The development of clean, sustainable energy systems is one of the grand challenges of our time. Most projections indicate that combustion-based energy conversion systems will remain the predominant approach for the majority of our energy usage. Moreover, gas turbines will remain a very significant technology for many decades to come, whether for aircraft propulsion, power generation, or mechanical drive applications. This book compiles the key scientific and technological knowledge associated with gas turbine emis-sions into a single authoritative source. The book has three parts: the first part reviews major issues with gas turbine combustion, including design approaches and constraints, within the context of emissions. The second part addresses fundamental issues associated with pollutant formation, modeling, and prediction. The third part features case studies from manufacturers and technology developers, emphasizing the system-level and prac-tical issues that must be addressed in developing different types of gas turbines that emit pollutants at acceptable levels.
Timothy C. Lieuwen is professor of aerospace engineering and executive director of the Strategic Energy Institute at the Georgia Institute of Technology. Lieuwen has authored one textbook, edited two books, written seven book chapters and more than 200 papers, and received three patents. He chaired the Combustion and Fuels Committee of the International Gas Turbine Institute of the American Society of Mechanical Engineers (ASME). He is also on the Propellants and Combustion Technical Committee of the American Institute of Aeronautics and Astronautics (AIAA), and he previously served on the AIAA Air Breathing Propulsion Technical Committee. He has served on a variety of major panels and committees through the National Research Council, Department of Energy, NASA, General Accounting Office, and Department of Defense. Lieuwen is the editor in chief of the AIAA Progress in Astronautics and Aeronautics series and is serving or has served as an associate editor of the Journal of Propulsion and Power, Combustion Science and Technology, and the Proceedings of the Combustion Institute. Lieuwen is a Fellow of the ASME and received the AIAA Lawrence Sperry Award and the ASME Westinghouse Silver Medal. Other recognitions include ASME best paper awards, the Sigma Xi Young Faculty Award, and the NSF CAREER award.
Vigor Yang is the William R. T. Oakes Professor and chair of the School of Aerospace Engineering at the Georgia Institute of Technology. Prior to joining the faculty at Georgia Tech, he was the John L. and Genevieve H. McCain Chair in Engineering at the Pennsylvania State University. His research interests include combustion instabilities in propulsion systems, chemically reacting flows in air-breathing and rocket engines, com-bustion of energetic materials, and high-pressure thermodynamics and transport. Yang has supervised more than forty PhD and fifteen MS theses. He is the author or coauthor of more than 300 technical papers in the areas of propulsion and combustion and has pub-lished ten comprehensive volumes on rocket and air-breathing propulsion. He received the Penn State Engineering Society Premier Research Award and several publication and technical awards from AIAA, including the Air-Breathing Propulsion Award (2005), the Pendray Aerospace Literature Award (2008), and the Propellants and Combustion Award (2009). Yang was the editor in chief of the AIAA Journal of Propulsion and Power (2001–9) and is currently the editor in chief of the JANNAF Journal of Propulsion and Energetics (since 2009) and coeditor of the Cambridge Aerospace Series. He is a Fellow of the American Institute of Aeronautics and Astronautics, American Society of Mechanical Engineers, and Royal Aeronautical Society.
Editors: Wei Shyy and Vigor Yang
1. J. M. Rolfe and K. J. Staples (eds.): Flight Simulation 2. P. Berlin: The Geostationary Applications Satellite 3. M. J. T. Smith: Aircraft Noise
4. N. X. Vinh: Flight Mechanics of High-Performance Aircraft 5. W. A. Mair and D. L. Birdsall: Aircraft Performance
6. M. J. Abzug and E. E. Larrabee: Airplane Stability and Control 7. M. J. Sidi: Spacecraft Dynamics and Control
8. J. D. Anderson: A History of Aerodynamics
9. A. M. Cruise, J. A. Bowles, C. V. Goodall, and T. J. Patrick: Principles of Space Instrument Design
10. G. A. Khoury (ed.): Airship Technology, Second Edition 11. J. P. Fielding: Introduction to Aircraft Design
12. J. G. Leishman: Principles of Helicopter Aerodynamics, Second Edition 13. J. Katz and A. Plotkin: Low-Speed Aerodynamics, Second Edition
14. M. J. Abzug and E. E. Larrabee: Airplane Stability and Control: A History of the Technologies that Made Aviation Possible, Second Edition
15. D. H. Hodges and G. A. Pierce: Introduction to Structural Dynamics and Aeroelasticity, Second Edition
16. W. Fehse: Automatic Rendezvous and Docking of Spacecraft 17. R. D. Flack: Fundamentals of Jet Propulsion with Applications
18. E. A. Baskharone: Principles of Turbomachinery in Air-Breathing Engines 19. D. D. Knight: Numerical Methods for High-Speed Flows
20. C. A. Wagner, T. Hüttl, and P. Sagaut (eds.): Large-Eddy Simulation for Acoustics 21. D. D. Joseph, T. Funada, and J. Wang: Potential Flows of Viscous and Viscoelastic Fluids 22. W. Shyy, Y. Lian, H. Liu, J. Tang, and D. Viieru: Aerodynamics of Low Reynolds Number
Flyers
23. J. H. Saleh: Analyses for Durability and System Design Lifetime 24. B. K. Donaldson: Analysis of Aircraft Structures, Second Edition 25. C. Segal: The Scramjet Engine: Processes and Characteristics
26. J. F. Doyle: Guided Explorations of the Mechanics of Solids and Structures 27. A. K. Kundu: Aircraft Design
28. M. I. Friswell, J. E. T. Penny, S. D. Garvey, and A. W. Lees: Dynamics of Rotating Machines 29. B. A. Conway (ed.): Spacecraft Trajectory Optimization
30. R. J. Adrian and J. Westerweel: Particle Image Velocimetry
31. G. A. Flandro, H. M. McMahon, and R. L. Roach: Basic Aerodynamics 32. H. Babinsky and J. K. Harvey: Shock Wave–Boundary-Layer Interactions 33. C. K. W. Tam: Computational Aeroacoustics: A Wave Number Approach 34. A. Filippone: Advanced Aircraft Flight Performance
35. I. Chopra and J. Sirohi: Smart Structures Theory 36. W. Johnson: Rotorcraft Aeromechanics
37. W. Shyy, H. Aono, C. K. Kang, and H. Liu: An Introduction to Flapping Wing Aerodynamics
Gas Turbine Emissions
TimoThy C. Lieuwen
Georgia Institute of TechnologyViGor yanG
Georgia Institute of Technology
edited by
Singapore, São Paulo, Delhi, Mexico City
Cambridge University Press
32 Avenue of the Americas, New York, NY 10013-2473, USA
www.cambridge.org
Information on this title: www.cambridge.org/9780521764056
© Timothy C. Lieuwen and Vigor Yang 2013
This publication is in copyright. Subject to statutory exception and to the provisions of relevant collective licensing agreements, no reproduction of any part may take place without the written permission of Cambridge University Press.
First published 2013
Printed in the United States of America
A catalog record for this publication is available from the British Library. Library of Congress Cataloging in Publication data
Lieuwen, Timothy C.
Gas turbine emissions / Timothy C. Lieuwen, Vigor Yang. pages cm. – (Cambridge aerospace series; 38) Includes bibliographical references and index. ISBN 978-0-521-76405-6 (hardback)
1. Gas-turbines – Environmental aspects. 2. Gas-turbines – Combustion. 3. Combustion gases – Environmental aspects. I. Yang, Vigor. II. Title. TJ778.L524 2013
621.43′3–dc23 2012051616 ISBN 978-0-521-76405-6 Hardback
Cambridge University Press has no responsibility for the persistence or accuracy of URLs for external or third-party Internet Web sites referred to in this publication and does not guarantee that any content on such Web sites is, or will remain, accurate or appropriate.
vii
List of Contributors page ix
Foreword by Alan H. Epstein xi
Preface xv
PART 1 OVERVIEW AND KEY ISSUES
1 Aero Gas Turbine Combustion: Metrics, Constraints, and
System Interactions . . . 3
Randal G. McKinney and James B. Hoke
2 Ground-Based Gas Turbine Combustion: Metrics, Constraints,
and System Interactions . . . 24
Vincent McDonell and Manfred Klein
3 Overview of Worldwide Aircraft Regulatory Framework . . . 81
Willard Dodds
4 Overview of Worldwide Ground-Based Regulatory Framework . . . 95
Manfred Klein
PART 2 FUNDAMENTALS AND MODELING: PRODUCTION AND CONTROL
5 Particulate Formation . . . 123
Meredith B. Colket III
6 Gaseous Aerosol Precursors . . . 154
Richard C. Miake-Lye
7 NOx and CO Formation and Control . . . 175
Ponnuthurai Gokulakrishnan and Michael S. Klassen 8 Emissions from Oxyfueled or High-Exhaust Gas
Recirculation Turbines . . . 209
PART 3 CASE STUDIES AND SPECIFIC TECHNOLOGIES: POLLUTANT TRENDS AND KEY DRIVERS
9 Partially Premixed and Premixed Aero Engine Combustors . . . 237
Christoph Hassa
10 Industrial Combustors: Conventional, Non-premixed, and
Dry Low Emissions (DLN) . . . 290
Thomas Sattelmayer, Adnan Eroglu, Michael Koenig, Werner Krebs, and Geoff Myers
ix Alberto Amato, Georgia Institute of Technology, Atlanta, Georgia, U.S.A.
Meredith B. Colket III, United Technologies Research Center, East Hartford, Connecticut, U.S.A.
Willard Dodds, General Electric Aviation Company, Cincinnati, Ohio, U.S.A. Alan H. Epstein, Pratt & Whitney Company, East Hartford, Connecticut, U.S.A. Adnan Eroglu, Alstom Power, Inc., Baden, Switzerland
Ponnuthurai Gokulakrishnan, Combustion Science & Engineering, Inc., Columbia, Maryland, U.S.A.
Christoph Hassa, German Aerospace Center, DLR, Linder Hoehe, Cologne, Germany
James B. Hoke, Pratt & Whitney Company, East Hartford, Connecticut, U.S.A. Michael S. Klassen, Combustion Science & Engineering, Inc., Columbia, Maryland,
U.S.A.
Manfred Klein, National Research Council, Ottawa, Ontario, Canada Michael Koenig, Siemens Energy Inc., Orlando, Florida, U.S.A.
Werner Krebs, Siemens AG, Fossil Power Generation Division, Muelheim an der Ruhr, Germany
Timothy C. Lieuwen, Georgia Institute of Technology, Atlanta, Georgia, U.S.A. Vincent McDonell, University of California, Irvine, California, U.S.A.
Randal G. McKinney, Pratt & Whitney Company, East Hartford, Connecticut, U.S.A.
Richard C. Miake-Lye, Aerodyne Research, Inc., Billerica, Massachusetts, U.S.A. Geoff Myers, GE Energy Company, Greenville, South Carolina, U.S.A.
Thomas Sattelmayer, Technische Universität München, Garching, München, Germany
xi When I first became interested in jet engines, smoke trails from the then ultramodern Boeing 707s were an arresting feature of that modern world. Ten years later, smoke was regulated and the U.S. Federal Aviation Administration had canceled the Boeing 2707 supersonic airliner program in the midst of growing environmental concerns. Back in the early 1960s, ground-based gas turbines were a very small business and concern for the environment was only minor. Over the five decades since the 707, the role of gas turbines in our society has greatly expanded, and con-cern regarding their emissions has grown even faster. Now, the electric power gen-eration gas turbine business has outgrown that of aircraft engines and emissions have become a market discriminator. Indeed, large fortunes have been won and lost on the basis of the emissions performance of land-based gas turbine engines. On the aero engine side, emissions performance is now featured in engine market-ing campaigns.
Combustion emissions might be thought an arcane topic. It is certainly complex. It is also of great importance to our society given the dominance of gas turbines for aircraft propulsion and power generation. There are three, basically indepen-dent, complicated problems associated with gas turbine emissions – the design of low-emissions combustors, the prediction of the effects of emissions on human health and the global environment, and the formulation of balanced and effec-tive policy and regulation. These challenges are important to three very different groups – technical folk, businesspeople, and policy makers and regulators. This book will be of interest to them all.
For the technical community, the science of how emissions are generated in a gas turbine combustor and their interactions with the atmosphere has always been a fascinating but challenging subject. The relatively recent concern for climate change has increased the complexity of the atmospheric science problem, especially for air-craft engines, from one mainly concerned with local air quality at low altitude to more complex interactions at the tropopause and in the stratosphere. During the last fifty years, design engineers have risen to the environmental challenge by realizing combustors with much lower emissions while at the same time significantly increas-ing reliability and life. One important aspect of combustor engineerincreas-ing, however, has
not changed over this time – we still do not have the technology needed to predict gas turbine emissions from first principles. The lack of first principles capabilities drives up product development costs and business risk.
Policy makers and regulators, who are not necessarily technical experts in the fields they regulate, face interesting challenges as well. These can be grouped into three general categories – technical, political, and diplomatic. Technical questions include, for example, consideration of currently unregulated emissions such as very small particulates and CO2, as well as the role uncertainty plays in resolving
con-flicting requirements such as NOx and CO2. Political challenges abound and include
issues such as how to best balance environmental protection with economic growth and how to balance local air quality with global climate change. Gas turbine emis-sions have also become a major diplomatic challenge. Aviation is the most interna-tional of endeavors, both in manufacture and operation. Most engines have parts and major subsections designed and manufactured in several countries. Aircraft take off and land in different countries thousands of times a day and so fall under the pur-view of more than one regulator. It is critical to the efficient operation of the world’s air transportation system that regulations be harmonized across the globe. This is the job of the International Civil Aviation Organization (ICAO), a branch of the United Nations with 189 member states. Getting 189 countries to agree on anything has never been easily or quickly achieved. The rise of climate change as a major world-wide issue with its attendant political and economic implications has only increased the complications of international rule making.
From the point of view of technical and policy folks, gas turbine combustor emissions bring fascinating challenges. For the business community, the fascination turns to dread. Why the dichotomy? The confluence of regulation and technical chal-lenge generates business uncertainty and risk, with financial penalties large enough to destroy a business. Manufacturers of ground-based engines are often contrac-tually responsible for the price of the electric power not produced if an engine is deficient. An engine that does not meet local air quality standards cannot be oper-ated, and may incur liabilities that dwarf the price of the engine. Manufacturers of aircraft engines face similar challenges; that is, until an engine meets emissions requirements, it will not be certified by regulatory authorities. Such engines cannot be legally shipped, and so the airplanes, which cost ten times more than the engine, cannot be delivered. Gas turbine development can cost up to two billion U.S. dollars, so long production runs are needed to amortize the cost. The business risk asso-ciated with emissions regulations is further amplified by the long-lived nature of the products. Engines typically have service lives of thirty years or more. Over this time span, emissions regulations usually change. Increased stringency can reduce the residual value of an engine, hinder sales, and even prohibit operation of engines in the field. Additional uncertainty is introduced by the degree to which regulations are not harmonized across political boundaries since niche markets cannot support high development costs. Thus, business planning for gas turbine emissions is a chal-lenge – and a concern.
These are hard problems. These are interesting problems. These are important problems at the confluence of engineering, regulation, and business. This book is the first to cover both the technical and regulatory aspects of gas turbine emissions. With chapters authored by some of the world’s experts in their respective fields, it has the breadth and depth to be of interest to all the stakeholders. It is valuable for experts in the field and informative for those just getting involved.
xv The development of clean, sustainable energy systems is one of the grand challenges of our time. Environmental and energy security concerns, coupled with growing energy demand, require us to increase, diversify, and optimize the use of energy sources while reducing the adverse environmental impacts of energy production, transmission, and use. In particular, we are confronted with four interacting issues: climate change, local air and water quality, energy supply, and energy security. Global warming has led to significant discussions about reductions of carbon dioxide emissions. Meanwhile, concerns about energy security and supplies for a growing uti-lization base are driving us to consider broader and more reliable energy resources. Finally, local air quality concerns are driving interest in other pollutants that lead to, for example, acid rain or photochemical smog, and that have additional implications for the management of power plant operations and emissions.
Gas turbines will continue to be an important combustion-based energy con-version device for many decades to come, for aircraft propulsion, ground-based power generation, and mechanical-drive applications. At present, gas turbines are a principal source of new power-generating capacity throughout the world, and the dominant source for air-breathing flight vehicles as well. Over the last decade, power generation from alternative sources, such as solar and wind, has significantly increased. Nevertheless, most projections indicate that the relative fraction of energy supplied by these sources will remain small, even several decades from now. These projections also indicate that gas-turbine-based combined cycle plants will continue to represent the majority of new power generation capacity. Moreover, as the supply of intermittent renewables grows, gas turbines will play an increasingly important role in stabilizing the electrical grid, where the supply and demand of electric power must match at every instant in time. The topic of gas turbine emis-sions, both traditional pollutants (NOx, CO, UHC, particulates) and CO2, is clearly
of significant interest.
In the aviation sector, emissions regulations continue to tighten. Climate change may lead the worldwide community to begin taxing carbon emissions for aircraft, and cloud formation associated with water vapor emissions continues to be an area of research. Particulate and NOx emissions can significantly influence local air quality
and can be controlled by appropriate combustor designs. Changes to engine cycles and pressure ratio to increase fuel efficiency, however, generally promote the pro-duction of emissions such as NOx, and, thus, maintaining safe, reliable, low-emission
aircraft engines is an increasingly important issue.
The present volume compiles the key scientific and technological knowledge associated with gas turbine emissions into a single authoritative source. The book consists of three parts. The first part provides an overview of major issues relat-ing to gas turbine combustion, includrelat-ing design approaches and constraints, at both the component and system levels, within the context of emissions. It also addresses approaches to meeting regulatory requirements. Important considerations for design optimization are discussed across all metrics of significance for gas turbine operation, including cost, safety, and reliability. The second part addresses funda-mental issues associated with pollutant formation, characterization, modeling, and prediction. This part treats aerosol soot precursors, soot, NOx, and CO. In addition, it
includes a chapter on emissions from gas turbines with significant levels of exhaust gas recirculation, or whose exhaust will be used for enhanced oil recovery or seques-tered in geologic formations; in these cases, the emissions-related concerns are quite different. The third part of this book presents case studies from manufacturers and technology developers, emphasizing the system-level and practical issues that must be addressed in developing different types of gas turbines that emit pollutants at acceptable levels. It is our hope that this book will provide a valuable resource to workers in this field, as a foundation both for scientists researching various aspects of gas turbine emissions and for technology developers who translate this funda-mental knowledge into products.
This book would not have been possible without assistance from many individ-uals. Peter Gordon encouraged this project and supported us throughout. Our assis-tant Glenda Duncan was a tremendous help . . . a great help in the numerous tasks associated with preparing the text. We owe a great debt of gratitude to Jong-Chan Kim for his enormous effort in editing figures and ensuring that the illustrations are of the highest quality. Dilip Sundaram deserves special appreciation for indexing the book.
3 1.1 introduction
The aircraft gas turbine engine is a complex machine using advanced technology from many engineering disciplines such as aerodynamics, materials science, combus-tion, mechanical design, and manufacturing engineering. In the very early days of gas turbines, the combustor section was frequently the most challenging (Golley, Whittle, and Gunston, 1987). Although the industry’s capability to design combus-tors has greatly improved, they remain an important design challenge.
This chapter will describe how the combustor interacts with the rest of the engine and flight vehicle by describing the relationship between attributes of the engine and the resulting requirements for the combustor. Emissions, a major engine performance characteristic that relies heavily on combustor design, will be introduced here with more detail found in following chapters. The wide range of operating conditions a combustor must meet as engine thrust varies, which is a major challenge for combustor design, will also be described. Last, the relationship between combustor exit temperature distribution and turbine section durability will be discussed.
1.2 Overview of selected aircraft and engine requirements and their relation to Combustor requirements
Aircraft gas turbine engines have been used in many different sizes of aircraft since their introduction in the 1940s. Small aircraft such as single-engine turboprops use engines of low shaft horsepower, which are of small physical size. Business jets and smaller passenger aircraft may use turbojets or turbofans with thrust in the range of several thousand pounds, usually with two engines per aircraft. The other extreme includes four-engine aircraft with turbofan engine thrusts as high as seventy thousand pounds and very large twin-engine aircraft with thrust per engine in the one hundred thousand pound class. These thrust designs are also physically very large, with fan diameters over 100 inches. In all of these applications, the engine system imposes a common set of requirements upon the combustor, as summarized in Table 1.1.
Constraints, and System Interactions
Randal G. McKinney and James B. HokeAs shown in Figure 1.1, these requirements are interdependent. Years of design and development within the industry have produced successive designs that improve upon all of the requirements concurrently. Although emissions are the focus of this text, each of these other requirements interacts with the emissions constraints and will be introduced briefly.
1.3 Combustor effects on engine Fuel Consumption
Gas turbine engines are Brayton cycle devices. An ideal version of such a cycle com-prises isentropic compression, addition of heat at constant pressure, and isentropic expansion through the turbine. Figure 1.2 is a simplified schematic of the effect of such a cycle on the pressures and temperatures in the engine. In real engines, all of the processes incur some loss of performance versus the ideal, manifested as a stagnation pressure loss in the combustor. Combustion systems incur pressure losses because of flow diffusion and turning, jet mixing, and Rayleigh losses during heat addition (Lefebvre and Ballal, 2010). However, at most power conditions, the effi-ciency with which the fuel chemical energy is converted into thermal energy is very high, typically greater than 99.9 percent. “Low” levels of 98 to 99.5 percent can be seen at low-power levels. In general, though, the combustion system is a small para-sitic effect on overall fuel consumption.
Table 1.1. Engine system-level requirements and supporting combustor characteristics Engine requirement Combustor characteristic
Optimize fuel consumption High combustion efficiency and low combustor pressure loss Meet emissions requirements Minimize emissions and smoke
Wide range of thrust Good combustion stability over entire operating range Ground and altitude starting Easy to ignite and propagate flame
Turbine durability Good combustor exit temperature distribution
Overhaul and repair cost Meet required combustor life by managing metal temperatures and stresses
Emissions
Stability Durability
Altitude relight
and starting Exit temperature
1.4 Fundamentals of emissions Formation
The pollutants emitted by engines that are of most interest are carbon monoxide (CO), unburned hydrocarbons (UHC), nitric oxides (NOx), and particulate matter
(PM or smoke). At low-power conditions, the inlet combustor pressure and temper-ature are relatively low, and reaction rates for kerosene-type fuels are low. Liquid fuel must be atomized, evaporated, and combusted, with sufficient residence time at high enough temperatures to convert the fuel into CO2. If the flow field permits
fuel vapor to exit the combustor without any reaction, or, if partially reacted to spe-cies of lower molecular weights, UHC will be present. If a portion of the flow field subjects a reacting mixture to a premature decrease in temperature via mixing with cold airstreams, these incomplete or quenched reactions lead to the production of CO, as detailed in Chapter 7.
At high power conditions, high air pressures and temperatures lead to fast reac-tions, with the result that CO and UHC are nearly zero. At these elevated tempera-tures, emissions of NOx and PM become more prevalent. NOx can be formed through
several processes, but the dominant pathway is thermal NOx, as described by the
extended Zeldovich mechanism, also detailed in Chapter 7. O = 2O N +O= NO+N N O = NO O N+OH = NO+ H 2 2 2 + +
The formation rate is exponentially related to the temperature in the flame, peaking near stoichiometric conditions. Thermal NOx emissions can be reduced by limiting
the time the flow spends at the high temperature and/or by reducing the maximum temperatures seen in the flame via stoichiometry control. Other NOx formation
Fan flow Core flow Fan Core Thrust Power to operate fan + some thrust
Compressor Core flow Combustor Turbine Pressure Temperature Pressure Temperature Figure 1.2. Summary of component characteristics.
mechanisms, such as NOx formed in the flame zone itself, are also described in
Chapter 7, but are negligible for aircraft engines.
When fuel-rich regions of the combustor flow exist at high pressures and tem-peratures, the formation of small particles of carbon can occur. These carbon par-ticles result from complex chemical processes and undergo multiple processes within the combustor such as surface growth, agglomeration, and oxidation prior to leaving the combustor, as detailed in Chapter 5. These particles pass through the turbine and exit the engine in the exhaust. When the concentration of the particles in the exhaust is high enough to be visible, as was often the case in early gas turbines, it is referred to as smoke or soot. Recently, the more general term particulate matter (PM) has been used to describe this emission. Modern engine smoke levels are invisible but still possess large quantities of very small soot particles and aerosol soot precursors (see Chapter 5) at the exhaust. Emerging research on the effect of PM on health and climate focuses more attention on measuring, modeling, and understanding the processes governing PM production.
These relationships between engine power conditions and emissions production lead to the behavior shown in Figure 1.3. As shown in the figure, levels of UHC and CO are highest at low power and drop quickly with increasing thrust. Conversely, NOx and PM increase with engine power and are typically maximized at maximum
power. Chapters 5 and 7 discuss these emissions formation processes in more detail.
1.5 effect of range of Thrust and starting Conditions on the Combustor
Flight gas turbine engines must provide a range of thrust and thrust response to power the aircraft mission. Missions vary depending on the aircraft application. Commercial aircraft and military transports have similar missions. Military fighters and other specialized aircraft can have very different missions because their use is not exclusively for the transport of payload between two points. Design require-ments are also very different for commercial and military applications. Military fighter engines are often designed for maximized thrust developed per unit weight so that the maneuverability of the aircraft is maximized. Military fighter engines also fly at a wide range of thrust throughout the flight envelope and must undergo fre-quent rapid thrust transients. Typically, commercial engines are designed for maxi-mum fuel efficiency per unit thrust. They fly at high altitude to achieve the best fuel efficiency and often do not have to endure the aggressive and numerous thrust transients of military fighter engines. Engine combustors must operate stably and efficiently over the full range of operating conditions, and must reliably relight if an engine shutdown or flameout should occur in flight.
1.5.1 Engine Mission Characteristics
A typical commercial engine mission consists of ground starting, taxi, takeoff, climb to altitude, cruise, deceleration to flight idle and descent, approach, touchdown,
thrust reverse, and taxi in. The extremes in combustor operating conditions drive the overall design approach. The combustor must meet performance, operability, and emissions metrics over the full range of operation. In order to do so, it must operate at the following extremes:
1. Minimum fuel-air ratio – This occurs during decelerations from high power to low power. Flight decelerations normally occur when descending from high alti-tude cruise and during approach throttle movements. They can also occur in emergencies. Minimum fuel-air ratio typically depends on the thrust decay rate, as the time response of the engine turbomachinery that governs the airflow is much longer than that of the fuel flow. Risk of weak extinction (flameout) is highest during decelerations.
2. Minimum operating temperatures and pressures – These occur at flight and ground idle conditions. Low pressure and temperature challenges combustion efficiency due to slower fuel vaporization and chemical kinetics.
3. High operating temperatures and pressures – These occur at takeoff, climb, thrust reverse, and cruise conditions. These conditions result in the bulk of NOx
formation and the most severe liner metal temperatures.
4. Ignition conditions – Ignition normally occurs on the ground but also occasion-ally in flight. Ignition is required at near surrounding ambient pressure and temperature. High altitude and extremely cold conditions are typically the most challenging to achieve ignition, flame propagation, and flame stabilization. These conditions lead to low temperature (−40ºF) and pressure (4 psia at 35,000 ft.) combustor inlet conditions.
Thus, the combustor design must meet the performance, emissions, and durability requirements at low- and high-power operations without compromising stability
0 10 20 30 40 50 60 0 20000 40000 60000 80000 100000 Thrust EI or Smok e Number HC CO NOx Smoke
and ignition. This requires favorable combustion fuel-air stoichiometry to meet requirements at all operating conditions. Two principal approaches have been used to achieve stoichiometry control in the industry. The first, fixed geometry without fuel staging, is the most common approach and is in the large majority of engines in service. These systems have all fuel injectors operating at all conditions. The second approach controls local fuel-air ratio through fuel staging. In these systems, not all fuel injectors operate at low power. This enables more active control of the local combustion fuel-air ratio.
1.5.2 Fixed-Geometry Rich-Quench-Lean (RQL) Combustors
Fixed-geometry combustors have been used in the gas turbine industry since its inception. Early designs used multiple cans in a circumferential array. The cans transitioned through an annular duct to the turbine (Figure 1.4a). Later designs used an annular duct geometry that allowed for reduced overall length and weight (Figure 1.4b). Annular combustors also have reduced liner surface area relative to can-annular combustors and therefore use less cooling. All designs use multiple fuel injectors to provide spray atomization and fuel-air mixing. Achieving good atomiza-tion and fuel-air mixing is critical for efficient combusatomiza-tion, low emissions, and good temperature uniformity into the turbine. Normally, the fuel is injected in the front end of the combustor and flow recirculation is created to provide a stabilization region for the combustion process. This is typically accomplished with air swirlers, which leads to vortex breakdown and flow recirculation. The stabilization zone promotes recirculation of hot product gases forward to the incoming fuel spray, thereby pro-viding a continuous ignition source and faster fuel droplet evaporation. Accelerated droplet evaporation is critical to high-efficiency combustion at low-power conditions, when low air inlet temperatures are insufficient to provide fast enough evaporation.
(a)
(b)
Figure 1.4. (a) Can-annular combustor (Pratt & Whitney JT8D-200); (b) RQL annular com-bustor (IAE V2500).
If continuous ignition is not provided at low power, the vaporization and reaction times can exceed the combustor residence time and flameout occurs.
The airflow distribution in a fixed-geometry combustor must be selected to achieve both low- and high-power performance requirements. Conditions at the combustor inlet vary significantly between low-power idle and high-power takeoff conditions. At idle, inlet temperature, pressure, and global fuel-air ratio are relatively low. At takeoff, the opposite is true (Figure 1.5). The operating temperatures and pressures are largely a function of the engine thermodynamic cycle; therefore the most significant parameter for the combustor designer to consider is the fuel-air ratio. Because air is introduced in stages along the length, the designer can tailor the airflow distribution to achieve key performance metrics. This creates a distri-bution in fuel-air ratio along the length of the combustor, leading to variations in local temperature as power level is adjusted. The difference in fuel-air ratio between high-power takeoff and low-power deceleration and idle conditions is critical because it determines the range of local fuel-air ratio in the front end of the combustor. For most modern gas turbines, the difference is large enough that the front end operates fuel rich (f/a > 0.068 for jet fuel) at takeoff conditions. Consequently, fixed-geometry combustors are referred to as rich-burning or rich-quench-lean (RQL) designs. This refers to the rich front-end fuel-air ratio that is diluted (quenched) by additional airflow in the downstream section of the combustor to reach the fuel-lean conditions at the combustor exit. The RQL-type design has several advantages and challenges, which are discussed later in this chapter.
As previously described, the challenges at low power are combustion efficiency and stability. The local fuel-air ratio in the RQL combustor front end at idle is designed to generate high recirculating gas temperatures (Figure 1.6). Therefore, the local fuel-air ratio should be near the stoichiometric (f/a ~.068 for jet fuel) fuel-air ratio to achieve high combustion efficiency. High combustion efficiency minimizes unburned hydrocarbon and carbon monoxide emissions that predominate at idle. Some increase in NOx emissions is generated by the hot front end, but emissions at
idle are not significant when compared to high power. By designing for near stoichio-metric conditions at idle, stability can be ensured at deceleration conditions, where minimum fuel-air ratio occurs. If the minimum fuel-air ratio during deceleration is
Pressure
Idle
Comb
ustor inlet condition
Thrust Take-off Idle Thrust Take-off Temperature
Steady state fuel-air ratio
Transient decel fuel-air ratio
not more than one-third below idle fuel-air ratio, the local fuel-air ratio in the front end is maintained above the weak extinction limit and flameout is avoided. Limiting of minimum deceleration fuel-air ratio is accomplished by the engine control and controls the maximum thrust decay rate for the engine transient.
At high-power conditions, the principal emissions challenges are NOx and
smoke. The RQL combustor axial temperature distribution at high power is depicted in Figure 1.7. The front end is fuel rich and consequently has lower flame tempera-tures. The dilution or quench region is characterized by peak gas temperatures as the fuel-rich mixture transitions through stoichiometric fuel-air ratio to the fuel-lean conditions at the combustor exit. In the front end, smoke is formed due to the com-bustion at fuel-rich conditions. Some of the smoke formed in the front end is oxidized in the high-temperature, oxygen-rich quench region. Thus, the front-end airflow level must be set with understanding of the formation and oxidation processes. The NOx
emissions are formed in both the front end and quench regions at high power. NOx
formation is exponentially a function of gas temperature, but also depends on the residence time at the local temperature. The highest rate of formation occurs in the quench region because it is the region where peak temperatures occur. However, time at peak temperature in the quench region is relatively short due to high mixing rates. In contrast, the formation of NOx in the front end is not negligible because it
has relatively longer residence time due to the flow recirculation. The presence of cooling flow in the front end also leads to NOx formation when it interacts with the
fuel-rich gas mixture.
CO HC
Threshold temperature
Gas residence time in combustor Compressor exit Turbine inlet CO consumed Mixing air “Quenches” reaction Combustion at near Φ = 1 Gas temperature
Recent advances have shown that substantial reductions in residence time and NOx can be achieved without compromising combustor stability and low-power
performance. Use of fuel injectors that produce small droplets uniformly dispersed within the airflow and rapid air jet mixing has enabled the residence time reduction. These advanced RQL combustor designs (Figure 1.8) have demonstrated NOx
reduc-tion of over 50 percent when compared to early annular combustors. They are also shorter and have lower volumes to reduce residence times. Reduced-length combus-tors are lighter and also have reduced surface area requiring film cooling. Advanced cooling schemes have been deployed to minimize NOx emissions and temperature
streaks into the turbines.
Overall, the RQL combustor has demonstrated excellent service history. Because it does not require complex controls to modulate fuel between injectors, it has dem-onstrated very good reliability. It also has inherently favorable stoichiometry for stability because the front-end airflow is minimized for NOx control purposes. The
front-end airflow is established as the minimum amount required for smoke control. If the fuel-air ratio range between high power and low power is large, the airflow required to control smoke can be larger than desirable for flame stability during decelerations. In these instances, the selected minimum transient fuel-air ratio must be raised to protect flight safety and reliability. In turn, raising the minimum fuel-air ratio limit increases the time required to decelerate the engine and can result in a safety risk during emergencies. If the deceleration time cannot be met with the revised minimum fuel-air ratio, then stability must be addressed by other means, such as by clustering fuel injectors provided with either more fuel or reduced airflow. This
NOx
Smoke
Threshold temperature
Gas residence time in combustor Compressor exit Turbine inlet Rapid NOx formation Rich combustion Φ ~ 2 Combustion at Φ = 1 as oxygen is added Gas temperature
zone remains above the weak extinction level locally and protects against flameout at worst-case deceleration conditions.
The critical challenges for the RQL design approach are smoke and liner dura-bility. As previously discussed, uniform mixing of fuel and airflow in the injectors can result in reduced smoke levels. When the fuel injector stoichiometry is fuel rich overall, the uniformity of the fuel-air distribution within the injector becomes crit-ical. A poorly mixed injector with a wide distribution will have regions that range from fuel lean to very fuel rich. The latter can produce the bulk of the smoke in the combustor. This occurs because the highest smoke generation often takes place in the most fuel-rich regions where there is sufficient residence time. Because the front end is designed with gas recirculation to achieve stability, these zones can produce smoke. Thus, the mixing and recirculation patterns are critical to smoke control.
The presence of fuel-rich and stoichiometric gases also introduces a liner dura-bility challenge. Because modern gas turbines operate at high temperatures and pressures, peak gas temperatures can exceed 4200ºF. Metallic liners have a practical temperature limit of <2000ºF for designs that meet typical durability life require-ments. Therefore, the liner must be cooled to prevent failure. Virtually all aero engine combustors feature hot side film cooling. Film cooling provides a protective layer of airflow on the liner surface that prevents convective heat transfer from high temper-ature gas. However, when fuel-rich gases in the front end interact with cooling, the film air provides oxidant for high-temperature combustion. Therefore, the presence of cooling air increases NOx formation in the forward portion of the combustor. In
the aft section of the combustor, cooling does not readily mix radially and there-fore decreases gas temperatures near the walls. The result is higher temperatures in Figure 1.8. Advanced RQL combustor (Pratt & Whitney PW1500 TALON X).
the midstream. The midstream peaked temperature profile also increases the hot-test streak temperature exiting the combustor. Therefore, aft cooling airflow affects the temperature profile and uniformity entering the turbines. Consequently, it is desirable to reduce cooling throughout the combustor. Improved liner designs have enhanced heat transfer efficiency, enabled emissions reductions, and strengthened turbine durability. The evolution of liner cooling designs will be discussed in a later section.
1.5.3 Fuel-Staged Combustors
Having discussed RQL approaches, we next consider fuel-staged combustors, which have seen limited use in commercial aircraft service. First-generation designs were introduced in the 1990s, and updated designs are scheduled for release in future engines. The overall approach in a fuel-staged combustor is to control the combus-tion stoichiometry through use of fuel injeccombus-tion in multiple locacombus-tions. Where the fixed-geometry RQL combustor injected fuel and air as uniformly as possible in the front end of the combustor, the staged combustor deliberately provides for mul-tiple airflow and fuel flow zones. The objective is to achieve fuel-lean combustion conditions for NOx reduction at high power. The fuel-lean conditions keep gas
tem-peratures low and virtually eliminate the highest temtem-peratures associated with stoi-chiometric conditions that exist in the RQL design.
The lack of fuel-rich and stoichiometric combustion creates two immediate ben-efits when compared to an RQL design. The first is that the fuel-lean flame produces very low levels of soot emissions. This means that carbon particulate emissions have the potential to be lower from fuel-staged combustors. Significant future efforts are required to characterize the full range of particulates emitted from both types of combustors (see Chapter 5). The second benefit is that the staged lean combustor requires less film cooling air for the liner. Because the lean reaction produces less soot, it is less luminous, resulting in reduced radiation heat load on the liner. Additionally, because the peak gas temperatures are lower, the convective heat loading is reduced. These factors allow for reduced liner cooling flux. This air can in turn be used for emissions control or to improve combustor exit temperature uniformity.
In a fuel-staged combustor, a large amount of airflow is mixed with the fuel at the injection point, so that fuel-lean conditions are achieved at high power with all fuel injectors flowing. The large amount of airflow and fuel-lean conditions pose a stability challenge at low power due to the fuel-air ratio lapse that occurs between high power and low power. To mitigate the stability risk, some of the fuel injectors are turned off at low power. This allows for the control of the combustion stoichi-ometry at idle to ensure high combustion efficiency. The zone that operates at low power is referred to as the pilot zone, and the high-power fuel injectors are referred to as the main zone. A difficult challenge for staged combustor designs is the tran-sition between operating with only the pilot at low power and all fuel injectors at high-power conditions. The transition often occurs at mid-power conditions such as approach thrust where fuel-air ratio, pressure, and temperature are not as high as
cruise, climb, and takeoff. Therefore, the local fuel-air ratio in the main stage may be unfavorable for efficient combustion at the lower temperatures and pressures. Consequently, more complex staging systems may be required where the main stage fuel injectors are turned on at different overall fuel-air ratios so that high efficiency is maintained. These fuel-air ratios are referred to as staging points. Initial designs were applied to engines with relatively low fuel-air lapse levels. These designs were operated with two fuel stages and a single staging point. More recent designs applied to engines with higher fuel-air ratio lapse may require more than a single fuel staging point to maintain staging efficiency.
Staging can also affect engine acceleration time from idle to higher power con-ditions. This is because of two factors. The first is the aforementioned combustion efficiency near the staging point. Lower efficiency results in reduced heat release and slower acceleration. The second is potential delay time to deliver fuel to the main fuel injectors. If some of the fuel flow is needed to fill fuel manifolds and fuel injectors, a delay occurs in the time to achieve combustion heat release and engine acceleration. Therefore, it is desirable to keep the main stage fuel system as filled as possible to achieve prompt acceleration when the throttle is moved. However, a full main stage fuel system is vulnerable to fuel coking. Fuel coking refers to the hard carbonaceous compounds formed in the internal passages of the fuel system when the fuel undergoes pyrolysis reactions when it is heated in the absence of air. Such compounds can block or reduce the flow of fuel through the main stage hardware. Coking is most common inside the fuel injectors because they are exposed to the high temperatures inside the diffuser casing. In the extreme, coking can limit thrust by limiting fuel flow. Most modern engines have idle air temperatures near or above the level at which significant coking occurs (400ºF). This air is in contact with the main stage fuel injectors containing the stagnant fuel. To prevent fuel coking, cool-ing and insulation features must be incorporated to prevent fuel from contactcool-ing passage walls over the critical temperature for coking. Some designs use the pilot fuel flow to cool the stagnant main fuel injectors. Other possibilities include using air pressure to purge the fuel from the most vulnerable areas.
A final challenge to the fuel-staged combustor designer is combustion instability. Combustion instability refers to temporal fluctuations in the heat release. Such fluc-tuations can be attributed to several mechanisms, typically involving excitation of natural fluid mechanic instabilities in the flow or fuel-air ratio oscillations. In the extreme, instabilities can damage hardware and result in engine damage and failure. All combustors have risk of instability, but staged lean combustors have been more prone to them. It is unclear if this tendency is related to differences in acoustic driving resulting from heat release distribution differences or to changes in acoustic damping as the combustor is modified for lean-staged operation (Lieuwen and Yang, 2005).
1.5.4 Ignition and Engine Starting
Gas turbine combustors are required to ignite on the ground and in flight. Ignition in flight is rare because it occurs after unplanned engine shutdown. The combustor
should ignite promptly after the fuel is turned on and provide efficient combus-tion to accelerate the engine to idle power. Delayed ignicombus-tion can cause excess fuel accumulation in the combustor and increased pressure pulses at light off. Increased pressure pulses can result in compressor stall that prevents engine acceleration to idle. On the ground and at low-speed flight conditions, the engine rotors are turned with a starter to provide airflow to the combustor for ignition and combustion. At higher-speed flight conditions, the ram airflow turns the rotor in a process referred to as windmilling. Ignition energy is typically delivered with a spark igniter. At least two igniters are placed in the typical annular combustion chamber to provide redun-dancy in the event of a failure. The spark produces plasma sufficient to initiate the combustion reaction. The ignited reactants must then be transported to an area where the reaction can stabilize and propagate to the other fuel injectors in the combustor. The same features that provide flame stability at idle and deceleration conditions are relied upon at sub-idle starting operations. The pressure at light off is usually near the outside ambient pressure because the rotors are not producing sig-nificant work. However, at higher flight speeds, the total pressure is typically slightly higher than ambient due to the stagnation effect. Temperatures at ignition are highly dependent on the thermal state of the engine. For the first start of the day on the ground, temperatures are usually only slightly higher than ambient. Altitude relight temperatures are highly dependent on the amount of time the engine has been shut down. For quick relight attempts less than a minute after shutdown, temperature at the combustor inlet can be greater than 200ºF. If the engine is shut down and windmilling for thirty minutes or longer, the air temperature is closer to the outside ambient.
Most commercial aircraft must meet requirements for both ground and altitude starting. The ground starting requirements include a range of ambient temperatures and airport altitudes. Typical ground starting ambient temperature requirements vary between −40 and 120ºF. Airport altitude requirements typically range between sea level and eight thousand feet. Altitude relight requirements are typically expressed on a flight envelope (Figure 1.9). There is a high-speed windmilling envelope and a lower-speed starter assisted envelope. The maximum altitude required for air start-ing depends on the aircraft. Commercial airliners normally require altitude relight capability of at least twenty-five thousand to thirty thousand feet. Business jets often require capability at thirty-five thousand feet because of their higher cruising altitude.
At the highest altitudes and in extreme cold, combustor ignition conditions can be very challenging. Pressures of less than five psia and temperatures below 0ºF are typical for an engine that windmills until cold. These conditions inhibit the atomiza-tion of fuel and vaporizaatomiza-tion of droplets. Low temperature and pressure also slow the reaction kinetics that promotes stabilization and propagation of flame. Therefore, design of the combustor should provide for three key features that enable ignition: a good fuel spray, a favorable airflow velocity, and the proper spark igniter location.
Small fuel droplets are critical to the formation of vapor necessary for igni-tion. Two types of fuel injectors are typically used: pressure atomizing and airblast
atomizing (Figure 1.10). The former uses high pressure to push fuel through a small orifice to generate the spray. The fuel can also be swirled prior to passing through the orifice to provide angular momentum that produces a spray cone. Airblast atom-ization uses the energy of the airflow to produce the spray. The fuel is typically deliv-ered to a cylindrical surface between two swirling airstreams. The cylindrical surface develops a thin film of fuel as a result of the action of the swirling inner airstream. As the thin film reaches the tip of the cylinder, the shear between the two airstreams atomizes the film into a spray. Airblast atomizer performance degrades as the air pressure drop decreases and should not be used if insufficient airflow is available to atomize the fuel. This occurs when windmill ignition is attempted at very low airspeed and when insufficient starter torque limits rotor speed in assisted starts. Often, airblast fuel injection systems will be supplemented with pressure atomizers in the locations where the igniters are placed. Such injectors that incorporate both pressure atomizing and airblast features are referred to as hybrid or duplex injectors. Increased fuel flow at the igniter locations can help achieve ignition. This additional nonuniform fuel flow is provided by upstream valves and is usually only present at low-power settings. Manifolds that deliver the fuel must be designed in such a way as to achieve the desired distribution of fuel.
Successful ignition also requires a favorable velocity in the region near the spark plug and the stabilization zone. Because most combustors are swirl stabilized, the recirculation of flow can be used to transport the ignited spark kernel to the sta-bilization zone. However, even a properly designed stasta-bilization zone can result in poor ignition characteristics. This problem can stem from two factors. The first is local velocities that are too high to sustain the reaction surrounding the spark. This results in the convective heat loss from the reaction kernel exceeding the heat
Mach number Altitude 0 0.2 0.4 0.6 0.8 1.0 1.2 0 10 20 30 40 Successful ignition No ignition x103ft
released by the reaction, quenching the reaction. This situation is created when there is insufficient volume and cross-sectional area in the stabilization zone for the quan-tity of airflow present. Therefore, care must be taken to ensure that local velocity does not exceed the flame propagation speed at light-off conditions. The other cause is improper igniter placement. If the igniter is placed in an area with flow direc-tion away from the recirculadirec-tion zone, the reacdirec-tion kernel can be carried out of the back end of the combustor. Igniters also must be placed in an area where the fuel spray provides sufficient local fuel-air ratio to achieve ignition. Conditions at igni-tion are often relatively high in overall fuel-air ratio because of low airflow levels, but wide variations exist in local fuel-air ratio. As a result, spark igniters are often placed at the downstream edge of the flow recirculation so that it receives robust fuel-air mixture from the conical fuel spray, but also provides reverse flow direction for stabilization.
1.6 Turbine and Combustor durability Considerations
The combustor has a significant impact on turbine durability and consequently impacts engine performance. The temperature distribution at the combustor exit affects the cooling airflow required to protect the airfoils and platforms in the tur-bines. This cooling airflow, in turn, reduces the engine performance by diverting flow from the mainstream so that it is not used to produce work. The cooling also causes mixing losses if it is introduced as low momentum film on the airfoils. Combustor film cooling is required for aero engines because the metallic liners are exposed to the high-temperature combustion process. Combustor cooling itself does not affect engine performance because it is added upstream of the turbines. However,
Air Air Fuel Primary fuel Secondary fuel
Airblast fuel injector
Duplex fuel injector in air swirler
combustor cooling does reduce the amount of airflow available to control emissions and mix out temperature streaks. Therefore, it is desirable to minimize the amount of cooling flow used.
As previously mentioned, the combustor exit temperature distribution has a large impact on the amount of turbine cooling required and, thus, the engine per-formance. Combustor exit temperature quality is normally described in terms of the radial average temperature profile and the hottest streak intensity. These are referred to as radial profile factor and pattern factor, respectively. They are typically described as nondimensional parameters:
radial profile factor = (Tra−T )/(Te e−T )i
where Tra is the average temperature at a given radial position, Ti is inlet temperature,
and Te is the mass averaged exit temperature. The pattern factor is given by:
pattern factor = (Tstr−T )/(Te e−T )i
where Tstr is the maximum temperature anywhere in the combustor exit annulus,
commonly called the streak temperature. The radial profile factor is determined as a function of radial position at the entrance to the turbine. The maximum pattern fac-tor occurs at only one spatial position in the combusfac-tor exhaust (Figure 1.11). Often this location is the result of random hardware variation and cannot be assumed repeatable from engine to engine. However, the radial distribution of pattern factor is also of interest to the turbine designer. Turbine static hardware (vanes and outer air seals) are impacted by the local gas temperatures while rotating blades are impacted predominantly by the radial average temperature profile because they rotate too fast for metal temperatures to respond to local effects. Thus, the static hardware cooling level is often set to protect against the hottest pattern factor streak, even though it occurs in only one place. It is useful to know the radial distribution of pattern factor so that static hardware cooling can be distributed more in the core region, where the hottest streak is likely to occur, and less near the walls, where the average tempera-tures are lower because of the effects of combustor cooling.
To achieve the target radial temperature profile and low pattern factor, the designer must control the mixing of fuel and air in the combustor. To achieve the low-est possible pattern factor, the designer would premix all of the airflow with the fuel at the combustor front end. This would produce a flat, uniform temperature profile at the combustor exit. However, considering liner cooling requirements, operability, and radial profile requirements, this approach is not practical for aero engines. In practical aero engine designs, cooling air is injected into the combustor in a way such that it provides a protective film near the combustor walls. As such, it generally does not mix readily with the other airstreams and the fuel in the combustor. Liner cooling is there-fore not effective at controlling pattern factor, but can be effective at providing cooler radial average profile near the inner and outer walls. The airflow not used for liner cooling is used to control radial profile shape and pattern factor. In RQL designs, the bulk of the non-cooling airflow enters through air jets downstream of the front end in the liner walls. In fuel-staged designs, most of the airflow is incorporated into the
fuel injection swirlers at the front end of the combustor. Therefore, the mixing pro-cesses to achieve uniform exhaust temperatures are quite different. In swirling flow mixers, multiple airstreams are often used to create shear layers that promote mix-ing. Fuel is injected into the airstreams so that it is dispersed and mixed with the air. Fuel injection is usually accomplished with jets, thin films, or pressure sprays. Swirling airstreams may be co-swirling or counter-swirling. Designers have successfully used both approaches. Counter-swirling airstreams produce the highest mixing rates, but result in low net swirl if not designed with unequal flow quantities. These principles are applied to both RQL and lean-staged designs because good fuel injector and swirler mixing is required for both design approaches.
Jet mixing is dependent on the arrangement, size, and upstream conditions influ-encing the jets. The penetration depth of a jet, Y, is proportional to the jet diameter, dj, and the square root of the momentum ratio, J:
Y d J~ j where the momentum ratio J is given by:
J = ( U )/( U )ρj j2 ρg g2
Uj and Ug denote the jet and cross flow velocities, respectively. Specific correlations
depend on the geometry of the duct and the arrangement of the jets (Lefebvre and Ballal, 2010).
The penetration of the jet can be controlled by the sizing, pressure loss, and upstream flow quantity. Efficient mixing of upstream gases also requires jet spacing dense enough to mix within the combustor length allocated. Therefore, the arrange-ment and size of the jets are critical to the spatial delivery of airflow to the critical regions where it is needed to mix temperature streaks and provide target radial aver-age temperatures.
Staged lean combustors may only use air jets to control radial profile shape, because the lean well-mixed front end delivers good pattern factor at high-power
10
0 20 30 40 50 60 70 80 90 100
Radial span (%) Radial profile factor Radial pattern factor
Maximum pattern factor
Temperature factor
conditions. For example, a row of smaller holes in the aft end of the combustor can effectively cool the inner portion of the radial distribution because of their limited penetration. At low-power conditions, staged lean combustors often have worse uni-formity because of the reduced number of fuel injectors operating. RQL combustors are more dependent on the jet mixing to deliver both the radial profile and pattern factor targets. Researchers have conducted numerous experimental studies to deter-mine the optimum jet arrangement for mixing flow in a duct. Holdeman found that the optimum arrangement was given by:
H = ( / )p h J
where p is the hole pitch, h is the duct height, and H is the characteristic parameter (Holdeman, 1993). For unopposed holes in a duct, the optimum value for H is 5 (Figure 1.12). This finding was for a uniform axial cross flow, but is often a good arrangement from which to begin optimization. For real combustors where the upstream flow is swirling, computational fluid dynamics analysis is useful to refine the distribution. Rig testing is required to determine the maximum pattern factor because CFD calculations typically cover a single fuel injector sector and thus do not provide random effects.
Combustion imposes two different types of heat loads on the liner. The first is radiation from the flame to the surfaces. The second is the convective effect of hot gases contacting the liner cooling films. The convective load can result in film tem-perature above metal temtem-perature in some areas of the combustor. The radiation flux is given by:
q = 0.5(1+ )( Tr w g g T ) 4
g w 4
ε ε − α
where εw is the wall emissivity, εg is the gaseous emissivity, αg is the gas
absorptiv-ity at the wall temperature Tw, and Tg is the radiating gas temperature. The gaseous
emissivity is dependent on the flame luminosity and gas temperature, which in turn depends on the combustion stoichiometry. Rich combustion tends to produce highly luminous soot. Therefore, the forward section of a RQL combustor tends to have more radiant heat load than a staged lean combustor, which produces very little soot. The midsection of a RQL combustor produces peak gas temperatures as the stoichiometry transitions from fuel rich to fuel lean. This region of a RQL also has higher gaseous emissivity than a lean-staged combustor does.
P
h
Convective heat load on the liner is dependent on the local gas temperature and velocity and its interaction with the cooling film. The effectiveness of the film is crit-ical to maintenance of acceptable metal temperatures because the film temperature is a key driver in the heat flux:
q = h(Tc film−Tmetal)
where qc is the convective flux, h is the convective heat transfer coefficient, Tfilm is
the film temperature, and Tmetal is the liner surface metal temperature. Film
tempera-ture is dependent on the local gas temperatempera-ture and film effectiveness. Film effec-tiveness depends on the nature of the film (slot flow, discrete holes, etc.) and the ratio of the cooling momentum to the mainstream flow momentum. The momentum ratio is referred to as the blowing parameter. In zones where the mainstream flow momentum is low and the blowing parameter is high, the cooling film effectiveness is reduced. This occurs most commonly in the front end of the combustor. Note that for equivalent film effectiveness, the RQL design will have higher film temperature because it reaches higher peak gas temperature than the staged lean combustor.
Cooling strategies for the outside of the liner wall vary widely with design (Figure 1.13). Outside cooling is important because it balances the hot side heat flux. With high backside cooling effectiveness, higher hot side heat flux can be tolerated. Initial liners made use of simple louvers that created slot films on the liner hot side and had minimal heat transfer on the outside of the liner wall. The louver length was determined by the distance that film effectiveness could be maintained. Louvered designs later evolved to incorporate more effective backside cooling strategies for the louver lips. All continuous ring louver liners fail because of thermal fatigue cracks. The cracks result from high thermal stresses on the full ring hoop. As oper-ating temperatures increased, more effective liner designs were required because
Combustor liner history
Sheet metal liner
Double pass liner
Floatwall liner
Film Louver lips
Louver lips with backside cooling Film
