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`.JEPPESEN®
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`A BOEING COMPANY
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`T
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`X T 8 0 0 K
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`UTC-2011.001
`
`GE v. UTC
`Trial IPR2016-01301
`
`

`

`ii
`
`Library of Congress Cataloging-in-Publication Number 91-14478
`
`10001375-006
`
`© 1997, 2001 , 2002, 2010 by Charles E. Otis, M. Ed. and Peter A. Vosbury, M. Ed.
`All Rights Reserved
`
`ISBN 0-88487-553-9
`9-780884-875536
`
`UTC-2011.002
`
`

`

`•
`
`CHAPTER 3
`Turbine Engine Design and Construction
`•
`TURBINE ENGINE ENTRANCE DUCTS
`The flight inlet duct (the point where atmospheric air
`enters the aircraft) is normally considered a part of the
`airframe, not of the engine. Nevertheless, it is usually
`identified as engine station number one. Understanding
`the function of the flight inlet duct and its importance to
`engine performance is a necessary part of any discussion
`on gas turbine engine design and construction. Figure 3-1
`shows a variety of aircraft flight inlet ducts at different
`locations.
`
`The design features of gas turbine engines are var(cid:173)
`ied, and engines in the same power classification often
`seem to have little or no resemblance to each other.
`Determining which design is optimal for a particular
`application or why an engine looks the way it does is
`sometimes difficult for the following reasons:
`
`• Many engine designs are proprietary. and manufactur(cid:173)
`ers are often reluctant to provide specific information
`about their engines.
`
`• Many designs that do not appear to be the optimal are
`in fact best-suited for the engine and the aircraft on
`which it is intended to be installed. A compromise of
`designs for operation over a wide range of altitudes
`and power factors is common.
`
`• Many designs depend on the prior experience of the
`manufacturer. Manufacturers might remain committed
`to their proven developments rather than changing to
`newer ones.
`
`PRINCIPLES OF OPERATION
`The flight inlet duct of a turbine engine must furnish
`a uniform supply of air to the compressor for the engine
`to enjoy stall-free compressor performance. The flight
`inlet duct must also create as little drag as possible. Even
`a smaU discontinuity of airflow can cause significant
`efficiency loss, as well as many otherwise inexplicable
`engine performance problems. To ensure that the flight
`inlet duct delivers air with minimum turbulence. it mu t
`
`Figure 3-1. Common engine Inlet location
`
`UTC-2011.003
`
`

`

`3-2
`
`Aircraft Gas Turbine Powerplants
`
`be maintained in as close to new condttJOn a~ pos~.ible . lf
`repairs to this component are neces-.ary. expertly in tailed
`flush patches are mandatory to pre\ent drag. Moreover,
`the use of an inlet cover should be u ed to promote· clean(cid:173)
`liness and to prevent corrosion and abra~ion.
`SUBSONIC FLIGHT INLET DUCTS
`Subsonic flight inlet ducts, such as lhose fo1Und on
`business and commercial jet aircraft. are of fixed geom(cid:173)
`etry and have a divergent shape. A diverging duct pro(cid:173)
`gressively increases in diameter from front to back. ac;
`shown in Figure 3-2. This duct i.~ sometimes refe:rrell to
`as an inlet di ffuser because of its effect on pressure. Air
`enters the aerodynamically contoured inlet at n,mbient
`pressure and <.tarts to diffuse. arriving at the compressor
`at a ~lightl y increased ~tatic pressure. U ually. the air i~
`permitteu to diffuse (incre~ in )Italic pres~ure) in the
`front portion of the duct and to progress ar a fairly con(cid:173)
`stant pressure past the engine inlet fairing (also called the
`inlet center body) to the compre!.sor. In thjs man1ner, the
`engine reccivcc; its air witb mjnimal turbulence and at a
`more uniform pressure.
`As the aircraft approachec; 1ts desired cruising speed.
`the increased inlet pressure adds significantly to the
`mass airflow. At cruising speed, the compressor reaches
`its aerodynamic design point and produces its optimum
`compression and best fuel economy. At rhis point. the
`fljght inlet, compressor, combustor, turbine. and tailptpe
`are designed to work in concert with each other. lf any
`section does not match the other-. (for example. hecau~c
`of damage, contan1ioarion, or ambient conditions). engine
`performance will be affected.
`The turbofan inlet is similar in design to the I!Urbojet
`inlet, except that only a portion of the air it p~rovidcs
`to the fan is <;ent into lhe core of the engine, with the
`remainder being accelerated out of the fan discharge duct.
`Figure 3-3 :.hows two common airflow arrangement\.
`Figure 3-3A i!> a fuJl duct design uttlized on luw- anll
`medium-bypass engines, and the other. Figure ·\-3B, is
`the short duct design of a high-bypass turbofan. The long
`ducting configuration in Figure 3-3A reduces surface
`
`Figure 3-3A. Turbofan low- and medium-bypass ratio
`Figure 3-38. Turbofan high-bypass ratio
`
`drag of the fan discharge air and enhances thrust Many
`of the older high-bypas~ engines cannot take advantage
`of lhis drag reduction concept because of the exces&ive
`weight associated with the wide diameter of a long duct.
`With the emergence of new lightweight materials and
`designs. however, newer generation engines can take
`advantage of thi!. drag reduclion concept.
`
`RAM PRESSURE RECOVERY
`As mentioned in Chapter 2. when a gas turbine engtne
`is operated in place on tbe ground. it has a negative pres(cid:173)
`sure within its inlet because of the high velocity airllow.
`As the aircraft moves forward, a condition known as ram
`pressure recovery takes place. This is the point at which
`pressure in-.ide the inlet rerum<; to ambient value. Ram
`compression is the re!:ult of ram velocity and diffusion of
`the airflow. That is. as air spreads out radially, it slows
`down axially, nnd pressure increases accordingly.
`Th~ aircraft inlet. while stationary, will not generally
`achieve I 00 percent duct recovery. which means the air
`pressure leaving the inlet will be lower tban rhe air pres(cid:173)
`sure entering. If ambient pressure is 14.7 pounds per
`square inch ab)lolute (p.s.i.a.), pressure at the compressor
`inlet will be slightly less than 14.7 p.c;.i.a. However, as
`the aircraft moves forward on the ground for takeoff, ram
`
`(B)
`
`INLET
`CENTER ---1~~
`BODY
`
`SUBSONIC
`GAS FLOW
`
`REDUCED VELOCITY
`INCREASED STATIC
`[ )
`PRESSURE
`
`Figure 3-2A. Divergent subsonic Inlet duct
`Figure 3·28. Divergent duct effect on airflow
`
`UTC-2011.004
`
`

`

`Turbine Engine Design and Construction
`
`3-3
`
`compression occurs and the pressure at the compressor
`inlet will eventually rerum to ambient value. Thi!> point
`is generally reached in the average inlet duct at an aircraft
`speed of Mach 0.1 to Mach 0.2.
`In Figure 3-4, note the gauge readings changing from
`a negative to a positive value as the aircraft goes from
`ground static condition to flight condition. As the aircraft
`moves faster in flight, the inlet produce more and more
`ram compression. The engine rakes advantage of this
`conrution by a corresponrung increase in compressor
`pressure ratio, creating greater thrust with less fuel.
`
`Use the following formula to calculate ram compres(cid:173)
`sion (ram pressure ratio) at any flight mach number:
`
`Whe re: y(gamma) = 1.4 (speci fie heat)
`M = Mach number
`r. 1. 2 = Constants
`~ = Pt2
`Pam
`Ps
`Cp
`_L =
`y-1
`Cp - Cv
`
`(See Appendix 8)
`
`For example, consider a businesl> jet that is traveling
`at Mach 0.8 f1ight speed at an altitude of 31.000 feet. Use
`the formula to calculate the pressure ratio of engine inlet
`pressure to ambient pressure.
`
`1.4
`
`:! = [1·c·~- 1
`
`X 0.82 JlA-1
`~ = [1 + (0.2 X 0.64 )]3·5
`Ps
`~ = 1.524
`Ps
`
`This calculation makes obvioul> that very high-speed
`aircraft create
`ignificant inlet ram compression. For
`example, the supersonic Concorde airliner, at a cruise
`speed of Mach 2.2. produced a ram compression of L0.7
`to I.
`The reasoning underlying the use <Jf the Pt!Ps formula
`is as follows:
`ln actuality, the Pt!Ps formula represents Pt2/Pam,
`which is the pre~sure total at the engine face divided by
`pressure ambient. to give the inlet compression ratio. The
`formula becomes applicable because Pt2 is e:.sentially the
`same value as Ptl , pressure total m the lip of the flight
`inlet, which is expressed in the formula as Pt.
`
`Figure 3-4. Ram pressure recovery
`
`GAUGES (+)
`
`A<:. a ir moves down the inlet. assuming I 00 percent
`efficiency, the total pressure at which the air entered does
`not change. The changes occur in the static pressure and
`ram pressure components of total pressure-as the air
`move!- Uuough the inlet diffuser toward the engine, tatic
`pressure increases and ram pre sure decreases, but total
`pressure remains the same.
`When density in pounds per cubic foot and velocity in
`feet per second are known. calculate the total presl.ure in
`the inlet as follows:
`Total pressure
`
`Ram pressure (Q) + Static pressure (p)
`a+ P
`
`1/2 p V 2 (flow density~) + p
`g
`
`Pl
`
`Pt
`
`Where:
`
`Q = l/2 p V2 (flow density~) g
`p = lb~ft ·•t gravity constant
`V = ftlsec (in the inlet)
`p = lbs/in 2 (l>tatic pressure)
`
`For exllmple, consider lin ai1p/ane at a11 altirude of
`25,000 feet that is cruising at 550 mile~ per hour (806
`feet per second). The static pressure in rhe .flight inlet is
`5.-15../ pound!. per square inch and the densiry. derived
`from a standard altitude chart (see AppendLr 7). is
`.034267 pound per cubic foot.
`1. What i:. the total pressure (Pt) in the inlet?
`2. What is the inlet pressure ratio (Cr)'?
`Solution to que!>tion I:
`Q = 1/2 p V2
`(flow density~)
`
`Q = 112 ( 0.0342671bs/:
`32.16 ft/sec
`
`3 J( 806ft )
`
`sec
`
`2
`
`Q = 1/2
`
`0.034267 lbs.
`tt?
`
`X
`
`sec.2
`8062 tt.2
`2
`32.16 ft.
`sec.
`(Continued on next page)
`
`X
`
`UTC-2011.005
`
`

`

`3-4
`
`Aircraft Gas Turbine Powerplants
`
`Q = 346 lbs/ft 2
`Q = 2.40 lbs/in 2
`Pt = Q + p
`Pt = 2.40 + 5.454
`7.851bs/in2 (psi)
`
`PI
`
`Solution to question 2:
`Cr = 7.85 + 5.454
`Cr = 1.44 to 1 (inlet pressure ratio)
`
`SUPERSONIC INLET DUCTS
`All !>Upersonic aircraft require a convergent-divergent
`inlet duct-either tixed or variable. A supersonic trans(cid:173)
`pon., for example. is configured with an inlet that slows
`the airflow to subsonic speed at the face of the engine.
`regardless of aircraft speed. Subsonic airflow into the
`compressor is required if the rotating airfoils are to
`remain free of shock wave accumulation. which is detri(cid:173)
`mental to the compression proces~.
`To vary the geometry. or shape. of the inlet a mov(cid:173)
`able restrictor is often employed to form a convergent(cid:173)
`divergent (C-D) configuration of variable proportion.
`The C-D shaped duct becomes necessary in reducing
`supersonic airflow to subsonic speeds. Bear in mind that
`at subsonic flow rates, air flowing in a duct acts as an
`incompressible Liquid, but at supersonic tlow rates air is
`compressed to the point of creating the familiar shock
`wave phenomenon.
`Figure 3-5 depicts a fLxed geometry (nonadjusrable)
`C-D duct in which supersonic airflow is slowed by air
`compression and shock formation at its throat area. After
`its speed is reduced to Mach-I. the airflow enters the sub(cid:173)
`sonic diffuser section where velocity is further reduced
`and its pressure increased before entering the engine
`compressor. Some military aircraft. designed to fly as
`fast as Mach-2, use this type of inlet. However. a fixed
`geometry inlet is not always operationally feasible for a
`variety of reasons associated with the stagnation pressure
`effect of supersonic inlets, but a full discussion of that is
`beyond the scope of this book.
`
`CONVERGENT
`SUPERSONIC
`DIFFUSER
`
`....
`
`An inlet shock wave is similar to shock waves com(cid:173)
`mon to aircraft wings and other airfoils. A shock wave
`is an accumulation of sound wave energy, or pressure,
`developed when the wave, trying to move away from an
`objecl, is held stationary by the oncoming tlow of air.
`One useful aspect of a shock wave is that airOow pass(cid:173)
`ing through the high-pressure shock region !)lows down.
`[Figure 3-6]
`
`OBLIQUE SHOCK WAVES
`WHEN Mn IS HIGHER THAN 1
`
`SUPERSONIC AIRFLOW
`NORMAL SHOCK
`ATMn z 1.0
`
`Figure 3-6. Shock wave formation
`
`Figure 3-7 c;;hows an example of a supersonic diffuser
`type of inlet. which provides a mean~ o f creating both a
`shock wave formation to reduce air velocity and a vari-
`
`SHOC~ ~' ::lAM';
`--+
`WAVES
`,, \
`SUPERSONIC ~ ~
`CONDITION
`DUMP
`SPILL VALVE OPEN
`VALVE
`TO VENT EXCESS
`OPEN
`AIRFLOW
`
`EDGE RETRACTED-THROAT AREA INCREASED
`
`~ ___.
`SUBSONIC ~-.....
`COHDmON
`DUMPVALVE
`SPILL VALVE OPEN
`USED AS SCOOP
`TO PREVENT
`TO INCREASE
`TURBULENCE
`AIRFLOW
`
`Figure 3·5. Supersonic convergent-divergent Inlet
`
`Figure 3-7. Supersonic airplane movable wedge Inlet
`
`UTC-2011.006
`
`

`

`Turbine Engine Design and Construction
`
`3-5
`
`able convergent-divergent shape to meet various flight
`conditions from takeoff to cruise. Air velociry drops to
`approximately Mach 0.8 in back of the final shock wave.
`and then to approximately Mach 0.5 by diffusion (spread(cid:173)
`ing out radially). In high-speed flight (for example. Mach
`2 to Mach 2.5). multiple shock waves form as the air
`!lows through the inlet duct. The shock waves that form
`in the front part of the inlet are known as oblique, and
`although they slow the air velocity down, the velocity
`is still supersonic. Where the inlet changes shape from
`convergent to divergent (in an area known as the throat
`or waist), a final normal shock wave forms, wl:tich low(cid:173)
`ers the air velocity to subsonic. The diverging part of the
`inlet duct then acts as a typical subsonic diffuser.
`The movable wedge design shown in Figure 3-7
`shows the various wedge positions and functions of con(cid:173)
`vergence, divergence. and shock wave formation. It also
`shows a spill valve used to control inlet air at high speed
`versus low speed. Many high-performance aircraft have
`either an excess or a deficiency of mass flow in vari(cid:173)
`ous engine operating environments, and they require an
`onboard computer that monitors the inlet conditions and
`manages the position of the inlet's movable components.
`
`BELLMOUTH COMPRESSOR INLETS
`Bellmouth inlets, which are used primarily on heli(cid:173)
`copters, are converging in shape and provide very thin
`boundary layers and ~.:orrespondingl y low losses in pres(cid:173)
`sure. This type of inlet produces a large drag factor, but
`its low speed drag is outweighed by the high degree of
`aerodynamic efficiency it provides.
`During calibration, engines on ground test stands aJso
`use a bellmouth inlet, sometimes fitted with an anti(cid:173)
`ingestion screen. Duct loss is so slight in this design that
`it is considered to be zero. Engine performance data, such
`as engine trimming for rated thrust, are obtained while
`using a bellmouth compressor inlet. [Figure 3-8J
`Aerodynamic efficiency and duct loss are illustrated
`in Figure 3-9. Notice that a rounded leading edge [Figure
`
`Figure 3-B. Bellmouth compressor inlet (with screen)
`
`ROUNDED
`\'\ ~JRIFICE
`STAllCOR~
`0 0
`LOW APPROACH
`--ii;i;i;i;ia• 'O
`o _Q_
`VELOCITY (V~a)
`STREAM AREA
`100% OF
`ORIFICE AREA
`
`(A)
`
`(B)
`
`Figure 3-9A. low-velocity entry flow through round-edge
`orifice
`Figure 3-98 . Low-velocity entry flow through sharp-edge
`orifice
`
`3-9A] enables the airstream to use the total inlet cross(cid:173)
`section wl:tile a sharp-edged orifice [Figure 3-9B] greatly
`reduces the effective diameter.
`
`COMPRESSOR INLET SCREENS, SAND,
`ANDICE SEPARATORS
`The use of compressor inlet screens is usually limited
`to rotorcraft, turboprops, and ground turbine installations.
`This might seem peculiar given the appetite of all gas
`turbines for debris such as nuts, bolts, stones, and so on.
`Screens have been tried in high subsonic flight engines
`in the past, but icing and screen fatigue failure cau ed so
`many maintenance problems that the use of inlet screens
`has for the most part been abandoned.
`
`When aircraft are fitted wit.h inlet screens for protec(cid:173)
`tion against foreign object ingestion, they can be located
`intema!Jy or externally at either the inlet duct or the
`engine compressor inlet. [Figure 3- I OA and 3-1 OB J
`Often these separators are removable at the discretion
`of the operator. fn the sand separator shown in F igure
`3-lO.B, inlet suction causes particles of sand and other
`small debris to be directed by centrifugal loading into a
`sediment trap.
`
`Some sand and ice separators employ a movable vane
`that can be extended into the inlet airstream. This causes
`a sudden rum in the engine inlet air, and sand or ice par(cid:173)
`ticles continue out undeflected because of their greater
`momentum. The movable vane in this installation is oper(cid:173)
`ated by the pilot through a contml handle in the cockpit.
`[Figure 3-J OCl
`
`UTC-2011.007
`
`

`

`3-6
`
`Aircraft Gas Turbine Powerplants
`
`(C)
`
`ENGINE AIR INLET SCREEN
`
`Figure 3-10A. Helicopter inlet
`Figure 3-108 . Sand and Ice separator
`Figure 3-10C. Sand and Ice separator (Inertial)
`
`ENGINE INLETS AND GROUND EFFECT
`Some gas turbine engine inlets have a tendency to
`fonn a vortex between the ground and the flight inlet.
`The suction creating the vortex is strong enough to lift
`water and debris, such as sand. small stones, nuts. and
`bolts. from the ground and direct it inw the engine. cau!>(cid:173)
`ing serious compressor erosion or damage. [Figure 3-1 11
`This is e pecially true on wing pod-installed engines tl1at
`are mounted with low ground clearance, as seen on many
`of the newer high-bypass turbofan-powered aircraft. To
`alleviate this problem, some inlets have been redesigned
`to be slightly out-of-round or flattened at the bottom to
`reduce ground effect.
`Earlier aircraft used a vortex dissipator (also known as
`a blow-away jet}. To dissipate the vortex. a small jet of
`
`Figure 3-11. Water vortex during test cell run up
`
`compressor discharge air is directed at the ground under
`the inlet from a di~charge ooZt.le located in the lower
`part of the engine flight cowl. The system is generally
`activated by a landing gear switch. which opens a valve
`in the line between the engine compressor bleed port and
`the dissipator noz7le whenever the engine is operating
`and weight is on the main landing gear. [Figure 3-121
`
`ACCESSORY SECTION
`The engine-driven external gearbox is the main unit
`of the accessory section. Accessory units essential to the
`operation of the engine, such as the fuel pump. oil pump,
`fuel control. and starter. and components such as hydrau(cid:173)
`lic p1.1mps and generators. are mounted on the main acces(cid:173)
`sory gearbox. The accessory section is typically located
`below the compressor section or at the rear of the engme.
`[Figure 3-13A and 8]
`The gearbox is often driven by a radial drive shaft that
`connects the main or auxiliary gearbox to a bevel gear
`
`Figure 3-12. Vortex dissipater
`
`UTC-2011.008
`
`

`

`Turbine Engine Design and Construction
`
`3-7
`
`(A)
`
`Figure 3-1 3A. Accessory gearbox location- six o'clock position
`Figure 3-138. Accessory gearbox location-rear
`
`system driven by the main rotor shaft. On some installa(cid:173)
`tions, an auxiliary gearbox is employed to drive the main
`gearbox. This arrangement permits the gearbox to be
`placed so that the envelope size of the engine can be kept
`to a minimum. [Figure 3-14AI The main gearbox can also
`be located at the front or rear of the engine if the inlet
`or exhaust locations accommodate it. ln rare instances.
`the main gearbox is located at the top of the engine in
`the area of the compressor. Some models of the CFM56
`turbofan engine have the gearbox mounted on the side.
`Each accessory and component drive pad is designt:d
`to provide a gear reduction from compressor speed, as
`needed. [Figures 3- 148 ]
`The system of seal drain tubes shown in Figure 3-15A
`connects to each drive pad and is normally routed to the
`bottom of the engine cowling to drain away fluids that
`present a lire hatard. The leakage is generally minute
`and present:. little problem al. it leaves the drain point
`into the atmosphere. Each individual drive pad is a point
`of potential oil leakage. The e fluids are classed as waste
`Ouids and include fuel from the fuel control or fuel
`pump, engine oil from the main oil pump or scavenge
`oil pump. and hydraulic oil from the hydraulic pump. Oil
`can also leak from !he gearbox through drive shaft seals
`of the components mentioned above and into the drain
`tubes. The aiJowable leakage rate of the various fluids
`is listed in the manufacturers' maintenance instructions
`and is typically in the range of 5 to 20 drops per minute.
`depending on the source of the leak.
`Figure 3-1 SB shows tl1e speed of each accessory. The
`direction that each accessory rotates is determined by
`tbe number of teeth on the gears and whether they mesh
`outside-to-<>utside or i n~ide-to-outside. When a large
`gear drive~ a '>mall gear. the ~mall one turns faste r and
`experience!> a Loss in torque. If the teeth mesh outside(cid:173)
`to-omside, the direction of rotation i reversed. In Figure
`3-158. for example, the compressor bevel gear with 47
`teem drives the radial bevel gear with 35 teeth. The radial
`bevel gear turns in the opposite direction and faster by a
`
`DRJVE
`SHAFT
`
`AUXILIARY
`~RBOX
`
`GEARBOX
`
`--- r
`I SHAFT
`I
`I I
`I I MAIN
`
`I
`
`; (OCATIONOF
`I CONNECTING
`
`I
`
`(A)
`
`(B)
`
`TYPICAL SMALL ENGINE GEAR RATIOS
`(Input 100% N2. 37,500 RPM)
`RATIO
`0.293:1
`0.167:1
`0.103:1
`0.112:1
`
`STARTER/GENERATOR
`FUEL PUMP
`OIL PUMP
`TACH GENERATOR
`
`RPM
`10,988
`6,263
`3,863
`4,200
`
`Figure 3-14A. Main and auxiliary gearbox arrangement
`Figure 3-148. Typical small-engine gear ratios
`
`UTC-2011.009
`
`

`

`3-8
`
`Aircraft Gas Turbine Powerplants
`
`......_ ___ AUXILIARY
`GEARBOX
`
`100% N2 =14,460 RPM
`
`4'0
`
`Figure 3-15A. Main accessory gearbox location
`
`COMPRESSOR
`BEVEL GEAR __ _.--
`47TEETH
`
`/
`RADIAL
`BEVEL GEAR___,/'
`35TEETH
`
`STARTER
`BEVEL GEAR
`25TEETH
`
`RADIAL
`BEVEL
`GEAR
`25TEETH
`
`HORIZONTAL
`~~~~·t.---- BEVEL GEAR
`37TEETH
`
`GENERATOR
`GEAR SHAFT
`55 TEETH
`
`Figure 3-158 . Typical large-engine gear train
`
`55 TEETH
`
`UTC-2011.010
`
`

`

`Turbine Engine Design and Construction
`
`3-9
`
`factor of 47/35. The alternator gear shaft rums the same
`direction as the gear driving it because it is meshing out(cid:173)
`ide teeth to inside teeth. It also turns faster by a factor
`of 67/28 because it has 28 teeth and the gear driving it
`has 67 teeth.
`A secondary function of many main gearboxes is to
`provide a collection point for scavenged oil before being
`pumped back to the oil tank. This arrangement permits
`splash-type lubrication of many internal gears and bear(cid:173)
`ings inside the gearbox. However, modem engines do
`not typically locate the main oil supply in the accessory
`gearbox. Instead. a separate oil tank is normally used.
`COMPRESSOR SECTION
`The compressor section houses the compressor rotor
`and works to supply air in sufficient quantity to satisfy
`the needs of the combustor. Compre sion results when
`fuel energy of combustion and mechanical work of the
`compressor and turbine are converted into potential ener(cid:173)
`gy. Compressors operate on the principle of acceleration
`of a working fluid followed by diffusion to convert the
`acquired kinetic energy to a rise in pressure. The primary
`purpose of rhe compressor is to increase the pressure of
`the air mass entering the engine inlet and discharge it to
`the diffuser and then to the combustor section at the cor(cid:173)
`rect velocity. temperature, and pressure. [Figure 3-16)
`The problems associated with these requirements are
`evident when considering that some compressors must
`increa~e air flow to a velocity of 400 to 500 feet per sec(cid:173)
`ond and raise its static pressure perhaps 20 to 30 times in
`the space of only a few feet of engine length.
`In early compressors. which were less efficient than
`current technology, a given amount of work input pro(cid:173)
`duced air at a lower pres ure and at a higher temperature.
`To improve on laminar air flow over hundreds of small
`
`airfoi ls at high velocity and pressure, compressors have
`undergone constant development through the years to
`achieve optimum efficiency. CUJTently. efficiency is said
`to be in the 85 to 90 percent range. As outlined in Chapter
`2, compressor efficiency i!l based on the principle of
`maximum compression with the lea t temperature ri~e.
`Laminar flow minimizes friction-induced heat in the air.
`
`A secondary purpose of the compressor section is to
`supply engine bleed air to cool hot section parts. pressur(cid:173)
`iLe bearing seals, and supply heated air for inlet anti-icing
`and fuel system beat for deicing. [Figure 3-17 J Another
`secondary purpose is to extract clean pressurized air for
`aircraft uses unrelated to engine operating requiremenu..
`This air is usually referred to a!. cuMomer service bleed
`air. and common uses for it include aircraft cabin pres(cid:173)
`suritation. air conditioning systems. and pneumatic start(cid:173)
`ing. Customer service bleed air can be shut off for a ·•No
`BleeJ" takeoff on some aircraft, with the bleed air that
`the aircraft needs being drawn from the auxiliary power
`unit. This enables the aircraft's main engines to produce
`more thrust during takeoff because bleeding air from
`the engines robs them of power. On '>Orne newer aircraft
`such as the Boeing 787, customer service bleed air is not
`drawn from the engine. lnstead, standalone compressors
`driven by electric motors provide the necessary air. The
`aircraft's engines must till provide power to drive tbe
`electric motors, because the electricity comes from tbe
`engine-driven generatOrs. but this method is potentially a
`more efficient source of power.
`
`CENTRIFUGAL FLOW COMPRESSORS
`The centrifugal flow compressor. sometimes referred
`LO as a radial outflow compressor. is the oldest compre~­
`sor destg.n but is still in use today. Many smaUer engines,
`
`--------------
`LEGEND
`INLET AND BY-PASS AIR ~
`
`COMPRESSION AIR ~
`COMBUSTION AND ~
`EXHAUST AIR
`
`Figure 3-16. General Electric CF-34 airflow diagram
`
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`
`

`

`3-10
`
`Aircraft Gas Turbine Powerplants
`
`TURBINE COOUNG &
`BEARING SEAL
`AIR INLET
`PRESSURIZATlON
`FAIRING
`7th STAGE
`- - - - - - - - - - TAKEOFF
`
`AIRFRAME
`CUSTOMER AIR
`FROM 10th
`STAGE
`
`FAN
`SPINNER
`
`,.--------
`1 ........ _
`--------
`
`Figure 3·17. Bleed air distribution
`
`as well as the majorily of auxiliary gas rurbine power(cid:173)
`plants. use thb design.
`[n a centrifugal flow engine. the compressor receive:.
`air at the center of the impeller in an axial direction ami
`accelerates the air outward by centrifugal reaction to it'
`rotational speed. The air then expands into a divergent
`duct-formed by the shape of the lliffuser vanes-called a
`diffuser. and, acting in accord with Bernoulli's Principle.
`as the air 'iprcads out, its speed slows. which causes the
`~tatic pressure to build. l Figure 3-IBAJ The pressure
`and velocity graph in Figure 3-188 :.how!> tb:u velocity
`changes from an increasing to a decreasing value as pres-
`ure rises to it~ designed level at the exit of the diffuser
`vanes. The compressor efticiency graph in Figure 3-180
`shows that in a centrifugal flow compressor, efficiency
`drops off after approximately I 5: I pres~ure ratio, but the
`axial flow compre!'.SOr maintains its efficiency to a much
`higher level of pressure ratios. So as pressure ratio relate~
`to mass airflow and rhrust of a rurbojet or turbofan engine
`Cor horsepower of a rurboprop or turboshaft engine), the
`axial flow compressor is required where higher power is
`needed for larger or faster aircraft.
`The primary components in a centrifugal fl ow com(cid:173)
`pressor assembly are the impeller rotor. the diffuser, and
`the manifold. [Figure 3- I 91 The impeller is usually made
`from an aluminum or titanium alloy and can be either sin(cid:173)
`gle- or llual-sided. The diffuser provides a divergent duct
`in which the air spreads out. slows down, and increa~es
`in static pressure. The compressor manifold distribute.,
`the air in a turbulent-free condition to the combustion
`section.
`
`~ ENGINE ANTl-ICE AIR
`
`~ AIRFRAME CUSTOMER SERVICE AIR
`17'7771 BEARING SEAL PRESSURIZATlON
`~ AND TURBINE COOUNG AIR
`
`The single-sided impeller shown in Figure 3-20A
`benefits from ram effect and less turbulent air entry,
`making it well-suited tO many aircraft application!.. The
`single-stage. dual-sided impeller design in Figure 3-208
`results in a narrower overall engine diameter and high
`mass airflow, making it a favored del>ign in many flight
`engines in the past. However. this design does not take
`advantage of the fuU benefit from ram effect because the
`air has to rurn radially inward from a plenum chamber
`into the center of the impellers.
`Attainable compressor pressure ratios are about the
`~arne for both types of single-c;tage impeller types.
`However, as seen in Figure 3-21, having more than two
`stages of single-entry type is considered impractical. The
`energy lol>s to the airflow that must slow when making
`the turns from one impeller to the next. the added weight,
`and the drive . haft power extraction all offset the benefit
`of additional compre~sion with more than two stages.
`The most common centrifugal compressor in use
`is the single-sided type in either one or rwo stages. It
`is most often used in small turboshaft. turboprop. nnd
`turbofan engines. lc is not used in large gas turbine
`engines because it imposes a significant limitation on
`mass airflow. For example, the Pratt and Whitney I 00
`serie~ turboprop has two separate!)' rotating centrifugal
`compressors plus a power turbine and b described a:. a
`three-shaft. two-spool engine.
`Centrifugal compressors have gained a resurgence of
`use. and recent developments bave produced compreS!>Or
`pressure ratios as high as I 0: I from a single centrifugal
`
`UTC-2011.012
`
`

`

`Turbine Engine Design and Construction
`
`3-11
`
`(A)
`
`(8)
`
`OUTLET
`
`INLET
`
`(D)
`
`100%
`
`====:.,.:;-:_::-__ ....:..___,
`
`AXIAL FLOW
`
`I
`CENT. FLOW
`
`--..,
`'
`
`' , ,
`\
`
`CE
`OF
`MF
`PI
`RC
`E I
`SE
`SN oc
`RY
`
`PRESSURE RATIO
`0 ~.-----------.-----------r---
`15
`40
`
`Figure 3-18A. Airflow at entry to diffuser
`Figure 3-188. Pressure and velocity changes through a centrifugal compressor
`Figure 3-18C. Spoke theory of accelerating air
`Figure 3-180. Compressor pressure ratio versus efficiency
`
`(A)
`
`(B)
`
`COMPRESSOR
`MANIFOLD
`
`Figure 3-19. Components of a centrifugal flow
`compressor
`
`Figure 3-20A. Single-entry, single-stage Impeller
`Figure 3-208. Dual-entry, single-stage Impeller
`
`UTC-2011.013
`
`

`

`3-12
`
`Aircraft Gas Turbine Powerp/ants
`
`lltREE SttAfT T\JRBOPROP
`
`(A)
`
`AIR
`
`(B)
`
`Figure 3-21A. Single-entry, two-stage, d\lal-centrlfugal flow compressor arrangement-PW-120 turboprop
`Figure 3-218. Compressor and turbines of PW-1 20 t urboprop
`
`compressor. Previously, only ax.ial flow compressors
`could anain this level of compression. The main advan(cid:173)
`tage of a centrifugal compressor over its axial counterpart
`is its horter length.
`The tip speed of centrifugal impellers reaches approxi(cid:173)
`mately Mach 1.3. Radial airflow, however. remains
`subsonic. The pressure within the compressor casing is
`capable of preventing airflow separation at low super(cid:173)
`sonic rotor speeds and causing a high energy transfer to
`the airflow.
`
`The advantages of the centrifugal flow compressor
`over the axial flow compressor are as follows:
`
`• High pressure rise per stage-up to I 0:1 for a single(cid:173)
`stage and 15: I for a dual-stage
`
`• Good efficiency (compression) over a wide rotational
`speed range from idle to full power
`
`• Simplicity of manufacture and low cost
`
`• Low weight
`
`• Low starting power requirements
`
`Disadvamages are as follows:
`
`• Large frontal area for a given mass airflow
`
`• Impracticality of more than two tages because of the
`energy losses between stages
`
`AXIAL FLOW COMPRESSORS AND FANS
`The axial flow compressor is so named because air(cid:173)
`flow and compression occur parallel to the rotational
`axis of the compressor. Axial compressors are classified
`by the number of rotating sections, called spools. At one
`lime, the word spool was used only when describing an
`axial flow compressor. but many manufacturers currently
`
`UTC-2011.014
`
`

`

`Turbine Engine Design and Construction
`
`3-13
`
`use spool when describing a centrifugal flow compressOJ
`as well.
`
`TYPES
`The three types of axial flow compressors are single(cid:173)
`spool. dual-spool, and triple-spool. Smaller engines tend
`to use a single-spool compressor. and larger engines use
`dual- and triple-spool compressors because the additional
`spools deliver higher compression and mass airflow. The
`single-spool compressor was