The stator-blade angles in the first four stages of a 16-stage axial-flow compressor were increased in order to decrease the angles of attack of these stages, and thereby to improve part-speed performance. The performance of this modified compressor was compared with that of the same compressor with original blade angles

Hatch, James E.

Medeiros, Arthur A.

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NACA RM E52B15	
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
RESEARCH MEMORANDUM
EFFECT OF VARIOUS BLADE MODIFICATIONS ON PERFORMANCE OF A
16-STAGE AXIAL-FLOW COMPRESSOR
III - EFFECT ON OVER-ALL PERFORMANCE CHARACTERISTICS
OF INCREASING STATOR-BLADE ANGLES IN INLET STAGES
By Arthur A. Medeiros and James E. Hatch
SLDOE4,Ry
The inlet stages of a 16-stage compressor were unloaded by
increasing the stator-blade angles by the following amounts; guide
vanes 90 , first, second, and third stage stators 100 , and fourth stage
stators 50 . The performance of this modified compressor was compared
with that of the same compressor with original blade angles. The peak
efficiency of the modified compressor was higher at all speeds up to
90 percent of equivalent design speed. The maximum efficiency of the
modified compressor occurred at 85 percent of equivalent design speed
and was 82 percent. This was 12 points higher than the maximum effi-
ciency of the compressor with original blade angles, which occurred
at an equivalent speed of 90 percent of design, and 3 points higher
than the peak efficiency of the compressor with original blade angles
at the same speed. The increases in peak efficiency attained by this
modification were from 10.5 to 7 points at equivalent speeds from 30
to 80 percent of design. The peak efficiency was decreased about
1/2 point at 90 percent of equivalent design speed and 4 2 points at
100 percent of equivalent design speed®
The surge pressure ratio was increased at all speeds up to 90 per-
cent of equivalent design speed and decreased at 90 and 100 percent of
equivalent design speed. The weight.flow was increased at equivalent
speeds up to and including 75 percent of design and was decreased at
higher speeds. The maximum reduction in weight flo g
 was 13.5 percent
at design speed. The changes in surge pressure ratio and flow were
such that the surge line was about the same for both configurations at
speeds up to 75 percent of equivalent design speed, whereas at higher
speeds, the surge line for the modified compressor occurred at a higher
pressure ratio for a given weight flog.
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2	 CONFIDENTIAL	 NACA RM E52B15
INTRODUCTION
All the blade rows of a multistage axial-flow compressor are usu-
ally set to operate at the angle of attack for maximum efficiency at
design speed so that optimum performance is attained at design condi-
tions. At speeds below design, the decreased energy addition leads to
high volume flows and low angles of attack in the exit stages. The
high volume flows in the exit stages limit the weight flow which can be
passed through the compressor and force the inlet stages to operate at
angles of attack higher than those for which they were designed. The
deviations with speed from the angles of attack for maximum efficiency
in both the inlet and exit stages are therefore large and result in law
over-all efficiencies at low speeds. This mismatching of the stages
naturally becomes more severe as the number of stages and over-all
pressure ratio are increased. A peak efficiency of 50 percent at
30 percent of design speed is reported in reference 1 for a 16-stage
compressor, although the design-speed peak efficiency is 80 percent.
If the blade rows of a compressor were set to operate at angles of
attack for maximum efficiency at some speed below design, the deviation
with speed from these optimum angles would be decreased and better
part-speed performance could be anticipated. Some loss in design-speed
performance may be incurred. Some such compromise may be necessary,
however, particularly in high over-all pressure ratio compressors, to
obtain a compressor with part-speed characteristics that will allow a
good acceleration margin for the engine in which the compressor is to
be used.
The method for achieving such a compressor design would be to set
the inlet stages to operate at angles of attack laver than those for
maximum efficiency and increase the blade camber of these stages so
that design pressure ratio is attained at design speed and flow. Con-
versely, the blades in the exit stages would have a decreased camber
and increased angle of attack. Because the middle stages tend to oper-
ate at a constant angle of attack at all speeds, these stages could be
set to operate at the angle of attack for maximum efficiency. These
stages could also be highly loaded because of the small range of angle
of attack at which they operate, and the number of stages required for
a given over-all pressure ratio would thereby be decreased.
Changing rotor-blade angles and blade cambers was not a feasible
method of altering the stage operating point in this investigation
because of fabrication and assembly difficulties. Therefore, the stage
operational point was varied by resetting the stator blades upstream
of the stage in which the changes were required. The effect on over-
all performance of both increasing and decreasing stator-blade angles
in the exit stages 30 from the original settings is presented in
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NACA RM E52B15	 CONFIDENTIAL
	
3
references 2 and 3. Some improvements in part-speed efficiency and
surge characteristics were reported when the angles of attack in the
exit stages were increased by decreasing the stator-blade angles 30.
The performance at design speed was unchanged by this modification.
Comparison of the three configurations indicated that optimum perform-
ance obtainable in this compressor with stator-blade resetting in the
exit stages was approached with the stator-blade angles decreased 30.
In this investigation also conducted at the NACA Lewis laboratory
and reported herein, the loading in the inlet stages was decreased by
increasing the setting angles of the inlet guide vanes and stator
blades in the first four stages. The blade angles were arbitrarily
changed by a relatively large amount to obtain a large effect on over-
all performance and thereby to determine if an intermediate change
might be advantageous. The inlet-blade angles were increased by the
following amounts: guide vanes 9 0, first, second, and third stage
stators 100, and fourth stage stators 5 0 . In order to determine the
effect on performance of changes in the inlet stages independently of
other blade angle changes, the blades in the fifth to sixteenth stages
were set at the original angles.
APPARATUS AND PROCEDURE
The test installation used for these runs is similar to that
described in reference 1 except that a 15,000-horsepower-drive motor
was used. The average inlet Reynolds number, relative to the first
rotor, was approximately 350,000 for the runs reported herein and in
references 2 and 3. Over-all performance data were obtained at equiva-
lent speeds of 30 to 100 percent of design over a flow range at each
speed from maximum to surge.
DISCUSSION OF RESULTS
The over-all performance of a 16-stage axial-flow compressor is
presented in figure 1 with inlet stage blade angles increased the fol-
lowing amounts: guide vanes 90, first, second, and third stage sta-
tors 100 , and fourth stage stators 5 0 . The performance of the same
compressor with original blade angles (reference 2) is presented for
comparison.
The speed at which the maximum efficiency occurs was decreased by
the modification from 90 to 85 percent of equivalent design speed. The
maximum efficiency of the modified compressor is 82 percent. This
efficiency is 12 points higher than the maximum efficiency of the com-
pressor with original blade angles, and 3 points higher than the peak
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4	 CONFIDENTIAL	 NACA RM E52B15
efficiency of the compressor with original blade angles at the same
speed. At speeds below 85 percent of design speed, the increases in
peak efficiency with the modified compressor are large, namely, 10.5,
7.5 1 7.5 1
 10.5, and 7 points at speeds of 30, 50, 65, 75, and 80 per-
cent of equivalent design speed, respectively. At equivalent speeds
higher than 85 percent of design, the peak efficiencies were decreased.
The peak efficiency was decreased about 1/2 point at 90 percent of
equivalent design speed and 4j points at 100 percent of equivalent
design speed.
At speeds up to and including 75 percent of equivalent design
speed, the weight flow was increased by the modification. This
increase indicates that in the compressor with original blade angles,
the inlet stages, at these speeds, were operating at angles of attack
greater than those required for maximum pressure ratio; decreasing the
angle of attack by resetting the stator blades therefore increases the
pressure ratio in these stages and allows more weight flow to be passed
through the choked, or negatively stalled, exit stages. At higher
speeds, the decreased angles of attack in the inlet stages of the modi-
fied compressor resulted in a lower pressure ratio for there stages.
The decreased pressure ratio in the inlet stages will not permit the
same weight flaw to be passed through the exit stages and, therefore,
the weight flow at equivalent speeds of 80 to 100 percent of design
was lower for the modified compressor. The maximum reduction in weight
flaw was 13.5 percent at 100 percent of equivalent design speed.
At equivalent speeds up to and including 75 percent of design, the
surge pressure ratio and weight flaw were both increased such that the
surge line of these speeds is about the same for both compressor con-
figurations. At 80 and 85 percent of equivalent design speed, the
surge pressure ratio was increased while the weight flaw was decreased;
a large shift of the surge line, at these speeds, to a higher pressure
ratio for a given weight flow resulted. The surge pressure ratio and
weight flaw were both decreased at equivalent speeds of 90 and 100 per-
cent of design. Because the weight flaw decrease predominated, a shift
of the surge line to a higher pressure ratio for a given weight flaw
still occurred at these speeds.
If, in addition to resetting the inlet stages to operate at angles
of attack for maximum efficiency at some speed below design, the blade
cambers were increased so that design pressure ratio at design speed
was still attained, the reduction in weight flow at design speed would
not be incurred, and the penalty in efficiency might not be so severe.
In addition, the improvement in performance at the intermediate speeds
might be greater because of the increased available pressure ratio in
the inlet stages with the greater camber.
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NACA RM E52B15	 CONFIDENTIAL	 5
SUMMARY OF RESULTS
The stator-blade angles in the inlet stages of a 16-stage axial-
flow compressor were increased by the following amounts: guide vanes
9o , first, second, and third stage stators 10 0, and fourth stage stators
50 , and produced the following results when compared with the same com-
pressor with original blade angles:
1. The maximum efficiency of the modified compressor occurred at
85 percent of equivalent design speed and was 82 percent. This effi-
ciency was 12
 points higher than the maximum efficiency of the compressor
with original blade angles, which occurred at 90 percent of equivalent
design speed, and was 3 points higher than the peak efficiency of the
compressor with original blade angles at the same speed.
2. At speeds below 85 percent of equivalent design speed, the peak
efficiencies were increased 10.5, 7.5 1 7.5, 10.5, and 7 points at
speeds of 30, 50, 65, 75, and 80 percent of equivalent design speed,
respectively..
3. At equivalent speeds of 90 and 100 percent of design, the peak
efficiencies were decreased by the modification. The peak efficiency
was 1/2 point lower at 90 percent of equivalent design speed and 4 1
points lower at 100 percent of equivalent design speed.
4. The weight flow was increased at equivalent speeds up to and
including 75 percent of design and was decreased at higher speeds. The
maximum decrease in weight flow was 13.5 percent at 100 percent of
equivalent design speed.
5. The surge pressure ratio was increased at all speeds except 90
and 100 percent of equivalent design speed at which it was decreased.
The changes in surge pressure ratio and weight flow were such that the
surge line was about the same for both configurations at speeds up to
75 percent of equivalent design speed, whereas at higher speeds the
surge line for the modified compressor occurred at a higher pressure
ratio for a given weight flow.
Lewis Flight Propulsion Laboratory
National Advisory Committee for Aeronautics
Cleveland, Ohio, February 11, 1952
CONFIDENTIAL
6	 CONFIDENTIAL	 NACA RM E52B15
1. Medeiros, Arthur A., Guentert, Donald C., and Hatch, James E.:
Performance of J35-A-23 Compressor. I - Over-All Performance.
Characteristics at Equivalent Speeds from 20 to 100 Percent of
Design. NA.CA RM E50,117, 1951.
2. Medeiros, Arthur A., Hatch, James E., and Dugan, James F., Jr.:
Effect of Various Blade Modifications on Performance of a 16-Stage
Axial-Flow Compressor. I - Effect on Over-All Performance Charac-
teristics of Decreasing Twelfth through Fifteenth Stage Stator-
Blade Angles 30 . RM E51LO3.
3. Hatch, James E., and Medeiros, Arthur A.: Effect of Various Blade
Modifications on Performance of a 16-Stage Axial-Flow Compressor.
II - Effect on Over-All Performance Characteristics of Increasing
Twelfth through Fifteenth Stage Stator-Blade Angles 3 0. RM E52A10.
CONFIDENTIAL
NACA RM E52B15	 CONFIDEMAL	 7
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Figure 1. - Effect on over-all performance of increasing inlet guide vane and stator-blade
angles in first four stages of 16-stage compressor.
CONFID NTTAL
,NACA RM E52B15	 CONFIDENTIAL
RESEARCH MEMORANDUM
EFFECT OF VARIOUS BLADE MODIFICATIONS ON PERFORMANCE OF A
16-.STAGE AXIAL-FLOW COMPRESSOR
III - EFFECT ON OVER-ALL PERFORMANCE CHARACTERISTICS
OF INCREASING STATOR-BLADE ANGLES IN INLET STAGES
a . A-114
Arthur A. Medeiros
Aeronautical Research
Scientist
James E. Hatch
Aeronautical Research
Scientist
Approved:	 N1 X..^
William A. Benser
Aeronautical Research
Scientist
Robert 0, Bullock
Aeronautical Research
Scientist
/Y.Oscar W. /Schey
Chief ' Compressor and
Turbine Research Division
pr
CONFIDENTIAL
a
NACA RM E52B15	 CONFIDENTIAL
Compressors - Axial Flow
	
3.6.1.1
Medeiros, Arthur A., and Hatch, James E.
Abstract
The stator-blade angles in the first four stages of a 16-stage
axial-flow compressor were increased in order to decrease the angles
of attack of these stages, and thereby to improve part-speed perform-
ance.
The performance of this modified compressor was compared with that
of the same compressor with original blade angles.
Restriction/Classification Cancelled
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Effect of Various Blade Modifications on Performance of a 16-Stage Axial-Flow Compressor. III - Effect on Over-All Performance Characteristics on Increasing Stator-Blade Angles in Inlet Stages

Effect of Various Blade Modifications on Performance of a 16-Stage Axial-Flow Compressor. III - Effect on Over-All Performance Characteristics on Increasing Stator-Blade Angles in Inlet Stages

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