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MAE 4261: AIR-BREATHING ENGINES
Gas Turbine Engine Combustors
Mechanical and Aerospace Engineering DepartmentFlorida Institute of Technology
D. R. Kirk
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COMBUSTOR LOCATION
Military
F119-100
Commercial
PW4000
Combustor
Afterburner
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MAJOR COMBUSTOR COMPONENTS
Compre
ssor
Tu
rbine
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MAJOR COMBUSTOR COMPONENTS
Key Questions:
Why is combustor configured this way?
What sets overall length, volume and geometry of device?
Compre
ssor
Tu
rbine
Fuel
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COMBUSTOR EXAMPLE (F101)
Henderson and Blazowski
Fuel
Compressor
Turb
ine
NG
V
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VORBIX COMBUSTOR (P&W)
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COMBUSTOR REQUIREMENTS
Complete combustion (hb 1)
Low pressure loss (pb 1)
Reliable and stable ignition Wide stability limits
Flame stays lit over wide range of p, u, f/a ratio)
Freedom from combustion instabilities
Tailored temperature distribution into turbine with no hot spots Low emissions
Smoke (soot), unburnt hydrocarbons, NOx, SOx, CO
Effective cooling of surfaces
Low stressed structures, durability
Small size and weight
Design for minimum cost and maintenance
Futuremultiple fuel capability (?)
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CHEMISTRY REVIEW
OHmnCOOmnHC mn 22224
478.4
1
m
n
s
22222
478.3
278.3
4N
mnOH
mnCONO
mnHC mn
Stoichiometric Molar fuel/air ratio Stoichiometric Mass fuel/air ratio
General hydrocarbon, CnHm(Jet fuel H/C~2)
Complete oxidation, hydrocarbon goes to CO2and water
For air-breathing applications, hydrocarbon is burned in air
Air modeled as 20.9 % O2and 79.1 % N2(neglect trace species)
Complete combustion for hydrocarbons means all C CO2and all H H2O
2878.3324
12
mn
mns
Stoichiometric = exactly correct ratio for complete combustion
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COMMENTS ON CHALLENGES
Based on material limits of turbine (Tt4), combustors must operate below
stoichiometric values
For most relevant hydrocarbon fuels, s~ 0.06 (based on mass)
Comparison of actual fuel-to-air and stoichiometric ratio is called equivalence ratio
Equivalence ratio = f = /stoich
For most modern aircraft f~ 0.3
Summary
If f= 1: Stoichiometric
If f> 1: Fuel Rich
If f< 1: Fuel Lean
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VARIATION OF FLAME TEMPERATURE WITH
FlameTem
perature
Flammability LimitsStill too hot
for turbine
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WHY IS THIS RELEVANT?
Most mixtures will NOTburn so far away fromstoichiometric
Often called Flammability Limit
Highly pressure dependent
Increased pressure, increasedflammability limit
Requirements for combustion, roughly f> 0.8
Gas turbine can NOToperate at (or even near)stoichiometric levels
Temperatures (adiabatic flame temperatures)associated with stoichiometric combustion areway too hot for turbine
Fixed Tt4implies roughly f< 0.5
What do we do?
Burn (keep combustion going) near f=1 withsome of compressor exit air
Then mix very hot gases with remaining air tolower temperature for turbine
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SOLUTION: BURNING REGIONS
Air
Compressor
Turbine
f ~ 1.0
T>2000 K
f~0.3
Primary
Zone
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COMBUSTOR ZONES: MORE DETAILS
1. Primary Zone
Anchors Flame
Provides sufficient time, mixing, temperature for complete oxidation of fuel Equivalence ratio near f=1
2. Intermediate (Secondary Zone)
Low altitudeoperation (higher pressures in combustor)
Recover dissociation losses (primarily CO CO2) and Soot Oxidation
Complete burning of anything left over from primary due to poor mixing High altitudeoperation (lower pressures in combustor)
Low pressure implies slower rate of reaction in primary zone
Serves basically as an extension of primary zone (increased tres)
L/D ~ 0.7
3. Dilution Zone (critical to durability of turbine) Mix in air to lower temperature to acceptable value for turbine
Tailor temperature profile (low at root and tip, high in middle)
Uses about 20-40% of total ingested core mass flow
L/D ~ 1.5-1.8
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COMBUSTOR DESIGN
Combustion efficiency, hb= Actual Enthalpy Rise / Ideal Enthalpy Rise
h=heat of reaction (sometimes designated as QR) = 43,400 KJ/Kg
34 tt
Rb
P TTQ
cf
h
General Observations:1. hb as p and T (because of dependency of reaction rate)
2. hb as Mach number (decrease in residence time)
3. hb as fuel/air ratio
Assuming that the fuel-to-air ratio is small
hm
TmTmmc
f
tatfaP
b
34h
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COMBUSTOR TYPES (Lefebvre)
Single Can
Tubular
or Multi-Can
Tuboannular
Can-Annular
Annular
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COMBUSTOR TYPES (Lefebvre)
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EXAMPLES
CAN-TYPERolls-Royce Dart
ANNULAR-TYPEGeneral Electric T58
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EXAMPLES
CAN-ANNULAR-TYPE
Rolls-Royce Tyne
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CHEMICAL EMISSIONS
Aircraft deposit combustion products at high altitudes, into upper troposphere and
lower stratosphere (25,000 to 50,000 feet)
Combustion products deposited there have long residence times, enhancing impact NOx suspected to contribute to toxic ozone production
Goal: NOx emission level to no-ozone-impact levels during cruise
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AFTERBURNER (AUGMENTER)
Spray in more fuel to use up more oxygen
Main combustion can not use all air
Exit Mach number stays same (choked Mexit= 1) Temp
Speed of sound
Velocity = M*a
Therefore Thrust
Penalty:
Pressure is lower so thermodynamic efficiency is poor
Propulsive efficiency is reduced (but dont really care in this application)
As turbine inlet temperature keeps increasing less oxygen downstream for AB and
usefulness decreases Requirements
VERY lightweight
Stable and startable
Durable and efficient
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RELATIVE LENGTH OF AFTERBURNER
Why is AB so much longer than primary combustor?
Pressure is so low in AB that they need to be very long (and heavy)
Reaction rate ~ pn(n~2 for mixed gas collision rate)
J79 (F4, F104, B58)
Combustor Afterburner
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INTRA-TURBINE BURNING
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BURNER-TURBINE-BURNER (ITB) CONCEPTS
Improve gas turbine engine performance using an interstage turbine burner (ITB)
With a higher specific thrust engine will be smaller and lighter
Increasing payload
Reduce CO2emissions
Reduce NOxemissions by reducing peak flame temperature
Initially locate ITB in transition duct between high pressure turbine (HTP) and low
pressure turbine (LPT)
Conventional
Intra Turbine Burner (schematic only)
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SIEMENS WESTINGHOUSE ITB CONCEPT
Tt4
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UNDERSTANDING BENEFIT FROM CYCLE ANALYSISFrom Turbojet and Turbofan Engine Performance Increases Through Turbine Burners, by
Liu and Sirignano, JPP Vol. 17, No. 3, May-June 2001
Conventional Intra Turbine Burner
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2 additional burners 5 additional burners
UNDERSTANDING BENEFIT FROM CYCLE ANALYSISFrom Turbojet and Turbofan Engine Performance Increases Through Turbine Burners, by
Liu and Sirignano, JPP Vol. 17, No. 3, May-June 2001
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Continuous burning in turbine
UNDERSTANDING BENEFIT FROM CYCLE ANALYSISFrom Turbojet and Turbofan Engine Performance Increases Through Turbine Burners, by
Liu and Sirignano, JPP Vol. 17, No. 3, May-June 2001