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Flight Control Actuation Systems for Fighter Aircrafts:
Philosophy of Design, Development & Challenges in Manufacturing and Testing
K S Nagesh, Ravi Jai Prakash
Aeronautical Development Agency, Bangalore
ABSTRACT
Actuation Systems form a vital cog in any Aircraft
Flight Control System, providing the motive force
necessary to move Aircraft’s Flight Control
Surfaces. Performance of Flight Control System
Actuators has a significant bearing on overall
Aircraft’s performance. High maneuverability of the
Fighter Aircraft is dictated by high performance
Flight Control Actuators, which in turn can lead to
difficult design, control and manufacturing
problems in their own right.
Broadly, Aircraft Flight Control Actuators are
divided into two classes, viz. Primary and Secondary
Actuators. Primary Actuators, are Safety Critical
Actuators while Secondary Actuators, are Mission
Critical Actuators.
State of the art Control concepts such as Active
Control Technology, Control Configured Vehicles,
Relaxed Static Stability, etc., resulted in highly
unstable modern day Combat Aircraft with highly
enhanced performance and agility. This
sophistication in Flight Controls led to an even
greater reliance on primary flight control surface
availability to the extent that many modern combat
aircraft could not be controlled without the
continued operation of the Primary Flight Control
Surfaces. Hence, Primary Flight Control Actuators
are mandatorily built with Hydraulic and Electrical
redundancy, to comply with Fail-Op, Fail-Op and
Fail-Safe philosophy. Failure in Secondary
Actuators can be tolerated, in the sense, that, while
they do not jeopardize Aircraft’s safety, they’ll
result in restrictions on Flight Envelope thus
curtailing Mission capabilities of Aircraft.
With the above introduction, this paper presents,
some details on latest technologies / trends in design
of Flight Control Actuators (both Primary and
Secondary), viz. Direct Drive valve, Servo Valve
Technologies which form heart of Flight Control
Actuation Systems. Effort is also put in, in this
paper, to cover the following aspects:
i. Techniques and Philosophy of Design for
incorporating hydraulic and electrical redundancy,
viz. Rip-stop design, quad redundant electrical
drives, etc.,
ii. Incorporation of highly reliable, robust
feedback devices such as LVDTs, RVDTs
iii. Manufacturing challenges in terms of
machining complex Hydraulic Control Modules
which house Hydro-Logic, Control Valves with sub-
micron accuracies, fabrication of non-linear Disc
Springs, Selective Assembly Techniques, etc., and
iv. Qualification, Acceptance and Flight
testing, to prove conformance to given SQR
Keywords: Direct Drive Valve, Actuators, Tandem,
Magnet, Main Control Valve, Bypass Valves, Linear
Motor
1. INTRODUCTION
Contemporary Aircrafts using Fly by Wire
technology for Flight Control Systems normally
employ two types of Actuators / Actuation System,
viz. Electro Mechanical and Electro Hydraulic.
While considerable research has been going on
towards realization of More / All Electric Aircrafts,
even today, Electro Hydraulic Actuation Systems
form a major chunk of Actuation Systems used for
Flight Control Surfaces, in particular owing to their
high power density.
Nevertheless, Electro Mechanical Actuation
Systems are used for low power and mission critical
applications such as weapon delivery, Brake Chute
deployment, even on Secondary Control Surfaces as
back up devices, etc…Also, considerable research
has been put in across the globe, towards
development of Electro-Hydro Static Actuators
(EHA), and Electric Backed Hydraulic Actuators
(EBHA). But it is still a considerable way ahead
before these systems can be solely deployed for
Control Actuation Systems in Military and Civil
Aircrafts.
Electro Hydraulic Actuation Systems, are broadly
classified into two varieties, viz. Electro Hydro
Servo Valve (EHSV) based and Direct Drive Valve
(DDV) based, depending upon the fluid metering
elements used. EHSV based systems are chiefly
used in 4th generation aircraft platforms like F-16,
Su-30, Mirage 2000, Airbus A320; whereas DDV
based systems are employed in 4+ & 5th generation
platforms like B-2 Bomber, Eurofighter Typhoon,
LCA-Tejas, F-18 (E/F), F-22, PAK-FA, Airbus
A380, Boeing 747, X-47B to name a few.
In the foregoing, design philosophy of Flight
Control Actuators / Actuation System in LCA is
briefly given below:
LCA-Tejas, is a 4.5 generation Fighter Aircraft. It
is designed to be statically unstable with a time to
double >150ms. Flight Control System (FCS) of
LCA is a state of the art Digital Fly-By-Wire Control
System built to meet PLOC requirements of one in
one million flight hours, with different failure modes
of operation, viz. Fail Op/Fail Op/Fail safe.
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To meet the above operational FCS requirements,
LCA has fundamentally two kinds of Flight Control
Surfaces, viz. Primary Flight Control Surfaces and
Secondary Flight Control Surfaces. Being built on
Delta planform, 04nos Elevons and 01no Rudder
make up Primary Flight Control Surfaces while,
06nos Leading Edge Slats (LES) and 02nos
Airbrakes make up Secondary Flight Control
Surfaces.
All Primary Control Surfaces in LCA are actuated
with the help of DDV based Dual Tandem Electro-
Hydraulic Actuators, while LES are actuated with
the help of EHSV based Simplex Electro-Hydraulic
Actuators. Airbrakes are actuated by Electro-
Selector based Actuation System as their functional
requirement is only two position deployment – Full
Extension and Full Retraction.
Being present state of the art technology,
considerable portion of this paper talks about the
design philosophy and development challenges in
DDV based Electro Hydraulic Actuators, starting
with its evolution from EHSV based Electro
Hydraulic Actuators.
2.PRIMARY FLIGHT CONTROL
ACTUATORS:
Primary Control Surface Actuators (Rudder and
Elevon) of LCA, are designed to operate from two
hydraulic sources to meet Failure mode operational
requirements. These actuators are dual tandem type
with a redundant hydraulic control module, in
construction. Actuator Servo Control employs an
electrically quad-redundant linear DDV. Loop
closure electronics are housed in Digital Flight
Control Computer (DFCC).
2.1 CONSTRUCTIONAL FEATURES:
Hydraulic schematic of DDV based Tandem
Actuator commonly used for Primary Flight Control
Surface Actuation System, is shown in Figure-1.
Reference numerals are used for indicating different
functional units of the Actuator. System#1 as shown
is dual hydraulic and quad electrical redundant. This
means that each of the Tandem Actuators 4 and 5 is
powered by independent hydraulic system and all
the electrical functional units have 04 coil windings
operated through independent sources of electrical
power. Actuation System essentially consists of the
following functional units:
i) Direct Drive Valve (DDV)
ii) Solenoid Operated Valve
iii) Bypass Valves
iv) Accumulator
v) Directional Control Valves and
vi) Tandem Actuator
Fig:1 Hydraulic schematic of DDV based redundant Actuation System
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Fig:2 Direct Drive Valve
A DDV 26 consists of a Linear Force Motor (LFM)
32 operated dual tandem Main Control Valve
(MCV) 31, with a position feedback device - Linear
Variable Displacement Transducer (LVDT) 33.
Figure – 2shows a typical DDV. LFM is a linear
motor which in response to a DC signal from the
Flight Computer, positions the spool of the tandem
MCV to make desired hydraulic inter connections.
The position of MCV is fed back to the Flight
Computer by LVDT coupled to the other end.
2.1.1 Linear Force Motor
Linear Force Motor as shown in Figure –
2principally, consists of two cylindrical permanent
Samarium Cobalt magnets having equal pole areas.
These magnets are held within fixed iron, which in
turn is slidably encapsulated in moving iron. Moving
Iron and Magnet Housing combine to form the
desired air gap as shown. Tubular Support, made of
material having high strength to weight ratio is
peripherally coupled to the Moving Iron. This
combination of Moving Iron and Tubular Support
form reciprocating mass in LFM. The Reciprocating
Mass is centred with the help of Diaphragms on
either side. Reciprocating Mass and Magnets are
enclosed by four independent electrical coils. A
return ring made of ferromagnetic material retains
and transmits the lines of flux toward the air gap for
generating high force output.
2.1.2 Main Control Valve (MCV)
Dual Tandem MCV as shown by Figure – 2consists
of a spool, sleeve and push rod. This is the key
element of whole of the Actuator. It is responsible
for symmetrical metering of fluid flow to both the
rod end and tailstock sections of Tandem Actuator 4
and 5 (Figure – 2). A push rod fit to the LFM
transmits motion from LFM to sleeve. Spool is the
key metering element with lands all along its length.
Sleeve encloses the spool and contains rectangular
metering ports matching critically with the lands on
the sleeve.
2.1.3 LVDT: A quad LVDT, with fine tracking
accuracy with all 04 channels, is used for sensing the
spool position of MCV.
2.2 PHILOSOPHY OF DESIGN
This section gives a peep-in into the general design
philosophy / guidelines in the design and
development of Flight Control Electro-Hydraulic
Actuator. The inputs from Aerodynamics, Flight
Mechanics and Control Law teams, in terms of
following are necessary: -
i. Hinge Moment / Stall Load
ii. No-load slew rate / deflection rate
iii. Deflection boundaries of Control
Surface
iv. Band width (Gain and Phase
requirements)
v. Modes of operation
vi. Electrical and Hydraulic Redundancy
requirements
vii. Reliability requirements, etc.
Electro-hydraulic Servo Mechanisms have the
capability of transmitting high power at quick
response. In Servo Synthesis, it is necessary to
determine the specification of the hydraulic servo,
such as the Supply pressure, and flow capacity into
the hydraulic flow control servo valve, and the
actuator size, so that the power requirement satisfies
the desired dynamic characteristics.
In general, a force versus velocity chart is employed
to make sure that the specification is suitable. When
the chart is drawn for the economic design, a drive
characteristic curve of the hydraulic actuator should
effectively enclose the load locus with as little
overlap as possible. This exercise conforms the
Actuators’ specifications to start with the design of
Servo Valve and Actuator.
A typical Actuator design algorithm is captured
below:
The Actuator Stroke is determined from the
installation data of the Actuator under consideration.
Fig:3 Kinematical Schematic of Installation of
Actuator
The Actuator Piston area ‘A’ is
𝐴 =𝑆𝐹 ∗ 𝐻𝑀𝑚𝑎𝑥
𝑟𝑒𝑓𝑓 ∗ ∆𝑝
Where
SF : Safety Factor
HMmax : maximum hinge moment
reff : effective lever arm
Δp : pressure difference at actuator piston
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The pressure difference at the actuator piston has to
be selected carefully taking into account the delivery
pressure of the hydraulic pump at realistic flow
conditions together with the pressure losses in the
valve and those in the supply and return lines.
Contemporary Aircrafts use hydraulic pressures
ranging from 3000PSI to 8000PSI, with most
common applications lying in the range of 3000 to
5000PSI.
Flight Control Actuators are normally operated in
closed loop control, which invariably needs a
feedback device sitting inside the Actuator for Ram
Position feedback. Hence, Piston Inner Diameter (d)
is controlled by this consideration of type and size
of feedback device used.
From the Actuator Piston Area (A), Piston Inner
Diameter (d) and in conjunction with standard
elastomeric seals size, Inner Diameter of Cylinder
(D) is arrived at.
Non-linearity such as damping, friction, etc., are
also captured and accommodated during the above
phase of design.
Pressure drop in the Servo Valve and flow at
maximum deflection rate Qmax of the surface are the
design values for the opening area of the ports in the
servo valve Aopening:
𝐴𝑜𝑝𝑒𝑛𝑖𝑛𝑔 = 𝑄𝑚𝑎𝑥/(𝐶𝑑√(2∆𝑝𝑠𝑒𝑟𝑣𝑜)/2𝜌) )
For a preliminary design, the valve spool stroke s
and the Valve Spool Diameter (d) can be calculated
from standard values d/s, keeping in mind that
𝐴𝑜𝑝𝑒𝑛𝑖𝑛𝑔 = 𝜋𝑑𝑠
Hence,
𝑆 = √𝐴𝑜𝑝𝑒𝑛𝑖𝑛𝑔/(𝜋𝑑
𝑠)
As concerns selection of Servo Valve, whether to go
for DDV or EHSV, there are several criteria. When
considering, single channel non-redundant
applications, EHSV becomes predominant choice
for typical aerospace installation, where weight and
envelope constraints are major limitations. For
those applications requiring increased redundancy
within a single actuation package due to the critical
nature of its control function, the use of a direct drive
control valve offers major advantages in reduced
complexity, total envelope size and weight.
2.2.1 DDV vis-à-vis EHSV
Important Control Valve characteristics which need
to be considered while considering either EHSV or
DDV, are as follows:
- Positional control of the metering spool in
the null region
- High pressure gain, coupled with valve
threshold and / or free play
- Measurable free play or backlash within the
valve control loop
- Valve operating force
Specific advantages of DDV over EHSV:
- reduced internal leakage
- reduced sensitivity to fluid contamination.
- significantly lowered Null shifts,
associated with temperature and pressure
variations
- Spool positional control available prior to
valve pressurization.
Specific advantages of EHSV over DDV:
- requires minimum electrical input power
- favorable weight and package size
- does not require closing an external
electrical feedback loop around the valve
spool
2.2.2 Salient features of DDV:
i) High Dynamic response
ii) Redundancy built into the system coupled
with lesser no. of components
iii) High Chip Shear Force
iv) Efficient i.e. low null leakage, pressure
drops and threshold current
v) Linear input current v/s output flow
vi) Excellent dynamic characteristics @ null
2.2.3 Design philosophy of DDV:
DDV design chiefly involves design of two sub-
systems, viz. Linear Force Motor and Spool Valve
(Main Control Valve).
Design of LFM in turn involves design of several
components, viz. Permanent Magnet, Pole Piece,
coil, Armature and Airgap. Some important
considerations in the design of LFM are: -
- Magnetic Circuit Analysis
- Demagnetization
- Balance at Null Current
Design of Spool Valve involves design of Spool,
Bush and Porting System. Some important
considerations in the design of Spool Valve: -
- Port opening configuration
- Flow Force analysis
- Pressure-drop in Flow passages
Design algorithm of DDV, typically, is as follows:
- Engineering Concept model for LFM,
Spool Valve and LVDT
- Identification of Physical Laws for LFM
and Spool Valve
- Obtaining Control Coefficients for Spool
Valve
o Steady State Flow Force
Coefficients
o Transient Flow Force Coefficients
- Identifying Objective Functions and
optimization
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o Chip Shear Force, Physical
Envelope and Weight
o Constraints based on Steady-state
requirements and valve
bandwidth
- Development, testing and Validation
Sample calculations in the design of Spool Valve:
- Porting System:
𝑊𝑝 = 𝑛𝑝𝑑𝑠 sin (𝜃𝑝
2𝑛𝑝
)
𝐴𝑝 = 𝑊𝑝𝑥
- Steady State Flow Force
𝐹𝑠𝑓 = 2𝐶𝑑𝑝𝐶𝑣𝑝(𝑃𝑠 − 𝑃𝑏)𝑊𝑝𝑐𝑜𝑠𝜃𝑗 . 𝑥
Steady state flow force coefficient-
𝐾𝑠𝑓 =𝜕𝐹𝑠𝑓
𝜕𝑥= 2𝐶𝑑𝑝𝐶𝑣𝑝(𝑃𝑠 − 𝑃𝑏)𝑊𝑝𝑐𝑜𝑠𝜃𝑗
- Transient Flow Force
x1 = distance between s1-c1 & s2-c2
x2 = distance between c1-R & c2-R
𝐹𝑑𝑚𝑝 = 𝐶𝑑𝑝𝑊𝑝(𝑥2 − 𝑥1)√𝑃𝑠 − 𝑃𝑏
𝜌�̇�
Damping Coefficient-
𝐶𝑚 =𝜕𝐹𝑑𝑚𝑝
𝜕�̇�= 𝐶𝑑𝑝𝑊𝑝(𝑥2 − 𝑥1)√
𝑃𝑠 − 𝑃𝑏
𝜌
Dynamic Equations:
a. Spool Valve
𝐹𝑒𝑚 = 𝑚𝑒�̈� + 𝐹𝑑𝑚𝑝 + 𝐹𝑠𝑓
𝐾𝐹𝐶 + 𝐾𝑚𝑥 = 𝑚𝑒�̈� + 𝐶𝑚�̇� + 𝐾𝑠𝑓𝑥
a. Linear Force Motor
𝑒𝑉 = 𝑅𝑐 + 𝐿𝑑𝐶
𝑑𝑡+ 𝐾𝑏�̇�
2.3 CHALLENGES IN MANUFACTURING
AND ASSEMBLY OF DDV:
2.3.1 CHALLENGES IN MANUFACTURING:
Fig:4 Diaphragm spring
i) Fabrication of diaphragms
ii) Machining of brittle magnetic material
iii) Null Edge Grinding of Spool and Sleeve
iv) Achieving sub-micron Cylindricity
between sleeve and spool over ~200mm.
v) Achieving 2μm squareness for all the
metered ports, in reference to a single plane
vi) Handling of high power Samarium Cobalt
Magnets andassemblywith required airgaps
Fig:5 Main Control Valve Assembly
2.3.2 CHALLENGES IN ASSEMBLY:
DDV is a Dual Tandem Valve. This necessitates,
simultaneous and identical metering of fluid to
corresponding actuator chambers to avoid force
fight near null region. Therefore, precise hydraulic
nulling of MCV coupled with quad electrical nulling
of LVDT and LFM are crucial procedures to ensure
Sleeve
Spool
Push Rod
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synchronous operation of Tandem Actuator. DDV is
said to be in null when the algebraic sum of currents
passing in quad coils of LFM produce equal pressure
gains on either half of Main Control Valve (MCV).
i) Hydraulic nulling of MCV with LFM:
This is done at LFM and MCV assembly
stage without LVDT. Referring to Figure – 2, with
input hydraulics ON, a low amplitude sinusoidal
signal is applied to the LFM. Pressure difference
between Control Ports C1-C2 and C3-C4 are plotted
against the input LFM current. The valve hysteresis
yields two occurrences of minimum differential
pressures. This is made symmetrical about zero
point on LFM current axis, iteratively, by means of
shims between LFM and MCV. A sample plot is
shown below.
Fig:6 Hydraulic nulling of DDV
ii) Electrical nulling of MCV & LFM with
LVDT:
After hydraulic nulling, LVDT is assembled to
MCV. The same low frequency sinusoidal signal is
applied and LVDT output versus LFM current is
plotted. The acquired hysteresis plot (Figure - 7) is
made symmetrical about zero point on LFM current
axis by means of shims between LVDT fixed
winding and the movable probe assembly.
2.4 SOLENOID OPERATED BYPASS VALVES
Figure -1, also show two solenoid operated valves
(SOV) 39 and 40 used for piloting bypass valves
(BV) 50, 60 and 70. While, combination of SOV 39
& BV 50 and SOV 40 & BV 60, aids in normal mode
of operation of the Actuator. In the event of failure
of both hydraulic systems, the combination of SOV
39, 40 and BV 60, aids in damped bypass mode
operation.
2.4.1 SOLENOID OPERATED VALVE (SOV)
SOV is a 3/2 type Directional Control Valve. Cross
sectional view of SOV is shown in Figure – 8. It
consists of push rod assembly with steel ball closing
the inlet port (from hydraulic system, under non-
operational condition) under the influence of an
opposing spring. 04 independent electrical coils
enclose push rod assembly.
Fig:8 Solenoid Valve
Fig:7 Electrical nulling of DDV
2.4.1.1 Principle of Operation
When a DCvoltage is applied to quad electrical
coils, the push rod gets pulled inside compressing
the spring and interconnecting inlet port to pilot port.
The return port is used to drain out leakage flow.
2.4.1.2 Manufacturing challenges in SOV
i) Realisation of ~Φ1mm precision steel ball.
ii) Machining of hard material push rod with an end
diameter of ~0.5mmover a length of ~2mm.
2.4.2 BYPASS VALVE
A typical bypass valve shown in Figure-9, consists
of a sliding spool pressed against spring on one side
and port for interfacing with control line of SOV on
the other.
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Fig:9Typical Bypass Valve
2.5 ACCUMULATOR AND DIRECTIONAL
CONTROL VALVES
Figure – 1, shows a piston type accumulator 69. It
is used to replenish lost fluid to the actuator
chambers in damped bypass mode of Actuator.
Directional Control Valves 67 and 68 allow free
flow of fluid in one direction and restricted flow in
the other thereby ensuring damping and eventual
safe ejection of pilot.
2.6 TANDEM ACTUATOR (HYDRAULIC
JACK)
Figure – 1shows Dual Tandem Actuators 4 and 5
connected to each other by means of a common
piston 6 and retainers 10. A Centre Dam 9 is used to
divide the Tandem Actuators. A quad LVDT
assembly 14 is contained inside the hollow section
of piston 6. Fixed quad winding assembly of LVDT
is fastened to the cylinder, whereas probes 16 are
connected to piston 6 in the manner as shown
inFigure – 1. The rod end assembly 23 can be hinged
to the control surface and tail stock to the airframe,
by means of spherical bearings 22.
2.6.1 Manufacturing Challenges in Tandem
Actuator
i) Machining of pistons of
~500mmwhilemaintaining overall cylindricity and
concentricity in the order of 10μm.
ii) Chrome plating and honing of Piston inner
diameter with high l/d ratio ~10.
2.6.2 SALIENT FEATURES ofTANDEM
ACTUATORS (Hydraulic Jack):
i) Lesser seal friction.
ii) High Reliability (for manned aircrafts)
a) Hydraulic redundancy.
b) Multilayered sealing using backup rings.
c) Minimal plumbing joints.
d) Rip stop design.
iii) Improved efficiency of operation
a) Lesser inter system leakage.
b) Lesser pressure drops.
c) Less frictional forces.
iv) Optimized envelope & weight
2.7.1 REALISATION CHALLENGES
In order to contain all the sub systems and the
complicated interconnections for redundancy
management, a control manifold is used. The
manifold shown by Figure–11, houses50-60 criss-
cross holes oriented in different angles.
Design challenges in Control Module are
fundamentally optimization of Wall thickness and
weight while accommodating whole hydraulic logic.
a) Isometric 3D View
b) Criss Cross holes for inter connection
Fig:10 Control Manifold
Machining of this component itself is an engineering
challenge with over 70~80 stages of settings to
achieve >1000 critical dimensions. This component
can be done only on 5 axes CNC machines.
2.7.2 QUALIFICATION TESTING AND
CHALLENGES:
Primary Flight Control Actuators are subjected to
stringent Environmental Tests before they are
certified for fitment onto aircraft. MIL STD 810 is
the reference standard for most of the environmental
tests, as applicable for aerospace applications.
While MIL STD 810 gives general guidelines, it is
the Environmental Map of each Aircraft
Programme, that gives exact environmental
requirements, what Line Replaceable Units (LRUs)
in Aircraft have to conform to.
Some of the critical qualification tests Actuators are
subjected to, while Actuators are operational, are
listed below:
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- Vibration Test
- Shock Test
- Pressure impulse test
- High Temperature
- Low Temperature
- Thermal Shock
- Endurance Cycling
- Functional Shock
- Arrestor Shock
- Acceleration
Each qualification test is preceded and succeeded by
a set of more than ~60 Acceptance Tests to capture,
degradation in the performance of Unit Under Test,
if any.
2.7.2.1 VIBRATION TEST
Vibration causes loosening of fasteners,
intermittency in electrical contacts, seal
deformation, component fatigue, cracking and
rupturing and excessive electrical noise in circuits.
Vibration tests are performed to verify that the
equipment will function in and withstand the
vibration environment
2.7.2.3 ACCELERATION TEST
This test is conducted to assure that equipment can
structurally withstand the steady state inertia loads
that are induced by platform acceleration,
deceleration, and maneuver in the service
environment, and function without degradation
during and following exposure to these forces.
.
2.7.2.4 THERMAL SHOCK
The test consists of stabilizing the actuator in a
chamber at the low temperature extreme i.e. -40 °C,
transferring the unit to a test chamber at high
temperature extreme i.e. 135°C within one-minute
time period, stabilizing and then returning the unit to
the chamber at the lower temperature extreme (-40
°C) again within the 5-minute time interval, several
times based on requirement.
2.7.2.5 HIGH TEMPERATURE STORAGE
AND PERFORMANCE
This test consists of stabilizing the actuator ambient
to 85 °C @ inlet oil temperature of 135 °C. Actuator
is subjected to all performance tests at these
temperatures.
2.7.2.6 LOW TEMPERATURE STORAGE AND
PERFORMANCE
This test consists of soaking the actuator at -54 °C
inside a thermal chamber for at least three hours.
After stabilization, all performance tests are
conducted at inlet oil and actuator ambient
temperature of -40 °C.
2.7.2.7 PRESSURE IMPULSE TEST
This test is conducted in accordance with SAE ARP
1383 standard. This test is an accelerated fatigue
test, wherein endurance of test item against
cumulative fatigue calculated from Miner’s Rule is
assessed.
2.7.2.8 ENDURANCE CYCLING
This test consists subjecting Actuator to ~5million
loaded cycles. This test needs to be conducted on a
Test Rig having Servo Loading capability. These
cycles are performed by repeating a Single Loop
Schedule (SLS), a combination of specified cycles
of different loads and strokes, @different
temperatures up to 135°C
3.0 SECONDARY FLIGHT CONTROL
ACTUATORS:
Leading Edge Slat (LES) Actuators and Airbrake
Actuators make up Secondary Control Surface
Actuators of LCA. These are considered secondary
as they are not safety critical but mission critical.
Accordingly, the redundancies built into Secondary
Actuators are also Simplex hydraulically and duplex
electrically, thus leading to only two modes of
operation, viz. Normal Mode and Fail-Safe mode.
These actuators are built of double-acting hydraulic
cylinders and are controlled by EHSV in case of LES
Actuators and by means of an Electro-Selector
Valve in case of Airbrake Actuators.
Actuator Servo Control in case of LES Actuators
employ an electrically duplex-redundant EHSV.
Loop closure electronics are housed in Digital Flight
Control Computer (DFCC). In case of Airbrake
Actuators, there are no Servo Control electronics as
it is only a two position Actuator actuated by means
of an Electro-Selector.
There is a duplex Linear Variable Differential
Transformer (LVDT), a position feedback device, fit
in the Hydraulic Jacks of both LES and Airbrake
Actuators. Feedback from LVDT is used for Loop
Closure in case of LES Actuators, while it is used
only as an indication of the position / deployment of
Actuators in case of Airbrake Actuators.
The following sections deal principally with LES
Actuators.
3.1 CONSTRUCTIONAL FEATURES:
A sample Secondary Control Surface Actuation
System is shown in the following sketch.
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Fig:11 Hydraulic schematic of EHSV based
redundant actuation system
This Actuation System, fundamentally, consists of
an EHSV, Hydraulic Cylinder (Jack), and Control
Module. While Control Module in case of
Secondary Actuators is not as complicated as in case
of Primary Actuators, there are special features built
into the Control Module such as Thermal Relief
Valves, Manual Release, etc., to address specific
failure modes of operation.
3.2 Electro Hydraulic Servo Valve (EHSV):
The following figure shows a typical nozzle-flapper
EHSV.
Fig:12Simplified diagram of EHSV
EHSV, uses a very fine nozzle stage for enabling the
movement of second stage spool metering element.
The Torque Motor is driven by differential current
input to two coils and develops torque, proportional
to input current, on the armature. The first stage,
called hydraulic amplifier, converts angular
displacement armature to unbalanced back pressure
of two nozzles. These changes in back pressure are
felt on both ends of the spool, effecting desired spool
movement. Movement of the Spool is fed back to
Armature by means of Feedback Wire connecting
Spool to the Armature. Thus the feedback spring
converts the Spool position to a force signal, which
is fed back to the torque motor. The Spool moves to
the position where the torque fed back by the
feedback spring balances the torque due to the input
current.
EHSV is spring biased, to retract. This necessitates
a bias current to be continually applied, to keep
EHSV @null. This provision of retract biasing
enables LES Actuator to retract automatically EHSV
torque motor fails, provided hydraulic power is
available.
3.3 Other features in LES Actuators:
There are two Thermal Relief Cum Check Valves.
These valves control hydraulic oil flow from EHSV
to the two chambers of Actuator’s Hydraulic Jack.
These valves allow flow from the valve only when
the applied hydraulic pressure exceeds a minimum
pressure called shuttle pressure. Whenever supply
pressure drops below the shuttle pressure, the valves
disconnect Actuator from EHSV, thus rendering
Actuator hydraulically locked. This causes the LES
to hold its last position when a hydraulic failure
occurs.
Thermal Relief Cum Check Valves open whenever
chamber pressures inside a hydraulically locked
actuator rise to higher levels (Under conditions like
excessive aerodynamic load on a locked actuator).
Check Valves bleed excessive pressure to prevent
the Actuator from structural and thermal damage.
Then, there is a Manual Release Valve. This valve is
a hand operated valve to inter-connect both the
control ports. This allows flow of hydraulic oil
between extend and retract chambers of the
Hydraulic Jack whenever the piston is hydraulically
locked.
3.4 Design and Sizing of EHSV:
Fundamental design requirements for Servo Valve
are as follows: -
- Max. Supply Pressure
- No load rated flow
- Max. null leakage
- Rated current
- No load flow gain
- Spool travel
Some of the critical steps involved in the design and
sizing of EHSV are as follows:
- Sizing of Torque Motor
Design of Permanent Magnets
Design of Feedback Wire Assembly
- Sizing of Hydraulic Amplifier
- Design of Spool Valve
Sizing of Control Port
Sizing of drain orifice
This paper does not go into the details of the design
of EHSV.
3.5 CHALLENGES IN ASSEMBLY
1. Assembly of EHSV is the most challenging
in case of LES Actuators. Some of the
challenges in EHSV assembly are as
follows:
a. Nulling of spool
Page 10 of 10
b. Finer adjustment of distance
between Nozzle and Flapper
within few μm
c. Iterative procedure of adjusting
air gaps between armature and
pole pieces.
d. Charging of magnets using
magnetizer.
3.6 REALISATION CHALLENGES OF LES
ACTUATOR
1. Hard Chrome Plating of blind inner
diameter of the order of ~40mm, in
hydraulic jack.
2. Lapping of nozzle in Thermal Relief Check
valve ensuring zero leakage.
3. Drilling of very high l/d (~60) holes in
Control Module
Vital components of Servo Valve are the Valve
Body, Spool, Nozzles, Orifice and Feedback Wire
Ball Assembly. Challenges in machining of these
parts is given below:
Machining of Valve Body
Machining of Spool
Machining of Nozzle and Orifice
Feedback Wire Ball Assembly
Resistance welding of a ball Φ0.8mm to Feed
wire.
3.7 QUALIFICATION TESTING AND
CHALLENGES INVOLVED:
Qualification testing of Secondary
Actuators / LES Actuators also entail all the
qualification tests explained under Primary
Actuators section. Qualification tests such as
Vibration Test, Acceleration, Shock, etc., do not
have any change between that of Primary Actuators
and that of LES Actuators. However, tests such as
Pressure Impulse Test, Endurance Test, etc., are
lighter in intensity, not in terms of loads or pressures
encountered by Unit Under Test, but in terms of no.
of cycles of load.
Hence, all the challenges associated with
qualification testing of Primary Actuators apply
equally good to all Secondary Flight Control
Actuators.
4.CONCLUSIONS:
While, the content covered in this paper
talks about, only the outline in Flight Control
Actuators / Actuation Systems design, realization,
testing and challenges involved. ADA has
successfully imbibed complex technologies
involved in Servo Valves. These developments have
seen Indian private industries to participate in such
high precision flight critical applications. The skill
set developed in this country has helped these
industries greatly.
LCA Programme has already seen certain
spin-offs of this capability for several upcoming
Aircraft programmes, while ISRO has used this
technology for launch vehicle programmes on a
large scale.
5. REFERENCES:
PATENTS
1. Control actuation system for aerospace vehicles
and a method thereof; B B Das, KS Nagesh and KS
Anand Kumar; 3653/CHE/2011, dtd.: 24.10.2011
2.Improved Direct Drive Valve; K V Simon, K. S.
Nagesh and K S Anand Kumar
3.Improved Linear Force Motor; Sreenivasan, K. S.
Nagesh and K S Anand Kumar
STANDARDS
4. Actuators: aircraft flight controls, power operated,
hydraulics, general specifications; SAE ARP 1281.
5. Electro hydraulic flow control servo valves; SAE
ARP-490 Rev D
TECHNICAL PAPERS / BULLETINS
6. Flight control actuation system for the B-2
advanced technology bomber; W S Schaeler, L J
Inderhees and J F Moynes; MOOG technical
bulletin 153, dtd.: 23.04.1991
7. Reducing complexity in fly by wire flight control
actuators; B S Lyle, General Dynamics SAE
Technical Paper 851752, 1985, Published: 1985-10-
01, DOI: 10.4271/851752
8. Proceedings National Workshop on Aerospace
Servo Systems, June 25-26, 1998
9. Direct Drive Valve based Dual Tandem Electro
Hydraulic Servo Actuation System Working
Principles and Challenges in Manufacturing by K.S.
Nagesh in DRDO Technology Spectrum, May 2014,
pp.03-09, ISSN 2348-5809
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