A major source of the noise generated by high bypass ratio aircraft fan engines is the unsteady pressures (forces and moments) acting on the fan blades. These unsteady pressures are the result of the interaction of the blades with the wakes of upstream blade rows, inlet flow distortions, and inlet turbulence. An experimental system was developed which will permit the measurement of the unsteady pressure distributions; a series of experiments were conducted to investigate these effects, and the results are compared with predictions from existing theoretical analyses

Henderson, R. E.

English

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Produced by the NASA Center for Aerospace Information (CASI) 
https://ntrs.nasa.gov/search.jsp?R=19750024309 2020-03-22T20:38:01+00:00Z
SEMI-ANNUAL. PROGRESS REPORT
(March 1975 to September 1975)
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NASA Grant NGR 39-009-275
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E '	v o INVESTIGATION OF THE UNSTEADY PRESSU:'E
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b N Robect E. Henderson
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The Pennsylvania State
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University Park, PA	 16802
October, 1975
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INTRODUCTION
A major source of the noise generated by high bypass ratio aircraft
fan engines is the unsteady pressures (and, hence, forces and moments)
acting on the fan blades. These unsteady pressures are the result of
the interaction of the blades with the wakes of upstream blade rows,
inlet flow distortions, and inlet turbulence. To attempt reduction of
the level of generated noise, -it is necessary to know how the unsteady
pressures, forces, and moments on the fan blades vary as a function of
blade geometry and operating conditions - rotational speed, solidity,
stagger angle, steady load, stage spacing, etc. The purpose of the
research conducted under NASA Grant NGR 39-009-215 is to develop an
experimental system which will permit the measurement of the unsteady
pressure distributions .,n the blades of an axial flow turbomachine, to
conduct a series of experiments to investigate these effects, and to
compare the results with predictions from existing theoretical analyses.
This report is the fourth report of prog ess under NASA Grant NCR
39-009-275. Tne effort during this period has concentrated on the
fabrication of an instrumented sta!Lor blade, its assembly with the pitran
sensors, conducting an initial series of tests, and the reduction and
analysis of test data.
TNSTRUMENTATION TO MEASURE UNSTEADY PRESSURES
In the development of an instrumented blade for the measurement of
chordwise unsteady pressure distributions in an axial flow fan stage,
there are a multitude of factors which must be considered. These include:
the selection of a sensor; the mounting of the sensor in the blade; the
distribution of sensors along the blade chord; the dynamic calibration
of the mounted sensor; and analysis of the time-varying signal. By
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initially concentrating; on the development of a system for a stator
blade, the additional problems which arise when considering a rotating;
blade, i.e., centrifugal effectca and signal transfer from a rotating
to stationary reference frame, are avoided.
After a review of commer.-j-.1y available dynamic pressure trans-
ducers, the STOW LABORATORIES, INC. PITHAN differential pressure
transducer was selected as the sensor to be employed in this program.
The advantages of this sensor are its size, .559 cm diameter; its
sensitivity, 2.0 volts maximum output; and an overload tolerance of
nearly 700 percent. The major disadvantage of the PITRAN is the drift
which occurs in the output signal because of ambient temr_rature
fluctuations. This problem can be corrected, however, by proper con-
ditioning of the signal with an electronic feedback system provided by
STOW LABORATORIES.
The blades of the AFRF test stage h.:ve a chord length of 15.24 cm,
a maximum thickness of 1.524 cm at the 33% chord position, and a span
length of 15.24 cm. To improve the spatial resolution of the measured
pressure, each sensor is mounted interna'.ly facing a cavity with con-
necting taps to both surfaces of the blade, Figure 1. In this way, the
unsteady pressure or each surface can be determined by alternately
scaling one surface, or the differential pressure can be recorded by
testing with both taps open. The cavity/surface tap method is more
advantageous than the use of flush-mounted sensors because of the diffi-
culty in fabricating a truly flush sensor. This is particularly true
in the case of a curved surface. The design of the cavity/sensor mount-
ing arrangement, has been discussed in previous Semi-Annual Reports [1].
The final configuration of one of the instrumented stator blade is
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shown fit 	 2. This blade has t;Ix chordwise pressure taps localcd
at the 5, 15, 30, 40, 50, and 75 percent chord positionsi wllh opening'
to both surfaces of the airfoil. The surface taps are distributed
radially over a 2.54 cm segment centered at the mean ra :us of the blade
in order to minimize flow interference from the upstream tiurface taps.
Two additional instrumented blade-, have been constructed wi`h pressure
taps at the 2, 40, and 80.8 p% , rcent chord positionsf. 'These brides art-
restricted to measurement of the unsteady pressure on altern: ,.te surfaces	 «
.1	 i.e., one blade contains taps to the pressure surface and the other
blade contains taps to the suction surface, because of the complexities
in fatricating a tubs- passage from the surface to the cavity near the
leading and trailing edges. The intent in utilizing these three blades
is to provide pressure measurements at eight chordwise positions from
which the minimum required number and location of surface taps necessary
to provide an accurate neasurement of : the unsteady pressure distribution
can be determined. The instrumented rotor blade will then be designed
employing this optimized number and location of pressure taps.
EXPERIMENTAL SET-UP AND DATA REDUCTION
The inns .tmented stator blade shown in Figure 2 was installed in
the AFRF at a location approximately two chord lengths downstream of
the Stage rotor. The rotor configuration consisted of 12 rotor blades
having the same symmetrical cross section as the stator, and a stagger
angle of 45 degrees at the mean radius. This rotor is designed to
operate at zero incidence along its entire span in a uniform flow, thus
producing no energy transfer to the fluid. However, viscous effects
within the blades introduce some f3uld turning, approximately 3.0 degrees,
at the design condition.
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The interaction of the wakes frvin the upstrearn rotor with the
instrumented stator result in unnLeady pressure fluctuations at a funda-
mental frequency equal to the product of the number of rotor blades and
the rotor -atational speed. The exact description of the wakes at the
stator inlet was obtained by analyzing the output of a hot-film probe
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and a wedge-type yaw probe located at the duct mean radius. The mean
flow angularity was first determined from the yaw probe output. The
hot-film probe was then aligned perpendicular to the absolute mean flow
direction, and the instantaneous wake velocity distribution was then
recorded on magnetic tape in the form of the hot-film. output. Simul-
,	 taneously, Lhe output of the PITKAN pressure transducers and the output
of a photocell which emitted a single pulse for each revolution of the
test rotor were recorded on magnetic tape.
Variations in the stator disturbance flow field were obtained by
operating the stage rotor at a fixed rotational speed with three differ-
ent throughflow velocities. At the rotor mean radius, these conditions
correspond to rotor inlet incidence angles of 0, 5, and 10 degrees.
In addition to the dynamic calibration (1) of the cavity/sensor
combination, each sensor was statical..y calibrated to determine its
sensitivity (psi/volt). This was accomplished using a micromanometer
which allows a precise pressure lovel to be set. One side of the
differential sensor was then instantaneously subjected to this pressure
while the other side was held at atmospheric pressure. The time history
of the sensor output was then recorded on an oscillograph, which allowed
the maximum voltage output to be determined.
Data analysis at each test condition was based on a total of 500
samples of the taped sensor or hot-film output which were digitized and
50
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phase-lock averaged, uning the photocell output as a trigger . l:ac•h
data sample contained 512 points with a time increment betwi-en point,,i
of 0.0002 sec. Tie timr duration of the digitizing process, 0.1024
sec, exceeded the rotor revolution period in all cases. The portion of
this averaged signal which corresponded to the time for one rotor revo-
lution, starting with the photocell pulse as time zero, was then Fourier
analyzed to define the sensor/film output -- both m ig,n i tude and phar {e
angle -- associated with the frequencies which correspond to integer
multiples of the rotor blade passing; frequency. The usual purpose of
phase-lock averaging; is to extract a periodic signal that is obscured
by the presence of noise. The signals analyzed here were not of this
type; however, the process was used to assure elimination of the effects
of response of random sources of excitation, such as mechanical vibra-
tion and turbulence.
EXPERIMENTAL RESULTS
An initial Bevies of measurements have been completed with the
instrumented stator blade having six pressure transducers/taps located
at the 5, 15, 30, 40, 50 and 75 percent chord positions. Measurement;
were conducted of the time variation of pressure on each side of the
blade with the opposite surface rap blocked, and of the differential
pressure with both surface caps open. At the present time, these data
are only partially ar:alyz!d and r. easur,mments at the 2 and 80.h percent
chord locations have not been conducted. As a result, it is not
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possible, at this time, to specify an optimum chordwise distribution of
surface taps, or to present a comparison of the measured data with
theoretical predictions. The results have den,onstrated, however, that
the PITRAN pressure transducers do function properly in the cavity
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mounting arrangement employed here, and that a satisfactory response to
the rotor wakes can be obtained.
Consider a test condition in which the rotor blades operate at zero
►ngle of incidence. If the signal from the hot-film probe is averaged
far 1-, 100-, and 500 sums, the wake velocity distribution as a function
of time appears as shown in Figure 3. These data demonstrate the
significance of phase-lock averaging in the separation of a random and
periodic signal. They also show the existence of significant random
fluctuations of the wake velocity. These random velocity fluctuations
result in the broadband noise present in a fan's radiated noise spectrum,
while the periodic velocity fluctuations give rise to the pure tone
noise.
The magnitude of the periodic velocity fluctuations is quite small,
t2.8 percent of the mean volocity. The currespondi.ng variation in
sensor output voltage on one side of the blade as a function of time
after 500 gums is shown in Figure 4 for the 5, 15, 30, and 50 percent
chord surface tap locations. The output of each sensor must be multi-
plied by the sensor sensitivity, which is different for each location,
in order to _rive at the static pressure fluctuation as a function of
time. While the velocity fluctuations and, hence, the pressure flue-
tuations are very small, the sensor system is capable of reproducing
the signal. Experiments at smaller rotor-stator spacing would result
in a larger velocity fluctuation and, Fence, larger pressure fluctuations.
FUTURE EFFORTS
The experimental data obtained to date with the instrumented stator
d e monstrate that the sensors and their cavity/surface tap mounting
arrangement are functioning properly. Further experiments will be con-
-8-
Ducted in the near 1 uture, in which i he unsteady pressurer ► at the 2
and 80.8 percent chord locations will be determined. By combining these
results with the measurements at the intermediate locations, the mini-
mum number of sensors and their locations, which are required to ade-
quately define the unsteady pressure distribution, will be determined.
This information will then be employed in designing an instrumented
rotor blade.
f
Additional experiments with the instrumented stator blade are
scheduled to be conducted in November. As a result of the initial tests
several modifications will be made to the instrumented blade and its
associated electronics to improve the quality of the data and the
required test time. These modifications included: (1) the use of
0-ring seals around the "ITRAN sensor to replace the washer/vacuum
grease seal shown in Figure 1, (2) the determination of the PITRAN
sensitivity aitor it has been mounted in the blade and (3) the procure-
ment of an aoditional signal conditioning package thus eliminating the
need for switching the electronics from sensor to sensor during a test.
References:
(1] Semi-Annual Progress Repo t (March 1974 to September 1974) on
NASA Grant NGR 39-009-0275, October, 1974.
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Investigation of the unsteady pressure distribution on the blade of an axial flow fan

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