SENECA COLLEGE
CFM56-3 Turbofan Engine
Description
David Aubuchon, John Campbell
8/3/2016
FLP-700
PROFESSOR LABIB
Table of Contents
Section
Page
1.0 Introduction........................................................................................................... 3
1.1 Overview and Main Parameters....................................................................... 4
2.0 Compressor Section.............................................................................................. 6
2.1 Low Pressure Compressor............................................................................... 8
2.2 High-pressure Compressor.............................................................................. 14
3.0 Combustion Section.............................................................................................. 19
4.0 Turbine Section.....................................................................................................
4.1 High Pressure Turbine.....................................................................................
4.2 Low Pressure Turbine......................................................................................
4.3 Low Pressure Turbine Frame...........................................................................
4.4 Turbine Cooling and Clearance Control System.............................................
22
22
26
27
28
5.0 Engine Shaft and Bearing Assembly................................................................... 32
6.0 Exhaust....................................................................................................... 34
6.1 Thrust Reverser................................................................................................ 34
7.0 Starter.................................................................................................................... 37
8.0 Fue1 System...........................................................................................................
8.1 Fuel System Operation.....................................................................................
8.2 Fuel Distribution Description..........................................................................
8.2.1 Fuel Pump..............................................................................................
8.2.2 Fuel Filter...............................................................................................
8.2.3 Wash Filter.............................................................................................
8.2.4 Servo Fuel Heater...................................................................................
8.2.5 Main Oil/Fuel Heat Exchanger .......................................................
8.2.6 Main Engine Control (MEC) ................................................................
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38
40
40
40
40
41
41
41
9.0 Lubrication System...............................................................................................
9.1 Supply Operation.............................................................................................
9.2 Scavenge Operation................................................................................
9.3 Venting Operation..................................................................................
44
44
45
46
10.0 Ice Protection............................................................................................
48
References...................................................................................................................
50
2
1
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1.0 Introduction
The CFM International CFM56-3 is a high-bypass, dual rotor, axial flow turbofan aircraft
engine. The CFM56-3 is a derivative of the CFM56 engine family that produces a thrust range of
18,500 to 34,000 pounds-force. The CFM56-3 produces 20,000 to 23,500 pounds-force depending
if it is mounted on the 737-300 or 737-400 respectively. The initial design of this engine first ran in
1974 and first flew in 1977 on the McDonnell Douglas YC-15 as a candidate in the United States
Air Force’s Advanced Medium STOL Transport competition. (Aviation Week & Space Technology,
1977) The CFM56 won the competition and became the new engine for the KC-135 tanker fleet as
well as the primary engine for the Boeing 737 and DC-8. Since then, the CFM56 has been upgraded
extensively and is now the most used high-bypass turbo-fan engine with over 800 million flight
hours. Currently, the CFM56-3 is used by the Airbus A320, A340, the Boeing 737, and KC-135R
Stratotanker.
Figure 1: McDonnell Douglas DC-3 with CFM56 engines.
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1.1 Overview and Main Parameters
The CFM56-3 is a dual-shaft engine with a thrust rating between 18,500 and 23,500 lbs
and compression ratio of 27.5-30:1. This particular variant has a bypass ratio of 6:1 with a 60inch fan diameter, and a dry weight of 4,300 lbs. Its dual-shaft design consists of a fan and
booster (low pressure compressor), high pressure compressor, annular combustion chamber, and
a high and low pressure turbine section. The two shafts respectively connect the low and high
pressure sections using a five bearing system (3 roller, and 2 ball bearings). The high pressure
turbine (HPT) spins at 15,000 revolutions per minute (rpm) at 100% N2 and the low pressure
turbine maintains 5,200 rpm at 100% N1. There are three gearboxes in the CFM56; they are the
inlet, transfer, and accessory gearbox. Due to its large bypass ratio, and fan diameter, the engine
was considered to be one of the most efficient in the market. The seven main components of the
CFM56-3 are listed below and illustrated in Figure 2.
1. Four stage booster: Single-stage fan and three-stage axial compressor
2. The high pressure compressor: nine-stage axial
3. The combustion chamber: annular axial flow type
4. The high pressure turbine: single-stage
5. The low pressure turbine: four-stage
6. The Exhaust: long cold type (separate and mixed properties), reversible
7. Accessory drive: 3 gearbox
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Figure 2: CFM56 schematic overview of engine components.
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2.0 Compressor Section
The CFM56’s compressor section adds energy to the airflow and converts it into high total
pressure in preparation for the combustion chamber. The compressor section is also responsible for
supplying bleed air for multiple other purposes such as cabin pressurization, and turbine blade
cooling. A compressor’s general function is simple in theory but complex in design. Mechanical
energy from the turbine section is transferred by a shaft to the rotor where it adds kinetic energy to
the airflow. The increase of kinetic energy is then transferred into potential energy (static pressure)
by the stators. The stator blades are designed as small diffusers and they convert the kinetic energy
into potential energy by decelerating the air.
The compressor section consists of a single stage fan, a three stage low-pressure
compressor, and a nine stage high-pressure compressor. The fan and three stage low-pressure
compressor are driven by the low pressure turbine (LPT) and the high pressure compressor is
driven by the high pressure turbine (HPT). A diagram of the CFM56’s compressor section is
shown in Figure 3.
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Single Stage Fan
LP Compressor
HP Compressor
Figure 3: Outline of compressor sections in the CFM56-5A
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2.1 Low Pressure Compressor
Secondary Airflow
Primary Airflow
Figure 4: Fan disk and low pressure compressor/booster section.
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The axial low pressure compressor (LPC) is also defined as the “Booster Assembly” for the
CFM-56. It consists of a single-stage fan rotor and a 3-stage axial booster which is mounted by
cantilever at the rear of the fan disk. The cantilever mount system is outlined in Figure 4. The
LPC section consists of the following components:
1. Spinner front and rear cone
2. Single stage outlet guide vane in the secondary airflow
3. Fan disk and fan blades
4. Four stage booster stage stator assembly in primary airflow
a. One booster inlet-guide-vane
b. Three booster rotor stages
c. Three booster stator stages
5. Twelve variable bleed valves
6. Booster stator vanes
The booster section is driven by the low pressure turbine (LPT) and provides two separate
airstreams denoted as primary and secondary (Figure 4). The primary air flows through the fan and
booster assembly where it is compressed for the high pressure compressor (HPC). The secondary
airflow generated by the 23.5in fan blades is ducted around the outside of the engine. The fan rotor
consists of 36 longitudinally dovetail fixated titanium alloy blades and provides approximately
80% of the total thrust generated by the engine. At static take-off power, the CFM56-3 has a bypass ratio of 6:1. This means that there is six times more secondary airflow than primary. The
equation for by-pass ratio is shown below.
By-pass ratio () = (secondary mass airflow) / (primary mass airflow)
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The fan blades are titanium alloy presumably for its high tensile strength and weight saving
characteristics. Each blade has a “mid-span shroud” that interlocks it with the blades on either side
creating a system of interlocked fan-blades. This is used as a “Band-Aid” solution to counter
aerolastic phenomena known as blade-tip flutter. Flutter is a positive feedback cyclic motion
caused by flexible cantilevered structures. The shroud increases the fan blade’s overall stiffness
reducing the ability of the tip to flex and generate flutter. The individual shroud is shown in Figure
5 and the shroud system is illustrated in Figure 6.
Figure 5: Fan blade schematic diagram showing the shroud and dovetail
blade root.
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The LPC uses a drum-spool type compressor assembly for sturdiness, and weight saving
advantages. The drum-spool assembly is a hollow, cylindrical piece of titanium alloy that is
connected to the fan disk. This hollow drum is denoted as the “booster spool” and is illustrated in
Figure 6. The drum type was used because it is ideal for a small number of stages and is also more
solid and rigid than disc-type. The compressor blades are fixated using circumferential dove-tail
fixation which is illustrated in Figure 7.
Figure 6: Fan and booster assembly.
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The low pressure booster blades are attached to the booster spool using three ring
assembly’s which corresponds to the three stages of low pressure stages. Each of these rings has a
varying number of blades because of the changing spool diameter and blade width. Stage 2 has 70
blades, stage 3 has 74 blades, stage 4 has 70 blades, and stage 5 has 55 blades. The blades are
approximately 3.5 – 4.2in long and are installed using circumferential dovetail fixation using
locking lugs to hold them in place. A depiction of this assembly is shown in Figure 7. Dovetail
fixation is used because it provides the best strength-to-weight characteristics by optimizing the
area of stress contact between the support and blade base. It is understood that  = F/A where ()
is stress, F is force, and A is effective area. By increasing the area (A), the overall stress value can
be reduced.
Figure 7: LPC circumferential dovetail fixation using lug locking system.
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The LPC’s stator assembly is composed of 5 stages of vanes, and inner and outer shrouds.
Their main objective is to convert absolute air velocity (C) into pressure by working as diffusers.
They are also responsible for redirecting the air to an ideal angle for the next rotor stage in order to
avoid blade stalls. The first stator can also be considered an inlet guide vane (IGV) for the LPC
because it adds a circumferential component (CU) to the absolute airflow (C) to avoid supersonic
blade-tip stalls for the first stage of compression. Stage 1 has 106 vanes, stages 2-3 have 124
vanes, stage 4 has 116 vanes and stage 5 has 90 vanes. In comparison to the HPC, this group of
stators does not include variably actuated stators because they are less susceptible to blade stalling.
A diagram of the LPC’s stator assembly is in Figure 8.
Figure 8: LPC stator vane assembly.
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2.2 High-pressure Compressor
Variable Stator Vane (VSV) assembly
Figure 9: High-pressure compressor cut-away.
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The high-pressure compressor (HPC) assembly (Figure 9) is designed to compress the air
further for combustion preparation. The overall structure of the HP compressor is constant internal
diameter, 9-stage, high speed, drum-disc, axial design. It is understood that compression ratio ()
and energy transferred to the air decreases with every stage because compressed air is more
difficult to compress, creating diminishing compression across the length of the system. The
benefits of having a constant internal diameter system include the idea that blades are longer than a
constant external diameter would be, and thus less losses are induced by the boundary layer
thickness. It also has a smaller size element allowing more volume for the secondary airflow and
bleed air distribution. The smaller size also allows more space for the variable stator vane
assembly which are configured across the first 4 stages of the HPC shown in Figure 9. The overall
compression ratio () of the CFM56-3 is 27.5:1. The components of the HPC consists of:
1. The compressor rotor
2. The compressor front stator
3. The compressor rear stator
4. Variable stator vane assembly
5. 4th, 5th, and 9th stage bleed air ducts
The HPC’s stator assembly is split into two sections: the variable stator vanes (VSV) from
stage 1-4, and the fixed stator vanes from stage 5-9. This setup is illustrated in Figure 10. The
variable stator vanes are made of steel alloy and are designed to optimize the relative airflow (W)
across the rotor blades when the engine is operating at varying RPMs. This leads to optimum
compression ratios outside of the engine’s design. They are actuated hydro-mechanically through
use of a bellcrank assembly shown in Figure 10.
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Figure 10: HPC variable and fixed stator vane assembly.
Figure 11: 9-stage CFM56-5b HPC drum-disc assembly.
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The HPC uses a titanium alloy drum-disc assembly (Figure 11) to facilitate sturdiness
across the 9-stage system. A drum-disk configuration is generally used for compressors with 5
stages or more. They offer greater support by combining the disk and drum so that the shaft’s
critical rpm can be increased as the shaft length increases. Nonetheless, using this assembly
increases the engine’s overall weight. The drum is attached to a disk in stage-3 which transfers the
energy from the high-pressure shaft. Figure 11 illustrates the drum-disk setup whereby stages 1-2
and 4-9 are drum assembly and stage 3 is disk assembly. The spool itself is driven by the titanium
alloy rotor shaft shown in Figure 10. To optimize airflow sealing, labyrinth seals (Figure 10) are
machined into the spool, which face soft-material structures on the stator assembly. The soft
material is used as a defense against blade tip wear.
The rotors in stages 1-3 use the longitudinal dovetail fixation and stages 4-9 utilize the
circumferential dovetail fixation. The spools for both of these setups are depicted in Figure 10
The longitudinal dovetail fixation is used in stages 1-3 mainly for ease of maintenance because the
longer blades are more susceptible to breaking. Stages 4-9 use circumferential dovetail fixation
because it eliminates the need for installing extra components to prevent axial movement of the
blades.
A portion of the air flowing through stages 4-9 is extracted for engine cooling and aircraft
usage (Figure 12). The air is pulled from these stages because it is compressed and also slightly
heated. The air, also known as “bleed air”, is redirected to multiple engine components for cooling
purposes and the aircraft cabin for pressurization and climate control. The 4th and 9th stage bleed
air is used for high and low pressure turbine cooling and the 5th stage is used for the aircraft cabin.
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Labyrinth Seals
Rotor Shaft
Figure 10: Spool assembly for the HPC.
Figure 12: Bleed air ports in the HPC.
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3.0 Combustion Section
Figure 13, 14: Combustion chamber cut-away.
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The combustion chamber uses an annular assembly and is installed between the HPC and
HPT nozzle shown in Figure 13, 14. Its main purpose is to mix and evaporate fuel before it
combusts in the pressurized air supplied by the compressor section. The annular configuration
(Figure 13, 14) is used by a majority of modern jet engines because of its lighter design, and
efficient homogenous airflow through the chamber. The homogenous airflow through the
combustion chamber helps to avoid cyclic loading on the high-pressure turbine which increases
their longevity – something that used to be problem while using can-type combustion chambers.
The combustion chamber’s main components shown in Figure 15, 16 consist of:
1. Swirl fuel nozzles (x20)
2. Swirl vanes
3. Outer cowl
4. Inner cowl
5. Outlet guide vane
6. Igniter plugs (x2)
The fuel vapour is discharged in the combustion chamber by 20 swirl vane fuel nozzles.
The swirl vane nozzles in this particular chamber were designed to create faster vaporization and
mixing of air and fuel by “swirling” the air within the combustion chamber. This helps to facilitate
faster fuel mixing causing a more efficient burn and increases the engine’s overall fuel economy.
The surface of the combustion chamber is cooled by a layer of air from the HPC bleed-air
and flows around the outside of the chamber which is depicted by the blue arrow in Figure 15.
There are two igniters in the annular combustion casing and are located at the 4 and 8 o’clock
positions looking in from the front of the engine.
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Figure 15: More detailed illustration of annular combustion chamber.
Figure 16: HPC airflow into the combustion chamber.
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4.0 Turbine Section
The primary task of the turbine section is to extract energy from the hot section gas flow
after the combustion chamber in order to drive the compressor rotors through the spool shaft. Some
of the energy extracted by the turbines is used to power engine accessories through the accessory
drive and gearbox located in the compressor section in the front of the engine. This engine has two
turbine rotor systems corresponding to the respective spools; the high pressure turbine (HPT)
drives the HPC and the low pressure turbine (LPT) drives the LPC. Turbine stages consist of a
nozzle (stator), which is fixed, followed by a turbine rotor which moves with its respective spool.
The turbine nozzle directs the gasses, generally slightly into the centrifugal direction from axial (in
the direction of the rotor) in order to optimize energy transfer from the gas to the turbine rotor
blades. The kinetic energy of the gasses is then transferred to the turbine rotor blades which
delivers it to the front of the engine through shaft torque. All the rotors in the turbine are of the
reaction type and have a convergent channel between the blades.
4.1 High Pressure Turbine
The HPT consists of a single axial stage which is placed right at the exit of the combustion
case as shown in figure 17; the nozzles are mounted within the end of the combustion case. The
nozzles are internally cooled by secondary combustion air entering the vane compartments through
inserts in the inner and outer ends of vanes and exiting through the vane leading and trailing edges.
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shroud
Figure 17: Low Pressure turbine location and blades
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Figure 18: High Pressure Turbine Nozzle
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The HPT rotor (figure 19), is a single stage, air cooled turbine. The HPT rotor drives the 9
HPC rotors and is directly connected to them by a bolted flange to form what is essentially a single
core drum-disk type rotor. Individual blades are fixed to the rotor disc through a double surface
longitudinal fir-tree fixation. The blades can be replaced independently if required. The blades are
internally cooled by a mixture of secondary combustion air and compressor discharge air that
enters through the blade dovetail and exits through holes in the front sides and trailing edges as
shown in figure 17. The rotor casing has a shroud with a thermal response matched to the rotor to
provide tip clearance control and structural stability. An air cavity between the shroud and the
combustion case directs mixed 5th and/or 9th stage high pressure compressor bleed air onto the
Figure 19: High Pressure Turbine Rotor
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shroud support and the outer surface (backside) of the shrouds. This cooling air maintains closer
clearances between the shrouds and the rotor blades. More detailed Information on the HPT
clearance control system is explained at the end of this section.
4.2 Low Pressure Turbine
The LPT consists of 4 stages placed axially right after the HPT (figure 20). All blades have
longitudinal dove-tail fixations. The first nozzle is part of the HPT shroud. The 3 other nozzles are
fixed by the LPT case. An air cavity between the 1st stage nozzle support and the combustion case
(figure 17, 22) directs 5th-stage high pressure compressor (HPC) bleed air through the nozzle vanes
for cooling. After passing through the vanes, the air pressurizes and cools the cavity between the
aft side of the HPT rotor and the forward side of the LPT rotor. The LPT case (figure 20) is cooled
by fan air, to help minimize the gap between rotor blades and stator. The 4 stage axial type LPT
rotates within the LPT casing; the rotors are supported by the LPT drum-disc which is connected to
the LPC through a shaft within the HPT drum. The four stages drive the fan and 3 consecutive LPC
rotors.
LPT casing
Stators 2-4
Figure 20a: LPT casing, frame, stators (2-4) and rotors 1-4, dove tail fixation
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4.3 Low Pressure Turbine Frame
The turbine frame (figure 20) is a major structural assembly at the rear of the engine.
The turbine frame consists of:
-
Inner hub which contains the bearing #4 and #5.
-
Outer casing: lt provides mounts for attaching the rear of the
engine to the airplane strut.
-
Struts: Twelve struts are arranged in a slanted position relative to
the hub to pro-vide the turbine frame with correct bending
stiffness. These struts are hollow and provide passage for the
following:
+ Aft sump oil scavenge tube (No.5 strut)
+ No.4 and 5 bearings oil supply tube (No.6 strut)
+ Aft sump overboard
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4.4 Turbine Cooling and Clearance Control System
Bleed air from different parts of the engine is used to cool the HPT disk, blades and nozzle
vanes, the LPT disk and nozzle vanes, and the turbine case. The HPT also has a clearance control
system (CCS) which uses cooling air from the HPC section to optimize HPT blade clearance and
exhaust gas temperature control, allowing its maximum performance.
The HPT nozzles are internally cooled by secondary combustion air which enter the vane
compartments through inserts in the inner and outer ends of vanes. The cooling air exists through
film cooling holes in the nozzles’ walls at the leading and trailing edges (figure 17).
The HPT rotor blades are internally cooled by a mixture of secondary combustion air and
9th stage HPT discharge air (CDA) (figure 21, 22) that enters through the blade roots and exits
through holes in the front sides and trailing edges (figure 17).
The LPT stage 1 nozzle is cooled by stage 5 HPC air. The air is channeled straight through
the nozzles into the core of the engine from which is cools the discs and roots of the LPT rotors 1-3
as well as actively seals the gaps between the stators and the rotors. The 4th rotor stage is also
cooled in the same way by stage 5 HPC air that comes from sump air vent (figure 21).
The LPT casing is cooled with fan discharge air (FDA) through channels which flow
around within the casing (figure 21). This is part of the clearance control system which keeps the
dimensions between the case and the rotors to a minimum such that performance is optimal.
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The CCS insures maximum steady state HPT performance and minimizes exhaust gas
temperature (EGT) transient overshoot during rapid change of engine speed. The HPT rotor casing
has a shroud with a thermal response matched to the rotor to provide tip clearance control and
structural stability. This is helped by an air cavity between the shroud and the combustion case
which directs mixed 5th and/or 9th stage HPC bleed air onto the shroud support and the outer
surface (backside) of the shrouds (figure 17, 22).
The System consists of:
- HPT Clearance Control Valve: This schedules which mixture is used depending on the
steady state operation; idle RPM uses 9th Stage, take-off RPM 9th Stage, climb RPM uses
9th and 5th stage and cruise RPM uses 5th Stage.
- HPT Clearance Control Timer: This controls the operation of the HPTCCS valve during
the first 182 seconds of engine operation if N2 RPM is 96% or greater during take-off
condition. The timer will be activated if: aircraft on GRD and N2 RPM >94%.
The normal schedule will be interrupted for180 sec and the following cooling air schedule
will be performed: 0-8 sec no air, 8-152 sec 5th stage, 152-182 sec 9th and 5th stage, 182
sec 9th stage (normal schedule).
- HPTCCS Timer Lockout Solenoid Valve:
The valve prevents timer operation after take-off through an air sensing relay (weight on
wheel switch).
- HPT Shroud Manifold Air Tubes: channels air from compressor to turbine.
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Figure 21: Engine Air Schematic
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Figure 22: Hot Section Air Flow
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5.0 Engine Shaft and Bearing Assembly
Figure 23: Diagram indicating the bearing, sump, and shaft locations with respect to high and
low pressure.
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The engine rotors are supported by 5 bearings where No. 1 is the most forward and No. 5 is
the most aft. The bearings are located within two dry sumps in the fan and turbine frames. There
are two types of bearings: ball, and roller. Ball bearings absorb the axial and radial loads
throughout the length of the engine shaft and roller bearings only absorb radial. There are two
roller bearings located in the fan and turbine casing, and two ball bearings in the fan casing. An
illustration of their locations is in Figure 24.




The No.1 and No.2 bearings hold the fan shaft
No.3 holds the front of the HP shaft
No.4 bearing holds the rear of the HP shaft
No.5 holds the rear of the low-pressure turbine shaft (LPT)
Figure 24: Shaft and bearing locations in the CFM-56-5b
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6.0 Exhaust
The unmixed exhaust system receives and discharges the fan and turbine air through
separate propelling nozzles to the atmosphere at a velocity and direction to produce the required
thrust.
Turbine Exhaust:
The turbine exhaust system provides an efficient exit for the turbine exhaust; it is fixed and
designed to create optimal thrust at the design point. The turbine exhaust consists of hot combusted
gases exiting the low pressure turbine at high velocity and provides 22% of the engine’s thrust. The
fan exhaust is channeled around the hot section exhaust such that noise is reduced due to slightly
reduced shear.
Fan Exhaust:
The fan exhaust is high velocity exhaust exiting the fan or first stage compressor. The fan exhaust
provides 78% of the total forward thrust. The direction of the fan exhaust can be reversed during
landing by the thrust reverser to produce additional braking power for the airplane on the ground.
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6.1 Thrust Reverser
The thrust reverser system reverses the direction of the fan exhaust during landing and
braking of the airplane. Stowed, the thrust reverser provides a smooth surface for the fan exhaust to
produce thrust. Deployed, the thrust reverser redirects the fan exhaust to produce reverse thrust.
The thrust levers in the flight compartment initiate forward or reverse thrust. The thrust levers are
connected electrically to the thrust reverser control module which controls the thrust reverser
hydraulic actuators. The thrust reverser control system synchronizes the deployment of the thrust
reverser halves. The thrust reverser position indicating system provides visual indication to the
flight compartment of thrust reverser position.
Figure 25: Exhaust / Thrust Reverser
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Figure 26: Thrust Reverser Schematic
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7.0 Starter
The engine start System provides the means of rotating the engine N2 compressor on the
ground or in flight to an rpm range at which engine start can be attained. The starter for this engine
is pneumatic; it uses high pressure air to turn a shaft which turns the N2 LPC spool. The engine
starter converts compressed air pressure into rotational mechanical energy sufficient to accelerate
the engine to starting speed. The starter is a lightweight, single stage, axial flow, turbine air motor.
The starter begins rotating the N2 compressor of the engine when compressed air is supplied to the
starter. Rotation of the N2 compressor establishes airflow through the engine inducing rotation of
the N1 compressor. Fuel and ignition are supplied by advancing the start lever to the IDLE detent
when the N2 compressor has reached 20% N2 rpm. Engine light up occurs shortly after the start
lever movement. A clutch disengages the starter if it exceeds 20% N2 to prevent burning the
starter. The air supply for starting the engine can be obtained from the auxiliary power unit (APU),
from a pneumatic ground cart via service connection or through cross bleed air from an operating
engine.
Figure 27: Pneumatic Starter
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8.0 Fuel System
The fuel distribution system is a hydro-mechanical system which provides and controls fuel
flow to the engine combustor. The system also powers and schedules:
- Variable stator vanes
- Variable bleed valves
- High pressure turbine clearance control system. To maintain maximum engine
performance within: stall margin, rotor speed, compressor discharge pressure, turbine
temperature limits.
The fuel also cools the engine lubricating oil.
8.1 Fuel System Operation
Fuel from the aircraft fuel tank enters the engine through a fuel pump inlet (-35 psi), it is
then pressurized through the low pressure centrifugal boost pump and flows through the fuel/oil
heat exchanger and the fuel filter. Fuel then flows through a high pressure fuel gear-pump (870 psi
at takeoff), then through a wash filter, and into the Main Engine Control (MEC) unit. Some of the
fuel is extracted before entering the MEC through the pump wash filter to be heated by the servo
fuel heater and then supplied to the MEC; this provides clean and ice-free fuel for servo operation.
The fuel pump has a higher fuel flow capacity than the fuel and control system requires, thus the
fuel flow is divided in the MEC into metered flow and bypass flow. The bypass fuel is ported back
to the inlet side of the fuel/oil heat exchanger, and metered fuel from the MEC flows through the
pressurizing valve, the fuel flow transmitter, the fuel manifold, and fuel nozzles into the
combustion chamber.
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Figure 28: Fuel System Schematic
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8.2 Fuel Distribution Description
8.2.1 Fuel Pump
The fuel pump is located on the accessory gearbox (AGB) aft face between the horizontal
drive shaft housing and the lubrication unit at the 8 o'clock position (figure 30). The fuel pump
pressurizes and circulates fuel within the fuel system. The fuel pump contains a LP centrifugal
boost stage that delivers a boost pressure to the HP stage to avoid pump cavitation followed by a
HP stage positive displacement type gear-pump. For a given number of revolutions, it delivers a
constant fuel flow regardless of the discharge pressure.
8.2.2 Fuel Filter
The fuel filter is found between the heat exchanger for fuel/oil and the high pressure stage
of the fuel pump. The fuel filter stops ice and unwanted particles in the fuel; this is so that they
cannot cause damage to the high pressure stage of the fuel pump and the MEC. If the filter
becomes clogged; a bypass valve releases the fuel to the high pressure stage of the fuel pump. The
fuel filter has an integrated differential pressure switch which senses and activates the fuel filter
bypass warning System, alerting the pilots of impending fuel filter bypass resulting from a clogged
fuel filter.
8.2.3 Wash Filter
The wash filter is more fine than the fuel filter and prevents contaminants, larger than 60 microns, from
entering the MEC servo controls. Should the filter clog, the bypass valve relieves the fuel to the MEC servo controls.
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8.2.4 Servo Fuel Heater
The servo fuel heater is attached to the aft side of the fuel/oil heat exchanger, located
between the transfer gearbox and MEC. The servo fuel heater raises the fuel temperature to prevent
ice from entering the control servos inside the MEC.
8.2.5 Main Oil/Fuel Heat Exchanger
The main oil/fuel heat exchanger is attached to the fuel pump. lt is located at the 9 o'clock
position on the fan case. The main oil/fuel heat exchanger is a tubular type consisting of the
housing, removable core and core access cover. The purpose of this device is to cool the oil while
heating the fuel which promotes the evaporation in the combustion chamber. The fuel portions of
the heat exchanger contain a bypass valve. This permits fuel to bypass the core in the event of a
blockage (>26psid).
8.2.6 Main Engine Control (MEC)
The MEC is basically a speed governor which senses engine speed (RPM) and adjusts the fuel
flow as necessary to maintain the desired speed set by the thrust lever. The main engine control is
hydro-mechanical and operates using fuel operated servo valves. The MEC performs the following
functions:
1. Controls engine speed by metering fuel to the engine fuel nozzles during all modes of
operation. Fuel not used by the MEC is returned to the fuel pump downstream of its low
pressure element. The MEC also uses fuel from the fuel pump as a hydraulic medium.
2. Automatically schedules fuel flow to maintain the thrust lever speed setting and establishes
the maximum safe fuel flow limit under any operating condition. As conditions vary the
limits vary according to predetermined acceleration and deceleration schedules. In order for
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the MEC to determine the schedules, certain parameters such as compressor discharge
pressure (CDP), 9th stage compressor bleed pressure (CBP), compressor inlet temperature
(CIT) and core engine speed (N2) must be sensed. The MEC senses these parameters and
also amplifies and computes them to establish acceleration and deceleration limits for fuel
flow. The computed limits are compared with actual fuel flow and are imposed when the
limits are approached.
Figure 29: Main Engine Control Map
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Figure 30: Accessory Drive Diagram with Fuel Pump and MEC
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9.0 Lubrication System
The engine oil system is a self-contained, center vented and recirculating type system. The
oil system to provides lubrication and cooling for the engine main bearings, radial driveshaft
bearings and gears and bearings in the transfer gearbox (TGB) and accessory gearbox (AGB).
The oil system consists of:
- oil storage system,
- oil distribution system and
- oil indicating system
9.1 Supply Operation
The oil tank provides storage of oil for continuous distribution by the supply system. There
are four positive displacement pumps on a single shaft driven by the AGB. Oil flows from the tank
to the supply pump in the lubrication unit on the AGB. The pressurized oil is then pumped through
the oil supply filter to the main bearings, radial driveshaft and gearboxes. The supply pump
incorporates a pressure relief valve that diverts the oil flow to a scavenge pump in the event of
abnormal operating conditions. The pressure relief valve opens when the pressure downstream of
the supply pump exceeds 305 psi. If the supply filter becomes clogged, the flow can divert through
the bypass valve. The bypass valve opens when the pressure drop reaches 17.4-20.3 psi across the
oil supply filter. A clogging indicator pops up before a filter bypass condition. When the pressure
drop across the supply filter reaches 11.6-14.5 psi, the magnets are forced apart and the indicator
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pops up, becoming visible in the glass inspection bowl. A bimetal spring prevents actuation at low
operating temperatures.
9.2 Scavenge Operation
After distribution, oil is returned to the lubrication unit from three sumps. The forward
sump services the No. 1, No. 2 and No. 3 main bearings. The aft sump services the No. 4 and No. 5
main bearings. The gearbox sump for the AGB also collects oil through an external tube from the
TGB. The lubrication unit contains a scavenge pump for each sump. The oil is drawn
through one of three magnetic chip detectors (MCD) in the lubrication unit and is pumped through
the scavenge oil filter to the main oil/fuel heat exchanger. When the scavenge oil filter becomes
clogged, the flow diverts through the bypass valve. The bypass valves start opens when the
pressure drop reaches 36.3-39.2 psi across the scavenge oil filter. The clogging indicators pop up
before a filter bypass condition. When the pressure drops across the scavenge reaches 28-34 psi,
the magnets are forced apart and the indicator pops up, becoming visible in the glass inspection
bowl. A bi-metal spring prevents actuation at low operating temperatures.
Before returning to the engine oil tank, oil passes through the servo fuel heater and enters the main
oil/fuel heat exchanger perpendicular to the fuel flow. A drain is provided at the forward and aft
bearing compartment for possible oil leaking past the stationary air/oil seal. The forward seal drain
exits through the 8 o'clock fan frame strut. The aft seal drain exits through the 6 o'clock turbine
frame strut.
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9.3 Venting Operation
The bearing sumps and gearboxes are interconnected to collect oil vapors before air/oil
separation and venting. The oil tank vent and the TGB/AGB sump are connected to the forward
sump. Vapors from the forward and aft sumps pass through rotating air/oil separators into the main
shaft center vent tube to be vented out the exhaust. The separated oil is returned to the sumps. The
oil is cooled in the main oil/fuel heat exchanger. Fuel enters the cylindrical core through the fuel
inlet in the housing. It flows the length of the core through half the core tubes. The fuel flows
around a baffle in the core access cover and returns through the remaining core tubes. The oil
circulates around the fuel tubes in the core, transferring heat to the fuel by convection and
conduction. The cooled oil exists back through the servo fuel heater.
Figure 31: Oil System Component Map
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Figure 32: Oil System Schematic
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10.0 Ice Protection
The function of the System is to maintain ice free inlet cowl surfaces during flight and
ground operations. The System consists of ducting, a dual butterfly pressure regulating and shutoff
valve, pressure tap and switch downstream of the valve, an anti—icing distribution spray ring and
an exhaust port located at the bottom of the nose cowl. The anti-icing system is located on the right
side of the engine. The controls and indications are in the flight compartment.
The inlet cowl anti-icing System is a thermal system using hot bleed air from the 5th and
9th stages of the HPC section. The valve is operated by the switch on a panel within the cockpit. A
blue light indicates valve position and an amber light indicates an overpressure condition of 65 psig
or an over temperature condition of 440°C in the anti-icing ducting.
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Figure 33: Thermal Anti-Ice (TAI) Schematic
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References
Aviation Week & Space Technology. (1977, Feburary 21). YC-15 Enter New Flight
Test Series. Aviation , p. 27.
CFMI Customer Training Services. (2002, January). CFM56 Basic Engine.
Cincinnati, Ohio, USA. Retrieved from www.scribd.com.
International, C. (2003). CFM56-3 Line Maintainance Manual. Cincinnaty, Ohio,
USA.
Lufthansa Technical-Training. (2000, April). Training Manual B737-300/400/500.
Frankfurt, Germany.
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