Equation(A3.3)says that for a high cycle efficiency, the pressure ratio of the cycle should be increased. Figure A-7 shows the history of aircraft engine pressure ratio versus entry into service and it can be seen that there has been a large increase in cycle pressure ratio. The thermodynamic concepts apply to the behavior of real aerospace devices Trent 890 Trent 77 cF6-80c2A8 CF6-80E1A4 CF6-80C2A8 M565c RB211524D4 CF6-50A RB211-22 CFM56-5B TF39-1 CF6-6d JT9D-70 CFM56-3C JT9D-3A Spey 512Spey 512-14 JT8D-219 BD- Spey 505- JT8D-17 Tay 611 Tay 651 占10 0 2000 Year of certification Figure A-7: Gas turbine engine pressure ratio trends (Janes Aeroengines, 1998) Muddy points When flow is accelerated in a nozzle, doesn t that reduce the internal energy of the flow and therefore the enthalpy? (MP 1A.5 Why do we say the combustion in a gas turbine engine is constant pressure?(MP 1A.6 Why is the Brayton cycle less efficient than the Carnot cycle? (MP 1A.7) still within the system boundary? (MP 1A(( g in the exhaust outside tsg-that Does it matter what labels we put on the corners of the cycle or not?(MP 1A.9) u the work done in the compressor always equal to the work done in the turbine plus Is vork out(for a Brayton cyle)?(MP 1A.10) L.A. 4 Gas Turbine Technology and Thermodynamics The turbine entry temperature, T, is fixed by materials technology and cost. (If the temperature is too high, the blades fail. )Figures A-8 and A-9 show the progression of the turbine entry temperatures in aeroengines. Figure A-8 is from Rolls royce and Figure A-9 is from Pratt whitney. Note the relation between the gas temperature coming into the turbine blades and the blade melting temperature1A-8 Equation (A.3.3) says that for a high cycle efficiency, the pressure ratio of the cycle should be increased. Figure A-7 shows the history of aircraft engine pressure ratio versus entry into service, and it can be seen that there has been a large increase in cycle pressure ratio. The thermodynamic concepts apply to the behavior of real aerospace devices! Figure A-7: Gas turbine engine pressure ratio trends (Jane’s Aeroengines, 1998) Muddy points When flow is accelerated in a nozzle, doesn’t that reduce the internal energy of the flow and therefore the enthalpy? (MP 1A.5) Why do we say the combustion in a gas turbine engine is constant pressure? (MP 1A.6) Why is the Brayton cycle less efficient than the Carnot cycle? (MP 1A.7) If the gas undergoes constant pressure cooling in the exhaust outside the engine, is that still within the system boundary ? (MP 1A.8) Does it matter what labels we put on the corners of the cycle or not? (MP 1A.9) Is the work done in the compressor always equal to the work done in the turbine plus work out (for a Brayton cyle) ? (MP 1A.10) 1.A.4 Gas Turbine Technology and Thermodynamics The turbine entry temperature, Tc , is fixed by materials technology and cost. (If the temperature is too high, the blades fail.) Figures A-8 and A-9 show the progression of the turbine entry temperatures in aeroengines. Figure A-8 is from Rolls Royce and Figure A-9 is from Pratt&Whitney. Note the relation between the gas temperature coming into the turbine blades and the blade melting temperature. 1960 0 10 20 JT3D Conway 508 Conway 550 JT8D-17 JT8D-219 CFM56-3C PW4084 PW4168 CFM56-5B CFM56-5C4 CFM56-2 JT8D-1 Tay 611 Spey 555 Spey 512 JT9D-3A JT9D-70 TF39-1 RB211-22 RB211-524D4 CF6-50E PW4052 CF6-80C2A8 CF6-50A CF6-6 JT9D-7R4G Spey 512-14 Spey 505 Tay 651 GE90 Trent 890 Trent 775 30 40 1970 1980 Year of Certification Overall Pressure Ratio (OPR), Sea Level, T-O 1990 2000 CF6-80C2A8 CF6-80E1A4