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RM L54I02

NACA E FILE

RESEARCH MEMORANDUM

CASCADE INVESTIGATION OF A RELATED SERIES

OF 6-PERCENT-ТШСК GUIDE-VANE PROFILES

AND DESIGN CHARTS

By James C. Dunavant

Langley Aeronautical LaboratoryLangley Field, Va.

CLASSIFICATION CHANGEО ТОDECIASSJF

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTON

November 26, 1954

CONFIDENTIALT62- 25959

https://ntrs.nasa.gov/search.jsp?R=19930088445 2018-09-09T11:48:44+00:00Z

NACA RM

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

CASCADE INVESTIGATION OF A RELATED SERIES

OF 6-FERCENT-THICK GUIDE-VANE PROFILES

AND DESIGN CHARTS

By James C. Dunavant

SUMMARY

The trend to higher weight flovs in axial-flow compressors pointsup the need for improved inlet guide vanes having high operating limits.A new blade series was designed to operate at near choking inlet Machnumbers and to have near maximum critical Mach numbers by employing highaerodynamic loading of the profiles in the leading-edge region and rela-tively straight trailing edges. Low-speed cascade tests of four bladesections at solidities of 0.75, 1.00, and 1.50 were made for each sectionover a wide range of angle of attack. From these tests, design anglesof attack were selected and are used in a method of determining thecamber for guide-vane turning angles as high as 50°. All turning anglesmeasured at low drag conditions are plotted by the carpet-plotting methodand from interpolation the necessary camber is readily obtainable forconditions other than those tested. A number of simple computationswere made to estimate the high-speed characteristics of the blade series.

INTRODUCTION

In attempting to obtain the maximum possible thrust per unit frontalarea, compressor designers have increased the weight flow and hence theaxial Mach number entering the compressor as much as possible. The axialMach number is too often limited by choking in the guide vanes. Thishas been avoided in some cases by eliminating the guide vanes anddesigning the first rotor for axial entering flow. However, there isstill a need for guide vanes in compressors including high-flow designssince rotor efficiency usually decreases with increasing rotor tip Machnumber in the transonic range. By using guide vanes the rotor tip Machnumber can be lowered for the same weight flow and rotational speed.

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NACA БМ

At axial inlet Mach numbers of 0.5 to 0.7, the guide vanes will Ъе

required to operate at supercritical speeds and at the higher turning

angles the exit velocities will be near sonic. Guide-vane design,

therefore, must consider the presence of shocks on the blades and the

possibility of choking of the flow in the guide-vane passages.

Guide vanes which turn the flow efficiently and correctly have been

designed from two-dimensional cascade data using corrections for the

change in axial velocity across the blade row such as found in refer-

ences 1 to 3- In. the past, guide-vane design has largely consisted ofthe use of either circular-arc mean lines with arbitrary thickness

distributions or the airfoil data obtained by Zimney and Lappi (ref. k).

An example of the former is reference 5 which presents a method of

designing circular-arc sheetmetal guide vanes which for many applica-tions is satisfactory. However, none of these methods is entirely

suitable for operations at the speeds considered necessary herein.

Moreover, while the low-turning-angle data of Zimney and Lappi aresystematic, their higher turning-angle data are not systematic and

cannot be interpolated to other cambers.

The purpose of this investigation was to develop a related series

of profiles designed specifically for guide-vane application at highsubsonic speeds. This paper presents design information for a new series

of 6-percent-thick blade sections from results of low-speed cascade tests

of these sections over a range of turning angles up to 50°, solidities

of 0.75, 1.00, and 1.50, and design lift coefficients from 0.6 to 2Л.

The analysis .of results includes extrapolation of the pressure-

distribution data by the method of reference 6 to get an indication of

the high-speed characteristics, including the critical and choking Mach

numbers.

This paper also presents some new turning-angle data for the NACA

65-(12)10 guide vanes, for which data were originally presented by

Zimney and Lappi. These new data were obtained with improved cascade

testing techniques.

SYMBOLS

A area

с blade chord

Cd section drag coefficient

С 7 camber, expressed as a design lift coefficient of isolated airfoil

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NACA RM L5IH02 CONFIDENTIAL

Cw wake momentum difference coefficient

M Mach number

p static pressure

P total pressure

PR resultant pressure coefficient: difference between local upperand lower surface pressure coefficients

q dynamic pressure

s tangential spacing between blades

S pressure coefficient, —ll

V velocity

x chordwise distance from blade leading edge

у perpendicular distance from blade chord line

a angle of attack, angle between inlet flow direction and bladechord

3 air angle between flow direction and perpendicular to blade row,deg (see fig. l)

a solidity, chord to spacing ratio (c/s)

8 flow turning angle, deg

Subscripts:

1 upstream of blade, row

2 downstream of blade row

Т throat

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BLADE SECTION DESIGN

For high weight-flow compressors, the axial velocities entering theguide vanes may exceed the critical speed of the blade profile and insome cases the exit velocities may even be sonic. Entering Mach numbersthat produce sonic exiting velocities are illustrated in figure 2(a) asa function of turning angle for nonviscous one-dimensional flow throughcascades of infinitely thin blades. Figure 2(b) shows the enteringMach number for which the blade passage chokes with no turning as afunction of blade thickness and solidity.

The basic requirement of a high speed quide-vane profile is goodaerodynamic performance over a range of solidities and turning angles formaximum speeds before choking. This means that (l) premature choking ofthe flow in the guide-vane passages must be avoided, and (2) airfoilprofiles must be shaped to avoid shock waves so as to minimize shock-boundary-layer interaction effects and thereby avoid separation andkeep wake losses low. Only with data obtained in tests of profileswhich meet these requirements can the flow into the first stage of thecompressor be properly matched.

To arrive at suitable guide-vane profiles, a mean line and thicknessdistribution system of reference 7

was followed. In selecting a mean

line and hence the chordwise loading, the change in stream velocitythrough the passage was the primary consideration. In guide-vane appli-cations the entering velocity is lower than the leaving velocity and,therefore, the leading-edge region can be more highly loaded than thetrailing-edge region and thereby delay supersonic velocities on the bladeupper surface. This type of loading accomplishes most of the turningearly and permits the use of lightly loaded, nearly straight trailingedges. The lightly loaded trailing edges are desirable for less sensi-tivity of turning angle to variations in boundary layer, Reynolds number,turbulence, or surface finish.

The mean line used in this related series of guide-vane profiles isa combination of the a = 0 and the a = 1.0 mean lines of reference 7«The a = 0 mean line concentrates the loading near the leading edge. Toremove the reflex curvature near the trailing edge which is characteristicof the a = 0 mean line, the uniform loading of the a = 1.0 mean linewas added in the proportion of h- to 6. The resulting mean line is desig-nated the Alj.Ko' mean line. This designation follows a system developed

specifically for compressor and turbine profiles. (See the appendix ofref. 8.) In this system, the mean lines from a = 1.0 to a = 0 areidentified by letters of the alphabet for each increment of 0.1, and thesubscript indicates the proportions of the individual lift coefficients.Thus, the АьК> mean line is the a = 1.0 mean line (A) with a

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MCA RM L54I02 CONFIDENTIAL 5

Ст = ОЛ combined vith the a = 0 mean line (K) with C? = 0.6.

The ordinates for this combined mean line for Cj =1.0 are given in

table I.

To obtain high critical and choking speeds, it is necessary to havenot only a thin blade section, as seen from figure 2(b), but also a thick-ness distribution with the maximum thickness located well forward topermit the highest possible Mach number before choking, since turning theflow from the axial inlet direction contracts the passage toward the rear.These requirements are satisfied by the NACA 63-006 airfoil thicknessdistribution (ref . 7) whose maximum thickness is located at 35 percentchord. The ordinates of this basic section were modified to give athicker trailing edge by fairing a straight line from the trailing-edgeradius, equal to 1/10 the maximum thickness, tangent to the upper andlower surfaces. Ordinates for this profile are given in table II. Theairfoil series evolved is designated the NACA бЗ-СС^-А^Кб^Об guide-vane

profiles .

APPARATUS AND PROCEDURE

Apparatus

The test apparatus was the high-aspect-ratio cascade tunnel describedin reference 9> modified by reducing the test section width from 20 inchesto 10 inches. Five-inch chord blades were used giving an aspect ratioof 2.0. Seven blades were used in the cascade except in the case of thelowest solidity tests for which only five blades could be fitted into thetunnel. The side walls in the entrance to the test section contained anupstream boundary-layer suction slot. Solid test section side walls wereused in all blade tests since reference 9 has shown that wall boundary-layer buildup through cascades accelerating the flow is insufficient torequire the use of porous walls.

Five-inch chord models were constructed of four blade cambers,CI

Q = 0.6, CI

Q = 1.2, CI

Q = 1.8, and CI

Q = 2Л. Two of the seven-

blade sets were constructed of plastic and two of metal. The centerblade of each cascade contained a row of 22 static-pressure orificesdistributed over the upper and lower surfaces at midspan.

Tests

The upstream flow angle was measured approximately one chord lengthupstream of the cascade at the tunnel center by means of a removablenull-type probe. The downstream angle was determined by averaging eight

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angular measurements made approximately one-half chord length downstreamat the midspan of the blades. Two measurements were made behind each ofthe four centermost passages. These measurements were made in a regionin' which no total pressure loss had occurred.

Blades of each camber were tested at three solidities, a = 0.75,1.00, and 1.50. Blade cambers were tested through a range of angles ofattack in k° angle-of-attack increments, the limits of the angle-of-attack range tested being determined by drag rise (drag coefficientvalues greater than twice the minimum drag) occurring at very high andvery low angles of attack.

Turning angles were measured for a typical guide-vane cascade atReynolds numbers from 260,000 to 100,000 to determine whether criticalReynolds numbers existed. Variation of the turning angle with Reynoldsnumber in the range tested was 0.7°; hence, the test results for practicalpurposes were assumed to be unaffected by Reynolds number.

Because of the large variation in pressure ratio across the variouscascades and the changes in upstream area, the blower was not able tosupply either the pressure or the volume required to maintain a constantentering velocity for all the tests. Hence, the tests were run at max-imum output of the blower and the resultant Reynolds number based on theupstream velocity and the 5-inch chord varied among tests from a minimumof 208,000 to a maximum of 29 ,000.

Blade forces were obtained using the method of reference 10. Normalforces determined by two methods, momentum change and integrated pressuredistribution, were used as a check on testing technique accuracy. A max-imum variation of 5 percent was permitted for acceptable tests except forconditions of low lift where a difference in coefficients of 0.05 waspermitted.

RESULTS AND DISCUSSION

Aerodynamic Characteristics

Pressure distributions.- Low-speed cascade pressure distributions andblade section characteristics are shown in figures 3 to 14. As would beexpected from the mean-line design, the pressure distributions show highestvelocities over the upper surface as well as highest loading in the leading-edge region. Generally, the velocity along the upper surface decreasesfrom the leading-edge region to the trailing edge. Variations in solidityand camber change these characteristics only in altering the overallloading. An exception to this occurred in the test of the NACA 63-(2УЦ.К£)Обsection at a solidity of 1.5 where highest turning angles of the tests were

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NA.CA RM

measured. In this instance the exiting velocity approached that of the

peak leading-edge region surface velocity; hence the velocity and pres-

sure along the upper surface are almost constant. (See fig. 1 .)

For many pressure distributions the pressure recovery over the

upper surface was accompanied by a small laminar bubble located betweenthe kO- and 70-percent-chord stations. The bubble usually moved toward

the leading edge as the angle of attack was increased. Flow over the

leading edge was evidently laminar, separated at the bubble (indicatedby the flattening of the pressure distribution) and reattached as a tur-

bulent boundary layer (indicated by sharp pressure rise immediately fol-lowing the flattened separated region). Typical examples of this can

be seen in the pressure distributions of figures 6, 7, and 9-

Turning angles.- Essentially uniform variations of turning angle Qwith angle of attack a can be seen in the blade-test figures 3 to 14.

Slopes of the curves for Q against a are approximately 1.0 exceptfor the test for a = 0.75, CJ

Q = 0.6 in which the slope was signifi-

cantly less. The data show that the largest increases in Q (at equal

values of a) for a given increase in Cj occur at the highest solidity.

For the lowest solidity (0 = 0.75) 9 is increased only about 1° in

increasing CIQ from 1.8 to 2 Л but increases by 5° at the highest

solidity (a = 1.5)- (Recent unpublished test results of an annularpassage with guide vanes designed using these data showed good agreement

with the design values of turning angle.)

Some effect of aspect ratio is present since the tests were made

with solid walls and are not exact two-dimensional tests. Figure 15

shows a comparison of turning angles obtained for the NACA б5-(12Аю)-Ю

blade section at aspect ratios of 1.0 and 2.0. Lower turning angles areproduced by the aspect-ratio-1.0 cascades for several reasons. The

velocities induced by the trailing vortices due to reduced lift in the

boundary-layer region as well as flow of the wall boundary layer onto

the suction surface have a greater effect in decreasing the turning

angle at the aspect ratio 1.0 condition. Also, increases in the axial

velocity due to the boundary layer which are greater at aspect ratio 1.0have the effect of decreasing the turning angle. A more detailed discus-

sion of this appears in reference 9-

The turning-angle data of Zimney and Lappi for the б5-(12Аю)Ю

blade sections (ref. ^) are shown also in figure 15- The discrepancybetween the reported test results and recent tests is attributed toimproved cascade testing techniques used in the recent tests most impor-

tant of which was the addition of wall slots for boundary-layer removal

upstream of the blades.

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NA.CA ЕМ L51*-I02

Separation and drag.- Qualitative observations for separation weremade Ъу probing the passage with a small tuft. During the tests noseparation of the flow was observed except as stall occurred at highand low angles of attack. At the low angle of attack the flow separatedat the very leading edge of the lower surface. However, the flow reat-tached near the 50-percent-chord station. The constant pressure regiontypical of this stall can be seen just behind the leading edge on thelower surface in figures 9(a) and 12(a). At high angles of attack,separation, when occurring, started on the upper surface at about thequarter-chord point and the flow did not become reattached. No attemptwas made to record these tests due to the unsteadiness of the flow andthe difficulty of obtaining accurate turning-angle measurements whenthe blade wakes extended completely across the blade passage.

Low drag values were obtained over a 10° to 18° range of angles ofattack. In this range, a design angle of attack was selected, by amethod which will be discussed later; the drag coefficients at designangle of attack are shown in figure l6(a). For equal design turningangles, lowest drag coefficients occur at the highest solidity as shownin figure l6(a). However, when drag is totaled for a number of bladesto indicate all profile losses as in figure l6(b), minimum losses wouldbe obtained using the lowest solidity tested (0.75) except for highturning at which a solidity of 1.00 would produce lower profile losses.Although lowering the solidity may reduce total blade profile losses,increases in secondary flows and Mach number effects that may accompanya decrease in the number of guide vanes may affect adversely the per-formance of subsequent blade rows. While profile losses themselvesconstitute only a fraction of the total losses of a blade row there areindications that the blade surface boundary layer affects the secondaryflow losses and hence blade profile losses may contribute more to totalloss of a blade row near the blade end regions than the measured profilelosses would indicate. Setting a lower limit on guide-vane solidity isillusory as the limit would vary with the particular application. However,it appears that guide vanes may be designed with solidities of 0.75 andlower in some applications.

Design angle of attack.- To make ;the data obtained more usable forthe design of entrance guide vanes, a design angle of attack was selectedfor each test condition. Each angle selected from the various pressuredistributions is the highest angle of attack before the pressure distri-bution formed a sharp peak velocity at the leading edge and also beforedrag rise occurred. The design angle of attack is shown in figure 17for various turning angles and solidities plotted for accurate inter-polation by the carpet-plotting method described in reference 11. Noabscissa scale is given for figure 17; however, turning angle and solid-ity values are plotted so that along a line of constant turning anglethe smallest horizontal division is an increment of solidity of 0.005.

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NA.CA ЕМ

Similarly, along a line of constant solidity the smallest horizontal

division is an increment of turning angle equal to 0.2°.

In figure 18, test results of angle of attack, turning angle, solid-

ity, and camber variations are plotted by the method described in refer-ence 11. The limits of the angle-of-attack ranges plotted were the angles

of attack at which twice the minimum drag occurs. Using this plot, camberis readily determined for intermediate values of solidity, turning angle,

and angle of attack. Camber for design angles of attack at solidities

other than 1.5, 1.0, and 0.75, is obtained by entering figure l8 with the

design angle of attack determined from figure 17 and the turning angle.

In figure 18, lines of constant turning angle and constant angle of

attack are plotted to a horizontal scale with camber as the vertical

scale. Each solidity forms a 6,a carpet plotted so that points of

equal turning angle and angle of attack on the three 9,a carpets are

spaced to a horizontal solidity scale. Hence, to interpolate for camber

when given turning angle, angle of attack, and solidity, the point of

given turning angle and angle of attack is determined on the 9,a car-

pet for the three solidities 0.75, 1-0; and- 1-5- The curve through these

three points represents the variation in camber for solidities from 0.75

to 1.5 for the given turning angle and angle-of-attack combination.Location of the point representing the given intermediate solidity isfound by measuring from a point of known solidity, turning angle, and

angle of attack a horizontal distance equivalent to the increment of

solidity. For this carpet plot the smallest horizontal subdivision isequal to a 0.002 increment of solidity. The ordinate of the point thus

located on the faired line is the desired camber.

Prediction of High-Speed Performance

Choking Mach numbers.- Although no high-speed cascade tests were

made of the blade sections, a number of simple computations were made to

estimate some of the high-speed characteristics of the blade series. The

values estimated were the critical and choking Mach numbers. All of these

analyses are based on the assumption that the low-speed turning angle would

be obtained at all speeds up to choking. Although "this assumption is notprecisely true, the variation in turning angle is expected to be smaller

than for the blades of circular-arc camber lines of reference 12 which

increased a maximum of about 2° at choking.

Choking in blade rows is determined by the minimum passage or throat

area, given for this blade series in figure 19(a). For fair comparison

of throat areas of the three solidities, throat areas are plotted against

the low-speed design turning angles. The maximum entering Mach number

determined from the ratio of throat and upstream areas and isentropic

flow relationships is given in figure 19(b). For low design turning

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ю •.:..:: : : '•• ^содаиштадв*1 "•• NACARML5^io2

angles where the maximuin flow area is restricted by the maximum bladethickness, higher maximum entering Mach numbers can be obtained by usinglow solidities as indicated in figures 19(a) and (b). However, forturning angles of 25° or greater, the maximum entering Mach number whichcan be obtained is essentially the same regardless of solidity since theminimum area, then at the trailing edge, varies little although a finitetrailing-edge thickness was used. Figure 19(Ъ) shows that for the designturning angles the choking entering Mach number is near the maximum obtain-able for the resulting low-speed turning angle.

Critical Mach numbers.- The entrance Mach numbers associated withthe first occurrence of sonic velocity on the blade surface for the bladesections at design angle of attack were estimated using low-speed cascadepressure distributions and correcting for the effects of compressibilityas in reference 6. Figure 19(c) shows the critical Mach number for bladesections at design angle of attack. For low design turning angles dif-ferences in critical Mach number for the various solidities are small;however, for high design turning angles significant differences are shown.Although exceeding the critical speed cannot be taken as a positive indi-cation of increased losses, speeds up to critical are expected to show noappreciable change in losses. The effect of high Mach numbers on the low-speed pressure distribution is to shift the loading toward the trailingedge. This effect is insignificant for low turning angles; however, forhigh turning the changes are pronounced. The low-speed and extrapolatedpressure distributions of a typical high turning angle cascade are shownin figure 20.

SUMMARY OF RESULTS

A series of guide-vane blade sections was designed to have thosequalities believed desirable for guide vanes such as high aerodynamicloading in the leading-edge region, relatively straight trailing edges,and low maximum thickness. From low-speed cascade tests and analysis ofperformance at high speeds the following blade section characteristicsare shown:

1. Low drag coefficients were obtained over an angle-of-attackrange of 10° to 18°. No turbulent separation of the flow was detectednear the selected design angle of attack.

2. The relatively straight trailing edges of the blade series areeffective in guiding the flow at the trailing edges since the variationof turning angle with angle of attack is uniform.

J.'High loading of the mean line in the leading-edge region produceshigh loading in this region for most cascade conditions; however, at the

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•

NACA KM L5 I02 CONFIDENTIAL ц

highest of turning angles, pressures along the upper surface are almostconstant.

k. At the design angle-of-attack condition for turning anglesgreater than 25° the choking area corresponded to an entering Mach numbervhich is near the maximum obtainable for the resulting low-speed turningangle. For lover turning angles the maximum entering Mach number islimited by the blockage due to maximum blade thickness.

Langley Aeronautical Laboratory,National Advisory Committee for Aeronautics,

Langley Field, Va., August 19,

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12 CONFIDENTIAL NACA RM

REFERENCES

1. Kahane, A.: Investigation of Axial-Flow Fan and Compressor Rotors

Designed for Three-Dimensional Flow. NACA TN 1652,

2. Beatty, Loren A., Savage, Melvyn, and Emery, James C.: Experimental

Investigation of Flow Through Three Highly Loaded Inlet Guide Vanes

Having Different Spanwise Circulation Gradients. NACA RM L52D25a,

1952-

3. Lieblein, Seymour, and Ackley, Richard H.: Secondary Flows in Annular

Cascades and Effects on Flow in Inlet Guide Vanes. NACA RM E51G27,

1951.

Ij-. Zimney, Charles M., and Lappi, Viola M.: Data for the Design ofEntrance Vanes From Two-Dimensional Tests of Airfoils in Cascade.

NACA WR L-188, 19 5. (Formerly NACA ACR L5G18.)

5. LieЪlein, Seymour: Turning-Angle Design Rules for Constant-Thickness

Circular -Arc Inlet Guides Vanes in Axial Annular Flow. NACA TN 2179,

1950.

6. Dunavant, James C., and Erwin, John R.: Investigation of RelatedSeries of Turbine-Blade Profiles in Cascade. NACA RM L53G15, 1953.

7. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.:Summary of Airfoil Data. NACA Rep. 82^ > 19 5- (Supersedes NACA

WR L-560.)

8. Erwin, John R., and Yacobi, Laura A.: Method of Estimating the

Incompressible-Flow Pressure Distribution of Compressor Blade

Sections at Design Angle of Attack. NACA RM L53F17, 1953-

9. Erwin, John R., and Emery, James C.: Effect of Tunnel Configuration

and Testing Technique on Cascade Performance. NACA Rep. 10l6, 1951.

(Supersedes NACA TN 2028.)

10. Herrig, L. Joseph, Emery, James C., and Erwin, JohnR.: SystematicTwo-Dimensional Cascade Tests of NACA 65-Series Compressor Blades

at Low Speeds. NACA RM L51G31, 1951.

11. Felix, A. Richard: Summary of 65-Series Compressor-Blade Low-Speed

Cascade Data by the Use of the Carpet-Plotting Technique. NACA

RM L5 Hl8a, 195*4- .

12. Lieblein, Seymour, and Sandercock, Donald M.: CompressibilityCorrections for Turning Angles of Axial-Flow Inlet Guide Vanes .

NACA TN 2215, 1950.

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NACA EM L51H02• • • •• •

• • • • •• • •• • •(

COmENTIAL 13

OlABLE I.- COORDINATES FOE A^Kg MEAN LINE

[clo = i.o]

2.Or

- .

PP 1.0 -

о 100

X

0

.51.252.55.0101520

25303540

55055бо6570758о85

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2.3081.642.912

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• • • •>• • *• • <• • • • •

I* ••• •CONFIDENTIAL NACA EM L54I02

TABLE II.- THICKNESS DISTRIBUTION COORDINATES FOR NACA 63-006

AIRFOIL WITH TRAILING EDGE THICKENED (t/c = 6 PERCENT)

[Stations and ordinates given in percent of chord]

20 r

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.7910

L. E. radius: 0.297Т. Е. radius: 0.6

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NAG A RM L54I02 CO№*£DENTIAL 19

4.8

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20 40 60 80 100 О 20 40 60 80 100 О 20 40 60 80 100Percent chord Percent chord Percent chord

(a)a = OOe; 6=3.8°. (b) a =4.0°, 6 = 7.8°. ' (c) a = 8.0°; # = ll.l°.

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(d) a = 12.0° -,0=14.9°

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(e)a=l6.0°-, 0= 18.8°.

Figure k-.- Blade-surface pressure distributions and blade sectioncharacteristics for cascade combination, P]_ = 0°, a = 1.0, andNACA 63-(бА1

4.К5)Об blade section.

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20 NACA RM IffrlQZ

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, CONFIDENTIAL

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at J3- = 0°.

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NACA 65-(2 X5)06 blade section, a = 32.5°; a = 1.5.

CONFIDENTIALNACA-Langley - 11-26-54 - 300

•> ••• • •*•

CONFIDENTIAL

research memorandum - nasa · research memorandum cascade investigation of a related series of...

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