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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