MAE 4261: AIR-BREATHING ENGINES Gas Turbine Engine Combustors Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk

COMBUSTOR LOCATION Commercial PW4000 Combustor Military F119-100 Afterburner

MAJOR COMBUSTOR COMPONENTS Turbine Compressor

MAJOR COMBUSTOR COMPONENTS Fuel Combustion Products Turbine Air Compressor Key Questions: Why is combustor configured this way? What sets overall length, volume and geometry of device?

COMBUSTOR EXAMPLE (F101) Henderson and Blazowski Fuel Turbine NGV Compressor

VORBIX COMBUSTOR (P&W)

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 Future – multiple fuel capability (?)

CHEMISTRY REVIEW 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 Stoichiometric Molar fuel/air ratio Stoichiometric Mass fuel/air ratio Stoichiometric = exactly correct ratio for complete combustion

COMMENTS ON CHALLENGES Based on material limits of turbine (Tt4), combustors must operate below stoichiometric values For most relevant hydrocarbon fuels, ys ~ 0.06 (based on mass) Comparison of actual fuel-to-air and stoichiometric ratio is called equivalence ratio Equivalence ratio = f = y/ystoich For most modern aircraft f ~ 0.3 Summary If f = 1: Stoichiometric If f > 1: Fuel Rich If f < 1: Fuel Lean

VARIATION OF FLAME TEMPERATURE WITH f Still too hot for turbine Flammability Limits

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

SOLUTION: BURNING REGIONS Turbine Air Primary Zone f~0.3 f ~ 1.0 T>2000 K Compressor

COMBUSTOR ZONES: MORE DETAILS Primary Zone Anchors Flame Provides sufficient time, mixing, temperature for “complete” oxidation of fuel Equivalence ratio near f=1 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 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

COMBUSTOR DESIGN Combustion efficiency, hb = Actual Enthalpy Rise / Ideal Enthalpy Rise h=heat of reaction (sometimes designated as QR) = 43,400 KJ/Kg General Observations: hb ↓ as p ↓ and T ↓ (because of dependency of reaction rate) hb ↓ as Mach number ↑ (decrease in residence time) hb ↓ as fuel/air ratio ↓ Assuming that the fuel-to-air ratio is small

COMBUSTOR TYPES (Lefebvre) Single Can Tubular or Multi-Can Tuboannular Can-Annular Annular

COMBUSTOR TYPES (Lefebvre)

EXAMPLES CAN-TYPE Rolls-Royce Dart ANNULAR-TYPE General Electric T58

EXAMPLES CAN-ANNULAR-TYPE Rolls-Royce Tyne

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

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

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)

INTRA-TURBINE BURNING

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)

SIEMENS WESTINGHOUSE ITB CONCEPT Tt4

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

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 Continuous burning in turbine 2 additional burners 5 additional burners