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Copy,· . RM L53Gl

RESEARCH MEMORANDUM

INVESTIGATION OF A RE LA TED SEPJES OF

TURBINE-BLADE PROFILES IN CASCADE

By James C. Dunavant and JohnR. Erwin

Langley Aeronautical Laboratory Langley Fie Id tVa.

This material contain,s in!nrmaUOn of !he ~ Ja ..... TIlle 18. m.a.nner to an unauthoI1z.ed person 1s

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON

December 7, 1953

https://ntrs.nasa.gov/search.jsp?R=20050080790 2018-05-03T06:52:24+00:00Z

NACA RM L53G15 ~h._.-...... ¢S_ NATIONAL ADVISORY CQ!.'rr.nTl"ZE FOR AERONAUTICS

• . . RESEARCH rillJIORANDUI.J:

INVESTIGATION OF A RELATED SERIES OF

TURBINE-BLADE PROFILES IN CASCADE

By James C. Dunavant and John R. Erwin

An application of airfoil design methods was used to design series of related turbine-blade profiles to satisfy the conditions of inlet flow anele and turning angle encountered in the usual range of turbine operation. A series of blade profiles applicable to most turbine blading requirements and a secondary series with particular reference to impulse conditions were desir,ned. Five blade sections from these series ranging in mean-line turning angles from 650 to 1200 were tested in low-speed cascade tunnels. From low-speed test results optimum blade angles of attack were selected at each test condition. The induced angle and the deviation angle of the flow were determined from the low-speed data. If these angles are known for the solidity and inlet angle of an application, the necessary camber is specified. A method of predicting high-speed pressure distributions from low-speed cascade test results is presented to extend the usefulness of the low-speed data. Sample high-speed tests of two of the five blade sections were made at Mach numbers up to tile critical value. The results indicated satisfactory flow conditions in all of the olade passages tested.

INTRODUCTION

Current methods of selecting turbine blade sections are based largely on empirical rules derived from experience with steam and gas turbines. In many applications satisfactory results have been obtained but no system­atic method of designing suitable turbine blade sections has been published. In the case of closely spaced, highly cambered blades such as are encoun­tered in turbine practice, the theoretical determination of the blade shapes is a difficult problem. A number of theoretical methods of varying degrees of approximation have appeared (for example, refs. 1 to 5).

2 CO~ NACA RM L53G15

For isolated airfoils it is possible by theoretical means to derive the pressure or velocity distribution about a given profile. On the basis of this theory a simple procedure has been developed (ref. 6) in which the theoretically determined basic thickness distribution and mean camber line are superimposed to determine the shape of the isolated pro­file corresponding to any desired pressure distribution.

The ,purpose of present investigation was to determine a related blade series that would have satisfactory aerodynamic performance throughout the ranges of inlet angle and turning anGle encountered in current aircraft gas turbine practice. In this work, the same procedure used in the isolated-airfoil case of superimposing a thickness distribution on a mean camber line has been followed. By this method, an infinite number of blade profiles covering a wide range of cambers can be designed. The basic thickness distribution and the shape of the camber line for blade series were determined by cut-and-try procedures involving tests in a low-speed cascade tunnel and graphical flow studies with the aid of a wire-mesh plotting device (ref. 7).

In order to verify the performance of blade series developed by this method, five profiles designed for widely differing turbine operating con­ditions were tested in cascade tunnels. Two of the profiles were investi­gated at speeds up to the critical; however, choking flows were not obtained. From low-speed test results, optimum blade operating conditions were selected at each test condition. These data are used as a basis of a nlethod which is presented for determining blading configuration for given design values of turning angle and inlet air angle.

A

c

flow area

blade chord

pressure coefficient,

SYMBOLS

p - Pm ~

pressure coefficient at low speeds

camber, expressed as design lift coefficient of isolated airfoil

coefficient of total force parallel to mean velocity

CO~AL

..

.,

UACA Rf\~ L53G15 c~ 3

CWl

Fc

H

ML

P

PR

q

s

S

t

V

w

x

y

a..

~

., e

p

cr

wake momentum di~~erence coer~icient

ratio o~ wake ~orce coef~icient to integrated total-pressure loss in wake

Mach number

mean-line length

pressure

resultant pressure coe~ficient; dir~erence between local upper­and lower-sur~ace pressure coe~ricients

dynamic pressure,

tangential spacing between blades

pressure coerricient,

blade thickness

velocity

wake width

chordwise distance rrom blade leading edge

perpendicular distance from blade chord line

angle or attack; angle between entering ~low direction and chord

air angle measured ~ram axial (see rig. 1)

ratio or specific heats

turning angle

density

solidity, chord spacing ratiO, cIs angle between leaving ~lov direction

line direction

change in direction leading edge

trailLig-edge mean-

streamline at the

4

LID

tic

Subscripts:

1

2

c

m

o

t

w

cr

CO~ NACA RM L53G15

ratio of blade forces perpendicular to mean velocity and forces parallel to mean velocity

maximum thickness, percent chord

upstream of blade row, undisturbed stream

downstream of blade row

camber

mean passage

stagnation

thickness distribution

wake

critical, condition at first attainment of sonic velocity on blade surface

BLADE SECTION DESIGN

General Design Criteria

The same procedure used in the isolated-airfoil case of superimposing a thickness distribution on a mean camber line has been used to obtain turbine blade profiles. The development procedures used for airfoil mean lines and thickness distributions are given in references 6 and 8, respec­tively. Blade series were designed to have satisfactory aerodynamic per­formance throughout the ranges of inlet angle and turning angle encountered in current aircraft gas turbine practice. For the purpose of designing blade series, flow conditions were divided into two groups requiring dif­ferent blade sections to approach optimum performance. One group corre­sponds roughly to a reaction turbine condition in which the flow is accel­erated through the cascade; blades designed particularly for this condition are classified as the primary series. The other group corresponds roughly to an impulse turbine condition for which the entering and leaving veloci­ties are nearly equal; blades designed particularly for this condition are classified as the secondary series. Two dimensionally, the reaction con­dition occurs when the turning angle is greater than twice the inlet flow angle and the impulse condition occurs when the inlet and exit angles are equal.

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NACA RM L53G15

Four criteria were believed to be desirable for turbine blade sections:

(a) Turbine blades should have low drag and no separation of the flow from the blade surface.

(b) Blade sections should have critical Mach numbers closely approaching the maximum values for the particular combination of inlet and turning angles to delay shock fOL-wation in the passage.

(c) The passage flow area should decrease from leading to trailing edge. The minimum area in the blade passage for a given blade spacing should be as large as possible without exceeding the leaving-flow area downstream of the passage.

5

(d) Whenever a penalty in over-all performance would not be incurred, a mean line and thickness distribution yielding nearly straight upper­surface trailing edges for all cambers is desirable since the leaving­flow direction for such sections is usually determined within close limits.

Primary Series

Figure 2 gives the theoretical maximum entering Mach number based on one-dimensional isentropic flow at choking flow for inlet angle and turning angle combinations. For blade sections having high reaction, and hence low maximum entering Mach number, highest critical Mach numbers will be obtained with blade sections having the aerodynamic loading con­centrated in the low Mach number leading-edge regions rather than in the trailing-edge regions where the flow velocity is high. At high speeds, surface velOCity distributions (for reaction conditions in particular) are modified by the effects of compressibility so that the loading in the trailing-edge region increases relative to the loading in the forward region. For these reasons, the flow can be turned quite rapidly in the forward region of the blade passage without exceeding the local velocity of sound. High aerodynamic loading of the blade mean line in this region is desirable to produce a given total lift or turning angle with the low­est maximum surface velocity and minllnum losses. This also permits selec­tion of a mean line and thickness distribution yielding nearly straight upper-surface trailing edges for all cambers.

With these considerations in view, blade sections of arbitrary mean line and thickness distribution for typical inlet- and turning-angle con­ditions were laid out and compared. Low-speed velocity distributions were obtained for the more promising sections by means of the wire-mesh plotting device (ref. 7) and the low-speed ca.scade tunnels. A metl!ou of extrapo­lating these low-speed velocity distributions to high-speed velocity dis­tributions was developed in order that the high-speed pressure distributions

6 ~ C . ............-- NACA RJ'.1 L53G15

and critical Mach numbers of various sections could be compared. This method of extrapolating low-speed cascade tests results is given later. This process was used to select a single basic mean line and thickness­distribution shape for the primary series.

Secondary Series

In the ideal impulse case, the critical entering Mach number would approach 1. 0 for high solidities. However, at both low and high solidities practical blade thicknesses prevent the maximum critical Mach number from being closely approached because the average velocity through the passage necessarily remains high and large velocity differences across the passage are needed to produce the required lift. With a uniform average velocity tlrrough the passage, highest critical Mach numbers will be obtained by distributing the loading along the entire chord thus curving the mean line near the trailing edge. Therefore, the desire for straight trailing edges was compromised to distribute the loading more evenly along the chord and obtain higher critical Mach numbers.

A development similar to that of the primary series was carried out to obtain a secondary blade series more suitable for the particular requirements of impulse and low-turning blade operating conditions.

Mean Lines

Figures 3 and 4 give ordinates of the basic mean-line shapes in terms of Cl = 1.0. The designation for these mean lines is given by the system

o used in reference 9. In the present development, the free-stream lift coefficient C1 is a meaningless parameter used only for convenience to

o relate the ordinates of figures 3 and 4 to the design value of turning angle and inlet angle. (See fig. 1.) The relation is

8 + 68 + 68 induced deviation

tan-l C1

f-dYc\ _ tan- l C1

(dyc) O\dxc)0.5 o\dXc 95

where (dYc ) and (dYc) are the slopes of the mean line at dxc 0.5 dxc 95

0.5 percent chord and 95 percent chord for C1 = 1.0. All quantities in o

this relation are known except C1 and the values of b/a. A reasonable o

aEsumption for the values of 68 can be made on the basis of test results to be described later and thus C1 can be calculated by a simple

o

c

•

NACA RM L53Gl5

trial-and-error solution. The blade mean-line coordinates are obtained from the values of figures 3 or 4 by the following relation:

Thickness Distribution and Ttlickness

7

Thickness distribution coordinates for the blade series are tabulated in figures 3 and 4 for a 20-percent-chord maximum thickness and conversions to other maximum thicknesses are made by multiplying the ordinates of the uncambered airfoil section by the ratio of the desired thickness to the tabulated thickness. Position of the maximum thickness for the primary series is at 20 percent of the chord. For the secondary series the posi­tion of maximum thickness is shifted rearward to 40 percent of the chord near the position of the maximum blade passage area (where the flow has turned to axial). The value of maximum thickness for any particular blade configurations was selected so as to be consistent with the desires to have the area where choking will occur a maximum and to eliminate large flow area increases (diffusion) within the blade passages. Variation of the blade maximum thickness within these limits is deemed to have little effect upon the flow.

On ordinary airfoils where the camber is low, the thickness distribu­tion is laid out on perpendiculars to the mean line above and below the mean line at stations distributed in proportion to the chord length. How­ever, in the case of a highly cambered turbine blade section, the mean line may be as much as 30 percent greater in length than the chord line. !n this case, the thickness is distributed along the curved mean line on perpendiculars stationed in proportion to the mean-line length. This method, illustrated in figure 5, was used on the blade sections of this paper.

The leading-edge radius varies as the square of the thickness ratio which is defined as t/ML in percent for the cambered airfoil. Therefore, the leading-edge radius for the cambered airfoil I£R/ML is obtained from the leading-edge radius given in figures 3 or 4 I£R/c by the following relation:

~ = !:E!.~\2 ML c \20 J

For the secondary series thickness distribution; the leading-edge radius determined from the thickness-distribution equations of reference 8 was unsuitable and an appropriate value (see fig. 4) was selected. The trailing-edge radius varies directly with thickness.

c

8 NACA RM L53G15

TWO-DIMENSIONAL CASCADE TESTS OF BLADE SECTIONS

Limits for the range of turbine-blade operating conditions were determined by a study of many turbines used in aircraft engines. These limits are defined by the equation 8 = ~l + 550 t200 for values at ~l from 00 to 650 . Average values of solidity were chosen to be 1.5 and 1.8. The investigation was confined to these limits. Blade sections were tested over a range of angles of attack at several inlet angles to obtain the optimum angle of attack and turning. The optimum angle of attack was determined from the blade-surface pressure distribution at the leading edge and consideration of wake losses. The method of selection of this optimum value is discussed later.

Blade-Section Models

The blade sections were cambered with mean lines having camber angles arbitrarily 50 greater than the desired turning angle. This overturning of the mean line was an allowance for the induced flow angle at the nose. The induced flow angle or turning of the flow ahead of the blade row in the direction opposite to the turning angle is caused by the circulation about the blade and is partially counteracted by the interference between "blades. Previous cascade tests and flow plots had indicated 50 to be an average value of this angle. The 600 and 750 turning-angle sections were designed with maximum thicknesses of 10 percent of the chord and the 900 and 1050 sections, with maximum thicknesses of 15 percent of the chord.

The blade section for the 1200 turning condition was designed using the secondary series since turbine blades operating with turning angles as high as 1200 necessarily would be impulse or very close to impulse. The mean-line turning of this section was 1200 ; no allowance was made for induced flow angle since flow plots indicated this angle to be negligible at 600 inlet angle. Maximum thickness of this section was made 25 per­cent of the chord to prevent the passage from diverging excessively at the center. A small irregularity in the lower-surface shape occurred ~ecause of the extreme thickness of the section. This irregularity was removed by using a circular-arc fairing. These blade sections are shown in figure 6.

Cascades Investigated

The following table presents the conditions at which tests were made:

:2D

NACA RM L53G15 9

Desired Blade section Solidity a = 1.5 and Solidity 0=1.<3 and

turning angle, camber angle, inlet air angle ;31 of - inlet air angle 131 of -9, deg 9c, deg

00 ,,,,0 7~O 1...-0 _0 __ 0 1:;>° 20° 25°

n I n ~I -60° ~./ .JU 'l-) U .1U 30~ 135~ 40v 45v

60 65 a a a a a a

75 80 a a a la I b

a,b b b a,b b

95 a I a a a b a,b b b a,b

I I

110 a a a a liJ \ 105

120 120 q I aLov-speed tests.

bHigh-speed testS.

Low-speed tests, identified by the letter a were conducted over a large range of angle of attack. High-speed tests, identified by the letter b were conducted at a single angle of attack or through a very narrow range of angle of attack.

Low-Speed Tests

Description of equipment.- The test facilities used in this investi­gation were the Langley low-speed cascade tunnels described in refer-ence 10. One major modification had been made in the larger cascade tun­nel by reducing the test section width from 20 inches to 10 inches. All tests were made using solid walls in the test sections. Blade chords were 6 inches. The highest blade surface velocity attained in these tests was 206 feet per second.

Test Reynolds numbers.- As the pressure available from the blower was limited, the same dynamic pressure or Reynolds number could not be obtained at all test conditions due to the large variation in pressure drop across the blade row. Hence, the tests were run at near maximum output of the blower. Reynolds numbers based on the chord and upstream velocity showed wide variation with test conditions and in some cases were as low as 120,000 which is within the ordinary range of critical Reynolds numbers at which high losses occur in cascade. Low Reynolds numbers were obtained on blade configurations having large pressure drops across the blade row. Hence, the upstream velocity was the lowest velocity in the cascade and the flow velocity increased rapidly in the blade passage. The Reynolds number based on upstream velocity under such conditions is therefore of questionable value. By basing the Reynolds number on the average of t.he upstrea.'TI and do;;nstreaiTI velot.:i ties, a more sui table index is obtained. For this assumption, the range of Reynolds number for the low-speed tests was from 320,000 to 500,000.

C

10 NACA RM L53G15

Reduction of data.- Nondimensional force coefficients based on the upstream dynamic pressure Ql were calculated from pressure distribution,

turning angle, and wake-survey data. The method used in calculating blade force coefficients is given in reference 11. Comparison of the force coefficient perpendicular to the chord calculated from turning-angle data and the force coefficient from the integrated pressure distribution was made as an accuracy check on the tests and these data agreed within 5 per­cent. The CWl and CDl values presented are, respectively, the wake

momentum difference coefficient and the coefficient of the component of the total force on the blade parallel to the mean velocity. The ratio of forces perpendicular and parallel to the mean velocity LID was calculated for blade comparison.

High-Speed Tests

Description of eQuipment.- Tests were made of two of the blade sec­tions in the 7-inch high-speed cascade tunnel at the Langley Laboratory. A diagrammatic sketch of the tunnel is shown in figure 7. The 40-inch­diameter duct is used as a settling chamber. Two screens are contained in this duct to insure uniform flow ahead of the tunnel inlet fairing. The tunnel upper and lower floors are attached to the inlet fairings. Both floors are adjustable in the vertical direction. The horizontal adjustment of the floor ends is made by telescoping the floor plates. The tunnel side walls consist of large circular assemblies which are rotated independently of the side-wall inlet fairing and the floors. Each side wall contains a protruding slot connected to a suction pump for removal of the boundary layer. The test section incorporating seven 4.2-inch-chord blades is located approximately 7~ inches downstream of

the protruding slots. Boundary layer on the floors is bled off by dis­charging the flow between flexible floor extensions and the upper and lower blades to the atmosphere. Suction chambers connected to an exhauster are installed to remove this flow when the upstream pressure is insuffi­cient to discharge the flow to the atmosphere.

Due to limited blower pressure ratio, a diffuser was connected to the cascade downstream of the blades to increase the pressure drop avail­able across the cascade. A straight aluminum plate was attached to the trailing edge of the lower blade and sealed at the blade trailing edge and at the walls. A second plate consisting of a 24-inch straight sec­tion followed by a diffuser section was connected to the trailing edge of the top blade. Linkages kept the straight sections parallel and pro­vided a means of positioning the floors.

Total pressure and total temperature were measured upstream in the inlet duct. Static-pressure taps were located in the side walls between the protruding slot and the blade row. Flow angle upstream was measured

C

NACA RM L53G15 11

with a removable probe at three points on a line parallel to and just behind the slot. Downstream static-pressure taps were arranged in a row parallel to the blade row on the side wall and in rows parallel to the flow direction on the floors. A 11-tube total-pressure rake was inserted through a slot in the 24-inch straight section of the upper floor to measure the loss in total pressure caused by the center blade.

Test procedure.- The inlet air angle relative to the cascade of blades was set by rotating the circular side-wall plate containing the blade row. The angle of attack of the blades was set by means of an inclinometer.

Uniformity of upstream flow conditions was determined from three angular measurements with the removable yaw probe, upstream wall static pressure, and static pressures on the leading edges of the second, fourth, and sixth blades which contained special orifices for this purpose. Flow conditions upstream were altered by adjusting the floor positions and by controlling the boundary-layer bleed-off on the floors and walls. Down­stream flow conditions were adjusted to obtain equal static pressure on the side walls and floors by changing the floor angle. The leaving-flow direction for a particular cascade configuration was assumed to be parallel to the floor direction when equal static pressures were obtained on the floors and walls. Insertion of the total-pressure rake into the flow caused changes in the blade-surface pressure distribution as well as in the upstream and downstream Mach numbers. Therefore, all test data, except for the total-pressure loss, were recorded with the rake removed. Loss measurements were made at identical blower speeds and tunnel set­tings even though rake interference caused differences in the static pressures on the walls and floors downstream.

Tests were made for a range of speeds including critical blade speed. ChOking flows were not obtained due to limited blower pressure ratio. Average Reynolds numbers for the highest speeds calculated in the same manner as for the low-speed tests were between 1,200,000 and 1,500,000.

Reduction of data.- Normal-force coefficients were calculated from the integrated pressure distribution and from differences of the momen­tum and pressure upstream and downstream for each test. A comparison of these values made to determine testing accuracy showed a maximum discrep­ancy of 8 percent; generallY, the discrepancy between values was much less than the maximum of 8 percent. An additional check was made to determine whether the boundary-layer build-up through the cascade was sufficient to prevent the tests from being considered two-dimensional. The two-dimensional mass flow through the blade passage at the center of the cascade was calculated from upstream conditions and compared with a similar mass "flow calculated at the dowIlstream rake position. This com­parison showed the mass flow downstream to be between 3 percent less and 5 percent greater than the upstream iii The wall boundary layer,

12 NACA RM L53G15

therefore, was small and the tests were concluded to be very nearly two-dimensional.

A wake force coefficient waS calculated using the method given in reference 12 for determining drag-coefficient values. Additional assump­tions were made in applying this method to a cascade in that the down­stream conditions were assumed to be free-stream conditions and the static pressure at the rake was equal to the static pressure far down­stream. Wake force coefficients were calculated by this method; however, values presented are based on the dynamic pressure upstream of the cas­cade; hence

lw Po - P02

= Fc dw o qlc

DISCUSSION

Turning Angles

Results of the low-speed tests are shown in figures 8 to 30. The turning angle increases in almost direct proportion to the angle of attack. In figure ll(f), however, the measured turning angle does not show a comparable increase for an increase in the angle of attack at leaving-flow angles near 740. With high exit flow angles, the straight portion of the convex surface no longer forms a guided region at the trailing edge. The downstream shift of the straight trailing-edge por­tion relative to the adjacent blade, especially in the lower solidity tests, is evidently the cause for the turning-angle variation. From this, it may be expected that separation of the flow from the trailing surface will occur for leaving-flow angles greater than 740 but leaving-flow ali.gles in the neighborhood of 740 may be considered the normal design llmi t. A nonlinear variation of e with CL is observed in figure l8( f) at low angles of attack. In this case, the lower surface velocity peak near the leading edge of the very highly cambered airfoil and the following increasing pressure gradient may have produced laminar separation over the concave surface since the wake forces show some increase at this condition. The turning angle decreases slightly at this condition because of the secondary boundary-layer flow along the wall from the high-pressure lower surface to low-pressure upper surface. The resulting thickened boundary layer on the upper surface results in a greater flow deviation from the mean-line direction.

The high-speed test results presented in figures 31 to 49 showed no significant change in turning angle with Mach number. Choking flows were not obtained so the changes in turning angle to be expected at su~ercritical

CO.TIAL

..

-.

NACA RM L53G15 13

pressure ratios (ref. 13) were not investigated. Turning angles measured for the high-speed tests varied with angle of attack in a manner similar to the low-speed tests. Measured turning angles for the Mach number range tested were consistently between 10 and ~ greater than corresponding low­speed test measurements. Since no variation of turning angle with Mach number was observed, this discrepancy is not attributed to effects of varying Reynolds numbers. In the high-speed tunnel, the leaving flow is assumed to be parallel to the downstream floor plates. The upper-floor boundary layer is ~ormed along a blade lower surface since the floor is a continuation of a blade lower surface just as the lower floor is a con­tinuation of the upper surface of a blade. Hence, because of the dif­ferent boundary layers, the flow between the downstream floors is not necessarily exactly parallel to the floors as was assumed in obtaining the turning angles presented. For subcritical pressure ratiOS the leaving­flow direction for turbine blading incorporating guidance is determined within close limits by the trailing-edg~ direction.

For convenience in determining the effect of blade shapes and inlet air angle on turning angle from low-speed test results, the deviation or leaving-flow angle relative to the mean-line direction at the 95-percent­chord station is plotted in figure 50 against blade camber angle and in figure 51 against leaving-flow angle. These figures present the variation of the deviation angle for a particular angle of attack selected as an optimum design condition. The method of selecting des ign angles is dis­cussed in a subsequent section. In figures 50(a) and 51(a), the high value of deviation angle for the Bc = 1100 blade section at an inlet

angle of 450 is believed to be a result of laminar separation on the lower surface. This point, therefore, is excluded in analyzing the deviation-angle variation. The deviation angle for this blade section is ~ssumed to vary in a manner similar to that shown in figures 5O(b)

Jand 51(b) for the same blade section at a higher solidity. General trends indicated by figures 50 and 51 are increasing deviation angle with increasing inlet air angle for constant camber. The deviation angle increased with increasing blade camber for constant air inlet angle or constant exit flow angle. The low values of deviation angle indicate that the straight trailing edges satisfactorily guided the leaving flow.

Wake Losses

Low speed.- Trends of the variation of wake losses with changing blade conditions are indicated in the low-speed test results shown in figures 8 to 30. Wake measurements are presented as coefficients deter­mined from losses based on the upstream dynamic pressure. Blade-section efficiency is indicated by the LID values plotted. At high reaction conditions the large difference between the entering velocity on which the coefficients are based and the leaving velocity causes high drag and

14 NACA RM L53G15

lift coefficient values but the LID value is relatively unaffected. Generally, LID ratio increased with increasing camber for a constant inlet flow angle. The lower solidity tests produced higher LID ratios since increased lift was obtained with only slightly higher drag values.

High speed.- Values of CWl measured in the high-speed tunnel at

very low speed agree with values measured in the low-speed tunnels within the accuracy of the testing apparatus. Wake momentum difference coeffi­cients increased slightly with increasing Mach number up to critical speed. Tbis, however, is not an indication of any deterioration of the flow because a wake momentum difference coefficient based on downstream rather than upstream dynamic pressure shows no significant change in value witbin the Mach number range investigated.

Wake losses do not indicate any strong effects of separation at speeds up to the critical speeds and from the standpoint of wake losses the blade series are consi~ered satisfactory.

Surface Pressure Distributions

Low-speed tests.- The primary blade series has high curvature near the leading edge with nearly straight trailing edges. In tests of the primary blade shapes, the velocities along the straight, unloaded trailing­edge portions were only slightly greater than the leaving-flow velocities indicated by the dash at 100-percent chord on pressure distributions. For the operating condition of large velocity increases across the blade row, the trailing-edge surface velocity, although only slightly greater than the leaving-flow velocity, was the maximum blade surface velocity. Typical examples of this condition are shown in figures 22, 25, and 28. For oper­ating conditions producing little or no velocity increase across the blade row, the highly loaded leading edge and the relatively high entering velocity caused upper-surface velocities near the leading edge to be far greater than the leaving velocity as seen in figures 10, 13, and 16.

In the low-speed tests, detrimental effects due to high velocities near the upper surface leading edge were not evident. No separation of the flow from the upper surface was detected in these tests for normal operating conditions. On the lower surface high velocities occurred near the leading edge of highly cambered blades at low angles of attack. The steep positive pressure gradients followin~ the velocity peaks sometimes caused laminar separation of the flow from the lower surface. For example, the dasbed pressure distribution of figure 29(a) shows the resulting separated region just behind the nose on the lower surface. This separa­tion resulted in a reverse flow region in the blade passage near the lower surface. Secondary boundary-layer flow along the wall resulted in a thick­ened boundary layer on the blade upper surface and, hence, a lower turning angle. Increasing the stream turbulence by inserting a 1/2-inch mesh •

C~IAL

NACA RM L53G15 15

screen upstream of the blade row greatly reduced the extent of the separated region. This method was used to obtain the tests shown in figures 17, 18, 29, and 30. The constant-pressure region typical of laminar separation is not detectable on these pressure-distribution plots nor was arq flow rever­sal indicated in the surveys made in the blade passage with a small tuft. However, in the low angle-of-attack tests of figures 18 and 30, -the drag values are higher than at higher turning angles.

Pressure distributions from tests of the highly cambered secondary blade section are shown in figure 30. M::>re uniform distribution of the aerodynamic loading (fig. 5) along the chord resulted in a blade shape having a curved trailing edge and nearly uniform pressure along the upper surface at test conditions producing negligible velocity increase across the blade row. M::>re nearly ideal surface velocities than for the reaction blade series are obtained for this particular condition in that higher entering Mach numbers are expected before sonic velocity occurs on the blade surface. Also, higher flow!;> "Would. be obtainable before choking occurs since the curved trailing edge allows larger minimum areas for equal leaving-flow angles.

High-speed tests.- The blade-surface pressure distributions measured during the high-speed tests -were f?imilar in shape to the low-speed distri­butions although the pressure coefficient S increased with increasing speed, especially for the higher velocity regions. Hence, for reaction conditions the loading shifted rearward with increasing Mach number. Examples of this may be seen in figures 33, 34, 42, and 43. Radical changes in the pressure distribution occurred as blade surface speeds exceeded sonic speed and shocks formed in the passage. In figure 46(b), the effect of a shock on the pressure distribution can be seen as the pressure coefficient decreases abruptly in the region near the 2O-percent­chord station. At most test conditions, sonic velocity was exceeded only slightly due to the limited blower pressure ratio.

As in the low-speed tests, separation was not observed for these blade series within the range of conditions in which the blades would be expected to operate. In this respect, therefore, the blade series are considered satisfactory.

Critical Mach Numbers

Entering Mach numbers are limited by choking of the blade passage; however, losses in blading can become significant before choking occurs. Limiting entering Mach numbers based on losses cannot be established from the high-speed tests since the test velocities were limited by the avail­able air supply. Entering Mach numbers in excess of critical were obtained without excessive wake forces in high-speed tests such as those of figures 38, 40, and 42. These tests did not cover the range of

NACA RM L53G15

discharge Mach numbers of sonic and above, and blade operating character­istics at these speeds are unknown.

Representative test values of critical speeds are given in figure 52. Critical Mach numbers calculated by the method of applying compressibility effects to low-speed pressure distributions given in a subsequent section are shown i~ figure 52 to differ little from test values. Critical Mach numbers approach the theoretical maximum Mach number for the cascade com­bination when operating in a reaction condition. However, as impulse con­ditions are approached, maximum Mach number increases more rapidly than the critical Mach number. Theoretically, at an impulse condition the flow could be diffused and turned with infinitely thin blade sections to yield critical Mach numbers equal to 1.0. Practically, however, the amount of diffusion possible is limited and the blade thickness must be necessarily large; hence, critical Mach numbers at impulse conditions are much less than the maximum. The critical Mach number for a secondary-series blade designed specifically for impulse condftions with uniform loading along the chord is higher than that of the primary blade section, however, as can be seen in figure 52. In order to estimate the critical speed of the secondary blade, compressibility effects were applied to the pressure distribution obtained from the wire-mesh flow plot of this blade section.

Critical Mach numbers near maximum were obtained at reaction con­ditions. However, for impulse conditions blade sections with straight trailing edges have lower critical Mach numbers than blade sections with camber in the trailing-edge region. Therefore, with respect to the criti­cal Mach number the blade series also may be considered satisfactory.

Application of Compressibility Corrections to Low-Speed Tests

Data can be obtained from low-speed cascade tests with a minimum of equipment, time, and computation. The data are of value for direct use and for comparing different blade sections intended to accomplish a given task. In most turbomachines, however, the blades operate at high speeds so the effects resulting from the compressibility of the flow must be con­sidered. This is particularly true of aircraft-engine turbine blading, for the velocities leaving the blades are usually near sonic. A method is required, therefore, for correcting low-speed cascade pressure distri­butions for the first-order effects of compressibility in order to esti­:nate the high-speed pressure distribution. No rigorous theoretical method of accomplishing this correction is available.

The semi-empirical procedure about to be described is derived from an approximate method used to estimate compressibility effects for isolated airfoils. The Prandtl-Glauert theory for thin airfoils in a free stream leads to a formula relating the pressure coefficient Cp at a stream Mach

number Ml to the incompressible pressure coefficient CpO

CO~

3D NACA RH L53G15

In the present method, it is arbitrarily assumed that this formula can be used to compute the compressible pressure coefficient at a station

17

in the turbine blade cascade from the low-speed value at that station by replacing ~ with a mean value taken across the passage at the station

in question. The mean stream Mach number ~ll is deterrn..ined from kno .... u

values of the passage area Am, the entering stream tube area Al' and

the entering Mach number Ml from the one-dimensional isentropic flow relationship

The pressure coefficients Cp and CpO are also based on mean passage

condition obtained from the areas Al and Am.

As the potential flow in the two-dimensional passage is perpendicular to an equipotential line across the passage at any station, the effective flow area Am will be determined by the length of this line. Such a line may be approximated with fair accuracy by a circular arc which is perpen­dicular to both passage boundaries. The correctness of this assumption of area was checked on flow plots made with the wire-mesh-flow plotting device described in reference 7. No exact construction can be given for these circular arcs since the position and hence the slope of one end of the arc is not Imown. To facilitate the determination of a suitable arc, a transparent template consisting of a series of arcs of varying radii was made. This method is illustrated in figure 53. Equal arc lengths on the template were marked so that passage width could be read directly on the templa te after selecting the proper arc. A planchette, a small square of transparent plastic having two lines at right angles and a scale of inches inscribed on it facilitates selecting the appropriate arc and reading the corresponding arc length.

TYPical results of this method of extrapolation of low-speed data are compared in figure 54 with the equivalent high-speed pressure distribution obtained in cascade tests. Good agreement of pressures is noted in this case for the entire blade surface. Critical entering Mach numbers obtained by means of this extrapolation show good agreement with test results as shown in figure 52. At Mach numbers greater than critical the method can­not be used rearward of the shock location. In cases where the flow pat­tern is altered by viscous effects at high speeds, the method will obviously not give correct predictions.

18 NACA RM L53G15

Selection of Design Camber

The results of the low-speed tests at constant inlet air angle per­mit the selection of design blade operating conditions. Interpolation of results between blade shapes and test conditions is made possible by the geometric similarity of the mean camber lines. Thus, by determining the optimum angle of attack at each inlet air angle and solidity the blade camber is specified for a desired turning angle at any inlet air angle. To obtain design blade cambers for other conditions, an inter­polation between both inlet air angle and turning-angle conditions can be made.

The upper-surface velocity in the trailing-edge region has been observed to be consistently slightly higher than the leaving-flow velocity. Since the leaving-flow velocity will be a fixed value for a given design, then the surface velocity in the trailing-edge region is also fixed leaving only the flow in the leading-edge region a variable. Hence, in selecting the optimum angle of attack at each test condition of air inlet angle and solidity, the criterion used was the flow at the leading edge of the airfoil. The. effect of circulation around the blade causes a turning of the airstream in the direction of the blade lift just ahead of the leading edge; this change in direction of the stagnation streamline

II n at the leading edge is referred to as induced flow angle ~i d d' n uce The value of the induced angle was obtained experimentally for the blade shapes tested in the low-speed cascade tunnel. Measurements were made by using closely spaced surface pressure orifices near the blade leading edge. In effect, the blade was rotated until a pressure equal to full total pressure was obtained at the orifice located at the end of the mean line and the orifices on either side indicated equal pressure drop. The absolute accuracy of such a measurement as this is low and is estimated to be only within ±3° since it depends on the accurate positioning of the pressure orifices.

The resulting induced angle values are shown plotted in figures 55 and 56. The most apparent trend was the decrease of induced angle with increasing inlet angle. The variation with solidity was of smaller magni­tude and almost constant over the entire test range. The induced angle is largest at low inlet air angles where the interference of other airfoils in the cascades is least. For the purposes of camber selection a mean value of the induced angle is presented as a function of only the inlet air angle and the solidity as shown in figures 55 and 56. B,y overcambering lJlade sections by the value of the induced angle, the effects of the induced flow are taken into account. Thus, total blade camber or mean­line turning angle will be the sum of the turning angle, induced angle, and the angle of flow deviation from the mean line exiting angle interpo­lated from figure 50 or 51. Design angle of attack of the section will

dyc - /).8

5 C -1 0 induced . , Lo- . be the mean-line slope:

NACA RM L53G15 C'" 19

SUMMARY OF RESULTS

The investigation conducted to obtain related turbine blade profiles for conditions of inlet-flow angle and turning angle encountered in the usual range of aircraft gas turbine operation yielded two series of airfoil-type blades. A primary series was developed for reaction condi­tions and a secondary series was selected to produce more nearlY optimum results at impulse and low-turning conditions.

Relating blade sections by this design procedure allows interpolation of blade test results and design of blade sections with greater certainty of flow conditions within the blade passage. A simple design procedure is given whereby blade profiles from these series can be readilY ~btained for any design condition in the range of current interest. Satisfactory blade performance is indicated by the cascade tests of five blade sections from these series for widelY varying conditions which produced the fol­lowing results:

1. Separation was not observed for the blade series within the range of conditions in which the blades would be expected to operate.

2. Guiding the flow by means of straight trailing-edge portions is effective in maintaining a fixed turning angle up to critical speed.

3. Near-maximum critical Mach numbers were obtained at reaction con­ditions. However, for low reaction conditions blade sections with straight trailing edges have lower critical Mach numbers than blades with camber in the trailing-edge region.

4. A semi-empirical method for extrapolating low-speed pressure dis­tributions was found to produce resultE which were in good agreement with high Mach number tests. The method cannot be used in cases where the flow pattern is altered by viscous effects or shocks.

Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Langley Field, Va., JulY 14, 1953.

-

20 NACA RM L53G15

REFERENCES

1. Mutterperl, William: A Solution of the Direct and Inverse Problems for Arbitrary Cascades of Airfoils. NACA WR L-81, 1944. (Formerly NACA ARR L4K22b. )

2. Goldstein, Arthur W., and Jerison, Meyer: Isolated and Cascade Air­foils With Prescribed Velocity Distribution. NACA Rep. 869, 1947.

3. Weinig, F.: Flow Patterns of a Compressible Gas Through Turbine Cas­cades When Shock Waves Are Avoided. Tech. Rep. No. F-TR-2141-ND, Air Materiel Command, U.S. Air Force, May 1947.

4. Alpert, Sumner: Design Method for Two-Dimensional Channels for Com­pressible Flow With Application to High-Solidity Cascades. NACA TN 1931, 1949.

5. Costello, George R.: Method of Designing Cascade Blades With Pre­scribed Velocity Distributions in Compressible Potential Flows. NACA Rep. 978, 1950.

6. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.: Summary of Airfoil Data. NACA Rep. 824, 1945. (Supersedes NACA WR L-560.)

7. Westphal, Willard R., and Dunavant, James C.: Application of the Wire­Mesh Plotting Device to Incompressible Cascade Flows. NACA TN 2095, 1950.

8. Stack, John, and Von Doenhoff, Albert E.: Tests of 16 Related Air­foils at High Speeds. NACA Rep. 492, 1934.

9. 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.

10. Erwin, John R., and Emery, James C.: Effect of Tunnel Configuration and Testing Technique on Cascade Performance. NACA Rep. 1016, 1951. (Supersedes NACA TN 2028.)

11. Herrig, L., Joseph, Emery, James C., and Erwin, John R. : Systematic Two-Dimensional Cascade Tests of NACA 65-Series Compressor Blades at Low Speeds. NACA RM L51G31, 1951.

12. Heaslet, Max A.: Theoretical Investigation of Methods for Computing Drag From Wake Surveys at High Subsonic Speeds. NACA WR W-l, 1945. (Formerly NACA ARR 5C21.)

CO~

NACA RM L53G15 co.- 21

13. Hauser, Cavour H., Plohr, Henry W., and Sonder, Gerhard: Study of Flow Conditions and Deflection Angle at Exit of Two-Dimensional Cascade of Turbine Rotor Blades at Critical and Supercritical Pressure Ratios. NACA RM E9K25, 1950.

Axial

e

(")

.

'.'-'"

'

0,\ , ~

~

Figure 1.- Turbine cascade notation.

IV [\)

t ~

~ ;J:>

~ ~ V1 \jI o I-' V1

1.0

I~;=oa I I I (j,=~Oo

I I

1\,8, = 15° 1\,8, = 30° f3, = 45° Impuls e

\

.8 ~ Q) .0 E j c:

.s::. .6 0 0 ~ C'I c: .... Q)

1: .4 Q)

E j

E ';< 0 ~ .2

00

~ \ \ '\ \ \ ~ ~

'\ \ \ ['... l\,

'" I~ \ "I-

~ ~ ~

'\

~ ~

'\

~ - -

\ \ \ \ \

\ \ 1\ 1\ \ \

!\ '\ \

\ !~ [\ '\ \ \ I'" 1'\

" ~ \ 1\ ~ ~" ~ '1\, \ Max

"" reac ......; --~

mum ion

J

20 40 60 80 100 120 140 160 Turning angle, B, deg

Figure 2. - Maximum entering Mach number for cascade combinations of inlet angle and turning angle.

~~7

~ a ~

~ t-t

\.J1

~ I-'

\.J1

I\.) \..N

24

Yc

2.0

00 50 100 X

21== I ---I 100 0 50

Xc

A3K7 mean-line coordinates

for C10

- 1.0

o 0 0.5 0.397 1.25 0.636 2.5 1.426 5.0 2.359 10 3.689 15 4.597 20 5.217 25 5.623 30 5.852 35 5.936 40 5.897 45 5.753 50 5.516 55 5.200 60 4.814 65 4.367 70 3.870 75 3.326 80 2.746 85 2.133 90 1.465 95 0.801

100 0

dyc

(dxc )0.5 - 0.6574

dyc (dxc )95 • -0.1602

NACA RM L53G15

Thickness distribution coordinates t/c = 20 percent

(Stations and ordinates given in percent of chord)

Xt Yt

0 0 1.25 3.469 2.5 4.972 5.0 6.916 10 9.007 15 9.827 20 10.000 25 9.699 30 9.613 35 9.106 40 8.594 45 7.913 50 7.152 55 6.339 60 5.500 65 4.661 70 3.846 75 3.067 80 2.406 85 1.830 90 1.367 95 1.101

100 0

LER/c • 4.407

TER/ c • 1. 000

Figure 3.- A3K7 mean line, and thickness distribution, and coordinates

for primary turbine-blade series. For mean-line designation system, see reference 9.

NACA RM L53G15 25

2.0

0

7::~ 0

~ X

Yc 2:L- -20 I ~ 0 So 100

0 So 100 Xt

Xc Thickness distribution r-

mean-line coordina tP-s nates

B1E111 tic .. 20 percen· for C1

• 1.0 Q (Stations and ordinates given

Xc Yc in percent of chord)

0 0 Xt Yt

0.5 0.336 0 0 1.25 0.718 1.25 2.583 2.5 1.253 2.5 3.282 5.0 2.128 5.0 4.w 10 3.480 10 5.001 15 4.506 15 5.952 20 5.291 20 6.995 25 5.862 25 8.063 )0 6.261 30 9.025 35 6.514 35 9.727 40 6.642 40 10.000 45 6.652 45 9.725 50 6.551 50 9.009 55 6.342 55 8.016 60 6.016 60 6.908 65 5.551 65 5.848 70 4.983 70 5.000 75 4.332 75 4.3l2 80 3.680 80 ).624 65 2.824 65 2.936 90 1.953 90 2.248 95 0.948 95 1.560

100 0 100 0

~c) LER/c .. 3.300 (-dxc 0.5 - 0.5657 'lER/c • 1.000

~c ~ (dx~)95 --0.2017

Figure 4.- B,E,I, mean line. and thicknpgg distribution, ~~d coordir~te5 ~ ~ ... '

for secondary turbine-blade series. For mean-line designation system, see reference 9.

C

(")

, ~ ~ I Xc

radius

Yc

·Yt

-Yt

Figure 5.- Method of blade construction.

~

Trailing-edge radius

f\) C\

~ ;I>

~ t-t \Jl \.>I Q f-' \Jl

NACA RM L53G15

Primary turbine blade section, 9c - 65°, tic s 10 percent

Primary turbine blade section, 9c = 80°, tIc ... 10 percent

Primary turbine blade section, 9c ... 950, tic - 15 percent

Primary turbine blade section, 9c ... 110°, tic • 15 percent

Secondary turbine blade section, 9c - 120°, tic = 25 percent

Figure 6.- Profiles of the five blade sections tested.

27

Flexible

Adjustable floors Side-wall circular plate

Total-pressure probe Downstream static-pressure orifices

() Total-pressure rake

40" diameter upstream duct

Downstream floor plates

Protruding slot

upstream static-pressure orifices Ciffuser section

4

Return duct ~

Figure 7.- Cross-sectional view through high-speed cascade tunnel.

J\) co

()

~ :t>

~ t-t \J1 \jj o I-' \J1

NACA RM L53G15 29

s

4

-3 /'t ~ r--<>,.

2 ~ on

I 17 r-s

2

( ~,.... ,

~

0

o Upper surface D Lower surface

IA r-- ~r-.

~

.n' V

~ ~ ;J

"U:.. r-J

v

&41 II II I I 4.8

d\ S

32

~

~ ~

V r/

V 1.6

o ~ -20 0 20 40 60 80 100

Percent chord

o -2 0 20 40 60 80 100

Perc.tt chord Cb) a,=3S.2"; 8=49.3~

-20 0 20 40 60 80 tOO Percent chord

CO) a, =29.2·; 8= 42.3": (cl a,= 4,r; 9 = 55.S·.

6.4 J6

if\.

-u- )-C

"0... ::::.-

4.8 '2

/

;5 /

s 32

....cr" -0.., ~ r--I b'-'"

s 8

rI ,/

r/ 1.6

,/ :/

4

~ o ~ j,.,A:I

~ o -20 0 20 40 60 80 ,00 -20 0 20 40 60 80 ~

Percent chord Percent chord

Cd) a,a 47r; 8" 61.2·. Cel a, "S3.2·i 8=66.9":

Figure 8.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 0°, 0 = 1.5, 9c = 65°.

30

68

64

60

56

52

8, deg

48

44

40

36 24

-P"

7' g

-

08

o Cw,

¢ Cd,

~k. 0

vr ~ ~ V t7

-----V V

7 /

/ p-

/0 / ...('y ---p-

Jr-~ n- l..------ -V

II 28 32 36 40

a" deg

V V

V / V

~ ~

A I----V-0--

u-

44 48

NACA RM L53Gl5

/ V

V I--A..

A

~----V

Lft

~ 52

.16

.14

.12

.10

.04

.02

o 56

(f) Section characteristics; arrow shows design angle of attack.

Figure 8.- Concluded.

c

160

140

120

100

80

L o

60

40

20

a

•

NACA IM L53G15 31

s

o Upper surface

4

(~

i '", \ \ "'" ~ ( ,

~

~ ~ 1...0...

~, ""1~r-

s /.6 1.6

s 2

,Y V ~ ..n-l ~ no p-O-"' ~ ~ ~ .8 .8

i r ~ ~

,I o o o ~ -20 0 20 40 60 80 /00 -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord Percent chord Percent chard

s

Ib) a,= 322-; 8= 44.~ Icl a, =382-; 8=52.0-.

6.4

4B

,~ "\

'0...., ~ I'--n.

32

~ t-~

r .,.,....

1.6

o ~ -20 0 20 40 60 80 100

Percent chord Cd)a,·46.e-. 8 =60.8-.

s

6.4

\ '\

4B

~ i>--o.. ~ ~ r--

32

V .L ~

1.6

o V -20 0 20 40 60 80 100

Percent chord

Ie) a," 53.2"'; 8 .. 67.2"'.

Figure 9.- Low-speed pre~sure distributions and section characteristics for the cascade combinations ~1 = 15°, cr = 1.5, Bc = 65°.

32

68

64

60

56

52

8, deg

48

44

40

36 24

I o 8

[J Cw,

<> Cd,

Jj,. ~ ,D

17 ...;s- --z-> ~ ,/

7J'

V /

/ ~ V

/

V >f k:: ~ j-----i..I

1 ----

28 32 36

CO~

V 1/

V /

/ /

./~ j., ---~

~ ~ 1-0"

~ ~

40 44 48 a"~ deg

NACA RM L53G15

1/ V

p- -n...

-l--<>-

-n-

~ 52

.16

.14

.12

.10

.04

02

o 56

(f) Section characteristics; arrow shows design angle of attack.

Figure 9.- Concluded.

- ---- ----------------

160

140

120

100

80

L o

60

40

20

o

iD ..

•

NACA RM L53G15 33

s

o Upper surface o Lower surfoce

:1 H IIII :1 N IIII :\ k 111II \ \ s \

"<\

, \ i

I ~ I

~ ~

s 1.6 1.6 2

"1~ f'--o....

I "0...., j ~t-

~ ~t-.8 .8

~ ~ ~

j,....o-' i

nr::

: ~ I ~ ~ ~

.~

r-I Ijl ~ o o o

-20 0 20 40 60 80 100 -20 0 20 40 60 80 100 -20 0 20 40 60 80 100 Percent chord Percent chord Percent chord

(0) a,· 322; 8 = 44.1°. (b) 0,=37.8"; 8· SO.5. (c) 0,= 43.8"; 8=56.8°.

s

4

1 \ \

3

\ 1\ ~ ~

2

~ ~ 1 ,....rr trr ~

r vr -20 0 20 40 60 80 100

Percent chord (d) 0,= 46.8·; 8=59.7°.

S

6.4

4B

I~ \

32

~ i'"--n.. -0...., ~-

1.6

~ -u

~ o -20 0 20 40 60 80 100

Percent chord (e) a, = 53.2· i 8 ·66.2~

Figure 10.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 30°, cr = 1.5, 8c = 65° .

C01llllt

68

64

60

56

0, deg

52

J ?

48

"" '" 44 --0-.

40 28

;:J

I 08

o CWI

¢Cdl ,6b.

0

V ;.

V /

I? /

~ V -z..J

/ /

V V 1/

V i'J i{

...L -L-

Y

it 32 36 40 44

ai, deg

/-V

/ va /

'-' -...

/'

-X ~

48

NACA RM L53G15

r{ V

/ V

r-------~

- L--o-~

~ 52

.14

.12

.10

04

02

o 56

(f) Section characteristics; arrow shows design angle of attack.

Figure 10.- Concluded.

c~

140

120

100

80

L o 60

40

20

o

NACA RM L53G15 35

s

o Upper surface

6.4

.f>o-..

l I ~"\, ~~ ~ 1'--0.

..nO :;if fo;

4.8

I J I , V

I J ./

~

j s

2 32 j

V

s 32

I V I r/ ! ~ V i

/ ---'~

1.6 ~

/ 1.6

I 1;-'OOl iJU'

o o ,pcP w

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord -20 0 20 40 60 80 100

Pereen! chord Percent chord

(0) a, = 32.4°.; 8 = 50.6". (b) a,' 38.4°; 8 = 57.5°. (c) II, = 44.4°; 8 = 63.3".

16

!~ ~ ~ 16

12 Q

f I 12

.rf3 ~ ~ r-o-..

s 8

? / .y? !

s 8

;rP J ~ /

V 4 / /

4

.rlJ""' It-lf' ~ o I~ ~

,.4 o

-20 0 20 40 60 80 100 -20 0 20 40 60 80 100 Percent chord Percent chord

(d) a, = 47.4°; 8= 65.7°. (e) 11,= 56.4"; 8=73.7°.

Figure ll.- Low-speed pressure distributions and section characteristics :for the cascade combinations /31 = 00 , a = 1.5, 8c = &;0.

88

84

80

76

72

68

8, deg

/ /fi

64

60

56

" 52

I;{ 48

32

o 8

[J CW1

o Cd l ~J...

0

l::,.

~

/' V r---I---- V "'/ V V ~

V ~/ II /

<>

V i P V V

/ / L tp' lY

L V

/ / V V

V 1/ 9'

/ -V Q V -v- V ~ 0

L.r

l 36 40 44 48

ai' deg

/! V 1/ ~/ II <> ~ 1\

~

V 11\ /

V 0 0

52

/ l! If

/

/ ~

fU

~ k

I~

~

NACA RM L53G15

.20

d

lL 18

16

.14

'"

.12

Z V

Co

l"-.06

.04

.02

200

180

160

140

120

100

L o

80

60

40

20

~ 56

o 60

o

(f) Section characteristics; arrow shows design angle of attock.

Figure 11.- Concluded.

co~

NACA RM L53G15 37

s

o Upper surface

Po. ""'" Y). u P"" ~ '~

s '-'--< ~ -

J I

.~

r-o... ! J

I V

s 2 2 32

/ V

V V

I l.-rr' V-

i i

~ y-D'

00nr

/' V'

nrr V 1.6

'r£P pr i

I T"

o o o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Pereent chard -20 0 20 40 60 80 100

Percent chord (eJa,=53.7°; 8=72.5-.

Percent chord (0) a, = 41.4-; 8 = 59.5~ (bl a,=47.4°; 8=65.7°.

s

6.4

i.,.-,

00 ~ >-0..., "0...

4B

32 1

I 1.6 /

/' lboo' ~ o

-20 0 20 40 SO 80 100 Percent chord

(dlo,=56.7-; 8=75.2°.

s

16

12

8 1,,-0..

rP' 'V--( i)-..o.,

Ivory IY'-'

/ rr'

,/ 4

o ...oQl po--t ~ -20 0 20 40 60 80 100

Percent chord

(el a, = 63.S0; 8 = 81.8-.

Figure 12.- Low-speed pressure distributions and section characteristics ~or the cascade combinations ~l = 15°, cr = 1.5, 9c = Boo.

~-

38

84

80

76

72

68

0, deg

64

60

56

52 40

?

Jf

-<>--0-

o B o CW1

0 Cd l L:. .b.

0

.,/ ~ ,,-/'

/ / ~

n---

Il 44

s:

/ ~ ~ W- I::, ,...-

./ V A

/

.--£) ~ - -<.;r

" r1

~

"'I..J-

48 52 ai' deg

NACA RM L53G15

.16

l7 /

.14

r

// .12

1/ /

V L--- -IS

..()..--

........-

56

.10

~

~ <>

IV 1/\

/ ~

('/ p v 7

.04

V L--r= .02

~ o 64 60

180

140

120

100

80

L o

60

40

20

o

(f) Section characteristics; arrow shows design angle of attack.

Figure 12.- Concluded.

NACA RM L53G15 39

s

o Upper surloce

4

i bb " ~ I fbo. I I

~ ! 2 ~

~ t'-n..

s 2

s ~ &, !~ ~

2

! ~ ~ f- "':::;1-

./ ~ I ~

yO'

in. 1,-0-1 i I,.-cr'

.-<jr:r' ~ ~

II o [pOU 1'-" fbcP jJ-'

o o -20 0 20 40 60 80 100

Percent chord -20 0 20 40 60 80 100

Percent eIIord -20 0 20 40 60 80 100

Percent Chord

(0) 11,=45.2"; 8 =59.4°. Ibl a, = 49.0"; 8= 66.2°. Ie) a, =55.0°; 9=72.6°.

6.4

4B

s 32 1\

q ~ Po.-.

fo-i ~I-1.6

....u-' IrP

o ~ II-'-'

I~ -20 0 20 40 60 80 100

Percent chord

s

6.4

4B

~ 32 \

~ ~

-"-V 1,6

o L.JP ~Ir'

-20 0 20 40 60 80 100 Percent c:lIord

(.) a, = 66.So, 9 : 84.r.

Figure 13.- Low-speed pressure distributions and section characteristics :for the cascade combinations ~1 = 300 , C1 = 1.5, 9c = PlJo.

40

88

84

80

76

72

9, deg

68

64

60

56 40

o 8

[J Cw,

<> Cd,

1::.. .I.. 0

/ ;I

/ /1{

/' VA

t-'-'" -/ -----, t'--- ~

V V V"""

jJ L-----' tv 17 h--------!--D

Sl/ .--I

1,0"

t 44 48 52 56

ai' deg

/ /

V -V

-U

60

NACA RM L53G15

Ip

V 1/

/Y

C> V t::r

~ ~IJ-

~ 64

.16

.14

.12

.10

.04

02

o 68

160

140

120

100

80

L o

60

40

20

o

(f) Section characteristics; arrow shows design angle of attack.

Figure 13.- Concluded.

NACA RM L53G15 41

s

o Upper surface o La .. er surface

~I I I I I I I 16\ \ I I I I I 12

I 12

I ,

I 8 s

8

I J:X)--(

I I\r)("

,.pd ~ -u-., ~

I c ,,-rI I

4 / .p- ~c-

,..,...;: ? 4

\ ~ ~

o ~

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord Percent chard (a) a,=50.6·i 8=755. (b) a,=54.6·; 8=78.8·.

16

12 12

rj s

8

IP"" I"o-c ~ rP f-

P Ii

)

s 8

4 p:? )

.;I ~;e1

4

o ~ V (,.pq ~ o

s

16\ I I I I I I 2

8

,/ IP'"' i"o--< ~r ';:Pd 1

< Jf' V

i,..d' 0 -20 0 20 40 60 80 100

Percent chord (c) a, =57.0· i 8 = 8f.90.

r r--o... "Q,. -

f

j V

V ,...c'

-20 0 20 40 60 80 100 -20 0 20 40 60 80 100 Pwcent chord Percent c:IIor d

(d) a,=6U-; 8"85.6~

Figure 14.- Low-speed pressure distributions and section characteristics for the cascade combinations ~l = 15°, a = 1.5, 9c = 95°.

8, deg

92 I I I

09 -y L------r--- 0 CW1 / /

L-------r--- <> Cd I _ /

~ L 6L

r- 0 ,J3 b=-:;~ 6 / ~ - ~V--~

~----- / ~ r 2~7 <> _ -~ '~L L:i 0

./ -v / 0 ~~ ~t::::",_ /' ~

88

84

80

~r 0 _ ~

~ ,-76

~ ~ I J

68 7~ 52 56 60 64 at, deg

.10

.08

.06 Cw, a

Cd l .04

.02

o

ef) Section characteristics; arrow shows design angle of attack.

Figure 14.- Concluded.

200

160

120

L -o

80

40

o

.j::"' I\)

~ » ~ t"i \Jl VI

~ \Jl

NACA RM L53G15

o Upper surface o Lower surface

41 I I I I I 1 4/ 10-1 I I I I 41 I~ I I I I I • fOoc'Jn -~Po,.., '~~ !

, II % i 'tu

q

~ ~ 3 3 3

i \"u.. i ~ -

~ '7 f-

~ l-

I s s S

2 2 2

i d

/ I ..if'

VU

;I i ~ V ..r::J-""' lPor V ~ I 1

i p

~ Jr:T'

o o o -20 0 20 40 60 80 100

Percent cllor d -20 0 20 40 60 80 100

Percent chard -20 0 20 40 60 80 JOO

Percent chord

(0) 0/=51.3°; 8=75.2°. (b) 0/=54.1°, 8=78.1°. (cl a/ = 60.3°; 8 = 84.4°.

4 6.4

t'\ ...n. "'"' ""-(

~ ~ I-

3 !p:>-<

V '-"< ~ ~

4B

J V

s 2 /

I s

32

~ .......

l-4 .-

,(

/ 1.6

r ~ o !fno ~ o -20 0 20 40 60 80 /00 -20 0 20 40 60 80 100

Percent chord Percent chord

(d) a,·64 ... ,8=88.4~

Figure 15.- Low-speed pressure distributions and section characteristics ror the cascade combinations ~1 = 30°, a = 1.5, 8c = 95°.

44

96

92

88

84

8, deg

80

76

J o cWI

<> Cd l ~J...

0

6

V----------- /\

~ ~ /

/ V

V -A. 7

~ V Jf 0

1 52 56 60

ai' deg

/

J~ V

-

64

NACA RM L53G15

/'

// va ~

L.l.

~ ~

rY

"'LI

~ I I

68

.12

.10

.08

02

o 72

(f) Section characteristics; arrow shows design angle of attack.

Figure 15.- Concluded.

120

100

80

60

L o

40

20

o

NACA RM L53G15 45

o Upper surface

:I IIII11 3

i iR R R-i

1~

I ,~ "

5 5 2 2 ~

I' \ ~

5 2

, , i I

~

~ R e-

I I ,

I ,..l~ -~ ~ ~ f-

...Q'" If ~

0 ~ rm- ~ ru

1'\..U"

o o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord Percent chord Percent chord Cal a,=5L6°; (}=74.So. Cb) a,=54.6°; (}=77.9°. Ie) a,=6Uo; (}=84.8°.

4

3

R 3 ,\

'\ 2 \

~ ""'

s ~ .~ lao...,

s 2

'""'-c iu.., Il-<...

I V W

u..., ~ f-

~v ~ rr

01 ~ ........ i--O'"

~ IJ"T o -20 0 20 40 60 80 100

Percent chord -20 0 20 40 60 80 100

Percent chord

Cd) a,-64"'; (}=87.9°.

Figure 16.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 45°, 0 = 1.5, 9c = 95°.

,--

96

92

88

84

8, deg

80

76

72 48

o B

o cw,

<> Cd, /:i!:.

0

~ ~

ft

V ../' / /',

.JL L /

fJ

52

.....,

/ /

V

56

Li ~ P"

:

IV

A

V

L.I

60

ai' deg

~

~ ~

/ / ~ ~ F

l.-----' 'V -~ lIT'

~ I ~

64 68

.12 120

.10 100

.08 80

.06 ~60 Cw,

8 -1 L Cd, D .04 -----' 40

.02

D 72

20

(f) Section characteristics; arrow shows design angle of attack.

Figure 16.- Concluded.

+" 0\

(")

~ ~ t:-t \Jl

~ I-' \Jl

NACA RM L53G15 47

s

o Upper surface

6.4 6.4 o Lo-e, surface

6.4

I~ 0-0.

"'r,

! 48

I ~ ~I J ~f-

48 I I ~

-"i-

I J 4B

.i ~ >-0...., .JJ:rP' p....,.,

I ~ -'f-/

32 ·rfl iP /

If / s s c? /

I"' / 32

I I / I ~! /

1.6 V

ip.. '/ 1.6 /

V 1.6

I ~ ,.--u

I o 11, ~

.,.P o

-~ roo--'

,/'

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent cl1ard -20 0 20 40 60 80 100

Percent cllord Pere.ent chord (bJ 12,=605°; 8=94.00. (c) a, =63.5° • 8= 97.00.

\6

12

d ~

s B I ~f-

f / s

4 V J

sfP 1/ ./

I nr.n. n..A V a -20 0 20 40 60 80 100

Percent chord

\6

~

T ~i-

if 12

d I j I

8

IT II d

~IP-' / " ..D""

.t' o -20 0 20 40 60 80 100

Percent chord (a) a, = 73.0"; 8 = 105.0".

Figure 17.- Low-speed pressure distributions and section characteristics for the cascade combinations 131 = 30°, a = 1.5, 9c :::: 1l00. Wire screen in throat.

108 o 8

o Cw,

104 <> Cd,

Ll,L o AS L

100

96 r-

8, deg

92

88 )

84 52

/ V'-"

/ A V

fA/

L /

...u--r _____

1 56 60

L V

/

~ V

/

A .--/ v

...---

64

a" deg

~t--., ~ ~ /' / '7

/ ? / V

r! J V /

<! I

/ / ~

JV ~

V

~ I

68 72

I

I

I I

.12

./0

.08

.06 Cw, a

Cd, .04

.02

o 76

(f) Section characteristics; arrow shows design angle of attack.

Figure 17.- Concluded.

120

'00

80

60

L o

40

20

o

+" co

~ o > ~ t-l \Jl

~ f-' \Jl

NACA RM L53G15 49

o Upper surfoce

-"V( ~ '""I\..,

~ "'::::,-

s 1.6 ~ ~

/ s

1.6

/ r\, V

.8,....i -+I+-'t--+--I---+-:lr---l .8 V

Irq .,J? .8

~r ~ Ibn-l Ir'" O~~~~-L-~~~ -20 0 20 40 60 80 100

Percent chord

o -20 0 20 40 60 80 100

Percent chord ?2O 0 20 40 60 80 100

Percent chord

(0) a/= 58.4·; 9 =88.00• (bI a, =64.7"; 8 = 95.9°. (c) a, = 67.7"; 9 = 99.50

.

32 6.4

~

'too pi" '., "-0. ~t-

2.4 4B

,,0 ~

/ /

5 1.6 rf ~ -

tool ,.0 /

5 32

/ /

n rI .8

/

;£ 1.6

)JY, poet-"' o \f "ft

"UL' o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord Percent chord

(d) a/"74.6·; 9= 106.4~ (e) a, = 79.S"; 9 = 111.5".

Figure 18.- Low-speed pressure distributions and section characteristics for the cascade combinations ~l = 45°, a = 1.5, 9c = llOo. Wire screen in throat.

50

112

108

104

100

B, deg

96

92

88

84 56

\

,

./

, o 8

o CWI

<> Cdl

Ak 0

~ ~

,"-~

[b

~ <:

~ Vc

1/ £:.

---Ir' V ~

60

r". <> ~ k/ / ~

V 0

~

"" ~ V

11 64

/ /

1/ /0

~ ---- e L---

l---V

~

r--. 0

68 72

a" deg

1/ 1/

() [).

.---~ <> [).

0 0

NACA RM L53G15

14

/ .12

.10

~

.04

L.

02

140

120

100

80

L D

60

40

20

~ o 80

o 76

(f) Section characteristics; arrow shows design angle of attack.

Figure 18.- Concluded.

C

..

NACA RM L53G15 51

s

o Upper surface o lower surfoce

41 1 1 1 1 I 1 ,P..., V1\L1 1 1 &41 I 1 I I I I 3 r u...,

"a.. 4.8

~ P--o.

3

,..p...

! ~ "-0..

~ ~

2 V

fi V"'

s 2

s """"""I "'0..., ~

,.,...... :/ 32

I .-d' r

i

! r / r

./ V

1.6

!

II o ~

o r o -20 0 20 40 60 80 100

Percent cllor d -20 0 20 40 60 80 100 -20 0 20 40 60 SO 100

Percent chard Percent chord

(0) a, = 29.2°; 8 = 43.So. (b) a=34.2"; 8=48r. (e) a,= 40.2°; 8= 55.1°.

6.4 16

4.8

0- h-o cY ~ :u- >-..a.

/ 12

s 32 /

I

..()..

.,/ -0- .J">.

~ / s

S

1.6 I

~ ,/

V 4

o V I..-!r' / o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord

(cf)a,=47.2"; 8=62.0".

Percent chord

Figure 19.- Low-speed pressure distributions and section characteristics for the cascade combinations ~l = 0°, a = 1.8, 8c = 65°.

C

(")

68

64

60

56

8, deg

52

48

i I

! 44

I

I 40 24

I .4 ~ o e

t:,. ~ V [J cWI

~ V ~ <> Cd l /

~ V Llb. (', 0

L L V/ V

/' cI

~ /' V 6 /

t:,. V V V

/

V / V ~ /

V L

/V ----

-0--

L( V v

/ ..-0-----0

~ / I-Y ---"0.. / b- b--.JY V p-- ,., --n---jo-- ~~

-M I~ / -0- ~

/ 1 ~ I i

28 32 36 40 44 48 52

a" deg

(f) Section characteristics; arrow shows design angle of attack.

Figure 19.- Concluded.

./4

.12

.10

.08 CW1 a

Cdt .0

04

02

o 56

140

120

100

80

~ 60

40

20

o

VI I"\)

~ ~

~ t-t VI \>I Q I-' VI

NACA RM L53G15 53

s

o Upper !I.'ffoce o Lo.er surface

:I IJ IIII J H IIII :I NJ III i I 'Q. ~ '\ 1"0-I ~ I ~ r-o...

~ ""-< ~r-

s 1.6 1.6

s ~ ~-

~ ,...a

.n--'

1.6

I ~ -eP i Ir-D"" rF--., ~

jt-O""" .8 .8 V .8 ,

I i

;J If' I o o

p o

-20 0 20 40 60 80 100 -20 0 20 40 60 80 100 Percent chard

-20 0 20 40 60 80 100 Percent chord Percent chord

(0) o,=26.~; 8=39.S-. (b) 0,=33.0°;8= 46.9". (c) 0, =39.2° ; 8· 53.6°.

s

4

1\ '"\

3

'\tA-. ....,~

t::2:. 2 !,.-if

~ I"P

/ I I,/"'

V-~

-20 0 20 40 60 80 100 Percent chord

(d)0,"45.o-; 8=58.9".

s

6.4

4B

1\ ~

t"'-9-v..,

r--o.. ~ r-

32

~~ Ir""'"'

If" 1.6

L

o ~ -20 0 20 40 60 80 100

Percent dlard

(e)o,·5I.JO; 8"66.~.

Figure 20.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 15°, a = 1.8, 9c = 65°.

C~

8, deg

68

64

60

56

52

48

44

40

36 24

NACA RM L53G15

.16 160

D

o 9 / C cWI

/

'/ <> Cd l /

.14 140

A J... /V 0

Va .12 20

I/o V

.10 00

V /

l----&--V V --- "-"-./' l::,-r--.

.Lx V 1/ """-V ~ L..l

~

/ / ...(';

04

/ y -----f-<>-

-II

P' --- r;: ..('

~ .A- -Q-u LI

02 20

1 ~ I

44 48 32 36 40 28 52 0 o

ai' deg

(f) Section characteristics; arrow shows design angle of attack.

Figure 20.- Concluded.

C

NACA RM L53G15 55

s

o Upper surface o Lower surface

32

Po

\ 2.4 :1 N IIII :1 KIIIII

1\ I ~ i i

I.S 1\ ~

s I.S ~

~

s I.S

I '0.-, 1 ~f--.8

~ ~t-.8

---. ~f-

.8 I r ~

~ ~

L r o o o ~ -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Pen:ent chord -20 0 20 40 60 80 100

Percent chord Percent chord (0)0,=35.2"; 8 =48.8". (b) 0,=382°; 8= 51.S". (da,= 442"; 8=58.0".

s

4

3 !\ 1\

2 \ f'<\

1'0-K::::

~ ..",-

o W -20 0 20 40 so 80 100

Ptrcent chord

(d)a,=47.2"; 8=60Jr.

s

6.4

4.8

K I~ \

32

'0-., r-o-~

~r ~ jl-cr

I.S

V ~ o -20 0 20 40 60 80 100

Percent chord (e) a, =532·; 8=S7.1·.

Figure 21.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 30°, a = 1.8, 9c = 65°.

68

64

60

56

8, deg

52

48

44 32

NACA RM L53G15

.12 I

/ 0 8 [J Cw, / <> Cd,

A 1:. / 0

.10

/ .08

/ /~

/ / v-a-r-----.

~~ L--------- jCS / p -<->

/ ~

/ f)

~ P J'..

A ~ i< --- L---cr M '-' -U- r-'

.02

1 ~ I o

36 40 44 48 52 56 ai' deg

(f) Section characteristics; arrow shows design angle of attack.

Figure 21.- Concluded.

C~

120

100

80

60

1:. o

o

20

o

) NACA RM L53G15 57

s

o Upper surface o Lower surface

6.4

~ ~ ~ , f ~ l-

I r '-4S

..

12

:1111111 I

/

l! 32 B

.r..

,.p f7' ~ -

s s Iff"' vr

~

V j B

;I i a/'(1 I

1.6 ir fJ· ,;I

4

/~ ~ /

/

rt" V

4

~ o I pr

,...... 0 1 'pD d

[J..IU-o -20 0 20 40 60 80 100

Percent chord -20 0 20 40 60 eo 100

Ptn:enf c:IIard -20 0 20 40 60 eo 100

Percent c:hord Ie) a, = 44.4·; 8 = 63S·. fb) Q, =47.4·; 9=66.2·. Ie) a,= 50.4·~ 8=69.00.

20 32

15 H:).

If' -v...,. p I' ~""""

24

s 10 ! /

~ I ,/ /

i j s

16

5 V j

/ /

d Ir'

/

~~ l? B

o ~ ?

!"II (

..,n-l ~ o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Parc:ent chord Percent c:hord

!dl a,.56.4·~ 9=74.9" Ie) a,= 59.4·; 8=77.5"

Figure 22.- Low-speed pressure distributions and section characteristics for the cascade combinations ~l. = 00 , a = 1..8, 0c = Boo.

88 I

o B C CW1

84 <> Cd,

1\ ~ L 0

\ \ 80

1\

76 \ \

~ '" /).

72

0, deg

68 ~

£ ~ 64 L VI

I).}

/ 60 II /

lif

~ 56 44

V

/ V

48

~ ~ / V ~

d V V

52 ai' deg

d

V /

11 j

I / r ~

V 6 rr

I--- ./ V

iW" l

I

/ V

~ T

NACA RM L53G15

.16 160

.14

.12

.10

.04

.02

140

120

100

80

L -o 60

40

20

o 56

o 60

: (f) Section characteristics; arrow shows design anole of attack. d.

L ___ _

NACA RM L53G15 59

5

o Upper surfote

41 I I I I I I 3

~ ~

I~ ilu-n ~ ~ I'--f)

4

3

o La .. r surface 4

IOn )...

"-' JO.., h.,.

~f-3

-u....c ~ -

I A' 2

--el-/

.r/ I('

5 2 /

V 5

2

~ ~ V V

I LtJ V ! iPtD

lrif I paD W

o I o o -20 0 20 40 60 80 100

~rcent chord -20 0 20 40 60 80 tOO

P_t clIard -20 0 20 40 60 80 tOO

Percent chord

(a)a/~39.o-; 8=58.2".

5

6.4

4B

32

1.6

o -20

(b)a,=45.00; 8:642°. (c) a/" 48.0-; 8= 67.2-.

In. boo. "tJ JP=1 ~ e-

li )lP

J-vr

boP lID'

o 20 40 60 80 tOO ~rclllt chord

(d) 0,= 540"; 8 =72SO.

5

6.4

IP" -' f--P

fA>o k'1 j 4B

( V 32

I j~

./ 1.6

./

\JP jU"

o -20 0 20 40 60 80 100

P.cent chord (e) a,=60.o-; 8=79.1°.

Figure 23.- Low-speed pressure distributions and section characteristics ~or the cascade combinations ~1 = 15°, 0 = 1.8, 9c = 80°.

80

76

72

68

8, deg

64

60

56 36

I I )D 0 8 / 0 Cw, L <> Cd, V

.L.. L A 0 V L V V V

/ / ---~ 1\ V-~

----- V ~

V L 6 ~

I..-LY V VO r/ L V ~ V L ./

J .A V f--D-

v r v

/ ~ rr-II

V fJ

~

11 ~ I I

40 44 48 52 56 60 ai' deg

(f) Section characteristics; arrow shows design angle of attack.

Figure 23.- Concluded.

I I

.12

.10

.08

.06 Cw,

8

Cd, 04

02

640

120

100

80

60

1:. o

40

20

o

0\ o

~ ;J>

~ t-i \.n \)l Q I-' \.n

NACA RM L53G15 C~ 61

5

o Upper surface o Lo.e, surface

~ ~ 32 32

fh 2.4 \ 2.4 ~ 2.4 .~

~ .. I -u.

~ ! ~ >--0.

1.6

'oon .~

~ ~

5 1.6

~ ~

"-..

~

5 1.6

r-o--- ..-JY .... .8 ,.-if'

,..-IT

.-n-l .8 IY' vf .8

, fba v ,QOCi ~ f\p pr'

.i o [I !

o -20 0 20 40 60 80 100 $0 0 20 40 60 80 100

Percent chard -20 0 20 40 60 80 100

Percent chord Percent chord

(0) Q,=47.5°; 8=65.4°. (b) Q,=50.4°; 8=68So. (e) Q,=52.4°; 8=70.9°.

32

'\ 6.4 '-r--

\ 2.4 , 4S

5 1.6

-~

~

2 b..

\h 5

32

.8 JY' ~

..rOi r ~ ~ ~ ~ V

.if' 1.6

o r \PO tfT

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord Perce nt chor d

(d) Q,=54.4°·, 8=733°.

Figure 24.- Low-speed pressure distributions and section characteristics ~or the cascade combinations ~l = 30°, a = 1.8, 9c = Boo.

C

62

80

76

72

68

8, deg

64

60

56 40

---6--

-0-

}f

I o 8

Cl Cw,

<> Cd, J:::.L

0

---./Y

:...---

~ /

/ ~

u-

44

/

V /

V V

V~ ~ fY\. --.......

A

1-----0-<7

I-r ret .-----u-

1 48 52

ai' deg

NACA RM L53G15

.12

Jf /

/ k(

.10

/ V

.08

.06 CW1

8

-!J.-----V Cd,

fLS

..... A V

fa--

~ I I

56

.04

A

"V

-n- .02

o 60

120

100

80

60

L 0

40

20

o

(f) Section characteristics; arrow shows design angle of attack.

Figure 24.- Concluded.

C

.,

NACA RM L53G15

s

o Upper surface o Lower suri_

18

? IP' ~

1" ' ........ r-;f /

4B 12

r II rI ..n.

rl '" -u..,

..0-<1'-..

pi ~I-/

s s 8 8

/V ,/

Irflo. r/

...eli fF ;I-

I lA' ~ ~

,;I~ 4 1.6 4

~ ....

~itlo ,.-"-' D'l

V r o o o

-20 0 20 40 60 90 100 1Wc:ent chord

20 20406090 Pen:eIIt cllcrd

lbIa,"53S·, 8 -79.".

100 -20 0 20 40 60 80 100 Peae"t chord

[a) a,w52.l"i 8· 77.'r. [e) a,-57.1-. 8 "8t.8~

16 18 .r.

p ~

~

r:r ro... ~

9 ~-

I

? / 12 12

p ''''-0 /

!J II f f

I

s s 8 8

4 I~ / 4 lei I

r/ p"

~ ~ kfG ~ o o -20 0 20 40 60 90 100 -20 0 20 40 60 90 100

Percent chord P.cent cIIord

(tI) 11,· ... 1· • 8· 85:"'. Ie) a,· &iU", 8· 88.6".

Figure 25.- Low-speed pressure distributions and section characteristics' for the cascade combinations ~1 = 15°, C1 = 1.8, 9c = 95°.

64

96

92

88

84

8, deg

80

76

72 48

I I 0 8

0 CWI

<> Cd l

A b. 0 ~

V l/ /~

-. ~ /

V~ ~ _/ /

/ V V

/

V / v u lY

./ /

/ V V

t--a ---/ ~ V

./

1 52 56 60

ai, deg

V

7 II

Jy [7

tI--"

V .,

i/t5

NACA RM L53G15

.12

.10

.08

.02

120

100

80

60

J.. o

40

20

~ I o o

64 68

(f> Section characteristics; arrow shows design angle of attack.

Figure 25.- Concluded.

D

•

NACA RM L53G15

5

o Upper surface o Lower surface

4

I~ "'Co. ~

3

~ .....

I lb. I

; 2 ~ P-o..

V 5

2 5 ( I 2

.I ! n ,....--I ~

! ~tlq ~ ~

l.4 I

~ ~ I 17

A>. lr,oJ Ii o (1

T'

" 1:1

-20 0 20 40 60 80 100 -20 0 20 40 60 80 100 Percent chord

-20 0 20 40 60 80 100 Percent chord Percent chord

(tI) 0, = 54.0·, 8· 79.(1'. (c) 0, = 6l6°, 8 = 86.7°.

5

4

.reP ~

3

ifrP u", r'u.

/ /

2

I I

~ n.-,. .?

n ~ -20 0 20 40 60 SCI 100

Percent chord

(d) a,a 63. ... ; 8=88.6.

5

6.4

L"o--<

1/ ......... ~e-

4B

~ I 7 32

J )1

1.6

V ~ tP

o -20 0 20 40 60 80 100

Percent chord (el a,' 67. '!1'; 8= 925'.

Figure 26.- Low-speed pressure distributions and section characteristics ~or the cascade combinations ~l = 30°, cr = 1.8, 9c = 95°.

66

100

96

92

88

8, deg

84

80

76

72 48

o 8

o CWI

<> Cd,

A 1... 0

/ l/

-< ~ ) ('

--4 1.

52

V ~

/\ "'-7

V V V

P i/

/ V

<...T

"1..J'

... 56 60

ai' deg

/J P<

J'-.

-U

NACA RM L53G15

+ I

1;5

/ 1 v II p

v V fY

V

A

'\z '-' -----I

~\ ~

I I

64

.14

.12

.10

.04

.02

140

120

100

80

1... o

60

40

20

o

<f> Section characteristics; arrow shows design angle of attack.

Figure 26.- Concluded.

•

NACA RM L53G15

s

o Upper surface o L_r sur10ce

3.2 3.2 3.2

Vb q

1\

2.4 f\. ~

2.4 V\ ~

2.4

1\ , 1.6

l I~ ~

s 1.6 I' vo...,

""-.

s 1.6

:P.. ~ ~ "1 ...0-' ~ ~

OJ""

.8 ~ N:h.

~ ~ ~

b.,o.A ~

.8 ~

~ ~ .....

~

.8

~

o o o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord (0)0,=53.5-;8 =77.7".

3.2

2.

5 I. 6

8

r~ \ ,

'q...

n. ./ yU'

t r

PMent chord Perclnt c:IIord (b) a, = 56S-, 8=80.9°. (c) a, t 61.S0; 8= 86.3°.

3.2

~ ~

2.4

s "'-'00. ~

~ ~ '0..; ~

1.6

./ l---cP

~

rI ...-a" .8

r l'-'

0 -20 0 20 40 60 80 100

o -20 0 20 40 60 80 100

Percent chord Percent c:IIor d

(810,a1O.5";8= 95.6°.

Figure 27.- Low-~eed pressure distributions and section characteristics for the cascade combinations ~1 = 45°, a = 1.8, Bc = 95°·

68

96

I---

92 I---

88

84

0, deg

80

76

72 52

d /

j5 "0.-"-U

1 I 0 8

0 CW1

<> Cd l L

Ll. 0

/' ~ V lei

/

V

--~ r--r0-

56

/ ~

/ v ~

V /

Vi A- J--LS

~ ~

~ A A ____ V

..n ~ ~

u~

t 60 64

a" deg

/ If

V

P v

r\---t:r I--

A r>--v

~ b-

~ I

68

NACA RM L53G15

.12

.10

.08

.02

o 72

120

100

80

60

1.:. o

40

20

o

(f> Section characteristics; arrow shows design angle of attack.

Figure 27.- Concluded.

C

NACA RM L53G15

s

o Upper surface D Lower surface

::111 Uil :111 ml :1111111 r:f / r< -

! J' / ! i ~ / I

lP I ri

~~ J ..r>-

rfY ""': r--

s 32 32

5 8

I

I I I

fu:l rI

( I lOrn ~

/ / ~ .rI'

1.6 1.6 4

I U ~ I

1/ iIoo--' <IA-ro ~ I~ o o o

-20 0 20 40 60 80 100 Percent chord

-20 0 20 40 60 80 100 Percent chord

-20 0 20 40 60 80 100 Percent chord

(a) a,=605"; 8=94.6~ !b) a,=63.5"; 8=97.8~ (cl a,=65.5°; 8=100.l~

16 16

12 p V<>--< ~

! / 12

-n.

~ /

5 8 /

/ II 5

8

l? V c# p'

4

p /

~x:fT 1I 4

<no.-. ~ rcn..r o -...." "d

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Per cent chord Perce nt char d

(ell a,=67.5°·, 8= 101.6~ Ie) a,=720"; 8=1053°.

Figure 28.- Low-speed pressure distributions and section characteristics for the cascade combinations ~1 = 30°, a = 1.8, 9c = 110°.

70

116

112

lOS

104

8, deg

100

96

92

88 56

I I

0 9

c CWI

¢ Cd,

A J... 0

9'

....l:J.. /~ ~ C-4 ~

V 1/ /

~ 60 64

d 7

I V V /

=A-. . A ......

n [7

j..--"

68

~~

/ II

/ y

..".A tr V

~ I I

72

NACA RM L53015

. 14 140

.12 120

.10 100

.OS 80 Cw, a L

Cd, 0 . 06 60

. 04 40

(f) Section characteristics; arrow shows design angle of attock.

Figure 28.- Concluded.

C

NACA RM L53G15

32

2.41---M~~-+-+---t--; 2.

s s I. 6

8

OL--L_L-~_~~_~ 0

o Upper surface e lower surfoce

koJ,..,

~ ~ ~

n

\ ........

'\ ~ 1'-"-'"

71

32

lin -un. lOf),..,

JJ<-

~

2.4

s ~

~ • 1.6

/ V

n /' .8

\ V ~

o -20 0 20 40 60 80 100 -20 0 20 40 60 80 100

Percent chord -20 0 20 40 60 80 100

Percent chord Pereent cIIord

(b) a,= 61.4°; 8= 94.3°. (cl Q, = 67.4°; 8 = 101.1°.

32 32 ~ ~ ~

I,d)Q ~ -<:

Co: p-<T '\ -r-

2.4 \ 1/ I '-bd I

2.4

/ /

s 1.6 J

j

s 1.6

~ .8 ,/

.8

(\ V ,..,0.., l;i :wu IbO'

o -20 0 20 40 60 80 100

o -20 0 20 40 60 80 100

Percent chord Percent chord

(e)o, s 76.0"; 8'" 109.9°.

Figure 29.- Low-speed pressure distributions and section characteristics ~or the cascade combinations ~l = 45°, a = 1.8, Bc = 110°. Solid line indicates wire screen in throat. Dashed line indicates no screen.

c

72

112

108

104

100

8, deg

96

92

88 56

1 o 8

C CW1

<> Cdl Ll .I...

0 •

/ l/

~

-z:...:. V

IA L I

V 1 -to-/

ref ~ r---- k-l.

60 64

/ 1/

/ /

I

1\ 1'-'

j\ --<>

---L>v \L -- <>

0

-0

1 68 72

a" deg

NACA RM L53G15

.12

/

!7 p-

.10

.OS

~ ~ ~

V

1- .02

~ 76

o 80

(f) Section characteristics, arrow shows design angle of attack.

Figure 29.- Concluded.

120

100

80

60

J.. o

40

20

o

)D NACA RM L53G15 73

s

o Upper surfoc:e o Lower surface

32

2.4 J IIIIII J IIIIII -Do

[)Q.J ruu-

"u.. 1.6

a. vor- -0-,

l'h s

1.6 ~ >-0, ""C:fJ' IV" ~

s 1.6

~ .. ~ ! b iJIn.." n. ~

B Q Li ~ Inn-. no V

.8 .,P

rtba ~ ~

.8

o o o ~ -20 0 20 40 60 80 100 -20 0 20 40 60 80 /00 -20 0 20 40 60 80 100

Percent chord Percent chard Percent c:fIord

(a) 1I,=69.0·~ 8=115.3" (bl 11,=72.0-; 8=1195" (c) a,~75.0-~ 8=122.5-.

s

32

I.

)...n

6 V ~ .~ ~ l'

1/ 8 /

IQh ..0-' ~

~ .........

0 -20 0 20 40 60 80 100

Percent chord (d) 11,=78.0-; 8-125.8-.

s

"-C

2. b-.o-, I'--..-.

I "\ -

1.6 6 I I~ ~ 1/

.8 j

d V'

r ~ 0 -20 0 20 40 60 80 100

Percent chor d

ee) a,=8LO·; 8=128.9"

Figure 30.- Low-speed pressure distributions and section characteristics for the cascade combinations (31 = 600 , C1 = 1. 8, Bc = 1200 • Wire screen in throat.

74

132

b '---

c 128 c---<>

124

120

8, deg

116

112

108

104 60

6

81

Cw,

Cdl 1.. 0

64

l i _\

\

~

'\ \

~ L \ Vb LI'-,

V , /

.-/

~ ~

~

68

/ ;I

/ / /

j ~ V

/

~ Il- n.

V~ '-'

.,. 72 76

ai' deg

NACA RM L53G15

.~

/

~ '" ~

~ Li ~ w.......

~ 80

.14

.12

.10

70

60

50

40

!... o

30

.04 20

.02 10

o 84

o

(f) Section characteristics; arrow shows design angle of attack.

Figure 30.- Concluded.

~

4.8

4.0

3.2

2.4

s

1.6

.8

o

r If

yt IT , .

01

.-V ~

~ / ~

V V

V /

V ,....A ./

~ ,.....

20 40 60 80 P.rcent chord

•

o Upper surface o Lower surface

12

h.. 10

i

V 8

i

6 i s 1 4

i

i

2 !

100 o

-- Scr

V If

::.l

~ ",.. L ~

V ~/

V IY"

V ty"

I

~

- Il--"" ~ Il:: -1

o 20 40 60 80 100 Perc.nt chord

(o)M, = .24!S,M2 a.!S07; CWI a.0166: 9=68.21'. (b)M , ·.319; M2 a .801i Ow a .0164;9=68.2°. I

. Figure 31.- High-speed pressure distributions for cascade combinations ~1 = 10°, ~1· 48.4°, ec • 80°, a = 1.8.

~ :x:-

i

~ V1

-J V1

4.8

4.0

f\ 3.2

()

2.4

s

1.6

.8

r ?

o

o Upper surface o Lower surface

12

1 / r--< ""--(

~ ~ 10

)---

j

I J 8

// /

y' Scr-

6

s / V V L .--' 4

/ ...J v~

.........c V

20 40 60 80 Percent chord

(0) MI =.289; M2 :: .561, 0 = .0202; 8 = 70.5°, wI

2

10 o o

~ p---L V

V V V V

..-I It-

~ ~ II-" J:)

o 20 40 60 80 Percent chord

(b)M, =.374; M2 = .953; Ow ".0228~8=70,2°. I

Figure 32.- High-speed pressure distributions for cascade combinations ~l = 15°, al = 50.4°, 8c = 80°, a = 1.8.

100

-..::J 0'\

~ !l>

~ t"i VI

~ VI

..

4.8

f').

4.0 Lt"

3.2

2.4

s

1.6

.8

./

I\tf o o

o Upper surface o Lower surface

12

~ >-~ )..

~ yo

~ ..J. ~ ).-.., :::::. 10

/

/ / - ~Scr

8

/ /

/ /

~ V ~[)..

V y i ./

~ V V /'

/'

6

s

4

/ ~ ~

./ Y

V V

.-I

2

.-~ ........ - - --- -_ .... ----- -L-~~f1" ~

-~

20 40 60 80 100 o o 20 40 60 80 100

(a> M, ... 196;

Percent chord

M2 " .394; Ow ... 0186; 8 = 73.30•

I

Percent chord

(b) M, '.315; M2 • .908; ° ·.0231; 8~73.2°. wI

Figure ;;.- High-speed pressure distributions for cascade comhinations ~1 = 15°, a1 = 5;.4°, 9c = 80°, 0 = 1.8.

~ :x:-

i t-I V1

~ V1

::j

12

10

8

6

s

4

2

o

o Upper surface o Lower surface

12

10

Scr

8

/ ~ ~

Ir ..... 6---' ) ,...

,..-/ /'

L / / V L /'

/ L

(

6

s

4

/ V

V V

~

~ V

L V 2

D ~ ~

~ _J

60 80 100 o 20 40

~ ~ jr--1 r--' .J.- _1 __ - __ o o 20 40 60 80 100

Percent chord Percent chord (0) MI = .186i M2 = .406; G • .0184; 8 = 76.00.

WI eb) MI =.282; M2 = .883; G = .0213;8=76.~ w,

Figure 34.- High-speed pressure distributions for cascade combinations ~1 = 15°, ~1 = 56.4°, 9c = Soo, a = 1.8.

-..l co

~ o :x>

~ ~ VI

~ I-' VI

4.S

4.0

'~ 3.2

2.4

s

1.6

.S

-----

~y o o

«

o Upper surface o Lower surface

12

10

-- 0 tr"" II--.. ">---.-~ S

j

V /

V

V V

6

s

4

L V V Ir""

~

2

20 40 60 so 100 o

Percent chord

(0) M, = .224; M2 ".402; C =.0162; 8= 78.1°. w,

-r- Scr

~ -/

~ ~ V j V

l/ ,../'

.IV V-b----£~ ~y ~ . - '-- ~ __ L. .. _

o 20 40 60 Percent chord

80

(b) M I ".340, M2 " .906. CWI" .0219; 8" 78.3.

100

Figure 35.- High-speed pressure distributions for cascade combinations ~l = 200'~1 = 58.4°, 9c = 80°, a = 1.8.

~ &: 12 t"l

~

-.l \0

4.8

4.0

t:\

{'" 3.2

2.4

s (

1.6

.8

p-C

:1 o )

o Upper surface o Lower surface

4.8

!

I

I

4.0 A I ~

l "' ~Scr I b

~ "'(F---<

~~ 3.2

~ ~

"'-(~

~ :>.--. ~ ~

~

~ / ---

'-'

'7 V V

~ ;;:.:....-

/ Iv-"

2.4

s

1.6

-r----.r/ ,...-r r

~

V V (

.8

20 40 60 80 10 o ~I o o 20 40 60 80 100

Percent chord Percent chord (a) MI = .239; M2 = .308; CWI = .0152;8 = 68.S. (b) MI =.474; M2 = .743; CWI =.0196; 8=68.8°.

Figure 36.- High-speed pressure distributions for cascade combinations ~l = 25°, ~l = 48.4°, 9c = 80°, cr = 1.8.

CX> o

~ (") ;t>

~ t-I Vl

~ f-' Vl

4.B

4.0

3.2

1\ 2.4

5

1.6

.B

p

11 o

CI

o Upper surface o Lower surface

4.B

4.0

3.2 !\ \ Scr

~ f\ ~ "--.; J........

'r~ >-

Y

\ ~ ~ ~ I-i p.....,

~

~ (

-~ y-'"

-"' ~

2.4

5

1.6

r----~ ~

.I~ ~ fI""

V y.--'

.--I

.~ ~ b----'

.B

P ~ ~

,. 20 40 60 BO 100

Percent chord

o o 20 I

40 60 BO Percent chord

(a) MI ... IB6j M2 = .2IB; Ow = .0153: 8=71.0°. I

(b) MI = .499; M2 " .655; Ow, =.0175 i 8 =71.00,

Figure 37.- High-speed pressure distributions for cascade combinations ~l = 30°, ~l = 50.4°, 9c = 80°, a = 1.8.

I

100

~ &;

~ I:"i

~ V1

t='

~

4.B

4.0

3.2

(') ~

2.4 (

s

1.6

.B

o V )

o Upper surface o Lower surface

4.B

4.0

3.2

~ 2.4

~ "-~ '-----< ~

~

2

S

1.6

.-f h .....

----~

,~ /

.8

tr------£

20 40 60 BO 10 o o Percent chord

(a)M I =.265; M2 = .330;Cwl

= .0153; 8=73.4°,

()

V '(

p-;

II o

l\

\ Scr

~ ---. h ~ t---.,

~

L.------' V V

h--'

.,/

/ V

~

~ - -<-

20 40 60 BO 100 Percent chord

(b) MI = .500; M2 = .732; C = .0IBO~8=73.B~ WI

Figure 38.- High-speed pressure distributions for cascade combinations ~l = 30°, a l = 53.4°, Be = 80°, a = 1.8.

OJ I\)

~ :x:-

~ t"i \Jl \.N

~ \Jl

4.8

4.0

(f\

\ 3.2

2.4 (.

s

1.6 i

.8

j'tf"' -- -- _ .. _-o C)

o Upper surface o Lower surface

4.8

4.0

'\ -Scr 3.2

\ ~ I

:'-,

II"" ~ P-I

~ P---< """' ~ ......-V

2.4

s ~ i

~

~ V -1.6

p..--l-I /

~ r---' ~

.8

20 40 60 80 100 o

Percent chord

(a) MI = .180i M2 • .233; CWI • .0157;8=77.20.

C"'"

lP' ~ o

~ p

. L

V~ v

/ V

~ V

~~ ,

20 40 60 80 Percent chord

(b) MI =.485; M2=.832j C = .0178; 8· 77.30. wI

Figure 39. - High-speed pressure distributions for casca.de combinations ~1 = 30°, ~1 = 56.4°, 9c = soo, a = 1.8.

100

~ f2 t-t

~ I-' \J1

CP \.N

4.8

4.0

(")

1\ \

3.2

2.4

s

1.6

.8 P

~ o o

o Upper surface

o Lower surface 4.8

4.0

~

1 1 r-Scr 3.2

\ \ ~

--< I---< ~ -....

;;

l~ ~ ~ - r-

------, -<~~

~ V 16.------:

2.4

s

1.6

-i~ --'

..--I 1-

V V V f-

/

~ ~ I~

.8

~

I - -60 20 40 80 100

V ~ o o 20 40 60 80 100

Percent chord Percent chord (a) M I =.272; M2 = .322; C = .0145; 8 = 78.2°.

wI (b) MI =.504; M2 =.659; CWI

= .0173;8=78.8".

Figure 40.- High-speed pressure distributions for cascade combinations ~l = 35°, ~ = 58.4°, Bc = 80°, cr = 1.8.

~

~ ;J>

~ t"i \Jl

~ I-' \Jl

4.S

4.0 I

yo-3.2

2.4

s

1.6 (

.D-

II .S

II o CI

/ .0--,

0-... ~

>J-Cf g '""" 0--:

"V

iD...

~ 0-

/ /

cr ./'

"V- 'W

20 40 60 SO Percent chord

/'

o Upper surface o Lower surface

12

I 10

~ -'

/ S

j 6

s

4

< (

2

100 o

Scr Ir. I,.., irI

~~tv IV ~

l~ j

~ ~ l! V

.0-

L .LJ

~ ~.

V PJ tw ~?

~'---

o 20 40 60 so 100 Perc.nt chord

(a)M I -270jM2 "·509j OWl - .0156,8=S3.2°. (b)M I ·.364; M2 =.797; CWI

=·0202j8=S2.2°.

Figure 41.- High-speed pressure distributions for cascade combinations

~1 = 25°, ~ = 56.1°, 9c = 95°, a = 1.8.

~ S;

i t:-t V1

~ V1

0:> V1

4.8

4.0

()..

P 3.2

o 2.4

5

1.6

p-

If .8

~ o o

o Upper surface o Lower surface

4.8 if~ ~ D

I scr- ---

\ ,if' --

..(). k-\. ir. lo- 0-~

'~ r-o...

4.0

~ y

-I-'"' ('\.

1 :> '0 f-d

L 3.2

~

/ tr-u

V / V

~ fD

/ L

V

2.4

5

1.6

~ "u

-0... ""-ro-

P til /v J - -0---

.8

I , 80 100 60 20 40

~ ~ 1 _ I I , o

20 o 40 60 80 100 Percent chord Percent chord

(0) MI " .236; M2 ".384; CW1

".0162;8 = 84.8~ (b) MI " .409; M2 ".943; CW1

".0219; e" 84.7°.

Figure 42.- High-speed pressure distributions for cascade combinations

~1 = 30°, ~1 = 58.1°, 9c = 95°, a = 1.8.

~

~ ~ t;-i

(;J (;) I-' \.J1

4.S

4.0

-0. ,

3.2

2.4

s

1.6

.S r Il o o

~ I.n

V ~ ~ J"ooroo ~..r(

/ .D

/ /

/ ./

~ ~ :a ,a-

-- -- ---- ----- ~~-- ---_._-

20 40 60 SO Percent chord

o Upper surface o Lower surface

4.S

4.0

.C\

~ 3.2

/ /

.d 2.4

s

/.6

.S

~

100 -o

1/ fV" -~ ru.-~ ~ IV"

/ ~ ~

I 0-0 ~ kr-d I? J

~ / V

~ V

~ ~

~ ../ r ~ ~ IT

if "~~ ~--~ ~.-.- ------ -<---o 20 40 60 so 100

Percent chord (a) MI = .276; ~ •. 526; OWl ·.0Is7;8= S7.7~ (b) MI = .362; ~ = .SI3; OWl • .0212; 8 = S 7.So.

Figure 43.- High-speed pressure distributions for cascade combinations ~1 = 30°, a1 = 61.1°, 9c = 95°, a = 1.8.

~ ~ I:-i \.n

~ I-' \.n

OJ .....J

4.8

4.0

~ 3.2

()

2.4

s

1.6

.8

? lJ o o

o Upper surface o Lower surface

12

I.n--- D.-....... p- 10-....... h ~

V I~

~ [off

10

V / /

8

V ~

V V

V

Scr f--.,/' rv -v Lo.

'""" V

fO" /

b--O" V /

P--V" k{ .~

6

5

4

./ V ~ ro--- f-CT

./ if

1 V !.A-J

I~

2

i V n ~ J--

~ ~

, , o 100 60 80 40 20 o 20 40 60 80 100

Percent chord Percent chord

(0) MI = .235; ~ = .468; Cw, = .0198; 8 = 90.6~ (b) MI = .341; ~ =.967; CWI

= .0295; 8= 90.4.

Figure 44.- High-speed pressure distributions for cascade combinations

~1 = 30°, ~1 = 64.1°, 6c = 95°, a = 1.8.

co co

~ :x>

~ t"i Vl \.>I Q I-' Vl

4.8 I

4.0

,~ b-3.2

2.4

s

1.6

.8

? ~ o 10

o Upper surface o Lower surfoce

V fa..- In. -.. """ r.

rcr ~ '-' jV" ~

) ,/' b"

/V ~

/ ~

if ~

ro---1Y' ..0'"

20 40 60 80 100 Percent chord

(a)Ml a .284; M2 -.512; Ow -.0186;8=94.3. 1

12

10

8

6

s

4

2

o

ru-~

~

V o

Scr 1"1.

.,../ if' v-

~ ~ .0-0" i.J /

~ fU

V ~ .. f.cJ

---In. ~ r-' ~~ '\..J'

I

20 40 60 80 100 Perc.nt chord

(b)MI =.387; M2 =.902; Ow ... 0225i8=92.9~ I

Figure 45.- High-speed pressure distributions for cascade combinations

a1 = 35°, ~ = 66.1°, 9c = 95°, a = 1.8.

e

~ §Z

~ t"l

~ \JI

0:> \0

4.8

4.0

3.2

.n () "u

2.4

5

I.S

n.

/ .8

II o (

o Upper surface o Lower surface

4.8

4.0

3.2

'a. ~ 2.4

~ -u......... ~ -n..

~ ~ 0.. .n.

5

I.S

.., I.., ~ I~ I~

~ ..0-.0--r-'

h_ ~

.8

20 40 SO 80 10( o Percent chord

(o)M 1 = .223;M2 =.238; CW1

=.0"lie=83.S~

?

Il V

I 7

o

t<1

\ 1

1'0-0. I~ lr\. --.. ro::cr

p--.... !o. ~ IV"

~ I.n. V \ .u-- -

~ ..0-----

0- ~

~ 1

20 40 60 80 100 Percent chord

(b) M 1 = .632; M2 =.785; C = .0219; e = 83.8. wI

Figure 46.- High-speed pressure distributions for cascade combinations

~1 = 40°, ~1 = 56.1°, 9c = 95°, cr = 1.8.

\D o

~ ()

» ~ ~ \J1 IJ,j Q f-' \J1

4.8

4.0

3.2

2.4 i\

s

1.6

I"

.8

L I

o o

o Upper surface D Lower surface

~ ~ iQ.....

ro---. ~ ~

....... '().

~ ./

~ M. :r

,~

N-!. ~ :--r

20 40 60 80 100 P.rc.nt chord

(0) M, • .336j M2 • .317; Ow • .0153; 8 .. 85.3~ I

4.8

4.0

3.2

2.4

s

1.6

.8

o

I I

V\ '\ t--Ser

I~ ~

, . lo... ~ '0...._

to-

r t1:::l ~ h./

;:l "- 'n .n--10--r-'

~

~~ .- - ~--- --~~ ~ -~. --~ I I

o 20 40 60 80 100 P.rtlnt chord

(b) M, • .&1 •• M2 ".539; Ow, •. 0156; 8 = 95.'·.

Figure 47.- High-speed pressure distributions for ca.clde combinations ~l = 45°, ~ -= 58.1°, 9c • ~O, a • 1.8. '

~ n » i t"'I V1

~ V1

'i9

4.8

4.0

3.2

o

2.4

s

1.6

.8

o

o Upper surface o Lower surface

4.8

4.0

3.2

-r\ P\ \ ~ r--Scr

2.4

~ ~ r-o.. ....,----P---P---- _r-.

I~ '0. f'o-t~ b -to--

~ -0 ,-... "

s

1.6

IV--J f ~ 1.0--.0-

!.Lr

...,........-ro-

In Ln. V '- ~ ~

f-Q-

Hr--j-o-.8

1/ IL.J

I ---L- _____ -- - --

60 80 100 40 o 20 1/ ~

I I - n.. ---o o 20 40 60 80 100 Percent chord Percent chord

(0) M 1 = .328; M2 = .329; CW1

=.0126;8= 87.6. (b) MI =.556; M2 = .565; C =.0155;8= 87.5~ wI

Figure 48.- High-speed pressure distributions for cascade combinations

f31 = 45°, CL1 = 61.1°, 9c = 95°, C1 = 1.8.

'rB

~ o :x>

~ t-' Vl 'VI

~ Vl

•

4.8

4.0

3.2

< 2.4

s

1.6

I

.8

o

o Upper surface o Lower surface

4.8

4.0

~ 3.2

\

!\ ~

1'0. -v--~ tn..

'-'---~o h /'

2.4

s

1.6

~cr ~ fcJ""

........-

! -0......, ..c:r-- ..0-

LJ'" ..n--

.8

r ----.~~ '---

40 60 100 20 80 o o Percent chord

r\ \ I Scr

c

I~ =--v-.

r ~ ~ ~ o 20

f.o. -0--. ro-, to,

j.o--r'-'

.0--V

40 60 Percent chord

ro- l..r\. 0J

~ L

~ 1..0-V

~~ J. I

80 100

(a) M.=.412; M2 =.459; Cwl

=.0146j8=90.80 • (b)M, =.549jM2 =.629;Cw1

= .0193j8=91.0°.

Figure 49.- High-speed pressure distributions for cascade combinations

a1 = 45°, ~ = 64.1°, 9c = 95°, a = 1.8.

~ f2

~ V1

~

0' G)

10

8

"0 .. 6 c o

:0:: o '> 4)

o

CD <l

0' G) "0

C 0

:0:: 0 .:;; G)

c Q:)

<l

4

2

o

6

4

2

o 50 60

NACA RM L53G15

I /3/ 1 deg

0 0 [] 15

<> 30 A 45

~

~r- - -

~ r--- ---:--......... ~ --- "\

<V v--£ r--.-

(~ .....-

---~) ---(C) (j = 1.5.

I

.,--f---~ -~ .---< .....-

<V ----~ ~ ,..,.... )

[t:=: / v

(b) (j = 1.8. ~ I I I

70 80 90 100 110

Be ' deg

Figure 50.- Deviation of the leaving-flow direction from the mean-line direction at 95-percent-chord station for blade camber.

•

NACA RM L53G15

Q)

<l

c::: o ;: o .;: Q)

o Q)

<l

0

I 8

6

"t

2

0

6

4

2

o 10

~/=4p· A

_-+-_ ,-~

{j,=30· r--... --.... .... ~~ /1,= I~ ~

~ --- j;>--10::::: :::0:::::

~ r--------0 ~ f3, =~ ----- ~-

(0) 0-=1.5. I

f3,:: 45° \

v

.-n I---!--

f3, =3C1! V -.. t--

(~ f3,=15~ ~ ~ r--- r--o::: ~ !,=o:> r--

(b) 0- :: 1.8. I I I

20 30 40 50

P2,deg

95

I Be, deg

065 080 <>95 4.110

I

f=:::::: ~ r- v

~ p

~ t-?

V r----

~

/

~ 60 70

Figure 51.- Deviation o~ the leaving-~low direction ~rom the mean-line direction at 95-percent-chord station ~or leaving-~low angles.

I: t-'iibi ..

1.0 '8,

II

M1cr

I Bc = 950 -:

) !I Secondary blade section, a, = 56.6° : J e ~ 880 ---"1

'I i ~ !

~ I

V V I //J ~ I

V V /V ~ ~ ;:-::;:"/ 8. = 95° L;/ L C

V"~ ~~Jr ~ Primary blade section, a, = 61.10 L2r. / " l I I 1 B ~ 860

v/ ~ I I I I I. I ~i'" ldV Theoretical maxImum Mach number

LV:/" __ 0- --High-speed tests

V \ <1/ . Bc = 80· ° _ -<>--Low- speed tests (corrected) ~ Primary blade sectIon, a, = 53.4 -----6- Flow plot (corrected)

11 I I B~73° ~ I Il J J I J ~----------I.....-

.8

.6

.4

.2

~ 00 10 20 30 40 50 60 70 80 ~

f3" deg

Figure 52.- Critical Mach numbers for blade sections. cr = l.8.

§! t-l \J1

~ \J1

3D

•

NACA RM L53G15 C~ 97

/

.l"igure 53. - Measurement of flow-passage area for compressrbility correc­tions to low-speed tests.

-......­C~

()

• ,",; t:-<

s

i¥~, -

4.01

8c = 80° -

o Low-speed test

~--I""r-+-~~S 0 High-speed test, M,= .500 -cr <> Low-speed test

3.2

extrapo.lated, M, = .500-

2.4r---~-+~~~~~~--~---4~~~~--~--~--~

-:-~--~-~ ~

1.6rrrl-t--t-+-t-P~~LJ

.81t=t=U~~---U-g'

II

Q WA" o 20 40

Percent chord

60 80

~ 1

Figure 54.- Comparison of high-speed-test pressure distribution and pressure distribution extrapolated from low-speed blade test.

100

\0 CD

~

f; > ~ t:-< \Jl \.)J o .... \Jl

20

~ D

16

C' Q) 12 "'0 ,.

1:1 a:. u =1

1:1 E~

8 Q)

<l

4

o

Be. deg o 65

~ [P [J 80 ~ <:> 95 ~ ~

AIIO (0

~ ~

~~ ~ ~ L~~

~ ~,

~ '~

~7 I

10 20 30 4 - 50 60 /3" d eg

Figure 55.- Measured values of induced angle at low speed and mean induced angle for solidity of 1.5.

---.. o

~ ~ ~ t;-i VI VI c;:l f-' VI

\0 \0

20

16 D

0' ~ Q) ~ 12 ...

] ~ 0 Q)

I 0 :::3 ~ C

Cb 8 '-.~ <J ,',

J "

4

z > n > r. ., " '!:'. ~

0 , -';' -;' uo .., .. .. '"

T

Be, deg 065 JJ 80 095 AIIO ~120

<p ~ ---rh

~ ~

(~

~ ID <I> ~L~

< ~ ~ ~

~ ~ < ~ ~ ~l~

~ ~

I - -

13" deg ~.o

Figure 56.- Measured values of induced angle at low speed and mean induced angle for solidity of 1.8.

I-'

8

~ ~ t-i Vl \>I q I-' Vl