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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