super critical airfoils must study

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*GPMRT P*11411

DESIGN AND PERFORMANCE EVALUATION OF

SUPERCRITICAL AIRFOILS FOR AXIAL FLOWCOMPRESSORS

UNITED TUCHNOLOGISS CORPORATION

PimP & MOW Afh rfOmra 0,JlUnmhm OWNi

P.0. am 1Wmt Iolm lia . No 33M

JUNIt I

SFimi lR 0 4 = 7 N

j~a.m IAPPROVEDOR PUBLC RELEASE;DIU1TRIUUTION U01LESITED

PREPARED PIR

UU1PARTMEWr OF THE NAVYNAVAL AIR SYSTEMIS CO-ANO

79 06 2


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UNCLASSIFIED

SECURITY CL.ASSIFICATION OF THIS PAGE (When Date E n t e r e d ) .

REPORT DOCUMENTATION PAGE BEFORE COMPLETING FORMT. REPORT NUMBER 2 GOVT ACCESSION NO. 3. R E C I P I E . T- S CATALOG

NUMBER

FR 11455'

4. T IT L E ad Subtitle COVR

;UPERCRITICAL AIRFOILS FOR VIAL '%.-VR-FORMING -)M. ORT NUMBERLOW .MPRESS0 s. r-- -

7.. CONTRACT OR GRANTNUMBER x)

S t e p z ( _ _ , H . E, -S, f. y, , 19-77-C 4

Hobbs D.E./ 1 S ~ l 7C 04~. PRFORIINORAI .- Ia. PROGRAM ELEMENT.

PR O JEC T. TASK.... AREA & WORK UNIT NUMBERS

Prat t & Whitney Airc ra f t Gro . -7Government Products Div is ion - - P.O. Box 2691 ' /

West Palm Beach, Flor ida 3340211 CONTROLLING OFFICE NAME ANO AOORESS 12. R ' O RT OAT-

Naval Air Systems CommandFebruary 1979

Washington, D.C. 203611. NUMBER42F DAGES

42t4 . MoNITORING AGENCY NAME AOORESSXIl diflterent trot C o n t r o l i n g Office IS. SECORITY

CLASS. (of thle r e p o r t )

Un cl as s i f i ed. ...... .iA15a. DcECL ASLi FIC ATION/ DOWNGRADING

SCHEDULE

14. DISTRIBUTION STATEMENT (of t id Re p o r t )

Approved for p u b l i c l ease ; distribution unlimiLed.

17. DISTRIBUTION STATEMENT (o1bm

IS. SU PPLEMEN TA RY N O T E S

'11. KEY WOROS (Continue on rover * s i d e if neecessae'v and i d e n t i fy byblock number)

S u p e r c r i t i c a l Airfoil Tu r b u l e n t Boundary Layer Compressor

Transonic Cascad e Tunnel Laminar Boundary Layer Shock le s sAirfoils

Fan E x i t S t a t o r Light Pen Scope

Computer Des ign

20. ABSTRACT C on inue an revere* oid If oceemary and Identify y block ,num,ber)

Pratt & Whitney A i r c r a f t has deve loped an a na ly t i c a l d e s i g n procedure

for produc ing supercritic l cascade airfoils s tisfying p r a c t i c a l

aerodynamic and structur l r e q u i r e m e n t s . The purpose o f t h i s r e s e a r c h

is to demons t ra te the p o t e n t i a l fo r s u b s t a n t i a l e ff i c i e nc y

improvement : in compressor s tages i nc o r po r a t i ng these airfoil d e s i g n s .

DO 1JAN73 1473 EDITION OF I NOV45 IS OBSOLETE SSIFIED

SJAN72 IA VCLSSIFICD C-URITY CLASSIFICATION OF THIS PAGE When Da ta E~ntered)


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UNCLASSIFIED

59CUMTY CLAI5SFICATION OF THIS ,AGE( n, D t a . .. ...

20

? The midspan sec t ion of a fan exi t s t a tor was selected for the designapplication. An a i r fo i l was designed, fabricated, and t es ted in at ransonic cascade tunnel. Test condi t ions were varied over a widerange of inlet flow angles and inlet Mach numbers to determine designand off-design performance. Test point information included in le t andexi t condi t ions and a i r fo i l surface pressure d i s t r i bu t i on .

Analysis of t e s t data taken a t the design point condi t ionssubstantiates a good match between measured and design shocklesssurface Mach number distribution, a l1,w loss resulting from anat tached suction side boundary layez, and an exi t flow angle withinone degree of design specificatiorn, . Thus design goals weresuccessfully met. Analysis of t es t resul t s for off -des ign condi t ionsconfirmed that the cascade also operates eff icien tly over a wide rangeof in le t angles and Mach number. The val id i ty of the Pra t t & WhitneyAircraf t supercr i t i ca l design procedure has been demonstra ted .

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UNCLASS IFIEDSEugWqiT CL _AfFICATiOW Of THIS P AGIE uhe Data Entered1)


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TABLE OF CONTENTS

Section Page

Table of Contents 1

List of Illustrations 2

1.0 INTRODUCTION 5

2.0 BACKGROUND 6

2.1 Development of Supercritical Design Methods 6

3.0 DESIGN 8

3.1 Design Procedure 8

3.2 Design Description 10

4.0 PERFORMANCE EVATUATION 11

4.1 Cascade Airfoil Fabrication 11

4.2 Description of Test Facility 11

4.2.1 Tunnel Operating Condition~s and Operation 12

4.2.2 Instrumentation 13

4.3 Testing Procedure and Data Acquisition/Reduction

System 134.3.1 Testing Procedures 13

4.3.2 Data Acquisition/R eduction 13

4.4 Test Results 14

5.0 ANALYTICAL SIMULATIONS 16

6.0 CONCLUSIONS 17

Illustrations 18

References 37

Appendix - Data Table 39

List of Symbols, Abbreviations, and Subscripts 41

Distribution List 43


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LIST OF ILLUSTRATIONS

Figure Page

I (a-re) Schlieren Photographs Showing Cascade FlowField for a Range of Increasing Inlet Mach Numbers(see Reference 1). (f) Schematic of the High-LossCondit ion shown in (e). 18

2 Korn Supercritical Cascade Geometry andAerodynamic Data 19

3 Supercritical Pi rfo i l Aerodynamic Design Requirements 20

4 Supe rc r i t i c a l Cascade Geometry and Aerodynamic DesignCondit ions 21

5 Cilercri t ical Cascade Design Surface Mach Number

Distribution 22

6 DFVLR Transonic Cascade Test Faci l i ty 23

7 Schematic of DFVLR Transonic Cascade Test Section 24

8 Comparison of Design Mach Number Distr ibut ionand the Measured Tebt Data. 25

9 Cascade Performance - Loss versus in let Mach numbera t design in le t angle. 26

10 Cascade Performance - Loss versus in let Mach numberat +7 degrees

incidence. 2711 Cascade Performance - Loss versus inlet Mach number

at +5 degrees incidence. 28

12 Cascade Performance - Loss versus in let Mach numberat +3 degrees incidence. 29

13 Cascade Performance - Loss versus in let Mach numberat -3 degrees incidence. 30

14 Cascade Performance - Loss versus inlet Mach numberat -5 degrees incidence. 31

15 Caecade Performance - Loss versus in let Mach numberat -10 degrees incidence. 32

16 Cascade Performance - Loss as a function of cascadeinlet angle or incidence for various upstream Machnumbers. 33

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LIST OF ILLUSTRATIONS (Cont. 'd)

FigurePage

17 Cascade Performance - Flow turning and exi t flowangle

as a funct ion of in le t angle for various upstream Mach

numbers.34

18 Results of Analytical Simulation of Near Design Test

Condition (Point 72)35

19 Results of Analy t ical Simulation of Ten Degree

Negative Incidence Test Condit ion (Point 38)36

3

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1.0 INTRODUCTION

This report discusses the details of the design and the performanceevaluation of a practical supercritical cascade. The purpose of thiswork is to demonstrate that potentially cubstantial improvements in

compreosor efficiency can result from employment of analyticallydesigned, shockless supercritical airfoi ls . Conventional compressorairfoi ls operating in the transonic regime typically exhibit highlosses, which are principally combinations of shock losses andshock-induced boundary layer separation lossea. Recent Pratt & WhitneyAircraft experience in the development and test of supercriticalairfoi ls in cascades has demonstrated the superiority of these airfoi ldesigns over conventional compressor airfoil designs in terms of bothreduced losses and increased incidence range. This contract providesfor the design, tes t , and data analysis of a supercritical cascade forapplication in high technology, future generation compressors of gasturbine engines.

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2.0 BACKGROUND

2.1 Development of Supercr i t ical Airfoi l Design Methods

Supercr i t ical a i r f o i l s are a class of t ransonic a i r f o i l s which operate

with subsonic in let and exi t flow velocit ies and with embedded regionsof supersonic flow adjacent to the a i r fo i l surface. The term supercr i t ica l refers to the presence of velocit ies in the flow f i e ldwhi.h are above the cr i t i ca l or sonic speed. His torica l ly, progreesin the design methods for t ransonic a i r fo i l s has severely laggedmethods used to design ful ly subsonic or supersonic a i r fo i l s . This la gis pr imar i ly due to the mathematical d i ff i cu l t i es in so lv ing th eequations which model the t ransonic flow f ie ld . Without th efundamental ab i l i t y to compute the velocit ies on the a i r fo i l surface,the well-developed low speed design techniques employing boundarylayer theory have been of no value.

The ear ly knowledge of a i r fo i l s in the t r ansonic regime was derivedfrom wind tunnel experiments on subsonic or supersonic designs. This

type of experimentation provided an understanding that aerodynamicdef iciencies of these designs were caused by the strong normal shockswhich terminated the embedded supersonic region. For i sola teda i r fo i l s , th is shock caused a rapid increase in drag and a reduct ionof lift as the approach Mach number increased through the highsubsonic range. In cascades, th is shock produced the analogous effec tsof increased to tal pressu re loss and reduced flow turning. Thesefea tures o f the flow f ie ld for a NACA 65 s e r i e s cascade are shown inthe sch l ieren photographs in Figure 1 from the work of Dunavant e t . a l .(1).

In 1965, a resurgence of in te r es t in developing improved supercri t ica ldesign methods resul t ed from Whitcomb's now-famous supercri t ica li so lated a i r fo i l experiment in the NASA Langley eigh t - foo t t r ansonic

tunnel. Whitcomb's exper imen tal ly developed a i r fo i l demonstrated theexistence of shockless supercri t ica l flow f ie lds (2). The shocklessfeature made the flow en t i r e ly i r ro t a t i cna l and thus amentable tomodelling with the potent ia l equation.

1. Dunavant, J. C. et. al. High Speed Cascade Tests of the NACA 65- 12A 1 0 )10 and NACA 65-(12 A2 18 b)10 Compressor BladeSect ions , NACA RML55, 108.

2. Whitcomb, Richard T., and Clark, Larry R., An Airfoi l Shape forEff icien t F l igh t at Supercr i t ical Mach Numbers, NASA TMX-1109,May 1965.

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These tes t resu lts provided the motivation for the development of apract ical transonic cascade design procedure by Prat t WhitneyAircraf t during 1976 and 1977. The current work was undertaken inNovember 1977 to demonstrate th is new design system for pract icalappl ica t ions to the compressor.

3.0 DESVIIN

3.1 Design Procedure

The object ive of the design process is to find an a i r fo i l shape thatsa t i s f i e s a re la ted set of aerodynamic and s tructural requirements.For s upe rc r i t i c a l cascade a i r fo i l s , these requirements can be br i e f lystated as follows.

1. The cascede miust produce a spec i f i ed ex i t veloci ty andflow angle for a specif ied in let veloci ty and f lowangle. The flow turning implied by these cond i t ionsshould be eccomplished with a minimunm of to ta l pressureloss through the cascade.

2. The cascade gap-to-chord ratio should be maximized(consistent with aerodynamic requirements on peaksuct ion side Mach number) to reduce engine weight andcost.

3. The cascade airfoi l must have a thickness and camberd is t r ib u t io n capable of sus ta in ing the aerodynamicpressure forces. Leading and t r a i l i ng edges must becapable of sus ta in ing erosion and foreign object damage.

To achieve these aerodynamic requirements, it is cur rent ly bel ievedthat the surface veloci ty d is t r ib u t io n should have cer taincharac ter is t ics . These charac ter is t ics are l i s ted below and summarizedin Figure 3.

Supercr i t ical cascade a i r fo i l character is t ics

1. A cont inuous acce lera t ion to the peak Mach number on thea i r fo i l suct ion surface, to avoid premature laminarboundary layer separat ion, or t r an s i t io n before the peak.

2. A peak Mach number less than 1.3 to avoid boundary layersepara t ion induced by a severe shock l-ave-boundary layerin teract ion , should a shock develop at off-designcond i t ions .

3. A continuous dece lera t ion froru the peak s - i t i o n surfaceMach number to the t r a i l in g edge, mainta.. ng a

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is of the type described by Lieblein (I1) or Stewart (12). Thetransonic flow solu t ion uses a second order accurate, f in i tedifference scheme to solve the two-dimensional, ful ly non- l inear,poten t ia l flow equation. The analys i s also includes a quasithree-dimensional capabi l i ty to account for the stream tube heightvar iat ions through the cascade. This feature is essent ia l for accurate

modelling of rea l i s t i c flow fields. With th is analysis , supercri t ica lf low fields can be computed fo r nea r ly all cascad es o f p rac t i ca lin teres t . The accuracy of th is solu t ion has been shown (9) to be

equ ivalen t to the Garabedian, Korn and Bauer cascade solution. Theboundary layer analyses program of McNally includes the compressible,in tegra l method of Cohen and Reshotko for laminar layers and thecompressible, i r t eg r a l method of Sasman and Cresci for turbulentlayers. Trans i t ion between the initially laminar layer and theturbulent layer is assumed to occur at the locat ion of a computedlaminar boundary layer separa t ion . Smoothing o f the boundary layerc ha ra c te r i s t i c s through t rans i t ion is used to avoid d i s c on t inu i t i e s inthe displacement thickness.

3.2 Design Descr ip t ion

The part icular design appl icat ion selected for this cont rac t work isthe midspan sect ion of a fan exi t s tator. The Mach numbers, f lowturning, and s ta t i c pressure r ise requirements are represen tat ive of

advanced gas turbine engine conf igura t ions and ref lect a design goaltypica l of current compressor a i r fo i l technology. The selected designis also cons i s tent with the requirements of the mean sec t ion of theex i t s t a to r of the 1600 f t / s ec t ip speed fan, prev iously designed andtested under NASA contract. The existence of this fan r ig and itspreviously measured performance base line using a conventional statorprovides future opportunity for an economical demonstration of asupercr i t ica l stator in a real turbomachinery environment. The statordesign point aerodynamic requirements for inle t flow angle, inle t Machnumber, and flow turning are taken from Table 10.10 of the NAS3-10482cont rac t report (13).

11. Liebl ien, S., and Rouderbush, W. H., Theoret ical Loss Relationsfor Low Speed Two-Dimensional Cascade Flow, NACA TN 3662, March0956.

12. Stewart, W. L., Analysis of Two-Dimensional Compressible FlowLoss Characteristics Downstream of Turbomachine Blade Rows inTerms of Basic Boundary Layer Character ist ics, NACA IN 3515, July1955.

13. Sulam, D.H., M.S. Keenan, and J.T. Flynn, Single-Stage Evaluationof Highly-Loaded Multiple-Circular-Arc Rotor High-Mach-NumberCompressor Stages, NASA CR-7264, PWA-3772.

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The performance requirements of the s t a to r design include an in le tMach number o f 0.76, in let flow angle of 46.8 degrees, and a flowturning o f 43.2 degrees to an axia l discharge angle. The design has anexit Macb number of 0.529, and a spanwise stream tube area contract ionof 1.124.

The design parameter se lec t ion phase of the design i t e ra t ion wasperformed using simple modif icat ions of the Korn supe rc r i t i c a l cascadea i r fo i l and resul ted in a choice of gap-to-chord r a t i o of 0.7. Th ecomputed peak suct ion surface Mach number was 1.27 and the di ffus ionfac tor was 0.34. This initial design was used to s t a r t the f inaldesign i t e ra t ion resu lt ing in the f i na l cascade design (see Figure 4).A p lo t of an ind iv idual blade contour and the corresponding des ignsurface Mach number d i s t r i b u t i o n is presented in Figure 5. Smoothnessq u a l i ty o f the computed Mach number di s t r i bu t ion is dependent on th ecomputational grid size employed in the t ranson ic analysis and ona i r fo i l contour varia t ions with smaller than reasonable manufacturingtolerances. The large Mach number varia t ions a t the t ra i l ing edge are

due to the computation of an invisc id s tagnat ion po in t at the tr ilingedge. This stagnation point w il l not appear in the rea l flow becauseof strong viscous effects.

4.0 PERFORMANCE EVALUATION

The performance evaluation required fabricating the airfoi ls , ter t ingthe cascade in a transonic tunnel, and analyzing the test data todetermine the extent to which the design objectives were met.

4.1 Cascade Airfoil Fabrication

Ten cascade a i r f o i l s of 69.85-,m (2.75 in . ) chord an' 167.64 mm (6.6in . ) span were manufactured from a s teel al loy. Surface contourto lerances of +0.05 mm (+0.002 in . ) were specified and contourinspect ions made with a New England A ir fo i l Tracing Machine a t threespon locations on each a i r f o i l . All ten a i r fo i l s were well with inspecif ied tolerances. Three a i r f o i l s were selected for s ta t ic pressureins t rumentat ion on the basis of th i s inspection. A to ta l of sevena i r f o i l s were selected for in s ta l l a t ion in the cascade tunnel .

4.2 Descr ip t ion of Test Fac i l i ty

The t ranson ic cascade f a c i l i t y used in t h i s inves t iga t ion was bu i l t byNACA-Langley in the early 1950's and t ransferred to the DFVLR(Deutsche Forschung and Versuchsanstalt fur Luft and Raumfahrt) inPorz-Wahn, West Germany in 1963 (14). This fac i l i ty was used to test

14. Starken, H., Breugelmans, F. A. E., and Schimming, P., Inves t iga t ion o f the Axial Veloci ty Density Ratio in a HighTurning Cascade, ASME Paper 75-GT-25, ASME Gas Turbine Conferenceand Products Show, Houston, Texas, March 1975.


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the original Korn supercritical cascade design. Figure 6 shows th ecascade test section as it is currently installed at the DFVLR. Aschematic cross-section to the test section is shown in Figure 7.

4.2.1 Tunnel Operating Conditions and Operation

The transonic cascade tunnel is a closed loop, continuously runningfacility with variable nozzle and variable test section height. Thistunnel is unique in its extensive endwall boundary layer controlsystem and in its substantial vacuum capacity, capable of removing asmuch as 50 percent of the total tunnel flow through various bleed

systems. These provisions are essential for controlling the secondaryf lows induced by the strong gapwise pressure gradients of the

upercritical cascade. The dry air is delivered by a set ofcompressors with a total installed power of 5000 KW 3728 Hp) and a

mass flow of up to 20 kg/sec (44 lbm/sec). The tunnel test sectionheight is variable between 150 mm 5.9 in.) and 450 mm 17.7 in.), andthe span is 167 -m 6.6 in.). These dimensions yield an aspect ratioof 2.4 for a airfoil chord of 69.85 n 2.7 9in. . The tunnel Reynolds

number can be varied from 4 x 105 to 4 x 10 by changing th esettling chamber pressure. The nozzle is half-symmetrical with avariable shape adjustment for the upper half and a variable height

adjustment for the flat lower half. This arrangement provides an inlet

Mach number range from 0.0 to 1.4. The inlet Mach number is uniformover the test section height within +1.5 percent of the maximum, andthe turbulence level in the inlet is 0.5 (+0.2) percent for an inletMach number of 0.3 to 0.7.

The upper wall cf the test section is slotted and equipped withauction to improve the inlet flow conditions. Additional suctioncapability is available for boundary layer removal ahead of the teatsection (slots in the sidewalls reaching from the nozzle top to thebottom wall) and through chordwiae slotted cascade endwalls within the

airfoil pack. The tunnel top endwall flow is separated from thecascade core flow by a tailboard system hinged at the trailing edge ofthe outermost cascade airfoil. The tunnel bottom endwall flow iscontrolled by a movable endwall flap. A lower tailboard is not used.

4.2.2 Instrumentation

The inlet and exit static pressure was measured by static taps on thecascade wall. For this cascade, the upstream measurement plane wasaxially 25 mm (0.984 in.) from the leading edge plane. The exitmeasurement plane was 22 mm 0.866 in.) from the trailing edge plane.The inlet air angle was measured at the same gapwise locations forthree consecutive airfoil channels. The downstream traverse probe

measured a combination of pressures, providing both total an6 staticpressures as well as flow angle for the complete wake traverseinformation. The traverse system was operated in a stepping modeduring loss measurements to ensure sufficient time for the

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probe- t ransducer system to accurately measure the to ta l pressure,s ta t ic pressure, and flow angle in portions o f the wake with s t rongflow gradients .

The cascade a i r f o i l s were heavily instrumented with surface pressuretaps. The a i r f o i l thickness requ i red the instrumentation of th reeseparate a i r f o i l s rather than j us t one. One a i r f o i l was pressuretapped chordwine on the suct ion surface, a second on the pressuresurface, and a th ird was instrumented spanwise on both surfaces atonly a few chordwise locations. The pressure and auction instrLmenteda i r fo i l s were arranged in cascade to record the flow with in a comonpassage.

4.3 Test ing Procedure and Data Acquisit ion/Reduction

The demonstrat ion of the accuracy of the supercri t ical design sys temthrough comparisons with measured cascade data requ i red prec ise

control of the experimental setup to ensure thAt the desiredaerodynamic conditions were close ly matched. The terting procedure anddata acquis i t ion necessary to achieve the required accuracy iredescribed in Sections 4.3.1 and 4.3 .2 .

4.3 .1 Testing Procedures

The in let angle of the cascade relat ive to the in le t duct was set toproduce the correct in le t flow angle (angle accuracy of 0.5 degree orless is possible). The in le t Mach number was then se t at the des iredvalue, and, to ensure that the measured performance was obtained underperiodic flow condit iona, several procedures were followed. Tunnelperiodic ty was checked by comparing the in le t flow angle es measuredat three idjacent gap positions and by observing the inlet and exits ta t ic prkssure d is t r ibu t ions displayed on manometer boards located inthe t e s t cill control room. The exit travers ing probe was operated ina continuously running mode, enabling three ad jacen t blades to betraversed in a reasonable length of time. The resulting wake profileswere displayed on a p lo t t e r to check a i r f o i l wake shape consistency.In addition, wake t raverses a t various spanwise loca t ions weremeasured to guarantee tha t an acceptable portion of the a i r fo i l wasoperating under two-dimensional conditions. Adjustments to obtainperiodic flow conditions were then performed on the movable bot tomendwail, flap, upper tailboard, and on the extensive boundary layercontrol vacuum system. Further adjustments in upstream pressure werealso required to maintain the desired inlet Mach number.

4.3.2 Data Acquisition/Reduction

The complete infolmation for each test point was recorded by acomputar-controlled automatic data acquis i t ion system and stored onmagnetic tape. This information included plenum conditions, in lets ta t ic pressures, in le t flow angles, a i r f o i l surface pressuredi s t r i bu t ions , and downstream wake traverse data.

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is nei r ly constant &t 0.02 up to the design po in t (M upstream 0.73,M leading edge a 0.76). At a Mach number 0.03 greater than the designpoint, the loss doubles to a velue of 0.04. This Mach number range isadequate for th is appl ica t ion , since th is fan s tator mean sect ion hasan

in le t Mach numberof 0.8 or less throughout

i t soperating

range.

Loss data for off -des ign inlet flow angles with varying in le t Machnumbers are shown in Fi tures 10 through 15. These in le t anglescorrespond to incidence angles of +7, +5, +3, -3, -5, and -iO degrees.These dUtp shown in Figure 16, fu l ly define the useful operat ingrange of the cascade. The measured upstream Hach number has been usedon a] of these f igu res . As in the case of the design point ; numericalsimuiations of off -des ign conditions suggest that average Mach numbersfor the upstream and leading edge are s l ight ly different . Test anddesign values, then, should be carefu l ly re la ted if precisecomparisons are required. At various arngles of incidence, numericalsimulations were made for tes t conditions having both low totalpressure lo sses and high in le t Mach numbers. The simulations indica ted

that the Mach number increase of 0.03 gave the most sat isfactoryresu l t . Thus, the loss bucket for the design Mach number of 0.76 wouldbe an interpola ted l ine on Figure hu at an in le t Mach number of 0.73.A s i g n i f i can t range of posi t ive and negative incidence around th edesign in let angle is presen t before losses increase to proh ib i t ivelevels.

Figure 17 depicts flow turning and exi t flow angle versus in let flowangle. Exit flow angle was nearly constant , ranging from approximately89 to 91 degrees. This nearly constant exit- angle over all testcondit ions is a s i g n i f i can t accomplishment of th is a i r fo i l design.This ai r exi t angle consistency is part icular ly important for the

bypass port ion of the fan s ta tors , which must tu rn the flow to anax ial discharge angle to maximize fan duct th rust .

The tes t resu l t s indica te tha t the supercri t ica l a i r fo i l has met itsdesign requirements and has operated eff icien tly over a wide range ofaerodynamic conditions. The tes t data taken a t the design pointcondit ions ind icate (1) a good match between measured and designsurface Mach number dis t r ibut ions , (2) an at tached suct ion sideboundary layer resu l t ing in low loss, and (3) an ex i t angle with in onedegree of design speci f icat ions . These resul ts ind icate tha t theshockless flow f ie ld , combined with firm control of the suct ionsurface diffusion rate, resu l t s in an attached boundary layeressen t i a l to good aerodynamic performance. In conclusion, theseresu l t s demonstrate the val idi ty of th is design method for fu turedesign work.

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6.0 CONCLUSIONS

1. A procedure developed at Prat t & Whitney Aircraft has beenused to design a supercri t ica l cascade to sa t i s fy practical

requi rement s.2. The cascade operated with low tota l pressure loss at the

design point and achieved the required flow turning withinless than one degree.

3. The design surface Mach number di s t r i bu t ion was closelymatched In the test . The a i r fo i l appeared to be shock-free andthe suct ion side boundary layer remained attached throughout 'ts diffus ion to the t rai l ing edge.

4. Off-design performance in terms of Mach number range,incidence range, and flow turning was very good over a broadrange of operat ing condit ions.

5. The flow exit angle was maintained to within +1 degree ofdesign (axial) for upstream Mach numbers ranging from 0.4 to0.75 and over a range of i n l e t incidence angles from -10 to +7degrees.

6. The loss r ise character is t ics at the design in let flow anglewere successfully controlled to provide a 0.03 margin in irletMach number before the measured loss doubled.

7. An analyt ica l simulaLion o f the t e s t design point cond;cionsprovided excel len t agreement between computed and measuredsurface Mach number d is t r ib u t io n s .

8. Analytical simulat ion of a wide varie ty of off-design testcondit ions fur ther subs tant ia tes the accuracy of the flowmodelling procedure for supercr i t ical t ransonic flows.

9. Analytical simulat ions suggest tha t an approximate increase inaverage Mach number of 0.03 exis ts between the upstreammeasuring plane and the a i r fo i l leading edge plane. It isprobable that th is effec t in the DFVLR tunnel is caused bysidewall boundary layer growth, which reduces the duct area byabout 1.6 percent between these planes for th is cascade.

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(a,, Ni o k ) M U b 1

beledory ,:,yet

Figure 1 (a-e) Schlieren Photographs Shoving Cascade Flow Fieldfor a Range of Increasing Inlet Macb Numbers (seeReference 1). (f) Schematic of the High-Loss Conditionshown in Ce)

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GEOMETRIC DATA

C = 2.80 in.r/C = 1.195

Sas = 5 DEGREES f bx = 2.37 in.

AERODYNAMIC DATA

N1 : 78M = 0.48hl = 43.0 DEGREES02 = 68.0 DEGREESTURNING = 25.0 DEGREESAVDR 1.03

X, Engine Axis

Figure 2 Korn Supercritical Cascade Geometry and Aerodynamic Data

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AT TACHED BOUNDARY LAYER

MACH WAVESWAVES

CONTINUOUS ACCELERATION1.4 -- TO BOUNDARY LAYER TRANSITION POINT

F-".----PEAK ACH NUMBER LESS THAN 1.3

cz 10 -SOINICCc . C O N T I N U O U S DECELERATION TO TRAILING

EDGE WITH LOW BOUNDARY LAYER SKIN0.8 FRICTION

f 0.6-

NEARLY CONSTANT SUBSONIC0 2 MACH NUMBER ON PRESSURE SURFACE

0 1 1 . 1 1 10 0.2 0.4 0.6 0.8 1.0

X/AXIAL CHORD

Figure 3 Supercr i t ical Air fo i l Aerodynamic Design Requirements

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GEOMETRIC DATA

iceas =74.25 DEGREES

ax s

AERODYNAM1I. DATA

M. 0.753N1 = 0.529

= 46.8 DEGREES

- -amp.___ TURNING = 43.2 DEGREESX, EW AMAVOR = 1.124

Figure 4 Supercritical Cascade Geometry and Aerodynamic DesignCond it ns


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kDESIGN CONUITIONSMI = 0.763

1.4 M = 0.529= 46.8 DEGREES

02 = 90.0 DEGREESTURNING = 43.2 DEGREES

1.2 AVDR = 1.124

1.0

0.8

v 0.2 0.4 0.6 0.8 1.0

, X/AXIAL CHORD

Figure 5 Supercritical Cascade Design Surface Mach NumberDistribution

U4


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SLOTTED SUCTION

FLEXIBLE ENDWALLOZZL LOCATION OFS[ ]__ : ' S TAT IC TAPS

FLOW.. ;.. -- .... :,.UPPER

I TA,;LBOARD

COMBINATIONSLOTS FOR SLOT / FLAP PROBESIDEWALL LN IO

BOUNDARYINJECTION

LAYER SUCTION

Figure 7 Schematic of DFVLR Transonic Cascade Test Sect ion

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1.4 DESIGN CONDITIONS

M1 = O.763M2 = 0.529

1.2 01 = 46.8 DEGREES1 2 = 90.0 DEGREESTURNING = 43.2 DEGREES

0 AVOR = 1.124

1.0

U S

0.8

0 0 0 0,

0.4

SOLID: DESIGN

M2 SYMBOL: TEST DATA FOR POINT 72

0 0.2 0.4 0.6 0.8 1.0

X/AXIAL CHORD

Figure 8 Comparison of Design Mach Number Distt ibution and theMeasured Test Data.

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0.20

0 18 -

0.16 -

0.14 -

CASCADE LOSS 0.10 -_________

APT00

P T1 -p 0.8I_ _ _ ___

0.06

0.04 - -PTV0

0.02 - ~PT. 72

0.4 0.5 0.6 0. 0.8 0

UPSTREAM MACHNUMBERTEST THEORETICAL

DESIGN DESIGN

Figure 9 Cascade Performance -Loss versus inlet Mach number atdesign inlet angle.

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0.20

0.18 -

0 16

0.14

0.12

CASCADE LOSS 0.10

PT - 0.08 -PTI PI

0.06

0.04

0.4 0.5 0.6 0.7 0.8 0

UPSTREAM MACH NUMBER

Figure 11 Cascade Performance - Loss versus inlet Mach numberat+5 degrees incidence.

28


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I

I

0.20

0.18 -

0.16

0.14 -

0.12

CASCADE LOSS 0.10P T

0.08

T1 0 06

-

0.04

0 02 -PT.42

0

0.4 0.5 0.6 0.7 0.8 0.9


Figure 12 Cascade Performance - Loss versus inlet Mach number at

+3 degrees incidence.

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0.20 -

0.18 -

r ~~~~~0 16 - ____ ____ ____ ____

v14 -

12

CASCADE LOSS 0.1

P TI -P 1 0.08 - __ _____

0.06 - _ _ _ _ _ _ _ _ _ _ _

0.04 -_____

0.02 PT. 47

0.4 0 5 0 6 0 7 0 8 0 9

UPSTREAMMACH NUMBER

Figure 13 Cascade Performance - Loss versus inlet Mach number at-3 degrees incidence.

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7.

0.20

0.18

0.16

0.14

0.12CASCADE LOSS T 0.10

0.08

0.06 _

0 .0 4. -

0 .0 2 - C - _ _ _P_ 3 2

0.4 0.5 0.6 0.7 0.8 0.9


Figure 14 Cascade Performance - Loss versus inlet Mach number at-5 degrees incidence.

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0.20

0.18

0.16

0.14 rCHOKE

0.12

CASCADE LOSS 0.100.08

PTI - P1

0.06

0.04

0.02 - PT. 38

0.4 0.5 0.6 0.7 0.8


Figure 15 Cascade Performance - Loss versus in le t Mach number at-10 degrees incidence.

32

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0.20 _ISYM. MI UPSTREAM

0.18 Q 0.4 ENGINE0o.s AXIS

0.16 - 0.7

I0 0.75S 0,14I-

0.12

0.10-

wa

0 08

o.06-\.04-

0.02

58 56- 54 52 50 48 I46 44 42 40

DESIGN

UPSTREAM INLET ANGLE, 1

I I II I I I I I I-10 -8 -6 -4 -2 0 +2 +4 +6 +8

INCIDENCE ANGLE

Figure 16 Cascade Performance - Loss as a function of cascadeinlet angle or incidence for various upstream Machnumbers .

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52

SYM MI UPSTREAM5o- 0 0.4

0 0.6

4 8 - - 0.7

0 0.75

46-

S 4 4 -

422Sz 42 - 00 0z

S40

38

36_ENGINE

34AXIS

32

r. 86 -

W 92--

58 56 54 52 50 48 46 44 42 40DESIGN

UPSTREAM INLET ANGLE,31

Figure 17 Cascade Performance - Flow tu rn ing and exi t flow angleas a f u n c t i o n o f inlet angle fo r va r ious ups t r eam Machnumbers.

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TEST CONDITIONS

M = 0.7354M u 0.5604

Sx= 47.0 DEGREES-= 0.84 DEGREES

0 IURNING - 43.84 DEGREESAVDR = 1.1734

~ E d3 .0

0.8

Symbol: Test data

0.2

S.1I I I I I I I I I0 0.2 0.4 0 1 0. 1.0

X/axial chord

Figure 18 Results of Analytical Simulation of Near Design TestCondition (Point 72)

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TEST CONDITIONS

M1 = 0.65661.4 M2 = 0.5682

01 -57.0 DEGREES02 - 90.04 DEGREES

1 2 TURNING a 33.04 DEGREESAVDR = 1.08780

1,0 0

/0S0.6 0U0

m

U

,

S0 .6

Solid: CalculationSymbol: Test data

0.2

Ol I I . I II .I I I I

o 0.2 0.0. 0. 1.

X/axial chard

Figure 19 Results of Analytical Simulation of Ten Degree Negative

Incidence Test Condition (Point 38 )

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REFERENCES

1. Dunavant, J. C. et. al. High Speed Cascade Tests of the NACA 65 -(12A10)10 and NACA 65-(12 A2 18 b)10 ompressor BladeSect ions , NACA RML55, 108.

2. Whitcomb, Richard T., and Clark, Larry R., "Aa Airfo i l Shape forEff icien t Fl ight ar Supercr i t ical Mach Numbers, NASA TMX-1109,May 1965.

3. Bauer, Francis, Garabedian, Paul, and Korn, David, Supercr i t i ca lWing Sections, Lecture Notes in Economics and MathematicalSystems, Vol. 66, Springer - Verlag, New York, 1972.

4. Bauer, Francis, Garabedian, Paul, Korn, David, and Jameson,Antony, Supercr i t ical Wing Sections II Lecture Notes inEconomics and Mathematical Systems, Vol. 108, Springer - Verlag,1975.

5. Bauer, Francis, Garabedian, Paul, and Korn, David, Superci r i t ical

Wing Sections III," Lecture Notes in Economics and MathematicalSystems, Vol. 150, Springer - Verlag, New York, 1977.

6. Korn, David, Numerical Design of Transonic Cascades, ERDAResearch and Development Report COO-3077-72, Courant Inst. Math.Sci., New York Univ., January 1975.

7. Stephens, Harry E., Application of Supercr i t i ca l Airfo i lTechnology to Compressor Cascader: Comparison of Theoretical andExperimental Resu l ts , AIAA Paper 78-1138, AIAA Fluid and PlasmaDynamics Conference, Seat t le , Wash., July 1978.

8. Ives, David C., and Liutermoza, John F., Analysis of TransonicCascade Flow Using Conformal Mapping and Relaxation Techniques,

AIAA Journal, Vol. 15, No. 5, May 1977, pp. 647-652.

9. Ives, David C., and Liutermoza, John F., Second Order AccurateCalcu lat ion of Transonic Flow Over Turbomachinery Cascades, AIAAPaper 78-1149, AIAA l1th Fluid and Plasma Dynamics Conference,Seat t le , Wash., Ju ly 1978.

10. McNally, W. D., Fortran Program for Calcu lat ing CompressibleLaminar and Turbulent Boundary Layers in Arbi t rary Pressu reGrad ien ts , NASA TN D-5681, May 1970.

11. Lieblien, S., and Rouderbush, W. H., Theoretical Loss Relationsfor Low Speed Two-Dimensional Cascade Flow, NACA TN 3662, March1956.

37


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WN 'a 0 C4NN

* 4 .4, N 1 ~~ ~ 4 6 o--C4

14 - 0N .0 0 N .4, 3 N N 0 % ~ .4 NA

4p, o

A0 N - ft--

N-6. 46 * % - - -C ' - 3 3 3 0

0% . . .

-u - 30 IV 10 10 1 0 3 3 3 3 C

W. a

4. a. - - - - - -

4 .

c ~6L,

~t . N 0 00 0 3 3 0 0 C 0 0 0 C 3A

~3


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-0 10T 4

4L Co

N0v

. . ~ 0 40 N 0

400

I~ ~ ~ g1 d0N

CCV

40-


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LIST C7 SYMBOLS, ABBREVIATIONS, AND SUBSCRIPTS

jnbols and Abbreviations

A stream tube area

a sound speed

AVDR axial veloci ty densi ty ra t io = stream tube in le tarea/ex i t area, AI/A2

bx a i r fo i l axia l chord

C irfoil chord

Cp s t a t i c pressure r i s e coeff ic ient(PS2-PS|) /(PTI-Pl)

DF diffus ion fac tor

- (1 - W /W1 ) + T/C W co P W2 cOs 2 j2 Wl

M Mach number w W/a

P static pressure

PT t o t a l pressure

Re Reynolds number based on in le t ve loc i ty and chord = WC v

U, V ve loc i ty components along x, y

W veloc i ty

x, y rec tangu la r coordinates in axial and tangentialdi rec t ions

loss coeff ic ien t = (PTZ-PT1)/(PTI-PI)

S9 surface chord angle

a i r angle measured from tangential

flow turning, p2-pl

cascade gap

kinematic viscos i ty

41


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Washington, D.C. 20361, Attn: AIR-5360

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