combustion chamber thereby causing high temperature rise. ⢠If combustion ... Comparison between turbojet,ramjet,scramjet and rocket .... horsepower with forward speed. ... once it goes out of the engine body it is not of any use for thrust production. ... â¢For jet engines with reheat or afterburning, the fuel consumption would.
Air-Breathing Engines Unit 1
Shiva U
1
Aerospace Propulsion II
Faculty
Shiva U Asst. Prof. (Sr. Scale)
Shiva U
2
Syllabus • Classification, operational envelopes; Description and function of gas generator, turbojet, turbofan, turboprop, turbo shaft, ramjet, scramjet, turbojet/ramjet combined cycle engine; Engine thrust, takeoff thrust, installed thrust, thrust equation; Engine performance parameters, specific thrust, specific fuel consumption and specific impulse, thermal efficiency, propulsive efficiency, engine overall efficiency and its impact on aircraft range and endurance; Engine cycle analysis and performance analysis for turbojet, turbojet with afterburner, turbofan engine, turboprop engine.
Shiva U
3
Shiva U
4
Operational Envelopes Each engine type will operate only within a certain range of altitudes and Mach numbers (velocities). The approximate velocity and altitude limits, or corridor of flight, within which airlift vehicles can operate. The corridor is bounded by a lift limit, a temperature limit, and an aerodynamic force limit.
Shiva U
5
Shiva U
6
Shiva U
7
Shiva U
8
Air Breathing Propulsion • Air breathing propulsion systems use oxygen in atmospheric air to burn fuel stored on the vehicle • Turbojet • Turbofan (High BR, Low BR, Afterburning) • Turboprop • RAMJETS • SCRAMJETS
Shiva U
9
Air Breathing Propulsion: Gas Turbine Systems Gas Generator • The basis of turbojet, turbofan, and turboprop propulsion is the gas generator • Supplies high-temperature, high-pressure gas • Stand alone, most of the energy of this device is used to drive turbines • Turbine rotational energy is converted into electricity
Shiva U
10
Air Breathing Propulsion Turbojet
By adding an inlet and a nozzle a turbojet can be constructed • Gas generator still supplies high-temperature, high-pressure gas • Some of the energy of this device is used to drive turbines and auxiliary systems • Most of the energy in the high-temperature, high-pressure gas is allowed to flow to the nozzle • Nozzle accelerates flow to high velocity to impart thrust Shiva U
11
Air Breathing Propulsion: Gas Turbine Systems
Shiva U
12
Air Breathing Propulsion: Gas Turbine Systems
Shiva U
13
Air Breathing Propulsion: Gas Turbine Systems
Shiva U
14
Air Breathing Propulsion: Gas Turbine Systems
Shiva U
15
Air Breathing Propulsion: Ducted Systems
Shiva U
16
Air Breathing Propulsion: Ducted Systems
Shiva U
17
Air Breathing Propulsion: Ducted Systems
Shiva U
18
Air Breathing Propulsion: Ducted Systems
Shiva U
19
TURBOJET Air sucked in through the inlet diffuser Compressed and used to burn the fuel in the combustor Combustion products used to drive the turbine Exhaust through the nozzle to generate jet propulsion
Shiva U
20
Shiva U
21
LIMITATIONS At higher Mach numbers Fuel consumption increases Moving parts do not contribute to engine power ° Share of
compressor at Mach 1 = 50 % Mach 2 = 15 % Mach 3 = 04 %
Moving parts causes losses High temperatures (around 3000 K) are produced Compressor blades cannot withstand that temperature No such high temperature withstanding blade material exists Compression created by speed is enough to keep engine process
Shiva U
22
RAMJET At speeds above Mach 3 a passive intake can compress the air due to ramming effect (without use of compressor) for subsonic combustion in the combustion chamber
Shiva U
23
Mach number decreased and point b kept constant TSFC becomes high Larger size of engine heat added increased and point d kept constant increase in maximum temperature
material properties of engine walls Shiva U
24
PERFORMANCE
Experimental Conditions Inlet temp = 220 K Cp = 0.24 kcal/kg-k γ = 1.4 Shiva U
25
FUELS USED Gaseous Fuel Ramjet * eg. hydrogen Liquid Fuel Ramjet * kerosene , synthetic hydrocarbon fuel eg. US made RJ1, RJ4 ; French CSD07T , CSD15T
Solid Fuel Ramjet * polymers loaded with metal particles like Mg ,Al or B eg. Polyether , polyester , polyurethane
Shiva U
26
ADVANTAGES Able to attain high speeds up to mach 5
No moving parts so less wear & tear and minimum losses Reduced weight and smaller engine Lighter and simpler than turbojet Higher temperatures can be employed
DISADVANTAGES Bad performance at lower speeds Needs booster to accelerate it to a speed where ramjet begins to produce thrust Higher fuel consumption Maximum operating altitude is limited High temperature material required Shiva U
27
SCRAMJET Supersonic Combustion RAMJET In ramjet supersonic speed of air is reduced to subsonic speeds in combustion chamber thereby causing high temperature rise. If combustion is done at supersonic speed temperature rise could be avoided. Achieving supersonic combustion is the ultimate challenge Dwell time in the combustor is low
Shiva U
28
SUPERSONIC COMBUSTION Major Issues
# Proper mixing # Ignition # Stable combustion
Flight Mach no. is 6 to 10 Inlet Mach no. is 2 to 4
Blockage caused by injection and heat release generates a “shock train” intense mixing and combustion with large gradients in flow properties & chemical composition in the axial,radial and circumferential directions Divergent combustor adds to proper mixing and ignition and to compensate for the pressure rise Shiva U
29
Schematic diagram of a scramjet engine
Shiva U
30
Shiva U
31
Hyper-X flight trajectory Shiva U
32
Comparison between turbojet,ramjet,scramjet and rocket
Shiva U
33
Combined cycle engines • Hypersonic propulsion systems can be categorized as liquid- and solid-fueled rockets, turbojets, ramjets, ducted rockets, scramjets, and the dual-combustion ramjet (DCR ). All existing hypersonic systems use either liquid or solid rockets as their propulsion system. • Ramjets and scramjets can operate efficiently at supersonic and hypersonic speeds, but there tend to be limitations to the range of Mach numbers over which they can operate. • For instance, the need to have sufficient compression in the inlet ordinarily requires that the ramjet engine operate supersonically. The inefficiencies of slowing the flow down to subsonic speeds makes the ramjet difficult to use for speeds exceeding Mach 5. Shiva U
34
• Scramjets can be used above approximately Mach 5 but below that there is in general insufficient energy in the captured airstream to enable efficient combustion in the supersonic combustor. • Both the ramjet and scramjet must be coupled with some additional form of propulsion (for missiles, this is chiefly a rocket) to accelerate the vehicle to its “take-over” Mach number. • To overcome these limitations, combined cycle engines have been developed.
Shiva U
35
Engine cycles for hypersonic vehicles Shiva U
36
• A subsequent version of the X-43B has had to be fully multi-use, and its main objective was to test the hybrid power plant type RBCC (rocket-based-combinedcycle), TBCC (turbine-based-combined cycle) and AAR (air-augmented rocket). • The engine type is TBCC turbines placed separately over a high-speed ram and part has its own entrance channel and nozzle. Most comprehensive and most complex the engine AAR. • It has all kinds of drive concentrated in a single flow channel. During take-off of the rocket engine flows into the combustion chamber, additional fuel in excess of oxygen, thereby increasing the tension of almost 50%. After reaching a speed of Mach 2, the rocket motor shuts down and re-activation occurs only in the absence of atmospheric oxygen in orbit. Wiring engine AAR and the principle of its operation is shown next slide. RBCC engine is its principle similar to AAR, but can operate in a wider speed range by changing the geometry of the engine. The advantages of such a drive train are clear. After minimal adjustments feasible in flight, can work in rocketry, Ramjet, maximum or jet mode. Shiva U
37
Shiva U
38
Combined cycle engines- Turbo-Ramjet • A simple combined cycle is a turbojet (or turbofan)/ramjet in which a secondary flow bypasses the core turbojet and participates to produce thrust in an afterburner. • As the Mach number increases, typically beyond M = 3, the afterburner transitions to operation as a ramjet while the turbojet maximum cycle temperature is reduced to maintain an acceptable load on the rotating machinery while maintaining the airflow path open to contribute to thrust generation in the afterburner. • The main issue in this configuration is, evidently, the matching of the core flow with the bypass flow to avoid reversed flow on any of the sides. Additional operational difficulties derive from the broad bypass ratio range during acceleration and deceleration and the thermal management of the moving parts during high enthalpy flight. Shiva U
39
Shiva U
40
Shiva U
41
Engine issues for hypersonic airbreathing propulsion systems
Shiva U
42
Total Temperature Rise with Increasing Mach Number in Trans-atmospheric Flight Shiva U
43
Shiva U
44
Shiva U
45
Thrust needs to be created for all flight regimes of the aircraft: • • • • • •
Take-off – normally maximum thrust Climb –reducing from maximumthrust Cruise –normally minimum thrust Manoeuvres – variable thrust Acceleration & Deceleration - variable Descend – Low thrust
• Landing – Less than maximumthrust
Shiva U
46
Jet Engine Performance • It is seen that engine thrust is proportional to the mass flow rate through the engine and to the excess of the jet velocity over the flight velocity. • The specific thrust of an engine is defined as the ratio of the engine thrust to its mass flow rate. The specific thrust is
Shiva U
47
• Because the engine mass flow rate is proportional to its exit area, A5/m depends only on design nozzle exit conditions. • As a consequence, F/m is independent of mass flow rate and depends only on flight velocity and altitude. • Assigning an engine design thrust then determines the required engine-mass flow rate and nozzle exit area and thus the engine diameter. Thus the specific thrust, F/m, is an important engine design parameter for scaling engine size with required thrust at given flight conditions. • Another important engine design parameter is the thrust specific fuel consumption, TSFC, the ratio of the mass rate of fuel consumption to the engine thrust • Low values of TSFC, of course, are favorable. The distance an aircraft can fly without refueling, called its range, is inversely proportional to the TSFC of its engines. Shiva U
48
Specific thrust may be written as :
• For a reasonable positive value of specific thrust to be achieved, Either Ve>Va i.e. substantial acceleration through the engine needs to be accomplished, • or pe>pa i.e a substantial pressure (static) residual (at exit face) inside the engine are required to be achieved.
Shiva U
49
Lect-3
The thrust relation shown in the last slide is of general nature and is valid for cases where a residual exit static pressure exists in the exhaust flow. If it is assumed that the expansion in the nozzle is completed to Pa , and hence the 2nd term, pressure thrust, can be neglected. Thus net thrust is:
Then Fn = 𝒎(Ve - Va) For a net thrust Fn, the thrust power may be written as:
THP = Fn.Va
Shiva U
50
The basic thrust equation indicates that as forward speed Va increases it is necessary to increase either the mass flow, or exit velocity Ve , or both, in order to hold the thrust, F, constant. The near-constant thrust characteristics at any altitude a desirable and attractive feature of jet engines (flat rated engines) A near-constant Fn results in almost direct increase in thrust horsepower with forward speed. This characteristic of turbojet engines exists well up into the high subsonic speed range, and with a properly designed inlet diffuser, extends into the supersonic range.
51 Shiva U
Lect-3
• In the supersonic region afterburner equipped engines enable large increase in thrust with Mach number. • This is possible from M=0 to M>3.0. At high flight speeds (Mach 1 and above), an appreciable proportion of the air compression is accomplished by inlet diffuser ram effect. • In fact, at a Mach number well above 3.0, due to enormous ram effect it is economical to dispense with the compressor, and hence, also the turbine. • This has given rise to ramjet engines. 52 Shiva U
Overall Efficiency Overall Efficiency Thermal Efficiency
Propulsive Efficiency
Shiva U
53
Lect-3
The propulsive efficiency η can be defined as the ratio of the p
useful propulsive energy or thrust power (F.Va) to the sum of that energy and the unused kinetic energy of the jet. This is the kinetic energy relative to the earth, and may be written as:
m.( V e Va )2 / 2 It then stands to reason that this unused exit kinetic energy is a waste energy and, once it goes out of the engine body it is not of any use for thrust production. Although inlet diffuser provides aerodynamic pre- compression of air, it also produces ram drag.
54 Shiva U
Shiva U
55
Shiva U
56
Lect-3
The propulsive efficiency ηp can be written as
.Va.(Ve Va) m ηp 2 (V V ) e a m.Va.(Ve Va) 2
2 Ve 1 Va
ηp is also known as Froude efficiency. From the above equations it is evident that : • Fn is maximum when Va = 0, (Take off) but ηp =0 • ηp is maximum when Ve/Va = 1, when thrust is zero.
57 Shiva U
Lect-3
•The propulsion efficiency is a measure of how well the propulsive
device is being used for propelling the aircraft. •It is different from the efficiency of energy conversion. • The efficiency of energy conversion is given by
η
energy
Ve2 Va2 m. 2 mf .Qfuel
Where,
and Qfuel m are the fuel mass flow and its heating value respectively
The denominator refers to the energy released by burning of fuel 58 Shiva U
Lect-3
The overall engine efficiency is given by
m .V .( V V ) a e a η η. η O p e f .Q fuel m
59 Shiva U
Lect-3
•At supersonic aircraft speeds the ram drag is also high.
•Moreover, at supersonic flight speeds it is difficult to design an air intake to efficiently handle the air
flows required by the engine,
which may have an afterburner for thrust enhancement.
•This requires matching the intake characteristics to the engine over a wide operating range of flow conditions as well as altitudes. •At hypersonic flight speeds this prompts us to look
at ramjet for
thrust generation. Since there are no rotating components e.g. compressor /turbine, matching the inlet to the engine is simplified. 60 Shiva U
Lect-3
•Fuel consumption for turbojet and other jet engines is normally presented in terms of thrust specific fuel consumption (TSFC). •The thrust specific fuel consumption varies with engine rpm, Mach number and altitude generally a minimum at the tropopause at 80-90 % rated rpm. •For jet engines with reheat or afterburning, the fuel consumption would be quite high, and SFC would show up as high value. In such operation sheer thrust requirement outweighs the high SFC. •Turbo-props have lower SFC. This fact has prompted development of Prop-fans .
61 Shiva U
Specific fuel consumption Available thrust is usually quoted in kN or lbs. Fuel efficiency is usually quoted as a specific fuel consumption
SFC can only be directly compared at a specific flight condition, usually a nominal cruise condition. Expressed in kg/N-hr or mg/N-sec Actual computation of fuel mass flow and net thrust would vary from one kind of jet engine to another Shiva U
62
Specific impulse • Specific impulse is defined as the thrust (N) divided by the fuel weight flow rate (N/s). The resulting measure is usually quoted in seconds and defines the weight fraction that is necessary to give a particular delta V for a rocket or range for an aircraft with a given lift to drag ratio. • For a jet engine the specific impulse can be determined from the specific fuel consumption. The product of SFC and Specific impulse is one. The conversion factor between SFC (mg/Ns) and Specific impulse (s) is 102,000mg/N (1E6mg kg-1 /9.81N kg-1). A high bypass turbofan engines have cruise SFC around 15mg/Ns, and takeoff SFC of 8mg/Ns. Shiva U
63
Engine
Specific impulse of propulsion technologies SFC Specific impulse (mg/Ns) (s)
Turbofan (Takeoff, M0.1) Turbofan (Cruise, M0.9) Turbofan (with Afterburning, M1.5) Solid rocket (including oxidizer) LH2LO2 rocket (including oxidizer)
Energy Density (MJ/kg)
7.5
13,600
43
15
6,800
43
30
3,400
43
408
250
3.0
227
450
9.7
Shiva U
64
Shiva U
65
Engine cycle analysis and performance analysis for turbojet, turbojet with afterburner, turbofan engine, turboprop engine.
Shiva U
66
Lect-7
Gas Turbine Cycles • Gas turbine engines operate on Brayton cycles. • Ideal Brayton cycle is a closed cycle, whereas gas turbines operate in the open cycle mode. • Ideal cycle assumes that there are no irreversibilities in the processes, air behaves like an ideal gas with constant specific heats, and that there are no frictional losses.
Shiva U
6 7
67
Brayton cycle • The Brayton cycle was proposed by George Brayton in 1870 for use in reciprocating engines. • Modern day gas turbines operate on Brayton cycle and work with rotating machinery. • Gas turbines operate in open-cycle mode, but can be modelled as closed cycle using air- standard assumptions. • Combustion and exhaust replaced by constant pressure heat addition and rejection.
Shiva U
68
Ideal Brayton cycle The Brayton cycle consists of internally reversible processes:
four
– 1-2 Isentropic compression (in a compressor) – 2-3 Constant-pressure heat addition – 3-4 Isentropic expansion (in a turbine) – 4-1 Constant-pressure heat rejection
69 Shiva U
Ideal Brayton cycle P
qin 3
2
Isentropic
Isobaric
T
3
qin 4
2
qout 1
qout
1
4 v
s
Brayton cycle on P-v and T-s diagrams 70 Shiva U
Actual/Real Brayton cycle • Actual Brayton cycles differ from the ideal cycles in all the four processes. • The compression process and expansion processes are non-isentropic. • Pressure drop during heat addition and heat rejection. • The presence of irreversibilities causes the above deviations.
71 Shiva U
Actual/Real Brayton cycle Pressure drop T
3
qin
2s
2a
4a
4s
qout 1 s
Actual Brayton cycle T-s diagram 72 Shiva U
Lect-7
Ideal cycle for jet engines • All air-breathing jet engines operate on the Brayton cycle (open cycle mode). • The most basic form of a jet engine is a turbojet engine. • Some of the parameters of a jet engine cycle are usually design parameters and hence often fixed a priori: eg. compressor pressure ratio, turbine inlet temperature etc. • Cycle analysis involves determining the performance parameters of the cycle with the known design parameters.
Shiva U
7 3
73
Ideal cycle for jet engines Combustion chamber/burner Diffuser
a
1
Compressor
2
3
Turbine
Nozzle
4 5
6 7
Afterburner
Schematic of a turbojet engine and station numbering scheme Shiva U
Prof. Bhaskar Roy, Prof. A M Pradeep, Department of Aerospace, IIT Bombay
5
74
Lect-7
Ideal cycle for jet engines The different processes in a turbojet cycle are the following: • a-1: Air from far upstream is brought to the air intake (diffuser) with some acceleration/deceleration • 1-2: Air is decelerated as is passes through the diffuser • 2-3: Air is compressed in a compressor (axial or centrifugal) • 3-4 The air is heated using a combustion chamber/burner
75 Shiva U
75
Lect-7
Ideal cycle for jet engines
• 4-5: The air is expanded in a turbine to obtain power to drive the compressor • 5-6: The air may or may not be further heated in an afterburner by adding further fuel • 6-7: The air is accelerated and exhausted through the nozzle.
76 Shiva U
76
Lect-7
Ideal cycle for jet engines 4
T
5
3
7
2 a
s
Ideal turbojet cycle (without afterburning) on a T-s diagram 77 Shiva U
77
Lect-7
Ideal cycle for jet engines • Afterburning: used when the aircraft needs a substantial increment in thrust. For eg. to accelerate to and cruise at supersonic speeds. • Since the air-fuel ratio in gas turbine engines are much greater than the stoichiometric values, there is sufficient amount of air available for combustion at the turbine exit. • There are no rotating components like a turbine in the afterburner, the temperatures can be taken to much higher values than that at turbine entry.
78 Shiva U
78
Lect-7
Ideal cycle for jet engines 6a
T 4
5, 6
7a
3 2 a
s
Ideal turbojet cycle with afterburning on a T-s diagram 79 Shiva U
79
Lect-7
Turbofan engine • Propulsion efficiency is a function of the exhaust velocity to flight speed ratio. • This can be increased by reducing the effective exhaust velocity. • In a turbofan engine, a fan of a larger diameter than the compressor is used to generate a mass flow higher than the core mass flow. This ratio mcold / m hot is called the bypass ratio. • Turbofan engines have a higher propulsion efficiency as compared with turbojet engines operating in the same speed range.
80 Shiva U
80
Ideal turbofan engine Diffuser
2’
3’
7’
Combustion chamber/burner Turbine Compressor Primary nozzle
Fan
a
Secondary nozzle
Lect-7
1
2
3
4
5
6
7
Schematic of an unmixed turbofan engine and station numbering scheme 81 Shiva U
81
Lect-7
Ideal turbofan engine 2’
Diffuser
3’
Nozzle
Combustion chamber/burner Turbine Compressor
Fan
a
7’
1
2
3
4
5
6
7
Schematic of a mixed turbofan engine and station numbering scheme 82 Shiva U
Lect-7
Ideal turbofan engine • The different processes in an unmixed turbofan cycle are the following: • a-1: Air from far upstream is brought to the air intake (diffuser) with some acceleration/deceleration • 1-2’: Air is decelerated as is passes through the diffuser • 2’-3’: Air is compressed in a fan • 2-3: Air is compressed in a compressor (axial or centrifugal)
83 Shiva U
Lect-7
Ideal turbofan engine • 3-4: The air is heated using a combustion chamber/burner • 4-5: The air is expanded in a turbine to obtain power to drive the compressor • 5-6: The air may or may not be further heated in an afterburner by adding further fuel • 6-7: The air is accelerated and exhausted through the primary nozzle. • 3’-7’: The air in the bypass duct is accelerated and expanded through the secondary nozzle.
84 Shiva U
Lect-7
Ideal turboprop and turboshaft engines • Turboprop engines generate a substantial shaft power in addition to nozzle thrust. • Turboshaft engines, generate only shaft power. These engines are used in helicopters. The shaft power is used to drive the main rotor blade. • In a turboprop engine, the advantages and limitations are those of the propeller. • Both turboprops and turboshafts have applications at relatively lower speeds.
85 Shiva U
Lect-7
Ideal turboprop and turboshaft engines Propeller
Compressor Combustion chamber/burner
Nozzle
Propeller pitch control Gear box
Compressorturbine
Power turbine
Schematic of typical turboprop engine 86 Shiva U
Lect-7
Ideal turboprop and turboshaft engines • Turboprops and turboshafts usually have a free-turbine or power turbine to drive the propeller or the main rotor blade (turboshafts). • Stress limitations require that the large diameter propeller rotate at a much lower rate and hence a speed reducer is required. • Turboprops may also have a thrust component due to the jet exhaust in addition to the propeller thrust. • In turboshafts, however, there is no thrust component due to the nozzle.
87 Shiva U
Lect-7
Ideal turboprop and turboshaft engines • In turboprops, thrust consists of two components, the propeller thrust and the nozzle thrust. • The total thrust of a propeller is equal to the sum of the nozzle thrust and the propeller thrust.
88 Shiva U
Lect-7
Ideal turboprop and turboshaft engines 5
6 7
P05
h
05
Δh
Δh
P06 06
Pa
7 s
Enthalpy-entropy diagram for power turbineexhaust nozzle analysis 89 Shiva U
Real cycle for turbojet engines Combustion chamber/burner Diffuser
a
1
Compressor
2
3
Turbine
Nozzle
4 5
6 7
Afterburner
Schematic of a turbojet engine and station numbering scheme Shiva U
3
90
Real cycle for turbojet engines • The different processes in a turbojet cycle are the following: • a-1: Air from far upstream is brought to the air intake (diffuser) with some acceleration/deceleration • 1-2: Air is decelerated as is passes through the diffuser • 2-3: Air is compressed in a compressor (axial or centrifugal) • 3-4 The air is heated using a combustion chamber/burner
Shiva U
91
Real cycle for turbojet engines • 4-5: The air is expanded in a turbine to obtain power to drive the compressor • 5-6: The air may or may not be further heated in an afterburner by adding further fuel • 6-7: The air is accelerated and exhausted through the nozzle.
Shiva U
92
Real cycle for turbojet engines T
4 5 3 7 2 a
s
Real turbojet cycle (without afterburning) on a T-s diagram
Shiva U
93
Real cycle for turbojet engines 6a 4
T
7 5, 6
3
2 a
s
Real turbojet cycle (with afterburning) on a T-s diagram Shiva U
94
Real cycle for turbojet engines • Afterburning: used when the aircraft needs a substantial increment in thrust. For eg. to accelerate to and cruise at supersonic speeds. • Since the air-fuel ratio in gas turbine engines are much greater than the stoichiometric values, there is sufficient amount of air available for combustion at the turbine exit. • There are no rotating components like a turbine in the afterburner, the temperatures can be taken to much higher values than that at turbine entry.
Shiva U
95
Real cycle for turbojet engines • For calculating the fuel flow rate required to achieve a temperature of T6a, we carry out an energy balance similar to that of the combustor. • The total fuel flow rate, f, is equal to the sum of the fuel flow rates in the main combustor and the afterburner. f = f1 + f2 • Where f1 is the fuel flow rate in the main combustor and f2, the fuel flow rate in the afterburner.
Shiva U
96
Real cycle for turbofan engines • A turbofan engine can have different configurations: Twinspool, three-spool, and geared turbofan. These may be either unmixed or mixed. • Cycle analysis of a turbofan can hence be slightly different depending upon the configuration of the engine. • We shall now carry out an real cycle analysis of an unmixed twin-spool turbofan engine.
Shiva U
97
Real cycle for turbofan engines Diffuser
2’
3’
7’
Combustion chamber/burner Turbine Compressor Primary nozzle
Fan
a
Secondary nozzle
1
2
3
4
5
6
7
Schematic of an unmixed turbofan engine and station numbering scheme
Shiva U
98
Real cycle for turbofan engines • The total thrust developed by the turbofan with two separate unmixed streams will consist of thrust due to primary nozzle and that due to the secondary nozzle. • Fn= Fn(primary nozzle) + Fn (secondary nozzle)
Fn mH (1 f )Vex V m H (Vexf V ) assuming (Pe Pa ) Ae to be negligible.
Shiva U
99
Real cycle for turbofan engines • The cycle analysis procedure will need to be slightly modified depending upon the turbofan engine configuration. • The differences in the various configuration arise because of the number of spools and turbine-compressor/fan arrangements as well as mixed and unmixed exhausts. • If the turbofan is of a mixed configuration, then, we will have to calculate the temperature at the nozzle entry from enthalpy balance of the two streams.
Shiva U
10 0
Real cycle for turboprop and turboshaft engines Propeller
Compressor Combustion chamber/burner
Nozzle
Propeller pitch control Gear box
Compressorturbine
Power turbine
Schematic of typical turboprop engine Shiva U
10 1
Real cycle for turboprop and turboshaft engines • Turboprops and turboshafts usually have a free-turbine or power turbine to drive the propeller or the main rotor blade (turboshafts). • Stress limitations require that the large diameter propeller rotate at a much lower rate and hence a speed reducer is required. • Turboprops may also have a thrust component due to the jet exhaust in addition to the propeller thrust. • In turboshafts, however, there is no thrust component due to the nozzle.
Shiva U
10 2
Real cycle for turboprop and turboshaft engines 5
6 7
P05
h
05
Δh
Δh
P06 06
Pa
7 s
Enthalpy-entropy diagram for power turbineexhaust nozzle analysis Shiva U
10 3
Real cycle for turboprop and turboshaft engines
Lect-10 • Δh is the enthalpy drop in an ideal isentropic power turbine and exhaust nozzle. •
is the fraction of Δh that would be used by an isentropic turbine.
• The propeller thrust power,Fn, prV, is
Fn, prV pr g PT h m
or, Fn, pr
prgPT h m
V pr propeller efficiency,g gear box efficiency,
PT power turbine efficiency Shiva U
10 4
Lect-10
Real cycle for turboprop and turboshaft engines • The exhaust nozzle thrust, Fn , Fn m(Vex V ), where, Vex
2(1 )nh
• Thus, the total thrust is given by, F Fn, pr Fn
pr g PT h m V
Shiva U
m( 2(1 )n h V )
10 5
Engine overall efficiency impact on aircraft range and endurance
Shiva U
106
Lect-3
Typical thrust generation capability of small aircraft engines of similar power
107 Shiva U
Lect-3
Typical propulsive efficiency of small aircraft engines of similar power
108 Shiva U
Lect-3
•Because Ve is considerably greater than Va, mainly at low flying speeds, the efficiency is much lower than that attainable with a propeller.
•Since propeller efficiency drops off rapidly at higher Mach numbers (>0.7) there is a speed where jet propulsive efficiency exceeds that of a propeller. •Because the overall efficiency of a turbojet is lower, the Mach number at which the overall efficiency of a turbojet equals the overall efficiency of a prop-jet engine is more than the Mach number at which their propulsive efficiencies are equal .
109 Shiva U
Lect-3
Jet Engine Thrust Characteristics
110 Shiva U
Lect-3
Jet Engine SFC Characteristics
111 Shiva U
Shiva U
112
Shiva U
113
Shiva U
114
