1w7 Turbines

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Unit level: 3Issue Date31/05/2007

Outcome 1 AIRCRAFT GAS TURBINE ENGINES

Core code: Group AWeek 7 83

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TURBINES

CONTENTS

7.1 INTRODUCTION

7.2 TURBINE OPERATING ENVIRONMENT

7.3 GENERAL CONSTRUCTIONAL FEATURES7.4 TURBINE WHEEL

7.5 NOZZLE VANES

7.6 TURBINE DISC

7.7 TURBINE STAGING

7.8 TURBINE TYPES

7.9 TURBINE AERODYNAMICS 7.10 TURBINE BLADE OPERATING LIMITATIONS

7.11 THERMAL CYCLE/FATIGUE

7.12 CREEP

Unit Name Unit No Unit value: 1


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7.1 INTRODUCTION

The turbine assembly converts part of the energy of the working fluid into mechanical work.In the turbo-jets, approximately three-quarter of the energy in the gases is converted intoshaft horsepower to drive the compressor and accessories. The remaining energy isconverted into velocity in the propelling nozzle of the exhaust unit to produce jet-thrust. Inthe turbo-props, turbine is designed to extract almost all of the energy from the gas (about90%), leaving a small amount to produce jet thrust (about 10%).

The turbine assembly is one of the most highly stressed parts in the engine. Not only must itoperate at temperatures of approximately 928°C (1800°F), but it must do so under severecentrifugal loads imposed by high rotational speeds of over 6000 rpm for small engines to8000 rpm for the larger ones.

7.2 TURBINE OPERATING ENVIRONMENT

Turbine has to work in the most stringent environment in terms of temperature, centrifugalforce and gas loading. The problems that are common to turbine are:

• Thermal Fatigue

• Heat cracking

• Over temperature burns

• Creep

• Deformation

Figure 7.1 shows some of the turbine blade damages as a result of the exposure of theturbines in the worst environment of temperature and loading.

The limiting factor of how much power will be allowed to generate by a designed engine isthe material of the turbine that will withstand the working environment. Considerableresearch has been carried out with the object of devising improved blades for this purpose.Three main trends have been revealed:

• Development of new alloys with still-higher, safe operating temperatures.

• Application of non-metallic materials, principally ceramics

• Use of blade cooling



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Figure 7.1: Turbine damage due to high temperature & loading



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7.3 GENERAL CONSTRUCTIONAL FEATURES

The turbine section comprises basically of two main elements, the turbine wheel and nozzlevanes.

A stage consists of a set of nozzle guide vanes/nozzle diaphragm followed by a row of rotorblades. The row of blades is installed on the periphery of rotor disc and that of nozzles onthe casing by proper design attachment. Blades and Nozzle guide vanes have requiredfeatures in the design of roots and tips for attachment to respective discs/drum and statorcases. They have proper reactive or impulsive or combination reactive-impulsive shape andcurvature.

Figure 7.2 illustrates turbine rotor and stator assembly.

Blades and nozzle guide vanes in the rear stages (in the LPT section) are longer andconsequently prone to vibration. So they are constructed with shrouds (called shroudedturbines) to have a better torsional rigidity when installation is completed.

For efficient cooling of blades, turbine blades and nozzle guide vanes are fabricated withcooling provisions.

7.4 TURBINE WHEEL

The turbine wheel consists of contoured blades attached to a disc. This in turns is attached to

the power-transmitting shaft of the engine. The jet gases leaving the combustion chamber isaccelerated and directed by the stationary nozzle vanes to act upon the turbine wheel, turningit at very high speed. The turbine wheel then extracts energy from the hot gases byexpanding them to lower pressure and temperature. High stresses are involved in thisprocess, and for efficient operation, the turbine blade tips may rotate at speeds up to 1300ft/s. The continuous flow of gas to which the turbine is exposed may have an entrytemperature between 700 and 1200°C and may reach a velocity of 2000 ft/s in parts of theturbine.

Root designs of turbine blades are similar to those of the compressor blades; fir tree andserrated dovetail being the widely used methods. Figure 7.3 shows some of the rootdesigns and tip designs of turbine blades. Note that in the tip, some blades, specially thelonger blades at later stages may have tip shrouding; such turbines having such tip-designsare called ‘shrouded turbine’.

In the ‘fir tree’ designs, blades are attached to the disc by means of the 'fir tree' root to theserrations in the disc; such attachment allows for different rates of expansion between thedisk and the blade while still holding the blade firmly against centrifugal loads. The blade isfree in the serrations when the turbine is stationary and is stiffened in the root by centrifugalloading when the turbine is rotating. The blade is kept form moving axially either by rivets,locking tabs or devices, or another turbine stage.



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Figure 7.2: Turbine Rotor and Stator and drives



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The actual area of each turbine blade cross-section is fixed by the permitted stress inmaterial used and by the size of any holes, which may be required for cooling purpose. High

efficiency demands thin trailing edges to the sections, but a compromise has to be made soas to prevent the blades cracking due to temperature changes during engine starting andstopping.

Some turbine blades are open at the outer perimeter, whereas in others a shroud is used.The shroud acts to prevent blade-tip losses and excessive vibration. Distortion under highloads, which tend to twist the blade toward low pitch, is also reduced. The shrouded bladehas an aerodynamic advantage in that thinner blade sections can be used and tip losses canbe reduced by using a knife edge or labyrinth seal at this point. Another advantage is thatshrouding dampens the blade vibration permitting the use of reduced blade chords, and thisresults in a shorter and lighter engine. Shrouding, however, requires that the turbine run

cooler or at a reduced rpm because of the extra mass at the tip.Blades are forged from highly alloyed steel and are passed through a carefully controlledseries of machining and inspection operations before being certified for use. Many enginemanufacturers will stamp a 'moment weight number on the blade to retain rotor balancewhen replacement is necessary.

7.5 NOZZLE VANES

This is a stationary section consisting of a plane of contoured vanes, concentric with the axis ofthe turbine, and set at an angle to form a series of small nozzles which accelerate and dischargethe gases from the combustion chamber onto the blades of the turbine wheel. For this reason,

the stationary vane assembly is usually referred to as the turbine nozzle, and the vanesthemselves, are called nozzle guide vanes. Turbine-nozzle area is a critical part of the turbinedesign. If too large, the speed of the discharging gases will be too low and the turbine will notoperate-at its best efficiency. If too small, the nozzle will have a tendency to 'choke' undermaximum thrust conditions. The jets of escaping gases which are formed by the dischargenozzle area directed against the rotating turbine blades in a direction which enables the kineticenergy of the gases to be transformed efficiently into mechanical energy by the rotating turbinewheel.

The nozzle guide vanes are of aerofoil shape, the passage between adjacent vanesforming a convergent duct. The vanes are located in the turbine casing in a mannerthat allows for expansion. Nozzle vanes may be either cast or forged. Some vanes aremade hollow to allow a degree of cooling using compressor delivery air to reduce highthermal stress and gas loads. In all cases the nozzle assembly is made of very hightemperature, high-strength steel to withstand the direct impact of the hot, high-pressure, high-velocity gas flowing from the combustion chamber.

The vanes are located in the turbine casing in a manner that allows for expansion.Nozzle vanes may be either cast or forged. Some vanes are made hollow to allow adegree of cooling using compressor delivery air to reduce high thermal stress and gasloads. In all cases the nozzle assembly is made of very high temperature, high-strengthsteel to withstand the direct impact of the hot, high-pressure, high-velocity gas flowingfrom the combustion chamber. Figure 7.4 shows attachment of turbine nozzles.



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.

(a) Shrouded turbine rotor blades

(b) The fir-tree method o f attaching a turbineblade to the disk allows the blade to be loose whenit is cold, but it becomes rigid at operating

temperature

(b)

(a)

Figure 7.3: Turbine blade root and tip designs & attachment to disk

Figure 7.4: Turbine nozzles



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7.6 TURBINE DISC

The turbine disc is a machined forging with an integral shaft or with a flange on to whichthe shaft may be bolted. The disc also has around its perimeter provision for the attachmentof the turbine blades.

To limit the effect of heat conduction from the turbine blades to the disc a flow-of cooling airis passed across both sides of each disc. The disc or wheel is a statically and dynamicallybalanced unit of specially alloyed steel usually containing large percentages of chromium,nickel and cobalt. After forging, the disk is machined all over and carefully inspected usingX-rays, sound waves, and other inspection methods to assure structural integrity.

7.7 TURBINE STAGING

To produce the necessary driving torque for the compressors and accessories, the turbine mayconsists of several stages, each having one row of stationary nozzle guide vanes and one row ofrotating blades. The number of stages depends on whether the engine has one shaft or two andon the relation between the power required from the gas flow, the rotational speed at which itmust be produced and the diameter of the turbine permitted. However, with the advent ofhigher compression ratios, the tendency in recent years has been to increase the number ofstages.

When the turbine has more than one stage, stationary vanes are inserted between each rotorwheel and the rotor wheel downstream, as well as at the entrance and exit of the turbine unit.Each set of stationary vanes forms a nozzle-vane assembly for the adjacent turbine wheel. Theexit set of vanes serves to straighten the gas flow before passage through the jet nozzle.

The number of shafts varies with the type of engine. High compression ratio engines usuallyhave two shafts, driving high and low pressure compressors. On high bypass ratio fan enginesthat feature an intermediate pressure system, another turbine is interposed between the highand low pressure turbines, thus forming a triple-spool system. On some propeller-or shaft-powerengines, driving torque is derived from a free-power turbine. The shaft driving the propeller orthe output shaft to the rotor blades of a helicopter, through a reduction gear, may bemechanically independent of other turbine and compressor shafts.

Shaft RPM, gas flow rate, turbine inlet and outlet temperature and pressure, turbineexhaust velocity, and the required power output must all be given consideration. Ifthe engine is equipped with a dual compressor, the problem is more complex thanever, since the turbine also must be dual or 'split'. In this event, the forward part ofthe turbine, which drives the high pressure compressor, can be single-stage becauseit receives gases of high energy directly from the burner, and turns at higher RPMthan the turbine for the low pressure compressor. The gases have expanded inpassing through the . forward part of the turbine, and by the time that the gasesreach the rear part of the turbine, which drives the low pressure compressor,considerably more blade area is needed if proper work or energy balance is to be



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maintained. To accomplished this, a two-stage (or multi-stage) turbine is employedfor the second part of the turbine.

To be capable of supplying sufficient power for the compressor, the turbine must bedesigned so that the gases have a high expansion ratio. This, in turn, results in alarge temperature drop of the gases passing through the turbine, and a cool turbineexhaust if the engine is equipped with an afterburner without exceeding thetemperature limit of the construction materials used is the afterburner.

7.8 TURBINE TYPES

7.8.1 General : Turbine blading is divided into three categories: (a) Impulse, (b) Reactionand (c) Combination impulse-reaction.

In practice, the combination impulse-reaction type blading, which consists of an impulsesection at the root gradually changing to the reaction at the tip, is used in modern engines.

In the impulse type of turbine blading, there is no change in pressure between the rotorinlet gas and the rotor exit gas, but there is a large deflection of the gases. The nozzle guidevanes are shaped to form passages, which increase the velocity and reduce the pressure ofescaping gases. The high velocity generated in the turbine nozzle guide vanes impinges onthe moving rotor blades. In the reaction type of turbine a pressure change takes place asthe gases flow through the rotor blades resulting in an equal and opposite reactive force.The reduction in pressure and the increase in velocity of the gases are accomplished by theshape of the passage between the rotor blades. In most practical designs, the turbine is acombination of both of these two types and is known as a reaction-impulse turbine.However, all turbine designs have one important principle in common; that of expanding ahigh-pressure and high-temperature gas to lower pressures and temperatures. See Figure7.5.

The mass airflow from the combustion chamber is delivered evenly to the turbine nozzles.The nozzles serve to accomplished two functions; that of accelerating the gases, and ofdeflection the gases to a specific angle in the direction of turbine wheel rotation.

7.8.2 Impulse Turbine: In a pure impulse turbine, the entire pressure drop occurs in

the stator-nozzles; the-pressure in the rotor is the same throughout. The density of the fluidremains approximately constant in the rotor passages, and the area of the passages isessentially constant, from the entrance to the exit. Ref: Figure 7.6.

The function of the rotor blades is to deflect the fluid, reducing the velocity in the rotationaldirections and perhaps gain some in the opposite direction. In doing so -the blades exert aforce on the fluid to change its momentum and the fluid exerts a corresponding reactiveforce on the blades: This reactive force acts about the axis of rotation to form a torque andthe rotor turns continuously.



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Figure 7.5: Comparison between (a) impulse and reaction blading;(b) impulse-reaction and impulse blading



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Figure 7.6: Impulse turbine ro or bladet

Figure 7.7: Reaction turbine rotor blade



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7.8.3 Reaction Turbine: In a pure reaction turbine, the nozzle guide vanes haveparallel air passage- which only serves to direct the gas onto the rotor blades at a desiredangle. But in practice, the air passage of the nozzle of a reaction turbine is slightlyconvergent, so as to give a slight increase in exit velocity.

The rotor blades form convergent duct passages which accelerate the gas and give rise to areactive force to drive the turbine. The reactive force is derived from the acceleration of gasthrough the turbine blades. Ref: Figure 7.7.

The area of the stator passages is approximately constant, whereas the area of the rotorpassages decreases from entrance to exit.

As a result of the pressure drop, the velocity increases in the blades, becoming a maximumat the exit. The immediate accelerating force is the pressure gradient (exactly as in thenozzles of an impulse turbine). This force is transmitted to the blades so that themomentum of the fluid is really changed by a force exerted on the blades just as in theimpulse turbine. In both types, the effective force acting on a blade is physically thedifference of pressure on the two sides times the blade area.

7.8.4 Reaction-Impulse Turbine: In a reactive blading, a certain amount of impulseforce is always present. This together with the reaction force on the blades forms a resultantforce which acts in the plane of rotating to drive the turbine. Ref: Figure 7.8.

7.9 TURBINE AERODYNAMICS

The characteristics of gas flow -through the turbine may be shown conventionally by meansof a gas flow diagram. Figure 7.9, graphically represents the blade and passage contour.The direction of gas flow and the magnitude of the velocity is indicated by the length of thearrow. Velocities relative to the rotating blade are shown by dashed-line vector.

The gas flow enters the nozzle blade tip at a moderate velocity V1, and is acceleratedthrough the nozzle to an exit velocity V2. Note the deflection from the axis by the nozzle.Since the rotor blades are rotating, the tip speed and direction can be indicated by vector

V3. The direction and magnitude of the gas velocity entering the rotor blade V4, will be thevectorial difference between V2 and V3. This is a relative velocity and represents the angle

of gas entrance relative to the rotor blades. If there should be insufficient deflection anglefrom the nozzle, the rotor speed would be limited to a lower value than desired. At a designRPM, the gas deflection out of the nozzle must be such as to provide a gas flow into therotor blade at the optimum angle of attack. Thus the blade angles of both rotors and statorsrelative to the case must be determined together and for some design condition. Thedistance between blades must be held to minimum to reduce engine length and weight, butit must provide running clearance and enough space to attenuate wake effects which mightinduce vibration.



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Figure 7.8: Impulse-Reaction turbine rotor blade

Figure 7.9: Impulse-Reaction turbine rotor blade



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Assuming flow to be axial at the root as well as at the tip, the flow angles must be altered tocompensate for the lower speed of the blade root V6. To obtain the greatest rate ofmomentum change for impulse, a high degree of direction change o€ the gas flow throughthe rotor blade must be maintained. To accomplished this, a greater deflection through thenozzle is provided, with the lower root speed V3, the rotor blade entrance velocity V4, willbe greater than at the tip and its angle of attack will be increased.

A similar situation exists with the gas flow out of the rotor blades and through the exhaustduct. In order to obtain approximately axial flow down the duct, the rotor blades mustdeflect the exit gas velocity V5, sufficiently to compensate for desired rotor RPM. If theangle of exit flow does not properly match the desired rotor speed, the exhaust duct flow,

V7, would be deflected off the axis and result in a swirl flow instead of axial flow down theduct.

7.10 TURBINE BLADE OPERATING LIMITATIONS

7.10.1 General: Turbine blade is subjected to large forces due to gas load as well asrpm along with varying temperatures with very high peaks. This stringent workingenvironment has made the turbine susceptible to:

• Thermal Fatigue: Repeated heating and cooling (thermal cycle) of the materialaffect the physical properties of material and the blade lose strength developingcrack and failure.

• Thermal shocks: When turbine blade is heated rapidly to its peak operatingtemperature, it causes uneven temperature distribution and as a result of it, severethermal stresses are developed. These thermal stresses or shocks cause ultimatelythermal cracking and failure.

• Creep: Elongation due to creep continues to increase with working hours andpermissible tip clearance finishes. Further operation may result in failure due tocontact with the casing.

• Vibration failure

7.11 THERMAL CYCLE/FATIGUE

High temperature operation affects the life of all the components in the engine by varyingdegree. The maximum temperature of a gas turbine plant cycle occurs at the entry of thehigh-pressure turbine section. Each time engine is run and shut down, turbine bladematerials are thermally expanded and contracted. This cyclic reversal temperature condition(which is thermal cycle) leads to development of thermal shock or thermal stress or thermalfatigue. This condition made the turbine blades life –limiting and must undergo shop visit for



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repair/overhaul after specified hours of operation at which it completes designed thermalcycle. As the turbine exceeds its life in terms of specified number of thermal cycle, it fails

due to thermal fatigue.

7.12 CREEP

7.12.1 General: If a metal is loaded, it deforms. But the original dimension is regainedupon removal of the load, provided the stress is within the yield point of the metal. But, dueto continual application of the load, as well as other factors like temperature, originaldimension may not regained 100%, rather there may be continual deformation of metal withtime until fracture. This continual deformation under normal loading is called CREEP.

Typical creep curve is shown in Figure 7.10 that shows that the stretching of the partoccurs in three phases.

7.12.2 Creep for turbine blades: Creep of the turbine rotor blades is the result ofthree factors: (a) RPM, (b) Blade Mass and (c) Gas Temperature

As (b) is fixed, any variations in rpm and/or gas temperature will affect the creepcharacteristics of the blade. See Figure 7.11.

Turbine creep, a gradual permanent increase in blade length or disc diameter, with time,leads eventually to a failure of the blade or rubbing of the blade tip against its casing. Thetime elapsed before failure depends on the load applied and the temperature.

There is, in general, a fairly rapid initial increase in blade length or disc diameter, followedby a long period during which the increase is approximately linear with time. Finally there isa period during which the increase of blade length is rapid, leading to total failure. Naturally,in practice, this last condition should never be experienced in the life of an engine. However,in the event of severe overspeeding and high temperatures this excessive creep wouldoccur. A considerable decrease in the life of the blades occurs for an increase in bladetemperature of only 20°C above the normal maximum operating temperature.



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Figure 7.10: Creep curve

Figure 7.11: Creep variations


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