Turbo Machinery Dynamics Part 2 pdf

Figure 2.10 conceptually illustrates a typical arrangement, where the condi-tions correspond to a ideal shock-free operation, b supersonic flow up to inlet lip lowed by subsonic flow beh

increasing the bypass ratio from 0 to 5 Ideally the engine must have a fan pressure ratio ofabout 3, but this would call for two or more fan stages In practice, only one fan stage isused to lower weight and fan noise, causing the core jet velocity to be much higher than thefan duct discharge velocity Military engines such as TF-39 for C5A transport planes do nothave such restrictions, have one-and-a-half fan stages, and hence a bypass ratio of 8 is used.Note that the best value for the bypass ratio of a given aircraft also depends on the engineand fuel weight, noise, and installation drag.

Turbofan engines are ideal for subsonic applications Supersonic conditions call for anafterburner, where the combustion of the mixed fan and gas generator exhausts occurs The

fan pressure ratio will then be subject to the conditions p t6 = p t3 , or p f = p c p t , and t f = t c t t

If aft combustion raises the mixed streams temperature to q a, the velocity at exhaust also isthe same Thrust expression then becomes

compari-p f varies with M0for the given matching items and a = 1, but in reality a also varies with

M0 The thrust ratio of afterburning (or thrust-augmenting) to non-after-burning case islarge, which has merit if the requirement is for a subsonic cruise followed by a supersonicburst in speed

Turboprop engines have similarities with the turbofan version, but have a considerablyhigher bypass ratio and consequent propulsive efficiency A number of qualitative differ-ences exist between them The propeller does not have a diffuser, so the tips are exposed tothe airflow that combines the velocity of aircraft and the blade’s own peripheral tip speed.Hence, the propellers reach sonic speed at the tip even at medium flight speeds Aircraftdriven by turboprop engines are generally limited to Mach 0.6, primarily because the pro-pellers tend to cause high noise levels and are inefficient during supersonic operation.Implementation of the propellers on the engine is through a gearbox

Thrust variation occurs with changes in altitude and speed in aircraft engines Theexpression for thrust relies on total air mass flow through the engine, which changes withthe flight Mach number, atmospheric density (hence altitude), and flow conditions withinthe engine High-bypass engine thrust decreases with increases in the flight Mach number,

as seen in Fig 2.6, for different bypass ratios Also, if takeoff requirements control the ing of the engine, turbofan and turboprop engines will experience exponential decay inthrust with altitude

siz-When a regenerative heat exchanger is added to a turboprop to withdraw heat from bine exhaust and transfer it to the air entering the combustor, reduction in specific fuel

−

−

− 0

0

0 0 0 θ

FIGURE 2.5 Turbofan engine with afterburner—variation of thrust (upper), specific impulse (lower) with Mach number.

consumption is obtained But the high weight of the component rules out its usage wherehigh compression capacity is required

In the preceding discussion the turbine inlet temperature q twas not changed in order to

evaluate the impact of other cycle parameters In Eq (2.10) as q tincreases, thermal ciency, thrust, and specific impulse increase in turbofan engines, because jet velocity is

effi-reduced by increasing q t for a given bypass ratio This can be observed by holding u e /u0constant as q tis changed Thrust per unit of airflow then becomes

Specific heat c v in the compressor rises due to the increase in temperature, causing g=

c p /c vto fall Greater variations may be expected in the combustor because of the sharperrise in temperature and the formation of H2O and CO2 The application of tabulated values

1(γ ) θθ /

F

u u e

( / ) (0 ) 0

0

1+α =  −1

FIGURE 2.7 Turbine inlet temperature and specific impulse: turbofan

first stage of the compressor are also prominent The net result of these losses is to call formore energy input than for an ideal isentropic compressor Compressor and turbine effi-

ciencies h c and h tare displayed in Fig 2.8 for a range of polytropic efficiency values ical of many turbofan engines Losses in the compressor heat the air, requiring more energy

typ-to be put in, hence h c < hpolytropic On the other hand, turbine losses make energy available for subsequent expansion, and so h t > hpolytropic

Burners are subject to losses arising from combustion and pressure drop Incompletecombustion of a mixture of air and fuel results in the formation of CO and soot

Characterized by burner efficiency h b, it defines the change in enthalpy flux from theburner’s inlet to its exit and is divided by the energy content of the fuel flow Heat value of

the fuel, h, also plays a role, the lower value being more appropriate for gas turbines since

water leaves the combustor in the form of vapor Loss of stagnation pressure due to viscouseffects, and to a limited extent by the addition of heat, will be represented in an expression

for the burner stagnation pressure ratio p b

When expansion in the exhaust nozzle is correct, it causes the flow to expand ically to ambient pressure within the nozzle When the flow is not fully expanded, as mayhappen if the nozzle pressure ratio is large enough to cause the exit velocity to exceed Mach1.0, further expansion occurs outside the nozzle without producing the corresponding

isentrop-thrust The pressure ratio p e /p0is controlled by the geometry of the nozzle and by the ratio

of stagnation and ambient pressures

Expressions for thrust and specific impulse observe modification when losses areincluded in the analysis Using the station numbers in Fig 2.1, the thrust from the fan flowbecomes

High-bypass turbofan engines with one fan have a limit of 1.6 for the fan pressure ratio

In Fig 2.9, gains in thrust and specific impulse can be obtained by using a fan pressure ratio

of 2.0 with a fan bypass ratio below 8, since the core jet velocity is larger than the fan jetvelocity in the given range of bypass ratios Effects of changes in the compressor pressureratio, compressor and fan polytropic efficiency, and turbine efficiency can be observedfrom similar comparisons with the aid of graphs

The performance of an engine system depends on the characteristics of the inlet, fan, pressor, combustor, turbine, and exhaust Thermodynamic calculations given in the previ-ous section do not relate to the shape, size, and form of the parts The behavior of individualcomponents, on the other hand, is determined by their mechanical characteristics and lim-iting factors

com-Interaction between the fluid and the surface it flows over, thermal effects, and tural integrity define the limits of operation of a component, and hence of the engine sys-tem Dynamic forces are created by the gas flowing through the designed shapes of passagewalls, shock waves, and boundary layers Gently varying cross sections in the channelscontrol the flow velocity along the axis and, to a lesser extent, perpendicular to the axis.Shock waves introduce step changes in pressure, temperature, and velocity; a drop in thestagnation pressure of the bulk flow over solid surfaces thus exerts considerable changes insupersonic flows In the proximity of end walls of passages between blades and nozzles,

g dm dt

a F

dm dt ga f f

4 0

4 0

0 4

4 0 0

0 4

F

dm dt u

u u

u T

p p c

8 0 8 0

0 8

8 0 0

0 8

the vane velocity must change rapidly from the mainstream value to zero due to the no-slipcondition Fluid behavior in the region near the walls is controlled by pressure and viscousshear forces because the momentum is negligible The boundary layer theory may be used

to define a correction factor for the passage and to predict the time when flow separationoccurs When the flow separates from a wall, the diffusing effect of the downstream por-tion of the passage is lost, and the walls no longer control the flow Exertion of fluid vis-cous shear stresses on passage walls also causes transfer of thermal energy between thefluid and the wall (Kerrebrock, 1992)

An inlet (or diffuser) performs the task of bringing air from ambient to conditionsdemanded by the fan or the compressor while efficiently capturing the flow over a widerange of free-stream Mach numbers Diffusers for a subsonic flight considerably differfrom designs for a supersonic flight to decelerate the airflow to subsonic levels Mach num-

ber M2at the fan is determined by the rotor speed and the air temperature, and is the largest

at high altitude and full engine speed Requirements are most significant when the aircraft

takes off at full speed and high T0, but the variation in M2is not large from subsonic cruise

when T0and rotor speed are low A reduction of 20 percent from takeoff to high subsoniccruise may be generally expected

An internal compression diffuser takes the form of a convergent-divergent channel wheresupersonic flow is reduced in speed, by a series of weak compression waves, to sonic velocity,

then down to subsonic condition Pressure recovery is enhanced at high M0by taking tage of the fact that a series of weak shocks produce much less loss of the stagnation pressurethan one strong shock, and may be used to advantage in the external compression inlets toform an oblique shock A combination of internal and external compression through an

advan-FIGURE 2.9 Thrust and specific impulse in turbofan engines with losses.

oblique shock with internal compression inside the lip provides an added benefit of reduceddrag from the cowling Axisymmetric diffuser designs are popular for pod-mounted engineinstallations Figure 2.10 conceptually illustrates a typical arrangement, where the condi-

tions correspond to (a) ideal shock-free operation, (b) supersonic flow up to inlet lip lowed by subsonic flow behind it, (c) flight Mach number increased to critical so that a normal shock stands just at the lip, and (d) back pressure adjusted so that the shock stands

fol-at the throfol-at are shown in the diagram Flow areas are denoted by A with approprifol-ate

sub-scripts

Figure 2.11 provides off-design performance characteristics of the simple internal pression inlet described in Fig 2.10 due to varying airflow mass When flow is supersonic

com-up to the throat, the shock moves downstream into the divergent portion of the throat This

mode of operation is shown for M0= 3 For lower flow Mach numbers, the bow shockremains in front of the inlet The most advantageous operating point is just above critical

as marked by the circles in Fig 2.11 External compression inlets do not face complicationsfrom starting and stopping, and hence their behavior is simpler Characteristics for an exter-nal compression inlet are shown in Fig 2.12

The temperature ratio across a fan or compressor stage depends on the tangential Mach

number of the rotor, M T , axial flow Mach number, M b or M a, and flow geometry as dictated

by the blade configuration Hence, parameters such as t sfor the stage can be correlated as

a function of M T and M a This also implies that the stage efficiency h sis a function only of

M T , M a and flow geometry, and the stage pressure ratio p s = p s (M a , M T) if the Reynoldsnumber is large enough, in the range of 3 × 105, based on blade chord Figure 2.13 illustratesthe performance characteristics of a fan stage without inlet guide vanes for high axial Mach

FIGURE 2.10 Inlet arrangement (a) isentropic diffusion for M < 1, (b) operation with shock ahead of lip,

1 1

numbers A low tangential Mach number helps minimize noise; supersonic tip speed yields

a higher pressure ratio Interesting items may be observed in the maps At a given speed, asthe mass flow reduces, pressure ratio rises until it reaches a bound where the flow becomesunsteady, as indicated by the stall line Pressure ratio is virtually unchanged by the mass

flow at low corrected speeds due to the absence of inlet guide vanes As N √q increases, the

constant speed characteristics become steeper

Multistage compressors consist of a number of stator and rotor stages placed in series on

a single shaft Hence they operate at the same mass flow and speed Flow area reduces gressively in the stages in order to maintain axial flow velocity, and may be accomplished

pro-by reducing the tip radius, increasing the root radius, or both simultaneously The reducedtip radius decreases the tangential Mach number, reduces air temperature rise, and lowersthe pressure ratio of downstream stages The blade height is shorter when the root radius isincreased, so tip clearances are more difficult to control Stress in supporting disks below theblades also rises Performance map of an HP ratio compressor is shown in Fig 2.14.The discussion on compressors relates to turbines in several ways, except for two spe-

cial items: (a) increased gas temperature at turbine inlet and (b) falling pressures as flow

progresses through a turbine, as opposed to rising pressures in a compressor High gas perature leads to lower tangential Mach numbers for turbine blades than for compressor

tem-FIGURE 2.11 Off-design performance of internal compression inlet (Kerrebrock, 1992).

blades, relatively easing the aerodynamic problems Also, the falling pressure in turbinesthins the boundary layers to reduce separation concerns.

Compressor efficiency has a stronger impact on the overall engine system than turbineefficiency Still, high-bypass turbofan engines rely heavily on turbine efficiency As notedbefore, an increased turbine inlet temperature improves thermal efficiency, but generousprovisions for cooling must be made As a consequence, the previous definition of effi-ciency needs modification The ratio of actual turbine work to the total airflow, includingcooling and ideal work that would be attained in expanding that flow through the definedpressure range, defines the turbine efficiency Cooling flow may also be assumed toexpand through the same pressure differential as the primary flow Cooling flow impactsthe turbine efficiency in multiple ways Cooling air exiting from the blades causes a higher

FIGURE 2.13 Highly loaded fan stage performance map—tangential Mach

0.96 (upper); tangential Mach 1.5 (lower).

level of drag It also suffers a pressure loss while traversing the cooling passage, so it has

a lower stagnation pressure when mixed with the main downstream flow for a given t t.Note that the entropy of the total flow increases due to the heat transfer from the hot pri-mary flow to the cooling flow

Empirical representation of turbine efficiency in terms of corrected parameters follows

on similar lines as that for a compressor The corrected speed N/ √q indicates the tangential Mach number, where q is the inlet stagnation temperature divided by the standard reference temperature and N is the physical rotor speed Corrected weight flow Wd/ √q represents the

axial Mach number Figure 2.15 provides details of the performance characteristics for atypical, single-stage 50 percent midradius reaction turbine Because gas mass flow is

mostly independent of speed for p t > 2.5, parameter (Wd/√q)/(N/√q) is used for the

abscissa All speed characteristics then compress into a single curve, so the turbine mass

FIGURE 2.14 Compressor performance map.

flow characteristic is similar to that for a choked nozzle Choking for a simple nozzle is similar to that for a turbine due to the extraction of energy by the rotor.

The operating performance of an engine depends on the characteristics of its individualparts Interaction between the major component modules establishes their suitability whenfunctioning as an assembly in an engine system The objective behind matching compo-nents calls for the application of constraints that result from the special needs of modules.Performance maps of the fan, compressor, turbine, inlet, and exhaust nozzles determined

in the previous section serve as the basis for matching, from which the performance of theresulting assembly may be predicted

Consider the performance maps of Figs 2.14 and 2.15 for a compressor and a turbinemounted on a shaft, with a combustor placed in between to form the gas-generating module

A turbojet engine will need an inlet and an exhaust nozzle associated with the core gas erator Placement of a fan, with or without a separate fan turbine, creates a turbofan engine.The power turbine permits the use of a larger fan operating in its own favorable speed range.Using the station numbers provided in Fig 2.1, matching of the compressor, combustor, andturbine modules implies satisfaction of the following relationships (Kerrebrock, 1992)

W2= (1 + f )W1or

(2.22)

W1c pc (T t7 − T t1) = W2c pt (T t2 − T t3)or

(2.23)

f, T t2 , and T t7are related by

(2.24)

Turbine nozzles are usually choked at full power Then (W2/A2d2)√d2has a unique value

as determined by the turbine nozzle geometry Equation (2.22) may then be expressed in

terms of p t7 /p t1 as a function of (W1/A1d1)√d1and T t2 /T t1

(2.25)

p p

f A A

W A

T T t

t t

7 1

1 2

2 2 2 2

1 1

1 1 2 1

(1 )( )(

θδ

hf

c T

T T

T T

p t t t t t

1 2 1 7 1

= −

1

3 2

1 2 7 1

−T T t = +c f c T T T T − 

t pc pt t t t t

A A t t t t

=

Assuming a value of T t2 /T t1= 6.0, 90 percent turbine efficiency, and using the pressor performance characteristics of Fig 2.14, the pumping features of a gas generatormay be represented as shown in Fig 2.16 Stagnation of a pressure ratio of 3.4 and tem-perature ratio of 5.2 occurs at 100 percent corrected speed.

com-Matching of a nozzle to the gas generator is considered next The size of the exit nozzle

A nmay need adjustment to ensure that the mass flow through the nozzle equals that of the

turbine, or W n = W1(1 + f ) Hence

(2.26)

This equation establishes the ratio of areas between the nozzle and the compressor Also, if

A n /A1is fixed, the corrected speed at which it operates then becomes a function of T t2 /T t1,which becomes the single control variable dependent on the fuel flow rate

Matching of spools is more complex, because inlet conditions to the HP compressor will

be dependent on the corrected speed of the LP compressor Pressure ratio of the LP turbineand core turbine exit pressure control this item Matching of the LP spool follows on thesame lines as shown in Eqs (2.21) to (2.25), except for the substitution of a gas generator

in place of the combustor

To illustrate, consider a turbofan engine with separate fan and core nozzles Station ber 2.5 is added in Fig 2.1 between the high- and low-pressure turbines and number 0.5 justaft of the fan in the core stream Corresponding to Eqs (2.21) and (2.22) the expressions are

A A

2.5 2.5

0.5 1 1 2.5 0.5(1 )

A A n

n n n

t t t

t n

δ θ = +(1 ) δ11 1θ131

3 1 1

FIGURE 2.16 Gas generator pumping characteristics of 3.4 and temperature ratio of 5.2 occurs at 100 percent corrected speed.

Because of the bypass, W1/W0.5= 1 +α Either the fan nozzle area or the pressure ratio

of the LP compressor and the nozzle area of the LP turbine must be known The powerexpression then takes the form

(2.29)

Overall engine mass flow must also match that for the inlet The corrected speed

(N c/√q0.5) for the core and the corrected weight flow (W0.5√q0.5/d0.5) of the compressor are

determined by T t2 /T t0.5 for the fixed nozzle, so the inlet must provide a variable M1at theengine face

A turbojet is considered to determine the overall performance for convenience Thrust

Axial and radial centrifugal compressor types find wide acceptance in aircraft engines.Smaller engines can easily take advantage of the centrifugal impeller’s requirement of areduced flow area at the inlet to obtain a pressure ratio as high as 5:1 in one stage whileoperating at nearly 80 percent efficiency The rotating impeller’s mechanical energy is used

to create centrifugal forces in the air stream Engines are also designed with a row of two

to four axial stages followed by a centrifugal stage Two impellers turning at a high speed,exceeding 30,000 rpm, may also be employed in an opposed arrangement to increase theflow capacity (Hunecke, 1997)

Figure 2.18 shows the impeller of a single-stage radial compressor Air enters the eye

of the impeller axially, then goes through a 90° rotation to exit at the periphery into the fuser to go through another 90° turn, before discharging into the manifold Adequate pre-cautions are needed to ensure that flow separation does not occur in the flow passage and

dif-to avoid the associated losses The diffuser transforms the high flow velocity indif-to sure head through the gradually increasing cross sectional area of the passages within.Most engines, however, rely on the axial compressor for air pressurization to handle the

dm dt g

A p h

A c T f

F

A p W A hf

c T

n

p t n

p p

p p M

A A

p p

n e

1 2.5 0.5 1

−T T t = ++f c c T T T T − 

t

pc pt t t t t

α

large mass flow rate necessary for high levels of compression and thrust generation Sinceflow direction is mostly axial, turning of flow direction is eliminated Also, the consistentexternal dimensions in an axial flow machine help hold down the aerodynamic drag Axialcompressors rely on the principle of lowering stream velocity to raise pressure, and henceare more susceptible to flow fluctuations.

A considerably larger number of parts go into the building of an axial compressor, thusincreasing the complexity of the design and manufacturing cost Major components aresupport frames at the front and rear ends, an external casing with stator vanes attached to

it, and a bladed rotor Figure 2.19 provides details of an axial compressor module.The front support frame in a turbofan engine is made of 8 to 12 radial struts attached atthe inner radius to a hub and to a case at the outer radius The hub provides accommoda-tion for bearings for the fan and compressor rotors Bypass engines are provided with anintermediate splitter ring to provide the inner flow path for air going through the fan duct.Bearing housings are attached to the frame Since load-carrying requirements of the LProtor are high, especially in the event of a blade loss, two bearings are provided at the frontend, one of them close to the fan A housing for the HP compressor rotor’s bearing is

FIGURE 2.17 Performance characteristics of JT3D engine at 35,000-ft altitude (Kerrebrock, 1992).

trifugal compressor.

FIGURE 2.19 Axial compressor module of General Electric J79 engine (Hunecke, 1997).

37

ation Fixed stage vanes are inserted as an assembly in T-shaped grooves machined in thecasing.

Since the engine’s front mount in bypass engines is located at the compressor case ward flange, engine thrust forces will be reacted at this point for an eventual transfer to thepylon and aircraft wing Note also that the mount arrangement is mostly at the top of thecasing, so thrust force along the engine centerline will set up bending moments in the cas-ing Some compressed air is nearly always bled and transported through manifolds forcabin pressurization, deicing, and other assorted tasks

for-A combination of a drum and disks is the most common form of rotor construction, with

a shaft segment at both ends to carry the bearings Disks may be attached at either ity of the drum by spacer rings Rotor blades are twisted in order to get the right flow inletangle along the full length Variation in magnitude of the inlet angle arises because theblade has a much lower tangential velocity at the root than at the tip, while the axial flowvelocity must be maintained along the full face The blade height reduces gradually as theairflow progresses through the stages, in nearly the same proportion as the pressure rises.Blade roots may be of the dovetail shape, or of the fir tree type if the loading is consider-able Attachment of the blade to the disk must be firm, but still permit room for growth dur-ing operation Longer blades tend to move a little in the seat, but lock under the action ofcentrifugal forces during operation Axial entry of the dovetail in the disk or drum grooves

extrem-is common in forward stages, but circumferential insertion extrem-is favored for the latter-stageblades Blade retainers may be in the form of a key or a fitted bolt

The rear frame of a compressor is required to transfer the compressed air stream to thecombustor, and so its configuration depends on the burner system type to be used in theengine The cross-sectional area is made to increase gradually in the direction of the flow

in order to reduce stream velocity and increase pressure The center portion of the framehouses the gas generator bearings for absorbing axial and lateral direction loads Framestruts must also accommodate tubes for oil lubrication and venting If the primary enginemount is located at the compressor rear frame, as is the case in many combat aircraft, thestructural integrity of the frame under the action of engine-operating loads is of signifi-cance

The operation of the engine during cruise is set by the equilibrium operating line.Acceleration of the engine is more crucial since more fuel is being injected and burned inthe combustor Since the turbine inlet temperature consequently rises, engine componentsdownstream of the compressor are momentarily accepting less airflow This will have theeffect of raising the compressor pressure ratio and discharge pressure, and can be detri-mental to the compressor if the spike in the generated power causes the compressor tosurge Thus, a margin is necessary between the operating and surge lines to prevent thecompressor from surging, and is usually kept at 20 percent But a generous surge margin

may not be achievable over the full operating range of the compressor At low correctedengine operating speed steady state operating line will likely come near the surge contour.Remedial action in the form of modulated airflow by adjusting the stator vanes or by bleed-ing midstage compressor air should help alleviate the situation Lapses in the surge marginare likely to take place either at very low or high corrected speeds, which in turn are caused

by very low or high temperature at the compressor inlet

Of particular interest in high-bypass turbofan engines are the fan blades GeneralElectric’s CF6-6 engine fan uses 38 titanium blades to generate a considerable portion ofthe total thrust Weight and stiffness (and hence natural frequency) control of the bladesmay be achieved by the simple expediency of drilling holes in the tip (as shown in Fig 1.2).Due to the large fan diameter, the root of the blade may be operating in the subsonic regime,but the tip peripheral velocity may exceed the speed of sound during maximum thrust con-ditions A circular profile for the tip section of the airfoil aids in meeting the more rigorousdemands of the supersonic flow, but conventional airfoil shapes are employed at the hub toserve the intended purpose

Air drag can be limited by reducing the outer diameter of the fan, and hence a smallerhub diameter is desirable A varying distribution of energy along the length of the blade,reaching a maximum at the tip, aids in achieving efficiency and stall margin targets A keyconsideration in the control of vibrations, dynamic stresses, and operating life is the bladelength Midspan shrouds promote structural integrity of the blades, and are also advanta-geous in the event of a bird strike However, the shrouds come with a penalty in efficiency,since they act as an obstruction to the airflow Erosion from rain, hailstone, and ingestion

of ice are other factors to be kept in mind in the design of fan blades Close matching ofdimensions, both in the blade and in the mating disk dovetail, is of special significancebecause of the high centrifugal loads The problem becomes even more acute since theblades must permit easy removal in the field Fan disks must be designed to withstand axial,radial, and tangential direction loads arising from the aerodynamic and centrifugal forces

on the blades Titanium has proved to be a reliable material for the disk because of itsadvantageous strength and weight characteristics The fan stage may be followed by one tofour LP compression stages, located just aft of the split in the air stream for bypass engines.Figure 2.20 provides details of the fan module in a high-bypass turbofan engine

Axial turbines are the norm in aircraft power plants because of the higher mass flow bility Besides meeting mechanical loads during power generation, turbine componentsmust operate in a high-temperature environment Turbine design will be controlled by thecompressor’s pressure ratio, quantity of energy that is to be extracted from the hot gases,rotor speed, and maximum permissible diameter Turbine stage designs tend to be based on

capa-a mixed combincapa-ation of constcapa-ant pressure capa-and impulse principles Selection of capa-appropricapa-atematerials and cooling techniques is of fundamental importance in a successful turbinedesign (Hunecke, 1997)

The rotor assembly consists of disks, blades, and shaft segments at both ends At thecompressor end, the shaft is of a conical design to permit attachment with the shaft drivingthe fan rotor (see Fig 2.21) A splined joint with a lock nut is favored for the transmission

of power from the turbine to the fan The disks are bolted together, with spacers placed inbetween the disks The varying thickness in the disks permits a safe distribution of stressescaused by the blade centrifugal forces at the periphery, with the hub axial length being two

to three times the length at the rim Bore sizing is crucial in that hoop directional stressespeak in the region, and must also permit the placement of a vent tube through the center hole

Blades are inserted in machined slots and are held in place either by clamps or retainingplates, with passage for cooling air directed toward holes machined in the base of theblade’s dovetail.

A number of cooling methods are used to maintain metal temperatures well below those

of the gas Cooler air from the compressor flows through the inside of blades and vanes,exiting through holes drilled in the leading and trailing edges Thermal barrier coating helpsincrease the resistance of the base metal to elevated gas temperatures, while also providingprotection from corrosion Provision of cooling air to the rotating blades is at best a diffi-cult proposition Figure 2.22 illustrates the case for a gas generator turbine Contamination

of cooling air by dust particles poses a serious problem in that the exit holes may getclogged, thus creating hot spots and premature metal fatigue In this respect aircraft enginesface an entirely different situation when compared with industrial land-based gas turbinesprovided with intake manifold filters to keep out the offending contaminants One solutioncalls for cooling air to go through sharp turns before entering a swirl separator in which dustparticles are whirled in the direction of shaft rotation, thus causing the particles to fly out

FIGURE 2.20 CF6-6 Engine fan disk (National Transportation Safety Board, 1990).

due to centrifugal action Cooling air then proceeds through traps to separate the particlesbefore entering the passages inside the blade.

Specially designed seals are required to prevent leakage of hot gases in the confinedspace where a shaft meets a stationary component Relative thermal growth between thecomponents aggravates the problem Labyrinth seals have been expressly developed toresolve the problem (Fig 2.23) A number of grooves are machined into the rotor, separatedfrom one another by a ridge with a knife-edged crest The mating location on the station-ary part has a flat cylindrical surface Knife-edges cut into the stator’s rub surface to pro-vide the sealing When hot gases leak past the gap, throttling reduces the pressure from oneknife-edge to the next At the last location the throttled pressure equals the pressure on theaft side, thus effectively reducing the leakage rate to zero The method may be applied atthe blade tips as well as at the inner interstage locations Blade shrouds are designed with

FIGURE 2.21 GE79 Turbine rotor (Hunecke, 1997).

Splines

Disk

Rotorblades Rotor

Rear stubshaft

TurbineshaftShaft from compressor

Connecting screw

the cutting tip to control leakage flow Softer honeycomb materials are more suitable in theturbine casing.

The horizontally split casing design facilitates servicing of turbine components On theother hand, General Electric’s CG6-80C engine uses a 360° casing, thereby eliminating anumber of parts required for joining the upper and lower halves at the horizontal flange.The turbine module is assembled by stacking the nozzle vane and rotor disk assemblies insequential order The removal of the complete turbine module and its replacement withanother takes care of the field-related problems, with the module then serviced at the fac-tory Bore scope holes are conveniently placed, usually for each stage, on the casing shellfor inspecting the internal parts of the turbine Thus, if an engine is experiencing excessivevibrations or out-of-range exhaust gas temperatures, the internal inspection with a lightprobe facilitates the diagnosis of the problem

The control of radial clearances between blade tips and the casing is achieved by ing cooler air on the external surface of the case in order to shrink the casing Air is directedfrom holes in tubes placed around the case in the plane of the rotor stages In the activeclearance control system, generally employed for the gas generator turbine, temperaturesensors and valves are employed to control the flow of the cooling air during periods ofvarying temperature operation Passive cooling systems for the LP turbines merely blowcooler air on the shell The expense of installing sensors and valves is not required in thelatter approach, but the cool air is being continually robbed from the system In eithercase performance gains are achieved, because leakage of hot gases past the blade tips isminimized

An aircraft engine is typically placed in a nacelle housing for protection from the ment, to reduce aerodynamic drag by providing a streamlined surface and to reduce noise

environ-A large number of appurtenances in the form of tubes, linkages, and accessories areattached to the casing, and the nacelle protects the components during the flight Nacelleaerodynamic technology arguably offers less opportunity for refinement than does rotatingaerodynamics Numerous ideas for reducing installation losses or drag are available, but thecost implications of the changes must be considered Possible influences on fuel burn and

FIGURE 2.23 Labyrinth seal.