R.J.H. Wanhill, A.J.A. Mom, H.J.C. Hersbach, G.A. Kool and J.A.M. Boogers. ABSTRACT. The NLR has two high velocity burner rigs for oxidation and hot ...
NATIONAAL LUCHT- EN RUIMTEVAARTLABORATORIUM NATIONAL AEROSPACE LABORATORY NLR THE NETHERLANDS
NLR TP 89152 U
NLR EXPERIENCE WITH HIGH VELOCITY BURNER RIG TESTING 1979- 1989
R.J.H. Wanhill, A.J.A. Mom, H.J.C. Hersbach, G.A. Kool and J.A.M. Boogers
DOCUMENT CONTROL SHEET ORIGINATOR'S REF.
SECURITY CLASS.
NLR TP 89152 U
Unclassified
ORIGINATOR
National Aerospace Laboratory NLR, Amsterdam, The Netherlands TITLE
NLR experience with high velocity burner rig testing 1979 - 1989
PREPARED FOR
a special issue of High Temperature Technology
AUTHORS
DATE
pp
R.J.H. Wanhill, A.J.A. Mom, H.J.C. Hersbach G.A. Kool, J.A.M. Boogers
890428
31
ref 7
DESCRIPTORS
Heat resistant alloys Protective coatings Corrosion tests Hot corrosion Test facilities
Jet engines Turbine blades Combustion chambers Repair
ABSTRACT
The NLR has two high velocity burner rigs for oxidation and hot corrosion testing. A large, custom-built rig was commissioned in 1977 to enable actual components to be tested. A smaller rig for specimen testing was purchased in 1986. Experience with these rigs over the last decade is reviewed.
217-02
NLR TECHNICAL PUBLICATION TP 89152 U NLR EXPERIENCE WITH HIGH VELOCITY BURNER RIG TESTING 1979 - 1989 by R.J.H. Wanhill, A.J.A. Mom, H.J.C. Hersbach, G.A. Kool and J.A.M. Boogers
This paper has been prepared for a special issue of High Temperature Technology (31 pages in total)
Division: Struct~ures and Materials Prepared: RJHW/
Completed
Approved: RJHW/
Typ.
890428
Order number: 043020 WS
-2TP 89152
.If""'
9
~
NLR EXPERIENCE WITH HIGH VELOCITY BURNER RIG TESTING 1979 - 1989
R.J.H. Wanhill, A.J.A. Mom, H.J.C. Hersbach, G.A. Kool and J.A.M. Boogers
ABSTRACT
The NLR has two high velocity burner rigs for oxidation and hot corrosion testing.
A large,
custom-built rig was
commissioned in 1977
to enable
actual components to be tested. A smaller rig for specimen testing was purchased in 1986. Experience with these rigs over the last decade is reviewed.
Keywords: turbine components; turbine coatings; repair welding and brazing; hot corrosion, oxidation; burner rig testing.
INTRODUCTION
The previous generation of aircraft based in the Netherlands experienced severe corrosion problems both in engines and airframes. At first these problems were attributed solely to the presence of environmental pollutants (chiefly S0 2 and nitrogen oxides) and seasalt particles in the atmosphere, and to the aggressive acid precipitation that resulted. Engine washing and cleaning
procedures
were
introduced,
but
it
became
evident
that more
corrosion resistant materials and coatings were required.
In 1975 the NLR was requested by the Royal Netherlands Air Force (RNLAF) to investigate the factors contributing to engine corrosion problems and to construct rigs for comparative testing of engine materials and coatings. We found
that
the
corrosion problems were
due not
only
to
environmental
factors but also to aircraft usage (mission type). This knowledge was used 81
-3TP 89152 to construct a
large burner rig capable of
flight-by-flight
simulation.
This rig was commissioned in 1977. In addition, a smaller rig was purchased in 1986: this is also capable of flight simulation testing.
In this paper we shall review the NLR's experience with these two burner rigs
during
the
last
decade.
Current
and
future
programmes
will
be
mentioned also.
THE BURNER RIGS
Burner Rig # 1 For Component Testing
A schematic of
the
large burner rig
is
shown in
figure
1 •. Air with a
pressure up to 1.45 bar is delivered by a series of centrifugal compressors to the cold air and hot gas sections. In the hot section fuel is injected and ignited with air in two J-79 burner cans. The combustion gases are then mixed in a baffled chamber to obtain a homogeneous flow.
At this stage
pollutants can be injected into the hot gas stream as required.
Just downstream to the exit nozzles are holders for non-rotating cascade assemblies of blades and vanes. The cascade holders can be moved rapidly between the hot and cold streams and also to intermediate positions. This movement is controlled by a process computer, which also controls the fuel and mass flows and pollutant injections. In turn the process computer is programmed
so
that
temperature
variations
representative
of
flight-by-
flight conditions are obtained for the cascade assemblies.
A technical specification for the rig is given in table 1. The gas stream can be controlled to any temperature between 650 °C and 1070 °C with Mach number up to 0.75. Temperature variations in the hot gas stream exiting the
81
-4TP 89152
...-'
9,)ttt:..
nozzle are typically ± 5 °C at 1000 °C. In view of the large nozzle exit area (200 mm x 50 mm) such temperature variations are very small.
Burner Rig # 2 For Specimen Testing
A schematic of the smaller burner rig is shown in figure 2. This is a test facility designed by Pratt and Whitney (PWA) and manufactured by Becan Ltd. The basic unit of this rig is a laboratory combustor whose mixing and flow characteristics and combustion chemistry closely approximate those of gas turbine combustors.
The
laboratory
systems
combustor
consists
of
two
and an instrumentation ring.
diffuser and a
burner
liner.
Fuel
sections
with
independent
The primary section is
is
injected via a
air
an annular
pressure-atomizing
nozzle and homogeneous combustion is achieved by a strong swirl-stabilized recirculation
zone.
In
the
secondary
section
the
combustion
gases
are
cooled by dilution with air from secondary jets. The instrumentation ring is
air
cooled and enables
control of
the
igniter,
a
flame monitor. and
access ports for injection of pollutants and erosive particles.
Downstream to the hot gas and cold air exit nozzles is a rotating specimen holder which can be moved rapidly between the hot and cold streams and also to
intermediate
positions.
This
movement
is
controlled
by
a
process
computer, which also controls the combustor, and enables flight-by-flight conditions to be simulated.
A technical specification for the rig is given in table 1. Its capabilities are similar to those of the large burner rig, except that a much higher gas temperature particles
81
can be
into
the
obtained and hot
gas
it
stream.
is
also
The main
possible
to
advantage
inject erosive of
the
smaller
-5TP 89152
.r'
9
,)itt:..
burner
rig
is
its much
lower
fuel
consumption:
this
is
offset
by
the
smaller test area.
HIGH TEMPERATURE MATERIALS AND COATINGS INVESTIGATIONS 1979 - 1989
Overview
Table 2 lists the burner rig investigations done by the NLR over the last decade. Most of these investigations involved testing of actual components in
the
large
burner
rig.
All
the
investigations
included
thorough
macroscopic and microscopic evaluations with extensive optical and scanning electron
metallography.
In
what
follows
the
main
aspects
of
each
investigation will be discussed.
FlOO Coating Evaluations
During introduction of the FlOO engine (F-15 and F-16 aircraft) in Europe a joint
NLR-RNLAF-PWA programme was
done
to
identify potential
corrosion
problems. Part of this programme was flight simulation burner rig testing of first and second stage turbine blades and first stage vanes coated with the standard PWA 73 pack aluminide and various alternative coatings. Four mission
types
frequencies
were
are
defined.
Their
shown in figure 3.
temperature Exact
service
profiles
and
temperature
relative simulation
(turbine inlet temperature, TIT, ~ 1425 °C) could not be achieved, and so the first stage blades and vanes were not air cooled. However, the actual metal temperatures, with the exception of slightly lower military temperatures
for
the
in-service temperatures.
first
stage
blades
and
corresponded
to
In addition, the times at cruise conditions were
shortened considerably to reduce test durations.
81
vanes,
(MIL)
-6TP 89 15 2
.rt"'
9
~
JP 4 fuel was used for this programme. Pollutants were injected to simulate the
West
European
environment
and
aircraft
usage.
S0 2
injected continuously in the hot and cold streams. NaCJI.
(5
wt.ppm)
was
(17 wt.ppm)
was
injected in the hot gas stream only during the air-to-ground missions, see figure 3.
Table 3 summarises the test programme, which began with several 150 hour burner
rig
runs
corresponding
to
594,
7 55
and
656
simulated
flights,
respectively, for the first stage blades and vanes and second stage blades. The standard PWA 73 coating performed better that the alternatives on the first stage vanes and second stage blades. However,
this coating was the
least corrosion resistant on the first stage blades.
Therefore a second,
longer test run was done with six different coatings on first stage blades up to a total of 1238 simulated flights.
The macroscopic appearances of the first stage blades after the second test run are shown in figure 4 (note the shorter test duration for some blades). It would seem that the PWA 270, RT 22A and BB coatings behaved much better than the others. But on the basis of inspections during the test run and detailed macroscopic
and microscopic
examinations
after
testing
concluded that the coatings should be ranked as follows [1]:
Coating PWA 73
81
Life in simulated service test (hours) 100
BB
< 150
PWA 275
< 150
PWA 273
150
RT 22A
150
PWA 270
275
it
was
-7TP 89152
.rt""
9
...~'
The PWA 270 coating is indicated to be much superior. Despite this,
the
standard PWA 73 coating has not been replaced yet. This is because the time between overhauls for current versions of the F100 turbine is relatively short. During overhaul the coatings are renewed as a matter of course, or else
blades
(and
vanes)
are
replaced
for
reasons
other
than
coating
degradation.
J85 Coating Evaluations
In the late 1970s GE changed the coating-material combination for J85 first stage turbine blades from Misco MDC-1 pack aluminide on Rene 100 to Codep B pack aluminide on Rene 80.
Since service experience had shown the Misco
MDC-1 coating to degrade seriously within 800 hours, the NLR was requested to evaluate the Codep B coating and, later on, several possible alternative coatings.
This was done in two burner rig test programmes,
table 4. All
blades were subjected to realistic temperature variations corresponding to ten
mission
types.
The
"salt
flights"
were
air-to-ground
missions
comprising 4 % of the simulated flights.
A preliminary investigation showed that both S0 2 and NaCi were essential for coating degradation similar to that in service. On this basis the first programme was done with 5 wt.ppm
so2
injected continuously in the hot and
cold streams, and 8 wt.ppm NaCt injected in the hot gas stream during the "salt flights". 1800
The test duration was
simulated flights.
blades
after
testing
Macroscopic
indicated
that
150 hours,
corresponding to about
and
microscopic
examination
the
corrosion resistances
of
the
of Misco
MDC-1 and Codep B coatings were similar.
The second programme was done with 17 wt.ppm NaCi injected in the hot gas stream during the "salt flights". The increased salt injection as compared
81
-8.rt"'
9
~
TP 89152 to the first programme was considered necessary to ensure degradation of all
the
coatings.
Also
the
test
duration was
increased
to
270 hours,
corresponding to about 3000 simulated flights. A classification of coating - and base metal -
attack is
given in table 5.
The RT 22A noble metal
aluminide coating was definitely superior. The performance of the RT 100 NiCoCrAtY
overlay
was
disappointing
in
comparison
to
the
results
PWA 270 in the F100 programme. The reason is coating quality:
for
the RT 100
was unevenly distributed, varying in thickness from 10 - 100
~m,
contained
problem with
significant
porosity [2].
This
is
a
well-known
and also
plasma sprayed coatings.
Although the second programme showed RT 22A to be superior to Codep B, the increased coating costs were considered to outweigh potential benefits in service. Thus the J85 first stage turbine blades continue to be coated with Codep B.
CF6-50 Coating Evaluations
A joint NLR-KLM investigation showed that CF6-50 first stage turbine blade life was
limited by degradation of the external Codep B pack aluminide
coating and corrosion attack in the uncoated cooling passages [3]. This led to
a
project
collaborative
within effort
the
COST 50
sponsored
by
Round III the
programme,
Commission
of
which the
was
a
European
Communities.
The
COST 50
project
consisted
of
two
coating
evaluation
programmes,
table 6. In the first programme the standard Codep B external coating was compared with several possible alternative coatings. turbine
blades
were
subjected
to
realistic
CF6-50
temperature
first
stage
variations
corresponding to six flight types based on GE data, figure 5. However, the
81
_,
9./I'
-9TP 89152 blades
were
internally
not
cooled.
The
duration
test
was
300
hours,
equivalent to 1200 simulated flights. Remarkably, it turned out that the Codep B coating had the best corrosion resistance under these conditions.
In view of this result, the second programme took Codep B as the baseline external
coating.
Seven
new
CF6-50
externally and internally coated as tested
without
internal
cooling
first
stage
turbine
listed in table 6.
using
two
blades
were
The blades were
simplified
versions
of
the
flights in the first programme. Repeated blocks of these simplified flights were
applied
as
shown
in
figure 5.
The
test
duration was
270 hours,
equivalent to 1367 simulated flights.
The
results
of
classification
the
of
second
coating
programme
performances
external coating combination was Codep B coating was MDC 43.
Taking
into
account
are is
given
shown
in
+ Sermaloy the
in [4].
The
table 7.
final
The
best
J. The best internal
desirability,
from a
cost
viewpoint, of simultaneous external and internal coating, the best overall coating system was Codep B
In
any
event,
+ C 30.
significant
improvements
in
external
and
internal
corrosion resistance were demonstrated by the COST 50 project.
hot
To date,
however, an improved coating system has not been applied to CF6 first stage turbine
blades
for
service.
Other
problems
that
limit
various
engine
component lives currently have higher priorities.
FlOO Braze and Weld Repairs
The most urgent problem arising from F100 turbine service has been the cracking of first stage vanes. These are complex and expensive components made
81
from
DS MarM200
+ Hf and
coated
with
the
standard
PWA 73
pack
-10TP 89152
.rt""
9
_,it!.
aluminide.
Repair is feasible owing to
times of replacements.
the high cost and long delivery
The original repair process involved welding the
cracks shut - as far as possible - and local patch repair of the coating. This process has several limitations, including the inability to repair all cracks
and
a
high
scrap
densification healing
rate.
An
alternative
repair
called
diffusion
(DDH) was developed by the Chromalloy Corporation.
This repair requires the following steps:
• stripping the entire coating • cleaning in a gaseous fluoride atmosphere to remove all corrosion • welding large cracks shut • filling all other cracks with a braze material that at the brazing temperature melts, diffuses into the base metal and resolidifies • hot isostatic pressing (HIP) to densify and further homogenize the repaired areas • recoating.
Compared to repair welding the DDH process enables many more components to be
salvaged
and
there
is
improved
structural
integrity.
The
NLR
was
therefore requested to evaluate the two repair processes with respect to simulated service performance in the large burner rig. repaired and new vanes were tested
Using JP 4 fuel,
with internaZ aooZing for 1900 simulated
flight cycles that in the latter part included intermittent injection of pollutants, figure 6.
Full results are reported in [5]. DDH repaired vanes performed almost as well as new vanes with respect to both coating degradation and base metal cracking.
Weld
repaired
vanes
confirmed by PWA full-scale approved for service.
81
were
engine
much tests,
inferior. and
These
results
were
the DDH repair was
then
-11TP 89152
.If""
9
~
Thermal Barrier Coatings (TBCs)
Until recently the FlOO combustor and augmentor liner have been protected
+ Ni-5 % At
by a TBC consisting of a magnesium stabilised zirconia topcoat bondcoat.
This coating has
limited life owing to
spalling.
The NLR was
requested to compare the thermal shock and long term oxidation resistances of
this
coating with
zirconia topcoat topcoat
+
two
possible
alternatives:
a
magnesium
stabilised
+ NiCoCrAtY bondcoat, and an yttrium stabilised zirconia
NiCoCrAtY bondcoat.
Using the smaller burner rig two test programmes were carried out, table 8. The
TBCs
were
applied
by
air
plasma
spraying
onto
the
substrates.
Simplified flight-by-flight temperature profiles were used, figure 7. These were closely followed by the specimens on the rotating holder.
Results of the test programmes are fully described in [6]. The Zr0 2 /Y 2 0 3
+
NiCoCrAtY coating was clearly superior and has now replaced the Zr0 2 /Mg0
+
Ni-5 % At coating for RNLAF service. However,
the tests also showed that
the TBC and substrate degradation mechanisms were very similar and depended on test type and duration as shown in figure 8. In particular, substrate oxidation was strongly time-dependent, beginning relatively early in the long term oxidation tests.
This makes
it worthwhile
to
investigate
the
potential usefulness of TBCs containing an additional layer acting as a barrier to inward diffusion of oxygen [7].
DISCUSSION
To date the NLR's experience with high velocity burner rig testing has been directed mainly to ad hoa problems This situation reflects - besides
81
rather
than research investigations.
the urgency of service problems -
the
-12TP 89152
.rr"
9
~
absence
of
a
jet
engine
industry
in
the
Netherlands.
However,
the
increasing emphasis on collaborative programmes, especially within Europe, is likely to encourage our participation and that of others in more general research
activities.
In
this
respect
we
envisage
a
two-tier
approach,
whereby several testing variables (e.g. material and coating combinations) are
examined
using
specimens
in
the
smaller
burner
rig,
followed
by
stringent selection for component testing in the large rig.
A prerequisite for successful collaborative programmes is agreement on test techniques. Burner rig testing is no exception, and we strongly endorse the VAMAS initiative to establish standard test procedures. In our view, there should be standards to define the environmental conditions (fuel chemistry, injected pollutants and erosive particles) for high temperature oxidation and types I and II
sulphidation.
These standards could be used for both
general research and ad hoc problems.
A more difficult problem is to establish standards for temperature profiles and mission sequences on a flight-by-flight basis. This is because not only must agreement be reached on the flight histories to be simulated, but also different
engine
components
experience
different
temperature
profiles
during the same mission sequence. We suggest a flexible approach whereby standards are developed to act as guidelines for actual test conditions. Even so, flight simulation testing is essential.
81
-13-
9·""
,)1::.. '
TP 89152
CURRENT AND FUTURE INVESTIGATIONS
Actual and potential NLR burner rig test programmes over the next few years include
• specimen
tests
in
the
smaller
burner
rig
to
enable
direct
comparisons with component tests in the large rig • coating evaluations, barrier
additions
including improved MCrAiY overlays;
to
TBCs;
local
(patch)
repairs;
and
diffusion relative
performances of coatings on unrepaired and repaired components • coating
behaviour
under
type II
(low
temperature)
sulphidation
conditions
• FlOO 4000 cycle turbine component testing.
This list is not complete: based on our experience over the last decade, we expect service problems to require additional testing.
81
9.)t!.·""'
-14TP 89152 REFERENCES
1.
Mom,
A. J. A. ,
and Hersbach,
H. J. C.
"Performance
of
high
temperature
coatings on F100 turbine blades under simulated service conditions", Materials Science and Engineering, 87 (1987) pp. 361-367.
2.
Mom,
A.J.A.,
Boogers,
J.A.M.
and Hersbach,
H.J.C.
"Burner rig
test
behaviour of various coatings applied on J85 1st stage turbine blades", NLR Technical Report 84075, National Aerospace Laboratory, Amsterdam, August 1984.
3.
Mom, A.J.A. "Codep B degradation on CF6-50 1st stage turbine blades as a
function
of
service
life",
NLR Technical Report
83027,
National
Aerospace Laboratory, Amsterdam, March 1983.
4.
Mom, A.J .A.
and Boogers, J .A.M.
"Simulated service test behaviour of
various internal and external coatings applied on CF6-50 first stage turbine blades", Applications,
High Temperature Alloys for Gas Turbines and Other
Part
II,
D.
Reidel
Publishing
Company
(1986)
pp.
1245-1264.
5.
Mom, A.J .A., Madhava, M., Kool, G.A. and Dean, M.
"Evaluation of DDH
and
simulated
weld
repaired
F100
turbine
vanes
under
service
conditions", Advanced Joining of Aerospace Metallic Materials,
AGARD
Conference Proceedings No. 398, Advisory Group for Aerospace Research and Development (1986) pp. 21-1 - 21.9.
81
-15TP 89152
.If"'
9,Itt-
6.
Boogers, J.A.M., Wanhill, R.J.H.
and Hersbach, H.J.C.
"Thermal shock
and
ceramic
NLR
oxidation
89085,
Publication
1989.
To
resistance
be
of
National
published
in
an
Aerospace AGARD
coatings", Laboratory,
Conference
Technical
Amsterdam,
Proceedings
March
on High
Temperature Surface Interactions.
7.
Burgel, R. and Kvernes, I. "Thermal barrier coatings", High Temperature Alloys
for
Gas Turbines
and Other Applications,
Publishing Company (1986) pp. 327-356.
81
Part
I,
D.
Reidel
TABLE 1: Technical specifications of the NLR burner rigs
SPECIFICATIONS
LARGE RIG (# 1)
SMALLER RIG (# 2)
440 kW
127 kW
"' 1.5 kg/s
"' 0.4 kg/s
"' 3.0 kg/s
"' 0.8 kg/s
e COMPRESSOR FACILITY e MASS FLOW - hot section at 1000 °C - cold section
e
TOTAL AIR PRESSURE
1.45 bar
;::£
;::;;
1.45 bar
I
oo-0'\
~
e MACH NUMBER e
GAS TEMPERATURE
;: ; 0.75 600 °C - 1070 °C
e GAS TEMPERATURE VARIATION ACROSS NOZZLE AT 1000 °C e
PRESSURE VARIATION ACROSS HOT NOZZLE
e
RESPONSE TIME 1000 °C + 300 °C
e TEST AREA e
INJECTION OF POLLUTANTS
500
oc ± 5
%
± 5
6s 200 mm x 50 mm - S0 2
,
NaCt, etc.
I
0.2 - 0.7 1650
Vl N
oc
oc
± 5 %
lOs 50 mm diameter - S0 2 , NaCt, etc. - erosive particles
TABLE 2:
Overv~ew
of NLR high velocity burner
INVESTIGATION
r~g ~nvestigations
COMPONENTS/SPECIMENS
!
1st stage blades
F100 TURBINE COMPARATIVE EVALUATIONS OF EXTERNAL COATINGS
2nd stage blades 1st stage vanes
TIME PERIODS
BURNER RIG
1979 - 1981 1979 1979
J85 1st stage turbine blades
1979,
1984
COMPARATIVE EVALUATIONS OF EXTERNAL AND INTERNAL COATINGS
CF6-50 1st stage turbine blades
1982 - 1985
EVALUATION OF BRAZE AND WELD REPAIRS
F100 1st stage turbine vanes
1984 - 1985
# 1
H 1-cJ
THERMAL SHOCK AND OXIDATION RESISTANCES OF THERMAL BARRIER COATINGS
SPECIMENS: Hastelloy X, HS 188
1987 - 1989
I
oo~
~
# 2
ln N
"-I
TABLE 3: FlOO coating evaluation programme
COATINGS PWA 73 PWA 273 PWA 275 BB RT 22A PWA 270
TYPES
APPLICATION METHODS
Pack aluminide (inward diffusion) Pack aluminide (outward diffusion) Gas phase aluminide Rh-noble metal aluminide Pt-noble metal aluminide NiCoCrAiY overlay
Ai-Si pack process Ai pack process Ai gas phase process Rh electroplate + Ai pack process Pt electroplate + Ai pack process Physical vapour deposition (PVD), electron beam
COATINGS COMPONENTS
MATERIALS 1ST TEST RUN (150 HOURS)
1ST STAGE BLADES 1ST STAGE VANES 2ND STAGE BLADES
2ND TEST RUN (250-300 HOURS) I
DS* MarM200 + Hf DS MarM200 + Hf DS MarM200 + Hf
PWA 73,BB,RT 22A,PWA 270 PWA 73,BB,RT 22A PWA 73, PWA 270
PWA 73,PWA 273,PWA 275,BB,RT 22A,PWA 27C
Vl N
METAL TEMPERATURES (oC) COMPONENTS MILITARY (MIL) 1ST STAGE BLADES 1ST STAGE VANES 2ND STAGE BLADES * DS
Directionally Solidified
1050 1050 983
co-
\.OCO
CRUISE
IDLE
813 805 830
590 562 446
I
TABLE 4: J85 first stage turbine blade coating evaluation programmes
PROGRAMMES COATINGS
TYPES
APPLICATION METHODS I
MISCO MDC-1
Low activity pack aluminide
Slurry A£ 2 0 3 + AJL pack process
CODEP B
Low activity pack aluminide
AJL pack process
Cr-AJL (1)
Low activity pack aluminide
Cr-AJL pack process
RB 505
Low activity pack aluminide
Cr-AJL pack process
RB 505 + Cr
Low activity pack aluminide
Cr-AJL pack process with Cr enrichment in top layer
PWA 73
• •
II
•
• • • • • • • • • • •
- NaCJL in hot gas stream during "salt flights" (see text)
8 wt.ppm
17 wt.ppm
- Test duration
150 hours
270 hours
High activity pack aluminide
AJL-Si pack process
RT 22A
Pt-noble metal aluminide
Pt electroplate + AJL pack process
SERMALOY J
AJL-Si slurry coating
Slurry AJL-Si + indiffusion treatment
RT 100
NiCoCrAJLY overlay
Low pressure plasma spray
- Monolithic Rene 80 blades (no internal cooling required)
•
- JP 4 fuel
•
- 5 wt.ppm S0 2 continuously in hot and cold streams
H
t;j
oo-
1.0 1.0 Vl N
I
TABLE 5: Classification of J85 first stage turbine blade coating performance (second programme)
CORROSION ATTACK COATING RANKING LEADING EDGE
RT 22A RB 505 + Cr PWA 73
Decreasing
CODEP B
Corrosion
RT 100
Resistance
RB 505 SERMALOY J Cr-AR. (1)
•
• • • •• • •• •• •• ••
• •• •• •• •• •• ••••
•••
Superficial attack
e e e Coating inner layer attack
TRAILING EDGE
••
CONVEX SIDE
• • • •• •• ••• •• •••
Coating outer layer attack
e e •• Substrate attack
CONCAVE SIDE
• • • •• •• ••• •• ••••
t-3 1-d
I
OON
~0
I Vl
N
TABLE 6:
CF6-50 first stage turbine blade (Rene 80) coating evaluation programmes ' APPLICATION METHODS
TYPES
COATINGS
Low activity pack aluminide Pt-noble metal aluminide 2-step low activity pack aluminide NiCoCrA9.,Y overlay High activity gas phase aluminide High activity gas phase aluminide At-Si slurry coating High activity gas phase aluminide High activity gas phase aluminide High activity gas phase Cr-At
CODEP B RT 22A C1A LC0-22 HI 275 c 30 SERMALOY J MDC 43 PULSE ALUMINIDE Cr-At (2)
A9., pack process Pt electroplate + A9., pack process Cr-A9., pack process Low pressure plasma spray A9., gas phase process A9., gas phase process Slurry At-Si + indiffusion treatment At gas phase process At gas phase with pulsed argon pressure Cr-At gas phase process
I
PROGRAMME I
OON
"' -
PROGRAMME II
I
EXTERNAL COATINGS CODEP B RT 22A C1A LC0-22 LC0-22 + PULSE ALUMINIDE
-
TEST CONDITIONS
EXTERNAL COATINGS
INTERNAL COATINGS
6 flight types Jet A-1 fuel 7 wt.ppm so2 continuously 20 wt.ppm NaCt in flight F4 Test duration 300 hours (equivalent to 1200 flights)
CODEP B CODEP B CODEP B + CODEP B + CODEP B + CODEP B + HI 275
Cr-At (2) c 30 SERMALOY J MDC 43 PULSE ALUMINIDE HI 275
C 30 SERMALOY J MDC 43 PULSE ALUMINIDE
TEST CONDITIONS -
2 flight types Jet A-1 fuel 7 wt.ppm so2 continuously 20 wt.ppm NaCt in flight F2 Test duration 270 hours (equivalent to 1367 flights)
Vl N
.
TABLE 7· Classification of CF6 - 50 first stage turbine blade coating performance (second programme)
CORROSION ATTACK COATING RANKING LEADING EDGE Decreasing Corrosion Resistance of External Coatings
It
CODEP B CODEP B CODEP B CODEP B HI 275 CODEP B
+ + + +
TRAILING EDGE
••
SERMALOY J C 30 PULSE ALUMINIDE MDC 43
CONVEX SIDE
CONCAVE SIDE
•• ••• ••• ••• ••••
• ••• ••••• •• •••• ••••
• ••• •• •• •••
•••• •••• ••••
•••• ••••
••••
COOLING CHANNEL CORROSION ATTACK COATING RANKING 1
2
3
4
5
6
7
8
I CP N \ON
Decreasing Corrosion Resistance of Internal Coatings
•••
It
MDC 43 c 30 PULSE ALUMINIDE Cr-A£ (2) HI 275 SERMALOY J
Superficial attack Coating outer layer attack Coating inner layer attack •••• Substrate attack
•••
•••• •••• •••• •••• •••• ••••
I
COOLING
••••• ••••• •••• ••••• •••
•• •• •••• •• •••• ••••
•• •••• • ••
••• ••• •••••• ••
••• •• •••
• •• •••
•••• ••• •••• •• • • ••••• •••• •••• •••• ••••
CHANNELS~
¥:~· 1
• ••••• •••
5
~~~ 8
I \J1 N
.
TABLE 8 · F100 combustor and augmentor liner TBC evaluation programmes [6]
PROGRAMMES
TBC COMBINATIONS
Zr0 2 /Mg0
SPECIMEN CONFIGURATIONS (DIMENSIONS IN mm)
SUBSTRATES
ICYLINDRICAL TUBE '
+ Ni-5 % Afl
I
TEST CONDITIONS
- Simplified flight temperature profiles
Hastelloy X Zr0 2 /Y 20 3 + NiCoCrAR-Y
COATING
THERMAL
10
SHOCK Zr0 2 /Mg0
r[-·f·l· -· -·-· -··3: 57.5 1:12.5.,j
+ Ni-5 % Ai
1..
:I
80
HS 188 Zr0 2 /Y 2 0 3 + NiCoCrAR-Y
II
LONG TERM OXIDATION
Zr0 2 /Mg0
+ Ni-5 % AR-
Zr0 2 /Mg0
+ NiCoCrAiY
Zr0 2 /Y20s + NiCoCrAH
I HS 188
·{I
I.
1.of=
~
FLAT STRIP
II
I
:I
57.5 80 COATING
- JP 4 fuel - Thermal shock test duration 56 hours maximum (equivalent to 1736 lowhigh-low thermal cycles) - Long term oxidation test duration 200 hours maximum (equivalent to 200 lowhigh-low thermal cycles)
I (X) N
1.0
w I
\..n
N
PROCESS COMPUTER
FLIGHTS
- - - - BLADE TEMPERATURE --BRACKET POSITION (INPUT)
AIR--
MIXING CHAMBER TEMPERATURE
BRACKET POSITION (OUTPUT)
AIR--~
I CON
1.0.;::-. I Ul N
BURNER CANS------'
EXIT NOZZLE COLD STREAM
EXIT NOZZLE HOT STREAM
440 mm
Fig. 1
Burner rig #1 for component testing
MASS FLOW FUEL FLOW FUEL TANK
SPECIMEN POSITION
POLLUTANTS
PROCESS COMPUTER
POLLUTANT INJECTION
FLIGHTS
SPECIMEN TEMPERATURE
EXHAUST NOZZLE
0
0
0
SPECIMENS
0
L-;::::::~sPECIMEN HOLDER
I
I
~OVEMENT
lIIIII
OF SPECIMENS FROM HOT TO COLD NOZZLE
SECONDARY AIR
PRIMARY AIR
I
nnnnn
COLD NOZZLE
111 I l l I l l I II I IJI IJ I I l l 1 11 l1 II II11I 1 II 1I1 1JI ----.1.1-'.L~
1
L
--.,--11 ~
1
11
r
''I-I...' II II II
PRESSURE TANK
Fig. 2
Burner rig #2 for specimen testing
I
II 'I Li
I 00 N 1.0 VI VI N
I
-26TP 89152
AIR COMBAT/AERO'S (40%)
MIL CRUISE IDLE 1-------~
AIR-TO-GROUND MISSION (20%)
X X X XXXXXXXXX X X -INJECTION OF SEASALT DURING20s
MIL CRUISE IDLE 1 - - - - - '
COMBAT PROFILE MISSION (30%)
MIL CRUISE IDLE 1-----'
GENERAL FLYING (10%)
MIL CRUISE IDLEI----1
0
90
30 ---ENGINE RUN TIME (minutes)
Fig. 3
81
Temperature profiles and occurrence frequency of the selected missions for FlOO turbine component testing [1]
-27TP 89152
PWA 73
250 TEST HOURS
PWA 273
250TEST HOURS
PWA 275
250 TEST HOURS
BB
300TEST HOURS
RT 22A
300TESTHOURS
PWA 270
300 TEST HOURS
Fig . 4
Coated FlOO first stage turbine blades after the second burner rig test run
PROGRAMME I
PROGRAMME ll
RANDOMIZED BLOCKS OF 200 FLIGHTS OCCURRENCE FLIGHT MAXIMUM TEMt't11HIURE FREQUENCY TYPE (oC) (%)
REPEATED BLOCKS OF 15 FLIGHTS
Fl
946 979 1006 1023 1029 1033
F2 F3 F4 F5 F6
7F1 + F2 + 6F1 + F2
15 35 30 10 5 5
SALT INJECTION (475 s) DURING INITIAL 53 TEST HOURS
hHTF4 .,1 I,.
SALT INJECTION (60s)
I
1..
..I
I
SALT INJECTION (238 s) AFTER 53 TEST HOURS
(X)N \.O(X)
V1 N
FliGHT CYCLE F1
900 s
QL--------------------------------TIME
Fig. 5
225 s
FLIGHT CYCLE F2
3900 s
QL-------------------------------------~)~------------
-TIME
Temperature profiles and occurrence frequency of the selected flights for CF6-50 first stage turbine blade testing
I
NORMAL DESIGN COOLING PARAMETERS
REDUCED COOLING CONDITIONS
0-900 CYCLES Tmax=985° C
901-1300 CYCLES T max= 1035° C
1301-1900 CYCLES Tmax =1050° C 1301-1500 CYCLES
1501-1900 CYCLES
7 ppm
so2 , 45 s/cycle
20 ppm
NaCI, 45 s/cycle
7 ppm
so 2 , 45 s/cycle
10 ppm
NaCI, 20 s/cycle
1000
cooling rate 985°C-- 300°C 90°C/s
cooling rate 1035°C -300°C 105°C/s
cooling rate 1050°C -300°C 90°C/s
I 00 N \.0 \.0
U1 N
400 heating rate 65°C-800°C
heating rate 55°C-800°C
160°C/s
heating rate 75°C- 800°C
165°C/s
175°C/s
200
0 60s
Fig. 6
45 s
60s
45 s
45 s
35 s
TIME (s)
Temperature profiles and pollutant conditions during 1900 simulated flight cycles for repaired FlOO first stage turbine vanes [5]
I
ONE SEQUENCE =60 MINUTES 30 CYCLES
THERMAL SHOCK
~
~I
30 MINUTES HOLD
r--1'--,
oc
I I
t
I
450
I
\
'\
~ H "d 00
I
w
\.00 \J1 N
-TIME
1050940-
n
8 MINUTESr----1
oc
I
LONG TERM
OXIDATION~
t
100-
' ONE SEQUENCE
60 MINUTES
-TIME
Fig. 7
Temperature profiles for the TBC evaluation programmes [6]
l>l
I
TYPE OF TEST
TBS ANO SUBSTRATE DEGRADATION
O.UALITATIVE ASSESSMENT OF DEGRADATION OCCURRENCE AND SEVERITY
TOPCOAT CRACKS
I SLIGHT
THERMAL SHOCK AND LONG TERM OXIDATION
MODERATE
SEVERE
BONDCOAT OXIDATION
BOND COAT /TOPCOAT CRACKS
I
TOPCOAT SPALLING
"----------'> I
_,
oow
__
.___
BONDCOAT CRACKS THERMAL SHOCK
1.0 -
_,)
SLIGHT
SUBSTRATE OXIDATION
TOPCOAT MICROSPALLING LONG TERM OXIDATION
I
) SLIGHT
MODERATE
SUBSTRATE OXIDATION
-TEST DURATION Fig. 8
TBC and substrate degradation mechanisms according to test type and duration [6]
SEVERE
lJl N
