Lewis, George W , Jr

English

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SECURITY INFORMATION
1
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RM E52C27
RESEARCH MEMORANDUM
EXPERIMENTAL INVESTIGATION O.F AXIAL-FLOW COMPRESSOR
INLET STAGE OPERATING AT TRANSONIC RELATIVE
INLET MACH NUMBERS
II - BLADE-COORDINATE DATA
By George W. Lewis, Jr.
Lewis Flight Propulsion Laboratory
Cleveland, Ohio
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON
S)* q~~ June 6, 1952
https://ntrs.nasa.gov/search.jsp?R=19930087219 2020-06-17T09:30:43+00:00Z
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ADVISORY COMIXTEE FOR AERONAUTICS
RESEARCHJ41MOMIIDUM
INVESTIGATIIM (X AXE3L-FLOW CCIWRWSOR
IItE3’T STAGE OPERATINGAT TRANSONIC REIATIVE
IZ!ZIXTMACHlWK81ZRS
II - BIADE-COORD- DATA
By George W.. Lewis, Jr.
.EWMARY .—.
Supplementary blade-design data including rotor- and stator-blade
coordinates and mean-line sngles at the leadlng and the trailing edges ----
are present ed for a tram onic axial-f low compressor inlet stage.e
INTRODUCTION’?
The over-all performance of a transonic axial-flow ccxnpressor inlet
stage operating at rotor relatlve inlet Mach nuuii)ers up to 1.1 is
reported in reference 1. In order to supplement the general description ..-
of the blade design contained in reference 1, specific- blade-coordinate ““ -
data for rotor and stator blades are presented.
BLADE-COOBIHNATE DATA
.
Rotor. - Although the rotor-bhd.e profiles described in reference 1
were designed along ccmical sur3?aces in order to closely represent the
t~e stresm13ne flow, it was desirable for fabrication purposes to
express the blade profiles in rectangular coordinates in planes perpen-
dicular to the stacldng radius.
.-
Rectangular blade coordinates at four
.-
reference locations were’ obtained frcm the design conical coorMnates by
means of graphical fairing. A sketch of the rotor-blade layout and hub-
contour coordinates is shown in figure 1. Rotor-blade design coordi-
nates at four refereqce locations are given in figure 2. Leading- and
trailing-edge radii vtiiedfrozn 0.015 to 0.020 inch. L -.
bspection of the fabricated blade-profile shapes with an opticsl
m comparator reveshd good agreement with the submitted blade-coordinate
data. Measured ratios of maximumblade thickness to chord length were
.-
0.048 at the tip and 0.075 at the hub. Blade leading- and trailing-edge.,
—
2 NACA RME52C27
mean-line angles as obtained from comparator layouts are presented in
figure 3. The rotor-blade mean-line angles..were determined along
approxhate streamlines from blade Wet to outlet based on the” assu&
tion of equal radial-streamline spacing. The measured rotor-blade-inlet
angles of figure-3 are approximately 2° hi~er than the values given in
figure 8(b) of reference 1, which were desired design values. The smsll
discrepancy is attributed to the inaccuracies of the transformation proc-
ess from conicsl to rectangular coordinates_. The true rotor-incidence
angles tie, therefore, correspondingly smaller thfi the vslues given ir”
reference 1.
The coordinates of a profile at the hub in a plane parallel to the
blade base (section G-G) as measured with an optical comparator are given
in figure 4 to Indicate .the form of the llade profile abng the hub .
streamline. ‘
Radial clesrance between the rotor-blade tip and the outer casing
varied from 0.023 to 0.026 inch and the axial clearance between the
rotor and the stator at the hub was 1.75 inches.
Stator. - The stator blades were constructedby linear fairing
between two circfiar-arc sections at the hub and the tip. A sketch of
the subsmic stata blade is shown in figure 5 and rectangular coordi-
nates for the hub-and the tip sections are given in figure 6. The
inner casing had a constant radius acrms the stator (5.186 in.) and a
radial clear~ce of approximately 0.020 inch. Measured stator-b~de
mean-line sngles at the leading and the tra~ing edges are gitien in fig-
ure 7. Measured stator maximumthickness ratios were 0.065 at the hub
and 0.084 at the tip. Leading- and trailing-edge radii varied from 0.015-
to 0.020 inch.
Lewis Flight Propulsion Laboratory
National Advisory Committee for Aeronautics
Cleveland, Ohio
1. Lieblein, Seymour, Lewis, George W., Jr., and Sandercock, Donald M.:
Experimental bvestigation of an Axial-Flow Canpressor Inlet Stage
Operating at Transonic Relative Inlet Mach l?tiers. Part 1 - Over-
All. Performance of Stage with T%ansonic Rotor smd Subsonic Stators
up to Rotor Relative Mach Nmiber of 1.1. NJICARM E52A24, 1952.
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G = -.120
a = 47° .09
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2.015
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.080
.126
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.196
.,217
.228
.229
.219
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.169
.128
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.015
x
o
.035
.250
.500
.750
l.oao
1.250
1.500
1.750
2.000
2.250
2.500
2.760
3.0.00
3.201
3.216
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.079
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— ,, .. ... . , ><
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*-F =.1.621 “..
G= .172
a.~”ot
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F = 1.562
G= .081
o,=23° 151 -
F’= 1.596
G= .272
U,= z30 15v-
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.183
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. 4al
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1.000
1.200
1.400
1.600
1.800
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2.867
2.882
x YL,
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0
.025
.049
.071
.089
.104
.115
.123
.127
.126
.121
.111
.087
.078
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.023
0
.015
—
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.015
0
.015
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.141
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.421
. 3s7
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.160
.223
.268
.296
.309
.308
.293
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.159
.072
-.050
-.200
-.371
-.432
-.455
2.061
.046
.078
.103
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.098
.067
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-.096
-.173
-.256
-.344
-.435
-.%66
-.455
0.015
0
.058
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.179
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.400
.600
. Soo
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1.2C0
1.400
1.600
1.8cI0
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2.41KI
2.600
2.800.
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3.127
3.142
.400
.600
.800
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1.200
1,400
1.6cQ
1.800
2.000
2.203
2.400
2.600
2.8oo
2.883
3.007
,., . . . . .
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Figure 2. - Coordinatesfmrotcm-blade bectlonsoffigure 1. (AlldimensionsIn inches
exceptwhere noted.) — . . .-—
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NACA RM E52C!27
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““l r
v Rotor-blade section on ,
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conic stream surface. ...—.
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Trailing edge
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Radius, in.
Figure 3. - Rotur-blademean-line angles.
,. . . .... ---
..
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NACA RM E52C!27
Section G-G
x
o
.025
.200
.400
.600
.800
1.000
1.200
1.400
1.600
1.800”
2.000
2.200
2.400
2.600
2.800
3.000
3.025
3.050
3.085
o
.020
.102
.170
U.221
.260
.290
.313
.325
.328
.324
.312
.288
.257
.213
.148
.048
.034
.017
-.017
o
-.026
-.002
.031
.056
.072
.083
.090
:096
.098
.Q95
.087
.072
.057
.035
.009
-.017
-.019
-.024
-.017
Axis
.—
Figure 4. - Coordinates for measured profile at hub psrallel
.,
.-
.—
.-
.-
to blade base obtained with optical c~arator. R~dius
at leading edge is 4.824 inches and radius at trailing edge “-- –
edge is 5.394 inches. (All dimensions in inches.)
—
.—
.
.,
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.
.
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=
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.-
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NACA RM E52C27
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B
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3.355
——
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5.20 R
Figure5. - 8ubmnic statcrbl.de. .(All dimensions
in inches.)
7
.. ..—
8 ~ NACA RM E52C27
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Flow
/
/’
/
Section B-B
A’ / LStacking point
l?.= 1.605
G= .150
a= 28.2°
x
o
.100
.350
.6Q0
.850
1.100,
1.350
1.600
1.850
2.lcO
2.350
2.600
2.850
3.100
3 J210
o.o15
.058
.129
.185,
.227
.259
.278
.286
..278
.261
.230
.187
.132
.062
.015
‘L
0.015
.002
.005
.009
.O11
.013
.014
.016
.014
.014 ,
.033.
.009
.006
.002
.015
.-
.
. . . .
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Axis
“v” ““
Section C-C!
F= 1.315
G= .148
a= 33.70 -
x
o
.100
.350
.600
i850
1.100
1.350
1.600
1.850
2.100
2.350
2.600
2.630
,
. .
‘u ‘L
3.015 0.01.5
.057 .007
.123 .028
.173 .043
.208 .054
.230 .062
.233 .064
.224 .060
.200 .053
.160 .040
.107 .024
.038 .CQ2
.015 .015
Figure 6. - Coordinates for stator-blade secti.ans of figure 5.
(All dtiens ions in inches except where noted.).
—
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—. =.-
.—- .—*._
.-
.—
, $_
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3.
c
NACA RM E52C27
Flow
Axis
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9
Stator-blade section
.. ----- ----
—.
6 7 8 9
Figure 7.
Radius, in.
-. Stator-bl.ade mean-line angles.
NACA-LU@ey-6-6-52-37S


Experimental investigation of axial-flow compressor inlet stage operating at transonic relative inlet Mach numbers : blade-coordinate data

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