Embed Size (px)
Citation preview
DESIGN AND CONTROL OF A VARIABLE GEOMETRY TURBOFAN WITH AN
INDEPENDENTLY MODULATED THIRD STREAM
DISSERTATION
Presented in Partial Fulfillment of the Requirements for
the Degree Doctor of Philosophy in the Graduate
School of The Ohio State University
By
Ronald J. Simmons, M.S.
* * * * *
The Ohio State University
2009
Dissertation Committee:
Professor Meyer Benzakein, Adviser
Professor Richard Bodonyi
Approved by
Professor Jeffrey Bons
Professor Jen-Ping Chen
Adviser
Professor Nicholas J. Kuprowicz Aerospace Engineering Graduate Program
Distribution Statement A: Unlimited Distribution.
Cleared for Public Release by AFRL/WS Public Affairs
Case Number 88ABW-2009-1697
The views expressed in this article are those of the author and do not reflect the official policy or position
of the United States Air Force, Department of Defense, or the U.S. Government.
ii
ABSTRACT
Abstract
Emerging 21st century military missions task engines to deliver the fuel efficiency of a high bypass
turbofan while retaining the ability to produce the high specific thrust of a low bypass turbofan. This study
explores the possibility of satisfying such competing demands by adding a second independently modulated
bypass stream to the basic turbofan architecture. This third stream can be used for a variety of purposes
including: providing a cool heat sink for dissipating aircraft heat loads, cooling turbine cooling air, and
providing a readily available stream of constant pressure ratio air for lift augmentation. Furthermore, by
modulating airflow to the second and third streams, it is possible to continuously match the engine‟s
airflow demand to the inlet‟s airflow supply thereby reducing spillage and increasing propulsive efficiency.
This research begins with a historical perspective of variable cycle engines and shows a logical
progression to proposed architectures. Then a novel method for investigating optimal performance is
presented which determines most favorable on design variable geometry settings, most beneficial moment
to terminate flow holding, and an optimal scheduling of variable features for fuel efficient off design
operation. Mission analysis conducted across the three candidate missions verifies that these three stream
variable cycles can deliver fuel savings in excess of 30% relative to a year 2000 reference turbofan.
This research concludes by evaluating the relative impact of each variable technology on the
performance of adaptive engine architectures. The most promising technologies include modulated turbine
cooling air, variable high pressure turbine inlet area and variable third stream nozzle throat area. With just
these few features it is possible to obtain nearly optimal performance, including 90% or more of the
potential fuel savings, with far fewer variable features than are available in the study engine. It is
abundantly clear that three stream variable architectures can significantly outperform existing two stream
turbofans in both fuel efficiency and at the vehicle system level with only a modest increase in complexity
and weight. Such engine architectures should be strongly considered for future military applications.
iii
Dedication
Dedicated to my beloved bride Bonnie.
iv
ACKNOWLEDGMENTS
Acknowledgments
I wish to express thanks to my adviser, Professor Meyer Benzakein, and the entire dissertation
committee for creating plentiful intellectual challenges, providing emotional support, and offering an
almost inexhaustible supply of patience. You recognized potential in this aging student long before I did
and cultivated a desire to live up to your expectations.
Furthermore, I would like to acknowledge a number of consummate professionals at the Air Force
Research Laboratory (AFRL). Mr. Jeffrey Stricker, Mr. Tim Lewis, Mr. Chris Norden, Mr. Jed Cox, and
Mr. Greg Bruening for their insight into variable cycle engine operation and research guidance. It is likely
that this research would have been helplessly adrift without your steadfast direction.
Additionally, I would like to recognize Dr. Tom Curran of Universal Technology Corporation for
his research into the history of variable cycle engines. To Mr. Jim Felder, Mr. Scott Jones, Mr. Tom
Lavelle, and Mr. Scott Townsend of the NPSS support group at NASA Glenn research center, you have my
most sincere thanks; without your tireless efforts this research would not have been possible.
Finally, I would like to express my most sincere gratitude to my family for their support
throughout this process. To my wife Bonnie, may God richly bless you for cups of late night coffee and
inspirational words after a demoralizing test. To my children who have been a continuous motivation to
me, I pray that God fill your heart with dreams and the faith to achieve each of them. Most of all, I wish to
thank the Lord for seeing me through this course of study and working every hindrance for good (Romans
8:28); to God be all praise, honor and glory forever.
This work was supported by the Air Force Research Laboratory, Propulsion Directorate, Turbine
Engine Division, Engine Integration and Assessment Branch, Wright-Patterson AFB, OH. The U.S.
Government is authorized to reproduce and distribute reprints for Governmental purposes notwithstanding
any copyright notation thereon.
v
VITA
August 16, 1965 ............................ Born – Harvey, Illinois
1988 ............................................... B.S. Aeronautical Engineering, US Air Force Academy
B.S. Astronautical Engineering, US Air Force Academy
1990 ............................................... M.S. Aeronautical and Astronautical Engineering, MIT
1991-1994 ..................................... Assistant Professor of Astronautics, US Air Force Academy
2006-Present .................................. Doctoral Student, The Ohio State University
PUBLICATIONS
Research Publication
1. R. J. Simmons, J.E. Cox, N.J. Kuprowicz “System level benefits of a turbofan propulsion system
equipped with an independently modulated auxiliary stream.” 56th
JANNAF Propulsion Meeting, Boston
MA, (2008).
FIELD OF STUDY
Major Field: Aerospace Engineering
vi
TABLE OF CONTENTS
Page
Abstract .......................................................................................................................................................... ii Dedication ..................................................................................................................................................... iii Acknowledgments ..........................................................................................................................................iv VITA ............................................................................................................................................................... v List of tables ................................................................................................................................................. vii List of figures .............................................................................................................................................. viii Nomenclature .................................................................................................................................................. x
Chapters: ......................................................................................................................................................... 1
1.0 Introduction .............................................................................................................................................. 1 1.1 Theoretical framework .......................................................................................................................................... 2 1.2 History of variable cycle engines .......................................................................................................................... 9 1.3 Vision missions................................................................................................................................................... 11
2.0 Computational framework ...................................................................................................................... 15 2.1 Numerical simulations ........................................................................................................................................ 15 2.2 Study engines ...................................................................................................................................................... 17 2.3 Controlling the double bypass engine ................................................................................................................. 21 2.4 Spillage drag ....................................................................................................................................................... 23 2.5 Aft body drag ...................................................................................................................................................... 26 2.6 Fuel use calculations ........................................................................................................................................... 27 2.7 Objective function and nested optimization ........................................................................................................ 30 2.8 Searching a discontinuous design space with numerous local minima ............................................................... 32
3.0 Results .................................................................................................................................................... 37 3.1 Termination of flow holding ............................................................................................................................... 37 3.2 Changes in component efficiencies with variable architecture ........................................................................... 44 3.3 Reduction in spillage drag .................................................................................................................................. 46 3.4 Reduction in aft body drag .................................................................................................................................. 49 3.5 Lift augmentation ............................................................................................................................................... 51 3.6 Heat sink capacity of third stream ...................................................................................................................... 54 3.7 Optimal tactical mobility mission variable cycle engine .................................................................................... 57 3.8 Optimal subsonic long range strike mission variable cycle engine ..................................................................... 66 3.9 Optimal supersonic strike mission variable cycle engine.................................................................................... 73 3.10 Variable features with greatest impact on performance ..................................................................................... 80 3.11 Recommended variable features for tactical mobility mission .......................................................................... 82 3.12 Recommended variable features for long range strike mission .......................................................................... 89 3.13 Recommended variable features for supersonic strike mission .......................................................................... 95
4.0 Conclusions and recommendations ...................................................................................................... 102 4.1 Viability of variable cycle engines.................................................................................................................... 102 4.2 Recommended future research .......................................................................................................................... 106
Bibliography ................................................................................................................................................ 109
vii
LIST OF TABLES
List of tables
Page
Table 1. Study engine characteristics .......................................................................................................... 17
Table 2. Rudimentary double bypass engine flow control .......................................................................... 22
Table 3. Optimal tactical mobility variable geometry at 4000 ft, Mach 0.0, 95o F design point ................. 58
Table 4. Optimal tactical mobility BPR & nozzle settings at 4000 ft, Mach 0.0, 95o F design point .......... 59
Table 5. Optimal tactical mobility variable features at 4000 ft, Mach 0.4 cruise ........................................ 60
Table 6. Optimal tactical mobility variable features at 35000 ft, Mach 0.8 cruise ...................................... 60
Table 7. Tactical mobility aircraft parameters............................................................................................. 64
Table 8. Tactical mobility mission performance ......................................................................................... 65
Table 9. Optimal subsonic LRS variable geometry at 0 ft, Mach 0.0, 95o F design point ........................... 67
Table 10. Optimal subsonic LRS BPR & nozzle settings at 0 ft, Mach 0.0, 95o F design point .................... 67
Table 11. Optimal subsonic LRS variable feature settings at 500 ft, Mach 0.7 cruise .................................. 68
Table 12. Optimal subsonic LRS variable feature settings at 40000 ft, Mach 0.8 cruise .............................. 68
Table 13. Subsonic long range strike aircraft parameters.............................................................................. 71
Table 14. Subsonic long range strike mission performance .......................................................................... 72
Table 15. Optimal supersonic strike variable geometry at 55000 ft, Mach 2.5 design point ........................ 73
Table 16. Optimal supersonic strike BPR & nozzle settings at 55000 ft, Mach 2.5 design point ................. 74
Table 17. Optimal supersonic strike variable feature settings at 30000 ft, Mach 0.5 loiter .......................... 74
Table 18. Optimal supersonic strike variable feature settings at 50000 ft, Mach 2.2 cruise ......................... 74
Table 19. Supersonic strike aircraft parameters ............................................................................................ 75
Table 20. Supersonic strike mission performance ......................................................................................... 78
Table 21. Maximum variation in component area during each vision mission ............................................. 80
Table 22. Sub optimal tactical mobility mission performance ...................................................................... 88
Table 23. Sub optimal subsonic long range strike mission performance ...................................................... 94
Table 24. Sub optimal supersonic strike mission performance ..................................................................... 97
Table 25. Performance summary................................................................................................................. 104
viii
LIST OF FIGURES
List of figures
Page
Figure 1. Notional variable cycle engine ........................................................................................................ 3
Figure 2. Propulsive efficiency as a function of mach and specific thrust ..................................................... 6
Figure 3. External airflow at maximum power design point .......................................................................... 6
Figure 4. External airflow at part power operation ......................................................................................... 7
Figure 5. Effect of inlet spillage on subsonic SFC (Johnson 1996)................................................................ 7
Figure 6. Effect of inlet spillage on overall engine drag; engines sized at 2.5Mach, 55000ft ........................ 8
Figure 7. MOBY schematic, three spool double bypass engine (Johnson, 1996) .......................................... 9
Figure 8. Double bypass variable cycle engine (Johnson, 1996) .................................................................. 10
Figure 10. Mobility vision system and tactical mobility mission profile ...................................................... 12
Figure 11. Long range strike vision system and subsonic mission profile .................................................... 13
Figure 12. Supersonic strike vision system and supersonic mission profile .................................................. 14
Figure 13. Nested optimization loops ............................................................................................................ 16
Figure 14. 2000 State of the art and advanced turbofan NPSS component based architecture ..................... 18
Figure 15. Double bypass variable cycle NPSS component based architecture ............................................ 19
Figure 16. Two dimensional illustration of double bypass engine ................................................................ 20
Figure 17. Three dimensional illustration of double bypass engine .............................................................. 20
Figure 18. Subsonic coefficient of spillage drag vs. airflow ratio ................................................................. 23
Figure 19. Supersonic reference spillage drag vs. Mach number .................................................................. 24
Figure 20. Supersonic spillage drag vs. airflow ratio .................................................................................... 24
Figure 21. Tactical mobility spillage drag as a function of throttle setting ................................................... 26
Figure 22. Aft body drag coefficient as a function of exhaust area ............................................................... 27
Figure 23. Computation of objective function within nested optimization ................................................... 31
Figure 24. Grid based search algorithm with one grid refinement ................................................................ 33
Figure 25. Gradient based search algorithm with one grid refinement .......................................................... 34
Figure 26. Creation of new generation via a genetic algorithm ..................................................................... 36
Figure 27. Internal airflow variations with unrestricted flow holding ........................................................... 38
Figure 28. Bypass ratio changes with unrestricted flow holding ................................................................... 39
Figure 29. Fan duct Mach number & pressure drop with unrestricted flow holding ..................................... 40
Figure 30. LPC efficiency and pressure ratio changes with unrestricted flow holding ................................. 41
Figure 31. Nozzle throat area variations with unrestricted flow holding ...................................................... 42
Figure 32. High cruise power hook with and without unrestricted flow holding .......................................... 42
Figure 33. Low cruise power hook with and without unrestricted flow holding ........................................... 43
ix
Figure 34. Expected change in component efficiency from design point to cruise ....................................... 44
Figure 35. Unrealistic change in component efficiency from design point to cruise .................................... 45
Figure 36. Variable cycle reduction in spillage drag at tactical mobility mission cruise points .................... 47
Figure 37. Variable cycle reduction in spillage drag at subsonic LRS mission cruise points ....................... 48
Figure 38. Variable cycle reduction in spillage drag at supersonic strike mission cruise points ................... 49
Figure 39. Variable cycle reduction in aft body drag at tactical mobility mission cruise points ................... 50
Figure 40. Variable cycle reduction in aft body drag at subsonic lrs mission cruise points .......................... 51
Figure 41. Variable cycle reduction in aft body drag at supersonic strike mission cruise points .................. 52
Figure 42. Fan pressure ratio during approach and landing .......................................................................... 53
Figure 43. Tactical mobility power hook during approach and landing ........................................................ 54
Figure 44. Variable cycle third stream air flow during assault landing ......................................................... 55
Figure 45. Theoretical heat sink capacity of third stream, subsonic LRS high altitude cruise ...................... 56
Figure 46. Fuel optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise....................... 62
Figure 47. Fuel optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise..................... 63
Figure 48. Tactical mobility range and payload for study engines ................................................................ 66
Figure 49. Fuel optimal Subsonic LRS control and performance, 500 ft Mach 0.7 cruise ............................ 69
Figure 50. Fuel optimal Subsonic LRS control and performance, 40000 ft Mach 0.8 cruise ........................ 70
Figure 51. Subsonic long range strike range and payload for study engines ................................................. 72
Figure 52. Fuel optimal supersonic strike control and performance, 30000 ft Mach 0.5 loiter ..................... 76
Figure 53. Fuel optimal supersonic strike control and performance, 50000 ft Mach 2.2 cruise .................... 77
Figure 54. Supersonic strike loiter and payload for study engines ................................................................ 79
Figure 55. Supersonic standoff range and payload for study engines ........................................................... 79
Figure 56. Effect of reduced variable features on tactical mobility mission fuel .......................................... 82
Figure 57. Effect of fixed primary nozzle throat on tactical mobility mission fuel ....................................... 83
Figure 58. Sources of improved variable cycle efficiency in tactical mobility mission ................................ 84
Figure 59. Sub optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise ....................... 86
Figure 60. Sub optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise ..................... 87
Figure 61. Sub optimal tactical mobility range and payload ......................................................................... 89
Figure 62. Effect of reduced variable features on subsonic LRS mission fuel .............................................. 90
Figure 63. Sources of improved variable cycle efficiency in subsonic LRS mission .................................... 91
Figure 64. Sub optimal subsonic LRS control and performance, 500 ft Mach 0.7 cruise ............................. 92
Figure 65. Sub optimal subsonic LRS control and performance, 40000 ft Mach 0.8 cruise ......................... 93
Figure 66. Sub optimal subsonic LRS range and payload ............................................................................. 94
Figure 67. Effect of reduced variable features on supersonic strike mission fuel ......................................... 95
Figure 68. Sources of improved variable cycle efficiency in supersonic strike mission ............................... 96
Figure 69. Sub optimal supersonic control and performance, 30000 ft Mach 0.5 loiter ............................... 98
Figure 70. Sub optimal supersonic control and performance, 50000 ft Mach 2.2 cruise .............................. 99
Figure 71. Sub optimal supersonic strike loiter and payload for study engines .......................................... 100
Figure 72. Sub optimal supersonic standoff range and payload for study engines ...................................... 101
x
NOMENCLATURE
Nomenclature
Ao Inlet stream tube area (in2)
A8 Nozzle throat area (in2)
A9 Nozzle exit area (in2)
A10 Aircraft aft body area (in2)
Ac Inlet capture area (in2)
AGL Above Ground Level (ft)
BPR1 Bypass ratio 1, mass flow of third stream / mass flow of LPC
BPR2 Bypass ratio 2, mass flow of 2nd
stream / mass flow of HPC
Cd Aircraft coefficient of drag
Cd AB Aft body drag coefficient
Cd spill Spillage drag coefficient
CD Overall aircraft coefficient of drag
CL Overall aircraft coefficient of lift
CFG Gross thrust coefficient
Cp Specific heat at constant pressure (BTU/lbm oR)
D Overall aircraft drag (lbf)
D AB Aft body drag (lbf)
D spill Spillage drag (lbf)
F Gross thrust (lbf)
Fn Net thrust (lbf)
g Gravitational acceleration (32.2 lbm ft/s2)
HPC High pressure compressor
HPT High pressure turbine
HPTB High pressure turbine blade
HPTN High pressure turbine nozzle
L Aircraft lift (lbf)
LPC Low pressure compressor
LPT Low pressure turbine
LPTB Low pressure turbine blade
LPTN Low pressure turbine nozzle
Mo Freestream Mach number
MSL Mean Sea Level
xi
fm
Mass flow rate of the fuel (lbm/s)
om Mass flow rate of air (lbm/s)
N Physical speed (%)
Nc Corrected speed (%)
OBPR Overall bypass ratio, (mass flow in 2nd
+ 3rd
streams) / mass flow of HPC
OPR Overall pressure ratio
p Total pressure (lbf/in2)
P Engine mechanical power (hp)
PSTD Sea Level, standard day pressure (14.7 lbf/in2)
Heat flux (BTU/s)
Q Fuel lower heating value (18,400 BTU/lbm)
r Range (nm)
R Gas constant for air (53.3 ft-lbf/lbm-oR)
S Aircraft surface area (ft2)
t Time (s)
To Free stream air temperature (oR)
T3 High pressure compressor discharge temperature (oR)
T41 High pressure turbine rotor inlet temperature (oR)
TSTD Sea Level, standard day temperature (518.7 oR)
TSFC Thrust specific fuel consumption (lbm/hr lbf)
Vo Vehicle velocity (nm/hr)
Ve Nozzle exhaust velocity (ft/s)
VABI Variable Area Bypass Injector
w Aircraft weight (lbf)
wf Fuel weight (lbf)
Fuel flow rate ( lbf/s)
W23 Mass flow entering the low pressure compressor (lbm/s)
W25 Mass flow entering the high pressure compressor (lbm/s)
Wc Engine demand corrected airflow (lbm/s)
descW Engine demand corrected airflow at design point, military thrust (lbm/s)
Inlet guide vane setting, used to vary component inlet area
Ratio of specific heats
o Overall cycle efficiency
p Propulsion efficiency of the jet
th Thermal efficiency of the gas generator
Atmospheric density (lbm/ft3)
1
CHAPTER 1
Chapters:
INTRODUCTION
1.0 Introduction
Early jet engines operated with a single flow stream which provided high levels of specific thrust,
but offered poor fuel efficiency. In the 1960‟s, two-stream turbofans were introduced to improve
propulsive efficiency by reducing the exhaust velocity and thereby reducing specific fuel consumption.
Over the past forty years turbofans have undergone evolutionary changes yielding significantly higher
bypass ratio engines; however, the basic cycle architecture has remained unchanged. This research will
explore the benefits of adding a second bypass stream, hereafter referred to as the third stream, to the basic
turbofan cycle architecture. When coupled to an intelligently managed variable architecture, this engine is
capable of maintaining engine airflow as power is reduced; and through this process, overall bypass ratio is
increased, effective fan pressure ratio is reduced, and an increase in propulsive efficiency is realized. This
innovative design promises both high specific thrust at military power and high efficiency at part power.
In addition to efficiently producing thrust this third stream, with a single stage of compression,
could be used for a variety of purposes. For example, it could provide a cool heat sink for dissipating
aircraft heat loads. As this would reduce heat transfer to the fuel system, it could eliminate the problem of
coking from hot fuels. It could also be used to cool air bled from the rear of the compressor; such cooled
cooling air could then be used to cool the last stages of compression and each of the turbines. This would
enable an increase in overall pressure ratio, a corresponding increase in compressor exit temperature, and a
higher turbine inlet temperature without the creation of new materials; such improvements would offer a
noticeable improvement in specific fuel consumption (Bruening, 1999). Additionally, by independently
modulating the flow through the second and third streams, it is possible to match the engine‟s demand for
airflow to the inlet‟s ability to supply airflow over a wide range of thrusts. This would offer a dramatic
reduction in spillage drag to supersonic aircraft operating at part power (Johnson, 1996). Furthermore, this
2
third stream could provide a stream of air at a constant pressure ratio independent of the throttle setting.
This air could be used for lift augmentation enabling high lift at low speeds for Extremely Short Takeoff
and Landing (ESTOL) operations (Carr, 1986). This research will investigate optimal operation of variable
cycle engines, assess the uses of the third stream air listed above, conduct cost/benefit analysis, and make
recommendations for implementation in future weapon systems.
1.1 Theoretical framework
In any engine design there are several competing performance parameters, among these are
specific thrust and fuel efficiency. High engine specific thrust is desirable for a myriad of military mission
segments including short takeoff and landing, supersonic flight, combat maneuvering, intercept, and rapid
response to time sensitive targets. Typically, specific thrust is maximized with a traditional turbojet or a
low bypass turbofan. Unfortunately, the competing demands of long duration cruise, loiter, reduced
exhaust gas temperature, noise reduction, and minimum operating costs dictate a much different engine
cycle, the high bypass turbofan. This cycle accelerates a much larger volume of air to relatively lower
velocities thereby, maximizing propulsive efficiency and minimizing fuel use (traditionally measured in
Thrust Specific Fuel Consumption, TSFC).
Maximizing the competing design parameters of specific thrust and fuel efficiency with a single
architecture began as an interesting academic exercise, and is presently evolving into operational necessity.
This pursuit began with the 1959 Supersonic Commercial Air Transport, the subsequent 1963 Super Sonic
Transport (SST), and the 1990 High Speed Civil Transport programs all of which were initiated with the
goal of creating a commercially viable supersonic transport plane. Representative of these programs was
the ambitious Boeing 2707 300 passenger, Mach 2.7, 4000 mile range aircraft (Simonsen, 2004). To be
profitable, this aircraft would require an engine capable of both efficient long duration subsonic flight,
traditionally a high bypass turbofan, and long duration supersonic flight, traditionally a much smaller cross
section turbojet. Military aircraft of this period which operated over a wide range of Mach numbers further
illustrated the need for engines with high efficiencies across their entire operating envelope.
The engines targeted to accomplish these demanding missions are called Variable Cycle Engines
(VCE). Such cycles traditionally make use of geometry changes to create high efficiency during cruise and
3
high specific thrust when the mission dictates. Note that the engine architecture changes are not typically
limited to the vane and mixer areas in the engine core; but rather, they extend to engine system variablilities
including inlet and exhaust areas (see figure 1). To fully explain how a given VCEs promise to provide
both high propulsive efficiencies and high specific thrust, it is necessary to define each of these parameters
in more detail.
Figure 1. Notional variable cycle engine
Engine cycle efficiency is usually expressed as the product of thermal and propulsive efficiency.
thpo
Where: o is the overall cycle efficiency
p is the propulsion efficiency of the jet
th is the thermal efficiency of the gas generator
This product representation of overall efficiency is very helpful in understanding how one might
improve engine performance. For example, propulsive efficiency is a measure of the effectiveness with
which the propulsive duct is propelling the aircraft. Specifically, propulsive efficiency is defined as the
ratio of aircraft thrust power to the engine mechanical power required to generate this thrust. Assuming a
negligible fuel flow rate and a perfectly expanded nozzle this can be expressed as (Mattingly, 1987),
0
02
VV
V
PowerMechanicalEngine
PowerThrustAircraft
e
p
Where: Vo is the aircraft velocity
Ve is the nozzle exhaust velocity
Modulated Cooled Cooling Air
Variable Area Compressors & Turbines
Separate Modulatable Third Stream
Substantial Cooling Air for Exhaust &
Aircraft Thermal Management
Constant Flow with Variable Fan Pressure Ratio
Adaptive
Core
Variable Core and Bypass Exhaust Nozzles
Variable Area Bypass Injector
Adaptive
Fan
Exhaust Inlet
4
Note that propulsive efficiency is maximized when the nozzle exhaust velocity is at a minimum. This is
why the most efficient turboprops and high bypass turbofans accelerate a large volume of air to a relatively
low velocity. As will soon be evident, one of the great strengths of a VCE is its ability to increase
propulsive efficiency at reduced power settings.
Thermal efficiency, on the other hand, is a measure of how well the rate of supplied energy is
converted into useful kinetic energy. Specifically, thermal efficiency is defined as the ratio of engine
mechanical power to the thermal energy input rate. Again assuming a negligible fuel flow rate when
computing engine mechanical power this can be expressed as (Mattingly, 1987),
Qm
VVm
RateInputEnergyThermal
PowerMechanicalEngine
f
e
th
2
0
20
2
Where: fm is the mass flow rate of the fuel
om is the mass flow rate of air
Q is the fuel lower heating value
Therefore, thermal efficiency is primarily a function of thermodynamic cycle improvement including
increased overall pressure ratios, combustion temperatures, and component efficiencies. While reasonable
improvements in each of these will be incorporated into this study‟s cycle, these incremental improvements
are not necessarily associated with VCEs.
One need look no further than the definition specific thrust to understand why it is impossible to
maximize both propulsive efficiency and specific thrust simultaneously with a conventional architecture.
0VV
mF
eo
Where: F is the gross thrust
Note:
Mass flow rate of fuel is considered negligible.
Only gross momentum thrust,eo Vm , and intake momentum drag,
oo Vm , terms are shown
here to illustrate the difficulty of simultaneously maximizing omF and
p ; other
effects (including spillage and aft body drag.) will be addressed in later sections.
Note that specific thrust is maximized when the nozzle exhaust velocity is at a maximum. This is why
aircraft engines in high performance aircraft are traditionally either turbojets or low bypass ratio turbofans.
5
At first glance it appears that high speed, high specific thrust applications are doomed to have
abysmally poor propulsive efficiency compared to their high bypass counterparts. However a closer
inspection reveals that as Mach number increases, high propulsive efficiencies can be achieved at higher
specific thrusts.
P
VF
PowerMechanicalEngine
PowerThrustAircraft op
oeoeo
ooeo
oeo
ooeo
VVVVm
VVVm
VVm
VVVm
2222
222
FVm
Vm
VVVm
Vm
oo
oo
ooeo
oo
oo
p
mF
V 211
1
If one were to evaluate this function at a normal cruise altitude of 36,000 ft or above, the atmospheric
temperature is relatively constant, and this equation further reduces to:
oo
p
mF
M 6011
1
Where: ooo TRMV
Mo is the free stream Mach number
R is the gas constant for air
To is the free stream air temperature
is the ratio of specific heats for air
60 is the constant from the speed of sound calculation above with units of slblb
f
m
This function for propulsive efficiency is plotted in figure 2 as a function of Mach number and specific
thrust. This figure shows that propulsive efficiency is a direct function of specific thrust, and that this
efficiency improves with reduced specific thrust at all speeds. However, the increase in propulsive
efficiency with decreased specific thrust is dramatic at low speeds, and is much less pronounced at higher
6
TS
FC
*T
SF
C*
Installed ThrustInstalled Thrust
TS
FC
*T
SF
C*
Installed ThrustInstalled Thrust
Mach numbers. For example, a propulsive efficiency of 0.8 requires a 12.5 specific thrust at 0.8 M;
however, the same efficiency can be achieved with a threefold greater specific thrust at 2.5 M.
Figure 2. Propulsive efficiency as a function of mach and specific thrust
If there were no other competing factors, one would design all engines to be high bypass turbofans
thereby minimizing the specific thrust and maximizing fuel efficiency. Unfortunately, this would result in
increased inlet cross section and produce a corresponding increased engine weight, drag and installation
losses; these losses would be particularly costly at high speeds. Furthermore, a reduced specific thrust
mandates an increased airflow for high thrust and, therefore, increased spillage drag at part power
operation. This last concept needs to be described in greater detail before one can fully appreciate the
motivation for variable cycle engines.
Spillage drag is a phenomenon which results from an engine operating away from the inlet‟s
maximum airflow, and traditionally maximum thrust, point. The engine inlet is sized to capture the
considerable air required by the engine for this high thrust design condition (see figure 3). Note that higher
Figure 3. External airflow at maximum power design point
M = 0.7 Mth = 0.7 M = 0.7 Mth = 0.7
0.0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1.0
0.0 0.5 1.0 1.5 2.0 2.5 3.0
p
Mach
F/mo=12.5
F/mo = 25
F/mo = 50
F/mo = 75
F/mo = 100
F/mo = 150
F/mo = 200
7
TS
FC
*T
SF
C*
Installed ThrustInstalled Thrust
TS
FC
*T
SF
C*
Installed ThrustInstalled Thrust
TSFC means that more fuel flow is required to produce a given amount of thrust. Notice in figure 3 that
the streamlines entering the engine are parallel and that all of the air captured by the inlet is ingested by the
engine. In a conventional engine, as the power is reduced, the airflow entering the engine is also reduced.
Unfortunately, the capture area of the inlet remains the same (see figure 4). The airflow which is captured,
Figure 4. External airflow at part power operation
i.e. slowed from the free stream velocity by the approaching inlet, but not ingested by the engine is termed
spillage (indicated by the red streamlines in figure 4). This spillage can result in significant drag especially
as the difference between current airflow and maximum airflow increases. As stated earlier, mission
requirements such as short takeoff/landing, combat maneuvering, or high Mach flight, can create missions
with a very high mass flow design point and extended duration flight at much lower mass flow
requirements. Figure 5 shows the significant increase in drag that can result from a high speed aircraft
operating at subsonic cruise with reduced engine airflow (Wa).
Figure 5. Effect of inlet spillage on subsonic SFC (Johnson 1996)
0.4 M = 0.7 Mth = 0.4 M = 0.7 Mth = 0.4 M = 0.7 Mth =
8
It is important to note that figure 5 only tells a portion of the story. While the flow mismatch is
greatest at the lowest power settings, this is also the point where the aircraft velocity is at a minimum.
Therefore, the greatest drag increase due to spillage is likely to be at an intermediate Mach number where
the product of the increased coefficient of drag and dynamic pressure is the greatest. Figure 6 illustrates
this increase in drag through the transonic region for a Mach 2.5 aircraft. If a variable cycle could maintain
airflow throughout this intermediate Mach region, a dramatic reduction in aircraft drag would be realized.
Such a variable cycle engine could be sized for maximum thrust and airflow at maximum Mach number,
thereby minimizing spillage at both the supersonic and subsonic cruise points.
Figure 6. Effect of inlet spillage on overall engine drag; engines sized at 2.5Mach, 55000ft
At this point one can appreciate the classic motivation for variable cycles. A well designed
variable cycle engine would offer an internal flow field tailored to any given mission segment. When high
specific thrust is required, the VCE sends a larger percentage of the air through the core and maximizes the
exit velocity. During reduced thrust mission segments, the VCE will port a larger percentage of the airflow
through the bypass stream(s) to maximize the propulsive efficiency. If a VCE can produce this reduced
thrust without decreasing the air flow, throttle dependent spillage drag can be effectively eliminated.
Certainly such an engine would offer increased off-design performance over a conventional non-variable
9
engine; unfortunately, this increased performance would come at the price of increased complexity, weight,
cost and an intricate control system.
1.2 History of variable cycle engines
Research into variable cycle engines has been ongoing since the 1959 Supersonic Commercial Air
Transport program highlighted the need for an engine which could be optimized for both supersonic and
subsonic cruise. Over the past five decades, no less than 25 different engine architectures were investigated
by the General Electric (GE) engine division alone (Johnson, 1996). For brevity‟s sake, only a few of these
architectures which inspired the double bypass mixed flow architecture used in this study will be addressed
here. Note that many of these systems showed great promise in overcoming the specified design
challenges; however, the cycles were too complex to be profitably pursued at the time.
In 1973 NASA reinitiated research into an efficient, low noise cycle capable of powering the
Mach 2.2-2.7 supersonic transport. At this same time the US Air force sought technologies to limit throttle
dependent installation losses (Johnson, 1976). The double bypass MOBY cycle depicted in figure 7 was
proposed by GE to meet these challenges. This aggressive cycle boasted three spools, three variable
Figure 7. MOBY schematic, three spool double bypass engine (Johnson, 1996)
turbines, three variable nozzles, two variable stator fans, and two bypass ducts one of which had a duct
burner (Johnson, 1996). This engine was capable of maintaining flow to 50% power by systematically
reducing the core and second bypass pressure rises as the throttle setting was reduced; the excess flow was
ported to the third stream thereby increasing propulsive efficiency. Core and intermediate spool speeds
10
were controlled by the intermediate and low pressure turbine vanes respectfully. While compressor op-
lines, the line which defines compressor air flow and pressure ratio throughout operation (see figure 34),
were controlled by the high pressure turbine vanes and bypass nozzle areas. This engine was quite
successful in minimizing spillage drag at part power and offered significant fuel savings over its
conventional counterpart. Unfortunately, overall system complexity prohibited further development at that
time (Johnson, 1996).
GE and other researchers noted the potential of a separated front and rear fan block with multiple
bypass ducts; if such a cycle could be created on a two spool architecture, the complexity might be reduced
enough to make this a commercially viable cycle. The first such simplification was the two spool double
bypass engine. Such architectures generally made use of diverter valves to create a single bypass engine
for maximum power and a double bypass engine for maximum efficiency at part power (see figure 8).
Notice that in this 1974 GE architecture the need for a static pressure balance at the mixing plane (just aft
of the low pressure turbine) was eliminated by a separate variable nozzle for each bypass stream as in the
MOBY engine. This considerable complexity was later reduced in 1976 by the introduction of a Variable
Area Bypass Injector (VABI) at the mixing plane just aft of the low pressure turbine. The VABI adjusts
the core and bypass inlet areas to the mixing plane and hence the total pressure of each stream. This double
bypass variable cycle provides the basis of the architecture proposed in this study.
Figure 8. Double bypass variable cycle engine (Johnson, 1996)
It would be only a few years before three stream engines were investigated throughout industry
and academia. However, the literature seems to focus on bypass streams that were preceded by diverter
11
valves. These valves, located aft of the front compressor block, were intended to divert flow to the third
stream at part power conditions. In some designs flow was either in the third stream (labeled first bypass in
the figure below) or second stream (labeled second bypass) exclusively and flow to the other stream was
interrupted entirely. Such a cycle is the selective bleed engine investigated extensively at both Canfield
and Chalmers Universities (see figure 9).
Bottom View: High pressure, maximum power mode
Figure 9. Selective bleed engine (Ulizar, 1995)
The promise of a two spool double bypass cycle offered a clear starting point for this research
project. However, the vision missions in this study require a continuous flow in the third stream for
cooling of the aircraft aft deck; this makes the two architectures above unacceptable. For this reason, the
best elements of the MOBY and double bypass architectures were combined in an effort to minimize
complexity and maximize performance. The double bypass architecture investigated in this study will
make use of a VABI at the mixing plane of the first bypass stream, thereby reducing nozzle complexity,
and a separate variable nozzle for the third stream, thereby enabling flow holding to below 50% power.
Further study engine details are presented in section 2.2.
1.3 Vision missions
This study investigated the feasibility of this variable cycle engine to effectively generate the high
specific thrust necessary for aggressive mission segments and the high propulsive efficiency necessary for
cruise and loiter. As performance improvements offered by adaptive cycles vary significantly across
operating conditions, missions with widely disparate flight segments are essential to evaluating overall
Intake LPC IPC HPC Comb. HPT LPT Mixer Afterburner Primary Nozzle
Chamber
Bypass Nozzle
First Bypass First Bypass
Core Flow Core Flow HP Shaft
HP Shaft
LP Shaft
LP Shaft
Core Flow Core Flow
Second Bypass
12
engine performance. This study will investigate the ability of the double bypass architecture to accomplish
three challenging vision missions: tactical mobility, subsonic Long Range Strike (LRS), and supersonic
strike.
The tactical mobility mission will assess the ability of a variable cycle engine to maximize both
subsonic cycle efficiency and to produce high thrust to weight for short takeoff and landing (see figure 10).
This AFRL defined vision mission, with four significantly reduced power cruise segments, should provide
significant insight into this architecture‟s ability to effectively vary flow paths and thereby maximize
efficiency. It should be noted that many other demands are placed on candidate engines by proposed
AFRL study vehicles. These demands include a readily available source of pressurized air for lift
augmentation down to 60% military thrust and a separate flow stream for aft deck cooling, aircraft thermal
load management, and to cool the aircraft hot section cooling air. For the purposes of this study, all
candidate engines must deliver 80 lbm/s of bleed air at a minimum of 1.9 pressure ratio during periods of
lift augmentation. Furthermore, variable cycle engines must maintain 15% of the total engine airflow in the
third stream for the thermal management needs described above.
Figure 10. Mobility vision system and tactical mobility mission profile
13
The subsonic Long Range Strike (LRS) mission will assess the ability of a variable cycle engine to
maximize subsonic cycle efficiency during extended high and low altitude cruise segments (see figure 11).
While this AFRL defined vision mission has a relaxed requirement for short takeoff and landing, it
demands a staggering 5000 nautical mile unrefueled range. The standoff radius, or distance from takeoff to
combat area, will serve as a figure of merit in this mission. As in the mobility mission, 15% of the total
engine airflow will be required in the third stream at all times for aft deck cooling, aircraft thermal load
management, and to augment cooling of the engine hot section.
Figure 11. Long range strike vision system and subsonic mission profile
The final mission was selected due to its long range supersonic and prolonged loiter segments.
Again, standoff distance will be used as a figure of merit in this mission (see figure 12). This mission will
task the VCE severely; it will require a high specific thrust engine for supersonic cruise and a highly
efficient engine for the prolonged loiter. Furthermore, the inlet which has been sized for the Mach 2.5 dash
requirement will spill air at a significant rate if the engine cannot hold air flow at both the reduced power
Mach 2.2 cruise and Mach 0.5 loiter conditions. As in previous two missions, 15% of the total airflow will
be required in the third stream at all times for aft deck cooling, aircraft heat rejection, and to augment
cooling of the engine hot section.
14
Figure 12. Supersonic strike vision system and supersonic mission profile
AFRL Propulsion Directorate studies indicate that no existing commercial engine can accomplish
these missions, nor can this proposed architecture without variable features. Therefore, the double bypass
mixed flow engine introduced above will require proper scheduling of adaptive features to maintain stable,
optimal performance throughout the flight envelope. A major research challenge was to establish a suitable
methodology to determine the optimal method of scheduling this VCE while minimizing the inlet, duct and
discharge flow losses associated with varying cycle geometry. The variable feature scheduling necessary to
maintain the cycle within temperature, speed, surge, and stall limits will be addressed in Chapter 3.
2.2 Mn / 50,000 ftFull weapons complement
2.2 Mn / 50,000 ft
Penetration
2.2 Mn / 50,000 ft
15
CHAPTER 2
COMPUTATIONAL FRAMEWORK
2.0 Computational framework
2.1 Numerical simulations
The Numerical Propulsion System Simulation (NPSS), a thermodynamic cycle analysis package
jointly developed by NASA and industry, was selected to conduct performance analysis in this study. This
code realistically models physical interactions within the engine with a component based, object oriented
cycle simulator. Its solver finds steady state and transient solutions subject to flow continuity, shaft power
balance, and user defined constraints. This cycle analysis package has been extensively verified against
proprietary tools which have been validated against existing engine performance data; comparisons
between these tools indicate that NPSS performance analysis is consistent with that predicted by legacy
tools at all operating conditions. The acceptance of NPSS as the US industry standard, the relative ease
with which data can be shared between programs and, the ability to make real time cycle architecture
changes makes this program ideal for this study.
Early in this study compressor vane and turbine nozzle settings were varied manually. However,
it soon became abundantly clear that just finding locally optimal solutions would require far too great a
time investment; a truly comprehensive search of the design space would require a more automated
process. For this reason the NPSS model was modified to communicate directly with Model Center®
, a
commercially available visual environment for process integration, and its integral optimization routines
via plug-ins. With this revision, NPSS output can be sent directly to an objective function which then
evaluates the suitability of a solution based on cycle efficiency and a myriad of constraints including
maximum temperatures, minimum surge margins, maximum spool speeds, minimum component pressure
ratios and acceptable component efficiencies. Then an optimization package, here a genetic algorithm,
generates new promising variable geometries for NPSS to evaluate. With this architecture in excess of
16
20,000 engine designs can be investigated per hour on a single processor PC and a more globally optimal
solution determined.
The next hurdle was error rejection by the NPSS software. As the very nature of a genetic
algorithm is to randomly seed the search space (defined by physical limits of variable features) ancontinue
searching the vicinity of more promising regions, many variable feature settings sent to NPSS will have no
possible converged solution. Much effort was expended to assure that NPSS always fails in a benign
manner in these instances. Specifically, the solver was modified to return to a previously converged
solution and to send a non-converged error flag to the objective function. Given these modifications
optimization runs could be completed even if several of the on or off design cases have no valid solution.
Once this was accomplished optimization loops could be nested to more fully search the on design
and off design space (see figure 13). It is important to note that the objective function seeks to minimize
mission required fuel load while ensuring a physically viable solution. In this process, spillage and aft
body drag are calculated and the engine is throttled until the desired installed thrust is obtained. As such,
the software is able to determine when it is more advantageous to cease flow holding at a power setting
greater than the point of interest and continue with a conventional power hook. The optimization
architecture is addressed further in section 2.7.
Figure 13. Nested optimization loops
Design Point
(NPSS Model)
Objective
Function
(OF)
Genetic
Algorithm
(GA)
Off Design Optimization Loop
Off Design Optimization Loop
On Design Optimization Loop
GA
NPSS OF
Off Design Point 1
GA
NPSS OF
Off Design Point n
17
2.2 Study engines
In order to evaluate the benefits of the double bypass variable cycle engine, two reference engines
were modeled. The first is a year 2000 state of the art (SOA) turbofan, and the other an advanced turbofan
with increased compressor exit temperature (T3), turbine inlet temperature (T41), inlet and exhaust
performance, component efficiencies and pressure ratios (see table 1).
2000 SOA Turbofan Advanced Turbofan Double Bypass VCE
% Adiabatic efficiency (Fan/LPC/HPC/HPT/LPT)
85 / na / 85 / 87 / 88 88.5 / na / 86 / 89 / 90 88.5 / 88.5 / 86 / 89 / 90
Cooling %W25* (HPTN/HPTB/LPTN/LPTB)
10 / 5 / 5 / 2 15 / 10 / 4 / 2 15 / 10 / 4 / 2
Modulated 15 / 5 / 2 / 1 **
Primary Nozzle CFG Pri. Nozzle Cooling CFG***
Fan Nozzle CFG
0.95 0.92 n/a
0.97 0.92 n/a
0.97 n/a
0.96
Inlet Flow Control (%W23) 0.0 2.5 2.5
Ram Recovery 0.95 0.97 0.97
T3max (0F) 1200 1400 1400
T41max (0F) 2940 3400 3400
VABI No Yes Yes
* Both the advanced turbofan and variable cycle utilize cooled cooling air
** Variable cycle modulates cooling at cruise power
*** 2000 SOA and advanced turbofans use film cooling of the primary exhaust nozzle
Table 1. Study engine characteristics
Notice that the increased maximum compressor exit temperature in the advanced core turbofan
and double bypass engines requires cooling of the HPC rear disk and the HPT blades. This is modeled by
bleeding air from the rear of the compressor and cooling it to 1150o F, which is 250
o F below T3max. The
heat removed from this flow is then introduced into the bypass stream. Finally, the cooled cooling air
reenters the flow path as HPT chargeable air. Note also that all three engines are assumed to be embedded
within the fuselage with appropriate pressure losses. The latter two engines make use of inlet flow control,
air bled from the LPC exit then injected into the inlet, to improve the ram pressure recovery.
A component based model of all three study engines was constructed in NPSS (see figures 14 and
15). Each engine was modified to accommodate the bleeds, heat exchangers, and other components
Figure 14. 2000 State of the art and advanced turbofan NPSS component based architecture
F
A
N
Amb0
Inlet
IFC
Front
Frame
Split 1
Comp
Duct
Burn 36
B41
B42
B45
B52
Leakage
Primary
Nozzle
Sink
Primary
Nozzle
D52
Mixer
B7
H
P
C
H
P
T
L
P
T
Fan
Duct
Fuel 36
Gear
Box
OB
Sink
Customer
Bleed
HP
Extraction
Aftr
Burn
Fuel AB
Split 2
Nozzle
Cooling
Sink
Nozzle
Cooling
Rear
Frame
Nozzle
Cooling
Duct
Wing
Sink
Wing
Blowing
Legend:
Air Flow path
Cooling Air Flow
Fuel Input
Shaft Connection
Nozzle Cooling (modeled as a separate nozzle)
Variable Components:
- Nozzle throat area
- Primary & secondary inlet areas to mixer
18
Figure 15. Double bypass variable cycle NPSS component based architecture
Legend:
Air Flow path
Bleed Air Flow
Fuel Input
Shaft Connection
Variable Components:
- Nozzle throat areas
- Inlet Area to Compressors & Turbines
- Primary & secondary inlet areas to mixer
Heat Out
F
A
N
Amb0
Inlet
Inlet
Flow
Control
Front
Frame
Split 1
Fan
Duct
Split 2
Comp
Duct
Fan
Nozzle
Burn 36
B41
Fan
Nozzle
Sink
B42
B45
B52
Leakage
Primary
Nozzle
Sink
Primary
Nozzle
D52
Mixer
B7
L
P
C
H
P
C
H
P
T
L
P
T
Fan
Duct 2
Fuel 36
Gear
Box
OB
Sink
Customer
Bleed
HP
Extraction
LPC
Duct
After
Burner
Fuel AB
Rear
Frame
B24
Heat In
Heat
Exchr
Wing
Sink
Wing
Blowing
19
20
necessary to accomplish the vision missions introduced in section 1.3. Notice that all yellow components
are variable; these will be exercised throughout this research to determine which are essential to optimizing
variable architectures. Notice the two conventional turbofans have separate „nozzle cooling‟ nozzles with a
reduced coefficient of gross thrust, CFG, as depicted in table 1. In the double bypass engine aft deck
cooling is performed by the fan nozzle; however, this configuration does not use film cooling and has a
slightly greater CFG.
So that comparisons from baselines to the adaptive cycle engine could be made effectively, all
study cycles were sized to a specified airflow for a given mission. The fan pressure ratio was then varied
until the required design point thrust was obtained. Finally, the HPC pressure ratio was varied until the
desired overall pressure ratio was achieved. As a result each engine was sized to the same inlet area,
airflow and installed thrust for a given mission. Furthermore, first order performance effects such as T3,
T41, and OPR were held constant for the advanced turbofan and double bypass engines. Performance
differences between these two engines are therefore strictly a function of the variable architecture. As one
would expect, the high specific thrust mission segments drives each engine to a relatively low bypass
configuration (see figures 16 and 17 for notional two and three dimensional flow paths).
Figure 16. Two dimensional illustration of double bypass engine
Figure 17. Three dimensional illustration of double bypass engine
21
2.3 Controlling the double bypass engine
Key to the performance of any variable cycle is the ability to increase flow in the bypass streams
as the engine is throttled back thereby increasing propulsive efficiency. Therefore technologies that limit
core air flow and promote movement of that flow into the more fuel efficient bypass streams are critical to
improved propulsive efficiency. This section will suggest one such mechanism for varying these flows and
briefly address the costs associated with such variations.
Control of this engine during flow holding is a relatively simple procedure. Corrected airflow is
held constant by operating the fan at 100% corrected speed throughout the period of flow holding. As the
engine thrust is reduced, flow through the core is reduced in three ways. First the inlet area to the low
pressure turbine (LPT) is decreased; this reduces airflow through the LPT and hence the engine core.
Second, the inlet area to the high pressure turbine is decreased which further reduces air flow through the
core. Finally, the inlet area to the low pressure compressor is decreased which reduces airflow to both the
core and second streams for a fixed low spool speed. The airflow that can no longer be accepted by the
core is sent to the second and third streams. Then the bypass ratio to the second stream adjusts to maintain
a static pressure balance while the variable area bypass injector located at the mixing plane adjusts to
maintain the desired secondary to primary stream total pressure ratio. Finally the fan and LPC operating
lines are maintained by adjusting the fan nozzle and primary nozzle throat areas respectively.
The creation of a robust and fault tolerant solver to perform the numerical analysis proved to be a
significant challenge. The first decision to be made was how to distribute the variable search space
between NPSS and the genetic algorithm optimizer in Model Center®. As the input and output of variables
at the plug in interface is the slowest element of the optimization process, it was decided to maintain
control of all flow continuity and work balance variables within NPSS. Therefore, the NPSS solver
maintains control of the primary nozzle throat area, VABI positioning, as well as the high and low spool
speeds. Furthermore, NPSS was given control of the LPC inlet guide vane through termination of flow
holding to maintain a minimum LPC surge margin and pressure ratio. This means that the remaining
component inlet areas and the fan nozzle throat area are varied by the optimizer in an effort to shift internal
airflow off design and determine globally optimal settings for any given flight condition.
22
As one would expect, the automated search of the design space locates many sets of vane angles
that do not return converged NPSS solutions. Much effort was expended in conjunction with experts at
NASA Glenn research center to identify and trap all the errors returned by NPSS during this
comprehensive off design space search. Furthermore, software was improved in NPSS to enable a rapid
restore of solver independent variables to a previously converged case in the event of solver non-
convergence or any other identified error. The release of NPSS version 1.6.5 incorporated each of these
changes and integrated the ability to modify on-design variables via Model Center®
plug ins; these
improvements have aided variable cycle modeling efforts throughout the modeling community.
To illustrate this engine control, table 2 shows the inlet guide vane settings as well as the primary
and fan nozzle throat areas for a notional mobility engine at design point and two flight conditions. Note
that the HPC vane angle is not varied in this study; instead a map with an embedded vane schedule is used.
As described above, HPT, LPT and LPC inlet areas are simultaneously decreased (smaller turbine inlet
areas are represented by smaller angles while smaller compressor inlet area is represented by larger angles)
as the thrust is reduced. Since the inlet air flow rate is held constant, air flow which cannot be accepted by
the core is sent to the two bypass streams. Notice that the bypass to the third stream increases tenfold and
the overall bypass ratio (OBPR) is nearly doubled at low altitude when flow holding to 52% power.
Table 2. Rudimentary double bypass engine flow control
23
The cost of increased propulsive efficiency can at times be unacceptable in terms of mechanical
complexity. Notice that the primary nozzle can have a threefold or greater increase in throat area when the
air flow rate is held constant during the entire power hook; however if flow holding is terminated at
approximately 80% power, the throat area variation can be cut in half and the specific fuel consumption
further reduced (see 35,000 ft cruise condition). This suggests that there are times when a minimal amount
of spillage drag should be accommodated in order to minimize both specific fuel consumption and
unnecessary mechanical complexity. Optimal engine control and the most favorable time to cease flow
holding will be addressed further in Chapter 3.
2.4 Spillage drag
Spillage drag is assessed during each mission segment in an effort to determine whether the
reduced overall drag helps justify the increased complexity of this variable cycle engine. To calculate
spillage drag, one must first determine a relationship for coefficient of drag as a function of corrected
weight flow. A notional AFRL furnished sharp edged, subsonic embedded inlet with the following drag
profile is used (see figure 18).
Figure 18. Subsonic coefficient of spillage drag vs. airflow ratio
24
This profile enables the calculation of drag coefficient for both tactical mobility and subsonic long range
strike missions as a function of inlet weight flow.
For the supersonic strike mission, the notional AFRL furnished Mach 2.5 sharp edge, embedded
supersonic drag profile is used (see figures 19 & 20). The inlet spillage drag coefficient for this mixed
Figure 19. Supersonic reference spillage drag vs. Mach number
Figure 20. Supersonic spillage drag vs. airflow ratio
0.00
0.01
0.02
0.03
0.04
0.05
0.06
0.5 0.7 0.9 1.1 1.3 1.5 1.7 1.9 2.1 2.3 2.5
Cd
sp
ill r
ef
Mach Number
Reference Inlet Spillage Drag Coefficient
0.00
0.05
0.10
0.15
0.20
0.25
0.30
0.35
0.40
0.3 0.4 0.5 0.6 0.7 0.8 0.9
C d sp
ill a
rea
Ao/Ac
Inlet Spillage Drag Coefficient
Mach 2.0
Mach 1.6
Mach 1.2
Mach 0.8
Mach 0.6
25
compression inlet is determined by summing the Cdspill area and Cdspill ref in these two figures. Note that the
stream tube to capture area ratio in Figure 20 is similar to the corrected weight flow ratio in Figure 18; both
of these ratios relate how closely the inlet is operating to its design airflow for a given Mach number.
As the corrected airflow numbers are output by NPSS for each of the points of interest, only the
inlet capture area must be calculated to determine the spillage drag at any given flight condition. This inlet
capture area is given by (Oates, 1978),
21
12
12
21
)1(5.01
R
gMM
T
P
A
Woo
std
std
o
c
Where: Ao is the inlet stream tube area
g is the acceleration of gravity at mean sea level
Pstd is sea level, standard day pressure
Tstd is sea level, standard day temperature
Wc is the engine demand corrected airflow
Inlet capture area, Ac, is then calculated with the value of Ao at military thrust, maximum cruise Mach
number and historical Ao/Ac ratios for inlets. Finally, the spillage drag can be determined as
cospilldspill AVCD2
21
Where: Ac is the inlet capture area
D spill is the spillage drag
Cd spill is the spillage drag coefficient. Note that this is the sum of Cd spill ref and Cd spill area
plotted above for the supersonic mission
is the atmospheric density
Given the process outlined above, one can now estimate the spillage drag for a given flight
condition. Figure 21 illustrates the increase in spillage drag as a function of percent thrust for the tactical
mobility mission advanced core turbofan. As the mobility platform operates at 60% military thrust at the
35,000 ft cruise point, the spillage drag here is only 1.9%. In this figure, the low altitude penetration curve
has a much lower slope, but it also has a much lower thrust requirement; here the percent thrust is only 32%
and the resulting spillage drag is 5.6%. While this mission has the lowest potential spillage drag, even here
26
it appears advantageous to flow hold as long as practical. Potential spillage drag in the other two vision
missions and conditions in which it is better to cease flow holding will be addressed further in Chapter 3.
Figure 21. Tactical mobility spillage drag as a function of throttle setting
2.5 Aft body drag
It was also hoped that aft body drag could be reduced by variable cycle architectures and that any
such reduction would help justify the cost and complexity of these advanced engines. For this reason, aft
body drag is calculated at each point of interest in a manner similar to spillage drag. These calculations
begin with a notional coefficient of aft body drag profile (see figure 22).
Notice that the coefficient of drag is a function of the ratio between aircraft aft body area, A10, and
nozzle exhaust area, A9. As this is treated as an aircraft re-engining program, a single A10 is used for each
mission of interest. Therefore, as A9 varies across the mission profile a coefficient of aft body drag can be
easily calculated. Then drag is easily determined as the product of dynamic pressure, coefficient of drag
and aft body area,
10
2
21 AVCD oABdAB
Where: D AB is the aft body drag
Cd AB is the aft body drag coefficient
0
2
4
6
8
10
12
14
16
18
20
0 10 20 30 40 50 60 70 80 90 100
% I
ncre
ase
in
Air
cra
ft D
rag
du
e t
o S
pil
lag
e
% Military Thrust
Tactical Mobility Spillage Drag
Advanced Turbofan, 4000 ft 0.4 Mn
Advanced Turbofan, 35000 ft 0.8 Mn
27
As aft body drag is typically a small percentage of vehicle thrust, very modest reductions in fuel
use are possible by minimizing this drag source. However, air flow holding in variable cycles did yield
modest reductions in A10/A9 and therefore reduced aft body drag; these results are presented in Chapter 3.
Figure 22. Aft body drag coefficient as a function of exhaust area
2.6 Fuel use calculations
In this study the basis of the objective function is required mission fuel; for this reason, a
simplified relationship between desired range and fuel required is sought. Deriving such a relationship
begins with the definition of fuel flow,
Where: t is time
w is aircraft weight
is weight of the fuel
is fuel flow rate
Therefore, the rate of change of fuel with respect to range rate can be expressed as
0.00
0.05
0.10
0.15
0.20
0.25
1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0
C d A
B
A10 / A9
Aft Body Drag Coefficient
Mach 2.0
Mach 1.2
Mach 1.0
Mach 0.8
28
Where: r is range
Substituting this expression into the definition of fuel flow and solving for range yields,
For jet engines fuel flow above is typically expressed as a function of the desired thrust,
Where: TSFC is thrust specific fuel consumption
is net thrust
Now if steady level flight is assumed, one can relate thrust to aerodynamic constants and vehicle weight as
follows,
Where: CD is the aircraft coefficient of drag
CL is the aircraft coefficient of lift
D is aircraft drag
L is aircraft lift
S is aircraft surface area
Note: Steady level flight assumption implies
fixed angle of attack, therefore both CL and CD are constant
altitude is constant, therefore is constant
thrust is constant, therefore TSFC is constant
unaccelerated flight, therefore L=w and Fn= D
29
Therefore, the integrand in the range equation above can be written as,
Substituting this into the range equation and solving yields,
This simple expression for range, commonly referred to as the Breguet range equation, offers a
means to rapidly calculate fuel burn for any given constant altitude mission segment. This straightforward
equation is an obvious starting point for this study‟s objective function; a function that is to be called tens
of millions of times during various optimization analyses.
A few additional assumptions are made when calculating mission fuel use. Each of these
assumptions tends to be very conservative in nature; that is to say that each minimizes the potential benefits
offered by an advanced, more fuel efficient engine. Therefore, actual performance of a three stream
adaptive engine will likely exceed that presented in this document. These assumptions include:
The vast majority of fuel burn is consumed in the level flight segments of each mission; therefore,
comparisons of fuel usage during these segments provide a reasonable and sufficient means of
evaluating the relative benefit of advanced cycles. This is conservative since reduced fuel use
realized during ground idle, taxi, takeoff, climb, descent, approach and landing would only
accentuate the benefits of advanced cycles.
Performance analysis is conducted as if this were an aircraft re-engining program; that is to say
that each of the engine designs is evaluated within a common airframe and at predetermined
installed thrust setting for each mission segment. Therefore structural weight, maximum fuel load,
30
and aerodynamic performance are all constant regardless of the engine being assessed. Again this
approach minimizes the benefits resulting from reductions in vehicle size, structural weight, and
thrust required that would be observable in a new aircraft program.
A 10% fuel reserve was assumed for each of the candidate missions.
Only the installation effects of spillage and aft body drag were considered in this analysis.
2.7 Objective function and nested optimization
Given the Breguet range equation and the assumptions above, it is possible to formulate an objective
function capable of rapidly assessing a candidate engine‟s potential to accomplish a particular mission.
The objective function used in this study is mission fuel required debited for a number of penalties
including: excessive component corrected speeds, unreasonable component efficiencies, insufficient
compressor pressure ratios, insufficient stall margins, and insufficient aft deck cooling. A severe penalty is
applied for clear physics violations such as failure of cycle to converge, fan diameters greater than
allowable, insufficient cruise thrust, excessive compressor exit temperature, and excessive turbine inlet
temperature. A more detailed block diagram of the objective function calculation within the nested
optimizations architecture (first introduced in Figure 13) is provided in Figure 23.
There are a few details in this block diagram that need to be discussed. First, both spillage and aft
body drag are accounted for when evaluating engine performance at each point of interest. As each of
these drag terms is fundamentally a function of the current throttle and variable architecture settings, the
NPSS solver must iterate at each point of interest until the installed thrust equals the desired cruise thrust.
When one considers the substantial benefits of minimizing these drag terms with a variable cycle, the
minimal computation time required for these calculations is easily justified.
Second, the cruise points of interest are evaluated in parallel NPSS instances. As such, each point
of interest must make an educated guess at the TSFCinst at the other locations in order to calculate an
objective function during its off design optimization loop. Fortunately this is not a problem since each off
design loop is seeking only to optimize performance at its particular point of interest. Furthermore, at the
end of each on design optimization loop, a common cost function is executed with the performance output
of each point of interest.
31
Design Point Optimization /
‘Fly’ to Desired Altitude and Mach Read in on design parameters from GA
Run design point & check performance
Compute a series of off design points until desired altitude & Mach is reached
Power hook to cruise thrust Throttle to desired thrust
Find spillage & aft body drag
Calculate installed Thrust (Fninst) Fninst = Fn - DAB - Dspill
Iterate until Fninst = desired thrust
Calculate installed TSFC (TSFCinst)
TSFCinst = / Fninst
Point of Interest (POI) Optimization Read in off design parameters from GA
Rerun POIs (in parallel NPSS instances)
Each POI estimates objective function; see calculation details below
(NOTE - TSFCinst for other POI assumed)
Off Design Optimization Loop
GA selects new Fan, LPC, HPT, LPT Is Genetic Yes inlet guide vanes (IGV) & Fan A8 settings Algorithm Generation (NOTE - settings different for each POI)
< Max allowed? Optimization progresses subject to limits
on vane & nozzle displacements as well as max generations without improvement
No
Calculate overall objective function If POI failed to converge or a constraint
violated, set objective function = 5 * 108
Guess fuel load required for mission
Find fuel reserve with Breguet range eq.
using the TSFCinst calculated at each POI
If fuel reserve ≠ 10%, modify fuel load & iterate until a 10% reserve is obtained
Objective Function = Mission Fuel Load + penalties for undesirable behavior
If fuel load does not equal max fuel load, find max standoff range and loiter time
Print optimal off design settings and performance data for this design point
On Design Optimization Loop
GA selects new IGV settings, Op lines, Is Genetic Yes corrected speeds & 3rd stream bypass ratio
Algorithm Generation Optimization progresses subject to limits
< Max allowed? on vane displacements, corrected speeds,
minimum third stream bypass ratio, and max generations without improvement
No
Optimization Complete
Figure 23. Computation of objective function within nested optimization
32
Finally, there are many engine designs that are simply incapable of performing stated missions
with the available fuel load. As mentioned in the assumptions above, required fuel load is always
calculated without increasing the aircraft size to accommodate fuel loads greater than the predetermined
max takeoff fuel. This approach is consistent with an aircraft re-engining program, and required fuel loads
greater than max takeoff fuel are simply understood to represent unacceptable engine designs.
Furthermore, maximum achievable high altitude cruise leg (standoff range) and loiter time, if applicable, is
calculated for each engine; similarly, ranges and loiter times less than those stated in the mission represent
unacceptable engine designs. These range and loiter figures of merit will be discussed further in chapter 3.
2.8 Searching a discontinuous design space with numerous local minima
The search methodology outlined in section 2.7 describes a nested genetic algorithm structure used
to determine the most optimal variable cycle engine for a given mission. The reason for this search
architecture is quite simple; if one has an engine capable of varying internal geometries, and hence internal
flows, it is highly probable that the variable feature settings would vary from the design point to each cruise
point of interest. Furthermore, the amount of flow variation from the core to the second and third streams
is a strong function of the design point selected. Therefore to be effective, engine optimization must
simultaneously investigate both the on design search space and the associated off design search space at
each cruise point of interest.
While the basic search architecture appears self evident, the selection of an appropriate
optimization method is a bit more difficult. To begin this process one must first understand the nature of
the on and off design search spaces. Both of these are replete with locations that violate either explicitly
stated design constraints such as minimum surge margin, maximum shaft speed, or maximum fan diameter,
or simply will not satisfy the most basic physical cycle requirements including shaft balance, mixing plane
pressure balance or maintaining subsonic flow in all ducts. While intelligently limiting the design variable
search range can mitigate some of these effects, there still exists a myriad of locations throughout the
search space for which no converged solution is possible. Unfortunately, the objective function response
surface not only has a number of unacceptable locations with essentially infinite cost, but it also abounds
33
with local minima (Millhouse, 2002). Therefore, the selection of a suitable search algorithm is absolutely
essential to making a reasonable exploration of this objective function and to drawing rational conclusions.
Each of the algorithms considered here begins with a specified range on each design variable.
Then this search space is discretized or seeded with a reasonable number of initial points. How one
proceeds from this initial state determines the character, efficiency, and reliability of the search method.
For example, a purely random search would arbitrarily sample the design space within the specified range.
With sufficient sampling locations this unstructured search would provide a reasonable understanding of
the space but would likely not return the global minimum of the cost function. Such a search would require
a large number of sample points, especially as the number of design variables increased, and therefore a
greater computational time than its structured counterpart; however, random searches do not typically fail
even in a completely random design space.
Although the objective function is remarkably complex in this problem, it is not random; for this
reason one might be compelled to explore a more structured search algorithm. The simplest such structured
algorithm is known as an enumerative or grid based search in which the search space in a divided into an
evenly spaced grid across the user defined range space. The cost function is then evaluated at each of the
intersections and a relative minimum is located among these initial points. At the lowest objective function
location, the grid is refined by reducing the spacing between grid points and the cost function is again
evaluated at each intersection. The process continues until a specified number of refinements has been
accomplished or until the grid size reaches a predetermined minimum, see figure 24.
Figure 24. Grid based search algorithm with one grid refinement
34
This rudimentary enumerative search is capable of finding a minimum value using a comparable
number of objective function evaluations as used in a random search. However if a cost function has many
local minima, the value returned by this method is not likely to be the global minimum. The easiest way to
mitigate this problem would be to increase the number of points in the original gridding thereby increasing
the probability that the global minimum is located. Unfortunately a finely gridded search significantly
increases the computation time required for even a modest sized multidimensional problem, thereby
rendering this methodology unacceptable for this study.
In an effort to avoid computationally intensive enumerative searches, researchers have developed
a family of algorithms known as calculus or gradient based optimizers. In their simplest form, these
methods begin from a seed point and attempt to find the minimum of an objective function by continuously
moving in the most favorable, or maximum gradient, direction. It is important to note that such methods
need not evaluate the partial derivatives of the cost function with respect to each design variable to be
successful; therefore, they can be used effectively on discontinuous objective functions. A rudimentary
gradient based search is illustrated in Figure 25. This search begins at point A and evaluates the objective
at each of the surrounding grid points x using a fixed step size. The search progresses to the lowest of these
objective evaluations B. This movement in the direction of the maximum gradient continues until a relative
minimum is located, here point D. The grid is then refined further and the process continues until a user
defined minimum step size is reached.
Figure 25. Gradient based search algorithm with one grid refinement
35
Gradient based methods require far fewer objective evaluations than random or grid based
searches described above and yield very good results for functions with a single minimum. However, these
methods often fail to return optimal solutions in functions with multiple local minima or when the initial
point is selected far from the global minimum (Goldberg, 1989). There are a number of modifications to
this basic architecture that can be employed to improve the probability that a more global minimum is
located in functions with several local minima. For example, one could begin with multiple initial points
and find local minima associated with each. Additionally, one could modify the gradient method to
incorporate random step directions, random step sizes or even the occasional movement in a direction with
a higher objective function (often called simulated annealing). These three modifications attempt to help
the algorithm move out of flat regions or local minima in which the search algorithm is currently „stuck‟.
With each modification, the computational time increases along with the probability that a more global
optimum would be located. Unfortunately even with the modifications outlined above, gradient based
searches are incapable of adequately searching the vast, discontinuous and noisy search space of this study.
Fortunately there is a search algorithm that is able to locate a global optimum with the regularity
of a fine mesh grid while requiring significantly fewer objective function calls. These search methods,
called genetic algorithms, begin by discretizing the search space between the allowable minimum and
maximum values. Then this space is seeded with a set of representative designs, called a population,
throughout this space. It is important to note that although this initial population is generally much smaller
than its traditional grid search counterpart, the very nature of a genetic algorithm yields a much higher
resolution than a grid search with the same gridding.
This increased resolution is accomplished by the unique nature by which subsequent points are
selected in a genetic algorithm. First each individual design in a population is described by a chromosome,
which is a series of ones and zeros that represents the design variable selection for that individual. The
objective function is evaluated for this and every individual in the initial generation. The most promising
designs are then chosen to reproduce in this generation; this process is called selection and determines not
only those who will reproduce but also the rate at which they will do so. During this reproduction,
elements of the parent chromosomes will pair together to form children with characteristics similar to the
parents. Finally, occasional random changes to the child chromosomes will be made in a process known as
36
mutation in an effort to recover genetic material lost in the selection or crossover processes. This process is
summarized for a notional two variable system in figure 26.
Figure 26. Creation of new generation via a genetic algorithm
While not as efficient as calculus based algorithms at solving problems whose objective has a
single minima, genetic algorithms have proven remarkably robust across a broad spectrum of problems. It
accomplishes this first by working with a population of points rather than a single point; this population,
which becomes increasingly well adapted with each generation, reduces the probability of reaching a false
minima. Second, the genetic algorithm uses objective function information only and not its derivatives nor
any other auxiliary information; this eliminates any susceptibility to discontinuities in the cost function and
makes them applicable to virtually any problem. Finally, genetic algorithms make use of probabilistic
transition rules to guide the search to more promising regions of the search space (Goldberg, 1989). These
unique features make a genetic algorithm well suited to search the immense, noisy, and discontinuous
search space presented in this study.
37
CHAPTER 3
RESULTS
3.0 Results
During the course of this research several million variable cycle engine designs were evaluated
using the optimization method outlined above. Through analysis of this data a great deal of insight into the
nature and potential of the double bypass engine was garnered. This chapter begins by summarizing some
of the major design, control, and computational lessons learned. Then an optimal variable cycle is
presented for each vision mission along with a schedule of variable features at each point of interest.
Finally, a study of the variable features with greatest performance enhancement is conducted and a sub-
optimal variable cycle is recommended for each mission which achieves superior performance with the
fewest variable features.
3.1 Termination of flow holding
Early in the research it became apparent that the double bypass VCE examined in this study was
remarkably adept at holding corrected airflow to very low power settings. While this is traditionally touted
as a significant benefit of variable cycles, one begins to wonder if indefinite flow holding truly yields the
minimum fuel use. Furthermore, prolonged flow holding could cause even more troubling aerodynamic or
mechanical problems. For these reasons some effort was expended in determining the most advantageous
time to terminate flow holding.
The following discussion is based on a notional tactical mobility engine operating at the high
cruise point of interest. The control of this engine has been modified to hold airflow until one or more of
the following physical limits is reached: compressor pressure ratio of one, no airflow in a duct approaching
a mixing plane, or supersonic flow in a duct. Therefore, the illustrations that follow do not represent the
optimal control of a variable cycle engine, but rather a depiction of the changes in internal flow and the
38
associated costs of indefinitely holding engine airflow constant. At the termination of this analysis, a more
appropriate time for termination of flow holding will be offered along with the associated enhancements in
cycle performance.
Figures 27 and 28 further illustrate the basic concepts of flow holding first introduced in Table 2.
Notice in Figure 27 that the overall engine airflow remains constant even though power is reduced. This is
primarily accomplished by closing the LPC inlet, HPT inlet, and primary nozzle throat areas while
simultaneously increasing the fan nozzle throat area (detailed optimal engine control is presented in
sections 3.5 thru 3.7) This discourages flow to the engine core and the second stream while promoting flow
to the third stream. Therefore, both the bypass ratio to the third stream and the overall bypass ratio increase
as power is reduced (see Figure 28). If there were no associated costs associated with these flow changes,
one would gladly accept the associated decrease in spillage drag and increase in propulsive efficiency.
Unfortunately, indiscriminately varying these internal engine flows does come at a price. As will soon be
evident, the performance costs associated with excessively changing internal flows will ultimately exceed
any propulsive benefit realized.
Figure 27. Internal airflow variations with unrestricted flow holding
0
50
100
150
200
250
60 65 70 75 80 85 90 95 100
Mas
s Fl
ow (l
b m)
% Military Thrust
Unrestricted Flow Holding, Airflow vs. Installed Thrust
Overall Airflow
Third Stream Airflow
Second Stream Airflow
Core Airflow
39
Figure 28. Bypass ratio changes with unrestricted flow holding
A dramatic rise in duct losses associated with prolonged flow holding became evident in the
earliest days of this study. Figure 29 plots the losses in the third stream duct just aft of the fan exit plane
(labeled fan duct in figure 15) against percent military thrust. Remember that as power is reduced, airflow
is increased in the third stream and, therefore, the Mach number in this stream increases for a fixed duct
size. As total pressure loss in a duct is modeled as a function of Mach number squared,
Where: p is the total pressure in the duct
even a modest change in duct Mach number can have a significant impact on total pressure. Figure 29
shows that as fan duct Mach increases from 0.25 to 0.56, the pressure loss increases fivefold and reaches
nearly 16%.
While this duct pressure drop is a very real effect, it can be mitigated in two ways. First, the third
stream duct can be slightly oversized at the design point. In other words the Mach number in this duct
could be chosen to be 0.15 rather than the 0.25 as depicted in this illustration. By doing so, this duct is
sized to the desired airflow and Mach number at the cruise points of interest where the bypass ratio is much
0.0
0.5
1.0
1.5
2.0
2.5
3.0
3.5
4.0
4.5
60 65 70 75 80 85 90 95 100
Byp
ass
Rat
io
% Military Thrust
Unrestricted Flow Holding, Bypass Ratio vs. Installed Thrust
Overall BPR
BPR1
BPR2
40
greater. Second, flow holding can be terminated when the propulsive benefit of increased bypass ratio is
surpassed by the duct pressure drop and other losses. This second option will be developed further below.
Figure 29. Fan duct Mach number & pressure drop with unrestricted flow holding
The next problem noted with indefinite flow holding was that the excessive variations in inlet
areas can have undesirable effects on component performance. Figure 30 provides an example of this
degradation in performance on the low pressure compressor. Notice that as the power is reduced the inlet
area is also reduced (depicted in the figure by an increase in IGV setting). While this does encourage flow
to the third stream, inlet area variations continuously reduce the pressure ratio and ultimately the
component efficiency.
There is a clear physical limit on LPC inlet area defined by the minimum allowable component
pressure ratio of 1.0; this limit occurs at roughly 60% power in this example. However, a more stringent
LPC minimum pressure ratio of 1.2 is enforced in this study to minimize the potential for flutter and any
associated component damage; this limit occurs at approximately 70% power in this example. A closer
examination of figure 30 reveals that there may be a point prior to either of these two pressure ratio limits
where reductions in inlet area, and hence flow holding, should be ceased. Notice that below 90% thrust the
LPC efficiency begins to drop at an ever accelerating rate. In fact by 70% power, or 1.2 LPC pressure
0%
2%
4%
6%
8%
10%
12%
14%
16%
18%
0.0
0.1
0.2
0.3
0.4
0.5
0.6
60 65 70 75 80 85 90 95 100
Th
ird
Str
ea
m D
uct
Pre
ssu
re L
oss
(%
)
Th
ird
Str
ea
m D
uct
Ma
ch N
um
be
r
% Military Thrust
Unrestricted Flow Holding, Duct Losses vs. Installed Thrust
Fan Duct Mn
Fan Duct Loss
41
ratio, the adiabatic efficiency is just 54%. As this drop in efficiency is also noted in other variable
components, one can easily understand how reductions in thermal efficiency will eventually offset the
benefits of improved propulsive efficiency associated with flow holding.
Figure 30. LPC efficiency and pressure ratio changes with unrestricted flow holding
The final concern with indefinite flow holding is excessive variations in nozzle throat area.
Figures 27 showed that as power is reduced, flow is moved from the core and second streams to the third
stream. In order to maintain the operating lines of the fan and LPC, the fan and primary nozzle throat are
changed accordingly (see Figure 31). While the variations depicted here are not beyond the capabilities of
current technology, smaller variations or even fixed nozzles are desirable from a cost and survivability
standpoint. The strategy for flow hold termination outlined below ensures that nozzle areas vary by less
than 100% from their design point area. A brief analysis of mission performance with fixed nozzles will be
presented in section 3.9.
Given the duct and component losses described above, one can easily envision a point in the
power hook where the relative propulsive efficiency improvements and spillage drag reductions are offset
by the increased duct pressure loss and thermodynamic efficiency reductions. If flow holding were
continued below this thrust setting, the fuel efficiency would actually be poorer than if a conventional, non-
0.00
0.20
0.40
0.60
0.80
1.00
1.20
1.40
1.60
1.80
0.0
10.0
20.0
30.0
40.0
50.0
60.0
70.0
80.0
90.0
100.0
60 65 70 75 80 85 90 95 100
Ad
iab
ati
c E
ffic
ien
cy o
r P
ress
ure
Ra
tio
LPC
IGV
Se
ttin
g
% Military Thrust
Unrestricted Flow Holding, LPC Schedule vs. Installed Thrust
LPC IGV Setting
LPC Efficiency
LPC PR
42
Figure 31. Nozzle throat area variations with unrestricted flow holding
flow holding, power hook was performed. Figures 32 and 33 show the performance of this notional
mobility engine at two different cruise points of interests. Notice that the high altitude flow holding power
hook shows an increase in installed TSFC at approximately 87% power (79% power at low altitude); and
Figure 32. High cruise power hook with and without unrestricted flow holding
300
400
500
600
700
800
900
1000
60 65 70 75 80 85 90 95 100
No
zzle
Th
roat
Are
a
% Military Thrust
Unrestricted Flow Holding, Nozzle Throat Area vs. Installed Thrust
Primary Nozzle
Fan Nozzle
43
below this value, a conventional power hook yields superior performance. It should also be noted that the
variable cycle power hooks presented do not show any improved performance at military power; this is
because the variable geometry in these power hooks has only been fully optimized at the cruise points of
interest. Nevertheless, these figures should prove sufficient to reveal the motivation for termination of
flow holding prior to the point of interest. Power hooks with optimized vane and nozzle settings at each
part power point are presented in sections 3.5 thru 3.7.
Figure 33. Low cruise power hook with and without unrestricted flow holding
Unfortunately, formulating a general rule for the location of this minimum in the flow holding
power hook is difficult as it is a strong function of the engine design and the current operating conditions.
For this reason, termination of flow holding is determined real time at either the minimum in the TSFC
curve or the minimum allowable LPC pressure ratio, whichever comes first; a conventional power hook is
then executed to the desired point of interest. It should be mentioned that the mobility mission presented
here has minimal spillage drag and therefore, tends to cease flow holding relatively early. It will later
become apparent that the supersonic strike mission, with a great potential for spillage drag at high speed
cruise, justifies a more prolonged period of flow holding.
44
3.2 Changes in component efficiencies with variable architecture
As noted in the previous section, component efficiency is a strong function of inlet guide vane
setting and component corrected speed, Nc (where ). Therefore, one would expect that
rotating component efficiencies would vary from design point to each cruise point of interest. Figure 34
illustrates this effect for a notional compressor at a given inlet guide vane setting. Notice that as the engine
is throttled from the design point A to the cruise point B, compressor efficiency increases by three percent.
This is a reasonable and expected increase in efficiency.
Figure 34. Expected change in component efficiency from design point to cruise
Early in this study it became evident that the on design optimizer was manipulating design point
location on component maps in an effort to maximize efficiency at the off design point. Figure 35 shows
how this might look on the same notional compressor map in Figure 34. Notice that as the engine is
throttled form design point C to cruise point B, component efficiency now increases by an unrealistic five
percent. This rather simplistic illustration only begins to describe a much larger problem that is further
exacerbated by scale factors and layered component maps.
To fully grasp this dilemma, one must understand that most engine models incorporate existing
component maps and scale them to the required design point mass flow, efficiency and pressure ratio. For
45
Figure 35. Unrealistic change in component efficiency from design point to cruise
example, the efficiency of design point A in figure 34 is 80% on the map; by multiplying this efficiency
with a scale factor of 1.075 a desired design efficiency of 86% is obtained. As this scale factor is also used
at all off design points, the efficiency of point B is scaled to 0.89; again, this 3% increase in efficiency from
design point to off design point of interest is reasonable. Looking at design point C in figure 35, the 78%
map efficiency would need to be scaled by 1.103 to achieve the desired efficiency of 86%. Using this scale
factor the off design point B efficiency jumps to nearly 92%; this 6% increase in efficiency is by no means
a reasonable excursion.
When one also considers that the compressor maps used in this study have multiple layers to
describe performance at different inlet vane settings, the problem becomes immediately obvious. The on
design genetic algorithm quickly determines that optimal off design performance can be achieved by
maximizing the efficiency scale factor at the design point, i.e. minimizing design point map efficiency,
thereby maximizing off design point efficiency as well. Therefore, design point vane angles, corrected
speeds, and operating lines are selected at locations that are completely unreasonable in order to maximize
these scale factors.
46
Two solutions to this problem are readily apparent. The first is to simply fix the design point of
each component at the intersection of the maps operating line, 100% corrected speed line, and on the inlet
guide vane fully open layer (as in Figure 34). While this does yield more reasonable efficiencies, it limits
the variable cycle‟s ability to vary flow while keeping surge margins, corrected speeds, and pressure ratios
within specified limits at each off design point. The second is to have the optimizer itself limit the scale
factors by placing an upper limit on off design efficiency. In this study, the cost function adds a penalty for
deviations in off design efficiency greater than 2.5% from the on design value; this is one of the penalties
for undesirable behavior described in figure 23. This penalty function proved effective in keeping
efficiencies reasonable while allowing internal air flow variations at all off design points of interest.
3.3 Reduction in spillage drag
As mentioned in the introduction, the prospect of matching an engine‟s demand for airflow to the
inlet‟s ability to deliver airflow is one of the classic motivations for creating a variable cycle engine. If one
is able to accomplish this inlet matching across a wide range of power settings and flight conditions,
spillage drag can be essentially eliminated and a corresponding reduction in fuel use realized. The results
presented in this section detail the spillage drag realized using the calculation methodology outlined in
section 2.4 for a variety of flight conditions and across all three candidate missions. Note that this data
does not represent the minimum possible spillage but rather that obtained for an optimized power hook in
which flow holding is ceased consistent with the logic presented in section 3.1 above.
If one assumes level steady state flight, i.e. cruise thrust is equal to aircraft drag, it is possible to
plot the percent increase in aircraft drag due to spillage throughout a power hook, see figures 36-38. These
results confirm that the three stream variable cycle is remarkably adept at reducing spillage drag at all flight
conditions and across the entire power hook. Furthermore, it is evident that there is a location in each
power hook where the benefits of reduced spillage are outweighed by increased duct losses, decreased
component efficiencies, or physical limits on variable features; at this point inlet matching is ceased.
Finally, one notes that spillage drag continually increases as engine thrust is reduced from military power,
and that the rate of increase is a strong function of dynamic pressure. For these reasons the optimizer notes
a greater incentive to hold engine airflow at higher airspeeds or when engine power is greatly reduced; this
47
will become increasingly evident in the subsonic long range strike and supersonic strike missions. These
issues will be addressed here while conclusions as to the overall cycle benefit will be reserved for later
sections.
Figure 36. Variable cycle reduction in spillage drag at tactical mobility mission cruise points
Figure 36 overlays the spillage drag curves for the advanced turbofan with those of the variable
cycle for the tactical mobility mission at the high and low altitude cruise flight conditions. In this particular
mission, the takeoff from a 1500 ft runway at 4000 ft pressure altitude on a 95o F day requires 24,000 lbf of
thrust per engine; this design point sizes these engines. Therefore, the engine is pulled back significantly at
both of the cruise points of interest. The 35,000 ft cruise location requires 63% military thrust and,
therefore, a 1.9% increase in aircraft drag is due to spillage is realized by the advanced turbofan. By
holding airflow to just 84% power, the VCE is able to eliminate the spillage drag at this cruise condition.
However, the 0.4 Mach low cruise requires just 32% military thrust and the variable cycle would incur too
great a thermodynamic efficiency loss by air flow holding to this point. Nevertheless, by flow holding to
82% the variable cycle is able to reduce the increased drag due to spillage from 5.6% to 3.2%. Thus by
holding airflow to approximately 80% power, the three stream variable cycle is able to reduce aircraft drag
by roughly 2% at each cruise point of interest.
0
2
4
6
8
10
12
14
16
18
20
0 10 20 30 40 50 60 70 80 90 100
% I
ncre
ase
in
Air
cra
ft D
rag
du
e t
o S
pil
lag
e
% Military Thrust
Tactical Mobility Spillage Drag
Advanced Turbofan, 4000 ft 0.4 Mn
Variable Cycle, 4000 Ft 0.4 Mn
Advanced Turbofan, 35000 ft 0.8 Mn
Variable Cycle, 35000 ft 0.8 Mn
48
Figure 37. Variable cycle reduction in spillage drag at subsonic LRS mission cruise points
Figure 37 shows an overlay of the advanced turbofan and variable cycle power hooks at the two
subsonic long range strike cruise flight conditions. Again these engines are sized for the takeoff condition
from an 8,000 ft runway at sea level 95oF day; this requires 30,000 lbf thrust per engine. For this design
point, both engines are operating at nearly 86% military thrust at high altitude cruise and neither produces
any increase in drag due to spillage. However, low altitude penetration requires only 30% power and the
advanced turbofan realizes a 15.3% increase in aircraft drag due to spillage. As stated earlier, this is in
large part due to the relatively high cruise speed, and hence increased dynamic pressure, of this penetration.
By flow holding to 78% power, the variable cycle is able to reduce the increase in aircraft drag to only
6.0%. This reduction of over 9% in drag suggests that a variable cycle would significantly reduce the fuel
required to accomplish this mission.
The supersonic strike mission spillage drag profile is illustrated in Figure 38 for the two cruise
points of interest. Here the engines are sized to produce 17,200 lbf thrust at Mach 2.5, 55,000 ft on a
standard day. Although this design point results in only a modest reduction in throttle setting at the two
cruise points of interest, the high speed cruise segment notes a very rapid increase in spillage drag as the
thrust is reduced. For this reason, the advanced turbofan suffers from a 16.2% increase in spillage drag at
0
5
10
15
20
25
30
0 10 20 30 40 50 60 70 80 90 100
% I
ncr
ea
se i
n A
ircr
aft
Dra
g d
ue
to
Sp
illa
ge
% Military Thrust
Subsonic Long Range Strike Spillage Drag
Advanced Turbofan, 500 ft 0.7 Mn
Variable Cycle, 500 Ft 0.7 Mn
Advanced Turbofan, 40000 ft 0.8 Mn
Variable Cycle, 40000 ft 0.8 Mn
49
the 72% military thrust high cruise flight condition. By flow holding to 76% the variable cycle is able to
reduce this to a mere 0.7% increase. Although spillage drag is nearly nonexistent at the 80% power low
altitude loiter condition, the variable cycle is still able to reduce the increase in spillage drag from 0.9% to
zero. As it was the supersonic commercial transport mission that first prompted the exploration of three
stream variable cycles, it should come as no surprise that the supersonic cruise segment of this mission
offers the greatest potential reduction in spillage drag. This substantial decrease in drag is certain to create
a corresponding reduction in mission fuel use.
Figure 38. Variable cycle reduction in spillage drag at supersonic strike mission cruise points
3.4 Reduction in aft body drag
It was suggested early in this study that an engine with a higher mass flow rate would necessarily
have a larger nozzle exit area for a given operating condition. Therefore, it follows that a variable cycle
engine which is flow holding would be able to fill the aft body area with exhaust better than its
conventional engine counterpart. Although aft body drag is typically much smaller in magnitude than
spillage drag, the potential drag reduction was deemed sufficient enough to warrant investigation. The
results presented in this section detail the aft body drag realized using the calculation methodology outlined
0
5
10
15
20
25
30
35
40
45
50
0 10 20 30 40 50 60 70 80 90 100
% I
ncr
ea
se i
n A
ircr
aft
Dra
g d
ue
to
Sp
illa
ge
% Military Thrust
Supersonic Strike Spillage Drag
Advanced Turbofan, 30000 ft 0.5 Mn
Variable Cycle, 30000 Ft 0.5 Mn
Advanced Turbofan, 50000 ft 2.2 Mn
Variable Cycle, 50000 ft 2.2 Mn
50
in section 2.5 for a variety of flight conditions and across all three candidate missions. Again, this data
does not necessarily represent the minimum possible aft body drag but rather that obtained for an optimized
power hook in which flow holding is ceased consistent with the logic presented in section 3.1 above.
Percent increase in aircraft drag due to aft body effects is plotted as a function of throttle setting in
figure 39 for the tactical mobility mission cruise points. Two conclusions are immediately evident from
this plot. First, the very small coefficient of aft body drag at subsonic speeds (reference figure 22) results
in a very modest increase in drag as thrust, and hence exhaust area, is reduced in a conventional engine.
Second, the variable cycle is more effective in filling the aft body area than its conventional counterpart.
While the reduction in drag is quite small, roughly a half percent reduction in drag is possible at each of the
cruise points of interest.
Figure 39. Variable cycle reduction in aft body drag at tactical mobility mission cruise points
The long range strike mission‟s aft body drag curves, given in Figure 40, follow the same trends
noted in the tactical mobility mission. Again, the variable cycle engine has a relatively constant aft body
drag across the entire power hook. While the high altitude cruise curves are essentially unchanged, the
higher dynamic pressure of the low altitude penetration increases the magnitude of the aft body drag and
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
0 10 20 30 40 50 60 70 80 90 100
% I
ncre
ase
in
Air
cra
ft D
rag
du
e t
o A
ft B
od
y
% Military Thrust
Tactical Mobility Aft Body Drag
Advanced Turbofan, 4000 ft 0.4 Mn
Advanced Turbofan, 35000 ft 0.8 Mn
Variable Cycle, 4000 ft 0.4 Mn
Variable Cycle, 35000 ft 0.8 Mn
51
hence the potential savings. The variable cycle was effective in reducing the aft body drag by roughly one
quarter percent at high cruse and one percent during the low altitude penetration.
Figure 40. Variable cycle reduction in aft body drag at subsonic lrs mission cruise points
The supersonic strike aft body drag profiles are different from the preceding two in both
magnitude and in nature, see figure 41. While the Mach 0.5 low altitude loiter aft body drag is essentially
constant across the entire power hook, the Mach 2.2 high altitude aft body drag rises abruptly for both the
advanced turbofan and the variable cycle. This is a direct result of the steep slope in the Cd curve at this
higher Mach number (see figure 22); therefore, even a modest reduction in exhaust gas area yields a
significant rise in aft body drag. Nonetheless, the variable cycle is able to achieve a modest 1.5% reduction
in aft body drag at high speed cruise while no significant change in aft body drag is realized at the slow
speed loiter condition.
3.5 Lift augmentation
An often touted benefit of variable cycles is that it can provide a readily available source of
constant pressure ratio air for lift augmentation. A skeptic might immediately respond that a conventional
0.0
0.5
1.0
1.5
2.0
2.5
3.0
3.5
4.0
0 10 20 30 40 50 60 70 80 90 100
% I
ncre
ase
in
Air
cra
ft D
rag
du
e t
o A
ft B
od
y
% Military Thrust
Subsonic Long Range Strike Aft Body Drag
Advanced Turbofan, 500 ft 0.7 Mn
Advanced Turbofan,40000 ft 0.8 Mn
Variable Cycle, 500 ft 0.7 Mn
Variable Cycle, 40000 ft 0.8 Mn
52
Figure 41. Variable cycle reduction in aft body drag at supersonic strike mission cruise points
engine could also provide a readily available source of pressurized air; however, this air would be much
more costly from an overall cycle efficiency standpoint. For example, a three stream variable cycle like
the one in this study could be configured so that the third stream air is pressurized by a single stage of
compression. By properly scheduling the variable features this source of pressurized air could be
maintained at a pressure ratio of 1.89, to permit choked flow through the discharge orifice, throughout the
period of lift augmentation. In contrast, a conventional engine would note a significant decrease in fan
pressure ratio as the throttle is reduced. This would require a conventional engine to use at least two stages
of compression when pressurizing air for lift augmentation. As such pressurization requires significant
energy extraction by the turbine and no appreciable thrust is produced by the lift augmentation system, the
fuel efficiency of the conventional engine would be appreciably worse than that of the variable cycle.
Each of these concepts is clearly visible in the following graphs. In figure 42 the pressure ratio of
the advanced turbofan second stream and the variable cycle third steam air is plotted as a function of
percent military thrust. As expected, fan pressure drops rapidly as the throttle is reduced in both the
conventional two stream engine and in the variable cycle optimized for fuel efficiency. Nonetheless, the
advanced turbofan with its two stages of fan compression is able to maintain the desired 1.9 pressure ratio
0.0
2.0
4.0
6.0
8.0
10.0
12.0
14.0
0 10 20 30 40 50 60 70 80 90 100
% I
ncre
ase
in
Air
cra
ft D
rag
du
e t
o A
ft B
od
y
% Military Thrust
Supersonic Strike Aft Body Drag
Advanced Turbofan, 30000 ft 0.5 Mn
Advanced Turbofan,50000 ft 2.2 Mn
Variable Cycle, 30000 ft 0.5 Mn
Variable Cycle, 50000 ft 2.2 Mn
53
thru 40% military thrust. What is more impressive is that the variable cycle is able to deliver the same
desired pressure ratio through 60% military thrust, as requested in the vision mission, with a single stage of
compression. This is accomplished primarily by decreasing the fan nozzle throat area as power is reduced
below 80% thereby back pressuring the fan. Unfortunately, this reduction in fan nozzle throat area
discourages flow to the third stream and is counterproductive from a propulsive efficiency standpoint
(optimal variable feature settings will be presented in sections 3.6 thru 3.8).
Figure 42. Fan pressure ratio during approach and landing
The question remains as to whether the performance of the less than fuel optimal operation of the
variable cycle exceeds that of the advanced turbofan. The throttle hooks in Figure 43 provide a clear
answer to this query. This figure reveals many of the variable cycle attributes outlined in this and previous
sections. First, reductions in spillage and aft body drag as well as improvements in propulsive efficiency
make the variable cycle more fuel efficient across the entire power hook. Second, use of single stage
compression air for lift augmentation further improves the efficiency of the variable cycle over the
advanced turbofan. Finally when the additional constraint of maintaining 1.89 pressure ratio air in the third
stream is added to this cycle, the two benefits above are reduced and the variable cycle efficiency begins to
approach that of the conventional two stream cycle.
54
Figure 43. Tactical mobility power hook during approach and landing
As stated earlier, the means of maintaining third stream pressure ratio was primarily to restrict the
fan nozzle throat area below 80% military thrust. As this does discourage mass flow into the third stream,
it is necessary to confirm that sufficient airflow remains to provide for both the lift augmentation and aft
deck cooling through 60% thrust. Figure 44 overlays the third stream airflow during the short field landing,
in which the 1.89 or greater fan pressure ratio is maintained, with the airflow demand. As expected, the
third stream flow decreases below 80% as the fan nozzle closes; however, the third stream airflow exceeds
demand throughout the approach and landing by 20% or more. The claim that a variable cycle can provide
a readily available source of near constant pressure ratio air for lift augmentation is easily justified.
3.6 Heat sink capacity of third stream
Modern aircraft systems generate profuse amounts of heat as a byproduct of their increasingly
complex onboard systems. Sources include advanced avionics, electric actuators, sensor suites, directed
energy weapons and electronic countermeasures just to name a few. Traditional thermal management
approaches shed this heat to the environment or use it to preheat the fuel. This approach is often frustrated
by supersonic flight in which ram air loses much of its cooling capability (Edwards, 2003). Furthermore,
55
the quest for increased fuel efficiency reduces not only fuel flow but also the heat sink capacity of the entire
fuel system. For these reasons additional heat sink capacity must be sought.
Figure 44. Variable cycle third stream air flow during assault landing
The variable cycle‟s third stream would seem to be an obvious location to exhaust this aircraft
thermal load. There are however a number of concerns with placing a heat exchanger in this bypass stream
including increased engine weight, pressure loss through the heat exchanger, duct sizing to accommodate
the heat exchanger, and additional hardware to carry the heat load from each source to the exchanger. As
each of these is beyond the scope of this research, the analysis here will simply concentrate on the capacity
of this stream to accept this heat load. Heat flux can be readily calculated with the equation,
Where: is the specific heat at constant pressure
is the mass flow rate of the working fluid
is the temperature increase of the working fluid
is the heat flux
56
Therefore, finding the theoretically available heat capacity of the third stream requires that one know the
mass flow rate of air, a single gas property and the maximum permissible temperature rise of the air.
As the mass flow and specific heat at constant pressure in the third stream can be readily found at
any given throttle setting and flight condition, the only remaining question is how much increase in air
temperature is allowable. To find the maximum heat capacity of this stream one should increase the air
temperature until a predetermined material limit is reached. The first such limit noted is the maximum air
temperature for effective nozzle cooling. Remember that bypass air is used to cool the nozzle aft deck in
both baselines and in the variable cycle (see figures 14 and 17). It is assumed that 15% of the total engine
airflow is sufficient to cool the aft deck at any flight condition, and hence at any given fan exit temperature.
Therefore, the upper limit for the third stream cooling air would be the air temperature of the advanced
turbofan second stream air at maximum fan pressure ratio and maximum airspeed. For the subsonic long
range strike mission this occurs at military power, Mach 0.7 low altitude penetration where the fan exit
temperature reaches 375 oF. Using this maximum temperature, the third stream heat sink capacity per
engine is plotted in figure 45 for the subsonic LRS mission at the 40,000 ft Mach 0.8 cruise condition.
Figure 45. Theoretical heat sink capacity of third stream, subsonic LRS high altitude cruise
57
In figure 45 it is clear that reducing power increases third stream heat sink capacity; this is a direct
result of increased third stream airflow and reduced fan pressure ratio. Note also that at this cruise
condition the compressor exit temperature is sufficiently low and cooling of the cooling air for the
compressor disk and turbines is not required; therefore, all of this heat sink capacity is available to dissipate
aircraft heat loads. Therefore, if only 13% of the heat flux capacity were utilized, a two engine aircraft
could effectively dissipate one megawatt of aircraft heat load at the 86% cruise power setting. This is
particularly appealing for thermal management as this variable cycle is most capable of accepting heat
loads at part power conditions where fuel flow is reduced and it is impractical to transfer significant heat to
the fuel system. This inherent capability to dissipate aircraft thermal loads makes a three stream variable
cycle very attractive for military applications.
The following sections will address how the effects of reduced drag and increased propulsive
efficiency combine to create an overall reduction in required mission fuel. During these discussions, it is
important to note which components vary, in what manner they vary, and to what extent flow is affected.
Such observations will provide the basis for subsequent results that attempt to maximize mission
performance with the minimum number of variable features.
3.7 Optimal tactical mobility mission variable cycle engine
The next three sections detail the most fuel efficient solutions to the three vision missions
determined by the nested genetic algorithm optimization routine. Note that the optimizing routine was
tasked to find a design point and two corresponding off design variable feature settings that minimized
overall mission fuel use without violating a myriad of constraints including maximum shaft speeds and
vane displacements as well as minimum component pressure and bypass ratios. This information returned
by the algorithm is far from an end in itself; however, it does provide a great deal of insight into how to
best design a variable cycle and which technologies are likely to be the most promising. Each of these
sections concludes with a much more detailed analysis of each off design cruise location and offers an
optimal schedule of variable features for efficient, stall free operation throughout a power hook.
To fully understand how a design point might be optimized, one must first appreciate what can be
controlled in these variable architectures and how each control affects engine operation. To alter internal
58
air flows the variable cycle engine has a number of tools at its disposal including the inlet areas to rotating
engine components, the variable area bypass injector primary and secondary inlet areas, and the two nozzle
throat areas. Theoretically, each of these can be continuously manipulated up to their mechanical limits.
However, it does not make sense to let an optimizer randomly pick each of these settings at either the
design or of design points; to do so would virtually always result in non viable cycles.
Therefore, one must begin by defining the variable features over which the optimizer would have
direct control, the permissible range of motion of each, and the relationships that define behavior of those
variable features that are not directly manipulated. In design point optimization the optimizer is given
direct control over thirteen settings; these are the inlet areas and corrected speed of each rotating
component, the compressor operating lines, and the bypass ratio to the third stream. NPSS then determines
the position of other variable features to satisfy continuity and user defined constraints. These include
calculating the second stream bypass ratio necessary to obtain the desired total pressure ratio at the mixing
plane, the VABI position necessary to achieve a mixing plane static pressure balance, and the primary and
fan nozzle throat areas necessary to pass the desired airflow. Note that the first order efficiency effects
such as engine air flow, overall pressure ratio, and turbine inlet temperature were held constant in all
engines. Again, NPSS maintained these constants by selecting the LPC pressure ratio to achieve desired
thrust, HPC pressure ratio to reach desired overall pressure ratio, and fuel flow to attain the desired turbine
inlet temperature.
The optimal rotating component design point settings for the tactical mobility mission are
displayed in table 3 along with the fixed baseline settings for comparison. As stated above, the optimizer
Table 3. Optimal tactical mobility variable geometry at 4000ft, Mach 0.0, 95o F design point
baselines VCE baselines VCE baselines VCE
Fan 0 2 1.00 0.94 100 96
LPC -- 5 -- 1.20 -- 104
HPC -- -- 1.13 1.18 100 95
HPT 120 110 -- -- 100 101
LPT 120 114 -- -- 100 101
op linevane setting * N c (%)
* HPC has an embedded vane schedule and cannot be varied by optimizer.
Fan and LPC vanes vary from 0 (maximum flow) to 100 (minimum flow)
Turbine nozzles vary from 120 (maximum flow) to 90 (minimum flow)
59
was given latitude to search the rotating component design space; this provided some additional freedom to
vary inlet areas off design without violating physics or user defined constraints. Despite this freedom, most
of the design point settings were consistent with conventional engine design; that is to say that 100%
component corrected speed, on the op line, and fully open guide vanes was roughly adhered to for most
components. The most notable deviations are noted in the turbine inlet areas and the LPC and HPC
operating lines.
Understanding these deviations provides insight into both the proper cycle design but also off
design operation. For example, the two turbine inlet areas are closed just slightly at the design point. By
doing so they are still capable of closing considerably off design, thereby moving flow to the second and
third airstreams, without incurring a large reduction in efficiency associated with full deflection of the inlet
nozzle. This also explains why the HPC operating line is dropped in the variable cycle; this lower
operating line increases the surge margin at design and ensures a minimum of 15% surge margin as the
HPT inlet area is varied off design. As the LPC surge margin is controlled with variations in the primary
nozzle throat area, minimum surge margin in this component is never an issue. Therefore, the lower
operating line was selected by the optimizer simply to maximize efficiency of this component off design.
Table 4 below displays the third stream bypass ratio selected by the optimizer as well as the
bypass ratios of the two baseline engines. As expected, the short takeoff requirement demands 24,000 lbf
of thrust per engine flat rated thrust to 4,000 ft 95oF. When this is coupled with a constraint on maximum
fan diameter of 60 inches, all three engines were driven to low bypass designs. One will also note that the
advanced turbofan and the three stream variable cycle have virtually the same on design overall bypass
ratio. The small difference is primarily caused by the disparity between the baseline nozzle film cooling
discharge coefficient and the third stream nozzle discharge coefficient; this will be discussed in greater
detail in sections 3.10-3.12.
Table 4. Optimal tactical mobility BPR & nozzle settings at 4000ft, Mach 0.0, 95o F design point
1 2 overall Fan Primary
Year 2000 state of art 1.44 -- 1.44 -- 820
Advanced turbofan 2.24 -- 2.24 -- 777
3 stream variable cycle 0.66 0.91 2.34 383 525
bypass ratio throat area (in 2 )
60
This engine design was then flown to both the high altitude cruise and low altitude penetration
points of interests. At these points two genetic algorithms independently optimized this engine for
minimum mission fuel and a common cost function evaluated overall performance. The resulting off
design parameters are summarized below in Tables 5 and 6. At these two discrete points a basic off design
control and the resulting improved performance of the variable cycle engine is evident. In each of these
two cruise points of interest the LPC inlet guide vane and the two turbine inlet nozzles are closed. Excess
airflow is moved to the third stream and the fan nozzle throat area is increased to accommodate this flow.
These changes are accompanied by very slight changes in the fan inlet guide vane area in order to optimize
efficiency at each cruise location.
Table 5. Optimal tactical mobility variable feature settings at 4000 ft, Mach 0.4 cruise
Table 6. Optimal tactical mobility variable feature settings at 35000 ft, Mach 0.8 cruise
These two tables clearly imply that the variable cycle indeed performed as intended. Notice that
the VCE specific fuel consumption is considerably lower than that of the advanced turbofan at both points
of interest; this implies an increase in bypass ratio typically associated with flow holding and therefore an
increase in propulsive efficiency. These suppositions are bolstered by the corresponding reductions in
spillage and aft body drag at both cruise locations which are also associated with flow holding. However,
this snapshot of data provides no information about how variable features vary from design point to the top
of climb or, from top of climb to cruise thrust. Such information is of great interest as the variable feature
settings necessary for stable engine operation during such transitions is a matter of some concern. Thus,
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 817 0.960 -- 4.6 0.7
Advanced turbofan 0 -- 120 120 -- 762 0.767 -- 5.7 0.8
3 stream variable cycle 11 53 91 104 727 389 0.643 80.6 1.4 0.4
Selected by Genetic Algorithm Computed
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 908 0.937 -- 0.5 1.0
Advanced turbofan 0 -- 120 120 -- 830 0.800 -- 2.0 1.3
3 stream variable cycle 4 48 92 92 744 439 0.724 83.6 0.0 0.6
Selected by Genetic Algorithm Computed
61
determining the most fuel efficient and stable variable geometry schedule throughout a power hook is
among this study‟s primary research focuses.
Given the optimization structure created in this study, the process of determining variable feature
settings that provide stable and fuel optimized operation through a range of power settings is a straight
forward, though somewhat cumbersome process. It is very similar to the off design optimization discussed
earlier in this thesis except that performance is fuel optimized at eleven discrete points, defined by10%
throttle increments, at each cruise point of interest. These continuously optimized power hooks provide a
reasonable resolution of geometry changes and offer sufficient insight as to how nearly optimal
performance might be achieved with fewer variable features.
Figures 46 and 47 illustrate both the variable feature settings, labeled adaptive feature schedule,
and the corresponding engine performance from military through 10% thrust at both tactical mobility cruise
points of interest. Before reaching any conclusions, it is first necessary to discuss the nature of the data
represented in these figures. First, all data was collected from independently optimized NPSS instances.
As the optimization routines were unaware of each other‟s conclusions, each search algorithm was free to
investigate the entire search space without bias introduced by adjacent power settings. Nonetheless the
adaptive feature schedule, and to a greater extent the engine performance, is reasonably smooth. Second,
the rate of adaptive feature movement is rapid during periods of flow holding, roughly 100-80% power in
these figure and, is typically followed by a period of slow or no movement. This observation is consistent
with the flow holding termination methodology introduced in section 3.1. That section details how flow
holding is terminated at the location where further manipulation of variable features to compel flow into the
third stream results in an increase in fuel consumption. Therefore, one would expect little internal
geometry changes immediately after the termination of flow holding.
There are several significant observations readily apparent from the adaptive feature schedules.
One immediately notes that there is little movement in the fan inlet area or in either turbine inlet area
throughout the entire power hook. While only small movements in the fan inlet guide vane are expected,
the lack of turbine nozzle movement is at first a bit troubling. A closer inspection reveals that a great deal
of turbine nozzle movement has occurred from the design point to the military power setting at both cruise
locations. This is a direct result of the substantial bleed airflow required for takeoff lift augmentation.
62
Adaptive feature schedule
Engine performance
Figure 46. Fuel optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise
63
Adaptive feature schedule
Engine performance
Figure 47. Fuel optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise
64
When the augmentation system is turned off, this bleed air is routed to the third stream by simultaneously
decreasing the turbine inlet areas and increasing the fan nozzle throat area. These changes and the
corresponding increase in third stream and overall bypass ratio are visible at military power in figures 46
and 47. One also notes that three components are consistently varied appreciably at both flight conditions;
these are the LPC inlet guide vane, the HPT inlet nozzle, and the fan nozzle throat area. This suggests that
these components have the greatest performance impact; this is investigated further in sections 3.9 - 3.12.
Engine performance in these figures is best understood by comparing variable cycle performance
to that of a conventional two stream mixed flow engine. First, the three stream variable cycle deviates from
its two stream fixed cycle counterpart by maintaining total engine airflow during the initial portion of the
throttle hook. During this portion of the throttle hook one notices a rise in the third stream mass flow and a
corresponding increase in both the third stream and overall bypass ratio. Then after flow holding is ceased,
continued geometry variations maintain the third stream airflow relatively constant; as a result, third stream
and overall bypass ratio continue to increase. The net result is a noticeable reduction in spillage drag and a
reduction in specific fuel consumption at all power settings.
Given the cruise point efficiencies above and a notional airframe, overall mission performance can
be determined for each study engine. Here a common airframe is used to assess all three engines so that
relative performance improvements can be determined, see table 7. Note that this makes the mission fuel
use and maximum range numbers more conservative than would be noted in a new aircraft program in
which reduced fuel weights would be accompanied by a corresponding decrease in structural weight and
decreased cruise thrust. Notice that the variable cycle engine is debited with a 10% increase in weight from
the comparable technology advanced turbofan for the additional bypass stream and variable components.
Table 7. Tactical mobility aircraft parameters
Maximum
gross
Aircraft
structure
Maximum
fuel load
Maximum
payload
Nominal
payload
Weights (lbf) 275,000 115,000 80,000 80,000 60,000
2000 SOA
turbofan
Advanced
turbofan
3 stream
VCE
Engine weight (lbf) 20,000 18,000 19,800
65
Mission analysis is summarized in table 8 below. This table illustrates that both the advanced
turbofan and variable cycle offer enhanced performance over the state of the art turbofan. Approximately
two thirds of the fuel savings is realized through thermodynamic improvements and is visible in both the
advanced turbofan and variable cycles. The remaining third is the result of reduced spillage drag and
improved propulsive efficiency; therefore, it is only visible in the variable cycle. Such fuel savings suggest
that a variable cycle might have a great deal to offer this mission particularly if most of this savings can be
realized with a minimum of variable features, this exercise is left for section 3.10.
Red text indicates negative numbers.
Table 8. Tactical mobility mission performance
The case for variable cycles becomes even more compelling when one converts the improved fuel
efficiencies into increased cargo capacity and an increase in overall mission range; figure 48 shows these
effects. While the baseline engine is incapable of performing the nominal mission without air refueling, the
variable cycle is capable of carrying 25% more cargo than required by the mission statement over the
nominal range. This could translate into an impressive 20% reduction in sorties for a given destination
payload requirement. The corresponding decrease in deployed aircraft, staged aircrews, support personnel,
maintenance requirements, and fuel costs would certainly be notable. Furthermore, the variable cycle
offers strategic ranges to this tactical aircraft. This means that this aircraft can carry a significant amount of
cargo into theater on its deployment leg thereby reducing the strategic airlift sorties required to transmit
cargo to the forward operating location. Furthermore, the variable cycle offers an increase in standoff
range thereby enabling basing well outside the combat zone; this increases not only the safety of this asset
and its support units, but also its long term sustainability. Each of these makes a compelling case for
variable cycles in a tactical mobility role; a role traditionally dominated by turboprops or fixed two stream
turbofans.
66
Figure 48. Tactical mobility range and payload for study engines
3.8 Optimal subsonic long range strike mission variable cycle engine
The variable cycle engine was also optimized for the subsonic long range strike mission. As this
mission is similar to the tactical mobility mission in its high cruise altitude and airspeed, one would expect
that the on and off design settings for the VCE are also alike. While this is a generally accurate assessment,
this mission does have a low level penetration that is twice as long and flown at roughly twice the airspeed
of its tactical mobility counterpart. This flight condition provides a greater incentive to minimize spillage
drag and, therefore, a corresponding increase in flow holding should be evident in the variable cycle low
altitude power hook.
The optimal rotating component design point settings for the tactical mobility mission are shown
in table 9 along with the fixed baseline settings for comparison. Again the optimizer has chosen settings
that are for the most part consistent with traditional design practices. Deviations from the fully open, on
the operating line and, 100% corrected speed are evident in only a few components. First, the HPC
operating line is dropped in order to maintain surge margins at off design conditions. Second, the LPC
operating line is dropped and the turbine inlets are closed slightly to allow off design variations in inlet area
without prohibitive reductions in component efficiencies.
67
Table 9. Optimal subsonic LRS variable geometry at 0 ft, Mach 0.0, 95o F design point
The engines in this mission are also sized by the takeoff requirement. Here, this two engine
aircraft is required to takeoff from a NATO standard runway on a 95oF day; this defines takeoff military
thrust as 30,000 lbf per engine. When this is coupled with a maximum fan diameter of 56 inches, these
engines are driven to very low on design bypass ratios, see table 10. Again, one notes that the advanced
turbofan and the three stream variable cycle have essentially the same overall bypass ratio.
Table 10. Optimal subsonic LRS BPR & nozzle settings at 0 ft, Mach 0.0, 95o F design point
Internal flow in this engine was again varied at each of the cruise points of interest until an overall
fuel optimal solution was determined; these optimal variable feature settings are given in tables 11 and 12.
The manner of flow manipulation is similar to that noted in the mobility mission. Inlet areas to both the
LPC and HPT are reduced at both points of interest which discourages flow into the engine core. Airflow
is then encouraged into the third stream by opening the fan nozzle throat area. Finally, the fan inlet guide
vane is fine tuned to improve its efficiency at both cruise points of interest. However, there is a notable
difference in the variable feature scheduling here; the LPT is moved very little from its design location at
either cruise location. Based solely on the magnitude of motion in these two missions, it would appear that
variations in the LPC, HPT and Fan A8 have the greatest impact on flow variations and therefore on
baselines VCE baselines VCE baselines VCE
Fan 0 1 1.00 1.01 100 99
LPC -- 5 -- 1.21 -- 99
HPC -- -- 1.07 1.17 100 96
HPT 120 112 -- -- 100 101
LPT 120 112 -- -- 100 100
vane setting * op line N c (%)
* HPC has an embedded vane schedule and cannot be varied by optimizer.
Fan and LPC vanes vary from 0 (maximum flow) to 100 (minimum flow)
Turbine nozzles vary from 120 (maximum flow) to 90 (minimum flow)
1 2 overall Fan Primary
Year 2000 state of art 1.15 -- 1.15 -- 697
Advanced turbofan 1.86 -- 1.86 -- 638
3 stream variable cycle 0.26 1.24 1.91 178 583
bypass ratio throat area (in 2 )
68
propulsive efficiency. While the data here is insufficient to corroborate this supposition, it does motivate
the more detailed analysis conducted in section 3.10.
Table 11. Optimal subsonic LRS variable feature settings at 500 ft, Mach 0.7 cruise
Table 12. Optimal subsonic LRS variable feature settings at 40000 ft, Mach 0.8 cruise
Minimizing spillage drag creates a great challenge during the low level portion of this mission.
The reason for this is twofold. First, the military thrust at 500 ft Mach 0.7 on a standard day is so great that
the penetration requires only 30% of the available power. Second the dense air and relatively high
penetration airspeed makes the dynamic pressure, and the corresponding spillage drag, much higher than in
the mobility mission. While flow holding to this power setting is possible with sufficient LPC PR at design
point, the cost in increased duct pressure loss and reduced component efficiencies makes this an
unacceptable solution. None the less, the variable cycle is able to reduce spillage drag nearly two thirds by
flow holding to 74% power. As the tactical mobility engine ceased flow holding at 81% thrust during low
level cruise, it appears that mission segments with increased potential for spillage justify an extended
duration period of flow holding. If this hypothesis is correct, a significant period of flow holding should be
noted in the supersonic cruise presented in section 3.9.
Once again the data in tables 11 and 12 suggest that the variable cycle is performing as expected.
Reduced spillage drag suggests that airflow is being maintained for a period of time as thrust is reduced.
Furthermore, reductions in specific fuel consumption over the thermodynamically identical advanced
turbofan imply that airflow is being encouraged into the bypass streams thereby increasing propulsive
efficiency. The optimal power hooks presented in figures 49 and 50 substantiate these suppositions. The
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 650 1.258 -- 13.0 2.7
Advanced turbofan 0 -- 120 120 -- 581 1.063 -- 15.2 3.0
3 stream variable cycle 13 47 90 110 422 458 0.886 74.4 6.0 1.8
Computed Selected by Genetic Algorithm
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 690 0.930 -- 0.0 0.9
Advanced turbofan 0 -- 120 120 -- 623 0.812 -- 0.0 1.0
3 stream variable cycle 10 37 92 104 238 532 0.749 85.7 0.0 0.7
Selected by Genetic Algorithm Computed
69
Adaptive feature schedule
Engine performance
Figure 49. Fuel optimal Subsonic LRS control and performance, 500 ft Mach 0.7 cruise
70
Adaptive feature schedule
Engine performance
Figure 50. Fuel optimal Subsonic LRS control and performance, 40000 ft Mach 0.8 cruise
71
selected adaptive feature schedule is consistent with the that observed in the tactical mobility mission.
While only modest movements are noted in the fan and LPC to fine tune efficiencies, other components
note rapid movements. For example, the HPT nozzle area moves to the fully closed position at military
power and remains there throughout the power hook. One also notes a rapid decrease in LPC inlet area
during the period of flow holding and a relatively constant setting thereafter.
These variable feature settings clearly produce the desired effects. Flow is encouraged into the
third stream as airflow through the core is diminished. When these vane and nozzle settings are coupled
with an ever increasing fan nozzle throat area, one notes a rise in third stream airflow during flow holding
and a reasonably constant airflow through the remainder of the power hook. The steady rise in bypass ratio
and the modest reduction in spillage drag yields reduced fuel consumption for the variable cycle at all
power settings and at each cruise location.
To fully understand the mission impact of this improved fuel efficiency, one must place each
engine in a notional airframe. Again a common aircraft is used, see table 13, and mission analysis is
performed for all three candidate engines. The results of this analysis, presented in table 14, though
impressive are not as dramatic as one might expect given the significant improvement in low level fuel
efficiency. This is the direct result of the mission description which logically defines the goal as increased
standoff range and not increased penetration range. Therefore the significant fuel savings realized during
the relatively short low level segment translates into a 16% increase in mission radius and a 15% reduction
in fuel use compared to the advanced turbofan engine.
Table 13. Subsonic long range strike aircraft parameters
As in the tactical mobility mission, this snapshot of mission performance does a poor job of fully
describing the overall system level benefits of a subsonic long range strike aircraft mated with this variable
cycle engine. To more adequately express what this variable technology has to offer, a plot of payload
Maximum
gross
Aircraft
structure
Maximum
fuel load
Maximum
payload
Nominal
payload
Weights (lbf) 225,000 82,500 112,500 30,000 20,000
2000 SOA
turbofan
Advanced
turbofan
3 stream
VCE
Engine weight (lbf) 10,000 9,000 9,900
72
verses range is offered in figure 51. Here one must remember that this aircraft is by definition a strategic
asset and therefore strategic ranges are paramount. The variable cycle offers an additional 1000 nm mile
range over the stated mission requirements; this can have a profound impact on mission generation and
sustainment especially in a prolonged conflict. For example, this range puts virtually half the world in
range using only internal fuel. Furthermore, assuming a reasonable fuel reserve for diversion or other
contingencies, a 6,000 nm radius mission could be accomplished with only two or three air refuelings. If
this represents the reduction of only a single refueling per sortie the savings in generated tanker missions,
aircrew and deployed support personnel would be substantial. This logistics tail associated with airborne
refueling is why fueling from a tanker to an airborne asset is often quoted as tenfold more expensive than
Red text indicates negative numbers.
Table 14. Subsonic long range strike mission performance
Figure 51. Subsonic long range strike range and payload for study engines
73
that loaded from a domestic airport. The analysis also shows that one third more payload is possible using
a variable cycle engine over the nominal mission range. Again the reduction in sorties required to deliver a
given payload would convey each the benefits outlined above and could also grant additional savings
offered by a reduction in aircraft procurement. These results suggest that the variable cycle architecture
presented in this study has a great deal to offer a subsonic long range strike platform.
3.9 Optimal supersonic strike mission variable cycle engine
The final mission analyzed is one that is typically considered a good fit for variable cycle engines.
In fact the supersonic transport programs of the past, with an extended duration supersonic cruise for
intercontinental flight and considerable subsonic cruise from coast-in to destination, were the original
motivation for variable geometry engines. For this reason, one would expect to see a significant reduction
in both spillage drag and fuel use at part power conditions. Before these can be adequately addressed, one
must understand the unique design point of this engine.
Unlike the previous two missions, this engine is sized by the thrust requirement at maximum
Mach number. As this aircraft penetrates enemy airspace at high altitude, it is considered desirable from an
operational perspective for it to have a dash capability of Mach 2.5. This high speed dash demands 15,300
lbf per engine at 55,000 ft on a standard day. This supersonic speed has a marked impact on the rotating
components at design, see tables 15 and 16. First one notices that the low spool speed is reduced greatly in
all engines while the core speed is maintained near 100%. This is to be expected as the air entering the fan
is much warmer in supersonic flight; hence the low spool speed ,and the overall pressure ratio, must be
Table 15. Optimal supersonic strike variable geometry at 55000 ft, Mach 2.5 design point
baselines VCE baselines VCE baselines VCE
Fan 0 0 0.90 1.10 75 71
LPC -- 20 -- 1.06 -- 75
HPC -- -- 1.08 1.13 96 90
HPT 120 119 -- -- 100 98
LPT 120 106 -- -- 88 76
Fan and LPC vanes vary from 0 (maximum flow) to 100 (minimum flow)
Turbine nozzles vary from 120 (maximum flow) to 90 (minimum flow)
vane setting * op line N c (%)
* HPC has an embedded vane schedule and cannot be varied by optimizer.
74
reduced to maintain the compressor exit temperature within material limits. Second, one notes that both
LPC and the LPT settings in the variable cycle are chosen at locations that allow for less movement off
design. If a corresponding decrease in movement is noted off design this would lend credibility to the
notion that the HPT and Fan nozzle A8 have the greatest impact on flow variation off design. Third, one
notes a dropping of the operating line in all three compressor sections; again this is used to maintain surge
margins within limits and to improve efficiencies off design. The high thrust requirement coupled with a
maximum fan diameter of 55 inches resulted in a relatively low bypass ratio at design point.
Table 16. Optimal supersonic strike BPR & nozzle settings at 55000 ft, Mach 2.5 design point
Off design optimization for this engine was conducted as before; however, there were some
unexpected results. A cursory glance at tables 17 and 18 reveals that the fan, LPC and LPT inlet areas
deviate little from their design point values to either of the cruise points of interest. Despite this, the
variable cycle is able to effectively eliminate the considerable spillage drag during high speed cruise
through judicious scheduling of the HPT inlet and fan nozzle throat areas.
Table 17. Optimal supersonic strike variable feature settings at 30000 ft, Mach 0.5 loiter
Table 18. Optimal supersonic strike variable feature settings at 50000 ft, Mach 2.2 cruise
1 2 overall Fan Primary
Year 2000 state of art 0.38 -- 0.38 -- 726
Advanced turbofan 1.05 -- 1.05 -- 612
3 stream variable cycle 0.34 0.63 1.26 238 626
bypass ratio throat area (in 2 )
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 745 1.010 -- 0.0 1.0
Advanced turbofan 0 -- 120 120 -- 611 0.875 -- 0.6 1.0
3 stream variable cycle 0 28 100 114 169 665 0.777 56.7 0.1 0.8
Selected by Genetic Algorithm Computed
Fan LPC HPT LPT Fan A8 Pri A8 TSFC Fl Hld % % Spill D % AB D
Year 2000 state of art 0 -- 120 120 -- 804 1.840 -- 17.0 9.3
Advanced turbofan 0 -- 120 120 -- 661 1.712 -- 19.2 8.3
3 stream variable cycle 6 25 102 114 252 707 1.355 72.2 0.7 6.9
Selected by Genetic Algorithm Computed
75
Fuel optimized throttle hooks at the high speed cruise and loiter points of interest are presented in
figures 52 and 53. As in the previous two missions, one notes that the HPT inlet nozzle closes as power is
reduced albeit a more gradual closure in this mission. This reduces flow to the core and the fan nozzle
throat area is increased to accept additional flow into the third stream. Furthermore, the fan inlet guide
vane makes only modest adjustments to improve fan efficiency off design as it did in the other missions
There are however a few notable differences in these figures. First, there is little movement in the LPT
inlet nozzle from design point to the points of interest and little movement throughout each throttle hook.
Again this gives credence to the theory that the Fan and LPT inlet areas have a smaller impact on fuel use
than some of the other variable features . Second, the LPC inlet makes a modest movement from the design
point to military power at each cruise location, but remains constant through the majority of the power
hook. For this reason a continual decrease in the primary nozzle area is not observed here as it was in the
previous missions.
The impact on fuel efficiency and spillage drag observed in figures 51 and 52 is similar in nature
to the other missions but different in magnitude. The reduction in spillage drag and specific fuel
consumption is really quite staggering during high speed cruise. This is accomplished by simply matching
engine airflow demand to the inlet airflow through 72% power and then allowing only minimal reductions
through 30% power. When this airflow holding is coupled with an increasing primary and fan nozzle area,
rising second and third stream bypass ratios are noted throughout the power hook.
Again these improvements in specific fuel consumption are better understood when placed in the
context of a notional airframe. Table 19 below shows the weights associated with the aircraft used in this
study. Like the subsonic long range strike airframe, that this aircraft requires a 50% fuel fraction to
effectively accomplish the loiter requirement.
Table 19. Supersonic strike aircraft parameters
Maximum
gross
Aircraft
structure
Maximum
fuel load
Maximum
payload
Nominal
payload
Weights (lbf) 250,000 90,000 125,000 30,000 20,000
2000 SOA
turbofan
Advanced
turbofan
3 stream
VCE
Engine weight (lbf) 15,000 13,500 14,850
76
Adaptive feature schedule
Engine performance
Figure 52. Fuel optimal supersonic strike control and performance, 30000 ft Mach 0.5 loiter
77
Adaptive feature schedule
Engine performance
Figure 53. Fuel optimal supersonic strike control and performance, 50000 ft Mach 2.2 cruise
78
Performance for the nominal mission profile is summarized in table 20 below. It is not surprising
that the year 2000 baseline again fails to achieve the stated mission objectives. However, one also notes
here that the persistent strike requirement with its eight hour loiter proves to be too much for even the
advanced turbofan. While there is some reduction in specific fuel consumption during loiter, this is
primarily the result of the significant spillage drag reduction and the corresponding decrease in fuel use
during the high speed penetration. Fuel savings on this leg is then translated into increased loiter duration.
Red text indicates negative numbers.
Table 20. Supersonic strike mission performance
As this mission uses both loiter time and standoff range as figures of merit, each is plotted against
payload in figures 54 and 55. Again some conclusions are immediately obvious. First, the variable cycle‟s
2.4 hour increase in loiter time over the year 2000 baseline is roughly equally split between thermodynamic
improvement, also noted in the advanced turbofan, and propulsive efficiency enhancements. To understand
just how this impacts cost on a system lever, one must consider the nature of this mission. This mission has
clearly been crafted to fulfill the requirement for a persistent airborne asset, holding in friendly territory,
and capable of responding to high value targets up to 500 nm distant in defended territory in less than half
an hour. Therefore, this increase in loiter time corresponds to one fewer generated mission every 24 hours,
or a 25% reduction in aircraft required. This savings is multiplied appreciably when one considers the
corresponding reduction in tanker sorties, fuel use, maintenance costs, and deployed aircraft, aircrew and
support personnel.
Second, one notes that the variable cycle increases standoff range by one third and, roughly 4/5 of
the increased range is due to improved propulsive efficiency and reduced spillage drag. This observation
requires some explanation as one would not expect such improvement at high speed cruise. Remember that
79
in this mission a Mach 2.5 dash capability was considered desirable. By designing the engine for this
increased airspeed, a reduced throttle setting is required for Mach 2.2 cruise; this creates the potential for
Figure 54. Supersonic strike loiter and payload for study engines
Figure 55. Supersonic standoff range and payload for study engines
80
both increased spillage drag or, with successful inlet airflow matching, increased overall bypass ratio.
Consequently, the variable cycle is able to deliver the high speed dash potential without sacrificing standoff
range. This improves the aircraft‟s ability to prosecute time sensitive targets from a distance, reduces its
exposure time in high threat environments, enables it to outrun many advanced threats, and grants it a
return to a distant safe haven.
3.10 Variable features with greatest impact on performance
The preceding sections discussed the most fuel efficient solutions returned by the optimization
routine for each vision mission. These solutions represent how one might design a variable cycle engine if
the number of variable features, along with the associated complexity and cost, were not an issue. But what
if a nearly optimal solution could be obtained with a much smaller subset of variable features? Identifying
the most promising technologies would focus component technology thereby reducing research and
development costs. But more importantly, it would likely reduce the unit cost of the operational engine and
sustainment costs over the engine‟s life span. Such a subset of vital technologies would be of great interest.
In the analysis above it is quite clear that the optimizer varies some components more than others.
Therefore, one might conclude that some components are more effective in varying internal flows and,
therefore, at improving fuel efficiency. In order to further investigate this supposition, maximum
component variations are presented in Table 21 for each of the three missions. On the far left of this table
map corrected weight flow, given in lbm/s, is shown at vane settings, rotational speeds and operating lines
consistent with those at points of interest. The right side of the table displays minimum and maximum vane
Red text indicates negative numbers.
Table 21. Maximum variation in component area during each vision mission
81
settings for the rotating components and the extreme fan and primary nozzle throat areas in inches. The
variations in component inlet area were then computed by assuming a constant flow per annulus area.
The decrease in rotating component inlet areas shown here represents the maximum swing from
on to off design. While these variations differ across the vision missions, a few conclusions are readily
apparent. First, the fan inlet area variation is minute in all three missions. Second, the LPC inlet area
variations are significant in the subsonic applications and virtually nonexistent in the supersonic mission.
Third, the HPT tends to show a greater variation than the LPT. Finally, the fan nozzle consistently varies a
greater amount than the primary nozzle. This data further corroborates the notion that some components
can be fixed with little impact on performance; however, it provides no insight into the ramifications of
fixing multiple components simultaneously.
As there are six variable features in this cycle, there are 62 subsets of variable features that could
be investigated. Performing a detailed analysis of each of these subsets across all three missions would be
computationally prohibitive using traditional methods. For this reason the objective function was modified
so that it would be able to determine the most promising subsets of variable features. This was
accomplished by introducing a step function penalty to the cost function for each variable feature that the
optimizer chose to vary. The magnitude of this step penalty was varied over several runs until a reasonable
number of variable feature subsets consistently surfaced.
The following sections will reveal the results of this sub optimal study and recommend a cycle
with reduced variablilities for each mission. There are a few things to note when reviewing this data. First,
it is important to note that the sub optimal studies conducted here do not reduce the engine weight as the
number of variable features is reduced. Therefore, engines with reduced variable features will require less
mission fuel than is reported here. Second, the fan nozzle throat area variation appears to be essential to
encouraging airflow into the third stream, and in no instance did the optimizer choose to fix this throat area.
Finally, the primary nozzle throat area can be fixed if one is willing to give up the ability to directly control
the LPC operating line. As a fixed primary nozzle is desirable from a reduced complexity perspective, this
will be investigated in all three missions. The most promising variable feature subsets can be seen in the
following sections.
82
3.11 Recommended variable features for tactical mobility mission
To fully appreciate these results presented in the next three sections, the analysis conducted with a
variable feature debited objective function will be discussed a bit further. In this analysis the optimizer
attempted to minimize the fuel use subject to all the constraints discussed to this point and a penalty for
each variable feature that it chose to manipulate. Therefore, if the variable feature penalty is large the
optimizer tends to converge on a fixed cycle. If this penalty is small, it converges on the optimal solutions
given above. By varying the penalty function, a consistent subset of variable features was noted across all
three vision missions. Note that the decrease in mission fuel use afforded by a variable fan nozzle was so
great that the fan nozzle was varied in every run except for the most ludicrous of variable feature penalties.
During this analysis it was impractical to communicate progress of each off design cruise point
optimization before it concluded. Such communication would greatly slow computations and would likely
be of little benefit prior to termination of the final generation. For this reason, the vastly different cruise
conditions often influenced the off design optimization routines to converge on different subsets of
promising variable features. The subsets of variable features that were investigated more rigorously were
those consistently returned by both of the off design optimization routines.
Figure 56. Effect of reduced variable features on tactical mobility mission fuel
83
A much more detailed analysis of each variable feature subset was conducted by manually fixing
the unused features and removing the variable feature penalty from the objective function. Therefore, the
optimization routines were free to move the remaining features up to their mechanical limits; the results of
this study are presented in figure 56. From this illustration it is evident that components such as the fan and
LPT can be fixed with very little impact on overall mission fuel use; in fact, fixing both of these increases
mission fuel by less than 1.5%. However, fixing the LPC comes at slightly higher price. What is most
striking is that fixing the HPT increases the variable cycle fuel use by roughly 5%; this is more than twice
the increase in fuel use caused by fixing the fan, LPC, and LPT combined. Two obvious questions arise
from this illustration. First, why was a fixed primary nozzle throat area not investigated? Second, why
doesn‟t the cycle with only variable nozzle throat areas converge to the advanced turbofan fuel use?
The first question is best answered with the engine control logic first introduced in section 2.3. It
was believed that the best method of maintaining the LPC pressure ratio and surge margin within defined
limits was to use to primary nozzle throat area to control, or more accurately fix, the surge margin after
flow holding was ceased. While this was effective in keeping the LPC within prescribed limits, it was
likely a bit too conservative. To allow for solutions with a fixed primary A8, the control was modified to
allow variations in LPC surge margin and pressure ratio subject to the minima described in section 3.1; the
Figure 57. Effect of fixed primary nozzle throat on tactical mobility mission fuel
84
results can be seen in figure 57.
Here a fixed primary nozzle was investigated at three different subsets of variable features. When
only the primary nozzle throat area was fixed, the optimizer was able to reduce overall mission fuel use by
an additional 1% (see table 22 for further details). A similar decrease in mission fuel use was noted in the
other two variable cycles with a fixed primary A8. The most promising of these solutions is a two variable
feature solution which utilizes only a variable HPT inlet and fan nozzle throat area. As both the advanced
turbofan and this variable cycle each has a single variable exhaust nozzle, this means that over 85% of the
potential fuel savings can be achieved by adding just one additional variable feature to the double bypass
architecture. While this is quite promising, the question still remains as to why the three stream cycle with
so few variablilities is so much more efficient than the advanced turbofan. This question can be answered
with the data in figure 58 below.
Figure 58. Sources of improved variable cycle efficiency in tactical mobility mission
In section 2.2 it was noted that the advanced turbofan and the three stream variable cycle were
created to be essentially the same engine thermodynamically. This was done in an effort to separate
thermodynamic efficiency improvements that would be visible in both the advanced turbofan and the
variable cycle from the propulsive efficiency benefits that would be only visible in the variable cycle.
85
There were two notable exceptions that need to be addressed here. First, turbine cooling flow is decreased
in the three stream architecture when power is reduced at cruise. This reduction in cooling flow is called
modulated cooling and is considered a variable technology in this study; for this reason it appears only on
the variable cycle. This graphic reveals that a 4.9% fuel savings is possible from this variable technology
alone. If modulated cooling yields similar savings in each of the other missions, investments in this
technology would certainly be warranted.
The second major difference between the advanced turbofan and the variable cycle is the manner
in which the aft deck cooling is introduced into the flow stream. In the advanced turbofan 15% of the
bypass air is used to cool the aft deck via film cooling. This film cooling suffers from a 0.92 coefficient of
gross thrust, CFG. As the entire third stream airflow is used to cool the aft deck in the variable cycle, film
cooling is not a viable option to reintroduce this large quantity of air into the exhaust stream. Therefore, it
was determined that this air would be introduced via a slotted exhaust as depicted in figure 17. Such a
system yields the clear benefit of a 0.96 CFG but it does so at the potential cost of increased radar cross
section. However, this increased radar observability is coupled with an inherent reduction in exhaust gas
temperature and hence reduced infra red heat signature. This trade was considered acceptable for the
purpose of this study. The improvement in CFG accounts for a 4.4% reduction in fuel use with minimal
impact on survivability.
The remaining 7.4% reduction in fuel use offered by the variable cycle is generated by a reduction
in spillage drag and improvements in propulsive efficiency. Nearly 5% of this improvement can be
realized with only variable HPT inlet and fan nozzle throat areas. This data indicates that the fuel
reductions are roughly equally split between improved fan nozzle CFG, modulated cooling, and variable
geometry effects in the tactical mobility mission. However, one expects that as the potential increases for
spillage drag in later missions a greater portion of the savings would be due to variable geometry.
As 86% of the potential variable cycle fuel reduction can be realized with only modulated cooling,
a variable HPT inlet and a variable fan nozzle throat, this is the sub optimal variable cycle proposed here.
Such a cycle offers a reduction in complexity, cost, risk, and maintenance over the fully variable cycle yet
captures the bulk of the benefits. The ability of this reduced variability cycle to effectively vary internal
flow is demonstrated in figures 59 and 60. The two adaptive feature schedule plots have been modified to
86
Adaptive feature schedule
Engine performance
Figure 59. Sub optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise
87
Adaptive feature schedule
Engine performance
Figure 60. Sub optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise
88
show the location of the two variable geometries at the takeoff condition design point. Notice that at both
cruise points of interest the HPT inlet area is closed significantly and the fan nozzle throat area is opened at
the military power condition. This condition was also noted in the fully variable engine and is a direct
response to turning off the lift augmentation system after takeoff and maintaining airflow during this
transition. At power settings below military power one notes a relatively constant airflow in the third
stream and a corresponding increase in bypass ratio. As in the fully variable cycle a reduction in spillage
drag and specific fuel consumption from both baselines is noted at all throttle settings.
Again this reduction in specific fuel consumption can be best appreciated in the context of an
airframe executing the vision mission. Table 22 summarizes the reduction in fuel use and potential
increase in range offered by the various variable cycle engines. Notice that although the fixed primary A8
solution has the same design point component settings as the optimal variable cycle, it offers slightly
improved mission performance due the control logic modifications presented above. Note also that the
reduced variability variable cycle still yields a 31% reduction in fuel use during the nominal mission or a
63% increase in mission radius over the year 2000 baseline. Remember that this performance is realized
without a reduction in engine weight corresponding to the reduction in variable features; therefore, these
figures are conservative.
Red text indicates negative numbers.
Table 22. Sub optimal tactical mobility mission performance
This improvement in performance over both baselines is accentuated in figure 61. Here the sub
optimal variable cycle yields only modest reductions in payload for a given mission range from the fully
variable cycle illustrated in figure 48. Therefore, the corresponding benefits noted in section 3.7 for the
fully variable architecture are observable here at a fraction of the complexity. It is important to note that a
89
very modest increase in engine complexity can yield considerable improvements in performance even in
missions traditionally considered a good fit for turbofan cycles.
Figure 61. Sub optimal tactical mobility range and payload
3.12 Recommended variable features for long range strike mission
Determining the sub optimal set of variable features for the long range strike mission progressed
in much the same manner as noted in the previous section. The one notable exception is that fixing the
primary nozzle throat area was investigated first. When it was determined that the modification to the
control noted above was beneficial in this mission as well, all sub optimal configurations were run with a
fixed primary nozzle throat area. As preliminary analysis conducted with a debit for variable feature use
suggested the same promising geometries, this study uses the same variable feature subsets identified in
section 3.11.
As table 21 noted little change in either the fan or LPT inlet areas, it was expected that only slight
deviations would be noted in overall fuel consumption when these geometries were fixed. Figure 62 not
only substantiates these predictions but further illustrates the relative insignificance of variations in LPC
90
Figure 62. Effect of reduced variable features on subsonic LRS mission fuel
inlet area. Notice that by fixing the LPC inlet mission fuel use increases by a mere 0.6%. While variations
in this area clearly have a larger influence on fuel use than the fan and LPT inlet areas combined, it can
hardly be considered an essential variable technology for this particular cycle architecture.
It is evident that variations in the HPT inlet area have the greatest impact on mission fuel use. In
fact this single variable component can yield a reduction in fuel use four times greater than that offered by
variations in the fan, LPC and LPT inlet areas combined. For this reason, the proposed sub optimal
variable cycle for the subsonic long range strike contains a variable HPT inlet area, variable fan nozzle
throat area and modulated cooling. This subset of variable features offers 95% percent of the fuel savings
noted in the fully variable cycle with only a fraction of the complexity.
Again one must inquire as to the source of fuel savings afforded by the three stream architecture
with only a variable fan nozzle. This engine depicted on the far right side of figure 62 offers a 12%
reduction in fuel use from the advanced turbofan, and it does this without the addition of variable geometry.
To address this issue the sources of increased efficiency were plotted in figure 63. Note that the sources of
improved efficiency are the same, but the impact of each has changed. For example, the subsonic LRS
variable cycle has a much lower design bypass ratio to the third stream than its tactical mobility
91
counterpart. For this reason a smaller percentage of the flow is in the bypass stream at any given point in
the mission profile. Therefore, by modulating cooling air in this cycle a larger percentage of the inlet flow
is impacted and fuel use is reduced by 5.6%. Similarly, the increased fan nozzle coefficient of gross thrust
effects a smaller portion of the air flow and creates only a 3.4% reduction in fuel use.
Figure 63. Sources of improved variable cycle efficiency in subsonic LRS mission
What is most interesting is the considerable impact of reduced spillage drag and increased
propulsive efficiency on fuel consumption in this mission. The introduction section suggested that variable
cycles would offer increasing benefits as engines were tasked to operate at reduced power settings and,
therefore, far from the inlet‟s maximum airflow rate. At such conditions variable cycle architectures could
match the engine airflow to the inlet‟s ability to deliver airflow thereby reducing spillage. Furthermore if
the additional airflow was encouraged into the bypass streams, an increase in propulsive efficiency would
also be realized. In this mission the variable cycle architecture begins to reveal its true strengths. Here,
fully half of the fuel savings is attributed to reductions in spillage and increases in propulsive efficiency.
The sub optimal power hooks illustrated in figures 64 and 65 verify that even with this reduced set
of variable features the three stream architecture is adept at constructively manipulating internal airflows.
Unlike the tactical mobility mission, the lift augmentation is not required at takeoff. For this reason a very
92
Adaptive feature schedule
Engine performance
Figure 64. Sub optimal subsonic LRS control and performance, 500 ft Mach 0.7 cruise
93
Adaptive feature schedule
Engine performance
Figure 65. Sub optimal subsonic LRS control and performance, 40000 ft Mach 0.8 cruise
94
gradual increase in fan nozzle throat area is noted throughout the power hook. When coupled with a
closure in the HPT inlet area, a continuous increase in airflow is observed in both bypass streams.
Furthermore, engine airflow is held to approximately 70% thrust at both cruise points of interest; this yields
a corresponding decrease in spillage drag. As always the improvements in specific fuel consumption are
best appreciated in the context of a specific airframe flying this vision mission. Fuel use for the nominal
mission and maximum high cruise radius is given in table 23.
Red text indicates negative numbers.
Table 23. Sub optimal subsonic long range strike mission performance
It is abundantly clear from this table and the depiction in figure 66 that the sub optimal variable
cycle suggested here performs virtually identically to the fully variable optimal architecture offered in
Figure 66. Sub optimal subsonic LRS range and payload
95
section 3.8. This three stream engine offers an impressive 32% reduction in fuel use over a year 2000 state
of the art engine for the nominal mission and nearly 15% savings over a two stream turbofan with
comparable thermodynamic cycle properties. Therefore all of the savings attributed to the fully variable
cycle in section 3.8 would also be realized in this sub optimal engine.
3.13 Recommended variable features for supersonic strike mission
It is in this mission that identifying a subset of variable features that provides nearly optimal
performance is most intriguing. This is because variable cycles in all their diverse and fanciful forms
originated in this classic mission with extended subsonic and supersonic segments. To identify the most
promising set of variable features for this mission would offer great insight into a truly perplexing problem.
Analysis began with the same potential subsets of variable features noted in the previous two missions.
Note that fixing the primary nozzle throat area and modifying the control logic accordingly did not yield a
reduction in mission fuel use; therefore, this component was fixed in selected variable feature subsets.
The data in table 21 shows that several variable components required very modest changes in inlet
area in the optimal solution. In fact the fan inlet, LPC inlet, LPT inlet, and primary nozzle throat areas all
Figure 67. Effect of reduced variable features on supersonic strike mission fuel
96
varied by 12% or less. This would suggest that once again the HPT inlet and fan nozzle throat areas would
have the greatest impact on overall engine performance. The data in figure 67 clearly substantiates this
claim. Notice that the fan, LPT and LPC inlet areas can be fixed with no discernable increase in fuel usage.
While fixing the primary nozzle throat area results in a 1% increase in fuel consumption, all four of these
variable features can be fixed with only a 1.6% increase in fuel use over the optimal solution. Again a three
stream variable cycle with only modulated cooling, variable HPT inlet area, and variable fan throat area
delivers nearly 94% of the fuel savings offered by the fully variable architecture.
As this mission offers the greatest potential reduction in spillage drag, one would expect that most
of the fuel savings was realized through extended flow holding which yields corresponding reductions in
spillage and increases in propulsive efficiency. The sources of increased overall efficiency illustrated in
figure 68 substantiate this claim. First one notes that the relatively low design bypass ratio to the third
Figure 68. Sources of improved variable cycle efficiency in supersonic strike mission
stream minimizes the fuel savings from the improved fan nozzle CFG but accentuates benefits from
modulated cooling; this is consistent with the results of the similarly low bypass ratio subsonic LRS engine.
While these two effects account for an impressive 8.3% reduction in fuel use, a staggering 16.7% reduction
in fuel use is made possible through reduced spillage and increased propulsive efficiency. The improved
97
performance offered here by the three stream architecture is truly remarkable and can be generated with a
minimum of variable features.
The now familiar movements of the HPT inlet and fan nozzle throat are depicted in figures 69 and
70. As power is reduced a corresponding reduction in HPT inlet area and a steady increase in fan nozzle
A8 is noted at both cruise points of interest. When this is coupled with flow holding, a steady increase in
bypass ratio and therefore propulsive efficiency is realized. It is interesting to note that this three stream
variable cycle is still able to effectively flow hold to 50% power with only a two variable geometries. This
is all the more impressive when one considers that the engine maintains all three compressors within
minimum pressure ratio and surge margin limits despite widely varying internal airflows and only a single
variable exhaust nozzle. Notice, however, that this high Mach cruise is the only location where such
extended duration flow holding is observed; this further corroborates the theory that while extended
duration flow holding is possible it is rarely advantageous.
This sub optimal engine was also evaluated in the context of a notional airframe flying the vision
mission; the results are presented in table 24. In this mission the benefits of the variable cycle are quite
remarkable. Even with the reduced number of variable features a 41% increase in loiter time and a 34%
increase in standoff range is realized. This can translate into a wide range of cost savings for this
persistent strike mission. Not only does this translate into one fewer sorties generated every 24 hour period
but it also translates into a reduction in fuel usage, maintenance hours, on station tankers, deployed ground
support, and a reduction is aircrew required.
Red text indicates negative numbers.
Table 24. Sub optimal supersonic strike mission performance
98
Adaptive feature schedule
Engine performance
Figure 69. Sub optimal supersonic control and performance, 30000 ft Mach 0.5 loiter
99
Adaptive feature schedule
Engine performance
Figure 70. Sub optimal supersonic control and performance, 50000 ft Mach 2.2 cruise
100
Further benefits are visible in figures 71 and 72. For example if the environment was target rich,
the variable cycle enables a 50% increase in payload while still maintaining a seven hour loiter capability
and a 4700 nm standoff range. Furthermore, the unprecedented supersonic cruise range gives a commander
tremendous flexibility. Assets can be transferred rapidly to a deployed location without cumbersome, and
by the way subsonic, tanker support. Furthermore, the nearly 6000 mi standoff range places literally half
the globe within a 4.5 hour flight time. Clearly variable cycle technologies offer a great deal to supersonic
strike aircraft.
Figure 71. Sub optimal supersonic strike loiter and payload for study engines
Notice that when a subset of features was sought that delivered nearly optimal performance with a
reduction in variable geometry, the same set of promising features was recommended for all three missions.
Clearly the most technologically challenging of these variable features to design is the variable high
pressure turbine inlet area. At this location engine temperatures are at a maximum so, any variability must
incorporate sufficient cooling air to maintain reasonable component life. Nonetheless, data to this point
suggests that research and development expenditures on this technology would yield the greatest return on
investment.
101
Figure 72. Sub optimal supersonic standoff range and payload for study engines
102
CHAPTER 4
CONCLUSIONS AND RECOMMENDATIONS
4.0 Conclusions and recommendations
4.1 Viability of variable cycle engines
This research began with the supposition that military aircraft would benefit greatly from an
engine capable of varying internal geometries to deliver both high specific thrust when dictated by the
mission and reduced specific fuel consumption at cruise. Such an engine would combine the traditional
strengths of turbojets and turbofans into a single cycle. As these engines, called variable cycles, have been
a matter of some interest for over 50 years this author can take no credit for this innovative idea.
To date the vast majority of research has focused on the tremendous benefits afforded by variable
cycles in supersonic applications; this is due in large part to the tremendous potential offered by these
engines in missions with both extended supersonic and subsonic cruise segments. Therefore when the
supersonic commercial transport began in the late 1950s, it stimulated an explosion in VCE research. This
period of research continued for roughly 30 years and produced such novel designs as the three stream,
three spool MOBY engine. While this engine was adept at holding airflow to 50% power and improving
propulsive efficiency by moving flow from the core to the second and third streams, its three variable
nozzles, three spools and duct burner made it too complex to develop further.
In contrast to previous efforts, this study sought to determine the applicability of a variable cycle
to a number of disparate emerging military roles including the tactical mobility, subsonic long range strike,
and supersonic strike missions introduced in section 1.3. Furthermore, the objectives of this study extended
beyond merely identifying improvements in uninstalled specific fuel consumption. Rather system level
benefits offered by the variable cycle were sought. Potential benefits include a source of cool air to act as
an aircraft heat sink, a readily available source of pressurized air for lift augmentation, and reductions in
103
inlet spillage and aft body drag offered by maintaining engine airflow at nearly constant levels through the
cruise points of interest.
To evaluate the applicability of variable cycle‟s to the three vision missions, a two spool three
stream architecture was identified, see figures 16-17. This engine, equipped with no less than seven
variable features, was capable of continuously modulating flow to the second and third streams throughout
the mission. This cycle was then modeled in NPSS, a component based object oriented cycle simulator,
and interfaced with optimization routines via integral plug-ins to Model Center®, see figures 13 and 15.
This nested architecture facilitated rapid determination of optimal off design variable geometry settings
through use of a common cost function and genetic algorithms running in parallel at each cruise location.
To determine relative improvements in performance two baseline engines were developed as well.
The first was a year 2000 state of the art two stream turbofan which represents technologies currently
fielded. The second was an advanced turbofan which mimics the same thermodynamic advances that were
incorporated into the variable cycle, see table 1. Therefore, performance differences between the year 2000
baseline and the variable cycle represent benefits would be realized through a new engine program.
Conversely, performance improvements noted between the advanced turbofan and the variable cycle are
the result of improved propulsive efficiency and reduced spillage and aft body drag.
Overall mission performance was investigated by placing each candidate engine in a fixed
airframe and evaluating fuel consumption at the cruise points of interest only. Such analysis is consistent
with an aircraft re-engining program and is considered conservative in its conclusions. That is to say that
the benefits realized by advanced engines are minimized for a given mission as reductions in fuel use are
not accompanied by a corresponding decrease in aircraft vehicle size, structural weight, reduced cruise
thrust, and corresponding further reductions in fuel use. Furthermore, fuel saving associated with reduced
fuel flow during extended ground idle, taxi, takeoff, climb, decent, approach and landing are also not
considered. Nonetheless, the performance improvements offered by the variable cycle are quite staggering.
In all three vision missions fuel use was reduced by a third and mission range was increased by one to two
thirds over the year 2000 state of the art, see optimal lines in table 25.
In order to understand how variable geometries, and hence internal flows, varied from design to
off design, optimized power hooks were generated at each cruise point of interest. This was accomplished
104
Red text indicates negative numbers.
Table 25. Performance summary
by running eleven off design assemblies in Model Center® and finding the minimum installed specific fuel
consumption solution at each throttle setting. These control and performance graphs proved to be most
enlightening. First, this analysis confirmed that as power was reduced engine airflow was held to some
extent in all three missions. During this period of flow holding variable geometries tended to make their
most rapid movements as airflow was increasingly routed to the second and third streams. As expected
each variable cycle was adept at increasing bypass ratio and hence propulsive efficiency, reducing spillage
drag, and improving installed specific fuel consumption. Second, some variable features were noted to
make very modest movements throughout either power hook. These results suggest that nearly optimal
performance can be achieved with a fewer variablilities and therefore a far less complex engine.
Therefore, the subset of variable features which returned the greatest benefit with minimal
complexity was sought. This process began by modifying the cost function to accommodate a penalty for
each variable feature manipulated by the optimizer. In this manner, the variable feature combinations to be
investigated were reduced from 62 possible to a much more manageable 16 promising solutions. Each of
105
these subsets was further investigated by manually fixing the identified components and executing a full on
and off design analysis. Contrary to initial expectations a consistent set of variable features emerged across
all three vision missions.
For this engine architecture, the use of just three variable features can yield an astounding 31-34%
fuel savings over the 2000 state of the art engine and a 12-17% fuel savings over the advanced turbofan in
across the vision missions, see sub optimal VCE lines in table 25. In an engine community that fights for a
1% improvement in efficiency annually, the fuel savings offered here by the three stream variable cycle are
truly astonishing. The three most promising technologies are modulated cooling, a variable high pressure
turbine inlet area, and a variable fan nozzle throat area. By cutting the chargeable HPT and all LPT cooling
air in half at cruise, a 5-6% reduction in fuel consumption is noted in all vision missions. The remaining
reductions in fuel use are primarily noted by reduced spillage drag and increased propulsive efficiency.
These are induced by simultaneously reducing the HPT inlet area, here a maximum of 20%, and increasing
the fan nozzle throat area by as much as 250%. Investments in variable high pressure turbines and
modulated cooling promises to pay remarkable dividends in variable cycle architectures.
The initial goal of this research was to identify, model and evaluate a three stream variable cycle
engine which incorporated a wide range of emerging technologies. Through this analysis it was hoped that
the most promising variable features could be identified and a practical architecture could be proposed that
would consistently realize superior performance to its advance two stream counterpart. In these areas this
study succeeded beyond any expectations. What is even more fascinating is how quickly the analysis tools
developed throughout this study have already impacted academia and industry. For example,
improvements to NPSS developed in conjunction with NASA Glenn Research Center have been
incorporated into NPSS releases 1.6.5 and later. These modifications, which greatly improve cycle
convergence and provide plug-ins for rapidly optimizing variable feature settings, have been of great
assistance to industry partners currently investigating variable cycles. Furthermore, study models have
been used by organizations within the Air Force Research Laboratory, AFRL, Propulsion Directorate,
AFRL Air Vehicles Directorate, and the Arnold Engineering Development Center to assist with their
independent research efforts. Finally, several members of the Versatile Affordable Advanced Turbine
106
Engines, VAATE, community have requested and received models and or optimization methods essential
to their ongoing research efforts.
4.2 Recommended future research
There are a number of future research topics suggested by this study. The first of these fall in the
category of exercising variable features available in this model but not investigated to this point. For
example, an adaptive high pressure compressor was not exercised in this study as no appropriate layered
map was readily available. If such a map were located with reasonable inlet area variation, its effectiveness
in this variable architecture could be readily investigated. The anticipated impact on core airflow is similar
to that offered by a variable HPT; however, large variations in HPC inlet area would likely result in a
noticeable drop in component efficiency and a lesser fuel savings than noted by similar HPT inlet area
variations.
Similarly, a partial augmenter could be exercised during mission segments that require high
specific thrust. What might seem unusual here is that augmentation is not as appealing in supersonic
mission segments where fuel economy is essential. However, this augmenter may prove to be very helpful
during takeoff in the tactical mobility and subsonic long range strike missions. During takeoff, partial
augmentation would assist in meeting the considerable thrust requirement without a corresponding increase
in engine airflow. This could translate into an engine with a reduced cross section and reduced weight.
Such an engine would also decrease the aircraft cross section, structural weight, aerodynamic drag and, for
an embedded application, aircraft length. Or for a given engine cross section, the engine could have a
higher design point bypass ratio and a corresponding improvement in fuel efficiency at all off design points
of interest.
Whichever option is chosen an augmented variable cycle would have a lower top of climb military
thrust and, therefore, operate at higher cruise power settings. This would tend to minimize variable cycle
benefits associated with increased propulsive efficiency and reduced spillage drag. To be comprehensive
such a study would need to investigate a range of partial augmentation and address the system level
impacts. Reductions in cruise fuel consumption would need to be balanced against the increased takeoff
fuel use, an increase in time to climb, and a corresponding reduction in cruise range. Furthermore, the
reduction in survivability resulting from the increased exhaust gas temperature and the potential increase in
107
radar cross section from two variable exhaust nozzles would need to be addressed. Such a study is by no
means trivial.
The second class of suggested research topics would be to investigate military missions not
covered in this study. These missions might benefit from a different set of variable features. For instance, a
high altitude sensor craft mission with an extended duration loiter would place demands on the engine
similar to that of the supersonic strike loiter or the long range strike high altitude cruise. However, onboard
sensors would create the additional demand of large power extractions at high altitude and generate
substantial amounts of waste heat. Such a craft might benefit from a variable inter-turbine burner. This
variable feature would not only enable the extraction of the additional power but it could also increase
thrust during the latter part of the climb. Also, this mission would provide an excellent framework to
investigate transient performance of the variable cycle during introduction and termination of high power
extraction. Finally this mission would offer an opportunity to investigate integrated and adaptive power
thermal systems which utilize a combination of fuel and third stream heat sinks.
Another area that could be investigated further is in the area of genetic algorithm search
optimization. This study expended a great deal of time creating engine models that converged over a wide
range of on and off design parameters. Also, much effort was exerted in restoring solutions in the event of
non convergence so that optimization could continue unhindered. However, minimization of search time
was limited to improvements in convergence criteria, step sizes, perturbation limits and reductions in
input/output to and from Model Center®. For the most part, recommended or default genetic algorithm
settings were used throughout this study. If the default settings resulted in unacceptable run times, a
tradeoff study was conducted which traded population size and number of generations against run time and
marginal improvement in solution. While this was sufficient to keep run times within acceptable limits,
further optimization of genetic algorithm parameters including crossover and mutation probabilities,
population size, number of generations, and maximum generations without improvement is warranted.
Another promising improvement in algorithmic efficiency is offered through the use of spatial statistics, or
more specifically kriging of the response surface. Previous studies suggest that kriging would substantially
reduce the number of objective functions calls required and return optimal designs in roughly a quarter of
the computational time (Millhouse, 2002).
108
This research also makes a compelling case for a variable inlet area high pressure turbine. While
this can be difficult to realize with conventional technologies, recent studies suggest that substantial
excursions in area can be achieved with the less conventional approaches of turbine nozzle camber
variation or fluidic control. Therefore, it is suggested that sufficient evidence exists to justify component
level research on adaptive high pressure turbines with inlet area excursions in excess of 25% closure. As
the desired effect is a reduction in airflow into the high pressure turbine, this component level research
might also wish to investigate variable area combustors. Variations in the combustor or in ducting just
downstream might eliminate much of the complexity associated with variable, cooled HPT nozzles and
deliver similar performance.
Finally, research should be conducted into commercial applications of variable cycles. While
there is little doubt that future supersonic transport aircraft would benefit from variable technologies, it is
suggested that subsonic commercial airliners could also realize a measurable benefit. Ever decreasing fan
pressure ratios have made variable fan nozzles appealing for future commercial engines (Cumpsty, 2009).
Furthermore, a modulated cooling system which halves the cooling flow to the high pressure turbine
promises a TSFC improvement of 2% (Cumpsty, 2009). By simply adding a variable HPT inlet area, one
could substantially vary flow to the core and realize numerous benefits. For example, one could increase
the specific thrust at takeoff. This would reduce takeoff length, increase climb gradient thereby reducing
noise in the vicinity of the airport and increasing the length of the more efficient cruise segment. Then by
modulating flow to the bypass stream, or streams, one could then return to a high bypass, lower fan
pressure ratio operation at cruise. As the aircraft burned fuel, cruise thrust required would be reduced and
the bypass ratio continuously increased. While the gains in the commercial applications will likely be
smaller, they will be multiplied over thousands of engines and millions of sorties annually; such savings in
fuel costs, reductions in emissions, and reduced noise are worth pursuing.
109
BIBLIOGRAPHY
Bibliography
Adibhatla, Shrider and Johnson, Karla “Evaluation of a Nonlinear PSC Algorithm on a Variable Cycle
Engine”, AIAA, Paper No. AIAA-93-2077, 1993.
Aleid, L. and Pilidis, P. “Variable cycle jet engines for a Mach 2.7 supersonic civil transport”, The
Aeronautical Journal, Paper No. 2352, pp. 31-36, 1997.
Armstrong, F.W. and Higiton, D.R. “Some Aspects for variable Cycle Propulsion Systems”, AGARD
Conference Proceedings No 205, Seine France, 1977.
Bowers, Douglas L. “Throttle Dependent Forces”, Air Force Wright Aeronautical Laboratories Technical
Report, AFWAL-TR-85-3055 (Approved for public release), 1985.
Brear, Micheal J., Kerrebrock, Jack L., and Epstein, Alan H. “Propulsion System Requirements for Long
Range, Supersonic Aircraft”, ASME Journal of Fluids Engineering, Vol. 128, pp. 370-377, 2006.
Brown, H. “Multi-variable Cycle Optimization by Gradient Methods”, AIAA, Paper No. AIAA-80-52,
1980.
Bruening, Greg B. and Chang, Won S. “Cooled Cooling Air Systems for Turbine Thermal Management”,
Paper 99-GT-014, 44th
ASME Gas Turbine and Aeroengine Technical Congress, Indianapolis, 1999.
Carr, John E. “An Aerodynamic Comparison of Blown and Mechanical High Lift Airfoils”, Proceedings of
the Circulation-Control Workshop NASA Ames Research Center, pp. 449-477, 1986.
Conners, T. R., Howe, D. C., and Whur, J. R. “Impact of Engine Cycle Selection on Propulsion System
Integration and Vehicle Performance for a Quiet Supersonic Aircraft”, AIAA-2005-1016, 43rd
AIAA
Aerospace Sciences Meeting and Exhibit, Reno, 2005.
Cumpsty, Nicholas A. “Preparing for the Future: Reducing Gas Turbine Environmental Impact”, GT 2009-
60367, IGTI Scholar lecture, London, 2009.
Davenport, W. R., Dixon, G. J., Fischer, A. Fledderjohn, K. R., Shearer, J. R., Steele, M. A., Stephenson,
D. W., and Weber, F. “Impact of Turbine Modulation on Variable-Cycle Engine Performance”, Air Force
Aero Propulsion Laboratory Technical Report, AFAPL-TR-74-109(Approved for public release), 1974.
Edwards, Tim. “Liquid Fuels and Propellants for Aerospace Propulsion: 1903-2003”, Journal of
Propulsion and Power, Vol. 19, No. 6, pp. 1089-1107, 2003.
Goldberg, David E. Genetic Algorithms in Search, Optimization, and Machine Learning, Addison-Wesley,
Reading, 1989.
Grey, D.E. “Study of Unconventional Aircraft Engines Designed for Low Energy Consumption – Pratt &
Whitney”, National Aeronautics and Space Administration Technical Report, NASA-CR-135065, 1976.
110
Grönstedt, Thomas. “Development of methods for analysis and optimization of complex jet engine
systems”, Ph.D. Dissertation, Chalmers University of Technology, 2000.
Grönstedt, U.T. and Pilidis, P. “Control Optimization of the Transient Performance of the Selective Bleed
Variable Cycle Engine During Mode Transition”, Journal of Engineering for Gas Turbines and Power,
Vol. 124, pp. 75-81, 2002.
Hill, Philip and Peterson, Carl. Mechanics and Thermodynamics of Propulsion 2nd
edition, Addison-
Wesley, Reading, 1992.
Johnson, J. E. “Variable Cycle Engines – The Next Step in Propulsion Evolutions?”, AIAA, Paper No.
AIAA-76-758, 1976.
Johnson, J. E. “Variable Cycle Engine Concepts”, AGARD Conference Proceedings 572, 1996.
Johnson, J. E. “Variable Cycle Engine Developments at General Electric: 1955-1995”, AIAA Progress in
Astronautics and Aeronautics, Vol. 165. Developments in High-Speed-Vehicle Propulsion Systems, 1996.
Johnson, James E.. et al. “Variable Cycle Gas Turbine Engines”, US Patent 4064692, 27 Dec 1977.
Koff, Bernard L. et al. “Modulating Bypass Variable Cycle Turbofan Engine”, US Patent 4080785, 28 Mar
1978.
Kuprowicz, Nicholas J. “The integrated Multi-Objective Multi-Disciplinary Jet Engine Design
Optimization Program”, Air Force Research Laboratory Technical Report, AFRL-PR-WP-TR-1999-2028
(Approved for public release), 1999
Latimer, R. J. “Variable Flow Turbines”, AGARD Conference Proceedings No 205, Seine France, 1977.
Mattingly, Jack D., Heiser, W. H. and Daley, Daniel H. Aircraft Engine Design, AIAA, New York, 1987.
May, Robert J., Tall, Wayne A. and Bush, H. Ivan “Potential Improvements in Engine Performance using a
variable Geometry Turbine”, AGARD Conference Proceedings No 205, Seine France, 1977.
McArdle, Jack G., Homyak, Leonard and Moore, Allan S. “Static and Transient Performance of YF-102
Engine with up to 14 Percent Core Airbleed for the Quiet Short-Haul Research Aircraft”, National
Aeronautics and Space Administration Technical Report, Paper 1692, 1980.
Millhouse, Paul T., “Improving the Algorithmic Efficiency of Aircraft Engine Design for Optimal
Performance”, MS Thesis, AFIT/GOR/ENY/98M-02, Air Force Institute of Technology, Wright Patterson
AFB, OH, 1998.
Münzberg H.G. and Kurzke, J. “The Pros and Cons of Variable Geometry Turbines”, AGARD Conference
Proceedings No 205, Seine France, 1977.
Murthy, S. N. and Curran, E. T., editors. Developments in High-Speed-Vehicle Propulsion Systems, AIAA,
Reston. 1996.
Nadon, Luc J. “Multidisciplinary and Multiobjective Optimization in Conceptual Design for Mixed-Stream
Turbofan Engines”, MS Thesis, AFIT/GAE/ENY/96D-6, Air Force Institute of Technology, Wright
Patterson AFB, 1996.
National Aeronautics and Space Administration, NPSS User Guide, Version 1.6.5, NASA Glenn Research
Center, Cleveland, 2008.
111
National Aeronautics and Space Administration, NPSS Reference Sheets, Version 1.6.5, NASA Glenn
Research Center, Cleveland, 2008.
Neitzel, R.E., Hirsckron, and R. Johnston, R.P. “Study of Unconventional Aircraft Engines Designed for
Low Energy Consumption – General Electric”, National Aeronautics and Space Administration Technical
Report, NASA-CR-135136, 1976.
Neitzel, R.E., Hirsckron, and R. Johnston, R.P. “Study of Turbofan Engines Designed for Low Energy
Consumption – General Electric”, National Aeronautics and Space Administration Technical Report,
NASA-CR-13503, 1976.
Payzer, R.J., “Variable Cycle Engine Applications and Constraints”, AGARD Conference Proceedings No
205, Seine France, 1977.
Perkins, G. M. “Three Stream Turbofan-Variable Cycle Engine with Integral Turbo Compressor”, Paper
AIAA-97-2104, 23rd
Joint Propulsion Conference, San Diego, 1987.
Salay, C. R. and Elliott, D. W. “Matching Engine and Aircraft Lapse Rates for High Speed Civil Aircraft”,
Journal of Aircraft, Vol. 33 No. 1, pp. 61-67, 1996.
Simonsen, Erick. “Boeing Historical Perspective, What Might Have Been”, Boeing Frontiers online,
Volume 3, Issue 6, 2004.
Strack, William C. and Shelby, Morris J., “The Challenges and Opportunities of Supersonic Transport
Propulsion Technology”, AIAA, Paper No. AIAA-88-2985, 1988.
Swan, W. C., Welliver, A. D., Klees, G. W., and Kyle, S.G. “Opportunities for Variable Geometry Engines
in Military Aircraft”, AGARD Conference Proceedings No 205, Seine France, 1977.
Thomas, Elmore V. et al. “Two Spool Variable Cycle Engine”, US Patent 4043121, 23 Aug 1977.
Ulizar, Inaki and Pilidis, Pericles. “The Handling of a Variable Cycle Engine: The selective Bleed
Turbofan”, Paper 93-GT-384, International Gas Turbine and Aeroengine Technical Congress, Cincinnati,
1993.
Ulizar, Inaki and Pilidis, Pericles. “Transition Control and Performance of the Selective Bleed Variable
Cycle Turbofan”, Paper 95-GT-286, International Gas Turbine and Aeroengine Technical Congress,
Houston, 1995.
Ulizar, I. and Pilidis, P. “Predicted performance characteristics of a variable cycle turbofan”, The
Aeronautical Journal, Paper No. 2190, pp. 263-268, 1997.
Vanderplaats, Garret N. Numerical Optimization Techniques for Engineering Design, third Edition,
Vanderplaats Research & Development Inc., Colorado Springs, 1999.
Varelis, A.G. “Variable Cycle Engine for Combat STOVL Aircraft” MS Thesis, Canfield University,
United Kingdom, 2007.
Woollett, Richard R. “Aerodynamic Performance of a Fan Stage Utilizing variable Inlet Guide Vanes
(VIGVs) for Thrust Modulation”, National Aeronautics and Space Administration Technical Report,
NASA-TR-83438, 1983.
Wood, Alan and Pilidis, Pericles. “A variable cycle jet engine for ASTOVL aircraft”, Aircraft Engineering
and Aerospace Technology, Vol. 69 No. 6, pp. 534-539, 1997.
112
Younosi, O., Arena, M. Moore, R., Lorell, M. Mason, J. and Graser R. “Military jet Engine Acquisition:
Technology Basics and Cost-Estimating Methodology”, RAND, Santa Monica, 2002.
