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Conception Mecaturbo 2013 [Mode De Compatibilité] Cours de conception aéronautique Introduction au dimensionnement mécanique et dynamique des turboréacteurs Reference : AERO0015-1 - MECHANICAL DESIGN OF TURBOMACHINERY - 5 ECTS - J.-C. GOLINVAL Dimensionnement mécanique des turbomaréacteurs des mécanique Dimensionnement 2 Principles of jet propulsion Comparison between the working cycle of a turbo-jet engine Air intake Compression Combustion Exhaust and a piston engine Air intake Compression Combustion Exhaust 3 Principles of jet propulsion For aircraft propulsion, the thrust is produced by the acceleration of the air passing through the engine. Low pressure (LP) Combustion Low pressure (LP) compressor chamber turbine Exhaust High pressure (HP) High pressure (HP) compressor turbine 4 Principles of jet propulsion For any flow section of the engine, the calculation of the thrust may be achieved using the following formula: Thrust = Pressure x Area of flow section + Mass flow rate x Velocity of flow Example of a typical single-pool axial flow engine Fcompressor Fcombustion Fturbine Fexhaust (85 kN) 150 kN 180 kN 25 kN 235 kN 205 kN Total thrust = 30 kN. Forward. 5 Brief review of the evolution of turbojet engines France 1921 GUILLAUME patent combustor axial-flow axial-flow turbine compressor At that time, these patents clearly described the air-breathing jet principle but were not executed in practice because of the high exhaust speed (e.g. 900 km/h) which was not compatible with the too-slow aerovehicle at that time (e.g. < 300 km/h in the early 1920s) 6 Brief review of the evolution of turbojet engines United Kingdom 1930 : patent of Frank WHITTLE’s turbojet engine combustor radial compressor stage exhaust nozzle Multistage, axial- flow compressor axial-flow turbine 7 Brief review of the evolution of turbojet engines United Kingdom 1937 : 1st test of Whittle’s engine , the W.U. (Whittle Unit) May 15, 1941 : 1st flight of the Wittle’s engine on the Gloster experimental aircraft Compressor Diffuser 1937 Wittle’s engine (WU) Turbine Helicoidal combustion chamber 8 Brief review of the evolution of turbojet engines United States • In September 1941, an agreement was signed between Great Britain and the United States so that the United States could have the engine W1X and a set of drawings of the Whittle W2B jet engine. • General Electric was chosen for jet engine development. • The W2B engine was built and tested on March 18, 1942, under the name GE 1-A. 9 Brief review of the evolution of turbojet engines Germany 1936 : development of an hydrogen turbojet demonstrator (He S-1) by Hans VON OHAIN in the HEINKEL company 1937 : design of the flight engine (He.S3) August 27, 1939 : 1st flight of the experimental He-178 airplane with jet engine He.S3B Jet engine Heinkel He.S3B This was the first flight of a turbojet aircraft in the world. 10 Brief review of the evolution of turbojet engines Germany First jet engine leading to mass production JUMO 004 turbojet engine Starter motor Control cone Cooling air duct Combustion chamber 11 Brief review of the evolution of turbojet engines Germany JUMO 004 development and production schedule Start of development Fall 1939 First test run Oct. 11, 1940 First flight in Me-262 July 18, 1942 Beginning of production Early 1944 About 6000 engines delivered May 1945 Engine life of about 25 hours (above the combat life of a German fighter !) Failure was due to the lack of turbine materials capable of withstanding the high stresses at high temperatures. 12 Brief review of the evolution of turbojet engines From 1940, design and construction of new turbojets increase drastically. Advanced concepts appeared very early after the principle of turbojet engine was discovered but their developments met two technological problems: • The complete lack of heat-resistant materials. • The blade air cooling technology. 13 Development process of the structural design Industrialization Testing Program Preliminary Certification Request Detailed Design Design For Review Proposal Program First Engine Launch To Turn milestone Definition of the thermodynamic cycle consistent with the specification and of the best structural architecture of the engine Today, the duration of this process is about 3 years (against ~ 6 years in the 90s) 14 Preliminary design Thrust requirement: for various operating points (limits of the flight envelope) Engine maximum power Altitude Aircraft structure maximum temperature Minimum (max Mach) aircraft Aircraft structure maximum pressure aerodynamic lift Speed Electric power requirement: increasing need in modern aircrafts (electric actuators, navigation systems) This has an impact on the engine architecture and on the structural design. 15 Challenges of turbojet technology Overall efficiency of a jet propulsion engine Overall efficiency η= η × η overall thermal propulsive Thermal efficiency Propulsive efficiency 16 Challenges of turbojet technology Thermal efficiency The thermal efficiency is defined as the ratio of the net power out of the engine to the rate of thermal energy available from the fuel According to the T-s diagram of an ideal turbojet engine, the thermal efficiency simplifies to η = − T0 thermal 1 T3 Thermodynamic cycle 3 4 2 1 e Temperature 0 0 1 2 3 4 e Entropy 17 Challenges of turbojet technology Propulsive efficiency The propulsive efficiency is defined as the ratio of the useful power output (the product of thrust and flight velocity, V0 ) to the total power output (rate of change of the kinetic energy of gases through the engine). This simplifies to thrust flight velocity η =F V 0 = 2 propulsive & + Wout V e V 0 1 total power exit velocity output 18 Challenges of turbojet technology If the propulsive efficiency is plotted versus the velocity ratio Ve / V 0 , it follows that high propulsive efficiency requires the exit velocity to be approximately equal to the inlet velocity. 100 90 80 70 propulsive 60 η 50 40 1 2 3 4 Ve / V 0 19 Challenges of turbojet technology To progress to the performance capabilities of today, two goals were (and still are) being pursued: 1. Increase the thermodynamic cycle efficiency by increasing the compressor pressure ratio. Trend in compressor pressure ratios Calendar years Compressor pressure ratio Late 1930 to 1940 3:1 to about 6:1 Early 1950 About 10:1 1950 to 1960 20:1 to about 25:1 2000 30:1 to about 40:1 20 Challenges of turbojet technology 2. Increase the ratio of power-output to engine weight by increasing the turbine inlet temperature Trend in turbine inlet temperatures °C Military Commercial Turbine Turbine inlet temperature Year 21 Challenges of turbojet technology Depending on the types of applications, different development goals may be pursued. Supersonic flight (military engines) Maximum thrust is sought by increasing the exit velocity (at the expense of fuel economy) and decreasing the engine inlet diameter (i.e. of the aerodynamic drag) Example SNECMA M88 military engine (used on the RAFALE airplane) 22 Challenges of turbojet technology Supersonic flight If the turbine inlet temperature reaches its limit, thrust augmentation by afterburning. Thermodynamic cycle Afterburner 3 6 Nozzle 2 4 1 Temperature 0 0 1 2 3 4 5 6 S Entropy 23 Challenges of turbojet technology Example of evolution of military engine performances Comparison between SNECMA M88 military engine (used on the RAFALE airplane) against ATAR engine (used on MIRAGE 3 airplane) ATAR M88 Power/weight ratio base x 2 Fuel Consumption base - 30 % Engine diameter base - 40 % Turbine entry temperature (°K) 1210 1850 Pressure ratio 6,1 24 Compressor exit temperature (°K) 600 880 24 Challenges of turbojet technology SNECMA M88 engine Disks ∆∆∆σσσ r = 1000 MPa 1580 °C ∆∆∆σσσθθθ = 1000 MPa θθθ = 650 °C locally 600 °C max Titanium Nickel base alloy superalloy 25 Challenges of turbojet technology Subsonic flight (commercial engines) A low thrust specific fuel consumption is sought by increasing the propulsive efficiency the principle is to accelerate a larger mass of air to a lower velocity. Solution: principle of the by-pass engine (called turbofan) Note : to keep a high thermal efficiency, the pressure ratio and the turbine inlet temperature must remain high. Drawback : the frontal area of the engine is quite large more drag and more weight result 26 Challenges of turbojet technology Solution: principle of the by-pass engine (called turbofan) Air flow rate ~ 5 to 7 times the air flow rate going through the gas generator 27 Challenges of turbojet technology Trend in thrust specific fuel consumption Single-pool axial flow turbojet Twin-spool by-pass turbojet Twin-spool front fan turbojet Advanced technology (high by-pass ratio) Propfan Year 28 Challenges of turbojet technology Evolution of turbojet engines to the technology level of today In conclusion, • new concepts or technological breakthroughs are rare; • advancements are rather due evolutionary improvements of the design To achieve good performances, parallel research and development effort were undertaken in areas such as in aerodynamics, aerothermics, acoustics, combustion process, mechanics, metallurgy, manufacturing, … Aim of the course Study the mechanical aspects of the design. 29 Mechanical challenges of turbojet technology Evolution of turbojet engines - mechanical aspects 1. Increasing of the compressor pressure ratio (r) In the early turbojets: r ~ 6:1 – 8:1 Increasing r « variable » geometry to adapt the compressor behavior to various regimes 30 Mechanical challenges of turbojet technology Solution n° 1 : concept of variable stator blades • Design of reliable airflow control systems • Prevention of air leakage at the pivots of the vanes at high pressures (temperatures). HP compressor Variable stator vanes 31 Mechanical challenges of turbojet technology Solution n° 2: concept of multiple rotors (r ~ 20:1 - 30:1) Example of a dual-rotor configuration Fan HPC HPT LPT CFM56 Advantages • Selection of optimal speeds for the HP and LP stages. • Reduction of the number of compressor stages. • Cooling air is more easily taken between the LP and HP rotors.
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