Contents I. Overviews I-1. History I-2. Applications I-3. Components II. Basic Thermodynamics and Fluid Flows II-1. Five Basic Principles II-2. Some Important Formula III. Cycle and Performance III-1. Ideal cycles III-2. Component Characteristics IV. Aerothermodynamics in Major Components IV-1. Compressor IV-2. Turbine IV-3. Combustor

Contents (Continued) V. Structure & Dynamics V-1. Blade Vibration V-2. Stresses on Blade VI. Materials and Failure Modes VI-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

I-3. Major Components 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

I-3. Major Components 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

I-3. Major Components 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

II-2. Some Important Formulas Adiabatic Process Stagnation Properties  Stagnation enthalpy  Stagnation temperature  Stagnation pressure Stagnation Properties in Adiabatic Process

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

Part III. Cycle and Performance

III-1. Ideal Cycle Gas Turbine Basic Cycle : Brayton Cycle Simple Shaft Power Cycle  Efficiency  Specific Work

Regeneration Cycle  Efficiency if T 5 =T 4,

Reheat Cycle  Efficiency  Specific Work

Reheat Cycle  Influence of temperature ratio to efficiency and specific ratio

IPCCombustor HPC HPTLPT Power Turbine Regenerator Exhaust Gas Intercooler Intake Air Water Fuel Output Shaft GAS-TURBINE WITH INTERCOOLER & 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 T S T S 3’ 2’ 4’ Ideal Brayton Cycle Real Brayton Cycle

Isentropic Efficiencies of Compressor and Turbine  Compressor isentropic efficiency  Turbine isentropic efficiency

Combustor Efficiency  Combustor Efficiency  Fuel-Air Ratio  Specific Fuel Consumption [kg/kwh]  Thermal Efficiency  Heat Rate [kJ/kWh] where W N is net work produced by the whole engine per unit mass of air [kW/kg], Q net,p is heat value, i.e., heat rate supplied by unit mass of fuel at constant pressure combustion process [kW/kg].

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

 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 C a1 =C a2 =C a Pressure ratio per stage

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

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

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...

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-up Acceleration Load Base 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  100 o F  10% increase in power 2.4% increase in thermal efficiency  Historically, 30 o F/year ( )  Resistance to the material selection : Fatigue(HCF/LCF), Creep, Corrosion Requirements & Considerations  Mechanical strength  Under 600 o F : Yield and endurance for low temperature  Above 600 o F: 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 : 200 o C/minute

Environmental Problems  High temperature oxidation  Occurs when nickel based superalloys are exposed to temperature greater than 1000 o F (538 o C).  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. 15 o C increase in blade metal temperature cuts 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