The Development of a Thermal Model to Predict Corrosion Propensity on Internal
Features of Turbine Airfoils
by
Daniel C Nadeau
A Project Submitted to the Graduate
Faculty of Rensselaer Polytechnic Institute
in Partial Fulfillment of the
Requirements for the degree of
MASTER OF ENGINEERING
IN MECHANICAL ENGINEERING
Approved:
_________________________________________
John Marcin, Project Advisor
Rensselaer Polytechnic Institute
Hartford, Connecticut
January, 2013
© Copyright 2013
by
Daniel C Nadeau
All Rights Reserved
Page: ii
CONTENTS
LIST OF TABLES ............................................................................................................. v
LIST OF FIGURES .......................................................................................................... vi
ACKNOWLEDGMENT ................................................................................................ viii
ABSTRACT ..................................................................................................................... ix
1. INTRODUCTION ....................................................................................................... 1
2. BACKGROUND ......................................................................................................... 2
2.1
GLOBAL ENVIRONMENTS ........................................................................... 2
2.1.1
RURAL .................................................................................................. 3
2.1.2
URBAN .................................................................................................. 4
2.1.3
INDUSTRIAL ........................................................................................ 4
2.1.4
MARINE / TROPICAL ......................................................................... 4
2.1.5
ATMOSPHERIC CONTAMINANTS .................................................. 5
2.2
THE BASIC GAS TURBINE CYCLE .............................................................. 6
2.3
COMPONENTS OF THE GAS TURBINE ...................................................... 9
2.4
2.3.1
COMPRESSOR SECTION ................................................................. 10
2.3.2
COMBUSTOR SECTION ................................................................... 12
2.3.3
TURBINE SECTION .......................................................................... 13
HIGH PRESSURE TURBINE (HPT) BLADE ............................................... 15
2.4.1
FUNCTION ......................................................................................... 15
2.4.2
DEGRADATION MODES.................................................................. 16
2.4.2.1 FOREIGN OBJECT DAMAGE (FOD) ................................ 16
2.4.2.2 CREEP ................................................................................... 17
2.4.2.3 FATIGUE .............................................................................. 18
2.4.2.4 OXIDATION ......................................................................... 19
2.4.2.5 HOT CORROSION ............................................................... 20
2.4.3
MATERIALS ....................................................................................... 20
Page: iii
2.4.4
PROTECTION MODES AND COOLING SCHEMES ...................... 27
2.4.4.1 COATINGS ........................................................................... 27
2.4.4.2 INTERNAL COOLING PASSAGES.................................... 30
2.4.4.3 FILM COOLING ................................................................... 33
2.5
HOT CORROSION ......................................................................................... 34
2.5.1
TYPE I HOT CORROSION (HIGH-TEMPERATURE) .................... 36
2.5.2
TYPE II HOT CORROSION (LOW-TEMPERATURE).................... 38
3. METHODOLOGY .................................................................................................... 39
3.1
PROBLEM DESCRIPTION ............................................................................ 39
3.2
SELECTED PART OVERVIEW .................................................................... 39
3.3
ENVIRONMENTS SELECTED ..................................................................... 41
3.4
CUT-UP INFORMATION .............................................................................. 44
3.5
THERMAL MODEL ....................................................................................... 46
4. RESULTS AND DISCUSSION ................................................................................ 48
4.1
ENVIRONMENT ANALYSIS RESULTS ..................................................... 48
4.2
CUT-UP ANALYSIS RESULTS .................................................................... 54
4.3
THERMAL ANALYSIS RESULTS ............................................................... 64
5. CONCLUSION.......................................................................................................... 70
5.1
FUTURE WORK ............................................................................................. 72
REFERENCES ................................................................................................................ 73
Page: iv
LIST OF TABLES
Table 1: Types of Atmospheric Environments [6] .............................................................. 3
Table 2: Corrosion Rates for Different Atmospheric Environments [6]............................. 3
Table 3: Typical Concentrations of Atmospheric Contaminates [6]................................... 5
Table 4: Steps for an Investment Casting Process for Turbine Airfoils .......................... 23
Table 5: Composition of various Nickel Based Superalloys [20]...................................... 40
Table 6: India Thermals: Predicted Temperature Ranges at Section Cuts ...................... 66
Table 7: Middle East Thermals: Predicted Temperature Ranges at Section Cuts ........... 67
Page: v
LIST OF FIGURES
Figure 1: The Brayton Thermal Cycle [23] ......................................................................... 8
Figure 2: Basic components of a turbine engine [7] ........................................................... 9
Figure 3: Engine Gas Characteristics [10] ......................................................................... 10
Figure 4: Compressor Gas Flow Characteristics [5] ......................................................... 11
Figure 5: Gas Turbine Combustor Schematic [24] ............................................................ 13
Figure 6: Turbine Gas Flow Characteristics [5] ................................................................ 14
Figure 7: Sample HPT Blade Picture [27] ......................................................................... 16
Figure 8: Foreign Object Damage on HPT Blades [18] .................................................... 17
Figure 9: Creep Damage on HPT Blades [9] .................................................................... 18
Figure 10: Temperature capabilities of several classes of alloys [3] ................................ 21
Figure 11: Investment Casting Process Flow Map [9] ...................................................... 25
Figure 12: Improvements in Casting Processes [5]........................................................... 26
Figure 13: Common Coating Processes [3]....................................................................... 28
Figure 14: A Schematic Depicting a Modern Day TBC Coating System [14].................. 29
Figure 15: Schematic of Common Cooling Techniques [13] ............................................ 30
Figure 16: Sample Trip Strip Configurations [13]............................................................. 32
Figure 17: A Sample Impingement Cooling Configuration [13] ....................................... 32
Figure 18: A Sample Pin-Fin Cooling Configuration [13] ................................................ 33
Figure 19: A Film Cooled Blade and Sample Cooling Hole Designs [1] ......................... 34
Figure 20: Effect of Temperature on Type I and Type II Hot corrosion [3] ..................... 36
Figure 21: Schematic diagram of the stages of Hot Corrosion [9].................................... 37
Figure 22: Schematic of Common Cooling Techniques [13] ............................................ 41
Figure 23: Map of Selected Cities in the Middle East [29] ............................................... 42
Figure 24: Plot of Airport Visit Data for the Selected India Engine ............................... 43
Figure 25: Map of Selected Cities in the India [17]........................................................... 44
Figure 26: Pictorial View of Cut-Up Analysis [2] ............................................................ 46
Figure 27: Comparison of PM2.5 and PM10 particles [16].................................................. 48
Figure 28: Residence Time for different particle radius sizes [16] ................................... 49
Figure 29: Annual Average PM10 Concentrations in World Cities [26]............................ 50
Figure 30: Annual Average PM2.5 Concentration in World Cities [19]............................. 50
Page: vi
Figure 31: Annual Average Concentrations of Ambient PM10 [22] .................................. 51
Figure 32: Trends of Ambient PM10 Concentration [22] ................................................... 51
Figure 33: Annual Average Concentrations of Ambient SO2 [22] .................................... 52
Figure 34: Annual Average SO2 Concentrations in World Cities [26].............................. 52
Figure 35: Trends of Ambient SO2 Concentration [22] ..................................................... 53
Figure 36: Magnetoscop Readings and Section Cut Locations [27].................................. 55
Figure 37: Internal cavity numbering & Pressure / Suction side labeling ....................... 56
Figure 38: Blade Cut-Up: India-A, Location: OD, Cavity 1 SS ...................................... 56
Figure 39: Blade Cut-Up: India-B, Location: OD, Cavity 5 PS ...................................... 57
Figure 40: Blade Cut-Up: India-C, Location: OD, Cavity 1 PS ...................................... 57
Figure 41: Blade Cut-Up: Middle East-A, Location: OD, Cavity 2 SS .......................... 59
Figure 42: Blade Cut-Up: Middle East-B, Location: ID, Cavity 6 PS ............................ 59
Figure 43: Blade Cut-Up: Middle East-C, Location: ID, Cavity 2 SS ............................ 60
Figure 44: Blade Cut-Up: Middle East-A, Location: OD, Cavity 2 SS - Measured ....... 61
Figure 45: Blade Cut-Up: Middle East-A, Location: ID, Cavity 2 SS ............................ 62
Figure 46: Blade Cut-Up: Middle East-A, Location: ID, Cavity 2 SS – Measured ........ 62
Figure 47: Blade Cut-Up: Middle East-C, Location: ID, Cavity 2 SS – Measured ........ 63
Figure 48: Blade Cut-Up: India-A, Location: OD, Cavity 1 SS – Measured .................. 64
Figure 49: Temperature Inspection Locations: Cavity 1 – Blue, Cavity 2 - Green ......... 65
Figure 50: India Takeoff Point Thermal Model Results .................................................. 66
Figure 51: Middle East Takeoff Point Thermal Model Results ...................................... 68
Page: vii
ACKNOWLEDGMENT
I would first like to thank Professor John Marcin for his guidance throughout the
duration of this project.
I also would like to express my appreciation to several
colleagues of mine at Pratt & Whitney for their assistance and advice on this project, I
appreciate the time they spent to teach and explain their knowledge – Mr. Mark Zelesky,
Mr. Jeff Levine and Mr. Mike Task. I would also like to thank Professor Brian Gleeson
and Tang Zhihong from Iowa State University for their assistance as well. Lastly I am
greatly thankful for the support and patience of my family, friends and fiancée during the
years of my graduate studies and the course of this project.
Page: viii
ABSTRACT
Modern day gas turbine airfoils are typically investment cast out of advanced
materials such as Single Crystal Superalloys that have increased melting temperatures
when compared to Equiaxed and Directionally Solidified Superalloys. These turbine
airfoils operate at gaspath engine temperatures which exceed the melting point of the
base alloy. Therefore, these airfoils are internally cooled and externally and sometimes
internally coated. Over the course of a jet engine’s mission, the internal cooling passages
of a high pressure turbine blade can experience a wide range of temperatures along with
different environmental contaminates such as salts that can lead to the development of
hot corrosion. Hot corrosion is a mechanism that can lead to rapid degradation of base
alloy chemistry leading to reduced environmental and mechanical stress resistance. The
goal of this project is to utilize a thermal analysis to predict the internal metal
temperatures experienced over the course of a jet engines mission. The internal metal
temperature prediction can then be used to determine the level of hot corrosion expected
for various features in various locations inside a turbine airfoil. An attempt was then
made to correlate these predicted temperatures to actual field run hardware for that given
mission. The world regions selected for this project were the Middle East and India with
the takeoff, climb, cruise, decent and landing mission points being analyzed for both
regions. Each regions mission analysis was based on environmental data including city
pairings for the flights, particulates within the atmosphere of the flights and temperatures
the engine is experiencing within flight. The environmental data collected was utilized
alongside data that was measured from actual hardware cut-ups that had traveled
missions in each of these two world regions. An exact correlation was not able to be
found between the temperatures predicted from the thermal model and the distress found
within the hardware; however, the project did provide a great step in the right direction
in achieving this correlation. The results from this project identified areas that needed to
be improved to better understand the correlation of metal temperatures to corrosion
zones. It is hoped that later versions of the thermal model can improve these areas and
eventually be able to predict locations within a high pressure turbine blade that may
corrode just by knowing the mission points and world region of operation.
Page: ix
1. INTRODUCTION
In both the aviation and power generation gas turbine industry, the performance
and durability of engine components are extremely important. In the aviation industry
these engine components are very important since the lives of the passengers within the
airplane are depending on the quality of the engines components. Whether the engine is
powering a single engine military jet fighter or a multi-engine commercial aircraft, the
safety of the passengers is always the top priority. Since the plane is relying on these
engines to stay airborne, it is essential for all aircraft engine producers to ensure that all
components are designed to the highest durability and are able to operate in various
environments since world environments are always changing. A major area of damage
and degradation within modern day gas turbines has been attributed to corrosion due to
the various operating environments.
Corrosion within an aircraft’s gas turbine engine can be as minor as a small
aesthetic problem or as large as a catastrophic structural or mechanical problem. Due to
this, corrosion has become a serious concern for the owners and operators of both
military and commercial aircraft. Around the world, aircraft can encounter and endure
various atmospheric conditions as well as diverse operating environments that can be
very aggressive. The gas turbine engines on these aircraft house particularly corrosive
environments where high temperatures and corrosive liquids and gases are constantly
present during operation. These engines operate under extremely high temperatures and
corrosion can be accelerated when the engine is exposed to the aggressive operating
environments of the aircraft. These environments have the ability to expose the gas
turbine engine to high temperatures, extreme temperature gradients, high pressures, large
stresses and the presence of oxidizing and corroding atmospheres that can include
ingested particulate materials. Therefore during engine design it is critical to understand
these environments when designing components. If materials are selected improperly,
the consequences can be extremely costly due to corrosion damage within the engine,
resulting in the grounding of an aircraft, higher maintenance and repair costs or in
extreme cases the complete loss of the aircraft and human life.
Page: 1
2. BACKGROUND
Gas turbine engines operate all over the world in military, commercial and
industrial applications. All these applications operate in different world regions and thus
each engine experiences different environments and global events over the course of
their life. To be able to study the effects of hot corrosion within a gas turbine engine, an
understanding is necessary of these operating environments, engine operations and
engine components. The following sections will breakdown the environmental factors
that can have an effect on a gas turbine engine as well as the key engine components
vital to the operation of the engine.
2.1 GLOBAL ENVIRONMENTS
All different environmental factors have the ability to influence the corrosion on
a material including the environments composition, pH levels, humidity, wind or water
currents and temperatures. All these factors exist in all different environment types
including atmospheric, fresh water, salt water and soil. For the purpose of this paper the
focus will be on atmospheric environments since this is typically the operating
environments of gas turbine engines. Whether in-flight or on the ground, aircraft are
constantly exposed to the atmosphere. The most corrosive atmospheric environments
are typically those containing significant amounts of pollution containing corrosive
agents such as sulfur dioxide, ammonia and smoke particles. Atmospheric environments
can vary throughout different regions of the world. Depending on the contaminants
present, these environments have typically been known to be categorized into rural,
urban, industrial and marine/tropical environments. A basic comparison of these
categories can be seen below in Table 1 and Table 2. Table 1 gives a brief description of
each category while Table 2 gives a general comparison of corrosion rates between
environments. A more detailed description of each category is found in the following
sections of this paper.
Page: 2
Table 1: Types of Atmospheric Environments [6]
Atmosphere Type
Description
Rural
● Generally the least corrosive
● Does not contain any significant amounts of pollutants
● Principal corrodants are oxygen and moisture content
Urban
● Similar to rural but with sulfur oxides (SOX) and nitrous
oxides (NOX) from vehicle and domestic fuel emissions.
Industrial
● Pollutants of sulfur dioxide, chlorides, phosphates, and
nitrates exist from heavy industrial processing facilities
● Special cases include contaminants of hydrogen sulfide,
hydrogen chloride, and chlorine which are highly
corrosive to most metals
Marine/Tropical
● Generally high corrosivity
● Characterized by chloride particles
● Deicing salts used in cold weather regions produce an
environment similar to marine
Table 2: Corrosion Rates for Different Atmospheric Environments [6]
Rate of Corrosion
2.1.1
Type of Environment
High
Tropical
Industrial
Moderate
Temperate
Suburban
Low
Arctic
Rural
Marine
Inland
RURAL
Rural environments typically are unpolluted by exhaust and sulfur containing
gases. These environments are usually sufficiently inland so that contamination and
high humidity from coastal waters are not present. In these areas the air quality typically
contains amounts of ammonia and nitrogen contamination from animal excrement,
fertilizers and high concentrations of diesel exhaust. Therefore these rural areas are
generally not considered very aggressive corrosion environments.
Page: 3
2.1.2
URBAN
Urban environments are found in highly populated areas generally having high
levels of automobile emissions and high rates of byproducts from combustion of
building heating fuels. Both of these conditions help to increase the levels of sulfur
oxide (SOX) and nitrogen oxide (NOX) concentrations within these areas. The corrosion
severity typically found in these environments is a function of pollution levels, humidity,
average temperature and equipment usage. These in turn depend on the population
density of the area, emission controls and pollution standards.
2.1.3
INDUSTRIAL
Industrial environments tend to be very diverse with the potential to produce a
variety of corrosive compounds. These environments can exist in different levels of
severity and different sizes yet have the same detrimental effect in the end. Sulfur and
nitrogen containing contaminants are most often linked to these environments. In these
areas the combustion of coal and fuel oils release sulfur oxides (SO2, SO3) and nitrogen
oxides (NOX) into the atmosphere along with other detrimental contributors such as
ammonia and its salts as well as hydrogen sulfide. In these areas many of these gases
accumulate within the atmosphere and return to the ground in the form of acid rain while
many industrial dust particles can be laden with harmful metal oxides, chlorides, sulfates,
sulfuric acid, carbon and carbon compounds. These particles, in the presence of oxygen, can
be highly corrosive and lead to many different forms of corrosion depending on the material
of attack. Industrial environments can sometimes be considered semi-industrial or severe
industrial, depending on the corrosivity of the atmosphere.
2.1.4
MARINE / TROPICAL
Marine environments tend to be considered the most sever environment in terms of
corrosion. These environments are found along coastal areas and are characterized by the
abundance of sodium chloride (salt) and sulfur compounds that are carried by sea spray,
mist, fog or winds. This sea spray is transmitted into the atmosphere by waves breaking on
the shoreline or by on-shore winds that contain tiny droplets of salt water that can be
transported many miles by ocean breezes. The aggressiveness of a marine environment
depends on the nature of the wave action at the surf line, prevailing wind direction, shoreline
Page: 4
topography, and relative humidity. The corrosion severity increases rapidly as the distance
from the shore decreases, however severe storms can carry salt spray far inland.
2.1.5
ATMOSPHERIC CONTAMINANTS
The primary sources of atmospheric contaminants come from the chlorides located
in the marine and industrial environments along with automobile pollutants. The presence
of these chloride salts in the atmosphere significantly increases the corrosion rates of most
metals. The other contaminates that have a large impact as stated previously are Sulfur
dioxide (SO2) and Nitrous oxides (NOX) found in industrial and urban environments. All of
these contaminates deposit onto the metal surfaces and react with oxygen and free electrons
from the metal surface producing varying increased corrosion rates. Table 3 below gives a
brief summary of typical concentrations of atmospheric contaminants and the region and
season they tend to be produced within.
Table 3: Typical Concentrations of Atmospheric Contaminates [6]
Contaminant
Region
Industrial
Sulfur Dioxide (SO2)
Rural
Sulfur Trioxide (SO3)
Hydrogen Sulfide
(H2S)
Ammonia (NH3)
Chloride (Cl-, Air
Sampled)
Chloride (Cl-, Rain
Sampled)
Season
Winter
Summer
Winter
Summer
-
-
Industrial
Urban
Rural
Industrial
Rural
Spring
Spring
Spring
Winter
Summer
Annual (Avg.)
Winter
Summer
Winter
Summer
Winter
Summer
Winter
Summer
Industrial Inland
Rural Coastal
Industrial Inland
Rural Coastal
Industrial
Smoke Particles
Rural
Typical Concentration
(μg/m3)
350
100
100
40
Approximately 1% of the
SO2 content
1.5-90
0.5-1.7
0.15-0.45
4.8
2.1
8.2
2.7
5.4
7.9
2.7
57 mg/l
18 mg/l
250
100
60
15
Page: 5
There are some major factors that can play a large role determining the impact of
corrosion rates of metals within these environments including Humidity, Rainfall, Wind and
Temperature. In general corrosion rates of metals will increase as the humidity increases
within an environment, while the corrosion rate may either increase or decrease due to the
amounts of rainfall within an environment. This is due to the addition of moisture into the
process and the electrolytes provided by this moisture. Wind plays a large role in the
distances that atmospheric contaminates are dispersed. The corrosivity of an atmospheric
environment is related to their relative distance and proximity to coastal (marine) and
industrial environments. Temperature can have a significant effect on the corrosion of
metals with the corrosion rate increasing as temperature is increased. This temperature can
play a large role in the form of attack that corrosion can take on, as the temperature increases
moisture can be evaporated on metallic surfaces and leave behind corrosive contaminates.
2.2 THE BASIC GAS TURBINE CYCLE
A gas turbine engine is an internal combustion engine that operates with the use
of the Brayton Thermal Cycle extracting chemical energy from a fuel and converting it
into mechanical energy utilizing air as the working fluid. In the Brayton Cycle the air is
at ambient temperatures at state 1 and compressed to a high pressure at state 2, fuel is
then added and continuously burnt to a firing temperature, state 3. The resulting highpressure, high-temperature air is expanded through a turbine to atmosphere, state 4. This
cycle can be seen depicted below in Figure 1. The higher the turbine inlet temperature,
typically called T4, for a given pressure ratio typically results in a higher power output
and efficiency of the engine.
The efficiency of the Brayton Thermal Cycle can be calculated by understanding
the conservation of energy (First Law of thermodynamics, Equation 1). This states that
the internal energy of a closed system is equal to the amount of heat supplied to the
system, minus the amount of work done by the system on its surroundings.
Understanding the plot in Figure 1 the efficiency of the Brayton cycle can be calculated
as follows:
(𝑞𝑖𝑛 − 𝑞𝑜𝑢𝑡 ) + (𝑤𝑖𝑛 − 𝑤𝑜𝑢𝑡 ) = ℎ𝑒𝑥𝑖𝑡 − ℎ𝑖𝑛𝑙𝑒𝑡
Equation 1
Page: 6
In this cycle the heat exchange can be expressed in terms of enthalpy. Treating
the working fluid as a perfect gas with constant specific heats the heat addition from the
combustor, qin, would be expressed as Equation 2, while similarly the heat rejected, qout,
is expressed as Equation 3.
𝑞𝑖𝑛 = ℎ3 − ℎ2 = 𝑐𝑝 (𝑇3 − 𝑇2 )
Equation 2
𝑞𝑜𝑢𝑡 = ℎ4 − ℎ1 = 𝑐𝑝 (𝑇4 − 𝑇1 )
Equation 3
Knowing equations 1, 2 and 3 the thermal efficiency of the Brayton Cycle can
now be expressed in terms of temperatures as seen in Equation 4.
𝜂=
𝑁𝑒𝑡 𝑊𝑜𝑟𝑘 𝑐𝑝 [(𝑇3 − 𝑇2 ) − (𝑇4 − 𝑇1 )
(𝑇4 − 𝑇1 )
=
=1−
𝐻𝑒𝑎𝑡 𝐼𝑛
𝑐𝑝 [𝑇3 − 𝑇2 ]
(𝑇3 − 𝑇2 )
Equation 4
Looking further the relationship between the different temperatures and pressures
can be examined. Knowing that process 1-2 and process 3-4 are isentropic and that P1 =
P4 and P2 = P3 the pressure and temperature relationships can be determined. These
relationships can be seen below in Equation 5, where the k value is the ratio of specific
heats Cp.
𝑇2
𝑃2
𝑃3
𝑇3
= ( )(𝑘−1)⁄𝑘 = ( )(𝑘−1)⁄𝑘 =
𝑇1
𝑃1
𝑃4
𝑇4
Equation 5
Finally the thermal efficiency of the Brayton Cycle can be seen in Equation 6.
This equation states that for a high efficiency, the pressure ratio of the cycle should be
increasing.
𝜂𝑡ℎ,𝐵𝑟𝑎𝑦𝑡𝑜𝑛 = 1 −
1
𝑟𝑝
; 𝑊ℎ𝑒𝑟𝑒 𝑟𝑝 =
(𝑘−1)⁄𝑘
𝑃2
𝑃1
Equation 6
Page: 7
Figure 1: The Brayton Thermal Cycle [23]
A gas turbine is typically made up of three main regions, a compressor to draw in
and compress the air; a combustor to add fuel to heat the compressed air; and a turbine to
extract power from the hot air flow. In a jet engine, cold air is pulled in through the
intake section and is delivered to the compressor section of the engine. In this section
the compressor will convert the mechanical energy into gaseous energy greatly
increasing the pressure of the air. From here the air is passed through the combustion
chamber where it is mixed with the fuel and ignited. This air/fuel mixture than expands
and releases heat accelerating through the turbine section of the engine. In the turbine
section the air/fuel mixture is converted back into mechanical energy by expanding the
hot, high pressure gasses into a lower temperature and pressure. This is done with the
use of stationary guide vanes directing the gases onto rotating blades. This direction of
gas onto the blades results in kinetic energy of the hot gases impacting on the blades and
Page: 8
rotating the blades and engine shaft. This air is then passed through to the last section of
the engine called the exhaust. This air passing through the exhaust creates thrust and
propels the jet engine, and aircraft forward. A sample cross section of a gas turbine
engine can be seen below in Figure 2.
Figure 2: Basic components of a turbine engine [7]
2.3 COMPONENTS OF THE GAS TURBINE
There are three major components to a gas turbine engine include a compressor,
a combustor and a turbine. Throughout the operation of a gas turbine engine the air
ingested will change velocity, temperature and pressure due to interactions with these
different sections and thus different energy changes will take place. Figure 3 below
shows a sample plot of these parameter changes and has been lined up with the different
sections of the engine; this shows the relationship each section has on velocity,
temperature and pressure.
The following sections of this report will give a better
overview of each of these major engine sections.
Page: 9
Figure 3: Engine Gas Characteristics [10]
2.3.1
COMPRESSOR SECTION
The compressor section of the engine is responsible for providing the turbine
section with the air that it will need in an efficient manner. In addition to supplying the
air, it also must supply this air at high static pressures. This is typically done in modern
day gas turbine engines with the use of a multi-stage axial compressor design. Each of
the stages will consist of a set of rotor blades attached to a rotating disk followed by
stator vanes attached to a stationary ring. The flow area between the compressor vanes
and blades always tends to be divergent and each stage incrementally will boost the
pressure of the air from the previous stage. The basic principle of the compressor is that
the rotor blades will convert mechanical energy into gaseous energy.
This energy
conversion is what greatly increases the total pressure (Pt) of the air flow; most of this
increase is in the form of velocity with a small increase in static pressure due to the blade
path divergence.
Page: 10
Each of the stator vanes of the stages will slow down the air with the use of the
divergent shape converting the accelerated velocity to a higher static pressure. These
stators are positioned at certain angles so that the exiting air is directed into the rotor
blades of the next stage at an angle that is the most efficient. The flow path of the blades
and vanes within this section has a decreasing cross-sectional area in the direction of
flow. This decrease in area reduces the volume of air as compression progresses from
stage to stage. As the pressure is built up by successive sets of blades and vanes, less
and less volume is required. Thus, the volume within the compressor is gradually
decreased. This compression process is repeated for each stage of the compressor,
Figure 4 below shows a two dimensional example of how the flow would pass through a
stage of blades and vanes in the compressor (Left) and the resulting pressure
characteristics that would be seen within a stage (Right).
Figure 4: Compressor Gas Flow Characteristics [5]
Along with the multiple stages of blades and vanes compressors also incorporate
inlet and outlet guide vanes located at the inlet and outlet of the compressor respectively.
These inlet and outlet vanes neither diverge nor converge. The purpose of these vanes is
to direct the inlet air to the first stage of the compressor at the most efficient angle as
well as “straighten” the air that will be provided to the combustor section of the engine.
Page: 11
Also in most engines small portions of air are typically bled out of this section to be
utilized as cooling air downstream in the hottest portions of the turbine section.
All the components utilized in the flow path of the engine section are shaped in
the form of airfoils to help maintain the smoothest flow of the air as possible. The
blades and stators of each stage are positioned at the optimum angles to help achieve the
most efficient airflow throughout the compressor.
A secondary function of the compressor is the supply the bleed air for various
purposes in the engine. The bleed air is taken from any of the various pressure stages of
the compressor. The exact location of the bleed ports depends on the pressure or
temperature required for a particular job.
The ports are small openings in the
compressor case adjacent to the particular stage from which the air is to be bled.
Varying degrees of pressure and temperature is available simply by tapping into the
appropriate stage. Air is often bled from the final or highest pressure stage because at
this point pressure and air temperature are at a maximum.
2.3.2
COMBUSTOR SECTION
Once the air leaves the compressor section of the engine it enters a diffuser
section through the outlet guide vanes on its way to the combustor. This diffuser is a
very divergent duct and the air enters this duct with a straight-line flow thanks to the
outlet guide vanes. The function of this duct is to convert most of the air velocity into
static pressure, as a result the highest static pressure and the lowest velocities of the
entire engine are found at the diffuser exit and the combustor inlet.
Once the air leaves the diffuser, it will enter the combustion section of the
engine, typically known as the combustor. This section has the task to control the
burning of large amounts of fuel and air. In the combustor gaseous or liquid fuels are
introduced by nozzles located around the combustion chamber.
The combustion
chamber is where a portion of the air from the compressor is introduced directly into the
combustion reaction zone while the remainder of the compressor air is introduced
afterwards to shape the flame and quench it to an acceptable firing temperature (T3).
This flow split is typically known as primary air and secondary air. The primary air is
used to support the combustion process while the secondary air is admitted into the
Page: 12
liners of the combustor. This secondary air is utilized to control the flame pattern, cool
the liner walls, dilute the temperature of the core gasses and provide mass. The cooling
air utilized in this section is critical due to the temperatures in the combustor being able
to reach levels as high as 3500°F. Towards the end of the combustor is a transition
section that has a convergent duct shape. This section accelerates the gas stream and
reduces the static pressure in order to shape the hot gases in preparation for the turbine
section with the desired velocity and temperature profiles. A basic layout/schematic for
a gas turbine combustor can be seen below in Figure 5.
Figure 5: Gas Turbine Combustor Schematic [24]
2.3.3
TURBINE SECTION
The turbine section of the engine basically acts like the compressor section in
reverse with respect to energy transformation. In this section, the turbine converts the
gaseous energy of the air/fuel mixture from the combustor back into mechanical energy.
The mechanical energy transformation is completed by expanding the hot, high pressure
gasses from the combustor to lower temperature and lower pressure gasses. Each stage
of the turbine is made up of a row of stationary vanes followed by a row of rotating
turbine blades. This is the reverse configuration when compared to the compressor. In
the compressor, the goal was to add energy to the gas with the rotor blades and then
convert to static pressure by the vanes. In the turbine, the goal is to increase the gas
velocity with the vanes and then extract energy with the rotor blades.
Page: 13
Like the compressor, the blades and vanes are of the airfoil shape to help provide
a smooth flow for the gaspath. As the air enters the turbine from the combustor, it is
accelerated through the first stage vanes due to the vane shape being a convergent duct
which converts the gaseous heat and pressure energy into a higher velocity gas flow.
Similar to the compressor, the vanes also turn the air flow so that it is directed into the
rotor blades at the optimum angle. As the air flows across the blades the gaseous energy
is converted to mechanical energy with the sacrifice of velocity, temperature and
pressure in order to rotate the turbine shaft. Figure 6 below shows a two dimensional
example of how the flow would pass through a stage of vanes and blades in the turbine
(Left) and the resulting gas characteristics that would be seen within a stage (Right).
Figure 6: Turbine Gas Flow Characteristics [5]
In the turbine section the blades and vanes have a gradually increasing span
resulting in each stage having an increased flow area when compared to the previous.
By having this design the difference in gas velocity is more equal across the whole
turbine section. If the flow area was not increased in this manor the pressure difference
would be across a particular stage, rather than the whole turbine, resulting in a reduced
amount of power recovered. The turbine section is typically separated into two different
sub-sections based on the pressure at the stages. These sections are the High Pressure
Page: 14
Turbine and the Low Pressure Turbine. Depending on the application these sections
could be a single stage or multiple stages having increasing flow areas.
Once the air is through the turbine section it will be discharged through the
exhaust section of the engine. Even though most of the gaseous energy up to this state
has been converted back to mechanical energy by the turbine a significant amount of
power still remains in the gas. This gas is then accelerated through a convergent duct to
help convert the gas into jet thrust.
2.4 HIGH PRESSURE TURBINE (HPT) BLADE
For this project, the main focus is a High Pressure Turbine Blade. This blade is
one of the most important parts within the engine and can experience some of the highest
temperatures when compared to other engine hardware.
2.4.1
FUNCTION
The function of the HPT blade, a sample part can be seen below in Figure 7, is to
extract power and energy from the hot gas stream exiting the engines combustion
chamber in order to drive the upstream compressor section of the engine. These blades
are required to be immensely strong and robust since each blade can operate at speeds
over 10,000 rpm and extract up to 750 horse power of energy. HPT blades tend to
average about 15,000 hours of service between engine overhauls depending on the
engine, operator and environments. These blades also do their work in the hottest part of
the engine core and must be able to function at peak performance in gas temperatures
reaching as high at 2900°F. The gas temperature entering the turbine is directly related
to the overall engines efficiency, the higher the gas temperature the more power that is
produced resulting in a more efficient engine. However the materials utilized for these
HPT blades limit the maximum usage temperatures at which they can operate within.
Thus different protection modes, cooling schemes, and alloy materials must be
developed to improve the engines efficiency. The following sections will cover the
different degradation modes experienced by HPT blades followed by the materials,
protection modes and cooling schemes utilized in their development.
Page: 15
Figure 7: Sample HPT Blade Picture [27]
2.4.2
DEGRADATION MODES
The operating conditions of gas turbine engines typically result in various
deterioration modes for gas turbine blades and, therefore, these parts typically have finite
operating lives. Under ideal loads the degradation of these components is primarily a
function of the operating temperatures entering and within the turbine. However the
engine is not always running at ideal conditions and thus different degradation modes are
introduced to the blades. Factors that can result in non-ideal conditions are corrosive
environments, over-firing of the engine and impact damage, these factors have the
possibility to result in accelerated component deterioration. Typical degradation modes
experienced by HPT Blades are as follows: Foreign Object Damage (FOD), Creep,
Fatigue, Oxidation and Hot Corrosion.
2.4.2.1 FOREIGN OBJECT DAMAGE (FOD)
Blade deformations can occur in the form of dents and scratches that can be
caused by foreign objects in the air stream of the engine. These objects can be ingested
by the fan at the inlet section of the engine or created by particles of metal and hard
carbon deposits from the combustion chamber section.
These dents and scratches
created by foreign objects have the ability to create stress concentrations in the blade
Page: 16
material that can lead to the initiation of fatigue, oxidation and corrosion processes.
Examples of Dents and Scratches on the blade surface can be seen below in Figure 8.
Figure 8: Foreign Object Damage on HPT Blades [18]
2.4.2.2 CREEP
Creep is a mode of inelastic material deformation occurring under sustained
loading at high temperatures for a period of time that will lead to permanent material
deformation. The rate of creep deformation is a function of the temperature, stress and
time. In a gas turbine engine there are a variety of different stresses applied on a given
component. In the case of a HPT blade the three largest factors that can lead to creep
deformation are centrifugal loading, operating temperatures and operating pressures.
Under these factors the blade material is weakened resulting in the acceleration of the
creep process. This process can eventually form voids that will lead to the initiation of
small cracks in the blade which can then grow in size and eventually lead to blade
failures. These failures can be as small as minor cracks or as devastating as a blade
cracking right in half, these failures are typically known as stress rupture. Another issue
that can arise from creep is the dimensional changes that can occur reducing its
aerodynamic efficiency and even lead to elongation and rubbing of the blade tip, which
in turn can lead to vibration and noise. The evolution of creep damage within a blade
can typically be represented by a creep curve consisting of three stages: primary
(transient), secondary (steady-state) and tertiary (unstable) which can lead to the rupture.
An example of a creep failure can be seen below in Figure 9.
Page: 17
Figure 9: Creep Damage on HPT Blades [9]
2.4.2.3 FATIGUE
Fatigue is the process of crack initiation and growth resulting from the creation
of alternating stresses during engine operation. Three distinct modes of alternating
stresses can be found in the operation of a gas turbine engine. These are Low Cycle
Fatigue (LCF), High Cycle Fatigue (HCF) and Thermal Mechanical Fatigue (TMF).
LCF is typically created with the alternating of stresses developed in the rotating
blades through the change in rotational speeds during engine operation, in particular the
engine starting up and the engine shutting down. LCF cracks are typically initiated in
regions where the developed stresses through rotation are the highest; these are areas like
the rims of rotors, turbine disks and blade to disk connections. This mode is typically
characterized by high amplitude low frequency plastic strains.
LCF is typically
prevented by the retiring of parts after a specified number of loading cycles (engine
startup and shutdown cycles).
HCF is typically generated due to the alternating stresses created in turbine
blades through the excitement of a resonant state through variations in gas impulse
Page: 18
loads. This mode is usually characterized by low amplitude, high frequency elastic
strains. Two main abnormal conditions lead to the creation of these types of variation,
the first being the creation of non-uniform temperature and pressure distributions across
the face of the turbine blade due to abnormalities in the combustion process. While the
second is the creation of gas velocity differences through abnormalities in the shape of
individual turbine vanes upstream of the turbine blades. HCF is typically prevented by
maintenance to maintain uniform combustion and uniform shape of the turbine vanes.
TMF is typically generated by the alternating stresses that are created do to the
change in gas temperatures resulting in different regions of the turbine blade being
heated or cooled at different rates when compared to other regions. This mode of fatigue
tends of occur in regions that are the thinnest in cross-section and where the greatest
thermal strains are located. In turbine blades, a rapid reduction in gas temperatures can
result in a thin trailing edge of the blade cooling faster than a thicker mid-span and
leading edge region. This difference in cooling rates can result in a difference in thermal
contraction between the trailing edge and the rest of the blade resulting in the creation of
stresses at the trailing edge region. TMF can also lead to its own form of cracking, when
compared to LCF and HCF, TMF cracks are more of a function of thermal gradients and
local distortions in blade shape. TMF is usually prevented by retiring the blade after it
has reached a set number of cycles and also limiting the severity of thermal strains.
2.4.2.4 OXIDATION
Oxidation is a phenomenon within the environment in which metals and alloys
that are exposed to oxygen at elevated temperatures will convert some or all of the
metallic elements of the metals into their respective oxides. These oxides can be both
beneficial as well as detrimental to the base alloy. To be beneficial, the protective oxide
scale must form and remain adherent, thus reducing the formation of further oxidation
and protecting the base metal. On the other hand most oxidation is detrimental to the
base alloy due to the oxide scales continually spalling off of the part, exposing fresh
metal for the next round of oxidation to attack resulting in progressive material loss.
This loss of material can have negative implications resulting in the loss of load-bearing
capabilities of the material that can lead to eventual component failure. Due to the
Page: 19
extreme operating temperatures that HPT blades operate within they are susceptible to
this high temperature oxidation. Oxidation problems and failures for HPT blades often
occur due to the formation of the brittle oxide surface layers on the external and internal
surfaces on the blades. These oxide layers can then lead to the premature fracture and
additional damage due to fatigue loadings.
2.4.2.5 HOT CORROSION
Hot Corrosion is an accelerated attack against the base metal of a HPT blade that
can occur due to the presence of contaminants in the hot gas environment. This type of
degradation mode can happen on both the internal and external surfaces of the turbine
blade. Some of the main elements that can cause hot corrosion include sulfur combined
with either sodium or potassium salts or independent vanadium and lead.
When
compared to clean environments, environments that contain these elements interfere with
the formation of protective oxide layers and thus, the rate of attack is accelerated. These
corrosive elements can originate through either the environment via the air intake or the
fuel of the engine in the combustion process. In research there are two main classes of
hot corrosion which are divided into the temperature ranges at which they occur. These
are classified as either Type I or Type II occurring between 1600°F to 2000°F and
1200°F-1500°F respectively. A more detailed section on this degradation mode can be
found later in this paper.
2.4.3
MATERIALS
Most modern day HPT blades are typically cast hardware created with the use of
different superalloys and the investment casting process. A superalloy is a metallic alloy
which has been developed to resist higher temperatures than basic metals, sometimes
until 80% of the absolute melting temperature is reached. All of these alloys tend to
have excellent creep, corrosion and oxidation resistances as well as good surface
stability and fatigue life. The main superalloys used in industry today are typically
Nickel, Cobalt and Nickel-Iron based due to their advancements in material properties
and survivability in harsh conditions within the turbine section of the engine, a
Page: 20
comparison of operating temperatures of these alloys to other alloys can be seen below
in Figure 10. For this project nickel based superalloys will be the focus.
Figure 10: Temperature capabilities of several classes of alloys [3]
Nickel-based superalloys are typically strengthened by either a solid-solution or
through precipitation hardening. Solid solution strengthened superalloys are used more
for applications requiring modest strengths like burners and combustors within gas
turbines. These components have particularly high-temperature corrosion resistance,
excellent fabricability and weldability, but lower mechanical strength. Precipitation
strengthened superalloys tend to be utilized in more demanding applications within the
gas turbine engines that requiring high-temperature strength and good corrosion and
creep resistance, these applications are typically turbine blades and vanes. [28]
Most nickel-based superalloys tend to contain 10-20% Chromium (Cr), up to 8%
Aluminum (Al) and Titanium (Ti), 5-10% Cobalt (Co) and small amounts of Boron (B),
Zirconium (Zr) and Carbon (C). However there are also some other common elements
that are added such as Molybdenum (Mo), Niobium (Nb) and Tungsten (W), all of
which play dual roles as strengthening solutes and carbide formers while Chromium and
Aluminum are necessary to improve surface stability. The major phases present in most
nickel based superalloys are as follows: the gamma (γ) matrix, gamma prime (γ’),
Page: 21
gamma double prime (γ’’), grain boundaries, carbides and topologically close-packed
phases. [9]
a) The Gamma (γ) Matrix - Is a continuous matrix that is a face-centered-cubic (FCC)
nonmagnetic phase that usually contains a high percentage of solid-solution elements
such as cobalt, iron, chromium, molybdenum and tungsten.
All nickel-based
superalloys contain this phase as the matrix
b) Gamma Prime (γ’) – This phase is the primary strengthening phase in nickel-based
superalloys. This is a coherently precipitating phase with an ordered FCC crystal
structure. This means that the crystal planes of the precipitate are in registry with the
gamma matrix. This close match in the matrix/precipitate lattice allows the gamma
prime to precipitate homogeneously throughout the matrix and have long lasting
stability. Gamma prime is also quite ductile and imparts strength to the matrix
without lowering the fracture toughness of the alloy. Aluminum and Titanium are
major contributors and are added in amounts and mutual proportions to precipitate a
high volume fraction in the matrix. This phase is required for high-temperature
strength and creep resistance within the superalloy.
c) Gamma Double Prime (γ’’) – In this phase nickel and niobium combine in the
presence of iron to form body centered tetragonals (BCT), which are coherent with
the gamma matrix, while including large mismatch strains. This phase provides very
high strength at low intermediate temperatures, but is unstable at temperatures above
about 1200°F.
d) Grain Boundaries – A film of gamma prime can be found along the grain
boundaries in stronger alloys and is produced by heat treatments and service
exposure. This film is believed to improve rupture properties.
e) Carbides – When carbon is added in amounts of about 0.02 – 0.2 wt% it combines
with reactive elements such as titanium, tantalum, hafnium and niobium to form
metal carbides. During heat treatment and service these metal carbides tend to
decompose and generate other carbides which tend to form grain boundaries.
Carbides in solid-solutioned alloys may form after extended service exposures.
These carbides all have a FCC crystal structure and tend to increase the rupture
strength of the alloy at higher temperatures.
Page: 22
f) Topologically close-packed (TCP) type phases – These are plate–like or needlelike phases that are generally undesirable, brittle phases that are formed during heat
treatment. The structures of these phases consist of close-packed atoms in layers
with relatively large interatomic distances one below the other. The characteristic
topology is generated by layers of close packed atoms being displaced from one
another by sandwiched larger atoms. The resulting structure is called topologically
close-packed (TCP), whereas the gamma prime phase is close-packed in all
directions and called geometrically close-packed (GCP). Superalloys containing
transition metals such as tantalum, niobium, chromium, tungsten and molybdenum
are the most susceptible to the formation of TCP phases. TCP phases tend to cause
lowered rupture strength and ductility.
The investment casting process is a casting method that is generally utilized
when complex components require precision and excellent surface finishes.
This
process is known for its ability to keep tolerances generally to ±0.003 in. with section
thicknesses as low as 0.05 in. or less while the resulting parts have excellent surface
finishes. The following table, Table 4, shows a summary of the 11 steps to this casting
process followed by an example flow map in Figure 11.
Table 4: Steps for an Investment Casting Process for Turbine Airfoils
Step
Step Description / Quality Impact
Die
Construction
- A die is created for the external shape of the wax pattern
- Another die is created for the internal passage ceramic core
- These are typically made out of tool steel and heat treated
Quality Impact:
- The die quality impacts how many parts can be made
- These dies have a limited life due to die wear
2
Wax Injection
- Wax is injected into the die for the wax pattern forming the shape of the desired casting
- For internally cooled parts, the ceramic core is placed inside the wax pattern prior to having wax injected
- The wax is then injected and it encapsulates the ceramic core with wax
Quality Impact:
- If there is any shift in the ceramic cores the wall thickness in the wax dies will be out of tolerance
- These errors in wall thicknesses can cause the cast parts to be unusable and considered scrap
3
Wax
Assembly
- An appropriate number of wax patterns are attached to a "tree" with a pouring cup, sprue and risers
Quality Impact:
- If the assembly of the tree has any geometry defects than issues can arise in the pouring process, resulting
in problems with the final castings
1
Page: 23
4/5
6
7
8
9
10
11
Slurry
Coating &
Stuccoing
- The wax tree assembly is then invested (coated) with ceramics of various sizes by dipping the assembled
tree in slurry, applying ceramic granules and then drying the assembly
- This is done under very controlled environmental conditions. The wet bulb and dry bulb temperatures are
closely monitored
- The ceramic granules are typically known as stucco and are applied so the mold can easily be broken from
the finished part. This is a refractory powder held together by a binder that disintegrates at firing
temperatures
Quality Impact:
- The cooling of the castings is directly related to the ceramic coating quality
- Uneven thicknesses can cause cracking during the cooling process
- Using the wrong stucco in the wrong dip steps can cause bad surface finishes
De-Waxing
- The invested tree is then dried after slurry and stucco coating and the wax is burnt out
- The remaining shell is then fired to sinter the ceramic particles together this results in a mold tree with a
series of individual article molds attached to it
Quality Impact:
- If firing is completed at the wrong temperatures or for the wrong durations the ceramic molds could break
resulting in the whole tree being unusable
Casting
- The liquid material is then poured into the ceramic mold often through a dross filter at specific conditions
for the given alloy
- The mold is then removed from the casting furnace
Quality Impact:
- If process times and/or temperatures are wrong casting defects can occur such as uneven cooling, uneven
fill and cracking
Shell
Removal
- The raw castings are then broken from the mold by mechanical and/or chemical means
- A caustic solution in an autoclave can also be utilized at elevated temperatures
Quality Impact:
- Care must be taken during removal since the castings could potentially break during a mechanical removal
of the shell
Cut Off
- Cut the sprues, gates and risers from the article and eliminate any residual gate material
Quality Impact:
- Care must be taken during cut off since the castings could potentially break during a mechanical cutting
operation on the casting
Finishing
- Clean and finish the article by removing excess casting stock and to meet customer requirements on the
surfaces using methods that may include grinding, polishing, blasting, and/or media finishing
- For parts with internal cores the internal ceramic core must be leeched out with a caustic solution. Excess
casting stock must be removed per customer requirements
Quality Impact:
- When leeching all chemical residual and caustic solution must be removed from all surfaces
- This can have an impact on part life at elevated temperatures
- Finishing small areas that is done by hand leaves room for operator errors
Inspection
- Inspect the finished part to the standards set by the customer by non-destructive means such as visual,
fluorescent penetration, radiography, ultrasonic (for wall thickness in hollow parts), dimensional gaging, xray diffraction and grain etch
Quality Impact:
- If any defects are missed the part may not perform to desired specifications resulting in poor customer
feedback
Page: 24
Figure 11: Investment Casting Process Flow Map [9]
Since the operating temperatures of HPT airfoils can be extreme, there is concern
that excessive creep can occur within these cast hardware. Realizing that turbine life
could be significantly challenged due to the high centrifugal stress levels and the
extreme gas temperatures improved casting techniques needed to be implemented. In
the mid 1950’s, Equiaxed cast turbine blades were introduced, offering much more
design flexibility for internal cooling configurations. However this method resulted in a
material structure that had a large number of grain boundaries with a random crystal
orientation.
These turbine airfoils generally undergo a great amount of centrifugal
loading and have the chance to elongate as the grains in these Equiaxed parts slip past
one another, potentially leading to a failure due to creep. To help prevent this grain
boundary slip and potential creep of turbine parts, a new casting method was developed
in the late 1970’s and is called the Directional Solidification process. In this method the
Page: 25
crystals are grown up the blade in order to align the maximum strength capability with
the direction of centrifugal stress, resulting in significantly fewer grain boundaries that
are all close to the radial direction. In this casting process, the casting tree is removed
from the furnace at a controlled rate such that the grains are oriented in the direction of
the cooling. To further improve the creep and material properties of these parts, the
Single Crystal process was introduced.
This process is based on the Directional
Solidification process; however, a single crystal is separated out in the beginning of the
process with the use of a helix shaped seed grain selector. This separated grain is then
grown radially from the blade root resulting in a structure that has no grain boundaries
and is aligned with the radial centrifugal stresses.
With the removal of all grain
boundaries the creep and material properties are further increased beyond what the
Directionally Solidified parts produce. Figure 12 shows a comparison in the grain
structure of these three casting processes.
Figure 12: Improvements in Casting Processes [5]
Page: 26
2.4.4
PROTECTION MODES AND COOLING SCHEMES
HPT blades are required to perform and survive long operating periods at
temperatures above their melting point. Various protection modes along with internal
and external cooling techniques are employed to bring down the temperatures of the
blade material to ranges below its melting point. Some of these techniques are the
application of coatings to the alloys, the design of complex internal cooling passages and
the addition of film cooling.
2.4.4.1 COATINGS
Coatings provide protection from the surrounding environment of the blade by
the application of a thin layer of material, either ceramic, metallic or a combination.
This surface layer is capable of protecting or inhibiting direct interaction between the
substrate and the surrounding damaging environment. The damage endured due to the
surrounding environment can either be metal recession due to oxidation or corrosion, or
a reduction in substrate mechanical properties due to the diffusion of harmful species
into the substrate alloy at these elevated temperatures. These coatings do not function as
inert barriers on the parts; instead they provide protection by interacting with oxygen in
the environment to form dense, tightly adherent oxide scales that inhibit the diffusion of
damaging elements. Due to this these coatings tend to by rich in elements like Al, Cr or
Si that readily participate in the formation of protective scales.
Modern day gas turbine blades have several different types of coatings and
coating processes available to protect the blades surface. However even with such a
wide selection most of these coating processes still have the same requirements. The
most important requirements for a coating are the oxidation and corrosion resistance
along with the adhesion to the substrate, phase stability and structural properties. Figure
13 below shows a basic tree of coating processes, as one can see there are many different
methods and thus an optimum method for an application all depends on that particular
application. When talking about a modern day HPT blade the most common coatings
are those following the Thermal Barrier Coating (TBC) process.
Page: 27
Figure 13: Common Coating Processes [3]
Modern day TBC’s are required to not only limit the heat transfer through the
coating but to also protect the engine components from oxidation and hot corrosion
[14].
When all coating processes are studied, no single coating composition appears to be able
to satisfy these multifunction requirements. As a result, a “coating system” has evolved
consisting of a metallic bond coat applied to the substrate followed by a ceramic top coat
applied to the bond coat. This coating system requires the bond coat and top coat to
work together to protect the engine component from all forms of degradation. A sample
schematic of a TBC coating system can be seen in Figure 14.
The top coat of the system is a thermally protective TBC layer with a low
thermal conductivity that is essential to maximizing the thermal drop across the
thickness of the coating. This portion of the system typically has a thermal expansion
coefficient that differs from the substrate to which it is applied. Therefore this layer
should have a high in-plane compliance to accommodate the thermal expansion
mismatch between the TBC and the underlying superalloy. In addition, the coating must
also be able to retain this property and its low thermal conductivity during prolonged
environmental exposure within the engine. For TBC applications, a porous, columnar
100-200 µm thick, yttria stabilized zirconia (YSZ) layer is typically preferred, as seen in
Figure 14. This layer may be applied utilizing either air plasma spray (APS) or electron
beam physical vapor deposition (EB-PVD). [14]
The second layer of this system is a metallic bond coat that is oxidation and hot
corrosion resistant and required to protect the turbine blade from environmental
degradation. This layer is required to remain relatively stress free and stable during long
term exposure and remain adherent to the substrate to avoid premature failure of the
Page: 28
coating system. It is also very important that this layer provide an adherent surface for
the TBC top coat. Normally, to help provide this adherence, a thin (<1 µm) protective
aluminum rich oxide is thermally grown on top of the bond coat, this layer is typically
know as the Thermally Grown Oxide (TGO) and can be seen in Figure 14. Since the
aluminum content of modern nickel based superalloys is not typically high enough to
form a fully protective alumina scale, an aluminum rich layer is applied so that the TGO
may form. A ~50 µm thick layer of either low sulfur platinum aluminide or MCrAlY
(where M is Ni or Co) is utilized for this purpose. This bond coat layer is typically
applied utilizing either low pressure plasma spray (LPPS) or pack cementation. [14]
Figure 14: A Schematic Depicting a Modern Day TBC Coating System [14]
Page: 29
2.4.4.2 INTERNAL COOLING PASSAGES
Modern day gas turbine blades incorporate various methods of internal and
external cooling. These methods utilize a variety of shapes and designs of internal cores
that cool the base alloy of the component from the inside, while other film cooling
methods are utilized to lay a protective barrier of cooling flow on the external surface of
the component. A sample schematic of a modern gas turbine blade with common
cooling techniques can be seen below in Figure 15. This schematic shows several
techniques utilized to cool a modern gas turbine blade. This particular blade consists of
a serpentine cooling passage lined with rib turbulators (trip strips), jet impingement is
used to cool the leading edge and pin-fin (pedestal) cooling along with ejection is used to
cool the trailing edge.
Figure 15: Schematic of Common Cooling Techniques [13]
Page: 30
As seen within Figure 15 the three main areas of a turbine blade, the leading
edge, mid body and trailing edge are all cooled internally with the use of different
cooling techniques. For mid body passages trip strip cooling is the most frequent
method utilized due to their capability to enhance the heat transfer of the internal cooling
passages. These trip strips are cast into the hardware typically on opposite walls of the
cooling passages to help increase the heat transfer around the passage. Heat is then
conducted from the pressure and suction side surfaces of the blade through the walls and
is transferred to the cooling flow passing internally through the blade. A few different
sample configurations of trip strip layouts can be seen in Figure 16. Impingement
cooling is a very aggressive cooling technique that is most commonly utilized near the
leading edge of airfoils where the heat loads are typically the greatest. In this technique,
impingement jets are utilized to strike the blade wall removing heat in a very effectively
method. As seen in Figure 17 the jet holes are placed perpendicular to the wall that is
going to be cooled so that the internal cooling air can be ejected through the jets and
onto the surface. Lastly, the tailing edge of the blade is generally cooled with the use of
pin-fin or pedestal cooling. This is due to the manufacturing constraints that can arise
due to the narrow trailing edge sections. This method is typically used to enhance the
heat transfer from the blade walls within the region, similar to the trip strip method. In a
pin-fin array, heat is transferred from both the smooth channel end wall as well as the
numerous pins within the passage. The coolant in this passage will flow around each of
these fins and is comparable to the flow around a single cylinder. As the coolant flows
around these pins the flow separates and develops wakes downstream of the pin, in
addition a horseshoe vortex is formed just upstream of the base of the pin wrapping
around the pin. This horseshoe vortex creates additional mixing within the passage
enhancing the heat transfer. A sample layout of a pin-fin design can be seen below in
Figure 18.
Page: 31
Figure 16: Sample Trip Strip Configurations [13]
Figure 17: A Sample Impingement Cooling Configuration [13]
Page: 32
Figure 18: A Sample Pin-Fin Cooling Configuration [13]
2.4.4.3 FILM COOLING
Film cooling is a widely utilized blade cooling technique that is a very important
component in the overall cooling of turbine blades. This process allows coolant to pass
from the internal cooling passages to the external blade surface with the use of a variety
of holes, typically known as “cooling holes”, placed within the blade. The injection of
this coolant onto the external surface of the airfoil allows for the creation of an insulating
layer of “film” of the coolant flowing along the external surface of the airfoil protecting
it from the hot mainstream gas. The basic idea behind the film cooling process is to
utilize the coolant from the internal passages to reduce the heat transfer to the wall. This
is accomplished since the temperature of the coolant flow is lower than the mainstream
gas temperature, thus reducing the mainstream gas temperature near the wall and
effectively reducing the amount of heat transfer to the blade. A sample blade utilizing
film cooling techniques can be seen in Figure 19 along with some 2-D samples of
cooling hole designs.
Page: 33
Figure 19: A Film Cooled Blade and Sample Cooling Hole Designs [1]
Unfortunately this method of film protection comes at a price to the engine. The
source of the coolant air is bleed-off air from the last stage of the compressor section of
the engine. This high pressure air bypasses the combustor section of the engine and thus
maintains a much lower temperature than the core turbine flow. This bled-off air
however is removed from the core mass flow and thus is removed from the overall thrust
of the engine. This must always be taken into account when designing both external
film cooling techniques as well as the internal cooling techniques covered within the
previous section of this report.
2.5 HOT CORROSION
When a superalloy component such as a HPT blade is operating at elevated
temperatures that are in the range of 1200°F to 2000°F one of the most common forms
of degradation that is experienced is hot corrosion.
Hot corrosion is a form of
accelerated oxidation of an alloy which occurs when a normally protective oxide scale is
degraded or destroyed by a salt deposit and unable to reform. This tends to happen in
high temperature combustion gases in the presence of certain impurities. In gas turbine
engine applications, hot corrosion problems are usually associated with the presence of
molten corrosive agents that contain sulfur, chloride and sodium. Sulfate salts tend to be
more prevalent in these environments however depending on the purity of the fuels used
fuels can also be a source of salt in the absence of an aggressive atmosphere. A
Page: 34
corrosive attack of the substrate will occur when these molten salts are able to defeat the
protective oxide scales such as aluminum oxide or chromic oxide.
Once an alloys surface has been partially or completely covered by the corrosive
molten salts, the conditions for severe corrosion can develop.
For a given set of
exposure conditions, the rate of attack for different alloys can differ by more than three
orders of magnitude. Each alloy corrosion rate varies due to the operating environment,
typically a given alloy corrosion rate depends on several variables such as: composition
and velocity of the atmosphere, composition and thickness of the salt deposit, rate of salt
deposition, operating temperature, thermal cycling and most important time.
The degradation of the aluminum oxide or chromic oxide can occur by several
mechanisms.
These include penetration and/or stripping of the scale by chemical
reactions by the salts with the oxide scale. This is typically known as fluxing and is a
reaction involving the dissolution of the oxide scales to form products such as
aluminates and chromates. A necessary condition for fluxing to occur is that the salt be
liquid or that the liquid be formed as a result of a solid-solid reaction. There are two
main types of fluxing that can occur, basic fluxing and acidic fluxing. Basic fluxing is
when the oxide scale on the alloy reacts with Na2O and dissolves in the salt as an anionic
species. While acidic fluxing is when the oxygen ion concentration of the melt is low
compared to the value required to maintain equilibrium in the detachment reaction of the
metal oxide.
As previously stated, hot corrosion is a complex process that is the result of high
temperature exposure to sulfur and salts. These sulfurs and salts can directly be related
to the global environments covered in a previous section of this report. Operation within
aggressive regions like marine and industrial will result in increased sulfur and salt
deposits and thus increase the chances of hot corrosion. The specific chemical processes
of hot corrosion however vary according to the temperature exposed to. It is common
for hot corrosion to be split into two different temperature ranges and called Type I and
Type II hot corrosion. Type I hot corrosion typically occurs at a higher temperature then
Type II hot corrosion. A basic schematic illustrating the effect of temperature on the
rate of hot corrosion damage can be seen within Figure 20. This double hump curve is
superimposed on the contribution due to oxidation to damage as well. The defining
Page: 35
temperature separating Type I and Type II hot corrosion typically is labeled as the
melting point, Tmp, of salt. Type I tends to occur above the Tmp while Type II occurs
near or below the Tmp but in the presence of small amounts of SO3.
Figure 20: Effect of Temperature on Type I and Type II Hot corrosion [3]
2.5.1
TYPE I HOT CORROSION (HIGH-TEMPERATURE)
High-Temperature hot corrosion also known as Type I generally occurs in the
temperature range of 1600°F to 2000°F.
However the high end temperature has
sometimes been set at 1850°F due to the molten salts becoming unstable above this
temperature and causing them to spall off the surface of the alloy and burn. This type of
hot corrosion generally proceeds in two stages: an initiation period exhibition a low
corrosion rate followed up by an accelerated corrosion attack. The initiation period is
typically related to the formation of protective oxide scales on the alloy while the
accelerated corrosion attack is related to the breakdown of the protective oxide scales. A
basic schematic of the stages of hot corrosion can be seen below in Figure 21.
Page: 36
Figure 21: Schematic diagram of the stages of Hot Corrosion [9]
Alloys that exhibit this type of hot corrosion typically undergo basic fluxing
during the initiation stage of the attack. The initiation stage begins with the consumption
of oxygen in the sulfate melt, this oxygen depletion leads to increases in the
thermodynamic activity of sulfur. During this initiation stage the sulfur diffuses through
the oxide scales to form sulfides in the alloy beneath the oxide scales. However the
composition or type of sulfides formed is different between components and depends on
the composition of the material. These may be sulfides of nickel, cobalt, aluminum or
chromium. Depending on the quantity of salt, temperature and the type of oxide scale
the duration of the initiation period may be only a few seconds. The disruption of the
oxide scale signals the transition to the propagation of the accelerated stage of attack.
For nickel alloys corrosion continues as long as sulfate deposition continues. Alloys
with chromium and aluminum allow the spread of hot corrosion through the oxidation of
sulfides in the alloy zone beneath the oxide scale. This sulfur-induced mode of attack is
sometimes referred to as sulfidation.
Cobalt binary and ternary alloys containing molybdenum or tungsten are not
susceptible to basic fluxing; however these are more susceptible to acidic fluxing. In
Page: 37
acidic fluxing oxides react with sulfates to produce strongly acidic conditions, this
occurs when the protective oxides are no longer formed on the alloy. This process is
self-sustaining when compared to basic fluxing because the decomposition of the salt is
not required; the molten salt becomes a solvent for the dissolution of oxides for all the
elements in the alloy. This results in very little alloy depletion and sulfide formation
ahead of the corrosion front.
2.5.2
TYPE II HOT CORROSION (LOW-TEMPERATURE)
Low-Temperature hot corrosion also known as Type II occurs due to acidic
fluxing over the 1200°F-1500°F temperature range. This acidic fluxing arises from the
inherent acidity of the as-deposited sulfates whose acidity is caused by the partial
pressure of sulfur in the combustion gas stream. This type of hot corrosion is typically
seen in industrial and marine turbine applications where fuel is not as clean as in
aerospace applications. The dirty fuels in the industrial and marine applications tend to
produce more sulfur dioxide and sulfur trioxide during the combustion process. The
presence of the sulfur-oxide gases causes the equilibrium between the sulfur oxides to
shift to higher sulfur trioxide pressures at lower temperatures.
This increases the
tendency for gas-phase acidic fluxing at lower temperatures.
The initiation phase of acidic fluxing involves a reaction with the oxides on the
alloy surface with solid sulfate deposits to form liquid sulfate solutions containing
aluminum, nickel, cobalt or refractory metal ions. However chromium oxide scales tend
to be much more resistant to this type of hot corrosion attack. This initiation is often
localized and characterized by pitting attacks associated with refractory metal carbides
that intersect the surface. This pitting corrosion tends to be observed on turbine blade
roots.
Page: 38
3. METHODOLOGY
3.1 PROBLEM DESCRIPTION
In modern day gas turbine design, it is hard to predict the corrosion that an
engine component may undergo throughout its given life. The idea behind this project is
to see how well corrosion in field run hardware correlates with predicted temperatures
during the design of the part. For the purpose of this project the areas of focus were
internal corrosion within a high pressure turbine blade. To do this, a HPT blade first
needed to be selected from a modern day engine that has known experience of internal
corrosion. Once selected, the known field distress of the blade due to corrosion can then
be correlated back to the predicted metal temperatures that were calculated during the
design process. These temperatures along with the known operating environments can
then be used to determine if there is any correlation between the internal corrosion and
the predictions.
The blade selected for this project is currently in an operational engine and
operates all over the world experiencing a variety of the different environments
discussed previously within this paper. For the purpose of this project, the selected
environmental regions to be analyzed are an India Environment and a Middle East
Environment. These environments were selected due to the known corrosion issues
within hardware operating within these regions. The blades selected experienced
significantly long periods of time in service operation. The aggressive exposure to
contaminants at long periods of time was postulated to provide and produce a worst case
scenario for hot corrosion. The selected blade also incorporates some of the cooling
schemes discussed preciously within this paper. The following sections will cover the
methodology used for this study and will give explanations of the selected part, proposed
environmental factor analysis, part cut-up analysis and proposed thermal model analysis.
3.2 SELECTED PART OVERVIEW
The part selected for this project is a High Pressure Turbine Blade currently
designed and produced by Pratt and Whitney. This blade is cast as a Single Crystal
Directional Solidified part utilizing an investment casting process and a nickel based
Page: 39
superalloy known as “Material A”. This material is comprised of various elements
including Cr, Co, W, Ta, Ti, Al and B with a melting temperature range of
approximately 2350°F-2450°F. This high melting temperature range shows why this
material is perfect fit for use within the turbine section of a gas turbine engine. Table 5
below shows a comparison between “Material A” and other alloy compositions.
Table 5: Composition of various Nickel Based Superalloys [20]
Since the selected blade for this project operates within the high pressure turbine
section of the engine, extreme temperatures are exposed to the external surfaces of the
blade. To help combat these temperatures, the blade design incorporates some of the
internal and external cooling schemes previously covered in this paper. The internal
cooling schemes utilized incorporate an internal core composed of three separate
circuits; a leading edge core, a mid body core and a trailing edge core. The leading edge
core utilizes impingement cooling along with trip strips throughout the core followed by
tip cap cooling ejecting out the trailing edge. The mid body core is a three pass
serpentine design that utilizes trip strips throughout the length of the core. Lastly, the
trailing edge core utilizes both trip strips and pedestal arrays for cooling along with flow
ejecting out the trailing edge. In all three cores, an internal diffusion coating is also
utilized to help protect the base alloy against corrosion damage within the part. Along
with these internal cooling schemes, external cooling techniques are also utilized to help
reduce the external temperatures of the blade. The two main methods utilized on the
external surfaces of the blade are a TBC coating system along with film cooling. A
basic schematic of the internal design of the blade can be seen below in Figure 22 with
the external cooling designs being very similar to the pictures in the previous sections.
Page: 40
Figure 22: Schematic of Common Cooling Techniques [13]
3.3 ENVIRONMENTS SELECTED
As previously stated in this report, the blades selected for this project have come
from an engine operating within either India or the Middle East. To help determine
different environmental factors within each of these environments that could be harmful
to both the engine and selected parts, a request was made to the engine operators for a
complete flight history for the two different engines. This flight history would provide
the different city pairings that the engine completed over it lifetime. This will supply a
more detailed region of operation to analyze for environmental data.
The company operating the Middle East engine was unable to provide the data
that was requested on the engine flight history. Due to this it was decided to select a few
Page: 41
different high air traffic city locations within the Middle East for this analysis. These
cities were selected based on highly populated regions and known major airports. The
locations of the cities selected for this analysis in the Middle East can be seen below in
Figure 23 and were as follows: Dubai United Arab Emirates, Cairo Egypt, Riyadh Saudi
Arabia, and Doha Qatar.
Figure 23: Map of Selected Cities in the Middle East [29]
The company that operated the engine in the India region was very helpful with
the request and was able to provide the full flight history of the engine and every city
pair that it had completed. This data was able to help narrow down the city locations
that would then be selected for this region. The breakdown of the data provided can be
seen below in Figure 24 with each airport IATA (International Air Transport
Association) code plotted.
Page: 42
Figure 24: Plot of Airport Visit Data for the Selected India Engine
From this data the cities selected were as follows: New Delhi (DEL), Mumbai
(BOM), Kolkata (CCU), Hyderabad (HYD) and Bangalore (BLR). All the selected city
locations can be seen below in Figure 25 on a map of India.
The analysis of the environmental data of these selected regions and city
locations was completed with the help of a working relationship between Pratt &
Whitney and Iowa State University. It was requested of Iowa State to put together a
comparison of air quality and pollution between the selected cities in the Middle East
and India and baseline this comparison to major cities in the USA and China. This
comparison will show the average concentrations of ambient SO2 as well as the
characterization and concentrations of particulate matter within these regions. This data
will help determine how corrosive the India and Middle East environments are in
comparison to other regions of the world.
Page: 43
Figure 25: Map of Selected Cities in the India [17]
3.4 CUT-UP INFORMATION
The next step of this project was to obtain the information from actual engine run
hardware. To do this 10 engine run parts have been gathered; 5 parts having operated
within the Middle Eastern Region and 5 parts having operated in the India Region.
These parts were then cut up and analyzed to see what type of corrosion, if any, is
experienced on the internal cores. To complete this analysis, the parts must first pass
Page: 44
through a variety of operations to identify the best areas to cut/analyze and also what
should be studied once cut. The operations to achieve these desired results are listed
below with a pictorial representation found following in Figure 26:
1) Macro images were taken of pre-cut state
 This allowed for comparison back to pre-cut state once the data is obtained
2) A Magnetoscop analysis was performed on all the parts to determine the best areas
to have the parts section cut for analysis
 A Magnetoscop is typically utilized in the determination of sulfidation in
aircraft turbine blades. This device takes permeability measurements and is
able to determine the areas where corrosion has occurred inside the blade
just by scanning the outside surfaces of the blade.
 This will allow for “high” corrosion readings/areas on the blade to be
marked, helping locate where the section cuts should be taken
3) Once the blades were marked for high Magnetoscop readings, six parts were
selected for section cuts
 Three parts from the Middle East and three parts from India
 Each of the six parts were section cut in the same two locations
 The cut locations were determined by the high Magnetoscop readings
 These cut locations also determined the cut locations for the thermal
analysis
4) Once the parts were section cut, each cut was mounted so that it could be analyzed
 Analysis
was
performed
under
magnification
so
that
accurate
measurements are recorded
5) The mounted parts were analyzed for the following:
 Corrosion Zones – Large areas / Small pitting locations
 Wall depletion at corrosion locations
 Internal Coating depletion
Page: 45
Figure 26: Pictorial View of Cut-Up Analysis [2]
3.5 THERMAL MODEL
A thermal analysis model was developed for this blade to understand the
temperatures experienced over a given mission. This thermal model was executed for
various mission points including takeoff, climb, cruising, decent and landing with each
of these points resulting in a range of thermal solutions for the blade. Each mission
point had different impacts on the blade and imposed different temperatures, pressures,
coolant flows and rotational speeds. These solutions showed the blade experiencing a
variety of internal and external predicted temperatures and helped to understand what
temperatures the internal cooling cavities are designed to experience. The mission
points selected are in line with typical missions experienced by the selected HPT blades
engine in both the Middle East and India regions.
The thermal model used was generated at Pratt and Whitney and utilizes a
boundary condition mapping software along with an ANSYS mesh to generate a thermal
solution. Within the model, reference surfaces have been developed for the internal
cavities, these surfaces were utilized to report back the predicted metal temperatures of
each cavity at each mission point. These reference surfaces were then section cut at the
same radial locations determined during the part cut-up analysis previously covered.
Page: 46
Once the section cuts were created, the predicted metal temperatures at each mission
point could be summarized at the locations of interest. These temperatures were then
compared to the areas of corrosion distress analyzed during the part cut-ups and see if
there is any correlation between the predicted temperatures and the distress experienced.
Page: 47
4. RESULTS AND DISCUSSION
4.1 ENVIRONMENT ANALYSIS RESULTS
Iowa State University was able to perform the requested analysis for most of the
cities asked of them in the India and Middle East regions. The first step of the analysis
that was completed was to define particulate matter and the difference between the two
major categories that they may fall within. The two main categories are PM2.5 and PM10.
PM10 typically consists of more course particles that are between 2.5 and 10 microns in
width that tend to originate from primary emissions like windblown dust, sea salt, fly ash
and road dust. PM2.5 particles on the other hand are finer and are typically less than 2.5
micros in width and tend to be largely formed from gases such as sulfates, nitrates and
ammonium. An example of size comparisons of PM2.5 and PM10 particles can be seen
below in Figure 27.
Figure 27: Comparison of PM2.5 and PM10 particles [16]
Page: 48
The atmospheric residence time of particulate matter is directly related to the size
of the specific matter in question. Ultrafine particles (< 0.1 microns) tend to coagulate
together while more coarse particles (> 10 microns) tend to settle out more rapidly.
Particles that are found in the 0.1 – 1.0 microns range, the prime range for sulfates, tend
to have the longest residence time within the atmosphere. This residence of different
particle sizes can be seen below in Figure 28 as a plot of Residence Time for different
particle radius sizes. From this it is seen that PM2.5 sulfates tend to reside for about 3 to
5 days within the atmosphere. Once the PM2.5 particles are within the atmosphere they
can be transported upwards to 1000km from the source of the originating gases. This
shows that a corrosive area/location does not need to be very close to a known corrosive
environment, the environment can expand into different regions and areas due to weather
patterns. An example of this is sea salts traveling inland to areas that are not close to the
ocean due to ocean winds.
Figure 28: Residence Time for different particle radius sizes [16]
When examined on a world wide scale, it was found that the regions with the
largest concentrations of PM10 were found in the Middle East and Southern Asia, Figure
29, while the regions with the largest concentrations of PM2.5 were found more in
Northern Africa, the Middle East, India and China, Figure 30. North America did not
Page: 49
tend to have as severe concentrations of PM10 and PM2.5 when compared to the rest of
the world.
Figure 29: Annual Average PM10 Concentrations in World Cities [26]
Figure 30: Annual Average PM2.5 Concentration in World Cities [19]
As seen in the above two figures, PM10 and PM2.5 concentrations are both high in
the Middle East and India regions. When broken down to the specific cities requested
data was able to be found on Delhi, Bangalore, Kolkata, Hyderabad, Mumbai, Ahu
Page: 50
Dhabi and Cairo. These findings were then compared back to New York and Beijing.
The highest concentration readings of PM10 were found in Delhi and Abu Dhabi; these
findings can be seen below in Figure 31. When looking at the change in concentrations
of PM10 between 1998 and 2010 Delhi and Mumbai have seen the largest increase
starting in the mid 2000’s, while Cairo and New York have stayed consistent and Beijing
has gradually decreased in concentration as seen in Figure 32.
Figure 31: Annual Average Concentrations of Ambient PM10 [22]
Figure 32: Trends of Ambient PM10 Concentration [22]
Along with particulate matter, sulfur oxides are a large contributor to corrosion
within gas turbine engines. Due to this, a similar study to the particulate matter study
above was completed for SO2 concentrations around the world and in the specific cities
Page: 51
requested, Figure 33 and Figure 34. From this analysis it is seen that there is a large
concentration of SO2 in the Middle East, specifically Cairo, and Eastern Asia. India did
not tend to have large concentrations of SO2 when compared to the concentrations of
PM10 within this region. When the trends of SO2 concentrations were examined between
1996 and 2010 it is seen that Delhi and Mumbai have consistently had lower SO2
concentrations than New York, while Cairo has had SO2 levels close to that of Beijing.
These SO2 trends can be seen below in Figure 35.
Figure 33: Annual Average Concentrations of Ambient SO2 [22]
Figure 34: Annual Average SO2 Concentrations in World Cities [26]
Page: 52
Figure 35: Trends of Ambient SO2 Concentration [22]
The engine history of blades operating within these regions of interest has shown
internal corrosion distress and after the analysis of these environmental factors it is
understood why. The review of this data leads to initial thoughts on the amount of
corrosion that is expected on the selected parts. The data shows that in the Middle East
PM10, PM2.5 and SO2 levels are very high leading to the expectation of internal corrosion
to some extent within these Middle East parts. The data for the India region also shows
high levels of both PM10 and PM2.5 however the SO2 levels are lower than the Middle
East. These levels would also lead to the expectation of some stage of internal corrosion
experienced on the blades from this region. Comparing the levels of contamination it is
expected to see more corrosion distress on the Middle East blades than the India blades
due to the higher SO2 levels. SO2 is very corrosive as previously covered within this
report, and the fact that the Middle East has higher SO2 levels would lead to the
increased corrosions expectation within these blades.
Due to the high levels of particulate matter in both regions the corrosion
generated could be due to the ingestion and deposit of this matter onto the internal cores.
As previously covered the internal geometry of this blade is complex leading to various
areas where particulate matter could collect. Serpentine turn passages, trip strip fillets
and pedestal fillets can be ideal locations for particulate matter to build up. These are
Page: 53
areas of interest when completing the cut-up analysis. When completing the thermal
analysis it is hopeful that temperatures are predicted within either the Type I or Type II
range. If these temperatures are predicted and knowing these environmental factors a
correlation may be able to be made between the corrosion experienced, the thermal
model and the region of operation.
4.2 CUT-UP ANALYSIS RESULTS
Once the ten parts were selected, six of them were separated out into two groups
to be cut-up; the other 4 parts were kept for additional studies. These six parts were then
split into groups based on the regions of operation: three from India and three from the
Middle East. These groups were then labeled with the following designations: India-(A,
B, C) and Middle East-(A, B, C).
The parts were then processed through the steps listed within the methodology
section of this paper. The first step was to perform the Magnetoscop analysis on each of
the six parts. This analysis showed that the best cut location for each of the blades was
at approximately 80% and 20% span. These spans were determined based on the three
highest Magnetoscop readings for each blade seen in Figure 36 below. These cut
locations were also utilized in the thermal analysis to determine the temperatures at these
locations (this is covered in the next section of the paper).
Page: 54
Figure 36: Magnetoscop Readings and Section Cut Locations [27]
The section cuts for each of the blades were then mounted in an acrylic button so
that they could be polished and the levels of corrosion examined. The same procedure
was completed for each of the six parts, with two cuts being taken resulting in 12 total
cuts (6 at 20% span – “ID”, and 6 at 80% span – “OD”). Once polished and examined at
increased magnifications, it became apparent that not as much corrosion was
experienced within the cores as expected. Each of the corrosion areas that were noticed
within the airfoil cavities was photographed and labeled based on its location within the
airfoil. The location within the airfoil is documented with each picture and can be
referenced on Figure 37. Figure 38 to Figure 40 show some good examples of the
corrosion witnessed, highlighted in red, on the parts that were run within the India
region.
Page: 55
Figure 37: Internal cavity numbering & Pressure / Suction side labeling
Figure 38: Blade Cut-Up: India-A, Location: OD, Cavity 1 SS
Page: 56
Figure 39: Blade Cut-Up: India-B, Location: OD, Cavity 5 PS
Figure 40: Blade Cut-Up: India-C, Location: OD, Cavity 1 PS
Page: 57
From the section cuts of India-A, the most severe corrosion witnessed is what is
shown in Figure 38 above. This area of corrosion is located in the first cavity of the
airfoil closest to the leading edge, typically a very hot region of the blade. Looking at
the corrosion itself, it is seen that there is some corrosion building up into the cavity as
well as some wall depletion into the base material.
Looking closely the internal
diffusion coating is still intact on the base material, however, is being attacked by the
corrosion. After inspection of the section cuts from India-B, it was determined that little
hot corrosion was evident. The picture shown in Figure 39 shows the start of some
corrosion in cavity 5 with some pitting. Lastly India-C shows a similar condition in the
leading edge cavity to that of India-A. This corrosion has progressed deeper into the
base material and, in some areas, has fully defeated the internal coating. Corrosion
buildup can easily be seen growing on the internal wall depleting the base metal of blade
wall thickness.
Figure 41 to Figure 43 show some good examples of the corrosion witnessed on
the parts that were run in the Middle East region. In these photographs, Middle East-A
shows a localized area of corrosion that has fully defeated the internal coating on the
cavity wall. Material buildup has started to take place in the corrosion area and the base
metal has started to be depleted. Middle East-B was similar to blade India-B where very
little to no corrosion was witnessed in the section cuts of the blades. Only small areas of
pitting on the internal coating were noted. Middle East-C showed an area of corrosion
attack that did not yet fully defeat the internal coating and attacked the base metal.
These results show that in the selected blades corrosion was evident; however, this
corrosion was not as large as what was expected due to the operating environments and
known engine history.
Page: 58
Figure 41: Blade Cut-Up: Middle East-A, Location: OD, Cavity 2 SS
Figure 42: Blade Cut-Up: Middle East-B, Location: ID, Cavity 6 PS
Page: 59
Figure 43: Blade Cut-Up: Middle East-C, Location: ID, Cavity 2 SS
To better understand the amount of material buildup and depletion in the areas of
corrosion, four parts were examined more in depth and measured. These parts were
Middle East-A (Two areas), Middle East-C (One area) and India-A (One area). During
examination three key measurements were observed and recorded. The first was the
approximate thickness of buildup in the corrosion area. The second was the approximate
remaining thickness, if any, of the internal coating. And lastly, the approximate
depletion of base metal was evaluated. Figure 44 below is for Middle East-A and is at
an increased magnification than Figure 41 above. In this picture it is seen that all of the
internal coating has been fully exhausted and the base metal is under attack. Measuring
from the approximated surface where the base metal should be, it is seen that the
corrosion has penetrated the base metal by about 0.0018”.
This also results in a
corrosion material buildup in the area of approximately 0.0010”.
Page: 60
Figure 44: Blade Cut-Up: Middle East-A, Location: OD, Cavity 2 SS - Measured
A second area from Middle East-A was examined on a more magnified level.
This area can be seen in Figure 45 and Figure 46 (magnified and measured) below.
From the measurements taken, it can be seen that some of the internal coating is still on
the base metal; however, the coating has been slightly defeated leaving about 0.0001”.
The penetration into the base metal is in the range of 0.0010” while the corrosion
buildup in the area is in the range of 0.0006”.
Page: 61
Figure 45: Blade Cut-Up: Middle East-A, Location: ID, Cavity 2 SS
Figure 46: Blade Cut-Up: Middle East-A, Location: ID, Cavity 2 SS – Measured
Page: 62
The more detailed measurement pictures for Middle East-C and India-A can be
seen below in Figure 47 and Figure 48 respectively. These are increased magnification
of previously shown Middle East-C (Figure 43) and India-A (Figure 38) pictures. From
the Middle East-C blade it is seen that the internal coating is still intact with a thickness
of about 0.0004”. The corrosion buildup on this coating however is in the range of
almost 0.0020”. However, it is hard to tell if this is all buildup on the surface of the
coating or if some of the base metal has also been defeated due to the coating. The same
scenario is seen in the India-A picture below as well. It looks like the internal coating is
still intact at a 0.0004” thickness and the corrosion material is building up on the coating.
Yet this could also be some base metal depletion as well.
Figure 47: Blade Cut-Up: Middle East-C, Location: ID, Cavity 2 SS – Measured
Page: 63
Figure 48: Blade Cut-Up: India-A, Location: OD, Cavity 1 SS – Measured
4.3 THERMAL ANALYSIS RESULTS
The next step of this project was to perform a thermal analysis for both the India
and Middle East regions. This was completed with the generation of a thermal analysis
model that could be run at various points along a given engines mission.
A sample
India and Middle East mission was obtained and the takeoff, climb, cruise, decent and
landing points were utilized within the thermal analysis. Each of the mission points
generated a 3D thermal model that was then section cut at two different locations,
approximately 20% and 80% span, which correspond to the same section cuts performed
on the actual engine hardware. Having the thermal and hardware section cuts as close as
possible to one another allows for a more accurate comparison between the two. A total
of 10 section cuts, 5 ID and 5 OD, were generated for each of the two regions. These
were then studied and compared back to the cut-up data discussed in the previous
section. For this comparison many different areas could be examined; however, to
simplify this for the purpose of this project only two locations were selected to be
studied and compared. The two locations chosen were the suction side (SS) walls of
Page: 64
cavities 1 and 2, these were determined based on the more detailed cut-up pictures
shown in the previous section. These pictures showed that corrosion was evident in
these locations and at levels that could be measured and examined. The two areas to be
examined at each of the section cuts can be seen below in Figure 49 shown on a sample
airfoil cross section.
Figure 49: Temperature Inspection Locations: Cavity 1 – Blue, Cavity 2 - Green
After review of all the 3D thermal models generated for the India region, it was
determined that the blades internal cores were operating within the temperature range of
Type II Hot Corrosion. Each of the ten section cuts were inspected and the predicted
metal temperature ranges for the two selected locations, Figure 49, were summarized as
seen below in Table 6. The temperature ranges within the table show that during the
takeoff, climb and cruise mission points the locations of interest are operating within the
Type II temperature range of 1200°F-1500°F. This is a large difference when compared
to the temperatures that the decent and landing points are operating within, these two
points are less than 1050°F or approximately 200°F below the lower limit of the Type II
Hot Corrosion range. From these thermal results, a conclusion can be made that the
corrosion that was seen on the engine hardware discussed in the previous section was
most likely initiated during either the takeoff, climb or cruise points of the mission. The
predicted temperatures show that the decent and landing mission points are far less
damaging thermally to the internal cores of the blade. This conclusion is supported
when the duration at each point of the mission is examined. The duration time of the
Page: 65
takeoff, climb and cruise points and an India mission shows that these three points
compromise about 78% of the operating time with the other 28% operated with the
decent and landing mission points. This supports that the internal cavities are exposed to
the hot corrosion temperatures for a majority of its operational time. A sample picture of
the India takeoff point thermal analysis can be seen below in Figure 50.
Table 6: India Thermals: Predicted Temperature Ranges at Section Cuts
India
Takeoff
Climb
Cruise
Decent
Landing
Units: °F
ID
Cavity 1 (SS)
1320°-1440°
1320°-1440°
1250°-1300°
< 1050°
< 1050°
OD
Cavity 2 (SS)
1440°-1520°
1320°-1440°
1250°-1300°
< 1050°
< 1050°
Type I Range ( 1600° to 2000° )
Cavity 1 (SS)
1300°-1500°
1350°-1500°
1250°-1450°
< 1050°
< 1050°
Cavity 2 (SS)
1450°-1500°
1400°-1500°
1250°-1450°
< 1050°
< 1050°
Type II Range ( 1200°-1500° )
Figure 50: India Takeoff Point Thermal Model Results
Page: 66
When the Middle East thermal models were examined it was seen that the
predicted temperatures have very similar temperature ranges as the temperatures in the
India models. These results also agreed with the India models showing that the blades
internal cores are operating within the Type II Hot Corrosion temperature range for the
majority of its operation. Similar to the India results the Middle East results were
summarized and the predicted temperature ranges were tabulated below in Table 7.
These predicted temperature ranges show that for the Middle East missions the takeoff,
climb and cruise points are operating within the Type II temperature range of 1200°F1500°F. While the decent and landing mission points are operated more than 200°F
below the lower limit of the Type II Hot Corrosion range.
Due to this a similar
conclusion for the Middle East thermal results can be made to the India results; that the
corrosion seen on the engine run hardware was most likely initiated sometime during
either the takeoff, climb or cruise points of the mission. Similar to the India analysis this
conclusion is supported with the examination of the duration at each mission point. The
duration time of the takeoff, climb and cruise points and a Middle East mission shows
that these three points compromise about 82% of the operating time with the other 18%
operated with the decent and landing mission points. This supports that the internal
cavities are exposed to the hot corrosion temperatures for a majority of its operational
time. A sample picture of the Middle East takeoff point thermal analysis can be seen
below in Figure 51.
Table 7: Middle East Thermals: Predicted Temperature Ranges at Section Cuts
Middle East
Takeoff
Climb
Cruise
Decent
Landing
Units: °F
ID
Cavity 1 (SS)
1300°-1500°
1350°-1340°
1200°-1300°
< 1050°
< 1050°
OD
Cavity 2 (SS)
1450°-1500°
1350°-1480°
1200°-1300°
< 1050°
< 1050°
Type I Range ( 1600° to 2000° )
Cavity 1 (SS)
1300°-1480°
1300°-1450°
1200°-1300°
< 1050°
< 1050°
Cavity 2 (SS)
1450°-1550°
1450°-1550°
1250°-1300°
< 1050°
< 1050°
Type II Range ( 1200°-1500° )
Page: 67
Figure 51: Middle East Takeoff Point Thermal Model Results
The thermal model temperatures did show that Type II hot corrosion can
potentially be predicted with the use of this type of thermal analysis. However it is
tough to fully predict the exact locations of corrosion based on this model without some
more data take from the engine run hardware. One piece of data that would be very
helpful would be to measure the metal temperatures that were reached on the engine
hardware within and around the corrosion locations. This data would help correlate the
predicted thermal results better to the engine run results; however this measured data
point was unable to be achieved within the timeframe of this project and is planned for
future work for better model calibration.
If this metal temperature data could be
achieved at various locations around the internal cores, a more detailed and calibrated
model could be generated for each region that could potentially better predict corrosion
zones along with the predicted metal temperatures.
During the review of the thermal results it was determined that service time was
a factor that was not considered in the thermal model that would have been very helpful
in its calibration to better predict corrosion areas. The generation, initiation and spread
of corrosion on a base metal is typically a factor of contaminates, temperature and time.
Page: 68
Knowing the regions of operation, the air contaminates can be determined and their
impact to hot corrosion can be considered with respect to their impact to hot corrosion.
To better understand these contaminates within the blade, a more detailed cut-up
analysis could be conducted to better study the composition of the corrosion buildup.
This would lead to a better understanding of what contaminates are involved within the
corrosion zones. Knowing these contaminates and the temperatures from the thermal
model, two of the three corrosion factors have been achieved. However, the predicted
temperatures from these thermal models were generated for a particular mission point
not the time of operation at these points. The last factor of corrosion is this time of
operation and it can play a large role in the creation, spread and penetration of the
corrosion within the part. If an area in question is held at a temperature in either the
Type I or Type II range for 30 minutes, it is typically going to develop less corrosion
damage than the same area held at the same temperature for 3 hours. This influence on
time would need to be built into the models for future studies; however, the thermal
models generated within this project are a good first step. This model was able to show
that the selected blades were operating within the Type II Hot Corrosion temperature
range for some, if not most, of the operating time. The next step is to better define this
time of operation at each mission point.
A last item that could also have an effect on the thermal models is the better
modeling of the external surface boundary conditions on the airfoil to match the field run
hardware. The thermal models within this report assumed full external TBC coating at
each of the five mission points within both regions. However, looking at engine run
hardware, it was determined that a majority of the blades at the end of their lives are
missing external TBC in some regions. For this particular airfoil, large regions around
the airfoil can be missing TBC after operation. If this TBC variation was built into the
thermal model, the external temperatures of the airfoil would increase. This in turn
would increase the predicted metal temperatures of the internal cores at these locations
and just might bump the predicted temperature range from the Type II to the Type I or
even oxidation temperature levels. This could show more variation in the predicting of
internal corrosion on these blades. The variation can also be region specific as well and
would need to be adjusted based on the inspected engine hardware prior to cut up.
Page: 69
5. CONCLUSION
It was originally expected to see large areas of corrosion within the blades from
both regions with the Middle East experiencing slightly more damage.
This was
expected based on the engine history within these regions, the environmental factors
within each region and the temperatures expected within the thermal model. However
after analysis minimal corrosion and little variation was seen between the thermal results
and cut-up analysis between regions. Minimal corrosion on the field run hardware was
assumed to be directly related to the internal coating operating correctly and doing what
it is designed to do. The environmental factors showed that both the PM10 and PM2.5
levels within these regions are much higher when compared to other areas of the world,
while the SO2 levels are either right on par with the rest of the world or at lower levels,
with the Middle East showing higher SO2 levels than India. This higher SO2 level is
what led to the expectation of higher corrosion levels in the Middle East blades. This
was not seen possibly do to minimal SO2 ingestion into the cores for both blades; a more
detailed analysis would be needed to support this. In the end the cut up engine hardware
went on to actually show very little differences between the two regions. Both regions
showed areas of corrosion but one was not more corroded then the other. Overall both
sets of engine hardware showed the same levels of distress and the thermal models
showed predicted metal temperatures within the same ranges between the two models.
After this project was completed within its entirety, it was determined that the
thermal model generated did not predict the levels and areas of corrosion to the accuracy
that was hoped for. The tool is still new and in its infancy, however this project was a
great first step in the right direction showing what areas of the model needed
improvements. The model was able to show the temperatures a given mission point and
whether or not if these temperatures were in the Type I or Type II range. However to
lack of an operating time variable within the model was a setback. If an operating time
factor was built into the model it may have been able to better predict particular
corrosion areas, or even a variation in temperatures for given mission points. It is hoped
that with more detail built into the operating times at each mission point that particular
corrosion zones will be able to be identified and predicted. To better predict these zones
more calibration would need to be implemented to the model. These calibration tools,
Page: 70
test and methods are related to external coatings, improved cut up analysis and internal
coatings.
Calibration of the predicted metal temperatures to engine run hardware was
difficult due to various issues that arose during the project’s completion. The first issue
that provided difficulty in the models calibration was the limited corrosion zones for
inspection within the part cut-ups. These limited zones were most likely due to the
internal coating on the part doing its job. This part had known corrosion issues within
the field run hardware in its early stages; these issues were corrected with the addition of
this internal diffusion coating. Due to this it was hard to correlate the thermal model
back to the engine run hardware. A second issue that provided a hardship was the
minimal variation placed on the external boundary conditions at each mission point. The
external TBC should have been adjusted to better match the field parts that were cut up,
this would have allowed for a more accurate prediction of the metal temperatures of the
blade. To do this a more detailed variation of TBC loss could be build into the model.
This variation in loss would allow the user to set parameters to spall off the TBC coating
at various points of the mission analysis. This would have a great impact on the external
boundary conditions of the blade and thus have a large effect on the internal cores.
These increased temperatures along with cooling air contamination could lead to
different internal temperatures and different levels of corrosion attack.
As previously stated, this project and model generation is a great first step for the
prediction of internal corrosion zones. The process illustrated and followed within this
project shows a step by step method of trying to predict hot corrosion within a given
turbine blade. This process shows that environmental factors can play a large role in the
creation of corrosion on and in a gas turbine blade, it is not fully temperature driven.
Before this analysis no real existing method was known for trying to predict particular
corrosion zones within a gas turbine blade. This method improved on previous thermal
analysis methods by allowing the model to be run at various mission points, not just one
particular point. This allowed for more models to be run and compared over the given
mission, allowing the user to analyze a wide range of temperatures for possible regions
of corrosion. This capability along with the environmental factors can help in the
prediction of hot corrosion. It is planned that this project will continue to be worked in
Page: 71
the future at Pratt and Whitney so that all the issues found during this project can be
corrected and implemented within the models. With more work and detail applied into
this analysis it is hoped that a model will be generated that can fully predict these zones
of internal corrosion. This would provide a powerful tool when designing new engine
programs. This would allow different regions to have specific parts designed for them;
less corrosive regions may not need internal coatings, resulting in less expensive parts,
while more corrosive regions will need internal coatings resulting in longer running
hardware. These types of decisions would be able to be made during the design phase,
rather than after parts have been field tested and shown damage.
5.1 FUTURE WORK
As previously stated within some areas of this report, there are some tasks that
were hoped to be worked however were not achieved yet still would be good to work in
the future. The first task would be the incorporation of a time factor into the thermal
model analysis. This will better help understand the dwell times of temperatures on the
base metal and may help better predict and understand the areas of corrosion. The
second task that should be worked in the future would be a more detailed cut-up analysis
on the actual engine run hardware. This increased detail would allow for a chemical
analysis to be performed on the cut ups that would show a better understanding of the
corrosion buildup composition. This would show what types of contaminants, whether
internal to the engine or from the external environment, have made their way into the
cores of the blade. This analysis wanted to be performed during this project but was
never achieved. The last future work task that would be valuable would be to perform
the entire analysis on a variety of other regions of the world. Some regions that would
benefit from this type of analysis would be North and South America, Europe, Africa
and East Asia. However this analysis should really be performed on all regions that
have air traffic and especially on engine programs that have known internal corrosion
problems.
Performing this analysis on different engine programs would allow for
different learning opportunities and information to be shared between engine programs.
Page: 72
REFERENCES
[1]
Bogard, David G. "The Gas Turbine Handbook." Airfoil Film Cooling. NETL,
n.d. Web.
[2]
Bornstein, Norman S. "Reviewing Sulfidation Corrosion - Yesterday and
Today." JOM (1996): 37-39. Print.
[3]
Bose, Sudhangshu. High Temperature Coatings. Amsterdam: Elsevier
Butterworth-Heinemann, 2007. Print.
[4]
Çengel, Yunus A., and Michael A. Boles. Thermodynamics: An Engineering
Approach. Boston: McGraw-Hill, 2002. Print.
[5]
Cervenka, Michael. The Rolls-Royce Trent Engine. 5 Oct. 2000.
[6]
Craig, Benjamin D., Richard A. Lane, and David H. Rose. Corrosion Prevention
and Control: A Program Management GuideA Program Management Guide for
Selecting Materials. Rep. 2nd ed. New York: AMMTIAC, 2006. Print.
[7]
Dahl, Jeff. "File:Jet Engine.svg." Wikipedia. Wikimedia Foundation, 23 Nov.
2012. Web. 2012.
[8]
DeLuca, D. P. Understanding Fatigue. N.d. PDF. United Technologies Pratt &
Whitney.
[9]
Donachie, Matthew J., and Stephen J. Donachie. Superalloys: A Technical
Guide. Materials Park, OH: ASM International, 2002. Print.
[10]
"FUNDAMENTALS OF GAS TURBINE ENGINES." Skybrary. N.p., n.d. Web.
<http://www.skybrary.aero/bookshelf/books/1621.pdf>.
[11]
Glage, Alexander. Nickel-based Superalloys and Their Application in the
Aircraft Industry. Thesis. UNIVERSITA’ DEGLI STUDI DI TRENTO, 2006.
N.p.: n.p., n.d. Print.
[12]
Gleeson, Brian. "High-Temperature Corrosion of Metallic Alloys and Coatings."
Materials Science and Technology. Verlag: WILEY-VCH, 2000. N. pag. Print.
[13]
Han, Je-Chin, and Lesley M. Wright. "The Gas Turbine Handbook." Enhanced
Internal Cooling of Turbine Blades and Vanes. NETL, n.d. Web.
[14]
Hass, Derek D. "Directed Vapor Deposition of Thermal Barrier Coatings."
Thesis. University of Virginia, 2000. Print.
[15]
"Hot Corrosion in Gas Turbines." High Temperature Corrosion and Materials
Application (#05208G). N.p.: n.p., n.d. N. pag. Print.
Page: 73
[16]
Husar, Rudolf B. Intercontinental PM Transport to North America: Sahara Dust.
N.d.
[17]
India [map]. 2012. Scale undetermined; generated by Dan Nadeau; using
“Mapquest.com, Inc”.
[18]
Józef Błachnio and Wojciech Izydor Pawlak (2011). Damageability of Gas
Turbine Blades – Evaluation of Exhaust Gas Temperature in Front of the Turbine
Using a Non-Linear Observer, Advances in Gas Turbine Technology, Ernesto
Benini (Ed.), ISBN: 978-953-307-611-9, InTech, DOI: 10.5772/29546. Available
from: http://www.intechopen.com/books/advances-in-gas-turbinetechnology/damageability-of-gas-turbine-blades-evaluation-of-exhaust-gastemperature-in-front-of-the-turbine-us
[19]
"NASA - National Aeronautics and Space Administration." NASA. N.p., n.d.
Web.
[20]
"Nickel Base Single Crystal." Nickel Base Single Crystal. C-M Group, n.d. Web.
[21]
Noordermeer, Jim. Gas Turbines for Cogeneration & Combined-Cycle. N.d.
Gryphon International Engineering Services Inc.
[22]
"Plot based on public databases and stemming from a P&W/ISU Joint Corrosion
Study, Courtesy of Dr. Zhihong Tang and Dr. Brian Gleeson, Iowa State
University, 2012”.
[23]
Shepard, Rob, P.E. Gas Turbine Technologies for Electric Generation. N.d.
PowerPoint. www.Neel-Schaffer.com.
[24]
SidewinderX. "File:Combustor diagram airflow.png." Wikipedia. Wikimedia
Foundation, 11 Jan. 2010. Web. 2012.
[25]
Sims, Chester T., N. S. Stoloff, and William C. Hagel. Superalloys II. New York:
Wiley, 1987. Print.
[26]
Sokhi, Ranjeet S. World Atlas of Atmospheric Pollution. London: Anthem, 2011.
Print.
[27]
Stolle, R. Conventional and Advanced Coatings for Turbine Airfoils. Tech.
München: MTU Aero Engines, n.d. Print.
[28]
"Superalloys: A Primer and History." Superalloys: A Primer and History. Web.
13 Dec. 2012.
<www.tms.org/meetings/specialty/superalloys2000/superalloyshistory.html>.
[29]
The Middle East [map]. 2012. Scale undetermined; generated by Dan Nadeau;
using “Mapquest.com, Inc”.
Page: 74