Hot corrosion life prediction methodology based on a combination of laboratory test data and field service turbine components, which show evidence of hot corrosion, were examined. Components were evaluated by optical metallography, scanning electron microscopy (SEM), and electron micropulse (EMP) examination

Fryxell, R. E.

Gupta, B. K.

English

NASA Technical Reports Server

N87-11193
EFFECTS OF SURFACE CHEMISTRY ON HOT CORROSION LIFE*
R. E. Fryxell
General Electric Company
and
B. K. Gupta
TRW, Inc.
This program has as its primary objective the development of hot corrosion
life prediction methodology based on a combination of laboratory test data and
evaluation of field service turbine components which show evidence of hot
corrosion. This program is divided into five tasks. Burner rig testing is
being performed by TRW. The discussion will describe the overall program with
a summary of results obtained in the first two tasks which have been completed.
Task I involves a comprehensive evaluation of six hot corroded turbine
components, having known operating history, to establish the degradation
mechanisms in the corroded areas. The six components selected are identified
in Table I. Four of these are the same component from engines of the same
model to afford maximum opportunity to make cause and effect correlations.
The high pressure turbine (HPT) components were obtained only after examining
blades from a large number of disassembled engines; hot corrosion is only
infrequently substantial in the HPT of engine models available for this
study. On the other hand, hot corrosion is more frequently encountered in the
low pressure turbine (LPT) since at the lower temperatures involved, corrosive
contaminants are more likely to accumulate in significant amounts and/or be
present a greater percentage of operating time. One LPT vane is included in
this study for comparison•
Evaluation of these components included optical metallography, scanning
electron microscopy (SEM), and electron microprobe (EMP) examination. Several
transverse sections of each airfoil were prepared using nonaqueous cutting and
polishing techniques.
;-._EC_r.DIt_ PAGE BLANK NOT
* Contract NAS3-23926
123
https://ntrs.nasa.gov/search.jsp?R=19870001760 2020-03-23T15:02:37+00:00Z
Results of the evaluation of the four CF6-50 Stage 1 blades are summarized in
Table If. Locations at which maximum degradation was observed in individual
sections are defined by % span (radlal)/% chord (distance from leading edge
(LE) to trailing edge). These observations are limited to the leading edge
and concave/pressure surface; the convex surface of such blades rarely suffers
significant environmental degradation. The listed temperatures refer to the
blade surface at the location of maximum degradation; these must be regarded
as approximate for two reasons: (i) the observed distress occurs over some
distance along the surface, and (2) assigned temperatures are calculated and
will, in general, change somewhat as the turbine ages.
Typical sulfidation was noted at some locations in all of these blades,
including development of a depletion layer (loss of aluminum) in the coating
or substrate alloy. Typical microstructures are shown in Figures i and 2.
However, sulfides were found in the most distressed areas in only two of these
blades. In the other two blades, although sulfides were found, the most
severely distressed regions did not contain any detectable sulfides.
This points up a primary difficulty in interpretation, namely that contaminant
ingestion occurs only when the aircraft is near ground level; during most of a
mission, simple oxidation is probably to be expected. Table II exhibits one
attempt to relate the observed degradation to operating history. Using flight
pattern information and the criterion given in footnote (a), total times were
estimated during which each engine might be considered vulnerable to
contaminant ingestion: total hours near coastal airports, and total hours (in
life of engine). One way of displaying this information is shown in Figure
3. There is in fact a trend showing increased hot corrosion with increased
low altitude operating time; the curvature is upwards as is not unreasonable,
i.e., above about 60-80 _m (nominal original coating thickness) the substrate
alloy is exposed. However, with this limited sampling, it is not possible to
make inferences concerning coastal versus total low altitude exposure.
Alternatively, the trends shown in Figure 3 may reflect total times at
takeoff/thrust reverse (higher) temperatures, a parameter which is clearly
related to total hours at low altitudes. This is shown in rate units (Figure
4). The highest rates correspond to the shortest lengths of mission for which
the percentage of time at higher temperatures (or lower altitude) is greater.
124
These are in fact interesting and instructive exercises, but remembering that
the present four blades were selected as rather extreme examples, the actual
amounts of corrosive degradation (Table II and Figures 3 and 4) do not define
possible curves which other examples should be expected to match. Further,
since oxidation and hot corrosion rates are competitive in some instances
(Table If), the cause and effect relationships with respect to environmental
degradation have that additional degree of complexity.
The trend shown in Figure 4 has, however, long been recognized in a
qualitative sense and hopefully continued quantitative evaluations coupled
with compilations of operating histories will lead to a reliable choice
between various possible correlations such as those depicted in Figures 3 and
4.
With respect to the ultimate blending of this type of information with the
burner rig studies to be performed in this program and application to llfe
prediction methodology, it should be pointed out that previous burner rig
corrosion tests at General Electric with very low salt ingestion levels (at
barely the threshold level for salt condensation) produced conventional Type 1
sulfidation but at rates indistinguishable from oxidation. This may relate to
the results in Table II, namely that oxidation and corrosion rates may vary
relative to each other, depending on the circumstances. Possibly as well
there are implications with respect to the design of somewhat more complex
burner rig tests; i.e., temperature cycling, intermittent salt ingestion,
etc.
Evaluation of the other two components listed in Table I revealed extensive
Type 1 sulfidation in the low pressure turbine vane but not in the J79 high
pressure turbine blade. Since these are different components/different engine
models, the observations cannot be factored into an operating history
correlation together with the other four components. Yet the structural
features of the degradation do contribute to baseline service related
information to be compared with hot corrosion burner rig tests which comprise
the remainder of this program.
125
Two of the three burner rig tests are to be conducted under identical
conditions as described in Table III for a maximum time of i000 hours (see
Figure 5). The first of these, Task II, provides a hot corrosion baseline for
the program alloys and coatings in the as-processed condition. Included in
this task are duplicate specimens of UT00 and Rene' 80, both uncoated and with
the following coatings:
pack aluminide RT21 (Chromalloy) and Codep (General Electric)
low pressure plasma NiCoCrAIY (Ni-23Co-18Cr-12AI-0.3Y)
At approximately 20 cycle intervals, specimens are visually examined,
photographed, and coil inductance measurements made in a series mode at 10MHz
with a multifrequency LCR meter as shown in Figures 6 and 7. These
measurements are patterned after studies at NASA-LeRC during the past several
years. Specimens are removed from the test when visual evidence of hot
corrosion is noted in three successive inspections. Additional specimens are
also exposed for lesser time periods. Specimen evaluation is performed as in
Task I.
A second identical hot corrosion burner rig test (Task IV) will be performed
with specimens which have been given a variety of aging treatments at II00C
(Task III) to determine the effect on hot corrosion behavior caused by surface
oxidation and/or interdiffusion between coating and substrate alloy.
Triplicate specimens of the above coated alloys will be aged under the
following conditions: isothermal inert atmosphere (vacuum) for i00 hours,
isothermal air furnace oxidation for i00, 300, 600 hours, one hour air furnace
cycles for i00 hours and cyclic burner rig oxidation for I00, 300 and 600
hours. One specimen of each material representing each of these conditions
will be metallographically evaluated and the remaining duplicates used in the
Task IV hot corrosion burner rig test.
The Task II burner rig test has been completed. Coil inductance measurements
are shown in Figure 8 for six uncoated specimens. The four long term
specimens were removed at the third inspection interval at which hot corrosion
was visually observed. Although Rene' 80 and UT00 survived about the same
126
lengths of time, based on this criterion, the coil inductances were widely
different, and metallographic measurements indicated a factor of two greater
corrosion rate for U700, in accordance with previous experience. Typical
microstructures with subsurface sulfides are shown in Figure 9 for one
specimen of each alloy.
Coil inductance measurements for coated specimens are shown in Figure i0 with
an expanded scale below 200 hours to show clearly the negative changes which
appeared at the first inspection, particularly for the aluminide coated
specimens. The overall changes are in all cases much smaller than for the
uncoated alloys (Figure 8) reflecting the corrosion protection offered by all
of the coatings. The four Codep coated specimens are shown in Figure i0 to
their time of removal and the greater coil inductance changes suggest a
greater corrosion rate; for all of these, coating penetration was confirmed
metallographically. The three RT21 coated specimens shown in Figure i0 ran
the full term i000 hours with only minor visual evidence of hot corrosion.
However, metallographic evaluation showed essential coating penetration in all
three, although clearly less extensive than for the Codep coated specimens.
One specimen of RT21 coated Rene' 80 (not shown in Figure i0) was removed at
697 hours and did exhibit coating penetration with additional corrosion of the
substrate to about i00 _m depth. All of the NiCoCrAIY coated specimens showed
extensive alloy depletion within the coating, occasional pits, but no complete
penetration was observed. Typical microstructures are shown in Figure Ii.
Subsequent to the completion of Task II, all the aging treatments in Task III
have been performed and Task IV is underway. At the conclusion of the Task IV
test, all the data generated in Tasks I through IV will be evaluated and an
empirical hot corrosion life prediction model based on these data will be
proposed. Also, recommendations will be made for other test parameters to be
evaluated, and evaluation methodology necessary to permit prediction of hot
corrosion life.
Task V will be a hot corrosion high velocity cyclic burner rig test designed
to check the validity of the proposed life prediction model. Duplicate
specimens of a maximum of four alloys and five coatings will be tested up to
127
a maximum of i000 hours and evaluated as in Tasks II and IV. The alloys and
coatings will be those from this program plus additional alloy-coating systems
selected by the NASA Project Manager.
The Task V experiment will be designed to test two aspects of the life
prediction model developed in Task IV. The ability of the model to predict
hot corrosion life under different rig test conditions than those previously
used in Task II and IV will be tested in one portion of the experiment. The
second portion of the experiment will test the ability of the model to predict
hot corrosion life of new alloys and coatings.
The ability of the model to predict hot corrosion life under different rig
test conditions will be evaluated using alloys and coatings previously tested
in Task II and Task IV. The hot corrosion tests for this portion of the
program will use one or more variations of the test parameters used in Task
II, IIl and IV. Test parameter variations that will be considered include the
following:
A different salt level
A different test temperature
Intermittent salt injection
A different sulfur level in the fuel
Additional or modified aging cycles
The specific testing parameters will be derived from the results of Task II
and Task IV, previous experience in evaluating field service hardware and
previous experience in conducting hot corrosion testing under a variety of
test conditions.
The ability of the model to predict hot corrosion resistance of new coatings
and alloys will be evaluated using alloys and coatings that have not been
evaluated previously in Task II and Task IV. The aging cycle will be chosen
from the results of Task IV.
128
Life prediction for both portions of the Task V experiment will be based on
the model developed in Task IV as well as extensive experience gained from
evaluating field service components and conducting hot corrosion tests.
Knowledge of the effect of variations in both engine operating conditions and
rig testing conditions will be used in predicting hot corrosion life under
different rig test conditions in the first portion of the experiment with the
same alloys and coatings tested in Task II and Task IV. With respect to
testing of additional alloys and coatings, use will be made of extensive
empirical correlations of composition and hot corrosion performance in
predicting hot corrosion llfe.
"; 129
Table I. Components Selected for Evaluation.
Number
Carrier Base Aircraft Component Total Hours
Southern Asia A3OOB (I) 2,694 1560
Northern Africa A3OOB (I) 4,641 2169
Southern A3OOB (I) 1,830 1471
Africa
Western Europe DCIO-30 (I) 2,625 610
South DC10-30 (2) 15,155 2987
America
U.S. Mllltary F4 (3) 2,700
(I) CF6-50 Stage I high pressure turbine blade, Codep coated Rene t 80.
(2) CF6-50 Stage I low pressure turblne vane, uncoated Rene' 77.
(3) J79 Stage 1 high pressure turbine blade, Codep coated Rene' 80.
% Takeoffs
at Coastal
Airports
67
77
50
46
72
Table II. Summary of Metallographic Evaluations and Operating History for
Four Stage I High Pressure Turbine Blades from CF6-50 Engines.
Total No. of Hours/
Carrier Base Hours Cycles Cycle
Southern Asia 2694 1560 1.72
Northern Africa 4641 2169 2.14
Southern Africa 1830 1471 1.24
Western Europe 2625 610 4.30
Hours <2000' El. (a)
Near
Coastal
AP Total
350 520
560 720
240 490
95 200
Location (b)
40/70-90
60/LE
60/80
50/60-90
80/50-80
IO/LE
10/60-80
60/90
10/60-80
20/60-80
Temp. " C
(Est.)
Max. Avg.
1090 820
980 740
1090 820
1090 820
I090 820
910 680
1020 760
I090 820
1020 760
I020 760
Maximum Depth of
Attsckj tan(c)
Total _r I000 Hours
100 37
50 20
200 75
230 50
230 50
50 27
75 40
180 I00
45 17
50 20
Type I
Sulfides
Yes
Yes
No
Yes
Yes
Yea
Yes
No
No
Yes
(a)_atimated, using 20 minutes as approximate total time per cycle; gate to takeoff, climb to 2000 ft.
(b)% Span/% Chord.
(C)Includlng depletion zone.
Table III Burner Rig Operating Conditions.
Specimen Temperature - 900" C _ 9" C
Test Cycle - I hour aC temperature followed by 6 minutes of forced
air cooling
Sodium Concentration - 0.5 ppm sodium (_i0%) in the combustion gases
introduced as aqueous NaCl
Combustion Air Preheat Temperature - 232" C _ I0" C
Specimens - Eight positioned equally on a 4.2 cm (1.64 inch) di_ecer
circle of a holder rotating at 600 rpm
Burner Nozzle Throat Diimeter - 2.54 cm (1.0 inch)
Burner Pressure - 1.0 psig
Nozzle Throat to Nearest Specimen - 4.45 cm (1.75 inch)
130
ORIGINAL PAGE IS
OF POOR
Figure i. Photomicrograph on the concave airfoil of a
Stage i high pressure turbine blade from a Northern Africa
CF6-50 engine. Sulfides are present in the Rene' 80 de-
pletion zone. Etched, 500X.
Figure 2. Photomicrograph at the leading edge on an airfoil of a
Stage i high pressure turbine blade from a Southern Africa CF6-50
engine. Sulfides are present within residual coating (arrows).
As-polished, 500X.
131
24C
20C
160
o_
120
_ 80
4O
0 o
Near Costa
I I I I I I I
100 200 300 400 500 600 700
Hours at < 2000 Feet Elevation (Estimated)
Figure 3, lloc Corrosion Degradation for CF6-50 Stage I Hlgh Pressure Turbine Blades
as Function of Time at Less Than 2000 Feet Elevation.
I
800
lOO 0
8o
6o
40
0
0
2O
Figure 4.
o
0 Oxidation
n Sulfidation
o
o
o
0
| I I I I
I 2 3 4 5
Hours/Cycle
Maximum Degradation Rates for CF6-50 Stage 1 High Pressure
Turbine Blades as Function of Average Mission Duration.
132
ORIGINAL PAGE IS 
OF POOR QUALin 
Figure  5. Burner R i g s .  
1. Burner 
2. Specimens on holder  r o t a t i n g  a t  600 rpm. 
3 .  Opt ica l  pyrometer 
4 .  Preheated combustion a i r  
5. S a l t  s o l u t i o n  spray nozz le  
133 
Figure  6 .  Inductance r i g / c o i l  w i t h  specimen core .  
r 
2-1/4 i n  
1 / 2  c i n .  
'I 
/ 8  i n .  OD x 1/32 I n .  Wall x 1-1/2 i n .  Long 
P l a s t i c  Form 
N o .  18 Gauge Enameled Copper 
Wire Wrapped Around P l a s t i c  
Form and Epoxyed t o  it  
S o l d e r  t o  Brass  Support 
P l a s t i c  Spacer 
5 /8  in .  0 . 0 3 2  i n .  Thick Brass Tabs 
Figure  7.  Inductance C o i l  and Support. 
134 
ORIGINAL PAGE IS
OF POeRQUALnY
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135
80.03
0.02
0.01
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-0.01
-0.02
ORIGINAL PAGE IS
OF POOR QUALITY
Codep/Rene '
\
'U700
/
NICoCrAIY (average of 4)
s
/'" ............................. _'_"RT21/U700
- /
!
I00
t _ RT211Rene* 80
/
/
,k ! I | •
200 w 400 600 800 1000
HOURS
Figure i0. Changes in Coll Inductance wlCh Hot Corrosion at 900C, Cyclic Exposure.
(a) (b)
Figure ii. Photomicrographs of specimens from burner rig
hot corrosion test. Type i sulfides are present. Etched.
(a) NiCoCrAIY coated U700, 1004.6 hours (500X)
(b) RT21 coated Rene' 80, 697.1 hours (200X)
136


Effects of surface chemistry on hot corrosion life

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