliminated burning, melting and warping
that occurred with some of the uncoated metal parts tested in this cooperative
effort. Some loss and cracking of the ceramic layer occurred in regions where
uncoated parts and specimens had previously melted or corroded. However,
metallographic examination showed no base metal deterioration of coated
parts. This demonstrated that coating loss will not necessarily result in
damage to metal parts. Recent tests at the NASA Lewis Research Center (ref. 5)
indicate that better coating system durability was obtained on a burner rig and
furnace specimens wh;,n the percent by weight of yttria was reduced to 6.2
or 7.9 in the zirconia ceramic coating and when the percent by weight of yttrium
was reduced to 0.15 or 0.35 in the bond coating. These ceramic and bond coat-
ings were not used herein.
PP
•	 6
4. REFERENCES
1. C. H. Liebert and F. S. Stepka , NASA Tech. PaQer 1425, 1979.
 
2. C. H. Liebert , S. Stecura , and J. E. Brown, Ind. Res. , (1976) 22.
3. R. C. Tucker, T. A. Taylor, and M. H. Weatherly, Plasma Deposited
MCrAIY Airfoil and Zirconia/MCrAIY Thermal Barrier Coatings, Union
Carbide Corporation, Indianapolis, 1976.
4. 1. Zaplatynsky, , NASA Tech. Memo 78942, 1978. 
5. S. Stecura , NASA Tech. Memo. 78976, 1978. 
4
'	 ^	 7
ORIMNAL FAGS 2
OF f4= QUALM
M.
Figure 1.- Coated sulid specimen for
corrosion protection and durability
tests in marine diesel fuel products
of combustion.
Figure 2. - Coated turbine vane Per 300 hours (10,800 cy-
clPs) in a military aircraft gas-turbine engine.i
•
.	 I
4
a	 ORIG NAL PAGE IS
OF POOR QUALITY
Flgure 3.- TBC coating patch on turbine plenum-vehicular engine,


Tests of NASA ceramic thermal barrier coating for gas-turbine engines

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