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MAE 5380: Advanced Propulsion

MAE 5380: Advanced Propulsion. GAS TURBINE PERFORMANCE Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk. AEROENGINE CYCLE ANALYSIS. Cycle Analysis [What determines the engine characteristics?]

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MAE 5380: Advanced Propulsion

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  1. MAE 5380: Advanced Propulsion GAS TURBINE PERFORMANCE Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk

  2. AEROENGINE CYCLE ANALYSIS • Cycle Analysis [What determines the engine characteristics?] • Cycle analysis is the study of the thermodynamic behavior of air as it flows through the engine without regard for the mechanical means used to effect its motion • We characterize components by the effects they produce • Actual engine behavior is determined by geometry; cycle analysis is sometimes characterized as representing a “rubber engine” • Main purpose is to determine which characteristics to choose for the components of an engine to best satisfy a particular need

  3. Chemical Energy Heat (Thermal Energy) Mechanical Power Mech. Power to GasFlow Thrust Power GAS TURBINE CYCLE ANALYSIS A gas turbine engine is an example of a thermal engine. A useful conceptual framework to analyze this is through the propulsion chain. Combustion Thermal Mechanical Propulsive The overall efficiency for the propulsion chain is given by:

  4. INDUSTRY TERMINOLOGY FOR OVERALL FIGURESOF MERIT [Pratt &Whitney]

  5. EVOLUTION OF THE AEROENGINE [Koff]

  6. SUBSONIC ENGINE SFC TRENDS(35,000 ft. 0.8 Mach Number, Standard Day [Wisler])

  7. Fuel Consumption Trend Fuel Burn JT8D PW4084 JT9D Future Turbofan PW4052 1950 1960 1970 1980 1990 2000 2010 2020 Year

  8. MAJOR TYPES OF THRUST PROPELLED AIRCRAFT [Walsh and Fletcher]

  9. CONCEPTS/TOOLS FOR ENGINE IDEAL CYCLE ANALYSIS • Concepts of stagnation and static temperature and pressure • Ideal gas equation of state [p = RT] and other attributes • One-dimensional gas dynamics • Thermodynamic laws (there are only two that we need!) • Behavior of useful quantities: energy, entropy, enthalpy • Relations between thermodynamic properties in a reversible (“lossless”) process • Relations between Mach number and thermodynamic properties (static and stagnation quantities) • Properties of cycles

  10. STAGNATION QUANTITIES DEFINED • Quantities used in describing engine performance are the stagnation • pressure, enthalpy and temperature. Stagnation enthalpy, ht , enthalpy state if the stream is decelerated adiabatically to zero velocity

  11. FOR A REVERSIBLE, ADIABATIC (I.E. ISENTROPIC) PROCESS

  12. ANALYSIS OF THE LINKS IN THE PROPULSION CHAIN Examine the magnitudes and behavior of the different efficiencies is greater than 0.99 for civil aircraft at cruise. It is set by the details of the combustion processes. We will take it to be unity in these lectures is a measure of the mechanical losses (such as bearing friction). It is also close to unity. The main figures of merit for gas turbine engines are the thermal and propulsive efficiencies. We can approximate the propulsion chain as We examine thermal efficiency first Thermal efficiency: Heat (actually fuel) is input. An airstream, , mass flow rate, passes through the engine. The mechanical work is the change in kinetic energy, which is, per unit mass,

  13. THERMODYNAMIC PROCESSES IN THE ENGINE • How should we represent the thermodynamic process in the engine? • It is cyclic (the air starts at atmospheric pressure and temperature and ends up at atmospheric pressure and temperature) • Consider a parcel of air taken round a cycle with heat addition and rejection. • Need to consider the thermodynamics of this propulsion cycle • To do this we make use of the First and Second Laws of Thermodynamics • We will review (one chart each) these concepts

  14. RECAP ON THERMODYNAMICS (I) First law (conservation of energy) for a system: “chunk” of matter of fixed identity E0 = Q - W Change in overall energy (E0 ) = Heat in - Work done E0 = Thermal energy + kinetic energy ... Neglecting changes in kinetic and potential energy E = Q - W ;(Change in thermal energy) On a per unit mass basis, the statement of the first law is thus : e = q - w

  15. RECAP ON THERMODYNAMICS (II) The second law defines entropy, s, by: Where dqreversible is the increment of heat received in a reversible process between two states The second law also says that for any process the sum of the entropy changes for the system plus the surroundings is equal to, or greater than, zero Equality only exists in a reversible (ideal) process

  16. REPRESENTING THE ENGINE PROCESS IN THERMODYNAMIC COORDINATES First Law: E = Q - W, where E is the total energy of the parcel of air. For a cyclic process E is zero (comes back to the same state) Therefore: Q (Net heat in) = W (Net work done) Want a diagram which represents the heat input or output. A way to do this is provided by the Second Law where ds is the change in entropy of a unit mass of the parcel and dq is the heat input per unit mass Thus, one variable should be the entropy , s

  17. Other variable: look at the control volume form of the first law, sometimes known as the “Steady Flow Energy Equation” For any device in steady flow Heat input Per unit mass flow rate: 2 1 Mass flow Device Shaft work q is heat input/unit mass wshaft is the shaft work / unit mass

  18. The form of the steady flow energy equation shows that enthalpy, h, h = e + pv = e + p/r is a natural variable to use in fluid flow-energy transfer processes. For an ideal gas with constant specific heat, dh = cpdT. Changes in enthalpy are equivalent to changes in temperature. To summarize, the useful natural variables in representing gas turbine engine processes are h,s (or T, s). We will represent the thermodynamic cycle for a gas turbine engine on a T,s diagram

  19. Gas Turbine Engine Components: • Inlet - Slows, or diffuses, the flow to the compressor • Fan/Compressor (generally two, or three, compressors in series) does work on the air and raises its stagnation pressure and temperature • Combustor - heat is added to the air at roughly constant pressure • Turbine (generally two or three turbines in series) extracts work from the air to drive the compressor or for power (turboshaft and industrial gas turbines) • Afterburner (on military engines) adds heat at constant pressure • Exhaust nozzle raises the velocity of the exiting mass flow • Exhaust gases reject heat to the atmosphere at constant pressure

  20. THERMODYNAMIC CHARACTERISTICS OF THE COMPONENTS(Ideal Components)

  21. THERMODYNAMIC MODEL OF GAS TURBINE ENGINE CYCLE[Cravalho and Smith] 4 3 2 5

  22. GAS TURBINES: VARIATIONS ON A CORE THEME [Cumpsty]

  23. GAS TURBINE ENGINE CORE [Cumpsty]

  24. SCHEMATIC CONDITIONS THROUGH A GAS TURBINE [Rolls-Royce]

  25. NOMINAL PRESSURES AND TEMPERATURES FOR A PW4000 TURBOFAN [Pratt&Whitney]

  26. COMMERCIAL AND MILITARY ENGINES(Approx. same thrust, approx. correct relative sizes) GE CFM56 for Boeing 737 P&W 119 for F- 22

  27. SOME TYPES OF GAS TURBINE ENGINES (I) [Rolls-Royce]

  28. SOME TYPES OF GAS TURBINE ENGINES (II) [Rolls-Royce]

  29. TYPES OF GAS TURBINE ENGINES (III) [Rolls-Royce]

  30. SCHEMATIC OF TURBOPROP ENGINE[Kerrebrock]

  31. AFTERBURNING ENGINES [Pratt& Whitney]

  32. STOVL CTOL UK/RN CV Lockheed Martin image, courtesy LMTAS

  33. Lockheed Martin Propulsion System Shaft-Driven Lift Fan Concept F119 Derivative Engine Roll Control Ducts Engine Nozzle (Up-and-Away Position) Allison / R-R Lift Fan Engine Nozzle (Vertical Thrust Position) [Davenport] image, courtesy PW

  34. Engines for JSF JSF F120 Engine F119 Derivative Low Observable Nozzle Low Observable Features Lightweight Structures Matured F119 Core Integrated Sub-Systems Enhanced Diagnostics Integrated Sub-Systems • Objectives: • Single Engine Safety • Affordability • Improved R&M Matured Control System w/Enhanced Diagnostics [Davenport] image, courtesy GE image, courtesy PW

  35. PRESSURE RATIO TRENDS [Jane’s 1999]

  36. AEROENGINE CORE POWER EVOLUTION: DEPENDENCE ON TURBINE ENTRY TEMPERATURE [Meece/Koff]

  37. IDEAL CYCLE ANALYSIS • Our objective is to express thrust, F, and thermal efficiency, (or alternatively ) as functions of • typical design limiters • flight conditions • design choices so that we can analyze the performance of various engines. Heat Addition T Expansion Compression Cooling

  38. TURBOJET ENGINE SHOWING STATIONS AND COMPONENTNOTATION[Kerrebrock]

  39. CALCULATION OF THRUST AND SPECIFIC IMPULSE IN TERMS OF COMPONENT PERFORMANCE • If the component performance is known, we can compute the thrust, specific impulse (or TSFC), thermal and propulsive efficiencies • The procedure will be shown here conceptually for a simple example a turbojet with an ideally expanded nozzle and the contribution of fuel mass to exit momentum flux neglected.

  40. METHODOLOGY • Find thrust by finding u7/uo (uexit/uo) in terms of q, temperature ratios, etc. • Use a power balance to relate turbine parameters to compressor parameters • Use an energy balance across the combustor to relate the combustor temperature rise to the fuel flow rate and fuel energy content.

  41. METHODOLOGY (II) • Write expressions for thrust and efficiency (Isp) • That is the easy part. Now comes the algebra...

  42. NOMENCLATURE  = total or stagnation pressure ratio across component (d, c, b, t, a, n)  = total or stagnation temperature ratio across component (d, c, b, t, a, n) , , ,

  43. IDEAL ASSUMPTIONS 1) Inlet/Diffuser: d = 1, d = 1 (adiabatic, isentropic) 2) Compressor or fan: c = cg-1/g , f = fg-1/g 3) Combustor/burner or afterburner: b = 1, a = 1 4) Turbine: t = tg-1/g 5) Nozzle: n = 1, n = 1

  44. THE ALGEBRA (I) • Working towards an expression for u7 / u0

  45. THE ALGEBRA (II) • Applying a similar procedure for the exit pressure • Equate this box to the previous box to get

  46. THE ALGEBRA (III) • Continuing on our path to find u7/u0 • Therefore

  47. THE ALGEBRA (IV) • Now relate temperature ratio across turbine to that across the compressor • This can be rewritten as • So ; or

  48. THE ALGEBRA (V) • One more substitution to write the temperature rise across the combustor in terms of t=TT4/To • Then substituting everything in • Thrust per unit mass flow (non-dimensionalized by the ambient speed of sound) as a function of design parameters and flight conditions

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