HIGH BYPASS RATIO TURBOFAN ENGINE App. J AIAA AIRCRAFT ENGINE DESIGN

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1 CORSO DI LAUREA SPECIALISTICA IN Ingegneria Aerospaziale PROPULSIONE AEROSPAZIALE I HIGH BYPASS RATIO TURBOFAN ENGINE App. J AIAA AIRCRAFT ENGINE DESIGN LA DISPENSA E E DISPONIBILE SU Prof. Ing. A. Ficarella antonio.ficarella@unile.it 1

2 NONAFTERBURNING, SEPARATE EXHAUST FLOW high thrust and low fuel consumption Mach < 0.9 2

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5 component,, efficiencies both the core m C and bypass m F airflows pass through separate nozzles total pressure ratio of the bypass air stream nozzle fixed convergent nozzle for both the core and bypass air streams pressure exit equal to the ambient pressure for unchoked nozzle greater that the ambient pressure when choked 5

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7 PARAMETRIC ANALYSIS uninstalled thrust with separate exhaust streams 7

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10 velocity and temperature ratios core airstream velocity and temperature ratios bypass airstream 10

11 freestream recovery maximum allowable turbine inlet total temperature T t4 afterburner 11

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13 INDEPENDENT VARIABLES fan pressure ratio f overall cycle pressure ratio c bypass ratio pressure ratio across the high-pressure compressor from the temperature ratios across the fan, the low-pressure compressor, and the high-pressure compressor are related to their pressure ratios the temperature ratios across the HP turbine and LP turbine obtained from power balances 13

14 the temperature ratios across the two cooling air mixing processes the pressure ratio across the HP and LP turbine are related to their respective temperature ratio and polytrophic efficiency the fuel-air ratio 14

15 the only unknowns for solution are the static pressure ratios P 0 /P 9 and P 0 /P 19 UNCHOKED FLOW the exit static pressure P e is equal to the ambient pressure P 0 exit Mach n. < 1 exit Mach n. M e determined using the compressible flow functions 15

16 CHOKED FLOW P te /P 0 obtained by the product of the ram and component for the respective airstream 16

17 THRUST SPECIFIC FUEL CONSUMPTION 17

18 SUMMARY OF PARAMETRIC ANALYSIS 18

19 EXAMPLE PARAMETRIC ANALYSIS fan pressure ratio 1.4, flight Mach n. 0.8, standard altitude 40 kft variation of the remaining two design variables 19

20 increasing either the bypass ratio or the compressor pressure ratio generally reduces the specific fuel consumption increasing the bypass ratio naturally reduces the specific thrust for high bypass ratio, increasing the bypass ratio no longer reduces the specific fuel consumption because the bypass and core velocities are so disparate for each compression ratio, there is a bypass ratio that gives the min thrust specific fuel consumption optimum bypass ratio * 20

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22 PERFORMANCE ANALYSIS off-design flight conditions and throttle settings 22

23 ASSUMPTIONS the flow areas are constant at stations 4, 4.5, 6, 16, 6A, 8 dry (AB off) the flow in choked at the high-pressure turbine entrance nozzle (4), at the low-pressure t. (4.5) and at the exhaust nozzle (8) the exhaust nozzle may un-choke at low throttle settings component efficiency and pressure ratio (burner, mixer, AB, exhaust) bleed air and cooling air fractions are constant power takeoffs are constant the air and combustion gases are modeled as perfect gas in thermodynamic equilibrium simplifying gas model: gases are calorically perfect upstream and downstream of the burner and afterburner 23

24 REFERENCING MASS FLOW PARAMETER (MFP) REFERENCING at any off-design point, a relationship between the two performances variables and - the constant can be evaluated at the reference point MASS FLOW PARAMETER calorically perfect gas 24

25 FOR HIGH-PRESSURE TURBINE VARIABLE SPECIFIC HEAT COOLED TURBINE nozzle throat stations just downstream of station 4 and 4.5 denoted by 4 and

26 the flow is adiabatic between 5 and 9 26

27 using referencing 27

28 UNCOOLED TURBINE i = ideal exit state 28

29 new relationships are required for the bypass ratio and the 4 exhaust nozzle dependent variables M 9, M 19, P 9 /P 0, P 19 /P 0 29

30 using referencing 30

31 the only unknowns for solution are the static pressure ratios P 0 /P 9 and P 0 /P 19 UNCHOKED FLOW the exit static pressure P e is equal to the ambient pressure P 0 exit Mach n. < 1 exit Mach n. M e determined using the compressible flow functions 31

32 CHOKED FLOW P te /P 0 obtained by the product of the ram and component for the respective airstream 32

33 SUMMARY OF PERFORMANCE ANALYSIS 33

34 EXAMPLE turbofan engine designed for a Mach n. of 0.8 at a standard altitude of 30 kft VSH (variable specific heat) gas model compressor pressure ratio of 30 ( cl =4, ch =7.5) fan pressure ratio of 1.5 bypass ratio of 8 PERFORMANCE VARIATION for full throttle operation with max compressor pressure ratio of 30 and max T t4 of 3200 R 34

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37 because the component performance curves break at about Mach n. of 4.5 at sea level, this engine has a theta break and throttle ratio (TR) of about 1.04 throttle ratio 37

38 dimensionless free stream temperature theta break point at which the control logic must switch from limiting c to limiting T t4 38

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