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83-GT-211
DEVELOPMENT OF CONTROLLED DIFFUSION AIRFOILS FORMULTISTAGE COMPRESSOR APPLICATION
A series of Controlled Diffusion Airfoils has beendeveloped for multistage compressor application. Theseairfoils are designed analytically to be shock free attransonic Mach number and to avoid suction surfaceboundary layer separation for a range of inlet condi-tions necessary for stable compressor operation. Theyhave demonstrated, in cascade testing, higher criticalMach number, higher incidence range, and higher loadingcapability than standard series airfoils designed forequivalent aerodynamic requirements. These airfoilshave been shown, in single and multistage rig testing,to provide high efficiency, high loading capability,and ease of stage matching, leading to reduced develop-ment costs and improved surge margin. The ControlledDiffusion Airfoil profile shapes tend to have thickerleading and trailing edges than their standard seriescounterparts, leading to improved compressor durability
INTRODUCTION
This paper reports the development of a new compressorairfoil type which is being used to replace most stand-ard series airfoils in Pratt & Whitney Aircraft advancedcompression systems such as the PW2037 high and lowpressure compressors. Standard series airfoils, suchas the NACA 400*, the NACA 65/Circular Arc, or theDouble Circular Arc airfoils, which have been used inmost current production compressors, were adaptations,for compressor application, of shapes which weredeveloped and optimized for isolated airfoil applica-tion, or were empirically derived shapes. Their devel-opment preceded the availability of analytical pro-cedures capable of predicting transonic cascade aero-dynamics.
Controlled Diffusion Airfoils are designed and optimizedspecifically for subsonic and transonic cascade appli-
cation. Through control of the diffusion on theairfoil suction surface, significant boundary layerseparation can be avoided over the entire range ofairfoil operation. For transonic applications, diffu-sion from supersonic to subsonic local surface velocitycan be accomplished without developing shock waves. Afamily of airfoils, capable of achieving these condi-tions, has been developed which can be described by asmall number of geometric parameters.
Experimental testing of a number of Controlled Diffu-sion Airfoil cascades, between 1978 and 1981, hasdemonstrated the capability of achieving low loss atelevated Mach number, increased incidence range atelevated Mach number, high loading levels, and thickerleading and trailing edges without performance penalty.These qualities can be utilized to achieve high com-pressor efficiency, fewer compressor blades and vanes,improved stability, improved durability and reduceddevelopment cost.
The development of this airfoil family required priortechnology improvements in analytical techniques andin cascade test procedures. Computer programs havebeen developed which are capable of efficiently com-puting transonic cascade aerodynamics with varyingstreamtube area and radius change. This capabilityhas been combined with viscous boundary layer analysesto predict laminar, transitional, and turbulentboundary layer development and potential separation.These programs have been supplemented by empiricalcorrections to analytical turning, supplied by an ex-tensive cascade test program. All of these capabili-ties have been combined in an extremely rapid computerinteractive direct design system, capable of designinga Controlled Diffusion Airfoil for a specific require-ment, and predicting its off-design performance.
*NACA 400 - NACA four digit airfoil with lccation ofmaximum camber fixed at 401.1 chord - NACA XLXX.
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DISCUSSION
Background
The abrupt increase in loss and decrease in turning ofa cascade of compressor airfoils, as incident Machnumber increases, is analagous to the transonic dragrise experienced by isolated airfoil sections. As Machnumber increases, local patches of supersonic flowappear on airfoil surfaces, usually terminated by shockwaves. The total pressure loss of the shock plus thelosses caused by shock induced boundary layer separa-tion combine to limit the iow loss Mach number rangeof a cascade of airfoils. Tc improve compressor effi-ciency, airfoil sections usually have been selectedhaving thinner leading edges at higher Mach numbers,in order to increase the transonic drag rise Mach number,Figure 1. However, this selection usually results indecreased useful incidence angle range (loss coeffi-cient less than .04) and decreased durability, Fig. 2.
Whitcomb, et al. (Ref. 1), in 1965, demonstrated ex-perimentally the existence of shockless supercriticalflows for isolated airfoil sections. By developingshapes which could diffuse from locally supersonicvelocities to subsonic velocities without shock waves,Whitcomb substantially increased the drag rise Machnumber of his supercritical wing sections over exist-ing airfoil sections of similar thickness. This cap-ability also suggested that these shapes could befurther optimized to reduce boundary layer drag, sincethe absence of shock waves would permit potential flowanalytical solutions to be employed in conjunctionwith standard boundary layer procedures. Bauer,Garabedian and Korn (Refs. 2, 3, 4) provided analyticaldesign procedures for supercritical wing sections which,through solution of the potential equation in the com-plex hodograph plane, permitted an airfoil shape to bederived from a specified shockless surface Mach numberdistribution.
Feasibility Demonstration
The current study was motivated by a desire to deter-mine if shockless transonic airfoils could be designed
65 CA
400 SERIES
CA
i/
LEADING EDGE CONTOURScA/
010
400 65: CA/
U
LL 0.050UT0
003 04 0.5 0.6 0.7 OR 0.9 CO
INLET MACH NUMBER
Figure 1 To achieve low loss as Mach number in-creases, standard series airfoils arechosen having thinner leading edges.
for compressor application, thereby permitting morerobust airfoils to be designed for higher Mach numbersor improving the compressor efficiency potentialthrough increases in the critical Mach number. In 1974,Korn (Refs. 5, 6) extended the previous isolated airfoildesign procedure to cascade airfoils. In a cooperativeprogram with Pratt & Whitney Aircraft, a supercriticalcascade section was designed by Korn in 1974. It wastested in the transonic cascade facility of the DFVLR(Deutsch Forschung and Versuchsanstalt fur Luft andRaumfahrt) in 1976 as reported by Stephens (Ref. 7).The test verified the shockless nature of the flowfield and the attached viscous boundary layer. Theairfoil exhibited excellent aerodynamic performanceat off-design conditions as well as at its designpoint, indicating that shocks developing at off-design
NOMENCLATURE
Symbols t airfoil maximum thickness
AVDR axial velocity density ratio - streamtube V velocityinlet height/exit height, Hl/H2
air angle measured from tangentialB airfoil chord
kinematic viscosityBX airfoil axial chord
solidity, B/TDF diffusion factor
V1cos 1 - V2cos ^,2(1 - V2/Vl) + (T/B)(
r cascade gap, pitch
2 V1loss coefficient = (PT1 - PT2)/(PTl - Pl)
H streamtube height normal to blade-to-bladeSubscripts
plane
M Mach number 1 cascade upstream inlet plane
P static pressureLE cascade airfoil leading edge plane
PT total pressureTE cascade airfoil trailing edge plane
Re Reynolds number based on inlet velocity and2 cascade downstream exit plane
chord = VB/x axial direction
2
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20 400._
wIC7 15
0
z10
Uz0UzL0 5
N7
003 04 0.5 0.6 0.7 0.8 0.9 1.0
INLET MACH NUMBER
Figure 2 Higher Mach number standard series air-foils have reduced incidence range.
conditions were relatively weak and did not significant-ly degrade the airfoil's performance. This was import-ant for compressor airfoil application since compressorairfoils must operate stably and efficiently over arange of incident Mach number and flow angles.
Although these test results demonstrated the feasibilityof supercritical compressor airfoil design, the Kornhodograph procedure did not represent a practicalcompressor airfoil design capability for several reasons;the airfoils designed tended to have undesirable struc-tural properties, the program could not compute at highsolidity, and it did not permit streamtube thickness
or streamtube radius to vary through the cascade, asoccurs for most compressor airfoils. In addition,since it was strictly a design procedure, it could notbe used to estimate the off-design capabilities of theairfoils designed. Recently, Sanz (Ref. 8) has reducedthe limitations on solidity of the Korn program.
First Practical Demonstration
Ives and Liutermoza, (Refs. 9, 10) developed a pro-cedure in 1975 which was capable of calculating tran-sonic cascade flows at practical solidities with vary-ing streamtube thickness, and, because it could calcu-late weak shock waves, could be used to calculate off-design performance. When combined with a boundarylayer calculation (McNally, Ref. 11) and a coordinatesmoothing routine in a computer interactive designprocedure, it provided a means of designing durable,structurally sound cascade airfoils capable of meetingshockless transonic design point conditions for turn-ing at minimum total pressure loss, as well as insuringa reasonable range of efficient, stable off-design per-formance. Criteria were then developed to describeairfoil surface Mach number distributions to bestachieve the aerodynamic design objectives (Fig. 3):
A continuous acceleration from the leading edgeto the peak Mach number on the airfoil suctionsurface, to avoid premature laminar boundary layerseparation or transition.
2. A peak Mach number less than 1.3 to avoidboundary layer separation which could be in-duced by a severe shock wave boundary layerinteraction, should a shock develop at off-design conditions.
3. A continuous, shock-free deceleration from thepeak suction surface Mach number to the trailingedge, maintaining a turbulent boundary layer witha low level of skin friction and avoiding separa-tion ahead of the trailing edge, and
A nearly constant subsonic Mach number distributionon the pressure surface.
CONTINUOUS ACCELERA I IONTO BOUNDARY LAYER TRANSITION POINT
The Ives & Liutermoza procedure was used to design afan exit guide vane section, under Naval Air SystemsCommand Contract N00019-77-C-0546 (Ref. 12), to pro-vide a demonstration of a supercritical cascade air-foil satisfying practical compressor aerodynamic andstructural requirements. The design requirements werethose of the mean section of the exit guide vane of a1600 ft/sec tip speed fan, tested previously underNASA Contract NAS3-10482 (Ref. 13). The design object-ives were inlet Mach number of 0.76, inlet flow angleof 46.8 degrees, turning of 43.2 degrees to an axialdischarge angle, and a streamtube contraction ratio(AVDR) of 1.124. The design iteration yielded a cas-cade having a solidity of 1.43 and a diffusion factorof 0.54, (Fig. 4). The airfoil was tested in theDFVLR transonic cascade facility in 1978. For compari-son, a multiple circular arc airfoil (MCA) (Fig. 5)was designed to meet the same aerodynamic requirements,resulting in a solidity of 1.64 and a design diffusionfactor of 0.51. It was also tested at the DFVLR cas-cade facility.
The design point surface Mach number distribution isdepicted in Fig. 6, incorporating the features des-cribed above. Superimposed is the experimental Machnumber distribution derived from static pressure tapson the airfoil surfaces corresponding to one gap, i.e.,the suction surface taps and pressure surface taps wereon consecutive airfoils. It can be seen that the air-foil closely matched its design Mach number distribu-tion. In this figure, and in all subsequent figures,inlet Mach number is adjusted from the upstream measure-
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/ B / 0EOME1RIC D41A
/ 6 2 /5 1 N
/- H 0 /0
_- x1 a2J
t B 0 01
Figure 4 Supercritical Fan Exit Guide Vane CascadeGeometry
II
10
08
06
M
W
Do t
oA 112 04 0.6 0.8 1.0
X BX
Figure 6 Supercritical Exit Guide Vane Design Experi-mental Mach Number Distribution CloselyAchieved Design Distribution
GEO ME'R^:, DAT A:
B 2 15 iN
B - 0 61
639
B - 0 0E
Figure 5 Comparison Multiple Circular Arc Fan ExitGuide Vane Geometry
ment station to the plane of the cascade leading edgeby accounting for streamtube contraction due to sidewall boundary layer growth. The analysis of the staticpressure data at several test points, used to arriveat this adjustment, is described at length in Ref. 12.The loss of the cascade as a function of inlet angleand Mach number is illustrated in Fig. 7. The airfoil
maintained low loss and excellent incidence angle rangeuntil its design Mach number was exceeded. The minimumloss as a function of Mach number is depicted in Fig.8, compared to that achieved for the Multiple CircularArc cascade. The supercritical exit guide vane sectionhad significantly lower loss up to its critical Machnumber. As illustrated in Fig. 9, the supercriticalexit guide vane achieved and maintained its design exitangle, with little deviation, over a wide range ofincidence angle and Mach number. The comparison MCAairfoil experienced significant underturning at designincidence angle and all positive incidence angles.Similarly, the incidence range of the supercritical exitguide vane section was significantly higher than theMCA section, until design Mach number was exceeded,Fig. 10. For this comparison, incidence range wasdefined by the airfoils reaching an absolute loss levelof .04 at each side of the loss bucL:et.
The supercritical fan exit guide vane test, besidesdemonstrating that a supercritical cascade sectioncould be designed to meet practical compressor airfoilrequirements, also demonstrated additional desirableaerodynamic and structural benefits. The cascade hadexcellent low Mach number performance due to theattached boundary layer design. It had excellent inci-dence range at all Mach numbers up to critical Machnumber, demonstrating that the design rules adoptedavoided strong shock wave boundary layer interactionsat off-design conditions. The airfoil had littledeviation in turning angle over its useful incidenceand Mach ranges, also due to its attached boundarylayer design. The airfoil's thick leading and trailingedges, relative to comparable high Mach number airfoils,promised improved durability. This successful testprompted the further development of airfoils obeyingthese design criteria for application to most compressorairfoil requirements from low to transonic Mach numbers.
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SYMBOL MLE-
016 • 0.43
0630.16
A 0 73
0.14 • 0.78
0.12
0.10
U 008
O 0.06
aoa ^^
0.02
58 56 54 52 50 48 1 46 44 42 40 38
DESIGN
INLET ANGLE, 13i
Figure 7 Supercritical Exit Guide Vane achieved lowloss and wide incidence range at all Machnumbers below critical.
XYMF30L MLE
O 0 4
❑ 063MOO
"; ^i O 1, ^' 9
Z /Q
u 4/1
SU PERCRIT',C ,L Ag2 EXIT GUIDE V l,^.F
58 5e 54 1)2 50 n8 1 46 94 12 40 38
INLET ANGLE, „I
Figure 9 Supercritical Exit Guide Vane maintained itsdesign exit angle over a wide range of in-cidence angle and Mach number. The compari-son MCA design had significant deviation atdesign incidence and positive incidenceangles.
Development of a Series of Controlled Diffusion Air-foils - Design
A series of eight airfoils of this type, designatedControlled Diffusion Airfoils "A" through "H" weredesigned and tested from 1978 to 1981, covering therange of airfoil requirements for an advanced multi-stage compressor. In addition, several standard seriesairfoils (NACA 400, NACA 65/CA, DCA), designed foridentical aerodynamic requirements, were tested fordirect comparison of their performance to the Con-trolled Diffusion Airfoils.
In order to design this large number of ControlledDiffusion Airfoil sections, and to provide a practicalengineering design procedure for the hundreds of sec-tions required for a typical high pressure compressor,improvements were made to the analytical procedures.The improved design calculation utilized the Caspar,Hobbs, and Davis finite area solution of the compres-sible potential equation, (Refs. 14, 15, 16). This
0.04 Q SUPERCRITICAL EXIT GUIDE VANE
!:: :
NiC
Z XU PE RU RITICAL
0.01 EXIT GUIDE VANE
004 0 0.6 0.7 1 0.8
DESIGNLEADING EDGE MACH NUMBER, MLE
Figure 8 Minimum loss of the Supercritical Exit GuideVane was significantly below that of theComparison Multiple Circular Arc design.
Q
w 1
(3
Q
U
OU
13
LEADING EDGE MACH NUMBER, MLE
Figure 10 The Supercritical Exit Guide Vane hadsuperior incidence range to the MCA airfoil.
procedure performs its computations in the physicalcascade plane and avoids the complexity of conformalmapping, used in the Ives program, and its attendantdifficulties at high solidity. The boundary layercalculation incorporated an improved laminar/turbulenttransition calculation (Dunham, Ref. 17), a laminarseparation and reattachment calculation (Roberts,Ref. 18), and a lateral streamtube contraction effectusing the Mangler transformation. The inviscidanalysis, the boundary layer calculation and empiricallymodified deviation corrections, were combined in anextremely fast computer interactive design system whichrequires only 13 CPU seconds on the IBM 3081 computerper cascade analysis.
Since the design method is a direct iterative procedure,speed of convergence is dependent upon the quality ofthe initially guessed airfoil contour. An analysis ofseveral converged designs permitted development of ageometric construction procedure which, for most cases,satisfies all the above aerodynamic and structuralcriteria with the initial airfoil contour. This con-struction synthesizes the meanline and thickness dis-tribution of a Controlled Diffusion Airfoil from a
5
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PROBETRAVERSE
PLANES
--
SLOTFLOW
distribution of circular arcs and ellipses. The para-meters for these curves are functionally related to theaerodynamic requirements. The current design procedureselects an airfoil, using this construction, and thenutilizes the analytical program to verify the requiredrange of off-design performance. Modifications aremade to the initial parameter selection if added in-cidence or Mach range is required.
ADJUSTABLEBE LLMOUTH
SECTIONCASCADE
ASSEMBLY WH HSLOTTED ENDWALLS
SCOOP FLOW / FOR BLEED FLOW
TAILBOARDS
Development of a Series of Controlled Diffusion A ir-foils - Cascade Test Procedur e
The eight Controlled Diffusion Airfoils designed usingthis procedure and their comparison standard seriesairfoils were tested in the United Technologies Re-search Center High Speed Cascade Tunnel from 1978 to1981 (Ref. 19). This facility (Fig. 11) has beenutilized extensively for compressor airfoil testingsince 1967, and is the latest in a series of cascadetunnels which have been used by Pratt & Whitney Air-craft to provide compressor airfoil design correlationssince 1944. This facility, which has test capabilitiessimilar to the DFVLR cascade tunnel, is an open circuittunnel which uses pressurized air exhausting throughthe test section to atmospheric pressure. It is capableof testing airfoils of four-inch span and two-inchchord over a range of inlet Mach number from 0.4 tochoke and over a Reynolds number range from Re = 400,000to Re = 800,000, satisfactory for most compressor air-foil requirements.
Tunnel side wall boundary layers are removed by perfor-ated plates upstream of the test section, and top andbottom wall boundary layers are removed by a scoop onthe top wall and slots on the bottom wall. Tailboardsare used to control periodicity within the cascade pack.Axial velocity density ratio, AVDR, is varied using asystem of side wall suction slots located within thecascade pack, aft of the minimum pressure region of theairfoil flow field, to avoid recirculation, Fig. 12. Acompletely independent suction system is used to removecorner boundary layers through slots at the intersectionof the airfoil suction surface with the side walls, toprevent spanwise flow of boundary layer fluid.
Airfoils are set in at constant stagger; the numberof airfoils depending on stagger, gap chord ratio andinlet angle. The entire cascade is supported betweendrums which are rotated relative to the inlet sectionto achieve a desired inlet angle. Test procedure isto vary Mach number and AVDR at each inlet angle, andto vary the inlet angle from choke to stall to com-pletely define the incidence and Mach number ranges ofthe cascade airfoil, and its response to streamtubecontraction.
Data acquisition includes an upstream traverse fortotal pressure and flow angle, a downstream traverseover six gaps for exit angle and total pressure, staticpressures on the side walls at the upstream and down-stream traverse locations and at intermediate axialstations, and static pressures on the airfoil surfaces,defining one gap. Data is reduced for each airfoilindividually to eliminate the effects of any residualnonperiodicity in the upstream conditions.
Development of a Series of Controlled DiffusionAirfoils - Cascade Test Results
PERFORATEDSIDEWALLS
INLET AIRFOR BLEED FLOW
DIRECTION
11FIXED
BE LLMOUTHSECTION
SIDEWALLQUADRANT
Figure 11 Schematic of the United Technologies ResearchCenter High Speed Cascade Tunnel.
Figure 12 A dual system of slots is used to controlAVDR and corner boundary layers.
Comparison cascade tests were run for a series ofdesign conditions intended to demonstrate the rangeof application for which Controlled Diffusion Airfoilscould offer improved aerodynamic and structural
performance.
Controlled Diffusion Airfoil "D" is compared to a Double
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INLET MACH NUMBER
CONTROLLED DIFFUSION AIRFOIL D
NACA 65. CAi30)09
1U
'_ 0.04
00)
00NO
0.02
■ CONTROLLED DIFFUSION AIRFOIL E
• NArA 60F
Circular Arc airfoil and to a NACA 65/CA profile cas-cade, both typically used for high transonic Mach numberrequirements. As can be seen in Fig. 13, the profileof the Controlled Diffusion Airfoil is considerably morerobust at leading and trailing edges than the DoubleCircular Arc airfoil and, at the trailing edge, the 65/ wCA airfoil. All three cascades were designed for aninlet Mach number of 0.7, inlet angle of 30 0 , flowturning of 13.6°, AVDR of 1.07, solidity of 0.933, andmaximum thickness 9% of chord. Their relative per-formance is compared in Figs. 14 and 15. At theirminimum loss incidence, Fig. 14, the Controlled Diffu-sion Airfoil had higher critical Mach number thaneither of the two comparison airfoils, despite its more Jrobust profile. In addition, it had superior incidence zangle range to the two comparison cascades at all Mach LLnumbers tested, Fig. 15.
CD
A comparison to a low Mach number airfoil type is illus-trated in Figs. 16, 17 and 18. A NACA 400 series air-foil is compared to Controlled Diffusion Airfoil "E".Both airfoils were designed for inlet flow angle of 30°,flow turning of 15°, AVDR of 1.07, solidity of 0.870,and maximum thickness 5% of chord. The ControlledDiffusion Airfoil, although designed for a high Machnumber requirement, 0.80, exhibited equivalent minimum
Figure 15 Controlled Diffusion Airfoil RD" had superiorincidence angle range at all Mach numbers.
0 10.3 0.4 0.5 0.6 0.7 0.8 09 1.0
INLET MACH NUMBER
DOUBLE CIRCULAR ARC AIRFOIL CAI30)09
Figure 13 Controlled Diffusion AirFoil I'D" has morerobust leading and trailing edges thanstandard series comparison airfoils.
Figure 14 Controlled Diffusion Airfoil I'D" had highercritical Mach number than the comparisonstandard series airfoils when compared attheir minimum loss incidence angles.
Figure 16 Controlled Diffusion Airfoil "E" had highercritical Mach number and equivalent loss atlow Mach number to the NACA 400 series com-parison airfoil.
U)
ftCD
00
zI
Uzo_Uz
LL
03CD
INLET MACH NUMBER
Figure 17 Controlled Diffusion Airfoil "E" had higherincidence range than the comparison NACA 400series airfoil at high Mach number andequivalent range at low Mach number.
loss and incidence range, even at the low Mach numbers(less than 0.5) for which the NACA 400 series airfoilwould be applied, (Figs. 16, 17). The relative loadingcapability of the two airfoils is illustrated in Fig. 18.
7
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For Mach numbers below 0.65, to avoid critical Machnumber problems for the NACA 400 series airfoil, theloading for each airfoil was varied by changing inci-dence angle and by applying suction to the cascadeendwalls, thus lowering AVDR, until the airfoil stalled.The data points obtained during this procedure wereplotted as loss versus 0 factor. For clarity, for eachairfoil the minimum loss data points, at each D factorlevel, were joined by a line and then the points abovethe line were eliminated. The remaining locus of mini-mum loss points, illustrated in Fig. 18, represents thelow loss loading limits for each airfoil. This is some-what analagous to an isolated airfoil lift to dragcurve. It can be seen that the Controlled DiffusionAirfoil was capable of sustaining a much higher loadingthan was the NACA 400 series airfoil.
• CON
• NAC
3
53N 0.03O
Z
LL 002
In
UO
0M
0L0.3 0.4 0.5 0.6 0.7 0.8
DIFFUSION FACTOR, DF
Figure 18 Locus of Minimum Loss - Maximum LoadingPoints. Controlled Diffusion Airfoil "E"had significantly higher loading capabilitythan the comparison 400 series airfoil.
Sinqle Staqe Riq Demonstration
As a demonstration of the applicability of ControlledDiffusion Airfoils to compressor requirements, a statorwas designed and tested for a single stage rig, in-tended to simulate the front stage of a multistagecompressor, under Contract NAS3-22008 (Ref. 20). TheMach number incident upon the stator ranged from 0.69at the tip to 0.88 at the hub, and design turningranged from 330 at the hub to 25° near the tip. Thestator was designed at an aspect ratio of 1.446 and hubsolidity of 1.429. In addition to standard rig per-formance instrumentation, surface static pressure tapswere placed on the stator at two span locations.
Data is presented for the 16% span section which wasdesigned for a Mach number of 0.85. Analysis of therig data indicated that, at its minimum loss point,this section operated at 1° positive incidence relativeto its design incidence, and that the streamtube con-traction from leading to trailing edge was 13% higherthan anticipated. This can be seen by comparing themeasured Mach number distribution to the design Machnumber distribution, Fig. 19. When the rig incidenceangle and streamtube contraction are used in theanalytical cascade calculation of Ref. 16, using
streamtube area change to simulate the effect of radiuschange, the analytical Mach number distribution closelymatches the experimental distribution, Fig. 20.
DE S I G N CONDITIONS
M1 = 0.850M2 0.531
BETA 1 = 42.882TURNING = 31.258AVDR = 1.0965
H1,HLE = 1.0000HLE!HTE = 1.0965HTE!H2 - 1.0000
• •■
■■
0.4 0.6 0.8
X/BX
Figure 19 Experimental Surface Mach number distribu-tion compared to design for a ControlledDiffusion Stator Section,
1 416% span from hub.
DATA ANALYSIS CONDITIO NS
• SUCTION SIDE M1 -0885• PRESSURE SIDE M2 -0616
12 b ,/ BETA 1 =4583
TURNING -'28.19
AVDR - 1.1631
H1!HLE -09398
HLE/HTE - 1.2376
HTE/H2 = 1.0000
i
0.6U •zzI■
3,6L /
0 02 0.4 0.6 0.8 1.0
x /Bx
Figure 20 Analyticai surrace Mach number distributionat rig inlet conditions compared to data fora Controlled Diffusion Stator section, 16%span from hub, 'at leading edge plane Mach 0.85.
1.4DATA
• SUCTION SIDE• PRESSURE SIDE
1.2
•1.0
0.8
Z I■SU
0.6 ■
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No indication of shock wave formation is evident inthe analysis, which would indicate a sharp discontin-uity if a shock were present. The boundary layer cal-culation, Ref. 11, indicated no separation. We concludethat there is also no experimental evidence for shockwave formation or boundary layer separation, based onthe agreement of the analytical and experimental Machnumber distributions. The 16% span section loss andexit angle are presented in Fig. 21. It can be seenthat, even at its high incident Mach number of 0.85,the section has substantial incidence range (over 4 0 )
and little change of exit angle with incidence.
A comparison is made between the performance of theControlled Diffusion Airfoil stator and a MultipleCircular Arc stator tested in the same rig under NASAContract NAS3-20809, ''Study of Blade Aspect Ratio ona Compressor Front Stage,'' Ref. 21. Although the twotests do not strictly constitute a back-to-back com-parison, since the axial location of the two stators,and hence the inlet Mach number, were not identical,by comparing the stators at equivalent inlet Machnumber, a reasonable comparison can be made. In Fig.22, the minimum full span losses of the two statorsare compared as a function of mass average inlet Machnumber, indicating that the Controlled Diffusion Air-foil stator had lower full span loss at every equivalentcondition through its design speed. Fig. 23 presentsthe spanwise distribution of loss for the two statorsat their minimum loss points. For this comparison,the Multiple Circular Arc stator is illustrated at 95percent rotor speed, to equalize its inlet Mach numberto that of the Controlled Diffusion Airfoil stator at100% rotor speed. The superiority of the ControlledDiffusion Airfoil stator is evident for the core flowregion.
70
Q 75
X
W 80
RM
0.05
Hzuo 04
cL) 003
C
002
50 48 46 44 42 40
INLET ANGLE, 31
Figure 21 Performance of Controlled Diffusion StatorSection, 16% span from hub, at Mach number0.85.
• CONTROLLED DIFFUSION STATOR
0.07 A MULTIPLE CIRCULAR ARC STATOR
H
0.06
05 10,So¢ 95
> 0 005 90 100Q U
03%N O 095 100
00.04
R-
Q0)
70 PERCENT DESIGN SPEED
0.03
0.55 0.60 0.65 0.70 0.75 0.80 0.85 0.90
MASS AVERAGE STATOR INLET MACH NUMBER, M1
Figure 22 Minimum Full Span Average Loss of the Con-trolled Diffusion Stator is less than thatof a Multiple Circular Arc Stator tested inthe same Single Stage Rig, at EquivalentInlet Mach numbers, through design speed.
D. 14
0 1 2
0 10
HZ
U 0.08
0,0 0.06
0004
0.02
o f 1 1 1 1 1 1 1 1 1
0 10 20 30 40 50 60 70 60 80 100
PERCENT SPAN FROM HUB
Figure 23 A CUN!Parisun of the Spanwise Loss Distribu-tions of the Controlled Diffusion Stator atits 100% Speed Minimum Loss point and theMultiple Circular Arc Stator at its 95%Speed Minimum Loss Point indicates that, atEquivalent Mach number, the ControlledDiffusion Stator had lower loss over theCore Flow Region.
Multistage Application
As a result of the performance benefits demonstrated incascade and rig testing, the Controlled Diffusion Air-foils, developed in this project, have been incorporatedin the Pratt & Whitney Aircraft Energy Efficient EngineHigh Pressure Compressor, developed under ContractNAS3-20646, and in the compressor of the PW2037 AdvancedTransport Engine, Fig. 24. Both compressors were testedduring 1981. The data available does not permit defini-tive evaluation of the Controlled Diffusion Airfoil per-formance in the multistage environment. However, bothcompressors achieved stable, efficient operation in theirinitial builds.
Li
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A new series of airfoils has been developed for multi-stage compressor application employing advanced tran-sonic cascade computational capability to control thediffusion rate on the airfoil suction surface. Ageometric construction has been developed which, formost applications, provides an initial airfoil shapecapable of achieving design criteria intended toprevent the development of shock waves or boundarylayer separation on the airfoil suction surface andprovide a range of efficient, stable operation.
Cascade testing has demonstrated the feasibility ofachieving shock-free transonic cascade operation, theability of Controlled Diffusion Airfoils to provideincreased incidence range, increased critical Machnumber and more durable airfoil shapes for high tran-sonic Mach number requirements, and their ability toprovide higher loading capability, and equivalent in-cidence range for low Mach number applications. Pre-vention of boundary layer separation provides reducedexit angle deviation and reduced sensitivity of devia-tion to incidence angle.
Single stage compressor testing has demonstrated thatControlled Diffusion Airfoils can provide efficient,shock-free performance, with wide incidence range andinsensitive exit angle deviation in a high Mach numbercompressor environment. Improved incidence range canprovide increased compressor surge margin and ease ofstage matching in a multistage environment, leadingto reduced development cost.
Controlled Diffusion Airfoils have been incorporatedin Pratt & Whitney Aircraft advanced multistage com-pressors.
ACKNOWLEDGMENT
The authors wish to acknowledge John F. Dannenhoffer,III, and David Spear for their contributions to theanalytical procedures, Joseph Lubenstein, Alan Ross,and Brian Robideau for development of the airfoil geo-metric design procedure, William Taylor and John Horanfor their experimental cascade test support, and RoyBehlke and James Brooky for the analysis of the singlestage test results. Part of this effort was supportedby NASC Contract N00019-77-C-0546, technical monitorDr. Max Platzer, and NASA Contract NAS3-22008, technicalmonitor Thomas Gelder.
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