ABSTRACT (EXECUTIVE SUMMARY)

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• ABSTRACT (EXECUTIVE SUMMARY)
• Engine Control System for the Advanced Light
  Helicopter (ALH)
• BY
• Mr. B.P. RAO Mr. K.M.BHAT L.RAJASHEKAR
• (ROTARY WING RESEARCH AND DESIGN CENTRE)
• This paper gives a brief overview of the engine
  requirements and hence its control system
  requirements for a modern helicopter based on
  flight requirement and design philosophy.
  Second part gives a brief insight into the
• Engine modules
• Basic control architecture
• Basic fuel system configuration
• Control system operation
• Indication and monitoring
• Control laws

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• INTRODUCTION
• ALH is integrated with two Turbomeca TM333-2B2
  Engines. These engines are controlled by
  dedicated single channel FADEC.
• The control system is designed to adapt the
  engine to the helicopter power requirements
  while remaining within defined limits. Basic
  principle is to maintain the output shaft rpm at
  the datum value at zero to maximum power level
  throughout the operating envelope of the
  helicopter.
• On TM 333-2B2 engine, the Engine Electronic
  Control Unit achieves this by controlling the
  fuel flow and inlet guide vane positions
  depending on the operating condition and power
  demand.

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• A typical power required vs speed for a given
  altitude and AUW (All Up Weight) is shown in Fig.
  1 along with the components of power
  requirements. The engine power limits of a
  typical helicopter are also shown.
• Fig.2 shows the Power required vs AUW and
  altitude (since the power and weight are
  normalised with density ratio) and the flight
  test validation of the same. Fig. 3 and 4 show
  the power and fuel flow with altitude for a given
  AUW. Fig. 5 shows the variation of power with
  bleed (power for the actual engine is kept
  constant even with bleed) at the expense of fuel.
  ALH is integrated with two Turbomeca TM333-2B2
  Engines. These engines are controlled by
  dedicated single channel FADEC.

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INDIVIDUAL COMPONENTS OF POWER FIG. 1
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FIG. 2
6
FIG. 3
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FIG. 4
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FIG. 5
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SINGLE ENGINE PERFORMANCE REQUIREMENTS AND
TESTS Helicopters have specific requirements
like constant rotor rpm under all operating
conditions, zero to maximum torque, autorotation,
g maneuvers, single engine new additional
requirements for multiengine helicopters etc.
Turbo-shaft engine have requirements like large
gas generators speed variation, minimum turbine
entry temperature and fuel flow function, etc. To
cater for all these requirements the principle is
to control fuel flow and air flow into the
engine. A fixed wing aircraft requires a landing
strip for take off and landing. In the event of
full engine failure, it requires a set of
procedures and capability (a high glide ratio for
all engine failure cases) and reduced speed with
sideslip for multiengine aircrafts, apart from a
proper landing strip.
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A helicopter can hover and it too has a defined
envelope and procedures for flight in case of
single engine failure.. The first is the H-V i.e.
the hover and speed regime for a safe recovery in
case of single engine or two engine failure. The
second is the take off and landing in case of one
engine failure for civil operation called as
CAT-A and CAT-B operation. The third is in the
event of all engine failure, the helicopter can
land safely (provided a safe spot is available)
in oblique descent or even in vertical descent
by autorotation. Here the potential engine of the
helicopter is converted into rotational energy of
the rotor to give a lift for a safe
touchdown. The first two single engine
operations are briefly described below.
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HEIGHT VELOCITY (H-V) DIAGRAM
For every helicopter, whether it is
single engine or multi engine, there exists
an envelope of initial velocities and heights
from which a helicopter cannot
land/flyway safely in the event of an engine
failure.
Height Velocity diagram consists of three main
points (as shown in Fig. 6), they are,

• High Hover Point (HHP)
• Low hover point (LHP
• Knee point
• Fig.7 shows the engine power variation (Actual
  Tests) during H-V Diagram Testing.

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FIG. 6
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FIG. 7
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CATA AND CATB PERFORMANCE
The Category A helicopter must be

• Multi engine
• Each engine should have separate controls and
  independent fuel system such that failure of one
  engine should not hamper the operation of other
  engine.
• It should have stay up capability. In the event
  of engine failure, it should have minimum rate of
  climb.

The Category B helicopters

• May be single or multi engine
• Need not have stay-up capability
• Normally certified for higher AUW and altitudes
  compared to CAT-A helicopters, as they have to
  land immediately in the event of engine failure.
• A typical CAT-A Take Off is shown in Fig.
  8 and the actual flight test
• demonstration is shown in Fig. 9.

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TYPICAL FLIGHT TAKE OFF PATH FOR CAT-A
OPERATION FIG. 8
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TYPICAL CATA FLIGHT DEMONSTRATION FIG. 9
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The main objective of this paper is the engine
control system which can be discussed in more
detail with an understanding of the main engine
requirements. Engine control system on ALH is
configured with FADEC (Full Authority Digital
Engine Controls) and associated controls and
sensors. The main functions of the FADEC are -
Automatic starting - Output Speed control -
Inlet Guide Vane (IGV) control - Automatic fuel
flow control - Various limitation function -
Auxiliary Power Unit (APU) mode control - Over
speed protection - Engine usage monitoring -
Maintenance/diagnostic aid - One Engine
Inoperative (OEI) Training mode
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  Installation requirements   The main functional
requirements of the installation are Constant
rotor rotation speed NR in all operating
conditions - Max torque limit C (usually imposed
by the aircraft transmission) - Complete engine
protection (N1 and N2 speeds, TET. acceleration
control ?N1/?t ...) - Good load sharing (in the
case of a multi-engine configuration).   Adaptati
on to requirements   To have a constant rotation
speed of the power turbine N2. the power supplied
by the engine is automatically adapted to the
demand. This adaptation is ensured by the
control. system which meters the fuel flow
injected into the combustion chamber so as to
deliver the required power (variation of the gas
generator N I rotation speed) while keeping the
engine within its operational limits.

• Power transmission
• The mechanical power supplied hy the engine. is
  used to drive the helicopter rotors through a
  mechanical transmission.
• This power drives
• The main rotor
• The tail rotor
• The main gearbox
• Twin engine configuration
• In a twin engine configuration, the engines arc
  installed at, (the rear of the main gearbox.
•  
• In a twin engine configuration, the engines arc
  installed at, (the rear of the main gearbox.
•  
• The power turbines of the two engines are
  mechanically connected to the main gearbox which
  drives the rotors (main and tail rotors).

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FUEL SYSTEM DESCRIPTION The fuel system
includes the following components Booster pump
(low pressure aircraft system) Low pressure pump
/ Alternator unit This unit includes a
centrifugal low-pressure pump and an alternator
to electrically supply the Digital Engine Control
Unit Fuel filter The fuel filter includes a
filtering element, a pre-clogging pressure
switch, a by-pass valve and a filter-clogging
indicator. High pressure ,pump and metering unit
This unit includes a gear type high pressure
pump fitted with a pressure relief valve. It also
has a metering unit which includes - A constant
?P valve A manual metering valve A fuel metering
unit (controlled by the DECU)  - A stop
electro-valve (of bi-stable type). It opens
during shut-down.

• Inlet guide vane actuator
• This actuator receives HP pump outlet pressure.
• Valve assembly
• The assembly includes
• - A start electro-valve for start injector supply
  during starting
• - A pressurising valve which gives fuel supply
  priority to the start injectors
• - A flow divider which gives supply preference to
  one of the main injectors
• A manual purge valve
• A main injector purge valve
• Fuel injection system
• - 4 start injectors, fitted around the combustion
  chamber casing
•  -12 main injectors. Pre-vaporising injectors
  installed at the rear of the combustion chamber
  casing
• .
•  Indicating devices
•  
• - Low fuel pressure switch

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 FUEL SYSTEM - OPERATION (1) This part deals with
the following operating phases prestart,
starting, normal operation, manual control and
shutdown.  Pre-start   - The LP and HP pumps do
not operate and there is no pressure in the
system The constant ?P valve is closed  The stop
electro-valve is in the "stop" position The start
electro-valve is in ventilation position The
pressurising valve is closed -The flow divider
is closed -The manual metering valve is in the
"neutral" position -The metering unit can be in
any position. During the electrical power up, the
metering unit is initialised.
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VARIABLE INLET GUIDE VANE SYSTEM- DESCRIPTION
OPERATION Description The variable inlet guide
vane system comprises a row of the vanes fitted
in the front compressor casing before the Ist
stage axial compressor. Each vane has a lever
connected to its outer end and each lever is
connected to a common actuating ring around the
outside of the casing The actuating ring is
connected by an actuating rod to the hydraulic
actuator The actuator receives fuel pressure from
the outletet of the HP fuel pump. It also has an
electrical connector through which it is
connected to the DECU.
Operation The angle of the variable inlet guide
vanes is determined by the DECU as a function N1
and T1 The DECU sends signals to the actuator to
vary the fuel pressure either side of a hydraulic
piston, thus actuating the rod and causing the
vane angle to vary.
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• IGV actuator
• Fuel injection system
• Start injectors
• Main injection system
• Electrical components
• Indicating system sensors
• Control system sensors
• Dedicated alternator
• Ignition unit
• Starter-generator
• Digital Engine Control Unit(DECU)
• This control unit controls and monitors the
  engine.
• Digital type
• Installed in the aircraft
• Serial data link (connection with the aircraft)

• Control system description
• The complete system includes aircraft components,
  engine components and the DECU
• Aircraft components
• Control components (analog and logic signals)
• Indicating components (instruments,lights.)
• DECU electrical supply
• Start and stop control logic
• Engine components
• Hydro mechanical components
• Fuel control
• LP pump
• Filter
• HP pump
• Metering unit (with manual control)
• Stop electro-valve
• Valve assembly
• Start electro valve
• Pressurizing valve
• Flow divider valve

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START
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CONTROL SYSTEM OPERATION (2) Functions ensured
by the electronic control system This part
mentions in a general way the main functional
electronic blocks Selection of the control mode
Start control N2 speed control Anticipation IGV
control Limitations (N1 ,torque..) N1 speed
control Flow limitation Control of the metering
valve Overspeed protection APU mode
Monitoring Data bus interfaces with the air craft
Fault detection and indication Cycle
counting ?N1 indication calculation Counting and
indicating of engine usage Engine operating hours
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Engine hours counting
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• CONTROL SYSTEM OPERATION (3)
• Starting
• This function includes the staring sequence ,the
  starting fuel flow control ,idling ,the
  transition from idle to flight and restart in
  flight
• Starting sequence
• The system ensures the cranking
  (starter),ignition
• ( ignition unit) and the fuel supply.
• Start is selected using the Stop Start Flight
  selector
• Stop (selection of engine shut down)
• Start (start control up to idle)
• Flight (normal control)
• In flight start
• The sequence is identical to a ground start ,but
  only permitted below 17N1
• It the start is selected above 17N1 the DECU
  will wait until N1 lt 17 before initiating the
  start.

• Starting fuel flow control
• During staring the fuel flow CH is metered so as
  to provide a rapid start without over temperature
• To this end, the fuel flow is controlled
  according to different laws
• Basic flow law as a function of T1 and residual
  t4.5 gas temperature.
• Starting flow law as a function of N1
  acceleration (?N1 /?t)
• Flow correction law as a function of t4.5 indexed
  proportional to N1.
• The elaborated fuel flow datum CH is used to
  control the metering valve via
• Choice datum
• Metering valve control,depending on the datum

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CONTROL SYSTEM OPERATION (4) Idle when starting
is completed ,the rotation speed stabilises at
idle (94 N2 if idle has been selected
) Transition from idle to flight This is effected
by moving the selector from start /training to
flight During the transition the torque and N2
acceleration are limited The transition is
completed when the system enters in to nominal
speed control Control functions Transition
control Speed control (N2 ,N1 ,limitation) Selecti
on of fuel flow datum CH Flow control Metering
valve control
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• CONTROL SYSTEM OPERATION (5)
• Control general
• Installation configuration
• The gas generator supplies power to the power
  turbine which is connected to the helicopter main
  rotor.
• Installation requirements
• Aircraft rotor speed (NR) almost constant in all
  operating conditions (because of the rotor
  efficiency)whatever the load applied.
• Max torque limitation (imposed by the mechanical
  transmission and the helicopter main gearbox).
• Power turbine rotation speed (N2) within given
  limits (in fact almost constant,as it is
  connected to the rotor)
• Limitation of the gas generator rotation speed N1
  
• Max N1
• Min N1 (to avoid engine flame out and critical
  speeds.)

-Protection against surge ,flameout,
overtemperature. Adaptation to requirements The
control system ensures the engine adaptation to
the requirements by metering the fuel flow CH
sprayed in to the combustion chamber. Thus, the
gas generator adapts automatically to the
requirements (N1 demand) to maintain constant the
power turbine rotation speed N2 whilst keeping
all the other parameters within determined
limits. This adaptation is illustrated by The
diagram W/N1,N2 whish illustrates the power W,the
max torque C and rotation speeds N1 and N2. The
diagram N1.N2 which illustrates the N1/N2
relation curve.
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• CONTROL SYSTEM OPERATION(6)
• Speed control general
• Control loop
• The control loop comprises essentially
• An anticipator linked to the helicopter
  collective pitch lever.
• A power turbine (N2) controller
• An N1 limiter for max and min N1,acceleration .
• An N1 controller
• A fuel flow limiter (CH) for max and min flow
  ,anti flame out,anti-surge.
• A fuel flow controller to control the fuel
  metering valve.

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Operating principle. In this type of control
system the position of the helicopter collective
pitch lever ,which represents the power required
,determines the basic N1 datum .This function
,which is called the anticipator, permits an
initial adaptation of the gas generator speed to
balance the power supplied with the power
required and thus maintain the N2
constant. Further more the anticipator supplies
an immediate signal for a load variation ,which
the detection time and provides a rapid reaction
of the control system. However this first
reaction is not sufficient ,as the power requires
depends on other factors. The basic datum is
modified by the N2 controller ,which is a
proportional controller ,after comparing the
difference between a datum ,the nominal NR and
the measured N2.thus the N2 ,and therefore the
NR, are maintained constant without static
droop. The N1 datum is thus elaborated as a
function of the anticipator and the N2 controller
The N1 datum is then limited in order to assure
certain functions such as rating stops,
acceleration and deceleration control,torque
limiting
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The N1 controller is integral and treats the
difference between the N1 datum and the actual
N1.It translates the difference in to a fuel flow
datum CH, in order to maintain the N1 constant
without static droop The fuel flow limiter then
modifies this datum in order to assure certain
protection functions such as anti surge ,anti
flame out etc. Finally the fuel flow datum is
treated to give a signal to the metering unit
which determines the actual fuel flow injected it
to the combustion chamber ,which determines the
operation of the gas generator ,particularly the
rotation speed N1
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CONTROL SYSTEM OPERATION (7) The control loop
comprises N2 control ,anticipation, N2 datum
selection ,limitations and N1 control. N2 (Free
turbine rpm) control The N2 controller is a
proportional type controller ,that is ,it has no
static droop .The controller treats the
difference between a datum N2 and the actual
N2. The N2 datum is calculated according to the
control mode idle ,flight, training flight. It
is also dependent on aircraft altitude. There is
an N2 Trim selector (Low-Normal-High) for the two
engine, which permits the pilot to select the NR
speed. Anticipation Load variations are
anticipated by a signal from a potentiometer
linked to the collective pitch lever. This signal
,XCP ,acts on the N1 datum. N1 limiter The N1
datum is limited to assure various limitations
(details on the following pages N1 control The N1
controller is an integral controller ,that is
without static droop.This controller treats the
difference between the datum and the actual N1
and elaborates the necessary fuel flow datum
CH Example of transitory control In the case of
an increase of load the control system responds
and increases the fuel flow CH.The N1 increases
in order to maintain N2 constant. APU mode A
cockpit selector permits the selection of APU
mode on No.1 engine only. Selection of APU mode
energises an electric clutch which allows the
engine to drive the aircraft electric generator.
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Torque limitation The N1 datum is also limited
to prevent overtorque of the helicopter main
gear box .These torque limits are also
calculated as a function of N2. Acceleration and
deceleration control During acceleration ,the
rate of change of the N1 datum is limited in
order to avoid surge caused by overfuelling.The
rate of change (?N/?t)is modified as a function
of P0 and T1 . The system includes a more
limiting recovery law which is used in the event
of the P3 signal becoming defective. During
deceleration,the rate of change of the N1 datum
is limited to prevent flame out during rapid
deceleration. Note The lowest demand (thermal
or torque ) will determine the limit.
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CONTROL SYSTEM -OPERATION (15)   Overspeed
  The DECU includes an optional power turbine
overspeed shut-down system.   This system will
shut-down the engine if the power turbine speed
reaches a certain limit.   This function is
ensured by a specific board which is separated
from the control boards (channels A and B).
  The overspeed logic includes two sub-Logics
  -One measures the channel A N2 signal,
compares it with the oyerspeed threshold and
supplies the negative of the stop electro-valve
in case of oyerspeed.   -The other measures the
channel B N2 signal, compares it with the
overspeed threshold and supplies the positive of
the stop electro-valve in case of overspeed.
  Thus, when both sub-logics detect an
overspeed, and only in this case, the engine is
shut-down.  
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  Re-arming   Once operated the system remains
in the overspeed condition until it is re-armed,
eitherby a push button . Cross inhibition The
system is designed to prevent both engines being
shut down by their overspeed systems. This is
achieved by a cross connection between the two
DECUs, when one engine overspeed system operates
it inhibits the overspeed function of the other
system. System indication  The overspeed system
provides self monitoring signal by means of
warning lights One light (OVSP MNTR) indicates
the situation of the system failure- armed or
dis-armed.  One light (ENG OVSP) indicates the
true overspeed situation and that the system has
operated.    Fault detection   A test circuit
simulates an overspeed in order to verify the
correct operation.  The test is selected by
pressing a push button.  After a test the system
is re-armed by depressing the re- arming push
button. During the test and after, the situation
of the system (failure -rearmed) is indicated by
the warning lights.  The unit with overspeed
test and rearming buttons is located in the rear
fuselage.  
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CONTROL SYSTEM -OPERATION (16) Operating
principle In this type of control system the
position of the helicopter collective pitch
lever, which represents the power required,
determines the basic N1 datum. This function,
which is called the anticipator, permits an
initial adaptation of the gas generator speed to
balance the power supplied with the power
required and thus maintain the N2 constant.
Furthermore the anticipator supplies an
immediate signal for a load variation, which
reduces the detection time and provides a rapid
reaction of the control system. However, this
first reaction is not sufficient, as the power
required depends on other factors. The basic
datum is modified by the N2 controller, which is
a proportional controller , after comparing the
difference between a datum the nominal NR, and
the measured N2. Thus the N2, and therefore the
NR, are maintained constant without static droop.

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    CONTROL SYSTEM. INDICATION AND MONITORING (1)
  The system ensures the indication of engine
parameters, performance indication, cycle
counting, rating exceedance time counting, fault
indication and maintenance aid.   Engine
parameter indication  Indication of N 1, N2,
TGT, Torque is provided direct from the engine to
the cockpit, to allow engine monitoring in the
event of total electronic failure and to check
that the DECU is maintaining the engine within
limits.   These indications are independent of
the electronic control.   ?N1 indication  The
system outputs the ?N1 to the cockpit indicator.
Counting functions  The DECU counts and
records in memory ,the N 1 and N2 cycles, the
engine hours run. The DECU is also provided with
a creep damage counter. This information can be
accessed through the RS 232 serial data link.  
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Engine
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• CONTROL SYSTEM -INDICATION AND MONITORING (4)
•  
• Maintenance aid
•  The system ensures the following functions
•  Fault detection
•  Fault isolation and identification of the
  component affected
•  Writing of a fault report containing information
  such as date, fault type, location
•  Transmission of the fault report to the
  helicopter system or to a computer (PC) for
  maintenance practices
•  Recording in the memory of the last fault
  report.
•  
• Data exchange with the aircraft and computer (PC)
  
•  
• The DECU exchanges data with the aircraft by
  means of a serial data link RS 232.
•  
• It can receive
• The engine serial number
• The request for the reset of the counters in
  memory
• The request for the reset of the overspeed usage
•  

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HELICOPTER
N1,N2
of the overspeed usage
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ELECTRONIC ENGINE CONTROL UNIT Function The EECU
controls and monitors the engine operation Main
characteristics Motorola/EFCIS processor EECU is
loaded with Flight critical Software, certified
according to DO 178B standard, Level A. Cards
installed in the EECU are modular. Each card is
separated depending on the functionality e.g, IGV
digital control card, IGV analog control card,
overspeed monitoring card etc. Mechanical
interfaces Installation through shock
mounts Electrical interfacesElectrical
connectors Redundant electrical supply 28VDC
helicopter and 115V AC engine alternator. RS232
interface Non volatile memory
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