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Turbine and Compressor
Design
Team:
Kevin Garvey
Alex von Oetinger
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Major Topics
Compressor and Turbine Design Cooling
Dynamic Surge Stall Propagation
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BackgroundHistory: First gas turbine was developed in 1872 by Dr. F. Stolze.
Gas Turbine EngineWhat does it do? Generates thrust by mixing compressed ambient air with
fuel and combusting the mixture through a nozzle topropel an object forward or to produce shaft work.
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How Does it Work? Newtons third law
For every action, there is an equal andopposite reaction.
As the working fluid is exhausted out the nozzleof the gas turbine engine, the object that theengine is attached to is pushed forward. In the
case of generating shaft work, the shaft turns agenerator which produces electrical power.
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How Does it Work? Cont.
Shaft
ExhaustGas
AmbientAir In
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Operation
Compressor is connected to the turbine via ashaft. The turbine provides the turning momentto turn the compressor.
The turning turbine rotates the compressor fanblades which compresses the incoming air.
Compression occurs through rotors and statorswithin the compression region. Rotors (Rotate with shaft)
Stators (Stationary to shaft)
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Types of Gas Turbines
Centrifugal Compressed air output is around the outer perimeter
of engine
Axial Compressed air output is directed along the centerline
of the engine
Combination of Both Compressed air output is initially directed along
center shaft of engine and then is compressed
against the perimeter of engine by a later stage.
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Example of Centrifugal
Flow
Intake airflow is being forced around theoutside perimeter of the engine.
CentrifugalCompressor
Airflow beingforced aroundbody of engine
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Example of Axial Flow
Intake airflow is forced down the center shaftof the engine.
MultistageAxial
Compressor
CenterShaft
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Example of Combination
Flow
Intake AirFlow
Axial Compressor
CentrifugalCompressor
Intake air flow is forced down the centershaft initially by axially compressor stages,and then forced against engine perimeter
by the centrifugal compressor.
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Major Components of
Interest Compressor
Axial
Centrifugal
TurbineAxial
Radial
Axial Compressor Centrifugal Compressor
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Axial Compressor Operation
A&P Technician Powerplant Textbook published by Jeppesen Sanderson Inc., 1997
Axial compressors are designed in a divergentshape which allows the air velocity to remainalmost constant, while pressure graduallyincreases.
Average Velocity
http://ffden-2.phys.uaf.edu/212_fall2003.web.dir/Oliver_Fleshman/3-15l.jpg7/29/2019 Turbine and Compressor Design
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Axial Compressor Operation
cont. The airflow comes in through the inlet andfirst comes to the compressor rotor. Rotor is rotating and is what draws the airflow
into the engine.After the rotor is the stator which does not move
and it redirects the flow into the next stage ofthe compressor.
Air flows into second stage. Process continues and each stage gradually
increases the pressure throughout the
compressor.
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Axial Compressor Staging
An axial compressor stage consists of a rotorand a stator.
The rotor is installed in front of the stator
and air flows through accordingly. (See Fig.)
www.stanford.edu/ group/cits/simulation/
http://www.stanford.edu/group/cits/simulation/http://www.stanford.edu/group/cits/simulation/7/29/2019 Turbine and Compressor Design
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Centrifugal Compressor
Operation
Centrifugal compressors rotate ambient air about animpeller. The impeller blades guide the airflow towardthe outer perimeter of the compressor assembly. The airvelocity is then increased as the rotational speed of the
impeller increases.
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Axial Turbine OperationHot combustion gasesexpand, airflowpressure andtemperature drops. This
drop over the turbineblades creates shaftwork which rotates thecompressor assembly.
Axial Turbine with airflow Airflow around rotor
Airflow through
stator
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Radial Turbine Operation
Same operationcharacteristics as axial flowturbine.
Radial turbines are simplerin design and lessexpensive to manufacture.
They are designed muchlike centrifugalcompressors.
Airflow is essentiallyexpanded outward fromthe center of the turbine.
Radial Flow Turbine
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Gas Turbine Issues
Gas Turbine Engines Suffer from anumber of problematic issues:
Thermal Issues Blade (airfoil) Stalls
Dynamic Surge
http://www.turbosolve.com/index.html
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Thermal Issues Gas Turbines are limited
to lower operatingtemperatures due to thematerials available for
the engine itself.
Operating at the lowertemperature will
decrease the efficiencyof the gas turbine so ameans of cooling thecomponents is necessaryto increase temperatures
at which engine is run.
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Cooling Methods
Spray (Liquid) Passage
Transpiration
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Spray Cooling
The method of spraying aliquid coolant onto the
turbine rotor blades andnozzle.
Prevents extreme turbineinlet temperatures from
melting turbine blades bydirect convection betweenthe coolant and theblades.
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Passage Cooling
Hollow turbine bladessuch that a passage isformed for themovement of a coolingfluid.
DOE has relatively newprocess in which excesshigh-pressurecompressor airflow isdirected into turbinepassages.
http://www.eere.energy.gov/inventions/pdfs/fluidtherm.pdf
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Transpiration Cooling
Method of forcing airthrough a porous turbineblade.
Ability to remove heat ata more uniform rate.
Result is an effusing layerof air is produced around
the turbine blade. Thus there is a reduction
in the rate of heattransfer to the turbine
blade.
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Blade (airflow) Stalls When airflow begins
separating from thecompressor blades overwhich it is passing as theangle of attack w.r.t. theblades exceeds thedesign parameters.
The result of a blade stallis that the blade(s) nolonger produce lift andthus no longer produces apressure rise through the
compressor.
Separation Regions
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Dynamic Surge
Occurs when the static (inlet) airpressure rises past the designcharacteristics of thecompressor.
When there is a reversal of
airflow from the compressorcausing a surge to propagate inthe engine.
Essentially, the flow is exhaustedout of the compressor, or front,
of the engine. Result, is the compressor no
longer able to exhaust as quicklyas air is being drawn in and a
bang occurs. http://www.turbosolve.com/index.html
Compressor
Inlet
Turbine Exit
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Dynamic Surge Effects
Cause: Inlet flow is reversed Effect: Mass flow rate is reduced into engine.
Effect: Compressor stages lose pressure.
Result: Pressure drop allows flow to reverse back intoengine.
Result: Mass flow rate increases
Cause: Increased mass flow causes high pressureagain. Effect: Surge occurs again and process continues.
Result: Engine surges until corrective actions are taken.
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Dynamic Surge Process
inm
outm
P
V
Surge Point,Flow
Reverses
No Surge
Conditio
n
CompressorPressure Loss
Occurs
Flow reverses
back intoengine
Corrective
ActionTaken
outm
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Axial Compressor Design
Assumption of Needs Determination of Rotational Speed
Estimation of number of stages General Stage DesignVariation of air angles
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Assumption of Needs
The first step in compressor design in thedetermination of the needs of the system
Assumptions: Standard Atmospheric Conditions
Engine Thrust Required
Pressure Ratio Required
Air Mass Flow Turbine inlet temperature
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Rotational Speed
Determination First Step in Axial Compressor Design Process for this determination is based on
assumptions of the system as a whole
Assumed: Blade tip speed, axial velocity, andhub-tip ratio at inlet to first stage.
Rotational Speed
Equation
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Derivation of Rotational
Speed First Make Assumptions: Standard atmospheric conditions
Axial Velocity:
Tip Speed:
No Intake Losses
Hub-tip ratio 0.4 to 0.6
Ut
350m
s
Ca
150 200m
s
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Compressor Rotational
Speed Somewhat of an iterative process inconjunction with the turbine design.
Derivation Process: First Define the mass flow into the system
is the axial velocity range from the rootof the compressor blades to the tips of theblades.
AUmdot
where U=1aC
1aC
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Axial Velocity Relationship
rr
a
t
ra C
r
rC *1
2
1
Radius to root of
bladerr
tr Radius to tip of blade
tr
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Tip Radius Determination
2
11
2
1
t
r
a
dot
t
r
rC
m
r
By rearranging the mass flow rate equation we canobtain an iterative equation to determine the blade tipradius required for the design.
Now Looking at the energy equation, we can determine theentry temperature of the flow.
p
a
c
CTT
2
2
101
22
2
11
2
00
UTc
UTc pp
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Isentropic Relationships Now employing the isentropic relation between
the temperatures and pressures, then thepressure at the inlet may be obtained.
Now employ the ideal gas law to obtain thedensity of the inlet air.
1
0
1
01
T
TPP
1
1
1
RT
P
Finally Obtaining Rotational
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Finally Obtaining Rotational
Speed Using the equation for tip speed.
Rearranging to obtain rotational speed.
Finally an iterative process is utilized toobtain the table seen here.
NrU tt 2
t
t
r
UN
2
Determining Number of
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Determining Number of
Stages Make keen assumptions Polytropic efficiency of approximately 90%.
Mean Radius of annulus is constant throughall stages.
Use polytropic relation to determine theexit temperature of compressor.
nn
P
PTT
1
01
020102
n = 1.4, Ratio of Specific Heats,Cp/Cv
is the pressure that the compressor
outputs
To1 is ambient temperature
02P
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Determine Temperature
Change Assuming that Ca1=Ca is the work done factor Work done factor is estimate of stage efficiency Determine the mean blade speed.
Geometry allows for determining the rotor bladeangle at the inlet of the compressor.
NrU meanm 2
a
m
C
U
1tan
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Temperature Rise in a Stage
p
ams
c
CUT 210
tantan
1
1
cos
aCV
This will give an estimate of the maximum possible rotordeflection.
Finally obtain the temperature rise through the stage.
2
2cos
V
Ca
Determine the speed of the flow over the blade profile.
Velocity flow
over blade V1.
DeflectionBlade_12
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Number of Stages Required
The number of stages required is dependentupon the ratio of temperature changesthroughout the compressor.
sT
TStages
0
ambTTT 2
is the temperature change within a stage
is the average temperature change over all the stagessT
T
0
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Designing a Stage
Make assumptionsAssume initial temperature change through
first stage.
Assume the work-done factors through each
stage. Ideal Gas at standard conditions
Determine the air angles in each stage.
http://images.google.com/imgres?imgurl=http://fromtheflightdeck.com/Stories/turbofan/index_files/image015.jpg&imgrefurl=http://fromtheflightdeck.com/Stories/turbofan/&h=188&w=264&sz=13&tbnid=3WP8tL-hn0UCMM:&tbnh=76&tbnw=107&hl=en&start=36&prev=/images%3Fq%3DCo7/29/2019 Turbine and Compressor Design
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Stages 1 to 2
Determine the change in the whirl velocity. Whirl Velocity is the tangential component of
the flow velocity around the rotor.
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Stage 1 to 2
Change in whirl velocity through stage.
12 www CCC
m
p
wU
Tc
C
11 tan aw CC Alpha 1 is zero at the first stage.
a
w
a
wm
C
C
C
CU
22
22
tan
tan
Compressor Velocity
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Compressor Velocity
Triangles
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Pressure ratio of the Stage
10
01
03 1
amb
sss
T
T
P
PR
9.0s
The pressure ratio in the stage can be determined throughthe isentropic temperature relationship and the polytropicefficiency assumed at 90%.
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Stage Attributes The analysis shows that the stage can be outlined bythe following attributes:
1.) Pressure at the onset ofthe stage.
2.) Temperature at the onsetof the stage.
3.) The pressure ratio of thestage.
4.) Pressure at the end of the
stage.
5.) Temperature at the end ofthe stage.
6.) Change in pressure
through the stage.Example of a single
sta e
V i i i Ai A l f
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Variation in Air Angles of
Blade Assume the free vortex condition.
Determine stator exit angle.
Then determine the flow velocity.
constrCw 2
13 tantan a
m
C
U
33
cos
mUC
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Air Angle TriangleAlpha 1 is 0 at
the inlet stagebecause there
are no IGVs.
Thus,
Ca1=C1,
and Cw1 is0
Note: This is
the whirl
velocity
componentand not a
blade
spacing!
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Red is
Green is
Blue is
Velocity Triangle
aC
aC
aC
V i i i Ai A l f
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Variation in Air Angles of
Blade Determine the exit temp., pressure, and density ofstage 1
Determine the blade height at exit.
Finally determine the radii of the blade at stator exit.
p
a
c
CTT
2
2
03 1
03
3033
T
TPP
3
3
3
RT
P
meanr
Ah
2
3
2
hrr meants
2
hrr meanrs
a
dot
C
mA
3
3
V i ti i Ai A l f
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Variation in Air Angles of
Blade Determine the radii at the rotor exit.
Determine the whirl velocities at the blade root andtip.
2
tstritr
rrr
2
rsrrirr
rrr
Note: That is the radius of the blade at the tip at rotor inlet.trirNote: That is the radius of the blade at the root at rotor inlet.rri
r
rr
meanmwrwr
rCC 22 tr
meanmwtwr
rCC 22
Note: because there is no other whirl velocity component in the
first stage.
22 wmw CC
Fi ll d t i th Ai
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Finally determine the Air
Angles
a
twtrt
a
mwmm
a
rwrrr
a
twt
a
mwm
a
rwr
C
CU
C
CU
C
CU
C
C
C
C
CC
2
2
2
2
2
2
2
2
2
2
2
2
tan
tan
tan
tan
tan
tan
Stator air angle at root ofblade
Stator air angle at middle ofblade
Stator air angle at tip of blade
Deflection air angle at root ofblade
Deflection air angle at middleof blade
Deflection air angle at tip ofblade
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Compressor Design ExampleDesign of a 5 stage axial compressor:
98.0
150
5.452
288
2262.0
2
sm
a
a
t
C
KT
KT
mrGivens:
Use this and chart to get
Rotational speed of engine.
Once rotational speed is found, determine mean blade tip speed.
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Example
s
mNrU
mrr
r
meanm
rtmean
6.2662
1697.02
KTTT amb 5.1642
Determine the total temperature rise through the first stage.
We are designing for more than just one stage, so we
need to define an average temperature rise per stage:
KStages
TTs 9.32
#0
Example (Air Angle
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Example (Air AngleDetermination)
2
0
1
12
1
1
55.126
0
64.60tan
w
m
sp
w
w
www
a
m
Cs
m
U
TcC
s
mC
CCC
CU
E l (Ai A l
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Example (Air Angle
Determination)
smC
V
C
CU
a
a
wm
21.205cos
03.43tan
22
21
2
15.40tan 212a
w
C
C
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Questions???