Handbook of Turbomachinery Second Edition Revised - Ventech!

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tur—turbine w—wall ?—mainstream (free-stream) condition TURBINE COOLING AS A KEY TECHNOLOGY FOR ADVANCEMENT OF GAS TURBINE ENGINES The performance of a gas turbine engine is strongly influenced by turbine inlet temperature, and substantial performance increases can be achieved with the maximum allowable turbine inlet temperature. From a performance standpoint, stoichiometric combustion with turbine inlet temperature around 2,000 8C (3,650 8F) would be a thermodynamic ideal because no work would be ‘‘wasted’’ to compress air needed to dilute combustion products. Therefore, the current industry trend pushes the turbine inlet temperature closer to the fuel stoichiometric temperature, especially for military engines. However, allowable component metal temperatures, even for the most advanced alloys and manufacturing processes, cannot exceed 930–980 8C (1,700–1,800 8F). To operate at gas temperatures well above this limit, a highly efficient component cooling system is required. Advancement in cooling is a major means to achieve higher turbine inlet temperatures, leading to improved performance and life of the turbine. Heat transfer is thus a very important design element for all parts of a modern gas turbine, but especially the combustor and turbine sections. In discussing the state-ofthe-art cooling design of the hot section it is important to keep in mind that the turbine designer is constantly under extreme pressures of development schedule, affordability, durability, and various other interdisciplinary constraints, all these strongly affecting the selection of a cooling design. Cooling Challenges for Constantly Increasing Gas Temperatures and Compressor Pressure Ratio Progress in modern high-efficiency and high specific power gas turbine engines is typically measured by the increases in operating temperature and overall compressor pressure ratio. The most common simple cycle engines with higher pressure ratios and correspondingly increased gas temperatures can achieve greater power for the same size and weight and better overall engine fuel efficiency. Engines with recuperators do not benefit thermodynamically from high compressor pressure ratio. Advanced alloys for turbine airfoils can safely operate at metal temperatures below 980 8C (1,800 8F) and alloys for discs and stationary structures are limited to 700 8C Copyright © 2003 Marcel Dekker, Inc.
(1,300 8F). But modern gas turbines operate at turbine inlet temperatures that are well above these limits. There is also a significant difference in the operating temperature between advanced aircraft turbines and industrial turbines. This is the result of fundamental differences in the required life, weight, fuel/air quality, and restrictions on emissions. For advanced aero engines turbine rotor inlet temperatures (TRIT) close to 1,650 8C (3,000 8F) and compressor pressure ratios around 40:1 are becoming a reality. A high specific power, which is a primary objective for this type of engine, is usually obtained along with high efficiency. Such extreme operating conditions inherently require frequent engine inspections and continuous health monitoring. For industrial engines the leading requirement is long-term durability without frequent inspections and overhauls. Typically, major industrial turbine components are expected to endure at least 30,000 hours between overhauls, with potential to be repairable such that the engine life can be extended to 100,000 hours. This compares to aircraft turbine component lives of only a few thousand hours. This factor, as well as the usual requirement of a compressor discharge pressure to be below the available gas pipeline fuel supply pressure, leads to a moderately high turbine rotor inlet temperature. The TRIT limit for a modern industrial gas turbine is established presently in the range of 1,260– 1,3708C (2,300–2,5008F). Figure 1 illustrates the recent history of continuously rising TRIT and compressor pressure ratios. This rising temperature trend results in the operating gas temperatures significantly exceeding acceptable material capability limits and requires the application of cooling in various engine hot section components, particularly in the components exposed to the high-temperature environment. Relatively cold air from the compressor discharge and, in some cases, from intermediate compressor stages is the traditional source to cool turbine components. After performing the cooling task, this air is discharged into the main stream. The cooling air discharged within each particular cooled stage practically does not perform any work on this stage before it is accelerated to the mainstream velocity. This leads to significant losses in engine performance. In summary, disadvantages of the open air cooling system include a relatively low cooling effectiveness, significant losses of work required to compress cooling air, and mixing losses that reduce aerodynamic efficiency of the turbine. The main advantage of the open air cooling system making it the most common for gas turbines is its simplicity, compared to a close loop cooling system. With compressed air pressure ratios for modern aero engines exceeding 30:1 and reaching 40:1, the temperature of the compressor discharge air approaches 650 8C (1,200 8F). This creates a significant difficulty in using this air to cool the turbine discs, stators, and vicinity of Copyright © 2003 Marcel Dekker, Inc.
Page 1 and 2:
D E K K E R Handbook of Turbomachin
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MECHANICAL ENGINEERING A Series of
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61 Computer-Aided Simulation in Rai
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127 Designing for Product Sound Qua
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Preface to the Second Edition The o
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Preface to the Second Edition Prefa
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Contributors Nathan G. Adams The Bo
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1 Introduction Earl Logan, Jr.*, an
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HISTORICAL BACKGROUND Earl Logan, J
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MECHANICAL AND THERMAL DESIGN CONSI
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Figure 1 Cross-section showing the
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Figure 2 Turboprop engine, TPE 331-
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Figure 4 Auxiliary power unit, APU1
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Durability. The mechanical and ther
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with time during flight. Such a typ
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experience and existing similar des
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adverse pressure gradient against w
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and its attachments. A substantial
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Figure 9 Typical high-pressure turb
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and the measurement of local heat-t
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due to cyclic operation. In additio
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the steady and transient conditions
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Table 3(a) Steps in a Three-Dimensi
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Figure 16 Schematic Goodman diagram
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occur due to particles of unburned
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pressures at critical points. These
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1. Ceramics are subject to ‘‘in
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with Transient Tests and Surface Co
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Overall engine performance goals tr
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Much of the complexity of turbomach
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End-wall boundary layers can also h
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pressure at the wake centerline. Th
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the same vaneless diffuser and shro
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from internal cooling passages, thr
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Further downstream, the wake was ev
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to its original position. This unst
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cooling film was found to trip the
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shroud static pressure and the exit
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stage in the design process allows
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Flow Physics Modeling Modeling of t
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passage to the next, assuming that
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The next step in modeling accuracy
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tend to be large, due to the necess
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Navier-Stokes equations are time-ma
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To satisfy these requirements, it i
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are discussed by a number of author
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source code, and to assign each seg
Page 95 and 96: selected locations in the flow path
Page 97 and 98: of the component designers must be
Page 99 and 100: Trends in Flow Modeling Capabilitie
Page 101 and 102: optimize the flow behavior. The eff
Page 103 and 104: 20. D. Eckardt, ‘‘Flow Field An
Page 105 and 106: 49. K. R. Kirtley, ‘‘An Algebra
Page 107 and 108: INTRODUCTION 3 Turbine Gas-Path Hea
Page 109 and 110: Figure 1 Maximum turbine rotor inle
Page 111 and 112: Figure 2 Intersecting S1 and S2 sur
Page 113 and 114: determination. Nothing in the simpl
Page 115 and 116: e well matched with the engine. The
Page 117 and 118: case different forms of the popular
Page 119 and 120: volume will be equal to the accumul
Page 121 and 122: Figure 9 Nusselt number vs. vane su
Page 123 and 124: Furthermore, to keep computer time
Page 125 and 126: Figure 12 Comparison of three turbu
Page 127 and 128: and small and cooled and uncooled t
Page 129 and 130: Figure 15 Miniature heat flux gauge
Page 131 and 132: Figure 17 Time resolved heat transf
Page 133 and 134: Figure 19 Surface pressure changes
Page 135 and 136: NONUNIFORM INLET FLOW All gas turbi
Page 137 and 138: Figure 22 Typical distributions of
Page 139 and 140: 4. Better instrumentation for heat-
Page 141 and 142: 24. R. H. Ni, ‘‘A Multiple Grid
Page 143 and 144: 4 Selection of a Gas Turbine Coolin
Page 145: Re L ¼ rVL=m—Reynolds number bas
Page 149 and 150: during engine operation are compati
Page 151 and 152: Figure 3 Airfoil cooling techniques
Page 153 and 154: expected to increase this limit to
Page 155 and 156: 1. Reduce the effect of main-stream
Page 157 and 158: P? is total inlet pressure. Tc=T? i
Page 159 and 160: thereby increasing both the skin-fr
Page 161 and 162: After four decades of advancement i
Page 163 and 164: correspondingly high coolant-side t
Page 165 and 166: component. Use of a hot cascade for
Page 167 and 168: nozzle vanes is limited by their ox
Page 169 and 170: Figure 9 Typical turbine hot sectio
Page 171 and 172: important factor to consider for tr
Page 173 and 174: Figure 10 Evolution of nozzle geome
Page 175 and 176: the exit of the combustor section f
Page 177 and 178: Figure 12 Inlet turbulence effect o
Page 179 and 180: in heat transfer as the Reynolds nu
Page 181 and 182: easonably well using correlations e
Page 183 and 184: in the section on blade cooling tha
Page 185 and 186: mass flow rate (e.g., M ¼ 1 for a
Page 187 and 188: holes. Cooling hole pattern is an i
Page 189 and 190: coolant flow with the main stream,
Page 191 and 192: above equation assumes a constant f
Page 193 and 194: Figure 15 Definition of film-coolin
Page 195 and 196: edistribution of the external heat-
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Three-Dimensional Effects The three
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In order to achieve the bulk metal
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coefficient is given by Nud ¼ hd=k
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Where G ¼ 2:24ðw=tÞ 0:1 ðReeÞ
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pressure drop. The 608/458 V-shaped
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Some of the studies have found that
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of the cooling air and the turning
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Figure 20 Impingement geometry defi
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correlated by Nu a d ¼ 0:63ðGd=m
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tangentially to the inner wall, beh
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The effectiveness of a suction surf
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vortex structure. Their studies sho
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Figure 23 Effect of preswirler pres
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component to the cooling air that c
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axial gap. A static pressure variat
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Weight, cost, and complexity constr
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Figure 28 Effectiveness of liner co
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thickness. A problem this poses is
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portion of the hole has been shown
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unacceptable engine performance pen
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analysis, in order to provide accur
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Figure 30 Application of PSP for fi
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thermocouple positioned close to th
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spray-painted black (Hallcrest, BB-
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Figure 32 Schematics of a hot casca
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the turbine static temperature (rot
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Figure 33 Experimental rig for inve
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their color when exposed to higher
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margins defined within each discipl
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them. The larger gas-path divergenc
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Figure 36 Blade-cooling selection d
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8. B. Barry, Turbine Blade Cooling:
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40. A. K. Tolpadi and M. E. Crawfor
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73. R. S. Abhari, ‘‘Comparison
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106. J. M. Owen and R. H. Rogers, F
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138. I. Egorov, G. Kretinin, I. Les
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integrity (stress levels) will have
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Figure 1 Instant entropy contours o
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Aerodynamic Interaction (Unsteady L
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Figure 3 Pitchwise time-averaged en
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The above considerations are all fo
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lades, the wavelength of the distur
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forcing and damping, and hence the
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frequency and its higher harmonics
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The two-cell pattern with a relativ
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Figure 9 Typical blade flutter boun
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field is generated to satisfy the t
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Figure 11 Instant static pressure c
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Figure 12 Calculated instantaneous
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Figure 14 Time-space static pressur
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one period Tp (neglecting change of
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It should be noted that the effects
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There are some further points to no
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Figure 17 Flow around an NACA-65 ai
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(11) for a simple case]: Where WA
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Regardless whether loosely coupled
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Several methods have been developed
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significant difference is introduce
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losses, but it may affect to a mini
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when we refine a mesh in an unstead
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8. W. S. Barankiewicz and M. D. Hat
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6 Fundamentals of Compressor Design
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Figure 3 AlliedSignal 331 turboprop
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each of the selected radii. The pro
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components of gas velocity upstream
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Observe that each of the curves for
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The conservation of angular momentu
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allowable static pressure ratio of
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Figure 7 Boundary layers and wakes
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component that is perpendicular to
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Figure 10 (a) Deflection of flow of
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increase in the downstream pressure
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Mach number at a compressor outlet
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Figure 12 Sketch of flow surface an
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useful shapes and orientations of b
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where G and oa are constant along a
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There is a good reason for this sit
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Projected Shape in Meridional Plane
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inherent flow range of the impeller
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compressors. Multistage centrifugal
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turbines. In these designs the high
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Corrections for Streamline Curvatur
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Preparation of Maps Once the losses
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that circumferential gradients meas
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Nastran, which was developed by NAS
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tests. One should take pains to pro
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e predicted in advance. Total press
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Cv Coefficient of heat at constant
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15. S. Lieblein, J. Basic Eng., Tra
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INTRODUCTION 7 Fundamentals of Turb
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Figure 1 Cross sections of generic
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espectively. Radial-inflow turbines
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Figure 4 Typical rotor blade shapes
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Figure 6 The expansion process acro
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the average for the gas flowing thr
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etween the velocity diagram and the
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gas dynamics relation p 00 p ¼ T 0
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Figure 9 Variations in turbine velo
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pressure, inlet losses are usually
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diffuser recovery, defined as Rp ¼
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different turbines easier, dimensio
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where and y ¼ T 0 in TSTD d ¼ p0
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sizing exercises where the details
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where Rth is the Reynolds number ba
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order to determine the overall turb
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The classical approach to satisfyin
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clearance. Since tip clearance repr
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turbine with an effective diffuser,
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atio ðp 0 in =pdisÞ is 3. The sta
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Automation of Calculations and Trad
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and for the rotor Zrotor ¼ ð0:1Þ
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the energy extracted and the rotor
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RADIAL-INFLOW TURBINE SIZING Differ
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interpolations from existing design
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Glassman [1], the optimum ratio of
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design fault in the radial-inflow t
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[27]. In order to avoid manufacturi
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The rotor exit critical velocity is
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The hub-to-tip radius ratio at the
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REFERENCES 1. A. J. Glassman (ed.),
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8 Steam Turbines Thomas H. McCloske
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Figure 3 Reciprocating steam engine
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This process began in the 1920s, al
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First Law of Thermodynamics The fir
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the amount of fuel burned and the p
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Figure 7 Rankine cycle temperature-
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The Mollier diagram is quite useful
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Figure 10 Theoretical Rankine cycle
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stages of the turbine is at a highe
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calculate the relevant thermodynami
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Figure 14 Steam leakage losses for
Page 459 and 460:
can reveal nozzle and/or blade eros
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Figure 17 Enlarged portion of the M
Page 463 and 464:
Traditionally, pressures have been
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Figure 21 Wetness thermodynamic los
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Figure 23 Radial steam turbine flow
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Figure 25 Percentage stage reaction
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The power output of the stage can b
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the LP turbines is split into paral
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Turbine Inlet Two stages in particu
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of 20%, which would cause excessive
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Figure 30 Typical low-pressure stea
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ductility and poor toughness, since
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‘‘tramp’’ elements to minim
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End Seals End seals or packing glan
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Drains Condensate can form during s
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Figure 32 Steam turbine blade roots
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Figure 34 Cross-section view of thr
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Figure 36 Interblade shroud and Tie
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Where The influence of notch sensit
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Similarly, LP turbine blades are su
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those made of steel. Note, however,
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esult, significant attention must b
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Figure 38 Centrifugal stresses on a
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Figure 39 Centrifugal bending stres
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onto the rotating blade. These can
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Vortex shedding from blade trailing
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Figure 42 Torsional coupled vibrati
Page 513 and 514:
Figure 44 Stall flutter of low-pres
Page 515 and 516:
Start-Stop Transients/Overspeeds La
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(4) ellipiticity, such as caused by
Page 519 and 520:
Figure 47 Campbell diagram (frequen
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Furthermore, in practice each manuf
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Figure 49 Modal diameters of a low-
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Figure 51 Three-dimensional computa
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Figure 52 Converging wedge and resu
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value as the operating value of uv/
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Lubrication Supply System Design De
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Figure 56 Steam turbine oil system
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The main oil reservoir is located b
Page 537 and 538:
Although a light oil does provide a
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temperatures are suspected as reaso
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There is no fixed or absolute value
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pivots incorporated in this bearing
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covers and doors. Pipe scale, rust,
Page 547 and 548:
Figure 63 Mechanical hydraulic cont
Page 549 and 550:
Figure 64 Steam turbine electro-hyd
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Figure 65 Steam turbine electro-hyd
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generator rotor moment of inertia h
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the turbine against overspeed. This
Page 557 and 558:
an additional electronic overspeed
Page 559 and 560:
Figure 68 Steam turbine protective
Page 561 and 562:
3. Need for better understanding of
Page 563 and 564:
REFERENCES 1. J. C. Zink, ‘‘Ste
Page 565 and 566:
No. TWDPS-1, The American Society o
Page 567 and 568:
58. A. V. Sarlashkar and T. C. T. L
Page 569 and 570:
81. IEEE Guide for Abnormal Frequen
Page 571 and 572:
Atlanta, GA, Oct. 18-22, 1992, PWR-
Page 573 and 574:
Performance Enhancement Program, he
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of the design process toward achiev
Page 577 and 578:
procedure capable of addressing mul
Page 579 and 580:
to achieve a specific target pressu
Page 581 and 582:
objective was to minimize the downs
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more closely represents the largest
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temperature by forming a thin film
Page 587 and 588:
such as the one described here. For
Page 589 and 590:
k is the thermal conductivity of th
Page 591 and 592:
through finite differences. Mathema
Page 593 and 594:
Figure 4 Comparison of blade cross-
Page 595 and 596:
integration of multiple disciplines
Page 597 and 598:
17. S. S. Talya, ‘‘Multidiscipl
Page 599 and 600:
INTRODUCTION 10 Rotordynamic Consid
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Figure 1 (c) AlliedSignal 331 turbo
Page 603 and 604:
Figure 2 Rotor system displacements
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Figure 3 Elliptic lateral whirl. av
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vibration (either forward or backwa
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Figure 6 The Laval-Jeffcott rotor.
Page 611 and 612:
are 1808 out of phase (i.e., the be
Page 613 and 614:
this graph is quite simple; however
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included on these types of graphs b
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on the shaft is zero. Thus, the ste
Page 619 and 620:
are uðtÞ ¼e zeott uo cos ott þ
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satisfies the inequality O < ot 1
Page 623 and 624:
Figure 16 Laval-Jeffcott rotor tran
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Figure 18 Whirl speed map: Laval-Je
Page 627 and 628:
Figure 20 Laval-Jeffcott rotor: ste
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Figure 22 Typical precessional mode
Page 631 and 632:
Figure 24 Examples of flexible disk
Page 633 and 634:
ackward modes, however, involve con
Page 635 and 636:
Figure 29 Steady unbalance response
Page 637 and 638:
assist in attenuating high vibratio
Page 639 and 640:
properly sizing the damper, a highl
Page 641 and 642:
Figure 31 Gas turbine schematic and
Page 643 and 644:
Flexible Disk If the frequency of a
Page 645 and 646:
motion for each subsegment to form
Page 647 and 648:
Figure 35 Whirl speed map—straddl
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Figure 37 Rotor-bearing-support str
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DESIGN STRATEGIES AND PROCEDURES Ro
Page 653 and 654:
Figure 38 Straddle mounted rotor. (
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Figure 39 Unbalance response: strad
Page 657 and 658:
Figure 40 Double overhung rotor—d
Page 659 and 660:
earing supports are reduced. The th
Page 661 and 662:
and also increases the critical spe
Page 663 and 664:
Examples of structure modes couplin
Page 665 and 666:
REFERENCES 1. D. Childs, Turbomachi
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section is intended to place today
Page 669 and 670:
pressure-fed propellant systems. Hi
Page 671 and 672:
normally be considered a very small
Page 673 and 674:
increase of required suction perfor
Page 675 and 676:
displaced by the F-1 oxidizer pump
Page 677 and 678:
exposure may be subjected to imping
Page 679 and 680:
the exhaust duct. Sized to create s
Page 681 and 682:
excess of 6,000 psi for consumption
Page 683 and 684:
lifetime. With very few exceptions,
Page 685 and 686:
using hydrocarbon fuel combustion p
Page 687 and 688:
can yield noticeable savings in req
Page 689 and 690:
Figure 3 Rocket engine system schem
Page 691 and 692:
Figure 6 Staged combustion system,
Page 693 and 694:
each stage. The term C0 represents
Page 695 and 696:
Figure 10 Some representative turbo
Page 697 and 698:
emembered that the second (and any
Page 699 and 700:
the greatest demands on turbopump p
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momentum exchange between working f
Page 703 and 704:
Figure 13 Arrangement of various tu
Page 705 and 706:
pressure levels. As long as combust
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Figure 14 Approxmate values of pump
Page 709 and 710:
liquid rocket engine systems, and e
Page 711 and 712:
above becomes a factor. This statem
Page 713 and 714:
We can then state (correctly) that,
Page 715 and 716:
Figure 15c Typical 2-D centrifugal
Page 717 and 718:
have seen that the flow rate throug
Page 719 and 720:
somewhat smaller magnitude of vecto
Page 721 and 722:
general indicator. Nss will give us
Page 723 and 724:
U ¼ rotor tangential velocity. By
Page 725 and 726:
and pump flow derived from the engi
Page 727 and 728:
In addition to shaft seals, shaft b
Page 729 and 730:
should be understood that the optim
Page 731 and 732:
stable operation is assured. In som
Page 733 and 734:
From Eq. (16) an impeller radial ou
Page 735 and 736:
Figure 18 Examples of centrifugal p
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tional to clearance distance) with
Page 739 and 740:
I would like to conclude this secti
Page 741 and 742:
generally result in lower pump hydr
Page 743 and 744:
Figure 15h Cross-section of a typic
Page 745 and 746:
solutions of Eq. (7) for both stage
Page 747 and 748:
Figure 15k H-Q characteristics for
Page 749 and 750:
Figure 15l Typical inducer configur
Page 751 and 752:
camber) is proportionally lower. Th
Page 753 and 754:
Figure 22 Single-stage impulse turb
Page 755 and 756:
the velocity ratio K, but in somewh
Page 757 and 758:
~V and ~n ¼ fluid velocity and con
Page 759 and 760:
Fig. 25. The nature of the flow in
Page 761 and 762:
Although shrouded centrifugal impel
Page 763 and 764:
is a parasitic device, whose flow r
Page 765 and 766:
apidly than does the inducer thrust
Page 767 and 768:
Figure 30 LH2 turbopump rotor mecha
Page 769 and 770:
eplacement of a failed bearing with
Page 771 and 772:
shall see, contemporary rocket engi
Page 773 and 774:
needs of the aircraft gas turbine a
Page 775 and 776:
was increased due to engine upratin
Page 777 and 778:
12 Turbomachinery Performance Testi
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istics. The more accurate the perfo
Page 781 and 782:
Inlet Distortion Most turbomachiner
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Gas Thermodynamic Properties Turbom
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ought to rest isentropically. Total
Page 787 and 788:
include not only the pressure chang
Page 789 and 790:
The circumferential placement of th
Page 791 and 792:
simply arithmetically averaging the
Page 793 and 794:
usually mounted in radial immersion
Page 795 and 796:
Figure 12(b) Radial flow angle prof
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Figure 15 Wake rake total pressure
Page 799 and 800:
Figure 16 Blade-tip pressure traces
Page 801 and 802:
Figure 18 Typical test rig schemati
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INSTRUMENTATION DESIGN CONSIDERATIO
Page 805 and 806:
Dynamic Pressure. Dynamic pressure
Page 807 and 808:
thermocouples. With care in constru
Page 809 and 810:
Figure 24 Typical two-dimensional f
Page 811 and 812:
Thermocouple lead wires should be i
Page 813 and 814:
Corrected Flow The flow rates used
Page 815 and 816:
pressure ratio [15]. It can be calc
Page 817 and 818:
SYMBOLS A Cross-section area Cd Dis
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INTRODUCTION 13 Automotive Supercha
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Figure 2 A Roots-type supercharger
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The supercharger was in routine use
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Table 1 Superchargers Versus Turboc
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1960s, turbocharging was used by US
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Roots Blower The Roots-type blower
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speeds. The sound is caused by the
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Figure 8b How the Ro-charger operat
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The vanes in many designs are preve
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Figure 11a Key components of the G-
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Figure 11c G-Lader supercharger is
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Figure 12b Schematic of centrifugal
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Figure 14b Components of the typica
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Turbine Design The turbine housing
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inertia of the wheel assembly. Mini
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Figure 18 (a) Turbocharger compress
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Theoretically this efficiency is: o
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The overall turbine efficiency incl
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adjust the spring force to vary the
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Figure 22b Two vanes are used to pr
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However, the temperature has to be
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air can reduce carbon monoxide and
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Figure 25 (a) Schematic of the COMP
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this plus the fact that the exhaust
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14 Tesla Turbomachinery Warren Rice
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press at the time of the invention
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can be made to simplify the equatio
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comparison of the results of variou
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y composing them of nested cones ra
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26. E. Bakke, ‘‘Theoretical and
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61. M. Piesche, ‘‘Investigation
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15 Hydraulic Turbines V. Dakshina M
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Table 1 Variables of Interest in Tu
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Figure 1 Variation of impeller shap
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Figure 2 Flow and velocity componen
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Figure 3 Schematic diagram of Pelto
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shown in Fig. 5(see KadambiandPrasa
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classified according to the directi
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The manufacturers of hydraulic turb
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The direction of the relative veloc
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Table 3 Variation of Some Quantitie
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denoted by hd. Thus the efficiency
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Figure 14 Draft tube setting. same
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REFERENCES 1. J. J. Fritz, Small an
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Handbook of Turbomachinery Second Edition Revised - Ventech!

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