Fis Book 3 Index: Airframe

FIS BOOK 3
INDEX
AIRFRAME
Chapter Title Page No.

1. Glossary of Terms 235
2. Aircraft Structure 243
3. Design and Construction 253
4. Structural Integrity and Fatigue 265
4A. Annex to Ch-4 Theory of Metal Fatigue 271
5. Hydraulic Systems 275
6. Pneumatic Systems 287
7. Powered Flying Control Systems 293
8. Automatic Flight Control Systems 303
9. Undercarriages 313
10. Braking Systems 325
Cabin Pressurization and Air Conditioning
11.
Systems
331
12. Fire Warning and Extinguisher Systems 339
13. Ice and Rain Protection Systems 347
14. Aviation Fuels 355
15. Composite Materials 369
16. Aero-Elastic Tailoring 381
17. Aircraft Abandonment 385
18. Oxygen Systems 399
235
CHAPTER 1
GLOSSARY OF TERMS
1. Aeroplane. A mechanically driven heavier than air aircraft (fixed wing) supported by dynamic
reactions of the air upon its wings.
2. Aerofoil. A surface designed to produce lift when driven through the air.
3. Aileron. A horizontal control surface usually hinged

to rear spar of main plane near wing tip for the purpose of
lateral (roll) control (Fig 1-1).
4. Aircraft. Any weight carrying device designed to be

supported by the air, either buoyant or dynamic.
5. Airframe. An aircraft without its engine/engines.

Fig 1-1: Aileron of HPT - 32
6. Air Brake. Any device primarily used to increase the air drag of an aircraft at will.
7. Alclad. It consists of duralumin sheet rolled on both sides with aluminium sheets used in
covering the fuselage main planes etc.
8. Amphibian. An aeroplane capable of taking off and alighting on either land or water.
9. Angle of Attack. The angle which the chord line makes with the relative airflow.
10. Anhedral. Downward inclination of the wings towards wing tip i.e. the downward angle the
wing makes with the horizontal when the aircraft is in rigging position.
11. Aspect Ratio. The span of an aerofoil divided by its average chord.
12. Attitude. Inclination of the axes of an aircraft relative to the horizon.
13. Auto Rotation. The tendency of an aerofoil, at high angles of attack, to rotate about the
relative airflow.
14. Axis. Imaginary lines about which the aircraft rotates.
15. Balance Tabs. Small surfaces hinged to trailing edge of the main control surfaces to reduce
aerodynamic loads and to assist the pilot in moving the control surfaces (Fig 1-1).
FIS Book 3: Airframe 236
16. Bank. The angle between the lateral axis of an aircraft and the horizontal.
17. Bonding. Connecting all metal parts of the aircraft to secure good electrical continuity and to
avoid the undesirable effects from Static Electricity.
18. Buffeting. An irregular oscillation of any part of an aircraft produced and maintained directly
by turbulent airflow.
19. Bulkhead. Any transverse partition in the fuselage which separates the two parts completely
from one another.
20. Camber. The curvature of a surface of an aerofoil.
21. Cantilever. A spar or beam fixed rigidly at one end only.
22. Centre of Gravity. That point at which the total weight is considered to act, irrespective of the
position of the body concerned.
23. Centre of Pressure. The point, usually on the chord line, through which the total reaction
may be considered to act.
24. Centre Section. The portion of the fuselage or hull forming a continuous structure with the
mainplane.
25. Chord. The distance between the leading and trailing edge of an aerofoil measured along the
chord line.
26. Cockpit. The portion of a fuselage or nacelle designed to accommodate pilot and crew.
27. Control. The intentional maneuvering of the aircraft into any desired position.
28. Control Column. The lever (or the pillar supporting hand wheel) by which the elevator and
aileron controls are operated.
29. Control Surface. A surface movable in flight, whose primary function is to govern the motion
of aircraft in pitch, roll or yaw.
30. Datum Line. A line fixed by the designer from which measurements are made when rigging
or truing the aircraft.
31. Dihedral. Upward inclination of the wings towards wing tip i.e. the upward angle the wing
makes with the horizontal when the aircraft is in rigging position.
237 Glossary of Terms
32. Differential Aileron . The aileron system whereby the up moving aileron will operate through
a larger angle than the down moving aileron.
33. Down Wash. The downward deflection of a stream of air by an aerofoil, or part of the aircraft.
34. Duralumin. It is an alloy of aluminium used mostly for the construction of aircraft structural
members.
35. Elevator. A horizontal control surface by which the upward or downward inclination of an
aeroplane in flight is controlled. The elevator is usually hinged to the trailing edge of the tail plane.
36. Empennage. The tail section of the aeroplane and

consists of a tail cone, fixed and movable surfaces (Fig 1-
2).
37. Fairing. Additions to any structure to reduce its

drag.
Fig 1-2: Vertical Stabilizer

38. Fin. Fixed vertical surface or surfaces forming part
of the tail unit by which the yawing stability of the aeroplane is affected.
39. Fineness Ratio. In a stream lined section, the ratio of maximum depth to length.
40. Flap. A movable auxiliary surface controlled by the pilot, for increasing effective camber of the
main plane.
41. Flutter. A violent, destructive vibration of the

aerofoil surfaces, caused by interaction of their inertia
loads, aerodynamic loads and structural stiffness.
42. Formers. Transverse members which are

used to give required shape to a fabric covered
fuselage or body on which the stringer rest and to
support the skin of monocoque structure (Fig 1-3).
Fig 1-3: Monocoque Fuselage
43. Form Rib. A non-structural member of an aerofoil whose purpose is to preserve the shape of
the aerofoil.
44. Frise Aileron. The aileron designed and pivoted so that when it is raised its nose projects
below the main plane. The nose does not project above the main plane when aileron is lowered.
45. Frames. Transverse structural members of the

fuselage supporting the longerons (Fig 1-4).
46. Fuselage. The main structure or body of most

types of the aircraft to which wings, tail plane, fin, rudder
and other surfaces are attached.
47. Glider. A non-power driven heavier than air

aircraft.
Fig 1-4: Cut Out of a Fuselage
48. Horn Balance. A localized balance area at the tip
of the control surface. This may be unscreened or screened by a surface in front (Fig 1-2).
49. Hull. The main structural watertight body of a flying boat which gives buoyancy and stability
on the water and enables it to take off and alight.
50. Jury Strut. A strut inserted to provide temporary support for a structure.
51. Keelson. A longitudinal member forming part of the main structure of a hull or float and
running internally along the bottom.
52. Incidence Angle. The angle which the chord line makes with the horizontal.
53. Lateral Axis. An imaginary line through aircraft's centre of gravity considered to be parallel to
line joining the wing tips.
54. Leading Edge. The forward member of an aerofoil.
55. Longerons. The principle longitudinal members of the fuselage which run from front to rear
and is usually supported at various points along its length by other structural members (Fig 1-4).
56. Longitudinal Axis. An imaginary line running fore and aft through the centre of gravity.
57. Mainplanes. The horizontal main wings of aerofoil section attached to the fuselage which
provides most of the lift to the aircraft.
58. Monocoque. A type of aircraft structure in which all the load is taken by the skin only.
59. Nacelle. The term applied to bodies enclosing crew and/or engines mounted on or between
the planes of an aircraft.
60. Normal Axis. An imaginary line, through the aircraft's centre of gravity which is vertical when
the lateral and longitudinal axis is horizontal.
61. Oleo Strut. Undercarriage member designed to absorb the landing shock by damping and by
preventing a rebound, thus providing a smooth landing.
62. Ornithopter. An aircraft which may obtain lift and propulsion by mechanical flapping of the
wings. No successful commercial type has been designed so far. (Leonardo da Vinci had designed
one in 1490 !!!)
63. Pitching. Angular motion about the lateral axis.
64. Propeller. A screw designed to convert the power developed by the engine into forward
‘thrust’.
65. Rib. A part of the aerofoil structure designed to carry compression and shear loads and to
give an aerodynamic surface the desired contour.
66. Rigging. It is the relative assembly checking and alignment of the aircraft.
67. Rigging Position. It is the position of the aircraft when the longitudinal and lateral axis of the
aircraft are in horizontal plane.
68. Rolling. Angular motion about the longitudinal axis.
69. Rudder. The movable vertical surface forming part of the tail unit by which the yawing of the
aeroplane flight is controlled (Fig 1-2).
70. Sailplane. A glider capable of sustained flight under the influence of atmospheric currents.
71. Semi-monocoque. A type of aircraft structure in which the skin also shares part of the load.
72. Service Ceiling. The height at which the rate of climb of an aircraft has a certain defined low
value (usually 100 ft/min).
73. Slat. An auxiliary aerofoil attached to the leading edge of the main aerofoil used to increase
the camber of the aerofoil when operated.
74. Slip Stream. The stream of air driven to the rear by the rotating propeller.
75. Slotted Aerofoil. An aerofoil having its two surfaces connected by one or more air passages
(or slots), whose primary object is to improve the airflow at high angles of attack.
76. Span. The overall distance from wing tip to wing tip.
77. Spar. A principal longitudinal member of a plane or of a control surface.
78. Stability. The quality possessed by an aircraft of returning to its original state of steady
motion after disturbance, without the use of controls.
79. Streamline. A body so shaped as to offer the minimum resistance to motion through the air is
said to be streamlined.
80. Stressed Skin Construction. Any construction which relies wholly or in part on a skin (with
or without reinforcement) to carry the main loads and provide stiffness.
81. Stringer. A stiffener which also assists the skin to carry direct load in the direction of its
length.
82. Strut. A structural member intended to resist compression in the direction of its length.
83. Sweepback / Sweep forward. The angle made by the perpendicular at root chord to the line
joining the quarter chord points of a wing.
84. Tail Plane or Stabilizer. An aerofoil placed at the rear end of the fuselage to balance and
provide longitudinal stability.
85. Tail Unit. The rear portion of an airframe, that includes tail plane, fin, rudder and elevators.
86. Trailing Edge. The aft edge of a streamline body or aerofoil.
87. Trimming Tab. A small tab attached to the trailing edge of a control surface, usually
adjustable in flight, to correct the trim of the aircraft under varying speeds and loads (Fig 1-2)
88. Truing. The adjustment of a component to its correct shape.
89. Under Carriage or Alighting Gear. The part of an aeroplane intended for its support on land
or water and to absorb the shock of landing.
90. Wash In. An expression used to denote the increasing angle of incidence towards the wing
tip.
91. Wash Out. An expression used to denote the decreasing angle of incidence towards the wing
tip.
92. Winglet. A small nearly vertical wing like surface, usually of aerofoil section, attached to the
wing tip. It is usually located rearward above the wing tip and is effective in reducing induced drag.
93. Wing spar. The principle spanwise structural wing member, designed to carry all the stresses
distributed through the whole area of the wing.
243
CHAPTER 2
AIRCRAFT STRUCTURE
Introduction
1. The airframe of a fixed-wing aircraft is generally considered to consist of five principal units,
the fuselage, wings, stabilizers, flight control surfaces, and landing gear. Similarly helicopter airframes
consist of the fuselage, main rotor, tail rotor (on helicopters with a single main rotor), and related
gearbox, and the landing gear.
2. The airframe components are constructed of a wide variety of materials and are joined by
rivets, bolts, screws, and welding or adhesives. The aircraft components are composed of various
parts called structural members (i.e., stringers, longerons, ribs, bulkheads, etc.). Aircraft structural
members are designed to carry a load or to resist stress. A single member of the structure may be
subjected to a combination of stresses. In most cases the structural members are designed to carry
end loads rather than side loads that is, to be subjected to tension or compression rather than
bending. Strength may be the principal requirement in certain structures, while others need entirely
different qualities. For example, cowling, fairing, and similar parts usually are not required to carry the
stresses imposed by flight or the landing loads. However, these parts must have such properties as
neat appearance and streamlined shapes.
Major Structural Stresses
3. In designing an aircraft, every square inch of wing and fuselage, every rib, spar, and even
each metal fitting must be considered in relation to the physical characteristics of the metal of which it
is made. Every part of the aircraft must be planned to carry the load to be imposed upon it. The
determination of such loads is called stress analysis. The term ‘stress’ is often used interchangeably
with the word ‘strain’. Stress is an internal force of a substance which opposes or resists deformation.
Strain is the deformation of a material or substance. Stress, the internal force, can cause strain. There
are five major stresses to which all aircraft are subjected:
(a) Tension. Tension is the stress that resists a force that tends to pull apart (Fig 2-1).
The engine pulls the aircraft forward, but air resistance tries to hold it back. The result is
tension which tries to stretch the aircraft. The tensile
strength of a material is measured in pounds per square
inch (psi) and is calculated by dividing the load (in
pounds) required to pull the material apart by its cross-
sectional area (in square inches). Fig 2-1: Tension
(b) Compression. Compression is the stress

that resists a crushing force (Fig 2-2). The
compressive strength of a material is also measured
in psi. Compression is the stress that tends to
shorten or squeeze aircraft parts.
(c) Torsion. Torsion is the stress that Fig 2-2: Compression

produces twisting (Fig 2-3). While moving the aircraft
forward, the engine also tends to twist it to one side,
but other aircraft components hold it on course.
Thus, torsion is created. The torsional strength of a
material is its resistance to twisting or torque.
(d) Shear. Shear is the stress that resists the Fig 2-3: Torsion
force tending to cause one layer of a material to
slide over an adjacent layer. Two riveted plates in
tension (Fig 2-4) subject the rivets to a shearing
force. Usually, the shearing strength of a material is
either equal to or less than its tensile or compressive Fig 2-4: Shear
strength. Aircraft parts, especially screws, bolts, and
rivets, are often subjected to a shearing force.
(e) Bending. Bending stress is a

combination of compression and
tension. The rod in Fig 2-5 has been
shortened (compressed) on the inside of
the bend and stretched on the outside of
the bend.
Fig 2-5: Bending
Fuselage
4. The fuselage is the main structure or body of the aircraft. It provides space for cargo, controls,
accessories, passengers, and other equipment. In single engine aircraft, it also houses the power
plant. In multi-engine aircraft the engines may either be in the fuselage, attached to the fuselage, or
suspended from the wing structure. They vary principally in size and arrangement of the different
compartments. There are three general types of fuselage construction: the truss type, monocoque
type and the semi-monocoque type.
5. Truss Type. A truss is a rigid framework made up of members such as beams, struts, and
bars to resist deformation by applied loads. The truss-framed fuselage is generally covered with fabric
245 Aircraft Structure
(Fig 2-6) and is usually constructed of steel tubing

welded together in such a manner that all members of
the truss can carry both tension and compression
loads. In some aircraft, principally the light, single-
engine models, truss fuselage frames are constructed
of aluminum alloy and may be riveted or bolted into
one piece, with cross-bracing achieved by using solid
rods or tubes. Fig 2-6: Warrant Truss
6. Monocoque Type. The monocoque (single

shell) fuselage relies largely on the strength of the skin
or covering to carry the primary stresses. The true
monocoque construction (Fig 2-7) uses formers, frame
assemblies, and bulkheads to give shape to the
fuselage, but the skin carries the primary stresses.
Since no bracing members are present the skin must
be strong enough to keep the fuselage rigid. Thus, the
biggest problem involved in monocoque construction is
maintaining enough strength while keeping the weight
within allowable limits. Fig 2-7: Monocoque Construction
7. Semi-monocoque Type. To overcome the

strength / weight problem of monocoque construction,
a modification called semi-monocoque construction
(Fig 2-8) was developed. In addition to formers,
frame assemblies, and bulkheads, the semi-
monocoque construction has the skin reinforced by
longitudinal members. The reinforced shell has the
skin reinforced by a complete framework of structural Fig 2-8: Semi-monocoque Construction
members. Different portions of the same fuselage

may belong to anyone of the three classes, but most aircraft are considered to be of semi-monocoque
type construction. The semi-monocoque fuselage is constructed primarily of alloys of aluminum and
magnesium, although steel and titanium are found in areas of high temperatures. Primary bending
loads are taken by the longerons, which usually extend across several points of support. The
longerons are supplemented by other longitudinal members, called stringers. Stringers are more
numerous and lighter in weight than longerons. The vertical structural members are referred to as
bulkheads, frames and formers. The heaviest of these vertical members are located at intervals to
carry concentrated loads and at points where fittings are used to attach other units, such as the wings,
power plants, and stabilizers. Fig 2-9 shows one more form of the semi-monocoque design now in
use.
8. The stringers are smaller and lighter

than longerons and serve as fill-ins. They have
some rigidity, but are chiefly used for giving
shape and for attachment of the skin. The
strong, heavy longerons hold the bulkheads and
formers, and these, in turn, hold the stringers. All
of these joined together form a rigid fuselage
framework. There is often little difference
between some rings, frames, and formers. One
manufacturer may call a brace a former, Fig 2-9: Fuselage Structural Members
whereas another may call the same type of
brace a ring or frame.
9. Stringers and longerons prevent tension and compression from bending the fuselage.
Stringers are usually of a one-piece aluminum alloy construction. Longerons, like stringers, are
usually made of aluminum alloy. By themselves, the structural members discussed do not give
strength to a fuselage. They must first be joined together by such connective devices as gussets,
rivets, nuts and bolts, or metal screws. A gusset (Fig 2-9) is a type of connecting bracket. The metal
skin or covering is riveted to the longerons, bulkheads, and other structural members and carries part
of the load. The fuselage skin thickness will vary with the load carried and the stresses sustained at a
particular location.
10. There are a number of advantages in the use of the semi-monocoque fuselage. The
bulkheads, frames, stringers, and longerons facilitate the design and construction of a streamlined
fuselage, and add to the strength and rigidity of the structure. The main advantage, however, lies in
the fact that it does not depend on a few members for strength and rigidity. This means that a semi-
monocoque fuselage, because of its stressed-skin construction, may withstand considerable damage
and still be strong enough to hold together.
11. Fuselages are generally constructed in two or more sections. On small aircraft, they are
generally made in two or three sections, while larger aircraft may be made up of as many as six
sections. Quick access to the accessories and other equipment carried in the fuselage is provided for
by numerous access doors, inspection plates, landing wheel wells, and other openings.
Wing Structure
12. The wings of an aircraft are surfaces which are designed to produce lift when moved rapidly
through the air. The particular design for any given aircraft depends on a number of factors, such as
size, weight, use of the aircraft, desired speed in flight & at landing, and desired rate of climb.
13. The wings of some aircraft are of cantilever design, that is, they are built so that no external
bracing is needed. The skin is part of the wing structure and carries part of the wing stresses. Other
aircraft wings use external bracings (struts, wires, etc.) to assist in supporting the wing and carrying
the aerodynamic and landing loads. Both aluminum alloy and magnesium alloy are used in wing
construction. The internal structure is made up of spars and stringers running span wise, and ribs &
formers running chord wise (leading edge to trailing edge). The spars are the principal structural
members of the wing. The skin is attached to the internal members and may carry part of the wing
stresses. During flight, applied loads which are imposed on the wing structure are primarily on the
skin. From the skin they are transmitted to the ribs and from the ribs to the spars. The spars support
all distributed loads as well as concentrated weights, such as fuselage, landing gear, and on multi-
engine aircraft, the nacelles or pylons.
14. Inspection openings and access doors are provided, usually on the lower surfaces of the
wing. Drain holes are also placed in the lower surface to provide for drainage of accumulated
moisture or fluids. Various points on the wing are located by station number. Wing station 0 (zero)
is located at the center line of the fuselage, and all wing stations are measured outboard from that
point.
15. In general, wing construction is based on one of three fundamental designs described below.
Modifications of these basic designs may be adopted by various manufacturers.
(a) Monospar. The monospar wing incorporates only one main longitudinal member in
its construction. Ribs or bulkheads supply the necessary contour or shape to the aerofoil.
Although the strict monospar wing is not common, this type of design modified by the addition
of false spars or light shear webs along the trailing edge as support for the control surfaces, is
sometimes used.
(b) Multi-spar. The multi-spar wing incorporates more than one main longitudinal
member in its construction. To give the wing contour, ribs or bulkheads are often included.
(c) Box beam. The box beam type of wing construction uses two main longitudinal
members with connecting bulk-heads to furnish additional strength and to give contour to the
wing.
16. The main structural parts of a wing are the spars, the ribs or bulkheads, and the stringers or
stiffeners, as shown in Fig 2-10.
17. Wing Spar. Spars are the principal structural members of the wing. They correspond to the
longerons of the fuselage. They run parallel to the lateral axis, or toward the tip of the wing, and are
usually attached to the fuselage by wing fittings, plain beams, or a truss system. Spars may be made
Fig 2-10: Internal Wing Construction
of metal or wood depending on the design criteria of a specific aircraft. Most aircraft recently
manufactured use spars of solid, extruded aluminum or short aluminum extrusions riveted together to
form a spar. As a rule, a wing has two spars. One spar is usually located near the front of the wing,
and the other about two-thirds of the distance toward the wing's trailing edge. Regardless of type, the
spar is the most important part of the wing. When other structural members of the wing are placed
under load, they pass most of the resulting stress on to the wing spars.
18. Wing Ribs. Ribs are the structural crosspieces that make up the framework of the wing.
They usually extend from the wing leading edge to the rear spar or to the trailing edge of the wing.
The ribs give the wing its cambered shape and transmit the load from the skin and stringers to the
spars. Ribs are also used in ailerons, elevators, rudders, and stabilizers. Various types of ribs are also
illustrated in Fig 2-11. In addition to the wing rib, sometimes called ‘plain rib’ or even ‘main rib’, nose
ribs and the butt rib are shown. A nose rib is also called a false rib, since it usually extends from the
wing leading edge to the front spar or slightly beyond. The nose ribs give the wing leading edge area
the necessary curvature
and support. The wing rib or
plain rib, extends from the
leading edge of the wing to
the rear spar and in some
cases to the trailing edge of
the wing. The wing butt rib
is normally the heavily
stressed rib section at the
inboard end of the wing
near the attachment point to
the fuselage. Depending on
its location and method of
attachment, a butt rib may Fig 2-11: Basic Rib and Spar Structures
be called a bulkhead rib or a compression rib, if it is designed to receive compression loads that tend
to force the wing spars together.
19. Honeycomb Construction. It is an efficient way of stabilizing / stiffening relatively thin

skins which are in close proximity (Fig 2-12). It is used on wing and tail trailing edges and access
panels on many aircraft types. The honeycomb material is usually manufactured from strips of
aluminium foil or resin impregnated Nomex cloth.
The strips are bonded together alternately into a
large block which can be machined to any desired
shape. After machining, the block is stretched, and
the strips deform into a matrix of hexagonal cells -
similar to a honeycomb. These may then be bonded
between two sheets of aluminium, or a composite
material such as Kevlar, to form a light stiff
sandwich. Sandwich panels are made using a
variety of skin materials including stainless steel,
titanium and CFC, and often with expanded foams
instead of honeycomb. Such high strength materials
Fig 2-12: Honeycomb Sandwich Panel
are used extensively in current aircraft primary
structures.
Nacelles or Pods
20. Nacelles or pods are streamlined enclosures used on multi-engine aircraft primarily to house
the engines. They are round or spherical in shape and are usually located above, below, or at the
leading edge of the wing on multi-engine aircraft. If an aircraft has only one engine, it is usually
mounted at the forward end of the fuselage, and the nacelle is the streamlined extension of the
fuselage.
21. An engine nacelle or pod consists of skin,

cowling, structural members, a firewall and engine
mounts. Skin and cowling cover the outside of the
nacelle. Both are usually made of sheet aluminum
alloy, stainless steel, magnesium or titanium.
Regardless of the material used, the skin is
usually attached to framework by rivets. The
framework usually consists of structural members
similar to those of the fuselage. The framework
Fig 2-13: Semi-monocoque
includes lengthwise members, such as longerons And
and stringers, and widthwise / vertical members, Welded Tubular Steel Engine Mounts
such as bulkheads, rings and formers. A nacelle or pod also contains a firewall which separates the
engine compartment from the rest of the aircraft. This bulkhead is usually made of stainless steel
sheet metal, or as in some aircraft, of titanium. Another nacelle or pod member is the engine mount.
The mount is usually attached to the firewall, and the engine is attached to the mount by nuts, bolts,
and vibration-absorbing rubber cushions or pads. Fig 2-13 shows examples of a semi-monocoque
and a welded tubular steel engine mount used with reciprocating engines.
22. To reduce wind resistance during flight, the landing gear of most high-speed or large
aircraft is retracted (drawn up into streamlined enclosures).The part of the aircraft which receives or
encloses the landing gear as it retracts is called a wheel well. In many instances, the wheel well is
part of the nacelle; however, on some aircraft the landing gear retracts into the fuselage or wing.
Empennage
23. The empennage is also called the tail

section and most aircraft designs consist of a tail
cone, fixed surfaces and movable surfaces. The
tail cone serves to close and streamline the aft
end of most fuselages. The cone is made up of
structural members (Fig 2-14) like those of the
fuselage however, cones are usually of lighter
construction since they receive less stress than
the fuselage.
Fig 2-14: Tail Cone
24. Other components of the typical

empennage are of heavier construction than the
tail cone. These members include fixed surfaces
that help steady the aircraft and movable surfaces
that help to direct an aircraft's flight. The fixed
surfaces are the horizontal and vertical stabilizers.
The movable surfaces are usually a rudder and
elevators. Fig 2-15 shows how the vertical
surfaces are braced, using spars, ribs, stringers,
and skin in a similar manner to the systems used
in a wing.
25. Stress in an empennage is also carried Fig 2-15: Construction Features of
like stress in a wing. Bending, torsion, and shear, Rudder and Vertical Stabilizer
created by airloads, pass from one structural member to another. Each member absorbs some of the
stress and passes the remainder to other members. The overload of stress eventually reaches the
spars, which transmit it to the fuselage structure.
Flight Control Surfaces
26. The directional control of a fixed-wing aircraft takes place around the lateral, longitudinal, and
vertical axes by means of flight control surfaces. These control devices are hinged or movable
surfaces through which the attitude of an aircraft is controlled during takeoff, flight, and landing. They
are usually divided into two major groups, the primary or main, and the auxiliary control surfaces.
27. The primary group of flight control surfaces consists of ailerons, elevators, and rudders.
These are similar in construction and vary only in size, shape, and methods of attachment. In
construction, control surfaces are similar to the all-metal wing. They are usually made of an aluminum
alloy structure built around a single spar member. Ribs are fitted to the spar at the leading and trailing
edges and are joined together with a metal strip. The ribs in many cases are formed from flat sheet
stock. They are seldom solid and more often the formed, stamped out ribs are reduced in weight by
holes which are punched in the metal. The control surfaces of some aircraft are fabric covered.
However, all turbojet powered aircraft have metal-covered surfaces for additional strength
28. The control surfaces previously described can be considered conventional, but on some
aircraft, a control surface may serve a dual purpose. For example, one set of control surfaces, the
elevons, combines the functions of both ailerons and elevators. Flaperons are ailerons which can also
act as flaps. A movable horizontal tail section is a control surface which supplies the action of both the
horizontal stabilizer and the elevators.
29. The secondary or auxiliary group of control surfaces consists of such members as trim tabs,
balance tabs, servo tabs, flaps, spoilers, and leading edge devices. Their purpose is to reduce the
force required to actuate the primary controls, to trim and balance the aircraft in flight, to reduce
landing speed or shorten the length of !he landing roll, and to change the speed of the aircraft in flight
They are usually attached to, or recessed in, the main control surfaces.
253
CHAPTER 3
DESIGN AND CONSTRUCTION
Introduction
1. This chapter is intended to be a brief introduction to the study of airframes from the designer's
point of view, including some of the general problems confronting him. As with most subjects a
knowledge of the steps which led to the present position is a great help in understanding current
problems. Flying machines obviously have changed enormously over the last hundred years. Wright
Flyer, which the Wright Brothers flew at Kittyhawk, was very different than the Apollo that went to the
moon, and a fighter ace of 1918 flew a very different aircraft than what his successors fly today.
Wire Braced Structures
2. If we look at the aircraft in Fig 3-1 we can have no doubt about the form of its construction.
The wings and the fore and aft structures carrying the covered surfaces were all made of rectangular
frames which were prevented from collapsing (or parallelogramming) by wires stretched from corner
to corner. Although the methods were not original,
there were two imaginative pieces of structural
thinking. Firstly, the idea that two wings, one
above the other, would make a lighter, stronger
structure than the type of wing arrangement
suggested by bird flight, and secondly, the idea
that a rectangle could be held in shape with two
light wires rather than with one much heavier
diagonal member like a farm gate. At that stage,
and for the next thirty years, the major structural
Fig 3-1: Replica of Wright Flyer (1903)
material was wood initially bamboo and later
mainly spruces, a lightweight timber with very straight grain and medium strength. Strangely enough,
balsa wood, which means so much to the model aircraft enthusiast, was not used during that period
but has been used much more during recent years as a filler, or core material, in flooring panels for
large aircraft. Wire bracing continued to be used as a major feature of aircraft construction for many
years.
Braced Monoplanes
3. The braced monoplane design is used almost exclusively for small high-wing aircraft, of which
the Cessna–152 is a typical example (Fig 3-2). The bracing struts, running from the fuselage to a
point about halfway along the wing, relieve the spars of much of their vertical load and anchor them in
torsion. A form of wing construction similar to the biplane is therefore possible, but the spars must be
deeper to resist the greater bending loads, and
consequently a thicker wing section is used. To
minimize the wing area and weight without an
excessively high landing speed, flaps are
generally fitted.
Unbraced or Cantilever Monoplanes Fig 3-2: Cessna-152
4. With the elimination of all external bracing, the wing structure needs to be much stiffer in
bending and torsion. This requires the deepest possible spars, but, as explained in the chapters on
aerodynamics, thick wing sections are unsuitable for high speeds. A compromise is made by tapering
the wings in elevation from a thin tip to a thicker root, where the stresses are greatest. To keep the
wing section approximately uniform the wing is also tapered in plan. Compared with a rectangular
plan form, this has the additional advantage of bringing the centre of lift nearer the fuseIage, so
reducing bending loads. However, excessive taper causes undesirable stalling characteristics.
Definitions
5. To avoid misconceptions of the terms used in this chapter, the following list of definitions is
included:
(a) Elastic Limit. When stress exceeds the elastic limit of a material, the material takes
up a permanent 'set', and on release of the load it will not return completely to its original
shape.
(b) Stiffness or Rigidity. It is the ratio of stress over strain (Young's Modulus).
(c) Hooke’s Law. The change of length is directly proportional to the change of force
(up to a certain value).
(d) Young’s Modulus. Different materials extend by different amounts when subjected
to the same load. So by relating stress and strain in the way suggested by Hooke’s law
provides us with an important identifying characteristic of the material. This characteristic or
the property of the material is called Young’s Modulus or the Modulus of Elasticity.
(e) Design Limit Load. The maximum load that the designer would expect the
airframe or component to experience in service is the Design Limit Load.
(f) The strength of a beam is proportional to its depth squared, and stiffness to its depth
cubed.
255 Design and Construction
OPERATING CONDITIONS FOR AIRFRAMES
Static Strength and Stiffness
6. Static strength requirements determine the design of a large proportion of aircraft structure.
They are specified by applying specific safety factors to the Design Limit Loads (DLLs) which result
from each design case:
(a) Proof Load. Proof load is normally 1.125 X DLL. When proof load is applied, the
aircraft structure must not suffer any permanent deformation and flying controls and systems
must function normally.
(b) Design Ultimate Load. The design ultimate load (DUL) is 1.5 X DLL. The structure
must withstand DUL without collapse.
7. Static strength is proved by loading a representative airframe to, first, proof load and then to
DUL in the critical design cases. Such static strength testing is carried out before the type is released
to service. Structural failure beyond DUL (as is usual) implies that the structural reserve factor is
greater than 1.0.
8. Aero-elastic effects (in particular flutter) often set limits to the maximum speed of fixed wing
aircraft in each configuration. Structural stiffness is normally checked during static and ground
resonance testing. A flight flutter investigation is included in the development programme. Flutter is a
violent, destructive vibration of the aerofoil surfaces, caused by interaction of their inertia loads,
aerodynamic loads and structural stiffness.
9. When structures incorporate fibre-reinforced composite materials, design calculations and

testing have to take account of additional environmental factors such as temperature, moisture uptake,
ultra-violet ageing and the potential effects of accidental damage.
10. Temperature, Corrosion and Natural Hazards. Airframes need to contend with
elevated temperatures from two causes viz local heating of the structure near the engines, heat
exchangers, hot gas ducts etc. and kinetic heating of the outside surface of the airframe at high Mach
numbers. At high TAS large increases in temperature is caused by the compression of air. The
forward-facing parts of an aircraft are being subjected continuously to this heat and take on the
temperature of the air with which they are in contact. For example, at 40,000 feet at the standard
temperature of -56.5° C and at a sustained speed of Mach 2 the temperature rise results in an
airframe temperature of about 120°C. This temperature is above the safe limits of the materials used
in erstwhile conventional aircraft construction. While the effects of the temperature rise on the external
portions of the airframe can be overcome by use of special metals such as titanium, the cockpit and
various electronic and hydraulic components have to be artificially refrigerated. In the absence of
special metals, stainless steel or a thicker aluminium alloy must be used to allow for the loss of
strength at the high temperature accompanying high speeds. At about 94°C, aluminium alloy loses
10% of its strength but at about 120°C, it loses 40%. This means that 40% more aluminium alloy
must be used on those parts of the airframe exposed any length of time to stabilized temperatures of
120°C, and this additional material increases the structure weight of the aircraft. Corrosive conditions,
caused by fluids such as water and general spillages, are a sizeable problem for airframe maintainers
and corrosion is exacerbated by damage to paint and other protective finishes. Some materials,
particularly certain aluminium alloys and steels are susceptible to stress corrosion cracking where
cracks grow in a stressed component from corrosion on the surface. New designs should, wherever
possible, exclude the use of materials known to be highly susceptible to stress corrosion cracking or
exfoliation corrosion and eliminate undesirable features such as water traps, dissimilar metal contact
or ineffective protective coatings.
11. Environmental Conditions. Operating conditions also affect material choice and
component design. In addition to the natural hazards of lightning and bird strikes, design attention
must be paid to the effects of flight in saline environments and flight through erosive and sand laden
atmospheres. Man-made hostile environments such as atmospheric industrial pollution also influence
the range of materials which can be used in airframe construction.
Material Requirements
12. The ideal properties of an airframe material would include: low density, high strength, high
stiffness, good corrosion resistance, high impact resistance, good fatigue performance, high operating
temperature, ease of fabrication and low cost. This list is not exhaustive, and needless to say, there
is no perfect material. Each material choice is a compromise which attempts to find the best balance
of the most important requirements of the component. There is always a mix of materials on any
particular aircraft type.
13. It is generally found that the stronger and stiffer an engineering metal is, the more dense it is.
Therefore the strength / weight and stiffness / weight ratios for the commonly-used aircraft metals are
similar. The increasing use of composite materials reflects their inherent advantages over metals.
First, within their physical capabilities, they can be readily engineered to meet any specific
requirement. Secondly, composite components can be designed and manufactured to complex
shapes. For example, the weight and size of a gearbox casing can be minimized by the use of a
skeletal composite structure in which each stress is reacted to by a specific feature of the structure,
thus eliminating surplus material. This is almost impossible to achieve with conventional metal
manufacture, although integral machining or chemical etching techniques can be used to manufacture
minimalized metal components of such simple geometric shapes as wing and skin panels.
14. For each metal there may be dozens of different specifications, each with its own
characteristics. The strength, hardness and ductility can vary greatly but the stiffness varies little. The
'specific' properties of a material, i.e. strength / weight and stiffness / weight ratios are important
because the designer is always seeking to minimize mass. High stiffness of a material is important
because we normally wish structures to deflect as little as possible. Even if two materials have similar
specific properties, the less dense material may be the better choice as it will have a greater volume
for a given load and the extra thickness is of advantage for components in compression or shear
(doubling the thickness quadruples the load at which a sheet will buckle). However, the space
available may, in some circumstances, restrict the use of the less dense materials.
Use of Materials
15. Since the first days of aviation, the manufacture of safe and effective aircraft structures has
demanded the use of the highest contemporary technology. Indeed, the majority of technological
advances can be attributed to aviation. Early aircraft were constructed by the most skilled craftsmen
using the best timber and other materials available. The need for more consistent and stronger
materials led to developments in aluminium alloys during the 1930s, although wood and canvas were
still widely used even during the Second World War. Arguably the first totally composite aircraft, the
DH 98 Mosquito, owed much of its high performance and subsequent success to the ingenious use of
a hard wood and balsa wood sandwich composite in its structure. Aircraft constructed since the war
have relied almost entirely upon aluminium, magnesium, titanium and steel, although the structures of
current aircraft, make extensive use of advanced composite materials. Several all-composite aircraft
are in production, and the trend towards their greater use is likely to continue in the present Century.
16. The most commonly used material in the current generation of airframes is still aluminium
alloy. Titanium alloy is used for structure adjacent to engines, heavily loaded fuselage frames and
items like flap tracks, which are subject to wear and not easily protected against corrosion.
Magnesium alloy is very rarely used except for small items like control-run brackets, and for helicopter
transmission cases. Steel is sometimes used for heavily-loaded parts like wing attachments and
undercarriage components, and of course for bolts and other hardware. Carbon Fibre Composites
(CFC) are now appearing on large passenger transports for components such as control surfaces,
fairings and even fins. The Harrier-2 has CFC wings, tail, rudder and front fuselage, and a similar
proportion of CFC seems likely to be used in the next generation of combat aircraft
17. Aluminium alloy is cheap, and is easy to produce, to machine and to fabricate. Pure
aluminium has good natural corrosion resistance but is very weak. The stronger aluminium alloys
have poorer corrosion resistance which is sometimes improved by cladding the alloy with a thin layer
of pure aluminium. In any case, they are normally protected with a paint scheme. Components like
wing skins, spars, ribs and fuselage frames are often machined by integral milling, where up to 90%
of the material is removed (Fig 3-3). Thinner skins may be chemically etched to remove the metal for
weight saving. Many small parts are

made by cutting from a sheet and
bending to shape. Final assembly can
be by adhesive bonding but is usually by
riveting and bolting. The very highest
strength aluminium alloys (with about 7%
zinc as the main alloying ingredient) are
often used for wing upper skins but they
have a relatively poor fatigue
performance. Slightly less strong, but Fig 3-3: Machined Skin Wing Construction
more commonly used, are the Duralumin
alloys with 4% copper. These alloys have to be heat-treated to produce the required properties and
this means that their service temperature is limited to 1200 C. If the material is accidentally heated
above this, by fire for instance, it may have to be renewed. Under development is a series of
aluminium alloys which contain about 3% lithium. These have about 10% increase in both specific
strength and stiffness over established alloys, and may well compete with CFC.
18. Magnesium alloy is not as easy to work as aluminium, although it machines and casts quite
readily. Its two major disadvantages are its reactivity (which means that it burns easily) and its
extreme susceptibility to corrosion. Magnesium must be protected extremely well against corrosion
and should be used only where it can be inspected easily.
19. Titanium is very expensive indeed owing to its natural scarcity and the difficulty of extraction
from its ore. Titanium alloy is not easy to form or machine but it can be welded and cast with care.
The advantages of titanium alloy are that it has a high specific strength, maintains its properties well
up to 4000C, and has very good corrosion resistance (it needs no protective coating). It is used for
fire-walls, helicopter rotor hubs and wing carry-through sections and many other applications.
Welding is generally done by the electron beam method.
20. Superplastic Forming / Diffusion Bonding (SPF/DB). SPF/DB is a relatively new process
which is being used at the moment to make production items, like a Tornado heat exchanger, out of
titanium. It is also being developed for future use on aluminium. At about 9000 C, titanium will deform
steadily under constant load to a strain of several hundred per cent, this is super plasticity. At the
same temperature, if two clean titanium surfaces are pressed together they will weld by atomic
diffusion across the joint line, this is diffusion bonding. SPF/DB combines the two processes in one
operation. SPF/DB can save cost, weight and production time compared to alternative methods and
can give a better finished product with less fasteners and excellent welds.
21. Steel. There are hundreds of different specifications of steel with widely varying properties.
Wide use is made of many different alloying constituents and various heat treatments. It is
comparatively easy to machine, form and weld, except for the very high strength and stainless
versions. Low alloy steels are used for fasteners and highly-loaded brackets. Steel has to be
protected against corrosion by painting or plating with cadmium or chromium. Stainless steels with
upwards of 12% chromium are used infrequently in airframes.
22. Non-metallic Composites. The term ‘composite’ is usually taken to mean a matrix of a
thermo-setting plastic material (normally an epoxy resin) reinforced with fibres of a much stronger
material such as glass, Kevlar (a synthetic filament) boron or carbon. Glass provides a useful cheap
material for the manufacture of tertiary components, whereas Kevlar, boron and carbon fibre
composites (CFC) are used in the manufacture of primary structures such as control surfaces or
helicopter rotor blades.
23. Advantages of CFC. The main advantages of CFC are its increased strength / weight and
stiffness / weight compared to the normal airframe materials. This results in lighter structures and
hence performance benefits, the usual weight saving being about 20%. CFC is very resistant to
corrosion and does not necessarily have to be protected. Complex shapes can be made more easily
than in metal. Large structures can be manufactured in one piece, thus saving machining and
assembly time. CFC has a very good fatigue performance, especially at the stress levels used in the
present generation of CFC structures
24. Disadvantages of Composite Materials. Current composite materials have inherent

drawbacks which limit their use, although continuing research and development will solve these
problems in due course. Almost all such disadvantages result from limitations of the matrix materials
used. The most relevant of these drawbacks are:
(a) Most composite materials are relatively elastic, and unless adequately reinforced they
tend to fail prematurely at such features as fastener holes.
(b) Their use is limited to ambient temperatures below about 1200C.
(c) Unless adequately protected, they tend to absorb atmospheric moisture.
(d) They have poor impact resistance, the main effect of which is poor resistance to
erosion caused by hail or sand. Leading edges of composite flying surfaces are frequently
fitted with a titanium or stainless steel sheath to overcome this problem.
(e) They are more difficult to repair than comparable metal structures. Curing the resins
during manufacture and repair must be carried out within a narrow range of temperature and
humidity conditions, and this requires the use of environmentally controlled facilities.
AIRCRAFT DESIGN
Combat Aircraft
25. Introduction. The main features influencing fuselage design for a combat aeroplane will be
power plant installation, fuel and undercarriage stowage and weapon carriage. Stealth technology
(the avoidance of a recognizable infra-red and radar signature) will play an increasing role in future
designs.
26. Area Ruling. In following the area rule to minimize drag, the design will aim at a low frontal
area with a smooth build-up of cross-section area over the forebody, canopy and wings, followed by a
gradual decrease over the afterbody and tail surfaces, the whole being spread over a length defined
by fuel volume requirements.
27. Powerplant Installation. The choice of the number of engines has an important effect on
fuselage shaping and directly influences the fuselage cross-section shape. The decision is largely a
policy one, influenced by airworthiness, cost and maintainability factors.
28. Intake Design. The number of engines will also dictate the disposition of intake ducts and
fuel tanks, again influencing fuselage shape. Intake design is governed by speed considerations and
engine size, since the intake must deliver air to the engine with minimum pressure loss, evenly
distributed over the engine face and free from turbulence. This is of particular importance at high
supersonic speeds where the air must be slowed to subsonic speed.
29. Weapon Carriage. Internal weapon stowage has the advantage that area ruling is
unchanged, whatever the aircraft role. Additionally, the external surfaces are aerodynamically clean.
However, during the life of an aircraft, weapon requirements are likely to be diverse and will alter
several times.
30. Stowage Points. External stowage is more adaptable to different roles and requirements,
and may be under the fuselage or under or on top of the wings. The pros and cons of alternative
methods of external stowage are shown in tabular form in Table 3-1.
31. Front Fuselage Structure. For most types of interceptor/attack/strike aircraft the fuselage
is a conventional stressed skin, stringer and longeron arrangement with frames and bulkheads. The
front fuselage, forward of the wing/fuselage joint, contains the small pressurized area of the crew
compartment. This is confined by a small front bulkhead, cockpit floor and a larger rear bulkhead.
The forward pressure bulkhead carries the weapons system radar or weather radar, whilst the rear
pressure bulkhead, which may slope, normally carries ejection seat attachment points, and may also
accept nose undercarriage mounting loads. The fuselage upper longerons, taking end loads due to
bending, may also carry the canopy rails and/or fixing points. This is the conventional front fuselage
layout.
Stability and Loads and Release and

Position Performance Accessibility
Control Flutter Jettison
Under Tandem Forward stores Generally free Usually good Generally
fuselage carriage destabilizing from flutter and satisfactory if
possible: directionally; flexibility pylon-
reduces drag of stores remote problems mounted
rear stores from C of G can
cause large pitch
disturbances
Under or Side-by-side Destabilizing in Generally Release of Usually good
over carriage gives pitch when in causes flutter: light stores
wing unfavourable front of a low requires big difficult; wing
interference tailplane. tanks to be flow causes
drag. Big stores Asymmetric compartment- pitch
sometimes carriage causes alized to control disturbances:
affect induced rolling moment - C of G. requires
drag adversely can limit g on pull Response to accurate
out jettison of store C of G
heavy stores control - can
can cause conflict with
severe loads flutter
consideration
Table 3-1: Features of External Store Positions
32. Wing/Fuselage Transfer. The manner in which the wing loads are transferred to the
fuselage will depend to a large extent upon engine and wing position.
(a) Wing Configuration. A low wing gives a shorter undercarriage and good crash
protection whereas a high wing can give a better aerodynamic performance. A mid-wing
position is attractive aerodynamically for supersonic performance.
(b) Light Loading. If the wing structure is lightly loaded it is possible to transmit the
loads to the fuselage through heavy root ribs and strong fuselage frames.
(c) Heavy Loading. With a heavier loaded wing structure it is usual to make the whole
centre section box continuous across the fuselage, thus avoiding transmitting primary wing-
bending loads to the fuselage.
33. Centre and Rear Fuselage. The volume aft of the rear pressure bulkhead is available for
the avionics bay, fuel stowage, engines and jet pipes. The strength and stiffness necessary for taking
tailplane and fin bending and torsional loads will depend upon the particular design layout, for
example one fin or two, or whether the aircraft is of tailless design. If deck landings are envisaged,
then the rear structure must be capable of accepting loads from the arrester hook.
34. Variable Wing Geometry. The design of a variable wing geometry aircraft invariably leads
to problems of cost and weight penalties which must be traded off against the better overall
performance gained. The wing carry-through box and the wing attachment pivots must be adequately
stiff, strong and crack resistant. Apart from the complexity of the wing sweep actuators there is the
problem of providing the mechanism necessary to maintain parallel pylons. For good aerodynamic
performance the seals between wing and glove must be effective over the full sweep range and allow
for any structural flexibility. The pivot bearings themselves are in a highly loaded region, yet they
must maintain low friction properties under dynamic loads in varying temperature conditions.
Transport Aircraft
35. Pressurized Fuselage. For a large pressurized fuselage the problem is complicated by the
presence of cut-outs from windows, doors and undercarriage stowage and by the interaction effects at
the wing/body junction. Adequate provision must be made against catastrophic failure by fatigue and
corrosion and the task is further complicated by safeguarding against accidental damage. In practice
the present trend is to design a fuselage in which structural safety is provided by the fail-safe concept.
In detail, all fuselages are greatly different in design, but pressure cabins have a number of principles
in common:
(a) Low stress levels due to pressurization loads.
(b) Cut-outs, such as doors and windows, are reinforced so that the fatigue life is at least
equal to the basic fuselage.
(c) Materials are chosen for good fatigue properties and the fuselage is provided with
crack-arresting features.
36. Fuselage Shape and Cross-section. The use of a parallel fuselage eases seating and
stowing arrangements. It also permits stretching of the fuselage, as development of the engine and
airframe proceeds, to extend the life of the basic design. Usually the parallel part of the fuselage is
continued as far aft as is consistent with tail clearance. There are four main factors which determine
fuselage size:
(a) Cabin length and width (i.e. usable floor area).
(b) Freight volume.
(c) Passenger and freight distribution to maintain the correct CG position (a serious
problem with random seat loading).
37. Wing Position. A high-wing aircraft generally has a lower floor to ground height than a low-
wing aircraft, which eases loading and may result in a reduction of weight. The overall height will be
less, which eases ground maintenance. However, a low-wing aircraft provides a better cushion in
case of a crash landing, whereas for a high-wing aircraft it is necessary to strengthen the main
fuselage frames to prevent collapse of the wing. The problems of main undercarriage attachment and
stowage are less for a low-wing aircraft.
265
CHAPTER 4
STRUCTURAL INTEGRITY AND FATIGUE
Introduction
1. Structural integrity is concerned with the structural airworthiness of aircraft. The objectives of
structural integrity policy are:
(a) Flight Safety. The principle aim is to prevent structural failure and to maintain the
airworthiness of aircraft. Statistically, the probability of losing an aircraft through structural
failure must not exceed 0.001.
(b) Operational Effectiveness. Aircraft must remain operationally effective with

minimal restrictions on the release to service. Front line availability should not be
compromised by the need for structural maintenance.
(c) Life Cycle Costs. Fatigue life consumption and structural maintenance workloads
must be kept low to minimize life cycle costs, particularly as many aircraft are retained in-
service beyond originally planned dates.
2. In high performance aircraft, fatigue is a primary threat to structural integrity. All aircraft
structures incur insidious fatigue damage in normal service due to the fluctuations and reversals of
loads which arise from flight manoeuvres, turbulence, pressurization cycles, landing, and taxiing. This
chapter deals primarily with the ways in which the fatigue problem is dealt with-in the Air Force aircraft,
but other threats to structural integrity which should be noted include:
(a) Overstressing. When loads in excess of design limitations are applied to an

airframe structure, damage may result beyond a level that would contribute incrementally to
fatigue. Overstressing can result in irreparable structural damage, necessitating withdrawal of
the airframe from service or even in catastrophic failure during flight.
(b) Operational Hazards. Operational hazards include bird strikes, ricochet damage,
and battle damage. Aircraft are designed to withstand specified levels of operational damage,
based on the statistical probability of occurrence.
(c) Corrosion. Corrosion is a chemical reaction occurring in metals exposed to water

or other contaminants which not only weakens structures by causing metal decay, but also
introduces stress concentrations that reduce fatigue life.
(d) Stress Corrosion Cracking. Stress corrosion cracking occurs when certain metal
alloys are subjected to permanent tensile stress in a corrosive environment. Such tensile
stress is usually residual, as a result of forging, interference fit, or misalignment during
assembly. The normal atmosphere is often sufficient to destroy the cohesion between
individual metal grains, and stress corrosion cracking propagates around the grain boundaries.
Metal Fatigue
3. When a load is applied to a metal, a stress is produced, measured as the load divided by the
cross-sectional area. The stress at which the metal fractures is known as its ultimate stress, which
defines its ability to resist a single application of a static load. However, if a metal structure is loaded
and unloaded many times, at levels below the ultimate stress, the induced stresses cause
accumulative damage which will eventually cause it to fail. This type of damage constitutes metal
fatigue, and it is important to note that eventual failure can occur at a stress level well below the
ultimate stress.
4. In metal, it is normally tensile loads that cause fatigue. Cracks almost always start on the
surface of structures where there are discontinuities of shape, such as fastener holes, screw threads,
machining marks, and sharp corners. Surface flaws caused by corrosion and fretting can also help to
initiate fatigue cracks.
5. Composite materials too are subject to fatigue failure, but the failure mechanism is different,
and is unlikely to pose comparable problems in service. Reinforcement materials have considerably
higher inherent fatigue strengths than conventional metal alloys, and fatigue is not a consideration in
composite materials at stress cycles below approximately 80% of ultimate stress. When fatigue does
occur in composites, its mechanism is a fracture of individual filaments of the reinforcing material.
Because composite materials are not homogeneous, such fractures do not precipitate cracking but
more gradual weakening of the whole material mass. Thus, whilst metal structures suffering fatigue
retain their design strength until cracking reaches a critical point when failure occurs very rapidly,
composite structures only gradually loose their properties. Such 'soft failures' as they are termed
provide ample evidence to be observed during routine inspection well before critical strength is lost.
Design Philosophy
6. There are basically two design philosophies used to counter the threat to structural integrity
from fatigue:
(a) Safe-life. Safe-life design, which is used for most combat and training aircraft, aims
to ensure that there will be no cracking of significant structure during the specified life of an
267 Structural Integrity and Fatigue
aircraft. Crack arrest is not a feature of safe-life design, and should there be a significant risk
that cracking has started, the structure is usually retired from service.
(b) Fail Safe or Damage Tolerant. The damage tolerant, or fail-safe approach,
requires that structures maintain adequate strength in service until planned inspection reveals
unacceptable defects. Fail-safe structures are characterized by crack-arrest features,
redundant load paths, and relatively long critical crack lengths. Fracture mechanics analysis
plays a large part in damage tolerant design, and the setting of inspection intervals. Most
transport type aircraft are designed to damage tolerant principles.
Establishing Fatigue Life
7. The relationship between various

alternating stress levels which can be
applied to a material and its endurance to
failure, can be established by testing a
large number of typical structural elements.
Fig 4-1 illustrates such a relationship in an
'S-N curve', which plots the applied
percentage of ultimate stress against the
number of cycles to failure. It can be seen
in this example that if the applied stress Fig 4-1: S-N Curve
was 80% of the ultimate stress, the
specimen could be expected to fail after 100 applications, but if the applied stress was reduced to
20% of the ultimate stress, failure would not occur until there had been 10 million applications. The
curve also shows that if the stress level in the structural element was to be increased by (say) 20% of
the ultimate stress (eg. from 30% to 50%), the life would be reduced by a factor of 10.
8. Annexure to this chapter considers the basic applications of S-N curves to the calculation of
fatigue life in structures. In order to apply fatigue formulae the design authority needs to know in
considerable detail how an aircraft will be used in service. The range and frequency of 'g' forces
which are likely to be applied during manoeuvres have the greatest influence on combat aircraft and
trainers, but turbulence, cabin pressurization, flying control movements, landings, vibrations, and
buffet are examples of other factors which must be taken into account.
9. To validate the theoretical life it is a requirement for Service aircraft that an early pre-
production airframe undergoes a full-scale fatigue test (FSFT), during which the structure is subjected,
in a test rig, to the load reversals expected to arise during typical operational sorties. Finally, a factor
is applied to allow for S-N curve scatter, and a fatigue life is declared for that particular aircraft.
Fatigue Life Conservation
10. One of the objectives of structural integrity policy given at para 1, is the minimizing of life
cycle costs, which can be achieved in part by conserving fatigue life. Clearly, the way in which an
aircraft is operated has a prime influence on the rate at which fatigue damage is accumulated.
Structures are designed to carry no more than the loads arising when aircraft are flown in accordance
with the prescribed flight limitations, and pilots are responsible for ensuring that the flight limitations
are not exceeded. A cockpit accelerometer provides a reasonably accurate means of monitoring the
loads to which an aircraft is subjected during manoeuvres, although the positioning of the instrument
away from the C of G may give rise to discrepancies with respect to the readings on the fatigue meter.
The degree of under or over-reading on a cockpit accelerometer will depend on the extent of this
displacement, its damping characteristics, and the rate of entry into or recovery from a manoeuvre.
Accurate cockpit readings and lower stress loadings on airframes are more likely when manoeuvres
are entered and exited smoothly. Inadvertent overstressing must always be reported so that
structural examination for damage is carried out.
11. Although high 'g' loading can be very wasteful of fatigue life, it is important to remember that
because metal fatigue is a cumulative process, most fatigue life consumption occurs during routine
sorties, at moderate 'g' levels, when aircraft are operated within the parameters laid down in the
aircraft flight manual. Fatigue limiting practices should consequently be observed at all times. Weight,
for example, is a critical factor. If an aircraft's weight is increased by 1%, the consumption of its
fatigue life may go up by as much as 5%. Unnecessary weight should not be carried, and sorties
should be so organized that the greatest degree of manoeuvring takes place towards the end of the
flight, when fuel weight is down.
12. In summary therefore, there are a number of fatigue life conservation measures which should
be the day-to-day concern of pilots, flight planners, and authorizing officers. These include:
(a) Ensuring that aircraft are flown with serviceable fatigue meters (if available).
(b) Keeping tight turns and other high 'g' manoeuvres to a minimum consistent with the
sortie operational requirements, and never exceeding aircraft flight manual handling and
clearance limitations.
(c) Structure-friendly handling into and out of manoeuvres.
(d) Carrying only the fuel actually required for sorties.
(e) Avoiding turbulence whenever possible.

269 Structural Integrity and Fatigue
(f) Minimizing close formation flying.
(g) Flying only as low, and for as long as, the task requires.
(h) Minimizing roller landings.
(j) Keeping down taxiing speeds, particularly over rough surfaces.

271
ANNEX TO CHAPTER 4
THEORY OF METAL FATIGUE
Introduction
1. This annex covers the basic theory of fatigue in metals. The concepts explained here are
those that are applied to the design, testing and monitoring of aircraft structures by aircraft designers.
As with any other aspect of aircraft design, the application of basic theory to practical problems
involves much added complication, so a clear understanding of the essential characteristics of fatigue
is important.
2. Reasons for Fatigue in Airframes. Fatigue is inevitable in any metal structure where
alternating tensile stresses are repeated many times. In order to obtain a high performance from an
aircraft, the mass of the airframe, as well as all other components, must be kept as low as possible
and therefore the stresses in it will be high. This is indicated by the fact that the maximum service
loads in structural components are often as high as 2/3 of their respective failure loads. The stresses
in an airframe will alternate in a pattern determined predominantly by the aircraft flight manoeuvres,
atmospheric turbulence, ground loads, cabin pressurization and thermal effects. It is mainly tensile
stresses that cause fatigue hence the proneness of wing lower skins and lower spar booms, for
example, to fatigue failure. The problem of high stresses is aggravated by the stress concentrations
caused by, for instance, fastener holes. At these places the local stress is several times the average
stress in the material and hence it is here that fatigue cracks start.
Stages of Failure
3. The progression of fatigue cracks can be split into three stages: initiation, crack growth, and
final fracture. The second and third stages are normally evident on a fatigue fracture surface. There
are often marks which indicate the way in which the crack front has spread and these are called
beach marks, or progression marks. The concept of S/N curves is used to cover the total process of
fatigue whilst the recently-developed science of fracture mechanics can be used for the crack growth
and final fracture stages.
4. Crack Initiation. Fatigue cracks almost always start at the surface of components, although
exceptionally they will start at sub-surface flaws. These cracks tend to start at fastener holes, screw
threads, machining marks, sharp corners and other geometrical features. For a round hole in a sheet
of metal under tension, the local stress at the side of the hole is three times the average stress, (i.e.
stress concentration factor = 3), and of course this area is where the cracks start.
5. Crack Moderation. Surface conditions created by corrosion and fretting will also help to
initiate fatigue cracks, and residual tensile stresses from manufacturing processes speed up the
fatigue process too.
6. Crack Growth. Once a fatigue crack has propagated to a length of, say, 1mm then the rate
of growth depends less on the conditions that initiated the crack and will correlate more with the
average stress in the surrounding material. With a large stress concentration factor the initiation
period is shorter than with a small stress concentration factor, but once a crack is established it grows
at about the same rate in both cases. If metal is being fatigued in a very corrosive environment then
not only will the crack initiation period be shorter than otherwise, but the subsequent rate of crack
growth will be faster because corrosion speeds up crack growth. Also, corrosion may enable fatigue
cracks to grow at low stress levels which would normally not cause fatigue at all.
7. Final Fracture. Final fracture will normally occur in a brittle manner (ie. with no plastic
deformation) as the result of a fatigue crack. It will happen when the crack reaches the critical crack
length for the applied stress. The failure does not depend on the average stress in the cracked
section reaching the ultimate tensile stress for the material, but on the conditions at the crack tip.
Miner's Rule
8. The cumulative fatigue damage caused by

varying amplitude loading cycles is normally
estimated using miner's rule. The appropriate S/N
curve, discussed in the previous chapter, is used
for the estimation.
9. Plotted on the S/N curve are three

different alternating stress levels S1, S2, S3
which are n1, n2 , n3 times respectively. In a real
calculation there may be several more stress
Fig 4-2: Miner's Rule - Use of S-N Curve
levels, depending on the overall loading. Miner's
rule states that the structure will fail when:
n1 n n
N1
+ N 2 + N3 + ….. = 1
2 3
Where, N 1, N 2, N 3 etc are the number of cycles to failure at the applied stresses. In a simple
example, taking just three stresses if we get for each 1000 flying hours:
S1 occurring 10 times - hence n1 = 10

S2 occurring 2000 times - hence n2 = 2,000
S3 occurring 200,000 times - hence n3 = 200,000
273 Theory of Metal Fatigue
And from S-N curve:

N 1 = 10 2 cycles
N 2 = 10 5 cycles
N 3 = 10 7 cycles
Then proportion of fatigue life used is given by:
n1 + n2 + n3 10 + 2000 + 200,000 = 0.05

=
N1 N2 N3 1000 100,000 10,000,000
Since this damage occurred in 1000 hours flying, the total safe life in flying hours (assuming the S/N
curve has been factored to allow for 'scatter') is equal to:
1,000 divided by 0.05, which is 20,000 hours.
Critical Crack Length (CCL)
10. The critical crack length is the length at which a crack propagates catastrophically to failure.
As has already been implied, it depends on the material and the average applied tensile stress. The
overall size of a component has only a secondary effect and so, for example, if the CCL was 100 mm
for a 1 m wide skin panel with a certain applied stress, then it would also be about 100mm for a 3 m
wide panel of the same material and applied stress.
11. All other things being equal, the higher the average applied stress on a component, the
shorter will be its critical crack length. In general, aluminium alloys tend to have relatively long CCLs,
titanium alloys shorter CCLs and high strength steels the shortest of all.
275
CHAPTER 5
HYDRAULIC SYSTEMS
Introduction
1. Hydraulic systems provide a means of transmitting a force by the use of fluids. They are
concerned with the generation, modulation and control of pressure and flow of the fluid to provide a
convenient means of transmitting power for the operation of a wide range of aircraft services. A
typical aircraft hydraulic system will be used for operating flying controls, flaps, retractable
undercarriages and wheel brakes. Hydraulic systems can transmit high forces with rapid, accurate
response to control demands.
Definition of Terms
2. The following definitions need to be understood:
(a) Pressure. Pressure is the force per unit area exerted by a fluid on the surface of a
container. Pressure is measured in bars, Pascals (Pa), or Newtons per square metre (N/m2).
1 bar = 100,000 Pa = 100,000 N/m2.
(b) Force. The force exerted on a particular surface by a pressure is calculated from
the formula: Force = Pressure x Surface Area.
(c) Fluid. A fluid is a liquid or gas which changes its shape to conform to the vessel
that contains it.
(d) Hydraulic Fluid. Hydraulic fluid is an incompressible oil. In aircraft systems, low
flammability oils are used, the boiling and freezing points of which fall outside operating
parameters.
Transmission of Force and Motion by Fluids
3. When a force is applied to one end of a

column of confined fluid a pressure is generated
which is transmitted through the column equally in
every direction. Fig 5-1 illustrates a simple
hydraulic system. Two cylinders are connected by
a tube. The cylinders contain pistons of surface
areas 0.01 m2 and 0.0001 m2 giving a piston area Fig 5-1: A Simple Hydraulic System
ratio of 100 : 1. The system is filled with hydraulic fluid and fitted with a pressure gauge. A force of
1000 Newtons (N) is applied to the small piston. The force will produce a pressure (P) in the fluid
which is equal to Force (F)
Piston Area (A)
ie. P = 1000 = 10,000,000 Pa or 10 MPa, which is 100 bar.

0.0001
The system pressure will appear on the gauge and will be felt on every surface within the
system. Thus at the large piston a force (F) will be exerted where:
6
F = P X A = 10 X 10 X 0.01 = 100,000 N.
The force applied at the small piston is therefore increased on delivery by the large piston by
a factor of 100, i.e. in the same ratio as that of piston area. This is sometimes referred to as force
multiplication.
4. If the small piston is now moved down

in its cylinder through a distance of 100 mm as
illustrated in Fig 5-2 the large piston will move
upwards through a distance inversely
proportional to the piston area ratio, i.e.
through 1 mm. The work done by the small
force is transmitted hydraulically and equals
the work expended in moving the greater force Fig 5-2: Relative Movement in a
through a smaller distance. Simple Hydraulic System
ie. Force 1 X Distance 1 = Force 2 X Distance 2

or 1000 N X 100 mm = 100,000 N X 1mm
This is the principle of the hydraulic lever and is the operating principle of any hydraulic
system.
Advantages
5. Hydraulic power has unique characteristics which influence its selection to power aircraft
systems instead of electrics and pneumatics, the other available secondary power systems. The
advantages of hydraulic power are that:
(a) It is capable of transmitting very high forces.

(b) It has rapid and precise response to input signals.
(c) It has good power to weight ratio.
277 Hydraulic Systems
(d) It is simple and reliable.

(e) It is not affected by electro-magnetic interference.
Although it is less versatile than present generation electric / electronic systems, hydraulic
power is the normal power source used in aircraft for operation of those aircraft systems which require
large power inputs and precise and rapid movement. These include flying controls, flaps, retractable
undercarriages and wheel brakes.
Principles
6. Basic Power Transmission. The basic principle of hydraulic power is covered at para 3. A
simple practical application is shown in Fig 5-3 which depicts a closed system typical of that used to
operate light aircraft wheel brakes (like the HPT – 32). When the force on the master cylinder piston
is increased slightly by light operation of the brake pedals, the slave piston will extend until the brake
shoe contacts the brake drum. This restriction
will prevent further movement of the slave and
the master cylinder. However, any increase in
force on the master cylinder will increase
pressure in the fluid, and it will therefore
increase the braking force acting on the shoes.
When braking is complete, removal of the load
from the master cylinder will reduce hydraulic
pressure, and the brake shoe will retract under
spring tension. The system is limited both by
the relatively small driving force which in
practice can be applied to the master cylinder
and the small distance which it can be moved. Fig 5-3: Simple Closed Hydraulic System
7. Pump Powered Systems. These limitations can be overcome by the introduction of a

hydraulic pump. Fig 5-4 shows a simplified pump-powered system in which a control valve transmits
pressure from the pump to the hydraulic jack. A ‘down’ selection at the valve causes hydraulic
pressure to be fed to the ‘down’ side of the jack and the pump will work to maintain system pressure
during and after jack travel. Fluid from the un-pressurized side of the jack will be pushed through the
return part of the system circuit back into a reservoir. When ‘up’ is selected, hydraulic pressure is
removed from the jack ‘down’ side and applied to the ‘up’ side. The fluid displaced by the subsequent
retraction of the jack is returned to the reservoir. Within the strength and size limitations of its
components, the force transmitted by the system is now effectively limited only by the pressure which
the pump can generate while the distance over which the jack can expand or contract is limited by the
volume of the fluid.
Fig 5-4: Simplified Pump-Powered Hydraulic System
Types of Hydraulic Systems
8. There are two types of hydraulic systems:
(a) Constant Delivery Type. In this type the pump is always delivering fluid whether or
not a service is being operated. Cut-out valves and an accumulator are used.
(b) Constant Pressure Type (live-line System). In this application the pump
incorporates a pressure-operated mechanism which causes the amount of fluid delivered by
the pump to reduce when the system pressure approaches a set figure, until eventually the
delivery ceases altogether, the pressure being stored in the lines. When a service is selected
and the system pressure falls, the pump again starts to deliver fluid until the pressure is
restored. No accumulator is necessary in this system which is generally known as the live-line
system.
Typical System
9. To maintain the integrity and reliability of hydraulic systems which power ancillary services
fundamental to aircraft airworthiness, power sources for the primary flying controls are duplicated. A
typical arrangement is for one of the sources to be dedicated to the primary flying controls and the
other to a wider range of services. In transport aircraft, a third hydraulic source is sometimes provided
to operate those systems not essential to flight such as undercarriages, brakes and doors. The
provision of a fourth power source for emergency use, and the cross coupling between sources,
maintain power to essential services even in the event of two power sources failing. Terminology for
this arrangement of systems varies from aircraft type to type; however, the source dedicated solely to
powering the flight control units is usually termed the ‘Primary System’, whilst ‘Secondary System’ is
used to describe the system providing flight control back-up and powering other services. ‘Utility’ or
‘Auxiliary’ is applied to the third system whilst the fourth is known as the ‘Emergency’ or ‘Back-up
System’. A schematic diagram for a transport aircraft hydraulic system is shown at Fig 5-5. The
function of components typical to most systems is described in the following paragraphs.
Fig 5-5: Typical Hydraulic Power System
System Components
10. Pumps. The majority of engine or motor driven pumps are positive displacement, rotary
swash plate types, having up to 10 axial pistons and cylinders contained in a barrel which is splined to
the drive shaft. Each piston

terminates with a ball and slipper (or
shoe). The slipper bears against the
swash plate surface, the angle of
which determines displacement and
direction of the flow relative to
rotation. As the barrel rotates the
distance between the swash plate
and pump body increases and
decreases throughout each
revolution. The pistons are driven in
and out of the cylinders, drawing in
fluid at low pressure at the open
end of the stroke and expelling it at
high pressure at the closed end, as
shown in Fig 5-6. Two basic
variations of this type of pump are
commonly used in aircraft systems;
one produces a constant volume Fig 5-6: Principle of a Swash Plate Pump
output, relying upon other
components in the system to control both pressure and volume, whilst the other is self regulating,
automatically varying its output to meet system demands.
Fig 5-7: Constant Displacement Pump and Control System

11. Constant Displacement Pumps. Fig 5-7 shows a constant displacement pump and the
associated components needed to control system conditions. Constant displacement pumps absorb
constant driving power whatever the output demand. When the pressure in the system reaches an
upper limit, a cut-out valve allows fluid to bypass the pressure line and flow back to the reservoir.
Because large volumes of high pressure hydraulic fluid are therefore constantly being circulated,
greater attention must be paid in system design to cooling the fluid to maintain it within design
temperature limitations.
12. Self Regulating Pumps. Although self regulating pumps are more expensive, and cost
more to maintain, they allow simplification of the total system and they are therefore more usually
chosen for Primary and Secondary systems. Fig 5-8 shows such a pump. Its operation is similar to
that of the constant displacement pump, but the angle of the swash plate is variable and is changed
automatically during operation by a device sensitive to system pressure. As the swash plate angle
varies, so does the stroke of the pistons and the output of the pump. Thus, when system pressure
drops as power demands on the pump are increased, the output of the pump is increased to match
the new demand. When system pressure increases, as all demands are satisfied, the pump output is
reduced, and the pump absorbs less power.
Fig 5-8: Self Regulating Hydraulic Pump
13. Hand Pumps. Some aircraft are fitted with a hand operated, positive displacement, linear
pump for use on the ground. (eg. Jaguar) Its operation is usually restricted to pressurizing systems
sufficiently for opening and closing doors and canopies, and for lowering and raising ramps. The
aircraft Auxiliary Power Unit or a Ground Power Unit is used if more extensive use of the hydraulic
system must be made on the ground.
14. Accumulators. As illustrated in Fig 5-7

hydraulic systems include an accumulator, the
purpose of which is to absorb shocks and sudden
changes in system pressure. A typical nitrogen filled
hydraulic accumulator is shown in Fig 5-9.
Compressibility of the nitrogen allows the
accumulator to absorb and smooth out the pressure
ripples caused by pump operation and also the
sudden changes in pressure caused by operation of
components such as jacks and valves. It also acts to
maintain pressure, to the limit of its piston
movement, when the pump ceases to operate. This
facility is used, for example, to maintain aircraft
parking brake pressure for long periods. When the
hydraulic system is not pressurized by the pumps the
gas pressure is typically 70 Bar (1000 psi).
Fig 5-9: Typical Hydraulic Accumulator
15. Reservoir. Hydraulic systems

require a reservoir in which the fluid
displaced when the servo jacks are retracted
is stored until required again. Obviously, the
capacity must be designed to accommodate
fluid displaced when all the system jacks are
retracted simultaneously. The reservoir
performs the secondary functions of cooling
the fluid and allowing any air absorbed to
separate out. The construction of a typical
reservoir is shown at Fig 5-10. Reservoirs
are usually pressurized either with nitrogen
or by system hydraulic pressure acting on a
piston. This pressure, of between 3 and 7
Bar, prevents the fluid foaming and provides
a boost pressure at the pump inlet. A relief
Fig 5-10: Construction of a Reservoir
valve is fitted to prevent excessive pressure
build up due to heating or system malfunction. The body of the reservoir may contain horizontal
baffles both to prevent fluid surging during aircraft manoeuvre and to promote de-aeration of returning
fluid.
16. Heat Exchangers and Temperature Warning Systems. As described later in para 27,
hydraulic system performance is adversely affected by the presence of either air or vapour absorbed
in or mixed with the fluid, and additional heat exchangers are usually included in high performance
systems to keep the fluid well below its vaporization point. Such systems also include temperature
sensors and warning systems to alert the crew if excessive temperature excursions do occur. For
normal fluids, such warning systems are activated at temperatures of about 100°C.
17. Filtration. To prevent fluid leakage and loss of pressure, the clearances between the
moving parts of a hydraulic component are minute, and the inclusion of even the smallest particles in
the fluid would cause damage to its precise surfaces. High levels of filtration are therefore applied to
the fluid. Several filters are included in most systems, so that each major component can be
protected from debris generated upstream of it.
18. Pressure and Thermal Relief Valves. The use of a cut-out valve to regulate the output
pressure of a constant displacement hydraulic pump was discussed in para 11. Because hydraulic
fluid is incompressible and mechanical damage can be caused to components if over pressurization
occurs, further pressure relief valves are situated at critical points in the system. They are frequently
termed ‘fuses’ because of this protective role, and they operate by balancing system pressure against
an internal reference spring. If system pressure rises above spring pressure, the valve opens
allowing fluid to escape into the system return pipes thus reducing pressure. The valve re-seals
automatically once system pressure returns to below the reference level.
19. Non-return Valves. There are areas in most

hydraulic systems in which it is necessary to allow fluid to
flow to a component but to prevent that fluid returning
along the same pipe. Non-return valves are used for this
purpose, and several are included in the system in Fig 5-5.
Such valves are similar in construction to relief valves, and
the principle of operation is shown at Fig 5-11. The valve
poppet is held closed by a weak internal reference spring.
Pressure of fluid flowing in the desired direction can readily
overcome spring force, and fluid can therefore flow
through the valve almost without restriction. If fluid
Fig 5-11: Principle of a
pressure upstream of the valve is reduced, the poppet
Non-return Valve
snaps closed to prevent a fluid flow reversal.
20. Control Valves. Both rotary and linear action control valves are used in hydraulic systems,
and each type is shown diagrammatically in Fig 5-12 and 5-13. Valve movement may be achieved by
mechanical, hydraulic or electrical means depending upon the application. The valves are invariably
of an ‘on’ / ‘off’ rather than a variable throttle type.
Fig 5-12: Linear Control Valve Fig 5-13: Rotary Control Valve
21. Jacks and Motors. Jacks translate hydraulic fluid pressure into linear mechanical
movement, as in the example illustrated in Fig 5-4. Part rotary motion is often achieved by causing
the jack to drive a connected crank in an arc, however, full rotary motion is achieved by using a
hydraulic motor. This operates on the reverse principle of the swash plate pump shown in Fig 5-6.
Hydraulic pressure is fed sequentially to the pistons arranged around the motor body, and these react
against the swash plate forcing it to rotate.
22. Instrumentation and Control. Compared to electrical systems, the instrumentation and
control of hydraulic systems are very simple. Cockpit instrumentation monitors system pressure, and
the aircraft central warning system usually provides warning of system pressure failure and system
over-heating. The crew are able to select manually an alternative system if one fails, although this
reversion can be automatic by operation of cross-system control valves sensitive to system pressure.
Sight glasses and gauges are provided in most reservoirs and accumulators so that fluid levels and
nitrogen pressures can be checked on the ground, whilst remote gauging systems are installed in
cases where these components are not readily accessible.
System Safety Features
23. Hydraulic systems and their components reach very high statistical levels of reliability.
Nevertheless, both military and civil aircraft design standards require that aircraft hydraulically
powered primary flying control systems must have a back-up with the capacity to provide continued
control for an indefinite period after failure of the primary system. They also require that secondary
systems, such as undercarriages and brakes, have back-up with capacity to operate them for one
landing. The provision of alternative power sources, system redundancy and emergency power is
made to meet these requirements.
24. System Redundancy. Alternative sources may include provision for the powered flying
control units of a control system to revert to manual control, or for other hydraulic sources to be
connected to the failed power system. For this purpose, hydraulically powered primary control
systems are powered by at least two hydraulic systems. The power systems are configured to be
totally independent of each other so that the failure of one, for whatever reason, does not jeopardize
operation of the other.
25. Emergency Power. Assurance that system operation can be continued for indefinite
periods after failure of one hydraulic pump, requires that two other pumps sources are provided. One
is usually a pump driven from the aircraft normal secondary power system. The other may be a pump
powered by an emergency source such as an Emergency Power Unit or a Ram Air Turbine. For
systems requiring only a limited duration of operation under emergency power, such as wheel brakes
and undercarriages, the stored energy of accumulators or ‘blow down’ nitrogen cylinders situated in
the system is used.
Limiting Factors
26. Several factors influence the effectiveness of hydraulic systems, and some of these are
expanded upon below. The adverse influence of such factors is minimized by careful design and
maintenance of the systems and selection of the most appropriate fluids. There is no ideal solution in
these cases, and the chosen solution is invariably a compromise between performance and the other
factors.
27. Temperature and Aeration. As hydraulic fluid nears its boiling point, fluid vapour and
absorbed air are given off and carried in the fluid. The presence of gas from this or any other source
introduces an unacceptable degree of compressibility into the columns of fluid in the system, causing
operation to become sluggish and erratic. In high performance systems preventive design features
such as reservoirs to prompt and contain the separation of gases from the fluid, and the provision of
adequate cooling, are backed by careful system maintenance to minimize the likelihood of air entering
the system.
28. Contamination. As discussed in para 17, contamination of fluid with even minute particles
will damage and degrade systems performance. Again, careful systems replenishment avoids this
problem, and adequate system filtration ensures that particles introduced into or generated by the
system are removed before they can be carried through the system into components where they will
cause mechanical damage. Many hydraulic fluids are also hygroscopic to a small degree. Again
careful system replenishment and routine monitoring of the fluid will minimize the possibility of water
absorption.
29. Flammability. Certain hydraulic fluids are highly flammable, and leaks or spillage presents
a significant fire risk. Non-flammable fluids are used almost universally in the systems of passenger
carrying aircraft, despite their being highly corrosive.
30. Hazardous Liquids. All hydraulic fluids are active solvents and many are also corrosive.
They are therefore hazardous to both aircraft surfaces and materials and to human beings. Non-
flammable fluids are particularly hazardous. Careful handling during maintenance is necessary to
avoid this problem.
System Health Monitoring and Maintenance
31. The maintenance activities carried out on hydraulic systems include first aid action to
disclose, contain and rectify component failure, and fluid monitoring used to observe overall system
health trends and to detect component degradation.
32. Filter Checks. As shown in Fig 5-5, filters are strategically placed throughout an aircraft
hydraulic system. A component failure may not immediately manifest itself as a system malfunction,
but routine inspection of the filter tell-tale devices will reveal that a failure has occurred. The filter will
also prevent debris migrating around the system to cause secondary failures. Maintenance action
can then be taken to restore and safeguard system integrity.
33. Fluid Monitoring. A systematic sampling programme of fluid contamination is carried out
on the majority of aircraft. The periodic chemical and spectral analysis of fluids serves to indicate
failure trends in particular components and the contamination and degradation of system fluid. Based
on these trends, timely component replacement can be taken, thus preventing eventual failure
occurring in the air and reducing repair costs.
287
CHAPTER 6
PNEUMATIC SYSTEMS
Introduction
1. The use of air as a medium to transmit energy and to do work offers many advantages to the
aircraft designer. Although some early applications of pneumatics have been superseded by
hydraulics or electrics, as technological advancement has overcome the initial disadvantages of these
alternative media, the inherent and unique advantages offered by the use of air and its main
constituent gases ensure that pneumatics will remain one of these three fundamental power
transmission media for aviation use into the foreseeable future. Unlike hydraulics and electrics,
pneumatic power is generated and stored in a number of different ways each relevant to the specific
end use, and it is therefore not appropriate to consider pneumatic power generation as a specific
topic. Instead, the principle characteristics of the medium, and the techniques and equipment
configurations used to exploit those characteristics for specific applications, are discussed in the
following paragraphs.
Unique Characteristics
2. The ready availability of high temperature, high pressure air as a by product of the propulsion
system, or even of aircraft forward motion, provides an extremely cost effective source of heat or
pressure energy. Systems which utilize such energy sources include cabin and cockpit pressurization
and heating, airframe and engine de-icing and the augmentation of flying controls. Similarly, air can
be cycled through a system and exhausted overboard after use, without penalty. Such ‘total loss’
systems are extremely space and weight efficient, and this factor influences the choice of air above
other energy transmission media which usually require to be contained in a closed circuit system for
technical or environmental reasons. Such ‘total loss’ air systems include engine starting and cabin
and equipment conditioning. Again, although air will support combustion, its properties are not
affected by temperature extremes, and it can therefore be used in power transmission applications
where high temperatures, fire risks or chemical reaction rule out the use of normal hydraulic fluids.
Pneumatic systems are therefore often used in engine nozzle and thrust reverser operating systems.
3. The ready compressibility of air offers both advantages and disadvantages for its use. The
advantages are that air can be compressed and used to store the resultant pressure energy either
long term for subsequent use or short term to absorb shocks or sudden changes in pressure levels.
However, because of this same compressibility, pneumatics are not suitable for use in control
systems requiring precise, rapid response movements.
Basic Generating Systems
4. The basic generating system as shown at Fig 6-1, consists of the following components:
(a) Compressor. It is normally a single-cylinder air-cooled compressor driven by the

engine or gear-box of the aircraft.
(b) Relief Valve. This is normally mounted close to the compressor and warmed by air
as a safeguard against low temperature conditions causing freezing of the other valves in the
system.
Fig 6-1: Basic Generating System
(c) Ground Charging Valve. This permits the storage cylinders to be charged from
ground supply sources when the aircraft compressor is not running.
(d) Anti-Freezer. This may be fitted in certain aircraft. It consists of methanol in a

container. The methanol is vaporized by the warm air from the compressor and circulates
through the rest of the system, lowering the freezing point of any moisture and permitting
satisfactory operation in very low temperatures.
(e) Oil and Water Trap. Moisture, which is always present in the atmosphere, is
precipitated in the oil and water trap, which is mounted at the lowest point in the generating
system. A drain valve is fitted at the bottom of the oil and water trap so that the collected fluid
may be drained away.
(f) Pressure Regulating Valve. As the compressor is generating air pressure all the
time the engine is running, when the storage cylinder is fully charged, the regulating valve will
by-pass the compressor output to atmosphere. A pressure-sensitive bellows connected to
the air bottle pressure operates the valve and an independent spring-loaded safety valve is
fitted. When the engine is stopped the pressure is prevented from leaking back into the
compressor by a non-return valve on the outlet side of the regulating valve.
289 Pneumatic Systems
(g) Storage Cylinder(s). The cylinder(s) act as a reservoir to store compressed air,
giving a reserve of power for short bursts of heavy service operation, or for emergency use in
the event of engine or compressor failure.
Typical Applications
5. The applications of pneumatics can be categorized under four main headings. These are:
(a) Pressure energy storage.

(b) Compression.
(c) Pressure energy transfer.
(d) Heat energy transfer.
Pressure Energy Storage
6. Undercarriage Blow-down Systems. Because hydraulic fluids cannot normally be

compressed, energy cannot be stored within simple hydraulic systems. However, this disadvantage
can be overcome by the integration of pneumatics into hydraulic systems. An application of such
Fig 6-2: Simplified Undercarriage Blow-down System

hydro-pneumatics is the undercarriage emergency blow-down system. A schematic diagram of such

a system is at Fig 6-2. In this particular example, release of high pressure air (nitrogen is normally
used to reduce the risk of a hydraulic oil fire) from the blow-down bottle enables the undercarriage
lowering system to be pressurized sufficiently to lower the undercarriage in the event of hydraulic
malfunction or failure. Other similar systems feed nitrogen directly into the hydraulic fluid. Although
such an arrangement avoids the disadvantage of pressurizing only a set volume of fluid (that
displaced within the accumulator by the gas), it does impose a considerable penalty during
subsequent fault rectification, because the gas is absorbed into the fluid necessitating replacement of
all of the fluid within the hydraulic system.
7. Fire Extinguishers and Liferafts. There are many other applications in which compressed
gases are stored for eventual use as an emergency or occasional energy source. Amongst the most
relevant are the use of nitrogen to pressurize engine fire extinguisher bottles for eventual use in
propelling extinguishant on to a fire, and the use of carbon dioxide stored with liferafts and life jackets
for subsequent release to inflate these items when they are required.
Compression
8. Shock Absorbers. Hydraulic systems are frequently configured to use the compressibility
of air to absorb shocks and sudden changes in system pressure. The system shown in Fig 6-2
includes a nitrogen filled hydraulic accumulator. The functions of the accumulator are to smooth out
any sudden changes in systems pressure caused by operation of components such as jacks and to
protect the system from sudden peaks in pressure which occur when system valves close. The graph
at Fig 6-3 shows the typical pressure variation in a system without an accumulator, whilst Fig 6-4
shows the comparable variation when an accumulator is used. Hydro-pneumatic shock absorbers,
based on a similar principle, are widely used in many undercarriages.
Fig 6-3: System Pressure Fluctuations Fig 6-4: System Pressure Fluctuations
Without Accumulator With Accumulator
291 Pneumatic Systems
9. Seal Inflation. The doors and canopies of pressurized aircraft require to be sealed
effectively, to maintain pressurization within the fuselage and to prevent the escape of unacceptable
volumes of conditioning air. The sealing of the irregular gaps between such doors and hatches and
their frames imposes a significant problem, and seals inflated by compressed air are often used in
such situations. The omni-directional force applied to such seals by low pressure air is ideal for such
applications, and the air can readily be tapped from the aircraft pressurization system.
Pressure Energy Transfer
10. Augmented (Blown) Lift Devices and

Flying Controls. Within the restrictions of
current aerodynamic knowledge and technology,
all VSTOL aircraft must be provided with devices
which impart energy to the surrounding air
stream to provide lift and control forces in the
absence of adequate forward air speed.
Purpose-designed VSTOL aircraft usually use
vectored lift/thrust systems which also provide
flight control forces. However, the use of
conventional aerodynamic devices enhanced for
STOL operation by ducting high energy air
streams over them is an effective alternative
solution. Fig 6-5 shows such a system in
Fig 6-5: Augmented Lift Devices
schematic form.
11. Starters. The abundant availability of high pressure air from gas turbine engines and APUs
allows its use for engine starting. This is achieved either by impinging upon the turbine directly, and
thus spinning up the engine, or more usually by driving a small turbine which is connected to the main
engine through suitable gearing
Heat Energy Transfer
12. Air Conditioning and Ice Protection. The compressors of most high performance gas
turbine engines are designed to produce volumes of air in excess of engine requirements. Such air at
high pressure and at temperatures up to 300°C is available through engine compressor bleeds, and,
as well as being used in cabin and cockpit pressurization systems, the air provides an effective source
of heat for air conditioning and for the ice protection of aerofoils and engine intakes.
293
CHAPTER 7
POWERED FLYING CONTROLS
Introduction
1. The level of aerodynamic forces needed to control the attitude of an aircraft is proportional to
the inherent stability of that aircraft and to the square of its speed. Thus, whilst the forces required to
control a low speed, well balanced aircraft may well be within the physical capabilities of the pilot,
those for a high speed or high stability aircraft will certainly not be.
2. The earliest power controls were of the power assisted type in which only a certain proportion
of the aerodynamic loads were fed back to the pilot, the hydraulic system overcoming most of the
load. In these systems, although the pilot did not have to provide all the force required, he retained the
natural feel of the controls, and the stick forces experienced increased as the square of the forward
speed. These power-assisted systems were eventually replaced by fully powered control systems, in
which none of the aerodynamic load on the control surfaces was fed back to the pilot. The only force
that the pilot need exert in the mechanical system, is the force required to overcome the friction etc.
all the necessary power being supplied by the aircraft's hydraulic or electrical system. However, a
pilot's assessment of the handling characteristics of an aircraft is very dependent on the control forces
that he has to apply, and thus it is necessary to incorporate "artificial" feel. This can be done in
several different ways, depending on the particular control force characteristics that are required. The
control system of an aircraft must be safe and reliable, and therefore a power control system must be
duplicated or capable of reversion to manual operation. Figs 7-1 to 7-4 shows four levels of control
from a fully manual system through aerodynamically assisted (servo-tab) and a power assisted
system, to a fully powered control system. Inputs are shown thus: Manual ,, Power .
Fig 7-1: Simple Manual Control Fig 7-2: Servo-Tab Assisted Control
Fig 7-3: Power Assisted Control Fig 7-4: Full Powered Control
3. Additional Features. The introduction of many advanced flight control concepts and
systems have been made possible by the use of powered flying controls in aircraft. Whilst such
features as stall warning (stick shakers) and stall prevention (stick pusher) devices can be integrated
into manual systems, they are more effectively installed in aircraft which are fitted with powered
systems. The use of such systems as auto-pilot, auto-land, fly-by-wire and fly-by-light, and
application of active control technology to neutrally stable fixed and rotary wing aircraft is totally
dependent upon the use of powered flying control systems.
Basic Requirements of Powered Flying Control Systems
4. Performance. A powered flying control system must perform to produce satisfactory

handling characteristics throughout the aircraft flight envelope. The system must therefore have the
appropriate power and range of movement needed to perform that task, whilst also being designed to
achieve a good power to weight performance. Powered control systems occasionally have
inadequate power for the task in hand, and this causes a condition known as “jack stall” in which the
servo jack cannot overcome the aerodynamic forces acting against it. The probability of such a
condition occurring on particular aircraft types is well known, and identification of the areas of their
flight envelopes where the phenomenon is likely to occur are clearly documented in the relevant
Aircrew Manuals.
5. Feedback. An essential feature of all powered flying control systems is that of a feed-back
loop capable of comparing the response of the system to that demanded by the pilot. Feedback to
the pilot in the manual and servo tab systems is accomplished automatically, because there is a
direct, fixed linkage between the pilot and the control surface. As the pilot moves his control, the
corresponding control surface moves by a similar amount. The pilot then can use visual or instrument
references to check that the aircraft has responded in the required manner to his control input. Thus
a complete feed-back loop is established, and Fig 7-5 shows such a loop in diagrammatic form.
Power assisted and fully powered systems require a similar feed-back loop. This is usually achieved
295 Powered Flying Controls
by a mechanical linkage which causes the powered flying control unit to drive until it reaches a
position relative to the pilot's input signal. Fig 7-6 highlights the feed-back linkage used in the simple
powered unit shown in Fig 7-4. In addition to
positional feedback, a pilot also requires to
receive a degree of feedback of flight forces.
Such forces are essential to provide the pilot
with tactile cues of the performance of his
aircraft during flight. In a manually controlled
aircraft, stick forces increasing as a square of air
speed give essential references to the pilot.
Such references are not fed back to the pilot
through a powered flying control, and methods
Fig 7-5: Positional and Rate Feedback Loop
of synthesizing feel are therefore incorporated.
Fig 7-6: Mechanical Positional Feedback
6. Accuracy. The powered flying control system must respond accurately to both the
amplitude and the rate of control input under all conditions of flight. Otherwise, the aircraft may be
endangered either by divergent oscillations being set up through the pilot over-compensating for
system inaccuracies, or by overstressing caused through too rapid a rate of response. The accuracy
of response is partly inherent in the power source used in the system, partly in the effectiveness of
feedback in the system and partly by the precision with which components of the system are
manufactured and installed.
7. Stability. Not only must the system respond accurately to the control input, but it must also
hold the control position and not deviate through spurious inputs caused by system errors. The
stability of a system is largely ensured by initially designing sufficient tolerance into the components,
although subsequent component maintenance of the highest order is required to ensure adequate
margins of continued stability. Deterioration in the condition of both mechanical and electrical
components and the inclusion of air in hydraulic systems are typical causes of degraded stability.
8. Irreversibility. The main objective of using powered flying controls is to off-load

aerodynamic forces from the pilot's flying controls. Similarly, it is essential that the effects of buffeting,
flutter and turbulence are also off-loaded. The inherent irreversibility of hydraulic and electrical
powered flying control units automatically ensures that this is accomplished. The likelihood of control
surface fluctuation must of course still be minimized by good design plus aerodynamic and dynamic
control balancing.
9. Safety and Reliability. Obviously, the reliability of its powered flying controls is paramount
to the safety of an aircraft. To provide the necessary real and statistical degree of reliability of such
controls, they are normally duplicated. In aircraft where flight loads would be within the physical
capability of the pilot, reversion to manual control in the event of system failure may be permissible.
Fig 7-7: Essential Features of a Powered Flying Control System
Typical Installation
10. Fig 7-7 shows the essential features of a typical powered flying control installation in
schematic form. Descriptions of its main components are included in the following paragraphs. The
system is based upon that used for the longitudinal control of a medium sized aircraft. It features an
autopilot pitch channel servo and an all-flying tail-plane used for trimming out the pitching moment
caused by use of the large span flaps often fitted to such aircraft. Elevator trim is provided
conventionally by an aerodynamic trim tab, although both this component and alternative integrated
trim devices are discussed in the following paragraphs. Lastly, the system utilizes conventional
cables and push / pull rods to transmit commands between the pilot's controls and the servo units. In
aircraft equipped for fly-by-wire or fly-by-light control systems, these items would be replaced by
electrical or fibre optic cables.
System Components
11. Powered Flying Control Unit. The basic features and operation of a hydraulic powered
flying control unit are shown at Fig 7-8 and 7-9. The unit is shown at rest in Fig 7-8 and in mid-travel
at Fig 7-9. Movement of the servo valve away from its mid-position occurs when the pilot moves his
controls, the servo valve allows high pressure fluid to enter and act upon the appropriate chamber of
the unit. The main piston remains stationary, and the whole body of the unit moves under the fluid
pressure, and its movement is transferred to the control surface. As the surface reaches its desired
position, the movement of the body in relationship to the stationary servo valve restores the valve to
its central position. The flow of hydraulic oil then ceases, and the unit is locked in its new position by
incompressible fluid trapped on both sides of the piston. This situation remains until a further control
signal from either the pilot or the autopilot causes the cycle to be repeated. By fixing the piston to the
aircraft structure and the unit body to the control surface, an automatic positional feedback is
achieved. If the roles of the two components were reversed, an additional linkage would be needed to
act as a feedback. Otherwise, the piston would travel to its extreme position whenever the servo
valve was moved.
Fig 7-8: Hydraulic Powered Flying Control Fig 7-9: Hydraulic Powered Flying Control
Unit at Rest Unit Moving
Artificial Feel Systems
12. The assessment of the handling characteristics of an aircraft is dependent on the magnitude
of the control movements and the control forces that the pilot experiences during flight. In an aircraft
equipped with non-powered controls, the stick forces increase in direct proportion to the square of the
forward speed, and this characteristic forms a major part of the 'feel' of an aircraft, providing valuable
tactile cues to the pilot on his aircraft's performance. In a perfectly irreversible fully powered-control
system, however, no such cues of aerodynamic loading are received by the pilot from the control
surfaces and, in order to avoid overstressing the aircraft and to provide a datum from which control
demands can be made, a system of artificial feel is normally built into the power control system.
Typical feel devices are described in subsequent paragraphs. The essential features of an artificial
feel system are as follows:
(a) Forces should increase as stick displacement is increased.
(b) The forces should be proportional to airspeed but ideally should reduce at high
subsonic speeds where the effects of turbulence are to reduce control effectiveness.
(c) To prevent overstressing in the longitudinal plane, feel forces proportional to ‘g’ forces
should be applied to the longitudinal controls.
(d) There should be a warning of the onset of a stall.
13. Spring Feel. The simplest type of feel system is one incorporating a spring. Movement of
the controls compresses or extends the spring and provides the pilot with a control force which is
dependent on the strength of the spring and the control deflection. The force is, however, independent
of the aircraft's forward speed, and this can result in heavy handling at low speeds and the ability to
overstress the aircraft at high speeds. For this reason, spring feel is normally confined to ailerons, or
used to supplement 'q' feel. The force characteristics of the simple spring system can be improved by
incorporating variable feel and a gear change mechanism, but this is at the expense of added
complications. In some installations, the spring is replaced by a torsion bar.
14. 'q' Feel. The 'q' feel system varies the force on the cockpit controls in direct proportion to
2
the aerodynamic loads on the control surfaces, i.e. proportional to ½ ρV or q. The system can be
supplied with either actual or simulated dynamic pressure as shown below:
(a) Simple ‘q’ Feel. In its simplest

form, the ‘q’ feel system links the cockpit
control to a piston which moves within a
cylinder (sometimes referred to as a ‘q’-
pot). The cylinder on one side of the
piston is supplied with pitot pressure, and
the other side with static pressure as
shown in Fig 7-10. By suitable Fig 7-10: Simple ‘q’ Feel System
arrangement of the linkages, movement of
the cockpit control is always opposed by dynamic pressure. Although simple, this system
requires a large cylinder (10-20 inch diameter).
(b) Hydraulic 'q' Feel. The bulk of the simple 'q' feel system can be overcome by
feeding the pitot and static pressures to either side of a diaphragm attached to a hydraulic
servo-valve (Fig 7-11).
The servo-valve provides
a metered hydraulic
pressure which is an
amplified value of the
dynamic pressure. This
hydraulic pressure is fed
to a small jack which
opposes movement of the
cockpit control. Fig 7-11: Hydraulic ‘q’ Feel
(c) Mach No Corrected Hydraulic 'q' Feel. At subsonic speeds, the hydraulic 'q' feel
virtually achieves a constant value of stick force per g. However, as high supersonic speeds
are approached, the
control surfaces lose their
effectiveness for a given
deflection due to the
effects of compressibility.
The Mach No. corrected
feel system prevents the
increase of stick force
with speed above a pre-
determined Mach No. by
Fig 7-12: Mach No. Corrected ‘q’ Feel
incorporating a Mach
capsule into the basic hydraulic feel unit described above. Fig 7-12 shows a schematic
arrangement of a Mach No. corrected feel system.
(d) Electrical' q' Feel. In the electrical 'q' feel system, airspeed information (i.e. pitot /
static) is used to vary the resistance of one arm of a Wheatstone bridge circuit. Any
imbalance in the bridge circuit drives an actuator, the position of which determines the degree
of extension or compression of a spring feel unit. Movement of the actuator restores the
balance of the bridge circuit.
15. V and V3 Feel. The 'q' feel system can be modified by the incorporation of a mechanical
system, i.e. various arrangements of non-linear springs, cams etc., so that the control loads
experienced by the pitot increase at a slower or faster rate than V2. The system can be designed to
produce loads approximately proportional to variations in V or V3. V3 feel is rarely used, but it has
been employed on the rudder circuits of some aircraft in order to prevent the application of large
control deflection, with corresponding large side-slip angles at high speeds.
16. g Feel. Longitudinal control forces can be modified by incorporating a bob-weight into the
control system. This produces an additional stick force which is independent of the forward speed of
the aircraft, but proportional to the amount of g applied. This increases the value of stick force per ‘g’,
and prevents the application of excessive amounts of ‘g’ during manoeuvres. Bob-weight systems are
often referred to as ‘g’-restricted or response-feel systems. Bob-weight systems suffer from inertia.
This means that a very rapid
application of excessive
control deflection can be made
before the aircraft, and
therefore the bob-weight, can
respond. A further effect of
this inertia is the undesirable
response of the bob-weight to
turbulence. To overcome
these problems ‘g’ feel is often
combined with a spring feel
system. Fig 7-13 shows such
an arrangement.
Fig 7-13: g Feel Unit
17. Additional Control Inputs. The use of powered flying control systems allow the integration
of additional control inputs such as trim adjustment, stall warning and automatic flight control.
Automatic flight control systems are discussed in a separate chapter.
18. Trim Adjustment. Manual control systems utilize fine adjustment of the main or ancillary
control surfaces to trim the aircraft flight attitude to compensate for centre of gravity displacement or
attitude variation at particular speeds or flap
settings. The trim systems allow the main
cockpit control forces to be minimized and the
control positions to be centralized. Although
some applications of powered flying control
systems retain the use of ancillary control
surfaces for trimming, many systems trim the
aircraft by small adjustments of the main control
system. However, adjustments to reduce feel
forces on the cockpit controls and to centralize
Fig 7-14: Feel Trim
them remain necessary if the pilot is to retain
references and cues during flight. A typical feel trim system is shown at Fig 7-14. Its principle of
operation is to zero the synthesized feel forces when the aircraft is correctly trimmed. A typical
position (datum) trim system is at Fig 7-15. Its principle is to apply trim adjustments to the aircraft
controls without the pilot's controls being moved away
from their neutral position.
19. Non-linear Response. Most manual control

systems contain a degree of non-linearity in their
operation. Small movements of the cockpit controls
when near to the extremes of their travel produce
larger control surface movements than will occur at the
centre of their travel. This is a desirable situation, Fig 7-15: Datum Shift Trim
allowing small precise control movements to be made in the critical centre of the control span and
large coarser movements to be made at the extremes. The use of powered flying controls offer the
opportunity to increase this non-linearity and thus increase the effectiveness of the control system.
The use of non-linear levers and cams, not acceptable in a manual system because of the variation in
control loads which they would impose, are permissible in powered systems to which variation in
control forces are less significant. Typical non-linear control systems are shown at Fig 7-16 and Fig
Fig 7-16: Differential Aileron Control Fig 7-17: Differential Effect
7-17.
20. Stall Warning and Prevention. The dangers inherent in stalling a high performance aircraft
have led to stall warning systems being fitted to most relevant aircraft. In their simplest form, they
consist of an electrical device which shakes the control column so that the pilot experiences cues
similar to a stall buffet. He is thus alerted to take the necessary corrective action. However, this is
not adequate for many transport aircraft, particularly those with a high tailplane configuration, and stall
prevention systems are often fitted to these aircraft. Typical systems such as that in Fig 7-7 include a
pneumatic jack which gradually imparts a nose down control input to the aircraft. The pilot can then
either accept and supplement the input or consciously over-ride it and take alternative measures to
avoid the stall.
21. Manual Reversion. As mentioned

above, reversion to manual control, in the event
of failure of a powered flying control system, is
an acceptable option for lightly loaded aircraft.
The system of reversion is usually a very simple
one and is often engineered to occur
automatically in the event of a hydraulic failure.
Fig 7-18 shows a typical system.
22. Multiplication. A more normal

arrangement for retaining adequate control in the
event of failure of a powered flying control
Fig 7-18: Hydraulic System with
system is achieved by multiplication of critical
Manual Reversion
components of that system. The most usual
methods of achieving this redundancy of critical components are shown at Fig 7-19 and Fig 7-20.
These include (Fig 7-19) split control surfaces each section having an independent powered control
unit, and (Fig 7-20) use of tandem jacks. All of these systems require an interlock within their
respective hydraulic jack servo valves to prevent a hydraulic lock occurring in the redundant jack and
thus locking the relevant control surface. The use of parallel jacks is widespread in aircraft in which
there is room to site such units adjacent to the control surfaces. However, for aircraft in which this is
not possible, particularly combat aircraft and helicopters, the tandem jack arrangement is normally
fitted.
Fig 7-19: Parallel Units Fig 7-20: Tandem Units

303
CHAPTER 8
AUTOMATIC FLIGHT CONTROL SYSTEMS
Introduction
1. The Problem. Since the first days of flight, the need to compromise between aircraft
performance and controllability of the aircraft has formed a central factor influencing specification and
design. The finite workload which a pilot can contend with, the speed of response necessary to
counter critical adverse changes in aircraft attitude and the span of concentration required to navigate
and fly over long periods, all had a considerable influence upon the design and configuration of
aircraft. Major compromises in performance had to be therefore accepted in order to retain
acceptable levels of controllability.
2. The Solution. The problems posed by workload, speed of reaction and fatigue have been
gradually solved by the development and subsequent evolution of automatic control systems. Such
systems augment the control applied directly by a pilot, whilst ensuring that he retains full command of
his aircraft. The systems initially comprised only automatic stabilization devices to counter gross
changes in aircraft trim. These were followed by automatic pilots which ensured stability in three axes
and that a selected heading and altitude was maintained. More recently, advances in technology
have made possible the combination of such individual systems, and this has led to the introduction of
fully integrated automatic flight control systems (AFCS). This step has in turn influenced the design
and configuration of aircraft, and it has allowed advances to be made towards achieving maximum
theoretical performance.
Functions Performed by an AFCS
3. Basic Mode of Operation. An AFCS or part-system operates on the following basic

principles:
(a) Compare actual response of an aircraft with that demanded by the pilot.
(b) Process the error between actual and required performance in order to generate a
correcting control command.
(c) Signal that command to the relevant aircraft control components.
(d) Implement that command by moving the relevant control surfaces.
(e) Monitor compliance with the original command by feeding back the actual effect of
the control input to the comparator.
These processes are represented in diagrammatic form at Fig 8-1. The following paragraphs
below describe the facilities which AFCSs or part-systems provide, whilst a more detailed description
of the components of an AFCS and the methods by which these function are included in subsequent
paragraphs.
Fig 8-1: Basic System Operation
4. Stabilization. The automatic stabilization of an aircraft in roll, pitch and yaw is a basic
function of all AFCSs. The simple part-systems, such as auto-stabilizers and stabilization
augmentation systems (SAS), act to stabilize the aircraft and maintain it in an attitude initially set up
by the pilot by applying up to three axes of control, however, they do not usually include the facility to
implement changes in attitude. Single or dual axis auto-stabilizers are installed in most aircraft which
Fig 8-2: Basic of a Single Axis Auto-stabilizer
have insufficient natural stability. In VSTOL aircraft for example the problems of maintaining stable
flight at low forward speeds is encountered and in helicopters the low values of longitudinal stability,
manoeuvre stability, and marked changes in dynamic stability that occur at different airspeeds have to
be over come. The basics of a typical single axis auto-stabilizer are shown in Fig 8-2.
305 Automatic Flight Control Systems
5. Altitude and Heading Control. All full AFCSs are capable of achieving altitude and
directional command and control of the aircraft. However, the simpler part-system auto-pilots include
only a facility to maintain height and heading initially set up by the pilot. In many such systems,
changes in both height and heading can be demanded by the pilot, and these systems are capable of
controlling aircraft attitude in order to achieve the commanded changes, for instance by flying the
aircraft through a controlled rate of turn. Such systems are simple, largely self-contained and
inexpensive. They therefore provide an extremely cost effective method of reducing pilot workload by
the augmentation of control during stable periods of flight. Fig 8-3 shows diagrammatically a three
axis auto-pilot with heading and altitude hold.
Fig 8-3: Simple Three Axis Auto-pilot
6. Automatic Throttle Control. In the long term, automatic maintenance or variation in

altitude requires control of engine power, and the auto-throttle facility provides such control. An auto-
throttle gives not only altitude control but also provides for automatic landings and speed control for
maximum endurance or other cruise conditions. Fig 8-4 shows an auto-throttle system in
diagrammatical form. The system monitors airspeed and pitch rate against datum parameters set
either by the pilot or as a product of an associated auto-land system.
Fig 8-4: Auto-Throttle System Diagram
7. Automatic Landing. The auto-land capability utilizes an ability to process the signals
received from external ILS and MLS facilities. As with the basic system depicted in Fig 8-1, this
system compares the actual aircraft landing profile detected from on-board sensors and ILS / MLS
signals with a programmed profile and makes appropriate corrections in attitude, direction and engine
power settings.
8. Automatic Compliance with a Defined Flight Profile. Micro-processors and associated

memory storage devices provide the capability for an AFCS to be programmed with details of required
flight profiles far in excess of those which can be accomplished within the sortie endurance of a
particular aircraft. Thus, by integrating the sub-routines of auto-land, auto-cruise, auto-pilot and auto-
stabilization as necessary, fully automatic flight control can be achieved. Much of the routine work
load of mission profiles can be automated, thus reducing the overall workload and releasing crew to
perform the more important non-routine work of combat or of transport missions. Similarly, the facility
for auto-hover in SAR and ASW helicopters greatly increases mission effectiveness.
Influence of AFCS on Aircraft Performance and Configuration
9. The above description of AFCS functions assumes that such systems are fitted to
conventional aircraft in order to improve handling or operational effectiveness. However, the full
integration of AFCS technology into ‘purpose designed aircraft’ enables many of the design
compromises previously necessary in aircraft performance to be avoided. Thus, use of an AFCS
allows the building and operation of much higher performance aircraft in which the AFCS performs the
core function of aircraft control, at the direction of the crew.
10. The size, and hence the structural weight, of the control surfaces fitted to conventional,
inherently stable, aircraft is dictated by the need to achieve maneuverability. Inherent stability in an
aircraft results in a balance between lift forces and aircraft weight such that tail plane forces act
downwards. This reduction of lift requires the wing to be larger, or at a greater angle of attack, which
leads to reduced aerodynamic performance. The use of fly-by-wire and fly-by-light systems and
Active Control Technology (ACT) enables the size of the tail plane balancing force to be reduced by
allowing the aircraft CG and CP to be placed closer together. Sensors and computer processors
perform a delicate balancing act between the moments generated by the wing lift and tail lift to give
the pilot an artificially stable aircraft with excellent maneuverability.
11. The full advantages of using ACT can only be realized by designing aircraft that utilize ACT to
establish specified control objectives. Such aircraft may be inherently unstable or have canard control
surfaces, such as EFA, and other advantageous aerodynamic and control features. Aircraft designed
this way are termed Control Configured Vehicles (CCV). One of the aims of CCV is to produce an
aircraft that achieves its specified performance with minimal control surface weight and drag. Once
the step of utilizing ACT to produce CCVs has been taken the aircraft’s control system must be
designed such that there is minimal likelihood of failure since the pilot may well be unable to control
the aircraft himself without the assistance of AFCS computers.
Fly-by-Wire and Fly-by-Light Systems
12. Definitions and Principles. The term ‘fly-by-wire’ was first coined to describe the control of
an aircraft by the pilot through electrical signals generated by movements of the pilot’s controls. Such
systems were introduced purely to obtain the advantages of electrical signalling over mechanical
signalling systems. Their introduction gave the opportunity to more easily integrate pilot control inputs
with other auto-pilot functions, and the terms fly-by-wire (FBW) and fly-by-light (FBL) are now used to
denote systems in which electrical signals generated by pilot control inputs are integrated with sensor
signals and consequently modified before being fed to the control surfaces. Thus, no direct or linear
proportional link remains between pilot and control surfaces. The severing of such a direct link allows
the AFCS to optimize aircraft performance in all flight conditions, leaving the pilot free to concentrate
upon the mission in hand. The following paragraphs examine particular types of control input
modification which can be realized by use of FBW or FBL systems.
13. Manoeuvre Demand Control Systems. In the systems described so far, control of attitude
is a basic function. However, in most high performance aircraft, particularly combat types, avoidance
of manoeuvre which will stall or over-stress the aircraft is a more major objective. Therefore, in
aircraft fitted with such systems sensors measuring acceleration and the rate of change of
acceleration are used to provide processor inputs, and the values of these parameters are compared
to limiting values permissible in the relevant attitude. Control correction signals therefore reflect
permissible manoeuvre rate limitations rather than purely the permissible attitudes.
AFCS Components
14. Overall System Requirements. Although the configuration of AFCSs will vary from aircraft
to aircraft, they all consist of the same basic modules. Fig 8-5 shows these modules in diagrammatic
form, and a description of the modules follows. Part-system devices, such as auto-pilots and auto-
stabilizers, consist of similar modules, although obviously the pilot performs many more of the
functions than he does in a full system. As an integral part of the aircraft flight control system, the
overall requirements of an AFCS are:
(a) Reliability. Reliability is ensured by the multiplexing (normally triplexing but

quadruplexing for FBW) of critical components and systems.
(b) Fail Safe. Where possible fail safe is ensured by careful design and attention to
failure mode analysis. This is often guaranteed by the system self checking the outputs of its
three components and disregarding one at wide variance with the other two.
Fig 8-5: Basic Modules of an AFCS

(c) Accuracy of Response. Response accuracy is achieved by careful design of the

system, and the inclusion of response parameter limits within its processing module.
(d) Output Stability. Output stability is also achieved by careful design of the system
and selection of its feedback characteristics. The use of adaptive feedback characteristics is
a normal method of varying system response in differing flight conditions. However, the
overall requirement for system stability limits the extent to which feedback can be varied.
15. External Inputs. The external control inputs to an AFCS originate from three sources:
(a) The initial flight profile demanded by the pilot. This input is often not relevant to part-
systems.
(b) Changes to attitude, course and altitude needed for operational or air traffic reasons
and interpreted and fed into the system by the pilot.
(c) Basic navigational information fed directly into the AFCS from ILS/MLS, VOR, GPS,
Inertial Navigation and Flight Director systems. Such information is admitted to the AFCS to
support the initial flight profile demanded by the pilot.
16. Sensors. To evaluate the difference between the demanded and achieved performance,
the AFCS processing unit requires datum information for all relevant parameters. This data is
obtained both from sensor inputs and from standard model parameter profiles. Model parameter
information is usually stored within the AFCS processor, but other datum information is provided
either as inputs from sensors provided specifically for this purpose or, more usually, as outputs from
similar sensors forming part of other discrete systems but coupled to the AFCS. Parameters for which
sensor data is required include:
(a) Pitch, roll and yaw rates using rate gyros.
(b) Pitch, roll and yaw attitude using vertical or heading gyros.
(c) IAS using an electrical pick off from the ASI.
(d) Sideslip using a lateral accelerometer or differential static pressure transducer.
(e) “g” measured by accelerometer.
(f) Altitude from barometric and radar altimeter sources.
(g) Track velocities from Doppler or GPS sources.
(h) Flight path information from VOR, ILS, Inertial Navigation and GPS sources.
(j) Position of the pilot’s controls from signal units incorporated in the controls.
17. Processing Unit. The processing unit is the module which performs the basic judgemental
process provided by the pilot in manual systems. Its functions are:
(a) Manipulating sensor information into usable and comparable signals.
(b) Comparing rate and positional sensors and feedback inputs, by using differentiation
and integration computing techniques.
(c) Establishing what degree of error exists between them.
(d) Calculating the amount of control response needed to correct the error, whilst
remaining within prescribed limits and to suit the handling qualities or scheduled flight path of
the aircraft.
(e) Initiating that control response by signalling movement commands to the appropriate
control surfaces.
The relatively simple processing needed for the operation of part-systems can be provided by
mechanical levers and linkages or by simple electrical bridge balance networks. However, full AFCSs
require use of the more powerful and versatile electronic processing capabilities of micro-chip
devices.
18. Signalling. The method by which movement commands are to be signalled between
processing unit and actuators, or direct to control surfaces, is selected during the design stage from
three available techniques:
(a) Mechanical analogue movements using rigid material linkages in the form of push/pull
rods or flexible cable loops.
(b) Electrical analogue or digital signals fed through conventional cables.
(c) Light analogue or digital signals fed through fibre optic cables.
Mechanical linkage systems are ideal for use with simple part-systems where the ease with which
modification and integration of signals, by using simple mechanical devices, offers considerable
advantage. However, mechanical systems have high friction and inertia, they are complex and
therefore expensive to install and maintain, and they are very vulnerable to disruption by damage or
loose articles. Electrical systems overcome many of these disadvantages, and their low density
permits duplication of signalling paths without incurring significant weight and space penalties. Light
systems offer even more advantages than electrical systems, although the generation and
subsequent decoding of light signals, which are in the form of modulated and pulsed beams,
obviously require additional operations to be carried out in the processing and actuation modules.
Fibre optic cables have a considerably greater data carrying capacity than do the equivalent electrical
items, and they are not prone to electro-magnetic interference from either on-board or external
sources. They are however prone to damage.
19. Actuators. AFCS actuators are powered flying controls, and as such are dealt with in the
previous chapter. The need for
actuators to respond rapidly and
accurately to signal inputs has
resulted in the elimination of all
types other than those powered by
hydraulics or electrics. The greater
power which can develop compared
to electrical units of equal weight
and size, plus their speed and
accuracy of response dictate that
they are used in a majority of
applications. Actuators in part- Fig 8-6: Pilot Input to System
systems may be fitted in series with
the manual control system.
Examples of installation, with a
series actuator are shown in Fig 8-6
and in Fig 8-7. In Fig 8-6, the spring
strut and autopilot servo move as a
single rigid rod when the pilot makes
a control input. In Fig 8-7, when the
autopilot servo moves, resistance of
the friction collar causes the spring
unit to compress thus absorbing
Fig 8-7: Autopilot Input to System
movement of the servo and avoiding
any movement in the control column. However, when the spring unit is fully compressed at the limit of
autopilot authority, further movement of the servo will cause the control column to move. The
advantages of such a system are:
(a) The actuator moves the control surfaces without also moving the pilot’s controls.
(b) The actuator acts as a rigid link in the control run, when it is not operating.
(c) The part-system has only limited authority over the control system, usually limited to
10% in helicopters and slightly more in fixed wing. Thus, if the part-system suffers a failure,
the pilot still retains a majority of the control range of movement.
20. Feedback. Position and rate feedback for the control loops of an AFCS system are derived
from system sensors monitoring the aerodynamic effect of command signals.
313
CHAPTER 9
UNDERCARRIAGES
Introduction
1. The undercarriage of an aircraft includes the wheels, tyres and brakes as well as the main
undercarriage leg components. It performs the essential function of providing an interface between
aircraft and ground during landing, take off, ground manoeuvring and whilst at rest. However, it is
completely redundant during flight, and therefore the design of an undercarriage is usually a critical
compromise between optimizing performance on the ground and minimizing weight and drag
penalties in the air. Examples of the extremes of this compromise range between the provision of the
minimum for a Remotely Piloted Vehicle - a detachable wheeled dolly for the aircraft to take off from
and a parachute to lower it safely after flight - to the more generally serviceable undercarriage which
allows the aircraft to be landed at high weights and on a wide variety of surfaces, and to be
manoeuvred rapidly and precisely between the runway and its dispersal area for replenishment, prior
to dispatch on further sorties.
Design Considerations
2. Principal factors which govern the design configuration of a particular undercarriage are:
(a) The aircraft’s role and its intended theatre of operation, for example, the requirements
for strategic aircraft operating from well founded airfields are considerably different from those
of tactical STOL aircraft intended to operate from semi-prepared strips.
(b) The configuration of the aircraft and its intended performance / cruise speed, for
example, high wing, high speed aircraft impose greater design problems than do low wing,
low speed aircraft and helicopters.
(c) The numerical factors, for example, landing speeds and weights, permissible length
of landing run and cross wind landing / take off capability all have considerable influence upon
undercarriage design.
Typical Configurations
3. The general design configuration for an undercarriage emerges from consideration of the
above factors:
(a) Physical strength of the components necessary to withstand landing, braking and
crosswind loads. The strength parameters are set out in defined design standards.
(b) Shock absorber performance capable of accepting the maximum intended sink rate of
the aircraft onto the ground, the type of surface over which the aircraft will taxi and the speed
of turning during taxi.
(c) Fixed (stronger, simpler and lighter) undercarriage or a retractable (less drag)
undercarriage.
(d) Streamlining and provision of undercarriage doors necessary to reduce drag during
flight.
(e) Dimensions of the ground track needed to provide stability during landing and taxi.
(f) Fuselage or wing space available for stowing and attaching the gear.
(g) The basic configuration for instance, the standard tricycle for good ground
maneuverability and stability, bicycle (with outriggers) for strength and relative ease of
stowage or tail wheel for simplicity and low cost. The most appropriate undercarriage for a
small helicopter may be a pair of skids, despite the complications which these impose upon
ground handling.
4. Fixed Undercarriage. On land airplanes, there are two basic classes of fixed gear
undercarriage: Main Gear with a nose wheel commonly called a tricycle gear, and main gear with a
tail wheel. There are several types of undercarriage in use for the main gear. These are used with
both the tail wheel and the tricycle gear configuration. They are split axle, tripod, single spring leaf
cantilever and single strut.
(a) Split Axle. It has the axle bent upwards and split in the
centre to enable it to clear obstructions on the ground (Fig 9-1).
This type is used on airplanes such as the Piper PA : 22. It is
suspended on shock cords wound around a fuselage member
which enables the whole assembly to spread when loads come on
it. A strut or tie rod is usually incorporated to brace the structure Fig 9-1: Split Axle
against side loads.
(b) Tripod. The tripod landing gear is illustrated in Fig 9-2

and is typical of airplanes such as the Aeronca and the Champion.
This gear consists of three members hinged so as to form a
triangle. Two of these are rigid. The third is an oleo leg, designed
to telescope and hence shorten its length when the load comes on
Fig 9-2: Tripod
the wheel. On landing, the whole assembly spreads outwards and
Landing Gear
315 Undercarriages
upwards until springs, rubber discs, or other devices take the weight.
(c) Single Leaf Cantilever. The Spring steel variant of

this type of main landing gear is used extensively on Cessna
aircraft. The gear consists of a single strap of chrome vanadium
steel bent to form the shape of the complete undercarriage
structure. It is attached to the fuselage in a cradle bulkhead by
bolts. It is capable of storing energy in initial impact, thereby
producing quite low load factors. Low maintenance, simplicity
and long service life characterize this gear type. On the Cessna Fig 9-3: Single Leaf
"Cardinal", a spring steel tubular gear (Fig 9-3) replaces the Cantilever
more familiar single leaf gear described here. The spring steel tubular gear has the same
characteristics as the single leaf type.
(d) Single Strut. This type is used, on several

modern, low wing, fixed gear airplanes such as the HPT–
32 and the Beech Musketeer. This gear consists of a
single leg or strut extending downward from its
attachment point on the main spar. The Strut usually
incorporates a hydraulic cylinder or rubber biscuits for
the purpose of absorbing the shock (Fig 9-4). Fig 9-4: Single Strut Gear
5. Retractable Gear. Retractable gears are made to retract or fold up into the wing or
fuselage in flight. The mechanical means and methods for accomplishing this are many and varied.
The wheel may fold sideways outwards towards the wing or inwards towards the fuselage. The latter
is most common on high speed military airplanes when the wing camber is shallow. On some multi-
engine airplanes the wheel folds straight back or forward into the nacelle and is left partly projecting in
order to protect the belly of the aircraft in the case of a wheels-up landing. Some retractable
undercarriages are made to turn through 900 as they travel up and so fold into the side of the
fuselage.
6. Most retractable undercarriage legs are cantilever, being a single oleo leg with no external
bracing. They are hinged at the top to permit them to fold. The means of retraction may be a hand
gear, electric motor, or motor-driven hydraulic pump. Where mechanical means are used, a hand gear
is also provided to allow for lowering the gear in an emergency.
7. Making the undercarriage retractable is a common practice with both the tricycle and tail
wheel configuration. In the case of tricycle gear, the nose wheel is also made retractable. In the case
of a tail wheel, however, because it is small and causes little drag, it is fixed.
Nose Wheel Vs Tail Wheel
8. The practice of placing a steerable third wheel forward of the main gear has found universal
acceptance in modern airplane design and is referred to as being a tricycle gear configuration. The
modern trend of the tricycle gear configuration by most manufacturers is the result of certain
advantages that this type of landing gear has over the tail wheel configuration. These advantages of
tricycle gear configuration are:
(a) Nose-over tendencies are reduced greatly.
(b) Ground looping tendencies are reduced.
(c) Visibility over the nose when taxiing, taking off or landing is superior due to the level
flight position of the airplane while on the ground.
(d) Greater maneuverability on the ground under high wind conditions due to the
negative angle of attack of the wings.
(e) Greater controllability on the ground in cross-wind conditions. Therefore, tricycle

geared airplanes can use single runway airports (which are becoming more numerous) with
greater safety in cross wind conditions than can tail wheel airplanes.
(f) A novice can usually learn to maneuver a tricycle geared airplane on the ground in
less time than he can master a tail wheel airplane. .
9. The landing gear configuration in which the third wheel is rearward of the main gear (i.e. at
the stern of the airplane) is referred to as a tail wheel configuration (old timers fondly call such
airplanes, tail-draggers). The tail-
wheel undercarriage dominated
aircraft design for the first four
decades of flight and is still widely
used on many small piston-engine
planes. The tail-dragger
arrangement consists of two main
gear units located near the centre of
gravity (CG) that support the
majority of the plane's weight. A
much smaller support is also located
at the rear of the fuselage such that
the plane appears to drag its tail,
Fig 9-5: Stable and Unstable Behavior of
hence the name. This tail unit is
Tricycle Gear vs. Tail-Dragger Gear
317 Undercarriages
usually a very small wheel but could even be a skid on a very simple design. However, the greatest
liability of this landing gear layout is its handling characteristics. This design is inherently unstable
because the plane's center of gravity is located behind the two main gears. If the plane is landing and
one wheel touches down first, the plane has a tendency to veer off in the direction of that wheel (Fig
9-5). This behavior can cause the aircraft to turn in an increasingly tighter "ground loop" that may
eventually result in scraping a wingtip on the ground, collapsing of the gear, or veering off the runway.
Landing a tail-dragger can be difficult since the pilot must line up his approach very carefully while
making constant rudder adjustments to keep the plane on a straight path until it comes to a stop.
Many tail-dragger designs alleviate these handling problems by fitting a tail-wheel that can be locked
instead of swiveling on a castor. Locking the tail-wheel helps keep the plane rolling in a straight line
during landing. Another disadvantage of the tail-dragger is poor pilot visibility during taxiing since he is
forced to peer over a nose that is tilted upward at a steep angle.
10. Tail wheel airplanes have advantages too. These are:
(a) The tail wheel has less parasite drag than a nose wheel due to its smaller size.
(b) The tail wheel is cheaper and easier to build and maintain.
(c) A broken tail wheel will not result in as much damage to an airplane as would a
broken nose gear.
(d) A tail wheel airplane can be more easily man handled on the ground and, because
the tail is lower than that of a tricycle-geared airplane, it fits into some hangar space more
easily.
(e) When using sand or gravel airports, the tail wheel aircraft will sustain less propeller
damage since the tips of the propeller are farther away from the ground and are less likely to
pick up loose objects, such as stones and debris.
(f) With constant use in rough fields, the tail wheel airplane is not as likely to sustain
airframe damage since it is the main undercarriage which takes the bulk of the load and the
shock when the airplane rides over depressions and irregularities on the ground The main
undercarriage (which hits the bumps first) is attached to a primary structure and is therefore
stronger and more rigid than a nose gear (which in the tricycle gear configuration is the first to
hit the bump) which is usually fastened to a weaker or non-primary part of the airframe. A tail
wheel will easily absorb bumps that may be severe enough to damage a nose gear.
On most modern airplanes, regardless of whether they have a fixed or retractable undercarriage, the
nose wheel and the tail wheel are steerable by the pilot’s controls.
Arrangement of Undercarriage Units
11. As the aircraft designs have kept on advancing, designers have been trying out different
arrangements of undercarriage to suit their requirements. Besides the tricycle and tail wheel type
some other undercarriage arrangements that have been used are described in the following
paragraphs.
12. Bicycle Gear. A relatively uncommon landing gear option is the bicycle undercarriage.
Bicycle gear features two main gears along the centerline of the aircraft, one forward and one aft of
the center of gravity. Preventing the plane from tilting over sideways are two small outrigger gear
mounted along the wing (Fig 9-6). The only real
advantage of bicycle gear is its lower weight and
drag than either the tail-dragger or tricycle
arrangements. Bicycle gears are also useful on
planes with very long and slender fuselages where
there is little room for more traditional undercarriage
arrangements. Unfortunately, bicycle gear are very
demanding on the pilot who must maintain a very
level attitude during takeoff and landing while
carefully managing airspeed. The pilot must also Fig 9-6: Bicycle landing gear
of Harrier
compensate for any rolling motion that could cause
the plane to land unevenly on one of the outrigger gear, and crosswinds are particularly difficult to
deal with. Because of these limitations, bicycle gears are generally limited to planes with high aspect
ratio wings that generate high lift at low angles of attack. Good examples of such planes are large
bombers with a narrow fuselage and large wingspan like the B-47. Another common application of
the bicycle undercarriage is aboard vertical takeoff and landing designs like the Harrier. Here, the
gear layout provides safety and stability in case of an engine failure during landing.
13. Single Main Landing Gear. This design is particularly simple, lightweight, and low drag
and may even include skids rather than wheels. This
simplicity makes the gear arrangement attractive for
use on light planes like gliders and sailplanes, but
the single main gear is generally impractical for
larger aircraft. Perhaps the best known application
of a single main gear arrangement was the U-2
reconnaissance plane. This aircraft had a single
large gear unit near the center of gravity plus a much
smaller tail-wheel. Two additional outriggers called
"pogos" were attached by ground crew to keep the Fig 9-7: Single Main Landing Gear
plane from tipping during taxi, but these were of U-2 Dragon
319 Undercarriages
removed prior to takeoff.
14. Quadricycle Gear. Quadricycle gears are

also very similar to the bicycle arrangement except
there are four main gear roughly equal in size and
mounted along the fuselage. Like bicycle gear, the
quadricycle undercarriage also requires a very flat
attitude during takeoff and landing. This arrangement
is also very sensitive to roll, crosswinds, and proper
alignment with the runway. The most significant
advantage of quadricycle gear is that the plane's floor
Fig 9-8: Quadricycle landing gear
can be very close to the ground for easier loading and
of the B-52 Stratofortress
unloading of cargo. However, this benefit comes at the
price of much higher weight and drag than bicycle gear. Quadricycle gear are sometimes used on
cargo planes, but probably the most well known example is the B-52 bomber (Fig 9-8). This aircraft
employs a cross between the quadricycle and bicycle arrangements since it has four main gear plus
two small outriggers near the wingtips.
15. Multi-Bogey Gear. A final variation that is

worth mentioning is the use of multiple wheels per
landing gear strut. It is especially common to place two
wheels on the nose strut of the tricycle arrangement to
provide safety and steering control in case of a tire
blowout. This additional tyre is particularly useful on
carrier-based aircraft where two nose wheels are a
requirement. Multiple wheels are also often used on
main gear units for added safety, especially on
commercial airliners (Fig 9-9).
Fig 9-9: Multi-bogey landing gear

16. When multiple wheels are placed on the same
The many, many, many, many,
gear unit, they are attached together on a structural wheels of the An-225
device called a bogey. The heavier the aircraft
becomes, the more wheels are typically added to the bogey to spread the plane's weight more evenly
across the runway pavement. In general, a plane weighing less than 50,000 lb (22,680 kg) has only
one wheel per main gear strut. Aircraft weighing up to 200,000 lb (90,720 kg) usually carry two wheels
per strut. On planes weighing up to 400,000 lb (181,440 kg), a four-wheel bogey is typical. Aircraft of
greater weight often carry four bogeys, each with four to six wheels. The best examples of multi-
bogey aircraft are large mega jets like the An-225. This mammoth cargo plane has seven pairs of
wheels on each main gear assembly plus four nose wheels, combining for a total of 32 tyres! Another
good example is the Boeing 747. The 747 is equipped with four main gear units, each with four-wheel
bogies, plus twin nose wheels so that the plane's weight is spread across 18 wheels.
Shock Absorbers
17. The shock absorber is the most complex component of the undercarriage. Its role is to
dampen the shocks of landing and taxiing and of movement over uneven runway pavements. Two
basic types of shock absorber are available. One utilizes the compressibility of oil at pressures above
700 bar to damp out shocks, whilst the other utilizes various combinations of oil and nitrogen under
pressure to provide damping.
18. Oil-Compression Shock Absorbers. The principle of the oil filled (liquid spring) absorber
is shown at Fig 9-10. On landing, movement of the leg is restricted by the slow rate at which oil is
able to pass through the damping orifices into the upper chamber of the liquid spring. If large shocks
are experienced, the oil remaining beneath the piston is compressed until its pressure exceeds the
loading of the piston non-return valve spring. At this point, a larger volume of oil is released round the
valve, thus damping out the larger landing shocks. On the recoil, oil is forced back below the piston
through the small damping orifices.
Fig 9-10: Principle of the Liquid Spring
19. Oleo-Pneumatic Shock Absorbers. The operation of oleo-pneumatic shock absorbers

(Fig 9-11) is as follows:
(a) Extended. As the aircraft becomes airborne the load on the shock absorber strut is
reduced to zero and consequently there is no force opposing the air pressure in the air
chamber. Thus, the separator piston is forced upward causing pressure to be transmitted
321 Undercarriages
through the liquid to the upper cylinder head, resulting in the plunger tube being forced
downwards until the piston reaches the stops.
(b) Static. When the aircraft is on ground its weight

compresses the strut. The separator piston is therefore
forced down by the liquid passing through the holes in the
piston to the plunger tube, until the air pressure is
sufficient to balance the load on the strut.
(c) Compression. When the aircraft lands the

piston and the plunger tube are forced up. The liquid
above the piston is forced through the holes in it causing
the clack valve to open, thus allowing the liquid to pass
through the holes in the valve to the plunger tube. The
separator piston is therefore forced downwards
compressing the air below it. This action absorbs the
landing shock.
(d) Extension. After the initial landing shock the air

pressure, being greater than that required to support the
load, drives the separator piston upwards. Thus the liquid
in the plunger tube is forced upward and the clack valve
closes leaving a single hole in the piston through which Fig 9-11: Oleo-Pneumatic
liquid can pass to the upper cylinder. This upward transfer Leg
of the fluid tends to force the upper cylinder head away from the piston, so the strut extends.
The restricted transfer of the liquid through the piston slows the extensions and damps the
rebound.
20. Variations on the oil / gas (oleo-pneumatic) absorber are shown at Fig 9-12. The combination
of oil and gas provides a more effective method of shock absorption, enabling a reduction in
component size and weight to be made for the same performance. Shock absorbers must be
designed so that they never reach full extension or closure under any operational load condition
otherwise the undercarriage will momentarily become rigid passing very high peak loads into the
aircraft structure. The nose undercarriage is subjected to a wider range of conditions than is the main
undercarriage, because of centre of gravity movement and pitching moments caused by braking
reactions. For this reason, two stage shock absorbers similar to that shown at Fig 9-12(d) are often
fitted to the nose to provide the greater required range of operation.
Fig 9-12: Oleo-Pneumatic Absorbers

Wheel Loading
21. Wheel loading (lb / unit area) is defined as the static load at take off on each wheel. Wheel
loading has to be considered in conjunction with runway strength. If the wheel loading is too high the
runway structure will be damaged and the tyre will sink into it. Consequently a designer must keep
these loads within the limitations of the existing runways. Further, It is necessary to classify both
pavements and aeroplanes in such a way that the load bearing capacity of the pavement can readily
be compared with the load exerted by the aeroplane. Pavement strengths are classified in various
ways as detailed in the following paragraphs.
Load Classification Number (LCN) System
22. The LCN is the ratio between two values, the first the 'standard value', is the load required to
produce a failure of a given surface when acting over an area of 530 in2 and the second value is the
load required to produce a failure on the same surface, but applied over a lesser specified area. The
ratio between these two values is expressed as a percentage and is known as the LCN. By
comparing the wheel loading of an aeroplane with the LCN of aerodrome pavement, it is possible to
determine whether the pavement is sufficiently strong for that particular aeroplane. The load exerted
by the aeroplane depends on:
(a) Aeroplane all up weight.
(b) Tyre pressure (and thus contact area).

323 Undercarriages
23. The LCN system was originally based on a minimum tyre contact area of 200 in2, with wheels
in single units. The introduction of increasingly heavier aeroplanes with their associated multiple
wheel units and higher tyre pressures i.e. smaller contact areas, complicated the original calculations
and, in order to obtain a simple figure on which comparisons could be made, the concept of
Equivalent Single Wheel Loading (ESWL) or Single Isolated Wheel Load (SIWL) was introduced.
24. Equivalent Single Wheel Loading (ESWL). The ESWL of a particular group of two or
more closely spaced wheels is the SIWL which, operating at the same tyre pressure as the wheels in
the assembly, produces critical effects on the pavement equivalent to those produced by the group of
wheels. There is a direct, though complex relationship between EWSL and LCN and conversion from
one to the other can be achieved by the use of a graph.
25. Load Classification Group (LCG) System. The LCG

system is a development of the LCN system and is intended to take LCN LCG
account of local LCN variations by placing pavement load bearing 101
I
capacity into groups each of which embraces a range of LCN values Upwards
as shown in Table 9-1. 76 – 100 II
51 – 75 III
26. Although the LCG system was evolved for classification of 31 – 50 IV
aerodrome pavements, it is convenient to group LCNs of aeroplanes 16 – 30 V
under this system, e.g. an aeroplane having a LCN of 70 would 11 – 15 VI
belong to Group III. Alternatively, knowing the aeroplane LCN, the 01 - 10 VII
appropriate pavement LCG requirement can be read from Table 9-1.
Conversion between LCG, LCN and ESWL can be made by using a Table 9-1: Correlation of
suitable graph. LCN and LCG
Pavement Classification Number (PCN) / Aircraft Classification Number (ACN)
27. The System. The CAN / PCN system provides a method of classifying pavement bearing
strength for aircraft above 12,500 lbs Maximum Total Weight Authorized (MTWA). The ACN is a
number expressing the relative effect of an aircraft load on a pavement for a specified sub-grade
strength. The PCN is a number expressing the bearing strength of a pavement for unrestricted
operations:
(a) ACN. The ACN is calculated taking into account the weight of the aircraft, the
pavement type and the sub-grade category. ACN values for aircraft (one for rigid pavements
and one for flexible pavements) are given in relevant manuals. These ACN values are to two
weights, one at MTWA and the lower weight for the APS or Operating Weight Empty (OWE).
If the aircraft is operating at an intermediate weight, the ACN value can be calculated by a
linear variation between the limits.
(b) PCN. PCNs are reported as a five part code. Apart from the numerical value of the
PCN, the report includes the pavement type (rigid or flexible) and the sub-grade support
strength category. Provision is made in the report for the aerodrome authority to place a limit
on maximum allowable tyre pressure, if this is a constraint and an indication of whether the
pavement has been evaluated by technical means or by past experience of aircraft use of the
pavement.
325
CHAPTER 10
BRAKING SYSTEMS
Introduction
1. Stopping an aircraft requires the rapid dissipation of large amounts of kinetic energy. The
energy is dissipated by conversion to heat energy in the wheel braking system and by being used to
do work against applied loads. Such loads include drag (from aerodynamic devices such as flaps and
spoilers) and opposing forces provided by reverse thrust devices or propeller reverse pitch. In some
cases, brake parachutes or external retardation devices such as arrester wires are also used to
absorb the kinetic energy. Typically, wheel brakes, aerodynamic devices and thrust reversers absorb
equal amounts of energy during a normal landing.
WHEEL BRAKES
2. Heat Dissipation. Temperatures of up to 1400 degrees Celsius are reached in high

performance braking systems. To prevent damage to the tyres and undercarriage structure, the heat
energy must be dissipated rapidly into the surrounding air. If this does not happen, as can be the
case after an aborted take-off and subsequent long taxi back to dispersal, the tyres can overheat and
burst, and brake fires are likely to occur. To prevent tyres bursting due to overheating, wheels are
fitted with fusible plugs which melt at a preset temperature. This allows the tyre to deflate at a steady
controlled rate.
3. Configuration. Most aircraft are equipped with hydraulically operated disc brakes, although
drum brakes are sufficiently effective for light aircraft. Both types can be operated either hydraulically
or pneumatically although nowadays hydraulic actuation is almost becoming universal.
4. Drum Type. The brake drums on the early wheel designs were pressed directly into the rim
of the wheel casting, consequently cooling was inadequate, resulting in excessive heat at the tyre
seat, and distortion and loosening of the drum. This led to the design of a wheel having a flanged
brake drum bolted to the wheel casting at a distance from the tyre seat. In addition a series of vanes
were required to circulate cooling air between the drum and the wheel rim. Construction and operation
of a typical inflated bag type (Fig 10-1) is as follows:
(a) Construction. The unit may best be described as a smaller non-rotating wheel
fitted inside the main wheel. Its hub is secured to the axle to prevent the whole unit from
rotating. The rim carries an inflatable rubber bag on which are seated a number of metal
backed brake shoes which are keyed to the brake unit casting to prevent rotation but which
are free to move radially outwards. To prevent the rubber bag from creeping these brake
shoes and to keep it free from dirt and grease, metal separators are placed between adjacent
shoes. The metal backs of the shoes have two lugs which pass through slots in the rim and
are engaged by a metal strap. The straps are spring-loaded to ensure that the brake shoes
are held away from the drum when the bag is deflated.
Fig 10-1: Inflated Bag Type Drum Brake Unit
(b) Operation. When the pilot operates the brake control the bag is inflated by air
pressure, causing the blocks seated on it to be pressed against the brake drum which is
rotating within the main wheel. The blocks cannot rotate so the resulting friction, which can be
varied by the pilot, retards and eventually stops the rotation of the wheel.
5. Disc brakes. These offer the advantages of higher surface area for contact between the
brake material and the rotating surfaces and larger capacity heat sinks to absorb the heat generated
during braking. High performance disc brakes are constructed as multiple stacks of discs made from
Fig 10-2: Wheel and Disc Brake Assembly

327 Braking Systems
carbon composites which are able to operate at the necessary temperatures. A typical multiple disc
unit consists of 4 or more rotors keyed to the inside of each main wheel, and 5 or more stators
assembled on to splines of each main undercarriage axle assembly. Fig 10-2 shows such a brake
assembly in situ. Operation of the brakes is usually through a single selection lever. Pedals attached
to the pilot’s rudder bar direct differential hydraulic pressure to the main wheel brake units to provide
steering. Hydraulic pressure operates either directly or through a servo system upon the brake units.
The pressure causes rotor and stator discs to be pressed together, and the resulting friction provides
a retarding force to the main wheels generating heat in the process. Rotor discs are usually
constructed in segments which allow a small amount of deflection to take place. This reduces
stresses and prevents the discs cracking.
Braking Control and Anti-skid Systems
6. Braking Dynamics. To minimize the

landing run, it is imperative that brakes apply
maximum retardation force without causing the
adhesion between tyre and pavement to be
exceeded thus causing the aircraft wheels to skid.
The point of skidding is dependent upon the
condition of the runway surface, the vertical load
which the aircraft tyres exert on the ground and
the retarding force applied by the brakes. The
mathematical relationship between vertical load
and retarding force is termed ‘µ’. Because vertical Fig 10-3: Relationship between
µ and Vertical Load on the Tyres
load is inversely proportional to the aerodynamic
lift acting on the aircraft, it follows that µ will
increase as aircraft speed decreases. Fig 10-3
shows this relationship for constant runway
conditions. Whenever braking force is applied to
the tyre, a degree of slip occurs between the tyre
and the surface of the pavement. This is defined
in terms of the difference between the rotational
speed of a braked wheel and the rotational speed
of a similar free rolling wheel. It is expressed as a
percentage. The value of µ varies with wheel slip,
and Fig 10-4 shows the relationship between
maximum available µ and wheel slip for varying
conditions on the same runway. Fig 10-4: Relationship between µ and
Wheel Slip for Varying Runway
Conditions
7. Control Systems. As can be deduced
from the above considerations, the application of maximum braking effort to minimize the landing run
requires the solution of complex dynamic equations, balancing braking forces with speed, weight and
runway conditions. Electronic sensing has permitted all phases of the braking process to be inter-
related and fail safe over-rides to be employed.
8. Anti-Skid Systems. Early mechanical anti-skid systems utilized the inertia of a small
flywheel to sense rapid changes of main wheel rotational speed such as occurs during a skid. On
sensing a skid, the systems reduced hydraulic pressure - thereby reducing braking effort and stopping
the skid. They reinstated pressure when skidding had reduced. The resultant cycling between
skid/no skid conditions caused the braking pressure to continuously pulse or modulate, and the
technique became known as brake modulation. Subsequent electrical systems used sensors to
measure wheel speed and compared the speed to a datum. The use of simple electronic processing
allowed a controlled profile of modulation to be achieved instead of the on/off characteristics of the
earlier mechanical systems, and considerable improvements in braking efficiency were achieved.
Modern anti-skid systems utilize control technology to vary not only the frequency of modulated
braking pulses but also their amplitude (pressure). Thus, the systems can maintain braking forces at
a level immediately below that which would cause skidding for all speeds and surface conditions. The
systems also hold the brakes off until after touch down and wheel spin up has occurred, and similarly
apply braking to spin down the wheels after take off and undercarriage retraction has taken place.
Thus, the systems can relieve the crew of much of the work load of brake management at the critical
periods of landing and take off.
TAIL PARACHUTES
9. These are also known as "Brake Parachutes" and are used to decrease the length of the
landing run. They are streamed from a point at the rear of the aircraft when the round out is completed
or after the wheels have touched the ground. In general, they produce enough drag to cause a
steady rate of deceleration varying from about 0.25 ‘g' to 0.35 g', depending on the particular
installation. Below 60 to 70 knots, the drag varying as the square of speed, falls to much lower figure
and tail chute drag becomes insignificant as compared to wheel brakes which can cope comfortably
with the inertia of the aircraft while maintaining the deceleration at 0.25 ‘g' or more. At high landing
speeds, if the brakes were required to produce the same rate of deceleration as the brake parachute,
dissipation of the heat generated would require an impossibly large mass of metal without burning out
both the brakes and tyres. The diameter of the parachute depends on the weight, size and the landing
speed of the aircraft. For aircraft having a landing weight of around 10,000 Ibs the "flying" diameter of
the parachute is from 6 to 8 feet. At a touchdown speed of 130 knots this gives a drag Of about
2500Jbs and a rate of deceleration of about 0.25 ‘g'. For large aircraft with landing weights about
100,000 Ibs the “flying" diameter of the brake parachute is about 35 feet. This produces at a
touchdown speed of about 150 knots, a drag of some 50,000 Ibs and an initial rate of deceleration of
about 0.35 ‘g'.
329 Braking Systems
10. On some installations provision may be made for automatic retraction of the brake parachute
when the speed falls to 60 to 70 knots. These parachutes can be used as often as necessary on
successive circuits and landings. On non-retractable installation a parachute is usually jettisoned at
the end of the landing run. At any time after the parachute has been 'streamed' it can be disconnected
or jettisoned by the pilot in an emergency. This applies both to the retractable and non-retractable
type.
AIR BRAKES
11. High speed aircraft, having comparatively high weights and low drag, tend to retain their
speed for a considerable time after the engine has been throttled back. Further having eventually
reached to desired lower speed, any slight downward flight path or increase in power causes an
immediate and appreciable increase in speed. An air brake is an integral part of the airframe, and can
be extended to increase the drag of an aircraft at will, enabling the speed to be decreased more
rapidly or regulated during a descent. Some aircraft lower the undercarriage partially or completely to
obtain the same effect.
12. Although, the area of the airbrakes on a typical fighter is small, considerable drag is produced
at high speeds. For example, an airbrake with an assumed CD of 1.2 and a total area of about 2.5
square feet produces a drag of about 5700 lbs. When opened at 500 knots at sea level, this figure is
indicative of the large loads imposed on an aircraft when flying at high indicated speeds. The
effectiveness of an air brake varies as the square of the speed and therefore at about 120 knots the
same air brake gives a drag of about 330 lbs only. A typical aircraft with air brakes ‘in’ and power ‘off’
takes ¾ minutes to decelerate to 150 knots while with air brakes ‘out’ this time is reduced to 1 minute.
13. Ideally air brakes should not produce any effect other than drag, although on some aircraft the
air brakes are designed to produce an automatic nose-up change of trim when extended. In practice,
however, the operation of most air brakes is accompanied by some degree of buffet, with or without a
change of trim. The strength of these adverse effects is usually greatest at high speeds, becoming
less as the speed decreases.
331
CHAPTER 11
CABIN PRESSURIZATION AND AIR CONDITIONING SYSTEMS
Introduction
1. The adverse physiological effects of altitude, rapid changes in altitude, composition of the
atmosphere and extremes of temperature and humidity necessitate that the environment in an aircraft
cabin or cockpit is carefully controlled within specified limits during all phases of flight. The limits
within which each parameter must be controlled and the methods adopted to achieve them are
outlined in this chapter.
Altitude
2. Unless breathing oxygen is normally available, the effective cabin altitude of an aircraft must
be limited to between 6,000 and 8,000 feet, whilst if oxygen is available but the use of pressure suits
is not desirable, the limit of cabin altitude is 25,000 feet. Further aspects of cabin pressurization are
discussed in subsequent paragraphs.
3. Rate of Change in Altitude. The permissible rate of change of cabin altitude is dependent
upon the general fitness and health of its occupants. The practicable limit for passenger carrying
aircraft is a maximum climb
rate of 500 feet / minute and
a maximum descent rate of
300 feet / minute, although
these figures may be more
than doubled for combat
aircraft. These parameters
are shown graphically in Fig
11-1, and aspects of the
pressurization system which
control the rate of cabin
altitude change are
Fig 11-1: Typical Cabin Altitude Profiles
discussed below.
Composition of the Cabin Atmosphere
4. To prevent the build up of carbon dioxide, water vapour, dust, fumes and odours, cabin
atmosphere must be changed continuously by a ventilation system. The rate of ventilating air flow is
dependent upon the volume of cabin space per occupant cabin (Complement density). The smaller
the space per crew member, the higher must be the air flow. In passenger aircraft, an air flow of
approximately 1.5 kg/minute is normally provided. This mass usually comprises 50% fresh air and
50% re-circulated air. The air flow is discharged into the cabin to create a general circulatory flow of
air, although higher speed air flows are provided for each passenger and crew member through
individual punkah louvres. In the more restricted volume of a combat aircraft cockpit, a flow of up to 5
kg/minute is usually provided. In normal flight conditions, 80% of this mass is arranged to circulate
directly around the crew, whilst the remainder is used for demisting and more general cockpit
ventilation. Ventilation is provided by the aircraft air conditioning and pressurization systems.
5. Temperature and Humidity. The

range of ambient temperatures and humidity in
which a crew can operate comfortably without
rapidly becoming fatigued is defined as the
Comfort Zone graphically depicted in Fig 11-2.
The air conditioning system controls both
temperature and humidity within the cabin to
remain within this zone. It can be noted that the
º º
dry bulb air temperature limits are 18 to 29 C
and the humidity range is 30 to 70 % RH.
Fig 11-2: The Comfort Zone
Pressurization Systems
6. Introduction. The ideal cabin altitude to be maintained in flight would be Sea Level.
However, fuselage structural strength considerations rule this out, and the more practicable maximum
altitude of 8,000 feet is normally achieved. In combat situations, the likelihood of battle damage
causing rapid depressurization
dictates that a much higher
cockpit altitude, normally
25,000 feet, be maintained.
However, many combat
aircraft systems are capable of
operating down to the lower
level of 8,000 feet when this is
operationally permissible. The
permissible difference
between the aircraft
operational ceiling and its
maximum cabin altitude is
defined as the differential
pressure which the fuselage Fig 11-3: Cabin Differential Pressures at Altitude
333 Cabin Pressurization and
Air Conditioning Systems
must be designed to tolerate. At a pressure altitude of 31,000 feet and a cabin altitude of 8,000 feet,
a differential pressure of 0.46 bar is imposed upon the cabin structure. The differential pressures
relating to other altitudes are depicted graphically in Fig 11-3. The aircraft pressurization system
comprises an air supply and a control system. The air supply, usually of conditioned air, must be
sufficient to maintain required cabin pressures, notwithstanding the normal small leakage of air from
the cabin and the deliberate dumping of air as part of the air conditioning cycle, and a system of
pressure control and safety valves managing both the cabin pressure and its rate of change within the
limits specified.
Degree of Cabin Pressure
7. There are three degrees of cabin pressurization adopted, viz:
(a) Hyperbaric. Maintaining ground level pressure within the cabin up to the aircraft
operational ceiling. In this the differential pressures are quite high (e.g. 13 psi at 50,000 feet).
(b) Isobaric. Maintaining a cabin pressure equivalent to an altitude of 8,000 feet with
oxygen carried for emergency use. In this the differential pressure at 50,000 feet would be
9.25 psi).
(c) Hypobaric. Maintaining a cabin pressure equivalent to an altitude of 25,000 feet. In

this type oxygen will be mandatory for crew and the differential pressure will be 3.75 psi at
50,000 feet)
8. Air Supply. The supply of air for use in pressurizing and conditioning the cabin or cockpit is
normally provided from the engines through an engine compressor bleed. Older aircraft types may
utilise separate, engine-driven compressors. The high pressure, high temperature supply is regulated
and conditioned before being fed into the cabin or cockpit. Aircraft equipped with APUs are usually
configured so that the air conditioning system can also operate using air supplied by the APU during
periods when the aircraft is on the ground. If, during an airborne emergency, the pressurization air
supply is suspended, air conditioning is normally maintained by the use of ram air. However, because
pressurization is no longer available, the aircraft must descend to an altitude below the normal safe
cabin altitude.
9. Control. Cabin pressure and its rate of change are controlled by the regulated release of air
to atmosphere through a discharge valve in the aircraft skin. Although the system can normally be
manually selected on or off, pressurization parameters are preset and automatically monitored and
controlled by the control system. In the event of a system malfunction or cabin differential limits being
exceeded, the control system is over-ridden by a safety valve which senses an unacceptable cabin
differential. The valve automatically opens to dump air, or to allow air to enter the cabin in the event
of too rapid a descent. Thus the cabin differential pressure is reduced to an acceptable level.
10. Introduction. The air conditioning system must be able to provide a supply of air sufficient
to satisfy ventilation and pressurization requirements and at the temperature and humidity necessary
to maintain cabin and cockpit conditions within the comfort zone defined earlier. The majority of
conditioning systems provide control and adjustment of air temperature and control of air humidity.
Some provide positive filtration of the incoming air, although the majority achieve a degree of filtration
only as a secondary function of the water extraction devices used for humidity control.
Temperature Control
11. The desirable temperature in flight in a pressurised cabin is about 20°C. Since this
temperature is reached after the air has been heated by the use of radio equipment, skin friction, solar
radiation, and heat radiated by the crew, it follows that the temperature of the air entering the cabin
should be well below 20°C. Since the temperature of the air supply coming from the engine
compressor casing may be as high as 350°C, the air is usually passed through an air-to-air heat
exchanger before entering the cabin. If the temperature drop obtained is still insufficient and
additional cooling is necessary, a cold air unit is installed in the air conditioning system.
12. Heat Exchanger. It is similar in action to a car radiator except that it is air, and not water,
which is cooled by the passage of ram air through the exchanger. The air to be cooled can be routed
to make four or more passes across the cooler. Heat exchangers are some times referred to as pre-
coolers or, where they are fitted between the compressor and turbine of cold air unit, they may be
termed as intercoolers.
13. Cold Air Units. The principle of the cold air unit is that when air is made to drive a
compressor by flowing through a turbine, the turbine extracts pressure and heat energy from the flow,
which therefore emerges from the turbine in an expanded form, i.e. at a lower pressure and
temperature. There are three forms of cold air units in service: Turbo Compressor, The Brake turbine
and Turbo fan type. A brief Description of each type is as follows:
(a) Turbo-Compressor Type. This is the first type of cold air unit used in earlier
aircraft and has been replaced by the other two types in the latest-aircraft. (Fig.11-4). An inter
cooler (heat exchanger) is installed between the compressor and the turbine. This intercooler
is ram air cooled. The compressor feeds charge into the intercooler, from which the cooled
air flows into the turbine, where it is expanded and cooled further. The compressor of the unit
is thus situated in the charge air circuit, upstream of the intercooler. On expansion of the
charge air, heat is removed by the turbine which drives the compressor but because the
compressor is in the charge air

circuit, the method of dissipating the
turbine energy differs from the other
methods described later. When cold
air is selected, charge air enters the
compressor and passes to the
intercooler, subsequently entering
the turbine volute and expanding
through the nozzle ring and turbine
into the supply ducting. The
temperature of the air entering from
the intercooler is too low to permit
further cooling by direct surface
heat-exchange, but further cooling is
obtained by the action of the turbine
on the charge air, where the heat
energy extracted by the turbine is
Fig 11-4: Turbo-Compressor
converted into mechanical energy. In
this form the energy is transmitted to the compressor where it is reconverted to heat energy,
returned to the airstream, and passed to the intercooler at a temperature at which direct heat
exchange is possible. All heat lost is finally dissipated to atmosphere through the intercooler.
(b) Brake Turbine Type. In this type (Fig 11-5), charge air from the engine compressor
passes through the ram-air cooled heat-exchanger to enter the turbine (through which it
e
x
p
a
n
d
s
)
,
t
h
i
s
Fig 11-5: Brake Turbine
results in a pressure drop and a considerable fall in temperature. Mounted on a common shaft
with the turbine is a centrifugal compressor which is driven by the turbine and which in
absorbing the mechanical energy of the turbine functions as brake. The air passing through
the compressor is obtained from and then discharged to atmosphere. The compressor is thus
external to the flow of charge air and has no other function than that of a brake to the turbine.
In operation the energy removed from the air by the turbine is used to drive the compressor.
A cold air unit of this type is used if the space available for installation is restricted.
(c) Turbo Fan Type. It is similar to brake turbine type, but is more bulky. In this type
(Fig 11-6), the turbine drives a centrifugal fan which is large enough to pass the cooling
airflow required by heat-exchanger. To obtain the maximum cooling effect, air is drawn
through the heat exchanger by the fan, then discharged to atmosphere. The air discharged by
the fan can not be used for further cooling purposes because its temperature is increased
considerably in passing through the pre-cooler and fan. The energy extracted from the
charge air by the turbine is used to drive the fan. Since the pre-cooler in this installation does
not depend on ram airflow for its operation the turbo fan can provide cool air for the cabins
while the aircraft is on the ground.
Fig 11-6: Turbo Fan
Humidification
14. At high ambient temperatures, high humidity retards body cooling and thereby causes rapid
exhaustion leading in extreme cases to heat stroke. Below 20% relative humidity discomfort is caused
by dryness of the throat and skin The humidity of the air can be increased by:
(a) The use of an atomized water spray.
(b) Injecting water into the cabin blower.

(c) Steam injection.
(d) The use of a venturi humidifier.
15. Air may be dried by chemical method, e.g. by the use of silica gel pack, or by water
extractors. Chemical drying is impracticable for air conditioning purposes, because of the bulk and
weight of the equipment required. Water is separated from the air stream either between the two
stages of cooling, or after cooling has been completed. Most water separators utilize momentum
separation techniques to remove the majority of water from the air stream. They comprise a bank of
swirl vanes or louvres or a coarse mesh filter through which the air must pass. As it does so, its
velocity and momentum are changed and any water held in the air coagulates into droplets which
separate from the main air stream and are ducted overboard. After passing through coolers and
water separators, the air is passed through mixing valves to be combined with hot moist bypass air
and recycled cabin air to form a resultant airstream at the temperature required for cabin conditioning.
339
CHAPTER 12
FIRE WARNING AND EXTINGUISHER SYSTEMS
Introduction
1. The design philosophy used in aircraft is first to prevent fire and secondly to provide adequate
fire protection. Protection is usually in the form of fire resistant materials used in the construction of
strategic systems and structures and fire retardent materials used in aircraft furnishings. However, in
those areas where a risk of fire remains, active aircraft fire protection systems are utilized. These
perform two basic and usually independent functions which are:
(a) Fire and overheat detection.

(b) Fire extinguishing.
2. The fire detection and overheat systems sense the presence of fire or excessive heat. They
employ area detectors in large fire zones and spot detectors for individual pieces of equipment. In
freight and passenger aircraft, they are often supplemented by smoke detectors positioned in the freight
bays, baggage holds and toilet compartments. In the case of fire, overheat or smoke, the systems
provide a visual and aural warning to the crew, identifying the area in which the problem exists. The fire
extinguisher system provides a capability for fighting airborne fires in specific major areas, typically the
engines and APU. Fire extinguisher systems invariably require crew intervention for their operation in
the air. However, in the event of a crash or crash landing, they may be activated automatically by
switches which close under high retardation forces or through airframe deformation. Aircraft are also
equipped with hand held extinguishers for use against small fires in internal areas and equipment.
3. The most common causes of aircraft fires are:
(a) Fuel leaks in the vicinity of hot equipment.
(b) Hot gas leaks from engines or ducting, impinging on inflammable materials.
(c) Electrical or mechanical malfunctions in equipment.
4. The initiating cause is usually equipment failure, although obviously damage incurred during
combat or a crash landing would provide ample additional cause. It follows therefore that the areas in
which fire protection systems are deployed should include the engine bays, the APU enclosure and
significant pieces of high energy equipment. Although the carriage of dangerous cargo in aircraft is
adequately legislated for, spontaneous fires can occur in freight and baggage holds. Access whilst
airborne may be possible to such areas, allowing the crew to enter them and fight the fires with portable
extinguishers.
Fire Detection System
5. The function of fire detection systems is to monitor designated areas or equipment for a rise in
temperature either at a higher rate or to a higher level than predetermined acceptable limits. This is to
provide a warning to the crew, and then to complete electrical safety circuits within the fire extinguisher
systems, to permit necessary system operation by the crew. Such safety circuits are provided to
prevent accidental operation of extinguishers, and they are over-ridden either by the detection systems
or by deliberate manual selection by the crew. The APU fire detection circuits are usually arranged to
automatically close down the equipment as part of their operation. It is important that detection systems
reset automatically when conditions return to normal not only to inform the crew that the problem has
receded, but also to be ready to react again if further overheating occurs. Two basic principles of
operation are used in detectors, either as simple electrical switches activated by the differential thermal
expansion of component metals, or as sensors in which temperature dependent changes in the
electrical resistance or capacitance are used to activate an electronic circuit. Smoke detectors are
devices which are sensitive to an increased presence of the chemical products of combustion in the
surrounding air. If smoke is detected, an alert is given to the crew, indicating the problem area.
6. Fig 12-1 shows the

construction of a typical heat sensitive,
self-resetting thermal switch. It
consists of a stainless steel barrel,
which has a high coefficient of thermal
expansion (expands more with an
increase in temperature), attached to a
mounting flange. Inside the barrel is a
spring bow assembly, which has a low
coefficient of expansion (expands less
with an increase in temperature). At
Fig 12-1: Thermal Switch
normal operating temperatures, the
contacts attached to the assembly remain open because the cylinder restricts the bow forcing its arms
apart. As the temperature rises, the cylinder expands more than the bow thus removing the restraint
upon its length and allowing the contacts to
spring together. This process is reversed
as the temperature returns to normal. The
device can be adjusted to operate precisely
at the required temperature.
7. Fig 12-2 shows the sensing

element used in the majority of area fire
detectors. The sensing element is in the
Fig 12-2: Continuous Wire Detection Element
341 Fire Warning and Extinguisher Systems
form of a wire about 2 mm diameter and is known as the continuous wire detection element, although
the term ‘Firewire’, an early trade name, is still widely used.
Fig 12-3: Installation of a Continuous Element System
8. The element can be relatively easily installed around the areas which require to be monitored.
Fig 12-3 shows such an installation comprising two separate systems. The system in Zone 1 monitors
the engine pod for fire, whilst that in Zone 2 monitors the jet pipe area for overheating caused by gas
leaks. Similar installations would also be used in an APU enclosure. Such systems are invariably
installed in continuous loops, as shown in the figure. The wire is vulnerable to damage caused by
vibration, resulting in reduction of electrical properties or actual fracture. The use of continuous loop
avoids the effect of the resultant open circuit, allowing the wire to operate normally and provide a fire
signal even when defective in part. A test device is included in the system to highlight the existence of
a fault and thus to allow timely rectification to be carried out.
Crash Switches
9. Crash switches operate either by sensing high retardation forces (typically in excess of 6 g) or
by the effect of structural deformation around them. They are usually installed in the undercarriage
bays or inside the belly of an aircraft. Inertia switches, sensing excessive ‘g’ forces, utilize sensors
based on either electronic or mechanical accelerometers. Fig 12-4 shows a pendulum switch. It has
the advantage over electronic accelerometers or mechanical free pistons of being omni-directional,
albeit only in a horizontal plane. The pendulum is suspended on a beam which allows it to swing
horizontally in any direction. Normally it is restrained from moving by a spring loaded lever below it.
However, if subjected to excessive horizontal deceleration the pendulum breaks away from its
restraining lever, allowing the lever to rotate and actuate a bank of electrical contacts in the fire
extinguisher circuits. The piston type of switch operates on a similar principle. A horizontal piston is
Fig 12-4: Typical Inertia Switch
restrained in its cylinder by a sprung lever. Under the effect of high horizontal deceleration forces, the
piston will overcome its restraint and move along the cylinder, allowing the sprung lever to rotate and
make a series of electrical contacts. Structural distortion switches are positioned inside the belly of the
aircraft. They are intended to operate during a crash landing when skin deformation will occur, despite
horizontal deceleration forces not being excessive.
Fig 12-5: Typical Engine and APU Systems Diagram

Fire Extinguisher Systems
10. Permanently installed fire extinguisher systems are normally provided to suppress fires in the
engine nacelles or bays, and the APU and heater enclosures of an aircraft. The fire extinguisher system
comprises selection switches sited with the cockpit engine control levers which activate the fire
extinguisher bottles adjacent to the engines or APU. Power is provided from the 28 V DC essential bus,
to ensure that the systems are always live. The systems are designed to deliver predetermined
volumes of extinguishing agent from the fire bottles to designated areas of the appropriate engine
installation. One fire bottle per engine is normally provided, and the systems of multi-engine aircraft are
invariably arranged so that extinguisher agent from each fire bottle can be fed to one or other engine.
Thus a “2 shot” system is provided, allowing the crew two attempts to extinguish a fire. This
arrangement is shown in the schematic layout of a typical system at Fig 12-5. The figure shows the pipe
and electrical interconnections necessary to provide the second shot capability. Shot 1 fires the left
engine bottle to the left engine or the right engine bottle to the right engine. Shot 2 fires the right engine
bottle to the left engine or the left engine bottle to the right engine. Following the phasing out of halons
(because of Ozone layer depletion), the extinguisher substance used in such systems is likely to be an
inert gas or a halocarbon agent which are in use due to their rapid knock down effect, especially in
confined spaces. When released from the system, the gas blankets the fire, purging oxygen away from
it.
11. The fire bottles comprise metal

spheres or cylinders containing the
extinguisher agent pressurized by
nitrogen typically at 40 bar. At this
pressure, the agent is in liquid form. A
range of bottle sizes is available to meet
different fire threats. A typical engine
bottle will hold a charge of 2.5 kg, whilst a
bottle of l kg would be used for smaller
applications such as the APU. Fig 12-6
shows in more detail the 2 shot bottles
used in the above system. The bottle has
two firing heads, each containing an
electrically actuated explosive squib
cartridge. When a head is fired, either by
crew selection or automatically by
operation of the crash switches, the
agent is propelled into the relevant pipe
gallery. It flows under pressure to the fire
area and issues through spray nozzles Fig 12-6: Detail of a 2 Shot Fire Bottle
vaporizing as it does so. The reduction of pressure and vaporization of the agent, as it is sprayed from
the system nozzles, cools the resultant gas thus enhancing its fire fighting capability by cooling the area
of the fire. As a safety device, each bottle is fitted with a safety disc. If excessive pressure builds up in
the bottle, the disc ruptures allowing the agent to vaporize and escape harmlessly overboard. Each
bottle has an integral pressure gauge which is read during each flight servicing.
12. This enables any failure or inadvertent discharge of the system to be detected before further
flight. Although the explosive squib cartridges have a limited effective life and require to be replaced
routinely, very little maintenance is required by extinguisher systems
Portable Fire Extinguishers
13. Hand-held extinguishers fitted to

most aircraft contain either carbon dioxide or
water / glycol mixture. Carbon dioxide is an
asphyxiant gas and must not be employed in
a confined space if the user is denied access
to oxygen or a ventilated area. The gas is
non-conductive and can be used against any
type of fire except, possibly, burning metals.
Water / glycol must not be used on burning
liquids or fire involving electrical equipment.
Fig 12-7 shows a typical example of each
type of fire extinguisher. When the handle of
the CO2 extinguisher is operated, the agent Fig 12-7 Examples of Portable Fire Extinguishers
flows into the horn, vaporizes and can be
directed on to the fire. As with the engine fire bottle, the portable extinguisher container is protected by
a bursting disc which ruptures if the internal pressure exceeds a safe limit. Contents of the water
extinguisher, including a glycol additive to prevent freezing, remain un-pressurized until the extinguisher
is operated. A small cylinder of carbon dioxide propellant is contained in the operating head of the
equipment. Initial action of operating the extinguisher punctures a diaphragm in this cylinder allowing
gas to pressurize the container. Further operation of the extinguisher then propels water on to the fire.
Cabin Protection
14. Research and development continue into means of safeguarding passengers and crew in the
event of a cabin fire on the ground. Although more relevant to civil passenger aircraft, equipment
improvements resulting from such research will be read across to military aircraft in due course. Two
main areas of research are being followed. One is the provision of smoke hoods or masks for
passengers and crew, to prevent smoke inhalation. The other is the provision of water mist inside the
cabin to cool it and to wash away smoke particles. The latter system requires considerable volumes of
compressed air and water to be pumped into the cabin through the aircraft air conditioning ducting. It
requires the provision of emergency ground equipment to achieve this, and its use would therefore be
limited to established airports. Current research in both areas has produced inconclusive results.
347
CHAPTER 13
ICE AND RAIN PROTECTION SYSTEMS
Introduction
1. The operation of military aircraft may necessitate flying in adverse weather conditions.
Provision must therefore be made to safeguard the aircraft against icing, the effects of which may
endanger performance and safety. The areas on an aircraft which are sensitive to ice formation are:
(a) Aerofoil surfaces.
(b) Engine intakes.
(c) Engine internal surfaces.
(d) Rotor blades and propellers.
(e) Windscreens.
(f) Instrumentation probes and vanes.
(g) Control hinges and linkages.
(h) Weapons and weapon carriers.
Principles of Operation
2. Ice protection systems are either active or passive. Active systems operate either by increasing
the temperature of local areas of the aircraft to above freezing point, or by chemically reducing the
freezing point of precipitation impinging upon the aircraft. Passive methods harness the momentum of
the main air stream to separate out precipitation and divert it away. Active systems may be further
categorized as either anti-icing or de-icing. Anti-icing systems prevent the formation of ice in critical
areas whilst de-icing systems work to remove ice which has already formed.
3. Many active systems will perform both functions, and the type used for each particular
application will depend both upon the sensitivity of a specific area to the effects of ice, and upon the
overall need to minimize aircraft weight and aircraft power consumption. Most aircraft utilize more than
one type of system, because of the wide range of requirements. De-icing systems tend to be lighter and
use less energy, but in certain areas the formation of any ice cannot be tolerated and therefore an
anti-icing system must be used. Such an area is the engine air intake. Any build-up of ice would
dramatically reduce its aerodynamic efficiency - thus affecting engine performance - whilst ice
dislodged by a de-icing system would be ingested risking an engine flame out and damage to
compressor blades. Both active and passive systems are used for intake anti-icing.
Fig 13-1: Aircraft Structure Anti-icing System
Ice Protection Systems
4. Thermal (Hot Air) Airframe. The majority of airframe structure anti-icing and de-icing
systems utilize hot air bled from the engine compressors. Fig 13-1 shows the airframe areas of a typical
medium transport, which may be protected by engine bleed air. Such systems sometimes allow air to
be bled from an APU for anti-icing use during critical periods of flight and this configuration allows
anti-icing to be used during an emergency landing, even though maximum power from the main engines
may be essential and therefore no engine
bleed air is available. The hot air is fed
through a system of selector valves, pressure
regulators and mixing valves which reduce
the pressure and temperature of the air to
operating levels. The controlled air is then
ducted through galleries to relevant areas.
Fig 13-2 shows the configuration of a typical
hot air bleed to a wing leading edge. Included
in the illustration is a temperature sensor
installed to activate system temperature
control valves and to provide a warning to the
crew if overheating occurs. Normal cockpit Fig 13-2: Typical Wing Leading Edge
indications include the temperature and Anti-icing System
pressure of air in the system.

349 Ice and Rain Protection Systems
5. Thermal (Hot Air) Engine. Hot air systems are also used for engine anti-icing. The
components which require protection include the inlet guide vanes and first stage compressor stator
blades plus the nose cone and structural support members within the intake. Engine anti-icing systems
are often an integral part of the engine and are independent of the airframe anti-icing system. Fig 13-3
shows such an arrangement.
Fig 13-3: Integral Engine Anti-Icing System
6. Thermal (Electrical). Although hot air systems offer advantages of simplicity and robustness,
electrical heating is widely used for anti-icing and de-icing systems when complex control arrangements
are needed or only small areas require to be heated. Electrical systems usually include heater
elements made from copper-manganese alloy either built up onto a backing material or deposited
(sprayed) onto the backing. Fig 13-4 shows both a built up and a deposited system.
7. The versatility which such manufacturing techniques offer, and the small cross section of the
resultant element makes them ideal for anti-icing of pitot heads, static vents and other probes and
vanes such as stall / angle of attack indicators. Heater mats are also used for de-icing helicopter rotor
blades and fixed wing propellers. The rotor blade application offers particular problems in system
control. A typical rotor blade has a span of 8 to 10 metres and a cord of 0.5 metres. The electrical load
required to heat such large areas continuously exceeds the power available in a helicopter, and
therefore the blades are de-iced by intermittent heating. However, the aerodynamic and dynamic
Fig 13-4: Electrically Heated Mats
Fig 13-5: Typical Fluid Anti-icing System

balance of rotor blades are critical to the controllability and airworthiness of the aircraft. Therefore,
blade heating must be programmed so that the ice build-up and subsequent break down occur
symmetrically, and the control system must protect against asymmetric failure of the heater mats.
Because of the resultant complexity and cost of helicopter rotor blade anti-icing, helicopters operating in
temperate or tropical areas are not normally equipped with blade anti-icing systems.
8. Chemical (Fluid) Diffusion. Chemical fluid systems are limited in use to anti-icing aerofoil
surfaces. The advantage of chemical diffusion methods is that they require only a limited power input.
The disadvantages are that they require replenishment after use and that they are difficult and
expensive to maintain and repair. Their principle of operation is shown in Fig 13-5. When anti-icing or
de-icing is required, de-icing fluid is pumped from a reservoir into porous surfaces which form the
aerofoil leading edges. The fluid diffuses through to the surface of the leading edges where it mixes
with any moisture lowering its freezing point. This prevents the formation of ice and causes existing ice
to break away. The effectiveness of chemical fluid systems is very dependent upon an even distribution
of fluid over the aerofoil leading edge. In turn, distribution is sensitive to the aerofoil angle of attack and
the resultant air-stream pattern along its top surface. Consequently, most aircraft equipped with
chemical fluid systems must be flown within a restricted speed band when the system is in use.
9. Momentum Separation. Momentum separation devices are used for anti-icing the engine
intake systems of helicopters and some ground attack aircraft and also to protect exposed control
system components. They are passive devices, and their principle of operation is to force the air stream
to make sharp changes in direction
and therefore of velocity. During
such changes, the higher momentum
of water particles because of their
higher mass causes them to
separate from the main air stream.
They can then be deflected away
from the intake or other critical area.
Fig 13-6 shows a common form of
momentum separation device called
the air dam or ‘barn door’ and its
principle of operation. Fig 13-6: Principle of the Air Dam Separator
10. Limitations and Effects. Because momentum separator systems interfere with the air intake
ram effect, their use is restricted to helicopters and other slow flying or piston engine aircraft which do
not harness the ram effect. There are many different configurations of separator ranging in complexity
from the most simple arrangement of positioning the engines with their intakes facing downstream so
that the intake air stream must turn through 180 degrees throwing water and debris clear to the
Aerospatiale Polyvalent (multi-purpose) intake shown in Fig 13-7a. Figs 13-7b and c show two more
commonly used systems. Fig 13-7b is an intake shield and 13-7c is the more ingenious wire grill or
Fig 13-7: Momentum Separation Device Configuration
basket. In non-icing conditions, the grill imposes little resistance to the air stream, but in icing conditions,
air passing through the grill speeds up and rapidly cools down causing water particles to freeze and
adhere to the mesh. Thus a shield of ice rapidly accumulates and protects the intake in much the same
way as does the conventional air dam. The design of the grill is such that ample surface area along its
sides will always remain clear of ice to allow sufficient air to enter the engine. The theoretical
disadvantage of the grill intake is that when the aircraft enters warmer air with a frozen grill ice will melt
and break away to enter the engine. In practice however, this problem is not significant. First, the mesh
size of the grill is selected during design to control the size of ice particles which do break away, and
secondly, such large changes in climatic conditions are seldom encountered or can be avoided - within
the normal sortie pattern of a helicopter.
Ground De-Icing
11. Active and passive anti-icing and de-icing systems are designed to become effective as soon
as the aircraft engines are started so that the aircraft can be protected from ice formation during the
critical take-off and subsequent climb-out phases of flight. However, if the aircraft has been parked in
the open in adverse conditions prior to start-up, significant accretions of ice, snow or slush may have
built up on the aircraft flying surfaces. Such deposits must be removed prior to flight, by the ground
crews. After physically removing the majority of such deposits, chemical fluid de-icing is used to
complete the task. This is achieved by the application of de-icing fluids in specially prepared thyxotropic
paste or gel form by the use of ground-based spray equipment. The effect of applying such a de-icing
gel is to melt any ice present and to prevent its reformation until the aircraft is airborne. The principle is
also sometimes used for the airborne anti-icing of unheated rotor blades on helicopters which must fly
for operational reasons in icing conditions. However, the effectiveness of the gel reduces during flight
as it is gradually thrown from the blades by centrifugal forces.
Ice Detection
12. Although significant flight hazards are posed by ice build-up, the fact that most active anti-icing
and de-icing systems consume considerable amounts of power preclude their use except when icing
conditions are actually encountered. Meteorological forecasts go much of the way to alerting crews to
the likelihood of entering icing conditions during a particular flight. However, such forecasts are not
Fig 13-8: Ice Detection Devices

always sufficient, and the crew must therefore remain alert to the need to activate the anti-icing and
de-icing systems at any time. The formation of ice on external visible projections would normally be the
first manifestation of having entered icing conditions. For this reason, the majority of aircraft are
equipped with flood lights aligned to illuminate relevant areas of the airframe which are visible from the
cockpit. Aircraft in which crew visibility is limited are often equipped with illuminated ice accretion
probes as shown in Fig 13-8a. These are positioned to be visible from the cockpit. In addition, most
aircraft are equipped with ice detection devices which either provide a positive alert or automatically
activate the anti-icing and de-icing systems.
13. Many different principles of operation are used in ice detection devices, but all either detect the
actual build-up of ice or the conditions in which a build-up will occur. Three different devices are shown
in Fig 13-8 b to d. The probe shown in Fig 13-8b contains a series of holes positioned in its leading edge
and a separate series in its trailing edge. The detector monitors pressure differential between the two
edges. In icing conditions, holes in the leading edge rapidly become blocked by ice. This causes a
change in the pressure differential. The change is detected, and a cockpit alert is activated. The device
in Fig 13-8c utilizes the change in resonant frequency of a probe which occurs when ice forms on it. The
probe is vibrated at its clean resonant frequency of about 35 kHz. The mass of any ice which forms on
the probe will reduce this resonant frequency, and the detector senses any significant frequency
change and activates the cockpit alarm. The device in Fig 13-8d works on the same principle as a wet
and dry bulb hygrometer, and it comprises two heated bulbs, one exposed to the air stream and the
other shielded, plus a simple outside air temperature (OAT) probe. The detector monitors the
temperature of the bulbs which are heated at a constant rate. When the exposed bulb is in a moist air
stream, it loses its heat to the surrounding air at a greater rate than does the dry shielded bulb. The
resultant temperature imbalance is detected. If the OAT is detected to be within the icing range,
contacts in the probe circuit close, and the alert system is activated.
Windscreen Ice and Rain Protection
14. Although de-icing fluid spray systems or hot air jets were utilized to de-ice the windscreens of
older aircraft, all current aircraft are fitted with electrically heated screens. The heating elements and
associated temperature control and overheat sensors are sandwiched in the glass laminations of the
screen. A thin film of gold is used for the heating element, and it is deposited directly onto glass.
Electrical connectors formed on the edges of the panel interface with the system electrical supply and
temperature controller. The heater systems serve both to de-ice and de-mist the screens.
15. At normal flying speeds, rain falling onto the screens is rapidly dispersed by the air flow.
However, to keep the screen clear during landing or during low speed flight, conventional, high speed
windscreen wipers are fitted to most fixed and rotary wing aircraft. The wipers are electrically activated
by the crew as and when needed.
355
CHAPTER 14
AVIATION FUELS
Piston Engines
1. Aviation piston engines are reciprocating engines, similar to motor car engines and use aviation
gasoline as a fuel. However, as failure of an aircraft engine through fuel problems is potentially
disastrous, safety dictates that aviation gasoline must conform to very rigid specifications.
2. Gasoline is a refined petroleum distillate, although production is considerably augmented by

synthetic processes. The composition of gasoline is suitable for use as a fuel in spark ignition internal
combustion- engines. Aviation gasoline is prepared from various selected grades of gasoline, blended
to give high anti-knock ratings, high stability, a low freezing point and an acceptable volatility. The
aviation gasoline is also known as AVGAS.
3. AVGAS consists of approximately 85% carbon and 15% hydrogen, and the atoms are linked
together in a form which characterizes the type of substances known as hydro-carbons. When mixed
with air and burnt, the hydrogen and carbon combine with the oxygen in the air to form carbon
dioxide .and water vapour. The nitrogen in the air, being an inert gas, does not burn or change
chemically, and serves to regulate combustion. It also helps in maintaining temperature during
combustion.
Gas Turbine
4. Some early gas turbine (jet) engines used aviation gasoline, but Whittle based his jet design on
kerosene (paraffin).
(a) Kerosene. The production of kerosene is limited to that obtained by normal distillation.
It soon became regarded as the most suitable fuel for gas turbines, commending itself on the
grounds of cost, calorific value, burning characteristics and low fire hazard. Kerosene is also
known as AVTUR and depending on type, has a typical boiling range of 150o to 280o C and a
freezing point not higher than -470 C. The US Service equivalent Fuel is JP8.
(b) 'Wide-Cut' Fuels. The quantity of kerosene that can be distilled from a given amount
of crude oil is limited, and this caused initial production limitations. As the jet engine has proved
to be not as fastidious as a piston engine, and capable of operating from any clean burning fuel
a wider distillation range of fuel was developed (known as 'wide-cut' fuels). These distillates are
produced by combining gasoline and kerosene fractions. Wide-cut aviation fuel is also known
as AVTAG and has a wider boiling range than AVTUR and a freezing point below -58oC.
AVTAG is interchangeable with the US designated fuel JP4. Wide-cut fuels present a greater
fire hazard than kerosene, due to lower temperature range of flammability, and higher vapour
pressure. AVTAG has, therefore ceased to be used by most operators and is now primarily
limited to emergency military use in some countries and for use in very cold climatic conditions
(freezing point of AVTAG is lower than AVTUR).
(c) High Flash Kerosene. Naval carrier operations produced a special requirement for
avoidance of vapour build-up within confined spaces. Higher density kerosene with a high flash
point (61°C as compared to 38o for normal density kerosene) was found suitable and is
specified as AVCAT.
(d) Other Special Requirements. Special performance aircraft have created

requirements for matching variants of jet fuel. Such cases include high speed/very high-altitude
aircraft and fuels needing exceptional thermal stability.
Density
5. The density of a substance is defined as its mass per unit volume. It is a measure of the
concentration of matter in a material. It is measured in kilograms per cubic meter (SI units) and can be
written in expression form as:
Mass of a substance
Density = (14.1)
Volume occupied by the substance
6. Fuel oils, which are mixtures, will have varying densities depending on how much of each
constituent is present in the mixture. Some typical densities are:
Water: 1,000 kg per cubic meter

Paraffin: 800 kg per cubic meter
Specific Gravity
7. The specific gravity (SG) or relative density of a substance is the ratio of its density compared to
that of water. So, for a given substance, the expression used is:
Density of substance
Specific Gravity = (14.2)
Density of water
Using the figures in paragraph 6, the specific gravity of paraffin is:
800 kg per m3
SG paraffin = = 0.80
1,000 kg per m3
357 Aviation Fuels
8. A knowledge of density and specific gravity is useful as it relates a given volume of a substance
to its mass without actually having to weigh it. For example, AVGAS has a typical density of 720 kg/m3
(this equals to a specific gravity of 0.72), and if a tank holds 2 cubic meter, the mass of fuel will be 720 x
2 kg = 1,440 kg.
9. By knowing the specific gravity of the fuel and one unit of volume, the mass can be worked out
by use of the conversion tables. For example, if the aircraft was refuelled with 3,000 litres of an AVTUR
with 0.80 SG, this would equate to 2,400 kg or 5,300 lbs of fuel.
10. The effect of varying SG is that:
(a) To refuel to a specified mass fuel load will require a greater volume of a low SG fuel
than of a high SG Fuel.
(b) If replenishing fuel tanks to full, then the loading of a low SG fuel will result in a lower
mass of fuel and therefore a reduced cruise range for that flight than if a high SG fuel had been
used.
Temperature and Specific Gravity
11. SG varies inversely with temperature, but as fuel is loaded by mass this wil only be significant if
full tanks are required. Should an aircraft be fuelled to “tanks full” with cold fuel and then be allowed to
stand in high temperatures, the expansion of the fuel will result in fuel being spilled overboard, through
a venting system.
Fuel Icing
12. Freezing Point. As jet fuel cools, the process will reach a stage where it will initially generate
a growth of wax crystals. Continued cooling will take the fuel to a frozen solid state. Should the
temperature rise the process will reverse. The freezing point of a jet fuel relates to one point in this
waxing process specially that temperature, in the ‘warming up’ process, at which waxy precipitates
disappear.
13. A certain amount of water is present in all fuels. The amount varies depending on the efficiency
of the manufacturer’s quality control and the preventive and removal measures taken during
transportation and storage. Some refuelling procedures require that, post-refuelling, fuel is allowed to
stand in order that water droplets can settle. The water gathers in the base of the tank, and can be
drained off through a water drain valve.
14. As fuel cools, the amount of dissolved water the fuel can hold is reduced. Water droplets are
then formed, and as the temperature is further decreased, these form ice crystals which can block fuel
system components.
15. In-large aircraft, the threat-of ice-build-up on fuel filters can be solved by using fuel heaters. In
most military aircraft, and certain civil aircraft the icing threat is solved by using di-ethylene glycol
monomethyl ether (di-EGME), a Fuel System Icing Inhibitor (FSII).
Fuel and Fire Hazards
16. Sources of Fire Hazard. With aviation fuels, there are three main sources of fire hazard.
These arise from:
(a) Fuel spillage with subsequent ignition of vapour from a spark, etc.
(b) Fuel spillage on to a hot surface causing self-ignition.
(c) The existence of flammable or explosive mixtures in the aircraft tanks.
17. Volatility. The first hazard depends on the volatility of the fuel. The lower the flash point, the
greater are the chances of fire through this cause. It is more difficult to ignite kerosene than to ignite
gasoline in this way.
18. Spontaneous Ignition. The second hazard depends on the spontaneous ignition
temperature of the fuel. In this respect gasoline has a higher spontaneous ignition temperature than
kerosene, but if a fire does occur the rate of spread is much slower in kerosene owing to its lower
volatility.
19. Fuel –Vapour / Air Mix. The third hazard depends upon the temperature and pressure in the
tank and the volatility of the fuel. Therefore, at any given pressure (or altitude), for any fuel there are
definite temperature limits within which a flammable fuel vapour/air mixture will exist. If the temperature
falls below the lower limit the mixture will be too weak to burn, while if the temperature rises above the
upper limit the mixture is too rich to burn. The limits vary with the chemical constitution of the fuel, and
reduce with altitude, so a general rule of thumb can not be given. In terms of combination of fuel/air
vapour mixture, a half-empty fuel tank presents a greater hazard than a full tank.
PISTON ENGINE FUELS
Properties of Aviation Gasoline
20. The five most significant properties of gasoline which influence engine design are as follows:
(a) Anti-knock value.

(b) Volatility.
(c) Vapour locking tendency.
359 Aviation Fuels
(d) Stability.
(e) Solvent and corrosion properties.
Anti-knock Value
21. The anti-knock value of a fuel is defined as the resistance the fuel has to detonation. It is
essentially a comparative and not an absolute figure, as the engine conditions under which the
detonation takes place are very important. A fuel which has a good anti-knock value is one that has
good detonation-resisting qualities compared with several other fuels being used under exactly the
same operating conditions.
22. Detonation. After ignition the flame normally travels smoothly through the combustion
chamber until the charge is completely burnt. The rate of burning may be as high as 18 meters per
second, which may seem very fast in view of the size of the cylinder but, nevertheless, it is steady.
Combustion is comparatively quiet, with a regular pressure rise and a steady push on the piston. When
detonation occurs, combustion begins normally, but at an early stage the temperature of the unburned
part of the mixture is raised so high that it ignites spontaneously, with a flame velocity in the
neighbourhood of 300 meters per second. The cylinder walls and piston receive a hammer-like blow
(knocking) giving rise to the characteristic pinking noise, familiar to motorists, though not audible in the
air because of propeller and other noises. The rate of pressure rise is too great to be accommodated by
movement of the piston, so that much of the chemical energy released is wasted as heat, instead of
being transformed into mechanical power.
23. Knock Rating of Fuels. Depending on their composition, fuels differ considerably in their
resistance to detonation. Highly rated fuels allow:
(a) An increase in compression ratio and hence in thermal efficiency, with a resultant gain
in economy and at the same time slightly increased power.
(b) An increase in permissible manifold air pressure (MAP) and therefore increased power.
(The power output of an engine is almost directly proportional to the weight of air consumed in a
given time and a higher MAP increases this weight).
24. It should be understood that these improvements apply only if the engine is designed or
modified to take advantage of the higher grade fuel. Such a fuel used in a low performance engine will
not give more power or greater economy but may on the other hand, cause fouling of the cylinders and
eventual mechanical failure.
25. Anti-Knock Additives. The anti-knock value of fuels can be raised by the addition of
anti-knock substances. The best known and most powerful of these is tetra-ethyl lead (TEL). This is
added to the fuel together with small amounts of an inhibitor (against gum formation) and ethylene
dibromide. The use of ethylene dibromide prevents the formation of deposits of lead oxide on the
combustion chamber, valves and sparking plugs. The presence of lead compounds promotes pitting of
valves, seats and sparking plug electrodes, for this reason the maximum amount of TEL that can be
added to a grade of fuel is limited by specification.
Octane Numbering and Fuel Grading
26. Before the advent of the more highly supercharged engines, the resistance to detonation of an
aviation fuel was expressed by its octane number. This rating system was based on the widely different
knock resistance of two pure spirits, iso-octane (excellent) and heptane (very poor). By degrading
iso-octane with heptane until the blend detonated in a variable compression engine under the same
standard conditions as the fuel under test, it was possible to classify that fuel by a number representing
the percentage of iso-octane in the test blend. Thus 87-octane fuel corresponded to a mixture of 87%
iso-octane and 13% heptane.
27. This system, however, took no account of the increase in knock resistance at high mixture
strengths, and for a very good reason. Although engines are supplied with a much weaker mixture
under cruising conditions than when developing high power outputs those using lower grade fuels do
not have to cope with markedly increased combustion pressure at a maximum output. Consequently,
the margin between the operating power of such engines, and the power as limited by detonation, is
smallest at weak mixtures and increases as the mixture is richened. The octane system, therefore,
specified weak mixture knock ratings only. With highly supercharged engines, however, combustion
pressures at maximum output are well above those at cruising powers, and it has become necessary to
specify knock rating for both rich and weak mixture conditions. Furthermore, as fuel with knock ratings
superior to iso-octane are now available, the rating of these fuels has become more involved,
necessitating the addition of tetra-ethyl lead to the reference fuels.
28. Only one grade of piston engine fuel is presently available for general distribution, AVGAS
100LL. (LL stands for ‘Low Lead’ – an unleaded version of AVGAS with anti-knock properties suitable
for modern piston-engine aircraft is yet to be introduced) Other grades of piston engine fuel may be
encountered in some locations. These are categorized by grade names consisting of two numbers, the
first being the knock rating for weak mixture conditions, and the second for rich mixture, e.g. Grade
100/130. Whilst weak mixture ratings are still measured in the same way as octane numbers, rich
mixture ratings are related to the maximum MAP that can be applied without detonation.
Volatility
29. A volatile liquid is one capable of readily changing from liquid to the vapour state by the
application of heat or by contact with a gas into which it can evaporate. The following properties of a
fuel are related to volatility: efficiency of distribution, oil dilution, ease of starting, carburettor icing and
361 Aviation Fuels
vapour locking tendencies. Some of these factors depend on the presence of low boiling and others on
the presence of high boiling fractions. Thus fuel volatility cannot be expressed as a single figure.
Storage Stability
30. The property of the fuel which if of interest here, is its tendency to form ‘gummy’ products in
storage. The term ‘gum’ here is applied to a colourless or yellowish sticky deposit which is sometimes
left as a residue when gasoline is completely evaporated. It may cause deposits in the intake manifold
and cause sticking of the inlet valves and any moving parts in the fuel system. Aviation gasoline fresh
from the refinery usually contains negligible amounts of gum, but when the gasoline is stored gum may
form. The degree of gum formation depends on the nature of the gasoline and the conditions of storage.
High atmospheric temperatures and exposure to air hasten gum formation. The formation is more rapid
in small containers like tins and drums than in large storage tanks owing to the greater ratio of surface
area to volume in the former case. Exposure to light may also cause gum to form more rapidly. Once
gum formation starts it proceeds quickly. Poor storage stability may also manifest itself with the
precipitation of white compounds in the fuel. This is not gum, but a lead compound from TEL. So long
as this is not excessive it is not in itself dangerous, but it usually indicates that something else is wrong.
Therefore when lead precipitation takes place the fuel should be viewed with suspicion, and none used
until it has been tested.
Solvent Properties
31. Unsaturated hydrocarbons are powerful solvents of rubber and some rubber-like compounds.
They also cause swelling of rubber, with resultant blocking of fuel lines, etc. Fuel pipes and systems
must therefore be manufactured from materials that can resist the solvent properties of gasoline.
Corrosive Properties
32. A small amount of sulphur is present in all aviation gasoline, and can cause corrosion as
described in para 42b.
GAS TURBINE FUELS
Properties of Aviation Gas Turbine Fuels
33. A gas turbine fuel should have the following properties:
(a) Ease of flow under all operating conditions.

(b) Quick starting of the engine.
(c) Complete combustion under all conditions.
(d) A high calorific value.
(e) Non-corrosive.
(f) The by-products of combustion should have no harmful effect on the flame
tubes, turbine blades, etc.
(g) Minimum fire hazards.
(h) Provide lubrication of the moving parts of the fuel system.
Ease of Flow
34. The ease of flow of a fuel is mainly a question of viscosity, but impurities such as ice, dust, wax
etc. may cause blockage in filters and in the fuel system generally.
35. Most liquid petroleum fuels dissolve small quantities of water and if the temperature of the fuel
is reduced enough, water or ice crystals are deposited from the fuel. Adequate filtration is therefore
necessary in the fuel system. The filters may have to be heated, or a fuel de-icing system fitted, to
prevent ice crystals blocking the filters. Solids may also be deposited from the fuel itself due to the
solidification of waxes or other high molecular weight hydrocarbons.
Ease of Starting
36. The speed and ease of starting of gas turbines depends on the ease of ignition of an atomized
spray of fuel, assuming that the turbine is turned at the required speed. This ease of ignition depends
on the quality of the fuel in two ways:
(a) The volatility of the fuel at starting temperatures.
(b) The degree of atomization, which depends on the viscosity of the fuel as well as the
design of the atomizer.
37. The viscosity of fuel is important because of its effect on the pattern of the liquid spray from the
burner orifice, and because it has an important effect on the starting process. Since the engine should
be capable of starting readily under all conditions of service, the atomized spray of fuel must be readily
ignitable at low temperatures. Ease of starting also depends on volatility, but in practice the viscosity is
found to be the more critical requirement. In general, the lower the viscosity, and the higher the volatility,
the easier it is to achieve efficient atomization
Complete Combustion
38. The exact proportion of air to fuel required for complete combustion is called the theoretical
mixture and is expressed by weight. There are only small differences in ignition limits for hydrocarbons,
the rich limit in fuels of the kerosene range being 5:1 air / fuel ratio by weight, and the weak limit about
25:1 by weight.
363 Aviation Fuels
39. Flammable air / fuel ratios each have a characteristic rate of travel for the flame which depends
on the temperature, pressure, and the shape of the combustion chamber. Flame speeds of
hydrocarbon fuels are very low, and range from 0.3 to 0.6 m/sec under laminar flow conditions. These
low values necessitate the provision of a region of low air velocity within the flame tube, in which a
stable flame and continuous burning are ensured.
40. Flame temperature does not appear to be directly influenced by the type of fuel, except in a
secondary manner as a result of carbon formation, or of poor atomization resulting from a localized
over-rich mixture. The maximum flame temperature for hydrocarbon fuels is roughly 2000° C. This
temperature occurs at mixture strength slightly richer than the theoretical ratio, owing to dissociation of
the molecular products of combustion which occurs at this mixture. Dissociation occurs above about
1400° C, and reduces the energy available for temperature rise.
41. Flame extinction in normal flight is rare in an otherwise serviceable engine. Most extinctions are
the result of engine mishandling or through excursions outside the permitted flight envelope. The type of
fuel used is of relatively minor importance. However, the wide cut fuels (AVTAG) are more resistant to
extinction than the kerosene (AVTUR) and engines are easier to relight using AVTAG. This is due to
the higher vapour pressure of AVTAG.
Calorific Value
42. The calorific value is a measure of the heat potential of a fuel. It is of great importance in the
choice of fuel, because the primary purpose of the combustion system is to provide the maximum
amount of heat with the minimum expenditure of fuel. The calorific value of liquid fuels is usually
expressed in mega joules (MJ) per kg. When considering calorific value, it should be noted that there
are two values which can be quoted for every fuel, the gross value and the net value. The gross value
includes the latent heat of vaporization, whilst the net value excludes it. The net value is the quantity
generally used. The calorific value of petroleum fuels is related to their specific gravity. With increasing
specific gravity (heavier density) there is an increase in calorific value per litre but a reduction in calorific
value per kilogram. Thus for a given volume of fuel, kerosene gives an increased aircraft range when
compared with gasoline, but weighs more. If the limiting factor is the volume of the fuel tank capacity, a
high calorific value by volume is more important.
Corrosive Properties
43. The tendency of a turbine fuel to corrode the aircraft’s fuel system is affected by:
(a) Water. Salt within water can cause corrosion. Dissolved water in fuel is which causes
corrosion is the dissolved water described in para 15 and 36. Salts can lead to corrosion of the
fuel system, which is particularly important with regard to the sticking of sliding parts, especially
those with small clearances and only small or occasional movement.
(b) Sulphur Compounds. Every effort is made to keep the sulphur content as low as
possible in aviation fuels. Unfortunately, removal of the sulphur involves increased refining
costs and decreased supplies and so some sulphur is therefore permitted. Sulphur can cause
corrosion in two ways:
(i) Corrosive Sulphur. The corrosive sulphur consists of sulphur compounds

such as mercaptans, sulphides, free sulphur, etc, which corrode the parts of the fuel
system, e.g. tanks, fuel lines, pumps, etc. It is detected in the laboratory by the
corrosive effect of the fuel.
(ii) Combustion of Sulphur. Sulphur and sulphur compounds when burnt in air
form sulphur dioxide and this, with water, forms sulphuric acid. The total sulphur is
estimated in the laboratory by burning the fuel, and thence by chemical analysis of the
products of combustion.
Effects of Combustion By-Products
44. Carbon deposition in the combustion system indicates imperfect combustion, and may lead to:
(a) A lowering of the surface temperature on which it is deposited, resulting in buckled

flame tubes because of the thermal stresses set up by the temperature gradients.
(b) Damage to turbine blades caused by lumps of carbon breaking off and striking them.
(c) Disruption of the airflow through the turbine creating turbulence, back pressure, and
possible choking of the turbine, resulting in loss of efficiency.
45. It appears that the carbon deposition depends on the design of the combustion chamber and
the aromatic content of the fuel (aromatics are a series of hydrocarbons based on the benzene ring).
The higher the aromatic content the greater the carbon deposits.
Fire Hazards
46. Fire hazards are covered in paras 16–19. As a general rule, kerosene needs to be at a
relatively high temperature to burn, and thus in cold climates is regarded as safer than gasoline, which
has a lower temperature range of flammability.
365 Aviation Fuels
Vapour Pressure
47. The vapour pressure of a liquid is a measure of its tendency to evaporate. The saturated
vapour pressure (SVP) of a liquid (i.e. the pressure exerted by vapour in contact with the surface of the
liquid) increases with increasing temperature. When the SVP equals the pressure acting on the surface
of the liquid, the liquid boils. Thus, the boiling point of a liquid depends on a combination of SVP, the
pressure acting on its surface and its temperature.
48. The SVP of aviation gasoline (AVGAS) at a temperature of +20° C is about 27 kPa absolute. It
follows, therefore, that this fuel boils at +20° C when the atmospheric pressure falls to 27 kPa. This
occurs at an altitude of about 35,000 feet (10668 metres). If the temperature of the fuel is higher, it will
boil at a lower altitude. All liquids have a vapour pressure although in some it is extremely small. These
small vapour pressures, however, become important at high altitudes.
49. Reid Vapour Pressure. The standard adopted for the measurement of vapour pressure of
fuels is the Reid Vapour Pressure (RVP). This is the absolute pressure determined in a special
apparatus when the liquid is at a temperature of 37.8° C. The maximum RVP allowed in the
specification of AVGAS is 48 kPa. This is designated a high vapour pressure fuel. AVTUR has an RVP
of 0.689 kPa and as such is a low vapour pressure fuel, as are all fuels with an RVP of 14 kPa or less.
Fuel Boiling and Evaporation Losses
50. At high rates of climb, fuel boiling and evaporation is a problem which is not easily overcome. A
low rate of climb permits the fuel in the tanks to cool and thus reduce its vapour pressure as the
atmospheric pressure falls off. However, the rate of climb of many aircraft is so high that the fuel retains
at its ground temperatures, so that on reaching a certain altitude the fuel begins to boil. In practice this
boiling has proved to be so violent that the loss is not confined to vapour alone. Layers of bubbles form
and are swept through the tank vents with the vapour stream. This loss is analogous to a saucepan
boiling over and is sometimes referred to as slugging.
51. Fuel losses as high as 20% of the tank contents have been recorded through boiling and
evaporation. The amount of fuel lost from evaporation depends on several factors. These are:
(a) Vapour pressure of the fuel.

(b) Fuel temperature on take-off.
(c) Rate of climb.
(d) Final altitude of the aircraft.
Methods of Reducing or Eliminating Fuel Evaporation Losses
52. Possible methods of reducing or eliminating losses by evaporation are:

(a) Reduction of the rate of climb.

(b) Ground cooling of the fuel.
(c) Flight cooling of the fuel.
(d) Recovery of liquid fuel and vapour in flight.
(e) Redesign of the fuel tank vent system.
(f) Pressurization of the fuel tanks.
(g) Using a fuel of low RVP.
53. Reduction of the Rate of Climb. Reducing the rate of climb imposes an unacceptable
restriction on the aircraft and does not solve the problem of evaporation loss. This method is, therefore,
not used.
54. Ground Cooling of the Fuel. This is not considered a practical solution, but in hot climates
every effort should be made to shade refuelling vehicles and the tanks of parked aircraft.
55. Flight Cooling of the Fuel. The use of a heat exchanger, through which the fuel is circulated
to reduce the temperature sufficiently to prevent boiling, is possible. High rates of climb, however,
would not allow enough time to cool the fuel without the aid of heavy or bulky equipment. At a high TAS,
the rise in airframe temperature due to skin friction increases the difficulty of using this method. On
small high-speed aircraft the weight and bulk of the coolers becomes prohibitive.
56. Recovery of Liquid Fuel in Flight. This method would probably entail bulky equipment and
therefore is unacceptable. Another method would be to convey the vapour to the engines and burn it to
produce thrust, but the complications of so doing would entail severe problems.
57. Redesign of the Fuel Tank Vent System. The loss of liquid fuel could be largely eliminated
by redesigning the vents, but the evaporation losses would remain. However, improved venting
systems may well provide a more complete solution to the problem.
58. Pressurization of the Fuel Tanks. There are two ways in which fuel tanks can be
pressurized:
(a) Complete Pressurization. Keeping the absolute pressure in the tanks greater than
the vapour pressure at the maximum fuel temperature likely to be encountered eliminates all
losses. However, with gasoline type fuels a pressure of about 55 kPa absolute would have to
be maintained at altitude and the tank would be subjected to a pressure differential of 45 kPa at
50,000 feet. The disadvantage is that this would involve stronger and heavier tanks, and a
strengthened structure to hold the tanks.
367 Aviation Fuels
(b) Partial Pressurization. This prevents all liquid loss and reduces the evaporation
losses. It involves strengthening the tanks and structure, and the fitting of relief valves.
59. Use of a Fuel of Low RVP. The disadvantage of kerosene lies chiefly in its limitations at low
temperatures. At temperatures below -47o C the waxes in the fuel begin to crystallize and may lead to
blockage of filters unless remedial measures such as fuel heating are introduced. Starting difficulties
under very cold conditions would also have to be solved. The ability to change to kerosene in hot
climates, would overcome many problems, but the limitations of this scheme are obvious. The ideal
solution would be to adopt one fuel that is suitable for universal use. Blended fuels of low RVP prove to
be the most suitable, as with them the evaporation losses are largely eliminated and the poor
low-temperature characteristics of kerosene are improved upon.
Fuel System Icing and Fungal Growth
60. All turbine-powered aircraft should use fuel containing Fuel System Icing Inhibitor (FSII) to
inhibit fuel system icing and fungal growth. If fuel containing FSII is not available, operation is permitted
for a limited period, provided that:
(a) The maximum period on fuel not containing FSII does not exceed 14 days, and is
followed by an equivalent period on inhibited fuel.
(b) The risk of ice formation is acceptable to the operational commander.
(c) Uplifts of non-inhibited fuel are recorded in the aircraft F700.
61. Present in all turbine fuels is a microbiological fungus called Cladasporium Resinae. This
fungus can grow rapidly in the presence of water and warmth, forming long green filaments which can
block fuel system components. The waste products of the fungus can be corrosive, especially to the
fuel tank sealing components. The inclusion of FSII in the fuel suppresses fungal growth.
369
CHAPTER 15
COMPOSITE MATERIALS
Introduction
1. A composite is a material which is made up of two or more distinct materials. Such

combinations exhibit the best properties of the individual materials and include other qualities that
none of the individual materials possesses. A familiar composite is concrete, which is basically made
up of sand and cement. Composites do occur in nature. e.g. in tree trunks, spider webs, and mollusk
shells. A tree is a good example of a natural composite, consisting of cellulose (the fibrous material)
and lignin (a natural polymer) forming the woody cell walls and the cementing (reinforcing) material
between them. Artificial composite materials consist of fibres (Reinforcement material) held together
by resins (Matrix material)
2. Composites are chosen for the manufacture of particular articles or components, primarily
because of weight saving for their relative stiffness and strength. As an example we can compare a
carbon fibre reinforced composite with its steel counterpart. The carbon fibre composite can be five
times stronger than 1020 grade steel while having only one fifth the weight. Aluminium (6061 grade)
is much nearer in weight to carbon fibre composite (though still somewhat heavier), but the composite
can have twice the modulus and up to seven times the strength.
Manufacture of Composites
3. The principle behind composite manufacture is the mixing of the reinforcement with the matrix
and the solidifying of the matrix using heat and pressure into the final shape desired. With the variety
of types of composites there are an associated variety of manufacturing methods. Some of them are
Wet lay up, Spraying, Resin transfer, Injection moulding, and Hot press moulding,
4. Many metal articles or components can instead be made from composites, but there are
important differences which mean that direct substitution should be made with care. Most
engineering materials are essentially isotropic, that is, they have the same properties such as strength
and modulus, in any direction. On the other hand, most composites will have very different properties
in different directions. This is because, although the matrix material is isotropic, the reinforcement is
not. Carbon fibres may be up to 100 times stronger under tension than they are in shear, and the
stiffness may differ in the two directions by similar ratios. The properties of the composite will reflect
the properties of the reinforcement, so that it can have greatly different properties in different
directions. Manufactured articles rarely require to be equally strong in all directions, and composites
can achieve this by particular arrangements of the reinforcement.
Types of Composite Materials
5. Composites are classified according to the type of Matrix material (usually referred to as the
Resin), and the type of Reinforcement material. For example - Polymer composites are plastics within
which there are embedded fibres or particles. The plastic is known as the matrix, and the fibres or
particles, dispersed within it, are known as the reinforcement. Because the size of the reinforcement
particles, or the type and length of the fibres, can be varied and because the directions in which they
can be placed within the matrix can also be varied, a very wide variety of properties can be achieved.
6. The matrix material (Resin) is the binder which transfers the load to the reinforcement. In
other words all loads are taken by the fibre and not the resin. The Matrix material is also the first line
of defence from environmental attack. It needs to be selected so that it can withstand the service
temperature (high or low) conditions, and it should not be affected by fluids (gases or liquids) with
which the composite will come into contact. There are four types of composites:
(a) Fibre Reinforced Composites. These are composed of fibers embedded in matrix.
Fibres are small in diameter and are pushed axially. They have very good tensile properties.
The Fibers must be supported to keep individual fibres from bending and buckling. The
combination of fibrous reinforcement in a matrix is the most common form of composite.
(b) Laminar composites. These are composed of layers of materials held together by
matrix. Sandwich structures fall under this category.
(c) Particulate Composites. These are composites in which metallic or non-metallic

materials are proportionately combined to form a useful material. They are composed of
particles distributed or embedded in matrix body. The particles may be powdered or flakes.
Concrete and wood particle boards are examples of this category.
(d) Hybrid Composites. Composites may be reinforced by suitable materials to form

hybrid composites.
Matrix Materials (Resins)
7. Many thermosetting and thermoplastic resins can be used for making composites but the
selection of the material depends on cost, manufacturing capability, mechanical and environmental
performance and availability. For structural composites the most common thermoset resins include
Polyester resins, Vinylester resins and Epoxy resins. For high temperature or environmental
resistance the most common thermosetting and thermoplastic resins include Polyimide resins,
Phenolic resins and Cyanate ester resins. For commodity or high volume products the resins are
often thermoplastic and include Nylon, Polycarbonate, Polyphenylene sulphide, and Polyamideimide.
371 Composite Materials
Reinforcing Fibres
8. The four reinforcing fibres in current extensive use for aircraft structures are:
(a) Glass fibres.

(b) Carbon (graphite) fibres.
(c) Aramid fibres (Aromatic Polyamid).
(d) Boron fibres.
9. The list is roughly in order of cost, with the cheapest materials at the top. The most common
fibres used are glass and carbon. Glass fibres are cheaper than carbon and come in various forms
that make them suitable for many applications however it has the lowest modulus. Carbon fibre is the
reinforcement material of choice for "advanced" composites. A major advantage of carbon fibres is
their higher fatigue resistance compared to glass or Aramid. Unlike these last two materials, carbon
fibres do not suffer from stress rupture. Aramid fibres have the highest strength to weight ratio and
therefore are more typically used in high value product areas where high-energy absorption is needed.
Aramid is broadly similar in tensile strength to glass fibre, but can have up to twice the modulus.
Toughness is an advantage where energy absorption is required, but, compared to carbon; it is lower
in compressive strength and has poorer adhesion to the matrix. It has high modulus but it is also
susceptible to the uptake of moisture. Aramid fibre properties depend on the structure used and can
be tailored for high toughness or high modulus. Boron fibres predate carbon fibres as high-modulus
reinforcement materials. However, boron has largely been replaced by carbon because of the high
cost of the former.
Fibre Orientation or Directionality
10. One of the strengths of composite materials is their ability to be tailored by orientating the
fibres to resist certain loads. Different methods for achieving this directionality are described below.
11. Unidirectional Laminates. The properties of unidirectional composites are quite different
from isotropic materials. Unidirectional materials are highly anisotropic and have exceptional
properties in the fibre direction and mediocre properties perpendicular to the fibre directions. There
are very few situations where composites are used purely in a unidirectional configuration. In most
applications there will be some form of loading away from the direction of the fibres. In this situation, it
is only the resin that resists this load which has no reinforcement. To resist these loads composite
structures are made by multi directional laminates.
12. Multi Directional Laminates. The bonding of individual unidirectional lamina or plies
together is used to form laminates. The laminae are oriented in a multi-directional manner to give the
effective properties required for the application. The order and direction of the lay-up can be
determined using lamination theory. There are certain rules that should be adhered to. If the lay-up
order is not symmetric about the thickness centre line, then the laminate will bend and/or twist when a
load is applied. If the directions of the lay up through the thickness has different number of
orientations in the plane then modulus and strength of the laminate will be different in the x and y
directions. A quasi-isotropic laminate is one that approximates isotropy by orientation of plies in
several or more directions in-plane.
13. Woven Composites. The main advantages of woven

materials are that they can be draped over or in non-flat moulds
while retaining some of their orthogonal properties. The weave
also acts to enhance their impact performance. The
reinforcements generally fall into the category of a cloth or
woven roving (Fig 15-1). There are several different patterns for
the weave including plain weave, which is the most highly
interlaced, a basket weave and satin weaves. The last has a
minimum of interlacing. Fig 15-1: Woven Composites
14. Striched Composites. A stitched reinforcement is constructed using a combination of

unidirectional reinforcements that are stitched together with a non-structural synthetic such as
polyester. Stitched fabrics offer greater strength and stiffness per unit thickness as compared to
standard woven rovings. While the stitching does have the effect of crimping the fibres, the crimping is
far less that with a woven roving. Additionally, reinforcements can be oriented along any combination
of axes.
15. Braiding. Braiding is a textile process that provides a high level of conformability, torsional
stability and damage resistance because of its very nature. Braiding differs from woven and stitched
fabrics in the method of interlacing. The interlacing provides a high degree of structural integrity
particularly in the hoop direction. For this reason braiding is used in many structural applications.
16. Random and Short Fibre Composites. Random reinforcement methods produce plies
with nominally equal properties in all directions (except through the thickness). This method of
manufacture is a very economical way to build up thickness, especially with complex moulds.
Mechanical properties such as strength and modulus are less than other reinforcement methods, but
are significantly higher than un-reinforced resin.
APPLICATIONS OF ALL ELEMENTS IN AEROSPACE SECTOR
Aircraft Applications
17. In aerospace the demands placed upon materials can be greater than in other areas, often
requiring a combination of light weight, high strength, high stiffness and good fatigue resistance.
Military aircraft were the first to use composites in significant quantities. The first applications were in
radomes and then in secondary structures and internal components. The modulus of glass, however,
is low compared with that of metals and it was not until the advent of boron and carbon
reinforcements that significant interest in terms of primary structures developed. Examples where
composites have been used extensively include: Airbus Industries A 320, Harrier AV-8B, European
Fighter Aircraft (EFA), Aircraft propellers, Helicopter Airframes, Helicopter rotor blades and Helicopter
rotor hubs. Table 15-1 depicts the use of composites in aerospace sector.
Item Typical Fibre Typical Resin

Fuselage –Military Carbon Epoxy
Fuselage – General Aviation Glass fabric Epoxy
Wings & Control surfaces Carbon Epoxy
Bulkheads & Floors Glass -
Engine Cowls – Carbon Fibre Reinforced Carbon Epoxy Bismaleimide

Fairings (Fuselage/Fin/Flap
Glass / Carbon Epoxy
Track/Landing gear doors, etc.)
Radomes Aramid / Glass Polyester
Rotor Blades Carbon / Glass Epoxy
Rotor Hubs Glass / Carbon Epoxy
Table 15-1: Application in Aerospace Sector
18. Aircraft Floor. Composite sandwich panels are commonly used for the construction of
aircraft flooring. The sandwich structure used typically consists of thin, high strength skins adhered to
a low density core. This construction leads to a high bending stiffness and strength at low overall
mass. The skins operate in nearly pure tension or compression and the bulk of the shear loads are
carried by the core. Core materials used include balsa wood and aramid paper (Nomex) or aluminium
honeycombs. The combination selected will depend on the application, for example aluminium alloy
skins on an aluminium honeycomb core might be used for the floor on a military transport aircraft
where concentrated loadings can be expected, whereas GFRP skins over an aramid honeycomb core
might be specified for a passenger aircraft floor. Safety regulations require the use of low smoke and
toxicity material in the construction of aircraft interior components.
19. The floor is designed to meet a number of requirements including stiffness and strength.
Stiffness is important for a number of reasons including passenger comfort when standing and
walking. The strength requirements are defined according to the worst of normal operating and crash
case loads. Consideration of strength is most important for concentrated loads applied to the skins
orto fittings. Specially designed fittings are normally used for introduction of high loads, e.g. a seat
attachment. The design requirements will typically vary according to the location (eg. a high traffic
entry area or under-seat region) and function of the panel but will usually form part of the aircraft
manufacturer's specification.
20. Aircraft Propellers. The

excellent fatigue performance of
composites is used to good effect with
propeller designs. Using a combination of
a unidirectional carbon spar and glass
cloth reinforced skins at (±45°) for
torsional resistance produces a very
effective component. Complex
aerodynamic shapes are more easily
manufactured with composites than with
metal and this provides added design
flexibility (Fig 15-2).
Fig 15-2: Aircraft Propeller
21. Aircraft Wing Skin. Aircraft wings are usually of semi-monocoque construction, with
relatively thin skins supported by spars, stiffeners and ribs. From the point of view of analysis, the skin
can often be treated as a plate supported at its boundaries. The primary loadings acting on the skin
are shear (arising from overall torsion of the wing) and tension/compression (arising from overall
bending of the wing).
22. In many cases the wing skins will be sized so that the wing structure meets torsional stiffness
criteria imposed by aeroelastic considerations. In order to maximise torsional stiffness, laminates with
a construction consisting predominantly of layers oriented at ±45° (relative to the axis of the section's
shear centre) are used. The torsional stiffness requirement usually dominates to the extent that the
wing torsional, i.e. plate shear, strength requirements are easily met.
23. Sandwich construction is favoured for some designs, particularly those approaching a pure
monocoque construction where the skins make a greater contribution to carrying the wing bending
loads. The use of sandwich construction allows a significant increase in skin buckling loads thereby
contributing to greater structural efficiency at increased loads. The designer will normally not allow for
buckling below ultimate load in this type of structure. While it may offer some benefits structurally, the
comparatively thinner facing material in sandwich construction makes for greater susceptibility to
impact damage.
24. European Fighter Aircraft (EFA). For the European Fighter Aircraft (EFA) currently under
development, the projected target for composites utilization is 35% involving the main wing, the
forward fuselage, and the fin and rudder. For any structure or subassembly it is likely that a
combination of materials will be used, each applied so that its individual set of properties can be used
to best advantage.
25. Harrier AV-8B. The Harrier AV-8B (Fig 15-3) is an example of a military aircraft using
advanced composite materials. The primary structural applications are the wing torque box and
control surfaces, horizontal tail and forward fuselage. Secondary structures are the gun and
ammunition packs, strakes, ventral fin, rudder, engine bay doors, nose cones and fairings. Twenty
five per cent of the airframe weight is fabricated from composite materials.
Fig 15-3: Harrier AV-8B
26. Helicopter Rotor Hubs. Conventional rotor hubs, i.e. the structure that connects the blades
and the main body of the aircraft, are very highly stressed and complex units which have a multiplicity
of bearings, seals and lubricators to allow blade movement whilst ensuring proper load transfer. Novel
designs using elastomeric composite materials with high levels of elastic deformation have resulted in
concepts which are essentially bearingless. There are several advantages including reduced
maintenance and drag, reduced parts count, lower weight and improved damage tolerance and
lifetimes. The centrifugal load from the elastomeric bearings is carried by a hub plate which is also a
composite material.
27. Helicopter Airframes. Helicopter airframes also provide opportunities for composites. In a
Franco-German programme a new aircraft is proposed which includes 80% composite viz. 24% CFRP
42% CFRP honeycomb and 12 % Kevlar honeycomb. Typical design features are:
(a) Frames and beams of Kevlar 49 carbon laminates.
(b) Panels of carbon and Kevlar sandwich construction with a Nomex honeycomb core.
(c) Carbon/ Kevlar sandwich structures for the underfloor for high energy adsorption.
(d) Landing gear frames of CFRP laminates.
(e) Carbon sandwich structures for the tail boom for stiffness and strength.
(f) Vertical and horizontal stabilizers of carbon and Kevlar laminates.
28. Helicopter Rotor Blades. In a similar way to propellers, composites have revolutionized
helicopter rotor blade design. As the blade rotates, pitch changes, which are necessary to balance lift
forces, cause very high levels of fatigue loading. Composite main rotor blades that utilize
unidirectional CFRP in the spar design have virtually unlimited life.
29. Space Applications. Use of composite materials in primary structures of major space
vehicles was based on the successful use of composites in missiles powered by solid-propellant
rocket motors. In terms of the innovative use of new materials space applications in many ways
provide more scope than the aircraft industry. For satellites the timescales from concept to part
manufacture can be as little as two years and the short product runs normally involved, the materials
element in the final cost is often relatively low. Also in many applications no other material is suitable
either for reasons of mass or thermal control. The mechanical design requirements for satellite
components are usually determined by launch conditions. Typically for conventional rocket launch,
loading would be 7 g in the axial direction with 1 g in the lateral direction, whereas for a space shuttle
launch the loading is more uniform with approximately 5 g in both directions. The natural frequency of
the structure, and hence its stiffness, will often control the mechanical aspects of the design. The
potential of vibration-induced coupling with the launcher is a significant concern. Once in orbit,
mechanical loads are comparatively low. Environmental conditions under some circumstances can be
arduous and severe thermal cycling can feature, as well as the effects of high vacuum and erosion
through atomic oxygen or micrometeroid impacts.
Advantages of Composites
30. Composite materials have exhibited the following properties which are of significant use in
aerospace industry:
(a) Weight savings are significant ranging from 25-45% of the weight of conventional
metallic designs.
(b) High resistance to fatigue and corrosion degradation.
(c) Due to grater reliability, there are fewer inspections and structural repairs.
(d) Directional tailoring capabilities to meet the design requirements.

(e) Improved dent resistance is normally achieved. Composite panels do not sustain
damage as easily as thin gauge sheet metals.
(f) Composites offer improved torsional stiffness.
(g) Thermoplastics have rapid process cycles, making them attractive for high volume
commercial applications that traditionally have been the domain of sheet metals.
(h) Unlike metals, thermoplastics have indefinite shelf life.
(j) Low thermal expansion can be achieved. Composite material can be tailored to
comply with a broad range of thermal expansion design requirements and to minimise thermal
stresses.
(k) Manufacturing and assembly are simplified because of part integration thereby
reducing cost.
(l) Close tolerances can be achieved without machining.
(m) Material waste is reduced and because composite parts and structures are frequently
built up to shape rather than machined to the required configuration, as is common with
metals.
(n) Excellent heat sink properties of composites, especially Carbon-Carbon, combined

with light weight have extended their use for aircraft breaks.
(o) Improved friction and wear properties.
(p) Ability to tailor the basic material properties of a laminate has allowed new
approaches to the design of Aero-elastic flight structures.
Aerospace - Degradation and Failure Mechanisms
31. The design of aerospace structures is often governed by stiffness criteria. There remains
however, a number of important degradation and failure mechanisms to be considered. A summary
(which does not cover all possible degradation and failure modes) is presented below.
32. Fatigue. Many aerospace structures are subject to dynamic loads that vary through
different cycles of normal operation. As with metal aerospace structures, fatigue is therefore an
important consideration in the design of composite structures for aerospace application. Fatigue
damage in aerospace composites usually initiates in the form of matrix micro-cracks or fibre-matrix
de-bonding which then grows in extent under the action of cyclical loads. Extensive experience in the
aerospace industry and elsewhere has shown that in carbon fibre laminates, of the types likely to be
used in practice, fatigue damage accumulation can be avoided by ensuring that the component is not
subject to strains above approximately 0.4%. Consequently, components subject to cyclical loading
will often be designed to a strain limit (this is often taken to be 0.4% but different companies may have
their own, sometimes product specific requirements).
33. Laminates with thick plies or with a concentration of plies in a particular direction all stacked
together tend to be more prone to damage initiation and accumulation. It is good design practice to
ensure an even distribution through the laminate of plies in different directions rather than grouping
together plies of a particular direction.
34. Impact. Aerospace structures may be subject to a range of types of impact during normal
operation, including bird strikes, rain, runway debris, sand and unintended impact during maintenance
(e.g. tool drops). For some parts of a structure impact may be the over-riding concern, e.g. high-
bypass jet engine main fan blades, helicopter rotor leading edges. In these cases the composite is
often protected from direct impact by surface layer of steel or titanium. This also serves to protect the
composite from erosion caused by rain or sand. In cases of severe impact damage there is usually
fibre breakage in the impact zone and an accompanying immediate weakening of the structure. This
type of damage would normally be readily identified, either at the time of its occurrence or in
subsequent inspections. Less severe damage affecting the matrix only is usually more difficult to
identify (damage of this type is commonly referred to as barely visible impact damage or BVID). Even
though not readily evident this damage can be quite significant internally and lead to a severe
reduction in matrix dominated properties. In addition damage initiated by impact can accumulate in
extent and severity through fatigue loading. Most structures will be subject to some damage of this
type in service hence it must be taken into consideration in the design. A major focus in the design of
aerospace structures is thus on ensuring damage tolerance. This imposes two requirements, firstly
that there should be redundancy in the structure and secondly that the laminate constructions used
should be tolerant of impact damage. Design of damage tolerant laminates follows similar good
practice to that mentioned above for fatigue design but may also consider factors such as the use of
impact resistant flexibilised matrices or the introduction of glass or aramid fibres in between carbon
fibre layers.
35. Environmental Effects. The operating environment for aerospace structures exposed them
to potentially damaging conditions. A variety of forms of environmental degradation are possible, the
more important of which include:
(a) Moisture Absorption. Laminates will tend to absorb moisture in high humidity
environments and this effect is exacerbated by high temperatures. Moisture absorption tends
to reduce the matrix strength and stiffness and can lead to degradation of the fibre to matrix
bond strength. It can also drive the development of micro-cracking which can grow by fatigue.
For this reason only matrix systems with known behaviour in so-called 'hot-wet' environments
are used in aerospace applications.
(b) Thermal Fatigue. Thermal stresses in composites should be considered on two

levels, firstly the global or structural level and secondly the micro-level. On the structural level
the stresses are controlled by differences in temperature and thermal expansion coefficient of
adjacent parts of the structure. On this level the designer has a good degree of control over
the level of thermal stress that may develop in operation through, for example, designing
adjacent parts of the structure to have similar coefficient of thermal expansion. On the micro-
level stresses are developed in the fibre and matrix as a result of their different coefficient of
thermal expansion. The difference between the fibre and matrix coefficients of thermal
expansion is inevitably large when carbon fibres are used and this means that significant
tensile stresses are developed in the matrix when the laminate is cooled far below its curing
temperature. Very low temperature operation (e.g. high altitude flight or in space) can lead to
micro-cracking of the matrix without any external load application. The designer of
components subject to thermal cycles whose minimum temperature is low relative to the
curing temperature will therefore carefully evaluate the fatigue behaviour of the laminates
considered.
Conclusion
36. The use of fibre reinforced composites has become increasingly attractive alternative to the
use of conventional metals for many aircraft components mainly due to their increased strength,
durability, corrosion resistance, resistance to fatigue and damage tolerance characteristics.
Composites also provide greater flexibility because the material can be tailored to meet the design
and certification requirements and they also offer significant weight advantages. Carefully designed
individual composite parts, at present, are about 20-30% lighter than their conventional metal
counterparts. While the all composite aeroplane may not be just round the corner, yet advances in the
practical use of composite materials should enable further reduction in structural weight of aeroplane.
The composite materials used in aircraft industry are generally reinforced fibres or filaments
embedded in resin matrix. The most common fibres are carbon, aramid, glass and their hybrid. The
resin matrix is generally an epoxy based system.
37. Composites also presents challenges. One area of some concern is the long term effect of
moisture absorption by composites. Another aspect which has attracted attention is the narrow data
base available with designers on real time experience with the use of composites. Presently,
predictions on the behaviour of composites are largely based on accelerated test cycling in
laboratories.
38. Observers, however point out that the challenges presented by composites are those which
can be overcome as more experience in their use is gained. Present indications point firmly in one
direction that is towards an increase in the use of composites increase in Aircraft manufacturing.
381
CHAPTER 16
AERO-ELASTIC TAILORING
Introduction
1. Aero-elastic tailoring means the tailoring of the structure of materials which are used in the
construction of aircraft components. By such tailoring the problem of divergence in forward swept
wings is eliminated. In respect of unswept and swept back wings Aero-elastic tailoring reduces shift in
the CP position under various conditions and improves flutter characteristics.
Divergence in Forward Swept Wings
2. When a forward swept wing flexes,

there is an increase in the section angle of
attack (Fig 16-1). This effect increases from
wing root to wing tips. When a swept wing
flexes under load all chord wise points at
right angles to the main spar (x-y) are
raised to the same extent. Thus points A
and B will be raised to the same extent with
respect to each other. Also points C and D
will be raised to the same extent with
respect to each other. However, as points C
Fig 16-1: Flexing of Forward Swept Wing
and D are closer to the wing tip, they will be
raised to a greater extent then points A and B. Hence, point C will be raised higher then point B. This
leads to an increase in the section angle of attack. This effect is the maximum towards the tips. This
phenomena of increase in angle of attack (and so the lift) when a forward swept wing flexes under
load, leads to divergence.
3. The problem of divergence was a serious one in swept forward wings and became a major
obstacle in their employment. Up to the advent of Aero-elastic tailoring, measures to increase
divergence speeds involved a heavy weight penalty and were unacceptable as a solution to the
problem. Aero-elastic tailoring however altered the entire scene.
Tailoring of Composites
4. In Aero-elastic tailoring, stiffness and strength qualities of a fibrous material are exploited.
Laminated composites consist of layers of material. Each layer consists of plies or fibres of a certain
thickness, strength and angular orientation with respect to each other. The angular orientation of the
fibres can be varied to suit any particular requirement. By selecting an appropriate angular orientation
of the fibres in each layer and an appropriate number of layers, a material of the desired strength can
be created. Also, the strength of a material can be made directional by appropriate fibre orientation. It
may be mentioned here that composites can be so tailored that they posses higher specific strength
and stiffness than conventional metals. By suitable fibre orientation a composite can also be made to
twist in the desired direction when it is compressed or stretched.
Orientation of Fibres
5. When a wing bends it also twists and deforms the skin and wing structure. For instance when
a wing bends upwards, it compresses the upper skin and stretches the lower skin. The deformation
of the skin wing structure can lead to wash in or wash out. To prevent divergence, wash in must be
reduced to the extent possible.
6. Fibre Orientation can be visualised from Fig

16-2. The 90º axes lie at an angle of 90º to the
reference axes. The 0º-90º quadrant is known as the
aft quadrant and the 90º to 180º quadrant is known
as the forward quadrant. Fibre orientation is referred
to as Q.
7. A Symmetrical laminate is that in which a

layer of material at some distance above a mid
surface reference line has identical ply thickness,
Fig 16-2: Fibre Orientation
angular orientation and material properties to a layer
at an identical distance below the reference line.
The term balanced means that for every material ply lying at some angle Q to a reference line, there
is an identical ply line at an angle of Q with respect to the line. In the case of an unbalanced
symmetrical laminate, bending of the wing will cause a change in the curvature of the surface as well
as twisting of the surface.
8. The orientation of fibres depends upon a number of factors. For instance, if a laminate is to be
designed with wing strength considerations in mind, only a small portion of the fibres will be available
for divergence tailoring.
9. In tailoring for preventing divergence in forward swept aircraft, fibre orientation in the forward
quadrant has been found to be beneficial. Only a slight orientation forward of the Q = 90º axis drives
the divergence speed to high values. This takes place because when wing flexes, it compresses and
twists the upper surface (Fig 16-3). When Q is little more than 90º (ie. towards the leading edge) the
twist is in anti clockwise direction on the right wing (and clockwise on the left wing). However, the skin
383 Aero-Elastic tailoring
cannot actually twist as it is rigidly fixed to the wing structure. Hence the force which wants to twist the
skin instead provides a torque to
twist the wing about its torsional
axis. In this case twisting is such
that the leading edge is lowered in
relation to the trailing edge. The
twisting thus reduces the section
angle of attack. In this manner, the
initial increase in section angle of
attack due to flexure is countered
and the problem of divergence is
overcome. A fibre orientation in the
aft quadrant is undesirable as the
Fig 16-3: Twisting of Wing
surface twisting will increase the
wash in.
10. In an un-swept wing with 90º < Q < 180º (towards the leading edge) the divergence speed
increases with increased in Q. Then flutter becomes the primary mode of instability. In aft swept wings,
divergence is critical with Q < 77º. Here with 77º < Q < 155º, the divergence speed increases with Q.
In forward swept wings, divergence is effectively counted in the range 90º < Q < 120º. In this range
however, the wing flutter is critical. For countering flutter the torsion stiffness of the surface has to be
increased. This involves fibre orientation Q = 140º (approx.). Hence, fibre orientations do matter for
controlling divergence and flutter, the difference being the most in swept forward wing. When the
forward swept is more than 30º. Here the designer has to choose a combination of fibre orientation to
get the best combination to counter both divergence and flutter.
11. The wind tunnel testing of a graphite epoxy wing has demonstrated two benefits of the use of
tailored composites.
(a) Higher divergence speeds due to the higher specific stiffness of graphite epoxy used
for the wing structure.
(b) Reduction in wash-in by fibre orientation in the forward quadrant.
12. Fibre orientation of Q = 110º reduce the outward (span wise) shift of the CP due to wing
bending in swept forward wings (by reducing the wash-in). This also helps in increasing the
divergence speed. In the case of un-swept and swept back wings the movement the CP about a
datum can be controlled by suitable fibre orientation. Here, the fibre orientation is such that surface
twisting takes place in the desired direction to prevent distortion of the structure.
Conclusion
13. By the use of an efficient composite structure, which is suitably tailored, the phenomena of
divergence in forward swept wing can be overcome. However, aero-elasticity is a multiple parameter
problem involving wing sweep, Mach No. and Aspect ratio interacting with the structure in a complex
manner. Designers as yet do not have a wide data base for aeroelastic tailoring. More development
programmes are needed to study aero-elasticity in respect of unrestrained forward swept wings and
for large size aircraft. The effects of under wing stress carriage also need greater study. However, the
present state of the art has enabled the successful flight of an aircraft with forward swept wings. In un-
swept and swept back wings aeroelastic tailoring can reduce the movement of the CP and improve
flutter characteristics.
385
CHAPTER 17
AIRCRAFT ABANDONMENT
Introduction
1. In an emergency it should be possible to abandon an aircraft with a reasonable chance of

success. Military aircraft operate in an extremely large flight envelope with speeds ranging from zero
to well over 1,000 knots, and operating heights from ground level to over 50,000 ft. Escape from
aircraft operating in the lower speed range may be possible solely with a parachute, but assisted
escape is necessary for those with a wide or high speed range. Escape systems are designed to aid
occupants to get clear of an aircraft and into an environment from which a normal parachute descent
is possible. This chapter gives a general description and the figures quoted are generic in nature.
Pilots must refer to the respective Pilots Notes and relevant orders in vogue for parameters relevant
to the aircraft which they are operating.
Conventional Escape
2. Up to about 200 knots, it is possible to escape unaided from an aircraft. Above about 200
knots it becomes extremely difficult and dangerous to escape unassisted. The situation is made
worse at any speed if the aircraft is subjected to ‘g’ forces, e.g. due to a spiral dive. The minimum
height at which the escapee must be clear of the aircraft is 1,000 ft AGL and it follows that the
decision to bale-out must be made at a greater height than this, depending on aircraft type (i.e.
difficulty of egress) and rate of descent. At the higher speeds, the slipstream across an exit makes it
difficult for the crew member to escape and if he does, it is possible that he will contact a part of the
aircraft.
3. Helicopter Escape. Most current helicopters are not cleared for the carriage of parachutes,
hence escape from them in flight is not possible. Most of the remainder are required to carry
parachutes when operating above 3,000 ft or on special test flights. Although helicopters generally fly
at speeds below 200 knots, there are many problems associated with clearing the aircraft. Some of
these are listed below.
(a) Incompatibility of aircrew equipment assemblies with exit path.
(b) Inadequate size of escape hatch or door.
(c) Failure of the helicopter to stay in a reasonably steady attitude after the cyclic stick
has been released.
(d) Difficulty in avoiding parts of the airframe after exit.

4. Assisted escape from helicopters is not provided. Even if the problems listed could be
overcome, survival without assisted escape is doubtful from less than 1,000 ft.
Methods of Assisted Escape
5. The current means of assisted escape from fixed wing aircraft is the Ejection seat. This has
provided many aircrew with a lifesaving facility, but as airspeeds increase, open seats do not give
adequate protection against the effects of wind blast.
6. Ejection seats have been developed to provide an entirely automatic escape facility, from
ground level upwards, within specified speed limits. The user is required only to initiate the firing
sequence, and thereafter all the required operations take place automatically. The firing mechanism
first clears the ejection path and then operates the ejection seat. As the seat rises, the seat systems
are activated by static rods, services are disconnected and emergency oxygen turned on. The feet
swings back from the rudder pedals due to inertia at the beginning of the upward travel of the seat on
its rails, and the legs are restrained close to the front of the seat pan by the automatic action of the leg
restraining cords. This prevents fouling of the lower part of the legs on the instrument panel and leg
flailing throughout the time that the occupant is exposed to high air blast while he remains in the seat
during the subsequent descent.
7. The drogue gun fires about 0.5 sec after the seat rises; the drogue bullet pulls the duplex
drogues from their housing. The drogues stabilize, and slow down the seat sufficiently to ensure safe
man / seat separation. This occurs 1.2 to 2.3 sec after ejection or as soon as it is safe (seat
deceleration less than about 4 ‘g’), or at a barometric height of 10,000 ft or 3,000 m. A 5,000 m
(16,400 ft) barometric capsule may be fitted to any aircraft flying over mountainous terrain.)
8. Some ejection seats are fitted with rocket packs, which are used to sustain the ejection
velocity provided by the cartridges in the ejection gun. The rocket is ignited as the seat leaves the
aircraft. The advantages of rocket assistance are:
(a) Less ejection force required as rocket increases velocity to 76 m/s (250 ft/s).
(b) Reduced acceleration due to the reduced ejection force of about 11 g, and reduced
rate of application.
(c) Increased trajectory, giving a zero-zero capability which is also beneficial in

conditions of low altitude and high sink rate.
Pre-ejection Considerations
9. Many aircrew have failed to survive emergencies which occur at an altitude sufficiently high
for a successful escape to be made simply because the decision to eject was taken too late. In single
387 Aircraft Abandonment
or twin-seat aircraft the decision must be made above the minimum safe ejection altitude (MSEA).
10. Minimum Safe Ejection Altitude (MSEA). It is generally accepted that ejection in straight
and level flight at 230 knots and 9,000 ft is the ideal. The rate of descent and aircraft attitude each
have an adverse effect on the MSEA, the rate of descent overriding the factor of aircraft attitude
except when very close to the ground. The following time increments are critical:
(a) Decision Time. This is the time taken for aircrew to evaluate the emergency and to
inform other crew members. The acknowledgement of orders also affects decision time.
(b) Crew Reaction Time. This is the time taken to react to the order to eject and to
operate the ejection seat.
(c) Time for Equipment to Function. This is the period from initiation of ejection until
the seat clears the cockpit.
(d) Time for Full Operation of the Seat. This is from seat initiation until the aircrew
member is descending vertically on a fully deployed parachute.
11. It is important that aircrew are aware of the fact that an ejection seat with a ground level
capability may have a minimum safe ejection altitude of several thousand feet when escape is
attempted in other than straight and level flight. The MSEA will be greatest in a high speed vertical
dive. A reasonable approximation is to allow 10% of the rate of descent. Other minima for particular
aircraft and situations may be published in Pilot notes and appropriate orders.
Ejection Drill
12. Ideally, the position which the individual adopts to carry out his task in reasonable comfort
should be that in which he can fire the seat without further adjustment, with a high probability of
successful uninjured escape.
13. Pre-ejection drills for individual aircraft are to be found in the appropriate Aircrew Manual, but
the ejection handle should in all cases be grasped firmly and pulled to the full extent of the operating
cable (25-120 mm or 1-5 inches). The pulling action will tend to place the body in an acceptable
ejection posture. If time permits, the harness and negative ‘g’ strap should be checked for tightness
(although the shoulder harness should not be over-tightened), and the head pressed against the
headrest.
Posture During Ejection
14. The posture of the body is extremely important in determining whether the aircrew will escape
uninjured and is directly related to the correct strapping-in procedures. If the back is correctly
positioned and supported during ejection by a correctly adjusted restraint harness, it can safely
tolerate the accelerations imposed on it by the 24.4 m/s (80 f/s) ejection guns. Poor posture could
result in injury even with the lower acceleration of rocket assisted seats.
15. The back is at its strongest, and thus more able to withstand loads such as those caused by
ejection, when it is in its normal position, i.e. straight when viewed from the front and slightly curved
like an elongated ‘S’ when viewed from the side. In the normal position, the back can withstand
accelerations of up to 30 g at a rate of application of over 300 ‘g’/s. If the back is bent or twisted, this
figure can fall to 9-14 g at rates considerably less than above. Poor posture may cause compression
fractures of the back, but only very rarely produce spinal cord damage.
16. The nature of the seat pack through which the ejection accelerations are transmitted, the
support afforded by the seat back, and the effectiveness of the restraining harness are of the upmost
significance in seat ejection. The user can only adopt and maintain a posture as good as the
equipment will allow:
(a) Personal Survival Packs. There is a variety of personal survival packs currently
installed as items of aircraft equipment in ejection seats. The main characteristics of these
packs are:
(i) Comfort. As well as ensuring good ejection characteristics, it is important

that the personal survival pack is correctly contoured and therefore comfortable. In
long range aircraft this factor is of great importance since the degree of comfort has a
marked effect on the efficiency of the user. Moreover, from the ejection point of view,
he is more likely to remain in the correct position on the seat if it is correctly shaped
and comfortable. Comfort is improved by a limited slip quality of the top surface,
since this permits small movements, when desired, to ease pressure points.
(ii) Shape. The pack should be of such a shape that it is located firmly in the
seat pan, but at the same time is capable of unhampered separation from the seat
during the process of escape. Its top surface should be shaped so that it encourages
the user to sit correctly in the best position towards the back of the pack.
(iii) Contents. The contents of the personal survival pack range from the bulky
rubber life raft and its accessories to small hard objects. These are contained in a
rigid box and the packing of the objects in the container is particularly critical to
ensure constant shape and ejection characteristics.
(b) Restraining Harness. Ejection seats are fitted with a combined harness system. It
is important that the harness system is adjusted correctly to ensure the maintenance of good
posture during the escape sequence. The location of the straps and harness fastening in the
optimum position, and the correct sequence of tensioning of the system, will restrain and also
maintain body position.
Post-Ejection Considerations
17. As the seat and occupant leave the aircraft they may be exposed to the following stresses:
(a) Wind Blast. When the seat clears the aircraft, the occupant is exposed to the ram
effect of the slip-stream. This is proportional to the IAS. At indicated speeds up to about 350
knots wind blast is not likely to cause injury. As speeds increase above 350 knots, there is an
increasing likelihood of injury unless appropriate restraint is provided. The upper limit for the
open seat appears to be about 650 knots.
(b) Sudden Deceleration. On entering the slip-stream the seat and its occupant
undergo a marked deceleration caused by the wind drag and higher the IAS, the greater the
deceleration effect. For a given IAS the maximum linear decelerations are not affected by
altitude. As the ejection altitude is increased, however, the deceleration time, is prolonged.
This is because for a given IAS, increased altitude results in a greater kinetic energy (higher
TAS) which takes longer to dissipate in the lower density. Ejection seats are provided with a
stabilizing system so that this deceleration is linear, otherwise an unstable system would
produce a variety of forces on the occupant. There are many factors which affect the drag
characteristics of the man / seat complex, so that it is not possible to lay down a maximum
IAS for safe ejection from the point of view of the deceleration effects. Assuming a maximum
safe peak linear deceleration of 35 ‘g’, it has been calculated that this might be experienced at
an IAS between 600 and 700 knots.
(c) Tumbling and Spinning. Unstable seats would tumble and spin and the high
acceleration loads could cause serious injury to the occupant, therefore seats are stabilized
by means of drogues. In most seats two drogues are used; a small one opening first, bringing
the seat into alignment with the relative airflow and pulling out a second, larger drogue. A ‘g’
stop is incorporated which prevents separation from the seat and deployment of the main
parachute canopy until the acceleration loads have been reduced to an acceptable level.
During stabilized free fall, spinning or swinging about a vertical axis can occur and this may
induce sensations of tumbling and the impression that the drogues have not deployed.
(d) Effects of Environment at High Altitude. If ejection occurs above the barometric
level of the automatic system, a fall to the set altitude occurs before man / seat separation
and the deployment of the main parachute canopy. This allows the seat to descend,
stabilized by drogues (e.g. from 50,000 ft to 10,000 ft in approximately 3.5 minutes.) The
reasons for this delay are:
(i) To prevent explosive opening of the main parachute canopy at an altitude

where the TAS would be high enough to cause severe tearing of the fabric and injury
to the escapee (elaborated in para 17e).
(ii) To keep the time spent at altitude to a minimum, as only a very limited
quantity of emergency oxygen is carried and, in the worst case, the oxygen mask
may have been lost due to the wind blast on ejection.
(iii) To keep the time spent in the low temperature regions of the atmosphere to
the minimum.
(e) Parachute Opening Load. The opening load of a parachute canopy depends on
many factors. These include its design and the length of its rigging or shroud lines, the
method of opening, the altitude and speed through the air at the moment of opening, the size
and design of the vent, the weight and porosity of the material, and air density and humidity.
Canopies used for emergency escape systems should be simple, compact, quick opening
and reliable. The principal factors which affect the opening load are terminal speed through
the air at the moment of pack opening and the altitude at that moment. Ejection seats fitted to
modern aircraft are equipped with the GQ aeroconical parachute. This permits parachute
opening at higher speeds and extends the performance envelope of the escape system. The
drogue system is designed to decelerate the seat / man system down to a safe speed for
parachute extraction and deployment. Release from the seat does not occur until the load on
the drogue system has reduced allowing the ‘g’-stop to release and altitude is lost so that the
barostat allows the time delay in the automatic release mechanism to run for its set period.
With rocket assistance and the need for earlier seat / man separation in low level ejections,
the g-stop has been removed from most modern ejection systems, but the time delay has
been slightly increased. The opening load of a parachute canopy is proportional to the TAS
(V2) and as the terminal velocity increases with altitude V2 also increases. Thus, parachute
opening loads are very much greater at higher altitudes, even at the minimum barometric
altitude of 10,000 ft. The opening jerk after seat / man separation from a seat stabilized by
the drogue system to 10,000 ft is very often described as severe. Opening the parachute at
higher altitudes, particularly after even a very short period of free fall, can be hazardous as
the load may produce physical injury or exceed the designed loads of the parachute or
harness system. Fortunately, parachutes tend to fail safe, in that deployment damage
relieves the excessive loading and still leaves sufficient canopy for a safe, and stable, descent.
Manual separation from an ejection seat and manual parachute deployment above barometric
altitude are therefore potentially hazardous. Seat systems are very reliable, certainly far more
reliable than the individual escapee who rarely has had previous experience.
Ejection Seat Malfunction
18. Action in the Event of Failure to Fire. Failure of the seat to fire is a very rare occurrence.
However, should it fail to fire at the first attempt, initial efforts should be directed towards obtaining a
normal ejection. The firing handle should be pulled again harder, and the alternative handle (if fitted)
can also be pulled. Jettison the cockpit canopy or if unsuccessful, open the cockpit canopy using the
normal aircraft system as the canopy jettison system may be the cause of the failure. If these actions
are unsuccessful, the situation should be re-assessed and the original decision to abandon
reconsidered, i.e. the actual emergency should be reviewed against the likelihood of a successful
manual escape.
Hazards of a Manual Escape
19. The seat and the associated aircrew equipment have been designed specifically for gun
ejection and the manual override facility is provided to overcome the failure of the automatic
separation which occurs once the seat is clear of the aircraft. It is not designed for escape from the
cockpit. It is therefore difficult, if not impossible, to leave a failed ejection seat unless conditions are
ideal. Moreover, the hazards of snagged straps and clothing etc. and of subsequent impact with parts
of the airframe after leaving the cockpit, cannot be discounted.
Manual Escape (Static Seats)
20. The actual drill for escaping from the aircraft varies with aircraft type and the make and mark
of the seat fitted. The appropriate drill is published in Aircrew Manuals. It should be borne in mind
that it is not a normal listed emergency drill, but rather a suggested course of action. The following
factors will have been considered when formulating the recommended procedure:
(a) The difficulty in freeing the parachute from its housing.
(b) The number and complexity of the attachments between the occupant and seat, all of
which must be freed before attempting to escape.
(c) The best method of actually getting out of the aircraft and the likelihood of snagging
of the straps, or being struck by parts of the aircraft during and after the escape.
21. The inverted fall-out escape is unlikely to be successful because freeing the parachute pack
from its housing requires considerable strength and agility, even under ideal conditions. High
airspeed, g loading, or loss, or partial loss, of control will increase the hazards.
EJECTION
Introduction
22. Ejection seats enable aircrew to abandon equipped aircraft safely in the most adverse
circumstances. The seats are fully automatic and very reliable. In principle it is rigidly constructed
metallic seat which is forcibly ejected from the aircraft by means of explosive charge. Aircraft in use in
IAF have different types of seats. The basic design of seats include an ejection gun, guide rail, seat
frame structure, adjustable seat pan, parachute container, drogue and face blind or seat firing control
called primary firing handle. A barostatic control unit is incorporated to initiate separation of seat and
pilot. The seat is stabilized by a drogue chute. An automatic leg restraint mechanism is also fitted.
Similarly arm guards are also provided on certain seats. The development and improvement of seats
is a continuous process and for this reason there are many types in use at any one time. This chapter
will give a general description of a typical seat and outline the principles of its operation by tracing the
sequence of events during an ejection.
23. General Description. The modern ejection seats have undergone a series of refinements
since its inception in l946. It is today a highly automated system that requires the occupant to only
initiate the firing mechanism to affect escape. Typically, the seat consists of a padded bucket, back,
and headrest. The seat is mounted on rails which guide the seat on its initial trajectory. Most seats are
propelled by rockets but the methods of restraint, seat separation, and chute deployment will vary
according to the various types of ejection seats. Generally, escape is initiated by pulling a firing
handle. In some seats a trigger within the handle then must be squeezed to initiate ejection. As the
ejection seat travels up the rails, a leg restraint system activates. The development of rocket
propulsion has produced the higher trajectory necessary to clear aircraft structures during high speed
escape as well as escape during low speed and zero-zero (zero velocity and zero altitude) ejections.
Seat stabilization gyros have been incorporated into recently developed ejection seats to cancel
asymmetric forces producing rotation and tumbling.
24. A personal equipment connector (PEC), on the left hand side of the seat pan, provides a
single action connection for the occupant’s mic / tel, oxygen supplies, and ‘g’-suit when applicable.
The seat portion is coupled to the oxygen regulator. A cover is provided to fit whenever the man
portion is not connected. The man portion of the PEC is part of the aircrew equipment assembly and
has an oxygen tube and mic / tel lead for connection to the occupant’s oxygen mask hose and helmet
mic / tel respectively. When released, either manually or during seat separation, the leg restraint
cords are also released.
25. The main oxygen system is connected to the seat via an automatic pull-off bayonet connector
behind the seat pan and then by pipe to the regulator / PEC. An emergency oxygen cylinder at the
rear of the seat feeds into the main supply line and has a release mechanism which is tripped
automatically by a striker during ejection, or can be operated manually at any time by operating a
control handle on the seat pan.
26. Prior to the ejection system launching, the canopy has to be jettisoned to allow the crew
member to escape the cockpit. There are at least three ways that the canopy or ceiling of the airplane
can be blown to allow the crew member to escape:
(a) Lifting the Canopy. Bolts that are filled with an explosive charge are detonated,
detaching the canopy from the aircraft. Small rocket thrusters attached on the forward lip of
the canopy push the transparency out of the way of the ejection path.
(b) Shattering the Canopy. To avoid the possibility of a crew member colliding with a
canopy during ejection, some egress systems are designed to shatter the canopy with an
explosive. This is done by installing a miniature detonating cord or an explosive charge
around or across the canopy. When it explodes, the fragments of the canopy are moved out
of the crew member's path by the slipstream.
(c) Explosive Hatches. Aircraft without canopies will have an explosive hatch.
Explosive bolts are used to blow the hatch during an ejection.
27. Operation. When a crew member pulls

either of the firing handles (main handle or the face
curtain down) on the ejection seat, it sets off a chain of
events that propels the canopy away from the plane
and thrusts the crew member safely out. From the
time the ejection handle is pulled, the exact amount of
time seat takes to clear the aircraft depends on the
seat model and the crew member's body weight.
Pulling the ejection handle on a seat sets off an
explosive cartridge in the catapult gun, launching the
ejection seat into the air. The acceleration imparted to
the seat must be controlled to within limits which are
physically safe for the seat occupant. As the seat rides
up the guide rails, a leg-restraint system is activated,
the seat main oxygen supply is disconnected and the
emergency oxygen supply is tripped on. These leg
restraints are designed to protect the crew member's
legs from getting caught or harmed by debris during
the ejection. An under seat rocket motor provides the
force that lifts the crew member to a safe height, and Fig 17-1: ACES II Ejection Seat
this force is not outside normal human physiological limitations. The seat, parachute and survival
pack are also ejected from the plane along with the crew member. Many seats, like Goodrich's ACES
II (Advanced Concept Ejection Seat, Model II), have a rocket motor fixed underneath the seat. After
the seat and crew member have cleared the cockpit, this rocket will lift the crew member another 100
to 200 feet (30.5 to 61 m), depending on the crew member's weight. This added propulsion allows the
crew member to clear the tail of the plane
28. Once out of the plane, a drogue gun in the seat

fires a metal slug that pulls a small parachute, called a
drogue parachute, out of the top of the seat. Fig 17-2
shows a parachute opening on a Martin-Baker ejection
seat during a test. The small parachute at the top is the
drogue parachute. This slows the person's rate of descent
and stabilizes the seat's altitude and trajectory. The
drogue gun has two separate cartridges, one mechanically
fired by the trip rod (which is attached to the aircraft) 0.6
seconds after the trip operates during ejection, the other is
gas fired by the barostatic time release unit (BTRU) or by
operation of the manual separation handle. A shear pin is
fitted through the top of the drogue gun barrel to retain the
piston until the gun fires.
Fig 17-2: Martin-Baker Ejection
Seat during a Test
29. The BTRU on the top right of the seat provides for
the automatic release of the drogue shackle, deployment of the personal parachute, and separation of
the occupant from the seat after ejection. The automatic sequence commences with the firing of a
cartridge in the BTRU. A time delay mechanism and a main barostat are able to delay the firing of the
cartridge, depending on ejection conditions:
(a) The mechanical time delay mechanism is triggered by the trip rod when the seat rises
on ejection. If the delay mechanism is unobstructed it allows the BTRU cartridge to fire 1.5
seconds after ejection.
(b) If the ejection takes place at altitude, the main barostat prevents operation of the time
delay until the man and seat have descended to 10,000 ft where a tolerable oxygen and
temperature environment exists. After ejection the drogue stabilized seat descends rapidly
with the occupant strapped in. At 10,000 ft the barostat removes its restraint on the time
delay, which is then free to operate.
30. After a specified amount of time, an altitude sensor causes the drogue parachute to pull the
main parachute from the pilot's chute pack. At this point, a seat-man-separator motor fires and the
seat falls away from the crew member. The person then falls back to Earth as with any parachute
landing.
Seat Limitations
31. Seats with rocket assistance may be used at ground level with no forward speed provided the
aircraft attitude is straight and level. Seats without rocket assistance may be used at ground level
provided the speed is 70-90 Knots, depending on the type of seat. The aircraft attitude should be
straight and level. If the aircraft is airborne and in a descent, the minimum height for ejection is
approximately 10 % of the rate of descent, ie. at 1,500 ft/min rate of descent the minimum safe height
for ejection is about 150 ft provided the wings are level. More detail for a specific seat / aircraft
combination is given in the appropriate Aircrew Manual.
Parachute Descent
32. It is necessary to adopt a correct body position prior to opening the parachute manually. The
best position is head out, torso slightly bent at the waist and feet held together. This position will
prevent the head being struck by parachute lift web and the entanglement of the rigging lines between
the legs. Following parachute deployment, sometimes, there are minor complications. Lift webs and
the rigging lines may be twisted. These can be un-twisted by manually spreading them apart. A
rigging line may be thrown over the canopy making two bulges. This can be connected by pulling
down the rigging lines of the smaller of the two bulges. Oscillation and pendulum like movement of
the body can occur prior to landing. These can be dangerous during landing and can be obviated by
pulling down on two adjacent lift webs and slowly releasing them to their normal position once
oscillation is damped. During descent, it is necessary to release the survival pack from its attachment
to the body.
Parachute Landing
33. High percentage of non-fatal injuries resulting from emergency escape occur on landing. It is
therefore important to know the basic rules on parachute landing. The landing position should be
assumed approximately 1000 ft. above the ground. Both arms should be stretched over head firmly
gripping the lift webs. The knees should be bent and other feet held together. The line of vision should
be directed at 45° angle to the ground and not straight down. The landing should be on the balls of
the feet. And the body rolled in the direction in which the parachute is moving. This distributes the
impact over a large area of the body. After landing, injuries can still be caused by dragging by the
parachute which must therefore be collapsed. The best method is to operate the quick release box
and to release the harness. Another method is to pull on the lower rigging lines which will cause the
canopy to collapse. At night, the landing position should be assumed from the moment the parachute
canopy is deployed. If landing on trees cannot be avoided, the arms should be folded in front of the
face and feet kept together. While landing in water pre-landing preparation is important. During
landing oxygen mask should be removed, the mae-west inflated on touching the water, and the
parachute canopy release operated immediately. As soon as this release has been operated, swim
upwind from it to avoid entanglement in the rigging lines.
34. The maximum rate of descent at impact, to avoid significant injury, is 8 m/s.
Human Factor in Delayed Ejection
35. Once the decision to abandon the aircraft is taken, the only action required is to fire the seat.
This does not take more than a second or two, yet the IAF has lost a number of experienced pilots,
because they waited too long to part company with the aircraft. The factors involved in such delays
are as follows:-
(a) Delayed Decision. Decision to eject involves recognition of emergency and coming
to a conclusion that no other procedure is compatible with safety and survival. Both
require thorough and repetitive training on ground.
(b) Delayed action.
(i) Fear of Unknown. The cockpit environment is known to the pilot and he
feels safe inside it. One may take a long time to come to the terms to leave it and
eject till it is too late.
(ii) Fear of Ridicule. One feels that he or she would be ridiculed on ground
after ejection that he / she could not recover the aircraft safely from an emergency
and they delay taking the decision of ejection till it is too late.
(iii) Past Success. Past success in bringing back an aircraft from an

emergency induces a confidence in the pilot that he would be able to bring the aircraft
out of this difficult situation also.
(iv) Saving the Civilian Population. At times while trying to save the people
below pilots delay their ejection and it proves to be fatal specially if the height is not in
hand.
(v) Professional Pride. This is a double edged sword, pride in a pilot is

required to be their but if it is not balanced then this pride results in a delayed ejection
specially by pilots with lots of experience.
(vi) Communication. This is a problem in trainer aircraft. Pre ejection drill has
to be covered in detail and followed to the core to ensure safe ejection.
36. Remember the aircraft is expensive but your life is more expensive.
399
CHAPTER 18
OXYGEN SYSTEMS
Introduction
1. Breathing ambient air on ascent to altitude produces a progressive fall in the partial pressure
of oxygen in the lungs (PO2). Above 8,000 feet the (PO2) will be at levels which are insufficient to
meet the body's requirements for oxygen and hypoxia will develop. This most serious of hazards
must be prevented in flight and one method of so doing is to provide an artificial pressure environment,
i.e. a pressurized cabin. The alternative method is to provide a source of added oxygen so as to
maintain the PO2 at ground level equivalent at all altitudes. In most military flying a highly pressurized
cabin (High Differential Cabin) is inappropriate for several reasons and so both the methods are
combined. The cabin is pressurized to a certain degree (Low Differential Cabin) and any short-fall in
oxygen required is met by a supplementary source in the aircraft.
2. The physiological and operational requirements for aircraft oxygen systems may be
summarized thus:
(a) Oxygen Concentration. Oxygen would be most simply and conveniently delivered
as 100% oxygen at all altitudes. This, however, has several disadvantages not least of which
are those of cost, weight and bulk; particularly since 100% oxygen is not required
physiologically until a cabin altitude of
34,000 feet is reached. Fig 18-1 illustrates
the concentration of oxygen required to
achieve this. In practice this aim is achieved
by providing an increase in inspired oxygen
concentration from ground level, until at
about 30,000 feet most oxygen systems are
delivering 100% oxygen. The delivery of
100% at 30,000 feet, rather than at 33,700
feet as theoretically required, allows a safety
margin. l00% oxygen will continue to
prevent hypoxia up to 40,000 feet but above
this altitude pressure breathing is required to
provide continued protection. Fig 18-1: Oxygen Required to
Maintain Ground Level Equivalent
(b) Adequate Nitrogen Concentration. Nitrogen must be present in sufficient quantity

to prevent the occurrence of Oxygen Ear or Oxygen Lung. Thus 40% nitrogen or more is
normally required in a breathing system until altitude/oxygen requirements dictate otherwise.
(c) Adequate Ventilation and Flow. The system must be capable of delivering up to
60 litres per minute along with instantaneous peak inspiratory flows of 200 litres per minute.
(d) Minimal Resistance to Breathing. Resistance due to valves and turbulent flow
throughout the system, caused by uneven surfaces, branches and changes in internal
diameters must be minimized to prevent disturbances to respiratory rhythm. Ideally, the flow
characteristics should be such as to produce no noticeable resistance to breathing.
(e) Temperature. The inspired gas should be within ± 5° C of cockpit ambient

temperature.
(f) Safety Pressure. Inward leaks around the face mask seal or from hose
connections must be countered. This is accomplished by providing a small positive
overpressure in the mask to ensure that any leaks are outward.
(g) Protection against Toxic Fumes and Decompression Sickness. A facility for
selecting 100% oxygen at any time and at any altitude is necessary in the event of toxic
fumes appearing in the cabin or when decompression sickness is liable to develop or has
done so (cabin altitudes above 18,000 feet).
(h) Indication of Supply and Flow. Indications of both supply and flow must be
available to the user at all times as a check of correct function.
(j) Evaluation of Integrity. Where possible fail-safe methods of operation should be

used (eg the crew member should be unable to breathe through the mask until it is correctly
connected to the rest of the system) together with the means to check emergency functions
(eg manual test of mask seal and pressure breathing facilities).
(k) Convenience. As much of the system as possible should be automatic and the
drills to cope with a failure should be simple. Failures must be immediately and clearly
indicated.
(l) Duplication. In aircraft with low differential pressure cabins there should be a back-
up system in the event of main system failure. There is no need for such an Emergency
Oxygen supply in aircraft with high differential cabins where the cabin itself provides the
primary protection against hypoxia and the oxygen equipment is only used if cabin
pressurization fails, or toxic fumes contaminate the cabin.
(m) Provision for High Altitude Escape. A separate emergency oxygen supply is
needed in aircraft fitted with ejection seats or from which bale-out is possible. This supply
401 Oxygen Systems
fitted either to the seat or to the personal parachute pack, is usually the same as the back-up
supply referred to at l above.
(n) Independence from Environment. The environment extremes sustained in flight

must not impair the performance of the oxygen equipment. This is particularly so in regard to
low temperatures, accelerations (aircraft manoeuvres and windblast on escape) and
atmospheric pressure changes.
3. Oxygen systems have been progressively refined over the years. The subject has become
increasingly complex and aircraft specific. Because of this the following account is necessarily of a
general nature.
4. In broad terms any aircraft oxygen system consists of two parts: a supply or store of oxygen
and a means of delivering it to the user (regulator, hose and face mask).
AIRCRAFT OXYGEN SUPPLY/STORAGE
General
5. Oxygen is most usually obtained from an on-board store which is replenished whilst the
aircraft is on the ground. Some systems however use the on-board generation of oxygen by
molecular sieve oxygen concentrators (MSOCs). Usually oxygen is stored either as a gas at high
pressure or as a liquid at low temperature.
6. Whatever the source, the gas supplied to the system must be of a certain high standard.
Thus, it must contain at least 99.5% oxygen, be odourless and virtually free of any toxic substances
(eg. the carbon monoxide concentration must be less than 0.002%). The maximum allowable levels
for various hydrocarbons are specified in relation to the type of storage system used since this will
influence the potential contamination hazard. To avoid the risk of ice formation at low temperatures
the water content must not exceed 0.005 mg per litre of oxygen at Standard Temperature and
Pressure (STP: 0° C, 760 mm Hg).
Gaseous Storage
7. In gaseous storage systems the oxygen is held in cylinders mounted outside the pressure
cabin. A typical gaseous storage system is shown at Fig 18-2.
8. The advantages of such a system are that it is relatively simple; oxygen is not lost by venting
when not in use and it can be used immediately after filling. However, the cylinders are bulky and
heavy and consequently, this system is unsuitable as a primary aircraft oxygen supply when weight
and space are at a premium.
Fig 18-2: Typical Gaseous Oxygen Storage System
Liquid Storage
9. The problems of weight and bulk are greatly reduced by storing oxygen as a liquid under low
pressure. Liquid Oxygen (LOX) vaporizes at -183° C at normal atmospheric pressure, each litre of
liquid yielding 840 litres of gaseous oxygen (NTP). Such systems therefore occupy about half the
space and are half as heavy as the high pressure gaseous systems. Between 3.5 and 25 litres of LOX
can be carried depending on aircraft type and crew requirements.
10. The double-walled insulated container, which is essentially a stainless steel vacuum flask,
with its control valves and connecting pipe work are collectively known as a LOX Converter. It is
divided into two parts: one is insulated and contains the liquid, the other is un-insulated and contains
the gas. The converter may be permanently mounted in the aircraft or be removable for rapid
replacement.
11. The converter is charged from a ground LOX dispenser. LOX entering container evaporates
and eventually cools the internal walls to -183° C. The container then rapidly fills with LOX.
12. The LOX system is compact, of low weight and the container will not explode if damaged.
Unfortunately evaporation and venting losses mean that the converter needs to be recharged at
frequent intervals. In addition LOX takes a long time to stabilize once in the converter and it may be
upset if the container is agitated, as for example, by aerobatics. For this reason combat aircraft
403 Oxygen Systems
require the addition of a stabilizing chamber which ensures that on charging the liquid in the container
is at a temperature at which its vapour pressure equals the normal operating pressure. LOX is prone
to contamination by toxic materials and great care must be taken to prevent the build-up of
contaminants.
Molecular Sieve Oxygen Concentrators (MSOC)
13. Most of the problems of LOX systems can be overcome by the onboard production of oxygen
by the pressure swing adsorption method, using a molecular sieve. A molecular sieve is a
synthetically produced porous material and if the pores are of a suitable size gas molecules are able
to pass through them. Generally the adsorption of a molecule depends upon its polarity and its size;
clearly if a molecule is larger than the pore size it cannot pass through the sieve. Careful design
ensures that if air is passed through the sieve under pressure most of its nitrogen content will be
adsorbed leaving behind an oxygen enriched product gas. (Because the adsorption characteristics of
argon are similar to those of oxygen the product gas will comprise 95% oxygen and 5% argon). Over
time, the sieve bed becomes saturated with nitrogen which needs to be purged to prevent it from
appearing in the product gas. Removal of nitrogen from the sieve is achieved by depressurizing the
bed to ambient pressure followed by back-purging with a portion of the product gas, a process known
as pressure-swing adsorption.
14. Limitations in the Use of MSOCs in Aircraft. Extensive in-flight experience has shown
the MSOC to be a very efficient filter of contaminants, including engine oil and hydraulic fluid
molecules from engine bleed air, as well as vapour, allowing oxygen to be concentrated in suitable
quantities. However, a separate gaseous supply is still required in case of engine failure or crew
ejection. Moreover, not all MSOCs are able to meet the requirement to provide the 100% oxygen
needed to protect against hypoxia following a rapid decompression from a cabin altitude at or above
20,000 ft so that a backup supply of 100% oxygen must therefore be supplied.
OXYGEN DELIVERY
General
15. From whichever source the oxygen derives, the easiest way in which it can reach the aircrew
is by a Continuous Flow System. Since the flow does not vary with the demand of the user, such a
system tends to be inefficient and wasteful. This disadvantage is more thoroughly overcome by
Pressure Demand Systems in which the flow of gas from the regulator varies directly with the
inspiratory demand of the user. In addition, the extra facilities required (Airmix, Safety Pressure,
Pressure Breathing etc) can be provided.
16. Most aircraft use pressure demand systems, and the principles underlying the design and
function of pressure demand regulators are essentially the same whether the regulators be panel-
mounted, man-mounted or seat-mounted.
Panel-Mounted Pressure Demand Regulators
17. The regulator consists of a demand valve, which incorporates a pressure reducing valve, a
breathing diaphragm and a lever mechanism. This is shown diagrammatically at Fig 18-3. When the
user breathes in, a fall in pressure in
the mask is transmitted to the
regulator where the reduction is
sensed by the breathing diaphragm.
The diaphragm moves inwards and
causes the lever mechanism to open
the demand valve. When the user
breathes out, pressure builds up in
the regulator as oxygen continues to
flow into it but is not demanded, the
diaphragm moves back and the
demand valve closes. The regulator
also includes refinements in the form
of Automatic Functions and Manual
Selections. Fig 18-3: Oxygen Pressure Demand Regulator
18. The Automatic functions are:
(a) Airmix. In order to deliver air which is progressively enriched with oxygen on
ascent, a venturi tube is fitted downstream of the demand valve. Opening into the venturi is a
passage linked to a chamber which incorporates an aneroid capsule and a non-return valve.
As oxygen flows through the venturi at high velocity a fall in pressure causes cabin air to be
sucked through the chamber and passage. Air, mixed with oxygen, is thus delivered to the
user: (Airmix). As altitude is increased the aneroid capsule expands, gradually closing off the
orifice and so reducing the amount of air mixing with the oxygen. There is a progressive
increase in the concentration of oxygen reaching the user until, at about 30,000 feet, 100%
oxygen passes to the mask, the orifice being completely shut.
(b) Safety Pressure. At cabin altitudes above 8,000 feet the risk of hypoxia as a result
of inward leaks in the system (especially with an ill-fitting mask) is prevented by Safety
Pressure. This is produced by applying a spring force of 2 mm Hg to the underside of the
breathing diaphragm. This opens the demand valve until an equal pressure is built up within
405 Oxygen Systems
the system to overcome the spring. The pressure within the mask is thus kept above ambient
throughout inspiration. The spring is prevented from acting on the breathing diaphragm by an
aneroid until the cabin altitude exceeds safety pressure height, a height which varies from
regulator to regulator.
(c) Pressure Breathing. Positive pressure breathing above a cabin altitude of 40,000
feet is achieved by applying a further spring force to the underside of the breathing diaphragm.
It is prevented from acting below 40,000 feet by a pressure breathing aneroid which encloses
the safety pressure capsule. At 40,000 feet the pressure breathing aneroid allows further
expansion of the inner aneroid and so a large force is applied to the diaphragm. This force is
related to cabin altitude by further gradual expansion of the pressure breathing aneroid. The
regulator will provide protection to an altitude of 50,000 feet at which time it will be delivering
30 mm Hg positive pressure to the user.
19. The indications of flow and contents are:
(a) Flow Indication. Tappings taken from both sides of the venturi (upstream and
downstream) allow the variations in pressure to operate a flow indicator.
(b) Contents Indication. A remote oxygen contents gauge is connected to the output
line of the cylinders or LOX container. The gauge is operated by oxygen pressure but is
calibrated in quantities expressed as a fraction of FULL.
20. The switches on the regulator are:
(a) ON / OFF Lever. The ON/OFF lever is normally wire-locked in the 'ON' position.
(b) Normal / 100% Lever. The Normal/100% lever allows 100% oxygen to be delivered
at all altitudes by blanking off the air entry port of the Airmix facility.
(c) Emergency / Press to Test Mask Toggle. The Emergency/Press to Test Mask
Toggle when deflected to right or left allows delivery of an additional 4 mm Hg pressure at all
altitudes, thus providing safety pressure (eg when toxic fumes are present in the cabin) or a
low pressure test of the mask seal, (mask toggle "up"). When pressed in it delivers oxygen
under a pressure of approximately 30 mm Hg and so provides a high pressure test of
connections and mask seal, (mask toggle down). This facility can also be used in flight to
attempt to blow debris off the mask inlet valve.
Emergency Oxygen (EO) Systems
21. A supply of emergency oxygen is available to each crew member should the main supply fail
(the EO is operated manually) or should ejection or bail-out be necessary (the EO is operated
automatically). Two principal forms of EO assembly are briefly described:
(a) Continuous Flow Emergency Oxygen Assemblies. In Continuous Flow EO

Assemblies oxygen is stored as a gas in a cylinder mounted on the ejection seat. It is
connected to the user via an oxygen flow regulator mounted on the cylinder head and a soft
rubber delivery tube. Once operated, the oxygen is supplied continuously at a rate of
approximately 12 litres (NTP) per minute initially (thereafter declining exponentially) and
provides a useful duration of about 10 minutes. The flow is modified by an Inlet Warning
Connector, which is fitted to the end of the mask hose, and serves to warn the user of a
disconnect in the main oxygen supply line.
(b) Demand Emergency Oxygen Assemblies. Oxygen for Demand EO Assemblies is

stored as a gas in cylinders mounted either on the ejection seat or, when for use by aircrew
without such seats, in the parachute pack. It has a contents gauge which is usually
connected directly to the cylinder, but in some cases it is mounted elsewhere on the seat in a
position where it is more easily seen by the occupant. Once initiated by the release
mechanism, the oxygen flows through a pressure-reducing head on the top of the cylinder
and thence, at a nominal pressure of 50 psi to a regulator. In the case of aircraft which use
man-mounted or seat-mounted primary regulators, the EO passes to the seat portion of the
PEC and thence to the primary regulator. This then controls a flow to the user with its own
delivery characteristics, including Safety Pressure and Pressure Breathing. The duration of
demand EO systems depends upon rate of the user: usually its duration is of the order of 10
minutes.
Passenger Oxygen Systems
22. The Ring Main System. In passenger carrying aircraft the primary protection against
hypoxia is cabin pressurization. The oxygen systems installed in such aircraft are designed to provide
emergency oxygen for the passengers and crew in the event of pressurization failure, or for
therapeutic purposes. Oxygen for these systems is usually stored as gas although liquid oxygen is
used in some aircraft. The high pressure supply is reduced by valves in the normal way before
passing to a ring main circuit for passenger supply or to the presssure demand systems usually fitted
on the flight deck for crew use. A Ring Main system is shown diagrammatically at Fig 18-4.
23. During normal flight, oxygen is supplied from the aircraft storage system to the passenger
oxygen regulator. In the event of cabin pressurization failure, and when the cabin altitude exceeds a
pre-set level (usually 10,000-14,000 feet) the regulator automatically raises the supply pressure to
approximately 80 psi (Emergency). This increased pressure activates a warning horn and its delivery
to the ring main operates an actuator in each mask presentation unit, causing the masks to ‘drop
down’ in front of the passengers to a position from which they can be applied to the face. A
continuous flow of oxygen at emergency pressure emanates from each mask, once its check valve is
released, and is maintained as long as the cabin altitude remains above 17,000 feet. When the
407 Oxygen Systems
aircraft has descended to a cabin altitude of less than 17,000 feet the control unit reduces the delivery
pressure to ‘Normal’. Flow is maintained at a reduced level and each mask then functions as a
demand type.
Fig 18-4: Passenger Ring Main System


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FIS BOOK 3

Chapter Title Page No.

3. Aileron. A horizontal control surface usually hinged

4. Aircraft. Any weight carrying device designed to be

5. Airframe. An aircraft without its engine/engines.

12. Attitude. Inclination of the axes of an aircraft relative to the horizon.

14. Axis. Imaginary lines about which the aircraft rotates.

20. Camber. The curvature of a surface of an aerofoil.

21. Cantilever. A spar or beam fixed rigidly at one end only.

36. Empennage. The tail section of the aeroplane and

37. Fairing. Additions to any structure to reduce its

Fig 1-2: Vertical Stabilizer

41. Flutter. A violent, destructive vibration of the

42. Formers. Transverse members which are

45. Frames. Transverse structural members of the

46. Fuselage. The main structure or body of most

47. Glider. A non-power driven heavier than air

54. Leading Edge. The forward member of an aerofoil.

63. Pitching. Angular motion about the lateral axis.

68. Rolling. Angular motion about the longitudinal axis.

77. Spar. A principal longitudinal member of a plane or of a control surface.

86. Trailing Edge. The aft edge of a streamline body or aerofoil.

88. Truing. The adjustment of a component to its correct shape.

Major Structural Stresses

(b) Compression. Compression is the stress

(c) Torsion. Torsion is the stress that Fig 2-2: Compression

(e) Bending. Bending stress is a

(Fig 2-6) and is usually constructed of steel tubing

6. Monocoque Type. The monocoque (single

7. Semi-monocoque Type. To overcome the

members. Different portions of the same fuselage

8. The stringers are smaller and lighter

Fig 2-10: Internal Wing Construction

19. Honeycomb Construction. It is an efficient way of stabilizing / stiffening relatively thin

21. An engine nacelle or pod consists of skin,

23. The empennage is also called the tail

24. Other components of the typical

25. Stress in an empennage is also carried Fig 2-15: Construction Features of

Flight Control Surfaces

DESIGN AND CONSTRUCTION

Wire Braced Structures

Unbraced or Cantilever Monoplanes Fig 3-2: Cessna-152

OPERATING CONDITIONS FOR AIRFRAMES

Static Strength and Stiffness

9. When structures incorporate fibre-reinforced composite materials, design calculations and

weight saving. Many small parts are

24. Disadvantages of Composite Materials. Current composite materials have inherent

(b) Their use is limited to ambient temperatures below about 1200C.

(c) Unless adequately protected, they tend to absorb atmospheric moisture.

Stability and Loads and Release and

Table 3-1: Features of External Store Positions

(a) Low stress levels due to pressurization loads.

(a) Cabin length and width (i.e. usable floor area).

(b) Freight volume.

STRUCTURAL INTEGRITY AND FATIGUE

(b) Operational Effectiveness. Aircraft must remain operationally effective with

(a) Overstressing. When loads in excess of design limitations are applied to an

(c) Corrosion. Corrosion is a chemical reaction occurring in metals exposed to water

Establishing Fatigue Life