turbinefundamentals-12889484277447-phpapp02.pdf

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GAS TURBINE DESIGN

FUNDAMENTALS


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GAS TURBINE DESIGN FUNDAMENTALS

Learning goals

This unit is dedicated to gas turbines andstudents are expected to gain knowledge and

understanding of:

Gas turbine theory,

Design fundamentals;

Practical considerations of gas turbines;

Gas turbine comparison.


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Gas turbines have been gradually evolved onthe dominant main propulsion and ship-service

prime movers for destroyers, frigates, cruisers

as well as the foil-borne engines for hydrofoilcrafts.


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A gas turbine is a rotodynamic machine whichuses a gas compression combustion expansion cycle. It differs from a reciprocating

internal combustion engine in that:1 - The compression and expansion is

performed using rotodynamic components

2 - The combustion takes place at constantpressure


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GT are characterised by: High power to weigth ratios;

Lower thermal efficiency;

High output shaft speed;

Better quality fuels;

High air to fuel ratios; High power to volume ratios;

High availability;

Lower exhaust gas emissions;


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Centrifugal

Turbine and

centrifugal

compressor


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The Centrifugal CompressorThe centrifugal compressor consists of an

impeller enclosed in a casing which contains the

diffuser.


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(Illustration Rolls-Royce Ltd., 1969)


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GAS TURBINE DESIGN FUNDAMENTALScompressors

(Reprinted with permission of copyright owner, United Technologies Corporation)


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GAS TURBINE DESIGN FUNDAMENTALScompressors


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GAS TURBINE DESIGN FUNDAMENTALSBurners

About 15-20% of the air from the compressor passes over swirl vanes as it enters the primary zone of the burner. Here

also the fuel is introduced through nozzles as a fine spray of droplets. The swirling air causes the good mixing

necessary to support rapid, high temperature combustion.


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The annular burner is well-suited for an axialflow compressor. It is shown in the air

distribution pattern in this type may involve

introduction of the compressor air in only the firsttwo zones. The tertiary zone may involve final

mixing only. The advantage of this type of

burner is that it minimizes size and weight

with a sound aerodynamic design.


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GAS TURBINE DESIGN FUNDAMENTALSTurbines

There are two basic types of turbines,comparable to the two types of compressors.Due to the sizable stresses involved, the radial

turbine is generally not suitable for the hightemperatures necessary in a gas turbine engine.Therefore, the axial flow turbine is the only type

that will be discussed here. (See slide n7)


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Turbines may be of the impulse or reaction typedepending on rotor blade design.


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(a) Impulse turbine rotor blades -The flow passages are of constant

cross--sectional area resulting in

essentially no flow speed,

pressure, or temperature change.

Those changes occur in thestationary blades (nozzles). The

turning of the flow causes the

rotor to move.

(b) Reaction turbine rotor blades -The blades act as nozzles to

accelerate the flow as pressure

and temperature decrease. These

processes take place in both the

stationary and moving blades.


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GAS TURBINE DESIGN FUNDAMENTALSGT PTO


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GAS TURBINE DESIGN FUNDAMENTALScomplex power system


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Calculations exercisesExample 1


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Example 1

A gas turbine unit has a pressure ratio of 10:1 and a

maximum cycle temperature of 700C. The isentropic

efficiencies of the compressor and turbine are 0.82 and 0.85respectively.

Calculate the power output of an electric alternator geared to

the gas turbine when the air enters the compressor at 15Cat a rate of 15kg/s.

Take Cp=1.005 kJ/kgK and = 1.4 for the compression

Take Cp=1.110 kJ/kgK and = 1.333 for expansion


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2

2

0

vhh +=

02,23,2301 hWQh ININ =++

( )12010212 TTChhW p ==

( )2323 TTCQ p =

( ) ( )

( )23

1243

23

3412 )(

inputheat

net work

TTC

TTCTTC

Q

WW

p

pp

=

+

==

The steady flow energy equation applies to each component of the turbine. Defining

stagnation enthalpy

one can analyse the compressor, for instance, using:

In the idealised cycle there is no heat transfer during compression and

expansion so (for instance) the specific work (per kg of fluid)

Similarly there is no work done in the combustion chamber so

The efficiency can then be calculated using


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3

4

2

1

1

2

1

T

T

T

T

P

Pk ==

=

( ) ( ) ( )( )

==

=

=

1

23

23

23

23 11111

rkTT

kTT

TT

kTkT

1

2

P

Pr=

1

3

T

Tt=

( ) ( )1243 TTCTTCW pp =

=

11

11

1

rrtTC

W

p

Let us assume the compressor and turbine are 100% efficient (no entropy rise) and define the temperature ratio using isentropic formulae as

Substituting in equation 4.3 we have

(4.4)

where ris the pressure ratio,

The specific work output can be calculated as a function of pressure ratio r and the non-dimensional peak temperature,

i.e. turbine inlet to compressor inlet temperature.

(4.5)

so

(4.6)


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Plotting these efficiency and work relationships with

pressure ratio from equations 4.4 and 4.6, we see that

efficiency rises with pressure ratio (figure 4.4)

for any given peak temperature t there will be somepressure ratio that produces the peak specific power

(figure 4.5).

At any given pressure ratio, increasing the peaktemperature (by injecting more fuel) increases the work

output.


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In practice only the simplest gas turbines, driving electrical

generators at constant speed, extract power directly from the gas

generator shaft as in Figure 4.2. When driving any other load a

separate power turbine is desirable:

Increases in load do not slow down the compressorand cause a drop in pressure ratio

The speed:torque characteristic, for a given fuel flow,

is much more stable (see figure 4.6) The starter can rotate the gas generator spool without

turning the load


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Calculations exercisesExample 2


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Example 2

A gas turbine takes air at 17C and 1.01bar and has a compression ratio of 8:1. The

compressor is driven by the HP turbine and the LP turbine drives a CPP via a gear

box.

The isentropic efficiencies of the compressor and turbines are respectively 0.80,0.85 and 0.83.

Determine the pressure and temperature of the gases entering the power turbine,

the net power developed by the unit per kg/s mass flow rate, the net work ratio and

the cycle efficiency of the unit.

The maximum cycle temperature is 650C

For the compression process take Cp= 1.005 kJ/kgK and gamma=1.4

For the expansion process take Cp=1.15 kJ/kgK and gamma=1.333

Neglect the fuel mass flow rate.


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The efficiency of an ideal simple cycle gas turbine ispurely a function of its pressure ratio. This has twoimplications:

Efficiency is poor at part-load, when the shaft speed

and pressure ratio is lower and one is closer to theself-sustaining point where all the fuel is used purelyto overcome component losses

When the effect of component losses is considered,we find that for any peak temperature there is somepressure ratio at which the efficiency peaks: addingfurther compressor stages will then reduce rather

than increase the efficiency.


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The heat exchange cycle overcomes some of thesedifficulties. The main result of inefficiency in a simple

cycle is that the exhaust is hot. Providing it is hotter than

the compressor exit temperature one can use a heat

exchanger to transfer heat from the exhaust to the air

before it enters the combustion chamber: a given turbine

entry temperature can thus be achieved with a lower fuel

flow than in the equivalent simple cycle


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GAS TURBINE DESIGN FUNDAMENTALSrecuperated cycle


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Efficiency increases with temperature ratioso the provision of sophisticated turbine

cooling systems is beneficial. Efficiency

also rises as pressure ratio is reduced but this

is at the expense of a drop in specific work sosome compromise must be found. Typically

heat exchange cycles operate with a pressure

ratio of 4 to 5 (compared with 11 to 30 for a

large simple cycle engine).

1

3

T

Tt=

The ideal cycle efficiency is then a

function of both pressure ratio and the

temperature ratio


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As a further refinement the WR-21 includes an intercooler to cool the air

between the LP and HP compressor stages. This leads to a rise in specificpower, since less turbine work is required to drive the HP compressor. By

itself the intercooler would lead to a drop in efficiency (heat is being wasted);

in a recuperated cycle, however, the lower HP compressor exit temperature

means that the exhaust gases passing through the recuperator can be cooled

further and there is a corresponding rise in efficiency.


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The final WR-21 novelty is that the power turbine has

variable throat area nozzle guide vanes. At low powersin a conventional engine the combustor exit temperature

must be reduced to limit the power; with a variable area

nozzle the power can be reduced by lowering the massflow whilst maintaining the temperature. Compressor

surge is avoided because the gas generator turbine,

seeing a higher back pressure, generates less power sothe shaft speed and compressor pressure ratio are

reduced (which does not have a severe adverse effect

on the efficiency since this is a recuperated cycle).


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GAS TURBINE DESIGN FUNDAMENTALSCompressor theory


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Compressor theory

There are basically two ways to analyse how

turbomachinery (compressors or turbines) works.1 - Trace the changes in temperature and pressure from

one blade row to the next using velocity triangles, in

which we consider flow within each frame of reference(stationary or rotating) to have constant total

temperature and pressure along a streamline

2 - By consideration of the overall power input (Eulerequation) resulting from the change in angular

momentum across the rotor.


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Compressor theory

( )

2 2

1 1

01, 01 01 1 12 tan2 2 2rel a

p p p

C V U

T T T U C C C C = + = +

1

01

,01

01

,01

=

T

T

P

P relrel

The rel suffix indicates that this is a stagnation

quantity in the rotating frame. (The total temperature as measuredby a thermocouple mounted on the rotor would be different to that

measured by a stationary one).


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Compressor theory

relrel TT ,01,02 =

( )2 2

2 202 02, 02, 2 22 tan

2 2 2rel rel a

p p p

V C UT T T U C

C C C= + = +

( )02 01 1 1 2 2tan tana ap

UT T U C C C

= +

relrel PP ,01,02 =

102 02

01 01

P TP T

=

In the absence of heat transfer

.

(4.8)

If we neglect frictional losses and changes in radius

and we can apply an isentropic relationship across the whole stage:


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Compressor theory

V1Air relative velocity

Ca1

Axial velocity component

C1Axial velocity component

U Blade velocity

, fluid angles


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Compressor theory


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Combustor

The calculation is based on the following assumptions

(figure 4.19): No pressure drop takes place across the combustor

i.e. burner pressure ratio P4/P5.

Heat addition takes place under constant pressurewith no work output

The specific heat capacity of flue gas leaving thecombustor is equal to that of hot air at the exit

temperature. Fuel used has got a calorific value of 42.7 MJ/kg

Use of the steady flow energy equation with no heat

loss to the surrounding and neglecting velocity andpotential heads.


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Combustor


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Combustor

)2

()2

( 52

555

...

4

24

44

.

zv

hmWQzv

hm ++=++++

55

..

44

.

hmQhm =+

05055

.

04044

.

TCmhmTCm pffbp =+

05054.

404044.

)1( TCmfhmfTCm pfbp +=+

05050404 )1( TCfhfTC pfbp +=+

Based on the following assumptions the general steady flow energy equation

Can be re-written as


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Combustor

04

05

0404

04

05

)1(

1

p

p

p

fb

C

Cf

TC

hf

T

T

+

+=

Burner temperature ratio Nomenclature:T04 = Stagnation temperature at

inlet to combustor

T05 = Stagnation temperature at

outlet from the combustorb= Adiabatic efficiency

hf= Calorific value of fuel

Cp04= Specific stagnation heat

capacity at inlet

Cp05= Specific stagnation heat

capacity at outlet

f= Fuel air ratio


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Combustor

Graph for estimating

the gasestemperature at the

combustor outlet for

a variety of fuels


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Turbines


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Turbine

As with the compressor, we can trace thevariation of temperature through the turbine

using velocity triangles.

( )2022

2

2

202,02 tan2

222ax

ppp

rel CUC

UT

C

V

C

CTT +=+=

1

02

,02

02

,02

=

T

T

P

P relrel

relrel TT ,02,03 =if uncooled



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Turbine

( )3,032

3

2

3,0303 tan2222

axp

rel

pp

rel CUC

UT

C

C

C

VTT +=+=

( )320103 tantan axaxp

CCU

C

UTT +=

Neglecting frictional losses and changes in radius relrel PP ,02,03 =

and we can apply an isentropic relationship across the whole stage:

1

01

03

01

03

=

T

T

P

P

02 01,

02 01

is T

isen s

T TWW T T

= =

The turbine isentropic efficiency:


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Raising the pressure ratio by adding more compressor stages

increases the efficiency but also raises the combustor inlettemperature: for a given metallurgical limit for the turbine entry

temperature (TET) or (TIT) turbine inlet temperature, this implies a

reduction in fuel: air ratio and hence on the specific work.



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Leading particulars

" leading particulars" characterize the engine so that

potential customers can tell at a glance whether the engine

might suit their needs. Additional factors can then beconsidered if the engine seems appropriate.



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Turbine blade cooling

Cooling is provided by:1 - convection inside the

blade

2 - impingement of air

jets inside the NGV3 - convection within film

cooling holes

4 - an insulating film of

air around the outside

the aerofoils.


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GAS TURBINE DESIGN FUNDAMENTALSTurbine inlet temperature can be an indicator of certain design

features of the engine. Higher inlet temperatures necessitate more

sophisticated blade and vane cooling mechanisms and more heatresistant metal components. With present technology, 980C

1100C is commonly the maximum for continuous use;

The engine rotor speed is of importance for applications whichrequire gearing to electric generators, compressors, pumps, or other

direct-drive components;

The type and number of compressor and turbine stages,pressure ratio, and air flow are mainly of informational interest. These

are rarely a determining factor in selection of an engine.

Heat Rate (HR) and/or Specific Fuel Consumption (SFC) are

often included in the engine description as a measure of engineefficiency.

GT alternator pack


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GT tandem alternator pack

GT compressor pack

GT marine propulsion pack


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Main components of a gas turbine


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GT maintenance


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GT main performance curves


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GT main performance curves



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COMPRESSOR CHARACTERISTICS

The most important compressor performance characteristics are the

pressure ratio, air flow, and rotational speed. The like-new unit hascertain physical capabilities which usually represent a maximum for

that design.

To characterize the compressor overall operating conditions would

involve an unrealistic number of tables and/or graphs.



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axial flow compressor map



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centrifuge flow compressor map


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Surge is a damaging process which should be avoided if at all

possible, and choke (maximum) flow represents a condition oflowered efficiency as concerns the compressor.


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GAS TURBINE DESIGN FUNDAMENTALS The following generalizations should be kept in mind when

evaluating compressor performance (at a given speed) with the aidof a map:

An increase in pressure ratio moves the compressor closer tosurge.

A decrease in pressure ratio moves the compressor towardmaximum flow (choke). For ambient temperature below 15C, theequivalent speed is greater than actual, and above 15C, it is lessthan actual.

An increase in pressure ratio is accompanied by a decrease inmass flow.


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Considering the fact that under full loadconditions, approximately 2/3 of the turbinepower goes toward running the compressor. Forthis reason, a 5% loss in compressor efficiencycan cause as much as 10% loss in overallefficiency!

Another possible source of inefficiency is the airfilter. Inlet air filters are generally used innon-aircraft gas turbines.

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