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RD-fl36 896 ROUGHNESS EFFECTS ON COMPRESSOR BLADE PERFORMANCE IN 1/2
CASCADE AT HIGH REYN..(U) AIR FORCE INST OF TECHWJRIGHT-PATTERSON APR OH SCHOOL OF ENGI. F J TANIS
UNCL ASSIFIED NOV 83 AFIT/GAE/AA/83D-23 F/G 20/4 N
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MICROCOPY RESOLUTION TEST CHART
NATKONAL BUREAU OF SIANDARDS-1963-A
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ROUGHNESS EFFECTS ON COMPRESSORBLADE PJEFORMANCE IN CASCADE
AT HIGH REYNOLDS NUMBER
THESIS
FIT/GAE/AA/83D-23 Frederick J. Tanis Jr.2Lt. U.S. Air Force
0.. iiC5-
DEPARTMENT OF THE AIR FORCE
AIR FORCE INSTITUTE OF TECHNOLOGY
Wright-Patterson Air Force Base, Ohio
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AFIT/GAE/AA/83D-23
THESIS
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~ROUGHNESS EFFECTS ON COMPRESSOR
BLADE P FORMANCE IN CASCADEAT HIGH REYNOLDS NUMBER
THESIS
AFIT/GAE/AA/83D-23 Frederick J. Tanis Jr.
2Lt. U.S. Air Force
.14;
Approved for public release; distribution unlimited ,
.94 -!
AFIT/GAE/AA/83D- 23
ROUGHNESS EFFECTS ON COIMPRESSOR
BLADE PERFORMANCE IN CASCADE
"* AT HIGH REYNOLDS NUMBER
THESIS
Presented to the Faculty of the School of Engineering
of the Air Force Institute of Technology
4Air University
in Partial Fulfillment of the
Requirements for the Degree of
Master of Science in Aeronautical Engineering
Frederick J. Tanis Jr.,B.S.
Second Lieutenant, USAF
November 1983
4 Approved for public release; distribution unlimited
-. PREFACE
I would like to thank my thesis committee members, Capt
Wesley Cox, Dr. Milton Franke and particularly my principal
advisor Dr. William C. Elrod for all the time they spent
working with me in order to get this study compleated. I
would also like to thank Maj. John Vonada for all of his
help in getting acquainted with the Cascade Test Facility
-i and keeping it alive throughout the course of this study. I
-am also very gratful for the the support of the members of
the AFIT Fabrication Branch and the technicians in the
Department of Aeronautics and Astronautics, with special
thanks going to Mr. John Brohaus and Mr. Harley Linville. My
deepest thanks go to my wife Melinda for her understanding
and support during my long hours away from home during the
course of this study.
Frederick J. Tanis Jr.
Nils GRA&I,:O DTIC TAB
t Unannounced 0Justif icat lOr
By-Distribution/
Availability Codes
Avil and/or4 Dist Speciali---i -i-% i
ii '-'----'
CO4TENTS
LIST OF FIGURES ................................ iv
LIST OF SYMBOLS ................................ vii
LIST OF TABLES ................................. ix
ABSTRACT. .................. .. . * . * ............. * * x
I. INTRODUCTION.... ... .... ..... . . .. . 1
OBJECTIVES AND SCOPE .............. 2
II. APPERATUS AND PROCEDURE.................. 4CASCADE TEST FACILITY.............o.. 4TEST SECTION. . o. .. . o. ........ ...... 4INSTRUMENTATION ..................... . 5HOT FILM SENSOR CALIBRATION........... 7EXIT PLANE TRAVERSER ................. 10BLADE ROUGHNESS CONFIGURATIONS........ 11DATA ACQUISITION AND PROCESSING ....... 12
III. DATA REDUCTION. ... ... ................. . 14EXIT VELOCITY ...... .... ..... ......... 14TOTAL PRESSURE LOSS................... 16TURBULENCE INTENSITY................... 19
IV. RESULTS AND DISCUSSION................... 20LOCATION EFFECTS.............. ........ 21ROUGHNESS MAGNITUDE EFFECTS........... 37REPEATABILITY............ ............. 48
V. CONCLUSIONS AND RECOMMENDATIONS......... 50CONCLUSIONS. ........ ......... ..... .... 50RECOMMENDATIONS. .......... ........ 50
BIBLIOG RAPHY.o.. ........ ........... .. ... 52
APPENDIX A: ROUGHNESS DEFINITIONS.............. 54
APPENDIX B: VELOCITY AND TURBULENCE INTENSITYPROFILES....................... 57
VI TA........................................... 98
.. , iii
" _ . - -,. -. . . . , ., .- . -. ,. . . .. . . - . .. . . . . . . .. . .-.... . . . . .
WIN
LIST OF FIGURES
Figure Page
1. Test S c i n. . . . . . . . . . . . . . . . . .. 5
2. Hot Wire CalibrationCurve..................... 9
3. Standard Data Window ............... ............. 104. Blade Roughness Configurations ................... 12
5. Roughness Configurations 2 and 3 ................. 21
6. Roughness Configurations 5, 7 and 8 .............. 22
7. Velocity and Turbulence Intensity ProfileConf#1 Eval#6: Smooth Baseline blade Traverse#l.. 26
8. Velocity and Turbulence Intensity ProfileConf#1 Eval#6: Smooth Baseline blade Traverse#2.. 27
9. Velocity and Turbulence Intensity ProfileConf#1 Eval#6: Smooth Baseline blade Traverse#3.. 2b
10. Velocity and Turbulence Intensity ProfileConf#l Eval#6: Smooth Baseline blade Traverse#4.. 29
11. Velocity and Turbulence Intensity ProfileConf#1 Eval#6: Smooth Baseline blade Traverse#5.. 30
12. Velocity and Turbulence Intensity ProfileTraverse#1 Conf#5 Eval#2: 20 micro-meter Ra
Roughness from L.E.-i/4Chord ..................... 32
13. Velocity and Turbulence Intensity ProfileTraverse#l Conf#7 Eval#l: 20 micro-meter RaRoughness from i/8-3/SChord ...................... 33
14. Velocity and Turbulence Intensity ProfileTraverse#1 Conf#8 Eval#l: 20 micro-meter RaRoughness from ./4-1/2Chord...................... 34
15. Velocity and Turbulence Intensity ProfileTraverse#l Conf#2 Eval#l: 27 micro-meter RaRoughness from L.E.-l/4Chord ..................... 35
16. Velocity and Turbulence Intensity ProfileTraverse#l Conf#3 Eval#l: 44 micro-meter RaRoughness from l/2Chord-T.E...................... 36
iv
'".-' 17. Loss in Non-Dimensional Total Pressurewith Ra Roughness ........................... . 39
18. Loss in Non-Dimensional Total Pressurewith Rtm Roughness .............................. 40
19. Loss in Non-Dimensionalpressurewith Rpm Roughness ........................... 41
20. Velocity and Turbulence Intensity ProfileTraverse#l Conf#1 Eval#6: Smooth Baseline blade. 44
21. Velocity and Turbulence Intensity ProfileTraverse#1 Conf#6 Eval#2: 3 micro-meter RaRoughness from L.E.-i/4Chord .................... 45
22. Velocity and Turbulence Intensity ProfileTraverse#1 Conf#5 Eval#2: 20 micro-meter RaRoughness from L.E.-1/4Chord .................... 46
23. Velocity and Turbulence Intensity Profile
Traverse#l Conf#2 Eval#1: 27 micro-meter RaA Roughness from L.E.-1/4Chord .................... 47
24. Arithmetic Average Roughness ................... 54
25. Three different surfaces with equal Ra ......... 54
26. Derivation of Rtm roughness .................... 55
27. Derivation of Rpm roughness .................... 55
28. Velocity and Turbulence Intensity ProfileConf#l Eval#7: Smooth Baseline blade ........... 58
29. Velocity and Turbulence Intensity ProfileConf#2 Eval#l: 27 micro-meter RaRoughness from L.E.-i/4Chord ................... 63
30. Velocity and Turbulence Intensity ProfileConf#3 Eval#l: 44 micro-meter RaRoughness from i/2Chord-T.E .................... 68
31. Velocity and Turbulence Intensity ProfileConf#5 Eval#2: 20 micro-meter RaRoughness from L.E.-l/4Chord................... 73
32. Velocity and Turbulence Intensity ProfileConf#6 Eval#2: 3 micro-meter RaRoughness from L.E.-/4Chord................... 78
33. Velocity and Turbulence Intensity ProfileConf#7 Eval#l: 20 micro-meter RaRughness from I/8-3/8Chord.................... 83
V
-- - - - - --.'.:.,.,:'. 4-..:f:, . : ,-.-..... ..., ..-. . .. .". - . ". . -.. - .• . - . ." .- . .
- - --. - - - -
~D ~ 34. Velocity and Turbulence Intensity ProfileConf#8 Evalil: 20 micro-meter RaRoughness from l/-/Cod.......... 88
35. Velocity and Turbulence Intensity ProfileConf#8 Eval#2: 20 micro-meter RaRoughness from i/4-1/2Chord .................... 93
bvi
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LIST OF SYMBOLS
_Symbol Name units
A Area in
C Chord in
Cp Specific Heat at ConstantPressure B/Ibm-R
dA DifferintialArea in
dY Differintial Length in
Ratio of Specific Heats NONE
L Sample Length micro-meters
m Mass Flow Rate ibm/sec
Nu Nusselt Number NONE
Non-Dimensional Total NONECPressure Loss
P Pressure lbf/in
P Mass Averaged Pressure lbf/in
Ra Arithmatic Average Roughness micro-meters
Re Reynolds Number NONE
Density ibm/c ub ic-f t
Rpm Average Maximum Peak toMean Roughness micro-meters
Rtm Average Maximum peak toValley Roughness micro-meters
s Scaling Factor NONE
T Temperature R
TI Turbulence Intensity percent
V Velocity ft/sec
V Corrected Velocity ft/sec
vii
-,
I Subscripts
1 Inlet
2 Exit
S Static
T Total
Abbreviations
AFIT Air Force Institute of Technology
CTF Cascade Test Facility
HP Hewlet Packard
vii
LIST OF TABLES
Table Page
I. PERFORMANCE DATA FOR ROUGHNESSLOCATION TSS.....................2 3
II. PERFORMANCE DATA FOR ROUGHNESSMAGNITUDE TESTS .................... 38
4-.x
ABSTRACT
"An experimental investigation of the effects of/~
roughness on compressor blade performance in cascade at high
Reynolds numbers was performed. A s-e-ve i'blade cascade of
NACA 64-A905 blades with a chord of 2 in.jand a-n aspect ratio
of 1 was tested in a linear cascade. The blades were mounted
with a stagger angle of 31 deg and arn angle of attack of 15
deg. acing between blades was 1.33 inehee which gave a
solidity of 1.5.
This study was divided into two parts; -first-to
determine the importance of the location of roughness on
blade performanceand .6cond to determine how the degree of
roughness affects blade performance. The velocity and
turbulence intensity profiles of the flow region downstream
of the center blade and the non-dimensional total pressure
loss parameterkC4 were used to evaluate blade performance.
Hot wire anemometery "a-used t& measureithe' velocity and
turbulence intensity profiles downstream of the blade. The
measured exit velocity and static pressure were used to
derive the exit total pressure.
Roughness located near the leading edge of the blade
caused the greatest impact on Oand the velocity and turb-
ulence intensity profiles. Changes in the magnitude of t-he'
roughness did not affect the blade performance until a
threshold was exceeded; then tih- performance decreased
rapidly.
A .x
I. INTRODUCTION
BACKGROUND
The wide operating range of todays military aircraft
engine has lead to a flow regime characterized by a high4%
Reynolds number in the last few stages of the high pressure44
ratio compressor. In this regime the loss in performance is
thought to be a function of roughness alone. This is in
comparison to the low Reynolds number flow where the
efficiency is a function of the Reynolds number and not as
heavily dependent on roughness. The lack of good cascade
data for use in specifing blade surface finishes inorder to
minimize the loss in total pressure across the blade row has
lead to increased research in this area. One performance
parameter that of particular use to the compressor designer
is the total pressure loss coeficient (Ref.15).
In recent years the Air Force Institute of Technology
in conjunction with the Air Force Wright Aeronautical
Labortories, AFWAL/POTC, has been studying the effects of
roughness on compressor blades at a high Reynolds number.
This study was initiated in order to characterize the
effects of roughness on the performance of axial flow
compressor blades in cascade. The ultimate goal would be to
find a technically quantifiable surface finish that will
optimize the thrust specific fuel consumption (TSFC) and the
__ maintenance time of the gas turbine engine. With this
accomplished the Air Force could save millions of dollars in
"4
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." reduced operating costs.
The research began in iL? 8 1 with the design and
construction of a cascade test facility at the Air Force
Instute of Technology's School of Engineering (Ref.2). The
following year a study of the roughness effects of
compressor exit guide vanes at a high angle of attack was
conducted. This study showed that there was minimal effect
of roughness on the blades because of a large separated flow
region on the suction surface of the blade. The losses from
the separation greatly over shadowed any losses due to the
*roughness (Ref.6).
OBJECTIVE AND SCOPE
The basic objectives of this investigation are asJ.
follows:
* "1. Locate the region on the blade
where roughness affects the loss
in total pressure the greatest.
2. Evaluate how the total pressure
loss coeficient varies with
changes in the magnitude of the
roughness.
The approach used to evaluate the effects that
.k roughness has on blade performance was to build a cascade of
roughened blades then perform a wake survey to evaluate
2
, _ .;. - - . --"- -" • -" ;" . " q - . : " -w -- -- --- - . " -: * " . , " . "
blade total pressure loss in non-dimensional form. In
addition to the total pressure loss coeficient the velocity
and turbulence intensity profiles of the blade wake will be
compared.
The surface roughness parameters chosen for this
investigation are the arithmatic averege roughness, Ra, the
average maximum peak to valley roughness, Rtm, and the
average maximum peak to mean roughness, Rpm (see Appendix A
for roughness definitions).
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- II. EXPERIMENTAL APPARATUS AND PROCEDURE
CASCADE TEST FACILITY
The cascade test facility (CTF) at the Air Force
Institute of Technology School of Engineering was used for
this investigation. This apparatus, which was described in
detail by Allison (Ref.2), consists of an air supply coupled
to a diffuser and stilling chamber with provisions for
mounting a 2 in by 8 in test section. The air supply for the
CTF consists of a 40 horsepower centrifugal blower with a
name plate rating of 3000 cubic feet per minute at 26 onces
of head. The blower brings cool air from outside and mixes
it with warm recirculated air in order to control the test
section temperature. The air is then filtered using an
electrostatic cleaner to remove particles that would be
.etrimental to a hot wire anemometer sensor. A flow
straightening system consisting of two fiber filters and a
honeycomb is located in the stilling chamber. The air is
then delivered to the test section with turbulence
intensities of less than 2%, and a Reynolds number per foot
in excess of two million.
TEST SECTION
The test section, as deplicted in Fig. 1, was a cascade
containing seven NACA 64-A905 blades with a chord of 2 in
and an aspect ratio of 1. The blades were mounted with a
stagger angle of 31 deg and an angle of attack of 15 deg.
4
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This produced a nominal turning angle of 22 deg. The spacing
between the blades was 1.33 in which gave a solidity of 1.5.
The cascade of blades was followed by a 13 in long
adjustable channel that could be used either as a constant
area or a diffusing channel. The top and bottom end walls of
the channel are fitted with static pressure ports that are
used in aligning the exit channel with the free stream
direction. When the end walls were parallel to the free
stream direction, the static pressure at the respective
static ports on the top and bottom end walls were in
agreement and the inlet velocity profile was uniform. In
addition there are static pressure ports in the side walls
of the channel to measure the local static pressure at each
traverse location. The inlet duct to the cascade is also
fitted with static pressure ports that are used to calculate
A ,he inlet velocity and check the profile during alignment of
the exit channel.
INST RUMENTATION
The CTF was instrumented with four pressure
transducers, two thermocouples, a hot wire anemometer and
numerous manometers. The pressure transducers were used to
measure stilling chamber(test section inlet) total, inlet.0
static, exit static, and ambient pressures. The
thermocouples were used to measure stilling chamber total
and control room ambient temperatures. The hot wire
anemometer was used to measure both the X and Y components
6
10-
of the velocity and turbulence (fluxuating velocity) in the
region downstream from the blades. The hot wire system
consisted of a TSI Model 1241-10 hot film X-sensor in
conjuction with two TSI Model 1050 constant temperature
anemometers. Manometers were used to monitor the static
pressure on the top and bottom end walls of the exit channel
to insure that this channel was properly aligned with the
free stream.
HOT FILM SENSOR CALIBRATION
The hot film sensors were calibrated using a
temperature compensation procedure that is described briefly
as follows. First the operating resistance of the sensor was
0 determined for temperature range of intrest assuming an over
heat ratio of 1.5 as recommended by TSI (Ref.16). Once the
operating resistance was determined the anemometers were
set-up and calibration data were recorded at various
velocities and temperatures in the expected test range.
Usually four temperatures and ten velocities were recorded
in the expected range. The following data were recorded at
each calibration point: 1) calibrator total pressure, 2)
calibrator total temperature, 3) exit static pressure
(ambient), and 4) voltage outputs from each channel of the
anemometer. This data provided the necessary information to
determine the velocity, adiabatic wall temperature, heat
transfer rate and mean boundary layer properties of the hot
film sensor. From this information the Reynolds number-.
'."
(based on the diameter of the hot film sensor and mean
boundry layer properties) and the Nusselt number could be
determined. The calibration curve for the sensor is a plot
of the Reynolds number raised to the 0.51 power (Ref.4)
versus the product of the Nusselt number with the ratio of
the mean boundary layer temperature and free stream static
temperatures raised to an exponent that would collapse the
calabration data for the various velocities and temperatures
into a straight line as seen in Fig.2. The exponent used to
collapse the data, Xo, was found by assuming a value and
running the data through a least squares curve fitting
routine until a good fit of the data was obtained. The
exponent was usually found to be approximately 0.03 in the
. range of velocities for this study. Once the exponent, Xo,
was found, the data were plotted and the slope and Y
intercept of the linear calibration curve were found and
stored for future use. In order to use the calibration curve
the voltage from the bridge circuit of the anemometer must
-. be converted to a Nusselt number then multiplied by the
appropriate temperature ratio raised to the Xo power before
reading the Reynolds number from Fig.2. The velocity is then
determined from the Reynolds number obtained from Fig.2
using the appropriate mean boundary layer properties of the
flow in the test section where the hot film sensor is
located. The above procedure was programed into the HP 3052
Data Acquistion System in order to reduce the time required
to calibrate the hot film sensors.
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" .. EXIT PLANE TRAVERSER
The CTF includes a computer controled -X-Y axis
traverser that positioned the hot film sensor in the desired
location in the region behind the blades to within 0.002 in
in the X and 0.001 in in the Y directions. The traverser
mount was integrated with the test section design so that
its X-axis could be aligned with the free stream direction
for the flow leaving the compressor blade cascade (See
Fiq.l). Once aliqned the entire flow field behind thecascade could be probed with the hot film sensor.
After some initial testing on a smooth blade, a
standard data window was established behind the center
blade. This data window is comprised of five traverses in
Cthe Y-direction at 1 inch intervals starting 0.25 in behind
the trailing edge of the center blade. Each traverse
,contained 133 points located at 0.01 in intervals beginning
0.6 in below the center blade (Fig.3). The length of each
tr3verse was 1.33 in which corresponded to a single blade
spacing.
4
CENTER
BLADE
Figure 3: Standard Data Window
10
5'. '" . :, - ' 'v ,\ '-' ' .. :, -- , . - ., . -, - -- .-- -. ,. - . - •. . - . . - , ,- .
BLADE POUG9NESS CONFIGURATIONS
The effects of surface roughness were studied in two
parts in order to determine the importance of location and
degree of roughness. In the first part a strip of roughness
was applied at discrete locations. The strip was 1/2 inch
N. wide and its leading edge was located at 1/4 inch intervals
along the chord (Fig.4) . The second part of the study used
blades with roughness applied from the leading edge to the
1/4 chord on the suction surface (Fig.4a). At this location
the magnitude of the roughness was varied from 0.5
micro-meters Ra for the unroughened blades to 27 micro-
meters Ra for the roughened blades. In both parts of the
study, roughness was only applied to the suction surface of
the blade where the boundary layer and wake region are
thought to be greatly impacted by small changes in surface
roughness.
The blades were cast from epoxy and then the surface
was roughened to the desired finish. The roughening was
accomplished by one of two methods. The first method was to
coat the blade, or a portion there of, with a very thin
layer of adhesive and blow on a fine grit. The other method
used was to sand blast the blade to produce cavities in the
surface. When applying the roughness extreme care had to be
excersized as not to alter the shape of the blade. The
majority of the blade configurations investigated were
o. roughened by the first method as a wider range of roughness
4. could be created.
U 11
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(A) (B) (C)
Figure 4: Blade Roughness Configurations
To measure the Ra, Rtm and Rpm values for the roughened
blades a profilometer manufactured by Rank Taylor Hobson was
used with the parameter and recorder modules attached. The
blades were measured for Ra, Rtm and Rpm at various
locations on the blade and an average of these measurements
was used to characterize the roughness for this
investigation.0DATA ACQUISITION AND PROCESSING
The acquisition and processing of the data were
accomplished using a Hewlet-Packard(HP) 3052A Automatic Data
Acquisition System which is integral to the CTF. The data
acquisition system consisted of a HP 9845B Computer which
controlled a HP 3455A Digital Voltmeter, a HP 3437A System
Voltmeter, a HP 3495A Scanner, and two !P 9895 Disk Drives.
Also integral to the system were a HP 9874A Digitizer, a HP
9871A Daisy Wheel Printer, and a HP 9872S Plotter for input
and output of data. The various software packages designed
for the system allowed for the maximun utilization of the
. systems capability. The data acquisition software allow the
experimenter to specify the number and location of the data
12
b'.
points to be taken. The computer would then direct the
traverser to move the hot film sensor to the desired
location. Once the sensor was in place the scanner would
step the voltmeter throuqh the various channels and record
the voltage outputs from the various transducers. The
voltages were then stored on magnetic storage disk as "raw"
data. The "raw" data were later reduced into usable form as
pressures, temperatures, and velocities and recorded on
magnetic storage disk as "engineering" data for use in
evaluating various performance parameters.
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III. DATA REDUCTION
The parameters used in evaluating the effect of
roughness on blade performance were the exit velocity
profile, non-dimensionalized total pressure loss coeficient,
and the turbulence intensity profile of the blade's wake.
The total pressure loss across the cascade was evaluated
from the pressure and velocity data obtained from the survey
across one blade spacing centered on the center blade in the
cascade.
EXIT VELOCITY
The exit velocity as measured by the hot wire
anemometer experienced small changes in magnitude with
changes in temperature and humidity of the outside air
-supply to the CTF. The changes were first noticed as shifts
in the velocity profiles of two tests of the same blade
configurations on different days. The shift was later
noticed when the exit total pressure calculated from the
measured exit velocity was greater than the measured total
pressure in the stilling chamber. The shift in velocity is
thought to come primarily from a change in resistance of the
hot wire probe support with a change in temperature, and a
change in heat transfer rate with a change in humidity.
Tests run using a single sensor probe support, which had a
lower resistance than the double sensor support primarily
used (0.05 vs. 0.16 ohms), did not show the same size shifts
14
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that had previously occured when the double sensor probe
support had been used. The 40 deg temperature increase in
the probe support's environment that occured when the sensor
was moved from the calibration (control) room to the test
section would cause an increase in the resistance of the
probe support as the temperature increased. Preliminary
experiments run to define how the resistance of the probe
support changed with temperature were unsuccessful. The
humidity problem has been studied by Larsen and Busch
(Ref.7) with the conclusions that humidity increases the
heat transfer rate between 0.5% and 2%, corresponding to an
increase in apparent velocity between 1 and 5% but no
quanitative relationships were found.
0 In order to correct the hot film velocity data a
continuity check was performed. If the calculated mass flow
entering the cascade based on pressure data differed from
the mass flow leaving the cascade based on the hot wire
data, a correction factor was applied to the exit velocity
data to force continuity through the cascade. The mass flow
rate exiting the cascade was evaluated by numericaly
integrating the following relation:
=fp:z v dy(i
Where fn is the exit mass flow rate, e is the local density
.". evaluated using the ideal gas law, V is the exit velocity as
measured by the hot wire and dy is the differential exit
15
area assuming unit width. The inlet mass flow rate was
evaluated using:
M 1, VA, (2)
Where fi, is the inlet mass flow rate, e is the local inlet
density, A, is the inlet, area and V, is the average inlet
velocity evaluated as follows with the known inlet total and
static pressures and total temperature.
* = 2C'I-(tI (3)
If the derived exit mass flow rate did not equate to
the inlet mass flow rate a scale factor was applied to the
hot wire velocity data as follows to force the equality.
-(4
V2 S~ (4
Where V2 " is the corrected exit velocity, s is the scale
factor that equates the exit mass flow rate to the inlet
mass flow rate, and V2 is the exit velocity as measured by
the hot wire. The scale factor s was evaluated by a
numerical trial and error scheme and was found to vary from
0.90 to 0.95.
TOTAL PRESSURE LOSS
The total pressure loss across the cascade was
16
calculated from the measured inlet total pressure and the
exit total pressure as calculated from the exit static
pressure and velocity. The total pressure at each data point
behind the cascade was calculated using the following
isentropic compressibility relation:
. P + j; T (5)
where V2' is the corrected exit velocity, P is the static
pressure as measured by the wall static pressure ports, and
Ts is the total temperature as measured in the stilling
chamber. The total temperature of the flow was assumed
constant through the test section.
The overall exit total pressure was calculated from a
mass-averaging of the total pressure of the individual data
.points in a single traverse. The mass-averaging was
accomplished with the following relation:
22 ( 6 )
where is the mass-averaged value of the total pressure, P.,
is the total pressure at a single data point in the traverse
being evaluated, 10 is the local density evaluated using the
• -- ideal gas law, and dA is the differential exit area. The
double integrals over the exit area were reduced to a single
17
"% ., -% - q q q . % ". ." .% " " % . . ." .. . . .% " % " . •. .• ..
g°'I
S"integral by assuming no var iations in the span wise
direction of the blade. For unit width of the blade, the
equation then reduces to:
R - _ V_ _ _ (7)
'2
The integrals were numerically evaluated using the
trapezoidal rule with dy being the step size in the
Y-direction (0.01 in).
Having calculated the exit total pressure, the loss in
total pressure across the cascade can be evaluated using:
= (8)IN
The total pressure loss can be non-dimensionalized by the
inlet dynamic head to give (Z, the desired performance
parameter.
APT
V 2 . (9)
where/0 is the inlet density evaluated using the ideal gas
law and V is the inlet velocity evaluated in equation 3.
The total pressure loss was calculated for each traverse and
left an average of all five traverses were taken as the total
pressure loss parameter in this study. It is realized thatV]
18
S S * . . . . . . . ..
-.
mixing losses occur in the wake but the error of the
instrumentation was greater than any mixing losses and an
average loss parameter was thought to be a better
representation of the blade losses.
TURBULENCE INTENSITY
The definition of turbulence intensity for this
investigation is the ratio of the RMS velocity (AC output of
the anemometer as indicated by a TRUE RAS voltmeter) to the
stream wise component of the local mean velocity (DC output
from the anemometer).
TT v , (10)0
Turbulence intensity would not normally be negative but
Pecause of the configuration of the X-sensor probe used to
measure the components of the velocity the two sensors allow
the orientation of the fluctuations to be determined. As a
result, positive value of the turbulence intensity indicates
that the fluctuations are occuring in a plane rotated in a
positive direction about the Z axes from the positive XZ
plane, and ,conversely, a negative Y turbulence intensity
indecates a negative rotation from the positive XZ plane
(Ref.17).
19
- .oJ
F7IV. RESULTS AND DISCUSSION
The objectives of the study were to determine the
importance of position and magnitude of roughness on the
loss in total pressure across a cascade of compressor
blades. Eight blade configurations were tested in the AFIT
CTF with sand type roughness applied at various locations
along the chord and with varying magnitudes. The cascade
used for this study did not use boundary layer removal
therefore the cascade could not be considered two
- dimentional according to Erwin and Emery (Ref.5). The
results being presented are good for making comparisons
between smooth and rough blades but should not be considered
design cascade data.
A. ' The non-dimensional total pressure loss coeficient, C0,
as defined below, was used as the primary means of
evaluating the blade total pressure losses.
Pr -( iY2epv 2
In addition, the velocity and turbulence intensity profiles
N of the blade's wake were studied to evaluate the effect that
roughness had on the wake. The surface roughness was
%..4 evaluated using a profilometer to measure Ra, Rtm, and Rpm
. values of the surface roughness (see Appendix A for
- .. *.*definitions of the above roughness parameters)
20
I .I
,. .- '', LOCAMION EFFECTS
Five blade configurations with roughness on por.tions of
the suction surface were tested to determine the importance
of the roughness location on the loss of total pressure
across the blade. Initially two blades were tested; one with
roughness from the leading edge to the 1/4 chord (Conf#2)
with an Ra value of 26.9 micro-meter, the other
configuration with roughness from the mid-chord to the
trailing edge (Conf#3) with an Pa value of 44.1 micro-meters
Figs.5a&b. The loss in total pressure across the blade, as
R: Conf#2 B: Conf#3
Figure 5: Roughness Configurations 2 and 3
measured by o, increased from 0.0963 for the blade with the
roughness near the trailing edge to 0.1037 for the blade
with the roughness begining at the leading edge. The blade
with the roughness begining at the leading edge thus
experienced an 11% greater loss in Q based on 0 of the
smooth blades. The increase in 0 occurred even though the
magnitude of the rough'ness was less on the blade
configuration with the roughness beginning at the leading
edge compared to the blade with the roughness near the
21
4 . .. -
trailing edge. These tests showed that rouihness located
near the leading edge, on the suction surface, is critical
to the performance of the blade. Further testing showed that
by moving the roughness location away from the leading edge
even small percentages of the chord caused a decreases in a.
Table I shows the decrease in c that accompanies the
movement of the roughness away from the leading edge. For
these tests blade configurations 5,7,&8 were used which
*, .- consisted of a 1/2 in band of 20 micro-meter Ra roughness
applied to the suction surface of the blade. The leading
edge of the strip of roughness was initially at the blade's
leading edge (Conf.5) Fig.6a but was progressively moved
away from the blade's leading edge in 1/8 chord intervals
(Confs.7&8) Figs.6b&c.
%LU7
R: Conf#5 B: Conf#? C: Conf#8
Figure 6: Roughness Configurations 5, 7 and 8
Moving the strip 1/8 chord from the leading edge decreased
a from 0.0973 for the blade with the roughness begining at
the leading edge (Conf#5) to 0.0918. Moving the roughness
__ back another 1/8 chord decreased 0 to 0.0870.
A cause for the high loss in total pressure when
22
W 01) trE0
4 I) 0 3 04-403 W 0 w 0
E .04 I4 U 'E
C.) I0 lam 0a rLoL43 (D 141 - 0 0 > to-~-
0f I U 0' 0-
04 -4-4 0->O4 u-
4 0 Uco r- m -w %D I44-- 10 43 In n %0 r E--4 IEr_1
m- IX I
8 1 r- O O m. '0m i- 1.3 1ir- 4 0 0 D CD
m i4 U) a;4C~ ~ C;J C; C; C; C;
.4. Ci)
to) I- -0
o) F- 0o m rI l n L41 0 01 0 0 0
w I4
u AA
E m~to Lnr c V-4 -
%.0 D-q I- N *- 4
P -4 CN r n r 0Oo2
" .. roughness begins at the leading edge of the blade is that
the size of the roughness is large in comparison to the
boundary layer thickness. On this premise, when the peaks
of the roughness extend through the laininar sublayer of the
boundary layer, a significant increase in the boundary layer
growth results. As the boundary layer grows along the blade,
the peaks of the roughness become submerged in the boundary
layer and their effects on the boundary layer decrease.
Similar effects have been seen on a flat plates where the
coeficient of skin friction decreased as the roughness was
moved away from the leading edge (Ref.1l). Schlichting
points out that the reduction in the coeficient of skin
friction is due to a reduction in the size of the roughness
in comparison to the local boundary layer thickness.
The effects of the location of roughness on the wake
down stream of the blade can be seen by inspection of the
velocity and turbulence intensity profiles in this region.
The measured velocity and turbulence intensity data for the
X and Y directions have been resolved into vectors with
magnitude and direction. The vectors were then plotted using
the point at which the data were taken as the origin and
drawing a vector in the appropriate direction. The solid
line vectors are the velocity vectors and the dotted line
vectors are the turbulence intensity vectors. The magnitude
of the vectors can be determined by measuring the length of
* the vector using the down stream position scale as a measure
of length and applying the appropriate scale factor. The
24
-v o- i.-. * ..
scale factors for the velocity and turbulence intensity
vectors are 540 ft/sec/in and 20 percent/in respectively. As
can be seen in the set of the velocity and turbulence
intensity profiles for the smooth blade (Conf.l),
Figs. 7,8,9,10,&11, both the velocity and turbulence
intensity profiles are fairly uniform in the free stream but
exibit large changes in magnitude and direction in the blade
wake region. The velocity profile exibits a small narrow
decrement accompanied by high levels of turbulence as seen
-. in the traverse located 1/8 chord from the blade's trailing
Oedge (Fig.7). As the traverse location moves down stream
(Figs.8,9,10,&ll) the velocity decrement becomes wider and
shallower accompanied by decrease of the levels of
0turbulence but this is accomoanied by a spreading of the
turbulence into the free stream. All the blade
gonfigurations tested exibited the same trends as the smooth
blades but with changes in the relative sizes of the
velocity decrement and turbulence levels. The following
dicussion of the effects of the location and magnitude of
roughness on the blade's wake will only deal with the
traverse located 1/8 chord behind the trailing edge of the
blade with the understanding that events similar to those of
the smooth blades also occur downstream of the roughened
blades. In order to verify that these trends do occur the
reader is refered to Appendix B which contains a complete
set of the velocity and turbulence intensity profiles for
each configuration tested.
25
.....................*'2..,.-...0 e.. .-
.U-
cr U1 0w.M
>0
0~
r--4 r4tta) I.
JiTl I~
0l 00-1!o
w il 0JpI
ill iii
* I * I p I . p I * c
sra see sil* GSS11E 6V6 6 1(6WHONI NOIJ.IS~d WW3NIS SSONiD
26
* b j - -. u
*Joil
etoll
' .* C'
~ ~II
J~ii1u el
gci m
wEr,4w) rOJ
zu 01~ W
i'i H I--11 U) 3Li ~ ,I'I,,,II~/
cc I~fI~IJJI~jaiI Ii0
I~l ii H
I ~ H I i Mi.1Pl itI I I
66H(63HN ISdW&9SSO
27'
*.*4~~il --4- *. -.- ...-. I I* IP. I I r
r ~ ~ ~ ~ ~ j -1.I IV'-- - ~ .- - ~ - -
. . . . . . . . . . . . . . . . . . . . . .. . . . . . .
zS
M~ 'I> IIi4 4U)E-
4JTM
G)Q
Jill 'Il 1 -II " f m ) -
z~~ ~ ~ ~ If1 1 ;x ) -Ii 1m1mK cnM0 L I
Li i I ~ il'~l~I'1'I'qiiI~ 11111 I i~ f E-4ljj0 4 1I9 ; fJf
.1 i iI ij
Y iN
cc~I~~t1~f
(SHdI NO-4S~ WUr-6 SOJi j
( L
*%p *a:
> hr
A0
"--4
~~~~®M11111 .- %'.f.--i.il:'l 11112'fl 1.. . . . . . . . . . .1111
in.
>LI CL U
u >- (Lz 4j >
-4- (0
4Z00
a w
I: r-4
WI r,~Ia-4
ieii iI [ Ii I
0 44
MSHONZ) NOI.LISOd WU361S. SSONO
29
7-.-77
Uoil
till
LO ice 5
0 10S
z. -w
U; Z 4. Q
0 Cl TI
00i Irv
0 to
* I I if -I! Il O l 0
in' Z m(SHN)NI.Id iJ3.SSO
Y T1 04
U". 30
............................... 0%0
. . . . . . . . . . . . .. . .'. .-.. ..- I
The velocity and turbulence intensity profiles for the
blade configurations where the roughness started" at the
leading edge and was moved away in 1/8 chord intervals
(Conf#5,7,&8) showed a large change in the blade's wake from
that of the smooth blade but did not show any differance
between themselves. This is seen by the large increase in
the depth of the velocity decrement and the increase in the
magnitude of the turbulence levels when Figs.12,13,&14 are
compared to Fig.7. When Figs.12,13,&14 are compared to each
other no significant differance can be seen. The velocity
and turbulence intensity profiles for the blade
configuration that has roughness begining at the leading
edge and extending to the 1/4 chord (Conf#2) and the one
with roughness from the mid chord to the trailing edge
(Conf#3) , Figs. 15&16 respectively, show a large change in
the blade's wake. When these profiles are compared Conf#2
exibits a larger velocity decrement than conf#3 but a
smaller leval of turbulence. A possible cause for the
increase in turbulance intensity, when roughness is applied
from the mid-chord to the trailing edge of the blade, is the
position at which the boundary layer becomes separated. When
the boundary layer experiences roughness after the mid-chord
it may immediately separate due to a unstabilzzing effect of
the roughness on the boundary layer in a unfavorable
pressure region. The separated flow over the blade would
*account for the high turbulence intensity levels. If the
boundary layer experiences roughness before the mid-chord,
31
* *1.7
:AL
cuN
4 ~..J ~ *-
,~ ->
LLm
00 F
w 0d
z-
U..-
QI- )tn a)
.41Z li li 1c 0 4
i laI i iiiii I i*i j
m E
CD C.)$
lit ~~fit
ii4L'L(S3HOD NO1IlSOd WW3N1S SS080
41' ,**
32
4)
*> U
0C-
00
1.4 4J fn
Z 0
*l I i
(S3H~N) NOI I WU11T,4
~~t,
33
FE . w-o - - . -
aS 0 U)lL
0 j
toW C4I
Lii~ INQ)
* 0 >
c- cfG
Li..li
U.E- I II
cc 0i~ ~'I i~~ II I ~ ~ i il ~ ~i1CE'4
ii'jI r
>
* * * * * 4
(S3H3)NI) NOIJISOd W3NIS SSONID
* . 34
CD
Cuj
$0 0ONi
U).> 0d
-4J 0n
(*UZ 44 .~ ij
0 ..............
o
j~~~4 m1 II~ILJ0
00
r Q)
r I wz 0.
1ff E- Ii
JL
a:1
1%* 35
-~~~~~~- - r . . . . , - .
-- ~ r--
---.... ~* -. - - .>
cu
* L40
AL HI
LLI 44L) '
12.4 WI
~t.to1!04
i' il m~ U( ) 4Jt
.~~. (S3H!D111 NOI..LI~ U~L NS~' !-.. ..Ih
'V I36E-
*in
.: :! A - .1
in a region where the flow is accelerating around the blade,
the boundary layer may not separate immediately dud to the
favorable pressure gradient but remain attached a along the
blade. The closer to the leading edge that the boundary
layer becomes detached, the greater the separation region
thus the larger the turbulence levels particularly if it is
unstable.
ROUGHNESS MAGNITUDE EFFECTS
• . Four blade configurations with roughness applied to the
first 1/4 chord were tested to determine the effects of
varying the magnitude of roughness on the loss in total
pressure across the cascade of compressor blades. The
results of these tests are oresented in Table II but can be
seen best in graphical form. Figures 17,18,&19 are plots of
.the increase in non-dimensional total pressure loss that
accompanied an increase in roughness as we hr(? ILy 7 , .ts,
an,! 7 im roughness parameters. Even though the size of the
,- values for the roughness parameters change from one plot to
the next the relative trend remains the same as can be seen
in Figs.17,18,&19. The sand type roughness used in this
study created such a uniform randomness that the method of
evaluating roughness did not affect the trend. For this
reason the following discusion will only deal with Ra values
of roughness. Rtm and Rom values of roughness could have
just as easily been used and would have led to the same
conclusions.
3714
J0 4
410 1I %D N 0)
I 04
CI CD 0 0 0UI c
wz N-41 P- Nr n :
rL0 -4-4r. E
0 - r I r-4 t-o '0 r-4 cotw Q*-e 4j 1- c r) r 0m 0 10 04 C 4 L 0 0c w-
E4 I. a * * G
43 o O 0 0 0II z C;C)t
0 ~ .-I 0nC2L m 0e Ai-'.
43 0 (aWI44'_404 '4 94 M
ca I r-4 %0 t4cq-4 40) -
E4 0.I
- ~I V *0
IPA A r-~38
U) 0f010iLI JZ toa
o2 0
'4. CU
-Aini 0
-~ L
Li
LL 0CE -'
L4-.
4'.I
-V.,-
U) in
M" MG9 W3WO
39
Cu M utu
I'- m' N U)
(45 F- @
oG
L)~ L) (i
I IE
0C'L CuC
0LJ
Z Hd
4.4J
-'. Iji ZO
a. r
0
U)
0G
SUB U 53WO
4440
*~- I- 4 - C4
CY L
M p Q U4> 44 42 -
4~C i'- CV S 0%.
c) -,) v
0)
LiiIT- D -
000 E
.1i)
0u0
xx
xL
.44
The change in total pressure loss across the cascade as
measured by o can easily be seen in Fig.17. As the roughness
magnitude varies from 0.5 micro-meter Ra for the smooth
blade (Conf#l) to 26.9 micro--meter Ra for the roughest
leading edge configuration (Conf#2) c changed from 0.0775 to
0. 1037 . For small values of Ra the change in a was
negligible, as the Pa value of the roughness increased the
non-dimensional total pressure loss, z, showed an increase.
When the z values for the blade configuration with a leadinq
edge roughness of 2.7 micro-meters Ra (Conf#6) is compaired
to that of a blade configuration with a leading edge
roughness of 19.8 micro-meters Ra (Conf#5) a 27% increase in
Itotal pressure loss is experienced. The percent increase is
based on the difference between 4 for the two tests in
comparison to z for the smooth blades (Conf#l) . A further
increase in roughness to a Ra value of 26.9 micro-meters
brings an additional 8.3% increase in non-dimensional total
pressure loss.
The increase in z with roughness is thought to be
caused by the peaks of the roughness extending through the
boundary layer and causing a significant increase in the
growth of the boundary layer. On this premise, when the Ra
value of the roughness was small the peaks of the roughness
did not extend through the boundary layer, thus the
insignificant change in r with roughness as seen in Fig .17.
As the magnitude of the roughness on the blade increases,
the peaks begin to extend through the laminar sublayer of
42.4
.*... . 4., . . . . .. . . ... . .
LI.-. ! .:'. the boundary layer and the loss in total pressure starts to
J:1
increase. The data seems to indicate that there should be
some Ra value of roughness at which point the peaks just
start to penetrate the laminar sublayer of the boundary
layer. At this point the blade's surface transitions from a
hydrodynamicly smooth surface, where the roughness is
submerged in the boundary layer and the performance is not a
function of the roughness, to a hydrodynamicaly rough
surface, where the peaks of the roughness extend through the
boundary layer and large losses in performance are
experienced with changes of roughness. This transition point
could not be investigated due to the lack of sand of the
necessary size. Similar affects of roughness on the total
pressure loss of compressor blades was found in a study by
Schaffler. Schaffler found that the polytropic efficiency of
a compressor decreased as the magnitude of roughness
increased past a critical value. At this point the surface
of the blades transitioned from a hydrodynamically smooth to
a hydrodynamically rough surface (Ref.10).
When the velocity and turbulence intensity profiles are
examined for configurations 1,6,5,&2 (Figs. 20,21,22,&23), an
enlarging of the wake region is noticed. These
configurations correspond to blades with roughness located
from the leading edge to the 1/4 chord with Ra values of
0.5, 2.7, 19.8 and 26.9 micro-meters respectively. When the
• i'~e., profiles of the first traverse of Conf#l (Fig.20) are
compared to those of Conf#6 (Fig.21) no significant
43
0 $4
Ld U4
C.0 0
'34 co
* S * 0
z .... .....
-c 0
Q 0Jill~
0 a4J 0o) vI 155 L
1 '!(!..~II I
Jil $4 %D
'UJ 4jt
i.a:.
>4
* ~ ~ ~ ~ ~ ~ ~ ~ ~ I Jill*. j ~ ' *.* ~ 4 .1 . .
,~ ******- ~*%**** ~I U '.*
cuu
00
if E-4 -4
> 4 -4zz
0- W
0
CLi.
H 11I !J l II
i E-1*iq~ l
(63H~~~~NI)a 44.I~ U3.SSO
cu
cuu.7.O.
d *En
0 a -0
zOb 0i4
Zk 0 4
z cc
0
I f (~1. 0 41 J~
F - I- I ~i ~ j 0
CE WIX.'. .
-LL 4Jj jIC t
II I i4~
IjI . .
>I., .
.7...ii I . i' I4
j~i~~iiH III -**~~ i~ i ll I4' I 4
6 IE 0 IS Ip E WE W~E(SHOD NOI11SOd WW3NiS SSOMO
46
La W .
in1
w 1.4 a
M~t
Cl U)
z
L4C~j 04 L
00z T LL1 J
II, jj III IJIj 4-(if 'jf i/ U~ 1 f Iu
~ I, 0
ifWillii
47I
-E't ST VM ISI SGIaSV> ............
-... .. . . . ... - . .. . .. - . . ... . .
a _.
, differences can be noticed. The velocity decrement in both
cases have the same general shape, characterized, by the
steep gradient on the pressure side of the wake (upper
v portion of the velocity decrement) and the gentle gradient
on the suction side (lower portion of the velocity
decrement) . The depth of the velocity decrement and the
magnitude and direction of the turbulence intensity vectors
are also the same. When configurations 5 and 2 (Figs.22&23)
are compared to configurations 1 and 6 a large difference in
their wake regions is noticed. Configurations 5 and 2, with
their larger magnitudes of roughness (19.8 and 26.9
micro-meters Ra respectively as comoaried to 0.5 and 2.11
micro-meters Ra for configurations 1 and 6), exibit a deeper
velocity decrement with larger gradiants on the sides than
configurations 1 and 6. The turbulence intensity vectors for
configurations 5 and 2 also show an increase in magnitude
over configurations 1 and 6.
.V The velocity and turbulence intensity profiles suggest
that as the roughness level increases there is an
insignificant change in the flow over to the blade until a
threshold is acheived. Once this threshold is passed a
significant increase in the boundary layer growth occurs and
a loss of performance is experienced.
REPEATABILITY
During the testing of the various blade configurations
4 lgreat care was taken to insure constant test conditions
48
. . . . . . . . . . . . . . . . .... . .. .. : ..- .. ...
during a run. It was found practically impossible to
duplicate test conditions from run to run because of 'changes
in the weather that affected the air supply's humidity and
temperature. These changes in the air supply primarily
affected the hot wire system by increasing the heat transfer
from the sensor and the resistance of the probe support.
These changes caused a uniform shift in the measured
velocity data from 3 to 7 percent. This increase in the hot
wire velocity data caused the calculated exit mass flow rate
to be greater than the inlet mass flow rate as derived from
the inlet static and total pressure data. The pressure
transducers having been very stable over the course of the
study were assumed correct and the inlet mass flow
calculated from the pressure data was therefore assumed
correct. To correct the calculated exit mass flow a scaling
".w factor was applied to the hot wire velocity data that caused
the derived exit mass flow to equal that of the inlet.
To estimate the total error of the system duplicate
tests were run on configurations 1 & 8 and comparied for
repeatability. After the mass flow correction factor was
applied to the measured velocity the average performance
parameter, 0, was within 3% for diplicate tests of the same
blade configuration as was the velocity data. The turbulance
intensity data differed by no more than 5% between the
. duplicate tests.
49
V. CONCLUSIONS A'ND RECOMMENDATIONS
CONCLUSIONS
The effect of roughness on the performance of
compressor blades was studied in a seven blade cascade of
NACA 64-A905 blades. The cascade used did not use boundary
layer removal therefore could not be considered two
dimensional according to Erwin and Emery (Ref.5). The
results of this investigation showed that roughness has
significant affect on the loss of total pressure across the
.cascade of blades but the data presented in this report
should not be considered hard fast design data.
Roughness affects the loss in total pressure across the
compressor blades as follows:
1. The greatest loss in total pressure occured with
xoughness located near the leading edge of the blade.
2. The loss in total pressure across the blade is
dependent on the degree of roughness only after a threshold
level of roughness is exceeded.
RECOMMEN DATIONS
A better understanding of how roughness affects the
loss in total pressure across compressor blades could be
accomplished if the boundary layer thickness of the blade
were evaluated and compared to the size of the roughness.
One possible method for evaluating the boundary layer
thickness would be to physicaly measure it with a hot wire
50
, A . .. . .. ... . .. .. . . .
anemometer. Before evaluating the boundary layer thickness a
new test section should be made that employs continuous
boundary layer removal along the side walls of the cascade.
Without boundary layer removal the cascade can not be
-+considered two dimensional according to the NACA report 1016
(Ref.5). In addition to the continuation of the present
investigation, a study of the effects that humidity have on
the heat transfer rate of hot film sensors and the effects
of a temperature change on the resistance of the hot wire
probe support should be initiated.
1'
i,051
BIBLIOGRAPHY
1. Abbott, I.H., von Doenhoff, A.E., and Stivers, L.S.,Jr., Report No 824, National Advisory Committee forAeronautics, p. 258.
2. Allison, Dennis M. Design and Evaluation of a Cascade-' -2 Test Facility, Unpublished MS Thesis. Wright-Patterson
AFB, Ohio: Air Force Institute of Technology, June1982.
3. Bammert, K. and Sandstede, H. "Influences ofManufacturing Tolerances and Surface Roughness ofBlades on the Performance of Turbines" ,Journal ofEngineering for Power (January 1976).
4. Bradshaw, P. An Introduction to Turbulence and ItsMeasurement, New York; Pergamon Press, INC., 1971.
5. Erwin, John R. and Emery, James C.,"Effects of TunnelConfiguration and Testing Technique on CascadePerformance", Report No. 1016 National AdvisoryCommittee on Aeronautics, P. 263.
6. Genovese, David T. Roughness Effects on CopressorOutlet Guide Vanes at H igh R o and Tigh AnleBAttack, Unpublished MS Thesis. Wright-Patterson AFB,M'Tio: Air Force Institute of Technology, December1982.
7. Larsen, Soren E. and Busch, Niels E. "On the HumiditySensitivity of Hot Wire Measurements".DISA InformationNo. 25: 4-5 (February 1980).
8. Perry, A.E. and Schofield, W.H. 'Rough Wall Turbulent
Boundary Layers", Journal of Fluid Mechanics, Vol.37(1969).
9. Pope, Alan. Wind-Tunnel Testing, New York; John Wileyand Sons, INC., 1954.
10. Schaffler, A. "Experimental and AnalyticalInvestigation of the Effects of Reynolds Number andBlade Surface roughness on Multistage Axial FlowCompressors", Journal of Engineering For Power,102:5-13 (Januarg§ -- --"
11. Schlichting, Hermann. Boundary-Layer Theory, New york;i". McGraw-Hill Book Co.,1979.
5• 52
.4
12. Smith, A. and Clutter, D. "The Smallest Height ofRoughness Capable of Affecting Boundary-LayerTransition", Journal of the Aero/Space Sciences,Vol.26 (April lT - - - - - -_
13. Stiles, R.J. Roughness Effects in Axial Compressors-AnEmperical Model-. UnpuETishe Report. WrTight-PattersonAFB, Ohio: Wright Aeronautical Laboratories, 1980.
14. Taylor Hobson. Surtronic 3. Operating Instructions.Leicester, Engla.--nT-T'lor Hobson, (undated).
15. Thomas, M.W. and Evans, D.C. General Electric GasTurbine Fundalmentals Vol II, Ufp'uf61fT7he-report,GenerlElit-F7c Co. h-Na e--O-io, July 1973.
16. TSI,General System Information For 1050 SeriesAnemometry,TSI Incorpora tecT,50-6-Caig-a- , St. Pual,Minnesota, 55164.
17. Vonada, John A. and Elrod, William C. Wake MixingInvestigation of Crenelated Trailing Edge Blades inSCasca e, Unp=W-' led report. Wright-Patterson AFB,7.!!,5i0Mir Force institute of Technology, October 1983.
53
, -. - . - - - -'- -"- --
..-.
."
APPENDIX A: ROUGHNESS DEFFINITIONS
Surface roughness is defined as the finer
irregularities in the repetitive or random deviations from
an intended surface contour (Ref.13). These random
deviations which look like a series of peaks and valleys are
most commonly characterized in terms of "Arithmetic Average
Roughness, Ra" as defined below.
lo- L~~
RIy(12)
Figure 24: Arithmetic Average Roughness
krithmetic Average Roughness does not completly
%characterize the surface for there could be several surfaces
with equal Ra but have different appearances and different
hydrodynamic characteristics Fig.25.
(a) Ra
(b Ra b
(C) V -' obbV .I Ra C
Figure 25: Three different surfaces of equal Ra (Ref.13)
. 5
_n 54
.=- - t " -" " *. %, *
For this reason two other roughness parameters will be
defined Average Maximum Peak to Valley, Rtm; and Average
Maximum Peak to Mean, Rpm.
'4 The Rtm roughness parameter characterizes the surface
by measuring the maximum distance between peaks and valleys
4 1."over several measuring lengths and then averages these
maximums for the total traverse Fig.26.
Am...,, t Ram.., Rnmj.. IA... " -t a m.,
SI .i "" ...I
Rtm =Rmax, Rmax - Rmax 3 Rmax. - Rmax(
J 5
Figure 26: Derivation of Rtm Roughness (Ref.14)
The Rtm roughness parameter defines the maximum average
*peak to valley distance, which is important when studing the
hydrodynamic character of a surface, but does not
distinguish between a surface primarilly of valleys or peaks
as in Fig 25b&c. A third roughness parameter, Average
Maximum Peak to Mean Roughness, Rpm is defined in Fig.27
similar to Rtm but only measuring the peaks of the
roughness. A, f ,oll
p . Rp, + Rp. + Rp, + Rp. + Rp.
Figure 27: Derivation of Rpm roughness
55
'-'"With the three roughness parameters Ra, Rtmn,and Rpm the
surface's hydrodynamic characteristcs can be adequately
-' def ined.
..
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VITA
Frederick Johnson Tanis Jr. was born on 29 July 1960 in
Vi: Red Bank New Jersey. After graduating from Middletown High
School South in 1978 he attended Lafayette College in Easton
Pennsylvania. He graduated from Lafayette College Cure Laude
with a Bachelor of Science degree in Mechanical Engineering
in May of 1982. He received his commision in the United
States Air Force Reserve through the Reserve Officer
Training Corps at Lehigh University in June of 1982. As his
first assignment Lt Tanis entered the Air Force Institute of
Technology rt Wright Patterson AFB in June of 1982.
Permanent Address:
500 Newman Springs Rd.
Lincroft New Jersey 07738
98
- .. . - 4 . * 4 V
o,. -. . '
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"..REPORT DOCUMENTATION PAGE. & REPORT SECURITY CLASSiFiCATION 1lb. RESTRICTIVE MARKINGS
UNCLASS IFIED2&. SECURITY CLASSIFICATION AUTHORITY 3. DISTRIBUTION/AVAILABILITY OF REPORT
2b. DECLASSIFICATION/OOWNGRAOING SCHEDULE Approved for public release;distribution unlimited
4. PERFORMING ORGANIZATION REPORT NUMBER(S) 5. MONITORING ORGANIZATION REPORT NUMBERIS)
AFIT/GAE/AA/83D-23
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School of Engineering AFIT/ENY
6c. ADDRESS (City. State and ZIP Code) 7b. ADDRESS (City. State and ZIP Code)
Air Force Instute of Technology
Wright-Patterson AFB, Ohio 45433
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PROGRAM PROJECT TASK WORK UNITELEMENT NO. NO. NO. NO.
11. TITLE Ilnclude Security Claaification)See Box 19
- , 12. PERSONAL AUTHOR(S)
Frederick J. Tanis Jr., B.S., 2d Lt, USAF13a. TYPE Of REPORT 13b. TIME COVERED 14. DATE OF REPORT (Yr., Mo.. Day) 15. PAGE COUNT
MS Thesis FROM _ TO _ 1o_ 1983 December 10816. SUPPLEMENTARY NOTATION O: 11w 14211 111-17.
1?. COSATI CODES 6S. SUBJECT WSPUjSaCuaktnI~i&@,JI necezaary and identify by block numb.r)FIELD GROUP SUB. GR. Cascade Testing, Compressor Blades, Roughness Effects21 1 High Reynolds Number, Total Pressure Loss
1. ABSTRACT (Continue an reuerme if neceuary and identify by blocl number)
Title: Roughness Effects on Compressor BladesPerformance in Cascade at High Reynolds Number
Thesis Chairman: Dr. William C. Elrod
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Willia C. 'lrod 513-255-3517 AFIT/ENY
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SECURITY CLASSIFICATION OF THIS PAGE
-'- An experimental investigation of the effects of roughness on compressorblade performance in cascade at high reynolds number was performed. A sevenblade cascade of NACA 64-A905 blades with a chord of 2 in and an aspect ratioof I was tested in a linear cascade. The blades were mounted with a stagger angleof 3t deg and an angle of attack of 15 deg. The spacing between the bladeswas t.33 in which gave a solidity of 1.5.
This study was divided into two parts, first to determine the importanceof the location of the roughness on the blade performance and second todetermine how the degree of roughness affects blade performance. The velocityand turbulence intensity profiles of the flow region down stream of the center
S.blade and the non-dimensional total pressure lossparameter, r, were usedto evaluate blade performance. Hot wire anemomete!y was used to measurethe velocity and turbulence intensity profiles down stream of the blade. The
* measured exit velocity and static pressure were used to derive the exittotal pressure.
Roughness located near the leading edge of the blade caused the greatestimpact on Q and the velocity and turbulence intensity profiles. Changesin the magnitude of the roughness did not affect the blade performanceuntil a threshold was exceeded then the performance decreased rapidily.
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