co x - nasa · tests were performed at a simulated mach number of 0.8, over a range of engine inlet...

N A S A T E C H N I C A L

M E M O R A N D U M

COinCOix

NASA TM X-3537

INLET REYNOLDS NUMBER

AND TEMPERATURE EFFECTS ON

THE STEADY-STATE PERFORMANCE

OF A TFE731-2 TURBOFAN ENGINE

George A. Bobula and Roy A. Lottig

Lewis Research Center

and

Lewis Directorate,

U.S. Army Air Mobility R&D Laboratory

Cleveland, Ohio 44135

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, 0. C. • MAY 1977

https://ntrs.nasa.gov/search.jsp?R=19770017194 2020-03-28T06:54:59+00:00Z

1. Report No.

NASA TM X-3537

2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle INLET REYNOLDS NUMBER AND TEMPERATURE

EFFECTS ON THE STEADY-STATE PERFORMANCE OF

A TFE731-2 TURBOFAN ENGINE

5. Report Date

May 19776. Performing Organization Code

7. Author(s)

George A. Bobula and Roy A. Lottig8. Performing Organization Report No.

E-8941

9. Performing Organization Name and Address

NASA Lewis Research Center andU.S. Army Air Mobility R&D LaboratoryCleveland, Ohio 44135

10. Work Unit No.

505-0511. Contract or Grant No.

12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

13. Type of Report and Period Covered

Technical Memorandum14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract

Effects of varying engine inlet Reynolds number index (0.75, 0.50, 0.25, and 0.12) and temper-ature (289 and 244 K) on a TFE731-2 turbofan engine were evaluated. Results were classifiedas either compression system effects or effects on overall performance. Standard performancemaps are used to present compression system performance. Overall performance parametersare presented as a function of low rotor speed corrected to engine inlet temperature.

17. Key Words (Suggested by Author(sl)

Turbofan engine; Reynolds number effects;Temperature effects; Steady-state perform-ance

18. Distribution Statement

Unclassified - unlimitedSTAR Category 07

19. Security Classif. (of this report)

Unclassified20. Security Classif. (of this page)

Unclassified21. No. of Pages

49

22. Price*

A03

' Foi sale by the National Technical Information Service,"Springfield. Virginia 22161

INLET REYNOLDS NUMBER AND TEMPERATURE EFFECTS ON

THE STEADY-STATE PERFORMANCE OF A

TFE731-2TURBOFAN ENGINE

by George A. Bobula and Roy A. Lottig

Lewis Research Center andU. S. Army Air Mobility R&D Laboratory

SUMMARY

An experimental investigation to determine the effects of engine inlet Reynolds num-ber and temperature on a moderately high bypass ratio turbofan engine, the TFE731-2,was conducted at the Lewis Research Center. Tests were conducted at a constant Machnumber of 0. 8, engine inlet Reynolds number indices of 0. 75, 0. 50, 0. 25, and 0.12,and inlet temperatures of 289 and 244 K.

Decreased engine inlet Reynolds number resulted in decreased corrected engineairflows and compression efficiencies, and increased bypass ratio, compression systemtemperature ratio, interturbine temperature ratio, turbine exit temperature ratio, andcorrected fuel flow. Decreased engine inlet temperature resulted in decreased cor-rected engine airflow, fan hub efficiency and corrected fuel flow, and increased com-pression temperature ratio, turbine exit temperature ratio, and interturbine tempera-ture ratio. Speed match was affected by Reynolds number only at the 244 K inlettemperature where a reversal in trend occurred.

INTRODUCTION

An experimental investigation was performed to determine the effects of engine inletReynolds number and temperature on the steady-state performance of a TFE731-2 tur-bofan engine. Variations of engine inlet Reynolds number and temperature which occurwith changes in aircraft flight condition affect gas turbine engine performance. Sucheffects on turbojet and turbofan engines are documented, for example, in references 1to 9. The literature concerned with these effects on turbofan engines is more limited,

however. An investigation into the engine inlet Reynolds number and temperature ef-fects on the TFE731-2 engine, which has several unique design features, can thereforecontribute to the understanding of turbofan engine performance.

The engine reported herein is a moderately high bypass ratio, two spool gas turbinewith a geared front fan. It has an axial low pressure compressor, a centrifugal highpressure compressor, and a reverse flow annular combustor.

Tests were performed at a simulated Mach number of 0.8, over a range of engineinlet Reynolds number index from 0.12 to 0. 75, and at engine inlet temperatures of289 and 244 K (520° and 440° R). Reynolds number index is defined as the ratio of thelocal Reynolds number at the flight condition to that at standard sea-level conditions.

APPARATUS AND PROCEDURE

Engine

The AiResearch Model TFE731-2 (serial number 7306 Build No. 2) turbofan enginetested in this program is a 15 600-newton (3500-lb) thrust, moderately high bypassratio, two spool gas turbine with a geared front fan. The 0. 556 gear reduction designalso results in the fan rotating counter to the low pressure compressor. The single-stage fan and four-stage axial low pressure compressor are driven by a three-stageaxial turbine. The single-stage centrifugal high pressure compressor discharges into areverse flow, annular combustor. The high spool is driven by a single-stage axial tur-bine which also drives the accessory gearbox. Fan and core flows discharge through

2 2coannular separated flow nozzles of areas 1262 and 617 cm (195. 6 and 95. 6 in. ), re-spectively. Figure l(a) is a schematic diagram depicting the airflow paths. Refer-ence 10 presents further details of the engine design.

Instrumentation

The instrumentation configuration for data reported in this text is shown in figure 1.Engine inlet temperature was measured upstream of the bellmouth at station 1. Totalengine airflow was based on the temperature measured at station 1.0 and a pressuresurvey at station 1.2. The majority of the instrumentation divided the compressionsystem into fan tip (stations 2 to 9), fan hub (2 to 22), low pressure compressor (22 to24), and high pressure compressor (24 to 3). The remaining probes were placed toevaluate overall engine performance (stations 5, 7, 8, 11, and 12).

Pressure, temperature, speeds, and other reported parameters were recorded onthe Lewis Central Automatic Digital Data Encoder (CADDE) (ref. 11).

Engine Installation

The engine installation in the altitude test chamber is shown in figure 2. The in-stallation was of a conventional direct-connect type, with the inlet bellmouth located ina plenum upstream of the engine chamber. The plenum was isolated from the test cham-ber by the front bulkhead (fig. 2(b)). Conditioned air was supplied to the inlet plenumto yield the desired engine inlet Reynolds number index and temperature. The air con-sumed by the engine was discharged through the exhaust nozzles and, with any test cellcooling air, passed into the exhaust collector (fig. 2(c)). The test chamber altitudepressure was controlled by valves downstream of the installation.

Test Procedure

During these tests, a constant ram pressure ratio PT 2/pALT of 1< 525' corre"spending to a simulated Mach number MQ of 0. 8, was maintained. (Symbols are de-fined in appendix A.) Testing was conducted at engine inlet total temperatures of 289and 244 K (520° and 440° R). At a given inlet temperature, inlet pressure was adjustedto achieve the desired engine inlet Reynolds number index while altitude pressure wasadjusted to maintain ram pressure ratio.

Operating line data were obtained at the following combinations of inlet temperatureTT 2 and Reynolds number index RNI2: TT 2 = 289 K, RNIg =0 .75 , 0.50, 0.25;TT' 2 = 244 K, RNI2 = 0. 50, 0.25, 0.12. The possibility of an effect of inlet tempera-ture was thus evaluated at RNI0 values of 0.50 and 0.25. The 244 K T™ 0 also per-

£t A y tt

mitted acquisition of limited data at 0.12 RNL, and of the higher corrected speed datafor 0. 50 and 0. 25 RNI2 without encountering turbine temperature limits.

RESULTS AND DISCUSSION

The first section of results presents the engine inlet Reynolds number index andtemperature effects on the fan, low pressure compressor, and high pressure compres-sor performance. These results are followed by the engine inlet Reynolds number indexand temperature effects on some of the engine performance parameters. All data pre-sented were referred to the highest power setting tested at the 0. 75 RNI2 condition.

Component Performance

Fan performance. - The fan tip and fan hub performance maps are presented in

figure 3. The open symbols represent the 289 K TT 2 data and the shaded symbolsrepresent the 244 K TT 2 data.

The data of figures 3(a) and (b) show that the fan tip and hub pressure ratios wereindependent of RNLj and TT 2 at a given corrected speed until RNI2 reached 0.12.At 0.12 RNL,, the fan tip pressure ratio fell 1.5 percent below the higher RNI2 data at95 percent reference fan speed.

The data of figures 3(c) and (d) show that for a given fan tip or fan hub pressureratio, corrected airflow decreased as RNL, decreased. The data indicate no signi-ficant TT 2 effect at 0. 50 RNI2; while at 0. 25 RNI2, the 244 K data was 1 to 1. 5 per-cent lower in airflow than the 289 K data. The same trend in corrected airflow occurredwith RNI2 and TT 2 at a given corrected speed (fig. 3(e)).

Fan tip and fan hub efficiencies (figs. 3(f) and (g)) generally decreased with decreas-ing RNI2. Fan hub efficiency also decreased with decreasing TT 2- The majority ofthe data indicated that near 103 percent of the reference corrected speed, all but the0.12 RNI2 fan tip efficiency data approached a value near 100 percent of the referencefan tip efficiency. The lower fan tip efficiency at 0.12 RNL, accompanied the drop inpressure ratio previously noted. The fan hub efficiency data showed a merging at highercorrected speeds, but the trend was not quite as strong in this case. Because of thegeneral drop in efficiency as RNI2 and TT 2 were lowered, it is to be expected thatmore turbine work would be necessary to maintain a given speed. This will be dis-played later in the text.

Low pressure axial flow compressor performance. - Low pressure compressorLPC performance is presented in figure 4. The data of figure 4(a), the LPC pressureratio-corrected speed relationship, showed no dependence on either RNL, or T™ <,.In figures 4(b) and (c), corrected airflow dropped as RNI2 decreased either at a givenpressure ratio or corrected speed, as previously noted in the fan. There was no effectof TT 2 on LPC airflow.

The major variation of LPC efficiency, shown in figure 4(d), occurred between0. 50 RNI2 and 0. 25 RNI2, the lower RNI2 data falling 2 to 2. 5 percent below thehigher RNL, data up to 100 percent reference corrected speed. Near 105 percentspeed, the efficiencies were equal.

High pressure centrifugal compressor performance. - The high pressure compres-sor HPC performance is presented in figure 5. Once again, the pressure ratio-corrected speed relationship (fig. 5(a)) was independent of RNL^ and TT 2- DecreasedRNL, at a given pressure ratio or corrected speed resulted in decreased correctedairflow (figs. 5(b) and (c)). No effect of TT 2 on corrected airflow was noted.

HPC efficiency plotted against corrected speed (fig. 5(d)) showed that lower RNI2

yielded lower efficiency above 97 percent reference corrected speed for the data avail-able. Instrumentation problems prevented determination of HPC efficiency at

0.12 RNIg and at 0.25 RNI2 with a TT 2 of 289 K. There was no effect of TT 2 onefficiency.

The presentation of inlet Reynolds number index and temperature effects on the in-dividual compression system components has indicated primarily a drop in both cor-rected airflow and efficiency as RNI« was reduced. Only the fan was affected by TT 2

variation, these secondary effects being a further drop in total corrected airflow and anadditional decrease in fan hub efficiency as Trp 2 decreased.

All components but the fan tip showed pressure ratio independent of RNLj. At0.12 RNL,, the fan tip pressure ratio fell below the higher RNL, data.

Effects on Engine Performance

The low rotor speed corrected to engine inlet temperature was chosen as the inde-pendent variable against which engine performance parameters would be plotted. Thischoice was made because the low rotor is related to both fan and core performance.Data comparisons will be made at a corrected low rotor speed of 95 percent of reference,unless it is noted otherwise. A tabulation is made of the comparison data in table I.

Compression system overall performance. - Figure 6 presents the overall perform-ance of the compression system. Since low rotor speed and fan speed are related by aconstant, normalized corrected fan speed is equal to normalized corrected low rotorspeed. Therefore, figure 6(a) is identical to figure 3(e). Figure 6(a) is shown to com-plete the presentation. The figures show that decreasing RNI2 from 0.75 to 0. 50 at289 K TT 2 decreased airflow by 0. 5 percent. A further decrease in RNL, to 0.25yielded a further drop in airflow of 0. 7 percent. TT 2 had essentially no effect at0. 50 RNI2, while at 0. 25 RNLj, the 244 K data was 1. 5 percent lower in airflow thanthe 289 K data. Part of this change in airflow may be due to changes in the engine matchwith TT 2 changes.

The core compression system overall pressure ratio P^ 3/P-j- 2 (fig- 6(b)) wasessentially independent of both RNI2 and TT 2 throughout the range of corrected lowrotor speed tested. However, with fan and compressor efficiencies dependent on engineinlet conditions, figure 6(c) shows the resulting effect on core compression system totaltemperature ratio. As RNI2 was decreased from 0. 75 to 0. 50, TT 3/TT 2 increased0. 5 percent. Decreasing RNI2 further to 0. 25 resulted in an additional rise in tem-perature ratio of 1.0 percent at 289 K TT 2-

At 244 K TT 2' the increase was 2-2 Per~cent between 0. 25 and 0.12 RNI2- It was also noted that the 244 K TT 2 data at 0. 50and 0. 25 RNL, were 1.2 percent and 0.9 percent higher in TT 3/TT 2 than the re-spective data at 289 K TT 2-

Energy requirements. - Figure 7 shows the effects of RNI2 and TT 2 on theenergy requirements within the engine. The lower compression efficiencies that

resulted with decreased RNI2 should require increased turbine work to operate theengine at the same corrected low rotor speed. Energy input to the engine, as reflectedin corrected fuel flow (fig. 7(a)), increased 3.2 percent as RNI2 decreased from 0.75to 0. 50 and interturbine temperature ratio (fig. 7(b)), TT 5/TT 2, increased 1. 6 per-cent, indicating an increased turbine energy level was required to maintain 95 percentcorrected low rotor speed.

Corrected fuel flow and interturbine temperature ratio also showed the effects ofinlet temperature variation. Corrected fuel flow decreased 3. 2 percent at 0. 5 RNI«and 6. 2 percent at 0. 25 RNI2 as TT 2 was decreased from 289 to 244 K. Interturbinetemperature ratio increased 1.2 percent and 3 percent for the same TT 2 changes.Aside from possible engine match changes previously mentioned, specific heat andspecific heat ratio variations with temperature through the engine are believed to be acontributing factor to these effects. The parameters used to refer speeds, fuel flows,airflows, and temperatures to standard conditions are derived from dimensional analy-sis and assume negligible specific heat effects. By basing correction parameters ondynamic similarity, specific heat variations can be accounted for (refs. 12 and 13). Ina report on this engine (ref. 14) the authors suggested modified correction parameters,both theoretically and empirically determined. These parameters, while offering abetter correlation of data at the same RNI2 but different TT 2 values, did not bringabout total agreement. With the limited inlet temperature variation of this test, newmodified correction parameters were not evaluated. It was only determined that therewas an effect of inlet temperature on some parameters.

Overall cycle performance. - Figure 8 presents the effects of engine inlet condi-tions on overall cycle pressure and temperature ratios. In figures 8(a) and (b) fan ducttotal pressure ratio PT jj/P-p 2 and engine pressure ratio PT 7/PT 2 are seen tobe independent of RNL, and TT 2. The fan duct total temperature ratio (fig. 8(c)) wasessentially independent of RNLj and TT ,,. Only the 0.12 RNL, data appeared todiffer, being at most 0. 5 percent higher than the rest of the data. The core engine totaltemperature ratio (fig. 8(d)) showed the result of operating at a constant corrected lowrotor speed, and therefore constant cycle pressure ratio, while decreasing RNLj. Thedecreasing component efficiencies and increasing heat addition, resulting in entropyincreases, yielded a higher final enthalpy state. Thus, as RNL, decreased from 0.75to 0.50, turbine exit temperature ratio TT ij/Trr, 2 rose 1.7 percent. DecreasingRNI2 further to 0.25 yielded an additional TT 7/TT 2 rise of 4.3 percent. The lowertemperature data showed similar trends.

The only parameter in figure 8 showing an effect of TT 2 was TT 7/TT 2- De-creasing TT 2 from 289 to 244 K at 0.50 RNI2 yielded negligible change; while at0.25 RNI2, the ratio increased 1.9 percent.

Core and fan stream matching. - The match between core and fan streams is eval-uated using bypass ratio (BPR) and also the speed match. Bypass ratio is the ratio of

the airflow bypassed through the fan duct to the airflow entering the engine core. The

core inlet airflow was determined from core exhaust gas flow, fuel flow, and cooling

flows. The core exhaust gas flow was based on measurements of exhaust gas condition

and nozzle model calibration test results. The model tests were performed by the en-

gine manufacturer and verified during the altitude-performance testing of reference 14.

Fan duct airflow was calculated as total engine airflow minus core inlet airflow.

Nozzle total-to-static pressure ratios are the forcing functions of core and fan noz-

zle corrected flows. Since the tests were performed at a constant PT 2/PALT' tne

variation of exhaust nozzle total-to-static pressure ratio was implied in figures 8(a)

and (b), the cycle total pressure ratios, for example,

PALT

There is, however, another effect not associated with engine inlet conditions revealed

in figure 9(a). Here it is seen that, over the engine operating range, the fan nozzle exit

external static pressure agrees with ambient pressure. The core nozzle exit external

static pressure, however, is seen to increase with power setting due to the effect of the

fan flow field on the core stream. The resulting nozzle total-to-static pressure ratios

are shown in figures 9(b) and (c). All nozzle performance figures showed no dependence

on engine inlet conditions.

At the test ram pressure ratio of 1.525, the exhaust nozzles were unchoked over

much of the engine operating range. The choked flow regions indicated in the figures

were determined from the nozzle calibration tests. Although the fan nozzle operated at

a higher pressure ratio than the core nozzle at the lower speeds and choked before the

core, figure 9(c) shows that the core nozzle pressure ratio increased approximately

twice as fast as that of the fan. The core nozzle choked soon after the fan. The result-

ing corrected core nozzle gas flow, presented in figure 9(d), is independent of engine

inlet conditions.

Bypass ratio, as seen in figure 10(a), can be expressed as

WABPR =

WA22

or, by rearranging as in appendix B, BPR can be represented solely in terms of param-

eters presented in previous figures as

BPR =

K

(EPR)

- 1

Bypass ratio increased 1.3 percent as RNI2 decreased from 0.75 to 0. 50 and 2. 1 per-cent as RNI2 decreased from 0.50 to 0.25. Examination of the behavior of the variousparameters in the second representation for BPR revealed that only TT 7/TT 2

WF/( t/07 6«>) were affected in such a way by decreasing RNL, so as to yield an in-crease in BPR. Examining the relative magnitudes of the effect of these parameters onBPR showed that the BPR increase was due mainly to increased TT 7/TT 2-

The speed match of figure 10(b) shows no dependence on RNI2 at 289 K TT 2- At244 K TT 2, however, effects of RNI2 were seen. The 244 K TT 2 data at 0. 50 RNI2

was 1. 3 percent higher in corrected high rotor speed than the 289 K data. The cold0. 25 RNI2 data fell 0. 5 percent below the cold 0. 50 RNI2 data, while the 0. 12 RNI2

data was 0. 9 percent above that at 0. 50 RNL,. This reversal is believed to be genuineas data recorded at a later date repeated the trend.

It is believed that the speed match trends were due to turbine performance andmatching changes. Sufficient turbine instrumentation was not available to verify this,however.

SUMMARY OF RESULTS

An investigation was undertaken to determine the effects of engine inlet Reynoldsnumber and temperature on the fan, compressor, and overall engine performance of aTFE731-2 turbofan engine. Inlet Reynolds number indices of 0. 75, 0. 50, 0.25, and0.12 were examined at inlet temperatures of 289 and 244 K (520° and 440° R), at a rampressure ratio of 1.525. The summarized results are as follows:

1. A reduction of inlet Reynolds number index RNI2 resulted in decreased fan andcompressor corrected airflow and efficiency at a given corrected speed.

2. Fan and compressor total pressure ratios were independent of RNL, to valuesof 0.25; and at 0.12 RNI2, only fan tip pressure ratio suffered a loss.

3. Fan airflow decreased with decreasing inlet total temperature TT 2, the effectbeing more noticeable at 0. 25 RNL,.

4. Fan hub efficiency decreased with decreasing TT 2-5. The compressor operating lines were unaffected by T™ 0.1,66. The overall fan/compressor total pressure ratio, engine pressure ratio, and

fan duct total pressure ratio were essentially independent of RNL, and TT 2 at agiven corrected low rotor speed.

7. Overall core compression system temperature ratio, interturbine temperatureratio, turbine exit temperature ratio, and corrected fuel flow increased as RNI2 de-creased at a fixed corrected low rotor speed.

8. Decreased engine inlet temperature resulted in increased core compressionsystem temperature ratio, interturbine temperature ratio and turbine exit temperatureratio, and a decreased corrected fuel flow at a given corrected low rotor speed.

9. Bypass ratio increased as RNI2 decreased but was independent of TT 2-10. Speed match was independent of RNI2 at 289 K TT 2- At 244 K TT 2, de-

creasing RNL> showed an initial drop in corrected high rotor speed, followed by a re-versal and increase in high rotor speed as RNI2 continued to drop.

Lewis Research Center,National Aeronautics and Space Administration,

andU. S. Army Air Mobility R&D Laboratory,

Cleveland, Ohio, November 19, 1976,505-05.

APPENDIX A

SYMBOLS

BPR bypass ratio (e.g., ratio of fan duct airflow at fan tip discharge to core engineairflow at fan hub discharge)

EPR engine pressure ratio (e.g., ratio of turbine exit total pressure to engine inlettotal pressure)

MQ simulated Mach number

NF fan speed, 0. 556 NL, rpm

NH high rotor speed, rpm

NL low rotor speed, rpmo

PS static pressure, N/cmn

PT total pressure, N/cm

RNI Reynolds number index, 6/(<p\/0)

TT total temperature, K

WA air flow rate, kg/sec

WF fuel flow rate, kg/hr

WG gas flow rate, kg/hr

6 ratio of total pressure to absolute pressure of NASA standard sea-level condi-tions

77 adiabatic efficiency

9 ratio of total temperature to absolute temperature of NASA standard sea-levelconditions

tp ratio of viscosity to viscosity of NASA standard sea-level conditions

Subscripts:

ALT simulated altitude

FH fan hub

FT fan tip

HPC high pressure compressor

LPC low pressure compressor

ref reference condition

10

2 fan inlet

3 HPC discharge

5 interurbine

7 turbine discharge

8 core nozzle discharge

9 fan tip discharge

11 fan nozzle inlet

12 fan nozzle discharge

22 fan hub exit and LPC inlet

24 LPC discharge and HPC inlet

11

APPENDIX B

BYPASS RATIO DEPENDENCY

While bypass ratio BPR was defined for this report as WAg/WA22, its calculationwas somewhat more involved. The calculations are detailed below:

WAfBPR =

WA22

WAg = WA2 - WA22

where WA2 was based on an airflow survey at station 1. 2 and WA22 was based oncore exhaust nozzle calibration information, fuel flow, and bleed flows.

The dependence of BPR on other parameters reported herein may be further de-veloped as follows:

WAQ WA9 - WA99 WA9BPR = 2- = £ ££ = £_ - 1

WA22 WA,22 WA22

1 + BPR =WA,

WA22

- K(WA7) = K(WG? - WF)

where K is an empirically determined constant accounting for bleed flows. Thus,

wA

1 + BPR =WA.

WA22K -I—± (WG7 - WF) K

12

1 + BPR =

(EPR)

K WF

13

REFERENCES

1. Denser, W. A.; Finger, H. B.: Com press or-Stall Problems in Gas-Turbine TypeAircraft Engines. SAE Paper 751, Apr. 1956.

2. Campbell, Carl E.; and Sobolewski, Adam E.: Altitude-Chamber Investigation ofJ73-GE-1A Turbojet Engine Component Performance. NACA RM E53108, 1954.

3. Wallner, Lewis E.; and Fleming, William A.: Reynolds Number Effect on Axial-Flow Compressor Performance. NACARME9G11, 1949.

4. NACA Conference on Aircraft Propulsion Systems Research. NACA, 1950.

5. Sanders, Newell D.; and Behun, Michael: Generalization of Turbo jet-Engine Per-formance in Terms of Pumping Characteristics. NACA TN 1927, 1949.

6. Walker, Curtis L.; Huntley, S. C.; and Braithwaite, W. M.: Component and Over-All Performance Evaluation of an Axial-Flow Turbojet Engine Over a Range ofEngine-Inlet Reynolds Numbers. NACA RM E52B08, 1952.

7. Seashore, Ferris L.; and Corrington, Lester C.: Component Performance ofJ71-A-2 (600-D1) Turbojet Engine at Several Reynolds Number Indices. NACARME56B14, 1956.

8. Campbell, Carl E.: Performance of Basic XJ79-GE-1 Turbojet Engine and ItsComponents. NACA RM E58C12, 1958.

9. Biesiadny, Thomas J.; Grey, Rudolph E.; and Abdelwahab, Mahmood: AltitudePerformance of a Low-Noise-Technology Fan in a Turbofan Engine With andWithout a Sound Suppressing Nacelle. NASA TM X-3385, 1976.

10. Steele, M. C.; and Roberts, F. L.: Highlights of the Design and Development of aModern Geared-Fan Jet Engine. SAE Paper 720351, Mar. 1972.

11. Central Automatic Data Processing System. NACA TN 4212, 1958.

12. Glassman, Arthur J., ed.: Basic Turbine Concepts. Turbine Design and Applica-tion. NASASP-290, Vol. 1, 1972, pp. 21-68.

13. Lancaster, O. E., ed.: Jet Propulsion Engines. Vol. XII, Princeton Univ. Press,1959.

14. Dixon, G. J.; Davenport, W. R.; and Steele, M. A.: Engineering Test Report ofAltitude - Performance Test Conducted at NASA Lewis Research Center onAiresearch Model TFE 731-2 Turbofan Engine. PE-8296-R, AiResearch Mfg.Co., 1972.

14

TABLE I. - CHANGES IN ENGINE PERFORMANCE

PARAMETERS FROM VALUE AT RNIg = 0. 75

AND NL/ f/02~ = 95 PERCENT (NL/ f^)ref

[+ indicates increases; - indicates decreases.]

Parameter

WA2 \^/62

TT,3/TT,2WF/(V^ 62)

TT,5/TT,2TT, 11/TT,2TT, 7/TT, 2BPR

NH/ Y/e^

Inlet Reynolds number index, RNLj

0.50 0.25 0.12

Inlet total temperature, TT 2, K

289

-0.5

+.5

+3.2

+1.6

0

+1.7

+1.3

0

244

-0.5

+1.7

0

+2.8

0

+1.7

+1.3

+1.3

289

-1.2

+ 1.5

+6.2

+4.7

0

+6.0

+3.4

0

244

-2.7

+2.4

0

+7.7

0

+7.9

+3.4

+.8

244

-5.2

+4.6

+6.2

+ 15.7

+.5

+ 18.3

+5.5

+2.2

15

LPC-, HPC-,( ,Fan reduction gears-x Fan-, \ \ rHPT

Combustor

/^LPT ^Fan nozzle

Core nozzle

12(a) Station locations and flow/path.

O Total pressureO Static pressure+ Total temperature:

Station 1.0Bellmouth inlet

Station L 2Airflow measurement

Station 2'Fan inlet

Station 9.0Fan tip discharge

Station 22Fan hub exit and LPC inlet

Station 24LPC discharge and HPC inlet

Station 3.0HPC discharge

Station 5.0Interturbine

Station 7Turbine discharge

Station 8.0Core nozzle discharge

Station 11.0Fan nozzle inlet

Station iaoFan nozzle discharge

(b) Individual station layouts (looking upstream).

Figure 1. - Schematic of TFFJ31-2 engine and instHrmentation.CD-12065-07

16

C-71-3593

(a) Installation in altitude facility.

Cooling air torus

•Front bulkhead

(b) Front bulkhead and inlet ducting.

Figure 2. - Engine installation.

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Open symbols denote 289 Kinlet total temperature

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20.50 60 70 80 90

Percent of reference corrected LPC speed, percent

(a) LPC pressure ratio plotted against corrected LPC speed.

Figure 4.-Low pressure compressor (LPC) performance at simulated Mach number MQ = 0.8.

26

120

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InletReynoldsnumberindex,RNI2

O 0.75D .50A .25O .12


Shaded symbols denote 2M Kinlet total temperature

30 'lO 50 60 70 80

Percent of reference corrected LPC airflow,, percent

(b) Operating lines.

Figure 4. - Continued.

110

ref

27


O 0.75Q .50A .25O -12

speed, percent

60 70

Percent of reference corrected LPC

(c) Corrected airflow plotted against corrected LPC speed.

Figure I*. - Continued.

28

>•uCHI

InletReynoldsnumber!ndex,RNI


Shaded symbols denote 2kb Kinlet total temperature

50 60 70 80 90 100

Percent of reference corrected LPC speed, percent (NL/i/6,9 IV ' V ZZ'ref

(d) LPC efficiency plotted against corrected LPC speed.

Figure A. - Concluded.

29

110

tou

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> denote 2kktemperature

:;:;h*

-

K :

:1:u.6080 90 100 110Percent of reference corrected HPC speed,

percent (NH/ 2»t )^ ref

(a) HPC pressure ratio plotted against corrected HPC speed.

Figure 5- ~ High pressure compressor performance at simu-lated Mach number M, 0.8.

30

1 10

pre

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80 90 100 110

Percent of reference corrected HPC

airflow, percent /WA^, 077/6.

(b) Operating line

Figure 5- " Continued.

ref

31

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InletReynolds

numberindex,RNI 2

O 0.75D -50A .25

Open symbols denote 289 Kinlet total temperature ;

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- — -.-----80 90 100 110

, Percent of reference corrected HPC

refspeed, percent

(c) Corrected airflow plotted against corrected HPC speed.

Figure 5- ~ Continued. £*

32

uc0)

<u0)

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Q.

C0)u1_(UQ.

110

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~~

/•:_^--:::(£):::

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80 90 100 110

Percent of reference corrected HPC

speed, percent (NH/ ")

(d) HPC efficiency plotted againstcorrected HPC speed.

Figure 5- ~ Concluded.

33

00

o

4)J3

3C

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34

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120

100

80

60

40

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40 50 60 70 80 90 100

Percent of reference corrected low rotor speed, percent

(b) Overall pressure ratio plotted against corrected low rotor speed.


110

ref

35

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roi_<uo. t-£ <0«) 1-*•>.—•.

CM

i- I—Q.I-

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<uu0)a.

InletReynolds

numberindex,

RNI



5050 60 70 80 90


(c) Overall temperature ratio plotted against corrected low rotor speed.

Figure 6. - Concluded.

36

140

r>4«o

LO"

coo1.

34-

I.8<uuca)i.<a

cVuI.01o.

120

100

80

60

40

20


O 0.75D .50A .25O .12


Shaded symbols denote 244 Kinlet total temperature

40 50 60 70 80 90 100


(a) Corrected fuel flow plotted against corrected low rotor speed.

Figure 7- ~ Energy requirements at simulated Mach number M. = 0.8.

110

37

1 iU .

i in

0 1 f\A ; . ; ; •

4-t ; ; ; ;

L. ::.:

3 : : : : :

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number • ' •:'. index

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A .257? 0 -12

;; Open symbols denote 289 K ; :;; ; ; ; ; ;;:; ; ; ; : • ; ; ; ; ; : ; ;•;

: Shaded symbols denote 2M K : ' • ' • ' • ' ' • • ' • ' - ' • ' • ' • ' • ' • ' ' • ' ' • ' • ' ' • ' ' • ' ' • ' • ' • ' • ' • ' ' • ' ' • ' • ' • ' •: inlet total temperature i ;:;; ;;;; ;;;; :;:; ;;;; ;;:; ;:;

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- -.-:'- TT- ~- — - — - -rr rrrr rrr- r^r

; • • ' •• • • • • ' t::-

46 50 60 70 80 90


100 110

(b) Inter-turbine temperature ratio plotted against corrected low rotor speed.

Figure 7. ~ Concluded.

38

o4J<Dl-

<Ul_3in

<O

O I-D Q-•o --.

C —ra

oC 4-"«) ci. o4) o««- 1-<u <u1_ Q.

c0)oL.(1)Q_


O 0.75D -50A .25O .12



50 60 70 80 90


(a) Fan Duct pressure ratio plotted against corrected low rotor speed.

Figure 8. - Overall cycle performance at simulated Mach number MQ = 0.8.

39

130

o

rai_

o>i.3in «-t/i o)0) v.

110

o> .C 0.

— --^O> I—.C4) t-

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70

50

30


O 0.75D -50A .25O .12



I

50 60 70 80 90 100 110

refPercent of reference corrected low rotor speed, percent ( NL/W6. \

(b) Engine pressure ratio plotted against corrected low rotor speed.


40

em

pe

ratu

re

rati

o,

ref

o

CM

0 1-3 1-

C .—(0t- 1-t-0)-—UC 4Ja) c.L. 1)0) O«4- t- Q/\(1)<1) "O1

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ymbols denote 289 Kt total temperaturesymbols denote 2kk

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— &— i — •" — — •—

|

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K

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60 70 80 90 100 "

Percent of reference corrected low rotor speed, percent [NL/./6, )V I* */ref

(c) Fan duct temperature ratio plotted against corrected low rotor speed.


41

o+JID

<UCL

0)

4-1 <*-._ VX l-<U^-x

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34J I-

UC *-»O C

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InletReynolds

numberindex,RNI


Shaded symbols denote 2M Kjinlet total temperature

50 60 70 80 90


(d) Turbine exit temperature ratio plotted against corrected low rotor speed.


42

CO

O

D

O

L.0)J3

3C

T)0)

Ein

4)Q.

NNO

C(D

[invd/(i1vd -' (paz j (euijou) ajnssajd OIJBJS [eujaixa

43

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100

90

80

70

60

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40k

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Percent of reference corrected low rotor speed, percent (NL/./62 j

(b) Core nozzle pressure ratio plotted against corrected low rotor speed.


44

120

no

win<Di~a.0)

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C —CO —

<O O-U >

i- cv <a4- o0) V.L. <U

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90

80

70

60

:-.-.:: :--± ::::::::-.: :g I:::::::::::::::::::: |;;|;;;;;;::;;; inlet ; ; ; : : ; : ; : ; ;

::::::::::::::: RNI2 :::::::::::

;:;;;:;;;;;:;;; o 0.75 ;;;;; ; ; ; ; : : ;::::::::::::::: D -50 ! : : : ; ; : : : : : :zle:;;;;;|; A .25 > ; ; ; : ; ; ; ; : ;

timm o .12 1 1 il II

zle..... ..-Open symbols denote 289 K :::::::::::::: inlet total temperature :::::::: :

' } ' • - | | j - • •

. . ^ .....i j i i l i . . . j

:: Core :;!":

::i!:ji i::::::::

! ; j ! Fan;;;;;;

50 60 70 80 90


100 110

ref

(c) Fan nozzle pressure ratio plotted against corrected low rotor speed.


45



Shaded symbols denote-244 Kinlet total temperature

1*040 50 60 70 80 90 100

Percent of reference corrected low rotor speed, percent (N

(d) Core nozzle corrected gas flow plotted against corrected low rotor speed.


110

ref

46

Q.OD

pe

rce

n

IDO.>-

JO

a)oc<ui_<uu-<u

Vf.

o4Jc<0o

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200

180

160

140

120

1 nn

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40

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InletReynolds

numberindex,RNI2

0 0.75a .soA .25O .12

Ben symbols denote 289 Kinlet total temperature

Taded symbols denote 244inlet total temperature

: ,' i .. - : :

• ;•

::: i:

'.' . . .'. :

'• ' :::

: - : . j .

- 1 -

... i .

70

. / ::^

i ' ' ; .

'.:.' ..".

\ •"%B

• j. -

• ; ; ' . : '..

. -.! '

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• ' : : -

• • T"r*2): . - : i ! i : :

• ( • •

^U :/fL

';

i .

,

• ' ' i

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& : • • ! . . : •

::: :'

: : • I / : -"1: .:

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r^

80 90 100

.....

-

'

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Kl_

• : ; . | ' . : .• • / K:;;: :;;

i -

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1

Percent of reference corrected low rotor speed, percent ( NL/«/6T)V '» 2/ref

(a) Bypass ratio plotted against corrected low rotor speed.

Figure 10. - Flow split at simulated Mach number MQ = 0.8.

47

•O<D<UQ.

O4JO

c .u0) l-C. <a0) Q.

c<uO1_<u

Q.

InletReynolds

numberi ndex,RNI

Open symbols denote 289inlet total temperature


40 50 60 70 80 90 100 110

Percent of reference corrected low rotor speed, percent (NL/«/6r-)V /V Z/ref

(b) Speed match; corrected high rotor speed plotted against corrected low rotor speed.


48 NASA-Langley, 1977 E-8941

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON. D.C. 2O546

OFFICIAL BUSINESS

PENALTY FOR PRIVATE USE $3OO SPECIAL FOURTH-CLASS RATEBOOK

POSTAGE AND FEES PAIDNATIONAL AERONAUTICS AND

SPACE ADMINISTRATION4S1

POSTMASTER : If Undeliverable (Section 158Postal Manual) Do Not Return

"The aeronautical and space activities of the United States shall beconducted so as to contribute . . . to the expansion of human knowl-edge of phenomena in the atmosphere and space. The Administrationshall provide for the widest practicable and appropriate disseminationof information concerning its activities and the results thereof."

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co x - nasa · tests were performed at a simulated mach number of 0.8, over a range of engine inlet...

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