Seoul National UniversityJune 29 - July 3, 1998 Fundamentals in the Gas Turbine Engine 1998. 6....

Seoul National UniversityJune 29 - July 3, 1998

Fundamentals in the Gas Turbine Engine

1998. 6.

Lecture Notes

Prepared by

Jeong-Lak Sohn, Dong Sub Kim and Sung Tack Ro


Contents

I. OverviewsI-1. History

I-2. Applications

I-3. Components

II. Basic Thermodynamics and Fluid FlowsII-1. Five Basic Principles

II-2. Some Important Formula

III. Cycle and PerformanceIII-1. Ideal cycles

III-2. Component Characteristics

IV. Aerothermodynamics in Major ComponentsIV-1. Compressor

IV-2. Turbine

IV-3. Combustor


Contents (Continued)

V. Structure & DynamicsV-1. Blade Vibration

V-2. Stresses on Blade

VI. Materials and Failure ModesVI-1. Materials

VI-2. Failure Modes

VII. Gas Turbine Development

VII-1. Flowchart for the Gas Turbine Development

VII-2. Development Organization


Part I. Overviews


I-1. History

1791 John Barber (UK) World’s first patent of the gas turbine engine (British Patent No. 1833) “A method of rising inflammable air for the purpose of producing motion and

facilitating metallurgical operation”

1903 Elling (Norway) World’s first gas turbine to produce power

1904 Stolze (Germany) 1905 Armengaud & Lemale (France) 1908 Holzworth (Germany) 1937 Whittle (UK)

World’s first jet engine (British Patent No. 347,206) W2/700 Nene, Tay (Rolls-Royce) Trent

J42 (Pratt & Whitney) PW4000

J31 (General Electric) GE90


I-2. Applications

Turbojet Engine Turboprop Engine

Turbofan Engine Turboshaft Engine

Types of Gas Turbine Engines


I-2. Applications

Aircraft Engines Commercial / military aircrafts Helicopters Missiles

Industrial Engines Power generations Mechanical drivers Marine / ground propulsion


I-3. Major Components

Three Major Components - Compressor - Combustor - Turbine



Compressor Situated at the front of the engine, Draws air in, pressurizes it, then delivers it into the combustion chamber. Two types of compressor design, centrifugal and axial flow.

Centrifugal compressor Axial Compressor



Combustor The air from the compressor passes into the combustion chamber where it is

mixed with the vaporized fuel sprayed from burners located in the head of the chamber.

The mixture is ignited, during the engine starting cycle, by igniter plugs located in the combustor.

Absorbs energy (heat) from fuel supplied from outside of engine Can, annular, tubular, cannular types of combustors

Can Type Combustor Cannular Type Combustor



Turbine Absorbs energy from the hot expanding gases leaving the combustor to keep the

compressor rotating at its most efficient speed and to produce required shaft power or thrust.

Axial and radial types of turbines

Axial Turbine

Radial Turbine


Part II. Basic Thermodynamics and Fluid Flows


II-1. Five Basic Principles

The First Law of Thermodynamics (Conservation of Energy)

Enthalpy

The Second Law of Thermodynamics

No engine can be more efficient than a reversible engine under the same conditions

Q

W

, , ,m h V z1 1 1 1

, , ,m h V z2 2 2 2

dS 0

h e p /

.

2

22

2

.

2

.

1

21

1

.

1 22Wz

VhmQz

Vhm


II-2. Some Important Formulas

Adiabatic Process

Stagnation Properties

Stagnation enthalpy

Stagnation temperature

Stagnation pressure

Stagnation Properties in Adiabatic Process

p / constant

h hV

T TV

C

p pV

p

0

2

0

2

0

2

2

2

2

20

2

11 M

T

T

)1/(20

2

11

Mp

p

)1/(120

2

11

M


Conservation of Mass

For one-dimensional flow in a pipe or duct,

Conservation of Momentum

For one-dimensional flow in a pipe or duct

Equation of State in Ideal Gas

min mout

m mout in 0

m m VAout in

p RT

0

dAnV

dAnVVF

inout

outin VmVmAAF

..


Part III. Cycle and Performance


III-1. Ideal Cycle

Gas Turbine Basic Cycle : Brayton Cycle Simple Shaft Power Cycle

Efficiency

Specific Work

/)1(

23

1243 11

s

rTTC

TTCTTC

uppliedheat

outputworknet

p

pp

4312 // pppprasratiopressureisrwhere

11

1 /)1(/)1(

1

rr

tTC

W

p

13 / TTTasratioetemperaturisTwhere


Regeneration Cycle Efficiency

if T5=T4,

C T T C T T

C T Tp p

p

( ) ( )

( )3 4 2 1

3 5

11r

t

( )/


Reheat Cycle

Efficiency

Specific Work

2 1 2

2

t c t c

t c t c

/

/

W

C Tt c

t

c

where c r

p 1

1

2 12

( ) /


Reheat Cycle Influence of temperature ratio to efficiency and specific ratio


IPC CombustorHPC

HPT LPT PowerTurbine

Regenerator

Exhaust Gas

Intercooler

Intake Air

WaterFuel

OutputShaft

GAS-TURBINE WITHINTERCOOLER & REGENERATOR


III-2. Real Cycle

Ideal vs. Real Brayton Cycle

2 2’ : Aerodynamic losses in compressor 3 3’ : Pressure drop in combustor 4 4’ : Aerodynamic losses in turbine

1

2

3

4

T

S

1

2

3

4

T

S

3’

2’

4’

Ideal Brayton Cycle Real Brayton Cycle


Isentropic Efficiencies of Compressor and Turbine

Compressor isentropic efficiency

Turbine isentropic efficiency

c

W

W

h

h

T T

T T

0

0

02 01

02 01

t

W

W

h

h

T T

T T

0

0

03 04

03 04

1/)1(

01

02010102

p

pTTT

c

/)1(

0403030403 /

11

ppTTT t


Combustor Efficiency

Combustor Efficiency

Fuel-Air Ratio

Specific Fuel Consumption [kg/kwh]

Thermal Efficiency

Heat Rate [kJ/kWh]

where WN is net work produced by the whole engine per unit mass of air [kW/kg],

Qnet,p is heat value, i.e., heat rate supplied by unit mass of fuel

at constant pressure combustion process [kW/kg].

bf

ff f

m

mwhere m mtheoretical

actual

theoretical actual

,

fm

mf

a

SFCf

WN

Work produced by the engine

Heat to the engine

W

fQN

net psupplied ,

SFC Qnet px ,


Cycle Performance Curves

Simple cycle Regeneration Cycle


Part IV. Aerothermodynamics of Major Components


IV-1. Axial Compressor

Elementary Theory Comparison of typical forms of turbine and compressor rotor blades

T-s Diagram

W mc T T mc T Tp p ( ) ( )03 01 02 01

s

T T

T T

03 01

03 01


Velocity Diagram

To obtain high temperature rise

in a stage ;

(1) high blade speed

(2) high axial velocity

(3) high fluid deflection in the rotor blade

Assuming that Ca1=Ca2=Ca

Pressure ratio per stage

)tan(tan

)tan(tan)tan(tan

tantantantan

21

010201030

21

.

12

.

12

.

2211

p

a

s

aa

ww

a

C

UC

TTTTT

UCmUCm

CCUmW

C

U

)1/(

01

0

01

03 1

T

T

p

pR ss

s


Design Process of an axial compressor(1) Choice of rotational speed at design point and annulus dimensions

(2) Determination of number of stages, using an assumed efficiency at design point

(3) Calculation of the air angles for each stage at the mean line

(4) Determination of the variation of the air angles from root to tip

(5) Selection of compressor blades using experimentally obtained cascade data

(6) Check on efficiency previously assumed using the cascade data

(7) Estimation on off-design performance

(8) Rig testing

Blade profile


Performance Curves

(a) Mass flow rate vs. pressure ratio (b) Mass flow rate vs. isentropic efficiency


IV-2. Axial Turbine

Elementary Theory

'0301

0301

)1/(

0301010

32

.

32

.

0301

.

0

.

3322

,/

11

)tan(tan)tan(tan

tantantantan

TT

TTwhere

ppTT

UCmUCmTTCmTCmW

C

U

sss

aapsp

a


Blade Profile Performance Curves


Blade Cooling

(a) Nozzle

(b) Rotor Blade

Impingement cooling

Convective cooling

Film cooling


IV-3. Combustor

Typical Combustion chamber

Pressure Loss

Pressure Loss Factor

Pressure Loss in the Combustor

1

2/ 01

0221

21

.2

0

T

TKK

Am

pPLF

m

2

01

01

.

01

0

2

pA

TmRPLF

p

p

m


Combustion Stability Loop

Methods of Flame Stabilization


Gas Turbine Emission

Effect of flame temperature on NOx emission

Dependence of emission on fuel/air ratio Diffusion vs. pre-mix burning

Pre-mixed combustor


V. Structure & Dynamics


V-1. Blade Vibration

Blade Vibrations Forced Vibration

Arises from the movement of the rotor through stationary disturbances such as upstream stator wakes, support struts, inlet distortions, or by forcing functions such as rotating stall.

Leads to high stresses and failure when the excitation frequency coincides with blade natural frequency.

Almost all the sources must be harmonics of the rotating speed of engine.

Flutter Arises by aerodynamic effects in the axial compressor. Occurs at frequencies that are not multiples of engine order and at different locations

on the compressor operating map.

Vibration Modes Natural Modes

Occur at characteristic frequencies determined by the distribution of mass and stiffness resulting from the variable thickness of the blade area.


Typical Vibration Modes Flap Modes

Torsional Modes

Disk Modes

The natural frequency or rotor vibration Reduced with increasing temperature Because of reduction in Young’s Modulus Increased at high speed Because of centrifugal stiffening

Rotor blade with 1F vibration mode

Rotor blade with 1T vibration mode


Typical vibration mode of a rotor in holographic image

Campbell Diagram

A design tool to estimate whether engine operates in resonance condition or not. Engine order : Excitation frequency Resonance condition : Coincidence of a natural frequency with exciting frequency


V-2. Stresses on Blade

Centrifugal Stress Centrifugal Tensile Stress

Limiting rotor tip speed and hub-tip ratio (i.e., blade length)

A factor in the hot section of gas turbines in conjunction with creep effects The maximum centrifugal stress occurs at the blade root.

Centrifugal Bending Stress Generated if the centers of gravity of shroud, foil, root are not located on the common radial axis.

Gas Flow Induced Steady State Stress Bending stress superimposed on the centrifugal stress Proportional to the aerodynamic loading on blade

Gas Flow Induced Alternating Stress Caused by stator vane wakes and wakes from support struts, etc...

2

222

max 12

2t

rt

bb

t

rr

bct r

rUANardr

a


Thermal Stress Blades are subjected to severe thermal stresses during transient conditions such

as startup and shutdowns. Typical thermal-mechanical cycle for a first stage turbine blade

Blade Failure due to Overspeed 25% overspeed 56% increase in resulting stress

Tension

Compression

Strain Metal Temperature

Warm-upAcceleration

LoadBase Load

Unload

Shut-down


VI. Materials and Failure Modes


VI-1. Materials

Typical Materials

Compressor- Blades/Vanes : Cr-Alloy, Titanium (Forging, Fabrication) - Discs : Ni-Alloy (Forging)- Cylinders : Cast iron, Titanium (Forging, Casting, Fabrication)

Combustor- Liner/Transition : Ni-Alloy, Hestalloy (Fabrication) - Casing : Steel (Fabrication, Casting)

Turbine- Blades/Vanes : Ni-Alloy (Casting) - Discs : Ni-Alloy (Forging)


Issues Related to the Material Selection High temperature material

100oF 10% increase in power

2.4% increase in thermal efficiency Historically, 30oF/year (1939-1979)

Resistance to the material selection : Fatigue(HCF/LCF), Creep, Corrosion

Requirements & Considerations Mechanical strength

Under 600oF : Yield and endurance for low temperature Above 600oF: Creep and endurance for high temperature

Corrosion resistance Low temperature Hot corrosion

Workability / Availability Casting/Forging Machining


VI-2. Failure Modes in Gas Turbine Blading

42% of engine failures are related to the blade-problems Failure Modes in Gas Turbine Blading

Low Cycle Fatigue - Compressor and turbine discs High Cycle Fatigue - Compressor/turbine blades & discs, compressor vanes Thermal Fatigue - Turbine vanes, combustor components Environmental Attack (Oxidation, Sulphidation, Hot Corrosion, Standby

Corrosion) - Hot section blades & vanes, transition pieces, combustors Creep Damage - Hot section blades & vanes Erosion & Wear Impact Overload Damage (Due to FOD, DOD or Compressor surge) Thermal Aging Combined Failure Mechanisms - Creep/fatigue, Corrosion/fatigue,

Oxidation/erosion, etc.


Fatigue High Cycle Fatigue (HCF)

Caused by aerodynamic excitations (Blade passing frequency) or by self-excited vibration and flutter.

Whereas fluctuating stresses may not be very high, the maximum stress at resonance can increase dramatically.

S-N (Stress vs. Number of cycles) curve

Low Cycle Fatigue (LCF) Occurs as a result of machine start/stop cycles. Associated with machines that have been in operation for several years. Minute flaws grow into crack which result in rupture. Predominant in the bores and bolt hoe areas of compressor and turbine disks which

operate under centrifugal stresses.

Thermo-Mechanical Fatigue (TMF) Associated with thermal stresses, e.g., differential expansion of hot section

components during startup & shutdown. Temperature variation in hot section blading : 200oC/minute


Environmental Problems High temperature oxidation

Occurs when nickel based superalloys are exposed to temperature greater than 1000oF (538oC). Nickel-oxide layer on the airfoil surface

When subjected to vibration and start/stop thermal cycles during operation, nickel-oxide layer tends to crack and spall.

Sulphidation A reaction which occurs when sulpher (in fuel) reacts with oxygen and attacks the base

metal. Particular concern when it is found in the blade root region or along the leading or

trailing edges, or under the blade shroud.

Hot corrosion Combined oxidation-sulphidation phenomena of hot section parts.

Standby corrosion Occurs during a turbine shutdown and as the result of air moisture and corrosive being

present in the machine. Blade fatigue strength is significantly reduced by corrosion.


Creep Occurs when components operates over time under high stresses and

temperature. Creep Curve

Creep-sensitive parts in engine Hot section parts and the final stages of high pressure ratio compressors. Mid span region of the airfoil which experiences the highest temperature. Disk rim region where high stresses and temperature can cause time dependent plastic

deformation.

15oC increase in blade metal temperaturecuts creep life by 50%.


Erosion/Wear Particulate Erosion

Compressor Particle size causing erosion : 5~10 microns Reduction in the surge margin can occur if the tips get severely eroded.

Hot Gas Erosion Turbine

Occurs when the cooling boundary layer on the blade surface breaks down even for short periods of time or cooling effectiveness drops.

The surface roughness of the blade contacted by the hot gas are subjected to high thermal stress cycles.

After several cycles, damage takes places and the increased roughness (erosion) worsens the problems.

First stage turbine vane

Combined Mechanism Corrosion reduce blade section size and drop the fatigue strength Erosion in the blade attachment regions reduce damping causing increased

vibration amplitudes and alternating stresses


VII. Gas Turbine Development


VII-1. Flowchart for the Gas Turbine Development


VII-2. Development Organization

Work Components

Engineering Organization

설 계 제 작 시 험

개념설계

기본설계

상세설계

치공구 제작

부품 제작

조 립

부품 시험

구성품 시험

엔진 시험

형상관리

개발총괄 설계 시험 /소재 공정

엔지니어링


형상관리

A 사업

B 사업

C 사업

개발 총괄

상세설계 /성능 제어 /공력 열전달 연소 /구조 동역학 /전기 전자

설계 시험 /소재 공정

엔지니어링

형상관리

개발 총괄 설계

설비기술 측정기술 시험기술

시험 /소재 공정

엔지니어링

Q C

생산관리

구매관리

제작관리

생산계획

치공구설계

생산기술

제작기술

선반

M illin g

판금

연마

용접

C o atin g

D eb urin g

가공 주조공장

제작

Design Organization

Test Organization

Manufacturing Organization

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Seoul National UniversityJune 29 - July 3, 1998 Fundamentals in the Gas Turbine Engine 1998. 6....

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