MAE 4261: AIR-BREATHING ENGINES
Gas Turbine Engine Combustors
Mechanical and Aerospace Engineering Department
Florida Institute of Technology
D. R. Kirk
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COMBUSTOR LOCATION
Commercial
PW4000
Combustor
Military
F119-100
Afterburner
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Turbine
Compressor
MAJOR COMBUSTOR COMPONENTS
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MAJOR COMBUSTOR COMPONENTS
Turbine
Compressor
Fuel
• Key Questions:
– Why is combustor configured this way?
– What sets overall length, volume and geometry of device?
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COMBUSTOR EXAMPLE (F101)
Henderson and Blazowski
Compressor
Turbine
NGV
Fuel
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VORBIX COMBUSTOR (P&W)
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COMBUSTOR REQUIREMENTS
•
•
•
•
•
•
•
•
•
•
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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
• Future – multiple fuel capability (?)
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CHEMISTRY REVIEW
•
•
General hydrocarbon, CnHm (Jet fuel H/C~2)
Complete oxidation, hydrocarbon goes to CO2 and water
m
m

Cn H m   n  O2  nCO2  H 2O
4
2

•
•
•
For air-breathing applications, hydrocarbon is burned in air
Air modeled as 20.9 % O2 and 79.1 % N2 (neglect trace species)
Complete combustion for hydrocarbons means all C → CO2 and all H → H2O
m
m
m


Cn H m   n  O2  3.78 N 2   nCO2  H 2O  3.78 n   N 2
4
2
4


Stoichiometric Molar fuel/air ratio
s 
•
1
m

4.78 n  
4

Stoichiometric Mass fuel/air ratio
s 
12n  m 
m

 n  32  3.7828

Stoichiometric = exactly correct ratio for complete combustion
4
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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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Flame Temperature
VARIATION OF FLAME TEMPERATURE WITH f
Still too hot
for turbine
Flammability Limits
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WHY IS THIS RELEVANT?
•
Most mixtures will NOT burn so far away from
stoichiometric
– Often called Flammability Limit
– Highly pressure dependent
• Increased pressure, increased
flammability limit
– Requirements for combustion, roughly f > 0.8
•
Gas turbine can NOT operate at (or even near)
stoichiometric levels
– Temperatures (adiabatic flame temperatures)
associated with stoichiometric combustion are
way too hot for turbine
– Fixed Tt4 implies roughly f < 0.5
•
What do we do?
– Burn (keep combustion going) near f=1 with
some of compressor exit air
– Then mix very hot gases with remaining air to
lower temperature for turbine
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Compressor
Air
Turbine
SOLUTION: BURNING REGIONS
Primary
Zone
f~0.3
f ~ 1.0
T>2000 K
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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 altitude operation (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 altitude operation (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

cP m a  m f Tt 4  m aTt 3
hb 
m f 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
cP
Tt 4  Tt 3 
f 
hbQR
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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-TYPE
Rolls-Royce Dart
ANNULAR-TYPE
General 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 don’t 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
J79 (F4, F104, B58)
Combustor
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)
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INTRA-TURBINE BURNING
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BURNER-TURBINE-BURNER (ITB) CONCEPTS
Conventional
Intra Turbine Burner (schematic only)
•
•
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 CO2 emissions
– Reduce NOx emissions by reducing peak flame temperature
Initially locate ITB in transition duct between high pressure turbine (HTP) and low
pressure turbine (LPT)
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SIEMENS WESTINGHOUSE ITB CONCEPT
Tt4
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UNDERSTANDING BENEFIT FROM CYCLE ANALYSIS
From “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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UNDERSTANDING BENEFIT FROM CYCLE ANALYSIS
From “Turbojet and Turbofan Engine Performance Increases Through Turbine Burners, by
Liu and Sirignano, JPP Vol. 17, No. 3, May-June 2001
2 additional burners
5 additional burners
Continuous burning
in turbine
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