ENGINE SELECTION: PERFORMANCE CYCLE ANALYSIS

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Università del Salento - FACOLTA’ DI INGEGNERIA INDUSTRIALE – FACOLTA’ DI INGEGNERIA INDUSTRIALE – BrindisiBrindisi Dipartimento di Ingegneria dell’Innovazione - Lecce

CORSO DI LAUREA SPECIALISTICA IN CORSO DI LAUREA SPECIALISTICA IN Ingegneria Aerospaziale ed AstronauticaIngegneria Aerospaziale ed Astronautica

PROPULSIONE AEROSPAZIALE IPROPULSIONE AEROSPAZIALE I

ENGINE SELECTION: ENGINE SELECTION: PERFORMANCE CYCLE PERFORMANCE CYCLE ANALYSISANALYSISCap. 5 AIAA AIRCRAFT ENGINE DESIGNApp. J AIAA AIRCRAFT ENGINE DESIGNApp. K AIAA AIRCRAFT ENGINE DESIGN

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Prof. Ing. A. [email protected]

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CONCEPT

1st STEP: finding an optimum engine for a particular application

it is time to determine an engine’s steady state performances THE OBJECT OF THE PERFORMANCE CYCLE ANALYSIS

IS TO DETERMINE THE ENGINE’S PERFORMANCE OVER ITS OPERATING ENVELOPE

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a parametric (design point) cycle analysis has been performed for the reference point engine – so-called reference conditions (subscript R):

engine SR, [F/m0]R (thrust specific fuel consumption, uninstalled thrust/mass flow rate)

engine components fR, fR (fan total pressure ratio, total enthalpy ratio)

flight conditions M0R, P0R, T0R (Mach number, pressure, temperature)

SEE APPENDIX H

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PERFORMANCE ANALYSISHIGH BYPASS RATIO TURBOFAN ENGINE - App. J AIAA AIRCRAFT ENGINE DESIGN

off-design flight conditions and throttle settings

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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

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REFERENCINGMASS FLOW PARAMETER (MFP)

REFERENCINGat 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

DETAILED KNOWLEDGE OF EQUATIONS IS NOT REQUIRED FOR THE FINALE EXAM

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FOR HIGH-PRESSURE TURBINE

VARIABLE SPECIFIC HEAT – COOLED TURBINE

nozzle throat stations just downstream of station 4 and 4.5 denoted by 4’ and 4.5’


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the flow is adiabatic between 5 and 9


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using referencing


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UNCOOLED TURBINE

i = ideal exit state


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new relationships are required for the bypass ratio and the 4 exhaust nozzle dependent variables M9, M19, P9/P0, P19/P0


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using referencing


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the only unknowns for solution are the static pressure ratios P0/P9 and P0/P19

UNCHOKED FLOWthe exit static pressure Pe is equal to the ambient pressure P0

exit Mach n. < 1

exit Mach n. Me determined using the compressible flow functions


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CHOKED FLOW

Pte/P0 obtained by the product of the ram and component for the respective airstream


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PERFORMANCE OF TURBINES WITH/WOUT COOLANT MIXERS

1st step is to analyze the behavior of high- and low-pressure turbine they are both deliberately designed to be choked the static pressure downstream the low-pressure turbine is tied to the mixer entrance

conditions the remainder of the engine performance analysis flows directly from this step

because the turbine provide the power for the fan and compressors and control the fan and compressor mass flow

throttle setting (Tt4)

turbine performance t, t primarily determined by efficiency and mass conservation

tH, tH remain constant for an uncooled turbine and calorically perfect gas these ratios can be considered constant for a cooled turbine + calorically perfect gas when a cooled turbine with variable specific heats is modeled, it is found that tH, tH vary

only slightly with engine operating conditions

tL, tL remain constant for an uncooled turbine and calorically perfect gas – for a turbofan engine having choked separated (unmixed) fan and core stream for a mixed flow turbofan, these ratios cannot be considered constant because tL must

modulate to maintain P6=P16 (Kutta condition)

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SUMMARY OF PERFORMANCE ANALYSIS

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ITERATIVE SOLUTION SCHEME

preferred iteration variables mixer bypass ratio ’ core entrance Mach number to the mixer M6

engine mass flow rate m0

M6 varies slightly over the off-design range

m0 has a small influence on cL

the reference point values serve as satisfactory initial estimates

reasonable initial estimate

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VARIATION IN ENGINE SPEED

the change in total enthalpy across a fan or compressor is proportional to the shaft rotation speed N squared

using referencing

for HP compressor

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SOFTWARE IMPLEMENTATION

Engine Test portion of AEDsys program reference point program ONX trust scale factor TSF

input engine’s thrust at sea level, static conditions (FSL) divided by the required sea level, static thrust (TSLreq)

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COMPONENT BEHAVIOR

off-design cycle analysis requires a model for the behavior of each engine component constant efficiency of rotating components and constant total pressure

ratio of the other components give answers adequate for preliminary design

COMPONENTS MAPS

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DIMENSIONLESS – CORRECTED PERFORMANCE

pressure and temperature made dimensionless by dividing each by their respective standard sea level static value

dimensional analysis of engine components yields many useful dimensionless or modified component parameters compressor pressure ratio, adiabatic efficiency, Mach number, ratio of

blade tip speed to the speed of sound

CORRECTED MASS FLOW RATE CORRECTED ENGINE SPEED

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effects of viscosity, Reynolds number, humidity, gas composition neglected

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FAN AND COMPRESSOR MAP

total pressure ratio, corrected mass flow rate, corrected engine, adiabatic efficiency

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COMBUSTOR MAPS

pressure loss performance vs corrected mass flow rate for different fuel-air ratios

efficiency of the combustor vs temperature rise or fuel-air ratio for various inlet pressure

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TURBINE MAPS

the flow through the turbine nozzle is a function of the turbine pressure ratio (only when not choked nozzle), inlet total pressure, inlet total temperature

total pressure ratio (expansion ratio) plotted as function of corrected mass flow rate and corrected mechanical speed

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for reasons of size, it is desirable that the turbine mass flow per annulus area be as large as possible choking + high pressure ratio all speed characteristics collapse onto a single line

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COMPONENT MATCHING

the off-design equations are silent with regard to the rotational speed of the rotating machines the rotational speed will, in general, be different from the design point it would seem that off-design equations set lacks some true physical

constraints this appearance is fortunately misleading

the main use of enforcing Nc=Nt is to provide accurate values of compressor and turbine efficiency

according to a typical turbine performance map, this machine can provide the same work for a wide range of N while varies slightly

THE TURBINE FLOW CONDITIONS CAN ADJUST THEMSELVES IN ORDER TO PROVIDE THE SAME AT DIFFERENT VALUES OF r

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consider single-stage, impulse, max work (no exit swirl), isentropic, constant height turbine – choked inlet guide vane and entirely subsonic flow in the rotor

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(1-t) is proportional to rm(v2R-v3R)

M2, M2R, M3R must increase to compensate for reduction in rm

the turbine provide the same t with slight changes in t for a wide range of Nc=Nt

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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 Tt4 of 3200 °R

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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

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dimensionless free stream temperature

theta breakpoint at which the control logic must switch from limiting c to limiting Tt4

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PERFORMANCE ANALYSISTURBOPROP ENGINEApp. K AIAA AIRCRAFT ENGINE DESIGN

off-design flight conditions and throttle settings

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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

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REFERENCINGMASS FLOW PARAMETER (MFP)

REFERENCINGat 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


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FOR HIGH-PRESSURE TURBINE

VARIABLE SPECIFIC HEAT – COOLED TURBINE

nozzle throat stations just downstream of station 4 and 4.5 denoted by 4’ and 4.5’


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power balance of the high-speed spool

The total pressure ratio of the compressor is determined from the component efficiency equation


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An expression for the mass flow rate of air through the turboprop engine at any conditions can be obtained for the case of a CPG when

The power produced by the low-pressure turbine at off-design is determined by its total pressure ratio

The total temperature ratio of the low-pressure turbine is related by the CPG component efficiency relationship


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work interaction coefficient for the propeller is obtained by a power balance of the low-pressure spool

Determination of the value of Cprop thus depends mainly on Tt4 (τλ) and τtL since ηprop, ηg, and ηmL, τm1, τtH, and τm2 are constant or essentially constant

Note that the flow is choked at the entrance to the low-pressure turbine (station 4.5) and is normally unchoked at the engine exit (station 9)


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The total pressure at station 4.5 is related to the flight condition and throttle setting by

the total pressure at station 5 is related to the nozzle operation by


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the only unknowns for solution is the static pressure ratios P0/P9

UNCHOKED FLOWthe exit static pressure P9 is equal to the ambient pressure P0

exit Mach n. < 1

exit Mach n. M9 determined using the compressible flow functions


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CHOKED FLOW

Pt9/P0 obtained by the product of the ram and component for the core airstream


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Equating the mass flow rate of air at stations 4.5 and 9 yields the following equation for the total pressure ratio of the low-pressure turbine in terms of the exit Mach number and the total temperature ratio of the low-pressure turbine


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Determination of the conditions downstream of station 4.5 requires an iterative solution


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The performance of the propeller can be simply modeled as a function of the flight Mach number

The equation for the Mach range of 0.7 - 0.85, given above, models the drop in ηprop experienced in this flight regime due to transonic flow losses in the tip region of the propeller


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SUMMARY OF PERFORMANCE ANALYSIS

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EXAMPLE

turboprop engine designed for a Mach n. of 0.8 at a standard altitude of 25 kft

compressor pressure ratio of 30

turbine enthalpy ratio of 0.6

ENGINE CONTROL: max compressor pressure ratio of 30, max Tt4 of 3200 °R, max Tt3 of 1600 °R

variation with changes in flight Mach n. and altitude – full throttle operation0break = 0.933

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sharp break at M0=0.1 due to the assumed prop

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S is very low compared with other turbine engine cycles

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EXAMPLE ENGINE SELECTION: PERFORMANCE CYCLE ANALYSIS

AAF

request for proposal WTO/S=64 lbf/ft2 TSL/WTO=1.25

mission analysis WTO=24000 lbf, S=375 ft2, TSL=30000 lbf

parametric analysis narrowed the range of engine design choice the search must focus on reduced fuel consumption

parametric sensitivity analysis led to the conclusion that the selection of the engine design point fan and compressor pressure ratios

should be from the high sides of their respective ranges the design point combustor and AB temperatures should be allowed to be drift

down from their limiting value

the high altitude and high Mach n. operational requirements require an engine with a high specific thrust to obtain a low frontal area for reducing drag afterburning, low bypass, mixed flow turbofan

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EXAMPLE ENGINE SELECTION: PERFORMANCE CYCLE ANALYSIS

AAF

request for proposal WTO/S=64 lbf/ft2 TSL/WTO=1.25

mission analysis WTO=24000 lbf, S=375 ft2, TSL=30000 lbf

parametric analysis narrowed the range of engine design choice the search must focus on reduced fuel consumption

parametric sensitivity analysis led to the conclusion that the selection of the engine design point fan and compressor pressure ratios

should be from the high sides of their respective ranges the design point combustor and AB temperatures should be allowed to be drift

down from their limiting value

the high altitude and high Mach n. operational requirements require an engine with a high specific thrust to obtain a low frontal area for reducing drag afterburning, low bypass, mixed flow turbofan

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CRITICAL FLIGHT CONDITIONS

first engine to be selected: baseline engine performance of this engine over the critical legs of the mission

CRITICAL LEGS high fuel consumption (high ) represents a boundary of the solution space extreme operating conditions

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MISSION FUEL CONSUMPTION

mission fuel usage plays a dominant role in the selection of the engine COMPUTER CALCULATED ALGEBRAIC ESTIMATE

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GETTING STARTED

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BASELINE ENGINE

reference point parameters

reference point performance

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OPERATING ENVELOPE – BASELINE ENGINE

it is important to first ascertain the Mach n. and altitude ranges over which the engine can operate at full throttle (Tt4max) – its full throttle envelope

an engine may not be able to operate at certain combination of M and altitude Kutta conditions (P6=P16) and power balance between fan + low-p.

compressor and the low-p. turbine (which fixes P6) must be satisfied simultaneously

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performance at partial throttle

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MISSION PERFORMANCE – BASELINE ENGINE

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THE SEARCH – baseline engine: engine 1

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the cycle pressure ratio was the first design choice varied as it was changed from 20 to 28 increasing amounts of fuel were

saved

the fan pressure ratio was the next design choice varied as it was changed from 3.9 to 3.3 increasing amounts of fuel were

saved

the bypass ratio is no longer an independent variable when to match total pressure at the mixer entrance

the total temperature leaving the engine during dry operation at the supercruise flight conditions decreases with increasing compressor pressure ratio and decreasing fan pressure ratio a low Tt6A is desirable to reduce the infrared signature of the aircraft

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fuel consumption is reduced by increasing cycle pressure ratio and engine bypass ratio – the reverse is true for the specific trust

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OPERATIONAL ENVELOPE AND MISSION PERFORMANCE

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SECOND REPRISE

now that the performance of the selected engine is know in terms of the fuel used, the effects of these changes in aircraft weight on aircraft performance can be updated

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although low fuel consumption is essential to achieving low WTO, it also results in increased values of the instantaneous weight fraction which increases the thrust loading TSL/WTO

increase in landing weight fraction requires a decrease in the wing loading or an increase of the max lift coefficient for landing

ENGINE SELECTION: PERFORMANCE CYCLE ANALYSIS

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