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1 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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Presentation on theme: "1 MAE 4261: AIR-BREATHING ENGINES Gas Turbine Engine Combustors Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk."— Presentation transcript:

1 1 MAE 4261: AIR-BREATHING ENGINES Gas Turbine Engine Combustors Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk

2 2 COMBUSTOR LOCATION Military F Commercial PW4000 Combustor Afterburner

3 3 MAJOR COMBUSTOR COMPONENTS Compressor Turbine

4 4 MAJOR COMBUSTOR COMPONENTS Key Questions: –Why is combustor configured this way? –What sets overall length, volume and geometry of device? Compressor Turbine Air Fuel Combustion Products

5 5 COMBUSTOR EXAMPLE (F101) Henderson and Blazowski Fuel Compressor Turbine NGV

6 6 VORBIX COMBUSTOR (P&W)

7 7

8 8 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 (?)

9 9 CHEMISTRY REVIEW Stoichiometric Molar fuel/air ratio Stoichiometric Mass fuel/air ratio General hydrocarbon, C n H m (Jet fuel H/C~2) Complete oxidation, hydrocarbon goes to CO 2 and water For air-breathing applications, hydrocarbon is burned in air Air modeled as 20.9 % O 2 and 79.1 % N 2 (neglect trace species) Complete combustion for hydrocarbons means all C → CO 2 and all H → H 2 O Stoichiometric = exactly correct ratio for complete combustion

10 10 COMMENTS ON CHALLENGES Based on material limits of turbine (T t4 ), 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

11 11 VARIATION OF FLAME TEMPERATURE WITH  Flame Temperature Flammability Limits Still too hot for turbine

12 12 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 T t4 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

13 13 SOLUTION: BURNING REGIONS Air Compressor Turbine  ~ 1.0 T>2000 K  ~0.3 Primary Zone

14 14 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 → CO 2 ) 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 ~ 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 ~

15 15 COMBUSTOR DESIGN Combustion efficiency,  b = Actual Enthalpy Rise / Ideal Enthalpy Rise –h=heat of reaction (sometimes designated as Q R ) = 43,400 KJ/Kg General Observations:  b ↓ as p ↓ and T ↓ (because of dependency of reaction rate)  b ↓ as Mach number ↑ (decrease in residence time)  b ↓ as fuel/air ratio ↓ Assuming that the fuel-to-air ratio is small

16 16

17 17 COMBUSTOR TYPES (Lefebvre) Single Can Tubular or Multi-Can Tuboannular Can-Annular Annular

18 18 COMBUSTOR TYPES (Lefebvre)

19 19 EXAMPLES CAN-TYPE Rolls-Royce Dart ANNULAR-TYPE General Electric T58

20 20 EXAMPLES CAN-ANNULAR-TYPE Rolls-Royce Tyne

21 21 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

22 22 AFTERBURNER (AUGMENTER) Spray in more fuel to use up more oxygen –Main combustion can not use all air Exit Mach number stays same (choked M exit = 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

23 23 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 ~ p n (n~2 for mixed gas collision rate) J79 (F4, F104, B58) Combustor Afterburner

24 24 INTRA-TURBINE BURNING

25 25 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 CO 2 emissions –Reduce NO x 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)

26 26 SIEMENS WESTINGHOUSE ITB CONCEPT T t4

27 27 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 ConventionalIntra Turbine Burner

28 28 2 additional burners5 additional burners 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


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