Full Scale 
Experimental 
Transonic Fan Interaction with a 
Boundary Layer Ingesting Total Pressure Distortion
 
 
 
 
Justin Mark Bailey
 
 
 
 
Dissertation submitted to the faculty of the 
 
Virginia Polytechnic Institute and State University 
 
in partial fulfillment o
f the requirements for the degree of
 
 
 
 
 
Doctor of Philosophy
 
In
 
Mechanical Engineering
 
 
 
 
Walter 
F. 

 
Clinton 
L. 
Dancey
 
Wing 
F. 
Ng
 
Kevin T.
 
Lowe
 
Alfred 
L. 
Wicks
 
 
 
 
 
November 
7
, 2016
 
Blacksburg, Virginia
 
 
 
 
 
 
Keywords: 
Experimental Engine Test
ing
, Distortion, Interaction
, Total Pressure, Boundary 
Layer Ingesting
 
 
 
 
 
Full Scale Transonic Fan Interaction with a 
Boundary Layer Ingesting Total Pressure Distortion
 
 
Justin Mark Bailey
 
 
A
BSTRACT
 
 
 
Future commercial transport aircraft will 
present 
more a
erodynamic 
architectures
 
to 
accommod
ate stringent design goals for 
higher fuel efficiency, reduced cruise and taxi NOx 
emissions, and reduced noise. 
Airframe d
esigns
 
most likely to 
satisfy the first goal feature 
architectures that lead to the formation of 
non
-
uniform
 
flow introduced to the engine
 
through boundary layer ingesting (BLI) inlets
, creating a different operational environment 
from which the engines were originally designed.
 
The goal of this study was 
to 
explore the 
effects such 
non
-
uniform
 
flow
 
w
ould have 
on the behavior and performance of a transonic 
fan in a full scale engine test environment. 
 
This dissertation presents an experimental study 
of 
the interaction between a full scale 
transonic fan and a total pressure distortion representative of 
a boundary layer ingesting 
serpentine inlet
. A five
-
hole pneumatic probe was traversed directly in front of and behind 
a fan rotor
 
to fully characterize the inlet and outlet fan profile.
 
The distortion profile was 
also measured 
at 
the aerodynamic interface
 
plane (AIP) 
with a
n SAE
 
standard total 
pressure rake, which has historically been accepted as the 
inlet profile to the fan. T
his 
provided a comparison 
between the present
 
work and current practice.
 
 
A
ccurate calculation of local fan performance metrics su
ch as blade loading, pressure 
rise, and efficiency
 
were
 
obtained
. 
The fan inlet measurement profile greatly enhanced the 
understanding of the fan interaction to the flow distortion and provided a more complete 
explanation of the fan behavior. Secondary flo
wfield
 
formation due to the accelerated flow 
redistribution directly upstream of the fan created 
localized 
bulk co
-
 
and 
counter
-
 
rotating 
swirl regions
 
that were found to be correlated with localized 
fan performance
 
phenomena
.
 
It was observed that the effe
cts of the distortion on fan performance 
were
 
exaggerated 
if 
the
 
assumed 
fan 
inlet profiles
 
were
 
data taken only at the AIP. 
 
The reduction in fan performance with respect to undistorted inlet conditions is also 
explored, providing insight into how such di
stortions can be compared to
 
baseline 
conditions. 
The dissertation closes with several recommendations for improving distortion 
tolerant fan design in the context of experi
mental research and development.
 
 
 
Full Scale Transonic Fan Interaction with a 
Bounda
ry Layer Ingesting Total Pressure Distortion
 
 
Justin Mark Bailey
 
 
GENERAL AUDIENCE ABSTRACT
 
 
 
Future commercial transport aircraft will 
present
 
more aerodynamic architectures to 
accommodate stringent design goals for higher fuel efficiency, reduced cruise 
and taxi NOx 
emissions, and reduced noise. Airframe designs most likely to satisfy the first goal feature 
architectures that lead to the formation of non
-
uniform flow introduced to the engine 
through boundary layer ingesting (BLI) inlets, creating a differ
ent operational environment 
from which the engines were originally designed. The goal of this study was to explore the 
effects such non
-
uniform flow would have on the behavior and performance of a transonic 
fan in a full scale engine test environment. 
 
Thi
s dissertation presents an experimental study of the interaction between a full scale 
transonic fan and a total pressure distortion representative of a boundary layer ingesting 
serpentine inlet. A five
-
hole pneumatic probe was traversed directly in front o
f and behind 
a fan rotor to fully characterize the inlet and outlet fan profile. The distortion profile was 
also measured at the aerodynamic interface plane (AIP), which has historically been 
accepted as the inlet profile to the fan. This provided a compar
ison between the present 
work and current practice. 
 
Accurate calcu
lation of local fan performance metrics 
were obtained. The fan inlet 
measurement profile greatly enhanced the understanding of the fan interaction to the flow 
distortion and provided a more
 
complete explanation of the fan 
behavior.
 
It was observed 
that the effects of the distortion on fan performance were exaggerated if the assumed fan 
inlet profiles were data taken only at the AIP. 
 
The reduction in fan performance with respect to undistort
ed inlet conditions is also 
explored, providing insight into how such distortions can be compared to baseline 
conditions. The dissertation closes with several recommendations for improving distortion 
tolerant fan design in the context of experimental resea
rch and development. 
 
Acknowledgements
 
 
 
iv
 
 
Acknowledgements
 
This body of work was completed over a 3
-
year span, but the design and construction 
of the research rig began 2 years before then and involved the work of many smart, 
generous, and dedicated individuals. I would lik
e to acknowledge their contributions here. 
 
Professor Walter O'Brien: Thank you for giving me the opportunity to contribute to the 
incredible research environment you have at the TurboLab. I count myself fortunate that I 
was able to do exactly what I wante
d for my graduate work 

 
get my hands dirty working 
with real jet engines. I may not get any recognition from the University for accumulating 
over 200 engine operating hours, but I should. I enjoyed working with you over the years 
and would enjoy doing so 
again. 
 
My committee 

 
Professor Todd Lowe, 
Professor 
Wing Ng, 
Professor 
Clinton Dancey, 
and 
Professor 
Alfred Wicks: Thank you for your support over the course of my research. 
Your insights and recommendations for improving this dissertation have been grea
tly 
appreciated and I included them in the document as best I could. Your assistance has helped 
me create a more well
-
rounded body of work that I could not have completed otherwise. 
 
Dr. Tony Ferrar: We were never given an instruction manual on how to buil
d a 
distortion tolerant fan research rig but you and I did alright figuring it out ourselves. Thank 
you for your dedication, focus, and commitment to quality that created a new standard for 
experimental research at the TurboLab. I would not have been able 
to complete my research 
without your efforts in overcoming the difficulties of starting this research. I enjoyed my 
time working with you over the years and I hope you carry a sense of pride knowing the 
legacy you pioneered at the TurboLab, and the foundat
ion you created that every 
generation afterwards benefits from. 
 
Dr. Bill Schneck: Your contributions to my research are incalculable. You were with 
Tony and I from day one, always willing to help in the extensive debugging procedures 
associated with an ex
periment of his magnitude: instrumentation, motion control, 

only a phone call away. Thank you for your efforts and suggestions over the years that 
helped make this re
search possible. I would not have collected a single data point without 
you. 
 
Chris Collins: Thank you for your efforts involving the construction of my research rig. 
You and I worked together on building the majority of the rotator and tunnel sections, 
se
questered in the lab day after day. Thank you for stepping in as our lab technician when 
we were in need of one. We made some great parts despite our inexperience, and those 
parts survived hundreds of engine hours between Tony and myself. I am proud of the
 
work 
we accomplished and I hope you are as well.
 
Acknowledgements
 
 
 
v
 
 
Gregg Perley: Your expertise and innovative solutions have been instrumental in 
improving my experiment when I took over as the lead researcher on distortion tolerant fan 
design. I enjoyed our technical dis
cussions over the years, but more so the off topic debates 
on life, film, and literature. Thank you so much for your input and contributions in helping 
me complete my research. Working at the lab was more fun with you around and I want to 
thank you for bei
ng an awesome person. 
 
Dustin Frohnapfel: Thank your hard work 
and assistance 
over the years in helping me 
complete my research. It takes at least two people to see our work come to fruition and you 
were with me anytime I needed. Thank you for your help in
: experimental design, 
component assembly, probe calibration, instrumentation install, alignment, data 
processing, engine runs, and everything else associated with our respective works. There 
are so many facets to consider when making quality measurements 
in the environment we 
work in but we tackled each one of them together and finished with one of the most unique 
and capable research facilities in the world. 
 

research unless you
 
were there with Tony and I, forced to move our probe by hand. It was 
a nightmare and took weeks to get all the data we needed. Your improvements to the probe 
motion control allowed me to expand my test matrix and complete far more tests than 
would have be
en possible otherwise. Your work on the probe calibration rig also allowed 
Dustin and I to calibrate the probe quickly and accurately. I would probably still be 
collecting data without your contribution. Thank you.
 

port technician ever! Thank you for working 
with me for six years and making sure I had everything I needed to complete my research. 

our talks getting to know you dur
ing my time at VT.
 
Everyone else who contributed to my research in some way: Dr. William Copenhaver, 
Dr. William Cousins, Todd Pickering, Steven Steele, Kevin Hoopes, James Lucas, Bill 
Hook, Bill Bradley, Bill Bryant, Joe Hall, Ali Alhamaly, Ahmed Aly, Jes
se Dollberg, Dr. 
Chaitanya Halbe. Thank you for your contributions and collaborations. 
 
My Mother and Father: Thank you for providing me with the foundation I needed to 
accomplish this work. You were only a phone call or short drive away and I always had 
y
our love and support to fall back on if I needed. None of this would have been possible 
without your guidance. I love both of you will all my heart.
 
My Sisters: Thank you for your support over my longer than expected stay in Virginia. 
I love you and hope m
y work instills you with confidence, to know that you can accomplish 

 
Table of Contents
 
 
 
vi
 
 
Table of Contents
 
Page
 
Chapter 1
 
Introduction
 
................................
................................
................................
...
 
1
 
1.1
 
Literature Review
 
................................
................................
................................
......
 
2
 
1.1.1
 
Stability
 
................................
................................
................................
.........
 
2
 
1.1.2
 
Performance
 
................................
................................
................................
..
 
3
 
1.1.3
 
Fan
-
Distortion Interaction
 
................................
................................
.............
 
3
 
1.1.4
 
Summary
 
................................
................................
................................
.......
 
3
 
1.2
 
Research Objectives
 
................................
................................
................................
..
 
4
 
Chapter 2
 
Experimental Methodology
 
................................
................................
..........
 
5
 
2.1
 
Experimental Investigation
 
................................
................................
.......................
 
5
 
2.1.1
 
Experiment Stations
 
................................
................................
......................
 
6
 
2.1.1.1
 
Station 0 and 1
 
................................
................................
...........................
 
6
 
2.1.1.2
 
Station 2
................................
................................
................................
.....
 
6
 
2.1.1.3
 
Station 3
................................
................................
................................
.....
 
7
 
2.1.1.4
 
Station 4
................................
................................
................................
.....
 
8
 
2.1.1.5
 
Station 5
................................
................................
................................
.....
 
8
 
2.1.1.6
 
Summary
 
................................
................................
................................
...
 
8
 
2.1.2
 
Inlet Conditions
 
................................
................................
.............................
 
9
 
2.1.3
 
Atmospheric Conditions
 
................................
................................
.............
 
10
 
2.2
 
Five Hole Probe
 
................................
................................
................................
.......
 
11
 
2.2.1
 
Design
 
................................
................................
................................
.........
 
11
 
2.2.2
 
Calibration
................................
................................
................................
...
 
12
 
2.2.3
 
Installation
................................
................................
................................
...
 
13
 
2.3
 
Measurement Grids
 
................................
................................
................................
.
 
14
 
2.3.1
 
Station 1 and Station 
3
 
................................
................................
................
 
14
 
2.3.2
 
Station 4 and Station 5
 
................................
................................
................
 
15
 
2.4
 
Modified Turbofan Engine
 
................................
................................
......................
 
15
 
Table of Contents
 
 
 
vii
 
 
2.4.1
 
Blade Analysis
 
................................
................................
............................
 
16
 
2.4.1.1
 
Single
-
per
-
rev Excitation
 
................................
................................
........
 
17
 
2.4.1.2
 
Metal Angles
 
................................
................................
...........................
 
17
 
Chapter 3
 
Data Reduction
................................
................................
............................
 
19
 
3.1
 
Data Acquisition
 
................................
................................
................................
......
 
19
 
3.1.1
 
Pressure
 
................................
................................
................................
.......
 
19
 
3.1.2
 
Temperature
 
................................
................................
................................
 
19
 
3.1.3
 
Fan Speed
 
................................
................................
................................
....
 
20
 
3.2
 
Data Collection
 
................................
................................
................................
........
 
20
 
3.2.1
 
Sample time
 
................................
................................
................................
 
20
 
3.2.2
 
Test Replications
 
................................
................................
.........................
 
21
 
3.2.3
 
Pressure Calibration
 
................................
................................
....................
 
22
 
3.3
 
Inlet Conditions
 
................................
................................
................................
.......
 
22
 
3.3.1
 
Bellmouth Static Taps
 
................................
................................
.................
 
23
 
3.3.2
 
Mass Flow Rate Calculation
 
................................
................................
.......
 
23
 
3.3.3
 
Mass Flow Rate and Fan Speed Consistency
 
................................
.............
 
24
 
3.4
 
AIP
 
................................
................................
................................
..........................
 
25
 
3.4.1
 
Total Pr
essure Consistency
 
................................
................................
.........
 
25
 
3.4.2
 
Static Pressure Consistency
 
................................
................................
........
 
26
 
3.5
 
Five Hole Probe
 
................................
................................
................................
.......
 
26
 
3.5.1
 
Pressure and Flow Angularity
 
................................
................................
.....
 
26
 
3.5.2
 
Total Temperature
 
................................
................................
.......................
 
27
 
3.5.3
 
Static Temperature
 
................................
................................
......................
 
27
 
3.5.4
 
Component Velocities
 
................................
................................
.................
 
27
 
3.6
 
Fan Performance Parameters
 
................................
................................
...................
 
28
 
3.6.1
 
Incidence and Devia
tion
 
................................
................................
.............
 
28
 
3.6.2
 
Blade Loading
 
................................
................................
.............................
 
30
 
3.6.3
 
Pressure Rise
 
................................
................................
...............................
 
30
 
3.6.4
 
Efficie
ncy
 
................................
................................
................................
....
 
30
 
Table of Contents
 
 
 
viii
 
 
3.6.5
 
Lieblein Diffusion Factor
 
................................
................................
............
 
30
 
3.6.6
 
Mass Flux
 
................................
................................
................................
....
 
31
 
3.7
 
Co
mparison Metrics
 
................................
................................
................................
 
31
 
3.7.1
 
Corrected Conditions for Fan Performance
 
................................
................
 
31
 
3.7.2
 
Data Presentation: Relative Difference from Undistorte
d
 
..........................
 
32
 
3.8
 
Uncertainty
 
................................
................................
................................
..............
 
33
 
3.8.1
 
Measured Parameter
................................
................................
....................
 
33
 
3.8.2
 
Five Hole
 
Probe Parameters
................................
................................
........
 
34
 
3.8.3
 
Calculated Parameters
 
................................
................................
.................
 
35
 
Chapter 4
 
Results and Discussion
 
................................
................................
...............
 
37
 
4.1
 
Consistent Inlet Conditions
 
................................
................................
.....................
 
37
 
4.2
 
Undistorted Inlet Flow Conditions
 
................................
................................
..........
 
37
 
4.2.1
 
Station 3 
-
 
AIP Inlet Profi
le
 
................................
................................
........
 
37
 
4.2.2
 
Station 4 and Station 5 
-
 
Fan Inlet and Fan Outlet Conditions
 
...................
 
38
 
4.2.2.1
 
Total and Static Pressure
 
................................
................................
.........
 
38
 
4.2.2.2
 
Tangential and Radial Flow Angle
................................
..........................
 
39
 
4.2.2.3
 
Total Temperature
 
................................
................................
...................
 
40
 
4.2.2.4
 
Ve
locity
 
................................
................................
................................
...
 
41
 
4.2.2.5
 
Incidence and Deviation
 
................................
................................
..........
 
42
 
4.2.3
 
Fan Performance
 
................................
................................
.........................
 
43
 
4.2.3.1
 
Blade Loading 






 
and Pressure Rise
................................
............
 
43
 
4.2.3.2
 
Blade Loading Sensitivity
 
................................
................................
.......
 
44
 
4.2.3.3
 
Diffusion and Efficiency
 
................................
................................
.........
 
45
 
4.2.4
 
Summary
 
................................
................................
................................
.....
 
46
 
4.3
 
Distorted Inlet Flow Conditions
 
................................
................................
..............
 
46
 
4.3.1
 
AIP I
nlet Profile
 
................................
................................
..........................
 
47
 
4.3.2
 
Station 4 and 5 
-
 
Fan Inlet and Fan Outlet Conditions
 
...............................
 
48
 
4.3.2.1
 
Total Pressure
 
................................
................................
..........................
 
48
 
4.3.2.2
 
Static Pressure
 
................................
................................
.........................
 
49
 
Table of Contents
 
 
 
ix
 
 
4.3.2.3
 
Total Temperature
 
................................
................................
...................
 
50
 
4.3.2.4
 
Tangential Flow Angle
 
................................
................................
............
 
51
 
4.3.2.5
 
Radial Flow Angle
 
................................
................................
..................
 
52
 
4.3.2.6
 
Upstream Flow Redistribution
 
................................
................................
 
53
 
4
.3.2.7
 
Incidence and Deviation
 
................................
................................
..........
 
53
 
4.3.3
 
Fan Performance
 
................................
................................
.........................
 
54
 
4.3.3.1
 
Blade Loading (





 
Distribution
 
................................
.....................
 
54
 
4.3.3.2
 
Blade Loading Orbits
 
................................
................................
..............
 
55
 
4.3.3.3
 
Pressure Rise Distribution
 
................................
................................
.......
 
56
 
4.3.3.4
 
Pressure Ri
se Orbits
 
................................
................................
................
 
57
 
4.3.3.5
 
Diffusion and Efficiency
 
................................
................................
.........
 
58
 
4.3.3.6
 
Compressor Map
 
................................
................................
.....................
 
59
 
4.3.4
 
Summary
 
................................
................................
................................
.....
 
61
 
Chapter 5
 
Conclusions and Recommendations
 
................................
...........................
 
65
 
5.1
 
Summary
 
................................
................................
................................
.................
 
65
 
5.1.1
 
Full Flowfield Measurements
 
................................
................................
.....
 
65
 
5.1.2
 
Fan Inlet Profile
 
................................
................................
..........................
 
66
 
5.1.3
 
Distortion Transfer
 
................................
................................
......................
 
66
 
5.1.4
 
Distortion
-
Fan Interaction
 
................................
................................
...........
 
67
 
5.1.5
 
Fan Performance Ground Tests
 
................................
................................
..
 
67
 
5.1.6
 
Local 
Performance Calculations
 
................................
................................
.
 
67
 
5.1.7
 
Secondary Flow Impact
 
................................
................................
..............
 
68
 
5.2
 
Conclusions
 
................................
................................
................................
.............
 
69
 
5.3
 
Recommendations
 
................................
................................
................................
...
 
69
 
5.3.1
 
Distortion Screen Calibration
 
................................
................................
.....
 
70
 
5.3.2
 
Fan Performance Ground Testing
 
................................
...............................
 
70
 
5.3.3
 
Distortion Tolerant Fan Design
 
................................
................................
..
 
70
 
5.4
 
Closing Statements
 
................................
................................
................................
..
 
71
 
Bibliography
 
................................
................................
................................
.....................
 
72
 
Table of Contents
 
 
 
x
 
 
Appendix A
 
Five
-
Hole Probe Calibration
 
................................
................................
...
 
75
 
A.1
 
Five
-
Hole Interpolation Maps
 
................................
................................
.................
 
75
 
A.2
 
Mach Number Dependency
 
................................
................................
.....................
 
76
 
Appendix B
 
Uncertainty Propagation
 
................................
................................
..........
 
78
 
Appendix C
 
Five
-
Hole Probe Analysis
 
................................
................................
.......
 
80
 
C.1
 
Structural
 
................................
................................
................................
.................
 
80
 
C.2
 
Modal
 
................................
................................
................................
......................
 
81
 
C.3
 
Time Response
 
................................
................................
................................
........
 
81
 
Appendix D
 
Fan Blade Modal Analysis
 
................................
................................
......
 
83
 
Appendix E
 
Fan Blade Geometry
................................
................................
................
 
84
 
Appendix F
 
Design Work I
nput
 
................................
................................
..................
 
86
 
Appendix G
 
Multiple Speed Data
 
................................
................................
................
 
87
 
G.1
 
AIP
 
................................
................................
................................
..........................
 
87
 
G.2
 
Station 4
 
................................
................................
................................
...................
 
87
 
G.3
 
Station 5
 
................................
................................
................................
...................
 
89
 
G.4
 
Fan Performance
 
................................
................................
................................
.....
 
91
 
 
List of 
Tables
 
 
 
xi
 
 
List of Tables
 
Table
 
................................
................................
................................
..............................
 
Page
 
Table 2.1. Modified JT15D
-
1 engine design parameters.
 
................................
.................
 
16
 
Table 2.2. Engine performance parameters at distorted test conditions.
 
..........................
 
16
 
Table 3.1. Pressure transducer specifications.
 
................................
................................
..
 
19
 
Table 3.2. Five
-
hole parameter RMS uncertainty.
 
................................
............................
 
35
 
Table 3.3. Measured and calculated parameter RMS uncertainty values given as a percent 
of the nominal value.
................................
................................
................................
.........
 
36
 
Table 4.1. Global performance parameters for distorted and undistorted cases.
 
..............
 
62
 
Table 4.2. Fan performance parameter comparisons between corrected outlet conditions 
and unmeasured fan inlet.
 
................................
................................
................................
.
 
64
 
Table A.1. Nume
rical and experimental RMS measurement standard deviation.
 
............
 
77
 
Table E.1. JT15D
-
1 fan blade geometry.
 
................................
................................
..........
 
85
 
 
List of Figures
 
 
 
xii
 
 
List of Figures
 
Figure
 
................................
................................
................................
.............................
 
Page
 

-
and
-

(Right).
 
................................
................................
................................
................................
 
1
 
Figure 1.2. Example of distortion impact on compressor surge ma
rgin [7].
 
......................
 
2
 
Figure 2.1. Schematic of experimental components and configuration.
 
.............................
 
5
 
Figure 2.2. Bellmouth schematic (left) and exper
iment photograph (right).
 
......................
 
6
 
Figure 2.3. Schematic (left) and photograph (right) of the distortion screen used in the 
experiment.
................................
................................
................................
..........................
 
7
 
Figure 2.4. Distorted (left) and undistorted (right) total pressure profiles.
 
.........................
 
7
 
Figure 2.5. Schematics (left and center) and photograph (right) of the AIP rake.
..............
 
8
 
Figure 2.6. Schematic overview of the experiment.
 
................................
...........................
 
9
 
Figure 2.7. Experimental operation line for distorted and undistorted cases.
...................
 
10
 
Figure 2.8. Five
-
hole probe schematic (left), fan inlet (Station 4) installation (center), and 
fan outlet (Station 5) installation (right).
 
................................
................................
..........
 
11
 
Figure 2.9. Pr
essure coefficient map (left) and calibration rig schematic (right).
 
............
 
12
 
Figure 2.10. Probe installation schematic (left) and photograph (right).
 
..........................
 
14
 
Figure 2.11. Measurement grids for Station 1 (left) and Station 3 (right).
 
.......................
 
15
 
Figure 2.12. Measurement grids for Station 4 (left) and Station 5 (center) with fan blade 
geometr
y schematic (right).
 
................................
................................
..............................
 
15
 
Figure 2.13. P&WC JT15D
-
1 turbofan engine.
 
................................
................................
 
16
 
Figure 2.14. Campbell diagram (left) with accompanying modal
 
frequency results at design 
speed (top
-
right) and deformation (bottom
-
right).
 
................................
............................
 
17
 
Figure 2.15. Metal angles of fan blade (left) and scan result deconstruction (right).
 
.......
 
18
 
Figure 3.1. Optical once
-
per
-
rev post process (left) and resolved fan speed (right).
 
.......
 
20
 
Figure 3.2. Good (left) and poor (right) sample calibrated press
ure signal during 60 second 
data collection.
 
................................
................................
................................
..................
 
21
 
Figure 3.3. Static calibration (left) and uncertainty (right) curves of the pressure 
transducers.
 
................................
................................
................................
.......................
 
22
 
Figure 3.4. Bellmouth static pressure tap measurement consistency.
...............................
 
23
 
Figure 3.5. Discharge coefficient (left) and discharge coefficient uncertainty (right).
 
....
 
24
 
Figure 3.6. Corrected mass flow and fan speed measurement consistency.
 
.....................
 
25
 
Figure 3.7. AIP total pressure measurement variability.
 
................................
..................
 
25
 
Figure 3.8. AIP static pressure taps measurement variability.
 
................................
..........
 
26
 
Figure 3.9. Polar coordinate velocity components.
................................
...........................
 
28
 
Figure 3.10. Schematic of rotor velocity triangles (left), and incidence (i) and deviation (

) 
angles (right).
 
................................
................................
................................
....................
 
29
 
Figure 3.11. Schematic representing th
e components required for calculating 
circumferential displacement.
 
................................
................................
...........................
 
31
 
List of Figures
 
 
 
xiii
 
 
Figure 3.12. Schematic representing the components required in calculating radial 
displacement.
 
................................
................................
................................
....................
 
32
 
Figure 4.1. AIP total pressure contours for varying speed undistorted inlet conditions.
 
..
 
38
 
Figure 4.2. Undistorted total (left) and static (right) pressu
re for Station 4 and Station 5.
................................
................................
................................
................................
...........
 
39
 
Figure 4.3. Undistorted tangential (left) and radial (right) flow angles
 
............................
 
40
 
Figure 4.4. Undisto
rted total temperature span
-
wise distribution.
 
................................
....
 
41
 
Figure 4.5. Undistorted axial (left) and velocity magnitude (right) span wise distributions
................................
................................
................................
................................
...........
 
42
 
Figure 4.6. Undistorted incidence and deviation angle.
 
................................
....................
 
43
 
Figure 4.7. Undistorted blade loading (left) and pressure rise (right) blade span 
distributions.
................................
................................
................................
......................
 
44
 
Figure 4.8. Outlet absolute flow angles (

5
) and inlet relative flow angles (

4
).
 
..............
 
45
 
Figure 4.9. Undistorted diffusion (left) and efficiency (right) blade span distributions.
 
..
 
46
 

 
................
 
47
 
Figure 4.11. AIP static pressure results.
 
................................
................................
...........
 
48
 
Figure 4.12. Distorted inlet (left) and outlet (right) total pressure profiles for 70% CFS.
 
49
 
Figure 4.13. Distorted fan inlet (left) and outlet(right) static pressure profiles for 70% CFS
.
................................
................................
................................
................................
...........
 
50
 
Figure 4.14. Distorted fan inlet (left) and fan outlet (right) total temperature profiles for 
70% CFS.
 
................................
................................
................................
..........................
 
50
 
Figure 4.15. Distor
ted fan inlet (left) and fan outlet (right) tangential flow angles.
 
........
 
51
 
Figure 4.16. Distorted fan inlet (left) and fan outlet (right) radial flow angles for 70% CFS.
................................
................................
................................
................................
...........
 
52
 
Figure 4.17. Distorted vector (left) and streamline (right) plots of the Station 4 in
-
plane 
velocity field at 70% CFS.
 
................................
................................
................................
 
53
 
Figure 4.18. Changes in incidenc
e (left) and deviation (right) angles relative to undistorted 
inlet conditions at 70% CFS.
 
................................
................................
............................
 
54
 
Figure 4.19. Station 4 tangential flow angle (left) and fan blade loading (right) at 70% CFS.
................................
................................
................................
................................
...........
 
55
 
Figure 4.20. Blade loading as a function of inlet flow coefficient (left) and tangential inlet 

 
............
 
56
 
Figure 4.21. Station 4 total pressure (left) and total pressure rise RDFU (right) at 70% CFS.
................................
................................
................................
................................
...........
 
57
 
Figure 4.22. Total pressure rise as a function
 
of inlet flow coefficient (left) and tangential 

 
.........
 
58
 
Figure 4.23. Local Lieblein diffusion factor (left) and 
local efficiency (right) under 
distorted conditions at 70% CFS.
................................
................................
......................
 
59
 
Figure 4.24. Pressure rise 

 
mass flux characteristic, all radii (left) and radially averaged 
distorted conditions (right).
 
................................
................................
...............................
 
60
 
Figure 4.25. Simulated compressor map (left) and scaled model data for the JT15D
-
1 
(right).
 
................................
................................
................................
...............................
 
61
 
Figure 4.26. Distorted and undisto
rted performance parameters trends for core and bypass 
flow
 
................................
................................
................................
................................
...
 
63
 
List of Figures
 
 
 
xiv
 
 
Figure A.1. Pressure coefficients for the five
-
hole probe calibration with experimental data 
(blue).
 
................................
................................
................................
................................
 
75
 
Figure A.2. Analytical and experimental pressure coefficient variability.
 
.......................
 
77
 
Figure C.1. Five
-
hole probe structural analysis: experimental setup (left), 30
°
 
section CFD 
model (center), and structural analysis deformation (right).
 
................................
.............
 
80
 
Figure C.2. Five
-
hole probe modal analysis.
 
................................
................................
....
 
81
 
Figure
 
C.3. Time (left) and frequency (right) of the pressure measurements from the five
-
hole probe.
................................
................................
................................
.........................
 
82
 
Figure D.1. Results of blade modal analysis.
................................
................................
....
 
83
 
Figure E.1. Fan blade geometry deconstruction.
 
................................
..............................
 
84
 
Figure F.1. Undistorted work input design distribution.
 
................................
...................
 
86
 
Figur
e G.1. Relative AIP Data from undistorted all speeds investigated.
 
........................
 
87
 
Figure G.2. Relative Station 4 total pressure from undistorted all fan speeds investigated.
................................
................................
................................
................................
...........
 
87
 
Figure G.3. Relative Station 4 static pressure from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
88
 
Figure G.4. Station 4 tangential flow angle difference fro
m undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
88
 
Figure G.5. Station 4 radial flow angle difference from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
88
 
Figure G.6. Station 4 incidence angle difference from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
89
 
Figure G.7. Relative Station 5 total pressure from undistorted for 
all fan speeds 
investigated.
 
................................
................................
................................
......................
 
89
 
Figure G.8. Relative Station 5 static pressure from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
89
 
Figure G.9. Station 5 tangential flow angle difference from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
90
 
Figure G.10. Station 5 radial flow angle difference from undistorted for all fan sp
eeds 
investigated.
 
................................
................................
................................
......................
 
90
 
Figure G.11. Station 5 deviation angle difference from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
90
 
Figur
e G.12. Relative blade loading difference from undistorted for all fan speeds 
investigated.
 
................................
................................
................................
......................
 
91
 
Figure G.13. Pressure rise for all fan speeds investigated.
 
................................
...............
 
91
 
Figure G.14. Lieblein diffusion factor for all fan speeds investigated.
 
............................
 
91
 
Figure G.15. Efficiency for all fan speeds investigated.
 
................................
...................
 
92
 
................................
................................
................................
................................
...............
 
 
 
Nomenclature
 
 
 
xv
 
 
Nomenclature
 
 
 
Symbol
/Acronym
 
Description
 
Units
 
a
 
Speed of sound
 
m/s
 
A
 
Cross sectional area
 
in
2
 

 
Absolute flow angle
, Probe relative radial angle
 
deg
 
(
°)
 
AIP
 
Aerodynamic Interface Plane
 
 
AWOS
 
Automated Weather Observing System
 
 

 
Blade relative flow angle
 
Probe relative tangential angle
 
deg (
°)
 
BDC
 
Bottom Dead Center
 
 
BPF
 
Blade Passing Frequenc
y
 
Hz
 
c
 
Blade chord
 
in
 
C
 
Absolute v
elocity
 
m/s
 
CFS
 
Corrected Fan Speed
 
%
 
CMF
 
Corrected 
Mass flow
 
Rate
 
lbs/s
 
CP
 
Pressure Coefficient
 
 
Ø
 
Diameter
 
 
EO
 
Engine Order
 
 
FAA
 
Federal Aviation Administration
 
 
FI
 
Fan Inlet
 
 
FO
 
Fan Outlet
 
 


 
Altitud
e
 
m
 
LDF
 
Lieblein Diffusion Factor
 
 
M
 
Mach Number
 
 
M
air
 
Molecular Weight of air
 
g/mol
 


 
Mass flow
 
rate
 
lbs/sec
 

 
Dynamic viscosity
 
kg/(m

s)
 
NIST
 
National Institute of Standards and Technology
 
 
NWS
 
National Weather Service
 
 

 
Rotational 
velocity
 
rad/s
 


 
Local station pressure at altitude
 
in
-
Hg
 


 
Sea level pressure
 
in
-
Hg
 


 
Total pressure
 
psi
 


 
Static pressure
 
psi
 



 
Parameter
 
 
R
 
Specific gas constant for air
 
J/(kg

K)
 
RDFA
 
Relative Difference From Average
 
%
 
RD
FT
 
Relative Difference From Target
 
%
 
RDFU
 
Relative Difference From Undistorted
 
%
 
Nomenclature
 
 
 
xvi
 
 
RMS
 
Root Mean Square
 
%
 
RMS
D
 
Root Mean Square 
Difference
 
%
 
s
 
Blade spacing
 
in
 
TDC
 
Top Dead Center
 
 

 
Tangential flow direction                
Circumferential Position
 
deg
 
(
°)
 
U
 
Blade speed
 
m/s
 
W
 
Relative velocity
 
m/s
 
x
 
x
-
coordinate direction
 
in
 
y
 
y
-
coordinate direction
 
in
 
z
 
Axial flow direction
 
in
 
Chapter 1
 
-
 
Introduction
 
 
 
1
 
 
Chapter 1
 
Introduction
 
In order to meet stringent requirements for next generation commercial aviat
ion
,
 
aircraft
 
will 
need to 
empl
oy more aerodynamic designs to 1) reduce noise, 2) improve 
landing, takeoff, and cruise 
emissions, and
 
3) improve fuel consumption
 
[
1,2]. 
These 
designs 
are expected to
 

-
and
-

architectures
 
into a mo
re hybrid or blended wing body (HWB, BWB) 
design
 
that wi
ll 
satisfy aircraft operation improvement goals
. A comparison between modern and proposed 
airframe architectures is 
shown in 
Figure 
1
.
1
. 
 
 
 
Figure 
1
.
1
. Modern aircraft 

-
and
-

 
(Left) and future blend
ed wing design 
(Right
)
.
 
A common feature of these HWB airframes is the integration of the engines with the 
aircraft
 
fuselage
. This relocation changes t
he inlet flow environment as the engines are no 
longer mounted below the wings where uniform inlet flow is ingested for the majority of 
the flight envelope. Mounting the engines in close proximity to the airframe, or embedding 
them within the airframe
,
 
res
ult
s
 
in 
non
-
uniform
 
(distorted)
 
inlet flow caused by the 
airframe body or inlet ducts
. In the interest of performance and efficiency, modern 
commercial transport engines are designed to tolerate 
only small levels of distortion,
 
such 
as during takeoff or st
rong crosswind conditions. 
Integrated propulsion systems will 
experience distorted inflow over the entire flight envelope and their interaction with such 
conditions will need to be studied to determine how 
engine performance is affected, and to 
support the
 
design of distortion
-
tolerant engines
.
 
To determine how well engines can withstand such an environment, three criteria will 
need to be evaluated in great detail; aeromechanics, stability, and performance. 
Although 
studies have shown that engine operation 
penalties are found in each one of these, the 
overall benefit of highly integrated airframe and propulsion systems 
can 
justify the 
operational challenges introduced to the engines [
4
]
. 
 
Chapter 1
 
-
 
Introduction
 
 
 
2
 
 
1.1
 
Literature Review
 
S
tudies have been 
performed 
on the evaluation of eng
ines 
subject to
 
non
-
uniform
 
inlet 
flows 
produced
 
by highly integrated engine
-
airframe systems. 
The three metrics that must 
be most considered are; aeromechanics, stability, and performance. Researchers can find 
background into aeromechanical fatigue caused
 
by distorted inflow through earlier 
publications 
[5,
 
6
]. As this work focuses on the impact of stability and performance, the 
literature review will investigate previous work on thes
e subjects in detail. An important 
component of the integrated engine
-
air
frame problem is the issue
 
of fan
-
distortion 
interaction, in which 
a coupling
 
effect between the distortion or distortion producing device 
and the fan
 
exists
. As will be shown later, these components act together as a single system 
that must be consi
dered 
when designing distortion
-
producing devices 
within the context of
 
distortion tolerant fan research.  
 
1.1.1
 
Stability
 
Stability is a qualitative measure of how well an engine can handle 
off
-
design
 
conditions. It is measured quantitatively 
via
 
surge margin
 
which 
represents how much the 
compressor is allowed to be aerodynamically loaded in 
off
-
design
 
conditions before
 
compression breakdown and flow reversal (surging)
. Higher surge margin 
represents
 
greater stability at lower performance while lower surge margin 
all
ows for
 
higher 
performance at the cost of reduced stability. 
An example of how distorted inflow affects 
stability is shown in 
Figure 
1
.
2
. 
The reduction in performance for equivalent engine speeds 
is accompanied by a reduced 
operat
ion range for 
off
-
design
 
conditions
.
 
Further 
background
 
can be found in previous publications [8
-
12]. 
 
 
Figure 
1
.
2
. Example of distortion impact on compressor surge margin
 
[7]
.
 
Chapter 1
 
-
 
Introduction
 
 
 
3
 
 
1.1.2
 
P
erformance
 
At constant speed a compressor will produce a greater pressure rise in lower 
mass flow
 
regions
 
but only so long
 
as the flow over the blades remains attached
 
[3]
. 
Once the 
aerodynamic 
loading on the bl
ade is too great
,
 
premature separation will
 
occur and
 
can 
lead to stall and eventually surge. Fans are 
typically 
designed to input constant work to 
the 
inlet 
flow to obtain uniform outlet conditions, maximizing thrust. Distorted inflow 
creates
 
regions of h
igh and low blade loading 
resulting in
 
non
-
uniform
 
work input
 
to the flow
 
and 
non
-
uniform
 
outlet flow that reduces 
efficiency and 
thrust output. Thus
,
 
to achieve the 
same required thrust
 
as uniform inlet flow
 
the fan must input more work to the flow [3,13]
. 
This forces the fan to operate at a higher speed and reduced efficiency, requiring more fuel. 
Although the 
overall 
aircraft system will be more aerodynamic and require less engine 
thrust, the 
non
-
uniform
 
fan outlet conditions must be studied to optimize 
flow conditions 
and maximize fan operating efficiency
 
and performance
 
while minimizing
 
fuel
 
consumption.
 
1.1.3
 
Fan
-
Distortion Interaction
 
Since the fan is moving its operating point along a constant speed line 
when 
subjected 
to distortion, the pressure rise vari
es for the flowfield. In the case of total pressure distortion 
profiles 
there
 
exist areas of high and low total pressure. The regions of low total pressure 
contain lower 
mass flow
 
and move the 
local 
operating point of the fan to a higher pressure 
rise. Thi
s resultant higher pressure
 
rise reduces the static pressure in that region upstream 
of the fan and drives a 
mass flow
 
redistribution 
from the AIP to the fan face
. Regions of 
higher 
mass flow
 
will 
migrate
 
towards regions of lower 
mass flow
 
and trend
 
toward
 
uniformity [14, 17]. 
 
 
The 
mass flow
 
redistribution results in secondary flow development of considerable 
magnitude [15, 16]. These secondary flow characteristics interact with the fan in a different 
manner than just the reduced 
mass flow
,
 
resulting in a 
more complex inlet 
profile
. 
Th
is
 
redist
ribution is found to be dependent
 
on the slope of the compressor characteristic [18].  
The secondary flowfield can be separated into radial and tangential components, where the 
tangential flow 
strongly
 
affects the 
bla
de turning, blade loading, and efficiency of the fan
. 
Although the total pressure distortion will not vary significantly between the distortion 
measurement plane and the fan face, the static pressure distortion will increase 
exponentially. The 
distortions 
in 
static and total pressure, as well as the radial and 
tangential flow angle
s,
 
have been found to be 
attenuated through the fan [16,19].
 
 
1.1.4
 
Summary
 
D
istortion tolerant fan
 
design
 
has been an ongoing area of research since the 
introduction of jet propulsion.
 
As the complexity of propulsion systems increase to meet 
the demands of improved fuel economy, durability, o
peration range, and performance, 
the 
accuracy and research capability must also improve to better design and validate 
Chapter 1
 
-
 
Introduction
 
 
 
4
 
 
components in experimental se
ttings. 
Readers
 
can find further information 
on 
distortion 
tolerant research and 
its 
evolution through previous publications [4, 13, 20]. 
 
1.2
 
Research Objectives
 
The previous research discussed has provided engine designers with valuable expertise 
in developi
ng 
more 
distortion
-
tolerant fan
s
. 
However,
 
many of the studies limit the 
complexity of the profiles to simple geometries. While this is fundamental in developing 
an understanding of fan response to 
non
-
uniform inlet flow, 
BWB and HWB 
airframes 
produce spec
ific distortion patterns that are more complicated
 
and
 
research is required to
 
understand how these 
specific profiles 
affect fan performance. 
 
 
The work presented here studies a total pressure distortion profile 
representative of 
a 
boundary layer ingest
ing
 
serpentine 
inlet
. The inlet of interest
 
that
 
produce
s
 
a combined 
total pressure and flow angularity distortion, but the components must be studied separately
 
to determine their contribution to the overall fan operation. What is unique about this work 
is t
he usage of a full size turbofan engine operating at 60%, 70%, and 80% design speed, 
where 80% corresponds to transonic tip velocities. Both fan inlet and fan outlet flowfields 
were measured using a calibrated five
-
hole pressure probe 
within 
0.5
 
blade chor
d
s
 
of the 
fan leading and trailing edges
 
to obtain all the 
flow
 
parameters necessary to characterize 
fan performance. 
 
 
This work will study the fan inter
action to this distortion profile in detail and will 
provide researchers and engine designers with a g
reater understanding of fan operation in 
the presence of total pressure distortion. 
In addition, the advanced 
measurement 
capabilities and techniques developed in this study 
can
 
aid
 
experimental researche
r
s in 
obtaining accurate and relevant results.
 
Final
ly, t
esting on a full scale engine
 
also provides 
insight into how the fan 
responds
 
to distortion on a system level, as isolated electric driven 
fan designs lack the 
low pressure turbine 
operation matching that 
exists in 
full
-
scale 
engine 
experiments
.
 
 
 
The
 
overall goal in making next generation aircraft architecture flight worthy is to 
design engines that are capable of handling the constant distortion 
produced by 
BWB and 
HWB 
airframes. 
This can 
be done via two ways; 1) 
increasing the operability of the eng
ines 
to such an extent that the distortion can be tolerated but at the cost of severely red
uced 
performance and efficiency, or 2) creat
ing
 
tailor
-
designed engines that are built to operate 
at maximum efficiency and performance at the distorted cruise condi
tions
 
with a large 
enough margin to tolerate 
off
-
design
 
conditions. The final product will likely be 
a
 
compromise of the two. 
 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
5
 
 
Chapter 2
 
Experimental 
Methodology
 
Obtaining 
accurate measurements on a full scale turbofan engine requires 
many 
custom
,
 
tailor
-
fit compo
nents to be designed
, built, calibrated, and installed. A basic 
overview of the equipment
 
and measurement protocol
 
utilized will be discussed here
. 
I
nterested readers are 
referred
 
to 
related
 
work that examines these components in further 
detail [20]. 
 
2.1
 
Expe
rimental Investigation
 
Full
-
scale experimental total pressure distortion ground tests were conducted on a 
modified Pratt & Whitney Canada JT15D
-
1 turbofan engine at the Virginia Tech 
Turbomachinery and Propulsion Laboratory in Blacksburg, Virginia. A schem
atic of the 
experimental setup is shown in 
Figure 
2
.
1
. 
 
 
Figure 
2
.
1
. Schematic of experimental components and configuration
.
 
The area
s of interest
 
are
 
divided
 
into 
six
 
stations
:
 
 
Station 0 
-
 
 
Ca
librated bellmouth inlet that funnels airflow from outside 
the test cell 
and 
into the experimental rig. 
 
Station 1 
-
 
 
Bellmouth measurement plane
 
with f
our static pressure taps equally 
spaced circumferentially around the tunnel. 
 
Station 2 
-
 
 
D
istortion screen
 
that 
produces
 
a total
 
pressure distortion 
representative 
of a 
boundary layer ingesting (BLI) serpentine inlet. 
 
Station 3 
-
 
 
AIP 
rake case 
with
 
60 local total pressure probes and six equally spaced 
circumferential 
wall 
static pressure taps. 
 
Station 4 
-
 
 
Fan inlet measurement plane located approximatel
y 1 inch upstream of the 
fan leading edge
 
(~40% chord).
 
Station 5 
-
 
 

of the fan trailing edge (~23% chord). 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
6
 
 
2.1.1
 
Experiment 
Station
s
 
Each of 
the
 
s
tation
s identified in 
Figure 
2
.
1
 
will be 
discussed
 
in further detail in the 
following subsections.
 
For further specifications on the installation procedures, 
the reader 
is referred to 
[20]
 
and 
[25]
. 
 
2.1.1.1
 
Station
 
0
 
and 1
 
The bellmouth 
inlet 
is composed 
of spun 
3003 Al
uminum
 
to obtain 
an ASME
 
standard 
curve profile. 
The bellmouth was calibrated at the TurboLab to obtain the 
bellmouth 

four 
static pressures 
obtained at 
Station
 
1
 
along with the discharge coefficient
, the 
mass flow
 
through the tunnel 
can be calculated. 
Station
 
1 is located approximately 1.5
Ø
 
(1.5 fan diameters) downstream 
of the bellmouth throat
 
[21]
. The bellmouth inlet sectio
n is comprised of the bellmouth 
inlet and a 2
Ø straight section immediately downstream. 
Once assembled these sections 
were never disassembled
, thus the 
calibration curve applies to the assembly
 
and not just 
the 
bellmouth. 
The bellmouth and accompanying dow
nstream sections 
were
 
treated as a 
single
 
assembly
. 
A schematic of the bellmouth and accompanying experiment photograph 
is shown in 
Figure 
2
.
2
 
 
 
Figure 
2
.
2
. Bellmouth schematic (lef
t) and experiment photograph (right).
 
2.1.1.2
 
Station
 
2
 
The 
distortion 
screen 
used in this experiment
 
was
 
designed to 
produce
 
a total pressure 
distortion 
representative of
 
a boundary layer ingesting (BLI) serpentine inlet.
 
The profile 
was modeled after the NASA In
let
-
A duct
,
 
from which the data can be 
referenced in 
[33]
. 
The losses for this profile are considered severe in the aerospace committee
 
[23]
 
and 
provided sufficient 
non
-
uniform
ities to 
significantly
 
alter 
fan
 
performance
. 
A schematic 
and 
photograph
 
of the distortion screen is shown in 
Figure 
2
.
3
. The distortion screen is 
located approximately 2.5
Ø downstream of the bellmouth throat. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
7
 
 
 
Figure 
2
.
3
. Schematic (left) and photograph (right) of 
the 
distortion screen used in 
the 
experiment.
 
Figure
 
2
.
4
 
provides
 
the profile 
shapes
 
for both the distorted and undistorted cases. 
The 
distorted profile has a signifi
cant total pressure reduction on the bottom half of the profile 
while the undistorted profile is uniform. 
The distorted profile under investigation has a 
high circumferential distortion and a mild radial distortion. The distortion intensity 
(maximum pressu
re deficit) for these tests is high and considered an extreme case
 
[33]
, 
[23]
. 
 
The distortion screen is bolted to the back of slew
-
ring bearing that allows 360
-
degree 
rotation such that the flowfie
ld can be mapped with 
circumferentially 
s
tation
ary 
instrumentation. Details regarding screen construction and implementation 
for both 
distorted and undistorted configurations 
can be found in previous publications 
[23]
.
 
 
Figure
 
2
.
4
. Distorted
 
(left)
 
and undistorted
 
(right)
 
total pressure profiles.
 
2.1.1.3
 
Station
 
3
 
The 
AIP 
is classically defined as the interface between the inlet and engine. This 
measurement 
s
tation
 
characterizes the disto
rtion profile 
as established by the Aerospace 
Recommended Practices 
albeit 
with a 50% increase (60 measurements as opposed to 40) 
in
 
total pressure measurement 
circumferential density
 
[24]
. 
In addition to the 60 probe total 
pre
ssure measurements there are six static pressure ports 
taps equally
 
spaced 
circumferentially around the inner wall. 
All measurements are assembled together
 
with
in 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
8
 
 
the AIP ra
k
e case and represent the AIP measurement plane.
 
Schematics and photographs 
of the 
AIP rake are shown in 
Figure 
2
.
5
. 
 
 
 
Figure 
2
.
5
. Schematics (left and center) and photograph
 
(right)
 
of the AIP rake.
 
2.1.1.4
 
Station
 
4
 
The fan inlet measurement plane 
was
 
located approxima
tely 1 inch upstream of the fan 
leading edge
, 
approximately 
0
.05
Ø
 
and 
40% chord. 
Since the 
fan
 
blade employs a 
backward lean and varying 
span
-
wise
 
chord length
,
 
a 
uniform
 
measurement plane with 
respect to the blade was not 
obtained
. The variation in axial 
measurement location with 
respect to the fan blade is 
0
.02
Ø and 
deemed
 
negligible
 
for the purposes of this study
. 
 
A pneumatic 
five
-
hole 
probe 
was radially traversed in this location by means of 
an 
electronically controlled traversing 
probe mount
. The fan 
case insertion point was sealed 
with 
a compression fitting that utilized a Teflon 
ferrule, allowing linear motion while 
providing a pressure seal
 
and damping
. 
The probe mount was designed
 
to ensure 
a 
normal 
traversing 
plane with respect to the fan case. 
 
T
he five
-
hole probe 
measures
 
five pressure
s
 
and one total temperature simultaneously. 
The five pressure
 
measurements
 
were
 
later post processed to obtain total pressure, static 
pressure, and flow angularity. 
 
2.1.1.5
 
Station
 
5
 
The fan outlet measurement plane was lo

the fan trailing edge, approximately 
0
.02
Ø and 
23% chord. 
Again, 
the variable 
span
-
wise
 
measurement 
distance from the tapered fan blade
 
trailing edge 
was considered negligible
. 
 
The five
-
hole probe was traversed in 
this location 
as was done at Station 4
, although 
constraints between the probe and engine
 
geometry
 
prevented near blade root 
measurements
 
as the reinforcement tube of the probe interfered with the splitter
 
(fixed 

een the bypass and core)
. 
The fan bypass was 
fully traversed and approximately 
20% of the core
 
outer flow path
 
was obtained.
 
A custom 
insertion mount was created for the outlet plane due to the axial
-
wise variation in 
fan case 
geometry
 
(fan case featured a
 
no
zzl
e
 
structure 

 
larger diameter at inlet, smaller diameter 
at the outlet)
 
and p
lanar traversing was 
verified during installation. 
 
2.1.1.6
 
Summary
 
An overall schematic of the experiment is shown in 
Figure 
2
.
6
. 
Station
s 4 and 5
 
were 
ac
cessed via a radial traverse that allowed motion during the experiment such that several 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
9
 
 
radial locations could be measured at each rotation of the distortion screen. 
 
The traverse 
and its components 
will be shown in greate
r detail in 
succeeding
 
sections.
 
 
Figure 
2
.
6
. Schematic overview of the experiment.
 
2.1.2
 
Inlet Conditions
 
Three separate inlet conditions were investigated 
for the distorted 
profile cases:
 
60, 70, 
and 80% CFS
 
(Corrected Fan Speed)
, with t
ransoni
c 
tip 
conditions 
obtained
 
at 
80% CFS. 
The distortion screen 
losses 
followed a
 
second order relationship with fan speed such that 
losses were 
relatively 
more severe at higher fan speeds. 
The

corresponded 
to three CMF

 
(Corrected Mass Flows)
 
b
ased on the engine operation line for 
the distorted 
screen
 
configuration as shown in 
Figure 
2
.
7
.
 
These three CFS

 
along with their 
corresponding CMF

 
were duplicated for the undistorted cases
 
and provided
 
details on 
how the dist
orted profiles interact with the fan 
in comparison to 
equivalent undistorted CFS 
and CMF conditions. 
Notice how higher fan speeds must be reached to obtain equivalent 
mass
 
flows: at 60% CFS only 3 lbs
/s
 
separate the distorted
 
case
 
from the undistorted 
equi
valent mass
 
flow
 
case
, while at 80% CFS
 
the discrepancy has grown to 7 lbs/s, a 
consequence of the increased relative losses through the screen for increasing fan speed. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
10
 
 
 
Figure 
2
.
7
. Experimental operation 
line for distorted and undistorted cases.
 
2.1.3
 
Atmospheric Conditions
 
Atmospheric conditions were an important consideration as the duration and 
complexity of data collection
, 
experiment 
construction
 
and
 
assembly, and engine run 
scheduling forced
 
tests to be co
mple
ted over a span of 
eight months.
 
This required 
the 
data 
to be normalized by atmospheric conditions at the time of collection such that 
standard 
conditions 
could be calculated and between
-
test comparisons 
could be made. 
 
Local atmospheric conditions wer
e obtained from the weather 
s
tation
 
located at the 
Virginia Tech Montgomery Executive Airport. The 
s
tation
 
was located less than one 
quarter mile from the test cell and contained NIST traceable
 
instrumentation monitored by 
the FAA. The atmospheric data
 
was
 
recorded every 20 minutes
 
and
 
available 
online 
from 
the NWS
 
(National Weather Service) in the form of lookup tables
 
that were interpolated to 
the time of collection.
 
The data provided had been corrected to sea level conditions, which 
was uncorrected 
durin
g post
-
processing 
to obtain the local atmospheric pressure at altitude 
using Equation 
(
2
.
1
)
, where 


 
is the local 
atmospheric 
station pressure at altitude
 
in in
-
Hg
, 


 
is the 
local atmospheric station pressure 
corrected 
to 
sea level 
in in
-
Hg
, and 


 
is 
the
 
station altitude
 
in meters
 
(649.7)
. 
 
 




















 
(
2
.
1
)
 
 
Pressure and temperature data was standardized using
 
Equation 
(
2
.
2
)
 
w
here Y 
represents 
a 
pressure or temperature
 
measurement
. 
A
tmospheric pressure (P
at
m
)
 
and 
temperature (T
atm
) are the local 
station 
conditions 
at alt
itude 
at the time of the data 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
11
 
 
collection
. S
tandard pressure (P
std
) and temperature (T
std
) are 14.696 psi and 59
°
F, 
respectively. 
The resultant parameter, Y
cor
, represents the corrected 
pressu
re or temperature
 
accounting for local atmospheric 
conditions at the time the measurement was recorded.
 
 
 









 
(
2
.
2
)
 
2.2
 
Five Hole Probe
 
The pneumatic five
-
hole probe utilized was a United Sensor

 
DAT design,
 
featuring 
five pressure ports housed 
with
in a prism head 
architecture
 
with a K
-
type thermocouple 
installed
 
below the ports. 
A schematic of the probe geometry as well as installation 
photographs is shown in 
Figur
e 
2
.
8
. 
A reinforcement tube is brazed to the sensing head to 
provide 
structural 
support. 
 
 
Figur
e 
2
.
8
. Five
-
hole probe schematic (left), fan inlet
 
(S
tation 4
)
 
installation (center), and 
fan outlet
 
(S
tation 5
)
 
installation (right).
 
2.2.1
 
Design
 
The probe was constructed 
of
 
316 stainless steel
, with
 
the overall
 
geometry 
parametrized to minimize deflection and deformation
 
during engine operation
. The 
resultant analysis
, outlined in 
Appendix C
,
 
found that a sensing head diameter of
 

a
 
reinforcement tube diameter 
of 

 
to be 
sufficient
ly rigid while minimizing the 
intrusiveness of the measurement
. 
A modal analysis
 
of the geometry
 
following the methods 
outlined in
 
Appendix C
 
provided
 
the
 
blade passing frequencies to avoid
 
during
 
testing
.
 
Appendix C
 
also includes an analysis of the time response of the probe. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
12
 
 
2.2.2
 
Calibration
 
Calibration of the probe w
as performed utilizing a custom
-
built system designed at 
Virginia Tech specifically for multi
-
hole probes. 
Figure 
2
.
9
-
right shows a schematic of the 
rig where air from a blowdown compressor is passed over the probe while it is rotated in 
both pitch and yaw to obtain pressure coefficient maps (
Figure 
2
.
9
-
left) for varying 
flow 
par
ameters. The probe was calibrated on a (
-
40
°
) 

 
(+40
°
) grid in 2.5
° increments, with 1° 
increments within a 10° x 10° window. Three inlet Mach numbers were evaluated (0.4, 0.5, 
0.6) with 2
-
5 replications completed depending on 
the 
grid and Mach number 
test
ed
. 
It was 
found that the Mach number had negligible effect on the calibration maps and was not used 
in the calibration map construction
, details of which can be found in
 
Appendix A
. 
All 
speeds were averaged to obtain the final
 
calibration 
map. 
A more d
e
tailed analysis of the 
calibration matrix and final interpolation map used can be found in 
Appendix A
. 
Further 
details on the setup and completion of the calibration can be found in previous publicati
ons 
[25]
. 
 
 
Figure 
2
.
9
. Pressure coefficient map (left) 
and calibration rig schematic
 
(right).
 
C
oefficients were calculated directly from the pressures measured at each pitch, ya
w, 
and
 
Mach number tested and took the form of Equation 
(
2
.
3
)
, with the coefficients only a 
function of the p
itch and yaw angle of the probe
. 
 
 








 
(
2
.
3
)
 
The pitch an
d yaw flow angle pressure coefficients were 
calculated
 
using Equations 
(
2
.
4
)
 
and 
(
2
.
5
)
, respectively, 
where the variable 


 
is a pseudo average
 
pressure
 
that is 
calculated using Equation
 
(
2
.
6
)
. 
 
 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
13
 
 
 
 
 















 
(
2
.
4
)
 
 














 
(
2
.
5
)
 
 












 
(
2
.
6
)
 
The t
otal and 
s
tatic 
p
ressure
 
coefficients are 
calculated
 
using Equations 
(
2
.
7
)
 
and 
(
2
.
8
)
, 
respectively. 
P
T
 
is the reference total pressure
 
in the calibration tunnel and P
S
 
is the 
atmospheric static pr
essure at the tunnel discharge. 
 
 














 
(
2
.
7
)
 
 














 
(
2
.
8
)
 
2.2.3
 
Installation
 
Figure 
2
.
10
 
shows a sc
hematic and photograph of the installation setup. 
The probe was 
inserted into the fan case by way of a radial traverse that 
was
 
electronically controlled 
outside of the 
test 
cell such that probe motion could occur during engine operation. The 
probe was fix
ed inside an alignment table that was used to 
null
 
the probe 
in the yaw 
direction 
inside the engine
 
prior to testing
. The alignment table was attached to the radial 


Radial positioning was verified 
during testing 
via the 
camera system shown in 
Figure 
2
.
10
-
right. 
Further details on the traverse system can be 
found in previous publications 
[25]
. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
14
 
 
 
Figure 
2
.
10
. Probe installation 
schematic (left) and photograph (right)
.
 
2.3
 
Measurement 
G
rid
s
 
2.3.1
 
Station
 
1 and 
Station 3
 
The static pressure taps at the 
Station 1 
bellmouth measurement plane
 
were l
ocated at 
45 degree offset
s
 
from TDC
 
and BDC
. 
The f
our taps were equally spaced circumferentially 
around the tunnel and are shown schematically in 
Figure 
2
.
11
-
left
 
in blue markers 
overlaid 
on the uniform 
bellmouth 
inlet profile. 
 
The AIP at 
Station 3
 
has 60 total pressure measurements and 
six
 
static pressure taps. 
These are shown schematically in 
Figure 
2
.
11
-
right, with the blue markers corresponding 
to the static taps and the black markers corresponding 
t
o the total pressure measurements. 
Only six static pressure taps were installed
,
 
equally spaced circumferentially around the 
tunnel wall starting at 
a 
30 degree offset from TDC. 
The AIP rake 
produced
 
5 radial 
measurements at the geometric ce
nters of equal 
area rings in 30
° 
increments around the 
annulus. 
The distortion screen was rotated in 15
° increments, which effectively doubled 
the circumferential resolution of both the static taps and rake total pressure measurements.
 
The screen rotation also significan
tly increased the 
number of 
repeated me
asurements 
which reduced the installation bias 
for both the probes and taps
. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
15
 
 
 
Figure 
2
.
11
. Measurement grids for 
Station
 
1
 
(left) and 
Station 
3
 
(right).
 
2.3.2
 
Station
 
4
 
and 
Station 
5
 
Due to the
 
varying
 
geometry
 
of the 
fan 
inlet and outlet 
planes for 
Station
 
4 and 
Station 
5
, respectively, obtaining equivalent 
measurement grid
s was
 
not feasible. 
Station
 
5
 
located 
the five
-
hole probe closer
 
to the fan
 
blades and since the blades
 
were forward leaning from 
root to tip
, the 
leading
 
edge probe to 
trailing edge 
blade clearance decreased with plunge 
depth. 
Station
 
5
 
also placed the probe close
r
 
to the splitter, creating an interference 
with
 
the probe reinforcement tube 
for 
large plunge
 
depths. Despite these issues
,
 
sufficient radial 
measurement density was obtained 
for Station 4 (
Figure 
2
.
12
-
left) and Station 5 (
Figure 
2
.
12
-
center). S
ignificant blade and engine geometry is 
identified
 
i
n 
Figure 
2
.
12
-
right. 
 
 
Figure 
2
.
12
. Measurement grids for 
Station
 
4
 
(left)
 
and
 
Station
 
5
 
(center)
 
with
 
fan 
blade 
geometry schematic (right).
 
2.4
 
Modified Turbofan Engine
 
All tests were 
conducted using a modified Pratt and Whitney Canada JT15D
-
1 turbofan 
engine
, a picture of which is shown in 
Figure 
2
.
13
. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
16
 
 
 
Figure 
2
.
13
. P&WC JT15D
-
1 turbofan engine.
 
At design condi
tions the 
transonic 
axi
-
centrifugal 
twin spool engine produces 2,200 
lbs
f
 
thrust
 
and further 
design parameters
 
are summarized in 
Table 
2
.
1
. 
Due to the 
strong 
distorted inlet conditions utilized
,
 
testing was comp
leted at sub design
 
conditions and the 
reduced
 
engine performance is shown in 
Table 
2
.
2
 
for all three distorted speeds tested
. 
 
Table 
2
.
1
. 
Modified JT15D
-
1 engine design parameters.
 
Fan Speed
 
16,000 RPM
 
Core Speed
 
32,000 RPM
 
Mass flow
 
Rate
 
73.1 lbs
m
/s
 
Fan Pressure Ratio
 
1.5
 
Bypass Ratio
 
3.3
 
Thrust
 
2,200 lbs
f
 
Fan Diameter
 
21 in
 
Number of Blades
 
28
 
Span Averaged Chord Length
 
2.74 in
 
Blade Tip Mach Number
 
1.3
 
Table 
2
.
2
. 
Engine performance parameters at distorted test conditions.
 
Corrected 
Fan Speed
 
60%
 
70%
 
80%
 
Mass flow
 
Rate
 
[lbs
m
/s]
 
39.8
 
46.9
 
54.6
 
Fan Pressure Ratio
 
1.18
 
1.26
 
1.35
 
Blade Tip Mach Number
 
0.78
 
0.91
 
1.04
 
2.4.1
 
Blade Analysis
 
For good reason
,
 
engine manufacturers are concerned about the level of distortion 
ingested by the engine
s
. While fatigue is of concern 
in
 
regards to durability, a primary 
concern is excitation. 
Fan blades passing through pockets of low pressure distortion will 
experience
 
an excitation
 
as the
y traverse regions of low and high pressure 
during each 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
17
 
 
rotation, placing
 
significant additional wear on the fan blade structure. 
This effect is further 
exacerbated for multiple
-
per
-
rev distortions. 
 
2.4.1.1
 
Single
-
per
-
rev 
Excitation
 
The disto
rtion pattern under study for this work was 
an intense, single
-
per
-
rev 
distortion, thus 
caution must be exercised to avoid 
exciting
 
blade resonating
 
frequencies
. 
Shown in 
Figure 
2
.
14
 
are the results of a modal analysis performed o
n the blade geometry 
to obtain the excitation frequencies
, which is detailed in
 
Appendix D
. These frequencies 
were then compared to the engine order excitations via a Campbell diagram. Only the 
second engine order poses a serio
us concern as it intersects with the 1
st
 
mode frequency of 
the blade at approximately 
61
% fan speed
, indicated in 
Figure 
2
.
14
-
left
. Care was taken 
during testing to ensure the engine avoided 
extensive operation at this condition, 
which 
was accomplished by reserving testing to 
cooler
 
days when the uncorrected fan speed 
would be 
lower
 
than 60% to match corrected conditions. 
 
 
Figure 
2
.
14
. Cam
p
bell diagram (left) wit
h accompanying modal
 
frequency
 
results
 
at 
design speed
 
(
top
-
right)
 
and deformation (bottom
-
right)
.
 
2.4.1.2
 
Metal Angles
 
To gain an understanding of the undistorted operating conditions of the engine it was 
necessary to obtain the metal angles of the fan blade. This was done by perfor
ming a 
three 
dimensional 
scan of the blade
, breaking up the span into discrete sections, and then 
constructing the camber line at each section to obtain the leading and trailing edge angles. 
Figure 
2
.
15
 
summarizes the process and 
results of this effort. 
A more detailed explanation 
of the process used, as well as a data table of the reversed engineered blade geometry, can 
be found 
in
 
Appendix E
. 
 
Chapter 2
 
-
 
Experimental 
Methodology
 
 
 
18
 
 
 
Figure 
2
.
15
. Metal angles of fan blade (left) and scan result deconstruction (right).
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
19
 
 
Chapter 3
 
Data Reduction
 
Due to the considerable amount of data required 
to complete
 
this study
,
 
special care
 
had to be taken
 
to ensure adequate data quality. This 
was
 
accomplished by ver
ifying the 
experimental setup and transducer performance 
prior to and following each experiment
. 
For these experiments, setup
,
 
inspection
,
 
and transducer calibration was performed before 
and after each test. Specifics on the methods utilized are described 
in the following 
sections. 
 
3.1
 
Data Acquisition
 
A National Instruments

 
(NI)
 
d
igital data acquisition system measured and recorded 
all d
ata obtained in the experiment, which included; voltage
-
output pressure transducers, a 
K
-
type thermocouple, and a voltage
-
o
utput optical sensor used 
for 
accurate
 
fan speed 
measurement
s
.
 
 
3.1.1
 
Pressure
 
Two separate pressure 
transducer 
models were utilized in this experiment to account 
for the varying pressure ranges recorded. 
The manufacturer specifics for 
these transducers
 
are
 
summ
arized in
 
Table 
3
.
1
. 
 
Table 
3
.
1
. 
Pressure transducer specifications.
 
 
Omega
®
 
PX139
 
Scanivalve
®
 
ZOC17
 
Full Scale Pressure Range
 
±
 
5
 
psig
 
± 15 psig
 
Accuracy
 
~ 
± 
0.32% Full Scale
 
± 0.0
8% Full Scale
 
 
The ZOC17 model was used to measure the five
-
hole probe pressures as well as three 
of the six AIP static pressures. 
All 60 total pressure measurements at the AIP, the remaining 
three static pressures at the AIP
,
 
and the four bellmouth stati
c pressures
 
were measured 

-
6251 
data acquisition card mounted inside a NI PXI
-
1044 chassis. 
Shielded wiring was used for 
both power and signal and tied into a common ground. T
he ZOC17 and PX139 had 
independent Agilent E3610A DC Power supplies
. The pressure transducer
 
outputs 
were 
sampled at 128 Hz. 
 
3.1.2
 
Temperature
 
A single temperature measurement was recorded via the K
-
type thermocouple 
embedded in the five
-
hole probe. The thermoc
ouple cabling terminated at a NI SCXI
-
1301 
isothermal block which connected to an NI SCXI
-
1102 thermocouple signal conditioning 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
20
 
 
unit and 
digitized in a
 
NI 6251 data acquisition card mounted in the chassis. 
The 
thermocouple 
output 
was sampled at 32 Hz. 
 
3.1.3
 
Fan
 
Speed
 
F
an speed 
was measured 
using an optical once
-
per
-
rev sensor system. A mirror
-
polished flat washer was installed on the 
tip of the engine 
spinner cone and the 
laser/encoder was mounted inside the tunnel wall of the experiment. A NI PXI
-
6251 data 
acqu
isition card measured and recorded the raw voltage output of the sensor and fan speed 
was 
calculated
 
during post processing. 
A threshold subroutine was applied to the raw signal 
(
Figure 
3
.
1
-
left) to obtain the rotational 
frequency
 
which 
was
 
later 
resolved
 
into fan speed 
(
Figure 
3
.
1
-
right). 
The o
nce
-
per
-
rev 
signal 
was sampled at 500 k
Hz. 
 
 
Figure 
3
.
1
. 
Optical once
-
per
-
rev post process (left) and 
resolved
 
fan 
speed (right).
 
3.2
 
Data Collection
 
The motion control and data acquisition was controlled via an automated LabVIEW

 
software program. Real
-
time data waveforms were monitored during the experiment 
such 
that any irregularities could be identified. 
In the event o
f
 
such occurrences
,
 
tests were 
immediately 
aborted
 
and restarted after correction
s were performed at the point in the 
sequence prior to malfunction. 
 
 
3.2.1
 
Sample time
 
A sample time of 
five
 
seconds was chosen for all tests. This duration was selected 
considerin
g 
several factors; previous experience, maintaining consistent inlet conditions, 
measurement grid density, 
test matrix size, 
and cost. Since these experiments require jet 
fuel to 
complete
, lengthy test times can quickly become expensive. The primary driver
 
for 
the sample time is the variation in engine operation condition over the course of a 
collection. The hydromechanical fuel control system onboard the engine
 
limits how fine 
the power settin
g on the engine can be adjusted,
 
t
hus during a data collection 
t
he fan speed 
can vary 
±
0.25%
 
as shown in 
Figure 
3
.
1
-
right
. Longer collection times would result in an 
improved average
,
 
but a reduction in grid density or experimental 
conditions
 
would 
be 
required
 
to keep the experiment
 
affordable
. 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
21
 
 
Figure 
3
.
2
 
shows a 

good

 
and 

poor

 
sample 
sixty
 
second data collection 
for port one 
of the five hole probe during an
 
experiment. As shown in both figures, the variability over 
the 
duration of the collection
 
can be significant
ly more than the systematic uncertainty, 
with 
more than double the magnitude for the 
poor collection. These large scale fluctuations 
are caused primarily by the 
unsteady
 
nature of the engine 
during 
operation. 
Good data 
collections
, like 
Figure 
3
.
2
-
left,
 
show the running average within systematic uncertainty 
bounds for the duration of the collection, while the poor collection
, shown in 
Figure 
3
.
2
-
right,
 
takes significantly longer for the running average to 
settle within the uncertainty 
bounds.
 
 
The majority
 
of the
 

where representative averages of the flow are obtained within five seconds, and thus a five 
second collection window was used for
 
each measurement.
 
For collections that were not 
well behaved, overall uncertainty was reduced via repeated experiments. 
For all data, a
 
sample average was taken over the 
five
 
second collection and treated as a single data point. 
The standard deviatio
n of 
the 
five second 
signal was carried as random signal 
uncertainty
 
through the overall uncertainty calculation. 
 
 
Figure 
3
.
2
. 
Good (left) and poor (right) s
ample calibrated pressure signal during 
60 
second 
data
 
collection
.
 
3.2.2
 
Test Replications
 
One of the most important comp
onents of experimental research and the scientific 
method
 
is result verification
; and the 
only way to accomplish this is through test replication. 
While one test result can be happenstance, and t
wo tests with similar results can be 
coincidence, three experiments with similar results is indicative of a pattern. Using this 
framework, a minimum of three tests were performed for each test condition examined. 
Entire tests or subsets of tests
 
that exhib
ited high variability between measurements
 
were 
repeated 
further
. This is particularly true of the undistorted cases, which were repeated ten 
to twelve times due to their lower 
time 
duration
 
(less than 5% of a distorted test)
. 
 
The number of test replicati
ons were factored into the overall uncertainty via the 
standard error
, 
as calculated
 
in 
Equation 
(
3
.
1
)
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
22
 
 
 








 
(
3
.
1
)
 
where



 
is the standard deviation of 
the
 
data sample vector of repeated with
 
measurement 
size, 
N
.
 
 
3.2.3
 
Pressure Calibration
 
Static calibrations were performed on all transducers b
efore and after each test. An 
example calibration curve is shown in 
Figure 
3
.
3
-
left 
involving
 
several transducers. The 
pressure transducer
 
measurements
 
were post processed by taking an average of the 
calibration curves between the
 
pre
-
 
and post
-
test
 
calibrations and 
applying
 
to all data 
collected. Accompanying all calibration curves is an uncertainty relationship (
Figure 
3
.
3
-
right) that is also applied to each data point collected and carried as a 
systemat
ic
 
calibration 
uncertainty throughout the data reduction process.
 
Details of the uncertainty methodology 
is described in Section 
3.8
.
 
A NIST
-
traceable Dwyer

 
digital manometer was used as the 
pressure 
reference. The calibration
 
pressure source was compressed 
nitrogen th
at was 
downregulated to the required 
input 
value. 
 
 
Figure 
3
.
3
. Static calibration (left) and uncertainty (right) curves of the pressure 
transducers.
 
3.3
 
Inlet Conditio
ns
 
For each test configuration investigated (fan speed and profile) discrete points at 
Station 4 and Station 5 had to be taken to obtain the full profile. This took approximately 
1.5 hours, in which case the engine fan speed had to be maintained as consist
ently as 
possible
 
to corrected conditions
. The following subsections describe how well 
th
ese 
conditions were maintained using Equations 
(
3
.
2
)
 
and 
(
3
.
3
)
 
as the parameter metric, where 
RDFA
 
represents the Relative Difference
 
From Average and 
RDFT
 
is the Relative 
Difference From Target. 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
23
 
 
 
















 
(
3
.
2
)
 
 


















 
(
3
.
3
)
 
3.3.1
 
Bellmouth
 
Static Taps
 
The bellmouth measurement plane at Station
 
1 utilized four static pressure taps equally 
spaced circumferentially around the tunnel. 
Figure 
3
.
4
 
shows the measurement consistency 
for each test case investigated. Higher fan speeds trended toward higher variability 
amongst
 
the measurements and distorted and undistorted cases show close agreement. 
Since the 
magnitude of the variation is w
ithin 0.55% RDFA for all cases, the measurement variability 
for the static taps was concluded to be sufficiently low
. 
 
 
Figure 
3
.
4
. Bellmouth static 
pressure 
tap measurement consistency.
 
3.3.2
 
Mass
 
F
low Rate 
Calcu
lation 
 
The mass flow rate 
was
 
calculated through an iterative procedure incorporating 
a 
series of three
 
relationships
 
between
 
the 

Re
)
,
 
mass flow
 
rate (


)
, and
 
bellmouth discharge coefficient (
C
D
).
 
The Re and 


 
are solved 
for 
using 
Equations 
(
3
.
4
)
 
and 
(
3
.
5
)
, respectively, and 
Figure 
3
.
5
-
left is used to obtain C
D
. 
For Equation 
(
3
.
4
)
, 

 
is 
the kinematic vi
scosity of air (8.9 x 10
-
4
 
kg/(m

s)) and 
d
 
is the tunnel diameter (21in). For 
Equation 
(
3
.
5
)
, P
T
 
and P
S
 
is the total and static pressure in the tunnel
,
 
respectively, 

 
is 
the specific heat 
ratio 
of air 
as
 
an ideal gas 
(
used under standard atmospheric conditions 
as 
1.4), 
M
air
 
is the molecular weight of air (28.97 g/mol), T
T 
is the total temperature in the 
tunnel, and 
R
 
is the specific gas constant for air (287 J/(kg

K)).
 
The 
data from 
Figure 
3
.
5
-
left was obtained from the bellmouth calibration perfo
rmed 
at the Virginia Tech TurboL
ab
, details of which can be found in 
[20]
. 
Figure 
3
.
5
-
right
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
24
 
 
shows the 
uncertainty inherent in the calculation of 
C
D
, 
which
 
is carried 
as
 
systematic 
uncertainty in the calculation of the overall uncertainty. 
Details of the uncertainty 
methodology is described in Section 
3.8
.
 
 
 













 
(
3
.
4
)
 
 




















































 
(
3
.
5
)
 
 
Figure 
3
.
5
. Discharge coefficient (left) and discharge coefficient uncertainty (right). 
 
3.3.3
 
Mass
 
F
low Rate and Fan Speed Consistency
 
The calculated 
mass flow
 
rate and measured fan speed were kept
 
as consistent as 
possi
ble over the course of 
each
 
test in order to validate the assumption of 
constant
 
inlet 
flow conditions 
between
 
measurement location
s
 
and 
test replications
. 
Figure 
3
.
6
 
shows the 
variability of the 
mass flow
 
r
ate (left) and fan speed (right) between these measurements 
for the different test conditions investigated. 
The results show higher variability for the 
distorted cases and consistent ranges across fan speeds. The variability between 
measurements is small
 
w
hen compared to the
 
parameter magnitude as indicated by the 
RDFA (
Figure 
3
.
6
-
left) and RDFT (
Figure 
3
.
6
-
right) values
, 
validating
 
the consistent inlet 
condition assumption. 
 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
25
 
 
 
Figure 
3
.
6
. Corrected mass
 
flow and fan speed measurement consistency.
 
3.4
 
AIP 
 
3.4.1
 
Total Pressure
 
Consistency
 
The variability in the total pressure measurements is shown in 
Figure 
3
.
7
 
for the 70% 
CFS case, with the d
istorted data plotted on the left and the undistorted on the right. 
The 
maximum RDFA
 
(
Equation 
(
3
.
2
)
) 
for each location is plotted and both the distorte
d and 
undistorted cases show in
significant variation
 
as indicated by the 0.12% maximum 
magnitude
. Thus
 
the inlet mass
 
flow variation shown from 
Figure 
3
.
6
-
left caused 
insigni
fi
cant var
iation on the pressure profile, 
further 
validating the 
consistent inlet 
condition consumption. 
Variability is higher in the distorted
 
case due to the larger 
pressure 
variation in the profile as well as the secondary flow inhere
nt from the flow redistribution, 
which will be examined in detail in 
Chapter 4
.
 
Previous literature exam
ine
s this 
phenomenon
 
in 
great
er
 
detail 
[23]
. 
 
 
Figure 
3
.
7
. AIP total pressure measurement variability.
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
26
 
 
3.4.2
 
Static Pressure
 
Consistency
 
Figure 
3
.
8
 
shows the static pressure variability a
t the AIP for the different test cases 
investigated. The distorted profiles 
produce
 
significantly higher variability and show a 
strong correlation to fan speed while the u
ndistorted profiles show consistent 
variation 
magnitude
 
across fan speeds
. 
The calcul
ated RDF
A variation between repeated tests cases 
is shown to be satisfactorily small. 
 
 
Figure 
3
.
8
. AIP static pressure taps measurement variability.
 
3.5
 
Five Hole
 
Probe
 
3.5.1
 
Pressure and Flow Angularity
 
The non
-
null
ing calibration
 
described in 
Chapter 2
 
allowed for the creation of pressure 
coefficient maps that 
were
 
used in the determination of flow angles and 
total and static 
pressure from experimental data. This was done by creating int
erpolation functions of the 
form given in Equation 
(
3
.
6
)
, where 

 
represents the 
pitch and yaw flow angles, and the 
total and static pressure coefficients
. 
 
 













          
 
(
3
.
6
)
 
The total and static pressures were
 
calculated 
using Equations 
(
3
.
7
)
 
and 
(
3
.
8
)
, 
respectively, where 


 
and P
1
 
are defined in Section
 
2.2
 
and 



 
and 



 
are 
calculated from Equation 
(
3
.
6
)
.
 
 
















 
(
3
.
7
)
 
 















 
(
3
.
8
)
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
27
 
 
Once the total and static pressures 
were
 
known, the flow Mach number 
was
 
calculated 
using Equation 
(
3
.
9
)
. 
 
 





















 
(
3
.
9
)
 
3.5.2
 
Total Temperature
 
The total temperature 
wa
s 
measured
 
directly from the K
-
type thermocouple embedded 
in the sensing 
head of the probe.
 
3.5.3
 
Static Temperature
 
The static temperature, which 
was
 
uti
lized to calculate the local flow velocities, 
was
 
calculated using Equation 
(
3
.
10
)
, where T
T
 
and T
S
 
are the total and static temperatures, 
respectively. 
 
 
    
 















 
(
3
.
10
)
 
3.5.4
 
Component Velocities
 
Once the static temperature 
was
 
known, the 
velocity magnitude
, 


,
 
was
 
calculated 
utilizing the speed of sound
, Equation 
(
3
.
11
)
,
 
and the Mach number
 
previously determined, 
as s
hown in
 
Equation 
(
3
.
12
)
. 
 
 






 
(
3
.
11
)
 
 




 
(
3
.
12
)
 
The velocity magnitude
 
was
 
then decomposed into polar coordi
nates as illustrated in 
Figure 
3
.
9
 
via coordinate transformations
, which follows from 
Treaster 
[36]
. This results 
in 
the calculation of the radial, tangential, and axial velocity using Equations 
(
3
.
13
)
 
-
 
(
3
.
15
)
, 
respectively
.
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
28
 
 
 
Figure 
3
.
9
. 
Polar coordinate v
elocity component
s
.
 
 









 
(
3
.
13
)
 
 













 
(
3
.
14
)
 
 













 
(
3
.
15
)
 
3.6
 
Fan Performance 
Parameters
 
Using the results obtained from Section 
3.5
, fan performance metrics 
were
 
calculated 
using classical turbomachinery relationships. The following sections
 
outline the equations 
used to obtain the p
arameters discussed in 
Chapter 4
. 
 
3.6.1
 
Incidence
 
and Deviation
 
Using the component velocities calculated in Section
 
3.5
, classical velocity 
triangles 
were
 
constructed 
using a mean
-
line analysis 
such tha
t the blade relative flow parameters 
could
 
be calculated, as shown in 
Figure 
3
.
10
-
left. 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
29
 
 
 
Figure 
3
.
10
. Schematic of 
rotor velocity triangles (left), and 
incidence (i) and deviation 
(

) angle
s (right)
.
 
Blade relative flow angles at the inlet and outlet to the rotor, 

1 

2
, 
were
 
calculated 
using Equation 
(
3
.
16
)
, where 
U
 
is the blade speed. 
 
 
 













 
(
3
.
16
)
 
The leading edge metal angle is subtracted from the inlet relative flow angle to calculate 
the blade incidence, as shown in 
Equation 
(
3
.
17
)
, which is a representation of how far 
off
-
design the leadin
g edge flow angle
 
is. 
Positive incidence is indicative 
of the
 
relative flow 
approaching more heavily on the pressure surface of the blade and negative incidence 
is 
indicative of the flow approaching more heavily on the suction surface. 
 
 













 
(
3
.
17
)
 
The trailing edge metal angle is subtracted from the outlet relative flow angle to 
calculate the blade deviation, as shown in Equation 
(
3
.
18
)
. This parameter represents the 
magnitude
 
of 
under turning or premature separation
 
(ne
gative values
)
 
and overturning
 
(
positive values
)
. Overturning is a 
rare 
occurrence 
and
 
is
 
usually the result of non
-
uniform
 
flows and non
-
standa
rd fan geometry
 
affecting the aerodynamics
. 
 
 














 
(
3
.
18
)
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
30
 
 
3.6.2
 
Blade Loading
 
Blade loading is a representation of how much turning the blade is required to perform 
in order to maintain a constant
 
work input across the rotor. This parameter is strongly 
dependent on the blade speed, as shown in Equation 
(
3
.
19
)
. 
Higher blade loading requires 
more output from the vanes, pushing them closer to their separation and stal
l limit
s
. 
 
 
































 
(
3
.
19
)
 
Assuming constant U and 
constant 
axial velocity C
z
 
through the rotor, 
Equation 
(
3
.
19
)
 
can be simplified to Equation 
(
3
.
20
)
. 
 
 





















 
(
3
.
20
)
 
3.6.3
 
Pressure Rise
 
The pressure rise across the rotor 
is calculat
ed by dividing the outlet local total pressure 
by the corresponding inlet total pressure, as shown in Equation 
(
3
.
21
)
. 
 
 








 
(
3
.
21
)
 
3.6.4
 
Efficiency
 
Combining the pressure rati
o and 
blade 
loading, the local efficiency 
was
 
calculated 
using Equation 
(
3
.
22
)
, where 

 
is the specific h
eat ratio for an ideal gas, 


 
is the specific 
heat of air at constant pressure
, and 





 
 



























 
(
3
.
22
)
 
3.6.5
 
Lieblein Diffusion Factor
 
The amount of turning possible across a blade
,
 
and thus the 
allowab
le work input
,
 
is 
limited by the amount of diffusion possible within a blade passage. 
The Lieblein diffusion 
factor, calculated in Equation 
(
3
.
23
)
, is the parameter 
most often
 
used in determining 
excessive diffusion. Typical compressor desi
gns limit the LDF to < 0.5
 
[28]
.
 
Figure 
3
.
10
-
left can be used a
s a
 
reference for the terminology in Equation 
(
3
.
23
)
. The parameters s 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
31
 
 
and c represent the blade spacing and chord, 
respec
tively
. The 
chord to spacing ratio, 
c/s, 
is termed the solidity. 
 
 


























 
(
3
.
23
)
 
3.6.6
 
Mass Flux
 
 
The local 
mass flux
, expressed as the 
mass flow
 
rate 
per unit cross sectional area, is a 
representation of the local 
mass flow
 
at a 
discrete point within the flow and 
was
 
calculated 
with 
t
he flow p
arameters obtained
 
from
 
the five
-
hole probe. 
Combined 
with 
the local 
pressure rise,
 
the mass flux can be used to
 
approximate
 
the fan performance characteristic 
which
 
is shown 
later 
in 
Chapter 4
. 
 
 



































 
(
3
.
24
)
 
3.7
 
Comparison Metrics
 
3.7.1
 
Corrected Conditions
 
for Fan Performance
 
Due to the strong secondary flow present for the distorted cases
 
a
nd that which is 
inherent in turbomachinery
, it was necessary to develop a means of tracking streamtubes 
between measurement planes: 
from Station 4
 
to Station 5.
 
This was accomplished by 
calculating circumferential and radial shift based on the flow veloci
ty in those respective 
directions. 
For example, 
Figure 
3
.
11
 
shows a graphical representation of the 
circumferential shift of a streamtube as it travels from Station 4 to Station 5. 
 
 
Figure 
3
.
11
. Schematic representing the components required for calculating 
circumferential displacement.
 
The streamtube travel path is broken into three sections:
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
32
 
 
 
1)
 
Measurement 
Station 4 to the fan 
blade 
leading edge
 
2)
 
Fan 
blade 
leading edge to the fan 
bla
de 
trailing edge
 
3)
 
Fan 
blade 
trailing edge to 
measurement 
Station 5
 
 
The duration each fluid particle spends in sections 1) and 3) is calculated by Equation 
(
3
.
25
)
 
while the duration through the fan is calculated similarly b
ut using the average 
relative velocities, 
w
4
 
and 
w
5
, in place of the axial velocity (
C
Z
) and the blade chord length 
in place of the axial length (
z
z
). 
 
 






 
(
3
.
25
)
 
The 
tangential
 
velocity compone
nt, 
C

, is then used in Equation 
(
3
.
26
)
 
to calculate t
he 
circumferential displacement, where 


 
represents a polar coordinate velocity. 
 
 







 
(
3
.
26
)
 
This procedure is 
performed for each 
section
 
and the overall displacement between 
Station 4 and 
Station 
5 is obtained with a summation.
 
A similar 
procedure
 
is performed for the radial displacement from Station
 
4 to Station 
5 with the exception of using the radial velocity c
omponent, 
C
r
, when utilizing Equation 
(
3
.
26
)
 
to calculate displacement
 
in place of the tangential velocity component, C

. A 
graphic representation is shown in 
Figure 
3
.
12
. 
 
 
Figure 
3
.
12
. 
Schematic representing the components required 
in 
calculating 
radial
 
displacement.
 
3.7.2
 
Data Presentation: 
Relative Di
fference
 
from
 
Undistorted
 


,
 
calculated in Equation 
(
3
.
27
)
 
where 
Y
 
represents a measured 
or calculated parameter. This descripto
r was used as a means of showing the parameter 
variation due to the distorted
 
inlet conditions
 
in relation to the undistorted conditions
 
at 
equal fan speed
, effectively removing the design nuances of the fan from the presented 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
33
 
 
results
. For the results disc
ussed in 
Chapter 4
, this parameter is typically used
 
in the data 
plots
, presented as a percent (%),
 
and abbreviated as RDFU. 
 
 














 
(
3
.
27
)
 
For flow angles, a difference calculation is performed between the 
distorted and 
undistorted cases, as shown in 
Equation 
(
3
.
28
)
. 
 
 








 
(
3
.
28
)
 
3.8
 
Uncertainty
 
The uncertainty for this study was 
determined through careful analysis of the 
calibration and measurement techniques utilized. Every measurement had a quantifiable 
uncertainty associated with it
 
that was carried throughout the analysis procedure, from 
transducer 
calibration to 
parameter 
calculation. Recall from 
Figure 
3
.
3
-
right, the 
uncertainty components associated with static transducer calibrations before and after ea
ch 
experiment. 
These uncertainty curves were established for each transducer and the 
uncertainty associated with each measured value was determined from
 
the voltage recorded 
at the time of measurement.
 
Combining these calibration uncertainties with the ran
dom 
uncertainties associated with repeated measurements, an overall uncertainty was obtained. 
 
3.8.1
 
Measured Parameter
 
The overall uncertainty for a measured parameter is given in Equation 
(
3
.
29
)
 
where the 
conf
idence interval used for the investigation was set at 95%. 
 
 































 
(
3
.
29
)
 

-
distribution
, 


,
 
was chosen as the
 
mean estimation
 
parameter
, with 

 
representing the number of observations. Equation 
(
3
.
29
)
 
utilizes the standard error
, 



, 
defined in Equation 
(
3
.
1
)
. The systematic error so
urces in the system, 


, represent 
measurement bias
 
due to a number of factors; calibration reference, 
linear data fits, 
electrical 
wiring
, pneumatic plumbing
, data acquisition,
 
signal conditioning,
 
etc. The 
random error sources, 


, represent fl
uctuations in 
the recorded data signal 
and 
include
 
the 
standard deviation of the transducer signal for both the 
static 
calibration and the 
experimental measurement signal. The 
random
 
error associated with the calibrations 
was
 
very small and could have been
 
considered negligible but 
was
 
kept throughout the study 
for completeness. 
Uncertainties obtained from Equation 
(
3
.
29
)
 
were obtained for each 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
34
 
 
unique 
measurement location
, where repeat measurements were acc
ounte
d for in the 
standard error, 



.
 
 
3.8.2
 
Five Hole Probe Parameters
 
The uncertainty associated with the five hole parameters has two additional 
components 
due to the 
non
-
nulling 
probe 
calibration 
performed
. 
This is shown in Equation 
(
3
.
30
)
, where 


 
is the 
transducer 
calibration 
uncertainty 
from the probe calibration
, 


 
is the interpolation uncertainty associated with the interpolation schema employed when 
using the calibration maps to determine the 
flow 
p
arameters, and 


 
is the experimental 
uncertainty calculated in Equation 
(
3
.
29
)
 
from the 
measured 
experimental
 
engine data. 
 
 















 
(
3
.
30
)
 
The 
transducer 
calibration uncertainty, 


, is calculated similarly 
for
 
the experimental 
engine data
 
as the probe calibration incorporates several replicated data sets to ensure data 
quality. Pre and post
-
test
 
transducer calibrations were performed for each probe calibration 
test and the uncertainties associated with each test were accounted for. The final overall
 
probe calibration
 
uncertainty is calculated using Equation 
(
3
.
29
)
. 
 
The interpolation uncertainty is calculated using a
 
jack
knife approach
, further details 
of which can be found in
 
Tomczak
 
[32]
. 

the interpolation schema util
ized during post processing. This is performed by first 
calculating a pseudo
-
value
 
five hole parameter (total pressure, static pressure, yaw angle, 
pitch angle), 




 
Equation 
(
3
.
31
)
. 
 
 

















 
(
3
.
31
)
 
I
 
represents the total number of calibration points used in the interpolation, which for a 
two dimensional interpolation is 4. The parameters 


 
and 




 
represent the parameter 

from the map, respectively. 
The interpolation uncertainty is then calculated using Equation 
(
3
.
32
)
. This procedure was performed for each unique
 
experimental data point. 
 
 

























 
(
3
.
32
)
 
Table 
3
.
2
 
summ
arizes the RMS
 
(Root Mean Square)
 
nominal 
uncertainties
 
for 
each 
term in Equation 
(
3
.
30
)
, where RMS is defined in Equation 
(
3
.
33
)
, where 
N
 
is the data 
sample length and 

 
is the parameter of interest. 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
35
 
 
 












 
(
3
.
33
)
 
For the cases of percent (%) nominal, the parameter of interest, 

, is a dimensionless 
relative difference between the uncertainty and the nominal value, as 
calculated using 
Equation
 
(
3
.
34
)
, where 

 
represents the dimensioned parameter of interest. 
 
 









 
(
3
.
34
)
 
As shown
 
in 
Table 
3
.
2
, the uncertainty associated with 
the 
engine experiments has the 
highest contribution to the overall uncertainty, followed by the probe calibration 
uncertainty, and finally the interpolation map. This table shows that the probe calibration
 
per
formed
 
was sufficiently accurate for the experiments conducted. 
 
Table 
3
.
2
. 
Five
-
hole parameter 
RMS 
uncertainty
.
 
[RMS % Nominal]
 


 
 


 
 


 
 


 
Total Pressure
 
0.04
 
0.02
 
0.60
 
0.60
 
Static Pressure
 
0.08
 
0.02
 
0.14
 
0.16
 
Yaw Angle
 
0.59
 
0.32
 
1.39
 
1.54
 
Pitch Angle
 
1.68
 
0.71
 
4.23
 
4.61
 
3.8.3
 
Calculated Parameters
 
The uncertainty associated with calculated par
ameters was 
determined
 
using a 
sequential perturbation technique
, with
 
details of the 
procedure 
found in 
Appendix B
.
 
Table 
3
.
3
 
summarizes the RMS
 
nominal val
ues of the calculated parameter uncertaintie
s
 
presented throughout the dissertation. 
 
 
Chapter 3
 
-
 
Data Reduction
 
 
 
 
36
 
 
Table 
3
.
3
. 
Measured and calculated parameter 
RMS uncertainty
 
values
 
given as a percent 
of the nominal value.
 
[
RMS 
%
 
Nominal
]
 



 
CFS
 
3.5
 
Mass flow
 
6.2
 
Bellmouth Static Pressure
 
0.9
 
AIP Total Pressure
 
2.4
 
AIP Static Pressure
 
1.3
 
Total Temperature
 
9.5
 
Incidence
 
2.9
 
Deviation
 
2.9
 
Blade Loading
 
3.4
 
Pressure Rise
 
8.7
 
Efficiency
 
16.4
 
Lie
blein Diffusion Factor
 
8.9
 
Ma
ss flux
 
9.4
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
37
 
 
Chapter 4
 
Results and Discussion 
 
4.1
 
Consistent Inlet Conditions
 
Since the probe and 
distortion 
screen 
were
 
traversed and rotated, respectively
,
 
to obtain 
the measurement field
,
 
it was important to ensure that consistent inlet conditions were 
maintained b
etween measurements. This was accomplished by 
continually monitoring and 
adjusting
 
the 
un
corrected fan speed 
as 
atmospheric conditions 
varied
 
over the course of a 
test. 
A similar procedure was performed between tests, which occurred on different days 
and d
ifferent atmospheric conditions. 
The uncorrected fan speed was calculated real
-
time 
using Equation 
(
4
.
1
)
, where 
N
cor
 
is the c
orrected fan speed (60, 70, 80%
), and 
T
atm
 
was the 
current local atmospheric temperature and 
T
std
 
was the standard reference temperature used 
in the investigation (59
°
F). 
Figure 
2
.
7
 
shows the consistency of such a procedure, which 
lead to the acceptable uncertainties listed in
 
Table 
3
.
3
.
 
 











 
(
4
.
1
)
 
For all 
test 
configurations and speeds 
investigated
, the corrected fan speed was kept to 
within 
±
 

r, and tunnel Mach number 
were kept 
to 
within 
± 1.2% of the target value. These variations represent
 
2 standard 
deviations of the sample population data. 
No 
distortion 
screen dependent 
variations
 
were 
found in any 
of the 
above 
parameters
 
mentioned. 
 
4.2
 
Undist
orted Inlet Flow Conditions
 
Undistorted fan performance condition
s were obtained for two reasons; t
he first 
was
 
to 
provide a means 
of
 
comparison
 
for the 
distorted conditions, 
and the 
second 
was to
 
gain an 
understanding how the engine performs in the undist
orted experimental rig utilized. The 
engine 
was
 
always 
operating in 
off
-
design
 
conditions, even in the undistorted inlet flow 
experiments
 
due to the pressure loss of the backing screen
. The bellmouth, accompanying 
tunnel sections, and AIP rake all contribu
te to losses in the overall system that must be 
accounted for. This 
was
 
done with duplicating the measurements made in the distorted tests
 
for the
 
undistorted conditions. The following 
sub
sections highlight the results. 
 
4.2.1
 
Station 3 
-
 
AIP Inlet Profile
 
The i
nlet total pressure profile
 
for three respective test speed
s
 
investigated is shown 
Figure 
4
.
1
. Low losses through the experimental rig ducting is shown for the 50% CFS test 
conditions, with increasing losses for increasing fan spe
ed
. 
A more pronounced boundary 
layer becomes prevalent in the 65% CFS case and is at its most severe at 85% CFS. 
Figure 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
38
 
 
4
.
1
 
shows a uniform inlet profile for the test conditions evaluated, providing a valid inlet 
profile to compar
e against for distorted conditions. AIP static pressure results are 
included 
later in the document in the 
distorted inlet flow discussion of 
Section 
4.3
. 
 
 
Figure 
4
.
1
. AIP total 
pressure contours fo
r
 
varying speed 
und
i
s
to
r
ted 
inlet conditions
.
 
4.2.2
 
Station 4 and 
Station 
5 
-
 
Fan Inlet and Fan Outlet 
Conditions
 
Undistorted data measured at Station 4 (fan inlet) and 
Station 
5 (fan outlet) are 
summarized in the following 
sub
sections. 
For 
c
onciseness, both stations are plotted 
together, with solid lines representing Station 4
 
(fan inlet)
 
and dashed lines representing 
Station 5
 
(fan outlet)
. 
Fan blade geometry
 
(splitter, part
-
span shroud, stiffener)
 
is 
add
itionally plotted for reference
 
as th
eir effects on the measurements
 
are distinguishable. 
 
4.2.2.1
 
Total and Static Pressure
 
The 
undistorted 
total and static pressure
 
distributions 
at Station 4 and 
Station 
5 
are 
shown in 
Figure 
4
.
2
, with the total pressure results shown on t
he left and the static pressure
s
 
shown 
on the right. Data is colored according to 
%
CFS and referenced 
in
 
the colorbar on 
the far right
 
of the plot
. 
 
The inlet total pressure 
of 
Figure 
4
.
2
-
left 
is 
well behaved across fan speeds; 
th
e only 
component of note is the drop in total pressure
 
radially outward beyond the stiffener for 
CFS beyond 72.9%, which correspond to transonic conditions. The subsequent increase in 
static pressure for similar conditions (
Figure 
4
.
2
-
right) supports this. Station 5 total 
pressure 
also 
shows a well behaved distribution, with increasing fan output pressure 
for 
increasing blade span.
 
The part
-
span shroud causes a disruption in the linearity of the 
distribution, and the 
mid
-
core 
regio
n 
shows a more aggressive rise in total pressure before 
reducing
 
as it approaches the splitter. Otherwise, the blade geometry effects on the
 
undistorted
 
total pressure distribution are 
small
. 
 
The static pressure distribution
 
of Station 5 
in 
Figure 
4
.
2
-
right 
is more varied than the 
total pressure
 
due
 
to the profound effect the splitter has on the upstream measurements. 
The measurement plane of Station 5 is directly upstream of the splitter as 
engine 
geometry 
constraints
 
prevented t
he plane from being 
further
 
upstream. 
With the total pressure 
distrib
ut
ion unchanged in this region (
Figure 
4
.
2
-
left), the reduced flow velocity as it 
approaches the splitter will cause an increase in the static pressure as shown.
 
Other than 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
39
 
 
the splitter effects
,
 
the static pressure distribution is similar to the total pressure, increasing 
in sensitivity with increasing CFS. The tip region of the static pressure distribution
 
shows 
a slight reduction, indicating 
higher velocity
 
flow
 
as the total pressure distribution is 
unchanged. 
 
 
Figure 
4
.
2
. Undistorted
 
total (left) 
and static 
(right) 
pressure for Station
 
4 and 
Station 
5.
 
4.2.2.2
 
Tangential and Radial Flow Angle
 
The undistorted tangential a
nd radial flow angles
 
are plotted in 
Figure 
4
.
3
. The 
tangential flow angle (
Figure 
4
.
3
-
left) at Station 4 has a small
,
 
bulk 
fan
-
wise pre
-
swirl 
component that does not vary significantly in the radial dire
ction. 
Near core flow will be 
he
avily influenced by the spinner
 
while 
bypass flow will be influenced by the blade 
rotation
, though a uniformity is present across the span.
 
Close agreement is observed 
between all CFS

. 
 
Station 5 tangential flow ang
le is more heavily influenced by the splitter and part
-
span 
shroud. The overall distribution of the tangential flow angle has a backward lean, with 
increased angle near the root and reduced near the tip
 
(this correlates well with the blade 
loading distribu
tion shown later in 
Figure 
4
.
7
)
. This is
 
required
 
to 
maintain constant 
near
-
work input across the blade span, 
requiring
 
more turning at lower radii. 
The splitter and 
part
-
span shroud cause a sharp rise in the tangential flow angle
, with the effects from the 
part
-
span shroud 
perturbating
 
further radially outward. 
Close agreement is observed 

 
The radial flow angles (
Figure 
4
.
3
-
right)
 
at Station 4 are nearly exclusively positive, 
ind
icating radially outward flow
 
toward the blade tip
. 
This is strongest near the spinner, 
where flow is navigating around 
the cone and 
moving into the annulus portion of the 
engine. 
The radial flow angle magnitude decreases with increasing blade span as this
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
40
 
 
radially outward momentum 
is dissipated through the remainder of the flowfield. 
A sharp 
rise in radial flow angle at the tip is likely due to the clearance between the casing and 
blade tip
, causing
 
a static pressure gradient that forces flow outward towar
d the case
, an 
effect supported in 
the Station 5 static pressure distribution of 
Figure 
4
.
2
-
right
. 
Although 

discrepancy for transonic conditions radia
lly outward from the part
-
span shroud, although 
the tip
-
most data point
 
shows close
r
 
agreement. 
 
The radial flow
 
magnitude 
at Station 5 is reduced 
from that at 
Station 4. Core flow is 
on average 
exclusively
 
axial, while the region directly above the splitt
er 
features a higher 
radial
 
component
, 
indicat
ing
 
flow moving 
out of the
 
core
 
and into the bypass. T
his effect 
is more
 
pronounced 
at lower speeds
, indicating a mismatch between the fan and core 
compressor at 
off
-
design
 
conditions.
 
Assuming the trend contin
ues for increasing CFS, the 
design condition of the fan should reflect near zero radial flow at this location. 
The splitter 
and 
part
-
span shroud entrap the flow in a pocket of radial outward flow, before decreasing 
radially outward beyond the shroud, where
 
it continually decreases before the sharp tip 
increase. Overall trends of both Station 4 and 5 for the radial flow ang
l
e show a backward 
lean distribution for increasing blade span,
 
a trend that is disrupted at the tip
.
 
 
Figure 
4
.
3
. Undistorted tangential (left) and radial (right) flow angles
 
4.2.2.3
 
Total Temperature
 
 
The total temperature distribution is shown in 
Figur
e 
4
.
4
. Station 4 shows a
 
constant 
inlet total temperature with the exce
ption of the near
-
tip rise, which is more pronounced 

. 
Station 5 
conditions show a similar trend to the total pressure of 
Figure 
4
.
2
-
left, with increasi
ng total 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
41
 
 
temperature with increasing blade span. The part
-
span shroud is the only portion of the 
blade geometry to significantly alter the distribution, 
producing
 
a higher total temperature 
at the fan outlet in this region
 
than the trend suggests
. 
 
 
Figur
e 
4
.
4
. Undistorted total temperature 
span
-
wise
 
distribution. 
 
4.2.2.4
 
Velocity
 
The axial velocity (left) and velocity magnitude (right) 
span
-
wise
 
distributions are 
shown in 
Figure
 
4
.
5
. 
Close
 
agreement of
 
axial velocity between Stations 4 and 5 are shown 
at span locations below the part
-
span s
hroud, while large discrepancies 
exist radially 
outward
. Both the splitter and 
part
-
span 
shroud 
cause
 
severe axial velocity depreciations
 
at Station 5
. A
xial velocity reductions at the tip are observed for transonic 
conditions
, a 
consequence of the shock formation due to the high blade speed rotation
 
rate
. 
 
The velocity magnitude distribution of 
Figure
 
4
.
5
-
right shows 
the accelera
ted flow 
profile from Station 4 to Station 5. Splitter effects at Station 5 are the 
primary
 
engine 
geometry components that appear to affect the distribution, as both the stiffener and part
-
span
 
shroud do not al
ter the distribution significantly. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
42
 
 
 
Figure
 
4
.
5
. Undistorted axial (left) and velocity magnitude (right) span wise distributions
 
4.2.2.5
 
Incidence and Deviation
 
The fan blade incidence and deviation is shown in 
Figure 
4
.
6
. The result
s
 
for Station 4
 
show a 
slight
ly varying
 
positive 
incidence
 
on the blades for all span locations measured 
with the exception of the tip
, indicating the flow is directed more towards the pressure side 
of the blade than an optimal design would predict
. 
Magnit
ude is shown to be speed 
dependent
 
for radii below the blade stiffener
. Station 5 shows a deviation of less than 5 
degrees for the majority of the locations and speeds investigated. It is interesting that the 
near
-
splitter and shroud show a
 
reduced deviati
on
 
from the other measurements
, with the 
splitter causing positive deviation, or overturning
.
 
Such an effect would not be possible in 
traditional cascade aerodynamics.
 
The stiffener appears to not have affected the 
distribution. 
The results 
agree well with
 

supporting a conclusion
 
that
 
the 
results are reasonable
 
[22]
. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
43
 
 
 
Figure 
4
.
6
. 
Undistorted incidence
 
and 
deviation a
ngle
.
 
4.2.3
 
Fan Performance
 
The baseline undistorted fa
n performance metrics provided information on the design 
behavior of the engine. This was especially useful 
in validating the 
m
easurements as well 
as determining
 
how the fan blade geometry affected such metrics. 
 
4.2.3.1
 
Blade Loading
 







 
and Pressure Ri
se
 
The 
undistorted 
blade loading 
distribution 
is shown in 
Figure 
4
.
7
-
left, with decreased 
loading for increasing blade span; a consequence of the 
near
-
constant work 
design 
input 
which requires more turning at lower blade sp
an due 
to the lower blade speed
.
 
Details of 
the work input distribution are 
given
 
in
 
Appendix F
. 
 
Close agreement between all speeds 
tested is observed. Differentiation from the linear backward lean trend 
is 
observed for the 
part
-
span
 
shroud, causing increased loading
 
at that location and the area radially outward 
from it
. The stiffener and tip region also move the trend away from its 
close linear 
relationship, though less so than the shroud. 
This unloading at the blade tip was done to
 
prevent separation 
[27]
. The 
splitter appears to have negligible effect on the loading
 
distribution
. 
 
The undistorted pressure rise distribution is shown in 
Figure 
4
.
7
-
right with increasing 
sensitivit
y with blade span for increasing blade speed. 
Only the part
-
span shroud shows a 
significant alteration to the distribution, causing a reduction in the pressure rise, an effect 
observed across all fan speeds tested. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
44
 
 
 
Figure 
4
.
7
. Undistorted blade loading (left) and pressure rise (right) 
blade span
 
distributions
.
 
4.2.3.2
 
Blade Loading Sensitivity
 
The blade loading 
calculation given in Equation 
(
3
.
20
)
 
can be rearranged for the outlet 
a
bsolute flow angle, 


, and the inlet relative flow angle, 


 
as shown
 
in Equation 
(
4
.
2
)
. 
 
 
























 
(
4
.
2
)
 
The 
fan 
blades 
are more sensitive to the inlet flow coefficient at higher radii due to the 
nature of the blade design. To maintain radially uniform work input, the turning required 
at higher radii is lessened due to the increased blade speed, U. The relative inlet flow a
ngle 
(

4
) is increased as well for increased blade speed and is illustrated via the velocity triangles 
in 
Figure 
4
.
8
-
right
. As 
Figure 
4
.
8
-
left
 
shows, 

2
 
does not vary as strongly as 

1
 
nor 
is it
 
monotonically increasing. Thus the proportionality constant to the inlet flow coefficient to 
Equation 
(
4
.
2
)
, 




, 
 
will correspondingly increase for increasing radii. 
 
Reduced relative inlet flow angle is observed for incr
easing fan speed, a trend which 
extends from the core to the stiffener, before it is disrupted. Although the trend is consistent 
for the majority of the blade span, there remains close agreement between the 
measurements. 
 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
45
 
 
 
Figure 
4
.
8
. Outlet absolute flow angles (

5
) and inlet relative flow angles (

4
).
 
4.2.3.3
 
Diffusion and Efficiency
 
The Lieblein diffusion factor for the undistorted cases is shown in 
Figure 
4
.
9
-
left
,
 
which follows a similar 
backward lean 
distribution 
as
 
the blade loading observe
d
 
in 
Figure 
4
.
7
-
left; 
decreasing for 
higher blade spans. The diffusion shows a higher sensitivity to the 
engine geometry than the loading however, with noticeable shifts in 
the overall trend
. Both 
the splitter and part
-
span shroud
 
cause large increases in diffusion from expected values. 
The effects of the stiffener on diffusion are slight, less so than the tip effects. CFS appears 
to affect the severity of the diffusion 
depending
 
on the location; reduced diffusion in the 
core for i
ncreasing CFS while the bypass shows increased LDF for increasing CFS. 
Without the splitter and shroud, the LDF should be below 0.5 for the entire bypass region, 
however these geometry components push the LDF close to 0.6 in both instances, and 
even 
higher
 
for 85% CFS. 
The core LDF exceeds the recommended design 
limitation
 
of 0.5, and 
along with the splitter and part
-
span shroud regions, are likely 
prematurely 
separating. 
 
The efficiency distribution is shown in 
Figure 
4
.
9
-
right, w
ith 
magnitude
 
strongly 
dependent on blade span.
 
No coherent structure for the distribution is apparent, though 
close agreement is observed for 
the majority of the test cases investigated. The 85% CFS
 
test case shows a higher diversion than the other speed
s
 
tested near the stiffener and tip 
region but follows closely with reduced speeds at lower 
blade span
.
 
The splitter appears to 
have negligible effects on the 
local 
efficiency, while the shroud and stiffener show 
reductions in the efficiency for regions nea
r them, with very severe reductions observed 
for the shroud. 
The radial averaged efficiency of all speeds investigated is just under 80%.
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
46
 
 
 
Figure 
4
.
9
. Undistorted diffusion (left) and 
efficiency (right) blad
e span distributions
.
 
4.2.4
 
Summary
 
Fan inlet and fan outlet measurements were obtained for undistorted inlet conditions to 
serve as a baseline standard upon which to 
compare the 
distorted measurements. 
The inlet 
and outlet profiles allowed 
for 
the calculation o
f fan performance metrics that will be 
utilized in a similar way
 
for the distorted tests
. The 
data presented
 
showed the impact the 
unique fan blade and engine geometry has on the radial distribution 
of the fan performance 
parameters 
and they will play a la
rge role in the way the fan interacts with
 
and responds to
 
the 
distortion
 
profile. 
 
4.3
 
Distorted Inlet Flow Conditions
 
Distorted data was obtained for 60, 70, and 80% CFS and compared directly with the 
equivalent
 
undistorted CFS. 
Unless otherwise stated, only
 
the 70% CFS 
results
 
will be 
presented in the body of this work. This was done
 
for brevity as the majority of the profile 
features and parameters are consistent across speeds.
 
Appendix G
 
shows the results 
presented in
 
Section
 
4.3
 
for all speeds investigated. 
On request, the full data set is available 
from the author. 
Recall that the distorted results are presented using the RDFU parameter 
define
d
 
in Equation 
(
3
.
27
)
 
in most cases, 
wit
h flow angularity and efficiency data presented 
as differences as calculated in Equation 
(
3
.
28
)
. This was done to remove the engine bias 
from the experiment and isolate the distortion effects. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
47
 
 
4.3.1
 
AIP Inlet Profile
 
The total pressure profile me
asured at the AIP
,
 
representative of a 
BLI serpentine inlet, 
results in
 
a 
distortion intensity more severe than
 
probable
 
inlet 
designs would 
employ
. 
Figure 
4
.
10
 
shows the intensities for all inlet speed
s
 
tested, with a minimum 9
1.
5
% 
recovery for 80% CFS.  
As can be see
n
, p
rofile shape is preserved for each fan speed
. 
The 
contour color schemas for the 60 and 70% CFS tests look identical to the 80% CFS when 
plotted on their relative scales. 
This is shown in
 
Appendix G
.
 
 
Figure 
4
.
10
. 
Distorted 
AIP tota

investigated. 
 
The static pressure profile at the AIP
 
exhibited similar characteristics between speeds, 
with the relative redu
ction in static pressure increasing in severity with increasing 
CFS, as 
shown in 
Figure 
4
.
11
. Minimum pressure occurs at BDC
 
(180
°
)
, corresponding to that seen 
in the total pressure profile
 
which is given to the right of 
Figure 
4
.
11
 
as reference
. 
Although 
the variation in static pressure is not nearly as significant as the total pressure, it is present 
and measureable, as 
Figure 
4
.
11
 
shows. 
The systematic uncertainty of the measur
ements, 


, is show
n
 
within the figure for reference. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
48
 
 
 
Figure 
4
.
11
. AIP static pressure 
results.
 
4.3.2
 
Station 4 and 5 
-
 
Fan Inlet and Fan Outlet Conditions
 
4.3.2.1
 
Total Pressure
 
The total pressure results for 
Station
 
4 and 
Stat
i
on 
5 are shown in 
Figure 
4
.
12
-
left and 
Figure 
4
.
12
-
right, respectively. 
The 
Station 4
 
total pressure profile is very similar to that 
measured at the AIP
. A RMS
D
 
(Root Mean Square 
Differ
ence
)
 
analysis between 
Station
 
4
 
and the projected 
Station
 
3
 
results onto 
Station
 
4
 
varies from 0.53% to 1.2%, 
for
 
60 to 
80% 
CFS
,
 
respectively.
 
RMSD
 
is calculated 
using Equation 
 
(
4
.
3
)
, with 
N
 
representing the 
data vector length and 


 
representing 
the data parameter of interest and 

 
the nominal 
value.
 
 

















 
(
4
.
3
)
 
This shows that not a significant amount of 
local or global 
total pressure 
variation 
occurs b
etween these 
Station
s, though there is evidence of mixing at the interfaces of the 
high and low total pressure regions (105
° and 265°). 
A strong total pressure distortion is 
present at 
Station
 
5
 
and
 
although not as extensive as at 
Station
 
4, 
it 
has increas
ed in
 
relative
 
severity, causing 
significant losses to persist 
and exacerbate 
through the fan. The large
,
 
low 
total 
pressure 
pocket
 
at 
Station
 
4
 
at BDC
 
has been radially reduced and is limited to the 
core and splitter regions of 
Station
 
5. 
Some areas show 
an increased total pressure as a 
result of the distortion but they are limited, and a global average reduction of 4.3%
 
from 
the undistorted case
 
was calculated
. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
49
 
 
 
Figure 
4
.
12
. Distorted i
nlet (left) and outl
et (right) total pressure profiles
 
for 70% CFS
. 
 
4.3.2.2
 
Static Pressure
 
The static pressure 
non
-
uniformity
 
becomes more prevalent at 
Station
 
4 than was 
observed at 
Station
 
3
 
(
Figure 
4
.
11
)
, and is shown in 
Figure 
4
.
13
-
left
. An increase in the 
static pressure distortion 
develops
 
as the profile approaches the fan
, and is of nearly equal 
relative magnitude from the total pressure distortion of 
Figure 
4
.
12
-
left
. A noticeable 
asymmetry 
is also
 
present
, and interestingly it appears to have shifted 
the profile 
in the 
counter
-
fanwise direction. This static pressure distortion governs flow redistribution 
which
 
will become clear in 
the 
flow angle
 
and vector plots
 
shown later.
 
 
Similar to the total p
ressure profile in 
Figure 
4
.
12
-
right, the 
Station 5 
static pressure 
results in 
Figure 
4
.
13
-
right show the Station 4 
non
-
uniformity
 
entering the fan
 
has been 
contracted radially
 
and limited to the core
. Ac
ute
 
regions of low and high static pressure 
variation 
still 
exist, particularly near the splitter and in the core flow region
s
. 
The entire 
flowfield exhibits a 
total reduction
 
in static pressure from undistorted conditions
 
of 4.9%.
 
 
Recall from 
Figure 
4
.
11
 
that the static pressure variation at the AIP showed only a small 
variation, 0.7%, as compared to the 4.5% of Station 4. This indicates the 
high 
secondary 
flow variability is 
not present at the AIP, but quickly develops as the fl
ow approaches the 
fan face. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
50
 
 
 
Figure 
4
.
13
.
 
Distorted
 
fan i
nlet (left) and outlet(right) static pressure profiles
 
for 70% 
CFS
.
 
4.3.2.3
 
Total Temperature
 
That total 
temperature profiles at Station
 
4 and 
Station 
5 are 
shown in 
Figure 
4
.
14
. The 
differences between the distorted and undistorted conditions for station 4
 
(
Figure 
4
.
14
-
left)
 
are negligible, varying 
approximately 
0.3% across the profile. 
Station 5
 
(
Figure 
4
.
14
-
right)
 
shows a larger variation, approximately 4%, with a distribution 
similar to that of the blade 
loading
 
discussed 
later
. 
 
 
Figure 
4
.
14
. 
Distorted fan i
nlet (left) and 
fan 
outl
et (right) total temperature profiles
 
for 
70% CFS.
 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
51
 
 
4.3.2.4
 
Tangential Flow Angle
 
The tangential flow angle is shown in 
Figure 
4
.
15
. The redistribution of the flow driven 
by the static pressure deficit shown in
 
Figure 
4
.
13
-
left creates a substantial tangential flow 
angle variation that varies with both circumferential position and radius. Large co
-
 
and 
counter
-
swirl regions of significant intensity
 
are present entering the fan. Larger tangential 
flow angles are pr
esent near the spinner, with reduced intensity 
radially 
out
wards
 
towards 
the tip. 
This radial gradient is stronger in the co
-
swirl region. 
The profile is slightly 
asymmetric with a larger counter
-
swirl region due to the higher static pressure
 
on that side 
of the profile (see 
Figure 
4
.
13
-
left)
. 
Co
-
swirl magnitude is higher due to 
spinner effects
. 
 
Station 5 of 
Figure 
4
.
15
-
right shows transmission of the
 
tangential
 
flow angles through 
the fan. Higher flow tu
rning with respect to undistorted conditions 
is seen 
exiting the co
-
swirl region and entering the counter
-
swirl region 
near
 
BDC. Reduced turning 
is observed 
exiting 
the 
counter
-
swirl region starting near 240
° for a large circumferential portion of the 
prof
ile for the core flow region, where the significant turning required 
near the blade root 
has become too excessive. Banded regions around 
the 
profile show the effect of the blade 
and engine geometry
 
(splitter, part
-
span shroud, and stiffener) appears to ent
rap flow 
exiting the fan.
  
The shroud
,
 
specifically
,
 
appears to reduce the amount of turning produced 
by the blades at 
nearly 
all circumferential locations.  
 
The tangential flow angle 
magnitude
 
has been reduced while passing through the fan
,
 
but the fan o
utlet conditions still present severe flow angle variation that will reduce the 
efficiency of the stator vanes downstream
, which are not in the proper orientation to 
efficiently turn this distorted fan outlet flow
 
in the axial direction
. This inefficiency 
will 
persist through the byp
ass duct, resulting in a system
-
wide reduction in operational 
efficiency. 
 
 
Figure 
4
.
15
. 
Distorted f
an inlet
 
(left) and 
fan 
outlet (right) tangential flow angles.
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
5
2
 
 
4.3.2.5
 
Radial Flow Angl
e
 
The radial flow angle
, shown in 
Figure 
4
.
16
,
 
is lower in magnitude
 
than the tangential 
flow angle at Station 4 but increases in severity through the fan
. 
At station 4
 
(
Figure 
4
.
16
-
left)
, the flow redist
ribution from the top of the profile to the bottom is shown with the high 
radially outward (positive) flow at the bottom of profile
 
centered around 165
°
. This 

sink 
point

 
is offset 
from BDC 
due to the 
non
-
axi
symmetric static pressure distribution 
upstream
 
of the fan
 
that was shown in 
Figure 
4
.
13
-
left
, an interesting development 
considering
 
the axisymmetric nature of the total pressure distortion entering the fan
. 
S
trong 
radially inward flow characteristics are shown in the right
-
h
alf of the profile below the 
part
-
span shroud
, centered at approximately 270
°
. This feature is also driven by the static 
pressure distribution which has a radial gradient in that region forcing flow radially inward.  
 
Station 5 of 
Figure 
4
.
16
-
right
 
shows the radial flow distribution after traversing the fan. 
The majority of the flow 
does not contain a strong radial component
, but a large 
circumferential band around the splitter radius shows a tende
ncy toward radially inward 
flow. T
his is caused by the 
static pressure gradient around the splitter
, creating
 
strong radial 
flow migration into the core of the engine. This indicates that the core compressor is heavily 
influencing the distorted flow development.
 
This thin band is the resul
t of a radial 
contraction and circumferential spreading of the large radial inward flow on the right half 
of the Station 4 profile. 
The radial outward flow sink point from Station 4 has been 
circumferentially shifted approximately 70
°
, which 
can be
 
calcula
ted from Equation 
(
4
.
4
)
. 
Since the 




 
ratio
 
is 
relatively constant
 
for increasing fan speed, the angular 
displacement is independent of 
CFS
. 
 
 
















 
(
4
.
4
)
 
 
Figure 
4
.
16
. 
Distorted f
an inlet (left) and 
fan 
outlet (right) radial flow angles
 
for 70% 
CFS
.
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
53
 
 
4.3.2.6
 
Upstream Flow Redistribution
 
Flow redistribution upstream of the fan face is the driving fo
rce for the high secondary 
flows measured
 
at Station 4
. The 
static
 
pressure gradient created by the distortion 
results 
in the formation of
 
large in
-
plane velocity components.
 
This is best shown in the plots of 
Figure 
4
.
17
.
 
Flow is
 
redistributing from the high pressure regions
 
in the top half of the 
profile
, 270
°
 
> 

 
< 
90
°
, to the lower pressure regions
 
at the bottom of the profile
, 90° 
< 

 
< 
270°, creating high tangential flow angles at 90° (co
-
swirl) and 270° (counter
-
swirl) an
d 
h
igh radial flow angles at 165
° (
tip
-
ward
). The non
-
symmetric nature of the of the 
secondary flow profile
 
is due to the offset, non
-
symmetric static p
ressure profile at the fan 
face, as
 
shown in 
Figure 
4
.
13
. This effect is most cle
arly seen in the sink point of 
Figure 
4
.
17
-
right  near 165°.  
 
 
Figure 
4
.
17
.
 
Distorted v
ector (left) and streamline (right) plots of the 
S
tation 4 in
-
plane 
velocity field
 
at 70% CFS
.
 
4.3.2.7
 
Incidence and Deviation
 
The change in incidence and deviation on the blades due to 
the 
distortion is show
n
 
in 
Figure 
4
.
18
, with incidence angle at Station 4 on the left and deviation angle at Station 5 
on the right
. The incidenc
e is heavily influenced by the location of 
the 
total pressu
re deficit, 
but the approximate
 
30
°
 
fan
-
wise 
offset of the positive 
incident 
region centerline shows a
n
 
increased sensitivity
 
toward counter
-
swirl.
 
Incidence is strongest when total pressure and 
co
unter swirl act in combination.
 
Reduced incidence near the hub around 90
° is in the area 
of strong co
-
swirl. 
 
Figure 4.18
-
right 
show
s
 
a 
negative to positive deviation swing in the 
majority of the 
profile and most heavily in the 
co
-
swirl region, beginning i
n the core and moving radially 
outward to the part
-
span shroud. 
The bulk of the co
-
and counter
-
swirl inlet flow regions 
show increased and reduced 
deviation
, respectively. 
Reduced 
deviation
 
is 
shown for the 
majority of the core region, with 
overturning pre
sent at
 
BDC at minimum total pressure
 
and on the left
-
side of the profile within the bypass
. 
Strong radial gradients are present 
due 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
54
 
 
to the fan blade and engine
 
geometry 
which heavily influence the 
propagation of the 
distortion through the fan
. 
The fan and
 
engine geometry also tend to collect the flow into 
bands around the annulus. 
 
 
Figure 
4
.
18
. 
Changes in
 
incidence (left) and deviation (right) angles
 
relative to 
undistorted inlet conditions at 70% CFS.
 
4.3.3
 
Fan 
Performance
 
Due to the complex nature of the distortion profile, a non
-
uniform fan performance 
distribution 
was
 
found for all fan performance metrics. 
Unless otherwise stated, figures are 
from the 70% CFS data sets as the profile distributions are independ
ent of fan speed, though 
the ranges will vary. 
Performance metrics for all speeds 
are
 
given in 
Appendix G
. 
 
 
4.3.3.1
 
Blade Loading
 
(






 
Distribution
 
The blade loading 
relative difference from undistorted conditions
 
is shown in 
Figure 
4
.
19
-
right
, with the 
tangential inlet flow angle at Station 4 shown for reference in 
Figure 
4
.
19
-
left. The loading 
shows a 
profile 
that 
is 
divided
 
into two distinct circumferential 
regions, 3
4
5
°
 
< 

 
<
 
1
3
5° where the loading is 
largely 
reduced, and 
1
3
5°
 
< 

 
<
 
3
4
5° 
where 
the 
loading is 
largely 
increased, particularly radially 
tip
-
ward
 
beyond the par
t
-
span shroud.
 
There 
exists a blending between these regions, but this trend applies to the majority of the 
profile.
 
 
The reduced loading 
area 
is within the region of high and transitional total pressure, 
but also 
co
-
rotating tangential flow (co
-
swirl). 
Th
e positive loading region of the profile is 
where the fan is exiting the low pressure region of the dist
ortion and being subjected to 
increased
 
pressure and increased counter
-
swirl. 
The increased 
counter
-
swirl 
will
 
aim to
 
increase the
 
aerodynamic loading
 
o
n the blades
. The 
higher magnitude
 
for
 
more highly 
loaded blades 
indicates that the counter swirl more heavily influences the loa
ding 
distribution, with a global increase of around 10%
. 
 
The heavy loading also makes the 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
55
 
 
blades more susceptible to premature
 
separation, which occurs in the 195°
 
< 

 
<
 
285° 
region. This is supported in the Lieblein diffusion factor calculation presented in Section 
4.3.3.5
 
 
Figure 
4
.
19
. 
Station 4 tangential flow angle (left) and fan blade lo
ading (right)
 
at 70% 
CFS.
 
4.3.3.2
 
Blade Loading 
Orbits
 
Due to the 
non
-
uniform
 
distribution of blade loading caused by the distortion, 
individual fan blade performance will vary as it completes a 
rotation
. This is shown in 
Figure 
4
.
20
 
wher
e the distorted and undistorted blade loading relationships are plotted 
together. The 
loading 
structure as a function of axial inlet flow
 
coefficient
 
is
 
shown in 
Figure 
4
.
20
-
left
, with good agreement 
between
 
all three speeds teste
d
. 
Data points are 
colored by circumferential position using the colorbar 
for reference on the far right of 
Figure 
4
.
20
 
and the Station 4 total pressure 
distribution
 
is 
shown for reference
 
in the bottom 
right
 
of 
Figure 
4
.
20
-
left
. Minimum inlet flow coefficient corresponds to 180
°
 
or BDC on 
the total pressure contour. The radial loading characteristic for the undistorted cases
 
is 
plotted as a dashed line for 60, 70, and 80% CFS
. 
 
Orbital paths for sample co
re and bypass flows are plotted for simplicity, though the 
orbital relationships ar
e preserved throughout the span 

 
higher loading sensitivity for 
increasing radius. Note 
that 
the maximum loading occurs at approximately 240
°
 
for both 
the core and bypass r
adii
, indicating 
that blade loading 
is more strongly influenced by 
tangential flow angle than inlet flow coefficient. Maximum loading also appears to be 
independent of radii, with all 
span
-
wise
 
locations peaking at approximately the sa
me 
circumferential lo
cations. A
 
typical behavior occurs at the part span shroud, as observed in 
the undistorted characteristic. 
 
The orbital paths are plotted as function of 
the 
tangential inlet flow 
coefficient, 
normalized by axial inlet flow in 
Figure 
4
.
20
-
right. The inlet swirl angle profile at Station 
4 is plotted in the bottom right of the figure for reference. As shown, the maximum loading 
occurs at approximately 240
°, the location of maximum counter
-
rotating swirl.
 
Minimum 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
56
 
 
blade loading occurs a
t approximately 135
°
, the location of maximum co
-
rotating swirl. 
Orbit structure is preserved across the blade span, with increasing sensitivity for increasing 
radius, though much less aggressive than with inlet flow
 
coefficient.
 
The lower the stage loadin
g
,
 
the more likely the flow will remain attached to the blades 
and ope
rate as efficiently as possible, with 
modern engines typically 
keeping the 
 
maximum 
blade loading to 0.75 
[30]
. 
While the undistorted profile show
s
 
blade loa
ding values 
exceeding this maximum in the core, the distorted flow will have large circumferential 
regions that exceed this limit at higher 
span
-
wise
 
locations. Also note the location of the 
undistorted operating point in relation to the orbits 

 
on the lo
wer 
spectrum of the 
blade 
loading orbital path. This means that the 
work input due to distortion is nearly exclusively 
higher in all areas of the profile 
and not limited to the undistorted region.
 
Despite nearly 
half the 
distorted 
profile 
being of 
high rec
overy 
that
 

loading is higher than the op
erating
 
point for the majority of the fan rotation
, indicating
 
the 
full annulus effect of localized distortions and the inability of the fan to recover from 
distortions as sever
e as 
the one
 
investigated
 
in this work
. 
 
 
Figure 
4
.
20
. Blade loading as a function of inlet flow coefficient (left) and tangential inlet 
flow ratio (right) for distorted and undistorted conditions
 
at all dis

.
 
4.3.3.3
 
Pressure Rise
 
Distribution
 
The 
total 
pressure rise
 
relative
 
difference from undistorted conditions is shown in 
Figure 
4
.
21
-
right, with the 
Station 4 
inlet total pressure plotted for reference in 
Figure 
4
.
21
-
left. 
The pressure rise distribution shows higher pressure rise in the lower pressure region 
of the distortion and 
a 
lower 
pressure 
rise in the higher pressure regions. It was shown in 
Figure 
4
.
12
-
right that the t
otal pressure uniformity at
 
the fan exit will have greater 
significance for lower pressure
 
fan
 
inlet regions. The severely reduced pressure rise radially 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
57
 
 
inward from the part
-
span shroud at BDC is due to the low 
total 
pressure exiting the fan in 
that regio
n
, caused by both excessive 
diffusion (supported in 
Figure 
4
.
23
-
left)
 
and low 
axial inlet velocity.
 
Note the 
rotated nature of the profile, approximately 15° in the fa
n
-
wise direction. 
 
 
Figure 
4
.
21
. 
Station 4 total pressure 
(left) and 
total pressure rise RDFU
 
(right)
 
at 70% 
CFS.
 
4.3.3.4
 
Pressure Rise 
Orbits
 
Total p
ressure rise orbits are plotted as functions of 
axial and tangential 
inlet flow 
coefficients
 
in 
Figure 
4
.
22
-
left and 
Figure 
4
.
22
-
right, respectively, for all three distortion 
speeds tested. 
The black dashed line is the undistorted conditions across the blade span 
with the circular black markers indicating the 
undistorted
 
fan speed
 
operation point and the 
diamond markers indicating the equivalent 
mass flow
 
conditions 
for
 
70% CFS. 
As shown, 
orbital structure is preserved for each speed but the range is increased
 
and shifted
 
for higher 
output values
. 
Well behaved
 
orbits can be seen in
 
the core region, with maximum pressure 
rise occurring at approximately 255
°
. Minimum pressure rise occurs at approximately 90
°
 
at the transition between high and low total pressure regions.
 
The majority of the bypass 
flow has maximum pressure rises near t
he minimum inlet flow coefficients at 
approximately 180
°
. Minimum pressure rise occurs at the location of maximum pressure 
at TDC. 
 
The discrepancy between the bypass and core maximum and minimum pressure rise 
locations is due to the significantly reduced 
outlet total pressure in the core, caused by 
increased fl
ow separation as a result of excessive diffusion
. This is supported 
by
 
the LDF 
calculation
 
(Equation 
(
3
.
23
)
 
and shown in Section 
4.3.3.5
)
, which exceeds 0
.65 in this 
region. Orbit
al
 
st
ructure
 
variation 
between the core and bypass 
is due to the increased fan 
speed at higher radii, where the relative difference between the axial inlet f
low and fan 
speed is increased and the pressure rise characteristic is mor
e sensitive
 
to changes in 
inlet 
axial flow. 
Notice how the core orbits contain both the equivalent fan speed and equivalent 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
58
 
 
mass flow
 
operating points while the bypass only co
ntains the equivalent fan speed, which 
indicates 
the core has 
a wider performance
 
range and more stable operation.
 
The pressure rise as a function of the inlet tangential flow ratio
 
for the 70% CFS
 
is 
shown in 
Figure 
4
.
22
-
right. For the bypass orbit, maximum pressure rise occurs at the 
minimum tangential flow 
ratio, slightly before BDC at 180
°
. Maximum pressure rise in the 
core corresponds to the location of maximum counter
-
swirl, at approximately 255
°
. 
Minimum pressure rise for both core and bypass occurs at approximately 105
°
, 
corresponding to the location of
 
maximum co
-
swirl.
 
The slope characteristic of the core flow orbit in 
Figure 
4
.
22
-
left suggests small 
pressure rise sensitivity to inlet flow for highly loaded blades, while the slope characteristic 
in 
Figure 
4
.
22
-
right suggests a more sensitive pressure rise distribution to inlet flow angles. 
An opposing
 
relationship exists for the bypass region
 
where a sharp sensitivity to inlet flow 
and a relaxed sensitivity to inlet flow angle. In summary: blade loadin
g affects the pressure 
rise sensitivity to inlet flow coefficient and inlet flow angle. 
 
 
Figure 
4
.
22
. 
Total p
ressure rise as a function of inlet flow coefficient (left) and tangential 
inlet flow ratio (righ
t) for distorted and undistorted conditions
 
at
 

.
 
4.3.3.5
 
Diffusion 
and Efficiency
 
The Lieblein diffusion factor
 
distribution
, calculated using 
Equation 
(
3
.
23
)
,
 
is shown 
in 
Figure 
4
.
23
-
left, displaying a simil
ar distribution to the blade loading profile of 
Figure 
4
.
19
-

aggressive
. 
Figure 
4
.
23
-
left
 
shows the overall LDF to be excessively
 
high in the 
more 
heavily loaded
 
regions, far beyond the 0.5 recommended 
value 
for non
-
separated flow
 
[29
]
. 
This 
suggests
 
the flow is separated in this region, but lower LDF 
elsewhere in the profile 
help offset
 
the separation
 
a
nd allow 
the 
blades to recover 
and 
avoid initiating a rotating 
stall phenomena and potentially fan surge. 
The 
difference in LDF
 
from undistorted 
conditions results in 
a localized decrease of 0.05 in limited areas to large area decreases up 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
59
 
 
to 
0.4
. The
 
near
 
entirely positive difference between conditions indicates that there is 
little
 
relaxation in diffusion across the rotor due to distortion. Even the high recovery region near 
TDC of the total pressure profile results in small increases in the diffusion coe
fficient, 
creating a more aerodynamically difficult environment 
nearly 
everywhere in the profile 
despite the limited extent of the distortion. 
 
 
The local efficiency of the fan is calculated following 
Equation 
(
3
.
22
)
, utilizing th
e 
l
ocal measured total pressure rise, the
 
inlet
 
total 
temperature
, and the work input to the 
rotor
. 
Although efficiency calculations in turbomachinery applications have been shown 
to 
carry large uncertainties [20], the 
reported 
trends and behaviors of the
 
pre
sent
 
measurements still provide valuable information on fan performance, as shown in 
Figure 
4
.
23
-
right
. 
Low efficiency values 
as a result of the counter
-
swirl 
is shown 
within
 
the 
majority of the bypass region for 
areas of high cou
nter
-
swirl (

)
. High 
efficiency values were obtained between the splitter and part
-


These blade geometry 
components appear to entrap efficiency 
values circumferentia
lly, creating
 
bands of 
consistent performance within the annulus. 
Efficiency characteristics are largely dependent 
on the inlet tangential flow angle; increased efficiency in co
-

and reduced efficiency in counter
-
swirl regions

 
The 
efficiency difference
 
compared to undistorted conditions ranges from a reduction of 30% to an increase in 15%, 
with a distribution 
nearly identical
 
to that shown in 
Figure 
4
.
23
-
right, with a global 
reduction
 
of 8.5%. 
 
 
Figure 
4
.
23
. 
Local 
Lieblein 
diffusion
 
factor
 
(
left
) and
 
local efficiency
 
(
right
)
 
under 
distorted conditions at 70% CFS
.
 
4.3.3.6
 
Compressor Map
 
A simulated performance map of the fan can be constructed ut
ilizing the experimental 
data collected. Although the 
mass flow
 
was not varied for the fan speeds under 
investigation, the 
mass flux
 
can be used to simulate
 
such a procedure. 
Figure 
4
.
24
-
left 
shows the pressure rise characteristic
 
as a function of the 
mass flux
, where the distorted 
conditions are shown in colored markers
 
and the undistorted in colored lines overlaid with 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
60
 
 
black markers
.
 
The colors correspond to CFS as shown in the 
colored 
label on the far right 
of the figure.
 
The di
storted pressure ri
se 
-
 
mass flux
 
characteristic follow
s
 
orbital paths like 
that shown in 
Figure 
4
.
22
-
left for each radial location along the 
blade 
span
, increasing in 
pressure rise for increasing blade span
.
 
To obtain a single ch
aracteristic 
line typically used for compressor maps, the distorted 
condition had to 
be 
averaged in two ways. The first incorporated a radial average, where 
all the data 
at each circumferential location was averaged at each radial measurement 
location acro
ss the blade span, 
resulting in a single pressure rise orbit, as shown in 
Figure 
4
.
24
-
right
. 
The final characteristic was obtained by sorting the radially averaged data
 
for 
increasing mass flux
 
and 
performing a two
-
point moving av
erage
, given in 
Equation 
(
4
.
5
)
 
and 
results 
shown in 
Figure 
4
.
25
-
left
. 
For Equation 
(
4
.
5
)
, 
Y
 
is the radially averaged data 
set, 
Z
 
is the two
-
point moving average, 
N
 
is the numbe
r of data points, and 
k
 
is the mass 
flux sorted index in data set 
Y
. 
 
 

















, 
for 
k
 
= 2:
N
 
(
4
.
5
)
 
 
Figure 
4
.
24
. Pressure rise 

 
mass flux characteristic, all radii (left) and radially ave
raged 
distorted conditions (right).
 
While the shape of the lines do no
t
 
replicate expected characteristic curves like 
Figure 
4
.
25
-
right
, they do 
indicate
 
the reduced performance expected 
from
 
distortion and its 
relation to undisto
rted 
performance
. 
If we recall from 
Figure 
1
.
2
, the distortion is expected 
to reduce the performance of the compressor for a given speed, and this reduction increases 
as de
sign conditions are approached 
[7]
. 
The results from 
Figure 
4
.
25
-
left show
 
the reduced 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
61
 
 
performance values, most specifically in the bypass
 
region
 
of the fan. 
The data from this 
investigation shows that the 
pressure
 
characteristic 
for the distortion 
lies appro
ximately 
midway between undistorted performance for equivalent fan speed (above
 
the distortion 
line
) and equivalent 
mass flow
 
(below
 
the distortion line
). 
This relationship can then be 
subsequently used to estimate what overcompensated 
fan operating
 
condit
ions are required 
to obtain the design condition performance
 
in the presence of distortion
, until the onset of 
stall.
 
 
Figure 
4
.
25
. Simulated compressor map (left) and scaled model data for the JT15D
-
1 
(righ
t).
 
4.3.4
 
Summary
 
The results presented in previous sections show the impact of total pressure distortion 
on 
fan 
performance 
for several 
engine speeds
. The reduced axial velocity upstream of the 
fan caused by the pressure deficit creates varying inlet 
conditions
 
integral to fan 
performance. The result
 
of this pressure deficit 
produced by the distortion screen results in 
downstream flow redistribution that 
creates
 
significantly high tangential and radial flow 
variations. This secondary flowfield distribution produ
ces varying fan performance 
profiles, whose trends are governed heavily by the total pressure and tangential inlet flow
 
characteristics
, as well as the 
design 
blade loading 
distribution
. 
 
Global 
fan 
performance calculations 
due to distortion
 
are
 
shown in r
elation to 
undistorted equivalent conditions in 
Table
 
4
.
1
 
The global average pressure ratio lies 
between the undistorted equivalent fan speed and equivalent 
mass flow
 
cases
 
(see Section 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
62
 
 
2.1.2
 
and 
Figure 
2
.
7
 
for reference)
, while the global averaged blade loading and efficiency 
show an 
average 
8.5% increase and a 11% reduction, respectively.
 
Table
 
4
.
1
. Global performance parameters 
for distorted and undistorted cases.
 
 
Fan 
Speed
 
Undistorted Equivalent 
Fan Speed
 
Distorted
 
Undistorted Equivalent 
Mass flow
 
Pressure Ratio
 
60
 
1.19
 
1.18
 
1.17
 
70
 
1.27
 
1.26
 
1.22
 
80
 
1.37
 
1.35
 
1.29
 
Blade Loading
 
 
60
 
0.47
 
0.51
 
0.46
 
70
 
0.47
 
0.52
 
0.47
 
80
 
0.49
 
0.52
 
0.48
 
Efficiency
 
[%]
 
60
 
78
 
69
 
79
 
70
 
79
 
70
 
79
 
80
 
79
 
73
 
79
 
 
Plots of 
local 
performance parameters as a function of circumferential position fo
r 
select core and bypass radii are
 
shown in 
Figure 
4
.
26
, along with the in
let total pressure 
(1
st
 
row) and the tangential inlet flow angle (2
nd
 
row)
 
for reference
. This plot configuration 
shows the variation of the performance metrics as the 
fan rotates through 
key distortion 
features.
 
The undistorted equivalent fan speed metric
 
or reference
 
is shown in the dashed 
blue line and the distorted parameter is shown in red. 
 
Pressure rise (3
rd
 
row) follows an atypical trend for the core flow, showing a closer 
matching with the tangential inlet flow angle than the inlet total pressure. 
The bypass shows 
a more expected trend, with the pressure rise inversely rela
ted to the inlet total pressure for 
as long as the flow can remain attached.  
 
Blade loading trends exclusively with tangential flow angle, decreasing in regions of 
co
-
rotating sw
irl and increasing in areas of counter
-
rotating swirl. 
Despite the relatively 
equal magnitude of the tangential flow angle
s
, unequal magnitudes of the blade loa
ding 
result from the distortion
 
that is heavily 
biased
 
to increased loading, which precedes the 
directional change of the flow angle by approximately 30° for both core and bypass regions. 
 
Efficiency in the core appears to 
be 
largely unaffected
 
by the trends of the inlet total 
pressure and tangential flow angle. 
The bypass region track
s
 
well with the
 
tangential flow 
angle, increasing for co
-
swirl and decreasing for counter
-
swirl. 
Both core and bypass 
regions show significant overall reductions in 
fan
 
efficiency. 
 
The Lieblein diffusion factor is inversely proportional to inlet total pressure, a 
conseq
uence of its dependence on the relative inlet flow angle to the blades which is 
directly tied to the 
inlet 
axial velocity. 
Lower axial velocity will tend the blades toward 
positive
 
stall, 
increasing the relative velocity to a larger percentage of the blade
 
speed, 
increasing the diffusion factor and reducing the blade stall margin. 
The bypass distribution 
favors the counter
-
swirl region, with peak values obtained at approximately 225°
, 
corresponding 
to 
the peak in blade loading
. 
 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
63
 
 
A 
few 
final note
s
 
referring 
to
 
Figure 
4
.
26
 
show
 
the full annulus impact of the distortion 
profile. Due to the severity and three dimensional nature of the distortion, a recovery point 
is never achieved for any of the fan performance parameters investigated. 
There are very 

performance settles back to undistorted equivalent fan speed
 
conditions. 
Blade loading and 
the diffusion factor are two exceptions
,
 
while the primary pe
rformance metric of fans, 
total 
pressure ratio, is constantl
y varying through a revolution.
 
Also, the loading and diffusion 
are nearly exclusively increased and decreased as a result of the distortion. While the 
pressure rise metric may be higher in some a
reas, this is 
a deceptive result as the efficiency 
and amount of work is reduced and increased, respectively. 
 
 
Figure 
4
.
26
. Distorted and undistorted p
erformance parameters trends for core and bypass 
flow
 
Table 
4
.
2
 
summarizes the fan performance metrics for measured inlet profiles, and 
corrected
-
uncorrected outlet conditions
 
as defined in Section 
3.7.1
. The bulk of the results 
Chapter 4
 
-
 
Results and Discussion
 
 
 
 
64
 
 
presented were for both mea
sured 
fan 
inlet profiles and corrected 
fan 
outlet conditions. The 
purpose of this table is to show the importance of 
measuring the fan inlet conditions such 
that accurate fan performance metrics can be calculated.
 
Contour plots of the different 
parameter c
alculations were not
 
plotted due to their strong similarity in shape to the 
corrected and inlet measured case. However, the data values varied significantly, especially 

Unmeasured
 

 
As
 
shown in
 
Table 
4
.
2
, the pressure ratio variation between 
cases
 
is negligible, as the 
pressure difference between Station 3 and Station 4 is very small, only 0.53% at 60% CFS 
and 1.2% at 80% CFS. Due to the high circumferential ex
tent of the pressure distortion at 
Station 4 and 
Station 
5, there is 
in
significant mismatching
 
between fan inlet to fan outlet 
comparisons, thus the global pressure ratios are similar. 
 
Blade loading and efficiency, which are both heavily dependent on the 
inlet tangential 
flow 
angle, show strong discrepancies between measured 
and 
unmeasured
 
fan inlet 
profiles. Relative differences between the calculated parameters are approxim
ately 8% for 
blade loading and 7
% for efficiency
, of nearly equal value as the dis
torted effect being 
investigated
. The discrepancies between loading and efficiency are not strongly dependent 
on fan speed. These results emphasize the importance of measuring the inlet flow profiles 
in order to calcula
te accurate performance results
, othe
rwise the effects of distortion can 
be 
exaggerated
. 
 
 
Table 
4
.
2
. 
Fan performance parameter comparison
s between corrected outlet conditions 
and unmeasured fan inlet. 
 
Fan 
Speed
 
Parameter
 
Corrected Outlet
 
Measur
ed Inlet
 
Uncorrected Outlet
 
Measured Inlet
 
Uncorrected Outlet
 
Unmeasured Inlet
 
60
 
 
Pressure 
Ratio
 
1.18
 
1.18
 
1.18
 
Blade 
Loading
 
0.51
 
0.51
 
0.55
 
Efficiency
 
69
 
70
 
65
 
70
 
Pressure 
Ratio
 
1.26
 
1.26
 
1.26
 
Blade 
Loading
 
0.52
 
0.52
 
0.55
 
Efficiency
 
70
 
71
 
65
 
8
0
 
Pressure 
Ratio
 
1.35
 
1.34
 
1.35
 
Blade 
Loading
 
0.52
 
0.53
 
0.56
 
Efficiency
 
73
 
71
 
66
 
 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
65
 
 
Chapter 5
 
Conclusions
 
and Recommendations
 
This dissertation presented
 
a 
full scale experimental study on the 
effects of 
total 
pressure
 
inlet distortion on fan performance.
 
A tota
l pressure distortion representative of a 
boundary layer ingesting serpentine inlet was introduced upstream of a modified Pratt and 
Whitney 
Canada 
JT15D
-
1 turbofan engine. Using a calibrated 5
-
hole probe
, flow 
parameters necessary to characterize the fan i
nlet and 
fan 
outlet flowfields were measured 
and 
results 
calculated. 
The results presented in this dissertation have highlighted the 
necessity of the extensive testing procedures required to accurately determine the effect 
total pressure distortion has on 
fan performance. Using the measurements and analyses 
previously described, the following sections summarize 
the conclusions obtained from this 
study as well as providing 
recommendations for future research and development. 
 
5.1
 
 
Summary
 
Principal conclusions a
ssociated with the following discussions are shown in italics.  
These conclusions will be summarized and collected at the end of this section.
 
5.1.1
 
Full Flowfield Measurements
 
The automated traverse system utilized in this experiment allowed data to be taken 
qu
ickly and repeatedly. The fine radial and circumferential control 
enabled
 
dense 
measurement grids to be obtained for all test cases investigated. 
Without this capability, 
full flowfield measurements would not have been possible and a sparse measurement gri
d 
would have 
been insufficient
 
to local
ly
 
characterize blade performance.
 
This dense 
measurement grid allowed sufficient data to be collected to characterize the full flow field, 
such that linear interpolations between data points could be 
accurately 
made.
 
This enabled 
accurate determination of the distortion effect 
distribution and global 
characterization. 
 
The results presented show the full annulus effect of the distortion despite its limited 
circumferential extent
. This 
non
-
symmetric
 
circumferential var
iation required a fine 
circumferential grid to accurately measure the distortion changes as the fan passed through 
the distortion. 
Fan blade and engine geometry characteristics 
thus require
 
a fine radial 
grid in order to accurately capture bypass and core 
flow behavior
. The part
-
span shroud 
and blade stiffener significantly 
changed the
 
fan performance and quantifying
 
this
 
required 
sufficient measurement density around 
both 
components
. 
 
Acceptable radial and circumferential density at the fan outlet also 
all
owed for
 
accurate 
determination of fan outlet angle distribution
. Such information would be required for 
optimal design of outlet stator vanes for maximum stage efficiency.
 
 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
66
 
 
5.1.2
 
Fan Inlet Profile
 
Total pressure distortion screens are typically calibrated in wi
nd tunnel configurations 
in the absence of 
downstream turbomachinery
 
where the distortion profile measured one 
diameter downstream of the screen plane is established 
as the fan inlet profile for future 
engine testing 
[7]
. In th
is configuration flow angularity is so low as to be considered 
negligible. These screens are then placed in fan testing environments where the screen is 
located
 
two
 
diameter
s
 
upstream of the fan inlet and the measured profile from the wind 
tunnel configura
tions 
is
 
expected to 
hold true
. 
While this has been shown to be a good 
assumption for lower speed flows in the case of total pressure, it is not a good assumption 
for higher speed flows
 
[23]
 
and ignores the considerable radial 
and tangential flow 
variation
 
entering the fan
. 
 
Utilizing this AIP plane in screen calibration tests and assuming the measured profile 
is applicable to fan performance tests is inaccurate. Proper characterization of the fan inlet 
profile requires calibrat
ion in 
the pe
rformance testing configuration
 
with
 
f
ull flowfield 
parameter measurements directly upstream of the fan.
 
Without this correct inlet profile, 
global work and efficiency calculations are 
exaggerated
 
and local performance trends and 
relationships
 
to the incoming profile are lost. 
It was found in this 
research
 
that the global 
pressure rise calculation was impartial to fan inlet measuring, but work and efficiency are 
strongly affected. 
 
Fan inlet measurements
 
also provide 
important 
detail
 
of the dis
tortion development
, 
shown in this dissertation as the 
substantial flow redistribution upstream of the fan that 
resulted in high tangential and radial flow components
.
 
The secondary flow profiles 
produced by the screen 
were
 
a result of the fan interaction 
with the distortion profile, 
creating a 
non
-
axisymmetric 
static pressure profile at the fan inlet and driving flow 
redistribution, 
creating non
-
axisymmetric
 
tangen
tial and radial flow variation.
 
This off
-
centerline convergence point for the secondary flowf
ield 
was
 
a result of the fan interaction 
as the total pressure distortion was axisymmetric. 
The spinner geometry also heavily 
influences the distortion development, particularly with core flow and co
-
swirl regions. 
The measured inlet profile will also allo
w for blade redesign, such that the optimal blade 
shape can be constructed that bests performs for the total pressure and secondary flow 
distortion it is subjected to. 
 
5.1.3
 
Distortion Transfer
 
The results obtained at the fan outlet, Station 5, showed 
that 
pres
sure and secondary 
flow variations
 
persisted
 
in the distortion profile.
 
While the 
radial flow angle variation 
was mitigated
 
through the fan, 
total pressure, static pressure, and tangential flow angles 
were exacerbated
 
in magnitude. 
The structure of the dis
tortion was also significantly 
altered; 
fl
ow phenomena entering the fan experienced large radial contractions and 
circumferential spreading. Much of the flow phenomena was banded in extrema regions 
due to the interaction of the distortion with the
 
blade ge
ometry; splitter, part
-
span shroud, 
and stiffener. 
 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
67
 
 
5.1.4
 
Distortion
-
Fan Interaction
 
Due to the coupled effect of the fan and distortion screen, the fan is
 
acting as a driver 
of the 
distortion development in a way that varies from traditional wind tunnel screen 
calibration. Varying work input and 
non
-
uniform
 
pressure outputs contribute to 
non
-
uniform
 
inlet pressure profiles that drive flow redistribution further upstream of the fan 
inlet. This flow redistribution creates secondary flow 
characteristics
 
that furthe
r com
plicate 
the distortion profile
.
 
 
Since the fan and screen form a coupled system, the total and static pressure distortion 
as well as the resulting secondary flow effects become 
inseparable.
 
Therefore
,
 
obtaining 
isolated total pressure effects on fan p
erformance 
is
 
not feasible, and 
secondary flow effects 
cannot be decoupled from the total pressure distortion in 
traditional 
fan performance 
measurement configurations.
 
 
5.1.5
 
Fan Performance Ground Tests
 
Fan speed was shown not to 
significantly 
vary 
the 
results
 
for the parameters and 
measurements
 
of interest
 
in this investigation
. 
Trends and 
distributions of the fan behavior 
in varying regions of the distortion were
 
preserved between 
speeds
. Thus,
 
lower speed 
distortion tests can be performed to reduce cost and 
risk, while preserving the flow 
phenomena 
behavior for speeds up to the engine stall limit. 
The
 
benefit of performing 
varying distortion speed
 
measurements would be for estimation of fan performance in 
relation to 
the 
compressor map. 
A minimum of two disto
rtion speed tests could be 
completed at reduced speed to obtain the operating line 
characteristic
, allowing estimation 
of the distortion 
mass flow
 
at 
varying
 
engine speeds. The losses at each engine speed could 
then be estimated by averaging the undistorte
d equivalent fan speed and equivalent 
mass 
flow
 
pressure rise. 
This estimation would be accurate at lower speeds
,
 
but unstable at near 
design conditions as the stall limit of the fan is rapidly approached as a consequence of the 
distortion. 
 
5.1.6
 
Local Performa
nce Calculations
 
This dissertation showed the radial and circumferential variations of fan performance
 
as a result of the 
total pressure
 
distortion. This was accomplished by having full inlet profile 
measurements to accurately calculate performance 
metrics
.
 
The strong association of 
blade 
loading
 
and efficiency with inlet tangential flow angle, and the strong association between 
pressure rise and inlet total pressure could not have been determined without the fan inlet 
profile. The behavior of the fan at va
rying positions within the distortion
 
could be 
determined and the sensitivity to each inlet flow parameter could be evaluated
. 
 
The local performance 
metrics resulted in
 
improved
 
global 
calculations
 
where 
estimated compressor maps could be constructed to s
how 
the 
increased operating 
conditions required to meet design conditions.
 
One such map was created, and it was 
observed the distortion performance was bounded between undistorted equivalent fan 
speed and equivalent mass flow conditions. To offset the dist
ortion degradation to 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
68
 
 
performance, however, 
the t
arget design speed 
would need to be
 
less than 
100%
. The 
distortion profile inv
estigated in this dissertation featured 
a 
maximum 
90% recovery 
at the 
AIP and 
resulted in a 15% reduction in the
 
engine
 
operating
 
range
 
(
global 
engine stall at 
85% CFS)
. 
 
Local performance calculations also show the range of performance metric variation
s
. 
For example; performance increases were observed in the low total pressure
 
and counter 
swirl regi
ons but
 
came at the cost of 
redu
ced 
efficiency
. Large efficiency reductions around 
30% occupied a substantial region of the profile distribution in areas of high counter
-
swirl.
 
Varying behavior between the core and bypass 
was observable through this data set
, 
providing a comparison for h
ighly loaded areas 
of the blade root to the less loaded areas of 
the blade mid
-
span and tip.
 
It was shown that the 
blade loading
 
affect
ed
 
the pressure rise 
sensitivity to inlet flow c
oefficient and inlet flow angle, an effect distinguished between 
the core
 
and bypass regions. The core 
output 
flow 
also showed varying magnitude and 
distribution
 
form the bypass
, necessitating 
varying and 
possibly
 
nonlinear
 
stator design 
 
Finally, 
local performance calculations 
show
 
the full annulus effect of localized 
distorti
ons
. For example; blade loading
 
due to distortion
 
was found to be 
nearly 
consistently
 
higher in all areas of the profile and not limited to 

undistorted

 
region. 
Despite nearly half the distorted profile being of high recovery and considered 


blade 
loading 
was
 
higher than the 
undistorted 
op point for the majority 
of the fan rotation, indicating the full annulus effect of localized distortions and the inability 
of the fan to recover from distortions as severe as 
the one investigated
.
 
5.1.7
 
Se
condary Flow Impact
 
The impact of a total pressure distortion of this magnitude on fan performance is severe. 
In a research effort where inlet duct architectures are designed to minimize fan efficiency 
losses to 1%, an 8% global fan efficiency loss for tot
al pressure distortions indicates 
research efforts need to be redirected to the components most susceptible to performance 
degradation as a result of next generation aircraft architectures. While pressure rise can be 
shown to be higher across the fan in so
me areas, this is a deceptive performance 
improvement as efficiency reductions and increased blade loading are observed 
almost 
everywhere in the profile
, a direct result from the secondary flow developed due to the total 
pressure distortion
. 
 
The counter
-
s
wirl developed in the distortion most heavily impacted the fan 
performance, contributing to increased diffusion that 
affected
 
the entire profile, resulting 
in a global blade loading increase of 10%. High recovery regions of the profile that would 
have thou
ght to have been undistorted showed
 
increased diffusion, evidence of the full 
profile 
effect
 
of localized distortions. Also, the fan responded unequally to the relatively 
equal inlet swirl angles magnitudes, creating a bias toward increased blade loading i
n the 
counter
-
swirl regions that was not offset by 
equal 
unloading in the co
-
swirl region
 
which 
resulted in the global loading increases and efficiency decreases. 
 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
69
 
 
5.2
 
Conclusions
 
The primary poin
ts of note from the previous section
 
are summarized below:
 
1)
 
Full 
flowfield measurements
 
allow 
for 
sufficient data to be collected to 
characterize the
 
fan response to the distortion;
 
allowing for 
A) 
accurate linear 
interpolation 
between
 
data points in both the radial and circumferentia
l 
directions, B) 
flow parameter dist
ributions to be created upstream and 
downstream of the rot
or, and C) fan blade geometry effects to be discerned
.
 
2)
 
Measuring the 
fan 
inlet profile 
is essential to ensure
 
that
 
performance
 
and 
efficiency 
are calculated correctly and that local trends and relat
ionships are 
accurately determined.
 
Significant differences were calculated for fan 
performance in the absence of fan inlet measurements. The fan inlet 
measurement plane 
also provides detail on the 
distortion development
 
and the 
substantial flow redistribu
tion upstream of the fan that resulted in high 
tangential and radial flow components
 
due t
o the total pressure distortion that 
created blade incidence swings of +/
-
 
10
°
. 
 
3)
 
Fan outlet measurements showed the persistence of the pressure and secondary 
flow var
iations in the distortion profile. While the radial flow angle variation 
was mitigated through the fan, total pressure, static pressure, and tangential 
flow angles were exacerbated in magnitude
 
4)
 
S
econdary flow effects cannot be decoupled from total pressure
 
distortion in 
traditional 
fan perform
ance measurement configurations
, and it is these 
primarily these 
secondary flow characteristics that drive the local fan response, 
not the total pressure. 
 
5)
 
The counter
-
swirl region of the distortion led to local effici
ency reductions of 
30%, blade loading increases of 60%, and 
diffusion factors exceeding 0.7. 
 
 
6)
 
Trends and distributions of the fan behavior in varying regions of the distortio
n
 
were preserved between speeds, indicating that
 
lower speed tests can be 
perform
ed to study distortion interactions to reduce cost while prese
rving the 
phenomena of interest.
 
7)
 
Local performance metrics resulted in improved global calculations
,
 
the range 
of performance metric variations, and the full annulus effect of localized 
distorti
ons.
 
5.3
 
Recommendations
 

collaborated with 
G
eneral 
E
lectric 
on a project that improved compressor aerodynamic 
efficiency that 
could offer 
2.5% 
savings 
in fuel burn 
[35]
 
for uniform inlet flow. If such 
extensive research effort is being pursued to improve modern engine designs, it is the 

 
strongest recommendation
 
that more research must be invested in the area of 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
70
 
 
distortion tolerant fan design, where t
his work showed global efficiency reductions 
can 
reach 
8% for distorted inlet flows characteristic of next generation airframe 
architectures
. 
This
 
future engine
 
operating environment create challenges that 2.5% 
efficiency 
increases 
will not offset, and fut
ure research needs to focus on the problems that cause the most 
performance loss. 
 
The following sections highlight several other recommendations obtained from this 
work. 
 
5.3.1
 
Distortion Screen Calibration
 
Recommended procedures for future investigators includ
e the following:
 
 

 
When possible, measure the fan inlet profile with equal measurement density 
as that performed for the fan outlet. Distortion profiles measured in wind tunnel 
configurations in the absence of fan effects will develop differently when place
d 
upstream of the fan.
 
The improved accuracy on performance calculations more 
than offset the difficulty and risk associated with the procedure.  
 

 
Use caution when attempting to perform isolated pressure and swirl distortion 
fan response tests. Total press
ure distortions inherently produce their
 
own 
secondary flow variations that can be of equal magnitude to the swirl distortion 
produced by the inlet that is being simulated. 
 

 
Include flow angularity measurements at the AIP to determine the severity 
of 
the s
econdary flow formation farther upstream of the fan face. This will indicate
 
whether the mixing of pressure regions creates secondary flow further upstream 
than the fan face.
 
5.3.2
 
Fan Performance Ground Testing
 

 
H
igh speed or near design 
engine 
tests are not nec
essary 
for character
izing
 
fan 
response to distorted inflows. This will reduce experiment costs
,
 
and lower 
speeds will allow for more intense distortions to be safely studied without fear 
of engine stall. 
 

 
Multiple speeds should be studied for evaluating th
e distortion over 
a
 
large 
portion of the compressor map. Such information will be useful 
in
 
estimating 
the fan performance at higher speeds and developing a model for the fan 
behavior with respect to the distortion. 
 

 
T
esting on a full scale engine
 
should b
e performed whenever possible, 
as 
isolated electric driven fan designs lack the 
coupled 
low pressure turbine 
operation that exists in full
-
scale engine experiments. 
 
5.3.3
 
Distortion Tolerant 
Fan Design
 

 
Although the part
-
span 
fan blade 
shroud on this engine did 
significantly alter 
the fan performance, the riblet
-
type feature near the tip of the blade showed far 
Chapter 5
 
-
 
Conclusions
 
and Recommendations
 
 
 
 
71
 
 
less departure from expected performance. Thus, a riblet
-
type feature could be 
considered for future
 
fan designs 
to provide aeromechanical support
 
without
 
suffering the significant losses in fan performance that part
-
span shrouds 
produce.
 

 
Future research should include a
djust
ing
 
the stator vane outlet profile p
osition 
based on the fan outlet flowfield for improved aerodynamic efficiency. 
 

 
Conduct research i
nto developing a fan model that will assist in designing
 
a 
distortion
-
tolerant fan
. The 
extensive data set generated through this research
 
can be used
 
as 
experimental 
val
idation. The recommendations and conclusions 
obtained from this work should be impleme
nted for the initial design. 
 
5.4
 
Closing Statements
 
The goal of this work 
was
 
to contribute to the body of research that has been pushing 
towards designing a distortion tolerant fan
 
for next generation aircraft architectures
. How 
that is accomplished has stil
l yet to be determined. Shrouds, placed in at the mid
-
span in 
the engine investigated here, or at the top of the fan blades create a more stable blade that 
increases the stall margin, allowing for more complex distortions
 
but at the cost of 
reduced 
perform
ance
,
 
increasing fuel consumption over a flight envelope.
 
A more coupled solution between inlets and fans could result, where fans are 
individually designed for maximum performance for specific inlets. Inlet and exit guide 
vanes are 
tailor
-
designed as well
 
to operate most efficiently for the distortion produced by 
the inlet at design conditions. This would be an expensive solution between airframe and 
engine 
designers,
 
but 
would 
allow for the greatest performance and efficiency benefit.
 
 
The work presented 
here shows that the performance and stability margin reductions 
are substantial for the severity of distortion investigated and will require 
extensive
 
redesigns on modern fans to power the aircrafts 
long
-
term 
effectively. The overall 
performance benefit ma
y offset the reduction in fan performance, but the durability of the 
engines will be significantly affected and this must be taken into consideration.
 
T
he solution
 
to performance and operability challenges
 
inherent in blended wing body 
architectures 
will i
nvolve several aspects mentioned above, and new ones gathered from 
quality full scale experiments. The current state of the art is just beginning to understand 
the impact these distortions have on the performance of the engines and more research is 
require
d before a 
complete 
solution can be found. The work presented here highlights some 
of the inaccuracies of current e
xperimental testing methodology
. U
tilizing the 
recommendations mentio
ned in the previous subsections
,
 
more accurate 
measurements
 
can 
be made,
 
leading to improved fan performance and flow development calculations and 
further characterize
 
distortions and their impact on fan 
performance
. 
 
 
 
 
 
Bibliography
 
 
 
 
72
 
 
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[36]
 
A. Treas
ter and M. Yocum. The calibration and application of 
fi
ve hole probes. 
Technical
 
Memorandum, (TM 78
-
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for Science and Engineering, Applied Research Laboratory.
 
 
Appendix A 
-
 
Five Hole Probe Calibration
 
 
 
 
75
 
 
Appendix A
 
Five
-
Hole
 
Probe Calibration
 
The 
following sections detail 
a subset of the five
-
hole probe 
analysis
, specifically the 
Mach
 
number dependency study and the final interpolation maps. 
 
A.1
 
Five
-
Hole Interpolation Maps
 
Final pressure coefficient maps for the five
-
hole probe are shown in 
Figure 
A
.
1
. These 
maps were used to calculate the five
-
hole probe parameters discussed in Section 
3.5
. The 
map structure is well behaved and uniform through the bulk of the pitch and yaw angles 
calibrated. More
 
variability is shown at the outer edges of the grid domain but as shown in 
the blue experimental data markers, the engine data was captured exclusively within 
± 
20
°
 
yaw and 
± 
15
°
 
pitch bounds. 
 
 
 
Figure 
A
.
1
. Pressure coefficients for the five
-
hole probe calibration with experimental 
data (blue)
.
 
Appendix A 
-
 
Five Hole Probe Calibration
 
 
 
 
76
 
 
A.2
 
Mach Number Dependency
 
The five
-
hole probe calibration included 
several tests performed at multiple Mach 
numbers:
 
0.4, 0.5, and 0.6. 
The Mach numbe
r dependency was determined 
by comparing 
the difference between the numerical standard deviation from the experimental. 
 
The numerical derivative was calculated for each five
-
hole pressure coefficient: total 
pressure, static pressure, yaw angle, and pitch 
angle. The pressure coefficient relationships 
are of the form given in Equation 
(
A
.
1
)
, where 


 
represents the standard deviation from 
the mean and 


 
represents variability in the final result due to error sources in the 
measurement
. 
 
 

























 
(
A
.
1
)
 
The numerical variation between for the five
-
hole probe parameters was estimated 
using a Taylor series expansion as shown in Equation 
(
A
.
2
)
, where 



 
is the 
measurement 
u
ncertainty of 

-
test transducer calibrations. 
 
 
 




















 
(
A
.
2
)
 
This procedure was performed for each pressure c
oefficient and compared to the 
standard deviation of the experimental data for
 
all three speeds tested and is shown in 
Figure 
A
.
2
. 
 
Appendix A 
-
 
Five Hole Probe Calibration
 
 
 
 
77
 
 
 
Figure 
A
.
2
. 
Analytical and exp
erimental pressure coefficient 
variability
.
 
Table 
A
.
1
 
summarizes the RMS standard deviations fo
r the numerical and experimental 
data sets. As shown, the experimental standard deviation between pressure coefficient 
measurements is 
comparable to the numerical results
 
for the 0.5 and 0.6 Mach number 
cases
, making them indistinguishable
. The 0.4 Mach number case shows a higher 
variability due to the relatively higher
 
uncertainty percentage of the nominal pressure 
values. 
Since the vary
ing speeds could not be distinguished experimentally, it was 
concluded that the Mach number effects on the probe in the ranges expected were 
negligible and calibration maps were constructed using an average of all three speeds and 
functions of pitch and ya
w angle only. 
 
Table 
A
.
1
. 
Numerical and experimental RMS measurement standard deviation.
 
 
Numerical
 
Experimental 
-
 
All Speeds
 
Mach Number
 
0.4
 
0.5
 
0.6
 
CP Total
 
0.4
 
0.13
 
0.09
 
0.12
 
CP Static
 
0.11
 
0.07
 
0.05
 
0.09
 
CP Yaw
 
0.44
 
0.26
 
0.18
 
0.2
0
 
CP Pitch
 
0.12
 
0.08
 
0.05
 
0.05
 
Appendix 
B 
-
 
Uncertainty Propagation
 
 
 
 
78
 
 
Appendix B
 
Uncertainty Propagation
 
The 
uncertainties associated with 
the 
calculated parameters detailed in Sections 
3.5
 
and 
3.6
 
were derived using the numerical 
sequential perturbation 
technique. This is 
summarized below and follows from
 
Figliola and Beasley 
[31]
. 
 
Beginning 
with a
 
relationship among 
an
 
independent variable, Y, and a finite number 
of de
pendent variables, X, that number from 1 to L
, such that:
 
 















 
(
B
.
1
)
 
Perturb each independent variable in the positive manner by adding its corresponding 
uncertainty, 


.
 
 






















 
 






















 
 

 
 




















 
(
B
.
2
)
 
Perturb each independent variable in the negative manner by subtracting its 
corresponding uncertainty, 


.
 
 






















 
 






















 
 

 
 




















 
(
B
.
3
)
 
Using a central difference schem
e, the dependent variable perturbation with re
spect to 
each independent variable uncertainty can be calculated for
 

. 
 
 
















 
(
B
.
4
)
 
 
The overall uncertainty can then be calculated by performing a root sum 
square for all 
dependent perturbations. 
 
 
Appendix 
B 
-
 
Uncertainty Propagation
 
 
 
 
79
 
 
 
















 
(
B
.
5
)
 
 
 
Appendix 
C 
-
 
Five Hole Probe Analysis
 
 
 
 
80
 
 
Appendix C
 
Five
-
Hole Probe 
Analysis
 
C.1
 
Structural
 
A geometric
 
parameterization was performed on the 
five
-
hole probe 
used throughout 
this 
research 
to maximize structural integrity while minimizing instrument intrusiveness. 
A brief description of the final probe dimensions are given in 
Section 
2.2
, a design result 
from the structural analysis performed. 
Figure 
C
.
1
 
details the design steps taken to achieve 
this
, utilizing Ansys
®
 
CFX and S
tatic 
S
tructural
 
analyses
. 
 
Figure 
C
.
1
-
left shows the probe in the fan inlet measurement configuration where the 
maximum amount o
f probe surface area is exposed to the flow
, representing the 
structurally 
weakest position 
of the probe during the experiment.
 
The probe behaves like a cantilevered 
beam, with a 
cylindrical 
support at the compression fitting mounted to the top of the fan 
case. 
To obtain the radial pressure distribution, a 30
° wedge CFD model was performed 
with the probe modeled as a fixed boundary
, with the inlet flow corresponding to the 
highest CFS tested (80%), and the spinner was modeled as a rotating boundary. 
The mod
el 
domain is shown in 
Figure 
C
.
1
-
center
. 
The resulting pressure profile on the probe 
was then 
input into the static structural analysis using a c
ylindrical support
 
at the top of the probe
 
where the compression fitting would be loc
ated.
 
A schematic of the analysis is shown in 
Figure 
C
.
1
-
right
.
 
Sensing head diameter and reinforcement tube diameter were selected 
based on this procedure. The tube wall thickness was a standard used for United Sensor

 
for their 
pressure probes and was a fixed parameter in the analysis. 
 
 
 
Figure 
C
.
1
. 
Five
-
hole probe structural analysis: experimental setup (left), 30
°
 
section 
CFD model
 
(center)
, and structural anal
ysis deformation (right).
 
 
Appendix 
C 
-
 
Five Hole Probe Analysis
 
 
 
 
81
 
 
C.2
 
Modal
 
A modal analysis was performed on the probe geometry to indicate any blade passing 
frequencies that should be avoided during testing. The results are summarized in 
Figure 
C
.
2
, with the 
modal analys
is shown in the top
-
right for increasing insertion length. The 
blade passing frequencies
 
(BPF)
 
of interest are shown in the bottom
-
right, with only the 1
st
 
mode of the 
6

 
of concern. Insertion lengths between 5 and 6 inches near 
80% CFS we
re observed with caution with no vibratory probe effects found. The PTFE 
ferrule compression fitting at the fan case insertion helped dampen the vibratory effects of 
the flow over the probe. 
 
 
 
Figure 
C
.
2
. Five
-
hole probe modal analysis
. 
 
C.3
 
Time Response
 
The time response of the five
-
hole probe pneumatic measurement system was estimated 
using a gated valve system in which a known pressure source was opened to
 
the 
measurement system 
by electron
ic control of a gate valve. 
The pressure input was placed 
on the probe, followed by 25 feet of pneumatic tubing, and terminated in the transducer 
input port. 
Synchronizing the gate valve with the data acquisition, an estimation of the 
response time of the 
pressure probe system can be obtained. 
 
The results of the test are shown in 
Figure 
C
.
3
 
with the time domain shown on the left 
and the frequency domain shown on the right. The standard score given on the ordinate of 
Figure 
C
.
3
-
left is calculated using Equation  
(
C
.
1
)
, where X is the recorded pressure data, 

 
is the population mean of X, and 

 
is the population standard deviation of X.
 
Appendix 
C 
-
 
Five Hole Probe Analysis
 
 
 
 
82
 
 
 
 






 
(
C
.
1
)
 
The results from 
Figure 
C
.
3
 
show that for a 0.25% error of the input pressure, the rise 
t
ime is approximately 0.5 seconds with a settle time of approximately 0.8 seconds.
 
The frequency response, shown in 
Figure 
C
.
3
-
right, is
 
obtained by performing a fast 
F
ourier transform of the pressure data, X. This calculation as s
hown in Equation 
(
C
.
2
)
, 
where 
n
 
is represents the sample length. 
 
 


























 
(
C
.
2
)
 
Where
 
 








 
(
C
.
3
)
 
Figure 
C
.
3
-
right shows a peak amplitude of approximately 0.5 Hz, ind
icating a two 
second response time for the measurement system utilized. These pressure results can also 
be applied to the Station 1 and Station 3 data as it utilizes a similar measurement 
methodology. 
 
 
Figure 
C
.
3
. Time (left) and frequency (right) of the pressure measurements from the five
-
hole probe. 
 
 
Appendix 
D 
-
 
Fan Blade Modal Analysis
 
 
 
 
83
 
 
Appendix D
 
Fan Blade Modal Analysis
 
The blade modal analysis mentioned in Section 
2.4.1.1
 
resulted from a modal analysis
 
performed in Ansys
®
 
using the geometry obtained from a 3D scan. 
The blade was modeled 
as titanium, with compression supports at the 
pressure and suction sides of the 
part
-
span 
shroud
,
 
and a fixed support at the blade root. 
A similar analysis was run with 
frictionless 
supports at the pressure and suction sides of the part
-
span shrouds but the results were 
nearly indistinguishable to the compression only support. The compression only support 
boundary condition provided the most realistic setup to experimenta
l conditions and was 
therefore used in the analysis. 
Figure 
2
.
14
 
is reproduced below for convenience. 
 
 
Figure 
D
.
1
. Results of blade modal analysis.
 
 
Appendix 
E 
-
 
Fan Blade Geometry
 
 
 
 
84
 
 
Appendix E
 
Fan Blade Ge
ometry
 
The JT15D
-
1 fan blade geometry was reverse engineered by way of a 3D scan. This 
geometry was then imported into a CAD software package to fix coordinate locations. 
Figure 
E
.
1
 
details the process used to deconstruct
 
the scan
 
geometry. The scanned 
geometry was imported in Unigraphics
®
 
and a common data coordinate reference was 
created at the intersection of the three planes (
Figure 
E
.
1
-
left), one that was normal to the 
fir
-
tree 
base front
 
face
, one th
at was bi
-
normal to the base front, and one that was tange
n
t 
to the bottom
 
of the
 
 
base. Using the bottom tangent datum plane as a reference, datum 
planes were created 
span
-
wise
 
along the blade, creating 2D curves at the intersections of 
the plane with the
 
blade (
Figure 
E
.
1
-
center)
. Finally, circles tangent to the pressure and 
suction sides of the intersections were created at the leading and trailing edges of the blade 
as shown in 
Figure 
E
.
1
-
right. A line
 
was created through the center 
of these
 
circles and 
extended beyon
d the end of the blade where it
s angle offset with respect to the axial flow 
direction
 
could be measured (

). 
 
 
Figure 
E
.
1
. Fan blade geometry deconstruction.
 
A summary of the blade geometry obtained from the process outlined above is given 
in 
Table 
E
.
1
. 
The solidity w
as calculated by measuring the chord of the blade at each 
intersection and dividing by the base width. The 
data 
in 
Table 
E
.
1
 
was used throughout the 
dissertation. 
 
 
 
Appendix 
E 
-
 
Fan Blade Geometry
 
 
 
 
85
 
 
Table 
E
.
1
. JT15D
-
1 fan blade geometry.
 
Slice 
Number
 
Span
-
wise Distance from
 
Engine 
Centerline to Slice [




]
 
Leading Edge 
Blade Angle [°]
 
Trailing Edge 
Blade Angle [°]
 
Solidity
 
1
 
0.44
 
43.3
 
5.3
 
2.3
 
2
 
0.47
 
43.8
 
1.4
 
2.3
 
3
 
0.50
 
46.9
 
3.5
 
2.4
 
4
 
0.52
 
49.1
 
12.3
 
2.4
 
5
 
0.53
 
51.7
 
19.1
 
2.4
 
6
 
0.56
 
49.6
 
27.0
 
2.4
 
7
 
0.59
 
49.8
 
31.8
 
2.5
 
8
 
0.62
 
50.5
 
36.5
 
2.5
 
9
 
0.68
 
51.7
 
41.1
 
2.5
 
10
 
0.72
 
52.0
 
44.0
 
2.6
 
11
 
0.76
 
56.7
 
47.1
 
2.6
 
12
 
0.80
 
59.8
 
52.0
 
2.7
 
13
 
0.84
 
59.3
 
55.6
 
2.8
 
14
 
0.90
 
61.7
 
57.3
 
2.9
 
15
 
0.95
 
68.8
 
61.7
 
3.0
 
16
 
0.98
 
73.1
 
63.5
 
3.1
 
17
 
0.99
 
81.7
 
NA
 
NA
 
 
 
Appendix 
F 
-
 
Des
ign Work Input
 
 
 
 
86
 
 
Appendix F
 
Design 
Work Input 
 
The undistorted blade loading
 
distribution given in 
Figure 
4
.
7
-
left was the result of the 
design requirements for near
-
constant work input on the flow. A constant work input 
creates a more uniform outlet flow that results in greater efficiency. The undistorte
d work 
distribution for the JT15D
-
1 is shown below in 
Figure
 
F
.
1
. The distribution shows some 
radial and geometrical variation
 
but trends toward uniformity. Variations are shown to be 
exacerbated at higher fan speeds. 
   
 
 
Figure
 
F
.
1
. Undistorted work input design distribution.
 
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
87
 
 
Appendix G
 
Multiple Speed 
Data
 
 
Given below is a collection of the results presented in Section 
4.3
 
for all of the speed
s 
investigated. 
The results in Section 
4.3
 
were for 70% CFS case for brevity and the 
remaining distortion conditions are given below for reference. 
 
G.1
 
AIP
 
 
Figure 
G
.
1
. 
Relative AIP Data 
from undistorted 
all speeds investigated. 
 
G.2
 
Station 4
 
 
Figure 
G
.
2
. Relative Station 4 total pressure 
from undistorted 
all fan speeds investigated.
 
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
88
 
 
 
Figure 
G
.
3
. 
Relative Station 4 
static
 
pressure
 
from undistorted
 
for all fan speeds 
investigated.
 
 
Figure 
G
.
4
.
 
Station 4 tangential 
flow angle difference from undistorted for all fan speeds 
investigated.
 
 
Figure 
G
.
5
.
 
Station 4 radial flow angle difference from undistorted for all fan speeds 
investigated.
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
89
 
 
 
Figure 
G
.
6
. Station 4 incidence angle difference from undistorted for all fan speeds 
investigated.
 
G.3
 
Station 5
 
 
Figure 
G
.
7
. Relative Stati
on 5 total pressure
 
from undistorted
 
for all fan speeds 
investigated.
 
 
Figure 
G
.
8
.
 
Relative Station 5 static pressure
 
from undistorted
 
for all fan speeds 
investigated.
 
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
90
 
 
 
Figure 
G
.
9
.
 
Station 5 tangential flow angle difference from undistorted for all fan speeds 
investigated.
 
 
Figure 
G
.
10
. 
Station 5 radial flow
 
angle difference from undistorted for all fan speeds 
investigated.
 
 
Figure 
G
.
11
. Station 5 deviation
 
angle difference from undistorted for all fan speeds 
investigated.
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
91
 
 
G.4
 
Fan Performance
 
 
F
igure 
G
.
12
. Relative blade loading 
difference from undistorted for all fan speeds 
investigated.
 
 
Figure 
G
.
13
. Pressure ri
se
 
for all fan speeds investigated.
 
 
Figure 
G
.
14
.
 
Lieblein diffusion factor
 
for all fan speeds investigated.
 
 
Appendix 
G 
-
 
Multiple Speed Data
 
 
 
 
92
 
 
 
Figure 
G
.
15
. Efficiency
 
for all fan speeds investigated.
 
