DESIGN AND CONTROL OF A VARIABLE GEOMETRY 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 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.

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 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 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 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 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.

ii

Dedication Dedicated to my beloved bride Bonnie.

iii

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 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.

iv

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

v

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 vi

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

vii

LIST OF FIGURES List of figures Page Figure 1. Notional variable cycle engine ...... 3 Figure 2. Propulsive efficiency as a function of mach and ...... 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. ...... 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. 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 viii

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 ix

NOMENCLATURE Nomenclature

2 Ao Inlet stream tube area (in ) 2 A8 Nozzle throat area (in ) 2 A9 Nozzle exit area (in ) 2 A10 Aircraft aft body area (in ) 2 Ac Inlet capture area (in ) AGL Above Ground Level (ft)

BPR1 Bypass ratio 1, mass flow of third stream / mass flow of LPC nd BPR2 Bypass ratio 2, mass flow of 2 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 o Cp Specific heat at constant pressure (BTU/lbm R)

D Overall aircraft drag (lbf)

D AB Aft body drag (lbf)

D spill Spillage drag (lbf)

F Gross thrust (lbf)

Fn Net thrust (lbf) 2 g Gravitational acceleration (32.2 lbm ft/s ) HPC High pressure compressor HPT High pressure turbine HPTB High pressure 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

x m Mass flow rate of the fuel (lb /s) f m m o 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 2 p Total pressure (lbf/in ) P Engine mechanical power (hp) 2 PSTD Sea Level, standard day pressure (14.7 lbf/in ) Heat flux (BTU/s)

Q Fuel lower heating value (18,400 BTU/lbm) r Range (nm) o R Gas constant for air (53.3 ft-lbf/lbm- R) S Aircraft surface area (ft2) t Time (s) o To Free stream air ( R) o T3 High pressure compressor discharge temperature ( R) o T41 High pressure turbine rotor inlet temperature ( R) o TSTD Sea Level, standard day temperature (518.7 R)

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)

Wc des 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 of the gas generator 3  Atmospheric density (lbm/ft )

xi

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 . This would offer a dramatic reduction in spillage drag to supersonic aircraft operating at part power (Johnson, 1996). Furthermore, this

1 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 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 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

2 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.

Constant Flow with Variable Separate Modulatable Variable Area Fan Pressure Ratio Third Stream Bypass Injector

Substantial Cooling Air for Exhaust & Aircraft Thermal Management

Inlet Adaptive Adaptive Exhaust Fan Core

Modulated Cooled Variable Area Variable Core and Bypass Cooling Air Compressors & Turbines Exhaust

Figure 1. Notional variable cycle engine

Engine cycle efficiency is usually expressed as the product of thermal and propulsive efficiency.

o  p th

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),

Aircraft Thrust Power 2 V0  p   Engine Mechanical Power Ve V0

Where: Vo is the aircraft velocity

Ve is the nozzle exhaust velocity

3

Note that propulsive efficiency is maximized when the nozzle exhaust velocity is at a minimum. This is why the most efficient 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),

m 2 2 0 V V Engine Mechanical Power 2  e 0  th   Thermal Energy Input Rate m f Q

Where: m f is the mass flow rate of the fuel

m o 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 , 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.

F  Ve V0 m o

Where: F is the gross thrust

Note:  Mass flow rate of fuel is considered negligible.

 Only gross momentum thrust, m o Ve , and intake momentum drag, m o Vo , terms are shown

here to illustrate the difficulty of simultaneously maximizing F m o 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 or low bypass ratio turbofans.

4

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.

Aircraft Thrust Power F V    o p Engine Mechanical Power P

m V V V m V V  V  o e o o  o e o o m 2 2 m o V V o V V V V 2  e o  2  e o  e o  m V m V  o o  o o m F o V V  2V m o Vo  2  e o  o  2 1   p    1  1  F   2 Vo  m o 

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:

1   p     1  1   F   60 M o   m o 

Where: Vo  M o  R To

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 lb 60 is the constant from the calculation above with units of m lbf s

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

5

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.

1.0

0.9 F/mo=12.5

0.8 F/mo = 25 0.7 F/mo = 50 0.6 F/mo = 75 p 0.5

 F/mo = 100

0.4 F/mo = 150

0.3 F/mo = 200

0.2

0.1

0.0 0.0 0.5 1.0 1.5 2.0 2.5 3.0 Mach

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

TSFC* TSFC*

TSFC* TSFC*

M = 0.7 Mth = 0.7

Installed Thrust Figure 3. External airflow at maximum power design point 6

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,

TSFC* TSFC* TSFC* M = 0.7 Mth = 0.4 TSFC*

Installed Thrust

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)

7

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

8 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 9 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 10 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

Bypass Nozzle

First Bypass First Bypass

Core Flow Core Flow HP Shaft

LP Shaft

Intake LPC IPC HPC Comb. HPT LPT Mixer Primary Nozzle Chamber

Second Bypass

Core Flow Core Flow HP Shaft

LP Shaft

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 11 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 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

12

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.

13

Penetration 2.2 Mn / 50,000 ft

2.2 Mn / 50,000 ft Full weapons complement

2.2 Mn / 50,000 ft

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.

14

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

15

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.

NPSS OF

GA

Off Design Point 1 Objective Genetic Algorithm Function (OF) (GA)

NPSS OF

GA Off Design Point n Off Design Optimization Loop

Design Point (NPSS Model) On Design Optimization Loop

Figure 13. Nested optimization loops Off Design Optimization Loop 16

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 85 / na / 85 / 87 / 88 88.5 / na / 86 / 89 / 90 88.5 / 88.5 / 86 / 89 / 90 (Fan/LPC/HPC/HPT/LPT) Cooling %W * 15 / 10 / 4 / 2 25 10 / 5 / 5 / 2 15 / 10 / 4 / 2 (HPTN/HPTB/LPTN/LPTB) Modulated 15 / 5 / 2 / 1 ** Primary Nozzle CFG 0.95 0.97 0.97 Pri. Nozzle Cooling CFG*** 0.92 0.92 n/a Fan Nozzle CFG n/a n/a 0.96

Inlet Flow Control (%W23) 0.0 2.5 2.5

Ram Recovery 0.95 0.97 0.97

0 T3max ( F) 1200 1400 1400

0 T41max ( F) 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

o o bleeding air from the rear of the compressor and cooling it to 1150 F, which is 250 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

17

Legend:

Air Flow path

Cooling Air Flow

Fuel Input

Shaft Connection

Nozzle Cooling (modeled as a separate nozzle) Wing Blowing Wing OB Variable Components: Sink Sink - Nozzle throat area - Primary & secondary inlet areas to mixer

Customer Bleed

Nozzle Fan Nozzle Nozzle Leakage Duct Split 2 Cooling Cooling Cooling Duct Sink

18 Fuel 36 Fuel AB

F H H L

Aftr Rear Primary Front A Comp P P Primary Amb0 Inlet IFC Frame Split 1 Duct Burn 36 B41 B42 B45 P B52 D52 Mixer Burn Frame B7 Nozzle Nozzle Sink N C T T

Gear Box

HP Extraction

Figure 14. 2000 State of the art and advanced turbofan NPSS component based architecture

Legend:

Air Flow path

Bleed Air Flow

Heat In Fuel Input Wing Blowing Wing OB Sink Sink Shaft Connection

Variable Components: - Nozzle throat areas

Customer Fan Fan Fan - Inlet Area to Compressors & Turbines Nozzle Bleed Duct B24 Nozzle Sink - Primary & secondary inlet areas to mixer

Fan Leakage Duct 2

Fuel 36 Fuel AB

19 F L H H L

Inlet Front LPC Comp After Rear Primary Primary Flow A Split 1 Split 2 Burn 36 B41 B42 B45 B52 D52 Mixer B7 Nozzle Amb0 Inlet Frame Duct P Duct P P P Burner Frame Nozzle Control Sink N C C T T

Gear Box

Heat Heat Out Exchr

HP Extraction

Figure 15. Double bypass variable cycle NPSS component based architecture

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

20

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.

21

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 22

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

23

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 Reference Inlet Spillage Drag Coefficient 0.06

0.05

0.04

0.03

d spill ref spill d C

0.02

0.01

0.00 0.5 0.7 0.9 1.1 1.3 1.5 1.7 1.9 2.1 2.3 2.5

Mach Number Figure 19. Supersonic reference spillage drag vs. Mach number

Inlet Spillage Drag Coefficient 0.40 Mach 2.0 0.35 Mach 1.6 Mach 1.2 0.30 Mach 0.8 Mach 0.6 0.25

0.20 d spill area spill d

C 0.15

0.10

0.05

0.00 0.3 0.4 0.5 0.6 0.7 0.8 0.9

Ao/Ac Figure 20. Supersonic spillage drag vs. airflow ratio 24 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),

1  1 2 W P 2  g  c std 2 1  1 M o 1 0.5( 1)M o    Ao T 2  R  std

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

2 D  C 1  V A spill d spill  2 o  c

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

25 it appears advantageous to flow hold as long as practical. Potential spillage drag in the other two vision missions and conditions in which it isTactical better to Mobility cease flow Spillage holding Dragwill be addressed further in Chapter 3. 20

18 Advanced Turbofan, 4000 ft 0.4 Mn Advanced Turbofan, 35000 ft 0.8 Mn 16

14

12

10

8

6

4

2

% Increase % in Aircraft Drag Spillage due to 0 0 10 20 30 40 50 60 70 80 90 100 % Military Thrust

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,

1 2 DAB  Cd AB   Vo  A10 2

Where: D AB is the aft body drag

Cd AB is the aft body drag coefficient 26

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 thereforeAft reduced Body aft Drag body Coefficientdrag; these results are presented in Chapter 3. 0.25 Mach 2.0 Mach 1.2 Mach 1.0 0.20 Mach 0.8

0.15

d AB d C

0.10

0.05

0.00 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0 A / A 10 9 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

27

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 28

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, 29

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. 30

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, rd Is Genetic Yes corrected speeds & 3 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 31

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

32 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 33

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

34

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 35 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.

36

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

37

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 benefitUnrestricted realized. Flow Holding, Airflow vs. Installed Thrust 250

Overall Airflow 200 Third Stream Airflow Second Stream Airflow

) Core Airflow m 150

100 Mass Flow (lb Flow Mass

50

0 60 65 70 75 80 85 90 95 100

% Military Thrust

Figure 27. Internal airflow variations with unrestricted flow holding

38

Unrestricted Flow Holding, Bypass Ratio vs. Installed Thrust 4.5 Overall BPR 4.0 BPR1 BPR2 3.5

3.0

2.5

2.0 Bypass RatioBypass 1.5

1.0

0.5

0.0 60 65 70 75 80 85 90 95 100 % Military Thrust

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

39 greater. Second, flow holding can be terminated when the propulsive benefit of increased bypass ratio is surpassed by the duct pressureUnrestricted drop and other Flow losses. Holding, This second Duct option Losses will vs. be developedInstalled further Thrust below. 0.6 18% Fan Duct Mn 16% 0.5 Fan Duct Loss 14%

0.4 12%

10% 0.3 8%

0.2 6%

4%

Third Stream Duct Mach Number Mach Duct Stream Third 0.1 2% (%)PressureLossDuct Stream Third

0.0 0% 60 65 70 75 80 85 90 95 100

% Military Thrust

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 40 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. Unrestricted Flow Holding, LPC Schedule vs. Installed Thrust 100.0 1.80

90.0 1.60

80.0 1.40 70.0 LPC IGV Setting LPC Efficiency 1.20 60.0 LPC PR 1.00 50.0 0.80 40.0

LPC IGV Setting IGV LPC 0.60 30.0 0.40 20.0

10.0 0.20 Ratio Pressure or Efficiency Adiabatic

0.0 0.00 60 65 70 75 80 85 90 95 100

% Military Thrust

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- 41

Unrestricted Flow Holding, Nozzle Throat Area vs. Installed Thrust

1000

Primary Nozzle 900 Fan Nozzle

800

700

600

Nozzle Throat AreaNozzle Throat 500

400

300 60 65 70 75 80 85 90 95 100 % Military Thrust

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

42 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.

43

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 44

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.

45

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 46 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. Tactical Mobility Spillage Drag 20 Advanced Turbofan, 4000 ft 0.4 Mn 18 Variable Cycle, 4000 Ft 0.4 Mn 16 Advanced Turbofan, 35000 ft 0.8 Mn Variable Cycle, 35000 ft 0.8 Mn 14

12

10

8

6

4

2

% Increase % in Aircraft Drag Spillage due to 0 0 10 20 30 40 50 60 70 80 90 100

% Military Thrust 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

o mission, the takeoff from a 1500 ft runway at 4000 ft pressure altitude on a 95 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.

47

Subsonic Long Range Strike Spillage Drag

30 Advanced Turbofan, 500 ft 0.7 Mn Variable Cycle, 500 Ft 0.7 Mn 25 Advanced Turbofan, 40000 ft 0.8 Mn Variable Cycle, 40000 ft 0.8 Mn

20

15

10

5 % Increase in % Spillage Aircraft Drag due to

0 0 10 20 30 40 50 60 70 80 90 100

% Military Thrust 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

o from an 8,000 ft runway at sea level 95 F 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

48 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 missionSupersonic fuel use. Strike Spillage Drag

50 Advanced Turbofan, 30000 ft 0.5 Mn 45 Variable Cycle, 30000 Ft 0.5 Mn

40 Advanced Turbofan, 50000 ft 2.2 Mn Variable Cycle, 50000 ft 2.2 Mn 35

30

25

20

15

10

% Increase % in Aircraft Drag Spillage due to 5

0 0 10 20 30 40 50 60 70 80 90 100

% Military Thrust 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

49 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. Tactical Mobility Aft Body Drag 1.8 Advanced Turbofan, 4000 ft 0.4 Mn 1.6 Advanced Turbofan, 35000 ft 0.8 Mn Variable Cycle, 4000 ft 0.4 Mn 1.4 Variable Cycle, 35000 ft 0.8 Mn 1.2

1.0

0.8

0.6

0.4

0.2

0.0 % Increase % in Aircraft Drag Aft due to Body 0 10 20 30 40 50 60 70 80 90 100

% Military Thrust 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

50 hence the potential savings. The variable cycle was effective in reducing the aft body drag by roughly one quarter percent at high cruse andSubsonic one percent Long during Range the Strike low altitude Aft Body penetration. Drag 4.0 Advanced Turbofan, 500 ft 0.7 Mn 3.5 Advanced Turbofan,40000 ft 0.8 Mn Variable Cycle, 500 ft 0.7 Mn 3.0 Variable Cycle, 40000 ft 0.8 Mn

2.5

2.0

1.5

1.0

0.5

0.0 % Increase % in Aircraft Drag Aft due to Body 0 10 20 30 40 50 60 70 80 90 100 % Military Thrust

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

51

Supersonic Strike Aft Body Drag 14.0 Advanced Turbofan, 30000 ft 0.5 Mn 12.0 Advanced Turbofan,50000 ft 2.2 Mn Variable Cycle, 30000 ft 0.5 Mn Variable Cycle, 50000 ft 2.2 Mn 10.0

8.0

6.0

4.0

2.0

0.0 % Increase in % Aft Aircraft Drag due to Body 0 10 20 30 40 50 60 70 80 90 100 % Military Thrust

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

52 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. 53

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,

54 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

55

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

56

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 57 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

vane setting * op line N c (%) 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

* 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)

Table 3. Optimal tactical mobility variable geometry at 4000ft, Mach 0.0, 95o F design point

58 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.

bypass ratio throat area (in 2 ) 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

Table 4. Optimal tactical mobility BPR & nozzle settings at 4000ft, Mach 0.0, 95o F design point

59

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.

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 -- 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

Table 5. Optimal tactical mobility variable feature settings at 4000 ft, Mach 0.4 cruise

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

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, 60 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. 61

Adaptive feature schedule

Engine performance

Figure 46. Fuel optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise 62

Adaptive feature schedule

Engine performance

Figure 47. Fuel optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise 63

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.

Maximum Aircraft Maximum Maximum Nominal gross structure fuel load payload payload

Weights (lbf) 275,000 115,000 80,000 80,000 60,000

2000 SOA Advanced 3 stream turbofan turbofan VCE Engine weight (lbf) 20,000 18,000 19,800 Table 7. Tactical mobility aircraft parameters

64

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. 65

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.

66

vane setting * op line N c (%) 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

* 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)

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.

bypass ratio throat area (in 2 ) 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

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

67 propulsive efficiency. While the data here is insufficient to corroborate this supposition, it does motivate the more detailed analysis conducted in section 3.10.

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 -- 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

Table 11. Optimal subsonic LRS variable feature settings at 500 ft, Mach 0.7 cruise

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 -- 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 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 68

Adaptive feature schedule

Engine performance

Figure 49. Fuel optimal Subsonic LRS control and performance, 500 ft Mach 0.7 cruise 69

Adaptive feature schedule

Engine performance

Figure 50. Fuel optimal Subsonic LRS control and performance, 40000 ft Mach 0.8 cruise 70 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.

Maximum Aircraft Maximum Maximum Nominal gross structure fuel load payload payload

Weights (lbf) 225,000 82,500 112,500 30,000 20,000

2000 SOA Advanced 3 stream turbofan turbofan VCE Engine weight (lbf) 10,000 9,000 9,900

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 71 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 72 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 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

vane setting * op line N c (%) 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

* 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)

Table 15. Optimal supersonic strike variable geometry at 55000 ft, Mach 2.5 design point

73 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.

bypass ratio throat area (in 2 ) 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

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.

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 -- 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

Table 17. Optimal supersonic strike variable feature settings at 30000 ft, Mach 0.5 loiter

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

Table 18. Optimal supersonic strike variable feature settings at 50000 ft, Mach 2.2 cruise

74

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.

Maximum Aircraft Maximum Maximum Nominal gross structure fuel load payload payload

Weights (lbf) 250,000 90,000 125,000 30,000 20,000

2000 SOA Advanced 3 stream turbofan turbofan VCE Engine weight (lbf) 15,000 13,500 14,850

Table 19. Supersonic strike aircraft parameters 75

Adaptive feature schedule

Engine performance

Figure 52. Fuel optimal supersonic strike control and performance, 30000 ft Mach 0.5 loiter 76

Adaptive feature schedule

Engine performance

Figure 53. Fuel optimal supersonic strike control and performance, 50000 ft Mach 2.2 cruise 77

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

78 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

79 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

80 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.

81

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

82

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 83 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.

84

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 85

Adaptive feature schedule

Engine performance

Figure 59. Sub optimal tactical mobility control and performance, 4000 ft Mach 0.4 cruise 86

Adaptive feature schedule

Engine performance

Figure 60. Sub optimal tactical mobility control and performance, 35000 ft Mach 0.8 cruise 87 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

88 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

89

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 90 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

91

Adaptive feature schedule

Engine performance

Figure 64. Sub optimal subsonic LRS control and performance, 500 ft Mach 0.7 cruise 92

Adaptive feature schedule

Engine performance

Figure 65. Sub optimal subsonic LRS control and performance, 40000 ft Mach 0.8 cruise 93 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 94 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

95 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 96 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

97

Adaptive feature schedule

Engine performance

Figure 69. Sub optimal supersonic control and performance, 30000 ft Mach 0.5 loiter 98

Adaptive feature schedule

Engine performance

Figure 70. Sub optimal supersonic control and performance, 50000 ft Mach 2.2 cruise 99

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.

100

Figure 72. Sub optimal supersonic standoff range and payload for study engines

101

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

102 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 103

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

104 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

105

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 106 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). 107

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 . Variations in the 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.

108

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 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.

109

Grönstedt, Thomas. “Development of methods for analysis and optimization of complex 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. 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.

110

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.

111

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.

112