University of Tennessee, Knoxville
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Exchange
Masters Theses
Graduate School
12-2010
Investigation of the Effects of Inlet Swirl on
Compressor Performance and Operability Using a
Modified Parallel Compressor Model
Nicholas Joseph Fredrick
[email protected]
Recommended Citation
Fredrick, Nicholas Joseph, "Investigation of the Effects of Inlet Swirl on Compressor Performance and Operability Using a Modified
Parallel Compressor Model. " Master's Thesis, University of Tennessee, 2010.
http://trace.tennessee.edu/utk_gradthes/795
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To the Graduate Council:
I am submitting herewith a thesis written by Nicholas Joseph Fredrick entitled "Investigation of the
Effects of Inlet Swirl on Compressor Performance and Operability Using a Modified Parallel Compressor
Model." I have examined the final electronic copy of this thesis for form and content and recommend
that it be accepted in partial fulfillment of the requirements for the degree of Master of Science, with a
major in Mechanical Engineering.
Basil N. Antar, Major Professor
We have read this thesis and recommend its acceptance:
Robert W. McAmis, Milton W. Davis
Accepted for the Council:
Carolyn R. Hodges
Vice Provost and Dean of the Graduate School
(Original signatures are on file with official student records.)
To the Graduate Council:
I am submitting herewith a thesis written by Nicholas Joseph Fredrick entitled “Investigation of
the Effects of Inlet Swirl on Compressor Performance and Operability Using a Modified Parallel
Compressor Model.” I have examined the final electronic copy of this thesis for form and content
and recommend that it be accepted in partial fulfillment of the requirements for the degree of
Master of Science, with a major in Mechanical Engineering.
Basil Antar, Major Professor
We have read this thesis
and recommend its acceptance:
Robert McAmis
Milton Davis
Accepted for the Council:
Carolyn R. Hodges
Vice Provost and Dean of the Graduate School
(Original signatures are on file with official student records.)
Investigation of the Effects of Inlet Swirl on Compressor
Performance and Operability Using a Modified Parallel
Compressor Model
A Thesis
Presented for the
Master of Science Degree
The University of Tennessee, Knoxville
Nicholas Joseph Fredrick
December 2010
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DEDICATION
This thesis is dedicated to my wife Hillary. I would not have been successful in
this or in all of my other endeavors without her constant support.
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ACKNOWLEDGMENTS
I would like to thank all those that helped and provided their support during the
preparation of this thesis. Dr. Rob McAmis provided his support during my studies at
UTSI and helped find and develop the topic of the thesis reported here. Dr. Milt Davis
provided the thesis topic as well as much needed guidance during the preparation. He
was a wealth of information and could be counted on to answer any question that arose.
I would also like to thank Dr. Alan Hale and Mr. Jason Klepper. Alan and Jason’s
knowledge of the modeling tools used in this thesis as well as their willingness to
provide support and answer questions were a great help to me. I would also like to
thank the Major Professor of my thesis committee, Dr. Basil Antar. His support during
the preparation of this thesis was a great help to me. I would also like to thank Arnold
Air Force Base and Aerospace Testing Alliance for providing time and resources to
complete this thesis
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ABSTRACT
Serpentine ducts used by both military and commercial aircraft can generate
significant flow angularity (inlet swirl) and total pressure distortion at the engine face.
The impact of inlet swirl on the engine performance and operability must be quantified
to ensure safe operation of the aircraft and propulsion system and to define installed
deficiencies.
Testing is performed over a wide range of flight conditions in the
propulsion system flight envelope in order to quantify these effects. Turbine engine
compressor models are based on experimental data which can be collected at a limited
number of discrete operating points. These models can be used as an analysis tool to
optimize the engine test plan and help during validation of the design.
The Dynamic Turbine Engine Compressor Code (DYNTECC) utilizes parallel
compressor
theory
and
compressor performance.
quasi-one-dimensional
Euler
equations
to
determine
In its standard form, DYNTECC uses user-supplied
characteristic stage maps in order to calculate stage forces and shaft work for use in the
momentum and energy equations. These maps are typically developed using
experimental data. These maps can also be created using characteristic codes such as
the 1-D Mean Line Code (MLC) or the 2-D Streamline Curvature Code. The MLC was
originally created to predict the performance of individual compressor stages and
requires greatly reduced computational time when compared to 2-D and 3-D models.
This thesis documents work done to incorporate the MLC into DYNTECC as a
subroutine. The combine DYNTECC/MLC was then used to analyze the effects of inlet
swirl on the fan performance and operability of the Honeywell F109 turbofan engine.
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The code was calibrated and validated using the F109 cycle deck. Additional code
validation was performed using experimental data gathered at the United States Air
Force Academy (USAFA). F109 fan maps were developed for various cases of inlet
swirl and results were presented showing shifts in corrected mass flow, fan pressure
ratio and fan stability limit.
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TABLE OF CONTENTS
1.0 INTRODUCTION ....................................................................................................... 1
2.0 LITERATURE SEARCH .......................................................................................... 10
2.1 SWIRL CHARACTERISTICS OF S-DUCT ENGINE INLETS ............................... 10
2.2 EFFECTS OF SWIRL ON ENGINE PERFORMANCE AND OPERABILITY ........ 12
2.3 CONCLUSIONS OF THE LITERATURE SEARCH .............................................. 18
2.4 PROBLEM STATEMENT ..................................................................................... 19
3.0 BACKGROUND....................................................................................................... 20
3.1 DYNTECC ............................................................................................................ 20
3.2 1-D MEAN LINE CODE ........................................................................................ 25
3.3 F109 TURBOFAN ENGINE .................................................................................. 29
4.0 APPROACH ............................................................................................................ 31
4.1 OUTLINE OF APPROACH ................................................................................... 31
4.2 INTEGRATION OF THE 1-D MEAN LINE CODE INTO DYNTECC ..................... 32
4.3 MEAN LINE CODE CALIBRATION ...................................................................... 34
4.4 VALIDATION OF COMBINED DYNTECC/1-D MEAN LINE CODE ..................... 38
4.5 PRESENTATION OF RESULTS .......................................................................... 41
5.0 RESULTS ................................................................................................................ 42
5.1 BULK SWIRL ........................................................................................................ 43
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5.2 PAIRED SWIRL .................................................................................................... 47
6.0 SUMMARY AND CONCLUSIONS .......................................................................... 55
7.0 RECOMMENDATIONS ........................................................................................... 58
LIST OF REFERENCES ............................................................................................... 59
VITA .............................................................................................................................. 66
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LIST OF FIGURES
Figure 1: Axial Compressor Velocity Triangle ................................................................. 2
Figure 2: Bulk Swirl ......................................................................................................... 3
Figure 3: Paired Swirl ...................................................................................................... 3
Figure 4: Parallel Compressor Theory Concept [15] ....................................................... 7
Figure 5: Compressor Characteristic Map ....................................................................... 8
Figure 6: Wellborn, Reichert and Okiishi Experiment Apparatus [16] ............................ 11
Figure 7: Secondary Flow Velocity Vectors at the AIP Down-Stream of S-Duct [17] .... 12
Figure 8: Swirl Pattern at the AIP Behind a Delta Wing Swirl Generator [19] ................ 14
Figure 9: Turning Vane Concept for Generating Bulk Swirl [21] .................................... 16
Figure 10: Swirl Chamber Concept for Generating Twin Swirl [21] ............................... 16
Figure 11: DYNTECC Compression System Control Volume Model [14]...................... 21
Figure 12: F109 Turbofan Engine ................................................................................. 30
Figure 13: F109 Rotor Relative Total Pressure Loss .................................................... 36
Figure 14: F109 Rotor Exit Deviation Angle .................................................................. 36
Figure 15: F109 Stator Relative Total Pressure Loss .................................................... 37
Figure 16: F109 Stator Exit Deviation Angle ................................................................. 37
Figure 17: F109 Cycle Deck Stall Line Rotor Diffusion Factor ..................................... 38
Figure 18: DYNTECC/MLC Predicted and F109 Cycle Deck Specified Fan Map
Comparison (No Swirl) .................................................................................................. 39
Figure 19: DYNTECC/MLC Predicted and USAFA Measured Fan Map Comparison (No
Swirl) ............................................................................................................................. 39
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Figure 20: DYNTECC/MLC Predicted and F109 Cycle Deck Specified Fan Stall Line
(No Swirl) ...................................................................................................................... 40
Figure 21: DYNTECC/MLC Predicted and USAFA Measured Fan Stall Line (No Swirl)41
Figure 22: Loss in Stability Pressure Ratio (ΔPRS) ...................................................... 42
Figure 23: DYNTECC/MLC Prediction of the Effect of Co-Swirl on F109 Fan
Performance and Operability ......................................................................................... 45
Figure 24: DYNTECC/MLC Prediction of the Effect of Counter-Swirl on F109 Fan
Performance and Operability ......................................................................................... 46
Figure 25: Loss in Stability Pressure Ratio (ΔPRS) as a Function of Bulk Swirl Angle 47
Figure 26: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl
and 10° Twin-Swirl on F109 Fan Performance and Operability at 53% Fan Speed ...... 49
Figure 27: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl
and 10° Twin-Swirl on F109 Fan Performance and Operability at 62% Fan Speed ...... 50
Figure 28: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl
and 10° Twin-Swirl on F109 Fan Performance and Operability at 71% Fan Speed ...... 51
Figure 29: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl
and 10° Twin-Swirl on F109 Fan Performance and Operability at 84% Fan Speed ...... 52
Figure 30: Twin Swirl Segment Diffusion Factor at Constant Fan Speed ...................... 53
Figure 31: Loss in Stability Pressure Ratio (ΔPRS) as a Function of Percent Referred
Fan for 10° Counter-Swirl and 10° Twin-Swirl ............................................................... 54
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NOMENCLATURE
α
β1
Swirl angle
Inlet relative flow angle
β2
Blade exit relative velocity flow angle
β2’
Blade exit metal angle
γ
δ
ΔPRS
ν1
ν1R
Ratio of specific heats
Blade exit angle deviation
Loss in stability pressure ratio
Blade inlet absolute tangential velocity
Blade inlet relative tangential velocity
ν2
Blade exit absolute tangential velocity
ν2R
ρ
σ
ω
Ϛ
A
a1
Blade exit relative tangential velocity
AEDC
AIP
AVDR
CFD
Cp
Arnold Engineering Development Center
Aerodynamic interface plane
Axial velocity density ratio
Computational fluid dynamics
Specific heat at constant pressure
DOE
Dr
Design of Experiments
Rotor diffusion factor
DYNTECC
E
e
Fb
Dynamic Turbine Engine Compressor Code
Energy function
Internal energy
Blade force
FX
HB
Axial force distribution acting on the control volume
Total enthalpy of the bleed flow
HPC
IGV
IMP
LPC
M1
High pressure compressor
Inlet guide vanes
Impulse Function
Low pressure compressor
Inlet Mach number
Density
Blade solidity
Blade relative total pressure loss
Cross section correlation factor
Area
Inlet speed of sound
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M1R
Inlet relative Mach Number
M2R
Blade exit relative Mach number
MFF1
Blade inlet relative mass flow function
MFF2
Blade exit relative mass flow function
MLC
PR
PR1
PRDS
Ps
1-D Mean Line Code
Stage pressure ratio
Undistorted stability pressure ratio
Distorted stability pressure ratio
Static pressure
Force of the engine case acting on the control volume
Ps1
Blade inlet static pressure
Ps2
Blade exit static pressure
Pt1
Blade inlet total pressure
Pt1R
Blade inlet relative total pressure
Pt2
Blade exit total pressure
Pt2R
Blade exit relative total pressure
Q
R
rhub
Rate of heat addition to the control volume
Gas constant
Radius at the hub
rtip
Radius at the blade tip
SAE
SW
TR
Ts1
The Society of Automotive Engineers
Rate of shaft work
Stage temperature ratio
Inlet static temperature
Tt1
Inlet total temperature
Tt1R
Inlet relative total temperature
Tt2R
Blade exit relative total temperature
u
U1
Axial velocity
Blade entrance wheel speed
U2
Blade exit wheel speed
USAFA
V1
United States Air Force Academy
Blade inlet absolute velocity
V1R
Blade inlet relative velocity
V2
Blade exit absolute velocity
V2R
Blade exit relative velocity
x
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Vm1
Blade inlet axial velocity
Vm2
Blade exit axial velocity
W
WB
Mass flow rate
Inter-stage bleed mass flow per distributed length
xi
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1. INTRODUCTION
Turbine engines generate thrust by compressing incoming air in the inlet and
compressor, mixing that air with fuel and igniting the air/fuel mixture in the combustor,
then expanding the high pressure and temperature air through a turbine and nozzle.
Though centrifugal compressors are used in smaller engines, most large military and
commercial turbine engines use axial flow compressors [1]. Axial flow compressors
compress the air by passing it through a series of rotating airfoils called rotor blades and
stationary airfoils called stator vanes. Just like aircrafts wings, rotor blades have an
angle of attack called the incidence angle. The incidence angle is the angle between the
velocity vector of the flow relative to the rotor blade and the camber line of the rotor
blade and is shown in Figure 1. Compressor performance and operability are affected
by variations in the conditions of the flow entering the compressor such as pressure and
temperature distortion and flow angularity. If the pressure ratio across the compressor
is raised high enough, or if the incidence angle of the rotor blades becomes too large,
flow over the rotor blades will separate from the suction surface resulting in stall [2].
The entrance to a turbine engine compressor is often referred to as the
aerodynamic interface plane (AIP). Flow angularity at the AIP can have both radial and
circumferential velocity components [3]. Of these two components, the circumferential
component of the absolute velocity vector (often referred to as swirl) has the strongest
effect on compressor performance because the angle of this component (swirl angle, α
in Figure 1) can affect the incidence angle of the rotor blade [3]. Most swirl patterns at
the AIP of a turbine engine can be grouped into two main types: bulk swirl and paired
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Rotor
Stator
Β2 ’
V2R
β2
v2
α2 Vm2
v2R
V2
U
Β1 ’
V1R
v1R
β1
α1
Vm1
v1
V1
V
Vm
VR
U
α
= Absolute Velocity Entering the Stage
= Axial Velocity Entering the Stage
= Relative Velocity Entering the Stage
= Rotor Blade Velocity
= Angle of the Absolute Velocity Vector (Swirl Angle)
v
vR
β
β’
β1 – β1 ’
= Absolute Tangential Velocity
= Relative Tangential Velocity
= Angle of the Relative Velocity Vector
= Blade Metal Angle
= Incidence Angle
Subscripts 1 and 2 represent Blade Row Entrance and Exit
Figure 1: Axial Compressor Velocity Triangle
swirl [3]. Bulk swirl is shown in Figure 2 and consists of a flow that is rotating either in
the same direction as engine rotation (co-swirl) or in the opposite direction of engine
rotation (counter-swirl). In the case of bulk swirl, the entire flow field at the AIP is
rotating in the same direction. Paired swirl, shown in Figure 3, is composed of two
vortices rotating in opposite directions. One vortex is a co-swirl vortex while the other is
a counter-swirl vortex. If the two vortices are symmetric, then the paired swirl is referred
to as twin swirl, else it is referred to as offset swirl. There are other types of swirl that
will not be addressed such as tightly wound vortices that for from proximity to the
ground or on the corners of the engine inlet.
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+
At Constant Radius
Swirl Angle
Co-Swirl
Compressor
Rotation
Swirl
0°
-
180°
360°
Circumferential Position
+
At Constant Radius
Compressor
Rotation
Swirl
Swirl Angle
Counter-Swirl
0°
-
180°
360°
Circumferential Position
Figure 2: Bulk Swirl
+
At Constant Radius
Compressor
Rotation
Swirl
Swirl Angle
Twin Swirl
0°
-
180°
360°
Circumferential Position
+
At Constant Radius
Swirl
Compressor
Rotation
Swirl Angle
Offset Swirl
0°
-
180°
Circumferential Position
Figure 3: Paired Swirl
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360°
Inlet guide vanes (IGVs) are used to change the angle of the flow entering the
compressor at the AIP. When operating properly, IGVs will add co-swirl (α > 0) to the
flow which will move the compressor away from stall. In the past it was not necessary
to simulate swirl at the AIP of a turbine engine during ground testing because of the
relatively straight inlet systems and the incorporation of IGVs into turbine engine
designs. Current aircraft designs may incorporate S-ducts with sharp bends into the
engine inlet systems in order to hide the engine face from radar waves. Investigations
have been performed in order to characterize the effects the S-ducts have on the flow
properties at the AIP of turbine engines. These studies have shown that flow separation
in the S-duct results in swirl at the AIP [4]. In some cases, the swirl can be severe
enough to cause flow separation on the IGVs which effects engine operability by
causing an additional loss in stability margin [5].
Engines without IGVs such as the RB199 in the Tornado fighter aircraft [6] or
high by-pass ratio turbofan engines are more sensitive to swirl at the AIP than those
with IGVs. This is because swirl will change the incidence angle of the first stage
compressor blades on the engine.
Engine operability problems arose during initial
development of the Tornado [6]. Despite extensive modeling and full scale inlet/engine
tests, during initial flight testing engine stalls were encountered in the engine to the left
of the pilot at subsonic flight speeds at high angles of attack and stalls were
encountered in the engine to the right of the pilot at supersonic speeds at angles of
attack approximately equal to zero. During the initial inlet/engine compatibility testing,
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only steady-state instantaneous total pressure distortion was measured at the AIP. The
distortion coefficients for the Tornado inlet were high but were within allowable limits.
These limits were derived from correlations of engine stall to instantaneous pressure
distortion using statistical methods and could vary from manufacturer to manufacturer
[6].
The Society of Automotive Engineers (SAE) has published a guideline to address
engine inlet total pressure spatial distortion, ARP 1420 [7].
This guideline is
supplemented by SAE Aerospace Information Report AIR1419 [8] and additional reports
have been issued by the SAE S-16 committee regarding inlet planar waves [9] and inlet
temperature distortion [10]. None of these reports address inlet swirl. Because of the
sensitivity to swirl displayed by engines without IGVs, the SAE S-16 committee is
currently developing an aerospace information report that addresses the topic. Along
similar lines, Davis, Hale and Beale [11] have made an argument for the enhancement
of ground test techniques in order to better simulate swirl at the AIP of an engine in
flight citing the effect of inlet swirl on engine performance and operability.
Several
methods for simulating inlet swirl in turbine engine ground tests are under development
at the Arnold Engineering Development Center (AEDC) [12, 13]. Evaluating the effect
of inlet swirl on turbine engine compressor performance during ground tests would
reveal engine performance and operability issues before initial flight testing which would
ensure safe operation of the engine/aircraft and avoid costly airframe and engine inlet
modifications.
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To fully understand the effects of swirl on turbine engine compressor
performance and operability, engine testing must be performed over the entire engine
flight envelope. The test points can be optimized using turbine engine compressor
models that are validated using experimental data at a limited number of discrete
operating points. Because these models are based on experimental data, they can be
used to optimize test points which in turn will improve the experiment.
Total pressure distortion and swirl at the AIP of a turbine engine are often
interdependant and areas of high swirl correspond to areas of low total pressure [3].
Compressor models can be used to decouple total pressure distortion and swirl such
that the effect of swirl alone on compressor performance and operability can be
evaluated. This would be very difficult to do in a test environment but is a relatively
simple input for a compressor model.
A type of compressor model often used in the study of inlet distortion is called a
parallel compressor model. Parallel compressor models sub-divide compressor control
volumes into parallel or circumferential segments that can have different inlet boundary
conditions. Each segment acts separately but in parallel with each other and exit to the
same boundary condition [5]. An illustration of the parallel compressor theory concept
is shown in Figure 4.
Parallel compressor models require source terms such as blade forces, shaft
work, bleed flows, and energy addition or subtraction due to heat transfer in order to
properly model the compression system [14]. These source terms are often presented
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COMPRESSOR MODEL
DISTORTION
PATTERN
LOW PRESSURE
HIGH PRESSURE
Figure 4: Parallel Compressor Theory Concept [15]
in the form of a compressor characteristic map, an example of which is shown in Figure
5.
For a given corrected mass flow and corrected speed, these maps supply the
pressure and temperature rise across the compressor. These characteristic maps are
typically developed using experimental data. If experimental data is unavailable, these
maps can be created using characteristic compressor codes such as the MLC or the 2D Streamline Curvature Code [15].
The compressor characteristic maps required to run the parallel compressor
models are an inherent disadvantage when attempting to model inlet swirl.
As
mentioned previously, compressor performance is dependent upon the incidence angle
of the rotor blades. The presence of swirl at the AIP will change the incidence angle of
the first stage rotor blades on turbine engines that are not equipped with IGVs. When
swirl is present, new compressor characteristic maps must be developed that include
the affects of swirl [5]. In the past, compressor characteristic maps with multiple speed
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Compressor Temperature Ratio
Compressor Pressure Ratio
Lines of Constant
Corrected Speed
Stability Line
Corrected Mass Flow
Figure 5: Compressor Characteristic Map
lines were developed for different cases of swirl. Should the inlet conditions to the
compressor change, data from the compressor characteristic maps were interpolated to
obtain the stage characteristic data at the desired inlet condition. The development of
the multiple characteristic maps increased the amount of time required to model swirl
while interpolating the data increased uncertainty in the final output.
Rather than develop multiple compressor characteristic maps outside of the
parallel compressor code for different cases of swirl, a faster and more efficient method
would be to develop the source terms for the desired test case using a function
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incorporated into the parallel compressor code. This would also increase accuracy of
the model output because stage characteristics would be developed by the MLC
internally for the case being studied eliminating the need for interpolation. In 2003,
Grady Tibboel integrated a one-dimensional compressor stage characteristics code into
a parallel compressor code called DYNTECC [15].
The objective of the research reported herein was to integrate the MLC into
DYNTECC as a subroutine. This work was based on the work performed by Tibboel.
The modified DYNTECC parallel compressor code used the MLC to determine F109
stage characteristic point-by-point in order to analyze the effects of various types of inlet
swirl on F109 compressor performance and operability.
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2. LITERATURE SEARCH
2.1 SWIRL CHARACTERISTICS OF S-DUCT ENGINE INLETS
The Tornado fighter aircraft encountered stall in both engines at different flight
conditions during prototype flight testing resulting from the swirl pattern induced at the
AIP by the inlet duct [5, 6]. Engine and airframe manufacturers as well as United
Kingdom research groups investigated the engine/airframe compatibility issues both
theoretically and experimentally.
Aulehla [6] compiled the main results of these
investigations and was able to draw conclusions from those results. Aulehla found that
the inlets on most conventional supersonic fighter aircraft generate some sort of swirl.
Through the investigations, it was found that bulk swirl could be attenuated using simple
solutions such as intake fences. It was found that the attenuation of paired swirl was
much less than that of bulk swirl using the intake fences because of the stable nature of
paired swirl. Flight testing revealed that engines without IGVs are sensitive to even
small changes in the magnitude of counter-swirl.
Guo and Seddon [4] performed experiments on a model s-duct intake with both
horizontal and vertical offsets in an effort to determine the characteristics of the flow
exiting the duct. The model was installed in the 7’x5’ wind tunnel at the Aeronautical
Engineering Laboratory of Bristol University. Guo and Seddon found that at high angles
of attack, flow separation occurred on the lip of the intake. This separation resulted in
the formation of strong bulk swirl at the AIP.
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Figure 6: Wellborn, Reichert and Okiishi Experiment Apparatus [16]
Wellborn, Reichert and Okiishi [16] used the experimental apparatus shown in
Figure 6 located at the NASA Lewis Research Center Internal Fluid Mechanics Facility
to investigate the characteristics of flow through a diffusing s-duct.
Unlike the
experimental apparatus used by Guo and Seddon, this apparatus was equipped with a
contraction nozzle to uniformly accelerate the flow in the settling chamber, ensuring
laminar flow into the s-duct. The results of the experiment showed that curvature of the
duct induced pressure driven secondary flows which resulted in paired swirl at the AIP.
This result was different than the result obtained by Guo and Seddon because of the
absence of flow separation at the entrance of the s-duct.
Loeper and King [17] used computational fluid dynamics (CFD) and Design of
Experiments (DOE) to investigate the effects of four S-duct CFD geometric parameters
as well as the S-duct entrance Mach number on inlet performance. The four geometric
parameters evaluated were offset distance, overall duct length, area expansion ratio
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and elliptical aspect ratio at the entry. The study revealed that strong vortices at the
AIP, an example of which is shown in Figure 7, developed in short length ducts.
2.2 EFFECTS OF SWIRL ON ENGINE PERFORMANCE AND OPERABILITY
2.2.1 Experimental Studies
After engine stalls were encountered during prototype flight testing of the
Tornado, Flitcroft, Dunham and Abbott [18] performed an experiment in order to
determine whether bulk counter-swirl at the engine AIP resulted in higher levels of flow
distortion entering the high pressure compressor (HPC). To perform the investigation,
Flitcroft, Dunham and Abbott measured the transmission of inlet flow distortion through
a representative military fan that was not equipped with IGVs. A gauze screen was
Figure 7: Secondary Flow Velocity Vectors at the AIP Down-Stream of S-Duct [17]
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used to generate a total pressure distortion while turning vanes were used to generate
bulk counter-swirl. The representative fan was a 3 stage low-hub/tip ratio research
machine. The flow distortion exiting the fan module was measured and compared to
the distortion entering at the AIP in order to determine whether the distortion was being
intensified or attenuated by the swirl. The results of the experiment showed that the
time averaged total pressure distortion entering the fan module was attenuated by as
much as 45% by the bulk counter-swirl. Though the average total pressure distortion
exiting the fan was attenuated by the swirl, the turbulence of the flow exiting the fan with
swirl was three times the level than the turbulence of the flow exiting the fan with total
pressure distortion alone.
In 1991, Pazur and Fottner [19] investigated the influence of inlet swirl distortions
on the performance of the low pressure compressor (LPC) of a LARZAC 04 turbofan
engine. The LARZAC 04 is a twin spool turbo fan engine with a two-stage fan that is
not equipped with IGVs [19]. Twin swirl at the AIP was generated using a delta wing
swirl generator with a 60° sweep.
The swirl generator was mounted 1.5 engine
diameters from the AIP. The swirl pattern at the AIP generated is shown in Figure 8.
Data was gathered with the swirl generator at various angles of attack. The swirl angle
was a function of the swirl generator angle of attack and would increase with angle of
attack. Swirl angle was independent of fan speed. The investigation performed by
Pazur and Fottner revealed that fan pressure ratio, mass flow rate and isentropic
efficiency decrease as swirl intensity (swirl angle) increases. Their investigation also
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Figure 8: Swirl Pattern at the AIP Behind a Delta Wing Swirl Generator [19]
revealed that no total pressure distortion exists at the exit of the LPC in the
circumferential direction. The effect of the decrease in pressure ratio and mass flow
rate is to shift the constant speed lines on the compressor map down and to the left.
Pazur and Fottener did not have the capability to throttle the LPC, so they developed a
formula to compute the constant speed lines with distorted flow on the compressor map
using the steady state operating points with inlet swirl distortion and compressor maps
developed with undistorted flow as a basis.
Schmid, Leinhos and Fottner [20] investigated the effects of a typical inlet
distortion pattern generated by the intake diffuser of an engine for hypersonic flight on a
LARZAC 04 turbofan engine. The subsonic section of the intake duct of the Hypersonic
Transport System Munich was modeled in ¼ scale and the swirl at the engine AIP was
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measured. The model revealed that the s-duct diffuser generated a strong bulk swirl
pattern at the AIP. A non-symmetric half delta wing swirl generator with one edge
parallel to the main flow equipped with a winglet to prevent the formation of a vortex
was used to simulate the bulk swirl pattern at the AIP of the LARZAC 04 engine. The
LPC map with co-swirl and counter-swirl was developed by throttling the LPC by
reducing the by-pass nozzle area. The investigation revealed that at high LPC speeds,
both co-swirl and counter-swirl lead to a decrease in surge margin, LPC pressure ratio,
and mass flow rate.
A delta wing swirl generator was used to generate the desired swirl patterns in
the two previously cited references.
Davis, Beale and Sheoran [21] detailed two
difference type of swirl generator currently under development at AEDC. Turning vanes
are under investigation for generating bulk swirl while a swirl chamber is under
investigation for the generation of twin swirl [21]. Turning vanes, shown in Figure 9,
resemble a set of inlet guide vanes and would feature variable blade incidence angle
and twist angle in order to provide remotely variable swirl angle [21].
The swirl
chamber, shown in Figure 10, operates in place of a bellmouth and forces the flow to
enter the inlet duct tangentially so that an internal circumferential flow is established
[21].
The swirl angle is changed by changing the entrance openings of the swirl
chamber. Both concepts have been analyzed using computational fluid dynamics and
functional prototypes are under development.
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Figure 9: Turning Vane Concept for Generating Bulk Swirl [21]
Figure 10: Swirl Chamber Concept for Generating Twin Swirl [21]
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2.2.2 Numerical Studies
Bouldin and Sheoran [3] used parallel compressor theory to evaluate the impact
of swirl on compressor performance and operability. The authors chose the Honeywell
ASE120 industrial power engine for their evaluation. Compressor maps for bulk co- and
counter-swirl as well as compressor maps for undistorted flow were developed and
input into the parallel compressor model. Using the compressor maps and parallel
compressor model, the authors were able to reproduce the effects of bulk swirl. The
authors were then able to use the parallel compressor model and maps to study the
effects of more complex swirl patterns such as twin and offset paired swirl. Though total
pressure distortion and inlet swirl show a strong relationship, the authors did not
perform any swirl evaluations that included total pressure distortion.
Davis and Hale [5] used a one-dimensional mean line stage characteristic code
to determine the effects of inlet swirl on blade row performance. The authors then used
these effects to develop scale factors which were input into the characteristic maps
already in the parallel compressor.
The authors then used this modified parallel
compressor model to evaluate the effects of twin swirl on compressor performance and
operability with and without total pressure distortion.
The authors performed their
evaluation on two different compression systems, the one stage NASA Rotor 1B and
the multi-stage High Tip Speed Compressor.
Gerard Welch [22] investigated the impact of bulk co- and counter-swirl as well
as twin swirl on compressor performance and operability using a steady-state, nonlinear
actuator-duct model. Total pressure distortion was also included in the investigation.
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The author used this model to determine the critical sector angle for swirl distortion and
related loss of stability pressure ratio to swirl descriptors.
Davis, Beale and Sheoran [21] used three different numerical methods to predict
the effects of inlet swirl on the fan performance of an F109 turbofan engine. The three
methods used were the one-dimensional mean line analysis, parallel compressor
analysis, and three-dimensional Euler analysis. The authors used the one-dimensional
mean line analysis and the three-dimensional Euler analysis to predict the effects of co
and counter bulk swirl on the fan performance. The authors used the results from the
mean line analysis to develop scaling factors that were used in the parallel compressor
analysis. The parallel compressor analysis was then used to evaluate the effects of co
and counter bulk swirl as well as twin swirl on parallel compressor performance.
2.3 CONCLUSIONS OF THE LITERATURE SEARCH
Spurred by the Tornado’s early engine/airframe integration issues resulting from
swirl at the engine AIP, the effects of S-duct geometry on the swirl pattern at the AIP
has become well understood.
Both experimental and numerical studies have been
performed in an effort to define the swirl pattern that results from the use of S-ducts in
engine inlet design. Despite the presence of swirl resulting from the use of S-ducts,
little actual research has been performed to determine the effect of swirl on compressor
performance and operability.
Parallel compressor theory is a numerical technique
sometimes used to evaluate the effects of bulk and paired swirl on engine performance.
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Different numerical techniques such as one-dimensional mean line analysis and threedimensional Euler analysis have been used to determine the effects of bulk swirl on
compressor stage performance and could be used to develop compressor stage
characteristics for use in parallel compressor models.
Parallel compressor theory
coupled with a one-dimensional mean line analysis is a method that can be used to
evaluate the effects of more complex swirl patterns on compressor performance.
Rather than generating multiple compressor maps for co- and counter- bulk swirl
and interpolating between those maps, the analysis time and complexity would be
decreased and accuracy would be increased by integrating the 1-D Mean Code into
DYNTECC as a subroutine. Operating as a subroutine to DYNTECC, the MLC can
determine compressor stage characteristics real time for any user specified inlet
condition. Work has already been performed in an effort to integrate the MLC into
DYNTECC as a subroutine [15]; however only total pressure distortion was modeled
using the combine DYNTECC/MLC. The code developed by Tibboel did not include
inputs for swirl and relied on correlation to determine rotor and stator loss and deviation.
2.4 PROBLEM STATEMENT
The goal of the work reported herein was to modify the combined
DYNTECC/MLC developed by Tibboel so that inputs for both paired and bulk swirl could
be provided by the user. The MLC would be used as a subroutine by DYNTECC to
determine compressor stage characteristics for parallel compressor segments with
swirl. The combined DYNTECC/MLC was to be used to analyze the effects of different
types of bulk and paired swirl on F109 fan performance and operability.
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3. BACKGROUND
3.1 DYNTECC
DYNTECC is a one-dimensional compression system model developed to
investigate the effect of external disturbances on the compression system stability limit
[14]. DYNTECC models the entire system as an overall control volume which is divided
into elemental control volumes that represent each stage (Figure 11). The inlet and exit
boundary conditions of the overall control volume are specified by the user. Either static
pressure or Mach number can be used as the specified exit boundary condition. Quasi
one-dimensional mass, momentum and energy conservation equations are solved using
a finite difference numerical technique for each elemental control volume. Source terms
supplied by stage characteristic maps are used to provide the momentum and energy
equations with stage forces and shaft work [23].
3.1.1 Governing Equations
The one dimensional forms of the mass, momentum and energy conservation
equations are the basis for the governing equations of DYNTECC.
The one
dimensional conservation of mass used by DYNTECC is as follows,
(1)
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Figure 11: DYNTECC Compression System Control Volume Model [14]
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where W B is the inter-stage bleed flow per distributed length, W is the mass flow rate, ρ
is the density and A is the area [14].
The one dimensional conservation of momentum used by DYNTECC is written
as follows,
(2)
where,
(3)
is the momentum impulse term in which u is the axial velocity and Ps is the static
pressure [14]. FX is the axial force distribution acting on the control volume and is
written as follows,
(4)
where Fb is the blade force and
is the force of the engine case acting on the control
volume [14].
The one dimensional conservation of energy equation is written as,
(5)
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Where SW is the rate of shaft work, Q is the rate of heat addition to the control volume,
HB is the total enthalpy of the bleed flow and E is the energy function [14]. The energy
function is written as,
(6)
where e is the internal energy [14].
3.1.2 Stage Characteristics
During normal operation, DYNTECC uses steady state stage characteristics
(Figure 5) to obtain stage pressure ratio (PR) and temperature ratio (TR) which provide
closure for the conservation equations. The characteristic maps provide stage PR and
TR as a function of W and rotor rotational speed.
For test cases between the
characteristic map constant speed lines, values for pressure rise and temperature rise
must be interpolated.
Knowing the stage pressure and rise, DYNTECC then backs out steady state
values for the axial stage forces (FX) and rate of shaft work using equations 2 and 5.
These steady state values are used during pre-stall operation; however, during dynamic
events such as rotating stall or surge, these steady state values are not correct. For
this study, this fact is not relevant because operation of DYNTECC was halted once fan
stall was achieved.
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3.1.3 Solution Technique
DYNTECC has an option to use an explicitly formulated numerical solver or an
implicitly formulated numerical solver.
The implicit solver is utilized for analysis of
steady state or slow transient events. The explicit solver is used to study transient
events such as stall and was used during for this study. A detailed description of the
explicit and implicit numerical solver used by DYNTECC as a solution technique can be
found on pages 329 and 330 in Reference 24.
3.1.4 Inlet Distortion Modeling
DYNTECC utilizes parallel compressor theory in order to model inlet distortion.
Parallel compressor theory divides the compressor into radial or circumferential
elements. These elements act independently from each other and exit to the same
boundary condition. An illustration of the parallel compressor theory concept is shown in
Figure 4. The compression system becomes unstable (stalls) when any of the elements
becomes unstable. DYNTECC was originally developed to model inlet total pressure
and total temperature distortion. During normal operation, DYNTECC utilizes the same
stage characteristic maps for all of the elements because the stage characteristic does
not change with inlet total pressure or temperature. The presence of swirl will change
the incidence angle of the rotor blades which in turn will change the stage
characteristics. In order to use DYNTECC to model swirl, stage characteristic maps
had to be developed that include the effects of swirl [21]. This method was used in Ref.
3, Ref. 5 and Ref. 21.
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3.2 1-D MEAN LINE CODE
The 1-D Mean Line Code is a compressible, one-dimensional, steady-state, rowby-row simulation that uses velocity diagrams (Figure 1), blade loss correlations and
deviation correlations to develop stage characteristics [25]. The MLC was developed in
order to understand blade-by-blade performance and to provide stage characteristic
maps for parallel compressor models when experimental data or detailed blade
geometry required to produce the stage characteristic maps were not available. The
MLC has four modes of operation, three of which use a Newton multi-variable method to
converge on a solution while the fourth method calculates the solution directly based on
inlet flow conditions, relative total pressure loss across the rotor blade (ω) and exit
angle deviation (δ) [25].
3.2.1 Governing Equations and Solution Method
The inlet total pressure (Pt1), total temperature (Tt1), Mach number (M1) and α are
known values input by the user. Using these values along with the isentropic relations,
inlet static pressure (Ps1), static temperature (Ts1) and speed of sound (a1) are
calculated. Velocity triangles (Figure 1) are then used by the MLC to determine the inlet
relative velocity (V1R), inlet relative flow angle (β1) and inlet relative Mach Number (M1R).
Isentropic relationships are then used to calculate the inlet relative total pressure (Pt1R)
and relative total temperature (Tt1R).
In order to calculate the total pressure and
temperature across the rotor, the relative mass flow function (MFF), relative total
temperature ratio, relative total pressure ratio and ratio of areas perpendicular to the
flow must be calculated [25].
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A detailed development of the MFF can be found in Ref. 1 and in Ref. 25. From
Ref. 25, the relative mass flow function entering the bladed region (MFF1) is written as
Equation 7.
(7)
The ratio of relative total temperature is derived from the first law of thermodynamics
and from the Euler turbine equation and is developed in detail in Ref. 25. From Ref. 25,
the relative total temperature ratio is written as:
(8)
where Tt2R is the blade exit relative total temperature, U1 is the blade entrance wheel
speed, U2 is the blade exit wheel speed and Cp is the specific heat at constant pressure.
Note that if the radius at the inlet and exit of the blade row is the same, then U 1 and U2
will be the same and the ratio of relative total temperature will be one.
The ratio of relative total pressure is developed in detail in Ref. 25 and is shown
below in Equation 9.
(9)
Pt2R is the blade exit relative total pressure and γ is the ratio of specific heats. The
relative total pressure loss, ω, is defined in Equation 10 [25].
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(10)
The relative total pressure loss must either be provided by the user from test data, or
can be obtained from the Hearsey Correlations [26].
A detailed discussion on the
source of the ω used in the analysis is located in Section 3.2.3.
The equation for the ratio of areas perpendicular to the flow is developed in Ref.
25 and is shown below in Equation 11.
(11)
The axial velocity density ratio (AVDR) is defined in Ref. 25 as:
(12)
where rtip is the radius at the blade tip and rhub is the radius at the hub. From Ref. 25, Ϛ
is a cross section correlation factor that accounts for boundary layer growth, the error
associated with using a 1-D representation for a radial distribution of inlet flow angle and
the error associated with neglecting the radial velocity.
The blade exit relative velocity flow angle (β2) is found using Equation 13 from
Ref. 25. In order to calculate β2, δ must either be provided by the user from test data, or
can be obtained from the Hearsey Correlations [26].
A detailed discussion on the
source of the δ used in the analysis is located in Section 3.2.3.
(13)
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β2’ is the blade exit metal angle. Both β2 and β2’ are shown in Figure 1.
The relative mass flow function at the exit of the blade (MFF2) can now be
calculated using Equation 14 from Ref. 25.
(14)
R is the gas constant. M2R is the blade exit relative Mach number and is calculated
using Equation 14. Knowing M2R, Pt2R and Tt2R, the isentropic relationships are then
used to calculate the blade exit static pressure and temperature. From these static
conditions, velocity triangles are used to determine the absolute blade exit total
pressure and total temperature [24].
3.2.2 Rotor and Stator Loss and Deviation
If experimental data or detailed blade and stator geometry are not available, ω
and δ are calculated using empirical correlations developed by Hearsey and are based
on cascade and machine experiments with NACA 65-series and double-circular-arc
airfoils [26]. Experimental data can be used to calibrate the loss and deviation which
results in the development of add-loss and add-deviation maps. These maps, along
with inlet flow conditions are then used by the MLC to determine the stage characteristic
(pressure ratio and temperature ratio) at the desired inlet condition.
During the course of research, additional capability was added to the MLC.
Rather than use the Hearsey Correlations with add-loss and add-deviation, an option
was added to the MLC that allowed for the look-up of rotor and stator ω and δ directly
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from maps rather than using Hearsey Correlations with add-loss and add-deviation.
The development of these maps is discussed in Section 4.2.
3.2.3 1-D Mean Line Code Integration into DYNTECC
Grady Tibboel integrated the MLC into DYNTECC in order to provide “real time”
source terms such as stage pressure and temperature ratio at a specified operating
condition. Rather than developing multiple maps, DYNTECC can utilize the MLC to
determine the source terms for any input swirl angle, inlet total pressure, or inlet total
temperature. DYNTECC passes the following information to the MLC: constants (R, γ,
gc), total pressure, total temperature, Mach number, mass flow, rotor speed, inlet and
exit areas, compressor exit static pressure. The MLC then determines the loss and
deviation using the calibrated add-loss and add-deviation maps and calculates the
stage PR and TR for that flow point. The calculation of the PR and TR for the given flow
point replaces the interpolation of PR and TR from a stage characteristic map. The PR
and TR are then passed back to DYNTECC.
3.3 F109 TURBOFAN ENGINE
The Honeywell F109, shown in Figure 12, is a high by-pass ratio turbofan engine
with a maximum thrust of 1330 lbf at sea-level static, standard day conditions. The
F109 has a single stage axial fan, a two stage centrifugal high pressure compressor, a
reverse flow annular combustor, a two stage axial high pressure turbine and a two stage
axial low pressure turbine. The F109 has a by-pass ratio of 5:1 and the by-pass flow
mixes with the core flow before exiting through a common nozzle [21]. The F109 is
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ideal for inlet swirl testing because the fan is not equipped with IGVs which will make it
more sensitive to the presence of swirl.
Universities like the USAFA and Virginia
Polytechnic Institute and State University (Virginia Tech) use the F109 for research
[27,28].
Figure 12: F109 Turbofan Engine
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4. APPROACH
4.1 OUTLINE OF APPROACH
The following is an outline of the approached used to meet the goals of the
research reported herein. The main bullets describe the task that had to be completed
in order to meet the goals of the research. The sub-bullets describe additional tasks
that had to be completed in order to complete the task described in the main bullet.

Integrate the updated MLC as a subroutine into DYNTECC.
o Develop stalling criteria for DYNTECC.
o Eliminate the need to use correlations to find rotor and stator
relative total pressure loss and blade exit deviation angle by
developing method to look those values up from a user provided
table.

Calibrate the MLC in order to provide F109 fan rotor and stator loss and
blade exit deviation angle.
o Use values of F109 fan total pressure and total temperature rise
from the F109 cycle deck in the MLC operating in calibration mode
to determine F109 fan rotor and stator relative total pressure loss
and blade exit deviation anle.
o Build 2-D look up tables containing values of relative total pressure
loss and blade exit deviation angle for both the F109 fan rotor and
stator (four tables total).
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
Validate the combined DYNTECC/MLC using data from the F109 cycle
deck and experimental data gathered at the USAFA in May of 2008.
o Validate the steady-state constant speed lines predicted by the
combined DYNTECC/MLC.
o Validate the stall line predicted by the combined DYNTECC/MLC.

Use the combined DYNTECC/MLC to analyze different types of swirl and
present results.
o Determine the change in F109 fan pressure rise and mass flow for
different types of swirl.
o Determine the effect of different types of swirl on F109 fan stability
limit.
4.2 INTEGRATION OF THE 1-D MEAN LINE CODE INTO DYNTECC
The integration of the MLC into DYNTECC as a subroutine was based on the
work of Grady Tibboel [15]. The version of the MLC used for Tibboel’s research was
replaced with a newer version. The newer version of the MLC included improved user
inputs capability and additional calibrations options. Because engine operability was
also going to be studied, a new stall criterion had to be developed for DYNTECC.
During normal operation, DYNTECC uses the stage characteristic maps to determine
when stall occurs. DYNTECC analyzes the slope of the stage characteristic, and when
the slope reaches zero, DYNTECC assumes that the stage has then become stalled.
When operating with the MLC, characteristic maps are not used and stall is not
defined in the MLC. This necessitated the modification of the MLC to output a
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notification when the stage reaches the stability limit. Rotor diffusion factor (Dr) is a
measure of velocity diffusion on the suction side of a rotor blade and can be correlated
directly with total pressure loss [1]. Diffusion factor can also be used as a measure of
blade loading [29]. Equation 15 from Reference 1 was used to calculate the rotor
diffusion factor.
(15)
The velocities in Equation 15 are shown in Figure 1. σ in Equation 15 is solidity, which
is the ratio of the blade chord length and the blade spacing.
The MLC was modified to calculate and output rotor diffusion factor. This task
was made easier by the fact that the MLC already calculated and output all of the
parameters necessary to calculate rotor diffusion factor. DYNTECC was modified so
that it compares the diffusion factor output by the MLC with a user specified stalling
diffusion factor and would quit executing indicating system stall.
The ability to look up rotor relative total pressure loss, rotor exit deviation angle,
stator relative total pressure loss, and stator exit deviation angle directly from tables
while by-passing the Hearsey correlations was added to the MLC. The option to bypass the Hearsey correlations was added because these correlations were developed
for a fan operating at near 100% speed. It is important to note that the F109 was not
the only application for the F109 fan, and subsequently, the F109 does not operate at
the 100% design fan speed.
The maximum fan speed the F109 operates is at
approximately 84%. When operating well below the maximum fan speed, the output of
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the Hearsey correlation was significantly degraded and the add-loss and add-deviation
outputs from the MLC are unrealistic.
4.3 MEAN LINE CODE CALIBRATION
The MLC was used in a standalone mode to develop the rotor and stator total
pressure loss and blade exit deviation angle tables. As mentioned in Section 3.2.1, the
MLC as four modes of operation, three of these modes are predictive and one of these
modes is a calibration mode. In calibration mode, stage inlet conditions, mass flow,
rotor speed, AVDR, and known stage characteristics are input by the user. The MLC
uses a package called Replicas to then solve for the corresponding rotor and stator
relative total pressure loss and exit deviation angles that corresponds to the user
specified inputs.
The F109 cycle deck was used to provide mass flow and stage characteristics at
the following fan speeds (percent of design fan speed): 20%, 45%, 60%, 70%, 80%,
85%, 90%, 95%, 100%, 105%, 110%, and 115%. The cycle deck was only able to be
used for calibration purposes because the F109 is equipped with a single stage fan. In
order to calibrate the MLC, PR and TR for each stage must be known. Engine cycles
decks normally only output overall fan PR and TR, and for multistage fans, this would
not be enough information for calibration purposes.
Initially, actual test data obtained at the USAFA was to be used to calibrate the
MLC. The USAFA data was obtained at only four fan speeds: 53%, 62%, 71% and
84%. The fan temperature ratio measured at the USAFA was less than the isentropic
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fan temperature ratio that corresponded to the measured fan pressure ratio and was
deemed invalid. Because of the lack of valid fan temperature ratio, there were not
enough inputs available to calibrate the MLC using the data obtained at the USAFA.
The pressure loss and exit deviation output from the MLC during the calibration
process was plotted as a function of incidence angle and inlet relative Mach number
and a surface fit to the output in order to aid in extrapolation and provide evenly space
grid points.
Two-dimensional tables for rotor relative total pressure loss, rotor exit
deviation angle, stator relative total pressure loss, and stator exit deviation angle were
built using the surface fit to the MLC output. The rotor and stator relative total pressure
loss and blade exit deviation angles output by the MLC as well as the surface fit used to
extrapolate and build the two-dimensional look-up tables are shown in Figures 13-16.
Total pressure loss across a cascade is a function of blade incidence angle and inlet
Mach number [29]. The rotor and stator relative total pressure loss and exit deviation
angle were treated as a function of inlet relative Mach number and blade incidence
angle because in the relative reference frame, rotating blades are treated as a cascade.
As mentioned in Section 4.1, the rotor diffusion factor was used to determine stall.
Figure 17 shows the rotor diffusion factor along the F109 cycle deck stall line as
calculated by the MLC. The average of the stalling diffusion at each constant speed line
in the F109 cycle deck was used as the DYNTECC user specified stalling diffusion
factor.
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Rotor Relative Total Pressure Loss, MLC Look-Up Table
1.2
1.2
1.1
0.25
1.1
0.25
0.2
1
0.2
0.15
0.9
0.15
0.1
0.8
0.1
0.05
0.7
0.05
0.6
0
0.6
0
0.5
-0.05
0.5
-0.05
Inlet Relative Mach Number
Inlet Relative Mach Number
Rotor Relative Total Pressure Loss, Data
1
0.9
0.8
0.7
8
10
12
Incidence(degrees)
14
-0.1
16
8
10
12
Incidence (degrees)
14
16
-0.1
Figure 13: F109 Rotor Relative Total Pressure Loss
Rotor Exit Deviation (degrees), MLC Look-Up Table
1.220
20
1.1
18
1.1
18
Inlet Relative Mach Number
Inlet Relative Mach Number
Rotor Exit Deviation (degrees), Data
1.2
1
0.9
0.8
0.7
16
16
1
14
14
0.912
12
10
0.8
10
8
0.7
6
8
6
0.6
0.6
4
4
0.5
0.52
2
8
10
12
Incidence (degrees)
14
16
0
8
10
12
Incidence (degrees)
Figure 14: F109 Rotor Exit Deviation Angle
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14
16
0
Stator Relative Total Pressure Loss, Data
Stator Relative Total Pressure Loss, MLC Look-Up Table
0.15
0.8
0.75
0.75
0.7
0.7
Inlet Relative Mach Number
Inlet Relative Mach Number
0.8
0.65
0.6
0.55
0.5
0.45
0.650.1
0.55
0.50.05
0.35
0.35
0
-20
-15
-10
-5
Incidence (degrees)
0
5
0.05
0.45
0.4
-25
0.1
0.6
0.4
-30
0.15
10
0
-30
-25
-20
-15
-10
-5
Incidence (degrees)
0
5
10
Figure 15: F109 Stator Relative Total Pressure Loss
Stator Exit Deviation (degrees), MLC Look-Up Table
0.8
0.8
0.75
14
0.75
14
0.7
0.713
13
0.65
12
12
0.6
11
0.55
11
0.510
10
Inlet Relative Mach Number
Inlet Relative Mach Number
Stator Exit Deviation (degrees), Data
0.65
0.6
0.55
0.5
0.45
0.4
0.35
-30
-25
-20
-15
-10
-5
Incidence (degrees)
0
5
10
0.459
9
0.4
8
0.35
8
7
-30
-25
-20
-15
-10
-5
Incidence (degrees)
Figure 16: F109 Stator Exit Deviation Angle
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0
5
10
7
Rotor Diffusion Factor
0.585
F109 Cycle Deck Stalling Diffusion Factor
0.58
Average Stalling Diffusion Factor
0.575
0.57
0.565
0.56
0.555
0.55
0.545
0.54
40
50
60
70
80
90
100
110
Fan Speed (%)
Figure 17: F109 Cycle Deck Stall Line Rotor Diffusion Factor
4.4 VALIDATION OF COMBINED DYNTECC/1-D MEAN LINE CODE
Before beginning the analysis of F109 fan performance with swirl at the AIP, it
was important to verify that the DYNTECC/MLC predicted fan performance was in
acceptable agreement with the F109 cycle deck and the experimental data obtained at
the USAFA over the entire range of operating speeds. Figure 18 is a comparison of the
constant speed lines generated by the combined DYNTECC/MLC and the steady state
speeds line from the F109 cycle deck used to generate the loss and deviation look-up
tables for the MLC. The DYNTECC/MLC prediction was within ±0.5% of the F109 cycle
deck for all speed lines, showing excellent agreement. Figure 19 is a comparison of the
DYNTECC/MLC predicted speed lines with constant
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105
Normalized Fan Pressure Ratio (%)
F109 Cycle Deck
100
DYNTECC/MLC
0.5%
95
0.5%
0.4%
90
NF = 85%
85
80
NF = 70%
75
NF = 60%
70
0.2%
NF = 45%
65
30
40
50
60
70
80
90
100
110
120
Normalized Corrected Mass Flow (%)
Figure 18: DYNTECC/MLC Predicted and F109 Cycle Deck Specified Fan Map
Comparison (No Swirl)
105
Normalized Fan Pressure Ratio (%)
Air Force Academy Measured Data
DYNTECC/MLC
100
N1 = 84%
95
1.9%
90
N1 = 71%
85
N1 = 62%
80
N1 = 53%
75
40
50
60
70
80
90
100
110
Normalized Corrected Mass Flow (%)
Figure 19: DYNTECC/MLC Predicted and USAFA Measured Fan Map Comparison (No
Swirl)
39
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speed lines measured at the USAFA. The DYNTECC/MLC prediction was within ±0.8%
for all speed lines except the 84% fan speed line, showing reasonable agreement. The
fan speed of the USAFA measured data was not a constant 84% during data
acquisitions and may not be a valid constant speed line.
Before evaluating the effects of inlet swirl on F109 fan operability, it was
important to first determine whether the choice of rotor diffusion factor as the stalling
criteria was valid. Figure 20 is a comparison of the DYNTECC/MLC predicted stall line
and the F109 cycle deck specified fan stall line. The DYNTECC/MLC predicted stall line
shows excellent agreement with the F109 cycle deck specified stall line at lower fan
speeds. At 84% fan speed, the DYNTECC/MLC stall line was ~1.4% higher than the
F109 Cycle deck specified stall
Normalized Fan Pressure Ratio (%)
105
100
95
1.4%
90
85
80
F109 Cycle Deck Stall Line
DYNTECC/MLC
75
40
50
60
70
80
90
Normalized Corrected Mass Flow (%)
Figure 20: DYNTECC/MLC Predicted and F109 Cycle Deck Specified Fan Stall Line
(No Swirl)
40
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Normalized Fan Pressure Ratio (%)
105
100
95
0.71%
90
85
80
Air Force Academy Measured Stall Line
DYNTECC/MLC
75
40
50
60
70
80
90
Normalized Corrected Mass Flow (%)
Figure 21: DYNTECC/MLC Predicted and USAFA Measured Fan Stall Line (No Swirl)
line. Figure 21 is a comparison of the DYNTECC/MLC predicted stall line and the
USAFA measured fan stall line. The DYNTECC/MLC stall line was within ±0.7% of the
USAFA measured fan stall line, showing acceptable agreement.
4.5 PRESENTATION OF RESULTS
The effect of the different swirl patterns modeled at the F109 AIP on the
compressor performance will be shown on a compressor map. The compressor map
has the ability to show the change in fan pressure ratio and mass flow at a constant
speed for each of the different cases modeled. The effect of the different swirl patterns
on F109 fan operability will be characterized by the loss in stability pressure ratio
(ΔPRS). ΔPRS is described in Reference 7 as the percent change in stability pressure
ratio between the undistorted and distorted stability limit point at a constant corrected
mass flow. ΔPRS is shown graphically in Figure 22 and was calculated using Equation
41
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16 [7] where PR1 is the undistorted stability pressure ratio and PRDS and the distorted
stability pressure ratio at the same corrected mass flow. A positive value for ΔPRS
indicates a loss in stability pressure ratio while a negative value for ΔPRS indicates a
gain in stability pressure ratio.
(16)
93
Undistorted Stability
Pressure Ratio, PR1
Normalized Fan Pressure Ratio (%)
91
Distorted Stability
Pressure Ratio, PRDS
ΔPRS
89
Distorted Inlet
Stall Line
Clean Inlet
Stall Line
87
85
83
Steady State
Operating Point, PRO
81
55
57
59
61
63
65
67
69
Normalized Corrected Mass Flow (%)
Figure 22: Loss in Stability Pressure Ratio (ΔPRS)
42
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71
5. RESULTS
Two different types of bulk swirl and one type of paired swirl were analyzed using
the combine DYNTECC/MLC. Because both DYNTECC and the MLC are essentially
one dimensional, only one value of swirl angle is chosen for each parallel compressor
segment in DYNTECC. In reality, swirl angle at the AIP changes with radius. The swirl
angle used by the combined DYNTECC/MLC is a notional average swirl angle for the
segment. For comparison to experimental data, the combined DYNTECC/MLC would
use a root mean square average of the swirl angle across the radius as the input.
5.1 BULK SWIRL
As discussed in Section 2.1, is has been found that bulk swirl can develop at an
AIP downstream of an s-duct at a high angle of attack. The bulk swirl is a result of flow
separation at the lip of the inlet. As discussed in Section 1.0, there are two types of bulk
swirl: co-Swirl and counter-swirl. Co-swirl and counter-swirl were modeled using the
combined DYNTECC/MLC at three increasing levels of swirl intensity: 5°, 10° and 15°.
These intensities were similar to the same intensities modeled in Ref. 21. These swirl
levels were evaluated at four different fan speeds: 53%, 62%, 71% and 84%. These
speeds correspond to the fan speeds where data was collected at the USAFA. Mach
number was used as the back boundary condition for the parallel compressor model.
Figure 23 shows the DYNTECC/MLC prediction of the effect of the varying levels
of co-swirl intensity on F109 fan performance and operability. As the intensity of the coswirl is increased, the fan pressure ratio at a constant referred mass flow decreases.
43
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The referred mass flow at the stall point for each fan speed also decreases as co-swirl
intensity is increased. At 84% fan speed with 15° co-swirl, the flow through the fan
begins to move into the choked region. The decrease in fan pressure ratio and referred
mass flow at fan stall resulting from the co-swirl is cause by the reduction in blade
loading.
increases.
The incidence angle of the rotor decreases and the amount of co-swirl
The reduction in rotor incidence angle causes the blade to become
unloaded, which decreases the fan pressure ratio but increases the fan stall margin.
The DYNTECC/MLC prediction of the effect of counter-swirl on F109 fan
performance and operability is shown in Figure 24. As the intensity of the counter-swirl
increases,
the fan pressure ratio at a constant referred mass flow increases. The
referred mass flow at the stall point for each fan speed also increases as counter-swirl
intensity is increased.
At 71% fan speed, the referred mass flow at fan stall with 15°
counter-swirl is almost the same as the clean inlet steady-state operating point referred
mass flow. At the highest F109 fan speed, 84%, the referred mass flow at fan stall is
greater than the clean inlet steady-state operating point. This means that the F109 fan
would stall if it encountered 15° counter-swirl at 84% fan speed.
The increase in fan pressure ratio and referred mass flow at fan stall resulting
from counter-swirl is caused by the increase in blade loading. The incidence angle of
the rotor increases and the amount of counter-swirl increases. The increase in rotor
incidence angle causes the blade to become more loaded, which increases the fan
pressure ratio and decreases the fan stall margin.
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105
Co-Swirl
100
Normalized Fan Pressure Ratio (%)
Swirl
Compressor
Rotation
95
5°
90
10°
Clean
NF = 84%
85
NF = 71%
80
NF = 62%
15°
NF = 53%
75
40
45
50
55
60
65
70
75
80
85
90
95
100
105
Normalized Referred Mass Flow (%)
Clean Inlet
5° Co-Swirl
10° Co-Swirl
15° Co-Swirl
Stall Point
Clean Inlet Stall Line
5° Co-Swirl Stall Line
10° Co-Swirl Stall Line
15° Co-Swirl Stall Line
Clean Inlet SS Operating Point
Figure 23: DYNTECC/MLC Prediction of the Effect of Co-Swirl on F109 Fan Performance and Operability
45
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110
Counter-Swirl
105
Normalized Fan Pressure Ratio (%)
Swirl
Compressor
Rotation
NF = 84%
100
95
15°
10°
5°
90
NF = 71%
NF = 62%
85
Clean
NF = 53%
80
75
45
50
55
60
65
70
75
80
85
90
95
100
105
Normalized Referred Mass Flow (%)
Clean Inlet
5° Coutner-Swirl
10° Coutner-Swirl
15° Coutner-Swirl
Stall Point
Clean Inlet Stall Line
5° Counter-Swirl Stall Line
10° Counter-Swirl Stall Line
15° Counter-Swirl Stall Line
Clean Inlet SS Operating Point
Figure 24: DYNTECC/MLC Prediction of the Effect of Counter-Swirl on F109 Fan Performance and Operability
46
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4
Speed
Speed
Speed
Speed
Less Stall
Margin
53% Physical Fan
62% Physical Fan
71% Physical Fan
84% Physical Fan
2
0
-2
More Stall
Margin
Loss in Stability Pressure Ratio, ΔPRS (%)
6
-4
-6
-20
-15
-10
-5
0
5
10
15
20
Bulk Swirl Angle (°)
Co-Swirl
Counter-Swirl
Figure 25: Loss in Stability Pressure Ratio (ΔPRS) as a Function of Bulk Swirl Angle
The loss (or gain) in stability pressure ratio is plotted as a function of bulk swirl
angle in Figure 25.
ΔPRS increases as the amount of counter swirl increases, a
decreases as the amount of co-swirl increases. The change in ΔPRS becomes more
pronounced as the fan speed increase.
5.2 PAIRED SWIRL
As discussed in Section 2.1, when flow separation is not present at the lip of an
inlet, a paired swirl pattern will develop at the AIP downstream of an s-duct. Twin
paired swirl was modeled using DYNTECC/MLC with two equal segments with 10° of
swirl in each segment. DYNTECC divided the fan into two equal parallel tubes. One
47
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tube had 10° of co-swirl applied while the other had 10° of counter swirl applied. The
10° intensity was chosen because that is the intensity that was modeled in Ref. 21.
Mach number was used as the back boundary condition for the parallel compressor
model. Figures 26 – 29 show the effect of 10° twin swirl on fan performance and
operability predicted by DYNTECC/MLC at 53%, 62%, 71% and 84% referred fan
speed.
The DYNTECC/MLC prediction for 10° co-swirl and 10° counter-swirl are
included in these figures for comparison.
DYNTECC/MLC predicts that twin-swill will have a combine co- and counter-swill
effect on the fan performance and operability. Like co-swirl, twin swirl lowers the fan
pressure ratio at a constant pressure ratio relative to the clean inlet case. However,
rather than increasing the stall margin, twin swirl decreases the stall margin like
counter-swirl. The fan pressure ratio predicted by DYNTECC/MLC for 10° twin-swirl is
approximately and average of the fan pressure rise for co- and counter swirl at the
same referred mass flow.
This makes sense, because half of the compressor is
operating at an elevated fan pressure ratio while the other half is operating at a lower
fan pressure ratio.
The overall fan pressure ratio will be an average of these two
separate compressors.
The referred mass flow at stall is much less with 10° twin-swirl than with a clean
inlet or with 10° co-swirl, but is slight higher than with 10° counter-swirl. The twin-swirl
test case becomes stalled once the diffusion factor of the parallel compressor segment
with 10° counter swirl reaches the stalling diffusion factor, shown in Figure 30. This
48
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Normalized Fan Pressure Ratio (%)
83
53% Fan Speed
82.5
82
81.5
81
80.5
80
79.5
79
78.5
78
45
50
55
60
Normalized Corrected Mass Flow (%)
Clean Inlet
10° Coutner-Swirl
10° Co-Swirl
10° Twin-Swirl
Clean Inlet SS Operating Point
Clean Inlet Stall Line
10° Counter-Swirl Stall Line
10° Co-Swirl Stall Line
10° Twin-Swirl Stall Line
Stall Point
Figure 26: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl and 10° Twin-Swirl on F109 Fan
Performance and Operability at 53% Fan Speed
49
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Normalized Fan Pressure Ratio (%)
88
62% Fan Speed
87
86
85
84
83
82
81
50
55
60
65
70
Normalized Corrected Mass Flow (%)
Clean Inlet
10° Coutner-Swirl
10° Co-Swirl
10° Twin-Swirl
Clean Inlet SS Operating Point
Clean Inlet Stall Line
10° Counter-Swirl Stall Line
10° Co-Swirl Stall Line
10° Twin-Swirl Stall Line
Stall Point
Figure 27: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl and 10° Twin-Swirl on F109 Fan
Performance and Operability at 62% Fan Speed
50
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Normalized Fan Pressure Ratio (%)
95
71% Fan Speed
94
93
92
91
90
89
88
87
86
85
60
65
70
75
80
Normalized Corrected Mass Flow (%)
Clean Inlet
10° Coutner-Swirl
10° Co-Swirl
10° Twin-Swirl
Clean Inlet SS Operating Point
Clean Inlet Stall Line
10° Counter-Swirl Stall Line
10° Co-Swirl Stall Line
10° Twin-Swirl Stall Line
Stall Point
Figure 28: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl and 10° Twin-Swirl on F109 Fan
Performance and Operability at 71% Fan Speed
51
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Normalized Fan Pressure Ratio (%)
104
84% Fan Speed
102
100
98
96
94
92
90
70
75
80
85
90
95
100
Normalized Corrected Mass Flow (%)
Clean Inlet
10° Coutner-Swirl
10° Co-Swirl
10° Twin-Swirl
Clean Inlet SS Operating Point
Clean Inlet Stall Line
10° Counter-Swirl Stall Line
10° Co-Swirl Stall Line
10° Twin-Swirl Stall Line
Stall Point
Figure 29: DYNTECC/MLC Prediction of the Effect of 10° Co-Swirl, 10° Counter-Swirl and 10° Twin-Swirl on F109 Fan
Performance and Operability at 84% Fan Speed
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Diffusion Factor
0.6
0.55
0.5
0.45
0.4
0.35
0.3
85
90
95
100
105
Normalized Corrected Mass Flow (%)
Co-Swirl Segment
Counter-Swirl Segment
Stalling Diffusion Factor
Stall Point
Figure 30: Twin Swirl Segment Diffusion Factor at Constant Fan Speed
result is different than the result obtained in Ref. 3, Ref. 5 and Ref 21. The model
utilized in Ref. 3 and Ref. 21 used static pressure as the back boundary condition,
whereas a Mach number was used as the back boundary condition for the research
reported herein. The difference in boundary condition may explain difference in parallel
compressor segment that stalled first.
Figure 31 shows DYNTECC/MLC predicted
ΔPRS as a function of percent referred fan speed for 10° twin-swirl and 10° counterswirl. Because of the reduced fan pressure ratio coupled with the increased referred
mass flow at stall, the ΔPRS increases faster with twin-swirl than with counter-swirl as
referred fan speed increases.
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10
10° Counter-Swirl
9
Loss in Stability Pressure Ratio, ΔPRS (%)
10° Twin-Swirl
8
7
6
5
4
3
2
1
0
50
55
60
65
70
75
80
85
90
Percent Physical Fan Speed
Figure 31: Loss in Stability Pressure Ratio (ΔPRS) as a Function of Percent Referred
Fan for 10° Counter-Swirl and 10° Twin-Swirl
54
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6. SUMMARY AND CONCLUSIONS
A new version of the MLC was incorporated into the parallel compressor model
DYNTECC as a subroutine. DYNTECC was able to use the MLC to calculate a pointby-point representation of the stage characteristics internally without the use of
characteristic maps. Both DYNTECC and the MLC were modified in order to determine
stage stall criteria.
The MLC was modified to output a rotor diffusion factor while
DYNTECC was modified to compare the rotor diffusion factor to a user specified stalling
diffusion factor. Additional modifications were made to the MLC in order to by-pass
blade relative total pressure loss and deviation correlation and look up those values
directly from a user provided two-dimensional table.
Rotor and stator relative total
pressure loss and exit deviation angle tables were developed using a standalone
version of the MLC operating in a calibration mode. Stage characteristics required for
developing the relative total pressure loss and blade exit deviation tables were obtained
from the F109 cycle deck.
The combined DYNTECC/MLC was validated by comparing clean inlet
predictions to the F109 cycle deck and well as clean inlet data obtained at the USAFA.
The DYNTECC/MLC predicted F109 fan pressure ratio was within 0.5% of the F109
Cycle deck fan pressure ratio and was generally within 0.8% of the USAFA measured
fan pressure ratio at the same referred fan speed and referred mass flow.
The
DYNTECC/MLC predicted clean inlet stall line was compared to the F109 cycle deck
stall line as well as the clean inlet stall line measured at the USAFA and showed
acceptable agreement with both.
55
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Varying intensities of bulk swirl and paired swirl at were modeled using the
combined DYNTECC/MLC.
Two cases of bulk swirl were modeled: co-swirl and
counter-swirl. Each of these cases was modeled at the following swirl angles: 5°, 10°
and 15°. In the case of co-swirl, DYNTECC/MLC model predictions showed that as the
swirl angle increases, the fan pressure ratio decreases and fan stability margin
increases. In the case of counter-swirl, DYNTECC/MLC model predictions showed the
opposite.
As counter-swirl intensity increases, DYNTECC/MLC predicts that fan
pressure ratio will increase while fan stability margin decreases.
Twin swirl was the only case of paired swirl modeled using DYNTECC/MLC.
DYNTECC/MLC predicted that twin swirl reduces fan pressure ratio and also reduces
fan stability margin. The loss in stability pressure ratio with 10° twin swirl was compared
to the loss in stability pressure ratio with 10° counter-swirl. For the same amount of
swirl, the loss in stability pressure ratio was much greater and increased at a higher rate
with paired twin-swirl compared to bulk counter-swirl.
Once rotor and stator relative total pressure loss and blade exit deviation tables
were developed, modeling different cases of inlet swirl with the combined
DYTNECC/Mean code was simplified because new stage characteristic maps did not
have to be developed for each case. Modeling a different case of swirl was as easy as
changing one or two DYNTECC inputs. No data has yet been gathered for the F109
with flow angularity at the AIP. An experiment is planned for the fall of 2010 at the
USAFA with a swirl generator developed by David Beale at AEDC. Because no data
56
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has yet been gathered with swirl at the AIP of an F109, the DYNTECC/MLC predictions
for various cases of bulk and paired swirl have not yet been validated.
57
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7. RECOMMENDATIONS
The work performed in this thesis helped further the development of a new tool
capable of modeling the effects of inlet swirl at the AIP of an axial flow turbine engine.
However, predictions made with the modified DYNTECC/Mean Line with swirl at the AIP
have not yet been validated using measured test data. As discussed in Section 2.2.1,
two different swirl generators were under development at AEDC. The swirl chamber
concept intended to produce a twin swirl pattern discussed in Section 2.2.1 has been
removed from consideration and a new concept is currently under development at
AEDC and is being lead by David Beale. The turning vane concept intended to produce
bulk swirl patterns discussed in Section 2.2.1 is still being considered. It is
recommended that further testing on the F109 with the above mentioned swirl
generators be completed so that the DYNTECC/MLC model can be validated.
Another recommendation would be to better understand the consequences of
using a one-dimensional model to predict multi-dimensional phenomena.
Average
compressor stage characteristics were used to develop the relative total pressure loss
and blade exit deviation tables, and mean line values of swirl were modeled using the
code. Swirl angle changes with radius, and different values of swirl at the hub and tip of
the compressor might have an effect on the compressor performance and operability. It
should also be noted that the combined DYNTECC/MLC is only capable of modeling a
single stage fan or compressor. The capability to model multiple stage compressors
should also be pursued.
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LIST OF REFERENCES
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1.
Mattingly, J.D., Elements of Gas Turbine Propulsion, American Institute of
Aeronautics and Astronautics, Inc., Reston, VA, 2005.
2.
Walsh, P.P. and Fletcher, P., Gas Turbine Performance, 2nd ed., Blackwell
Science and ASME Press, Fairfield, NJ, 2004.
3.
Bouldin, B. and Sheoran, Y., “Impact of Complex Swirl Patterns on Compressor
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Guo, R.W. and Seddon, J., “Swirl Characteristics of an S-Shaped Air Intake with
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Davis, M. and Hale, A., “A Parametric Study on the effects of Inlet Swirl on
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Simulations,” GT2007-27033, ASME Turbo Expo 2007: Power for Land, Sea,
and Air, Montreal, Canada, May 14–17, 2007.
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Aulehla, F., “Intake Swirl – A Major Disturbance Parameter in Engine/Intake
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7.
SAE S-16 Committee, “Gas Turbine Engine Inlet Flow Distortion Guidelines,”
ARP 1420, Revision B, Society of Automotive Engineers, February 2002.
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SAE Aerospace Information Report AIR1419, Revision A, “Inlet Total-PressureDistortion Considerations for gas-Turbine Engines,” May 1982.
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SAE S-16 Committee, “An Assessment of Planar Waves,” AIR5866, Society of
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SAE S-16 Committee, “A Current Assessment of the Inlet/Engine Temperature
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Davis, M., Hale, A. and Beale, D., “An Argument for Enhancement of the Current
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Beale, D.K., “Improving Information Productivity and Quality in Turbine Engine
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Beale, D.K., Cramer, K.B. and King, P.S., “Development of Improved Methods for
Simulating Aircraft Inlet Distortion in Turbine Engine Ground Tests,” AIAA 200261
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3045, 22nd AIAA Aerodynamic Measurement Technology and Ground Testing
Conference, St. Louis, MO, June 24-26, 2002.
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Hale, A. A. and Davis, M.W., “DYNamic Turbine Engine Compressor Code
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Tibboel, G.A., “Modification of a One-Dimensional Dynamic Compression System
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Wellborn, S.R., Reichert, B.A. and Okiishi, T.H., “An Experimental Investigation
of the Flow in a Diffusing S-Duct,” AIAA 92-3622, 28th Joint Propulsion
Conference and Exhibit, Nashville, TN, July 6-8, 1992.
17.
Loeper, K.M. and King, P.I., “Numerical Investigation of Geometric Effects on
Performance of S-Ducts,” AIAA 2009-713, 47th AIAA Aerospace Sciences
Meeting Including the New Horizons Forum and Aerospace Exposition, Orlando,
FL, January 5-8, 2009.
18.
Flitcroft, J.E., Dunham, J. and Abbott, W.A., “Transmission of Inlet Distortion
Through a Fan,” Engine Response to Distorted Inflow Conditions: Conference
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Proceedings of the Propulsion and Energetics Specialists’ Meeting (68 th),
AGARD, 1987.
19.
Pazur, W. and Fottner, L., “The Influence of Inlet Swirl Distortions on the
Performance of a Jet Propulsion Two-Stage Axial Compressor,” Journal of
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Schmid,
N.R.,
Leinhos,
D.C.
and
Fottner,
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Davis, M., Beale, D. and Sheoran, Y., “Integrated Test and Evaluation
Techniques as Applied to an Inlet Swirl Investigation using the F109 Gas Turbine
Engine,” GT2008-50074, ASME Turbo Expo 2008: Power for Land, Sea and Air,
Berlin, Germany, June 9-13, 2008.
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Welch, G.E., “Determination of Critical Sector Angle of Inlet Swirl Distortion using
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Exhibit, Cincinnati, OH, July 8-11, 2007.
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23.
Davis, M.W., “A Stage-by-Stage Post-Stall Compression System Modeling
Technique: Methodology, Validation, and Application,” PhD Dissertation, Virginia
Polytechnic Institute and State University, Blacksburg, VA, December 1986.
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Davis, M.W., Hale, A. A. and Klepper, J., “40 Years of AEDC Development,
Evolution and Application of Numerical Simulations for Integrated Test and
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Smith, S.L., “1-D Mean Line Code Technique to Calculate Stage-by-Stage
Compressor Characteristics,” MS Thesis, University of Tennessee, Knoxville, TN,
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Hearsey, R.M., Program HT0300, NASA 1994 Version, The Boeing Company.
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Falk, E.A., Jumper, E.J. and Fabian, M.K., “An Experimental Study of Unsteady
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Kozak, J.D. and Ng, W.F., “Investigation of IGV Trailing Edge Blowing in a F109
Turbofan Engine,” AIAA 2000-0224, 38th Aerospace Sciences Meeting & Exhibit,
Reno, NV, January 10-13, 2000.
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29.
Liblein, S., “Aerodynamic Design of Axial-Flow Compressors,” NASA SP-36,
Chapter VI, 1965.
65
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VITA
Nicholas Joseph Fredrick was born in Clarksburg, West Virginia, in August 1981
to Dr. George Fredrick and Mrs. Mary Ann Fredrick. He attended Simpson Elementary
School, Bridgeport Middle School, and Bridgeport High School in his hometown of
Bridgeport, West Virginia. In 2005, Nicholas graduated cum laude from West Virginia
University in Morgantown, West Virginia with a Bachelor of Science in Mechanical
Engineering and a Bachelor of Science in Aerospace Engineering. After graduating
from West Virginia University, Nicholas began work for Aerospace Testing Alliance at
the Arnold Engineering Development Center as a turbine engine analysis engineer.
Nicholas enrolled in the graduate program at the University of Tennessee Space
Institute in the spring of 2008.
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