Axial Compressor Deterioration Caused by Salt Ingestion

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Elisabet SyverudDepartment of Energy and Process Engineering,

NTNU,orwegian University of Science and Technology,

Trondheim, Norway

Olaf BrekkeRoyal Norwegian Navy,

Bergen, Norway

Lars E. BakkenDepartment of Energy and Process Engineering,

NTNU,orwegian University of Science and Technology,

Trondheim, Norway

Axial Compressor DeteriorationCaused by Saltwater IngestionGas turbine performance deterioration can be a major economic factor. An example iswithin offshore installations where a degradation of gas turbine performance can mean areduction of oil and gas production. This paper describes the test results from a series ofaccelerated deterioration tests on a General Electric J85-13 jet engine. The axial com-pressor was deteriorated by spraying atomized droplets of saltwater into the engineintake. The paper presents the overall engine performance deterioration as well as de-teriorated stage characteristics. The results of laboratory analysis of the salt deposits arepresented, providing insight into the increased surface roughness and the deposit thick-ness and distribution. The test data show good agreement with published stage charac-teristics and give valuable information regarding stage-by-stage performancedeterioration. �DOI: 10.1115/1.2219763�

ntroductionCompressor fouling causes 70–85% of the performance loss

ue to deterioration in gas turbines �1�. The degradation rate wille affected by site-specific conditions of the fouling and by engineperating conditions. A thorough review of engine performanceeterioration mechanisms and modeling is given by Kurz andrun �2�.From operating experience and literature the first stages of an

xial compressor are expected to be the ones most heavily fouled.eposits will have different characteristics depending on the na-

ure of the fouling. Dry particles in dry atmospheres are likely toeposit in different areas compared to sticky matters and oilyompounds. At high inlet humidity, the drop in static pressureuring acceleration will increase the dust adhesion on the blades3,4�.

To understand and reveal the degradation and restorationechanisms related to recoverable compressor deterioration, a

ystematic test campaign has been performed on a General Elec-ric �GE� J85-13 jet engine. The overall performance of the gasurbine and compressor stage characteristics were analyzed in de-ail. The tests reflect actual performance deterioration as found inffshore applications and provide valuable information regardingverall and stage-by-stage compressor performance deteriorationn a real engine.

The tests were part of a larger online water wash test program.he results from this program are reported in a different paper �5�hich should be seen in context with the present work.

est Facilities and Engine DescriptionThe GE J85-13 engine has been a successful engine in different

pplications since it first entered service in 1960. More than 6000f these engines are still in service worldwide. The Royal Norwe-ian Air Force �RNoAF� F-5 Freedom Fighters have two GE85-13 engines for propulsion. Because these aircrafts have beeneplaced by the F-16 aircraft, one GE J85-13 engine has beenade available to the Norwegian University of Science and Tech-

ology for use in this project. Engine testing was carried out at theNoAF’s test facilities at Kjeller, Norway. Kjeller is located in-

and, �25 km east of Oslo at an altitude of 119 m above sea level.The GE J85-13 is a compact, light-weight, single-spool turbojet

ngine. It has an eight-stage axial-flow compressor that is directly

Contributed by the International Gas Turbine Institute �IGTI� of ASME for pub-ication in the JOURNAL OF TURBOMACHINERY. Manuscript received October 1, 2004;nal manuscript received February 1, 2005. IGTI Review Chair: K. C. Hall. Paperresented at the ASME Turbo Expo 2005: Land, Sea and Air, Reno, NV, June 6–9,
005, Paper No. GT2005-68701.
ournal of Turbomachinery Copyright © 20

aded 03 Sep 2011 to 138.250.84.31. Redistribution subject to ASME

coupled to a two-stage turbine. The exhaust nozzle has a variablethroat area, and the engine is equipped with an afterburner. TheGE J85-13 is rated to deliver a minimum thrust of 2720 lb�12.1 kN� at 100% shaft speed without the afterburner and a mini-mum thrust of 4080 lb �18.1 kN� with full afterburner.

The eight-stage compressor has a pressure ratio of 6.5:1. First-and second-stage rotor blades are coated for corrosion protection.To handle off-design operation, the compressor has variable inletguide vanes �IGV� and bleed-off-valves at stages 3–5, as well as avariable exhaust nozzle. The variable geometry is controlled bythe throttle angle, but the timing is ambient temperature biased. AtISO conditions, IGV will be at maximum deflection and bleed-off-valves will be fully open below �81% corrected speed andfully closed at �96% corrected speed. At ambient temperaturesabove ISO conditions, the bleed-off-valves will close at lowerspeed settings. The nozzle is controlled by the throttle as long asthe engine is running below the exhaust temperature limit. Whenthe maximum exhaust temperature is reached, the exhaust nozzlewill open to increase the exhaust flow. This reversal in the exhaustnozzle schedule occurs close to the maximum throttle angle.

Bleed air is also used for internal cooling of the afterburneractuators, combustor liner, and turbine guide vanes. Bleed air foranti-ice is taken at the compressor discharge. The anti-ice airvalve was kept closed during all tests associated with this project.

Engine Model Data. Detailed knowledge of compressor stagegeometry is available through several sources providing a goodbasis for developing engine simulation models.

NASA carried out several tests on the GE J85-13 in the late1960s and throughout the 1970s, and reported compressor stagetemperature and pressure profiles �6�, individual compressor stagecharacteristics �7,8� as well as gas path geometry �9�.

Initial Engine Condition. The engine used in this test programwas preserved and stored after serving �300 h onboard an air-craft. The engine was not overhauled prior to testing.

To avoid disturbances to the airflow attributable to loose coat-ing, the coating on the second-stage rotor blades was removed byhand prior to the testing, using emery paper grade 320. The com-pressor inlet, inlet guide vanes, first-stage rotor, and stator werecleaned by hand using solvent, while the upper half of the casing,including stator vanes from stages 1 to 7 were cleaned using anultrasonic cleaning bath. In addition, online compressor cleaningwas performed prior to the testing using water at a flow of13.2 l /min, for a 30 second period at idle, 70%, 90%, 95%, and100% corrected engine speed. Details of the online water wash
system are given in a different paper �5�.
JANUARY 2007, Vol. 129 / 11907 by ASME

license or copyright; see http://www.asme.org/terms/Terms_Use.cfm

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ngine InstrumentationAdditional sensors were installed to provide more information

han is available from standard, test-cell instrumentation. The en-ine instrumentation is shown in Fig. 1.

The temperatures at stages 1, 3, 4, 6, and 8 were measuredsing a single resistance temperature detector �RTD� at each statorow. The entire 15 mm sensor length was immersed into the air-ow, giving a representative measure of the gas path bulk average

emperature. Because the sensors are unshielded, the velocity er-or will be significant in the temperature reading. The velocityrror was calculated with a temperature recovery factor of 0.6510�.

The static pressures at stage 5 and at compressor dischargeere measured at a single point on the circumference. The stage 5

tatic pressure was measured in the bleed channel.Engine throttle and nozzle position were recorded manually at

ach setting. Relative humidity and ambient temperature were re-orded manually throughout the testing and measured at the sameocation outside of the test-cell intake. Due to test-cell recircula-ion, the recorded compressor inlet total temperature �CIT� waslightly higher than the measured ambient temperature. The CITas measured using four sensors located at the engine inlet

creen.All instruments were calibrated prior to the test program, and

he measurement uncertainties were calculated based on methodsiven in the ASME performance test codes �11,12�.

To reduce the data scatter, a minimum of 60 data points �2 Hzampling rate� were collected at each setting. The average of theseeadings was taken as the steady-state data point.

ngine Degradation Method and EquipmentAccelerated engine deterioration was done through the inges-

ion of atomized saltwater. A similar method has been used by thenited States Navy �13�. The saltwater ingestion system is illus-

rated in Fig. 2. Saltwater was atomized in front of the compressornlet using four atomizing nozzles installed on a bracket on the

Fig. 1 GE J85-13 engine instrumentation

Fig. 2 Salt ingestion system

20 / Vol. 129, JANUARY 2007


online water wash manifold positioned 0.77 m in front of thecompressor inlet. This ensures complete atomization of the salt-water prior to the inlet guide vanes. The saltwater was gravity fedfrom a storage tank while air pressure was taken from the pres-surized air supply in the test cell.

The droplet diameter of the saltwater spray was measured as23 �m volume median diameter �VMD� in still air using a Mal-vern laser diffraction spray analyzer and applying the ASTM Stan-dard for calculating spray characteristics �14�. The measurementswere taken at the center of the fully developed spray at a distanceof 44 cm from the nozzles. The 23 �mVMD is higher than the�8 �m droplet size indicated as typical values after the intakefilters of marine gas turbine units �15�.

For all salt-deterioration tests, 3 l of saltwater were ingested ata flow rate of 0.2 l /min. The salt concentrations in the initial testswere 0.2% and 0.4% �by weight�. For the accelerated deteriorationtests associated with the online water wash trials, 3 l of saltwaterwas ingested with a salt concentration of 1% �by weight�. The saltused in the tests constituted �by weight� 99.83% sodium chloride,0.04% calcium, 0.03% magnesium, and 0.1% sulphione.

Test ProcedureThe engine base lines were established for steady-state opera-

tion at 12 operating points from 60% corrected speed to full load.Prior to establishing each baseline, the engine was run for 5 minat full load. Engine speed was then reduced to idle before beingincreased to the initial throttle setting. At each throttle setting, theengine was allowed to stabilize for 1 min before reading 60 datapoints. To prevent impact of hysteresis from the instrument orcontrol system, all throttle settings were established at increasingengine speeds. Engine base lines were recorded prior to degrada-tion, after salt degradation, and after each online water wash.

All salt degradation trials were run at 10 deg throttle angle�equivalent to 97.5% engine shaft speed� at constant nozzle posi-tion and with closed bleed-off-valves and fully open IGV.

To completely clean the compressor of salt deposits, a 5 minonline water wash was done at engine speeds varying from fullspeed to idle for two complete cycles using 17.6 l of water perminute �droplet size of 200 �mVMD�.

Salt DepositsThe salt deposits on the axial compressor were analyzed non-

destructively. Deposits on the stator vanes were analyzed at alaboratory. The deposits on the rotor blades were visually evalu-ated, because the rotor assembly could not be moved to the labo-ratory within the available timeframe.

Location of Deposits. The salt deposits were mainly foundalong the leading edge of the first four stages and on the pressureside of the stator vanes along the hub. The heaviest deposits werefound along the first-stage annulus, at the leading edge of thesecond-stage stator vanes and along the hub at the pressure side ofthe second-and third-stage blades and vanes.

Compared to the stator vane deposits, the deposits on the rotorblades were significantly smaller. The rotor blade deposits weremainly on stages 2–4. Deposits of a coarse structure were foundalong the hub, while deposits of a finer structure were found alongthe annulus. These deposits covered 35% of the second-stagepressure side and 20% of the third-stage pressure side.

For the fourth stage, the deposits were different from the pre-vious stages, occurring as a shimmering layer covering the entirepressure side surface, except for the 20% closest to the hub.Suction-side deposits were virtually nonexisting, except for sometraces found along the separation line on the first- and second-stage blades. Leading-edge deposits were found on stages 2–4.

Only immeasurable traces of salt �streaks� were found on thesuction side of the fifth- and sixth-stage stator vanes and along theleading edge of the fifth-stage rotor blades. The spanwise distance
between the streaks varied from 0.11 mm on the fifth-stage suc-
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ion side to 0.17 and 0.88 mm on the sixth-stage suction and pres-ure sides, respectively. No salt deposits were found on the sixth-tage rotor or on stages 7 and 8.

The salt deposits on the stator vanes at the top half of theompressor were measured by dissolving the salt in water andeasuring conductivity and chloride content. The salt contenteasured by conductivity correlated well with the salt contenteasured by chloride content. Chloride content was transformed

o a distribution of the salt deposits on the stator vanes as shownn Fig. 3.

In Fig. 3, the data are compared to results of an investigation ofompressor blade contamination carried out on the 16-stage com-ressor of the Nuevo Pignone MS5322R �16�. The deposits areiven as a percentage distribution to outbalance differences in theumber of vanes measured. The salt deposits measured for the GE85-13 include the deposits along the annulus; this is not the caseor the data obtained for the MS5322R.

In the MS5322R, the rotor and stator blades of the first stagead more deposits on the suction side. The deposited masses onhe blades aft of the first stage were approximately equal for thewo sides. More deposits were found on the stator blades than onhe rotor blades. It is suggested that this is due to a certain clean-ng effect on rotor blades from the influence of centrifugal forcesn the dirt particles �16�. The findings for the MS5322R corre-pond to the findings of this project, except that the salt depositsn the present study were mainly on the pressure side of the statoranes.

Deposit Thickness. The salt deposits were analyzed using labo-atory electron microscopes. Figure 4 shows the leading-edge de-osits close to the hub on the second-stage stator vanes. Theeavy leading-edge deposits are probably caused by the constant

Fig. 3 Deposit distribution on stator vanes

ig. 4 Salt deposits at the leading edge of the second-stage
tator vanes „at 6.5Ã magnification…
ournal of Turbomachinery


shaft speed during salt ingestion. The buildup is strong enough towithstand the increased turbulence due to changing incidence atlower speed settings, since only parts of the deposits were brokenoff by the airflow although the engine has been run at severalspeeds after salt deterioration. The leading-edge deposits weremeasured using the microscope and the results are given in Table1.

Mean Line Salt Crystal Characteristics. The grain size andsurface structure of the salt crystals along the mean line of the firstthree stators were analyzed from Struers replicas. Struers replicasenable three-dimensional replicas of material microstructure witha resolution as low as 0.1 �m �17�. In addition, the replicas col-lect loose surface particles that can be further analyzed with amicroscope.

Figure 5 shows the salt grains found on a 330�240 �m surfaceon the pressure side of the first-stage stator. In general, the suctionside had smaller, more homogenous grain sizes. The salt crystalson the pressure side were larger and more distributed; for stages 2and 3 they were caught in a thin film of liquid salt flowing in thedirection of the air flow.

The technical roughness is influenced both by the size andshape of the salt crystals as well as by the distribution of saltcrystals along the blade. The authors are fully aware of the diffi-culties involved when measuring technical roughness. The estima-tion of a technical roughness level from the two-dimensional viewgiven in Fig. 5 will only be approximate.

For all stages, the arithmetic mean diameter of the salt grains inthe sample is considered a measure of the grain size, k. The aver-age distance between the salt crystals is calculated based on thenumber of crystals found in the sample. The associated equivalentsand roughness, ks, is estimated based on the average distancebetween the salt grains and by comparing the ratio of grain size tograin spacing with the associated equivalent sand roughness forregularly arranged spheres �18,19�. The relative roughness is theratio of equivalent sand roughness to the blade chord length. Theresults are given in Table 2. As expected, the largest relativeroughness is found on the pressure side.

The small chord lengths of the GE J85-13 compressor make theengine sensitive to small grain sizes �in terms of relative rough-ness�. The equivalent sand roughness levels given in Table 2 im-ply less roughness than from a surface of emery grade 320 �20�.

Table 1 Leading-edge deposit thickness

Stator stage No. Leading-edge salt deposits

1 No deposits at leading edge2 500 �m at hub, 25 �m at annulus3 200 �m at hub, no deposits along annulus4 125 �m at hub, 25 �m at annulus

5–8 No deposits

Fig. 5 Salt grains on first-stage stator vane pressure side
along mean line „330Ã240 �m picture…
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ignificance of Salt Deposits on Compressor Perfor-ance CharacteristicsThe effect of Reynolds number on compressor performance is

ell known for compressors in new and clean conditions. TheSME test code on compressors �21� does not give any attention

o testing in deteriorated conditions. A study of Reynolds numberffects in centrifugal compressors shows how efficiency can beorrelated to increases in the friction factor with increased surfaceoughness �22�. Some research exists on the effect of increasedoughness in axial compressors: A study of profile loss due tourface roughness on NACA 65-12�06� profiles showed a 2% in-rease in the profile losses at a relative roughness �ks/chord� ofE-4 and a 10% increase at 50E-4 �2�. Another study showed a% reduction in the static pressure ratio at a relative roughness of0E-4, but the static pressure ratio is a strong function of theolume flow and blade velocity �20�. The effect of surface rough-ess on blade profile loss is quantified through a momentum thick-ess correlation �23�. Additional insight is given into the boundaryayer characterization of rough surfaces, although applied to newnd clean surfaces �19�.

The critical roughness Reynolds number, defined in Eq. �1�19�, is independent of the characteristic length of the compressor

Recritical =ksV

��1�

When the critical Reynolds number is above 90, the flow can bessumed hydrodynamically rough with turbulent attached flowi.e., with the salt crystals protruding into the laminar sublayer�.or these cases, the associated blade profile loss will dependolely on the surface roughness to chord level, and not on thehord Reynolds number �19�.

Applying Eq. �1� to the GE J85-13, a critical Reynolds numberf 90 implies a maximum equivalent sand roughness of 5 �m.able 2 shows the salt crystals have equivalent sand roughness

evels above 5 �m for all stages except the suction side of stagehree stator vanes. This implies that the blade loss, hence, thefficiency, of the deteriorated engine will be dependent on theoughness level and not on the chord Reynolds number. Thehange in stage efficiency should therefore be possible to calcu-ate based on the correlation for momentum thickness with surfaceoughness �23�.

Another aspect of the increased surface roughness is the changen deviation angle due to the thickening of the blade boundaryayer. A correlation for changes in the deviation angle due to sur-ace roughness is found from tests of a helicopter engine after500 h of flight �24�. The relative roughness levels reported weren the range of 10.9E-3 in the first stage to 0.7�10E-3 in the lasttage and at Reynolds numbers of 240,000. Typical Reynoldsumbers in the GE J85-13 are 200,000 to 600,000; however, theelative roughness level given in Table 2 is only one tenth of theevel reported for the helicopter engine �24�. Therefore, changes tohe deviation angle due to deterioration are considered negligible

Table 2 Salt crystal grain size and distribution at mean line

tatorvane

Mean grainsize, k��m�

Mean graindistance

��m�Equivalent sand roughness,

ks ��m�

Relativeroughness,ks / chord

1 PS 22 88 25 11E-41 SS 14 88 10 4E-42 PS 16 62 20 11E-42 SS 8 25 15 8E-43 PS 15 65 15 10E-43 SS 5 40 2 0.6E-4

or the salt-deteriorated GE J85-13.

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Engine Performance DeteriorationThe engine deterioration due to salt deposits is analyzed in the

following section. All overall engine performance parameters pre-sented are corrected to ISO reference conditions �25�. Variationsin humidity were not accounted for in the calculations. Datapoints measured at low speed settings were not included, becausethese include large unaccounted bleed flows.

Initial Salt Deterioration Trials. Figure 6 shows the resultsfrom the initial salt-deterioration testing, where salt was injectedin three separate tests without water wash in between. The figureshows the total mass of salt ingested in the tests. The overallengine performance deterioration is shown for intake depression,which is defined as the difference between the total and staticpressure in the bellmouth throat. The intake depression deterio-rates rapidly and is outside of the measurement uncertainty toler-ance even after 6 g of salt ingestion. The intake depression wasthe most sensitive parameter to compressor deterioration and wasfound to be the best method for condition monitoring of the en-gine, because it is less dependent of the engine control mode.

Overall Engine Performance Deterioration. The engine per-formance deterioration with 30 g of salt ingested is shown in Figs.7–10. For simplicity in the presentation of data, only two repre-sentative cases are shown. These cases were recorded on the sameday at identical compressor variable geometry and nozzle posi-tions. Because the testing was done at ambient temperatures abovethe ISO reference condition, the 100% corrected speed setting wasnever reached. At these high ambient temperatures, the bleed-offvalves are fully closed above 95% corrected speed.

Figure 7 shows the engine performance deterioration in termsof changes on the compressor equilibrium operating line. Thestatic-to-total compressor pressure ratio is given at stage 5 �station2.5� and at compressor discharge �station 3�; both parameters are

Fig. 6 Deterioration of intake depression

Fig. 7 Effect of deterioration on compressor operating line



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elated to the total inlet pressure. Deterioration shifts the compres-or operating line to a lower flow rate and a lower pressure ratio.imilar percentage changes to the compressor pressure ratio areound at both stations, indicating that the pressure deteriorationas at the front end of the compressor. Front stage fouling isnown to cause larger changes in the flow rate than in the pressureapability �3�, and the results of Fig. 7 are in agreement with this.

The surge line is not known for the engine; however, surge wasot encountered at any time during the test.

Figure 8 shows the change in exhaust gas temperature withhrust. The exhaust gas temperature is approximately nine percent-ge points higher for the deteriorated engine. The flattening of theeteriorated engine curve at the high end is due to exhaust gasemperature limitations in the engine control.

The variation in compressor isentropic efficiency with correctedngine shaft speed is shown in Fig. 9. This figure shows that

Fig. 8 Change in exhaust gas temperature with thrust

Fig. 9 Deterioration of compressor isentropic efficiency

Fig. 10 Deterioration of compressor mass flow



deterioration reduces the isentropic efficiency by approximatelythree percentage points along the operating line, and reduces theengine operating speed. In addition, the isentropic efficiency curveis flattened when the maximum exhaust gas temperature isreached.

The variation in compressor mass flow rate with engine shaftspeed is shown in Fig. 10. This figure shows the significant reduc-tion in air flow rate for the deteriorated engine. Again, the dete-riorated engine performance is shifted to a lower shaft speed;however, the mass flow rate is only slightly affected by the ex-haust gas temperature limitation. Mass flow rate is directly relatedto the intake depression which was found as the best parameter forengine condition monitoring in the initial salt deterioration trials.

Stage Work Coefficient DeteriorationThe compressor temperature measurements allow for a close

look at the stage work coefficient during salt deterioration. Thepresented test data are limited to engine operation above 95%engine shaft speed where the engine bleed-off-valves were closedand the IGV fully open.

The following assumptions are made when analyzing the testdata:

• The temperatures were measured at stages 1, 3, 4, 6, and 8,and a linear temperature distribution was used for the re-maining stages. The stage temperatures were measured inthe stator; however, the measurements were assumed to bevalid at the stator outlet plane.

• The temperature dependence of the specific heat at constantpressure and the ratio of specific heats are defined by afifth-order polynomial correlation for dry air �26� and aseventh-order polynomial for water vapor �25�. These poly-nomial correlations fit original data within 0.1% at the tem-perature range applicable to the GE J85-13.

• Air is assumed to be a thermally perfect gas, but with hu-midity included in the calculation of R, cp, and � as de-scribed in Appendix A. With an engine pressure ratio of6.5:1, this implies a 0.3% error compared to the results us-ing the Redlich-Kwong correlation for humid air mixtures�27� in the range of temperatures and pressures applicable tothe GE J85-13 axial compressor.

• The axial velocity is assumed constant from rotor inlet torotor outlet in a stage.

• The stage pressure rise is found from the measured poly-tropic efficiency.

• The stator deviation angle is assumed constant independentof incidence angle or corrected speed �9�. Increased devia-tion due to surface roughness is neglected as discussedabove.

• The discharge coefficient of the bellmouth is assumed to be0.98.

• The pressure loss in the IGV is assumed negligible. For thedata presented, the fully open IGV will not cause any turn-ing of the flow.

• Deterioration is assumed to increase the sidewall boundarylayers due to deposits on the first-stage annulus. The result-ing decrease in effective flow area is modeled through ablockage factor defined as the ratio of true flow area togeometrical flow area. Assuming no deviation of the eight-stage flow coefficient for the deteriorated engine comparedto clean conditions, the deteriorated engine blockage factorwas found to be 0.96. The blockage factor is assumed con-stant for all stages.

• The equations used to calculate stage performance are givenin Appendix B.

The test data are compared to the theoretical work coefficientbased on the engine geometry and the stage work coefficient
found in literature �7�.
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The theoretical work coefficient is defined from the Euler tur-ine equation, Eq. �2� �28�

�theoreticalT = 1 − ��tan�1� + tan�2�� 2�

here 1 is the exit angle from the preceding stator and 2 is thexit blade angle �relative� from the rotor.

The data published on the stage work coefficients are calculatedased on constant values for cp �1.005 kJ/kg K� and � �1.4� �7�.n this paper, these gas properties are dependent on gas tempera-ure and humidity. The difference will be largest in the aft stageshere the gas temperatures are at the highest. Still, the publishedata gives a good indication of the GE J85-13 stage performanceor a large range of flow coefficients.

Figures 11–14 show the change in the stage work coefficientith engine deterioration for stages 2–4 and 6. For stages 2–4,eterioration reduces both the stage work coefficient and the flowoefficient. This was expected for a degraded stage. For stage 6,

Fig. 11 Stage 2 work coefficient



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the data for deteriorated engine and clean engine overlay eachother. The same trend is found in all stages 5–8, indicating zerofouling in the aft stages. This agrees well with the observation ofonly minor traces of salt deposits in stages 5 and 6 and no saltdeposits in stages 7 or 8.

The measurements are in good agreement with the publisheddata �7� for stages 2–4. However, for stage 6, the measured stagework coefficient is higher than the published data. Because thecurrent analysis is based on the difference between the clean anddeteriorated condition, this deviation from expected stage workcoefficient is of no consequence. The bellmouth discharge coeffi-cient and the temperature recovery factor will influence on thestage work and flow coefficients. These factors are kept constantin the analysis.

ConclusionsThe test results presented in this paper show deteriorated per-

formance of the GE J85-13 jet engine after accelerated salt-deterioration tests undertaken at the test facilities of the RoyalNorwegian Air Force. Accelerated salt testing was found to be aneffective method to systematically deteriorate gas turbine com-pressor performance.

Deposited matter in an engine will depend on the nature of thedeposits �i.e., the material, the particle size, and the adhesive ca-pability of the deposits�. With the present testing, the salt depositswere found mainly in the front stages of the compressor, and thestators were found to have more deposits than the rotor blades.The surface roughness levels and applicable Reynolds numberswere found to give added profile losses but no significant in-creases in the stator deviation angle.

The overall engine performance shows the difficulties involvedin condition monitoring of deteriorated engines, because the dete-rioration causes a nonlinear shift in the degraded performance.This hampers the comparison of deterioration trends based on thecalculation of percentage deviation at constant corrected enginespeed.

Intake depression was found to be the parameter most sensitiveto compressor deterioration. In addition, intake depression wasless influenced by the engine control mode.

The test data show good agreement with published data onincremental stage work and give valuable information of stage-by-stage engine performance deterioration. The salt deteriorationcaused an increase in the sidewall boundary layers and significantdeterioration in the stage characteristics of the first four stages.

This paper provides valuable information on stage performancedeterioration used in the evaluation of test data from online waterwash tests �5�.

AcknowledgmentThe project has been supported by Statoil ASA, Dresser-Rand


Norwegian Operations and the Royal Norwegian Navy. The Royal



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orwegian Air Force is acknowledged for allowing access to theirest facilities and for allowing us to borrow a GE J85-13 for testurposes. Without the help of Hans K. Henriksen, a retired test-ell operator of the RNoAF, this project would never have comeo fruition.

omenclatureA � effective flow area �m2�cp � specific heat at constant pressure �kJ/kg K�C � absolute air velocity �m/s�

CIT � compressor inlet temperature �K�CIP � compressor inlet pressure �kPa�

Comb � combustor �Fig. 1�Compr � compressor �Fig. 1�

EGT � exhaust gas temperature �Fig. 1�GE � General Electric

IGV � inlet guide vanesISO � International Organization for

Standardization1

k � grain size of salt crystals ��m�ks � equivalent sand roughness ��m�m � mass flow rate �kg/s�

MW � molecular weight �kg/kmole�N � shaft speed �rpm�

NACA � National Advisory Committee forAeronautics

NASA � National Aeronautics and SpaceAdministration

p � pressure �kPa�PS � pressure side �concave� of blades/ vanesR � gas constant �kJ/kg K�r � radius �m�

R0 � universal gas constant �kJ/kg K�Re � Reynolds number

R.H. � relative humidity �%�RNoAF � Royal Norwegian Air Force

RTD � resistance temperature detectorSS � suction side �convex� of blades and vanesStg � stage

T � temperature �K�U � blade velocity �m/s�V � relative air velocity �m/s�

VMD � volume median diameterwar � Specific humidity �kg H2O/kg dry air�

x � molar fraction of component

reek Symbols � absolute air angle �deg� � relative air angle �deg�� air density �kg/m3� � efficiency� � kinematic viscosity �m2/s�� temperature recovery factor� � ratio of specific heats

�T=cp�Tt /Utip2 � stage work coefficient

�=Ca /Utip � flow coefficient

ubscriptsa � axial velocity component

amb � ambient conditionmix � mixture �humid air�poly � polytropic

s � static conditiont � total condition

tip � tip

1ISO reference conditions assume an ambient temperature of 288.15 K, a baro-
etric pressure of 101.325 kPa and a relative humidity of 60%.


VMD � volume median diameterw � tangential velocity component

2.1, 2.2,…, 2.7 � Compressor stage 1, 2,…, to stage 73 � Compressor discharge5 � Turbine discharge

Appendix A: Gas Properties for Humid AirGas properties for humid air are calculated from the following

correlations �25�

cp�mix = warmolarcp�H2O + �1 − warmolarcp�dry air�

�mix = warmolar�H2O + �1 − warmolar�dry air�

Rmix =R0

MWmix

MWmix = xH2OMWH2O + xdry airMWdry air

where warmolar=war�28.96/18.015�.

Appendix B: Stage Performance CalculationsConsider a stage where the rotor inlet, rotor outlet, and stator

outlet are designated as stations 1, 2, and 3, respectively. Theperformance of a single stage is calculated using the followingiterative sequence. The same sequence is used on all compressorstages.

�1� The following parameters are known at the rotor inlet: pt1,ps1, Tt1, Ts1, cp1, �1, and m.

�2� From the principle of conservation of mass and the assump-tion of a thermally perfect gas, the following parameters are cal-culated at the rotor inlet

�1 =ps1

RmixTs1, Ca1 =

m

�1A1

�3� The definition of total condition gives the absolute air ve-locity at the inlet

C1 = �2cp�mix�1�Tt1 − Ts1� , Cw1 = �C12 − Ca1

2

�4� The blade velocities at rotor inlet and outlet are given by theshaft rotation, where the radius, r, is the mean line radius forcalculation of velocity diagrams and the tip radius for calculationof stage work coefficient

U1 = 2�r1N

60, U2 = 2�r2

N

60

�5� To start the calculation loop, the static temperature at thestator outlet, Ts3, is given an initial value.

�6� The total temperature is calculated from the temperaturerecovery factor and the measured temperature, Tm

� =Tm − Ts

Tt − Ts⇒ Tt =

1/� �Tm − Ts� + Ts

1 − �

�7� All stage work is applied in the rotor row, hence, the totaltemperature will not change across the stator row

Cw2 =cp�mix�1�Tt2 − Tt1� + Cw1U1

U2

C2 = �Cw22 + Ca2

2

Ts2 = Tt2 −C2

2

2cp�mix�1

�8� Assuming constant axial velocity across the rotor row, Ca2
=Ca1
JANUARY 2007, Vol. 129 / 125


t

t

bo

vu

R

1

Downlo

�2 =m

Ca1A2

ps2 = �2RmixTs2

pt2 = ps2� Tt2

Ts2��mix/��mix−1�

�9� The stage pressure rise is found from the measured poly-ropic efficiency and assumed constant in all stages

pt3 = pt1�Tt3

Tt1� poly�mix/��mix−1�

ps3 = pt3�Ts3

Tt3��mix/��mix−1�

�3 =ps3

RmixTs3

�10� The stator exit velocity is found assuming zero deviation inhe stator blade exit angle, 3=0

Ca3 =m

�3A3

C3 = Ca3 cos�3�

�11� Then, the value for cp�mix at the stator outlet is updatedased on Ts3 and a new value for static temperature at the statorutlet is calculated

Ts3 = Tt3 −C3

2

2cp�mix�3

�12� The calculations are repeated from step 6 and on using thealue for Ts3 found in step 11, and the whole sequence is repeatedntil Ts3 converges.

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