STRUCTURAL STRENGTH OF AIRCRAFT

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• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP81-01043R002300220004-4
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• Original Classification: K
• Document Page Count: 140
• Document Creation Date: December 23, 2016
• Document Release Date: September 11, 2013
• Sequence Number: 4
• Content Type: REPORT

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lifirISUIT1011
STRUCTURAL STRENGTH OF AIRCRAFT
(Prochnosti Samoieta)
By S. N. Kan ?
STATE PUBLISHING HOUSE FOR THE DEFENSE INDUSTRY, MOSCOW
1955
Pages 155 - 208, 217 - 264
STAT
STAT
PREPARED BY
TECHNICAL DOCUMENTS LIAISON OFFICE
MC LTD
WRIGHT-PATTERSON AIR FORCE BASE, OHIO
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Table of Contents
Page
Chapter VIII- The Controls 1
Section 31. Principal Systems of Basic Control 4
Section 32. The Principal System of Control Cables 10
Section 33. Hydraulic Aids 15
Chapter IX - The Fuselage 23
Section 3i;.. External Loads of the Fuselage 25.
Section 35. Fuselage Design 29
Section 36. Design of a Semimonocoque Fuselage 38
Section 37. Design of a Truss-Type Fuselage 47
Section 38. Pressurized Cabins 48
Section 39. Ejection Seats for the Crew 59
Chapter XI - Landing Gear 71
Section 43. Basic Requirements of a Landing Gear 76
Section 44. Bicycle Type Landing Gear Diagram 78
Section 45. Landing-Gear Wheels 80
Chapter XII - Shock Struts 85
Section 46. Designation of Shock Struts 85
Section 47. Function of an Oleo-Pneumatic Shock Strut 88
Section 48. Influence of the Forces of Friction Produced
by the Packing Collar on the Work of the Strut 94
Section 49. Incorrect Fillingsof Shock Absorbers 98
Section 50. Shock Strut Design 103
Chapter XIII - Performance of a Landing Gear Energy
Diagram 111
Section 51. External Loads Acting on a landing Gear 112
Section 52. Cantilever landing Gear 117
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Section 53. Semicantilever Landing Gear 127
Section 54. Landing Gear with a Lever Suspension
of Wheels
ii
330
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CHAPTER VIII
THE CONTROLS
The controls* of aircraft consist of mechanisms, rockers, riushZandLpull rods,
etc., by which the pilot controls the rudder, ailerons, trim tabs, flaps, wing flaps,
brakes, the raising and lowering of the landing gear, the propulsive units, etc.
The entire complex of aircraft controls is easily divided into two groups. The
first group includes the control by ailerons, elevator, and rudder. This control is
called "basic control". The second group includes the remaining controls, i.e.,
control of the trim tabs, wing flaps, flaps, etc. which is called "supplementary or
auxiliary control".
In this way, using basic control, the pilot has the possibility of generating
or eliminating the moment of the airplane relative to its three axes in space. The
pilot himself is in a position to maintain or change the flight path.
Basic control, in turn, takes the form of "manual control" which is carried out
by manual action and "foot control", carried out by foot action.
Hand control includes control of ailerons and elevators, and foot control cov-
ers control of the control surfaces.
The concentration of control of the ailerons and elevators in one unit, as is
done in all modern aircraft, has been in effect throughout the history of aviation.
* According to the writings in the contemporary chapters of the widely used book by
esi?n of Aircraft Parts", Oborongiz, 1947.
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ly that control of the aircraft requires only a minor effort. The pilot, carrying
out basic control, must sense this control and must smoothly increase the pressure
proportionally to the deflection of the control surfaces from the neutral position,
in a direction opposite to that of the deflection. For this, the controls must have
iltendencv-to return to the neutral position under the action of aerodynamic---
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-!This system of control was worked out by the Russian inventor, Ullyaninyi in 1906.
2_J
To simplify control of the aircraft, the direction in which the-hands and feet
. are moved is coordinated with the instinctive motions of man. Thus, creating a
- diving moment of the aircraft, the pilot instinctively strives to move his entire
- body forward. Therefore, his stick (col) is also moved forward, which deflects
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the elevator downward. This leads to the obtainment of additional lift on the
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?horizontal tail surface; the lift produces a diving moment along the arm leading to
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the CG of the aircraft and reduced the ankle of attack of the wing.
If it is required to bank +urn thb aircraft, the Valet pushes the atiek
_sideways or turns the control wheel. This action leads to lowering of the aileron
on one wing and raising an the other wing, thus producing the necessary banking
1) 4 I
moment of the aircraft.
The action of the rudders, i.e., a turning of the aircraft relative to the ver-:
'tical axis, is accomplished by leg control. If the aircraft is to turn to the right',
the pilot depresses the right pedal with the foot. This causes the rudder to deflect
to the right, which leads to a generation of lifting force on the vertical tail eur-i
faces, effective toward the left. This force creates a moment relative to the verti:-
cal central axis of the aircraft, which turns the aircraft to the right.
The quality of this or any other system of control is determined primarily by
-.the rapidity of action and ease with which the pilot is able to maneuver the air-
- craft relative to any spatial axis. Controls which meet these requirements will re-
,
dude pilot fatigue and give him an opportunity to devote his entire attention to
fulfilling the basic mission. From the above, however, it does not follownecessari-
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The ease of control depends greatly on the friction in the control-cable hinges.
In order to reduce this friction, the control mechanisms in modern aircraft are fit-
ted with ball bearings. Friction in hinges, backlash, and deformation in the con-
trols produce a retardation effect. This effect results in the fact that between
the moment of application of pressure to the controls and the incipient deflection
of the control surfaces, a certain time will elapse, a measurable part of a second.
It is obvious that the occurence of this lag is harmful, especially in flight close
to the ground or in flight in bumpy air.
The design of the basic controls must meet certain requirements. These demands
are common to all types of aircraft and include: sufficient mechanical strength and
rigidity, light weight and resistance of projecting parts in contact with the air
flaw, convenience of use and long life, simplicity of production and repair, etc.
Concerning specific requirements, in relation to controls, the following basic
specifications should be mentioned:
1. The control rods must have such geometric dimensions that the frequency of
their vibration is outside the sphere of the revolutions and even double the rpm of
the propulsive unit (propeller, turbines, etc.) i.e., they must be out of the way of
resonant vibrations. The determination of the frequency of free vibration of the
control rods is described in Chapter XIV.
2. The design must guarantee the independence of action of the ailerons and el-
evators, i.e., the operation of the control stick in the longitudinal direction must
not cause deflection of the ailerons and an inclination of the stick to the side
must not deflect the elevators.
3. Possible deformations of the wing or the fuselage must not lead to fouling
of the lines or control mechanisms.
4. The design of the foot controls must allow for the regulation of the posit-
ion of the pedals at the level of the pilot.
5. The angles of inclination of the control mechanisms, the angles of inclina-
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tion of the controls, the loads on the control levers and control pedals, and the
ruggedness of the control lines etc. must satisfy the strength standards.
Section 31. Principal Systems of Basic Control
All modern aircraft have two accepted systems, differing only in the method of
controlling the ailerons. In one system, control of the ailerons is achieved by in-
clination of the lever (the stick) to the sides, and in the other control systems,
the ailerons are actuated by the rotation of the control wheel relative to the axis
which is applied to the column. For accomplishing this, the column can be displaced
4)11:Py_in_the plane of symmetry of the aircraft.
The first as well as the second control system may be either single or dual.
In the dual system there may be two sticks, two wheels, or a combination of a stick
and a wheel. Often the second stick is removable so as to relieve the copilot when
he is performing duties not immediately connected with piloting the aircraft.
The individual control mechanisms, levers, and tubes usually are installed un-
derneath the pilot's seat, in order not to impede the work of the pilot in the cabin.
This also protects ball bearings and hinges from damage.
The selection of the type of mechanism and distribution of control levers re-
lative to the stick depends on the layout of the cabin and wings. In a cabin placed
rearward relative to the wing, the aileron control levers are placed in front of the
stick. If the cabin is located ahead of the wing, the aileron control levers and
the elevator control levers are placed in back of the stick. For circumventing var-
ious obstacles, the elevator control levers sometimes are installed at the side of
the fuselage as are all other control cables.
Control by the Stick
In the case of a cabin located behind the wing, the ailerons are controlled by
a torque tube (1) (Fig.171a) having an axis of rotation ab with one radial and an-
other radially resisting ball bearing. On the tube a lever (2) is mounted, to which
a rod (3) for aileron control is attached. The stick (4) rotates in a longitndinAl
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r.
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direction relative to the axis cd and in a lateral direction, together with the tube
(1), relative to the axis ab. The front of the rod (5) for elevator control is at-
tached to the stick, and the rear to the rocker (6).
The independence of action of the rudders and ailerons is obtained by placing
the upper end of the rocker (6) on the extended axis ab. On deflecting the stick
sideways (for aileron control) the tube (5) describes the exterior of a cone so that
the cable to the elevator (7) is not displaced in a longitudinal direction. Moving
Fig.171 - Stick Control
a - Cabin located behind the wing; b - Cabin located in front of the wing
the stick only in longitudinal direction does not produce a rotation of the tube (1)
so that the aileron cable does not function.
If the cabin is located in front of the wing or directly below it (Fig.171b),
the aileron control lever is shifted toward the rear, behind the stick. The rod (3)
to the ailerons is underneath the rod (5) going to the elevators. To ensure inde-
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'pendence of function of the ailerons and rudders in cabins placed far ahead of the
wing, the control lever is designed in the form of a small triangle (21) with connect-
ing hinges on the aileron control rods.
The rod (5), at the connection point with the stick (4) and the rocker (6), is
provided with spherical joints or universal joints. For lower angles of deviation,
double-aligning ball bearings are used.
Fig.172 - Stick Control with Torque
Tubes
a) To elevator; b) To aileron
In some aircraft (Fig.173) the
lateral directions cd
rection, the stick is
In cases when the control cables to
the elevators do not pass through the
plane of symmetry (Fig.172), control is
transmitted by means of a torque tube,
for control of both ailerons (tube 1) and
elevators (tube 2). Independence of ac-
tion of the rudders and ailerons is a-
chieved if the rotation of the rod (3)
is laid through the axis Oa and the ro-
tation of the rod (2) and the stick
through the axis cd.
axes of rotation of the stick in longitudinal and
and ab are manipulated by elevation. In the longitudinal di-
rotated in relative to the axis cd and in a transverse direc-
tion, together with the tube (1) or the aileron controls relative to the axis ab.
The bracket (2) to which the stick is mounted, compensates through the cut-out in the
tube (1), for the gap in the lower end of the stick. The forward end of the tube (1)
carries the lever (3) of the aileron control. The independence of action of the rud-
ders and ailerons is achieved by placing the axis of the attachment joints of the
rod (4) to the lever (5) of the elevator control, on the axis ab or on its.extension.
Wheel Control
The control wheel (1) shown in Fig.1 is located on the upper part of the column
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(3) and is rotatable. The wheel is rigidly coupled to the cogwheel (2) which carries
a bicycle?type chain. To the chain are attached two cables (5) which, by turning the
wheel and thus passing the chains over the cogwheel, actuate the rods (or cables)
going to the ailerons. The column may be swung forward or backward on ball bearings
Fig.173 ? Stick Control with Transmission by
Raising the Axes of Rotation
(4) thus actuating the system of rods (or cables) of the elevator control.
In this mechanism, full independence of action of the rudder and ailerons is
maintained.
Foot Control
The basic elements of foot control are pedals and mechanisms, which connect
with the rudder control cables.
The dimensions of the pedals are determined by convenience of foot placement,
these dimensions usually being about 140 mm width and 280 mm length.
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The feet of the pilot move together with the pedals, through which action for-
ces are exerted both in the horizontal and vertical planes. For ease in operating
the pedals, the distance between them must be minimal. In different aircraft, this
spacing varies between 250 and 500 mm
and is determined by the type of manual
controls and the design of the pedal
mechanisms.
Displacement of the pedals in the
horizontal plane is achieved by means
of a hinged parallelogram (Fig.175) or
by linear float guides (Fig.176).
The hinged parallelogram is simple
in design. It consists 'of two cross
bars (1) and (2) which are attached
over hinges (Fig.175) to the longitudi-
. nal elements (3). This hinge mechanism
is actuated over the pedals (4), attach-
ed to the longitudinal elements (3).
The pedals can be regulated under-
neath the pilot's seat.
The fixed points of the entire me-
chanism are the hinges (5) and (6) and the cross bars (1) and (2).
To the cross bar (1), the bracket (7) is welded. To this a sectional cable
guide (8) is mounted, for controlling the rudders. The strap (9) serves for attach-
ing the feet of the pilot to the pedals.
Because of the design of the front sections of the hinged parallelogram, the
pedals are shifted gradually without turning. This in itself facilitates the con-
trol and prevents the feet of the pilot from slipping off the pedals.
Fig.174 - Wheel Control
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The design of the linear float guides (Fig.176) for progressive shifting of the
pedals can be obtained by the following system: The pedals are attached to the brack-
et (1) and, by means of the adjusting sliders (2), can be shifted freely along the
Fig.175 - Foot Control over a Hinged Parallelogram
linear guides (3). The pedals, over the bracket (1) and the sliders (2) connected
with the cable (4), control the deflection of the control surfaces. The sliders (2)
2
In this, the pedals attached to the le-
Fig.176 - Foot Control by
vers describe an arc of a circle, which
are interconnected by the cable (5)
over the pulley (6) to achieve an integ-
ral cable line.
The displacement of the pedals in
the vertical plane is usually accom-
plished with the help of rocking levers.
Linear Float Guides
changes their inclination. In some air-
craft, the pedals are connected to hinges or joints forming a sort of parallelogram,
which permits a displacement of the pedals without changing their inclination.
Figure 177 shows a system of lever attachment to the upper and lower grouping
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of the pedals. Here, the pedals are hinged to the levers, which are able to rock re-
lative to the fixed axis ab. The motion of the pedals is further transmitted over
the sector (1) to the cable (2). The sector (1) is connected to the cable (3) over
the pulley (4).
Fig.177 - Foot Control with Levers, Swinging in a Vertical Plane
Section 32. The Principal Systems of Control Cables
The cables of both hand control and foot control may be rigid, flexible, or mix-
ed. In rigid control cables, the forces are transmitted by means of tubes which
work by extension and contraction. Flexible control lines consist of cables, wires,
or belts which transmit forces only by extension. Finally, mixed control lines com-
prise portions of both rigid and flexible types.
The rigid cables (tubes) have the following advantages: Tubes, working with lit-
tle stress under compression, deform less and for that reason are more advantageous;
in addition, they do not contribute to the generation of self-oscillations (flutter)
of the wing or tail group (see Chapter XIV). Here, friction in the control lines
can be reduced to a minimum by having all joints work on ball bearings, which in ad-
klash. Rigid control lines result in longer life of the cIsTAT1s.
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With respect to elevator control, separate rods located in the wing can be used
with the possibility of controlling the aircraft by only one aileron.
In listing the drawbacks of rigid control cables, the difficulty of "by-passing"
equipment, complexity of design due to the great number of joints, and excessive
weight of the entire control system. In addition, it is of importance that no res-
onance of the tubes with the power plant occurs.
Flexible control lines, due to their design, have small dimensions, light weight,
and readily permit "by-passing" installations.
One drawback of flexible control lines to be mentioned is a low mechanical
strength, leading to deformations of the Gables which are harmful since they may gen-
erate self-oscillations (flutter) of the wing or tail unit. In actual operation, de-
spite the fact that they are pre-stretched when being installed, flexible control
lines are less long-lived, and considerable friction occurs in the transmission. Be-
cause of their constant bending on the pulleys, the cables wear out rapidly.
The possible systems of transmission of manual stick control are presented in
Fig.178. The control stick (1) (Fig.178a) in its function as a fixed point has a
horizontal joint (3) attached to the fork of the horizontal tube (2). This tube (2)
is supported by the radial bearing (4) and by the radial thrust bearing (5). Push-
ing the stick (1) forward and rotating the tube (2) actuates the aileron control le-
vers (6). Through this action the elevators remain in neutral, so that the upper
joint of the first rocker (11) coincides with the longitudinal axis of the tube (2).
Consequently, pushing the stick (1) forward causes the rod (7) to describe a
cone and the elevator to remain stationary. Deflection of the elevator is achieved
by pushing the stick (1) toward or away from the pilot. This produces a displace-
ment of the lower end of the stick either forward or backward and causes a displace-
ment of the rods (7), (8), (9), and (10) or (100 as well as of the lever (14) at-
tached directly to the elevator. These rods rest on the rockers (11), (12), and
(13). In presence of the control lever (14) below the rudder (dotted lines) the
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rocket (13) is replaced by an intermediate double-arm lever (15) in the coatrol
system.
In some instances of control transmission, the rocker (Fig.178b) is replaced by
guides (11). When using guides, the transmission system has less clearance.
The deflection of the ailerons (Fig.178c), as shown above, is achieved by push-
ing the control stick (6) forward together with the rods (17), (18), (19), and (20)
// // //
o
r.-40
/0' C=0=1
1425a) /27-1H==.77-T-%
7
C:=216m4
Fig.178 - Control Lines for Manual Stick Control
which rest on the rockers (22) and (23). The forward shift of the rods causes rota-
tion of the double-arm levers (24) and (25) which are connected to the rods (16) and
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The above is a consideration of the system with rigid controls. If the controls
are flexible (Fig.178d), the rod (7) is mounted to the double-arm lever (27), to
which cables, lines or, belts going to the elevators are attached. These are sup-
ported along their length by guides (29). Turning the control lines is done with the
help of the pulleys (32) (Fig.178c) but only in the case of cables. For deflecting
Fig.179 - Control Cables for Manual Wheel Control
the ailerons (Fig.178c), the sector (26) is used instead of the lever (6). Since
the cables in flexible control lines work only by traction, double-arm levers (27)
and (28), forming a locking chain, are attached to the elevators (Fig.178d) and the
ailerons (Fig.178f). In Figs.178e and 178g, a mixed system of control lines is
shown. In this, auxiliary elements are the double-arm lever (30) and the triple-
arm lever (31).
Control cables controlled by a wheel considerably reduce the stress on the ail-
erons, thus increasing the rpm of the wheel. In this case, full deflection of the
ailerons requires more time than does stick control. For this reason, wheel control
is not used in manuvering aircraft. The advantage of this type of control lies
only in the transfer of stress from the wheel to the aileron control cables. Ph?
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central control mechanism (Fig.179a and b) consists of an articulated tube (1) to
which the wheel (3) and the cogwheel (4) rotating relative to the axis (2) are at-
tached. Control of the elevators is carried out by the lever (5) attached to the
horizontal part of the tube (1). The tube (1) can be turned only relative to the
joints attached to the brackets (6) with the brackets (7). Aileron control here is
also carried out by cables attached to the column by a chain (8) placed over the cog-
wheel. Turning the wheel (3) rotates the cogwheel (4), thus actuating the cable
which, over a system of pulleys (9) and (10) (Fig.179b), goes to the ailerons. The
I \ / \
!-I \ 6)/ \
(4--
- 1 ?
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ii
Fig.180 - Foot Control Cables of the Hinged Parallelogram Varient
a - From the side; b - From above
pulleys (9) and (10) is so arranged that the cables pass through them along the ax-
is ab, which is the pivot of the column. This ensures independence of action of the
rudders and ailerons. The three-arm lever (11) ensures the transmission of stresses
from the cables to the rods (12) which are directly attached to the ailerons.
The control cables of the foot control can be rigid, flexible, or mixed. Flex-
ible and mixed cables are more common since they have less weight, require less
space, and permit easier installation and fitting. As described earlier, the pedals
of the centrol foot-control mechanism may be in the form of a horizontal parallelo-
gram (Fig.180) or may be displaced forward (Fig.181) along guides (2). In the first
instance, for reducing the load on the feet, the cables (1) (Fig.180) are attached
less to the lever arms than to the pedals (2). The rudder control cables are con-
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nected directly tothe.iW7o le'VWS-(5)-(Figaibi:d)-or to the sector (3)-(Fig.180a1b),
independent of_the.position of the rudder.iwith_respect to the fuselage.
In the second variant, the pedals (Fig.181) are connected to a lug (5) over a
,
cable (3) which turns on pulleys (4). Direct control of the rudder is achieved over
-..1the double-arm rocker (8), which is turned by the cables (7). The levers (10) of the
....irudder control are connected to the rocker (S) by rods. The kinematics of the rudder
1
.control are satisfied by only one rod (9).
1
_
, the following. Through the control elements, the pilot controls the hinge moment of
the rudders (ailerons) Mh which is well defined by the equation
a) 1-41
4 '3 5
4-- '&43)- I
1 /1
b\ ? ? ? ?
Fig.181 - Foot Control Cables of the Guide Variant
Section 33. Hydraulic Aids
Hydraulic aids or boosters take the form of special mechanisms which relieve
the pilot from excessive burdens of control. However, this does not eliminate the
feeling of controlling the aircraft.
The necessity of hydraulic boosters in aircraft control systems is explained in
U2
v
= ChOp
?
, --where ch coefficient of hinge moment of the rudder or aileron;
5 .
5 s a area of the rudder surface;
,
15
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p = air density;
v flying speed;
b = chord of the rudder or aileron.
At low speeds, the hinge moment Nh is lowered by changing the coefficient ch by
'producing aerodynamic balance of the rudders or ailerons (cf. Section 18). For ex-
ample, in axial aerodynamic balance, the hinge moment is reduced by displacing the
axis of rotation of the rudder along the chord backward, nearer to the center of pres-
sure. In heavier aircraft (increased S and b) and in high-speed aircraft (increas-
ed IT), it is impossible to facilitate control by aerodynamic compensation since un-
desirable overcompensation might occur. In addition, in flights approaching the
speed of sound, a sharp increase, in the hinge moment might take place, due to signi-
ficant change in the position of pressure of the control surfaces.
Hydraulic boosters in the control system permit meeting the requirements for
sensing and control at times when the pilot has to apply only slight pressure to the
controls.
Such boosters also make it possible to eliminate aerodynamic compensation, in-
cluding internal compensation, which is very important for high-speed airplanes with
thin airfoils.
It should be mentioned that hydraulic boosters also act as dampers for quench-
ing oscillations generated by rudder flutter.
In modern aircraft, an attempt is made to place the hydraulic boosters as near
the rudders and ailerons as possible so that the greater part of the control cables
can be relieved.
This leads to a lowering of structural weight and a decrease in
friction on the joints and pulleys of the control lines. To ensure independence of
function of the controls a self-supporting hydraulic system is usually installed.
The hydraulic boosters are applied easily and quickly according to the wishes of the
pilot, resulting in direct control of the aircraft. The presence of hydraulic boost-
ers does not disrupt control and does not cause a lag in the deflection of the rud-
. STAT
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_iders, which is done by the stick.
Hydraulic boosters of two types exist; reversible and nonreversible.
In the reversible system of hydraulic boosters, the deflection of the control
surfaces, basically, is produced by the hydro-system; a small portion (from 1/30 to
1/6) is accomplished with the help of pilot's forces which constitute the constant
percentage of the full necessary stress. The more common tractive forces for rudder
10
deflection and the portion contributed by the pilot, can be coupled by means of lev-
ers r special pistons of the hydraulic system. In the first case a hydro-system
_with mechanical coupling and in the second case one with hydraulic coupling is in-
18
volved.
The nonreversible system of hydraulic boosters is characterized by the fact
) _I
that the entire hinge moment is compensated by the hydro-system. Here, the sensing
of control is ensured by a spring, whose tension increases on deflection of the
_ .stick. Hence, independently of the flying speed, the pilot must apply the same
pressure to the control stick, which can change only with respect to the angle of
- deflection of the rudder.
Booster Diagrams
The principal diagram of a reversible hydraulic booster system with hydraulic
coupling is shown in Fig.182. The control system includes: the slide valve (3),
the booster (4), induction pipe and exhaust manifold, and the piping (9) and (10)
.with the cock (11).
The diagram shows the system in the neutral position, when the fluid from the
induction pipe flows back freely to the tank through the central groove in the cen-
ter piston of the slide valve, the piping (10), and the exhaust manifold.
In the working position, for example, while pushing the control stick (1)
,(forward), three slide-valve pistons, rigidly interconnected, are shifted forward
with the help of the control rod (2). This causes displacement of the fluid from
ugh the piping (9) and the cock (11) into the right half STAT
17
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.the booster cylinder (4) and into the left part of the coci-Of-ihe
cylinder (3). By thiS action, the pipe (12), over the slide valve, connects the .
a)
b)
10
2 8 a-
12
Fig.182 - Reversible HydrauliclUmiAr System, with Hydraulic Connection
a) Exhaust; b) Induction pipe; c) Center piston of valve
left part of the booster cylinder with the exhaust manifold. Consequently, the
'force P transmitted through the rod (5) to the control surface (6) will produce
pressure to the right on the booster piston (4) and pressure to the left on the
cylinder bottom of the slide valve.
(87)
where p is the pressure of the fluid into the induction pipe, in kg/cm2;
S is the area of the booster cylinder;
Sz is the area of the slide valve.
The control rod (2), tightly connected with the slide-valve cylinder (3), will
measure the force of pressure on the left cylinder.
pSL. (88)
On dividing the left and right-hand sides of eqs.(87) and (88), the following
relationship is obtained.
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18
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(89)
which shows that the greater the force exerted on the rod Pp the greater will be the
force PT on the hydraulic booster.
Fig.183 - Reversible System of Hydraulic Booster, with Mechanical Connections
a) Induction pipe; b) Exhaust
The force exerted on the control rod P overcomes the hinge moment of the con-
trol surface, deflecting the latter through some angle. This deflection ceases as
soon as the control stick is shifted.
In the same way, the actuation of the rod (2) and the slide-valve piston (3) is
stopped. The booster cylinder (4), actuating the slide-valve cylinder, closes the
induction pipe and thus stops the boosting.
By turning the cock (11) to the position (I), it is possible to control the
rudder without boosting. By turning the cock to position (II) it can be used for
locking the control surfaces, thereby preventing displacement of the booster pis-
ton (4).
The slide-valve piston (3) and the booster piston (4) are connected by rods
which slide along the guides (8).
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The booster cylinder (4) is rigidly attached by means of bolts (7).
Figure 183 presents the reversible system of hydraulic boosters with mechanical
connections. The auxiliary elements here are: the double-arm rocker (3), the hinge
attachment to the control lever (4) of the servorudder, the slide valve (5), the
cylinder which is rigidly attached to the rudder (6), the booster (7), the cylinder
which is hinged to the body of the aircraft, the pipeline system (9) and the induc-
tion and exhaust pipes of the hydraulic-booster system (8) with flexible hose. The
hydraulic-booster system is illusLraLed in tne neutral position, in which the fluid
from the induction pipe flows freely back into the tank through the central groove
of the slide-valve piston and the exhaust manifolds.
In the working position, for?example, on pulling the control stick (backward),
the control rod (2) displaces the slide-valve piston (5) to the left. By this, the
exhaust line is separated from the induction lines and the latter will dispense
? fluid under pressure into the right part of the booster cylinder (7) which causes
its piston to be displaced to the left. The fluid from the left part of the booster
cylinder will flow freely into the exhaust manifold through the pipe (9) and the
slide valve (5).
From the condition of equilibrium of the double-arm rocker (3) with the arms
(a) and (b) it follows that
Pp ,= b + a
PT a
b + a a
Consequently: Pp = Pm and PT = PP b + a
1 a
(90)
If the pilot stops pulling the stick, the slide-valve piston (5) will become
stationary and its cylinder, connected rigidly with the rudders, will shift to the
left and thus disconnect the induction pipe from the booster cylinder (7).
A consideration of the system of hydraulic boosters with mechanical connection
shows that it starts o .erating immediately, without jars, and has several advantagSTATub
?
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over a system of hydraulic boosters with hydraulic connections. One of these is the
relative ease of changing the arms of the double-arm rocker (3), giving the possi-
bility of changing the ratio (90). In the latter system (Fig.183), friction on the
booster cylinder (7) does not influence the force PT on the control lines. Besides
this, in operation of the aircraft, less rigid requirements as to tightening of the
slide valve (5) are possible.
The systems of nonreversible hydraulic boosters are shown in Figs.184 and 185,
resembling the systems shown in Figs.182 and 183. In Fig.1841 a displacement of the
stick (1) causes a displacement of the rod (2) and of the pistons of the slide valve
(3). As a result of thi4 the fluid from the induction pipe flows into the booster
cylinder (4) through the pipe (9) and the cock (11) where (in the booster cylinder)
the force P from the piston is exerted on the rod (5) going to the rudder. At this
time, the other cavity of the booster cylinder, through the cock (11), the pipe (9)
and the slide valve, connects with the exhaust manifold. The coordination of de-
flection of the rudder (6) by working the stick (1) is achieved by connection of the
booster cylinder (4) with the cylinder of the slide valve (3).
The pilot senses the force Pp exerted on the control stick by the springs (10).
The piston rods of the slide valve (3) and the booster cylinder (4) slide along
the guides (8).
The booster (4) is connected to the body of the aircraft by bolts (7).
A study of this system shows that shifting into the neutral position is accom-
plished by moving the control stick (1). In the same way, if the stick (1) is mov-
ed, the springs (10) will tend to bring it into a neutral position. This will be
accompanied by a displacement of the slide-valve piston (3), and the fluid from the
corresponding part of the booster cylinder will flow into the exhaust manifold.
In the system pictured in Fig.185, moving the stick (1) causes displacement of
? the rod (2) and of the pistons of the slide valve (5). This causes the fluid from
the induction pipe to pass along the pipe (9) into the booster cylinder (7) where
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pressure by the piston Pp is transferred to the control rods (4). At this time, the
other cavity of the booster cylinder, through the pipe (9) and the cylinder of the
Fig.184 - Nonreversible System of Hydraulic Booster, with Hydraulic Connection
a) EXhause; b) Induction pipe
slide valve (5), is connected with the exhaust manifold (8).
On the rod (2), leading to the slide-valve piston (5), no force or sensing by
means of the springs (3) is present. Here, as in the system in Fig.184, return to
Fig.185 - Nonreversible System of Hydraulic Booster, with Mechanical Connection
a) Exhaust; b) Induction pipe
the neutral position is ensured by moving the control stick (1) and also by the
presence of proportionality of the travel of the stick with the angle of deflection
STAT
of the rudder (6).
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intik
CHAPTER IX
ThT FUSELAGE
The fuselage is usually the base of support for many parts of the airplane:
the wirg, the tail group, the engine nacelle, and the landing gear. Besides this,
,===l2e 'lZw1;;;;Is ?????:.L'i "
*ow,/ ???? ? ?Aft
ak)
Fig.186 - Fuselage with Internal Power Plant
The construction is extended and thus absorbs the fuselage loading
a) Engine
the fuselage houses the crew, the fuel, armament, the power plant, equipment, and
other loads. Hence, the exterior loads of the fuselage are determined primarily by
the forces of mass action of parts within it and by the forces exerted by the in-
dividual parts of the airplane.
From an aerodynamic point of view, the fuselage is the parasite part of the
airplane, since it does not accomplish a lifting force and actually produces drag.
For this reason, the overall size of the fuselage must be minimal and its form
streamlined. In smaller machines, the overall size of the fuselage is determined by
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the size of the engine, the dimensions of the cabin, or the bulk of the armament.
In larger machines, the overall size of the fuselage is determined by the dimensions
of the cargo space, and the crew quarters. Practically speaking, the range of
transverse dimensions is 1 - 3 m or larger.
The installation of jet engines in the fuselage outside of its body raises the
question: should the engine be located within the power system of the fuselage
(Fig.186) or outside it (Fig.187). In the first instance, the fuselage is more
satisfactory from an aerodynamic and mechanical strength viewpoint. The rationality
of arrangement within the power system is demonstrated by the fact that, in the
power system in Fig.1861 the structure of the fuselage is larger in cross section
and, for that reason, better absorbs the external loads.
Fig.187 - Fuselage with Engine Outside the Power System
The structure is less extended and therefore less favor-
ably absorbs the fuselage loading
a) Engine
The length of the tail section of the fuselage is determined by the necessarily
great distance from the center of gravity of the airplane to the tail assembly for
insurance of sufficient stability and control of the airplane. This length fluc-
tuates within the limits of 2.5 to 3.2 lengths of the mean chord of the wing.
The presence of a landing gear with a nose wheel requires lengthening of the
nose section of the fuselage.
ST
Fuselages of high-speed airplanes are characterized by rounded cross sectiAT
vAlo.
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a
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Fuselages of jet aircraft have less free space. Consequently, because of the ?
decreased size of the wings, the fuel usually is placed in the fuselage. The fuel
capacity of jet aircraft is greater than that of piston-engine airplanes. In addi-
tion, on single-engine airplanes the engine is placed within the fuselage, taking up
a considerable portion of the available space. The air scoops greatly influence the
form and design of the fuselage. All this leads to the fact that on jet aircraft
there is less free space within the fuselage.
Section 34. External Loads of the Fuselage
Stresses produced by parts riveted to the fuselage (wings, tail section, power
plant) form the basic active loads of the fuselage. Besides this, the fuselage is
loaded by stresses exerted by the cargo and various units placed within it and by
the weight of its own structure.
The external loads of the fuselage, according to size, placement, and distribu-
tion, are determined by the basic existing norms of stability. The stability norms
demand reliability of the fuselage in all types of flight and in all landing situa-
tions.
The computable cases of the fuselage are denoted as follows.
A A" B C D liz 1-1"
A1, 1, P '1' f
11 .
These cases can be subdivided into two groups: symmetric and asymmetric loads.
The first six instances are already familiar (see Section 7. "Norms of
Stability") and belong to the group of symmetric loads, which are loads placed
symmetrically with respect to the surfaces of symmetry of the fuselage. The group
of asymmetric loads includes and Hlf) (Fig.188).
The case Hf represents the transverse stress on the nose section of the fusel-
age, due to loads placed within it (Fig.188a).
The case Hf represents the transverse stress on the tail section of the fusel-
age, due to loads of the vertical tail surfaces (Fig.l88b).
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?
In the general case of calculating the fuselage it is well to take into con-
sideration all forces coming in contact with it, i.e., the entirety of its equili-
brium, which is much too complicated.
tz
Fig.188 - Case of Asymmetric Load of
the Fuselage
a- Case Hf corresponds to the situa-
tion when the nose section of the
fuselage, under the effect of the force
of inertia of the mass, turns sideways.
b- Case H4 corresponds to the situation
f
when the vertical tail surfaces are de-
flected sideways by the load
1) Case Hn' . 2) Case Hz
f
bending.
For
simplification, it is possible, with a
sufficient degree of accuracy, to cal-
culate the fuselage in accordance with
its parts, dividing it into tail, nose,
and center section (Fig.189).
Such a division of the fuselage is
permissible since many computable cases
can be calculated only for individual
parts of the fuselage. For example,
the case Cf - nose dive - is related
only to the tail section. In the given
instance, the internal load of the
fuselage will be a load exerted only by
the horizontal tail surfaces. Conse-
quently, the tail section of the fusel-
age will work as a cantilever beam in
The case Hz (see Fig.188a) also refers to the tail section of the fuselage
which, in this instance, works in bending and torsion as a cantilever beam.
n
The case Hf is concerned only with the nose section of the fuselage, which in
this instance works in bending as a cantilever beam.
The cases AI" Af' and Ef refer to the entire fuselage.
The fuselage as a whole must be considered as a double-support beam. The
attachment bolts of the wing to the fuselage represent the supports.
Under symmetric load, the fuselage works in bending, and under asymmetric
in bending and torsion.
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Ph.t
mzst
2
17dive
2 (91)
We may establish the reference curves Q and Mb as for the usual double-support
cantilever beam.
Case Hz . Here the fuselage is calculated according to the load of the verti-
cal tail surfaces B.O.P (Fig.191). The load is often ascertained by calculation.
STAT
1;1..
1
alien-itata iir'cliteraining the
freehand curves of the lateral force
Fig.189 - Subdivisions of the Fuselage
For convenience in calculating the
;fuselage, it is subdivided into three
,parts: nose, center, and tail sections.
a) Nose section; b) Center section;
20 .
c) Tail section
liOraialer ind the tangent T of stress are- ---1
the bending moment Mb, and the torsion
moment M. With the help of the curve,
we may determine the stresses of a sec-
tion of the fuselage and the magnitude
of Q, Mb, and Mt of any section.
In this way, for the calculation
it is necessary not only to determine
the external loads but also to estab-
lish a reference curve of the lateral
forces as well as of the bending and
torsion moments for all computable
cases and to determine which sections may
cases, in order to define these computable
.be most endangered. If we are interested in the stability of only one section of '
the fuselage, there is no need for a reference curve and it is sufficient to compute
Q, Mi,, and Mt only for that section.
As an example, we will discuss the character of the reference curve for several
calculable cases.
Case Cf. In this case, it is necessary to calculate the fuselage according to
the load of the horizontal tail surfaces (Fig.190) which begins with a dive of the
aircraft [see eq.(83)]:
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J1111..1211?1_ JJ
Fig.190 - Case Cr Airplane in a Dive
a) Attachment of center panel to
fuselage; b) Reference curve Mb;
c) Reference curve Q
In carrying out the dive, the airplane
is calculated, in relation to the tail
section of the fuselage, from the load
of the horizontal tail surfaces Ph .t.
The tail section of the fuselage is a
beam, working on the fixtures of the
center panel to the fuselage.
Cd fl
Fig.191 - Work of the Fuselage in the
Case Hf
a) Reference curve Mb; b) Reference
curve Q; c) Axis of fuselage;
d) Reference curve Mt
In this case, the tail section of the
fuselage will work in bending and tor-
sion. The magnitude of the bending
moment for the fuselage section is
determined by generation of the force
Pv.t.
over the observed section to the
force Pvit. along the axis of the
fuselage. The magnitude of the torsion
moment is produced by the force along
the longitudinal axis of the airplane.
Fig.192 - Work of the fuselage in the Case HI;
a) Propeller; b) Tank; c) Engine; d) Reference curve Q; e) Reference curve Mb
Here it is necessary to calculate only the nose section of the fuselage in bending,
from the concentrated stress which are equally produced by the weight of each unit
(Anaine. nropeller. etc.) to the load factor n.
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determined according to eq.(85):
\
0.37
ma.
2
where V is the maximum speed of the airplane;
Sv.t
Sv.t. is the area of the vertical tail surfaces.
(92)
In the case Hz ' the fuselage works in bending and torsion. The torsion is pro-
duced by the activation of the upper axes of the fuselage.
Case H. As noted above, the case H; is related to the nose section of the
fuselage (Fig.192). Here, the nose section of the fuselage is calculated in bending
_ as a cantilever beam from the line of the concentrated inertia force Gagrn.
Section 35. Fuselage Design
From a consideration of the action on the fuselage by loads we see that the
fuselage works in bending and torsion. The design of the fuselage must be consider-
ed in relation to these deformations.
Depending on the type of power system, fuselages are of the truss type or gir-
der type. As to the material used in construction, mostly metal and less often wood
is used.
Truss-type fuselages, in turn, are divided into: rigid types, consisting of
truss elements capable of either expansion or contraction; and braced types, in
which the rigid diagonal rods are substituted by a pair of bracing struts capable of
handling only expanding loads. Besides this, there are truss-type fuselages in
which the nose and center sections consist of a rigid rod framework and the tail
section of a braced framework.
Girder-type fuselages have longitudinal (longerons, stringers) and transverse
framing (transverse trusses or bulkheads). The entire body is covered by a skin.
The Truss-Type Fuselage
The trt.iss-type fuselage (Fig.193) consists of a system of rads differing STAT
29
7 tq
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?
?
usually from the ideal spatial truss in that the connections at many angles are not
hinged but fixed.
Fig.193 - Design of the Truss-Type Fuselage
The truss-type fuselage consists of a three-dimensional system
made up of rigidly interconnected rods.
1- Longerons; 2- Strut; 3- Diagonal Bruce
The spatial truss of the fuselage consists of four plane trusses, two horizon-
tal and two vertical. Each two-dimensional truss consists of two booms joined to
each other by struts and diagonal braces*.
In addition to the four flat trusses mentioned above, within the fuselage there
are cross-wise connections. In Fig.193 there is shown a section of the fuselage
whose elements consist mostly of fixed rods. The individual transverse connections
may also be trusses or frames. In the case of truss connection, instead of one
rigid diagonal, two boom braces or wires are used (Fig.194) which always are
pre-stretched. At times of stress on the truss cell, one brace absorbs the strain
while the other is relaxed; during maximum stress, this yields completely so that,
according to calculation, it always takes only one of the two transverse booms,
i.e., the one which works on the given load of expansion.
The strength of the brace construction depends greatly upon the degree of
*Ribbon types are the common type of the horizontal and vertical trusses.
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,preliminary stretching of the ribbon bracing. The method of calculation and practi-
cal recommendations for the brace construction
b)
Fig.194 - Bracing Section of the
Truss-Type Fuselage
In the installation of a ribbon truss
instead of one diagonal two (more)
trusses are installed. In this instance,
the exterior force P stresses the truss
(1) for expansion, truss (2) for buck-
ling, and the upper boom for contraction.
a) Boom; b) Boom force; c) External
force; d) Bracing-wire force; e) Bracing
wire
were given by N.Ye.Zhukovskiy and his
student, V.P.Vetchinkin.
/3
The elements of the truss-type
fuselage are usually tubes. Some-
/
times they have open profiles. The
materials are steel and aluminum.
Often the Joining of the tubular
elements is done by welding and the
Joining of the open profile elements
by riveting.
The welded articulations are
accomplished in the following
manner. The tubes are welded to the
ribbon and welded along the engaging
flanges (Fig.195a). For greater
strength, the abutting Joints are
often reinforced by open corner
plates, with subsequent welding
(Fig.195b). If the metal truss has braces, the welded corner plates can serve as
mounting lugs (Fig.195c). In brace construction, the corner plates are sometimes
replaced by welded tubes (Fig.195d).
Assemblies of riveted construction are made with the help of corner plates.
Very often it is impossible to install diagonals or braces in the separate
cells, for example, in the cockpit. The lack of diagonals here may be compensated
by rigid Joints, 1.e., by a framework (Fig.196).
In order to achieve a streamlined form of the truss-type fuselage, it is fitted
with a special light-weight body which may be covered with fabric, laminated wood,
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or aluminum. Part of this body or individual pieces, for convenience of utiliza-
tion, may be of mixed construction.
Fig.195 - Design of the Welded Nodes of the Fuselage
The joining of the elements (rods) is accomplished by welding.
For increased strength, corner plates are often welded on.
a) Spacer; b) Vertical strut; c) Cross strut; d) Ribbon;
e) Corner plate; f) Bracing wire
Girder-Type Fuselages
Constructionwise, girder-type fuselages are more advantageous than truss types.
Possessing a better aerodynamic form, they have smaller outside dimensions and per,-
mit a more ready displacement of the useful load. In a military sense, girder-type
fuselages are preferable over truss-type fuselages. On receiving a hit in the
longerons or diagonals, the truss-type fuselage may become unfit for combat. Con-
versely, in a girder-type fuselage (monocoque) even numerous hits are not very
dangerous.
The girder-type fuselages (having ihe aspect of a beam) usually are of metal
construction. They may be subdivided into three types:
Longeron girder fuselage;
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Stringer girder fuselage;
Nacelle type fuselage.
Fig.196 - Cut-Out in the Truss-Type
Fuselage
In the production of cut-outs (cockpit,
doors, portholes, etc.) no diagonals are
installed in the cell. The lack of dia-
gonals here is compensated by rigid
Joints, i.e., by a frame.
a) Spacer strut; b) Boom longeron;
c) Rigid framework
often thicker than in the longeron fuselage.
The longeron fusela&e (Fig.197)
has longitudinal frames, consisting
of four strong longerons and a series
of light stringers. The elements of
the longitudinal frame are intercon-
nected by the transverse frame or
bulkheads. The power plant is cover-
ed by aluminum plates.
The stringer fuselage (Fig.198)
has longitudinal elements consisting
of a series of high-strength string-
ers, which are connected to each
other by frames. There are no long-
erons. The covering, aluminum is
The coque-type fuselage (Fig.199) consists of a relatively thick covering
attached to the inside frame.. '1%z:re are no stringers or longerons. Fuselages of
such type are found only in a few aircraft prototypes. In larger machines, because
of the possibility of loss of stability (buckling), the skin must be thick, which
exerts excessive stress on the structure.
At the present time, the basic type of fuselage is the stringer type, where all
elements of the longitudinal frame and covering can absorb bending moments. But if
the fuselage has many cut-outs (Fig.200), then to compensate the weakening of the
structure, caused by these cut-outs, the longitudinal frame is fitted with longer-
ons. In these instances, longeron construction of the fuselage is mandatory.
33
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,e?
Fig.197 - Design of a Longeron Fuselage
The longeron fuselage has four strong longitudinal elements known as longerons.
a) Longerons; b) Transverse trusses; c) Stringers
Structural Elements of the Girder-Type Fuselage
The metal stringers (Fig.201) are pressed aluminum profiles in the form of
angles (a), angles with bulb (b), Z-sections (c) or U-sections (d).
Longerons, in cross section, have the form of profiles (Fig.202) or tubes.
Longerons may be riveted to the covering (Fig.203).
The bulkheads (Fig.204) are made either of an entire profile, more often a
channel section, or from profiles and sheets riveted to each other. Reinforced
STAT
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bulkheads, as a rule, are riveted into a box section.
The joining of the longitudinal frame to the transverse frame (Fig.205) is done
Fig.198 - Design of a Stringer Fuselage
In the stringer fuselage all the longi-
tudinal elements (stringers) are roughly
of one section.
a) Light transverse truss; b) Stringer;
c) Heavy transverse truss
Fig.199 - Design of a Shell-Type or
Coque Fuselage
The shell-type fuselage has no longi-
tudinal frame; the body consists only
of bulkheads.
a) Bulkhead
c)
Fig.200 - Cut-Outs in the Fuselage
Cut-outs in the fuselage weaken its mechanical strength.
This weakening is compensated by longerons.
a) Gunner station; b) Entrance hatch; c) Machine-gun mount
with rivets. Usually, the bulkheads are provided with inside holes for passing the
stringers or longerons.
The joining of the stringers with the bulkheads is done through holes in STAT
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b)
c)
d)
Fig.201 ? Design of Stringers
Stringers consist of pressed aluminum profiles.
Fig.202 ? Design of Longerons
Longerons are usually profiles which, together
with the adjoining fuselage skin, form the
sections of the completed contour.
a) akin; b) Longeron
b)
Fig.203 ? Riveted Longeron
The longeron may be riveted
from several profiles, forming
the completed contour.
a) Longeron; b) Skin
Fig.204 ? Design of Bulkheads
The bulkheads are prepared from pressed profiles
or riveted from several bent profiles.
a) Fuselage skin
36
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.bulkhead itself or through a special box or angle plate.
general, the stringers are not joined to the
Fig.205 - Joining of Stringers and
Bulkheads
The bulkheads, for joining with the
stringers, are provided with holes. The
stringers are riveted to the bulkheads
with the help of reverse bulkheads or
special angle plates.
a) Stringer; b) Fold; c) Bulkhead;
d) Angle plate
on the butt ends of the longerons (Fig.206a);
age are assembled along the entire contour by
(Fig.206b) or fittings (Fig.206c).
In many designs, in
bulkheads.
The covering of the fuselage with
a sheet metal is most often done with
aluminum sheets. Every sheet is shaped
in accordance with the form of the
fuselage. The butt joint of the sheets
is lashed to the stringers and bulk-
heads. The possibility of using wooden
planking on a metal frame is not ex-
cluded.
The assembly of the individual
fuselage parts is carried out in vari-
ous ways: Longeron-type fuselages are
assembled by means of joints, laid out
the parts of the stringer-type fusel-
screws or bolts, by means of tapes
Fig.206 - Assembly Plan of Individual Parts of the Fuselage
In longeron design, the butt joints are prepared from four joints mounted
to the longerons. In the stringer design, the butt joining is done along
the entire contour of the fuselage.
1) Butt joint; 2) Longeron; 3) Screws; 4) Bulkhead; 5) Plate; 6) Stringer;
"1 vlfting; 8) Fabric strip; 9) Connection under tape STAT
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Section 36. Design of a Semimonocoque Fuselage
In its strength characteristics a semimonocoque fuselage is very much like a
monospar wing. Its energy diagram consists of longitudinal stiffeners, transverse
frames, and a stressed skin. External loads deform the fuselage in the same manner
as they deform the wing.
in a fuselage, concentrated forces are applied to the bulkheads. Concentrated
loads are transmitted through the bulkheads to the skin in the form of distributed
forces. These forces flow down the skin toward the wings and cause bending and
twisting of the fuselage. Thus in separate cross secLirms_ of the fuselage lateral
forces Q, bending moments Mb, and torsion moments Mt are present.
The bending moments Mb are absorbed by the longitudinal stiffeners and the
skin. Lateral forces and torsional moments are absorbed by the skin. As in the
energy diagram of a fuselage, i.e., the character of action of external forces, the
flow of stresses through structural elements, and the work of the fuselage elements
are similar in principle to the energy diagram of a wing, whereas the order of cal-
culations - the order of determining the loads in the fuselage elements - is similar
to that of the wing. Data necessary for these calculations consist of the diagrams
of lateral forces, bending and
fuselage elements.
Determination of Normal Loads
torsional moments (see Section 34), and the size of
In a longeron-type fuselage (Fig.207) the bending moment Mb is absorbed by
longerons in the form of compression and tension and to a lesser degree by stringers
and the skin. This is explained by the thinness of the skin, only slightly rein-
forced by stringers, and by the presence of cut-outs.
In order to simplify these calculations let us assume that both skin and
stringers do not participate in the work. In such a case, the total bending moment
Mb of the fuselage will be absorbed by longerons (Fig.207b). Axial stresses will
STAT
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also form in the booms
where 111 is the distance between longerons.
b
II
1
0
n
(93)
a) b)
Fig.207 - Action of the Bending Moment in a Longeron-Type Fuselage
The bending moment (a) is absorbed mainly by longerons in
the form of contraction and expansion (b).
1) Bending moment; 2) Longeron area; 3) Tensile stress below longeron
When the magnitude of the axial stress S is known, it is possible to find the
magnitude of the normal stress in longerons:
For the top boom
where Fv is the cross-sectional area of an upper longeron;
For the bottom boom
where Fn is the cross-sectional area of a lower longeron.
In a stringer-type fuselage (Fig.208a) the bending moment is absorbed by ISIWT
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stringers and skin. The mechanical strength of such a fuselage is limited by the
strength of the stringers.
a)
6)
"11
e)
Fig.208 - Action of the Bending Moment in a Stringer-Type Fuselage
Here the bending moment acts on both stringers and skin,
causing expansion and compressions in them.
1) Ssk = stress in the skin; 2) Sstr = stress in the stringer; 3) Total com-
pressive force; 4) S = total tensile force; 5) Arm pair
Forces which act separately on the skin and stringers in zones of tension and
compression (Fig.208b), form two resultant forces S (Fig.208c) which, in turn, con-
stitute a couple. Therefore the magnitude of S is determined by the formula
-11112?
"mar:
*
where 'mean is the lever arm of the couple.
-
In approximate calculations, it may be assumed that
9
H H
entin
in such a case
s.. 41_13 .
2
//
3
In order to determine the stresses, we start with the well-known law of stress
distribution through the depth of a beam (Fig.209), according to which the greatest
stresses will be found near the fibers at the extremeties while in the neutral STAT
(94)
(95)
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surface they diminish to zero. Thus, for our calculations we take the
cross-sectional area of the elements located in the zone of greatest compression and
tension, i.e., in the top and bottom arches.
Fig.209 - Stress Distribution Through the Depth of the Fuselage
High stresses occur in the arches of the fuselage and lower
stresses in the elements near to its neutral layer due to
the action of the bending moment.
a) Upper arch; b) Neutral layer; c) Lower arch; d)' = stress in the
co
upper arch; e)0 exp = stress in the lower arch
The common cross-sectional area of elements in one arch is equal to
Fatch= tifstr ? Fsk (96)
where n is the number of stringers in one arch;
fstr is the cross-sectional area of one stringer;
Fsk is the area of skin of one arch;
is the so-called reduction coefficient which shows what section of the skin
is working to its full capacity. In the zone of expansion, p = I (i.e.,
the entire skin is working to its full capaciW, while in the zone of
compression
308
cp 1.
b
Here b is the distance between stringers;
o is the thickness of the skin.
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Thus the normal stresses in stringers are equal to
a = =
str FAtib 2
3 - (it for Fsk (f)
(97)
In a true monocoque fuselage (without stringers), the bending moment Mb is ab-
sorbed by the skin alone. For full load performance it is necessary to secure an
adequate stability of the skin (to prevent buckling). All considerations made in
determining the strains and stresses in a stringer-type semimonocoque fuselage are
Fig.210 - Lateral Bending of the Fuselage
In case of lateral fuselage loading, strains and stresses are
calculated in the same manner as in symmetrical bending.
a) Mb = bending moment of cross section; b) P .t. = lateral load;
v
c) Axial stress on longeron (boom).
fully applicable to a true monocoque fuselage. Stresses in the skin caused by the
action of Mb are determined by eq.(97) in the following manner:
Al b
str " ?
1/Fsk
(98)
As a result it can be stated that fuselages of all three types work in a STAT
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similar manner and differ only in the degree to which different elements participate
in the absorption of Mb.
In a longeron-type semimonocoque fuselage moment Mb is absorbed by longerons
and the loss of stability (buckling) of stringers and the skin does not affect the
mechanical strength of the fuselage. The fuselage will fail only if longerons are
destroyed.
In a stringer-type semimonocoque fuselage, most of the moment Mb is absorbed by
stringers; therefore, their strength determines the strength of the fuselage and the
loss of stability by the skin is not dangerous.
In case of a lateral asymmetric loading of the fuselage (Fig.210), the line of
reasoning in determining strains and normal stresses is the same as before, with the
exception that, in all formulas, the quantity B (cross section) must be used instead
of H.
Determining Tangential Stresses
Tangential stresses T in the skin of the fuselage occur under the action of the
lateral forces Q and the torsional moments Mt.
Torque takes place in the case of asymmetrical loading of the fuselage. If the
fuselage is loaded symmetrically, Mt is absent.
First let us examine the process of determining tangential stresses caused by
symmetrical lateral forces Q (Fig.211). The force Q will be absorbed by the side-
walls of the fuselage, causing tangential bending stresses lb in them. Tangential
stresses in the skin of the top and bottom arches of the cross section are negli-
gible and can be disregarded.
In order to determine the tangential stresses produced by Q in the skin of the
fuselage and assuming that these stresses are evenly distributed, let us find their
magnitude by dividing the force by the sectional area of the skin portion on which
Q is acting. In a longeron-type fuselage (Fig.211a), the force Q is mainly absorbed
by the sidewalls, whose area is approximately equal to 2H1. STAT
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Ora 2.)
Fig.211 - Action of a Lateral Symmetric Force
The force Q acts on the sidewalls of the fuselage and produces tangential
bending stresses h? The force Q when applied to the tail unit, together
with the resultant of the internal forces R forms a couple - the bending
moment.
1) Q = resultant external force; 2) R = Q = resultant internal force lb
Therefore,
(100)
Now let us define how the tangential stresses are determined when a fuselage is
loaded asymmetrically. In this case, in addition to tangential bending stresses
(due to the action of the force Q), other stresses will be produced by the torsional
STAT
Therefore,
(,)
2//17,
(99)
where 6 is the skin thickness.
In a stringer-type and in a true monocoque fuselage (Fig.211b and c), the force
2
Q is absorbed mainly by the walls - H high. The area of these walls will be equal
3
2 4
to 2 -5 H8 -5 H.
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moment Mt. For example, let us take the case of asymmetric fuselage loading with a
? load acting in the direction away from the vertical tail surfaces (Fig.212), the
force P, The action of the couple Pv.t. and Q in the horizontal plane (action of
^ Mb) has been examined above. Here we are going to examine the action of the force
P
v.t. = Q which, due to the lever arm a, produces a torque
Mt Q ? a,
where the lever arm a is the vertical distance from the line of action of the force
to the middle of the cross section of the fuselage.
Pv.t.
Fig.212 ? Loading of a Fuselage, with the Load Acting Away From the
Vertical Tail Surfaces, with a Force Pv.t.
The force Pv.t? applied to a point A, together with the force Q
applied to the point 111 form a bending moment of the fuselage. The
force Py.t, applied to B together with the resultant of the internal
forces Rtb produce the torsional moment Mb Pv.L.a. The force
Pv.t.
is balanced by the stresses Ii), while Mb is balanced by the
stresses IL'
a) Arm of the bending moment; b) Arm of the torsional moment;
c) RI = Q = resultant force
lb
The lateral force Q in any cross section of the fuselage is absorbed by the
skin of the top and bottom arches and produces tangential stresses lb in it. The
STAT
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?
magnitude of lb in a longeron-type semimonocoque fuselage is determined by eq.(99)
where DI is the distance between longerons.
The tangential stresses lb produce a resultant force R1 = 0 which, together
with the force Pv.t., forms a couple which produces twisting of the fuselage
or
Mt = Qa,
Mt = P v.ta.
The determined torque will act on the entire contour of the cross section of
the fuselage and produce torsional stresses there. The tangential torsional stress-
es are determined by the formula
or
Mt
t 2 F(f
a
2F6
(102)
(103)
where F is the cross-sectional area limited by the contour of the fuselage.
All these formulas, with a single modification are also valid for the
2
stringer-type and true monocoque fuselages; only Di must be replaced by 7B, where 13
is the width of fuselage.
In order to determine the resultants of tangential stresses in the skin, an
algebraic addition of tangential stresses of bending b and of torsion 1 t must be
performed.
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r r r
b t'
Section 37. Desigp of a Truss-Type Fuselage
In its strength characteristics, a truss-type fuselage is similar to a semi-
monocoque fuselage, except that in this case the main strength element is not the
skin but the girders and diagonal braces. Therefore, the determination of external
forces and the process of plotting the loading diagram of Q, Mb, and Mt as the
initial data in our calculations, remains the same as a semimonocoque fuselage.
When the loading diagram and Lhe dimensions of the fuselage are determined, it
is possible to find the stress and strain in its elements.
The bending moment of the fuselage (Fig.213) will be fully absorbed by compres-
sion and tension of its longerons. The magnitude of the axial stresses in the
longerons (booms) is determined from the formula
Mb
Sboom =
2H
(104)
The lateral forces Q (Fig.214) produce axial stresses in the girders and dia-
gonal braces of the lateral plane trusses. The lateral force Q is evenly distribut-
ed between the two plane trusses:
If the angle a between the vertical girders and diagonal braces is known, we
can determine the axial stresses in the latter from the following formula:
Sb race
2 cos a.
(105)
The stress in a vertical girder according to the conditions of equilibrium of
the assembly is equal to the shearing force of a plane truss. In this case,
(
,)
S 2qirder ?
47
(106)
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A vertical loading of the tail unit will cause bending of the fuselage in the
horizontal plane and twisting.
Since the stresses produced by Q and Mt in the upper truss are cumulative, this
truss is usually reinforced by double
tie-rods or stronger diagonals.
Fig.213 - Action of the Bending Moment
The bending moment of the fuselage will
be fully absorbed by compression and
tension of the longerons.
a) Sboom = tension in the lower boom
Fig.214 - Action of the Lateral Force
?
The stress in the vertical girder is
equal to the lateral force of the plane
truss The stress in the diagonal
braces is equal to the lateral force of
the plane truss divided by cos x, i.e.,
Sbrace - 2 cos
a) Q = transverse force on fuselage;
b) Sbrace = stress in brdce; c) Trans-
verse force in plane truss
In order to increase the rigidity of the fuselage and to include all four plane
trusses in the resistance work to Mt, use is made of internal cross braces, diagonal
braces, or frames are employed inside the fuselage. These lateral elements perform
the functions of wing ribs, i.e., they distribute local loads over the entire
longitudinal assembly of the fuselage. More elaborate calculations demonstrate that
the load on these lateral elements is negligible.
Section 38. Pressurized Cabins
Designation of Pressurized Cabins
f jet aircraft, not only did the speed of flight but also AT
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altitude increase considerably. Thus, the problem of creating the necessary vital
conditions for the crew and maintaining its combat efficiency at high altitudes ac-
quires now more significance than ever before.
Atmospheric pressure decreases with altitude and so does the partial oxygen
pressure (the percentage of which, as that of other gases in the air, remains prac-
tically constant up to an altitude of approximately 2000 m). This leads to a short-
age of oxygen or anoxia; for flights at altitudes above 4500 in man needs an addi-
tional supply of oxygen. A harmful effect on the human body is also produced by the
low pressure of the surrounding atmosphere; pains in the joints, stomach, teeth,
etc., and attacks of altitude sickness (aeroemholism). These symptoms appear during
long flights at altitudes above 8000 m.
Vital conditions for the crew and passengers at high altitudes are created by
oxygen equipment, pressurized suits, and pressurized cabins.
Oxygen equipment is ordinarily used at altitudes above 4500 m. It adds some
oxygen to the inhaled air and just prevents symptoms of anoxia, but only below a
certain altitude. Prolonged flights at altitudes above 10,000 m, even with oxygen
equipment, are very difficult both because of the low atmospheric pressure and be-
cause of an inadequate partial oxygen pressure produced by the oxygen equipment.
Therefore, even though individual flights in open cockpits with oxygen equip-
ment at altitudes of the order of 5,000 m are known, i t must be considered that the
highest altitude for mass flights in open cockpits is 12,000 in and that, as a rule,
above 9000 - 10,000 m pressurized suits or cabins must be used which make it possi-
ble to increase the pressure on the body and create a higher partial oxygen
pressure.
Pressurized suits are complicated and hinder the movements of the pilot and
therefore did not come into wide use (they are used only for record flights).
Due to the_fact that a pressurized cabin is completely or partially insulated
from the surrounding atmosphere, it is possible to raise its oxygen content and
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1
increase the pressure (with respect to the surrounding atmosphere), thus creating
the necessary conditions for sustaining life and efficiency of the crew and passen-
gers.
The problem of temperature control may be solved by insulating pressurized
cabins and by installing heating equipment. This problem acquires special impor-
1
Lance in connection with high-altitude flights in regions of low temperature and in
connection with airplane heating at high speeds.
Flying in pressurized cabins at high altitudes is less tiring than
high-al ti tude fl ights w i Lh oxygen equi pment .
Pressurized cabins are the surest means toward securing the efficiency of the
crew of a modern airplane at high altitudes under combat conditions.
The idea of building pressurized cabins was first forwarded and proived in 1875
by the great Russian scientist D.I .Mendeleyev.
Types of Pressurized Cabins and How They Operate
In accordance with the means by which vital conditions are produced in pressur,-
ized cabins, these are divided into two main types: 1) pressurized cabins,
2) sealed cabins.
in pressurized cabins which are now the most widely used type, a blower con-
stantly blows air under pressure into the cabin, thus producing a certain super-
charging; at the same time, through certain valves which keep up the necessary pres-
sure in the cabin, the air is discharged into the atmosphere. As a result, a con-
tinuous ventilation of the atmospheric air throughout the cabin is achieved. Thus,
supercharging and ventilation are the characteristics of a pressurized cabin.
Pressurized cabins are sometimes called supercharged cabins. The percentage of oxy-
gen in the air blown into the cabin is the same as that in the outside atmosphere,
but the partial oxygen pressure is proportional to the absolute pressure in the
cal)] n.
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Let us now examine how a cabin is supercharged. It may seem that, in order to
create the most favorable conditions, it is necessary to keep up a normal atmos-
pheric pressure in the cabin at all altitudes. However, in case of a blowout at
high altitudes, due to puncturing of the cabin or to some other damage, the pressure
drop will be extremely sharp. At an altitude of 16,000 m in such a case the pres-
sure would diminish 10 times within a fraction of a second, which would lead to
serious physiological consequences for the crew members. Also, in order to sustain
a higher positive pressure in a cabin, it is necessary to reinforce its structure
considerably which, in turn, leads to a substantial gain in weight.
Therefore, the pressure produced in the cabin is such that it secures comfort-
able living conditions and that, at the same time, even a sharp pressure drop in
case of a blowout will be endured by the crew without injury.
Up to 2000 - 3000 m, the pressure of the surrounding atmosphere is ordinarily
maintained in a pressurized cabin (segment 1 of the curve in Fig.215). in a
pressurized cabin, this is accomplished by a suitable valve. The absolute pressure
thus attained is maintained in the cabin up to a certain altitude where the pressure
is raised so that it will be higher than that of the surrounding atmosphere (segment
2 of the curve). This positive pressure is usually equal to 0.25 - 0.30 atm and is
kept constant during the rest of the climb (segment 3 of the curve). Thus the cabin
altitude is lower than the airplane altitude. According to Fig.215, at an altitude
of let us say 15,000 m, conditions in the cabin will represent those at an altitude
of 7000 m.
However, even though the absolute pressure in the cabin will be adequate for
the crew, there will be a shortage of oxygen. Therefore, it is necessary to use
oxygen equipment in pressurized cabins above a certain altitude. Such oxygen equip-
ment has to be switched on in open cockpits at an altitude of 4000 m while, in
pressurized cabins, oxygen equipment must be put into use only at much greater al Li-
, at 9000 m).
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Ventilation of a pressurized cabin consists of constant circulation of fresh
air through the cabin and is measured by the volume of air flowing through the cabin
in unit time. Ventilation is used in order to maintain a certain concentration of a
given substance in the cabin air (oxygen, water vapor, carbon dioxide, etc.).
p utni
10.
cui
NL,
N. I
IN a
1
- 1 .
,
? 4
,
I 1
6 8 14 16 - Ili H Ma
Fig.215 - Character of Change in Atmospheric Pressure
With Respect to Altitude and Corresponding
Change in Pressure in a Pressurized Cabin
The cabin altitude is generally lower than the airplane altitude.
a) Pressure in cabin; b) Standard atmosphere
The general diagram of a pressurized cabin (Fig.21(a) includes l) feed system;
2) air release and pressure control equipment.
The feed system consists of a blower (engine compressor or a special cabin
supercharger) from which the air is blown through a check valve (4) into a manual
supply cock (2), with which the pilot can manually control the delivery of air from
the blower or stop it completely and connect the cabin directly with the Outside
atmosphere by opening a special scoop (depressurize the cabin and switch on venti-
lation with atmospheric air). From the cock (2), the air flows into the collector
(3) from where, through a number of small openings, it enters the cabin, blowing STAT
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over the canopy and the windshield. Air is heated by compression in the supercharg-
er and thus heats the cabin (completely or partially) and prevents fogging of the
pilot's enclosure. An an air-intake volume control (I) may be added to the
feel system, in order to control irregular supercharginp of the compressorduring
fluctuation in engine operation. A filter (5), containing active charcoal, silica
gel, cotton, and the like, may be used to clean the air or oil and decomposition
products (which are present when the cabin is supercharged by the engine compres-
sor).
The air release and pressure control equipment consists of pressure regulators
desipned for maintaining the given pressure in the cabin and of safety devices pro-
tecting the cabin from extreme internal and external pressures. The regulator or
the valve of permanent absolute pressure (6) remains open until a certain altitude
is reached, and the cabin freely comunicaLes wiLh the surrounding atmosphere (see
segment I of the curve in 1"ig.215). The valve (6), according to its adjustment,
maintains a permanent absolute pressure (see segment 2 of the curve in 1'ig.215);
when the required positive pressure in the cabin is reachol, the regulator for pen-
manent positive pressure (9), which functions on the prInciple of premoire varia-
tion, comes into operation and secures the segment i of the curve (see Fig.215).
Sometimes the valves (() and (9) are comhined into a sinple aggregate, a pres-
sure regulator (Pli). To protect. the cahin filnu excessive internal pressures in
case of failure of the pressure regulator, a connecting safety valve WO is insert--
el. This valve is adjusted to a somewhat greater pressure drop than that of the
valve (9). The pressure-variation check valve (8) allows atmospheric air to enter
the cahin when a vacuum starts forming there (during rapid descents) and thus pro-
tects the cabin from collapse due to externnl pressure. The pressure-variation
check valve (8) is often used as a manual control valve for the cabin pressure.
? Manual control is accomplished by changing the design of the valve accordingly and
)rovidin it with a hand wheel (7). In caae of necessity,
5i
iL is possihle to bleesTAT
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the cabin air into the atmosphere.
Pressurizol cabins may al so be suppl ed wi th heating equ i pment (if the tempera-
ture of the supercharged air is not ii ? c lently high), a i r-cool i ng equipment, a i r
fill. ems, humid I Ly control equ i pment, and general runt rol equ ij iment (control of a i r
I ntak e, a i r pressure, tempera tu re, humid i ty, purl ty, etc.).
Fig.216 - Principal Diagram of Cab i n Ventilation
The swi Leh ing diagram i tic lut I en the a i eat system i tse f ( intake,
supetvharging) and the al r-d t,charge system and pressu IP control .
I e i tig pressurized cabins, the technician must careful ly check the good
work ing cond i L ion of al I al Li Wile equ i !anent mid of the devices that. keep the cabin
3cal si e (Mei r i lure at Iiigh al L i Ludes may I ead t o crashes or even catas-
trophe.
In a seal ed cabin, air is pumped through an a i r purl Fi('ation plant where the
rospi rat, i Utl p rodue Ls (water vapor, carbon d inx ide) are absorbed . i muI taneous I y,
oxygen from tanks and air are con t I iniously added to the cabin atmosphere (as a com-
pensaL ion for the oxygen used 11 breath ng). Thus a certain posi Ii ve pressure i
created in the cab iii and i s ma in La i ned by va I ves which regu late the d I scharge of
air from the cabin.
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Thus, in this type of cabin not only a positive pressure (supercharging) but
also a constant ventilation of the cabin, regulating the concentration of the vari-
ous air components, is achieved.
The high percentage of oxygen in sealed cabins makes it possible to produce the
necessary partial oxygen pressure at lower cabin pressures (in comparison with a
pressurized cabin) but at the same time fire hazards are increased. One of the ad-
vantages of sealed cabins is that their functioning does not depend on the condi-
tions of the outside atmosphere. This fact makes it possible to use sealed cabins
at such altitudes where pressurized cabins can no longer be used. The duration of
flight in a sealed cabin is limited by the amount of oxygen in the tanks. Cabins of
this type are also more complex than pressurized cabins.
Energy Diagrams and Loads Acting on Pressurized Cabins
The loading of a pressurized cabin and its functions as a unit of power largely
depends on the degree to which it is included in the energy diagram of the fuselage.
A pressurized cabin may be built as a separate unit of power, which is install-
.
ed (suspended) inside the fuselage without being included into its energy diagram
(Fig.217a). Such a cabin is sometimes called a suspensed cabin. It is loaded with
the difference of internal and external pressures, and also with loads (forces of
inertia and the mass) of men and cargo distributed inside the cabin. Since the
loads of the fuselage are not transmitted to the cabin, its deformations due to such
loads is less which, in turn, makes the sealing of the cabin more reliable.
The shape of a suspended cabin when the fuselage is large does not depend on
its contours so that the cabin can be given a shape that is more advantageous in
relation to strength. It can then be of minimum size, which will be determined
only by the accommodation of the crew and the requirements of its work. This per-
mits a decrease in the cabin weight, but its structural material, not included in
the energy diagram of the fuselage, is not adequately exploited.
Cabins forming airtight sections of the fuselage (Fig.217b) are also widely
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used. These are completely included into the energy diagram of the fuselage. Such
cabins represent sections of the fuselage and accordingly, in addition to the loads
mentioned above, they also absorb aero-
dynamic loads distributed over the en-
tire surface of the cabin and the loads
of other parts of the al rcraft connect-
ed with the pressurized cabin (such as
wings, tail assembly, fuselage, etc.).
Therefore loading of a pressurized
cabin with a pressure difference is in-
cluded in the energy diagram of both
cabins. Under normal conditions, the
cabin is loaded with internal positive
pressures. However, if the aircraft
would descend rapidly, the cabin would
be subject to an external pressure.
Since a pressurized cabin, due to its form and structure, is more capable of with-
standing internal pressures, loading the cabin with excessive external pressures is
more dangerous. Such a condition may arise if, during a rapid descent, the valve
(9) (see Fig.216) fails to equalize the pressure inside the cabin with the pressure
of the surrounding atmosphere.
Canopies, due to the curvature of their surface and due to considerable speeds,
are loaded with large aerodynamic forces. These loads are usually directed toward
the outer side of the enclosure.
Fi p.217 - Diagram Showing Inclusion of
Pressurized Cabins into the Structure
of the Fuselage
Suspensed cabin (a) increases the reli-
ability of its sealing, since it does
not participate in the work of the
energy diagram of the fuselage; pres-
surized section of a fuselage (b) con-
stitutes a part of the fuselage.
Design of Pressurized Cabins
The design of a pressurized cabin must meet the requirements of strength,
rigidity, pressurization, heat insulation, survival, and must be free from fire
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hazards, etc.
A pressurized cabin, when loaded, acts first of all as a thin-walled vessel un-
der the action of internal or external pressures and secondly (not always) as a
thin-walled vessel subjected to bending and twisting in the energy diagram of the
fuselage.
The strength of a pressurized cabin is secured by making its construction
similar to that of the fuselage and by selecting a suitable assortment of profiles
for the power units.
The work of a pressurized cabin as a vessel has its own peculiarities, which
influence the choice of a rational form, its structure, and strength. As it is
well known, the most advantageous form for a closed vessel loaded with internal
pressures, in relation to its strength, is a sphere. But a spherical cabin cannot
be used in aircraft since it would be difficult to accommodate the crew in a cabin
of such a form, and since it does not correspond to the outline of the fuselage.
Consequently, there is a general trend to design pressurized cabins in the form of
cylinders with circular cross sections, closed on the ends with spherical or conical
end pieces. The transition from the walls of the cylinder to the end pieces must be
smooth, without sharp edges. If any sharp edges are present, the end pieces will
compress the walls of the cylinder in radial directions and it will then be neces-
sary to place reinforced frames at the point of junction. Even when the transition
from the end pieces to the cylinder is smooth, additional local stresses occur, and
the destruction of the cabin during strength tests by the use of internal pressure,
usually begins at the joint of the end pieces with the cylindrical part of the
cabin.
In the walls of a cabin of a cylindrical form with a circular cross section,
under the action of internal forces, only normal tensile stresses occur in the
longitudinal and transverse cross sections. But if the cabin has another form (and
this is often necessary, due to the design of the airplane) the cabin walls will STAT
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subjected to bending and therefore must be made thicker and more heavily reinforced.
All this tends to increase the cabin weight. Strength and weight requirements are
the most important factors that limit the use of positive pressures exce'eding 0.25 -
0.30 atm in noncylindrical cabins.
The walls of pressurized cabins are usually riveted from duraluminum sheets of
a rather small thickness (1.5 - 3 mm), reinforced by longitudinal and transverse
frames. Since canopies are loaded wi th large aerodynamic forces, careful attention
must be given during manufacturing and operating to their strength and to mechanisms
securing them to the cabin.
Adequate sealing of the cabins is one of the most important conditions that
make it possible to sustain the necessary positive pressure. This is achieved by
1) an airtight structure of the cabin and the canopy (sealing of riveted seams and
canopy connections); 2) making hatches and enclosures airtight; 3) sealing the out-
lets of rods, fillets, cables, electric wiring, etc. Failure of the airtight struc-
ture of the cabin or its partial inadequacy may lead to such air losses that the
work of the supercharger is unable to compensate them. Therefore, when operating
pressurized cabins, special attention must be given to the condition of the cabin
seal and to elimination of any defects.
The rigidity of the structure of a pressurized cabin is especially important .at
the outlets of movable leads, controll ing various units, since this makes the seal-
ing of these outlets more reliable.
A wider margin of safety for the crew working in a pressurized cabin is achiev-
ed by preventing explosive decompression. With an increase in altitude, greater
positive pressures in the cabin become necessary and a sudden blowout will be
accompanied by an even more dangerous pressure drop. If the cabin is shot through,
an explosive decompression may be prevented by a rapid increase in the amount of
air blown into the cabin by the supercharger or by impeding depressurization by
covering the walls of the cabin and the enclosure with special protectors hindersTAT
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the flow of air through the hole. This protector may simultaneously serve as a
heat insulator.
To prevent fogging and freezing of the canopy, it is heated with a stream of
hot air. The sections of the enclosure not heated by this air flow, for the same
reasons (and also in order to decrease heat losses), may be designed with double
windows, with desiccator cartridges inserted in the air gap between them.
Section 39. Election Seats for the Crew
Desi n of E ection Seats and the Principle of Their Operation
With the increase in flying speed, the problem of leaving the cabin of the air-
plane in case of emergency becomes more difficult. These difficulties increase for
the following reasons.
1) As the speed increases, the aerodynamic forces acting on the body of the
pilot as he bails out from the cabin, increase greatly in magnitude. For instance,
at a speed of 600 km/hr the body of the pilot, showing only halfway above the cock-
pit, is subject to an aerodynamic force of about 450 kg from the oncoming air
stream; the head alone is subject to an impact of about 60 kg. Therefore, bailing
out at high speeds requires considerable time and strength, which may prove to be
above the physical possibilities of the pilot.
2) With the increase in speed, the danger increases of colliding with the tail
assembly or some other part of the aircraft after the pilot has left the cockpit.
Therefore even if the pilot left the cabin successfully and on time, for instance by
climbing over the side of the cabin, there still remains the danger of colliding
with the tail assembly.
3) A high-speed air stream has a marked effect on the unprotected face and on
the lungs of the pilot.
For these reasons (the first two being the most apparent), bailing out from the
aircraft in the usual manner (bailing out directly from the cabin) is restricted to
STAT
flying speeds of 500 ? 600 km/hr. One of the means for getting clear of the air'-
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craft is a bail-out (ejection) from the cabin of the pilot's seat together with the
pilot himself.
1
vc
, 1
1111
b;-4 Ard
A)
Fig.218 - Principal Diagram of an Ejection Seat
During ejection, the canopy is automatically jettisoned
after which the ejection mechanism is actuated.
a) Seat; b) Footrest; c) Hand rail; d) Backrest; e) Headrest;
f) Ejection mechanism; g) Guide rail
Ejection may be performed upward through the open cabin enclosure or downward
through a special porthole in the fuselage. Ejection of the pilot's seat upward by
means of compressed air or firing of a cartridge is structurally more simple.
Figure 218 shows the principal diagram of an ejection seat. Two pairs of
rollers are fastened to the backrest of the seat. These slide over two guide rails.
The guide rails are placed on both sides of the backrest and are fixed to the bulk-
heads of the fuselage. The ejector consists of a cylinder hinged to the fuselage
with its lower end, and a piston with a rod. The latter is placed inside the cylin-
der, and the upper end of the rod is hinged to the backrest of the seat. During a
normal flight, the piston is locked inside the cylinder by a special retainer in STAT
der to prevent movement of the seat in the guide rails during negative acceleration.
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During ejection, the canopy is jettisoned automatically, followed by actuation
of the ejector mechanism. The cylinder is filled either with gases under high pres-
sure from firing of a cartridge or with compressed air from a special tank under a
pressure of 100 - 150 atm. Under the action of this high pressure, the retainer of
the piston inside the cylinder unlocks and the piston, pushed by the gases, entrains
the seat, together with the pilot, secured to it by his harness. The seat moves
along the guide rails for only a very short period of time (0.15 - 0.20 sec), but
the acceleration is great so that, by the time the piston leaves its cylinder, the
seat has acquired a very high velocity. With this velocity Vo, known as the initial
velocity, the seat begins its free flight relative to the airplane.
When the seat enters the air stream and experiences its braking effect, it
starts lagging behind the aircraft which continues its flight. In the air, a second
or two after the ejection, the pilot harness is unlocked, the seat starts moving
away from the pilot so that even a few seconds after ejection it is already possible
to open the parachute without running any risk.
To permit use of ejection at any altitude, the chute is provided with oxygen
equipment, which supplies the pilot with oxygen after he leaves the cabin. The use
of automatic devices for unlocking the pilot harness, opening the chute, etc.
greatly increases the reliability of functioning of the entire ejection system.
EnsurinigL the Necessary Tralpctory of the Seat
In order to ensure safety, the path of the seat should follow a trajectory
somewhat above the tail assembly. The actual magnitude*of this height depends on
the flying speed. With an increase in flying speed, the braking effect of the air
stream on the seat increases greatly (in proportion to the dynamic head) so that it
will lag behind the airplane to an even greater extent. Therefore, with an in-
crease in the speed of the aircraft Vc the path of the seat drops (if at the same
time the initial velocity of the seat Vo remains constant). This can be observed
STAT
in Fig.219, where examples of seat trajectories at different aircraft velocities Vo
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are plotted (the positions of the seat correspond to a time approximately 0.3 sec
after ejection).
Fig.219 - Relation Between Trajectory of the
Seat and Airplane Speed
The greater the speed of the aircraft V" the closer
will the trajectory of the seat approach the tail
assembly.
In addition to being influenced by the speed of the airplane, the trajectory of
the seat is also greatly affected by the initial velocity Vo. The higher Vo, the
higher will be the trajectory above the tail assembly (Fig.220). The necessary
initial velocity of ejection Vo is usually so selected that it secures the needed
elevation of the trajectory over the tail assembly. Thus, an increase in the air,-
craft speed Vc requires an increase in the initial velocity Vo. The initial velo-
city Vo for fighters with an indicated speed of 800 km/hr is 10 - 15 m/sec; if the
speed of the airplane is increased to 900 - 1000 km/hr, the required velocity Vo
may reach 20 m/sec.
At a constant aircraft speed Vc and initial velocity Vo, the path of the seeSTAT
may be elevated by increasing the weight G of the seat and by decreasing the drag
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:10.k the seat (which is proportional to ODF) and its lift usually directed downward
,(and proportional to CLF). Therefore, streamlined seats with a higher value of the
ratios _2_. and -g- are most favorable. This is particularly apparent at high
Cup CLF
. speeds.
The angle of inclination x of the seat (see Fig.218) has very little effect on
the height of the trajectory. The most effective magnitude for E is 5 - 200.
Fig.220 - Relation Between the Seat Trajectory
and Initial Velocity of the Seat
The greater the initial velocity of the seat Vo, the greater
will be the elevation of the seat path above the tail assembly.
In relation to the aspects of a given flight, it must be noted that, as far as
the height of the trajectory is concerned, a curvilinear flight is usually not more
dangerous than.a horizontal flight. This is due to the lower speed in a curvilinear
flight. In a dive, the distance by which the seat is separated from the tail assem-
bly is 15 - 20% greater than in a horizontal flight and, thus, is more favorable.
Acceleration During Ejection
The required initial velocity Vo is acquired by the ejected seat during a com-
paratively shdrt time which is equal to the power stroke of the ejector mechanism so
(see Fig.218). Therefore the seat moves in the guide rails at high acceleration.
This produces brief but large G-forces which act on the seat and the pilot. At a
given power stroke so, the overloads will be proportional to the required magnitusTKr
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of the velocity Vo. G-forces produced during ejection, tend to exert excessive
loads on the spine.
Figure 221 shows the maximum overloads n (the weight of the pilot is not taken
into consideration), which can be endured by man, depending on the duration and the
direction of their action. During ejection, the pilot can withstand brief overloads
28
24.
12
8
?
'd)
e)
Fig.221 - Maximum 0-Forces n Withstood by Man
The magnitude of ng depends on the duration and
direction of their action.
4 I sec
a) Tolerated overload ng; b) Spine-to-chest; c) Head-to-pelvis;
d) Pelvis-to-head; e) Duration of action
acting in the direction of head -.pelvis as high as 18 - 20 G and, in case of a
special construction of the seat (see later), even up to 28 G. Ejection seats are
designed for the average pilot, and therefore overloads during ejection must not
'exceed 18 - 20 G. Such G-forces, due to their brevity (0.1 - 0.2 sec), can be with-
stood by the average flight crew without harmful effect.
It could have been possible to decrease the overloads produced by ejection by
extending the power stroke so, during which the seat gains speed, but the stroke is
of the cabin and is ordinarily equal to 0.7 - 1.0 m. Witt?TAT
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such a power stroke and the overloads mentioned above, the initial velocity Vo,
necessary at high-speed flights, is reached.
rag
Acmor
gr mean
SI)
Fig.222 - Diagram Showing the Over-
loads n on the Seat During the
Piston Stroke so.
In ease of an unfavorable diagram (a)
the pilot will experience greater
accelerations n gmae
load n
gmax
coefficient 71 by the following relation
is connected with the initial
from where
Since the magnitude of the initial
velocity Vo depends on the average
acceleration in the process of ejec-
tion, it is a prerequisite to avoid any
fluctuations in the acceleration and,
therefore, avoid overloads (Fig.222a),
striving for a smooth increase and de-
crease of overloading throughout the
entire stroke of the piston (Fig.222b),
i.e., an attempt should be made at in-
creasing the
efficient of
full stroke.
magnitude of n (the co-
acceleration) during a
This would tend to de-
crease the maximum overload n ax dur-
ing ejection.
The magnitude of the maximum over-
velocity Vo, the power strokes and the
0,
g III .1 I 914. (is?
-
V n O= 9 r *.i"g max SO ?
(107)
Therefore, at a given velocity Vo it is possible to decrease the overload n
by increasing the power stroke so and the coefficient 71. But if it is necessTATto
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increase the initial velocity Vo without, at the same time, increasing the overload ?
.
n (for instance, in connection with an increase of the flying speed), this can
be accomplished by enlarging the power stroke so and the pressure coefficient n.
The product ng mean so (where ng mean =1 ng Max) is an important characteristic
of the ejector performance, and its magnitude is equal to the distance of the free
fall of an object during which the velocity Vo is attained. The influence of the
overload ng mean and of the power stroke so on the magnitude of the initial velo-
city Vo is exactly identical.
The Influence of the Air Stream During and After Ejection
Even
while the seat is moving along the guide rails, the face and body of the
pilot, as they emerge from the cabin,
f7
10
5
05
t sec
Fig.223 - Overloads Caused by Abrupt
Deceleration of the Velocity of the
Seat in the Air Stream
The G-forces reach their maximum at
the instant the seat leaves the guide
rails and then diminishes sharply.
are subjected to the
moving at the
air stream of
impact of the ai/r stream,
speed of flight. In an
such velocity the pilot
remains for only about 0.1 sec, until
the seat disengages itself from the
guide rails. After that, the movement
of the seat is greatly retarded and the
speed of the air jet diminishes. At
indicated speeds lower than 800 km/hr,
a direct impact of the air stream on
the face produces heavy loads which,
however, are tolerated without harmful
consequences or ill effects. At fly-
ing speeds over 800 - 850 km/hr, the
pilot's face must be protected during
ejection.
Due to the abrupt deceleration of the seat in the air stream, considerable
accelerations are produced in the direction of flight and in the corresponding
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STAT
forces of inertia. Forces of inertia, equal in magnitude to the aerodynamic forces
acting on the seat, produce heavy acceleration in the back-to-chest direction. This
overload reaches its greatest magnitude at the instant the seat clears the guide
rails and then diminishes abruptly (Fig.223). Near the ground, at a velocity of
900 km/hr, the magnitude of the overload due to the braking effect may be about 20G.
This magnitude of G-forces despite their brevity, is near the maximum (see Fig.221).
With an increase in altitude, the rarefaction of the air causes a reduction in the
forces that produce the braking effect on the seat and, accordingly, a decrease in
the overloading. Overloading due to the braking effect may be reduced by increasing
the weight G of the seat and by streamlining its form (decreasing the braking
effect), i.e., by increasing the relation
e,F
After the seat leaves the cabin, besides the phenomena just discussed, a sharp
deceleration of the upward movement takes place, produced by the gravity of the seat
and by the negative lift (directed downward). As a result, the overload which, dur-
ing the ejection, acted in the head-to-pelvis direction changes its sign to the
negative and, after the seat is ejected, acts in the pelvis-to-head direction.
These overloads usually are not great (about 40), and are brief (less than one sec-
ond), but, they act in the most unfavorable direction (see Fig.221).
After the seat leaves the cabin, it may start rotating about its lateral axis.
This is produced by the moment of the aerodynamic forces acting on the seat and,
also, by the initial moment of the eccentrically applied ejection forces, and by the
action of aerodynamic forces on the upper part of the seat during ejection. The
magnitude of the angular moment depends greatly on the height of the headrest. If
the height of this headrest is chosen correctly, the rotation is not great and does
not lead to unfavorable consequences.
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v
Ejector Seat Mounts
An ejection seat (Fig.218), as previously stated, consists of a seat, guide
rails, and the ejector mechanism.
An ejection seat is designed to assist in counteracting the heavy overloads
during ejection. For this reason, the seat is supplied with a footrest, hand rails,
and an inclined backrest with a headrest. Before ejection, the pilot raises his
feet, places them on the footrest, grasps the hand rails, and leans on the backrest
thus considerably relieving the spine of overloads. In order to decrease overloads
and to soften the impact, the seat, backrest, and headrest are covered with elastic
cushions. With the use of such devices the maximum overload, tolerated by man, is
increased to 28G with the duration of its action being equal to,, 0.015 sec.
Placing the feet on the footrest relieves the pelvis of large loads, prevents
catching the legs on some part of the cabin during ejection, and decreases the drag
of the seat and of the pilot on reaching the air stream. An inclination of the
seat, best for relieving the spine, is chosen. The pilot must strap himself to the
seat securely with the shoulder harness to prevent his body from leaning forward
during ejection under the action of inertia forces. At the same time, the shoulder
harness must not be too tight since this would cause an additional compression of
the spine.
To decrease overload during deceleration and to elevate the trajectory of the
seat, it is desirable to streamline its form.
Ejectors may be pyrotechnical (operated by explosion of a cartridge) or pneu-
matic (operated by a tank with compressed air). The advantage of a pyrotechnical
mechanism over the pneumatic type lies in its small weight and simplicity of design.
But pyrotechnical mechanisms have less stable operating characteristics: at low
temperatures the time of explosion is increased and, therefore, the initial vela-
city Vo decreases. Since the temperature changes in the cabin during flight are
STAT'
not great, this drawback of cartridges has no practical significance.
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Careful testing of all the elements of ejection seats and a control of the high
quality of their manufacture guarantee reliability of their functioning.
The ejection seats described above can be effectively used at velocities of
about 900 km/hr, at low altitudes. At greater altitudes, this maximum speed also
increases.
Possibilities of Ejection at Higher Flying Speeds
The me-Lhod of ejection described above is restricted to flying speeds of the
order of 900 km/hr for the reasons discussed below. If these causes could be
eliminated, this limiting speed could be increased.
1) On increase in the flying speed, the elevation of the seat trajectory above
the tail assembly becomes smaller. Elevation of the trajectory by increasing the
initial velocity Vo, without at the same time enlarging the overloads, demands an
elongation of the power stroke of the ejector so. Since the height of the cabin is
limited, the use of telescoping ejector mechanisms, together with an increase in the
coefficient T1 by using a cartridge train may be the solution to this problem.
An elevation of the trajectory may be achieved by increasing the weight of the
seat (at a simultaneous increase of the power of the cartridge) and by improving the
streamlining of the seat, i.e., by increasing the ratio G
clop
In aircraft with vertical tail surfaces, the elevation of the trajectory can be
decreased so that, with this type of tail assembly, the flying speed can be increas-
ed, which would allow ejection without increasing the initial velocity Vo and the
overloads. *
Ejection downward requires smaller initial velocities and, therefore, permits
a considerable reduction in the overloads; for this reason, ejection downward can be
used at much greater velocities than ejection upward.
2) At high flying speeds, the impact of the air stream on the face is painful.
This can be eliminated by the use of a special blind, pulled down in front of the
pilot's face during ejection and protecting him from the impact of the relativeSTAT
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stream.
3) With the increase in flying speeds, the overloads from the braing effect of
the air stream increases. The fight against such overloads is very difficult. Some
results may be obtained if the weight of the seat is increased and its streamlining
made more perfect (an increase in
clop
4) As the velocities increase, the rotation of the seat becomes more intensi-
fied. This may produce large additional overloads due to the centrifugal force.
In this respect, a forward rotation is the most unfavorable. Rotation of the seat
can be prevented by means of a small auxiliary parachute, attached to the upper part
of the seat and opening automatically as soon as the seat leaves the cabin.
In this manner, although increasing flying speeds make ejection of the crew
more difficult, these difficulties can be overcome. By making use of the measures
discussed above, which tend to decrease the difficulties of ejection at high speeds,
it is possible to employ ejection seats up to speeds near the velocity of sound.
Flying at supersonic speeds would require closed, detachable cabins, and also
discovery of an effective means for decreasing the speed of the aircraft.
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CHAPTER XI
LANDING GEAR
Landing gears serve for taxiing, for take-off runs, and for landing runs.
Also, the landing gear must absorb the kinetic energy of the aircraft during touch-
down.
Fig.232 - Landing Gear Diagram of a
Modern Aircraft
The main wheels are placed slightly
behind the center of gravity of the
aircraft, and the nose wheel is carried
far ahead.
The predominating landing-gear type
in modern high-speed aircraft is a re-
tractable tricycle landing gear with a
nose wheel (Fig.232). In this type,
main wheels are placed slightly behind
the center of gravity of the aircraft,
and the front wheel is carried far to-
ward the nose.
Retractable landing gears make it
a) Center of gravity of aircraft; possible to decrease the drag of the
b) Nose wheel; c) Main wheel aircraft in high-speed flights. At the
same time, they somewhat complicate the
structure and increase the weight. Landing gears with a nose wheel were used even
in the early years of aeronautical development (Ufimtsev's
spheroplane, 1909), with
the nose wheel having the function of an anti-nose-over device. Landing was per-
formed on the two landing gear wheels and the tail skid. After some time this
anti-nose-over device went out of use. This can be attributed to the fact thai -I -
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nose of the fuselage was built quite short and, therefore, the nose wheel was placed'
very near the main wheels. This caused frequent failures of the front landing gear
strut due to heavy loads. Moreover, this short distance led to rocking of the air-
craft and created the danger of wing-over (ground loop).
A diagram of a landing gear with nose wheel was revived in the Thirties, since
it had some advantages over a landing gear layout with a tail wheel or a tail skid.
The presence of a nose wheel permits
applying the brakes more fully in order
to shorten the landing run. Here there
is no danger of nose-over. The existing
landing gear types do not require a very
exact landing performance. In fact, no
matter what the landing speed may be, re-
actions of the main wheels will always '
rotate the aircraft so as to diminish the
angle of attack of the wing. This will
Fig.233 - Landing of an Aircraft with
Drift.
Since the direction of motion of the
aircraft does not coincide with the
plane of symmetry, the wheels of the
landing gear will be loaded with
lateral forces F.
a) Angle of crab; b) Center of gravity
of aircraft; c) F = force of ground
reaction; d) Direction of motion of
aircraft; e) Axis of aircraft
Fig.234 - Motion of the Aircraft
The forces of friction F and the force
of inertia N1 of the aircraft form a
couple which forces the aircraft to nose
over.
a) N = force of inertia; b) Center of
gravity of aircraft; c) Force of friction
lead to a decrease in lift, forcing the aircraft closer to the ground. A lower
value of lift during the run insures greater normal loads on the wheels, and, STAT
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therefore, greater friction forces. It is true, however, that the length of the run
is negatively affected by the fact that the head resistance of the aircraft is small
in this case because of negligible angles of attack.
Finally, statements are required on the presence of directional stability dur-
ing runs over an airfield. If, for some reason, the aircraft moves with a crab
(Fig.233), either in case of landing with a cross wind or if the landing gear
- strikes an obstacle asymmetrically, the lateral forces of the main wheels will de-
crease the angle of crab. .No lateral forces are created in the nose wheel since
the front strut is designed as the swiveling type, i.e., it can rotate freely. A
tricycle layout of the landing gear has two disadvantages.
1. When an airplane moves over a rough runway, the nose wheel surmounts ditches
and hillocks with greater difficulty than the tail wheel. This is due to the fact
that the friction force F and the inertia N (Fig.234) of the aircraft form a couple
which forces the nose down and causes the nose wheel to be pressed hard to the
ground.
2. When an aircraft runs along the runway at a high speed, self-oscillations of
the front strut may occur (shimmy phenomenon), whose physical aspect is described
below.
Self-Oscillations of the Front Strut of a Landing Gear
The front strut of the aircraft may always perform two independent types of
motion, i.e., it has two degrees of freedom (Fig.235). First, the wheel (since it
is able to swivel) can turn through some angle y about its vertical axis. Secondly,
the wheel (or, more correctly, the point of its contact with the ground) can deviate
from the direction of motion to a certain magnitude y. This motion is caused mainly
by the elasticity of the tire, and partially by the elasticity of the strut, and as
a result of crabbing of the aircraft nose. From here on, in order to simplify our
reasonings, we will consider only the elasticity of the strut.
The two indicated degrees of freedom create the possibility of transverse STAT
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VII
Fig.235 - Diagram of Self-Oscillations
of the Front Strut of the Landing Gear
Due to the presence of the initial angle
of drift, the wheel vibrates with re-
spect to the direction of motion of the
aircraft.
. a) Direction of motion of aircraft
III etc.),
oscillations.
Figure 235 shows a diagram of
kinematic transverse oscillations,
i.e., oscillations due to the lack of
possibility of the wheel' to turn in a
plane parallel to the ground (i.e.,
relative to the vertical axis), with-
out advancing motion. Moreover, it is
assumed that the wheel cannot have any
advancing motion in a direction per-
pendicular to its plane.
Let us assume that while the air-
plane moves along the x-x axis, the
wheel, under the action of external
forces, turns to make a certain angle
Ymax with it (position I). In such a
case, the wheel will tend to move at an
angle of Y with respect to the
motion of the aircraft. If, during the
first moment of motion of the turning
wheel (at y = 0), the plane of the
latter is normal to the ground (It), in
the consecutive moments (y greater
than 0) the wheel will tilt so as to
form a certain angle 8 (III).
In Fig.235, the positions of the
wheel, placed on the right side (I,
demonstrate its inclination with respect to the vertical (angle@ ). STAT
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I ,
Due to the presence of the angle e (position IP), the point of intersection of
the wheel axis with the plane of the ground, with respect to which the wheel is
'turning, will not be located at infinity, and the wheel while deviating from the
path x-x (enlarging of y) will turn so as to decrease the angle y. At a certain
inoment, the wheel will occupy the point of greatest deviation III Cr a' Yilax
6 iie) and the angle y will be equal to zero.
Starting with position III, as a result of the presence of the same angle 8,
the wheel will begin to approach the axis x-x, i.e., y and 6 will become smaller
and the angle y will increase in magnitude but only with a negative sign. At the
moment of intersection of the axis x-x by the wheel (position V) the angle 8 will be
equal to zero, and the angle y will have its maximum negative value. All further
positions of the wheel (VI - IX) represent a mirror image, on the left side, of the
first five positions.
It can be concluded from the above diagram -that, during self-oscillations, the '
front strut will execute harmonic oscillations.
Under real conditions, the angular distances y and the lateral shifts y will be
performed with variable speeds, and, therefore, with acceleration. Due to these
accelerations, the strut will be loaded with forces of inertia.
Moreover under real conditions, sideslip also occurs. Because of the presence
of inertia forces, elasticity of the tire, and possible sideslip, the self oscilla-
tions at high speed may proceed with an increasing amplitude (up to y ), which
might lead to detachment of the tire, breaking of the front strut, and to other
troubles.
In order to damp the oscillations of the front strut, a vibration damper is
mounted to it in which, with every turn of the wheel, a fluid flows through a small
aperture from one container into another. The energy of oscillation of the front
strut is thus transformed into heat, and as a result the vibrations are damped.
The most complete and thorough research data on self-oscillations of the fro] S.TAT
75
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.strut can be found in the book of Academician M.V.Keldish "Shimmy of the Nose
Wheel", which was honored with a Stalin prize.
Section 43. Basic Requirements of a Landing Gear
The basic requirements made on a landing gear are as follows:
1) Ensure free movement and maneuvering of the airplane on the ground. In
order to improve the maneuverability of the aircraft, the wheels are supplied with
brakes, and the nose wheel is made to swivel (braked wheels tend to shorten the
length of the runs).
2) Sufficient strength of the landing gear to satisfy all design cases which
are foreseen by the normal strength requirements.
3) Possibility of absorbing the kinetic energy of the tires and shock struts,
with relatively small overloading.
4) Minimum drug in flight, which is obtained by using a mechanism for retract-
ing the landing gear.
5) Absence of nose-over tendency, sufficient stability, andcontrollability of
the aircraft during its motion on the ground. All these requirements are satisfied
by a rational layout of the landing gear on the aircraft.
The layout of the landing gear is defined by the basic geometric parameters,
which are described in Fig.236: the stagger of the main wheels e, the wheel base b,
the track gage B, and the height of the landing gear h.
The stagger of the main wheels must be very small so as to decrease the loads
acting on the front strut. But at the same time the stagger must allow the plane to
take its parking position at any landing. This is insured by making
13, ? P.
For real aircraft, the following relationship exists:
- 0,1 ? 0,15.
b
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Since e
ing the
= hte, the ratio t) can be diminished by increasing the wheel base b, lower-
center of gravity h of the aircraft, and decreasing the angle 0.
Increasing of the wheel base b is
usually attained by elongating the nose
section of the fuselage and, partially,
by carrying the front strut forward.
Generally,
Fig.236 - Layout of the Landing Gear
Basic geometric parameters: e - Stagger
of the main wheels; b - Wheel base;
B - Track gage; h - Height of landing
gear.
------,
L 0,25 ? 0,35,
where L is the length of the fuselage.
Lateral directional stability and
controllability of the aircraft when
taxiing on the ground depend on the
track gage of the landing gear B. For
greater lateral stability and increased maneuverability of the aircraft on the
ground, greater track gage is deslrable; this will reduce swaying of the airplane,
permitting the use of softer shock absorbers. But the directional stability of an
aircraft with a wide track gage is impaired. This is explained by the presence of
the arm formed by the wheels with respect to the center of gravity of the aircraft
which, in case of a rough runway, will create large torsional moments. A wide track
gage is not advantageous from the point of view of the strength of the wing, as this
will cause excessive bending. In modern airplanes, the ttack gage fluctuates be-
tween 20 and 32% of the wing span. In multiengine aircraft the track gage is
usually determined by the position of the engine nacelles.
The distance of the center of gravity from the ground is determined by the
height of the landing gear. The height chosen for the landing gear is such that,
during a three-point landing of the aircraft, the wing has a landing angle of
attack. Sometimes the height of the landing gear is determined by the minimum STAT
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allowable distance of the lower tip of the propeller blade from the ground..
In the general case, any landing gear forms a three-dimensional energy diagram
with an axis to which the wheel is mounted. The shock absorber will be within this
energy diagram.
Section 44. Bicycle Type Landing Gear Diagram
High-speed jet aircraft have a relatively thin airfoil. This very often
creates unsurmountable difficulties if the main wheels of the landing gear have to
Fig.237 - Bicycle Type Landing Gear
The main wheels are placed under the fuselage.
a) Wing-tip wheel; b) Main wheel; c) Nose wheel
be retracted into the wings. The mounting of special bays or wells for the retrac-
tion of the landing gear impairs the aerodynamics of the aircraft.
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The disadvantages and difficulties of main-wheel retraction pointed out above
are eliminated in some aircraft by the use of bicycle landing gears. In this lay-
out, the main wheels are not placed (Fig.237) under the wings but under the fusel-
age, into which they are retracted. Therefore, the basic difference of this layout
from the usual one is the almost zero track gage of the landing gear. This last
fact deprives the aircraft of lateral stability during taxiing on the ground and
does not permit maneuvering of the aircraft with the help of the main-wheel brakes
of the landing gear.
Lateral stability of an aircraft with a bicycle landing gear is attained by
mounting relatively light-weight shock struts under the wings (wing-tip wheels,
Fig.237). Maneuvering of the aircraft on the airfield, especially at low speeds,
when the tail unit is not effective, is achieved with a controllable nose wheel
(controllable nose wheels are also encountered in ordinary landing gears).
It must be noted that lateral stability of an aircraft at high taxiing speeds
is present even without the wing-tip wheels, as a result of the action of the aero-
dynamic forces of the wing. Also, the bend of the wing makes the air-gap between
the ground and the wing-tip wheels large enough so that they will not strike a rough
runway. Wing-tip orienting struts make contact with the runway when the aircraft is
parked and when it moves at low speeds.
The presence of a bomb bay in the fuselage makes it necessary to mount the main
wheels of the landing gear in bombers far behind the center of gravity of the air-
craft. This increases the loading of the nose wheel considerably and does not pen-
mit the aircraft to move on its main wheels alone. Consequently, both take-off and
landing of an aircraft must be performed with the help of the nose wheel, in con-
junction with the main wheels.
In order to be able to use the cy max of the wing during landing and to per-
form the take-off at the most advantageous angle of attack, turning wings and exten-
sion landing gears are used. With a turn of the wing or extension of the landing
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gear the aircraft can move over the runway under different angles of attack.
Section 45. Landing-Gear Wheels
Landing-gear wheels ensure the roadability of the aircraft and absorb the shock
energy with their tires. The main elements of the wheels are the wheel rim and the
tire. The rim (Fig.238) is usually cast from an aluminum or magnesium alloy. It is
test together with the hub.
Fig.238 - Landing Gear Wheel
1 - Wheel rim with hub; 2 - Brake disk;
3 - Wheel axis; 4 - Removable rim;
5 - Brake jacket; 6 - Brake blocks;
7 - Brake chamber; 8 - Air intake pipe;
9 - Ball bearings
Fig.239 - Wheel and Tire
A wheel has two characteristic dimen-
sions: the outside overall diameter D
of the wheel and the diameter of the
cross section of the tire d.
a) Diameter of tire; b) Diameter of
wheel
For convenience in mounting and stripping the tire, one of the rims is remov-
able. Roller bearings are mounted inside the hub. All modern wheels are supplied
with brakes, whose disks are fixed on the axis; the steel brake-drum jacket is
pressed into the rim. Brakes may be of the following three types: tire brakes,
block brakes, and disk brakes. Brakes are mostly either pneumatically or hydraulic-
ally controlled. STAT
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Wheels 'have two characteristic dimensions (Fig.239): overall outer diameter D
of the wheel ana diameter of the cross section of the tire d.
Types of Wheels
Wheels in use at the present time may be subdivided into three types:
a) high-pressure wheels;
b) low-pressure wheels (Balloon type);
c) medium-pressure wheels (semiballoon type).
HiglIzaKsElt_ire wheels have a tire with a small diameter (d) and a high pressure
of the pre&iminary charge (about 8 atm). They can be used on airfields with hard
soil. Under these conditions, the wheel will have a small area of contact with the
ground. One of the defects of such wheels is that, due to the small volume of the
.tubes, they do not absorb shocks very well; moreover, they sink too much into swampy
ground, which in turn hinders take-off and may cause nose-over during landing.
Tires with superhigh pressure can be also grouped with this type. Their preliminary
air charge may be 10 or more atmospheres. High-pressure wheels because their over-
all dimensions are small, can be retracted more easily into an aircraft.
Low-pressure wheels have a larger tire diameter (d) and, therefore, a larger
volume of air, which helps them to absorb more kinetic energy. The roadability of
such wheels is greater than that of wheels with high pressure, because of low unit
pressure (less than 3 atm). Such tires are rarely used because of their large over-
all size, which makes their retraction into the aircraft during flight, quite diffi-
cult.
Medium-pressure wheels have all the good points and all the defects of the
first two types. They are used predominantly in our country. The pressure in the
tires of these wheels ???? 3-4 atm.
On wheels with small diameters, chamber brakes are generally used. Block
brakes are mounted on large wheels. Friction-disk brakes are used on some wheels.
The basic requirements of brakes are as follows:
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dependability of performance;
speed of braking and brake release;
simplicity and comfort of operation.
Tire Performance
A complete tire performance characteristic is given in the diagram showing the
load Pw, acting on the wheel, as a function of elongation of the tire 5 at different
values of the preliminary charge. Figure 240 shows a simplified diagram of this
dependence.
MOO
7000
5000
.5000
4000
Ma`
2000
1000 ?
a)
b)
20 40 110 10 100 120 140 IN 180
-
64171C.:
GT
Fig.240 - Wheel Characteristic D/d = 750/250
With the growth in the load adting on the wheel, deflation of the tire 6
increases (lowering of the wheel axis). The shaded area of the triangle
represents the work At, absorbed by the tire during its deflation. When
the tire has a preliminary charge of p = 3 and is fully deflated 6 =
170 mm; At max = 552 kg-m.
In the diagram oct is the deflation of the tire, which is desirable during
parking.
a) Pw kg - load on the wheel 750/250; b) 8 mm - deflation of the tire
It can be seen from this diagram that the load Pw acting on the wheel, in
STAT
relation to the deflation of the tire 61 changes according to a rectilinear law
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(from the pressure p). The hatched area of the triangle shows the work Ati, absorb-
ed by the tire during deflation. The work done is the product of force times dis-
tance. In our case, the force is the loading Pw of the wheel and the distance is
the deflation of the tire 6. Therefore, the work must be approximately equal to the
product
Pw
.6
?
This would have been true only if the force Pw would have been permanent in
value throughout the whole of the path 6. Since the loading of the wheel Pw
throughout the path a(deflation) differs from zero, the work done can be calculated
only by using the product of the mean value of Pw times 6. The mean value of the
load acting on the wheel will be equal to one half the greatest value of Pb,. For
this reason Ati is expressed in the following manner:
w
Ati = (115)
2
It may be seen from this diagram that, at a lower preliminary charge and at the
same deflation the tire will do less work. But if two wheels are loaded with an
equal load Pw the greater workAti will wi be done by the wheel, which will have a low-
er pressure. The greatest amount of work will be absorbed by the tire at its full
Ati max =
w max max
2
(116)
In this way, the work absorbed by the tire depends completely on the magnitude
of its deflation and the load acting on the wheel Pb,.
The main wheels of an aircraft are selected according to a standard, depending
on the deflation on parking act and the parking load Pwset. The deflation on park-
ing act is determined by the conditions of the life of the tires and prevention of
their detachment by lateral forces.
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According to the standard
0.2 - 0.25 for d,
6 ct
6 max = 0.7 d.
The pressure in the tires is so selected that, at a given deflation on parking,
the wheel load according to the standard would equal the real parking load.
The relation between wheel loading at maximum deflation of the tire to the
parking loading w max) is called the coefficient of load carrying capacity of the
Pw ct
wheel.
In order to make utmost use of the wheels on landing, it is necessary to use an
acceleration coefficient of operation neclose to the load carrying capacity of the
wheel. Then, in case of a rough landing, the greatest loading of the wheels will
equal Pw 1 and the tire will be deflated to such an extent, that any further in-
crease in this force will not diminish the volume of air in the tire.
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CHAPTER XII
SHOCK STRUTS
Section 46. Designation of Shock Struts
During the landing of an aircraft, at the instant its wheels touch the ground,
the path of motion of the center of gravity is somewhat inclined; in addition to the
horizontal speed component there is also a vertical component. The horizontal velo-
city component is damped during the landing run L due to the drag of the aircraft
and the forces of friction caused by the movement of wheels over the ground. The
vertical velocity component is damped during the path s, connected with the defla-
tion of the tire and contraction of the shock struts.
It can be demonstrated by simple calculations that the mass of the aircraft,
after touching the ground, undergoes vertical accelerations which are greater than
the horizontal ones. Therefore, the mass of the aircraft will experience large
forces of inertia, acting in a vertical direction.
Let us assume that the horizontal velocity component Vhor is equal in magnitude
to Vlalid
VhorP-4 Vland = 90 m/sec.
The vertical rate of descent V may practically become -2 misec.
But Vhor is damped during the path Lrun=1000 m, when Vy is damped during the
path sill m.
Therefore, the velocity ratio is STAT
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As a result, we find that the ratio of the paths is a magnitude many times
greater than the velocity ratio. This confirms the above statement as to the magni-
tude of the accelerations and the forces of inertia.
The shock absorbers of the landing
gear are designed to absorb the kinetic
energy of the aircraft during landing
and moving over the airfield.
The landing gear shock struts must
meet the following basic requirements:
1) A shock absorber, during its
power stroke, must absorb its share of
kinetic energy of the shock.
2) A shock absorber must perform
its work with small stresses. More-
over, the stresses in the shock absorb-
er must increase gradually, the great-
est stress coming at the end of a com-
plete stroke.
3) In order for the shock absorber to make the succeeding stroke, it is neces-
to limit the duration of the forward and reverse strokes. This time interval
not exceed 0.8 sec (maximum time for one stroke).
4) A shock absorber must have as large a lag as possible, i.e., the greater
of the absorbed work must be converted into heat, and must not return in
STAT
Vhor
: Vy = 50
:
2
=
25.
And the ratio of the path
Lrun :
s = 1000
:
1
=
1000.
,
5/7,12J
b)
Fig.241 - Diagram of Shock Strut
stroke
The diagram shows the stresses occurring
in the strut during its contraction.
The hatched area represents the work
done by the strut when fully contracted.
Curves 1, 2, and 3 correspond to differ,-
ent struts.
a) Maximum stress in strut; b) Stress
produced by contraction
sary
must
oart
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form of a reaction that would cause the aircraft to swerve.
? A full performance diagram of a shock absorber is given in Fig.241 which shows
changes in the load Pam acting on the shock absorber as related to its deflation s.
.
This diagram contains the following characteristic values: stress of pre-
:liminary contraction Pam? and the maximum stress Pam iai, which corresponds to the
maximum stroke of the strut s .
The area of the shaded portion of the diagram gives the magnitude of the
absorbed work (kinetic energy), since it represents the product of the mean stress
of the strut (force) by the stroke (path).
The curves of Fig.241, which belong to different shock absorbers, clearly
demonstrate the different character of change of loading of the strut as a function
of the deflation. The smoothest growth of loading of the shock absorber is shown by
curve 1, and the most abrupt by curve 3. But curve 3 is the one which absorbs most
work at the same stroke of the strut.
A shock strut must never work in the manner shown by curve 3, since the great-
est stresses there do not occur at the end of the stroke. As a result, large Over-
loads will occur during landings that are not too rough, and this would lead to a
premature wear of the structure.
On modern aircraft, landing gear shock absorbers consist of separate force
struts of the landing gear, which are so designed that they can greatly contract
under the action of external forces. Absorption of the shock takes place because of
contraction of the strut's length. Such struts may be constructed with rubber,
springs, or air as the shock-absorbing elements. At present, shock absorbers that
are a combination of pneumatic and hydraulic chambers are used. The fluid is highly
compressed in the air chamber and thus serves as an additional means of absorbing
the shock. Such shock absorbers are called oleo-pneumatic.
The skeleton diagram of this type shock absorption was first developed and
proposed by Prof. V.P.Betchinkin in 1921.
87
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?
Section 47. Function of An Oleo-Pneumatic Shock Strut
Pneumatic Lefi
In order to better comprehend the function of an oleo-pneumatic shock absorber
let us first examine the working principle of a strut of the old design (Fig.242).
e
Il
? , ,
' ? ' ,
4/ ?
Snaa 49 owl. orris
b)
Fig.243 - Diagram of a Pneumatic Strut
If there is no diaphragm, the shock
Fig.242 - Diagram of a Shock Absorber strut will work like a wheel tire with
the only difference that the curve AB
1 - Cylinder; 2 - Piston; 3 ? Diaphragm; will not pass through the origin of the
coordinates, due to the preliminary
4 - Packing; 5 ? Air; 6 - Liquid; contraction.
7 ? Filler plug a) Stress in the strut; b) Stress pro-
duced by compression
Let us imagine a strut charged with a fluid (6) and air (5), but without a dia-
phragm (3) in the cylinder. Under such a condition, the strut will work like a
pneumatic tire. The work, in this case, is absorbed by the compression of air. The
magnitude of the load Pam acting on the strut, with respect to the power stroke of
the piston s will change according to the curve AB (Fig.243), which is called a
polytropic curve.
In order to absorb more energy during a shorter power stroke of the leg, it is
necessary to create in it a preliminary initial pressure Po (segment OA). In
reality, during the power stroke of the leg, which corresponds to the given dia-
gram AB, the work will be expressed by the area OABD (shaded area). If the STAT
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?
preliminary pressure in the strut is equal to zero, the pressure diagram will have
the form of the curve OC; however, then in order to absorb the same amount of energy
as in the case AB, i.e., in order to absorb the area OABD, it is necessary to in-
crease the base by the magnitude of DE, i.e., to increase the power stroke of the
strut s. In other words, in case of soft shock absorption, the stroke of the strut
OE must be greater than in case of rigid shock absorption.
In case no diaphragm is used, the shock strut will work like a wheel tire.
Such a strut will have a small hysteresis, which does not satisfy one of the re-
quirements made on shock absorbe...s. In this design, the air first stores the work
and then gives it off again. The nature of the change of the working diagram de-
pends very little on the rate of contraction of the strut.
Hydraulic Leg
Let us examine the function
phragm
of a strut filled with a fluid and having a dia-
in the cylinder, but with an open filler plug (7) (see F1g.242). It is
that when the strut contracts,
pressure of the air (5), which
#lat.
.e ?
d)
--0
RIM
am
Fig.244 - Diagram of a Hydraulic Strut
The shaded area is the work done by
friction, which is absorbed by the strut
because of the flow of the liquid through
the opening in the diaphragm.
a) Small orifice; b) Work during power
stroke of strut; c) Large orifice;
d) viork during idling stroke of strut
89
clear
the
com-
municates with the outside atmos-
phere, will not change; therefore,
the strut will operate like a hy-
draulic leg. The whole of the
shock energy will be expended on
overcoming the hydraulic resistance
during the flow of this fluid
through the opening of the dia-
phragm. All the energy will be
converted into heat and the strut
will not return to its initial,
straight position. .Therefore sucsTAT
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a strut has a larger hysteresis than a purely pneumatic one.
Let us see what factors are influencing the diagram that shows the dependence
of the force Pam on the stroke of the strut s (Fig.244), for the case of a purely
hydraulic shock absorber.
The force that forces the fluid through the opening in the diaphragm depends on
the hydraulic resistance of the liquid. This resistance is proportional to the
square of the rate of flow of the liquid and inversely proportional to the area of
the cross section of the opening through which it flows. The initial and end velo-
city of the strut contraction are equal to zero; therefore, the force Pam, which is
necessary for strut contraction, is also equal to zero at the beginning and the end
of the stroke. As demonstrated by research, the rate of the strut contraction and
also the force Pam increase rapidly in the beginning in magnitude and then start to
decrease gradually. For this reason, the diagrams showing the dependence of Pam
on s have the form of the curves shown in Fig.244.
With respect to the dependence of the hydraulic resistance of the fluid on the
cross-sectional area of the orifice, it is expressed in the following manner; If
the cross-sectional area of the orifice is small, the curve of Pam as a function of
s is much steeper (curve 2). Therefore, here the stresses in the strut will be much
greater in magnitude and they will increase very rapidly. As the cross-sectional
area of the orifice in the diaphragm increases, the diagram will correspondingly
lower (curve 3). In modern aircraft, the area of the orifice is equal to about 2%
of the area of the piston. The form of the orifice also has an influence on the
magnitude of the stresses in the strut Pam. The viscosity of the liquid has an
effect of the same kind.
If there are forces which return the piston of the strut to its initial posi-
tion, some additional work will be spent on this. In such a case, Pam has a nega-
tive value (curve 4).
STAT
The character of the curve P as a function of s will change in accordance
am
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with the roughness of landing, on which the rate of the strut contraction depends.
In designing a shock strut, measures can be taken to influence the slope of the
curve P. This is achieved by changing the cross-sectional area of the orifice
with the help of a metering pin (Fig.245), which moves together with the piston
through the orifice in the diaphragm.
d)
Fig.245 - Metering Pin in a Shock
Strut
A metering pin changes the
cross-sectional area of the ori-
fice through which the liquid
flows and, thus, affects the shape
of the curVe 1 (Fig.244).
a) Stress of liquid; b) Diaphragm;
c) Cylinder; d) Pin; e) Piston
rie
a)
go Mil
I OA,
Fig.246 - Return Valve
During the forward stroke the valve opens
a larger orifice for the flow of fluid than
during the return stroke. Therefore during
the return stroke the path of the liquid
will be more inhibited than during the for-
ward stroke.
a) Forward stroke; b) Return stroke
The pin is chosen of such a shape that, at the beginning of the stroke when
the velocity of the piston is great, the area of the orifice
later, as the strut contracts and the velocity of the piston
would be larger, while
diminishes, the area of
the orifice would diminish too. In case of a rough landing, the stresses Pam will
increase more smoothly.
In order to achieve a smoother shock absorption, it is desirable to have a
larger hysteresis during the return
stroke. For this purpose, the strut must be
supplemented by an additional device, which would partially close the cross section
of the orifice during the return stroke. Such a device is a diaphragm with a
return valve (Fig.246).
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If such a return valve is employed, the curve (4) (see Fig.244) will follow a
lower path, as there will be a greater resistance to the flow of the liquid.
Combined Strut
Let us now imagine an ordinary strut, filled with a liquid and air at an
initial pressure pe The function of this strut will represent a combination of the
Fig.247 - Stress-Strain Diagram for an Oleo-Pneumatic Strut
a - Variation in stress in the strut during the stroke, both during forward
and return stroke; b - Curve 1 represents the forward stroke of the liquid
when the cross section of the orifice is small; curve II is produced by the
metering pin; curve III represents the return stroke of the liquid when a
return valve is used; and curve IV corresponds to the forward stroke of the
liquid in case of a soft landing.
a) Air pressure; b) Forward stroke; c) Return stroke
functioning of the two struts examined above, i.e., air (pneumatic) and oil
(hydraulic) struts. This operation maybe expressed by the diagram shown in
Fig .247a.
The first force (the air pressure) will vary both during the forward and the
STAT
return stroke according to the curve amb, with which we are acquainted from Fig.243
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(curve AB). The energy absorbed by the air during the forward stroke and liberated
during the return stroke is represented by the area Oambc0.
The second force (additional load) on the strut due to the flow of fluid
through the orifice in the diaphragm is known from Fig.244.
In order to obtain the curve representing the sum of pressures of the liquid
and air, it is necessary to add the pressures (ordinates) of the curve 1, 2, or 3
(see Fig.244) to the pressures (ordinates) of the curve amb. Then we will obtain
the cumulative curve anb, and the energy absorbed by the strut during the forward
stroke will be expressed by the area Oanbc0.
For the return stroke, the air pressure must follow the curve bma, and the
magnitude of the pressure exerted by the liquid (see Fig.244, curve 4) is traced
below the curve bma. This will give the curve bda, which will give us the cumula-
tive curve representing the pressures of both air and liquid during the return
stroke. The area anbda represents the energy converted into heat during the forward
and the return strokes. The energy not converted into heat will be expressed in the
diagram by the area Oadbc0.
In order to have the liquid flow through the orifice in the diaphragm and thus
perform work of friction, the pressures in the liquid and in the air chambers must
differ. The total force of the liquid pressure (ordinate 1 - 3, Fig.247b) is com-
posed from the force of air pressure (1 - 2) and the force of friction of the liquid
through the diaphragm (2 - 3).
We already know that it is possible to change the slope of the curve represent-
ing the loading of the hydraulic strut (see Fig.244), by ,changing the area of the
orifice in the diaphragm by mounting a metering pin or by using return valves. If
we make the cross-sectional area of the orifice smaller, the loading will increase
sharply in magnitude (curve I, Fig.247b). Such a strut (a more rigid one) makes it
possible to have more work done during a shorter stroke simax.
We can obtain a curve of the shape II by mounting a metering pin. Struts with
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such characteristics have smooth shock absorption during taxiing. But 'in order to
absorb the required amount of energy, the stroke of the piston must be greater.
If we introduce a device to impede the flow of the liquid during the return
stroke (see Fig.246), the energy absorbed by the liquid during the return stroke in-
creases (curve 4 in Fig.244 will become lower), and instead of the ordinates 2 - 4
(Fig.247b) we will obtain the ordinates 2 - 5, i.e., the characteristic of the re-
turn stroke will be in the form of the curve III. The amount of energy converted
into heat increases in such a case (area anbIIIa).
Therefore, in order to decrease the stroke of the piston it is necessary to
have a strut with a greater braking during the forward stroke (a higher diagram);
however, shock absorption in this case will be more rigid and will give greater
overloads during taxiing. At present, such struts are not used.
To ensure a smooth growth of the stresses during the forward stroke and a de-
crease of stresses in the shock absorber during short strokes, it is necessary to
have a strut that brakes during the return stroke but not during the forward stroke.
The greatest amount of shock energy is converted into heat during the return
stroke.
In case of a medium rough landing, the stroke of the strut and the work of the
fluid will be smaller (see curve IV in Fig.247b).
Section 48. Influence of the Forces of Friction Produced by the
Packing Collar on the Work of the Strut
Until now we have been examining the load acting on the strut Pam only as a
function of the force exerted by the air pressure in the air chamber and as a func-
tion of the force of the hydraulic resistance of the fluid in the orifice of the
diaphragm. But in reality, Pam depends also on the forces of friction produced by
packing collars and bearing axles. This can be found by a static test of the strut
(a slow contraction and extension of the strut), since then the rate of the strut
contraction is small; therefore, the fluid does not produce a braking effect and
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.cannot influence the curve
type shown in Fig.243, i.e
stroke.
Pam above s. Thus we would have to obtain a curve of thel
Op the same curve AB for the forward and the return
Fig.248 - Diagram of Stresses in a
Pneumatic Strut when the Friction of the
Packing Collar is Taken into Account
Under the action of friction of the pack-
ing collars and bearing axles, the
air-pressure curve is shifted upward
during the forward stroke and downward
during the return stroke.*
a) Forward stroke; b) Return stroke;
c) Curve of air pressure
of friction depends on the air pressure;
collar rings be pressed to the
However, experiments give
curve II (Fig.248) during the for-
ward stroke and curve III during the
return stroke, but not curve I which
corresponds to curve AB in Fig.243.
We see that the curve of an
actual strut will be shifted upward
with respect to the air-pressure
curve I during the forward stroke,
and downward during the return stroke.
This shift exactly corresponds to
the force of friction Ffr. But the
force of friction does not remain
constant through the entire length
of the stroke; it will increase dun-
ing the forward stroke and decrease
during the return stroke. This is
explained by the fact that the force
the greater (this pressure the more will the
surface of the cylinder.
Therefore, until now we were examining the performance of ideal struts which
do not have any friction forces produced by collar rings and bearing axles. In
reality, however, the characteristics of shock struts include the magnitude of
friction forces corresponding to the given position of the piston. At every instant
of the piston stroke, it must overcome air pressure and the force of friction. STAT
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Friction will produce work during both forward and return stroke; this work is irre-
versible, i.e., it constitutes an additional hysteresis. In Fig.248, work produced
by the forces of friction during the forward stroke is expressed by the area AbcB,
and during the return stroke by the area ABen.
Here the force of friction produced by packing rings is added directly to the
air pressure, while the friction produced by the liquid is absent because of the
slow movement of the strut. In order to obtain the characteristics of a strut under
actual working conditions, i.e., during a rapid contraction and expansion of the
strut, we must take curves anb and bda (see Fig.247a) representing air pressure and
liquid pressure and add to them the forces of friction both up and down, i.e.,
corresponding to the forward and reverse strokes. Then we will obtain the charac-
teristics of the strut (Fig.249).
Fig.249 - Diagram Showing the Stresses in an Oleo-Pneumatic
Strut, Taking the Friction Forces Produced by
Packing Rings into Consideration
Under the action of friction, produced by the packing collars and bear-
ing axles, the curve representing the stresses in the struts will be
shifted upward during the forward stroke and, downward during the return
stroke.
Here:
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Curve amb denotes the performance of the air during the forward and return
strokes;
Curve II is the performance of the liquid during the forward stroke;
Curve bda is the performance of the liquid during the return stroke;
Curve V is the performance of the liquid, including the friction forces pro-
duced by the collar rings during the forward stroke;
Curve VI is the performance of the liquid, including the forces of friction
produced by the collar rings during the return stroke.
At any instant of the piston stroke, it must overcome air pressure, hydraulic
resistance of the liquid passing through the diaphragm, and friction produced by the
packing rings and bearing axles.
The area zhzik (shaded) represents the hysteresis of the strut work during the
forward and return strokes. We see that the force of friction produced by packing
rings and bearing axles increases the hysteresis of work.
It has been established through tests that the friction produced by packing
rings and bearing axles comprises about 20% of the total force Pam of the strut.
For decreasing the length of the strut, an initial air pressure is creat,ed in
the strut, which according to the type of landing gear, varies from 25 to 100 atm.
The initial stress in the strut, friction forces not considered, will be equal to
Pa ni F P
Jr .
(119)
and the total initial stress in the strut, friction forces taken into consideration,
will equal
pa m = 1.2 Fppo, (120)
where F is the piston area in cm2;
pc) is the initial air pressure in kg/cm2.
is the maximum dynamic load on the strut.
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Therefore, in order to contract the strut, the external force, acting on the
strut, must be greater than 1.2 Fp . po.
Ordinarily the magnitude chosen for pc, is such that the product 1.2 Fp . pc)
would be less than the parking stresses. Then during parking of the aircraft, the
strut will be already somewhat compressed.
The relation of the initial loading to the parking compression is called the
coefficient of preliminary compression and is designated by the letter no.
Therefore,
no =
1.2 F p
p o
am.park.
(121)
In modern aircraft, no ms 0.6 - 0.8, and sometimes even smaller. The smaller is
the magnitude of no, the softer will be the absorption of shocks. Given no and us-
ing eq.(121), it is possible to calculate the necessary air pressure
Po
noPam. park.
1.2 FP
(122)
The principle of calculations involved in oleo-pneumatic shock absorbers was
first given by Prof. M.M.Shishmarev in 1930.
Section 49. Incorrect Filling of Shock Absorbers
The shock absorbers of an aircraft will work properly and will perform their
functions only if they are correctly filled. To charge the strut correctly, is to
fill it with the necessary quantity of fluid, so that the volume of air Vo will be
in accordance with the calculations, and the initial air pressure po will corre-
spond to the coefficient of preliminary compression no used in the calculations.
Let us analyze the functioning of shock absorbers when they are charged in-
correctly, i.e., in cases when there is less or more fluid in the strut than neces-
sary, and the initial air pressure does not correspond to the design pressure. The
entire analysis will be performed only qualitatively, taking note of the positE4T
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?11,
and negative sides of shock absorption during incorrect charging. Moreover we will
specify that the struts perform the work required by strength standards. Also, the
maximum stroke of the shock absorber s and the maximum stress in the shock
. strut Pam max must not exceed the magnitudes acceptable in a normally filled shock
absorber.
In order to simplify our reasonings, we will disregard the friction of collar
rings and bearing axles; this will not alter the qualitative result.
First Case
The amount of liquid in the strut is less than that required, and tha
initial air pressure po corresponds to the design pressure:
Obviously the volume of air Vo will be greater than required for normal work
of the strut.
am
0
....- ---- Fr /1
/ /
,.." / /
..." ,
Al
mcx
110..
Fig.250 - Stress Diagram for the Strut When Filled
With Less Fluid (by Volume) Than Required
In this case, the curves have a steeper slope. Shock absorption be-
comes softer but, at the given smax and Pam max, is not able to
absorb the required work.
a) Air pressure and forward stroke of the fluid at normal filling;
b) Forward stroke of fluid; c) Ain-pressure curve
Due to the increase in air volume, the curve representing its pressure
99
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(Fig.250 - solid curve) will be more sloping than in the case of normal charging.
Therefore, the stresses in the strut will increase more smoothly and the shock ab-
sorption will be softer. As a result of the limited stroke of smax (abscissa OM)
the maximum stress in the strut MN will be smaller than the stress MP, which is
equal to Pam and at the same time the strut will absorb less work (in magni-
tude). It is necessary to add to this that, due to the incorrect charging of the
strut, the wheel is less deflated and, therefore, also absorbs less work. This
smaller deflation of the wheel is explained by the fact that it is impossible to
apply forces to it that would be near Pw in magnitude, since these forces will
cause stresses in the strut, which will be close to Pam , which is not permis-
sible due to the limitedness of the stroke s .
As a result, when the -volume of air is increased, a shock absorber at a given
stroke s will not be able to perform the work required of it by the strength
max
specifications. The necessary work may be performed during the stroke s, which is
greater than s , which in turn may cause the breakdown of a shock absorber as the
strut would strike the stop.
Moreover, at certain instants of the piston stroke, the valve may happen to be
above the liquid level, which also creates abnormal working conditions of the strut.
Second (Opposite) Case
Shock absorber is filled with a greater amount of fluid than required,
and the initial air pressure 130 is equal to the design pressure:
In this case, the initial volume of air Vo is smaller than required, which
causes the air,-pressure curve to rise more steeply (Fig.251). This condition will
make the shock absorption more rigid since the stresses in the strut will grow
faster than when the strut is normally charged. Since the strength of the strut is
limited (by the permissible magnitude of Pam ), we will not be able to utilize
all of the permissible stroke smax. And, finally, the shock strut will absorb less
work. For this reason, we again will not satisfy strength specifications. STAT
100
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The total work will be absorbed when the values of Pam are greater than Pam max
which may lead to a breakdown of different parts of the aircraft.
Third Case
The volumes of both liquid and air correspond to the normal filling of
the strut, but the initial pressure of the air pc) is less than the
design pressure:
A decreased value of pc) lowers the magnitude of the stress on the strut, which
corresponds to the preliminary compression. Therefore, the shock strut, during
landing, will start contracting sooner - at smaller stresses (Fig.252), and the
curve representing the air pressure will slope more.
Fig.251 - Stress Diagram for a Strut
Filled with More Liquid Than Required
When the slope of the curve is
steeper, the stresses in the strut
increase more smoothly, which re-
sults in softer shock absorption.
In addition to this positive factor,
the limited stroke smax will result
in the fact that the greatest stress
in the strut will be smaller than
Pam max'
as in the first case (see
Fig.250), and will absorb less work.
Here, just as in the first
In this case, the curves are steeper. case, itis not permissible to
Shock absorption becomes more rigid and,
at the given smax and Pam max, is not able allow the tire to deflate corn-
to absorb the required work.
pletely; for this reason, the wheel
a) Forward stroke of fluid; b) Air pres-
will also absorb less work. As a
sure curve; c) Air pressure and forward
result shock absorbers (tire and
stroke of the fluid at normal filling
shock strut) are unable to absorb
the work required by the strength specifications. In a rough landing, a shock 1STAT
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occur at the end of the stroke of the strut, accompanied by a sharp increase in the
stress Pam.
Fourth Case
Volumes of liquid and air correspond to the required values, but the
initial air pressure p0 is greater than that corresponding to normal
filling:
Due to the greater value of pc), an increase in the preliminary compression of
the strut will occur. This will result in the fact that, during landing of the air-
craft, the shock strut will begin to contract at greater stresses (Fig.253), and the
air-pressure curve will be steeper.
Fig.252 - Diagram of Strut Performance When Charged with
Initial Air Pressure pc, Smaller Than Required
In this case, the curve has a steeper slope. Shock absorption becomes
softer but, at the given smax and Pam max, is not able to absorb the
required work.
a) Curves for air and forward stroke of fluid at normal filling;
b) Forward stroke of fluid; c) Air-pressure curve
As a result, the shock absorption will be more rigid since the stresses in the
strut will grow much more intensely during the stroke. Moreover, due to the limited
strength of the landing gear (limited by the permissible magnitude of Pam max), i STAT
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is impossible to utilize the permissible stroke s This latter fact will cause
the strut to absorb a diminished amount of work. As a result, the shock absorption
will not only be much more rigid but will also not satisfy the strength specifics-
tions with respect to the
c.?
awl
If
possibility of absorbing work.
Smax
O.+
Fig.253 - Diagram of Strut Performance
When Charged with an Initial Air Pres-
sure pc. Greater Than Required.
In this case, the curves lie higher.
Shock absorption becomes more rigid and,
at the given smax and Pam max' is not
able to absorb-Che requEied-imount of
work.
a) Forward stroke of fluid; b) Ain-pressure
curve; c) Curves for air pressure and for-
ward stroke of fluid at normal filling
On the basis of the four cases
we have just examined, we can con-
clude that shock absorbers must be
filled in accordance with the design
data a since only in such a case they
will ensure the design value of
maximum overload, when they are ab-
sorbing the greates work possible at
a given stroke smax.
Section 50. Shock Strut Design
Shock absorbers of the old
design (see Fig.242) made a long
piston stroke impossible, so that
the shock absorption was rigid (co-
efficient nE was large). The reason
for this was that, in order to make
the stroke of the piston longer, the
length of the strut had to be in-
creased; however, this length was quite great since the fluid occupied a consider-
able portion of the cylinder (vertically), and the diaphragm was placed inside the
cylinder.
Modern struts have no special volume for the fluid inside the cylinder. The
fluid is placed in the inner pocket of the piston rod. This change in structure
makes it possible to increase the stroke of the piston without changing the lenff
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Fig.254 - Shock Strut of the First Type
The characteristic feature of this strut
is that the packing is mounted on the
piston rod and the return valve is placed
on the bottom part of the piston.
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1C4
of the strut itself, to decrease the
initial pressure in the strut, and to
make shock absorption more elastic.
In such struts, a diaphragm is placed
inside the piston rad.
Shock Strut of the First Type
The strut consists (Fig.254) of
a steel cylinder (1), in which a
piston rad (2) is mounted so that it
can move freely inside the cylinder.
Sealing between the piston and cylin-
der is achieved by leather packing
collars (5). These are mounted in
the upper part of the piston rod and
are tightened by a bearing axle nut
(6). The packing gland (11) is
mounted in the control nut (10) of
the cylinder and is tied with a
stuffing gland nut (12). The return
stroke of the piston rod is limited
by the shoulder of the rod (7) and
the control nut (10), together with
a bronze journal box (9) and a shock
absorption ring (8). Inside the
cylinder a steel tube (3) (plunger)
is mounted, which forms the body of
the diaphragm. The lower end of this
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tube is lowered and positioned at the beginning of the piston pocket. On this same
end of the pipe, a slide valve (4) is mounted. This valve has one 8mm orifice run-
ning through its axis and four slanting side orifices of 5 mm each, as in the valve
shown in Fig.246. In the upper portion of the cylinder there is a filler pipe (15).
A steel ring (16) for fastening the wheel axle (17) is fixed on the lower end of the
rod.
The shock strut is filled with fluid and air through a filler pipe (15). The
liquid usually is an antifreeze mixture of glycerol and ethyl alcohol. The liquid
is placed in the upper portion of the pocket (13) of the piston rod (2), The level
of the mixture must be above the valve (4). Air occupies the volume (14).
The strut functions in the following manner: During the forward stroke, liquid
flows from the pocket (13) of the piston into the pocket (14) of the cylinder.
Liquid flows through the central and the four slanting orifices of the valve (the
valve will be lifted) and through the clearance in the collar (18), between the body
of the plunger (3) and the piston (2). During the return stroke, the valve lowers
into its position and the four slanting orifices are closed (see Fig.246). In this
case, the mixture will flow only through the central orifice of the valve and the
play of the collar. This results in braking of the liquid.
The strut has a limited possibility of controlling the velocity of the piston
return during the return stroke. The speed of the return stroke can be lowered by
making the orifices for the liquid flow smaller. But in reality, even at zero open-
ing of the valve, the strut will still be extended since the boost pressure of the
air in the cylinder will always create a force, which w4.1 be greater than the suc-
tion of the vacuum under the plunger.
Shock Strut of the Second Type
In a shock absorber of this type (Fig.255) the flow of liquid is also braked
mainly during the return stroke. The strut consists of a steel cylinder (1), in-
side of which, in the lower portion, a piston rod (2) is moving. Between the STAT
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ak) 44
b)
d)
1/
2/
22
/3
/0
/2
/3
/4
20
7
8
Fig.255 - Shock Absorber of the Second Type
?
It is characterized by the fact that the
packing is mounted on the cylinder and the
reverse valve is placed inside the piston.
a) Direction of piston stroke; b) Forward
stroke (compression); c) Return stroke
(expansion); d) Cross section through A-A
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cylinder and the piston, a permanent
packing is mounted at the lower por-
tion of the cylinder. It consists
of leather packing rings (3) and an
oil felt pad (8), tied with a nut
(7). The return stroke of the pis-
ton is limited by a stop bushing
(4), fixed on the rod, and a thrust
ring (5). Inside the cylinder, in
its upper portion, a plunger (9) is
mounted, whose lower end is lowered
into the pocket of the piston.
A nut (10) is fixed on the low-
er end of the plunger. It has a
central hole (11). Inside the plun-
ger a liquid-level control (15), is
fixed which is connected with a plug
(17). The technician can remove the
extra liquid from the strut with the
help of this pipe. On the upper end
of the piston rod, two permanent
bronze rings (12) and (14) are
mounted. These rings have a large
number of vertical holes around
their circumferences. In between
these rings, a movable ring valve
(13) is mounted, which has two holes
and a ring groove. Between the rod
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and the cylinder there is a pocket (20). This pocket is filled with the fluid.
During the forward stroke, an additional amount of the fluid enters this pocket
through the holes of the bronze rings (12) and (14) and through the radial gap be-
tween the valve (13) and the cylinder. The function of bearing axle of the piston
rod is taken over by the bronze rings (12) and (14) and by the bronze spacer (6).
The inner pocket of the rod (21), in its upper portion, into which the plunger en-
ters, is fashioned in the form of a cone. This changes the size of the orifice
through which the liquid flOws. A ring (18) for fixing the wheel axle (19) is weld-
ed to the lower end of the rod.
The shock absorber is charged with a liquid consisting of a mixture of glycerol
and ethyl alcohol through the filler plug (16). Charging of the strut with air is
.performed from the general inflation valve through the filler plug (16).
The strut operates in the following manner: During the forward stroke, liquid
flaws from the piston pocket (21) into the pobkets (20) and (22) through the orifice
(11) of the plunger, holes (12) and (14) in the bronze rings of the piston, and
through the annular clearance between plunger and rod, and between the floating ring
valve (13) and the cylinder. During the return stroke, the floating ring valve (13)
is pressed to the ring (12) and closes all its holes, except two. Then the mixture
flows slowly from the pocket (20) through these two holes in the rings (13) and (12)
into the pocket (22), from there through the orifice in the plunger (11) and through
the play in the ring between the plunger and the piston, into the piston pocket
(21).
The peculiar feature of this strut is that the piston area, on which the magni-
tude of the air pressure depends and which balances the external force of the strut,
is defined by the outside diameter of the rod. This makes it possible to increase
the piston stroke. In this case, there is a greater possibility to regulate the
Speed of the piston return during the return stroke by decreasing the cross section
of the holes in the valve (13). The level of the liquid during the entire period of
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the strut performance must be such that the ring (12) is always submerged in the
liquid. This prevents air from reaching the pocket (20) and avoids hydraulic shocks
during the return stroke of the strut.
Fig.256 - Shock Strut of the Third Type
It is characterized by the fact that the packing is mounted on the
cylinder and the reverse valve is placed on the lower portion of
the pin (plunger).
a) Forward stroke; b) Return stroke; c) Direction of piston stroke
Shock Strut of the Third Type
This strut (Fig.256) is a combination of the first two types and consists of a
steel cylinder (1) and a piston rod (2). The packing of both cylinder and rod is
permanent and consists of leather packing collars (3) and a tie nut (4), mounted to
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the lower portion of the cylinder.
The piston (9) is mounted to the upper end of the rod (2). It has an orifice
for free passage of the mixture. The upper pocket of the piston (10) forms a reser-
voir for the mixture. Inside the cylinder, a stationary pin (6) is mounted, whose
lower end carries a diaphragm with an orifice (7), together with a slide valve (8)
with two orifices. The cylinder has a filler plug (11).
A welded clevis with an axle for attachment of the wheel is mounted to the
lower end of the rod.
The strut operates in the following manner: During the forward stroke, the
mixture flows from the pocket (10) through the diaphragm orifice (7) and annular gap
between the valve (8) and the rod (2) into the pocket, which is filled by compressed
air. Noreover? the mixture from the cylinder will reach the circular pocket (5) be-
tween the cylinder and the rod through the hole in the piston (9). During the re-
turn stroke; the valve (8) lowers and closes all openings in the diaphragm, except
two. This achieves the braking effect.
The peculiar feature of this strut is that it possesses an additional possi-
bility to increase the piston stroke by lowering the level of the liquid, since here
it is not necessary that the piston (9) be covered by liquid.all the time. There
may be air in the space behind the piston (5), without causing any hydraulic shocks.
The same cannot be said on struts of the second type. But this strut, just as the
strut of the first type, does not permit complete control of the return speed of the
piston.
Conclusions. aperimental tests have shown that the best strut of the three
types examined above is the second one. A fluid of sufficiently high viscosity must
be used. It must also withstand cold weather during winter and low temperatures at
high altitudes (must be an antifreeze). A mixture of glycerol and alcohol is most
frequently employed. The amount of the liquid is controlled by the liquid-level
control pipe, by the filler pipe, or by a calibrating vessel. Ain-pressure contrsiwr
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is achieved either with the help of a pressure gage or by the position of the piston
rod while the aircraft is parked.
.
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CHAPTER XIII
PERFORMANCE OF A LANDING GEAR ENERGY DIAGRAM
The landing gear performs the coupling between the wheel and the aircraft. The
external loads acting on it are determined on the basis of applicable strength
specifications.
Fig.257 - Cantilever Landing Gear
1 - Shock strut; 2 - Wheel; 3 - Mechanism for retraction of
the landing gear
On modern aircraft, three basic types of landing gears are used: STAT
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a) Cantilever (Fig.257);
b) Senicantilever, with one or two inclined struts (Fig.258);
c) Lever suspension of wheels (Fig.259).
Fig.258 - Semicantilever Landing Gear
a - With one inclined strut: 1 - shock strut; 2 - landing gear
retraction mechanism; 3 - articulated cantilever.
b - With two cantilevers.
Section 51. External Loads Acting on a Landing Gear
The basic external loads exerted on the landing gear are the reaction forces of
the ground, generated during landing and taxiing of the aircraft. The ground effect
(Fig.260) completely balances the weight and inertia force of the aircraft mass
s) p
W ?1 4) W
where P is the load on the main wheel;
mw
is the load on the nose wheel;
n w
G is the aircraft weight;
112
(123)
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N is the inertia force of the aircraft mass.
Y = 0.75 G which is the lift of the aircraft during touchdown is equal to about
0.75 of its flying weight.
cr
Fig.259 - Landing Gear With Lever Suspension of the Wheels
a - Shock absorber inside the strut; b - Shock
absorber carried outside of the strut
a) Strut; b) Shock absorber; c) Wheel; d) Fork arm
Equation (123) may be rewritten in another form:
9/)
m w 1-Pnw
(124)
where n is the overload coefficient.
The magnitude of the ground effect and the line of its action may differ de-
pending on the quality of the landing, the character of the ground run, the rough-
ness of the runway, and the rigidity of its shock absorption. Ground-reaction
forces may be distributed throughout the elements of the landing gear in different
ways: on three points, i.e., on the main wheels and the nose wheel; on two poiSTAT
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and even on one point, which also depends on the character of the touchdown and the
movement of the aircraft on the ground.
b)
1c)
Nd)
Fig.260 - Loading of the Landing Gear
at Touchdown
The forces of ground reaction (external
loading of the landing gear) balance
the weight and inertia forces of the
aircraft mass.
a) Lift; b) Load on nose wheel, Pn w;
c) Load on main wheel P ; d) Inertia
mw
Fig.261 - Loading of the Landing Gear
on a Rough Airfield
The forces of the ground effect will
act at an angle to the horizontal,
i.e., vertical and horizontal compo-
nents of forces will act on the landing
gear.
In case of an aircraft landing on a smooth runway without the use of brakes,
the forces of the ground reaction will in the main act in a vertical direction.
During touchdown or taxiing on a rough runway, the ground effect will act at an
angle to the horizontal, i.e., the landing gear will be acted upon by the vertical
and horizontal components of forces (Fig.261).
Horizontal components of the reaction forces of the ground will also act during
the movement of the aircraft over a smooth runway and in cases of brake application
to the wheels (Fig.262).
When an aircraft makes ground contact with its wheels not locked, horizontal
forces are generated which include forces of friction of the wheels on the ground,
which tend to rotate the wheels. It must be noted that, at high touchdown veloci-
ties, these considerable friction forces cause rapid wear of the tires and greatly
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load the landing gear. For this reason, in some cases a special device is installed
for preliminary untwisting of the wheels before touchdown of the aircraft.
Fig.262 - Loading of the Landing Gear
When Brakes Are Applied
The horizontal components of the ground
reaction will act during the movement
of the aircraft over a smooth runway if
the brakes are applied.
Fig.263 - Loading of the Landing Gear
During a Touchdown With Drift
In this case, the landing gear will be
acted upon by forces of friction
directed perpendicularly to the plane
of symmetry of the aircraft.
a) Nn = Force of inertia; b) Force of
friction
During touchdown with a drift or in case of a"sharp turn during the ground run,
the aircraft will be subject to the action of horizontal components of the ground
reaction acting in a lateral direction (Fig.263). The outside wheel will always be
loaded more than the inner one, since its normal force is greater, and since the
horizontal components are the product of the normal force times the coefficient of
friction.
Normal forces are greater at the outside wheel since, under the action of
lateral forces, the aircraft will crab.
The landing gear must be strong enough to withstand all possible cases of load-
ing during its operation. The existing aircraft strength specification list a num-
ber of typical situations. The landing gear have sufficient mechanical strength to
resist these forces. The external forces are determined according to the existing
strength specifications.
Rough Touchdown
As a typical example, let us examine a case of a rough touchdown (El.g.).
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At the instant of touchdown (Fig.2(4),
possesses two velocity components:
Vhor is the horizontal component;
V is the vertical component;
while Vhor is a little greater than Vy.
Vhor
I -------
V
`11^11111to
A"
Fig.264 - Design Case El.g. for Rough
Landing
During a rough landing of an aircraft,
its vertical velocity component Vy must
be canceled during contraction of the
strut and deflation of the tire. There-
fore, because of the short path, large
accelerations must be in operation,
large forces of inertia, which means
large landing-gear reactions.
corresponds to greater landing-gear reactions
loads.
The overload coefficient will equal to
yr, w
11 e -
PM NI
previously mentioned, an aircraft
After touchdown, the aircraft
continues to run on the ground,
slowing down gradually and with
slight negative acceleration. Hori-
zontal inertia forces, also of a
small magnitude, correspond to
these.
But the vertical velocity com-
ponent V is damped within a very
short period of time (fraction of a
second). Therefore, here we have
large acceleration and inertia
forces of high magnitude, which
(P and Pn w) and to greater over-
Where P is the load on the main wheel during touchdown;
mw
0.42G is the load on the main wheel during parking of the aircraft.
For modern aircraft, nE may reach 2.6 - 3.5 and may even be larger.
(125)
At touchdown, the magnitude of Pm w depends on the magnitude of the accelera-
tion. Acceleration, in turn, depends on the stiffness of shock absorption. 1--riocr
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Strength specifications do not give the magnitude of ng, but the degree of roughness
of touchdown, i.e., they determine the vertical velocity components (Vi) of the air-
craft generated during a rough touchdown. In this manner, the kinetic energy A is
Igiven, which is determined by the well-known formula
A
GVv
2g '
where g = 9.81 m/sec2 is the acceleration of gravity.
(126)
The magnitude of the vertical velocity of the aircraft V for a rough touchdown
is assigned according to the weight of the aircraft and the magnitude of the landing
speed (Vy. is equal to about 2 - 4 misec).
In modern aircrafts, in order to obtain V = 3 m/sec, it is necessary to
approach the ground at a small angle, equal to about 5?. This will be a rough land-
ing.
S.N.Shishkin has published the results of experimental work on actual loading
of the landing gear in 1936. This work later served as a basis for planning the
strength.
Section 52. Cantilever Landing Gear ?
The landing gear shown in Fig.257, strictly speaking, is not a beam fixed at
one end, since its strut (Fig.265) has two supports. However, since the distance
between the two supports is very small, and in order to simplify the stress
analysis, we assume that the strut is of the cantilever type whose end is fixed
while an external force P is applied to the other.
Let us take a concrete case, for instance, a case of rough landing El.g. Here
the external load (design), which acts on the main wheel, is equal to
a.(14: f, (127)
where d 0.42 is the coefficient showing which portion of the aircraft weight G
STAT
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acts on one main wheel in parking; this coefficient is obtained as a
result of the resolution of gravitational forces of the aircraft
according to the lever law between main wheels and nose wheel;
f = 1.65 is the safety coefficient;
Ne = 2.6 - 3.5
is the overload coefficient; the exact value of n; depends on the
rigidity of shock absorption.
Fig.265 - Diagram of a Cantilever
Landing Gear
The landing gear represents a beam,
one end of which is fixed, while the
external force is applied to the
other end by means of the wheel.
(Fig.266) the axle of the wheel will bend as
ing moment will be at the point where the wheel is fixed to the axle, namely the
strut
The wheel, under the action of the
external load PE, is deformed and
transfers the load to the axle over two
roller bearings. We are not interested
here in the strength of the wheel since
the wheels are standard and their
strength is guaranteed by the manufac-
turer which produces them.
Let us resolve the force PE into
its components (Fig.265): One of them,
Pl, will act in the plane of the axle
and the strut, while the other, P2, is
perpendicular to this plane.
Action of the Force P1
Under the action of the force P1
a cantilever beam. The greatest bend-
Mb../3 a
1
where a is the lever arm of the force.
Tina fnroa P, an
the bending moment Mb = Pia are transferred from the axle STAT
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the struts so that in all the cross sections of the strut the bending moment Pia
will have a constant value. The diagram of Mb is shown in Fig.267. Through the
length e of the bending moments, Mb = Pia will be absorbed by the rod and cylinder.
If we draw the diagonal CD, we will divide the diagram of Mb into parts belonging
to the rod (the lower one) and the cylinder (the upper portion).
The force P1 will compress the rod of the strut with a force of
=
[OM
1.rd
where Frod is the cross-sectional area of the rod.
Fig.266 - Bending of the Axle
Under the action of the force Pl, the
wheel axle bends with a maximum-bend-
ing moment at the cross section near
the strut.
a) Wheel axle; b) Strut
Fig.267 - Bending of the Strut
Over the entire length of the strut, the
bending moment created by the force Pi
will be of the same magnitude. Through
the length e, this moment will be absorb-
ed by the rod and the cylinder simul-
taneously. The diagonal CD divides the
diagram of the bending moments into two
parts; one, referring to the rod, and
the other to the cylinder.
a) Cylinder; b) Piston
From the rod, this forte is transmitted to the mixture, and from the mixture to
the air. Air transmits this force to 'the cylinder bottom, from where the load'.STAT
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Fig.268 - Distribution of
Normal Stresses through
the Cross Section Under
the Xction of the Bending
Moment
Large stresses d occur in
the fibers which are far
from the neutral layer and
are small near the neutral
layer
..111
(6)
Fig.269 - Actual and Approximate Distribution
of Stresses in a Tubular Girder
When determining the stresses d in the tubu-
lar girder we consider it approximately to be
a double-belt beam with a height U=
Fig.270 - Loading of the
Upper Support
If we imagine the strut separated
from its upper support, it will
rotate under the action of the
moment Pia. Reaction in the up-
per support R acting on the lever
arm b hinders this rotation
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120
a) Neutral layer
2
3
a)
Fig.271 - Work Performed by the
Landing Gear under the Component
Force P2
Under the action of the force P2
the axle is bent, and the strut
is both bent and twisted
a) Torque diagram; b) Diagram of
strut bending moment; c) Diagram
of axle bending moment
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transmitted to the strut joint.
Due to the high pressure of the mixture and air, the cylinder will operate as a
vessel, subject to an internal pressure.
Under the action of Mb the strut is bent, and the individual fibers of the
strut and cylinder will be subject to normal pressures 0 (Fig.268).
The true distribution of the bending stresses through the cross section of the
pipe is described in Fig.269a. In determining the stresses 0, we approximately make
use of the law described in Fig.269b. By this we determine the stresses 0 in the
tube as if it was a double-ring beam with a height H and with an area of each ring
equal to F.
where we approximately assume that
and
Mb
HFP
a
F = 2tDd.
3
(128)
Reaction in the upper support (Fig.270) is determined from the equality of the
moments
P1 a = Rb,
from where
P1 a
R - (129)
Action of the Force P2
Under the action of this force (Fig.271), the axle is bent in the plane pe?,-.IAT-
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dicular to the strut and passing through the axle. Its greatest bending moment will
occur at the place where the axle is attached to the strut
Mb = P2 8"
For the strut, this moment is the torque Mt, which retains a constant value
through the entire length of the strut
Mt = P2 a = const.
(130)
Moreover, other bending moments caused by the same force P2 will act in the
cross sections of the strut. The Mb of the greatest magnitude will be found in the
joint
Mb = P22, '
Under the action of Mb = P2a in the wheel axle and of Mb = P21, normal stresses
will occur in the strut. The magnitude of these stresses is determined by the same
method as those caused by the force P1, i.e.,
Mb
6 = ? (131)
HF
Under the action of the torque Mt, lateral stresses will occur in the rod and
cylinder, whose magnitude is determined by the formula
Mt
T =
2F
(132)
n2
where F = II.' is the cross-sectional area of the tube circuit.
This eq.(132) is valid only for those sections of the strut length where there
is neither a two-piece cross beam nor a regular cross beam (Fig.272). In those
places where the cross beam is mounted, the torque will be transmitted by the bend-
ing of individual elements of the landing gear, in which normal stresses will occur.
STAT
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Cross-Beam Performance
If we would separate the cross beam at the joint k (Fig.272), a mechanism will
result, i.e., the rod will rotate inside the cylinder under the action of the torque
(see the lower position of the cross beam shown in dotted lines). From this, we can
Fig.272 - Work of a Two-Piece Cross Beam
When the Strut is Twisted
If we imagine the cross beam separated
at the hinge k, a mechanism will be
formed, which will be balanced on
application of forces of interaction T
of the two halves of the cross beam.
A)
Fig.273 - Bending of the Two-Piece
Cross Beam and Shear of the Connecting
Bat
The force T is transmitted to the rod
(cylinder). The two-piece cross beam
is bent and shearing of the connecting
bolt occurs.
a) Connecting bolt; b) Fulcrum;
c) Piston
conclude that the two-piece cross beam participates in the work of the strut under
the action of Mt = P2a. In practice it can be assumed that both halves of the cross
beam are tied together by a hinge. In this case, they can exert a mutual influence
only by means of a force since it is impossible to transmit a moment through a
hinge.
In order that the lower half of the cross beam does not move with respect to
the upper one, an interacting force T with a lever arm d, with respect to the axle
of the strut, must be in operation.
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The magnitude of the force T is determined from the condition of absence of
rotation
from which it follows that
Mt
(133)
Under the action of the force T, each half of the two-piece cross beam is bent
like a cantilever beam, which is fixed to the rod (Fig.273) or to the cylinder.
Fig.274 - Work of the Rod on the Fulcrum Section
The rod is bent under the action of the torque Mt
a) Shear forces - forces of interaction in rod; b) Rod diagrams
From the action of Mb normal stresses will occur in the rings of the two-piece cross
beam.
M
CT -=
IIF,
124
(134)
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where H is the distance between cross-beam rings;
Ft is the cross-sectional area of a cross-beam ring.
The maximum bending moment of the cross beam
Mb =TL
is transmitted to the rod or cylinder by shearing of the bolt. The shearing forces
R are determined from the equality of moments
from where
IL Rh,
(135)
The lateral stresses from shearing of the bolt are determined by dividing the
shearing force R by the cross-sectional area of the bolt Fb
I?
Tnwms?F
(136)
Let us examine the work of the rod on the fulcrum section. Let the rod (Fig.
274) intersect with a plane perpendicular to its axis and passing through the joint
of the cross beam k. Then, under the action of the force T, applied to the lower
portion of the two-piece cross beam, the cut portion of the rod will tend to move in
the direction of action of the force T. In reality there will be no such motion
since the cut-off portion must be balanced. Therefore, the acting force T will be
balanced by the internal force of the elasticity of the rod, which is located in the
plane of the cross section and causes shear. In this manner, the cross section of
the above-described rod will also be subject to interacting forces equal to T, under
whose action the rod and cylinder in the section under the two-piece cross beam will
be bent (see the diagram of Mb).
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Here the calculation of loads is performed separately for those caused by th-rAT
As a result of the analysis it can be concluded that the torque M = P2a at the
fulcrum section is transmitted by the bending of the rod, cylinder, and the fulcrum
itself.
Fig.275 - Loading of the Wheel During
a Front Impact
Here, unlike during a rough landing,
the component force P2 is directed
? backward.
a) Strut; b) Wheel; c) Force of front
impact
Fig.276 - Case of a Lateral Loading of the
Landing Gear
In this case, the wheel is loaded with two
forces: A force normal to the ground PE,
and a friction force PF, which lies in the
plane of the ground.
For a final evaluation of the strength of the landing-gear elements it is
necessary to sum up the similar stresses in the dangerous points of an element under
discussion, which are caused by simultaneously acting forces.
In exactly the same way must the analysis of the function of landing-gear
elements in other design cases be performed, if they differ from the just discussed
case only in the magnitude and direction of action of the external force. For in-
stance, in case of the wheel striking an obstacle during forward motion (Fig.275),
the component force P2 will be directed toward the rear.
Case of Lateral Loading of the Landing Gear
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force PE and for those caused by the action of the lateral force PF (Fig.276). The
action of the force PE has been discussed above. But the lateral force PF will bend
the wheel axle with a moment uniform throughout the entire length of the axle
Al .1--)Fr.
Under the action of the force P the entire strut will be subject to torque with
a constant moment throughout the entire length of the strut
Mt Ppc.
Moreover, the strut will be bent by the force PF. This bending moment will
vary throughout the length of the strut (see diagram of Mb).
Section 53. Semicantilever Landing Gear
Semicantilever landing gears (see Fig.258), if the cantilevers are removed, can
have one or two degrees of freedom. Two degrees of freedom are created by a
Cardanic joint of the landing gear.
Cantilevers are used to eliminate these degrees of freedom during touchdown and
moving of the aircraft on the ground, and to make the landing gear system more
rigid. To ensure retraction of the landing gear during flight, the cantilevers are
of the collapsible type.
Landing gear diagrams are designed with one cantilever if, during retraction,
the gear moves along the wing chord or along the wing span. If the gear moves at an
?
angle, i.e., simultaneously along the span of the wing and along its chord, two
cantilevers are used. In this case the strut can rotate around any of the two
axes.
gear.
Let us examine the work performed by the elements of a semicantilever landing
Let us take the case of a rough landing (Fig.277).
The external load does not depend on the diagram of the landing gear andsTkil
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be determined in the same way as in
a cantilever-type landing gear.
The axle and the lower portion of
Fig.277 - Performance of a Landing Gear
of the Semicantilever Type
The cantilever causes elongation of the
portion of the cylinder above the
cantilever nodal point
a) Cantilever; b) Cylinder; c) Piston;
d) Wheel
the strut, up to the cantilever nodal point,
will work in the same manner as they do
in a cantilever diagram. Only the port-
tion of the cylinder which is above the
cantilever nodal point will work in a
different manner since it will be sub-
ject to the stress of that cantilever.
In order to analyze the work of
the landing gear, let us substitute the
cantilever by the stress Scant acting
in it.
The design of the upper portion of
the cylinder is performed in two
stages:
First stage - the calculation of
action of the force PE is performed,
the landing gear being assumed as a
cantilever type;
Second stage - the action of the force acting through the cantilever is calcu-
lated. To do this, it is necessary to determine the stress in the cantilever.
Determination of Stress in the Cantilever
Since the cantilever is joined by a hinge, it causes only axial stresses. If
we imagine the cantilever to be dissected, the rigid diagram of the landing gear
will be turned into a mechanism, and the strut will rotate about the upper node
under the action of the moment M = Pia. In reality, there is no such rotation
since it is prevented by the cantilever. Therefore, the cantilever will be loaded
STAT
with a compression stress. The magnitude of this stress is determined from the
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STAT
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equality of moments (see Fig.277)
from where
P1 a = S cam r,
Scamtr ?
(137)
If there are two cantilevers, the order of determination of their stresses
remains the same:
it is necessary to write two equations of moments with respect to
the two axes of rotation of the strut.
Fig.278 ? Influence of a Cantilever On
Bending of the Cylinder
Force Pl bends the cylinder in one
direction, and the stress Scant in
another. Because of this, the upper
portion of the cylinder is unloaded.
to the liquid, and the latter will pass
force P1 to the nodal point of the strut
external pressure of air, the cylinder tends to burst.
In our example, the cantilever has no influence on the work of the strut due
to the force P2 (see Fig.275), since it lies in the plane of the strut and axle.
When the stress in the cantilever
is found, we can determine the internal
strain
a ScArd
?
Ficant
(138)
The upper part of the cylinder
will be unloaded of the action of the
force P1 (Fig.278, diagram Mb) due to
the cantilever.
The work of the strut, over its
entire length, due to the action of the
force P1 will be similar to the work of
a cantilever system; the rod being
compressed will transmit the force P1
it to the air; the air transmits the
through the cylinder bottom. Due to the
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But if the cantilever is carried beyond this plane, component forces are generated
due to stresses in the cantilever, which will be perpendicular to the plane ofthe
strut and axle; this may change the performance of the upper portion of the cylinder:
Since the cantilever is not at a right angle to the axis of the strut but at an
acute angle it to it, the cylinder will be loaded with an additional axial stress due
to the action of the component force Scant cos it. In our case, the cylinder in its
upper portion is also stretched.
Section 54. Landing Gear with a Lever Suspension of Wheels
The landing gear of jet aircraft has its own features. These are due primarily
to the absence of a propeller. The absence of a propeller makes it possible to use
shorter gear, i.e., to lower the aircraft and make it more stable during its run
over the runway.
The landing gear of jet aircraft generally uses a lever suspension of wheels
. (Fig.279). In this diagram, the shock struts are in a more favorable position
since, in this landing gear system, they are subject to compression only. This fact
eliminates possible detachment of the rod and piston from the cylinder and leakage
of the fluid; it allows raising the initial pressure in the strut and, therefore,
decreasing the overall strut dimensions. In the case of a landing gear with a lever
suspension of wheels, the shocks experienced by the latter due to the horizontal
components of forces become softer, i.e., the shock strut works not only from the
vertical but also from the horizontal forces acting on the wheel.
Retraction of the gear in high?speed aircraft, as already mentioned, becomes
mor
11411
' ficult due to the relatively small thickness of the wings. For this reason,
the la g gear is very often retracted inside the fuselage. It must be noted
that the fuselage also does not have much available space for retraction of the
gear. This explains why high?pressure wheels of small overall size are used in jet
aircraft. Such wheels assist a shock only slightly in softening shocks at touchdown
and, most important, impair the aircraft roadability. As a result, airfields wiSTAT
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a more impermeable soil are required.
A landing gear with a lever suspension of wheels may be of either of the canti-
lever or of the semicantilever type, in analogy with the diagrams examined above.
The characteristic peculiarity of this landing gear is the suspension of the wheels
on a special lever, which can rotate with respect to the joint, mounted on the
strut-knee. The shock absorber can be designed as a self-contained element (1 - 2)
(Fig.279) or it can be mounted inside the strut column (Fig. 280). In the first
case, both the rod and cylinder of the shock absorber are subject to axial stresses,
without being bent. This is explained by the presence of the double joints in the
assembly (1) and (2) (see Fig.279), which ensure transmission of only axial stress-
es. In the second case (Fig.280), because of the double hinge in the assembly (1),
the rod of the shock absorber is not bent, but the cylinder, which is the strut
column, is bent. For this reason, the shock absorber is in a more difficult posi-
tion than in the diagram of Fig.279.
Landing Gear With Shock Absorber Outside the Strut
The diagram in question (see Fig.279) consists of a strut (column) (4 - 6)
which is joined to the assembly (6) with a hinge and is free to rotate with respect
to the axis x - x. This degree of freedom ensures retraction and extension of the
landing gear with the help of the wheel-retraction mechanism, attached by a bolt to
the assembly (7). At take-off and landing, the assembly (7) remains stationary.
A lever (4 - 3) is hinged to the lower portion of the strut at the assembly
(4). This lever can rotate only in its plane with respect to the assembly (4), in
which the bolt is mounted on a large base. The freedom of the lever rotation is
limited by the shock absorber (1 - 2), which is double-hinged at the assemblies (1)
and (2), in order to prevent bending of the shock absorber.
At the assembly point (3) a semiaxis (3 - 5) is rigidly fixed to the lever.
The wheel is mounted to this semi-axis. It must be noted that instead of a semiaxis,
a semifork or a fork-and-axle could be Mounted on the lever.
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Below, as an example, we will study the work of separate landing-gear elements
during a rough landing, i.e., from the force Pw.
?
Er-
bw C
b)
Fig.279 - A Landing Gear With Lever Lifting Mechanism of
Wheels - Shock Absorber Outside the Strut
(a) Lifting cylinder; (b) Strut (main); (c) Shock absorber; (d) Fork
arm; (e) Semiaxle; Mb from force Pl; (g) Mb from force P2;
(h) Mt from force P1
Work of the Semiaxis. The force of the ground reaction Pw is directly applied
to the wheel and is transmitted to the semia.xis through the roller bearings. In
this case, the semiaxis (3 - 5) (Fig.279a) is bent like a cantilever beam. The
dia-
grams of the bending moments change according to a linear law with the maximum value
of Pwcat the point (3).
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Work of the Lever. The moment Pwc, together with the force Pw, are transmitted
to the lever (3 - 4). For convenience in calculation, let us resolve the force Pw
(Fig.279b) into its components Pi and P2. The force P2 acts along the axis of the
lever, while the force Pi is normal to P2.
?
Sam
p ? \
w
A
,'
Fig.280 - Landing Gear With Lever Lifting Mechanism of
Wheels - Shock Absorber Inside the Strut
(a) Lifting cylinder; (b) Mb from the Rime P1
The force Pi applied to the assembly (3) bends the lever in its plane, like a
simple beam with a cantilever. The supports are: the strut (4 - 6) with a fixed
assembly (4) and the shock absorber (1 - 2) with a movable assembly (1). It is not
difficult to find the reactions in the supports 114 and S. For instance, the STAT
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stress in the shock absorber (1 - 2) can be determined from the fact that the lever
is hinged at the assembly point (4)
From this it follows that
Plb= Sama.
Sam = P2
a
The diagram of the hgding mommts Mb for the lever. caused by the actiof
the force F11 is represented by a triangle (Fig.279b). The greatest moment will
occur at the point of assembly (1) of the shock absorber bracing and will correspond
to the value
P1d
Since the point of assembly (1) is not on the axis of the lever (4 - 3), the
left portion of the lever is slightly less bent.
As the force P1 is applied to the point (5) of the semiaxis, it will twist the
lever with a constant moment
Mt = P1 c,
which is fully transmitted to the strut through the assembly (4).
The force P2 will cause compressive stresses in the lever, while being trans-
mitted to the assembly (4) of the strut. Moreover, because of the lever arm c, the
lever (3 - 4) will be bent in a plane perpendicular to the plane of the lever
(3 - 4 - 1). The assembly (4) remains, as before, the seal. The bending moments
for individual cross sections of the lever over its entire length will be constant
(Fig.279b) and they will correspond to the value
3
P2 c.
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Strut Performance.' The external loads acting on the strut (4 - 6 - 7) will be
the forces transmitted through the assembly (4), and the stresses of the shock
absorber Sam, transmitted through the assembly (2). Under the action of the loads
just mentioned the strut is bent and axial deformations occur in it.
Let us examine the bend in the plane of the lever and the shock absorber
(Fig.279c). Here the strut is loaded with the forces Rif and S. The strut is
attached to the aircraft by the unit (6).
Through the segment (4 - 2) the bending moments change according to a linear
?1?1111M11.w.
law. They are equal to zero at the point (4). In section (2 - 6) they are constant
and equal to
Pwi,
which is clearly indicated from the general-view diagram of the landing gear.
The strut is additionally bent in the perpendicular plane (Fig.279d), being
loaded at the assembly with a concentrated moment
= wc.
This moment is the geometric sum of the magnitudes of the moments of the lever:
Plc and P2c (Fig.279b). The value of Pw can be easily obtained from the general
view of the landing gear (first projection).
The strut is bent in the perpendicular plane like a simple beam with a large
cantilever (6 - 2 - 4). The supports will be found in the hinge (6) and in the
unit (7) of the retraction mechanism. Therefore, throughout the length (4 - 2 - 6)
the bending moments will be constant, and throughout the length (6 - 7) they will
diminish until they will equal zero at the assembly (7).
This indicates that individual elements of the landing gear are deformed in
different ways. For instance, the lever is bent in two planes, being twisted and
compressed. In order to have any judgment as to the strength of the structure, it
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AM.
is necessary to determine the stresses caused by individual kinds of deformation,
to sum them up in a corresponding manner and to compare them with those which the
structure is able to withstand.
Landing Gear With the Shock Absorber Inside the Strut
This landing-gear diagram has the same elements (Fig.280) as the structure in
Fig.279. Here the column (4 - 6) is free to rotate with respect to the axis z - z,
and the retraction mechanism is arranged somewhat differently. The lever
(4 - 1 - 3) is attached to the column by means of a bracket at the assembly (4).
The leveFis forked on the end (1 - 3), At the end of this fork, the axle (3 - 3)
-ma*.
of the wheel is mounted. The shock absorber is placed inside the strut, whose rod
is attached to the lever at the assembly (1).
Let us examine the performance of separate elements of the landing gear under
the action of the vertical force Pw. We will proceed as before.
Axle Performance. The force of the ground reaction Pw is transmitted to the
axle (3 - 3) over roller bearings. In our case, the fork will be the support of the
axle. For this reason, the axle will be bent like a simple beam (Fig.280a). The
largest magnitude of the bending moment will occur in the middle portion of the
axle and will be equal to
Pw
4
Lever Performance. Let us assume that the lever and its fork (4 - 1 - 3),
consist of a single element. Then let us resolve the force Pw (Fig.280b) into its
components P1 and P2. The force P2 will create compressive stresses in the lever,
while P1 will bend it. The maximum bending moment will, as before, be equal to
P1d
which will occur near the junction (1) of the shock-absorber rod.
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The reaction in the assembly (4) will equal to
R, = P
4. w a
Since the forces P1 and P2 act on the lever symmetrically, no twisting of the
lever occurs.
The stress in the shock absorber is determined by the previous formula
Sam = P
1 a
411,110s...
.11.1??????..
Strut Performance. Unlike in Fig.279, only those forces will form the external
load of the strut (4 - 6 - 7), which are transmitted from the lever (3 - 4) through
the assembly (4); this assembly is carried somewhat toward the front. In this case,
the only force (Fig.280b) will be
R4 = pw a
Under the action of this force, the strut will undergo elongation and will be
bent by a bending moment of constant magnitude
Mb = R4 a.
since R4 = P14
a Mb = PW f.
We would have obtained the same result if we would have examined the landing
gear on the whole with the help of the general view in Fig.280.
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