os
THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS
345 E. 47th St., New York, N.Y. 10017
The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the
Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published
in an ASME Journal. Authorization to photocopy material for internal or personal use under circumstance not
falling within the fair use provisions of the Copyright Act is granted by ASME to libraries and other users registered with
the Copyright Clearance Center (CCC) Transactional Repotting Service provided that the base fee of $0.30 per page
is paid directly to the CCC, 27 Congress Street, Salem MA 01970. Requests for special permission or bulk reproduction
should be addressed to the ASME Technical Publishing Department.
Copyright ® 1995 by ASME
All Rights Reserved
95-GT-346
Printed in U.S.A_
GAS TEMPERATURE MEASUREMENT
IN THE HOT SECTION OF TURBINE ENGINES
George W. Tregay, Paul R. Calabrese, Mark J. Finney, and Kevin B. Stukey
Conax Buffalo Corporation
Buffalo, New York
ABSTRACT
An optical sensor system extends gas temperature
measurement capability in turbine engines beyond the present
generation of sensor hardware for production engines. The
sensing element incorporates a thermally emissive insert to
generate an optical signal proportional to the gas temperature at
the tip of the probe. The use of a sapphire lightguide allows
operation above the melting point of nickel based alloys.
Sensor development for aircraft turbines has included flight
hardware for use on the Fiber Optic Control System Integration
(FOCSI) Program sponsored by NASA Lewis Research Center.
The optical probe harness measured exhaust gas temperatures
in a General Electric F404 engine. Signals from four probes
were optically combined at a single detector assembly to
determine the average gas temperature. A comparison of
optical and thermocouple temperature measurements was
conducted during a preflight engine test.
The durability of the probe design has been evaluated in an
electric-utility operated gas turbine under the sponsorship of the
Electric Power Research Institute. The temperature probe was
installed between the first stage rotor and second stage nozzle
on a General Electric MS 7001B turbine at Houston Lighting
and Power Company. Two probes have been used in the field
test and they have a combined total of 4660 hours of operation
near 1600°F and 330 starts.
BACKGROUND
There is a continuing demand to increase the efficiency of gas
turbines and reduce their pollution emission. Performance has
improved with higher combustor temperatures and staged
combustion but these improvements require greater control of
the combustion process. Instrumentation is needed to directly
determine gas properties in the turbine region, yet historically
conventional temperature sensors have not had the durability
to survive in the hotter regions.
Designing thermocouple probes for high temperature service
faces difficulties in selecting materials (thermo-elements,
insulation and sheath) and processing techniques that do not
degrade when used above 1000°C (1832°F) in combustion gas
flows. For example, ASTM Committee E-20 (1993)
recommends an upper temperature limit of 982°C (1800°F)
for 24 gauge type K thermocouples in a sheath. This
committee points outs that these limits do not apply to
compacted mineral insulation, but identifies other constraints
such as the severe loss in strength of most potential sheath
materials at 871°C (1600°F). Wang et al (1991) studied the
long term stability of metal sheath thermocouples for turbines,
however their study only included temperatures up to 875°C
(1607°F). The Electric Power Research Institute sponsored
development of instrumentation for the hot section of turbine
engines. For a temperature probe inserted in the second stage
of a utility turbine, Brawley (1989) reports failures
encountered with an RTD probe. The authors of this paper
are not aware of thermocouple or RTD devices being used as
sensors on production turbine engines above 1093 ° C
(2000°F).
Using thermal emission from a hot material is an established
method of temperature measurement. Pyrometers are
commonly employed to measure the temperature of solids and
liquids but are not commonly used for gas temperature
measurement because the spectral emissivity of the gas is
strongly dependent on temperature and pressure (Jamieson et
al., 1963).
The Electric Power Research Institute study (Brawley, 1989)
cited above reported the successful operation of an optical
fiber thermometer immediately upstream of the second stage.
This sensor used a sapphire lightguide to relay thermal
Presented at the International Gas Turbine and Aeroengine Congress and Exposition
Houston, Texas - June 5-8,1995
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
radiation from the coated tip located in the gas stream to
outside the engine case. Details of this approach have been
described previously (Stewart, 1971 and Dils, 1983). The
temperature sensor described in this paper also used a sapphire
lightguide, however, the source of the thermal emission is
embedded inside the end of the sapphire rod to increase
durability. The probe housing uses aluminum oxide ceramics
which maintain most of their strength at temperatures up to
1371°C (2500°F) (Battelle Columbus Division, 1987) to
support the sensing element in the gas stream. The objective of
this ceramic based construction is to extend the measurement
range to values 278°C (500°F) hotter than the present
generation of thermocouple probes.
OPTICAL GAS TEMPERATURE SENSOR SYSTEM
The combination of an optical probe (sensing element
incorporated in a housing), a fiber optic cable, and an electrooptic signal processor constitutes a fiber optic temperature
sensor system. The theoretical basis and operation of the
system has been described previously (Dils, 1983; Tregay et al.,
1991; Finney et al.1992; and Tregay et al., 1993) and only a
brief description is presented here. The sensing element is
heated primarily by convection from the gas and, to a lessor
degree, by radiation from the gas and walls. The relatively
small diameter of the probe combined with the relatively low
thermal conductivity of sapphire causes the temperature of the
sensing element tip to approach the gas temperature. The
thermal radiation (visible and infrared) is produced by the
emissive material embedded in the tip of the sensing element is
in direct proportion to the temperature of the tip. A portion of
this thermal radiation is propagated down the sapphire
lightguide to the outside of the engine case. The optical signal
is then relayed to the electro-optic signal processor with a fiber
optic cable. The temperature of the probe tip is determined
from thermal radiation versus temperature calibration data
stored in the signal processor.
Sensing Element
Sapphire, the single crystal form of aluminum oxide,
possesses a melting point above 2000°C (3632°F), chemical
inertness in combustion flows, and optical transparency from
the near infrared to the ultraviolet. These properties combine
to make it particularly suitable as a lightguide for a combustion
gas temperature sensor. A thermally emissive source was
embedded inside the tip of a sapphire rod. Thermal radiation
generated by the emissive material is internally reflected at the
sapphire to air interface and propagated down to the opposite
end of the sapphire rod. Therefore, the sapphire rod served as
both a lightguide and a protective shroud for the emissive
material.
The sensing element was fabricated with a diameter of 0.06
inches. This small diameter combined with the high flow rate
of gas in the turbine engine results in a rapid response time. A
time constant of 1 second was measured in an atmospheric
combustor rig and projected to be less than 0.5 seconds in the
turbine region of an engine.
Probe Housing
An aluminum oxide tube supported the sensing element in
the hot gas stream; however, the section of the sensing
element with the emissive insert was exposed to achieve a fast
response time. The ceramic tube was supported outside the
hot gas path by a metal shell that attaches to the engine case.
Probes have been fabricated ranging in length from 79 to 340
mm. Vibrational damping techniques were incorporated when
required to reduce the mechanical stresses.
Optical Cable
An optical fiber transmitted the thermal radiation from the
sensing element to the electro-optic signal processor. The
glass fiber had a pure silica core 200 microns in diameter, a
doped silica cladding and a polyimide coating. The fiber was
contained in a protective cable that could be disconnected at
both the probe and signal processor. The cable construction
included layers of teflon and flexible stainless steel conduit to
resist crushing, a 75 pound pull, chaffing, and 278°C (500°F)
operating temperatures.
Signal Processor
The processor determined digital values of the optical
intensity in the spectral range of silicon and germanium
detectors, 400 to 1090 nm and 1090 to 1800 nm, respectively.
The ratio of these two intensities versus temperature was the
basis for the system calibration. The standard calibration
temperature range for the system was between 649°C
(1200°F) and 1371°C (2500°F), but could be extended to
1649°C (3000°F). Measurements from 649°C (1200°F) down
to 427°C (800°F) were possible by using only the germanium
detector signal; however, accuracy was reduced.
Optical Averaging
Multiple thermocouples are commonly employed to measure
an average temperature of the gas stream. Similarly, an
optical system was fabricated with four probes to produce an
optically averaged temperature of the gas stream. Four
optical fibers transferred the signal from each probe and were
terminated into a common connector at the signal processor.
In this arrangement, the detector assembly integrated the
optical signals generated from the four sensing elements and
the signal processor produced an optical average of the
measured gas stream temperature. If the difference in probe
temperature is less than 10°C (18°F) then the optical average
is nearly identical with the arithmetic average. As the
difference between individual probe temperatures increased
then the optical average will be biased to a temperature
slightly higher than the arithmetic average. The bias is
introduced because the thermal radiation increases nonlinearly
with temperature in a way that the hotter probes contribute a
greater proportion of the intensity. When operating in ratio
mode, the failure of an individual probe did not perturb the
system's ability to measure the average temperature of the
remaining probes.
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
High Temperature Stability
The durability of the sensing elements at high temperature has
been studied previously. As reported by Tregay et al. (191)
no significant change in optical intensity was observed after
several hours in a burner rig near 1500°C (2732°F) at Mach .37
and no visual changes were observed after several minutes in a
low velocity flame at 1900°C (3452°F).
Further stability studies have now been conducted. The
sensing element was calibrated for signal ratio versus temperature before and after a 100 hour temperature soak in a furnace
at 1371°C (2500°F). As illustrated in Figure 1, the signal ratio
from the insert was constant within the experimental accuracy.
produced no detectable errors. The components in the signal
processor are sensitive to ambient temperature but with
careful design they can be largely corrected. The residual
error from the signal processor is typically less than ±3°C
(±6°F) over the environmental temperature range of 25°C to
75°C (77°F to 167°F).
With extended operation inside a combustion turbine, the
probes have acquired a light tan deposit (presumed to be iron
oxide). While this could potentially influence the calibration,
no progressive shift in the values of the optical probes relative
to thermocouple at the same location were observed in a
previous study (Tregay et al., 1993). The average difference
between the optical probe and thermocouple in this test was
6°C (10°F).
RATIO
3.0
AIRCRAFT TURBINE APPLICATION
Before
2.5
Development History
After
The optical sensing technique was first tested on an aircraft
engine in 1991. The apparatus has continually improved with
lessons learned during the course of the engine test programs
listed in Table 1. This paper describes the most recent set of
hardware developed for the NASA Fiber Optic Control
System Integration (FOCSI) Program (Baumbick, 1993)
which is demonstrating the use of optical sensors for aircraft
and engine control. Under the propulsion portion of the
contract, General Electric Aircraft Engines has installed nine
optical sensors and electro-optic signal processing on a GE
F404 turbine engine (Poppel, 1994).
2.0
1.5
1.0
0.5
0.0
500
700
900
1100
1300
1500
TEMPERATURE (C)
FIGURE 1. CALIBRATION BEFORE AND AFTER
EXPOSURE TO 1371°C (2500°F) FOR 100 HOURS.
THE ESTIMATED UNCERTAINTY IN THE SIGNAL
RATIO IS ±0.01.
Accuracy
The accuracy of the sensor system has been estimated to be
±18°C (±25°F) relative to a reference thermocouple. The
components affecting accuracy have been discussed previously
(Tregay et al., 1991) and will be only briefly summarized here.
Sensors using optical fiber typically exhibit significant variation
in intensity due to connectors and environmental effects. This
system uses a ratio of intensity and consequently the temperature accuracy is only affected by the amount a parameter
perturbs the relative intensity of the optical signal in the two
spectral regions. Connecting and disconnecting the optical
cable typically influences the value by less than ±2°C (±4°F).
Cable temperatures up to 200°C (392°F) also have a less than
a two degree effect. Cable bends smaller than 50 mm (2 in.)
can influence the temperature by ±6°C (±11°F). Vibrational
acceleration up to 10 g's on the probe, cable and processor
Design Considerations
The optical sensor system, illustrated in Figure 2, was a
functional replacement for the four thermocouple harness used
to measure gas temperature on the lower half of the GE F404
engine. The two short and two long optical probes have the
same form factor as the thermocouple probes they replaced.
The optical cable was as rugged as the wire harness and
mounts with existing clips on the engine case. Four
individual fibers carry the thermal radiation from the probes
to the electro-optic signal processor which employed the
optical averaging described previously.
The size of the electo-optic signal processor module was
reduced to overall dimensions of 51 x 102 x 18 mm. The
calculated temperature was transmitted on a digital interface
as a 14 bit word with the processor's self-checking status
provided on 2 of the bits. One status bit indicated the status
of the operation of the processor itself. A combination of
hardware and software was used to set an error bit if the
signal processor did not cycle through its algorithm in a
specified time interval. The second status bit indicated
condition of the optical signal during ratio mode. After the
temperature is calculated, it can be used to check the
measured intensity against previously stored limits for the
expected intensity. If the intensity has been reduced by a
break in a fiber, for example, then it would be detected and
the optical error bit set.
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
TABLE 1. SENSOR DEVELOPMENT FOR AIRCRAFT TURBINE ENGINES.
TEST CONDITIONS
HARDWARE CONFIGURATION
Pratt & Whitney burner rig at 1500°C (2700°F) and Mach .37
(Tregay et al., 1991; Glomb, 1990).
Sensing element only.
GE F404 engine ground test for 50 hours (Baumbick, 1993;
Tregay et al., 1991).
Four free standing probes located at the turbine exit with
optical cable mounted on engine case. Signal processor
located off engine.
Pratt & Whitney F100 engine flight test in F-15 HIDEC
testbed aircraft at NASA Dryden Flight Research Facility for
numerous flights with temperature data recorded during
eight flights (Finney et al., 1992).
Single probe located inside the leading vane of the second
stage turbine with optical cable to a signal processor located
in a compartment adjacent to the engine.
GE F404 engine ground test for 250 hours (Tregay et al.,
1994).
Four free-standing probes located at the turbine exit with
optical cable mounted on engine case.
Signal processor located off engine.
Pratt & Whitney Joint Technology Demonstrator Engine
XTC65-2 testing (Spillman et al., 1994)
Sensing element replaces thermocouple in exhaust gas
sensor housing.
GE F404 engine ground test for equipment check out
(Tregay et al., 1994).
Four free-standing probes located at the turbine exit with
optical cable mounted on engine case.
Miniature signal processor located in electro-optics chassis
which mounts on engine during flight.
Flight tests conducted on F-18 System Research Aircraft at
NASA Dryden Flight Research Facility.
conducted separately for each component. The two optical
probe lengths resulted in two distinct resonance frequencies.
The fiber optic cable was supported by engine mounting clips
and mated to the optical probes and the electro-optic signal
processor during vibration testing. The signal processor was
powered during all testing, and no anomalies were observed in
the digital output signal.
Ground Test on F404 Aircraft Engine
Three complete sensor systems were fabricated for the
NASA FOCSI Program and shipped to GE Aircraft Engines.
The exhaust gas sensor hardware was mounted on a GE F404400 engine and ground tested at GE Flight Test Operations,
Edwards, CA. A series of tests were conducted on November
8-10, 1993, April 5-8, 1994, and April 19, 1994.
The thermocouple harness with four probes equally spaced
on the lower half of the engine was removed and replaced
with the optical sensor harness with four probes. The
optically averaged temperature measured with the four optical
probes was compared with the average temperature measured
by the four thermocouple probes that remained on the upper
half of the engine. The time history of an excursion from idle
to full power and back to idle is presented in Figure 3. These
measurements were in close agreement during steady state
with small differences observed during transients.
FIGURE 2. OPTICAL SENSOR HARDWARE FOR GE
F404 AIRCRAFT TURBINE.
Environmental Testing
The components of this optically averaged FOCSI system
(four probes, four-fiber cable and the electro-optic signal
processor) passed the environmental testing listed in Table 2 to
insure flight worthiness. Resonant vibration dwell tests were
4
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
TABLE 2. ENVIRONMENTAL TESTING FOR AIRCRAFT TURBINE.
Test
Parameter
Optical Probe
(39 mm Length)
Optical Probe
(147 mm Length)
Fiber Optic
Cable/Harness
Electro-Optic
Signal Processor
Resonance
Vibration Dwell
107 cycles at 1572 Hz
at 20 g's excitation
4 Hours at 589 Hz at
20 g's excitation
4 Hours at 21 Hz and
48 Hz at 0.05" peak
displacement
107 cycles at 1556 and
2000 Hz at 10 g's
excitation
Temperature
Cycle
(30 min/cycle)
<43 to >871 °C
(<200 to >1600°F)
25 cycles
<43 to >871 °C
(<200 to >1600°F)
25 cycles
-55 to 254°C
(-67 to 490°F)
25 cycles
-55 to 91°C
(-67 to 195°F)
25 cycles
Thermal Soak
(24 hours)
>816°C (>1500°F)
>816°C (>1500°F)
254°C (490°F)
88°C (190°F)
measure gas temperature in the turbine region of a General
Electric MS 7001B turbine. The probe was designed for
installation at the entrance to the second stage turbine using an
existing borescope port in the engine case and is illustrated in
Figure 4. For this application, a layer of high temperature
ceramic material is located between the sensing element and
alumina tube and between the alumina tube and the metal
tube. The purpose of these layers was to moderate the effect
of the turbine case vibration on the ceramic components.
The probe assembly was supported by a sealing gland at the
turbine casing and also by the shroud block within the turbine.
The metal portion of the assembly was designed to end
approximately 13 mm inside the shroud block. The combined
length of the ceramic support tube and the sensing element
reached approximately 38 mm into the hot gas stream.
TEMPERATURE (C)
1000
900
Thermocouple
800
700
600
500
400
---------: -
900
20
40
60
80
TIME (s)
FIGURE 3. TEMPERATURES MEASURED AT THE
TURBINE EXIT OF A GE F404 AIRCRAFT ENGINE.
THE ESTIMATED UNCERTAINTY FOR THE OPTICAL
SENSOR SYSTEM IS ±18°C (±25°F).
UTILITY TURBINE APPLICATION
Design Considerations
In 1983, the Electric Power Research Institute (EPRI) began a
research project to develop instrumentation for observing and
characterizing the hot sections of operating gas turbines
(Brawley, 1989). The temperatures measured by an optical
fiber thermometer provided useful information about transient
conditions in the turbine. EPRI decided to continue
development but with an emphasis on reducing risk to the
turbine. Specific objectives were the elimination of the water
cooling and metal parts in the gas path.
This paper describes the optical temperature system used to
Environmental Testing
Vibration, thermal shock, pressure, and frangibility tests
were conducted to experimentally demonstrate the ability of
the probe assembly to withstand the turbine environment.
This testing is summarized in Table 3. The vibration tests
were conducted with the probe mounted in a fixture that
simulated the turbine borescope port geometry. After
determining the resonant frequencies, the probe was subjected
to a resonance dwell test of 10 7 cycles at its first natural
frequency. The response level was continuously monitored
during the dwell test to ensure maximum stress condition. A
shift from 480 Hz to approximately 535 Hz was observed
after a "break-in" period. No significant changes in frequency
or response acceleration level were observed during the
remainder of the test. The unchanged frequency,
combined with a post-test visual inspection, indicated the
probe survived resonant dwell test undamaged.
The probe was then vibrated with the probe tip heated to
above 1232°C (2250°F). There was no significant difference
between hot and cold response levels and no observable
degradation of the probe. Stress at the base of the probe was
also measured and was found to be at relatively low levels,
approximately 25 MPa (3600 psi).
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
TABLE 3. ENVIRONMENTAL TESTING FOR UTILITY TURBINE.
TYPE OF TEST AND CONDITIONS
DURATION/RESULTS
Resonance vibration search from 10 to 2000 Hz. Excitation level of
5 g's, with the displacement limited to 0.050 inch peak-to-peak at
low frequencies.
First natural frequency 480 Hz.
Second natural frequency 1660 Hz.
Resonance dwell at 499-535 Hz with excitation level of 5 g's.
10 cycles.
The probe was vibrated with the distal end of the alumina
tube heated to approximately 982°C (1800°F).
Two 20-minute sweeps from 10-2000 Hz.
Approximately 1 inch of the probe tip was heated to
temperatures of 1232°C (2250°F) then removed to cool at room
ambient.
20 cycles.
Probe was pressurized with nitrogen and submerged in water.
No leakage was observed (i.e. bubbles) over 10
minutes.
FIGURE 4. PROBE FOR THE SECOND STAGE OF A
GE MS 7001B TURBINE.
Utility Field Test
The sensor system was installed at Houston Lighting and
Power Company on Unit Number 41. Unit 41 operated on
natural gas throughout this period. A single temperature probe
was installed between the first and second stages of the GE MS
7001B turbine using the borescope port in shroud block 23.
Approximately 2.4 m of the high temperature cable was located
inside the turbine compartment. A 24 m optical jumper cable
transmitted the signal through the remaining distance to the
control cab where the signal processing and data acquisition
equipment was located. The computer was programmed to
record temperature measurements for a five second period at
fifteen minute intervals.
On the first day of testing, the turbine ran for approximately
nine hours. During the initial portion of the test, measurements
-
-
were recorded for both the optical probe and the exhaust gas
thermocouple. The two probes exhibited a similar temporal
behavior but differed in the magnitude of the temperatures
illustrated in Figure 5. The difference in temperature is to be
expected since the optical probe was located upstream of the
second stage whereas the thermo-couple probes were located
down stream of the third stage. Note that the optical probe at
the second stage entrance measured a temperature overshoot
during the start that did not appear in the exhaust temperature
measurement. The optical probe survived the trip and restart.
The first optical probe was installed continuously from May
17, 1993 through February 15, 1994. The temperatures
measured with this probe after 1100 hours of operation are
illustrated in Figure 6. The probe was removed from the
turbine after 2200 hours and 160 starts. The probe, cable and
signal processor were checked and found to be functional.
Both the sapphire rod and alumina support tube showed no
visual indication of material loss. The optical cable also
appeared in satisfactory condition with the exception of the
connector at the probe end. The fiber had "pistoned" and was
extending out of the connector by approximately 0.02 inches.
Since the radiation from the sapphire rod overfills the optical
fiber, it is believed that the fiber movement did not affect the
indicated temperatures.
A second probe was installed on February 15, 1994 and
removed on November 30, 1994 after approximately 2460
hours of operation and 170 starts. Visually the sensing
element and ceramic support tube appeared in good condition.
SUMMARY
The temperature sensor described in this paper was
developed specifically to measure gas temperature in the
combustor and turbine sections of the engine. The
components exposed to the gas path were ceramic to achieve
greater strength than metal above 1093°C (2000°F).
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
TEMPERATURE (C)
1000
TEMPERATURE (C)
1000 r
Optical Probe Located
Upstream of Second Stage
900
900
800
800
700
Thermocopule Probe
Located at Exhaust
400
300
700
■A
500
4■
■■
800
Turbine Trip
and Rest
■
500
200
10
20
30
0
40
50
60
20
TIME (Minutes)
FIGURE 5. TEMPERATURES MEASURED WITH
OPTICAL PROBE IN THE SECOND STAGE AND
THERMOCOUPLE PROBES IN TURBINE EXHAUST.
THE ESTIMATED UNCERTAINTY FOR THE OPTICAL
PROBES IS ±18°C (±25°F).
The sensing element employed the thermal radiation naturally
generated at these temperatures. A sapphire lightguide with an
emissive insert provided a sensing element durable in a
combustion flow for short durations at 1900°C (3452°F), yet
small enough to achieve response time of less than one second.
The calibration stability of the sensing element has been
demonstrated over a 100 hour period in a furnace at 1371°C
(2500°F).
For aircraft engines, the sensor system was configured as a
multiprobe harness to determine average gas temperature.
Probe and cable durability have been demonstrated by vibration
testing at an excitation of 20 g's and included a dwell of 10'
cycles at the probe resonance. The signal processor has been
packaged in a 51 x 102 x 18 mm module to demonstrate the
miniaturization necessary for incorporation into an electronic
engine control. The electro-optic circuitry operated over an
environmental temperature range of -55 to 91°C (-67 to 195°F)
and was suitable for mounting on the engine. The optical
sensor system has demonstrated good agreement with
thermocouple probes on the opposite half of the GE F404
engine.
For utility turbines, the sensor system was installed between
the first and second turbine stage of a GE MS 7001B engine.
The probe penetrated 38 mm into the flow stream and
experienced the buffeting of the blade passing immediately
upstream.
40
60
TIME (Hours)
FIGURE 6. TEMPERATURE MEASUREMENTS AFTER
APPROXIMATELY 1100 HOURS OF OPERATION.
Temperature overshoots on start up were observed which
were not evident further downstream at the exhaust gas
temperature sensors. Two probes reached a combined
operating time of 4200 hours of operation near 871°C
(1600°F).
ACKNOWLEDGEMENTS
This work was financially supported by Electric Power
Research Institute, GE Aircraft Engines IR&D Program, a
subcontract under the NASA FOCSI Program, and Conax
Buffalo Corporation.
The authors wish to thank Gary Poppel, GE Aircraft
Engines, and the personnel at Houston Lighting and Power's
T.H. Wharton Plant for their help in conducting the test
programs.
REFERENCES
ASTM Committee E20 on Temperature Measurement and
Subcommittee E20.04 on Thermocouples, 1993, Manual on
the Use of Thermocouples in Temperature Measurement,
ASTM Special Technical Publication 470B, Chapter 3.
Battelle Columbus Division, 1987, "Engineering Property
Data on Selected Ceramics Vol. 3 Single Oxides," Metals and
Ceramics Information Center, MCIC-HB-07-Vol. 3, reprint.
Baumbick, R. J., 1993, Status of Fiber Optics Control
System Integration (FOCSI) Program, NASA TM 106151.
Brawley, G. H., 1989, "Gas Turbine Characterization and
Diagnostic Instrumentation Evaluation," EPRI Report AP6022.
Dils, R. R., 1983, "High Temperature Optical Fiber
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term
Thermometer," J. of Applied Physics, Vol. 54, pp. 1198-1201.
Finney, M. J., Tregay, G. W. and Calabrese, P R., 1992,
"Flight Testing a Fiber Optic Temperature Sensor," SPIE Vol.
1799, S pecialty Fiber Optics Systems for Mobile Platforms and
Plastic Optical Fibers, pp. 194-203.
Glomb, W. L., Jr., 1990, "Fly-by-Light Schemes Move into
Demonstration Stage," Laser Focus World, pp. 167-176.
Jamieson, J. A., McFee, R. H., Plass, G. N., Grube, R. H.,
and Richards, R. G., 1963, Infrared Physics and Engineering,
McGraw-Hill Book Co., Chapter 2.
Poppel, G. L., 1994, Fiber Optic (Flight Quality) Sensors for
Advanced Aircraft Propulsion , NASA CR-191195.
Powel, F. S., IV, 1990, "On the Leading Edge: Combining
Maturity and Advanced Technology on the F404 Turbofan
Engine," ASME Paper 90-GT 149.
Spillman, W. B., Jr., LaClair, R. D., Birdsall, J. C., Berthold,
J. W., III, and Tilstra, S. D., 1994, "Fiber Optic Sensors for
Fly-by-Light On-Engine Monitoring," SPIE Vol. 2295, Fly-byLight , pp. 142-155.
Stewart, J. E., Evans, D. E., and Larson, G. L., 1971,
"Remote Radiation Temperature Sensor," U.S. Patent
3,626,758.
Tregay, G. W., 1988, "Optical Fiber Temperature Sensor,"
U.S. Patent 4,794,619.
Tregay, G. W., Calabrese, P. R., Kaplin, P. L., and Finney,
M. J., 1991, "Optical Fiber Sensor for Temperature
Measurement from 600 to 1900°C in Gas Turbine Engines,"
SPIE Vol. 1589, Specialty Fiber Optics Systems for Mobile
Platforms, pp. 38-47.
Tregay, G. W., Calabrese, P. R., Finney, M. J., and Stukey,
K. B., 1994, "Durable Fiber Optic Sensor for Gas Temperature
Measurement in the Hot Section of Turbine Engines," SPIE
Vol. 2295, Fly-by-Light, pp. 156-163.
Wang, T. P., Wells, A. and Bediones, D., 1991, "5,000 Hour
Stability Tests of Metal Sheathed Thermocouples at Respective
Temperatures of 538°C and 875°C," ASME Paper 91-GT 182.
8
Downloaded From: http://asmedigitalcollection.asme.org/pdfaccess.ashx?url=/data/conferences/asmep/82304/ on 06/16/2017 Terms of Use: http://www.asme.org/about-asme/term