The objective is to develop heat flux sensors suitable for use on turbine airfoils and to verify the operation of the heat flux measurement techniques through laboratory experiments. The requirements for a program to investigate the measurement of heat flux on airfoils in areas of strong non-one-dimensional flow were also identified

Atkinson, W. H.

Strange, R. R.

English

NASA Technical Reports Server

N87-11186
DEVELOPMENT OF HEAT FLUX SENSORS IN TURBINE AIRFOILS
William H. Atkinson and Richard R. Strange
Uni ted Technol ogies Corporati on
Pratt & Whitney
This paper describes the work performed under contract NAS3-23529 "Turbine Blade
and Vane Heat Flux Sensor Development." The objective of this contract is to develop
heat flux sensors suitable for use on turbine airfoils and to verify the operation
of the heat flux measurement techniques through laboratory experiments.
The design of durable turbine airfoils that use a minimum amount of cooling air
requires knowledge of the heat loads on the airfoils during engine operation. Meas-
urement of these heat loads will permit the verification or modification of the ana-
lytical models used in the design process and will improve the ability to predict
and confirm the thermal performance of turbine airfoil designs. Heat flux sensors
for turbine blades and vanes must be compatible with the cast nickel-base and
cobalt-base materials used in their fabrication and must be capable of operation in
a hostile environment with regard to temperature, pressure and thermal cycling. In
order to not perturb heat flows to be measured, it is necessary to use the smallest
possible sensors.
During the first phase of this contract two sensor designs were identified that
met the established criteria for turbine airfoils. These sensors were the embedded
thermocouple sensor and the Gardon gage sensor. Both were fabricated into the air-
foi I wal I.
Figure l is a schematic of the embedded thermocouple sensor. This type of one-
dimensional steady-state sensor determines the heat flux by measuring the tempera-
ture drop across a thermal barrier. In this case, the airfoil wall acts as the ther-
mal barrier. The Gardon gage sensor is shown schematically in figure 2. This type
of steady-state sensor determines the heat flux from the temperature rise due to ra-
dial conduction from the insulated section of the hot side of the airfoil wall.
Both types of sensors were fabricated into turbine blades and vanes. The instal-
lation of the sensors required access to the inside wall of the airfoil. The blades
used for the heat flux sensor installations were two-piece bonded blades. The in-
strumentation was installed in one blade half. The two halves were then joined using
a 1255K braze material. For installation in turbine vanes, a window, through which
instrumentation was installed, was cut into the opposite airfoil wall. Figure 3
shows a vane being instrumented. After completion of instrumentation the window was
heliarc welded back in place. The airfoil wall was then smoothed to restore aero-
dynamic integrity. NASA has fabricated vanes by cutting them in half, installing the
instrumentation and then brazing the vane halves back together.
A calibration fixture, shown in figure 4, was designed and fabricated for the
calibration of airfoil-mounted heat flux sensors. This fixture was mounted below a
quartz lamp bank heat source, shown in figure 5. The fixture allowed positioning of
the airfoil so that the surface of the heat flux sensor was normal to the incident
radiation. The surface of the airfoil was coated with material having a known and
stable absorptance and emittance. The incident radiation was measured with a
reference Hy-Cal asymptotic calorimeter. The sensors were calibrated at a number of
49
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heat flux levels and over a range of sensor temperatures. The data was then normal-
ized to a temperature of 120OK. Figure 6 shows a typical calibration result for a
Gardon Gage sensor. This same apparatus was used for thermal cycle and thermal soak
tests as well as the calibration tests.
The sensors developed under the first phase of the turbine blade and vane heat
flux sensor program proved capable of measuring the heat loads on turbine airfoils
up to 1.6 megawatts/m L incident with an accuracy of +5 percent. They withstood
thermal cycling and thermal soak conditions expected in-a real gas turbine engine
environment. Sensors must be carefully installed in airfoils with complex internal
cool ing schemes. In such airfoi Is there may be regions where strong
non-one-dimensional heat flows exist. These non-one-dimensional flows make it dif-
ficult to obtain meaningful sensor calibrations. The full details of the sensor de-
velopment and laboratory testing are contained in references l and 2.
The second phase of this program demonstrated a variety of heat transfer meas-
urement methods on a simple test piece in an atmospheric pressure combustor rig. A
cylinder in cross-flow was chosen because this configuration has been tested exten-
sively and is well documented in the literature.
The combustor rig selected is a general purpose laboratory combustor shown sche-
matically in figure 7. The mixing and flow in the combustor closely approximated
those of many conventional gas turbine combustors, and the critical features of the
combustion chemistry were reproduced. The fuel was introduced at the front of the
combustor by a pressure atomizing fuel nozzle. The combustion was stabilized in the
forward section of the combustor by developing a strong, swirl-stabilized re-
circulation zone. The hot combustion gases were cooled by dilution jets located at
the back of the combustor. The primary and secondary jets have independent air sup-
plies to permit adjustment of the velocity and temperature profiles at the exit. The
ignitor, a flame monitor, and instrumentation ports were mounted in an instrumenta-
tion rig located between the primary and secondary sections. The 5 cm diameter ex-
haust could he operated at a temperature up to 1700K at Mach numbers up to 0.7 at
this maximum temperature.
The instrumentation cylinder selected was a 1.6 cm diameter tube made of
Hastelloy-X with a wall thickness of 0.15 cm. The diameter of the cylinder was a
compromise between a size large enough to permit installation of the instrumentation
and small enough to cause minimal flow blockage. During test the cylinder will be
cooled with channel-flow coolant in the tube. To obtain the required internal heat
transfer coefficient, the channel height was reduced by adding a center tube
concentric with the cylinder. The cylinder was prepared for instrumentation by
slicing a portion of the tube length in half. This allowed direct access to the
interior of the cylinder for the necessary Elox operations as well as installation
of instrumentati on.
Figure 8 shows a cylinder following the Elox operation but before installation
of the instrumentation. The sensor on the left will be a transient slug calorimeter;
the center sensor will be a Gardon Gage; and the sensor on the right will be an
embedded thermocouple sensor. Figure 9 shows the same cylinder after installation of
lead wires but before the interior wall has been smoothed or the Gardon gage cavity
has been filled with ceramic. Figure lO shows the completed installation, with the
center tube installed, prior to cylinder rewelding. The instrumentation is led out
through the center tube. Figure II is an end-on view of the cylinder following
rewelding of the cylinder halves.
50
Following construction, the heat flux sensors in the cylinder were calibrated
using the quartz lamp bank calibration facility. Figure 12 shows typical calibra-
tions for a Gardon gage sensor installed in a cylinder. All of the calibrations were
well-behaved. None of the steady-state sensors appeared to be affected by non-one-
dimensional flow problems.
Preparations are now underway to test the sensors in the cylinders in front of
an atmospheric pressure combustor. Initial tests are being conducted to characterize
the conditions at the combustor exit at a number of run conditions. Pratt & Whitney
is conducting tests to characterize the thermal and pressure profiles; both the spa-
tial and temporal variations of these quantities are being measured. In a parallel
effort, NASA Lewis Research Center is conducting Laser Doppler Velocimeter (LDV)
tests on a similar combustion rig. These tests are to characterize the turbulence
profiles behind the combustor. Testing will be done both with and without a flow
straightener. The preliminary test data looks good. The combustor has uniform ther-
mal profile at the exit, and relatively constant temporal variation and turbulence
Ievel s.
When the combustor characterization tests are complete, testing of the cylinder
behind the combustor will begin. The cylinder will be mounted 5 cm from the
combustor exit. Any free-stream instrumentation will be located upstream from the
cylinder, 2.5 cm from the combustor exit. Preliminary test data shows this will put
the cylinder in an area of uniform thermal and Mach number profiles, minimizing both
cylinder blockage effects and the effect of the cylinder on the free-stream
instrumen tati on.
The cylinder will be mounted on a traverse can. This will allow the various sen-
sor types to be sequentially moved into place behind the combustor, permitting mul-
tiple sensor types to be tested at the same conditions. All three sensors (the
steady-state embedded thermocouple sensor, the steady-state Gardon gage, and the
transient slug calorimeter) will be run over a range of test conditions. The result-
ing data should allow a good comparison between sensor types as well as between the
sensor results and the expected value (which will be calculated from the measured
free-stream condi tions).
A second cylinder will be tested under a limited number of conditions. That cyl-
inder will be a 1.6 cm diameter NiCoCrAIY tube with wall thickness of 0.48 cm. An
array of sputtered thin film thermocouples will be installed on this cylinder. These
will be run in front of the combustor rig in conjunction with a high response ther-
mocouple probe (developed under NASA contract NAS3-23154, reference 3). The heat
transfer coefficient is calculated from the ratio of the thermal variations of the
free-stream gas and the surface thermocouples.
The work done under this contract has yielded accurate durable sensors for use
on those sections of turbine airfoils where the heat flow conditions are relatively
uniform. It also will yield a data base of high temperature heat flux measurements
by several methods on a cylinder in cross-flow. In addition, the requirement for a
program to investigate the measurement of heat flux on airfoils in areas of strong
non-one-dimensional flow has been identified.
51
I •
o
•
REFERENCES
Atkinson, W.H.; Cyr, M.A.; and Strange, R.R.: Turbine Blade and Vane Heat Flux
Sensor Development. NASA Contract, NAS3-23529, August 1984.
Atkinson, W.H.; and Strange, R.R.: Development of Advanced High Temperature Heat
Flux Sensors. NASA Report, CR-165618, September 1982.
Elm•re, D.L.; Robinson, W.W.; and Watkins, W.B.: Dynamic Gas Temperature Meas-
urement System. Final Report, CR-168167. NASA Contract, NAS3-23154. May lO, 1983.
52
EMBEDDED SWAGE WIRE THERMOCOUPLES
COLD SIDE
_ 90 °
HO
ELECTRICAL SCHEMATIC EMBEDDED THERMOCOUPLE SENSOR
ALUMEL
Z CHROMEL
.¢.1// / // /111111(
HOT SIDE _ ALUMEL
1-2 = REFERENCE TEMPERATURE
1-3 = SENSOR OUTPUT
Figure 1 Schematic of the Embedded Themocouple Heat Flux Sensor
GARDON GAUGE
_'_ MATERIAL (FILL) \
_o,.wA:.
HOT SIDE
ELECTRICAL SCHEMATIC GARDON GAUGE SENSOR
ALUMEL
r ALUMEL
/I , CH.OM_L
-/ COLD SIDE / I IW//;_A)i l.l.'_,.."o,<'w_"<<_
1-2 = SENSOR OUTPUT
1-3 = REFERENCE TEMPERATURE
Figure 2 Schematic of the Gardon Gage Heat Flux Sensor
53
ORiGlNAL PAGE IS 
OF POOR QUALrrV 
Figure 3 Vane i n  P r  cess o f  Being Instrumented 
/ / /  / /  
\ MATE! C0a)ED REFLECTOR 
/ /  
r( QUARTZ HALOGEN LAMP 
- iv 
REFERENCE SENSOR 
F i  gure 4 Ca l  i b r a t i  on F i  x t u r e  
54 
Figure 5 Lamp Face o f  Quartz Lamp 
lsoor CRRDON GAUGE SERIRL NUHBER - G-BZT-IIC 
61 REF >- 1150 K 
I 
0 100 200 300 400 500 6 020 700 800 
HER1 FLUX TRRNSt4JTTED - K I L O W A T T S / M  
Fi gure 6 Calihration Curve - Gardon Gage Installed in Turbine Vane 
55 
FUEL 
NOZZLE 
PLATE 
INSTRUMENTED m 
PRIMARY 
SECTION 
INSTRUMENTATION 
RING 
EXHAUST 1 
NOZZLE 
PLATE I 
Figure 7 Schematic o f  Combustor Ri g 
Fi gure 8 Open Cylinder - Elox Completed 
56 
ORiGlNAL PAGE E3 
OF POOR QUALITY 
Figure 9 Open Cylinder - Lead Wires Installed 
Figure 10 Open Cylinder With Center Tube 
57 
ORIGINAL PAGE E3 
OF POOR QUALm 
Figure 1 1  End View of  Completed Cylinder 
_ _ _ _ _ _ - _  5% OF 1 M E G A U A T T / M ~  
00 
Figure 12 Calibration Curve - Gardon Gage Installed in Cylinder 
58 


Development of heat flux sensors in turbine airfoils

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