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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

MilitaryF119-100

CommercialPW4000

Combustor

Afterburner

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MAJOR COMBUSTOR COMPONENTSC

ompr

esso

r

Tur

bine

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MAJOR COMBUSTOR COMPONENTS

• Key Questions:

– Why is combustor configured this way?

– What sets overall length, volume and geometry of device?

Com

pres

sor

Tur

bine

Air

Fuel

Combustion Products

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COMBUSTOR EXAMPLE (F101)Henderson and Blazowski

Fuel

Com

pres

sor

Tur

bine

NG

V

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VORBIX COMBUSTOR (P&W)

7

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COMBUSTOR REQUIREMENTS

• Complete combustion (b → 1)

• Low pressure loss (b → 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

OHm

nCOOm

nHC mn 222 24

478.4

1m

ns

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 CO2 and water

• 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

2878.332

4

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 = = stoich

– For most modern aircraft ~ 0.3

• Summary

– If = 1: Stoichiometric

– If > 1: Fuel Rich

– If < 1: Fuel Lean

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VARIATION OF FLAME TEMPERATURE WITH

Fla

me

Tem

pera

ture

Flammability LimitsStill too hotfor turbine

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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 > 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 < 0.5

• What do we do?– Burn (keep combustion going) near =1 with

some of compressor exit air– Then mix very hot gases with remaining air to

lower temperature for turbine

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SOLUTION: BURNING REGIONS

Air

Com

pres

sor

Tur

bine

~ 1.0T>2000 K

~0.3

PrimaryZone

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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 =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 res)

– 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, b = Actual Enthalpy Rise / Ideal Enthalpy Rise– h=heat of reaction (sometimes designated as QR) = 43,400 KJ/Kg

34 ttRb

P TTQ

cf

• General Observations:

1. b ↓ as p ↓ and T ↓ (because of dependency of reaction rate)

2. b ↓ as Mach number ↑ (decrease in residence time)

3. b ↓ as fuel/air ratio ↓

• Assuming that the fuel-to-air ratio is small

hm

TmTmmc

f

tatfaPb

34

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COMBUSTOR TYPES (Lefebvre)

Single Can

Tubularor Multi-Can

TuboannularCan-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-TYPERolls-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

• 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 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)

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

Continuous burningin turbine