AVIATION INSTRUMENTS AND AUTOMATIC PILOTS

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP81-01043R001200220003-7
• Release Decision: RIPPUB
• Original Classification: K
• Document Page Count: 283
• Document Creation Date: December 22, 2016
• Document Release Date: October 23, 2012
• Sequence Number: 3
• Content Type: REPORT

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PliECISION MECHAlaCAL ITSTPUMF,VTS
T.I. Vilyavevf,tkaya
"1 A 11' "'4
1C. 1."4111.4 A.1.701,1TIC
4. '1 4 ' . ;)-T.
, ,
Authorized by the Yinistry of Defense Industry USSR
as Textbook for Technicums
STaTE PUBLISHIEG HOUSE
FOR THE DEFEESE IEDUSTRY
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A.
This book is a textbook intended for students of the instrument-building
technicums and has been compiled in accordance with the syllabus of the course in
"Instruments of Precision Yechanicg".
It sets forth the basic principles of operation of aviation instrunents and
autopilots, briefly describes the e1ents of desirn, and discusses th., questions
of the error of the instruments and the methods of elirinating such errors.
It rives an ides of the instruments that control the operation of the aircraft
engine as well as that of piloting and navic:ation instruments and automatic equip-
ment.
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PI.EFACE
Modern airplanes are equiped with instruments and automatic devices assuring
continuous control of the regime of flight and solving the complex problems of auto-
matic control and automatic orientation. The importance of instruments on aircraft
is increasing every year. At the present tire instruments have become one of the
most important factors determining the general technical level of aviation. The
development of aviation technology in turn has involved the improvement of the exist-
ing designs of aircraft instruments, the appearance of new designs, as well as the
utilization of fundamentally new methods of measuring various quantities, that had
not been previously in use. Aircraft equipment has undergone extensive quantitative
and qualitative modifications. The number of items has increase, and entirely new
forms of such equipment have appeared. 4;oetonatic devices for controlling flight
and for operation of power plant, radio equipment, devices for piloting and landing
aircraft under unfavorable meteorological conditions have been gaining over wider
use. In connection with the progress in the field of aircraft equipment the opera-
tional and tactical possibilities of utilizing aircraft have also expanded. The
further development of aircraft equipment and, in particular, of instrumentation is
proceeding along the lines of ever increasing automatization and increase in accu-
The present book has been written with respect to the syllabus of the course
in "Instrument of Precision liechanics" of Instrument building technicums and con-
tains materials on the section of the course entitled "Aviation Instruments and
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Automatic Pilots".
The purpose of the book is to give an idea of the instruments controlling the
operation of aircraft engines, the piloting navigation instruments, and automatic de-
vices. The book gives a description of the principle of operation of the individual
instruments and, in the must oeneral way, also presents the elements of design of
instruments and gives a considerable amount of space to the errors of aircraft in-
struments. The question or errors is considered in greater detail, since the tech-
nical instrument builder must have a distinct idea of the c&uses for individual er-
rors, of the methods for their total or partial elimination, and with respect to the
possibilities of a given method of meauurements of the possible accuracy of opera-
tion of the instrument.
The book does not give a description of instruments the, may be in question
for aviation or of instruments actually ir use but not typical for our modern USK
aircraft. Such instruments include: direct-current tachometers, tachometers with
rectifiers, etc. The automation of the aircraft engine is not considered. The book
does not include elements of calculation of instruments, question of installation,
disassembly and operation of instruments. In cases where the reader requires more
detailed study of some instrument (for example, in designing work) it will be neces-
sary to consult the book by D.A.Braslavskiy, S.S.Logunov, and D.S.Pel'por "The
Calculation and Design of Aircraft Instruments" (131b1.1) or one of the books given
in the Bibliography at the end of this book.
I express my sincere gratitude to G.O.Fridlender for his valuable assistance
with the book. I likewise express my appreciation to Ye.L.Veller, the editor of
this book, for a number of valuable comments made in reading of the manuscript.
The Author
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ILMODUCT1M
The most important units of an aircraft are as follows:
1. The body, the main design part of the aircraft which houses the crew, power
plant, equipment, and all units and devices forming a part of the design of the air-
craft.
2. The power plant, including the engine with the systems of fuel feed, cool-
ing, and lubrication, and also the propellers (for piston and turboprop engines).
3. The equipment, consisting of the instruments, mechanisms, units, and in-
stallations making it possible to control the aircraft to obtain optimum performance
of its mission. The concept of aircraft equipment includes the technical means al-
lowing:
a. control of steering of the aircraft, flying speeds, and operation of
the power plant, as well as mechanization and automation of these processes:
b. assuring a more complete utilization of the flight-technical means of
the aircraft in accomplishing its mission and increasing the safety of flight:
c. providing the most pleasant working conditions for the crew and passen-
The conditions of flight on modern aircraft, particularly on high-speed types
would be so complex and would require such acute and sensitive sensory organs of the
crew members, such an effort of memory and thought, such endurance, that in the best
case it would lead to extraordinary fatigue, and in the worst case it would be en-
tirely impossible to handle for the human organism, if technical equipment were not
called into action.
In the complex environment of modern flight, even at zero ground visibility,
the aircraft crew is rapidly able to solve, with the aid of instruments, the com-
plex problems of position fixing, of orienting the aircraft with respect to the
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ground, etc.
'Aircraft equip sent greatly simplifies the solution of many problems facing the
aircraft crew. For example, the intercom syetem of aircraft and the tneans for out-
side radio communication allow a norval conversation both between the individual
members and with crews of other aircraft or with the ground crew, despite the noise
of the power plants.
The systems of control and measuring instruments, widely used on modern air-
craft, and of transmission mechardsrs allow the pilot to take the decisicns required
for a given condition of flight and to iJmplement them without excessive efforts, by
using various power transmission mechanisms and drives. In a number of cases, the
latter task is considerably simplified by the use of appropriate automatic devices.
Thus, in spite of the progressive complexity of missions, the work of the crew
is increasingly facilitated through further development and improvement of aircraft
equipment.
The wide variety of modern aircraft instruments has been developed and improved
over a period of many years, in the process of improvement of other aircraft equip-
ment and in step with the continuous expansion and complication of the problems
solved during fli
t. Some instruments were in existence considerably earlier than
the first aircraft. For example, the magnetic compass and the methods of naviga-
tion by compass developed by the efforts of Russian scientists, such as Admiral I.F.
Kruzenshtern (1770-1846), Lt. I.P.Belavents (1829-1878), Academician I.P.De-Kolong
(18396-1902), and Academician A.N.Krylov (1863-1944), have found widespread applica-
tion in aviation. Russian aircraft designers were the first in the world to apply
the most advanced methods and instruments of the time with respect to navigation by
compass as well as the methods and instruments of astronomic orientation of the
aircraft, The Russian aircraft instrument building industry has always been, and
still is, at a very high level.
The first navigation instruments were designed by Russian aeronauts as early as
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1804, and the Academician Ya.D.Zakharov made a flight on a balloon equipped with a
comps, an instrument for determining ascent and descent and an optical telescope;
i.e., instruments allowing the route of the balloon to be plotted on the map.
The famous Russian designer A.F.Mozhayskiy, in designing his airplane, thorough-
ly laid out not merely the desifn of aircraft and engine, but also took into account
the purpose of the aircraft, i.e., he provided for equipment on board that was nec-
essary for the completion of practical flights. belying on the experience of navi-
gation and ship building, he installed on his aircraft bank-and-turn indicators,
altimeter, thermo.eters, a speedometer, and a compass. To him belongs priority in
the design of the entire set of aircraft equipment.
In the 1890-s and at the beginning of the 20th century, 1;ussian scientists de-
signed a series of instruments for navigation and piloting. Y.M.Pomortsev, in 1896,
4 44 Writtleart 41. ea, .1. 04 At. 4, A...a. re. 4
io71 4a, .4.WV:Z.Z, NO A NAL 11.-34.413..j...4 1/
the direction and angular velocity of clouds
and, in 1897, an instrument to determine the velocity and direction of motion of a
balloon, which was the prototype of modern sights that appeared considerably in
other countries.
In 1898, the famous Russian scientist K.E.Tsiolkovskiy was the first in the
world to propose the idea of an autopilot and to rive its working diagram.
The founder of radio engineering, the inventor of the first radio transmitter
and radio receiver, and the first person in the world to actually accomplish radio
transmission and radio reception (1895) and to discover the principle of radio loca-
tion, was the famous Russian scientist A.S. Popov.
The first heavier-than-air airplane in the world, the "Russkiy vityazt" con-
structed in 1913, was provided with a tachometer, clocks, altimeter, and compass.
The four,-engine aircraft "Illya Muromets", constructed in 1914, was equipped
with compasses, altimeters, speed indicators, clocks, and tachometers. On this
aircraft the method of navigation by compass was successfully employed for the
first time. At approximately the same time, A.r.auravehenko designed an anemomter,
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V.A.Slesartv a speed indicator (which was a few years ahead of the American speed
indicater uPioneer"), V.P.Vetchinkin, an accelerometer, etc. In contrast to foreign
designers, the designers of Russian aircraft always paid great attention to conven-
ient workine conditions for the crew and provided it with the equipment necessary
for flights. Russian designers developed a number of very interesting instruments
and devices, howevier, manufacture of these instruments in the necessary form and in
adequate number proved impos ible, since the development of aviation, and conse-
quently also of the aircraft-instrument building industry, was not properly sup-
ported by the ruling class.
The Russian aviation industry and aircraft instrument buil-line industry begat
their intensive development only after the October kevolution? In spite of the dif-
ficult conditions in the country at that time, the Soviet Government literally from
the first days began to build an aviation industry and an aircraft-instrument build-
ing industry. Already in 1919 a Soviet plant "Aviapribor4, turned out instruments
for the air fleets and in 1922 this plant changed to series production.
In 1923, this shop produced the following instruments: oil gages, gasoline
manometers, air thermometers, tachometers, altimeters, deflectometers, and speed in-
dicators. The plant had many designers and research instrument builders whose names
today are widely known (S.A.Nosdrovskiy, S.S.Tikhmenev, G.O.Fridlender, and others).
As a result of continuous close contact and work in collaboration with noted
Soviet scientists, pilots, and designers, instruments and automatic devices were
successfully used on aircraft and are still being created.
A Collective of Soviet instrument builders consisting of D.A.Braslavskiy,
Kachkachlyan and M.G.Elfkind has developed a number of instruments, including the
first gyromagnetic compass in the world.
A.A.Andronov, B.V.Bulgakov, S.E.Khaykin, as well as the young Soviet scientist
V.V.Solodovrikov, Ya.Z.Tsypkin and others, have had exceptional success in the theo-
ry of automatic control and the theory of gyroscopic instruments. Utilizing the
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theoretical work of the scientists, the Soviet designers have constructed and are
now constructing excellent models of instruments and automatic devices which. are suc-
cessfully.used in aviation, and which lighten the work of the crew and assure safety
The use of automatic devices is particularly widespread for piloting aircraft
(autopilots), automatic devices controlling aircraft power plants, and co7-u:ers for
automatic plotting of an aircraft course (automatic navii7ators).
Such intense development of the USSF. aircraft-instrument building industry has
been possible only on the basis of the success of Soviet scientist and desfE7ners in
the theory of regulation, the theory alkd design of various types of automatic de-
vices, electric measurement, F:yroscopic and other instruments.
Such work in the field of metroloy and in the
study of the errors of easur-
in instruments have been of particular importance.
The further improvement in the technical level
of our USSF aviation wifl con-
front Soviet instrument bIldlders with a number of problems
instruments and automatic devices and also with the 4ntroduf?-
tion of the most recent achievements of science and industry and technology into
. _ .
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CHAPV1Z I
PURPOSE MD APPLICATION OF AIRCEAFT INSTRUMEUTS
Depending on the purpose, the instrumental equipment of an aircraft is divided
into the following groups:
1. Instruments controlling the operation of the aircraft engines;
2. Piloting and navigational instruments;
3. Automatic devices controlling the operation of the aircraft engthes as well
as automatic piloting devices (autopilots).
Section 1. THEPOW PLANTE; of }ODEFJ AIRCRAFT and the IESTFUMENTS
WNTr:OLLING THEIP OPEIATICI:
nodern aircraft are equipped with piston engines (PE) turbojet engines (TJE),
jet engines (JE), or turboprop engines (TPE). At the present time piston engines,
air- or liquid-cooled, are still widely used. Jet engines are used on high-speed
aircraft. Turboprop engines whose appearance has been relatively recent, are begin-
ning to be used more and more. The necessity of using a propeller with such engines
limits their application; such engines are unsuitable for high-speed aircraft.
The number and types of instruments controlling the puer plant depends on the
special features and type of the power plant. However, as shown below, their no-
menclature varies only slightly.
The operation of such an engine is based on the conversion of the thermal ener-
gy of a Combustible mixture burned in the engine cylinders into mechanical work,
rotating the blades of a propeller, thus creating thrust.
For the exact setting and maintenance of the operating condition of an engine
it is necessary to:
1. Know the fuel supply on the aircraft; its control is accomplished by the
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2. To supply air and fuel to the engine in a definite proportion and under a
definite pressure. Indication is by means of the fuel gage and the vacuum gage
(with forced air feed of the engine, i.e., with supercharging used for maintaining
a definite ratio of oxygen to gasoline in the combustible mixture).
3. To assure the uninterrupted supply of lubricant to the friction parts of
the engine. In this case It is necessary to control not only the pressure, under
which the oil is supplied, but also its temperature. At a low temperature (below
10-15?C) the viscosity of the oil increases sharply, its rate of flow through the
pipe lines is dilninished, and its feed through channels of small cross section (for
example, to the engine bearings) is impeded. At high temperatures, the viscosity
o.f the oil decreases, it acquires fluidity, and adheres poorly to the clearance be-
tween the friction parts. At excessive temperatures, oil will burn, and the prod-
ucts of its combustion clog the
gages and oil thermometers.
4. Maintain the temperature of the cylinders and pistons within the allowable
During combustion of the fuel mixture the engine cylinders are strongly heated;
to avoid overheating, cooling is used. Depending on the method of dissipating the
heat, aircraft engines are conventionally divided into air-cooled and liquid-cooled
engines. In air-cooled engines, the temperature is checked by means of a cylinder-
head thermometer while in liquid-cooled engines, coolant thermometers are used.
Not only overheating but also overcooling of the cylinders is dangerous for
engines, since in this case the rate of combustion of fuel-air mixture is reduced.
An engine can lose its pickup, i.e., its ability to shift rapidly from one regime
of operation to another. The loss of pickup is particularly dangerous in landing
'when in some cases, the propeller rpm must be rapidly increased to prevent loss of
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To maintain the necessary pickup, the gasoline must be evaporated at a suffi-
'cient rate in the carburetor. The rate of evaporation depends on the temperature of
the carburetor, which is indicated by the carburetor thermometer.
5. Know the speed of the engine shaft. This value is measured by .the tachome-
The instrument indicating the composition of the fuel mixture is of great im-
portance in the operation. However, attempts to design instruments of this type
have not yet given the desired result.
The gas analyzers used for this purpose allow the composition of the fuel mix-
ture to be determined from the compostion of the exhaust i'as. The considerable er-
rors inherent in this instrument interfere with its widespread application.
TelLrbo'etandrel-ona-ines
The operation of jet engines is based on the reactive action of the jet of
gases formed by the combustion of fuel and expelled through a channel of small cross
section having the form of a nozzle. For the combustion of fuel in jet engines at-
mospheric oxygen is used (air-,!et engines) or special oxidizers (liquid-jet engine).
The peTformance of a jet engine is characterized by the rate of revolution of
the turbine, the temperature of the gases in the jet nozzle, the temperature and
pressure of the fuel and oil, the consumption of fuel and oxidizer, the temperature
in front of the turbine, the Each number at the entrance to the compressor, the
static pressure, etc. These are the same parameters that are measured in the opera-
tion of piston engines, but the limits of measurements of many of these quantities
are considerably wider for a jet engine. Since the instruments for controlling the
operation of a jet engine are designed with allowance for the peculiar features of
operation and for the range of operation, the possibility of using instruments
based on completely different principles, not used in control instruments for pis-
ton engines, is not excluded.
The instruments used to control the operation of a turboprop engine (TR) are
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analogous to the instruments used for jet engines.
Instruments and Automatic Devices for the Power Plant
For reliable and economic operation of an aircraft power plant as well as for
obtaining maximum thrust or power, it is necessary to provide under a)) conditions
of flight, the most advantageous regime of operation of the power plant and to con-
stantly check its operation by means of control instrument and automatic reculating
processes.
To facilitate the work of the crew during flight every effort is made to use
automatic equipment. The control instruments for operation of a power plant are
subdivided into the folio lno groups:
1. Instruments whose readings characterize the thermal reoime and condition of
the engine lubricant: the oil thermometer, coolant (or cylinder-head) thermometer,
the working-cos thenmeters (for :et engines) and the oil pressure gazes.
2. Instruments indicating the power or thrust developed by an aircraft engine:
vacuum gages, manometers, tachometers, thermometers, and gas analyzers.
3. Instruments indicating the fuel reserve and fuel consumption and the nil
reserve: fuel gases, flowmeters, oil gages.
The automatic regulators of the regime of operation of aircraft engines include
the following:
1. Automatic engine speed control.
2. Automatic coolant and oil-temperature controls and automatic cylinder-head
?temperature controls.
3. Automatic switches for supercharger speed.
4. Automatic boost pressure controls, etc.
Section 2. AIRCRAFT FLIGHT AND PILOTING INSTRUMENTS.
The Navigall.onal Regime of Flight.
Every flight is connected with the fulfillment of an assignment defined by the
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To maintain the regime of
librium of the aircraft in the
These tasks are performed
ld direction.
flight the ?pilot must continuously maintain the equi-
air and check its position with respect to the ground.?
by the aid of instruments. The instruments control-
ling the operation of the power plant allow the necessary performance of the air-
craft engine to be selected and maintained.
The piloting-navigation instruments make it possible to determine the position
of the aircraft and its speed. Accordine to the weather conditions, the pilot es-
tablishes and maintains the required navieational regime of flieht by the aid of
one or the other group of instruments.
Various causes, for example, eNsts, variation in thrust of the propellers and
other causes, ray change the position of equilibrium or cause a deviation from the
selected course. For this reason the aircraft crew rust be continuously able to
check the position of the aircraft in space and to restore it to the required atti-
tude.
A deviation of the aircraft may occur with respect to the xx, yy, and zz axes
(Fig. la).
Two systems of coordinate axes are differentiated: The moveable system, in-
variably connected with the aircraft coordinate system Oz, with the initial coor-
dinate in the center of gravity of the aircraft, is ca31ed the bound system of co-
ordinates, while the xx, yy, and zz axes art called the principal axes of stability.
The equilibrium of the aircraft with respect to the xx axis is called trans-
verse equilibrium and with respect to the zz axis, longitudinal equilibrium.
The longitudinal axis Ox, parallel to a wing chord, is directed forward and
lies in the plane of symmetry of the aircraft.
The normal axis Oy perpendicular to the axis Ox? lies in the plane of symmetry
of the aircraft. During horizontal flight of the aircraft, this axis is directed
upward.
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The transverse axis Oz is perpendicular to the plane of symmetry of the air,-
craft (positive values are measured on the side of the right wing). The planes of
the coordinates in the bound axes have the following designations:
Oxy: the plane of symmetry of the aircraft;
Oxz: the plane of the wings, or the principal plane;
Oyz: the transverse plane.
6)
Flg.l. Coordinate Axes of Aircraft
a- Oxyz system fixed relative to
the aircraft; b- Oixiyizi system
fixed relative to the ground;
1- Rudder; 2- Elevator; 3- Ailerons
4- Stabilizer; 5- Fin.
The axes (Fig.1, b) fixed relative to the
ground (the so-called ground axes) are se-
lected in the
nilowing way: the axis
is directed vertically upward from below.
The axes C1,1 and Olzi are located arbi-
trarily in a horizontal plane and include
an angle of 900. The origin of coordinate
is selected arbitrarily.
The position of the fixed axes with
respect to the ground axes, and consequent-
ly also the position of the aircraft with
respect to the ground, :Is determined by the
angles ;, y, and k (Fig.2). The angle Q
between the xx axis and the plane of the
horizon is called the angle of pitch. The
angle y of the rotation of the aircraft
with respect to its longitudinal axis is
tween the plane of symmetry of the air-
craft and the meridian (the line of inter-
section between the plane of the horizon and the plane of the geographical meridian)
is called the true course of the aircraft. The angle k is measured from the air
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meridian.
If, as the initial direction, the magnetic rather than the g ographic?meridian
is used,, the course is lalown as the ma?;n tic course instead of the true course.
Fig. 2. Angles Characterizing the Position of the Aircraft with hespect
to the Ground
gp Angle of pitch; y, Angle of bank; kl Aircraft course.
The longitudinal equilibrium of the aircraft depends on the angle of pitch,
since this angle varies with any variation in the angle of attack of the aircraft
(Fig.3) which, in turn, produces changes in the aerodynamic characteristics of the
aircraft.
The angle between the direction of the projection of the velocity vector onto
the plane of symmetry of the aircraft and the ming chord is called the angle of
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attack. As the wing chord we take either a line tangent to the lower surface of the
wing profile (tangent chord), or a line connecting the nose and tail of the profile
(internal chord). The ing profile
and their chords that are most common at the
w
present time are shown in
As already indicated* the angle of attack a affects the aerodynamic character-
istics of the aircraft, i.e., the lift Y and the drag Q (Fig.5)
Fig. 3. Angle of Attack of Aircraft
a - angle of attack,
V- Velocity vector of the air
V the aircraft speed in m/sec,
cy the coefficient of lift;
cz the coefficient of drag.
when the angle of attack increases, the coefficient of lift cv increases and
reaches its maximum value at a certain value of the angle of attack which is called
the critical angle (F!;.6). Further increase of the angle of attack leads to a
sharp decrease in c. and the lifting force. One of the conditions of equilf.brium
of an aircraft in horizontal,- rectilinear, and uniform flight
the weight of the aircraft G and its lifting force Y:
cySpV2
whence
Thus, the greater the value of Cy the sraller the speed necessary to maintain
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horizontal flight. Consequzntly, depending on the value of the speed V, an aircraft
wifl perfmm a rectilinear horizontal flight at different angles of attack.
The angle of attack, to a considerable extent, determines the speed necessary
Fig. 4. Typical Wing Profiles. Fig. 5. Aerodynamic Forces Acting on the
ab- Wing chords Wing
F.- Total aerodynamic force; Y- Lift-
ing force; Drag.
for maintenance of horizontal flight. Obviously, this is true also for any other
regime of flight. To each regime of flight there corresponds a definite minimum
value of the flying speed V at which the
AiN
aircraft is still able to raintain equili-
brium and maintain the assigned regime.
Thus, it is necessary to know the
aircraft speed V not only to calculate the
et
2 1 time of flight but also to naintain the
15 I
longitudinal equilibrium of the aircraft,
which depends to a considerable extent on
Fig.6. Relation Between the Angle of
the angle of attack.
Attack and the Coefficient of
Lift eyInstruments Indicating the iegime of 4t
The aircraft speed is judged from the
readings of the speed indicator. This instrument indicates the so-called air speed
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i.e., the speed of the.air raft with respect to the air. As will 'be pointed out la-
ter, this same instrument indirectly makes it possible to judge the longitudinal e-
quilibrium of the aircraft and consequently, the value of the angle of attack. The
flight altitude is indicated by the altimeter and the rate of ascent or descent by
the climb indicator.
The course of the ire ft is controlled by means df 'compasses and the turn in-
dicator. The latter indicates the presence
of deviations of the aircraft from recti-
linear flight, i.e. indicates changes in
course. The lateral equilibrium .of the
aircraft depends on the angle of bank. In
4 rectilinear flight, banking causes side-
slipping (Fig.?).
In curvilinear flight (Fir.8), for ex-
ample in turning, inertia forces* may cause
c - Angle o absolute bank. wing-over and slipping. If when the air-
Fig.?
f
^
r ,4
- Rectilinear Flight
a - Without bank; b - With bank;
craft executes a turn, the resultant of the
gravity and centrifugal forces, directed
along the apparent vertical, coincides with a straight line perpendicular to the
plane of the wings (cf.Fig.7 b), then a correct turn is being made. Houever, if
the apparent vertical deviates from this straight line, there will be an outside or
inside sideslip.
The lateral equilibrium of the airplane is checked by means of the bank indica-
tor.
The direction of flight can be determined either from visible landmarks or by
means of various compasses, or from airborne radio instruments. During the flight,
OP.OW * Ir7turning, the aircraft is subject to centripetal accelerations. In this
ease the centrifugal forces of inertia will act on the aircraft.
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.the crew of the aircraft &a constantly using instrum nts and estimating the position
of the aircraft fro their readings.
The quieting aircraft. instruments are particularly important in flight -without
visible external landmarks, 14hen the human prpaniam is subject to the action of.
Inertia forces, and the pilot easily loses track of the actual position of the air-
craft in space. In this case the Jape fection of the human organs of equilibrium,
is revealed. If, for example, the aircraft is flying without visible external land-
marks at a velocity V and makes an irregular left turn with a radius r and an
.lar velocity of turn tLi, then the centrifugal force F that forces the pilot toward
the right side of the ship is determined by tile equation F rr,t'Or mV W where ra is
the mass Of the plot.
?
*
orgelvvr
iERrtcAn.
Fig.8 - Aircraft Turn
a - regular left turn; lcrt turn with sideslip; r - radius.of.turn; w- - an-
gular velocity of turn; y - angle of ?absolute bank; yo - angle between arm rent'
and true vert ,
ical- yy.y angle of relative bank.
0
Since the pilot does not feel the turn and believes that the aircraft is in
rectilinear flight he reaches the conclusion that the aircraft is banking to the
right, and that it is necessary to straighten it out; in this case he will straight-
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en it out not with respect to the true vertical but to the apparent vertical
? (Fig.9). The power necessary for the turn exceeds the power 'necessary for recti-
linear flight, and therefore the aircraft begins to lose speed. If the aircraft has
a turn-and-bank indicator on board, the pilot will immediately discover the change
from rectilinear Alight to a regular or irregular turn and will know how to restore
the aircraft to the required attitude with respect to the ground.
The compass, altimeter, climb indicator, speed indicator, bank indicator, and
turn indicator allow the pilot to judge by indirect methods the position of the air-
carft with respect to the ground. The bank is determined from the bank-and-turn
indicator while the pitch is determined from the speed indicator and the climb in-
dicator.
If, at constant aircraft speed, the vertical speed changes, then the angle of
attack and the angle of pitch of the airplane also change.
But the indirect method of determining the position of the aircraft considera-
bly complicates piloting and places an excessive nervous strain on the pilot. In
addition, the speed indicator and variometer as indicators of the longitudinal posi-
tion of the aircraft give readings with considerable lag, while the readings of the
magnetic compass are unstable under bumpy flight conditions. For this reason, in-
struments free from the above faults were designed which make it possible directly
and accurately to determine the angles of longitudinal and lateral deviation and
? the course of the aircraft.
Such instruments are the gyroscopic instruments installed at the present time
on all aircraft.
netic compass etc. The work of the pilot on an aircraft equipped with these in-
struments in flight without visible landmarks is reduced to the observation of the
instruments and control of the aircraft rudders. This is purely mechanical, monot-
onous, but very fatiguing work.
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On an aircraft equipped with an automatic pilots the pilot Is considerably lenbs
_burdened and his work becomes less fatiguing. But the simplest autopilot can be used
only in rectilinear uniform horiiontal _aight; the more complex, so-called program
1
:autopilots, which control take-off land-
ing, or maneuver of the aircraft, are in-
stalled Only on special aircraft.
The maneuver aircraft Connected with
a Change of its speed in magnitude and di-
rection is called evolution. They in-
ciud
?
urningl diving, stunt flying, etc.
In evolutions, the aircraft- moves with
accelerations varying in magnitude and
direction, and consequently the forces of
inertia act both on it and on the airborne
instruments. The concept of overload is
k'
ordinarily used to characterize these
r- X ud
forces.
Overload (indicated by the symbol n)
is the ratio of the resultant acceleration
Fig.9 -Aircraft Turns a with which an aircraft is moving to the
a - Without bank; b and e acceleration of gravity g.
With bank by an angle:
(1.4)
High overloads can have a, disastrous effect on the organism of the crew, on
the Aircraft, and on the instruments. The degree of overload is checked by the
accelerometer.
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\
? Section 3.
AIRCRAFT NAVIGATION
The work of the drew in flying the assigned route and reaching its objective
is called aircraft navigation. On light aircraft designed for short trips this work
is performed by the_pilot. On medium and heavy aircraft which fly over considerable
distances, A qlnglA peron cannot do all the work; for this reason the pilot does
the piloting of the aircraft while the navigator handles its orientation. In many
cases the crew of an aircraft numbers 10 and more and includes several pilots, navi-
gators, a boamd engineer, a radio operator, etc.
ItirizMethodsofOrier-craft
For orienting an aircraft in the air the following methods are used:
a. The method of visual orientation, i.e., comparison of visible landmarks
(railroads, bridges, etc.) with the map on board the aircraft. Such orientation is
possible only when ground visibility is good.
b. Astronomical orientation reduced to the calculation of the position from
angles measured between the directions to selected heavenly bodies and the plane
of the horizon, allowing for the time of observation. Such orientation, which can
be effected only if the visibility of the heavenly bodies is good, is performed by
means of special optical instruments, for example, a sextant, optical sight, etc.
c. Radio orientation, reduced to the determination of the position of the
aircraft from the directions to ground radio stations or the distances to them, as
measured in flight. The method requires the existence of one or more ground radio
stations whose position is known and is carried out with the aid of special radio
instruments: radio compass, radio direction finder, etc.
Various methods of determining the position of the aircraft by means of radio
methods are known. One of these will be discussed below.
By using the readings of the radio compass (or other radi) instruments) it is
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easy to find the _angle a of the ground transmitter between- the direction of flight
to the radio station And the longitudinal axis of the aircraft. knowing thecourse
of the aircraft k (for example, from the readings of the magnetic compass), the
true radio bearing
on the earth's surface at which the true radio bearing has a certain value, we
obtain a line of possible aircraft positions (position line) but still not its
actual position.
By determining the value of the true radio bearing for two ground transmitters
whose position is known, we obtain two position lines. The point of intersection
of these position lines on the map will correspond to the position of the aircraft.
With such a method of orientation no radio transmitter with a directed emission
is required.
Many methods of radio orientation require a directive transmission. The pro-
cess of introducing radio engineering systems into the group of instrument equipment
Fig.10 - Determination of the Position of an Aircraft
a. - Geographical coordinates; To latitudes; X long-
itude; bo orihodromes and loxodromcs
of aircraft has still not been completed. There is no doubt that in the future
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these systems will acquire even greater importance. But it must be remembered that
the introduction of radio navigation instruments does not exclude the use of instru-
ments operating on other zinc.ples (mechanical altimeters, gyrocampasses, astron-
omical compasses, etc.)
This is explained in the first place by the danger of failure of the aircraft
radio equipment due to interference by ground radio stations, and, in the second
place, by the danger of being detected.
The operation of airborne radio equipment may reveal the aircraft and allow
its approach to be detected sometimes at very great distances.
d. The method of calculating the course, in which the position of the aircraft
is determined by calculating the value and direction of the segments of the course
already traveled by the aircraft from the take-off point. Under the conditions of
actual flight the aircraft flies along a certain curve connecting the initial and
final points of the route, i.e., both the longitude and the latitude of the place
wery during the time of flight (Fig.10a). In order to have the shortest possible
route, the aircraft must fly along the arc of the great circle between the initial
and final points of the flight course, called the "orthodrome", (Fig.10b). During
such a flight, the course of the aircraft varies continuously, since the orthodrome
intersects the meridians at different angles. This change of course is inconvenient
in operation, but on long-distance flights, such maneuvers result in a marked
shortening of the route and thus in a saving of fuel and time.
An orthodrome is divided into a series of segments within which the curvilinear
segments intersecting the meridians at different angles 01, 02, 03, etc. are re-
placed by segments each of -which intersects the meridians it meets at one and the
mine angle O.
The curve so obtained is called a loxodrome. The flight route is plotted in
advance; the proposed flight path is divided into separate sections in such a way
that, within the limits of each section, the direction of flight remains the same.
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4-,
If, during the flight the speed and time corresponding to each section are measured,
then the true distances traveled by the aircraft can be found and these distances
plotted on the corresponding scale on the map; in this way, the coordinates of the
aircraft at a given moment can be determined. The method of route calculation is
inconvenient in that it requires a very accurate determination of the course, speed,
altitude, and time of flight.-
Air-Navigation Control lnstrumcts
It is commonly known that to measure the course of an aircraft it is necessary
to know some fixed direction with respect to the ground, from which the calculation
is made. This-direction may be the geographic meridian, the magnetic meridian, etc.
The course reckoned from the magnetic meridian is called the magnetic course;
if
reckoned from the geographic merIdian, it is called the true course, while the anrle
between the geographic meridian and the direction of flii'ht (course), measured
clockwise from the northern direction of the meridian is called the tree course
angle (like the course this angle may be either true or magnetic). Usually the
course and the longitudinal axis of the aircraft do not coincide, since the speed
of the aircraft with respect to the ground (the ground speed) is the geometric sum
of two speeds, the speed of the aircraft with respect to the air (the airspeed)
and the speed of the air with respect to the ground (the wind speed). The true
course angle does not coincide with the true course of the aircraft and is equal
to the sum of the true course of the aircraft and the angle of drift w , i.e., the
angle between the longitudinal axis of the aircraft and the course, the angle of
drift due to the wind fw and the aerodynamic drift.wa (which arises, for example,
as a result of the unequal thrust of the propellers in multi-engine aircraft).
These existing methods of determining the angle of drift give the total angle of
drift, =4,14 4 W, and for this reason practical aerodynamic drift is not separate-
a
ly considered. The angle of drift is measured by the aid of special instruments,
for example, navigation sights, etc. In determining it it is necessary to know
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3,Y the altitude and duration of flight. The ang e of drift is considered posi-
tive if the aircraft deviates toward the right.
The course of the aircraft is determined by the aid of compasses.
It is more difficult to determine the speed of the aircraft than its course.
The dead reckoning method requires the knowledge of the aircraft speed with respect
to the ground, i.e., the ground speed.
Until now, no instruments indicating
the ground speed of the aircraft have
been designed.
Manometric speed indicators, widely
used in actual operation, indicate the
airspeed of the aircraft or in the best
case, the true air speed. A true air
speed indicator allows for the so-called
methodological error of the instrument,
which is manifested as a result of the
Fig.11 - The Navigational Velocity
variation in density of the air with
Triangle
height, while the air speed indicator
V - Air speed; U - Wind speed; W
does not allow for this error.
Ground speed; k - Course of aircraft;
The lack of an instrument indicating
0 - Direction of wind; 10 - Drift of
the ground speed makes it necessary to
? aircraft; a - Course angle
determine the value of this speed by nav-
igation instruments. The ground speed may be determined with a navigation sight,
:observing the rate of displacement of landmarks on the ground. In this case, the
altitude must be known.
The magnitude and direction or the ground speed may also be found by cOnstruc-
a navigational velocity triangle (Fig.11).
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alL11111Eallaal-Y212211Litalael2
One side or the navigational velocity triangle is the Vector of air speed of
Hthe aircraft V, another side is the vector of wind speed U, while the third side
representing the sum of these two Vectors gives the Vector of ground speed of the
aircraft W V U. The magnitude of the air speed of the aircraft is taken from
the readings of the true airspeed indicator or of the air speed indicator corrected
with. an air-navigation slide rule, making it possible to allow for the methodologi-
cal error of the. instrument. The wind-speed vector may be determined by the navi-
gation sight using any one of several methods, for example? measurement of the wind
from two angles or drift. It must be remembered that in determining the angles
of drift, the pilot is obliged to hold the assigned course and keep altitudes and
speed constant.
Fig.12 - Determination of the Wind
Vector U
lc, and k2 aircraft courses; 1/1,f2 an-
gles of drift on courses ki and k
respectively
The magnitude and direction of the
wind are determined in the following way:
Selecting some landmark on the ground and
observing it for a definite time interval,
the value of the angle of drift is deter-
mined, i.e., the direction of the course
is found.
The same is done by changing the
course of the aircraft by 40-500, thus
obtaining a new flight line of the air-
craft. On plotting the flight line on
the map the wind-speed vector is obtained
(Fig.12).
In determining the wind vector, the
navigator of the aircraft does not make the above-mentioned constructions on the map
but USe3 a wird triangle instead, which
allow him very rapidly and accurately to
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As already Mentioned, the wind triangle was proposed by A.N.Zhuravehenko at
the beginning of this century and up to now has been widely used in aviation to
obtain rapid and accurate solution of problems by the aid of the navigational tri-
angle or speeds thus making it possible to determine the speed and -direction of the
wind,. the course speed, the angle of drift, the course angle etc. i.e., to find
When the ground isInvisible, and when the Use of optical Sights is impossible,
the course angle Is determined by solving the navigational triangle. The course
angle, the angle of drift, etc. are determined by calculation if the direction and
speed of the wind are unknown (for example, from data obtained from the ground
sta-
tions).
The Calculated Position of the Aircraft
The aircraft crew at every moment of time must know the position of the air-
craft. This position is determined in terms of geographical longitude and latitude.
The longitude is reckoned from the Greenwich meridian to the west of it the western
longitude X w, and to the east of it the eastern longitude X e. The latitude is
reckoned from the equator (north and south latitude (pn,)? During flight along
the meridian only the latitude changes by the value
where tiSm is the path traveled by the aircraft along the meridian; and
R is the radius of the earth.
During flight along a parallel line, only the longitude changes by the value
where Sp ie the path traveled by the aircraft along the parallel;
r is the radius of a circle of the given parallel r R cos T1
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R is the radius of the earth;
T 4.4 the latitude of :the given parallel.
? In flight along an arbitrary trajectory the coordinates of the aircraft are
Continually changing.
Let us denote the latitude and longitude of the take-off point by? To and X0;.
current values of these coordinates by cp
and X.; the true course of the aircraft
by k; the direction of the wind by; the
w
wind speed by U; the time of flight by t;
the component of the course velocity along
1.Y4k
meridian by W,; and the component of
the course apeed along the parallel by Wn
(ig.13).
Then we may write
Fi 13 - Determination of the Calculated
Position of the Aircraft
V - Airspeed; W - Ground, speed;
- Wind speed; k - Course of air-
' craft;9 - direction_ of wind
V cos k + U cos Tw
W V sin k + U sin c)w
For an infinitesimal interval of
tine dt, the aircraft is displaced along
the meridian or the parAllel by the re-
In this case the coordinates of the aircraft vary in the following manner:
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_
ions with res
et to time we find
36? Weosk+Licos?Odt,
20? .
ti
360 is V s ink+ U co dt.
cos
? e
During the time of flight, the navigator periodically determines the coordin-
ates of the aircraft. This task. is considerably simplified if there is a navigation-
al coordinator on board the aircraft. This is an instrument .which automatically
solves the above equations and gives the values of X and T or an automatic navigator
(1. .5)
(16)
which solves the problem more fully
the map.
d plots the entire course of the aircraft on
.Navigt tonal and ?41Ot1FA Instruments
For the purposes of piloting and aircraft running, the following instruments
are used:
1. Speed indicators;
2. Compasses;
3. Altimeters;
4 Rate of climb indicators;
5 Gyro horizons;
6. Bank indicators;
7.. Turn indicators;
8. Navigation coordinators;
9 Mach-number indicator;
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10 Accelerometers.
In add tion in the majority of cases anautomatic pilot is installed on the
aircraft.
To assure reliability, many instruments are paired, for example in addition to
the magnetic compass a directional gyro a remote reading compass or a remote indi-
cating compass, a radio compass; etc, are installed on the aircraft simultaneously
with the magnetic compass.
Fig.14 - Instrument Board of an Aircraft
The instrument board of an aircraft with instruments arranged so that the crew
can conveniently observe their readings, is shown in Fig.14.
The central part or the instruMent panel is occupied by the in piloting and
navigation instruments: the speed indicator, turn indicator, and bank indicator,
,altimeter, rate or climb indicator, compasses gyro horizon, and clock. Since the
instrument boards are set Up in the zone of direct observation of the pilot, any
'increase in their dimensions would unavoidably lead to impairing the view. To
-Hpreserve free view, the instruments are compactly arranged their outside dimensions'
are reduced combination instruments are used, and finally the shape and design of
the instrument boards is carefully selected.
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Since the instrument panels are located at points of the aircraft where the
Heffect of vibrations is very considerable, while the instrument on these panels
are highly. sensitive to vibrations, a shock absorption for instrum nt panels is
-provided.
In individual cases, in addition to the instrument panel individual shock
absorbing must'be provided for one instruments, such a compasses.
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CHAPTSJ 11
BASIC REQUIRDENTS FOR AIRCRAFT EQUIPMENT
The development of modern aviation, the expansion and complication of the prob-
lems to be solved by aircraft crews in flight require an increase in the number of
aircraft equipment items installed on the aircraft. The number of objects in the
aircraft cabin requiring the attention of the aircraft crew is in the tens or even
in the hundreds. Some inttirents revire the performanc of complex calculators,
the use of special Tables, etc. For example, radio ecnipert and means for astron-
omic aircraft navigation).
All aircraft equipment including the instruments, is intended for the perform-
ance of various tasks under the specific conditions of flight, and must satisfy the
so-called tactical-technical requirements resulting from these problems and condi-
tions.
All tactical-technical requirements can be divided into functional, operation-
al, physical-technical, assembly, etc.
Section 1.
TACT IcAL-TEcliNICAL P3):1U111.1-aw?NTS
Each instrument must satisfy its purpose, for example, measuring instruments
;must measure definite physical values with the necessary accuracy, drives must de-
velop the necessary forces and Moments, radio equipment M118., operate in a definite
frequency range, etc.
:Whether aircraft instruments answer their purpose is determined by the nominal
Hvalues characterietic'for their physical-technical quantities and indexes, i.e.,
-4its rated characteristic S (nominal) and by the allowable deviations from these
L, 6
!characteristics.
The rated 'characteristic of any technical instWation is the quantity char-
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? 1,
acterizing its basic technical parameters corresponding t
its intended function.
For example, the rated characteristic of a radio instrument includes the range of
frequencies of the electromagnetic oscillations that can be used, the rated charac-
teristic of electrical' ins rumenta includes the value of the voltage, current, etc.
The selection of the rated characteristics must be done with great caution and care,
and the rated characteristics rust sattzfy the actual conditions of operation.
Otherwise the operation of a given instrument and of the aircraft units connected
with it will give great. difficulty, .he quality of performance or the mission will
be lowered, and there may even be danger to the aircraft and 'crew. At tne same
time, a selection of rated characteristics beyond the actual functional requirements
ay lead to cemplication of design, and to an increase in weight, dimensions, cost,
:etc.
The allowable deviations from the nominal are selected carefully, taking account
.of the service conditions. For example, it is often required of an instrument that
it exhibit minimum deviation from the nominal over only a portion of the scale in-
stead of over the entire scale. In this case, the remaining. range of the scale
may show somewhat higher reading errors.
Section 2.
PHYSICAL-TECHNICAL REQUIREMENTS
The physical-technical requirements to be met by aircraft instruments cover
'normal operation of the instruments under actual conditions of flight. In order
properly design andAnanufacture an instrument and to be able to compare various
designs of instrument and various methods of measurements and, of course, to select
the best ones,' and finally to operate the instruments properly, it is necessary to
study the conditions under which the instruments operate, and also the regularities
to which these conditions are subject. It is above all necessary to familiarize
,1,f0,1.1,11ePr9P.!7.ti.es of the atmosphere.
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. ? ???. . ????, ? .
The Atmoseee.nalszas_,ties_
'The atmosphere eurroundi7 the earth is divided into three layers: the tropo-
sphere, stratosphere, and the zone of rarefied gases. Since at the present time
flights are made mainly within the Units of the troposphere, it is particularly
:important to know its characteristics.
The lower layer of the atmosphere immediately adjacent to the earth's surface,
is called the troposphere. The troposphere is characterized by extensive horizontal
and vertical air currents. The parameters of the troposphere, i.e., the tempera-
ture, pressure, density, viscosity, etc. vary with height. The layer above which
the temperature remains COnStint is assumed to be the upper boundary of the atmo-
sphere. Since the atmospheric conditions vary substantially according to the sea-
son, weather, geographic location, etc., while the instruments and aircraft must
? function in different seasons and at various geographic points, it has been agreed.
that in calculations, designs, and tests or airplanes and their equipment, a certain
fictitious and conventional atmosphere with certain definite parameters, will be
used. This atmosphere is known as the standard atmosphere (SA). The standard
atmosphere gives an arbitrary law of variation in pressure
Po
temperature T, den-
sity p, y, etc, with the height, and also definite initial values of these quanti-
ties, po, To, po, yo, etc. corresponding to zero height.
The altitude with a corresponding initial temperature To = 150 -1-
n
initial pressure Po = 760 mr HE, mass density po - 0.125 kg-sec'/m' and gravimetric
density yo 1.225 kg/m3 is accepted as the level of the earth's surface, i.e.,
zero altitude H. For the upper limit of the troposphere a height equal to
11,000 m has been assumed.
Within the boundaries of the troposphere, the temperature varies according to
the law
273? = 288?C,
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where T is the absolute temperature at the height lit
is the temperature gradient, i.e., the change in te4prature per unit
height;
.., 0.0065 degrees/meter.
It is assumed in the Calculations that the temperature may vary within the
range from - 60 to 50?Q.
Within the limits of the troposphere the pressure varies according to the law
?
pI Po
where p is the pressure at the he 11;
R is the gas constant for air, equal to 29.27 - ?C.
By solving equation (11.2) for H, we find
Tor PH
s?.-- Po )
i;quation (11.2) is termed the standard barometric formula, while e .(11.3)
represents the standard hypsometric formula.
In the theory of aircraft instruments the so-called Laplace formula is tome-
.times _used. This has the form
mean ]
[Po
II 12 18,400 1 * log ?
273.PH
In contrast to the standard formulas, it is assumed in the derivation of the
Laplace formula that the temperature does not vary with height and is equal to some
mean value tmaan to + tH where to and tH are the temperatures at zero height and
2
at height H, respectively.
The values for the density of the height H are calculated by the formulas
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YR
For heights exceeding 11,000 m, the temperatur
varies by the law:
.5?C, and the pressure
(this formula has been obtained under the assumption of constant temperature in the
stratosphere).
It is assumed of the SA that the air is dry, meaning that its humidity is not
taken into account. In reality, the air contains a large quantity of moisture,
which causes the corrosion of the instrument parts, leads to leakage of electric
current in electrical instruments, creates water cushions in the bends of pipelines,
considerably distorting the instrument readings, and, finally, at low temperatures
may lead to icing, which is fatal not only for the instrument but for the entire
aircraft.
The variation in To p, y, and other parameters of the atmosphere, the variation
in timidity, etc. lead to errors in the instrument readings.
The function or the designers of an instrument includes the careful study of
the probable variations in the parameters of the atmosphere under service conditions
and prevention of possible errors in the instrument readings due to the influence
of these varying conditions or, at least, the minimizinv. of these errors.
Section 3.
OPERATIONAL REQUIREMENTS
The basic operational requirements to be met by aircraft equipment are simplic?
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ity and convenient control and use during flight.
This is accomplished as followst
1. By rational location of the instruments on the aircraft and on the instru-
ment panel. The instruments are grouped according to their purpose and are placed
in accordance with the requirements of a definite sequence of visual observations
and manual operations. In placing the 'instruments, cross vision and, more important,
reversed vision and motions must be 'avoided.
2. Protective blocking signalization, and automation. Even the most expedient
location of .the equipment and control levers and the use of automatic blocking and
signaling does not relieve the crew members of the need to follow the readings of
certain instruments. A reduction in the 5train on the crew is achieved by automa-
tion of the equipment.
3. Clearness and sharpness of the readings.
4. Convenience of approach, allowing periodic inspection, replacement of dam,-
aged instruments etc.
In calculating and operating instruments, it is necessary to remeMber the in-
fluence of vibration, accelerations, and the Iie.
Vibrations and Overloads
Vibrations, inclinations, and overloads in the presence of unbalanced parts
and units. may lead to extensive errors in the instruments. For this reason, indi-
vidual parts and units as well as the instrument as a Whole are balanced. In
addition special care is taken to avoid the possibility of coincidence of the
natural vibration frequency of an instrument with the frequencies of forced vibra-.
.tions that arise on the aircraft.
Al]. aircraft instruments are tested for resistance to vibration, .e., for ab--
:zence a- distortions in the instrupent readings under the influence of vibration,
and for Stability.:against vibration, by which we mean the trouble-free operation
,of. the instruments during -a certain definite period under an assigned Vibration.
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te.
The approximate values of the vibrational overload for various units of an air-
craft are given in Table 1. The instruments installed in the corresponding places
of an aircraft are calculated on the basis of this Table.
Table 1
Maximum Vibrational Overloads of Various Aircraft Units
Name of Unit
Aircraft engilie
Engine frame
Fuselage
Non shockproof instrument panel
Shockproof instrument panel
Vibrational Overload
2.5-4
1.5-2.5
0.6-1.5
Dust
For protection against dust which is raised in large quantities during take-
'off- or landing, careful hermetic sealing is .provided for the instruments*.
In addition, wherever necessary, special filters are installed (for example,
in the bearings of gyroscopic instruments).
11.
Section 4.
ASSIMLY REQUIRLAIENTS
The design of an aircraft and its equipment is intended to meet certain flight
:Missions. .These must include the realization of the closest possible design connec-
tion, Which is characterized in particular by the assembly requirements, namely: the
outside dimension and weight requirements, as well as matching the objects and sys-
tems of aircraft equipment with each other and with the design parts of the aircraft
Hermetic closure has a considerably broader purpose than just protection from
dust, for example, it isolates the internal cavity of an instrument from moisture,
isolates the instruments from the action of the atmosphere, etc.
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,as a whole (coMbination instruments connection of the autopilot and the gyroscople
sight, etc.).
The basic assembly requirement for aircraft instruments is the requirement for
minimum dimensions (still preserving convenience of use and reliability of opera-
tion). The total weight of aircx4ft equipment does not exceed 5-8% of the gro s
weight. The established 'ethods of calculations and design of aircraft instruments,
the properties of the materials used in these instruments, the requirement that these
instruments be given minimum dimensions, inevitably leads to a reduction in the in-
strument weight. revertheless, the designers and technologists, in designing and
manufacturing aircraft instr. .ts, must also strive toward maximum reduction in
weight, while still preserving reliability and convenience of operation.
In addition to the above requirements, the individual items of aircraft equip-
ment must also have the following features:
Remote reading, i.e., the instruments must include a device with which their
readings may be transmitted to an indicating system located at a place convenient
for observation or plotting. For example, the readings of all instruments control-
ling the operation of the aircraft engine must be transmitted to the instrument
panel of the aircraft; many elements of the navigational
t.L CI' rn
a N.
nf flight (the
course of the airplane, the air speed, etc.) are of interest not only to the pilot
but also to other members of the crew, and for this reason the values of these
quantities must be'simultaneoUsly indicated by instruments installed on the instru-
ment panels of individual crew members for example, pilot 'and navigator), etc.
The remote.connection is designed to transmit the instrument readings. -Since
each value of the quantity to be measured must correspond to a definite instrument
reading, * the distance transmission must assure the unambiguous correspondence of the
:values at input and output.
In the aircraft instrument building industry various systems of transmission
are used, operating on direct and alternating currents. The most widely used are
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the potentiometric remote reading and the induction remote reading with all their
lmodifications (selsyns magnesyns, etc).
Safety from fire, which is ensured by so d sign ng the instruments that they
'cannot cause a fire under ,any circumstance (this requirement is particularly import-
ant for instruments controlling the fuel-feed system of the engine).
Many of the above requirements are, in most cases, common to all instruments
(for example, elimination of te effect of vibrations, resistance to humidity, etc.),
while others apply only to some tye of instrument (for example, instruments
placed in the cabin do not have to be streamlined, etc.).
For this reason the requirements to be met by aircraft instruments are custom-
arily divided into general and spec al requirements. All aircraft instruments, with-
out exception, must meet the general requirements, while the special requirements
take account of the specific nature of a given instrument.
Section 5.
GENERAL ILT.)4UIRLFZENTS ?DR A BICRAFT IIST13.1317.7E! S
The general requirements are as follows:
1. Dependability and accuracy in operation are the basic requirements, on which
proper utilization of the flight characteristics of the aircraft and the safety of
flight depend to a considerable extent;
2. Convenience of operation, i.e., simplicity of observation, ease of handling,
simplicity of repair, installation, etc;
3. Minimum outside dimensions weight, and cost of the instruments, in accord-
ance with the Standards;
4. Trouble-tree operation of the instruments;
a) in the temperature range from - 60 to + 50?C;
b) under vibrations with a frequency ranging from 20 to 80 cps, and with an
amplitude at which the overload attains 1.5 G (this requirement refers to the in-
strument pointers);
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,L? 44.
under a humidity ranging from 0 to 100%.
Section 6.
ERRORS AND coRgmIgNs OF INSTRUMMTS
The errors and corrections are equal in absolute value, but have different
signs.
We will call the difference between the true value of the measured quantity
and the reading of the instrument the correction. In order to determine the true
value of the measured quantity from the reading of an instrument, the correction
must be algebraically added to the instrument reading.
No instrument readings are free from errors. According to the causes responsi?
ble, errors are divided into systematic and instrumental.
Systematic errors are errors caused by the method of measurement selected.
These errors appear as a result of using indirect methods of measurement in the
aircraft instrument construction in the majority of cases,
i.e.,
a method that does
not measure the quantity of immediate interest for the aircraft crew, but some other
quantity functionally related to it. This is explained by the fact that we still
have not learned to measure directly the majority of values subject to measurement.
-For example, if we desire to measure the altitude, we use a barometric altimeter or
a radio altimeter. In the former instrument,
instead of the altitude itself, the
pressure at this altitude is measured; in the second instrument, instead of the
Altitude, the time necessary for the passage of a signal sent from the aircraft to
the ground and back to the aircraft, where this reflected signal is picked up by a
-receiving device, is measured. Both the pressure at the altitude in question and
the time necessary for the passage of the signal transmitted by the radio altimeter
depend not only on the altitude itself, but alsoon the state of the atmosphere,
the season, etc.
It is impossible in the majority of cases to allow for all the factors that
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distort the value of the measured quantity; for this reason, the instrument readings
will necessarily include the error dependent on the method of measurement on which
the operation of the instrument is based. In Trking out the design of aircraft
instruments the selection of the method of measurement must be approached with ex-
treme care and great caution. It is necessary to consider the type of problems,
the given instrument, the function of the instrument, and its accuracy of operation
under various operating conditions etc. Having selected a method of measurement,
it is then necessary to detect all causes that could be a source of systematic
errors and as far as possible eliminate these errors in designing the instrument.
Since systematic errors are idue to the method of measurement, it follows that
these errors cannot be eliminated by careful preparation and calibration of the
instrument.
This, however, does not exclude the possibility of introducing into the design
of the instrument special compensating devices to eliminate or reduce ?hese errors
(for example, in the air speed indicator and the true air speed indicator). Errors
that depend on the quality of mannfacture of the instrument, the material used for
the parts and units, etc., are called instrument errors.
These errors may be eliminated by improving the quality of manufacture and
operation of the instrument.
Ensuring Reliabilitj and Accuracy of the Readings
As alreadypointed out, the error's of aircraft instruments may be reduced if,
:in designing, manufacture, and installation, the peculiar features of operation of
the instruments in question are taken into account.
Reliable operation of aircraft instruments under conditions of variable tem-
-peratures and pressures is ensured by the following:
a) By placing the instruments away from the sources of heat on the aircraft,
preferably in heated cabins, in which as constant a temperature As pos-
sible is maintained;
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'
b) by using individual heating, heat dissipation, or cooling;
c) by using materials of low sensitivity to temperature changes;
d) by using special devices to compensate the temperature errors and the
errors due to changes in pressure, Such devices are called cornpensator.
The moisture-proof and corrosion-proof operation or air raft instrument is en.-
sued by the following measures:
) By giving the instruments a form making it impossible for water to enter
the body of the instrument;
b) by using materials of the lowest possible hygroscopicity;
c) by coating the surface with water-impermeable and anticorrosion lacquers
and Paints;
d) by using special heaters and moisture collectors.
All parts of aircraft instruments and the instrument as a whole must have a
mechanical strength sufficient to ensure reliable operation under flight conditions,
where aerodynamic, vibrational shock, and other forces inevitably appear and lead
to a vibration and displacement of the pointers of the aircraft i-lstruments, and to
the generation of resonant vibrations in the elastic elements, fastenings, etc.,
and to an accelerated wear of bearings and shafts, to interference with control,
loss of hermeticity, etc.
Mechanical stability of aircraft instruments is attained by placing them at
points of minimum vibration, by using shock absorbers, by preventing resonance, and
by preventing self-loosening of attachments.
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cf.utrmi- III
INSTRUMETS CONTROLLING THE OPERATION OF THE A IRCRAFT ENGINE'
Se Lon 1.
GENAL DATA
Electr f eation of In truments Controlliet the Operation of the.
Aircraft Engine
The gasoline piston engine is an extremely popular type of power plant for the
modern aircraft. For this reason, the present chapter will consider mainly the in-
struments controlling its operation, i.e., manometer , thermometers, fuel gages,
tachometers, etc. As already stated, the nomenclature of control instruments for
other power plants used on aircraft, lifers only slightly from that now under con-
sideration.
In recent years, the control instruments for power-plant operation have been
largely electrified. This is explained by the fact that electrical instruments give
greater dependability of operation and are characterized by compactness, convenience,
and simplicity of installation, and are easily converted to distant-reading instru-
ments. This latter fact, in connection with the increasing size of many aircraft,
is becoming more and more important.
These instruments are fed by the board electric system of the aircraft. The
source of electric energy on the aircraft includes the aircraft generators driven
by the aircraft engine and storage batteries. The board electric system ordinarily
has a voltage of 27 v.
To maintain the voltage at a constant value at varying load in the circuit and
at varying engine speed, and to give a possibility of simultaneous operation of the
:generator and storage batteries as well as to prevent overloading, a junction box
is used. Its basic parts are: 1)-a voltage regulator; 2) a maximum relay to prevent
overloads; 3) a minimum relay to ensure parallel operation of the generator and the
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storage battery.
For installation of the entire electric equipment on an aircraft, special elec7
trio wiring is laid the total length of which may reach several tens of thousands of
meters.
In case of need, for example, to protect the radio installation from interfer-
ence a wiring is selected which, besides the rubber and cotton braiding- comprises
a braiding of copper wire. In installing electric instruments it is necessary to
provide for the protection of the wire from mechanical damage, the action of con-
densed moisture, heat, oil, etc, the absence of interference with radio reception,
i.e., the shielding of the instrument itself and its wiring, the observation of
fire-safety rules, low weight, small outside dimensions, and low cost.
Electric v,ethods of easu.rin Nonelectric Co2antities
Process in electrification of the instrument equipment was stepped-up consider-
ably in connection with the development of automatic aircraft equipment. Almost all
instruments installed on an aircraft for control of engine operation measure nonelec-
tric quantities by electric methods of measurement. In this case, the measured non-
electric quantity is either converted into the corresponding values of electric
parameters of circuits fed by an external voltage source, or is directly transformed
into an electric quantity.
Examples of a conversion of the former type are instruments based on the
changes in electric resistance (inductance capacitance, etc.) of electric circuits,
depending on the change of the parameter being measured. The resistance may vary
with the temperature or as a result of the displacement of a movable contact under
the influence of a change in pressure corresponding to the deformation of an elas-
tic element, etc.
Examples of a conversion of the latter type are instruments based on thermo-
electric piezoelectric and other effects.
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The Basic Elements of Electric Measurin Instruments for Control of
Aircraft Engine Operation
In the general case, an electric measuring instrument measuring a nonelectrical
quantity picks up the quantity to be measured and transforms it into another quanti-
ty which is convenient for remote transmission to the indicating part of the instru-
ment. For this reason such an instrument may be conceived as consisting of the
following parts:
a) A sensitive element represent ng a pi.rt of the instrument that picks up the
quantity to be measured and transforms it in o a displacement. The design of this
part of the instrument is to a considerable extent, determined by the character and
limits of variation in the quantity being measured.
The transformation of one quantity into another may take place not only in the
sensitive element, but also in special transducers the purpose of which consists in
the transformation of the level of the quantity to be measured or the type of energy.
For example, if the quantity being measured is the temperature, and its variation is
measured by the aid of an ohmic resistance, then the transformer or transducer will
transform the variation of thermal energy into a variation of electric energy. As
a result of the variation in temperature, the value of the resistance of an electric
circuit is also varied; consequently, with a constant source of emf, the value of
the current flowing through this circuit will also vary. It is precisely this vari-
ation in current that is recorded by the indicating instrument which is calibrated
in units of temperature.
? In the indirect methods of measurement, instead of the temperature, the measur-
ing instrument measures some quantity functionally related to the temperature, such
?as the variation in current. In this case, systematic errors are unavoidable, In-
!deed, the current may vary not only under the influence of a variation in tempera-
ture, but under the effect of other causes, e.g., fluctuations in the feed voltage.
b) A remote transmission which transmits an impulse, proportional to the quan-
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tity being measured to the indicating part of the instrument.
c) An indicator which reproduces the quantity being measured in the required
units. Most often DC electric measuring instruments, magnetoelectric galvanometers,
logometers, etc. are used as indicators.
d) A source of energy serving mainly to amplify the impulse received from the
object being indicated.
In many instruments the individual elements are combined and are installed in
a housing. In some cases certain parts may, be omitted altogether. For example, ir
instruments where the pLwameter being measured is directly transformed into an elec-
tric quantity, thtre is no need for an outside source of electric energy. We as-
sume that the electric instruments controlling the operation of an aircraft engine
consist of the following elements: a pickup (receiver), a remote transmission and
? an indicator.
? In this case we must not forget that all the above functions of an, electric
measuring instrument are preserved and that the individual elements are present in
? a concealed form. For example the function of a transducer ray be performed either
by the pickup or by the remote transmission.
The na-Aat variety of pickups designed to measure temperature, pressure, fuel
supply, etc. is striking.
The pickups of various instruments, even those designed for measuring one and
the same quantities differ substantially and must meet entirely different require-
ments (for example, pickups of resistance thermometer and pickups of thermoelectric
:thermometers).
At the same time, the indicators of various instruments designed for the meas-
uring of various physical quantities (temperature, pressure, fuel supply, etc.) are
characterized by great uniformity. Electric measuring instruments of the type of
galvanometers or ratiometers, are used as indicators for instruments controlling
the operation of the aircraft engine, electric measuring instruments, which
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are very widely used in measurement technology, are here employed.
ueh measuring instruments also are used:as indicators of many piloting-naviga
tional instruments, such as the potentiometric remote-reading compass, the radio
compass, etc.
Measuring instruments have very Much in common with respect to their design,
operating requirements to be met, installation, etc.; for this reason, the measuring
instruments used on aircraft will be discussed here separately, so that no later
reference to them is required.
Section 2.
t? ELECTRIC MEASIMIEG INSTRUMEUS FOR CONTROL OF POWER-PLANT OPEnATION
Magnetoelectric Measuring Instruments
The operation of these, instruments is based on the generation of a torque by
a the interaction of a permanent magnet and a current flowing through wires (the so-
called frame or coil) located in the field of this magnet.
In these instruments, either the frames or the magnets may be moving. The use
of instruments with moving miniature magnets has become possible only after the
creation of high-coercivity alloys, which provide a high torque at relatively small
dimensions of the moving magnet. A transitional stage is represented by instruments
with a fixed intra-frame magent. Since instruments with magnets inside the frame,
movable or fixed, are considerably more compact than instruments with horseshoe mag-
nets, they are becoming more popular than instruments with horseshoe magnets. This
applies both to galvanometers and to ratiometers. If the torque is dependent on
the current in a conductor, and the opposing moment is created by a special device
;(a coil, a permanent ma-net, or the like) then the instrument is called a galvanom-
eter. If the torque is dependent on the ratio of current in conductors, and the
opposing moment is created by an electic coil whose function can be taken over by
the frames, the instrument is called a ratiometer. A ratiometer can not have less
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than two frames.
NUDSISINSS121,52-20-E_XESMI14111
The design of a galvanometer with a horseshoe magnet (1) and the moving frame
) is Shown in Fig. 15. The inner core (3) provides a practically uniform radial
magnetic field with the inductance B in the gap between the magnet and the core.
The interaction of the conductor under the current (the frame) and the magnetic
field leads to generation of the torque M.
H11
i: ill
II
III
III
IN
-
Fig.15 - The Moving System of a
Galvanometer
1- Magnet; 2- Frame; 3- Core; - Bal-
ancing load; 5 and 6- Springs; lp- Ac-
a pivot 0.3-2 mm in diameter; the pivot is
tive length of frame; bp: Width of
made of special steel with jewel bearings
frame.
in precision instruments and bronze bear-
ings in less precise instruments (these are relatively seldom).
The form of the frame is determined by the purpose of the instrument. If the
The rotation of the frame is impeded
by the springs (5) and (6) which provide a
moment balancing the torque M, and serve
at the same time as current supplies (so-
called moving current taps). The balanc-
ing is effected by the counterpoise(4). To
set the needle of the instrument to the
zero division of the scale in the neutral
position, a special corrector is used
whose rotation changes the pont of attach-
ment of one or the hairs.
moment of inertia of the frame must bes mall i.e., if the instrument is intended
?
for recording rapidly varying phenomena, the frame is made in the form of narrow
lind long rectangles (in individual cases the frame may have the form of -a single
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conductor, in which case its natural frequency will be high and its moment of iner-
tia low). The frame represents a coil, most often of copper (temperature coefficient
t
0.004) or of aluminum (at 0.00423) wire, which is wound on the spider. The
spider is made of metal in cases where high damping of the instrument is necessary,
which is achieved by the action of eddy currents arising in the body of the spider.
Sometimes the frame is made without a si5ider to lighten the moving system.'
The magnets used are of the horseshoe type with special pole pieces, assuring
the necessary law of distribution of magnetic induction in the gap and allowing the
scale of the instrument to be expanded if necessary.
The magnets inside the fras,e of the instrument are given the form of cylinders
(Fig.16). The outside dimensions of such a system may be as much as 15 - 30 mm.
The value of the torque applied to the frame of the galvanometer may be found from
the following considerations: The force
of interaction F1 between the current in
the conductor and the magnetic field with
inductance Bo applied to one side of the
frame, may be found from the equation
Fig.16 - Design of Galvanometer with
Magnet Inside the Frame.
(ma)
where lp is the active length of the frame,
i.e., the part of the frame covered by the
radial magnetic field;
1.- Magnet; 2 - Iron ring; 3 - Armature;
4 - Outer magnetic circuit; 5- Movable I is the current in the frame
frame; 6 - Current feed
w is the number of turns on the
, frame. The same force acts on the other side on the frame, i.e., there exists a
couple F1 applied to the frame, the width of which is bp.
The torque applied to the frame is equal to
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i0 otp )/1:4.
The quantity S... it 1 b is called the active area of the frame. By substituting Sn
P P P
At an area Qp of the cross section of the frame within the boundaries of this
area the number of turns to be wound (w) depends on the size of these turns (9);
: the smaller the size, the greater the number of turns. Thus, to increase the torque
;live must attempt to reduce q. The area of the cross section of the framemav
p
be expressed by the equation
(m.3)
where K3 is the coefficient of filling of the frame by the conductor, allowing for
? the slack of the windings, the insulation., etc. and is equal to 0.7 - 0.8.
For known values of Qv k, and w, the value of q may be found from the expres-
and may be introduced in the expression for the resistance of a conductor
where is the specific resistance of the conductor material;
length.of the conductor'
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PIO
P7.
where 4-11-. Pi; is the electric power supplied to the instrument.
If the power supplied to the instrument is small (a very frequent case), it is
important to know (the conditions of maximum sensitivity of the instrument, i.e., the
conditions of obtaining the maximum value of the torque Y. for a given minimum power
in that case, it is precisely this value o. the Power P. that must be introduced
in the expressions for w, q, and M.
? Fig. 17 - Galvanometer Frameinthe Gap Fig. 18 - Relation of the Stationary
? of the Magnetic System. Moment of a Galvanometer and
the Value of the Air Gap.
6/2 - Gap between frame and magnet;
6 - full air gap; M - Stationary
? moment.
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U, in spite of this, the torque is still too low its further increase is a-
chieved by selection of the optimum value for the gap in the magnetic system. If the
'thickness of the cross section of the frame is h and the length is d (Fig 17), then
Qp dh
We shall denote the gap between the frame and the iron by
gap 8 = 8' 4. h.
By replacing h by 6 and 6', we obtain
, so that the full air
d ( (111. 5)
The quantity 6' cannot be taken too =all, since this may hinder the assembly of the
instrument and lead to the frame being placed beyond the pole pieces.
The expression for the moment n may be rewritten in the form
i.4
10 - 4130Sp rig 61
PP
'With increasing value of 6 the moment increases only to a certain lirnit after which
it then begins to decrease.
The relation l'itween the stationary moment and the value of the gap is Elven in
Piga& At 6 = 6', the moment M = 0. At 6 the magnetic induction in the gap
gradually decreases, which also leads to a gradual decrease of the stationary Y.
The turning moment developed by the frame is opposed by the moment Mi = cd de-
veloped by the hairs, where a is the angle of rotation of the moving system and c is
the coefficient of elasticity of the current conductor.
The position of the moving system is determined by the equation M2= 11 - ca .
the instant of eqrilibrium M2 0, t.e.,
;
kop
1 - --- 111 C
0 P0
46
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? Ji
Galvanometer Erro
and Preventive Methods
The galvanometer readings may be affected by the following:
4) Magnetic -and electric fields created by other instruments; their action. is
,neutralized by shielding.
bj Fluctuations in the feed voltage Of the electric power system of the aircraft,
of as much as ? 10% which ray lead to errors in the. instrument readings of the same
order. To eliminate these errors, the instrument is provided with a voltage etabil-
) Friction in the bearings and Unbalanced state of the individual unit and of
instrument as as a whole, which may lead to considerable errors.
d) Vibrations The influence of vibration on the position of the movable system
Is eliminated by the introduction of special dampers.
e) Fluctuations in the surrounding temperature leading to variation in the mag-
netic flux of the magnet and consequently to variations in the value of the mag-
netic induction Bo in the gap, to variation of the modulus of elasticity E, of the
resistance of the frame R, etc. i.e., leading to instrumental temperature errors.
The variations of certain of the aboveLquantities may be mutually compensated,
entirely or in part (for example, variation of the magnetic inductance in the gap is
partially compensated by variation of the modulus of elasticity); however, other va-
riations remain uncompensated and require the introduction of special compensating
devices.
The most widely used methods of temperature compensation for galvanometers are
follows:
1. The Method of additional resistance. The additional resistance Ra (Fig.19)
is connected in series with the galvanometer; this resistance is made of wire with
'a very low temperature coefficient (nickel, Constantan, etc). Such a resistance may
be considered temperature-independent. The current in the instrument, in this case,
7
'
\
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LI
RA + R
lrR R$ then,even considerable variation* of R will not cause great tempera-
ture errors, but the sensitivity of the instrument is considerably reduced.
2. The Eethod of additional resistance with a negative temperature coefficient.
The introduction of an in-series connected resistance with a negative temperature
coefficient i.e., the introduction of a so-called neutralizer, is the most modern
method of temperature compensation. A neutralizer produces almost no loss of sensi-
tivity of the instrument, since the quantity RD in this ease may be considered in-
significant. The compensation is effected by the selection of a law of variation of
Rd in accordance with the variation of R. The loss of sensitivity will be smallest
where the absolute value of the temperature coefficient e t of the neutralizer ex-
ceeds the value or the temperature coefficient of the material of which the re-
sistance of the galvanometer frame is made.
Such resistances which are used more
and more frequently, are rade of celite
( et 0.01), tellurium with silver
( Fit e 0.012), etc. Alloys exist in which
the absolute value of the temperature co-
Fig.19 - Additional Resistance Method
efficient of resistance is 10 to 12 times
Ra - Additional resistance; R Resis-
as great as the value of the temperature
tance of the instrument; I - Current
coefficient of the material of the frame,
flowing through the instrument; U
for example, silver sulfide. A study of
Feed voltage
this type of alloys vas first made in the
SR by Academician A.F.Ioffe.
3. Method of the thermosensitive shunt. Compensation of the error is afffcted
by selecting the resistanCe r of the shunt (Fig.20), in such a way that on any van-
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fl,tien in the temperature of the surrounding air, the current ip the 'frame will re-
-.mairt. constant.
The resistances h1 and R2 are made of Manganin or Constantan.
The resistance r is made of copper (at . 0.004) or nickel (at . 0.0052). With
this method of compensation, the sensitivity of the instrument is likewise consider-.
-ably decreased. The value. of the current I flowing through the instrument, is de-
mined by the relation
where I0 is the current of the voltage source;
r is the shunt resistance;
R is the resista4ce of the instmlents;
R2 is the additional resistance.
By varying the resistances la and
wide Units.
? 4. Lethod of the thernosensitive magnetic
magnetic compensation of a galvanometer.
The magnetic flux of the magnet in this
ease consists of three components: the
the instrument nay be regulated within
Fig.20 - Scheme of Temperature Compen-
sation by the Method of Ther-
mosensitive Shunt
Fig.21 Thermomagnetic Compensation
of Galvanometer
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,Iworking flux .0, the ux in the shunt (tosh and the flux of dispersion Sd. The mag-
netic induction in the gap is Bo. When the temperature is increased, the magnetic
induction Bo in the gap decreases, reducing the stationary moment cf. eq. (111 .2)
Ori the moving 8y8tem.
thermomagnetic shunt is made of material in which the resistance to the pas-
sage of magnetic lines of force is increased with increasing temperature (calmalloy,.
'alloys of iron, nickel, nickel chromium, etc). As the temperature increases, the magnetic
flux from the shunt is forced into the working gaps of the magnetic system of the
galvanometer, as a result of which temperature compensation is achieved.
The application of a thermomagnetic shunt involves a reduction in the magnetic
flux in the gap, and consequently also a 20-30 % reduction in the sensitivity of the
Anstrument.
Magpetoelectric Ratiometers
Such ratiometers represent a permanent magnet in the field of which are placed ? ?
_Htwo or more conductors (frames); a current is applied to the frames through special
.springs with very low elasticity. The value of the current applied to each of the
frames is determined by the nature Of the change in the measured parameter. ? The
;frames are connected in such a way that the direction ofthe current in them, and
consequently the direction of the torques arising under the action of the magnetic,
field on each of the conductors, is, different.
In ratiometers with two frate.s, the torques act in opposite directions to each
.other.: 'There are no special dvices in the instrument for creating a counter mo-
Anent. In the galvanometer, the counter moment, the value of which depends on the
iangle of rotation of the moving system is created by spring-filled current feeders.
---,In order to make the position of the moving system of the ratiameter stable, it is
--j,necessary that at least one of the torques acting on it be dependent not only on the
Jquantity being measured but also on its position. Without satisfying this re4uire-
ments a, stable position of the moving system of the instrument is impossible=
STAT
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.4;
?
A stable position may.be. attained, for example by selection of a certain law of
:distribution of the magnetic inductance 00 in the gap. In this case the frame be-
comes, as it were, an electrical spring.
The value of the moment M applied to the moving system of the instrument is de-
termined by the ratio M 111 - M2, where Mi and M2 are the moments acting on the
'first and second frames, respectively. These moments may be found fram the equations
m I
10- 4001 SI a will . alliB
M.) 10 - 30
?2
41 W,12
a9I2B01
where .):and I2 are the currents n thefirst and second frames respectively;
1
Bo' and 302 are the magnetic induction in the gap at the locus of the first
and second frame, respectively; for symmetrical frames,
a1 = a2 )1W1 P2
Consequently, in position of equilibrium, when M, . M2, we have
.f)B
?1 - 02
B?1
(III.10)
The values of the currents I and 12 vary it accordance with the variation of
the quantity being measured. To assure stable equilibrium of the moving system, the
_
_Anduction in the gap when the system is displaced must vary by a law inversely pro-
portional to the law of variation of the currents. Consequently, the frame with the
)1ighest current must rotate in such a way that the moment acting on it decreases
,..414ith decreasing magnetic induction in the gap. In this case, the moment acting on
theiraMe with the lower current oust increase as a result of the increase in mag-
netic induction in the PP04,
?
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For horseshoe magnets, the required law of distribution of magnetic inductance
in the gap may be obtained by experimental arrangement of the inner magnetic conduc-
tor (Fig 22) or by choice of the form of the pole pieces. In instruments with a
Fig.22 Ratiometers with Two Frames
A- With a horseshoe magnet; b. With a magnet inside the frame;a - Angle of
-rotation of the moving system; 6 - Angle between the frames; Current
flowing through the first frame; 12- Current flowing through the second frame
cylindrical magnet inside the frame the problem is simpler, since the induction in
-the gap of such a magnet is 44stril,wEckii by the cosine law
where 6 is the angle between the frames;
a is the angle characterizing the mutual position of the moving system and
The equation of equilibrium 111 112 now takes the form
Ilk COS 2 4A, cos (a + 4/30(cos a cos &? sin 251n 0.
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'11Gisa,
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Whence
1g2
(4, Cos ? I, sin 2 sin
cost?
/2
t, cos
slur,
The angle of rotation of the moving system m,
pression
tOn
r be calculated by the aid of the ex-
cos 5 -
T-2
sin 6
(III.12)
This equation shows that the value of the angle 6 between the frames affects
the value of the angle of rotation of the 7 ving system so that the angle 5 must
therefore be selected in such a way as to obtain the maximum value of the angle 6
while still maintaining stability of the instrument readrs.
merits, this angle is most often taken between 8 and
300.
In aircraft instru-
.Errors of Ratiometers and Methods o Eliminatinr Them
lqueuat4ons in Feed Voltage.
fixing the value of the angle 6, the position
?
of the moving system is determined by the ratio of the currents 11/12 in the frames
?which is practically independent of the fluctuations of the feed voltage.
Temzerature Errors. As in the galvanometer, a change in the temperature of the
. surrounding air leads to temperature ? errors in ratiometers. To compensate instru-
ment temperature errors, a neutralizer is sometimes connected in series with each
of the frames; but more often an additional resistance is introduced into the gen-
eral design of the instrument. ,This resistance is made of two materials,the re-
sistance of one of them remains almost constant with varying temperature (Constan-
,tan), while the resistance of the other varies considerably (copper). This method
'of'.compeneation is particularly widespread in bridge instruments.
A Comparison of galvanometers and ratiometers permits the following conclusion..
Ratiometers have an advantage over galvanometits in that their readings are
?Py,
53
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almost Independent of the fluctuations or the feed voltage, but the sensitivity of
galanometerf in considerably higher.
Both in ratiometric (two-frame) and galvnxnetrtc instruments, the angle of
rotation of the pointer, does not exceed 900.
To extend the scale of the instrument, specially shaped pole pieces are used,
which vary the law of distribution of magnetic induction in the gap; in this case
the axgle of rotation of the pointer may go as high as 270?.
Fig.23 - Three-Coil Ratiometer with
Moving Magnet
1 - Magnet; 2 - Coil; 3 - Damper;
4 - Screen
Fig.24 - Scheme of Arrangement of
Ratiometer Coil of Three-Frame
Ratiometer Coils on the Rotation
of the Moving System through 360?
To obtain a 3600 scale ratiometers with three frames are used. In some cases
the frames are made 'movable in other cases the magnets.
Figure 23 shows a three coil ratiometer with a moving magnet. To reduce its
weight, the miniature magnet is made from molded special magnetic powder. The re-
duction in the dimensions of the magnet inevitably leads to a reduction in the
magnetic induction in the gap. This is compensated by a considerable increase in
the outside dimensions and number of the windings of the frames which,' being fixed,
maybe made fairly heavy .and large. In its design, such an instrument is simpler
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'1 Yi10'?
, v r a,m,0
'
r- 7,no ? I
tilfZ";."""
t
4,1 4 A. 1 444 ?, ... I, ,..
ab
{ .1.2 ;:ri;,::, "Y'''' ti,h{.':.' .:b
1. t:.4,.-,,,, ,:tv. ,?.i, tr t ' 1 .,.!. 1... !.:11."
,?, elqui.1?. to
r
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sin 120 cos + cos 120 sin d
sin (a 120)] "
4.?ri (13 COS cos a
i(1.3- 12)
Equation 1I1.14) makes it possible to calcOAte the scale of the ratiometer.
:lhe instrument readings, as in the case of the two-frame ratiometer, depend on the
ratio of the currents in the frames.
The errors of three-frame ratiameters are similar in character and nature to
Hthe errors of two-frame instruments.
Section 3.
INSTRUMENTS FOR MEASURING PRESSURES AND TEMPERATURES
The Vacuum Manometer
The manometric instruments used a few years ago for checking the operation of
ia power plant have almost completely gone out of use, and for this reason these will
The only manometric 'instrument still widely used at the present time is the
-%vacuum gage, an instrument for measuring the pressure o, the fuel mixture in the
intake manifold of the engine. This instrument is used in engines with super-
harging where, in order to maintain a definite, composition of the fuel mixture,
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regardless of the altitude, forced air supply (supercharging) is used.
'The diagram of connection of the vacuum gage ia given in Fig..25, And the kine-
matic diagram of the differential vacuum gage in Fig.26.
The sensitive element of the vacuum gage co sist.of two sylphons .(2) and (3) -
_lwhich are rigidly attached to the hollow cylinder (1). The air is pumped out of the
sylphon (2) While the sylphol (3) is connected to the intake manifold (2) of the
engine.
The rigid center of the sylphon c3) is displaced as the pressure An it varies
under the influence of thepressure difference p - oil. The motion of the rigid
center is transmit ed over a transmitting mechanism to the instrument pointer (4).
Diagram of Connection of Fig.26 - Kinematic Diagram of Differ-
the Vacuum Manometer ential Vacuum Manometer
So long as the atmospheric pressure pH remains unchanged, the sylphon (2) is
6-inOt deformed, and only sylphon (3) operates. As the atmospheric pressure varies
vl_Ithe value of the force acting on sylphon (3) from outside also varies and the sen-
sitive element is deformed.
here-are,no-arrore in the instrument readings since, because of the variation
-1
lheciamospherie pressure and of the pressure in the body of the instrument the
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,s 1 on ( ) is also deformed; thus compensating the deformation of the sylph= (3).
The counterweight (5) is used to balance the instrument.
errors of the 1t3J Oi
As a result of the variation in the elasticity of the sensitive elements under
*the action of the sUrrounding temperature;so?called instrument temperature errors
are created in the instrument. In order to eliminate these errors, a bimetal tem,
perdture compensation is provided, the design of which was described in the discus-
sion on manometric navigational-piloting instruments (cf. Chapter IV, Section 1).
'Sometimes instruments with sensitive elements in the form of a corruEated box are
.used on aircraft. The design of such an ins,,rument is similar to that of the alti-
meter, 141-doh is described in Section 1 of Chapter V.
*MY
Fig.27 - Diagram of a Combination Instrument Operating on
the Principle of Resistance Measurement
-Pressure pickup; b - Temperature pickup; c Indicator
)
Electromechanical combination instruments are widely used on aircraft to meas-
!ure pressures and temperatures.
?The_principal,advantages of such instruments are simple des convenient
reliability, and compactness.
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-
1,
amppund Instruments for Measurin, Pressures or Tem eratures
? CompoundAnstruments whose indicator combines two to four instruments are wide-
ly
used; 'these have separate pickups for each quantity to be measured. Such instru
-ments lighten the work of the aircraft crew since they reduce the number of objects ,
to be observed.
The operation or electromechanical instruments for measuring pressures and tem-
peratures is based on the -variation of resistances in relation to the values meas-
ured.- A circuit diagram?of one of such instruments and an over-all view of the in-
dicator are given in it .27. The pressure p IS supplied to the elastic element (1)
-,having the form of a corrugated box or .sylphon.
If the pressure is being measured, then p represents the oul;ntitv measured
?(pressure of oil, gasoline, coolant, etc.).
T f
the '''?nqtrument is intended for
measuring temperature, then the pressure p must be a function of the temperature
loeing measured.
In this case the instrument must have a device for nicking ur the temtrrature
and converting it into pressure.
form of a thermoscartridge
Such a device may be made, for example, in the
filled with ligroin, the volume of which varies consid-
The inner cavity of the thermoscartridge communicates
'erably with the temperature.
element having the form, for example, of a corrugated box
As the temperature varies, the volume of the ligroin also varies causing a
0._deformation (or flexing) of the elastic element.
/1
?T
mechanism to the slide block (2), wiping a potentiometer. Each position of the
The deformation of the elastic element (1) is transmitted over the transmitting
iv slide corresponds to a definite voltage applied to the indicator, designed as a
44rn
galvanometer or ratiometer.
Very frequently a combination ratiometer-galvanometer
is used as indicator, in which case threeor four indicators are combined in a
(Fl..Z.? 270
Compound_instruments that measure orly the pressUre or only the temperature ,
59
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'syl on ( ) is also deformed, thus compensating the deformation of the sylphon (3).
The counterweight (5) is used to balance the instrument.
Errors of latun Marometer
As a result of the variation in the elasticity of the sensitive e1eents under
the action of the surrounding temperature, o-called instrument temperature errors
are created in the irstrument. In order to eliminate these errors, a bimetal tem-
perature compensation is provided, the design of which was described in the discus-
sion on manometric navieational-piloting instruments (cf. Chapter IV, Section 1).
Sometimes instruments with sensitive elements in the form of a corrugated to.: are
used on aircraft. The design of such an instrument is similar to that of the alti-
meter, which is described in Section I of' Chapter I.
1
P=1.(1)
Fig.27 - Diagram of a Combination Instrument Operating on
the Principle of Resistance Xeasurement
Pressure pickup; b - Temperature pickup; c - Indicator
Electromechanical combination instruments are widely used on aircraft to meas-
ure pressures and temperatures.
-1
The of such instruments are simple design, convenient_in-
,tstallation0, reliability, and compactness.
e
;
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prImA r ,1118,.41,
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lempund Instrume,nts for Measurinz PressuresAorTepa
Compound instruments whose indicator combines two to four instruments are wide-
used; these have separate pickups for each quantity to be measured. Such instru-
ments lighten the work of the aircraft crew 3ince
to be observed.
The operation or electromechanical instruments
peratures is based on the vz tion of resistances in relation to the values meas-
'
they reduce the number of objeCts
r measuring pressures and tem-
ured..
A circuit diagram of one of such instruments and an over-all view of the
dicator are given in
having the
g.27. The pressure p is
orm of a corrugated box or sylphon.
the pressure
in-
supplied to the elastic elemen (1)
Is being measured, then p represents the quantity measured
(pressure of oil, gasoline, coolant, etc.). If the instrument is intended for
:measuring temperature, then the, pressure p must be a function of the temperature
being measured.
In this case the instrument must have a device for picking
zirld converting it into pressure. Such a device
H'form of a thermoscartridge filled with ligroin,
TerablY,with the temperature.
;vidth an elastic
UD the ter-erature
may be made, for example, in the
the volume of which varies consid-
The inner cavity of
the thermoscartridge communicates
r___
ie7AJ41.4.1d41.r LLuv vine. Lis for e..,eImple, of a corrugated box.
As the temperature varies
the volume of the ligroin also varies causing a
42.,.deformation (or flexing) o, the elastic element.
1. .. .
4;6
The deformation of the elastic element (1) is transmitted over the transmitting
!mechanism to the slide block (2), wiping a potentiometer. Each position of the
'slide corresponds to a definite voltage applied to the indicator, design& as a
Iga1vanometer or ratiometer. Very frequently a combination ratiometer-galvanameter
y2?,4s used as indicator, in which case three
-1
e4 singl,p,body (Fig.27c).
or four indicators are combined in a
_ Compound instruments that measure or y the pressure or only the temperature
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are widely used. An example of such an instrument is the conbined diff rential m n-
oMeter with double or even triple indicators (Fi 28).
In addition to these instruments, bimetal thermometers whose pr..-ncipal part is
.a bimetal coil are also used for measuring tempertures. The deformations of the
coil corresponding to the variation,in temperature permit eetiir4ttng the temperature
(such instruments are used for measuring the temperature of the outs.,de air). Re-
sistance thermometers and thermoelectric th rmaneters are also in use.
ResistanceTherm Mete s
Resi tance thermometers are used,on aircraft for measurini:!, the temperature of
:air, oil, coolant, carburetor Mixture etc.
The design of. &la above-listed instruments .is the same and differs only within
the limits of measurement and the range of the individual. resistances. These in-
struments make use of the relation between the resistance of a conductor and the
:temperature, which obeys the law
no ( atti T)
(III.15)
where RT and Ro are the resistances at the temperatures T and To, respec-
tively, correlated by the expression
T
at is the temperature coefficient of resistance; for copper and nickel,
which are used in these instruments the coefficients are at . 0.004 - 0.00681
The instrument is arranged as a bridge circuit (Fig.29). One-arm, representing
the pickup, is placed at. the point of measurement and is made of a thermosensitive
naterial with a resistance RT, whose value varies by about 0.4% for each Centigrade
iof temperature increase. Three other resistances are mounted in the indicator.
1These are made of a material with a very low temperature coefficient (Constantan,
Alitiptir), Usually, two of these resistances (R) are given equal value 'while the
60
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2 MAJ-80 3 Enyx-so
Pickup Indicator
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,
L
*
third (RD is so selected that, at the equilibrium state of the bridge, equtjr
In one diagonal of the bridge, a galvanometer or ratiometer is connected (in this
particular case a ratiometer), and in the other diagonal, the feed source. The re-
sistance Rd steps down the voltage of the electric system of the aircraft to the re-
quired value. The resistance Ra is introduced to obtain currents of different value
II
Fig.29 - Ratiometric Resistance
Thermometer
RT- Resistance of the conductor; r-
sistance of the ratiometer frame; R
and RI- Resistances of bridge arms;
Rd- Additional resistances; Ra al*Ra27
and direction in the radiometer frames.
This resistance is made- of copper Rai and.
Xanganin Ra,), by which- means temperature
cmpensation is effected (Ra -Ra
In the equilibrium state of the
bridge, corresponding to a certain temper-
ature Toquile the equation RR]. = RETeguile
is valid; in this case, the current in the
ratioc,eter frame is the same in value but
differs in direction. Let us assume that
the value of this current is I. For any
other value of the temperature at the
) ?
Re-
point of measurement, the current 11 I+i
~INA\
WA" NFOS-SW
while the
Compensating resistances; Io- Current
r7".W.E1
4.11.Mass%o
the
current12 I i flows through
the other frame, where i is the current
generated by the change in resistance ET
used from ..roltage source; Ill 12 Cur-
of the receiver with any change in the
rent in the ratiometer frame; U- Volt-
? age supplying the feed source.
measured temperature. Depending on the
ratio of the current I to 12 the pointer
of the indicator assumes a certain posi-
tion. ,The scales of such instruments nev-
er have divisions exceeding 160?C. This is explained by the fact that the copper
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tances used in the receivers cannot operate at high temperatures, be-
cause of a change in their cha'racteristics as a result of oxidation. An e.xternal
view of the indicator of the resistance thermometer is shown in Fig.30a? and the cir-
Cony* ? fairtio
Rricoodithinicitaturirrodarrasairlfrnbir
emen t
ft trent pattl of base.
\\ aisambly of coodoctar
ftmetifelt4ii is base
I/t ? Nairtame AlOdi
11/
1/
ig
N. V
Asstimbiy ?'4.f.
arnmelimen I' at roar
part at base
Fig.30 - Electrical Resistance Thermometer
a- Indicator
I- Panel; 2- Piatiometer; 3- Scale; 14- Body; 5- Plug contact;
6, 7, and 00- Resistance.
b- Circuit diagram of resistances on indicator panel.
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,rnOr
cuit diagram of the resistances is given in 1ig.30?b. Figure 31 gives the diagram of
esistanee thermometer of the receiver. A copper wire, 0.05-0.07 mm in diameter,
with enamel insulation, is wound on a paper-Insulated hollow metal tube (2) which is
Fig. 31 - Diagram of Resistance Thermo- Fig.32 Contact Plate of Resistance
meter Pickup. Thermometer
1- Protective tube; 2- Inner tube; 1- Plate; 2- Protective tube.
3- Wire resistance
coated with lacquer and placed in the protective brass tube (1). During operation
of the instrument, a good thermal contact must be provided between the thermosensi-
Fig.33 - Circuit Diagram of Universal Dual Electric Resistance Thermometer
for Multi-Engine Aircraft.
tive resistance (3) and the medium whose temperature is being measured. To improve
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.this contact, a plate is sometimes used instead of .the tube .(2). Thisplate has the
form shown in Fig.32. The lugs of the plat (1) are pressed against the protective
,tube (2) In designing the instruments the possibility of change in the resistance
RT?due to heating by the current itself, must be allowed for and eliminated.
To measure the temperature of water, oil, air, etc. on twin- and four-engine
aircraft, a universal dual resistance thermometer is used (Fig.33). The pickups (1)
of the thermometer do not differ from the pickups of ordinary resistance thermometers
and are insta)led at the point whose temperature must be controlled.
The indica or has two independent ratiometers (2) and two scales graduated fror
50 to +150?C. The m ving magnet . installed within two pairs of coils set at an
angle of 120?. The electromagnetic field of both pairs of coils acts on the magnet.
At low temperatures the current is lower in the first pair of coils and higher in
the second pair. In this case, the action or the second pair predominates and de-
flects the magnet with the pointer downward. When the temperature increases, the
current increases in the first pair and decreases in the second pair, deflecting the
needle upwards.
The variable resistance is intended for regulating the sensitivity and for set-
ting the instrument to zero.
The indicator and pickup are linked to the conductor by means of plug connec-
tors.
Errors of the Resistance Thermometer
The cause of the systematic.terperatUre error is a heating of the resistance or.
Ithe thermosensitive elements by the current passing through it.
Instrument errors are caused by the following:
a) Influence of the temperature of the surrounding medium on the characteris-
tics of the elastic elements frame resistances, etc.
b) Wear, irregularity of the gaps, scale errors and errors due ,to unbalanced
'puts.
65
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n
-
The error from wear appears to be the result of the relatively high friction in
the axle journals, the hinge joints, etc. The error from the Ms is attributed to
the weakness of the hair selecting the gaps in the transmitting mechanism of the in-
strument.
The scale errors are the mechanism control errors and scale markings, i.e., dis-
parity between the angle of rotation of the pointer with the scale divisions.
The unbalanced state of the parts of the transmitting mechanism may cause a
change of the readings when the instrument is inclined and may be due to fluctua-
tions of the pointer under vibration.
c) Oscillations of the feed voltage.
According to the specifications for the instrument, the error in their reading
under constant vibration and normal temperature must not exceed ? 2.5% of the nomi-
nal scale value; at T .* 50?C, this error may go as high as * 5:7', and reach t 6% of
Che nominal value at T - 60?C.
The error of a dual resistance thermometer with ratiometric indicators does not
exceed 6?C over the entire scale.
Thermoelectric Thermometers
Thermoelectric thermometers are used for measuring high t&-14)eratures, for ex-
ample the temperature at the cylinder heads and in air-cooled internal combustion
:engines, the temperature of the exhaust gases in jet engines, etc.
The operation of thermoelectric thermometers is based on the utilization of the
hermal emf generated, in two cold junctions of disslar conductors when a temper-
--p.ture difference AT (Fig 34) arises at the soldering point. This phenomenon was
4'--ilirst discovered in 1756 by the Russian AOademician F.U.Epinus and later, in the
5.0-Jmiddle of the 19th Century, studied in detail-by M.P.Avenarius. In the USSR, P.1.
-
iBakhmetfyev, 1.1 Borman, and others have done and are still doing large-scale work
If- the temperature at; one end of the junction is constant, then the value of
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the thermoelectromotive force, for a given material of the thertnocouple, depends only
on the temperature of the other end.
In aircraft instruments, the temperature in the pilot cabin where the indicator
is installed and the so-called "cold" junction of the thermocouple is located i
Hot junction
Fig.34 Diagram of the Thermoelectric Thermometer,
1- Pickup disk; 2- Galvanometer; 3- Indicator
- Considered constant. Consequently, the instrtiThAnt readings depend on the?tempera-
Table 11
Material of Thermocouple
6 .
-- pick el-nichrome
-Chrome' alumel
HIron-constantan
-,Chromel-copel
1000 42
1000 40
500 27
500 49
up, of the 'Motu junction which is connected with the pickup of the instrument. The
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lmost widely used materials for thermocouples are given in Table 2.
At AT 300?C, the thermoeleetromotive force of a chromel-copel thermocouple
is equal to 20 my, and the thermoelectric current is directed from the positive
(chromel) electrode to the negative one ( opel). Since the thermoelectric currents
Fig.35 Forms of Thermoelectrodes
a- Thermocouple with junction to be screwed into a metal wall;
b- Thermocouple in form Of a disk; c- Sparkplug with thermocouple.
generated during the operation of the instrument are small, the measuring instrument
must possess a high sensitivity, and
this rAnAnn a ral.vanometer with an intra-
frame magnet is Used as the measuring device.
. The form of the thermoelectrodes may Vary widely; according to the point of in,
stillationJt may consist of a disk with two thermoelectrodes, two ends screwed into
-4a metal wall, etc. (Fig.35).
Errors of Thermoelectric Thermometers
Systematic error arises as a result of the inconstancy of the temperature of
-the "cold" junction and the resultant variation in thermoelectromotive force.
To eliminatethis error, either additional thermocouples creating a counter emf
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can be introduced, or the design of the indicator can be supplemented 'by a device to
its readings acoording to the value of the surrounding temperature. Such a
device might be a bimetal spiral (1) which as the temperature varies, changes the
position of the end of the hair (2); see Fig.,6. The moment of the hair is
Fig.36 - Bimetal Compensation of F4z.37.- Hair with Variable Point of
The Thermometer Attachment
1- bimetal coil; 2 - Hair a - angle of twist of the hair spring
due to change in value measured;
a - angle of deflection of hair spring
0
due to deformation of the bimetal
springs.
c (a -a); see Fig.37.
The instrument errors are caused by the variation in resistance of the galva-
nometer, conductors, et. with any variation in temperature of the medium surround-
-1
4-C?Ug the indicator. Compensation elf the temperature error is effected by means of a
'ILentralizer.
?J
etc.
In addition instrument errors also arise as a result of wear, of incorrect
?,Jrhe total error of the instruments TTsT - under constant vibration must not,
69
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-I
The Thermoelectric Thermometer for Measurinflthe Temperature of
-Exhaust Gases of a-Jet Engine.,
peculiarity of measuring the working gases in turbojet engines consist in the
'fact that the range of measurement is great and the temperature at various points of
110111???? ? 0111?14??
ammo -.???
? ?
:??????? ?
4
Fig.38. - Circuit Diagram of Thermoelectric Thermometer for Turbojet Engines
1.-?. Jet -nozzle; 2- Thermoelectric pickups; 3- Junction box; 4- Indicator;
Copper leads; 6- Thermoelectrode leads.
-it may differ. In order to form a correct idea of the temperature, it is measured
- at several points and the mean value of these measurements is then determined.
A circuit diagram of the thermoelectric thermometer is given in Fig.38.
At four points of the reaction nozzle, thermoelectric pickups (2) (hot junction)
are installed; the cold junctions of the thermocouples are placed in the junction
-box (3) where they are connected in series with each other.
The two free ends are
?connected to the galvanometric indicator (4) by the copper conductors (5).
A thermoelectromotive force proportional to the temperature of the point being
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imeasured is generated in each trtocouple. At the ends of the leads (5) connected
to the indicator the total thermoelectromotive force of the four thermocouples is ac-
tive. The indicator is calibrated in values of one fourth of the total temperature
of a single thermocouple and its readino correspond to the mean temperature of the
four points of measurements.
IV MI
4010/CArale
10,
Fig.39 - Bridge Compensation of the Variations in Temperature of the Cold
Ends of the junctions
The resistances R1 and ;-!.3 do not depend on the temperature;
The resistances R2 and R4 vary with the temperature.
The variation in temperature of the cold end of the junction leads to a syste-
matic temperature error which is compensated in one way or another. In some cases,
a bridge temperature compensation, whose diagram is given in Fig.39, is used.
Two arms of the bridge R1 and R3 are made of Yanganin whose resistance is prac-
tically independent of the temperature, while the other two arms R2 and R4 are made
of copper, i.e. of a material whose resistance varies with the temperature. At
zero temperature, all the resistances arc equal, and the bridge is balanced. A vari-
ation in temperature disturbs the balance of the bridge and produces an additional
--:(compensating) voltage in its diagonals, which is algebraically added to the thermo-
4
Helectromotive force of the thermocouple. The range of temperature measurements is
-Jfrom 300 to 90CPC. In designing such instruments, particular attention is given to
the corrosion resistance and electric strength of the pickup elements which operate
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--
Tr.
,
,at high temperatures; it is likewise very iaportant to ensure accuracy of measure-
ment of hiO temperatures. This is accomplished by selection of thermocouples of
'low sensitivity to low temper tures (below 300?C).
Section 4.
LSTRt.ETS FOR YEASULING?THE JANTITY AND COYSUYT IQ1: OF FUEL
-.12a4LItisUilata
Instruments designed for measuring the quantity of fuel it the aircraft tanks
are called fuel gages. The design of oil gages indicating the. Oil supply. on air-
craft' is similar to that of fuel .gage . The same arrangement ray also be used to?
measure the supply of oxidizer in liquid-)et engines, etc. Instrurents based sn
measuring the volume of the fuel by means of a float are widely used.
Since the fuel tacks are located at a considerable distance from the instrument
panel the instrument must be of the remote-reading.type. ,t the present time not
only fuel gages measuring the fuel supply in a single tank are used, but also in-
struments indicating the supply of fuel it all the gasoline tanks, the so-called
"summing" fuel gages.
The operation of remote-reading fuel rages is based on the transformation of a
nonelectric quantity, namely the height of the fuel level, into an electric quanti-
ty acting on a galvanometric or, rat4o.t.--.etric
A float fuel gage is designed on a bridge arrangement and consists of a pickup
and an indicator connected by a remote line (Fig.40).
The pickup (Fig.41,a) is installed in the fuel tank and consists of a float
Isubmerged in the tank,- whose lever (1) is displaced as the fuel is consumed. The
displacements of the lever are transritted to the slide (2), wiping the potentio-
4mtter (3), thus varying the ratio of the resistances ni and of the potentiometer
arms. The corrugated box (4), through which passes the rocker lever (5), imminents
penetration of fuel vapor into the potentiometer from the fuel tank, Thanks to the
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,
a- Operating mechanism of pickup;
I- Float lever; 2- Potentiometer slide
wire; 3- Potentiometer; 4- Corrugated
box; 5- Rocker arm
b- Pickup of float fuel gage in long
narrow tank
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elasticity of the box (4),-the lever (5) is displaced when the f oatjever (1) zhift3,
thust displacing the eliding contact of the potentionstors
The use of such a pickup is inconvenient in cases where the fuel tank has a
long narrow shape (F1g41,b).
The indicator of a float fuel gage is made in the for of a two-frame ratiometer'
,.(or of a galvanometer). The arrangement 61 the resistances on the indicator panel
As shown in Fig.42.
Fig.42 - Fuel Gage Indicator (scale removed).
1- Magnet; 2- Moving frames; 3- Instrument pointer; 4- Resistances
The fuel tanks of aircraft differ in shape, and for this reason the fuel gages
aerving an aircraft of a given type are assigned a definite mark; such a fuel gage
is suitable only for that particular type of aircraft. Depending on the shape of
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the tank, the shape of the pickup potentiometer also varies. To reduce
.or the external magnetic field on the instruments and to eliminate the
jthis instrument on others, a magnetic shield is used In the form of an
jplaced over the body of the fuel gage.
The diagram of the integrating fuel gage is given in Fig.43. The
#2111 are seledted in such a way that the variation in their
of the slide corresponds to
the fuel level in the tank. The
resistances
values with
instrument
set includes a switch used for change-over
of the fuel gage to the measuring of the
quantity of fuel in all tanks. In this
1, case, all the resistances R21, R211 and
R2111 are connected in series; when the
quantity of fuel is measured in separate
tanks, all pickups, except that installed
in the particular tank in which the quan-
tity of fuel is to be measured, are short-
ed.
Errors of Fuel. Gages
a) Systematic errors are caused by the banking and pitching of the aircraft as
Nkup
Fi 44 - Error due to Inclination of Tank
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, well as by the influence of accelerations of the aircraft, acting on the fuel in
the tanks. In inclinations when the tank is tilted together with the aircraft
(Fig.44), the fuel gage reading for one and the same fuel supply will differ accord-
ing to the position of the float. The smallest error corresponds to the case when
the float is in the middle of the tank. A similar influence on the error of the
instrument is due to accoleration, which change the fuel level.
The cause of systematic errors may also be any deviation in shape, dimensions,
installation of the tank from the design values.
b) Instrument errors of fuel gages may be divided into temperature errors,
scale errors, errors due to wear, irregular gaps, fluctuations in the feed voltage,
etc. These errors are compensated and eliminated by the same methods as those u',:.ed
in resistance thermometers. The total error over the entire scale of the instrument
does not exceed 7% of the nominal value.
Electric Capacitance Fuel Gages
Capacitance fuel gages, based on thele of an alternating-current bridge
(Fig.45) have recently come into wider use. The bridge is composed of the induct-
ance (1) transformer winding), the fixed capacitance (2) and tne variable capaci-
tance of the pickup (3), which represents a capacitar. The galvanometer (4) is
connected to a diagonal of the bridge across a rectifier.
The voltage in the bridge diagonal is amplified by an electronic amplifier.
The bridge is fed with higher-frequency AC (4404900 cycles) from a vacuum-tube
oscillator which, in turn is fed from the electric system of the aircraft through
r,
a vibrapack.
In the balanced state of the bridge which corresponds to the instant where
)--there is no fuel in the tanks (or when the tank is completely full of fuel), the
'1 bridge diagonal carries no current and the instrument pointer is at zero. Filling
-1-the-tank-with-fuel-(or-in the latter? case, consumption of fuel) leads to a change
.4
in the' of the capacitor, since the dielectric constants of the liquid
76
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well as by the influence of accelerat onsof the aircraft, acting on the fuel in
the tanks. In inclinations', when the tank is tilted together with the aircraft
(Fig.44), the fuel gage reading for one and the sAme fuel supply will differ accord-
ing to the position of the float. The smallest error corresponds to the case when
the float is in the middle of the tank. A similar influence on the error of the
instrument is due to acceleration which Change the fuel leVel.
The cause of systematic,errors may also be any deviation in shape, dimensions,
installation of the tank from the design values.
b) Instrument errors of fuel gages may be divided into temperature errors,
. scale errors, errors.due.to wear, irregular gaps, fluctuations in the feed voltage,
-etc. These errors are compensated and eliminated by the same methods as those lied
in resistance thermometers. The total error over the entire scale of the instrument
does not exceed 7% of the nominal value.
Electric Capacitance Fuel Gages
Capacitance fuel gages based on the use of an alternating-current bridge
(Fig.45) have recently come into wider use. The bridge is composed of the induct-
ance (1) transformer winding), the fixed capacitance (2) and the variable capaci-
tance of the pickup (3), which represents a capacitar. The galvanometer (4) is
connected to a diagonal of the bridge across a rectifier.
The voltage in the bridge diagonal is amplified by an electronic amplifier.
The bridge is fed with higher-frequency AC (400-1900 cycles) from a vacuum-tube
oscillator which, in turn is fed from the electric system of the aircraft through
a.vibrapack.
48--i In the balanced state of the bridge which corresponds to the instant where
"--1 there is no fuel in the tanks (or when the tank is completely full of fuel), the
bridge diagonal carries no current and the instrument pointer is at zero. Filling
-1-the tankvith-fuel (or in the-Iitter-case'consumption-of fuel)-leadfvto7a-change--
.
t 4 1
'-i-n-the-capacitance- of-the capacitor, since the dielectric constants of the liquid
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1
?
a 6
4- 2 _
i
pickups to be placed in a single tank and permitting a mean readings. The readings
f;
of the capacitance fuel gage depend only ,slightly on the bank or on the vibration
S
and air d !for. The change in the capacitance of the pickup (which increases as
the tank is filled with fuel) leads to an unbalance of the bridge. A current is
generated in the diagonal, of a value prOportional to the capacitance of the pickup
and, consequently, to the fuel level in the tank. The maximum deflection of the
pointer corresponds to a completely filled tank.
FAVAPPANWIR
1.????
Fig.45 - Circuit Diagram of Capacitance Fuel Gage
? Inductance; 2 - Corwtant capacitance; 3 - Variable
capacitance of pickup; 4 - Galvanometer
The pickup of the capacitance fuel gage shown in Fig.46, consists of a cylin-
drical capacitor with an inner electrode .(1) and an outer electrode (2)(in high and.
narrow tanks, the wall of the tank itself may be used as the outer electrode) be-
tween which there is an insulating layer. Between the insulating layer and the
outer electrode is the liquid whose level is being measured. An advantage of the
capacitance pickup is its simple design and its small size, allowing a number of
in the angle of pitch of the aircraft. In addition the readings of such instru-
ments are less affected by variations in the temperature of the surrounding air,
t
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since any change in temperature is acconpamied by a simultaneous'change in fuel
volume and in the dielectric constant.
'A ,avaeitance:pickup can be successfully used in ca es where the liquid is
electrically conductive or is chemically active (which is particularly important
for Uquid-jet engines in which the supply of oxidant must be checked).
One disadvantage of capacitance fuel gages lies in the influence of the ;:on-
n cting leads on the accuracy of measurement of volume; as a result, special
shielded conductors or special connection diagrams must be used.; Another disadvan-
tage is the relative complexity of the indicator.
The indicator of a capacitance fuel gage may be designed as a magnetoelectric
indicator connected across a rectifier, or as a directly connected ferrodynamic
indicator.
In recent times, clectrocapacitan7,e fuel gages based on the utilization of
, self-balancing AC bridges have come into use.
In such instruments, induction motors are used as indicators, operated by the
amplified voltage signals from the diagonals of the bridge.
Instruments 'Measuring Fuel Consption
An indication of fuel consumption is necessary to evaluate the effectiveness
of the operation of an aircraft engine, as well as the consumption of oxidizer in
jet engines. There are flow meters that determine the total consumption of liquid
--1
!
--i in kilogram or liters, the so-called integrating flow meters, and there are also
flow meters that determine the instantaneous hourly consumption of liquid in,
liters/hour.
The flow meters 11-14.1tA at present are based mainly, on one of the following
o4.
1. Measurement of' the hydrodynamic pressure of the liquid (orifice meters).
--Z-Dirtfertreatnremd-nt of the rate of flaw of liquid in a pipeline volum
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,
?
units of equal volume of liquid fed in succession rate
first method a throttling element, such as a Venturi tube a diaphragm,
or the like, is introduced in the stream
of liquid. The presence of this orifice
leads to the formation of a pressure dif-
ference depending on the velocity of the
stream. This permits calibration of the
measuring instrument in units of rate of
flow or of volume.
The manometer readings may be trans-s
mitted over considerable distances by
means of a remote connection.
In rate-of-flow and volume flow
meters a small vane is used as the sen-
sitive element. The rate of rotation of
f f
46?)
-n
the mixture of fuel vapor and air, re-,
co
spectively; h - Total height of the
,pickup cylinder; x.. Height of the
Fig.46 - Diagram of Pickup of
tance Fuel Gage
l Inner electrode; 2 - Outer elec-
trode; 3 - Insulation, 4 - Fuel;
e e
2' e3 - Dielectric constants
l'
of the
liquid, the insulating material, and
Declassified in Part Saniti?
?
the unloaded vane is proportional to the!
the rate of flow of the liquid. The rate
of rotation of the vane may be trans-
mitted by a remote connection to an in- -
dicating instrument, calibrated in units
of consumption.
If the rate of flow of the liquid
is a constant quantity or is known in
each case a measured volume of fuel
passes at each revolution of the vane.
The contact of an electrical chopper,
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sending pulses of electric current to an indicator, 'may be connected with the spin-
dle of the vane by means of a crimped spring. The volume of fuel passed is propor-
tional to the number of pulses.
An instrument determining the quantity of fuel
consumed, the so-called integrating flow :meter, consists of a pulse emitter of mag-
netic type and of an indicator-counter. Figure 47 givc.5 the diagram of the pulse
transmitter of a ragnetic integrating flow meter. The magnetic transmitter consists
of a vane enclosed in a tube.
The blades of the vane (1) are made
the vane rotates, an alternating emf is
of magnet steel and are magnetized. When
induced in the eoil of the pulse relay (3)..
The electric pulses received in the pulse relay are fed to a counter whose
are propor ional to the total number of pulses during the time of flight,
readings
.1 to
the volume aliquid that has passed through the measuring section. The pulse
magnetic flow meter may also serve for the measurement of instantaneous consumption
Fig.47 - Scheme of the Pulse Transmitter of the Integrating Magnetic Flow Meter
1 - Impeller; 2 - Indicator; 3 - Pulse relay
if a frequency meter determining the frequency of the pulses and calibrated in
'16_4
go
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units of instantaneous consumption, is used as an indicator.
In addition to the above-described types of flow meter there are a 'Argo num-
ber of various types; compound flow meters intended for simultaneous measurement of
the total and instantaneous consumption, are coming into wider use.
Section 5.
TACHOMETERS
Instruments designeC to measure thej.ate of rotation of an engine, turbine,
-et,C are called tachometer.
The electromagnetic ,tachometer, which is widely used in aviation, consists of
Fig.48 _Circuit Diagram of Electromagnetic Tachometer
- Stator winding of the pickup generator; 2 - Stator winding of the
synchronous indicator motor; 3 - Sensitive elements; 4 - Spiral spring;
5- Aluminum disc of magnetic damper; 6 - Scale; 7 - Pointer; 8 - Junction
of magnetic damper; 9- Magnetic 'junction of tachometer; 10 - Permanent
magnet of rotor of synchronous indicator motorl. 11 - Disk; /2 - Rotor
of generator
Ha pickup installed in the immediate proximity of the aircraft engine (or turbine),
--of an indicator mechanically connected with it and installed on the instrument pan-
?els of the pilot and flight engineer, stnelLof a system of conductors connecting the
lpi,cicup, -with the-indicator (Fig.4e):
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e
The pickup consists of a three-phase AC generator or the 4UG 1-48 type. The
rotor (12) i$ a permanent magnet of a high-coercivity alloy 'Alnico". The stator,
built up of sheets of transformer steel 0.5 mm thick, carries a four pole three-
'phase winding (1) in star connection. A three-pronged plug-connector is used as
.lead-out for the stator winding. The three-phase current of the pickup feeds the
synchronous electric motor of the indicator, whose rotational speed is proportional
to the frequency of the feed current, and therefore, to the engine rpm.
The winding (2) or the stator of the electric indicator irotor is analogous to
the winding (1) of the generator. The rotor of the indicator motor is of the com-
pound type consisting of two cross-shaped permanent magnets (10) made of a copper-
cobalt alloy witha high residual induction, and or a metal disk (11) made of non-
-magnetic metal. The rotor of the motor is shown in Fig.49.
When the magnetic field of the stator rotates rapidly, the magnet (2) is at
first unable to follow it, because of inertia. However, the rotating field induces
in the disk (1) currents that interact with the field, producing a mechanical moment
,directed toward the field of rotation. Under the action of this moment the rotor
-lbegins to rotate. When it approaches the synchronous speed, the magnet (2) will be
put into synchronous rotation. To facili-
tate the initial motion of the magnet at
the low rates of rotation, it is installed
on a bushing and is connected with the axle
by the spring (3) which allows the magnet
with the bushing to be rotated through 360,
with respect to the rotor shaft. The pur-
pose of the metal disk (1), rigidly con-
nected with the rotor shaft, is to facili-
t4te bringing the electric motor into syn-
ohronism at high rotational speeds and to,prevent the rotor from slipping when the
? dic,a,tor Motor
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speed of the tachometer pickup changes suddenly.
The output end of the shaft, of the electric motor carries rigidly attached,
the magnetic unit (9) of the tachometer (cf. Fig.48), consisting of two plates with
molded?in cylindrical permanent magnets installed in such a way that the opposite
poles of the magnets are opposite each other. The sensitive element is the aluminum
disk (3) with a low temperature coefficient of resistance, which is installed in the
air gap of the magnetic unit between the faces of the cylindrical magnets. The sen-
sitive element is connocted with the pointer (7), which is rigidly mounted to its
spindle.
When the magnetil unit (9) of thu tach mfter rotates together with the rotor
of the motor, eddy currents are produced in the sensitive element; their value is
proportional to the rate of rotation of the rotor of the indicator motor and conse-
quently also to the rate of rotation of the aircraft engine. The interaction of the
eddy currents with the magnetic field of the magnets, pressed into the plate of the
magnetic unit (9), produces a mechanical moment acting in direction of rotation of
the magnetic unit (9) and proportional to its angular velocity. If the magnet ro-
tates at a rate of n rpm, then the moment
-*here a is the coefficient of proportionality, acts on the sensitive element (3)
moment -is resisted by_moment Mn developed by the flat spiral spring (4), which
?
lis proportional to the angle of rotation of the moving system.
4.6
18 i
-14ere a is the angle of rotation of the moving system;
This means that, at .a:tiven rotational speed of the magnetic unit, the '
4
. ,
iand,the indicator pointer of the instrument will rotate through a.definite angle
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corresponding to the equality of the turning and opposing moments (an
the angle a or the indicator dial or the needle rotation proportional to the num-
ber of rpm of the engine shaft:
The tachometer is provided with a magnetic damper (8) which damps the oscilla-
tions of the moving System made analogous to the magnetic unit. The damping is
:affected by generation of eddy currents in the body of the aluminum disk (5). The
damper considerably facilitates reading of the instrument.
Fig.50 - Indicator of Dual Electric Tachometer
I - Synchronous electric motor; 2 - Gear wheels; 3 - Synchronoscope
Multi-engine aircraft use dual indicators designed on the same principle as
the above-described electreMagnetic tachotheter.
The indicator of the dual tachometer has two synchronous motors with a syn-
%_..ithronoun starter and two magnetic tachometers with pointers. The dials are rotated
JO_Iby means of geared couplings; one of the 4pindles is hollow (Fig.50).
The indicator is provided with a sychronoscope whose design. is shown by
itAC-PdMiP0,Pari% of the synchronocope is an asynchronous electric motor (1)
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'fed by the two tachometer pickups. The rotor (2) of the motor has a three-phase
winding and is fed from one of the pickups through contact rings and brushes. The
winding (3) of the stator is also of the three-phase type. This stator is f d from
the other pickup. The shaft of the rotorends in the indicator dick (4).
Fig.51 Synchronoscope
- Induction motor; 2 - Rotor; 3 - Stator winding;
4 - Indicator disk
The phases of the rotor and stator of the synchronoscope are so connected that
i
-Abe directions of their rotating fields are different. If the frequency (speed) of.
50J- ,
. 1
-both pickupe.is the same, then the magnetic fields of the rotor and stator of the
? -synthronoscope mutually compenate each other in frequency, there are no rotating
'.
';4 _.t,?,?
magnetic fields, and the rotor of the synchronoscope remains stationary. If the
STAT
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tr S',1
,Ake4.
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speed of rotation (and frequency) of one of the pickups is higher than that of the
other, then a resultant rotating field is formed with a velocity equal to the differ-
ence between the rates of rotation of the pickups. The rotation of the field takes
place toward the pickup with the higher rotational speed.
The resultant field impels the rotar and thus indicator disk. The rotation
of the disk may be observed through a slit on the indictor scale.
Errors of the Electron& netic Tachometer
Irtstrwenta.l Errors. The variation in the temperature of the ambient air
'causes a variation. in the electric resistance of the sensitive element in the mag-
netic induction in the gap of the magnetic unit, and also in the elastic properties
of the hair. With increasing temperature, the induction in the gap decreases, lead-
ing to a decrease in the turning moment. This leads to the appearance of a temper-
ature error compensated by a therm*sensitive magnetic shunt whose resistance in-
creases with the temperature; as a result, the magnetic flux passing through the
shunt decreases, thus leading to an increase in the induction in the gap.
The instrument errors due to friction wear, elastic hysteresis, as well as
to scale errors, etc. are the same as in other electric instruments.
The total error of the instrument at t 200C
Anzq not
exceed 35 rpm.
Electric, ferrodynamic, and other tachometers are also used in aviation, but
aU of. them are inferior in reliability of operation, convenience of use, cost, etc.
to magnetic tachOmeter.
The DC electric tachometer is a system consisting of a DC generator (a collec-
tor generator with a permanent magnet in the stator) and an electrically connected
indicator designed in the form of a magnetoelectric galvanometer.
These instruments have not found wide use because of the inadequate reliabili?
ty of the tachometer pickups, whose transient electric resistance varies with the
A
''---i-weal?of?the-brushee-and-the-fouling-of the collectors. leading to distortion of the
readings-4-- -
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The effort to increa3o the r1iabi1ity of operation of the tachometer has led
,to the design otelectric tachometers with AC g n rators which can be made without
sliding contacts (for example, a synchronous generator with a rotor in the form of a
-permanent magnet). Both the voltage and the frequen :* of such a generator are pro-
As an indicator for an AC tachometer the following may be ulied: 1) a high-
-sensitivity and precise magne.oelectric galvanometer connected across a rectifier;
and 2) a ferrodynamic AC galvanometer in which the permanent magnet is replaced by an
electromagnet excited by a special winding through which there flows a current of
the same frequency as in the frame. The use of an induction galvanometer is also
-possible.
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- Sanitized Copy Approved for Release 2012/10/23: CIA-RDP81-01043R001200220003-7
?
,r?-?
a,
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o:
CHAPTER IV
PILOTING- AVIGATIONAL INSTRUMENTS
SECTION 1.
BRIEF G 4"tl. INFORMATION ON V RANE INSTRUMENTS
For the measurement of altitude, speed, and vertical speed of flight, membrane
instruments are widely-used on aircraft. Their operation is based on the measure-
.
ments (by means of a.corrugated box, manometric or aneroid) of a certain difference
?
4-6-
03-4is no opening for the admission of air and all air has been exhausted, such a box
tis :ailed an aneroid*.
al
-0
t t #1
44U0 &NAM
44)
tg
Fig.52 - Elastic Elements of
Manometric Instruments
a - Aneroid box; pi e const;
b-- Mareum.'!Ixi-c bco:Pi const.
of pressure that is functionally connected
with the quantity being measured.
A manometric box consists of corrugated
membranes fused or welded at the ends
(Fig.52). Usually a pin, connected with the
region in which the pressure is being meas-
ured, is attached to the lower rigid (plane)
center.
The upper rigid center is connected
with a transmitting mechanism that moves a
pointer. Ahen there is a difference between the pressures inside and outside the
qmx, both membranes tend to expand if the pressure inside is less than that outside.
Since the lower rigid center is fixed, the displacement of the upper rigid
center is equal to the sum of the bending of both membranes.
If the inner cavity of the membrane box is hermetically sealed, i.e., if there
.j,/,,awY? --
_j* Sometimes gas-filled aneroid chambers are is
, _,????-?
? ,
..-
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The rela
f the. chamber and the effective pressure
f1extre of the ?membrane depends On the
aterial, 1cknes3, form, depth,
and number of corrugations. The charactekistics of the membrane may be rectilinear,
decaying or rising (Fig 53).
The deformation of the chamber is
transmitted over a transmitting me.;:hanism
to the in trument pointer, and minor
translatiOnal displacements (travel or
stroke) of the elastic elements are trans?
formed into measurable rotational motions
of the indicating pointer. In some instru?
.ents the number of soldered membranes,
tne number or -,lambers, is increased
charism is made in such a way as to
of the levers, the angles
use of standard scales with divisions
priwitdro diffimpowre ap
Fig.53 Characteristic of Corru-
gated F.embranes
I - Rising; 2 - Rectilinear;
3 - Decaying
allow regulation of the instruments by varying the length
--lbetWeen them, etc. Owing to this feature
24
---iplott0d... in advance is possible. Most often reciprocating,
-1geared transmiasing mechanisms are used in aircraft instrumentt3.
-la standardized mechanism is used, which simplifies manufacture, repair, and opera-
"
tion of the instruments.
The Barometric Altimeter
In flight, the aircraft crew must know:
1. The absolute elevation above sea level, corresponding to a pressure of
760 zmi lig and a temperature of 15?C.
2. The true altitude above the ground.
S4-.1
- rititr-relativerght -above a-certain place -selected in advance7i-for
Ithertake.off_or,landirts...place.i; .
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\\
? There are various methods and instruments for determining these heights. The
absolute and relative altitudes are measured by a barometric altimeter. The true
altitude is m asured by a radio altimeter.
The barometric altimeter is based on the use of the relation between pressure
and height, expressed by the standard barometric formula or the Laplace formula
(cf. Chapter II).
I-1
1
to 1 ?06
73 Ill Pu
The barometric altimeter consists of: a metallic barometer with an elastic
element in the for= of an aneroid chamber,, in whose inner cavity the residual pres?
sure amounts to 0.15 0.2 ram Hg while the pressure at the outer side is equal to
atmosphere. The deformation of the chamber is greatest on the ground, where the
atmospheric pressure is highest. On ascent to a certain height, the atmospheric
pressure drops, the chamber is relieved of load and tends to be displaced upward.
Figure)54, shows the kinematic diagram of the twin?pointer altimeters used at the
1
present time.
To increase the accuracy of the readings, the sensitive element Of the instru?,
.emBnt is made in. the form .of two aneroid chambers (1), whose inner cavities are iso?,
-lated from each other. With variation in altitude, the atmospheric pressure acting,
on the aneroid chamber from the outside also varies. The force of the atmospheric
40--- pressure is counteracted by the force of the elasticity of the chambers. To each
1.-"" atmospheric pressure there ? corresponds a aefinite value of deformation of the an?
feroid chambers. Any variation in atmosphilric pressure causes a variation in the
deformation of the chamber.
e 4
?vp
6
-----Therdeftrrstation- or, the aneroid- chambers is transmitted across a transmissing-
_
1
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mechanism to the instrulept pointer.. The transmitting mechanism consists of a com-
bination .of rocker and gear transmissions. The transmission of motion from the an-
eroid Chambers (1) to the toothed sector(6) isreffected by a reciprocating mechan-,
ism.
fig.54 -lanematic Diagram of the Twin-Pointer Altimeter
Aneroid chamber; 2 - Temperature compensation of the first kind;
Tie rod, 4 - Bimetal plate (temperature compensation of the second
kind); 5 - Shaft of a toothed .spctor; 6 - Sector; 7 - Gear; 8 -.Large
gear wheel; 9 - Gear; 10-- Shaft; 11 - Spring counterpoise; 12 - Rack;
13 - Base; 14 -.Scale of barometric pressure; 15 - Instrument scale;
16,17 Pointers; 18 - Gear transmission; 19 - Gear wheel; 20- Coun-
terpoise; iZI - Spring
The toothed sector (6) engages the gear (7) to whose shaft the large gear
1
-i wheel (8) is attached which in turn engages the gear (9). To the spindle of the
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? 1
,
. ? .
A
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? ? ? .
gear (9) the large pointer (16) is mounted.
Rotation of the gear (9) is transmitted at a gear ratio of 1:10, over the gear
transmission (18) to the small pointer (l7) which is mounted on a hollow shaft
through which the shaft of the gear (9) Passes. The
dreds, of meters and the small pointer thousands.
;
The aneroid chamber in the pickup mechanism is mounted on the rotating
base (13). This base may be rotated by the aid of the rack (12), whose rotation
large pointer indicates
is transmitted to the base (13) and the scale of the barometrie pressure (14). The
counterweight (20) LI used to balance the drive mechanism.
In order to avoid influencing the readings of the instrument, the weight of
the chambers them-elves muet be counterbalaned by the counterpoise (11) attached
by the aid of the spring (21).
The spring of the counterweight is designed to hold the connections together.
By rotating the shaft on which the spring is mounted, the initial position of the
rigid center of the aneroid chambers may be displaced. This makes it possible to
utilize the straightest part of the curve relating the stroke of the chambers to
the altitude.
The instrument is placed in a standard hermetic body of 80 mm diameter.
1;..e
- body communicates with the atmosphere over the air-pressure intake. The scale of
)
"
the instrument is graduated from 0 to 10,000 m.
Errors of the Barometric Altimeter
Systematic errors are caused by the use of an indirect method of measurements
tr--4 in the barometric altimeter, since this Instrument does not measure the altitude
tH itself but the pressure corresponding to a given altitude. The relation between
'---1 the pressure and the altitude is determied by eqs.(I1.3) and (II.4).
4
The altitude of flight H depends not only on the pressure corresponding to a
-cgiverr trititirte pirbut-alzo on the presInkre pc, on the ground and on-the disvtallyution
f
- "*.'- ? -? ???:, ^
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,
-
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40.
4.
71'7-
75,0"
I 4
or temperature tmean.
The deviations of these quantities from their calculated values, and the in-
-wean) etc. lead to sys-
tematic
of the relations definingthe value of H, pH, t-
tematic errors.
The principal systematic errors of the Altimeter are as follows:
I. Error caused by the variation in pressure on the airfield before take-off
of the aircraft. At a variation in pres.:ure on the airfield, the pointer of the
instrument shifts from the neutral position and the instrument shows a certain al
tude despite the fact that the aircraft has not taken off. This error 1.3 eliminated
by rotating the pointer of the instrument by means of a rack and pinion, with re-
spect to the fixed scale of the instrument, together with the entire mechanism,
sensitive elements, and barometric scaI
It must be remembered that the use of a rack and pinion for changing the
relative position of the pointer and scale and for setting the pointer to zero at
various lengths of the chamber stroke (because of the various values of pc), at
zero altitude, requires the Use of a scale that is uniform with respect to height.
With such a scale, at any altitude, one and the same angle of rotation az of
the pointer corresponds to a definite increment of height 6H; despite the fact that
the initial reading was displaced and the corrugated box shifted to a new position,
this will not lead to an error.
When the pinion is rotated, the barometric scale and the pointers indicating
the altitude are shifted toward each other, so that a loss of altitude will corre-
ponds to an increase in pressure.
After setting the pointers of the instrument to the zero division of the alti-
;44
Declassified in
7
tude scale; the -true value of the pressure po at he airfield is set on the pres-
sure scale. This is accomplished by use of the gear wheel (19), assuring the re-
quired ratio between the rate of rotation of the base of the instrument and that
-1 of the pressure scale.
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2 Error due to a variation in pressure at the airfield atter take-off. If,
at the moment of take-off the pressure at the airfield was pop the pressure at an
altitude H will be
and the reading of the instrument will correspond to this pressure. Let us assume
that, at the level of the airfield, the pressure has varied and has now become poi
this will lead to a variation in the value of the pressure pH at all rieights and
thus to errors in the measurement of the altitude of flight H. This error is elim-
inated by rotating the barometric scale to pressure values, corresponding to the
new value p. Together with the barometric scale, the entire mechanism and the
pointer of the instrument are rotated and a correction is introduced into the alti-
.meter reading. The value of this correction is calculated az follows:
If the calculated pressure on the ground is equal to 1)0, the actual value of
the pressure on the ground is Nov and the pres..=ure at the altitude of flight is
,equal to pH, then the altimeter will indicate the height in accordance with the
1 4' SOON
27
log h-, while the actual value of the height
Psa
log pc) - log pH
3. Error due to a change in the meanitemperature of the column of air teanto
thevaIue ragnitude of the error may be found from the -relations.
The t
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The correction in the instrument readings can be made by means of the rack.
and pinion.
Instrument Errors. An altimeter, like any other meMbrane instrument, has the
following characteristic instrument errors:
1. Scale errors due to imperfect adjustment of the mechanism and wrong scale
setting, i.e., errors due to the noncorre pondence of the angle
pointer with the divisions of the scale.:
2. Errors due to. friction, Irregular gaps, unbalanced parts, nonherretic in-
strument body, elastic afterwork, and hySteresis.
3. Temperature errors due to change in the elasticity of the sensitive element'
with any variation in temperature. The variation in dimensions of the parts of the,
of rotation of the
-J pickup mechanism may be disregarded, sinde such variations do net lead to substan-
tial errors. The temperature errors may reach as much as 3% of the instrument
readings. To eliminate these errors, a kinematic or a dynamic temperature compen-
sation is used.
Instrument temperature errors are inherent to all instruments with elastic
elements made of materials whose modulus of elasticity varies with the temperature i
1 ?
of the surrounding air. For this reasorOthe question of compensating these errors
is of great. interest. A11-.data on the illstrument temperature errors of altimeters
-
and on the method of their compensation is applicable to all other instruments with
elastic elements (speedometers, vacuum g4es, etc.)..
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???'.
A
4,0=4
.h2
errors is the chang in the modulu of elasticity E of the sensitive element with
variations in temperature. The law of variation in the modulus of elasticity may
be considered to be approximately linear, i.e., it may be assumed that the bending
of the sensitive element h varies according to the law
hero is the modulus of elasticity;
Eo is the value of the modulus of elasticity for T 0;
b. is the coefficient of proportionality;
T is the absolute. temperature.
Let us assume that, at -a given altitude, i.e., at a certain load on the sensi-
tive element, the pointer is rotated through a certain angle. With increasing tem-.
004? 4t. perature the modulus of elasticity do-
(Iv .3)
creases, and, in spite of the fact that
ii
hWicle
Fig.55 Bimetal Temperature Compen-
sation of the Second Kind
- Regulating screw; 2 =Bimetal
Plate; 3
or tOOthed-30CtOr;
ton 61:14 to 1-
- Tie rod
the load on the sensitive element remains
unchanged, its course will vary so that
the angle of rotation of the pointer will
also vary. The higher the temperature,
the greater will be the angle of rotationH
of the pointer at the same load. This
,variation in the angle of rotation of the,
:pointer may be considered an increment
in the transmission ratio of the instru-
ment.
In 1929, G.O.Fridlender proposed the
correction of the readings of membrane initrument3 by means of a kinematic temper?
AtArt?Mpertsattornotprryould vary the- tnnsmission factor of the inStriimeht 17 the
96
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2.;
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?E'
?S'eri;
1.4
necessary ratio. In a two-pointer altimeter:this is aecomplished by the bine al
plate (4), installed on the shaft (5) of the toothed sector (Cf. Fig.54)..
Such a compensation is called a bimetal temperature compensation of the second
kind. Its design is shown in Fig.55.
.The bimetal plate (2) .onsists of two Weldep, soldered, or fused metals with
different coefficients of temperature expansion.
A bimetal plate consisting of invar and steel is used in membrane aircraft
instruments:
With increasing temperature, the bimetal plate bends in such a way that the
metal with the higher temperature coefficient (steel) is located at the external
(convex)side.
This property of a bimetal plate is utilized to change the transmission factor
of the instrument. With increasing temperature, the rigidity of the chanter de-
creases, leading to an increase in the camber or curvature of the chamber under the,
same load, i.e., to a decrease in the altimeter reading.
To compensate the increase in camber, the transmission factor must be reduced.
31e?
- This may be accomplished by increasing the length of one of the driven arms or de-
creasing that of one of the driven arms of the transmitting mechanisms.
In an altimeter, a temperature compensation of the second kind is made on the
33--
-- driven arm. The increase in length of the driven arm, with increasing temperature,
-j is affected under the action of the curvature of the plate (2). With decreasing
42_
rnj
--lbe.installed as shown in Fig.551 i.e., the steel part of the plate must be placed
48-4
hilongside the shaft (3) of the toothed sector. The bimetal compensation, based on
co .
the driven arm, will completely compensate the error at only two Points of the
temperature the curvature of the chamber and the length of the driven arm decreases.
To reduce the transmission factor of the instrument, the bimetal plate must
scale': at all other points, the compensatifoills only partial.
In order to achieve complete temperature compensation over the entire scale,
'
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t:tt
t r tt.t,Att tyytt.rptt-nr
ate must be placed on the driving arm. However, for reasons of de-
this is not always possible. For example, in a barometric altimeter, a bimetal
cannot be mounted on a driving member.
A bimetal temperature compensation or the second kind, reducing the temperature
error over the entire scale of the instrument, may itself serve as a source of er-
ror. For example with an unloaded chamber, a variation in temperature will not
cause deformation of the sensitive element, while the arm of the toothed eector will
change in length. This leads to a change: in the transmission factor of the instru-
ment mechanism and cause errors in its reading. It is obvious that this error
will exist not only for an unloaded etate of the chamber but over the entire range
;of operation of the instrument.
A working diagram of the temperature compensation of the se:ond kind for vari-
ous values of the angle # betlin crank and piston is shown in Fig.56.
Under the influence of a variation in temperature by AT?, the driven arm re-
ceives an increment in length of Aa, and the shaft of the toothed sector rotates
'through an angle A a, whose value determines the error of the instrument reading.
_3In this case the tie rod 1 is rotated by the angle 65 0. The value of the angle Aa
?-ias indicated in Fig.560 may be determined from the relation
_1
...-4at wir-n- the angle is Aa For all values of different from every change
1-] 2
,
temperature, even for the unloaded state of the chamber, will produce a dis-
4e_'
-71paacement of the pointer.
_
It is impossible to satisfy the equation e in the unloaded state of the
chamber of a barometric altimeter. On this ground, at zero reading,of the instru-
ment, the load on the chatter reaches a m'fl.xiamm4 while its unleaded position corre-
'Sponds to an altitude at which the press 'e pH is equal to the residual pressure .
inside the aneroid chamber._ For this reaSon, we must assume that yi .1, which
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'leads to errors in the instrument reading due to the influence of the te perature
compensation of the second kind. To eliminate this error a temperature compensation
of the first kind is used. ;This type is installed on the upper rigid center of the
corrugated membrane thus causing any variation in temperature to affect the origin
,of the reading.
In calculating the kinematics of he instrument the operat on of both compen-
Fig.56 - Operation of a Bimetal Compensation of the Second
Kind at Various Values of the Angle
a - Angie # < 900; b - Angle # 900; c - Angle # = 90P
,1141 - Angle of rotation of the shaft of the toothed sector under
the influence of temperature change; ti a -- Increase
in length .of crank; LO - Angle of rotation of tie-rod.
__!sations is taken into account. Figure 58 gives the position of the crank drive of
the altimeter corresponding to a certain value of the height H.
;
-.: The camber hi of the compensation of the first kind is directed opposite to
the stroke of the chamber.
The Cuter h2 of the compensation of the second kind reduces (or increases)
-7-ithe length of the arm Of the toothed sector. The angles between the separatememb-
ibers Are taken in accordance with Fi8.580/ith the angle of rotation of the shaft
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of the toothed sector, d t rmini g the angle of rotation of the pointer, beingde-
Fig. 57 - Bimetal Temfierature Compensation
of the First Kind
Adjusting screw;
3 neroid box
rioted by?.
The distance AB = r is defined by the
relation
Fig.58 - Reciprocating Drive of
Altimeter
Angle of rotation of the shaft
of the toothed sector; 1 - Length
With decreasing temperature, the
elasticity of the box increases and its
deformation diminishes by the value mi; in
this ease the angles a and 's( vary. In
'der to keep the instrument readings con-
stant, it is necessary to hold the value
Of the angle a constant. In that ease
!eq. (IV.5) takes the form
of connecting rod (rocker); Arm of, r?m+hi= (a ?h2)cos ;1+1cos y IV.6)
toothed sector (crankshaft)
Let us subtract eq. (IV.5) from eq. (IV .6).
,,,,Jsineethe,,absolute_valucar the angle, y iz the altimeter is close to.zero,:while, the
jiintrements_ot_thitangle_are smn)1 it follows that the value of 1(cmy1- cos 0,
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/12 CQS t .
?
To find the Value of the "stroke -of the compensation of the first and second
dkind (hi, h2), eq (IV.7) i(s Solved for a Certain constant value of the temperature
,(when hi and h2 are constant) and for two assigned arbitrary values of the pressure
.:(usually for the pressure at ground level and the pressure at a certain altitude).
If the temperature and pressure are kmown, it follows that the values
After substitution in eq.(IV.7) we obtain
h2 COS at = tni ? hi,
(Iv.8)
h: cos 22 nts? ill.
.---After finding the values of hi and h2 from eq.(IV.8) we calculate the dimensions of
f
..'the bimetal plates, starting from the condition of obtaining their required deforma-i
A
--4tions. The dimensions of the plates are so selected that their stroke is somewhat
-'..-longer than the calculated stroke, i.e., we take
3E-4 ?
hi *to
h> h,.
The camber of the compensation of the first kind is detelmined by the equation:
/46
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"r""17,77,,i514+,044r,
vueivi
141tvare and b. are the thickness of the Layers, of the bimetallic plat ;
? al and a2 are the coefficients of linear temperature expansion.
The caniber of the compensation of the second kind is calculated by the equation
The regulation of the compensation of the first kind is accomplished by rotat-
it about the axis (Fig.57). To regulate the compensation of the second kind the
crew (1) is backed off or screwed in and,
at the same time, displaced along the axis
Eirp,
of the shaft (3); cf.Fig.55.
In the USSR the kinenatic type of
compensation is used considerably more of-
ten than the dynamic compensation. This
is explained by the fact that the regula-
tion of an instrument provided with such
Fig. 59 Power Temperature Compensation
a compensation is much simpler.
1?- Bimetal braces; 2 Xembrane box;
Dynamic Temperature Compensation.
The operatiag principle of the dynamic
'Ttemperature compensation is based on the force of action of the bimetal Clamp (1)
.lon-the sensitive element of the instrument (2) over the spreader pins (3); (Fig. 59).
?1_4
The displacement h of the rigid center of the elastic elements, due to the de?
The compeniation operates in the following way:
---Det-us-assume-thati-at-a-certain altitude the temperature of the surrounding
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'Pr
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.7#
! #
medium is eased while the atmospheric pressure remains constant. The boxis more
strongly compressed than before, and the instrument will give decreased readings.
11740 eliminate this error, the bimetal clamps must be flexed, under the influence of
the temperature change, in such a way that their force of compression against the
rigid center of the aneroid box is reduced, i.e., that the displacement of the rigid
center of the box at unchanged load also remains constant. The variation in the box
:at constant pressure, due to a change of its elastic properties with a change in the
temperature of the surrounding air, must be compensated by a change in the concen-
trated load from the tie rods.
With increasing temperature, the aneroid box, at the same atmospheric pressure
will be compressed morel,while the compression of the spreader pins becomes less. At:
decreasing temperature, the picture reverses.
The direction of curvature of the bimetal braces, their shape and thickness
the position and number of the spreader pins may be varied according to the require-
ymmts of design.
By using . carefully Astaiortod temnerkture compensation and by replacing the
.single-pointer instrument by a twin-pointer type, the accuracy of the altimeter
readings is considerably increased while the outside dimensions of the standard body,
. Of 80 tem Aiameter, need not be changed. The allowable scale errors of modern alti-
. _ ?
:meters at t 15?C amount to 30 m at the beginning of the scale and 200 m at the
0)-1
lend; at t = 50?C, the errors fluctuate within a range of 50 - 250 m and, at
1
t . 60?C, amount to 80 - 300m. However, such accuracy of readings is unsatisfac-
--,
-n
, !
,
Amry for modern aviation. The problem of creating an exact and reliable altimeter
.-3
44-4
icontinues to remain one of the most urgent problems of the aircraft-instrument
? _Jbuilding
Se I
-
Airspeed Indicators
4,--T110,41sAsaillY-gcci3tingin5trugents c? mit permit measuring_the..EMM :speed
?
the A.ReeA.,d,th, respect to the ground 'allowing for the winds instead, they indi-1
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-cate the speed of the aircraft with respect to the air.
The airspeed indication (Fig.60) consists of a pickup (I) installed outside the
, "'aircraft (for example, under the wing), which picks up the static and dynamic pres-
sure, and of an indicator II, installed in the cabin. The pickup consists of two
Fig.60 - Diagram of Aircraft Airspeed Indication
1 - Air-pressure pickup 11 - Indicator
1 - Side vents; 2 - Front vents; 3 - Static
pressure chamber; 4 - Full pressure chamber
---,chambers, for static and full pressure, respectively. In chamber (3) the pressure
equal to the atmospheric pressure. The air enters the static pressure chamber
'.through the orifice (1) and enters the chamber (4) through the orifice (2) at the
4.8-4
--iface of the pickup.
Let the pressure in the atmospheric layer at which the flight takes place be
-
!
)
''Ipi the flying.speed-V, the density of air in the atmosphere y, the pressure in the
c
, I
--Itotal-pressure chaMber P.2? the air velocity in it V2, and the density of the air in
-
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,the total-pressure chamber y2.
The pincip1e.of operation of the instruient is based on the Bernouilli equa-
tion, which has the fox
arv.:12)
17hp. level ng heitsz1 z2 can be negleCted. When contacting the face of the
tube, the-air stream is decelerated its velocity becoming equal to zero,
This pressure is called total
pressure, whilethe pressure P1. pH is called static pressure. Then, eq.(IV.12)
ray be written in the fo.
At flying speeds up to 400 kmAlour, the process of compression of the air in the
pickup may be considered isochoric, and the specific gravity of the air to be con,
stant, i.e.,
In that case,
At high speeds, the process of compression of the air is almost adiabatic,
80 that the Bernouilli equation now takes the form
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>el
.44 4-?,?
0 ) ?
so ving this equation for V and replacing Y by we find
where R is the universal as constant equal to 29.27 for air;
g is the acceleration of gravity;
k is the adiabatic index;
TH is the absolute temperature at the altitude H.
n./
The total-pressure chaMber is connected with the inner cavity of the sensitive
46
-:element of the indicator, while the statie-pressure chamber is connected with the
:body of the instrument.
In this way, the sensitive element of the indicator, consisting of a metallic
--imanometer, is subjected to the pressure difference Ap e? p2-p1 Under the action of
this Pressure difference the sensitive element, consisting of one or more corru-
gated boxes, is deformed. This deformation is transmitted over the transmitting
-)nechamism to the instrement.pointer.
Since the pressure difference Ap is correlated with the aircraft speed by
----eq.(1V.1/4), the scale of the indicator of the airspeed indicator may be calibrated
-lift units of speed.
'IA 1
Speed Indicator Pickup. As noted abdve? the speed indicator pickup (Pitot
-itube) consists of the static-pressure chaMber (5) and the total-pressure chamber (1);
-1(Fig.61). The Orifices (,) are placed at distances of four to five receiver diame-
,
-*era from the control point at a place where the air stream is not distorted. To
106
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.?.
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-icontrol icing, the pickup is provided with an electric heater. When the aircraft is
on the ground, the pickup is covered with a special cover. The indicator of the in-
-fltrument is manufactured in single-pointer and twin-pointer models)
Fig.61 7 Air Pressure Pickup
- Total-pressure chamber; 2 - Total-pressure tube; 3 - Open orifice of static
chamber; 4. Winding of electric heater; 5 - Static chamber; 6 - Static tube;
7 - Outer contact ring; 8 - Inner contact ring; 9 Insulating bushing; 10 - Cur-
rent conductor; 11 - Tube for electric conductor.
? The diagram of the two-pointer speed indicator designed by the Soviet designer,i
-41.G.Eltkind, is given in Fig.60. The range of measurements if from ?O to 1000 km/hr.
--Since the instrument has two pointers, its scale must necessarily be uniform. To
I ?
assure uniformity and increased sensitivity of the scale, the sensitive element is
-rade of two manometric boxes (3) and the flat spring (8) with variable elasticity.
. -Variable elasticity is obtained by means of screws which change the working length
7of a spring. With such a design of the sensitive element a characteristic, linear
---Iffith respect to speed, may be obtained.
-.I There is a transmission with a transmission ratio of 1:10 between the small and
--large pointer, so that the large pointer indicates tens and units of km/hr, while
lo-reduee-frietion and prolong the life all spindles rest on agate bearings.
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Fig.62 - Two-Pointer Speed. Indicator
1 - Rodv of instrument: 2 - Scale: 3 -Ilanometric corrugated boxes: 4 - Gear
transmission; 5 - Static pressure connecting branch; 6 - Total .pressure con-
necting stud; 7 - Bimetal temperature compensation; 8 - Variable elasticity
spring
Errors of the Speed Indicator
Sustematic Errors
1. Errors due to the Air Pressure Pickup. The form of the nozzle of the pick-
121), the form of the static orifices and their arrangement along the perimeter and
-1
-Ialong the generatrix of the pickup cylinder, the installation of the pickup on the
aireraft, its-arrangement-with-respect to the air currents encountered all have a
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-
:4
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_
considerable effect on the correctness of the total and .static pressure intake. The
errors due to inaccuracy of the pickup ,may go as_highas 30 kmihr. In recent years
1.1rork had been done, and is still going on, on designing the optimum type of air pres-
sure pickup. Original and varied designs of a few pickups have been worked out in-
-,eluding a few models of USSR pickups as w41 As those made by the firms Askania,
Pollsman, Pioneer, etc.,.
The best -characteristics are exhibited by the Soviet pickups, ir which the er-
Tor due to changes in the angle of flow of the jets .ranging from 00 to 25?, does not
--.exceed 0.4% while such error reaches 9 In the Kollsman receiver and 11.5 in the
Askania pickups.
Soviet designers have created pickups that operate well even at velocities close
to sonic. The operation of an air pressure pickup is characterized by the aerody-
coefficient k, which is equal to the ratio of the pressure drop sensed by the
-pickup to the calculated pressure drop, ranging from 0.96 to 1.2. This coefficient
-is introduced in the equation by which the flying speed V is determined;
v
211R g
th
hp
Ps
(IV.13a)
(111.114a)
The pickup is placed in such a position that its operation is unaffected by the
--propeller air stream and by the vortical mOtion of the air flowing around the air-
:
praft ?
2. Error due to the Influence of the Density of the Air. The operation of the
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41,
?
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, airspeed indicator
manometer.
The equation
1
ed on measuring the pressure diffe nc
V tz?A 2'!rAr
(A
,
Pe
P P1-P2
a metal
shows that the speed depends not only on p but also on pH and TH, i.e., on the de
sity of the air. For this reason a systematic error appears in the instrument read-
ings and is denoted as the error due to the influence of the density of the air.
As with the altimeter, this error is the result of the indirect method of measuring
speed. In order to obtain the value of the airspeed of an aircraft without this er-
ror, the navigator of the aircraft, using an aerial navigation slide rule, applies
- a correction to the readings given by the airspeed indicator by the aid of the equa-
,tion
-where V is the true airspeed;
tra
' Pa. if
p itT
Vinst is the speed shown by the speed indicator;
po, To are the calibration values of pressure and temperature;
pH and TH are the,pressure and temperature at the altitude of flight.
Instrument Errors of the Speed Indicator. 'Instrument errors of the speed indi4
cator are analogous to the' Corresponding errors of the altimeter.
In the submerged condition of the bo*, the angle (cf.Fig.56) is taken as,
equal to7T1 and .for this reason only a bimetal teMperature compensation of the sec-
ond kind is used in the instrument.
The allowable scale errors of the instrument at a temperature of + 15oC do not
1.10 STAT
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t
?
exceed km/hr, while at temperatures 4, 50?C and - 60?C these errors may reach
? 20 1?.
TrUe Airspeed Indicator. At present, instruments in which the correction for
the value of the density of the air in accordance with eq.(iv.) is automatically
:applied by changing the gear ratio of the instrument, are widely used. Instruments ,
showing the aircraft speed .with respect to the air, corrected to the value of the
,
density, are called true airspeed indicators, while the speed so shown is known as
the true airspeed.
The correction for pressure is usually applied by the aid of an aneroid box,
whose stroke depends on pH; the change in this stroke correspondingly chares the
length of one arm of the drive. The correction for temperature ray be applied by
means of a thermometer, automatically or manually. The variation in the texperature
of the surrounding medit.e. deforms the sensitive element of a thermometer. This de-
formation is converted into a change in the gear ratio of the instrument. Several
;different designs of such instruments exist. Figure 63 gives a diazram of a true
--,airspeed indicator with a carrier drive mechanism. The deformation of the manomet-
-
(1) 4c tranAmittnd over the shafts (2) and (1) to the toothed sector (8)
-Hand then to the' pointer (4). The change in the pressure is tompensated bY the
? ?
.ane-
roid box (5) which displaces the slide (9) along the carrier of the roller (2), in
- ;contact with it. Temperature compensation is achieved by the displacement of the
.-islide (10)along the carriers connected with the rollers (2) and (3).
-- The thermometer (6) displacing the slide (10) must be of the remote-reading
-- type and have a great steadying power. For this purpose a liquid thermometer should
-- be used..
-
-- At 'vet:feint the true airspeed indicators in use vary widely in design, and
their readings Pare corrected for the value of the density of the air, they include
anemometric instruments, wind vanes, and a whole series of manometric speed indica-
_ _ __?
- tors. The use of these instruments considerably lightens the work of the navigator.
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-r:
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,
,-!The pilot doea not
this instrument to
;a
y2.-1
iiirtegtlated- by the-toothed sector over the shaft- (12)
;17
seed a-true airspeed indicator; actually, cases in which he uses
udge the lift, it may evenbe harmful.. We shall explain this by.
the following example; The difference AT)
between the total and static pressure and
the lift Y is proportional to the density
of the air
to
isp
Fig.63 - Diagram of the Carrier of
a True-Airspeed Indicator
1 - Manometric box; 2,3 - Shafts;
4 - Indicator; 5 - Aneroid box;
6 - Thermometer; 7 - Support;
_T1,2
2g
2g (ria7)
For this reason, the change in the vel-
ocity V is a signal to the pilot of a
change in the lift Y.
The true airspeed indicator should
not react to variations in air density,
while the value of the lift, which depends
on it, does change. Thus the pilot cannot
judge the variation in the lift from the
readings of the true airspeed indicator.
On modern high-speed and high-altitude
aircraft, universal speed indicator of the
CSI type are installed.
Compound Speed Indicator (KUS).
Figure 64 shows the layout of a compound
8 - Toothed Sector; 9,10 - Slide blocks
speed indicator. The indicator hand (8)
the holder (14), and the--
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S TAT
4tea.
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shaft (15). The position of the box Aries accordinc to the density of the air*
ich corresponds to the given altitude HO causing an additional displacement of the
spindle (15)s and together with it of the sector (16) and the indicator hand (18).
This pointer indicates the airspeed corrected for the density of the air, i.e., the
true airspeed. with diminishinz density of the air, the aneroid boxes (20) expand,
Fig.A4 ttneTratic Diagram nf YES-1200
1 - Manometric box; 2 - Tie rods; 3 - Carrier; 4 - Clamp; 5 - Spindles;
6 - Sector; Gear; 8 - Pointer; ?9 - Scale-4 10 - Hair; 11 - Carrier;
12 - Tie rods; 13 - Sleeve; 14 - U-holder; 15 - Compensated spindle;
16 - Sector; 17 - Gear wheel; 18 - Pointer; 19 Air; 20 - Aneroid box
* The density of the air depends on pressure and temperature. The deformation of
the box is a function only of the pressure. The instruments of the temperature may
?ibe taken into account if it is assumed that the temperature varies with height ac-
cordingto,44. In that oases to each temperature will correspond its oVn2PreASure,
:and it nay_be considered that density is a function of pressure alone.
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and the arm between the tie rod
and the spindle (15) iu shortened. Because of
this ract, at one arid the slum displacement of the movable rigid center of the reeano-
metric box (1), the sectors (16) and (18) rotate through greater angle than the
sector (6) and the pointer (8).
In this way, the pointer (18), dndicatin6 the true airspeed, will always lead
the pointe. (8) indicatinc the airspeed.
By mean of the aneroid boxes (20)0 a correction for air density is applied,
i.e., for the value of the pressure 1.);4 and the temperature T;1.
In this way, the systematic errors in the ilt.73 are taken into account by the aid
of the aneroid boxes (20).
The instrument errors of the compound speed indicator are the same as in the
airspeed indicator. The principal instrumcr4t error is the teeeperature error. 7c
? eairinate it, a teerature compensation of the second kind must be used in the de-
? sign of the 'estrument. In the modern I'M, such a caeTensatioe is not used, since
the deformations of the boxes (1) and (20) under the influence of the temperature
compensate each other to a certain extent.
_ .
-Lath-N=0er iridiators
In addition to the true airspeed indicators, instruments showing the ratio of
the true airspeed to the speed of sound are also installed on aircraft. These in-
-,strumentz.a e?known as Each-number indicators :?
The velocity of sound does not vary with the pressure but it does depend on
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? If the value of the speed calculated by eq.(IV.14) is introduced into eq.(1-V.20),
*then the equation for determining the Mach number takes the following form
The Mach number indicators are calibrated in Mach numbers from 0.3, to 0.95.
Figure 65 gives a simplified kinematic diagram of the instrument.
The deformations 'of the manometric box (4) and the angle of rotation a, of the
spindle (1) are proportional to the quantitytip
I
k1.613
--where ki is the coefficient.of proportionality.
The angle or rotation et2 of the spindle (2) depends not only on the quantity to
I 6 I
--rout also on the deformation or the aneroid boxes (5). The angle a2 is largerat a
?:---bArger angle e and a shorter arm h, whose length depends on the deformation of thei
?4 -
50 '
' -I-aneroid box (5) under the action of the atmospheric pressure pH at the given
It
A
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I/
-3-
:r
Since h k3pti, it follows that
PH
k2, k3 are proportionality fa ors.
The angle a3 of rotation of the pointer (3) is equal to
k2 k3 k =k
P
HThus the angle of rotation of the pointer is determined by the relation
k
Pa
(IV.21) may be-repref7ented in the form:
Consequently,
p (, M7.5
?
Pm
a31(1 NI:?)3,;
r _i J.
(IV.22)
Powever, this means that to each Mach number there corresponds a single definite
angle of deflection of the instrument pointer, i.e., the scale of the instrument
-may be calibrated in equal numbers.
An exact determination of the speed of the aircraft, particularly the ground
--ppeed, is exceedinglyimportant.
? The lack of any ground speed indicator, and the errors inherent in the airspeed
iindicator considerably complicate the work of the crew. For this reason, designers
ientists are placing emphasis on the question of creating simple and reliable
:speed indicators and ground speed indicators.
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.4..
The rate of change of altitude, i0-4., the vertical speed of the aircraft Vy,
mined by an instrument known a a climb indicator.
Fig.65 - Kinematic Diagram of Mach Number Indicator
1 ? Axle whose angle of rotation depends on the flying speed,
2 - Axle whose angle of rotation is proportional to the flying
The operation of a cliMb indicator is based on the method of Measuring, by
Imeans of a manometer, the pressure difference Ap = p, pH, where pc is the pressure.
, ?
-linside a certain closed space, communicating with the atmosphere through a capillary
while pu is the Atmospheric pressure.
On the ground, and during prolonged horizontal flight the pressure in the man-
-ionatric box (2) and in the body (13) is the same, and the pointer (8) of the instru-
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Illent stands at zero. Every change in the vertical speed of flight means a change
_in..altitude..and, consequently, also in tht . pressure within the corrugated box,
in-
struizwnt are in existence in Which the hOGy.t.3 .connected with the atmosphere di-
,
-;,rectly,'and the corrugated lox .through a capillary. In this case an additional space-
Fig.66 - Diagram of the Climb Indicator
Capillary; 2 - Manometric box; 3 - Tie rod; 4 - Base;
5 - Spring plate; 6 - Rack and pinion; 7 - Scale; 8 - Glass;
9 - Carrier; 10 - Sector; 11 - Hair;. 12.- Gear; 13 - Casing;
14 - Support
a small tank), connected with the inner cavity of the corrugated box must be added
ito the design of the climb indicator. There is no basic difference in the opera-
._J
/"---tion of instruments of these two types.
When the aircraft climbs, the air emerges from the body of the instrument
I--ithrough the capillary (1); when the aircraft descends the air enters the body of
the instrument through the same capillary. In both cases the pressure pc of the
- instrument -cannot become equal to the atmOspherie pressure, and the pressure differ-
--ence-A-p-.. pa pH acts on the corrugated box (2). The deformation of the corru-
r1
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?
,
k' ?
, gated box is proportional to this pressure difference Ap. The rate of equalization
.of the preasure acting on the corrugated box depends on the capacity of the capillary
device to transmit air.
At a steady vertical rate of climb orldescent, when the altitude of the aircraft
:varies according to the law U Vyt, where t is the time of climb or descent, the
-
pressure difference Ap acting on, the box is determined by the relation
'
t ?
where D is the diameter of the capillary in m;
1 the length of the capillary in 11,4'
v the volume of the instrument case :in m?;
V is the rate of climb or descent in mize;
?:??
n ia the coefficient of viscosity of the air, kg-,ec/m2.
The viscosity of the air may, with
(IV.23)
a sliffirinnt decree of accuracy, be consid-
ered a linear function of the ambient temperature, i.e., we may take
Al n nT
(Iv.24)
-where n .. 0.62152 x 10-8 kg-sec/m2 deg.
44_,2
-- The pressure drop acting on the sensitive element of the climb indicator is
46_4
--very small. At rates of climb from 30 to 75 m/sec, the pressure drop is equal to
? 10 - 75 mm of water column.
.
r C4 The indicator part of the instrument Consists of a very sensitive manometer
044-
.;/ calibrated in units of vertical speed froM 0 to ?30 m/sec (or from 0 to ? 75 sec).
1
The special rack (6) is provided for setting the instrument at zero and for
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claanging the initial position of the s nsitive element. In diving and steep climb-
the vertical :3ped c:onsiderably excecL the calculated speec. In order to
irreversiblo deform tion of the box, the 5upport5 (14), which limit the s-
a ament of the box are provided.
One of the most important parts of the instrument is the capillary device (1).
In climb indicators designed to measure low velocities, this device is in the form
of a glass tube 0.45 mm in diameter and 50 mm in length.
If the instrument is designed for high vertical spc,et..1, 3uch a capillary is
? unsuitable. In this case a battery of Amilar capillari4,1s In an instru-
ment with a battery of capillaries, the climb indicator equation remains the
but the so-called capillary characteristic k is equa3 to
12.8niv
-where m is the number of capillaries in the battery.
'Errors of the Clilzb Indicator
Systematic Errors. The principal systematic errors of climb indicators are as
I. Lag;
2. Temperature errors;
3. Error due to change in the volume of the manometric box.
Lag. During 4 change in the vertical speed of the aircraft, the instrument
--readings lag behind the variations of speed. The lag in the instrument readings
46_
--remains marked until the values of the error become less than the threshold of
-4t_f
--sensitivity of the instrument.
The magnitude of the lag is determined by the time necessary for the reading
of the instrument to become the true read*ng. In modern instruments, the time of
54_
- lag is 4.- 6 sec. To reduce the time lag the diameter of the capillary must be
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'""-?
o
-I made
large as possible. IncreastIg theY diameter iimit the sensitivity o, the
initrument since a., comjderabie increase of the diameter of the capillary require's.
-4the use-of .corrugated boxes of excessive sensitivity.
Temperature Errors.
variation in the t
A variation in aititude i acc ,panied by a considerable
mperature of the outside air TH and in this case the tempera-
-ture Te in the instrument case also varies, but Tuch more slowly.
As a result of this, the process of ,cape of the air through the capillary is
nonisothermic and an error is created in the intrument. lf the temperature at
opposite ends of the capillary
climb indicator takes the form
anci
repectively, then the eouation of the
T
A p
The error due to the nonisothermic rature of the process may be found b,y
eouation
The variation in the
V e T
1T
1
(I )
temperature of the air in the instrument case leads to an
.additional change of pres:oare there 6 pc, that
instrument reading.
The change of the air
1 --flected in the temperature
.1
--of the air passing through
it introduces an error in the
temperature in the casing and in the atmosphere is re-
of the capillary, and consequently also in the viscosity
the capillary, which varies according to the law
'71where Tap - is the -temperature
,0- c
p.
of the air in the capillary.
These changes in temperature and pressure
'equations of the climb indicator.
If the temperature of the air at Cr!
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must be taken into account in the
ends of the capillary is denoted, in the
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1.
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41.
Declassified in
,genera
se by T1, and T2 respectively, the temperature of the air at a height of H
by TH? and the temperature in the capillary by Tcap, then the equation of the e1iirb
indicator takes the form
A p
Vy Ti Ic
ap
ir2 Tif
(iv. 28)
if the capillary device is placed deep within the case of the instrument and a brass
cylinder, in which the air is warmed by the temperature in the case of the instru-
ment., is installed in its path, then the influence of the noniSothermic nature of
the process .may be elininated even when the aircraft i descending and atmospheric
-air is rushing into the instrument; in this case T1, T2, Tcap alLbecome equal to
-
and the equation of the. indicator can be written in the form
A p
(111.29)
LC)
i.e., the error due to the influence of the ratio still remains in the readings
TH
of the instruments. This error can be eliminated by changing the transmission ratio
of the instrument in accordance with the temperature ratio , The temperature
TH
? compensation may be made in the form of a binetal plate, wtose stroke varies with
--the temperature ' in the instrument case, and which carries the plate of an aneroid
Lc
box, whose stroke varies with the pres ure Tic, this pressure in turn depending on
the temperature TH at the altitude of flight.
4f
It must be remembered that the use of an aneroid box with a linear character-
--istie relative to altitude, as a device for measuring the atmospheric temperature
admissible only on condition of a standard distribution of temperatures by height;
S.;
p -Hin this case it is assumed that the pressure in .he instrument case is a unique
A
liaunction of the temperature or the outer air.
rej 'ror due to. Change of Volume of the Sanometric Box.. When the vertical speed
f flight varies, the stroke of the box also varies and thus the volume 'of the in-
ease. This leads to a change in the characteristic k of the capillary.
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?The error thus caused does not exceed 0.02% of the instrument reading's.
The instrument errors of the climb indicator are the sarxe as thoe of other
nometric instruments.
The total error of the reading in a climb indicator calibrated to 30 m sec, is
about 1.5
Sec.
Section 2.
AIRCRAFT COMPASSES
Instruments for determining the course of the aircraft are called compasses.
These include compasses based on the utilization of various principles: magnetic,
gyroscopic, induction, astronomic, radio compasses, etc.
The simplest and, until today the most widely used type.), are the magnetic
compasses.
Marnetic Compasses
The operation of the magnetic coss is based on the interaction of tne
earth's magnetic field with a magnet in the instrument. The magnetic pole of the
earth is located near the geographic poles and is somewhat displaced with respect
to the It is considered by convention that the southern magnetism is concen-
trated in the northern hemisphere and the northern magnetism in the southern.hem-
,
isphere. The magnetic needle comes to rest with its north end pointing north, i.e.,
its north magnetic pole which conventionally possesses southern magnetism, is con-
sidered to be located near the north geographic pole, and the south magnetic pole ,
PA-1 near the south geographic pole. The angle between the geographic and magnetic
4r.L?,j
peridian is called'the -declination.
The declination varies with the geographic position, the season, etc. It is
t, :4
considered positive if the northern end of the needle is deflected to the east of
, the geographic meridian. The value of the declination is determined from special
1
cbarts- of -magnetic declinations on which places Afith the same magnetic declination
123
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I? .
are connected: by lines -called iisogons; on thee ae chart3 th
magnetic ralie3
are plotted. The magnetic field in alm63t all part:; of the earth, except the mag-
neticequator, i3 inclined to the horizon, and for this reanon the total fore, T
or errestrial Agl tiom, directed at an angle to the horizon, iay be reolved
into two'componontJ: a vertical component Y T An 0 and a horizontal compon
ent U T con 0 (Fig.67). The horizontal component H which citatili3hen the mag-
netic needle in a direction north-south the directing force of the magnetic
compan. The angle 0 between the total
force of terrcitria1 magnet15m and it..;
horizontal component i? 'ailed the in-
clination. A ,ompatiJ needle, freely ro-
tating about a horizontal axis, takes a
vertical poAtion at the magnetic pole and
a horizontal po5ition at the equator. A
magnetic needle set up from a point, when
placed in the northern her;-_lapnere, under
the action of the force developed as a
reult of the interaction between the
needle and the magnetic pole of the earth,
t..,nt_is to dip with it northern flld down-
ward. To eliminate this inclination, the
southern end of the needle is made heavier.
Fig.67 - Components of the.rthls
Magnetic Field
HH - Plane of the horizon; Pm - Mag-
netic pole; P1111 - Projection of the
magnetic pole onto the earthl mag-
restrial magnetism; H - Horizontal
In the southern hemispher, the northern
omponent; I - Vertical component,
end of the compass needle is made heavier.
0 - Dip
The magnetic compass (Fig.68) makes
it possible to determine the magnetic course of the aircraft. The principal part
of the co=pelos is the card with the magnet (1), resting on the steel point (3) on
an agate napphire hearing (4), attached to the column (5).
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The presence of magnets compels the card to align itself along the magnetic
lines of force of the earth, i.e., the card maintains an invariant direction with
respect to the earth.
The card is installed in the bowl (6), filled with liquid, which reduces the
? weight of the oard and helps to damp it oscillations.
The compaos card consists of the linb (7) which is a thin bras or aluminum
? disk (in compasses with a horizontal scale) or a cone (in compasses with a vertical
scale). In some zompasen, divisions from 0 to 3600 are marked on the disk. In
this case, the instrument is read off by the aid of a course marker rigidly attached
to the bowl (Fig.68a). The limb is connected with a hollow float (e), on which the
magnet (1)? and the damper (9), tearinE numerals, are attached. In this case, the
compass readings are read of from a scale rigidly attached to the bowl (Fig.68b).
The liquid in the bowl must have the lowest possible viscosity to prevent excessive
resistance to the motion of the card. The freezing point of the liquid must be
below ? 600C, and the boiling point above 50?C, while the density of the liquid
must not vary substantially with the temperature. Liquids that completely satisfy
these requirements have not yet been found; modern coir asses use licroin, whose
density varies considerably with the ambient temperature; this lcao:s to a change
e?f.
'
I P.
?
141..ft
dr" S 1
^f th? lin"" 2rA t^ the fee.mtien
- or to an increase in pressure at high tem'peratures.
4i_j
--Ocompreseion or overflow basin (10). At present, an overflow chamber consisting of
-7-1 an additional space is used into which air bubbles at lower temperatures and excess.
tr,t.eo
--iliquid at higher temperatures.
b?_
4.
0.14ci
low
To eliminate thene drawbacks, the instrument is provided with a compensating
The compensating chamber somewhat increases the volume of the compass and
thus ensures a wider range of compensatioh.
According to the purpose of the compass (for the navigator or for the pilot)
-1 the accuracy required for its readings differs. The navigator's compass, which is
!
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0
_
7t.
also called the master compass, must be more accurate. The pilot's compass is
called the course compaas. In such compasses the readings are made by the aid of
course mker which is attached in a fixed position with respect to the compass
a
bowl, and a movable card which takes the direction north-south (Fig.69).
There are many deign; of magnetic
cbmwseal but the layout iB the same in
all, with only the position and. form of the cale (vertical and horizontal scales)
or the system of reading differing.
The operation of all compasses follows the ,;,a2c 8eneral pattern: IT the mag-
netic needle is removed 'from its equilibrium position and then released, it will
return to original poAtion. The motion of the needle takes place under the
_influence of the following factors:
1. The turning moment M developed by the magneti card; the magnitude of
?? this moment may be found from the enuation
"turn rit T ml H sin w
? where M is the magnetic moment; M 2 ml;
m is the magnetic ma7s concentrated at one pole;
2 1 is the length of the magnet;
o is the angle between the direction of the magnetic lines of force and the
? axis? of the -magnet (a Variable quantity);
H is the strength of the earth's magnetic field.
At smn,11 angles of declination of the neelle, we may take sin 9 o; in that
-lease,
If,
Mturn MH 2 ml H (1)
1
2. The moment of resistance Mc of the liquid; at small angles 9, the magnitude
tof this moment may be found from the equation
f
X3 se
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6
Ft g.66 - Layout of the Parts of the Magnetic Compass
a - Diagram of compass; b - Compass card;
1 -'Magnet; 2 Pointer; 3 - Pivot; 4 - Bearing; 5 - Column; 6 - Bowl; 7 - Limb;
8., _Float; Damper; 10 - Overflow baein; 11 - Seale; 12 - Deviation of the
parta
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where K.is the moment of resifftanee of the liquid per un:it angular velocity of the
3. The moment of friction of the pivot on the 1 aring; this quantity is so in-
significant that it in usually negl cted.
4. The moment of the forces of inertia
where J is the moment of inertia of the card;
4) is the acceleration of the card (a variable quantity).
The equation of the natural vibrations of the moving system of the compass has
_the form
jT KTMH
A solution of eq.(IV.30) shows that, according to the ratio of K to MN, the
(Continuation of Fig,68)
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Fig.69 - External View of Navigator's Compass and
Pilot's Compass
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?
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compass card, in returning to the position of equilibrium (toward the meridian),
may execute oscillatory harmonic or aperiodic motions (Fig.70).
In modern aircraft compasses, a harmonic law of motion ia assumed for the
needle, and for this reason the compasses themselve are sometimes called periodic
compasses.
Characteristics of Compasses. The operation of a periodic compasa is character*
ised by a period
45
oscillation and by a damping decrement d; the latter represents
the ratio of the amplitudes of two .:iucces-
sive oscillations. The value, in d D is
called the logarithmic decrement of (tarp-
APEitiCAVC CURVE
ozievavic
cora
sec
Fig.70 - Character of Motions of
Compass Card on Return to Equili-
brium Position
compasses,
+4,-0
TorIV vco.viTe.s.ito
in g; the value of the damping decrement in
modern comFasses varies from 5 to 6.
In addition to these characteristics
we must also know the following data:
a. The damping time, i.e., the time
necessary for the return of a card after
deviation, to the meridian. In periodic
varies from 15 to 30 sec.
b. The entrainment, i.e., the angle through which the liquid entrains the card
when the-bowl is rotated together with the aircraft through 3600. The entrainment
depends on ,?rle shape of the card, the viscosity of the liquid, the rate of rotation,
etc. It is desirable to have the entrainment at a minimum value, since at high
entrainment, even an minor change in the course leads to errors in the instrument
readings. In modern compasses, for a 3600 rotation at a rate of 0.1 revolution
per second, the entrainment reaches 10-12? (at normal temperature).
c. The lag which depends on the friction in the bearing and is characterized
.by the angle at which a deviation of the card does not pass to the equilibrium
positions. Under conditions of flight with vibration, the lag does not exceed 1?.
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?
Errornass
The systematic errors of the compass comprise the deviation, as well an the
banking and turning errors.
Deviation of the Cozpass. Besides the gnetic field of the earth, the magnetic
field of the aircraft also acts on a compass installed in an iircraft. The earthts
magnetic field produces the so-called directional force, i.e., the force holding
the magnet c system to the direction of the ragnetic meridian. The magnetic field
of the arc rat, caused by various metallic aircraft parts, ray be considered as
consisting of two fields: a permanent field and a variable field.
The permanent magnetic field is produced by netals having a h ph coercive
force, which are magnetically hard (these are conventionally called hard iron). The
raolitude and direction of this field with
respect to the aircraft does not vary when
444svirm.
Atom/4w
CONAW
Mie/044141
Fig.71 - Deviations of the Mag-
netic eedle
Its course varies and, therefore, is
called permanent.
The variable magnetic field is pro-
duced by metals with a low coercive force
which are rapidly ragnetized and remag-
netized; i.e., metals that are magnetic-
ally soft, and are conventionally called
soft iron.
When the course of the aircraft
changes, both the direction and strength of this field change with respect to the
Aircraft, for which,reason it is called a variable field.
The aircraft magnetic field produces forces deflecting the magnetic system
from the direction of the magnetic meridian. The angle 6 between the direction of
t.agnetic meridian and the direction in which the magnetic system points under
,
tthe influence of the permanent and variable magnetic fields of the aircraft is
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called the deviation (ag. l). The deviation is''considered positive if the northern
nd of the magnet deviates toward the east,
The deviation of an aircraft is a variable quantity since the relative position
the restoring forces, i.e., the forces holding the needle in the position of the
magnetic meridian, and the forces causing
' ommcnow w the deviations, change with any change of
o nu ir
COMOWS8 gi
the aircraft course.
iwillimeoferkWerael If the horizontal component of the
a
100,001~ remanent magnetic field of the aircraft
4s al then the direction of the vector
reoresert4re this force males - certain
invariable angle a with the a)ds of the
aircraft, which angle is independent of
the aircraft course (Fig.72). At the same
tine the direction of this vector with
respect to the vector Ho representing the
horizontal component of the earth's mag-
netic field, does vary with the course of
the aircraft and causes a variation of the
magnitude and direction of the course,
producing the deviation Hi (Fig.73).
If the aircraft is set on a course k
4$0
a
Fig.72 - Action of the Cozponents of
the Permanent Field of the Aircraft
on the Magnetic Yeedle
H - Horizontal compoennt of the earth's
magnetic pole; H1 - Horizontal _component
of the aircraft magnetic pole; F Re-
sultants H and- Deviation;
1' k
k - Course; a - Angle between the vec-
tor Hi and the axis of the airplane
Ton which the vectors H and Hi coincide (Fig.73), then the deviation wil) be equal
to zero, since this produces a deviation of the force HI 0. If the aircraft
continues to turn clockwise then the combined forces H and H1 produce the result-
ant
F (cf. Fig.72). The magnetic needle of the compass points in the direction of
fthis resultant, 1.0.0 it indicates the compass meridian. The angle between the
compass meridian and the longitudinalaxis of the aircraft is called the compass
.STAT
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r'iater.v.wm.
-course, while the angle between the magnet c and compass reridians, i.e, the differ-
ence between the magnetic and compass courses, is called the deviation.
? In this case the frco Hi, produting?the deviation, varies in magnitude and
direction with respect to the aircraft. On resolving the force Hi in the direction
of the aircraft-axes xx and zz, we obtain the components and Hiz (Fig.74) whose
directions areconstant with respect to the aircraft, while- their maglitude varies
by the law
The magnitude of the deviation due to these forces is deterrined by thP
equations
C cos k
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Declassified in Part - Sanitized Copy Approved for Release 2012/10/23: CIA-RDP81-01043R001200220003-7
-;wher k i6 the course of the aircraft
1,1. and C 4re the?coeffi-ients of deviation duo to hard ion.
The forA ? -41 and H and the deviat-1,)n *oduced.durthr, two ?,W rotations
?1 ?
of the aircraft lead to zero and hane their sign twice. For this reason, the
the rpagnie
deviations arising under the influence of:?e permanent retic ,l Au
The dcvat5.ons , the aircrtft cora'*'; Ille to the oeaner.t .11,, -? '-cr ,,-,, nc
very large vaiue of the order of','' and N'Te).
in the desi0',of all colq si.hout
(.:xception, s.,pecd dei. ? 3-:n 4:73
a dovit.,110 cOmpttor, I- kn;orpo rtc and fualy -,-1.' ...A.1..4?, -t.,3 the
deviation &as to hard iron.
The variable 1:zr., ic field ,i-v--.:ner. ed bv tile'
isvi on soft iron (the ir.strl ,ints of hard '.iron and soft h, neLeoed sirce
as a result of it, force, art prd1.2,ced have the
froc hard iron CT the sez:dcircula: r3eviati
and the quadrant dedations A) and E.
The quadrant deviations vary acecrdirp to the law sin
cos 2, where k it .he course.
Since there ,f_s 7itt1e soft .1-on on an aircraft, the dev--
directio;n
the cmttaLt
P tc
no snecial
? usually small; theref ate thee do tor. are
corporated in the c,,:irass.
The deviation of a coLT
LtIS an aircraft :steS,e(i
courses.
and ?:,15?), simultancousl,y r 414, off rthe
difference between the ':.%aotic and compass course is the deviation
The curve of deviation :?r: expressed math=aticall
B sin k C cos D sin a E cos a
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AMMO
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F1g.75 - Compass Deviations
Position.of aircraft on various courses; b - Semi-
circular deviations; c - Quadrant deviations (D sin 21c)
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uhero is the deviation.
A is cocrricient characto Ling the pernanent deviation .due o 1tirn
induced in soft iron, and the error of settil u,
B,C ,tre coeffcients characteri.4in the semic!,reular deiaton dut to hard
!iron on the aircraft;
DpE are coefficients characterizin th ju4rant devJatons due to :met
induced in soft iron.
In c ses where the deviation of the co;:mas exceeds 10?, it :,rust be el:!ITA ,ated
reduced. The r x11:1111. value reached by scrcircuLr ,le,..iatior,n on the a'r-
craft, and for this reason :31)ccia1 tteitioi usually focused on its eliP,ina-.;ion
or reduction.
Elimination of Deviations. The perranent deviatirq.? characterized by the
Fig.76 - Deviation Compensator
1 - Body; 2 - Transverse spindles;
3 and 4 - Longitudinal spindles;
5 Compens ting magnets
Fi .77 - Principle of Operation of
the Deviation Compensator
a - Yinimm action of the zagnetic
field; b naximum action of the mag-
netic field; c - Mean position
'coefficient A, is taken into account by rotating the compass through the proper
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?
e with respect to the vertical axis.
The semicircular deviations, characterized by the expressions B sin and
C cos lc, are eliminated by a so-called deviation compensator, consistilig of two
pairs of permanent t-nagnets (Fig.76) which -produce moments equal in magnitud. but
? opposite in direction to the moments produced by the semicircular deviatiOns.
One pair of. magnets (longitudinal rapets) is arrange4 parallel to the longi-
tudinal axis of the aircraft. This pair Of magn desined to compensate the
force II . - Another pair of 1..'zigrlot,i3 (trannvert3e
1x ?
macnets) is arranred parallel
to the
zz axis of the aircraft. These nagnets are designed to compensate the :orce
lx
By -rotating the rollers (2) and (2) the :-)sition of the coLzpersatin;: magnets
can be changed, i.e., the action of their magnetic fields (Fil7:77) can be increased
.or decreased, and positions of the ragnets that keep the deviation to a minimum
can
be selected. Such a deviation compensator compensates only the semicircular devia-
tion. The quadrant deviation cannot be
elirAnated by permanent magnets. It may
be compensated by means of strips of .,-oft
11
top, or bottom of the com:pass (Fig.78).
iron placed symmetrically at the sides,
IIJ

""
The
cl!fisirant deviation is
not eliminated but allowance is made for
it. If it is necessary to eliminate it,
this can be done by means of soft iron
Fig.78 - Arracgment of Bars (2) of
bars.
Soft Iron Around the Corpass (1) to
After installing the compass on the
,
Eliminate the Quadrant Deviation
aircraft, the semicircular deviation is
first removed after which the operation of the corpass is periodically checked. The
'semicircular deviation is eliminated by the following procedure; The aircraft,
iwhile on the ground, is set to the course k 0; on this course the deviation
137
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is A k C E so that the Lel= containing sin h and sin 21: vanish. The total
:force acts in the direction of the right wing of the aircraft. To eliminate the
? deviation along the course k r 0 by the perranent magnets of the deviation compen-
? sators, the force 0113,s produced, equal in nagnitude to the force causing the devi-
ation on this course but opposite in direction, i.e., in the direction of the left
wing of the aircraft (Fig.79). In this case, not only the semicircular deviation
.1,00a MVO AIM. .111,
If,
ON*
Fig.79 - Elimination of Deviation on a F4.g.80 - Elimination of Deviation on a
Northern Course
Ct= At + Et +
Southern Course
Ct _ CI At + ET
1
?
but also the constant deviation A and the quadrant deviation E on the course k = 0
is eliminated; but for that, the deviation is increased on the course k = 1800.
iNhen the aircraft is rotated through 180, the deviation tik, defined by the equa-
tion . A C + Es is produced by the forces Al and Et acting in the direction of
-the left ming, and by the force CI, acting in the direction of the right wing. The
.force C1 I = At + Et + Ct, produced by the deviation compensator, acts in the direc-
tion of the left wing (Fig.80). The forces CI and - CI mutually cancel each other;
-consequent17,-the force 2 (At + Et) now acts toward the left and produces, the
devi-
STAT
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?
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ation 2A (A
The deviation
compen.
the deviation to the value A " A # E; in tl
is likewise equal to A h A +
If now the aircraft is again set to the
.r1
tor on the course k 1800 brings
s case, the deviation on .the course 0
course k 0 the devation will be
equal to .Ai( A .4- E- and, in character, will be -a quadrant and permanent deviation.
.With this order of work on the course k ,,, 0, as well as on all other courses, only
the quadrant constant deviation should r
IR star,,,,zo
r
XI
Rs ri r
I
komprommw
Fig.81 - Conditions Causing Banking Deviation
To cancel the deviation B cos k, the same operation is performed on the
courses 90 and 2700. After the semicircular deviation has been canceled, the re-
sidual deviation is determined on eight principal courses 0c, /45?, 500, 135?, 180?,
225?, 270?, and 315?, and eight equations of deviations A? ' A A A A
0 45 90' '
1?5180'
A
A are set up; in this equation, we find the value for which correc-
2250 A270' 315
tions have already been introduced into the instrument readings along any course.
During flight, the deviations may vary, for --exarple, when the relative posi-
tion of the equipment changes when the bomb load is released, etc. This is one
of the reasons why magnetic compasses are made remote-reading.
Banking Deviation. As already stated, the horizontal component H1 of terres-
trial magnetism gives a semicircular deviation while the vertical coefficient 2.
in rectilinear horizontal flight does not cause a deviation, since it acts in a
vertical plane and its projection on the plane of the coil is equal to zero. At
139
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4.3 4
longitudinal and transverse banking, with the aircraft in rectilinear flight., the
vertical component is inclined together with the aircraft, while the position of
the cops ,card rettains unchanged with respect to the plane of the horizon, i.e.,
the force fl. gives the projection 111 onto the plane of the card. The value of this
projection is determined by the equation.
where Y is the angle of pitch or bank. The force Es, during a longitudinal bank,
is directed along the xx axis and during a transverse bank along the zz axis
(Fig.81). From its character, the banking deviation will be semicircular just like
fig.82 - Turning Error on tht Course k 90?
Bank of compass card; e ?-? Inclination; T - Direction of
magnetic pole; Y.2 - Projection of the vertical component to
the direction of the axis zz; H2 - Projection of the horizontal
component -to the direction zz; zz - Transverse axis of the air-
craft; v - Turning error
that produced by the forces Hix and H. In most cases, the banking deviation is
not canceled, despite the fact that it sozetimes reaches considerable values.
banking deviation may be canceled by permanent magnets arranged in a vertical
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043R001200220003-7
41
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Turning Error. If, during -evolutions of an aircraft, for example in the case
of a turn not only the aircraft but aloo the compass card As banked with respect
to the plane of the horizon, then the vertical component of terrestrial macnetism
willno longer be perpendicular to the plane the card and will give instead the
rojection Yz acting in the direction
This direction depends on the angle of
bank 0 of
the compass card* (Fig.8). The comas card is then deflected from the
? direction of the magnetic meridian. The value of the error will depend on the
course of the airplane. On the course 0?, the :,,rojection Y, is directed along the
transverse axis of the aircraft. The horizontal component of terrestrial macnetisn
is projected onto the direction of the axis. mthe eorret,-ic comosition of
.he forces Y and H, we find the resultant H". The angle between this resultant
an the magnetic meridian will 5ive the value of the turnint errorv 7,
Ysn
tg =
II If
(1V../4
' - 1')
On the course 1800, the picture is analogous. On the courses 90? and 2700, the
Trojettion of the horizontal conponent of .terrestrial macnetism will act in the
ection zz.
In this case,
The force
AL,
H cos p
Y sin a is likewise projected onto the ads sz. From the geometric
caziTosition of these two forces, we find the value of the directing force of the
eompass. On the -:course 900 it is equal to 11" 5 2, sin H cos 0 and is directed
at,lor', the zz axis of the aircraft, coinciding with ,,he magnetic meridian.
g- The angle 0 of bank of the card does not always coincide with the angle Y of
bank of the aircraft.
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STAT
?".;.?'.??????"'?4
:71
7
4.
???-??..
-?-
,
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?
? -""'"' ' ?.?
?
??? 1,?
On replacing Z and H by T, i.e? by the total force
o ruly Write
7, sin T sin sin #1;
1.1 cos az:- T cos 0 cos 'tt;
II" rsin sin 1?cos 5 cos 41.
tri.f
terrestrial magnet
For fl . 900 - e;, the expression vanishes and H" - 0 '.e. in this case there
vill be no directing force in the co:2pass. If sin 6 > H cos 0, then the compass
needle Wfl1 point south. The angle 6 - 90? - a i called the crit cal angle of
At angles of bank less than the critical, on the courses 90 and 2700, .he mag-
netic system points north and, although in this case the directing force H" is de-
creased, there will be no error in the 4ns mrcent readings.
?o ro-
tate
angles of bank aabove the critical value, the nagnetic system tends 4'
.
tate through 1800; in this case there will be an error in the -instrument readings.
? The angles between the magnetic and compass courses is called the turnin7 error.
ThiN error bears this desigr.ation because it appears during turns of the air-
craft.
The analytic expression of the turning error has the :or
V st
-
If, in this expression, we replace ?7F by known values of the total force of
terrestrial magnetism T, the inclination 0, the bank 8 and the magnetic course kw,
-
we obtain the expression
k,=t km, --arc-tg [tg k cos
cos
The turning error on an arbitrary course of the aircraft is shown in Fig083.
? ?=. ? .?: r?-? ' ? .
?
142
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The instrument errors of, the compass are caused by lag, entrainment, inaccurate
manufacture, and unbalance of the compass card, consisting of a pendulum. Under the
AWN
de CORD
Fig.83 TurninE Error on an arbitrary Course
action of periodic disturbances (for example in bumpy flight), the mar;netic system
will oscillate about the equilibrium position.
? Rete-leadingCopasses
The presence of large ferromagnetic masses onairplanes, and the .inconstancy
of the magnitude and direction of the forces forrng these masses (as a result of
'the change in the position of the control levers, turrets, etc.) leads not only to
variation in the deviation, but also to 'a dependence of the deviation on factors
-not amenable to estimation. The impossibility of eliminating these factors has led ?
Ito the -development of remote-reading compasses in 'which the pickup, ordinarily con
? sisting of a magnetit.compase installed at the point on the aircraft !there the
143 STAT
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influence f the ferromagnetic ie in smallest (for example, in the tail of the
fuselage or in the wings of the aircraft), The compass readings through a remote
transmission are communicated to an indicator placed on the instrument board.
Among the remote-reading instrunents, the most widely used are the following:
remote reading gyromagnetic cmpass (GNC), rmote-reading potentiometric compass
pass (PPC)? remote-reading induction cozpass (iLIC), etc. (cf. infra).
zna4.440or r ..... .. .1
4(wicup 1 1 'I
reading 1
1
1 I /
m I t...; 1
1 1 1 i
1 1 1 I
/ 2
-, + 5
__
, ....- 1
r--/-_-1
1 ilk 0 i 1 1
ti 1 I 1
1 i 1
1 ;
? I
L????%. NW/. ???????? ???????? ????I??
Fig.84 - Principal Diagram of Remote-Reading Fot,...-tdometric
Compass
Stator; L Perm-
anent maglet-rotor; 5 Fo nter
Remote-Reading Potentimetric Compass. Figure 84 gives a diagram of the
remote-reading potentiametric compass.
The pickup consists of a magnetic compass with a ragnetic system which is
considerably more powerfUl than that of the ordinary magnetic compass and consists
of four magnets of's. length of 1 -110 mm and a diameter of d = 10 mm, with a
magnetic moment of 120,000 units.
At adeviation Of 900, such a system develops a moment of 2 G/cm while the
,magnetic system of an ordinary compass develops only 0.03 Mem. Such a great
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:
Fig.85 - Pickup of the Remote-Reading Potentiameteic Compass
- Compass card; 2 - Permanent magnet; 3 - Collectors; 4 - Stage with brushes; 5 - Body;
- Ring potentiometer; 7 - Bridge; 8 - Sylphon; 9 - Scale; 10 - Inspection hole; 11 - Gardanic
suspension; 12 - Damping; spring; 13 - Journal; 14 - Spring; 15 - Ring; 16 - Screw; 17 - Packing;
313 - Casing; 19 - Electric wiring
Declassified in Part - Sanitized Copy Approved for Release 2012/10/23: CIA-RDP81-01043R001200220003-7
noment is necessary to overcome the friction of the brushes on the potentiometer.
The three brushes (1), which are spaced at an angle of 120P are rigidly con-
nected with the magnetic system of the compass. They consist of three sliding
Fig.86 - Design of the Remote-Reading Potentiometric Compass Indicator
1 - Moving magnet; 2 - Spindle; 3 - Pointer; 4 - Bearing of spindle;
5 - Locknut for regulating longitudinal play of spindle; 6 - Permalloy
core of stator; 7 and 8 - Brushes; 9 -- Stator winding; 10 - Three contact
rings connected with the stator coils; 11 - Brushes; 12 - Screws for
attaching brushes; 13 - Plug contacts 14 - Scale; 15 - Plate; 16 - Ring
in which the inner case rotates; 17 - Flat annular spring; 18 - Fixed
sight; 19 - Rack; 20 and 21 - Gear drive; 22 - Inner case
current-collector contacts and slide along the potentiometer (2), attached to the
'compass case. The potentiometer is fed by the electric system of the aircraft. The
Magnetic system, together with the brushes, maintains a constant position in space.
The potentiometer varies its position together with the aircraft during its evolu-
STAT
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Declassified in
ton.
When the aircraft turns, the position of the potentiometer with respect to the
?magnetic meridian changes, and consequently it also changes with respect to the
brushes. In this way, depending on the course of the aircraft, the potential dif-
ference between the brushes of the pickup is changed.
? The voltage taken by the brushes from the potentiometer is supplied to the
three frames (the windings of the stator 3) of an indicator consisting of a ratio-
meter with a moving magnet. A ratiometer with moving frames may also be used. The
position of the moving system of an indicator depends only on the distribution of
currents in the winding, i.e., on the position of the brushes on the potentiometer,
with only one definite position of the moving system of the indicator corresponding
to each position of the brushes of the pickup. From the scales of the pickup and
Indicator, graduated from 0 to 360?, the course of the airtraft with respect to the
magnetif meridian is read off. The design of the magnetic pickup is given in Fig.85
and that of the indicator in Fig.86. To reduce the vibration, the instrument is
carefully shock-absorbed. The weight of the pickup is 5 kg and that of the indi-
cator about 1 kg.
The error of the instrument is of the same order as that of an ordinary mag-
netic compass; the lag is /10, the entrainment at an angular velocity of 1 rpm
J___s wt
exceed 12?; the damping time for temperatures from + 50 to - 60?C does not
emceed 20 sec.
The error of the remote transmission of the compass runs up to ? 2?.
Section 3.
NAVIGATIONAL COORDINATORS AND AUTOMATIC NAVIGATORS
The automatic determination of the geographic coordinates of a moving object
is possible in principle by utilizing the property of a gyroscope of keeping the
position' of its wads in space constant, regardless of the displacements of the
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c4,
'
Vt.40,s4,t\
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,P?
I.
_
moving objects and the earth's rotation, and also by calculating the path from the
variation in speed and course of the aircraft.
Gyro devices require exceptionally high accuracy, which is technically still
difficult to achieve; This forces us for the time being to give preference to in-
struments based on the method of path calculation. In this case the following
equations are continuously and automatically solved:
I
S
t
''..; 41.,;;;, I
e 21t it*'
0
ct's k(cos ) it:
r V .ku k+ dt,
?*ft)
where U is the windspeed,
B is the wind angle, i.e., the angle between the meridian and the wind vector.
The principal parts of the automatic navigator (Fig.87) are as follows:
A - The ground speed pickup. Since such an instrument does not exist at the
present time, the indicator part of a true-airspeed indicator is used instead. Some
instruments have devices that apply corrections for the wind speed. When the pick-
up of the true wind speed is used, the readings of the automatic navigator may show
an error whose value is smaller the higher the flying speed.
B Compasses.
C - Coordinator.
The airplane aircraft course or, more accurately, the values of sin k and
cos k are picked up by the compass, this is most often a magnetic compass) together
with the coordinators.
D Multipliers, which multiply the values of the airspeed by sin k, cos k,
?1-4 and likewise the wind speed U by cos 9B, and sin 9B (98 being the mind
_colt V
angle).
148
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?.?
ro,
fp,
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ow,
E summing device giving the sums:
tude.
V sin k U ?in 9B
V cos k U cos w
- Integrators making it possible to obtain the values of latitude and longi-
Latitude and longitude indicators.
RAesAWAN,
ibro#40
AN,47/04.41e44ri.cf
V sin le
so k
CAW/NA Tele
semd/wA45
Device
1./
stnn
Wel
?U.Sin
Wel
cos k
.40VAC17"/
?ewes
Ivor s n via ta...m.
mireeke."24,
ofirryptarR
AWACS - s
eiLec
coo/to/morale
0.0144;17ri
$aletfer Avres.earae
(MOWN foR
1/COS (IL
V COSI( 1.1 "-6 Nar;
Nujoof
U cow.
Cterirl
Fig 87 - Block Diagram of Automatic havigator
? Since, in practice, the aircraft must often fly in a direction laid out in
:advance, the position of the aircraft can be determined in an arbitrary system of
'coordinates (Fig.88).
-1 g
Knowing the values of L and D*and the points of take-off the value of the
graphic coordinates of the aircraft can be found.
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The idea of the automatic navigator is best realized in an instrument that
indicates the position of the aircraft directly on the map. Such automatic naviga-
tors are particu]arly convenient for short
flights.
0 biet tit
For long flights such an instrument
of
requires the use of special maps on which,
within the limits of each sheet, the scale
V
1 .
is kept constant while the lines of the
*
various courses (loxodromes) are recti-
Section L.
RADIO INSTRUMENTS
Modern radio instruments for deter-
mining the altitude, the distance traveled
and the location of the aircraft, as well
Fig.88 - Determination of the Position
as radar devices for determining the
of the Aircraft in Arbitrary Coordi-
ground speed, angle of drift, and wind
mates
speed, are all based on the principle of
L - Distance between initial and final
radio telemetry, developed by the great
points of flight; D - Deviation from
Soviet scientists A.I.Mandelishtam and
assigned route
? N D.Papaleksil and by many Soviet engin-
eers. As far back as 1932, the Soviet engineers D.A.Rozhanskiy and Yu.B.Kobzarev
developed and worked out the principles on which the operation of modern radio
'location instruments are based.
The principles of the determination of the direction based on radio methods,
were worked out by M.V.Shuleykin? A..N.Shchukin, H.Ye.Starik, and others.
Tadio-sotdes of the P.A.Mblehanov system are used for meteorological explora,-
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tion of the air.
Radio instruments are being more and more widely used. But their introduction
does not exclude the use of mechanical and electromechanical instruments, although
the latter are often considerably less accurate. This is explained by the danger
of interference in modern radio instruments, as well as by the impossibility of
using the operating principles of radio instruments in the design of certain instru-
ments (for example of absolute or relative altimeters). The fact that the opera-
tion of aircraft radio instruments depends on the operation of a radio transmitter
on the ground is a major disadvantage of such instruments.
Among radio navigation equipment, the following are most widely used: radio
compasses, radio semi-compasses; receivers of radio beacons, radio altimeters,
course, glide, and marker radio receivers, radio receivers of the hyperbolic and
circular systems of radio navigation, radio telemeters, radio automatic navigators,
etc.
Radio altimeters and radio compasses are very widely used in aviation.
Radio Altimeters
Aircraft radio altimeters solve one of the most complex problems of piloting
and navigation, the problem of determining the true altitude of flight. Knowledge
of the true altitude is necessary for proper landing, for flying over mountains, or
for determining the height over a given objective. The knowledge of the true alti-
tude is particularly important in blind flying (in fog, in clouds, at night, etc.).
According to the range of altitudes measured, radio altimeters are subdivided
into the following classes:
1 - Low-altitude altimeters, used for determining altitudes to 1500 in.
2 - High-altitude altimeters, for altitudes above 1500 in.
Low-Altitude Radio Altimeters. Figure 89 shows the operating principle of a
?w-altitude altimeter.- The transmitter installed on the aircraft generates electro-
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46`,M
magnetic oscillations whose mean frequency (of the order of 440 megacycles) can be
regulated by a modulator. The modulation frequency is of the order of 120 cycles.
The antenna of the transmitter continuously emits electromagnetic waves toward the
earth, and these waves, after being reflected from the ground, are received by the
antenna of the receiver.
The frequency of the emitted signal differs from the frequency of the reflected
signal by a quantity equal to the change in frequency of the transmitter during the
time of passage of the signal to the ground and back. By virtue of this fact, a
beat is produced in the detector of the receiver, in which the direct and reflected
signal are combined, As a result a low-frequency voltage is tapped from the detec-
4 . , ?
W
.., ..,
?
.4 :-
1V4iftr......----?..
--. .-
1
-1
Fig.89 - Operating Principle of a Radio Altimeter
? tor equal to the difference between the frequencies of the emitted and reflected
signals. -
This voltage is amplified by a low-frequency amplifier and is then fed across
an amplitude limiter to a frequency meter, where it is converted into direct current,
whose magnitude is directly proportional to the beat frequency.
An altitude indicator with a scale calibrated in meters is connected across the
!output of the frequency meter. The direct-current voltage received in the frequency
-meter does not depend with complete linearity on the altitude. There are methods
of transforming this actual relation into a law of direct proportionality. The
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.voltage received is fed to an electric differentiater whose output voltage is propor-
tional to the derivative of the altitude. Consequently, a voltmeter measuring the
voltage may be calibrated in values of vertical velocity, i.e., a clitb indicator
may be designed on this principle.
Figure 90 shows the processes taking place in the altimeter.
? The send line shows the variation in the signal frequency entering the receiv-
er directly from the transmitter; the broken line shows the variation in the frequen-
cy reflected from the ground. At constant altitude, the frequency of the second
signal varies by the same law as the frequency of the first signal, but with a lag
by the time T. The higher the altitude of flight, the greater is the shift in time
between these two curves and the greater the difference frequency between the direct
and reflected signals, which characterizes the altitude or the aircraft. This fre-
quency is called the beat frequency.
The beat frequency, that is the frequency of the low-frequency oscillation gen-
erated in the receiver is a constant quantity, except for short time intervals when
it drops to zero. At the detector output, a voltage of the difference frequency
P
f1 " - f2 - . L.106 AfF .
? 4
is generated, where Fp is the beat frequency in cycles;
Al' is the difference between the maximum and minimum frequency
of the transmitter;
F is the modulation frequency in cycles;
H is the flight altitude in meters;
a is the velocity of propagation of the radio wave;
k is the proportionality factor. ?
As a rule, the law-altitude radio altimeter has two measuring ranges and a
switch is provided for change-over.
The first range from 0 to 150 m is used in take-off and landing; the second
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range, to altitudes of 1500 m.
The accuracy of such an Inst. ent.Ie 5% i 2 m.
The use of a frequency-modulated altimeter for measuring high altitudes is im-
possible since in this ease we would require a very high-power which
would be capable of absorbing all the energy of the electric system of the aircraft.
High-Altitude Radio Altimeter. For measuring altitudes of flight above 1500 m,
effloGied
ri
&GNATS
r ne,r t
49041 -
AVM"
?
Low
pritlefety
4 fi
not*
==.
Fig.90 - Principle of Operation of Radio Altimeter
a - Variation in frequency of signals; b - Beat frequency
pulse radio altimeters are installed on aircraft. Such a radio altimeter consists
of a radio pulse transmitter and a radio receiver both in a single unit, of a trans-
mitting and receiving antenna (of symmetrical vibrators) and of an indicator unit.
-Figure 91 is a block diagram of a high-altitude radio altimeter. The powerful radio
transmitter with a narrow directional radiation its short pulses spaced at uniform
time intervals T with a strictly constant high-frequency repetition (the repeat fre-
:queney is 100 or WOO cycles, and the pulse period is 0.5 or 1 microsec). This reg-
ularity of generation is assured by using a special instrument called a thronizator.
The pulses so generated are fed into the modulator of a high-frequency vacuumr.
'itubet-oseillator. The-latter generates high-frequency pulses in the decimeter wave
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1band and radiates them into the lower hemisphere to the antenna.
The time intervals between the pulses are many times the duration of the pulses
'themselves. The pulses of the radiation are propagated with the speed of light, c.
In the intervals between thepulses,:the radio receiver operates and picks up
:the reflection of the signal from the ground, with a lag of time T equal to the time
'required for the pulse to reach the object and return to the receiver after reflec-
tion. The time T depends on the flight altitude H:
2H.
- 77'
cT
H m
24
(TV.35)
The direct proportionality between the time T and the altitude of flight allows
indica tcw
ft** Is
ametlier
Mft4110411
00,11144
MOUND
Orposmillhor
Comiemeor
int-y(144
The smith)* 4nte17na
Fig.91 - Diagram of High-Altitude Radio Altimeter
;the distance H to be determined by measuring the time T between reception of the
5pulses and reception of the reflected signals.
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Since the velocity of propagation of the pulses is very high (300,000 km/sec),
the time T so obtained is very ?small (1 microsecond corresponds to 150 m). At the
same time, the accuracy of measurement of these short time intervals must be very
high.
The receiver of a radio altimeter is constructed on the conventional superheter-
odyne principle. The signals, after detection, are amplified in an amplifier and are
then fed to the indicator.
The simplest and most perfected in-
strument for measuring time is the cathode,
ray tube, which is similar to the tubes
used in osc.illogaphs and television sets.
The diagram of such a tube is shown in
Fig.92. The electron gun (1) throws a
narrow beam of electrons on the screen,
which is covered by a substance that
fluoresces when electrons impinge on it.
Fig.92 Diagram of Cathode-Ray Tube
At the point of impact of the electron
VI.
win, __
? beam, a bright luminous spot is formed.
3,4,5,6 - Metal plates
If the direction of the electron beam is
-rapidly and periodically varied, the luminous spot will be rapidly displaced on the
::.screen, forming a luminous line on it. At a large number of such displacements per
second, taking place in oneand the Same order, the eye perceives the trace of the
'electrons on the screen as a stable image simi ar to the flashing of images., in-
visible to the eye, on the cinema screen. The electron ray passes between two pairs
? of metal plates, 3,4,5, and 6 which are arranged in mutually perpendicular planes.
If a positive voltage is impressed on one of the plates and a negative voltage on
the oppositeplate, the electrons are attracted by the positive plate and repulsed
by the negative plate. As a result, the electron beam and consequently, the lumi-
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nous spot mill be displaced in the corresponding direction.
In this may, by the aid of plates (3) and (4) the luminous spot can be ahifted
on the screen along the =taxis, and, by means of plates (5) and (6), along the yy
axis.
0
CIS C I 4 h.? ri .1 of rt.(' le 1 117 144 r
Alt
Rotectite astiliatitons eeci;vied by rectivrr
1
ro-r
-07
1
1
i
I I
I
I
I Ii
f 11
1
1
n
I
4
1
t
Fig.93 - Variation of Voltages on the Deflecting Plates of a Cathode-Ray
Tube in the Pulse Radio Altimeter
The alternating voltage Ux, varying by a sawtooth lam as shown in Fig.93a is
supplied to the plates (3) and (4). After each abrupt change in voltage, the lumi-
nous spot will be in the extreme left position. Theft, as the voltage Ux v4rie3
smoothly, the spot is displaced to the right. Vhen the voltage Ux becomes egaal to
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,
'zero, the luminous spot is at the center of the scale. On further variation of the
voltage, an ever increasing positive potential is created at the plate (4) and the
electron beam is displaced to the extreme right-hand position. In this position, the
voltage again abruptly changes its polarity to the opposite value, and the ray, being
then attracted to the plate (3), passes to the extreme left position, after which it
a - Cathode-ray meter; b - Circular-motion electron meter
again begins smoothly to move to the right, and so on.
The pulse signals of the radio transmitter are emitted each time the voltage
Ux undergoes an abrupt change. In the instrument, the signals of the transmitter
and receiver are rectified, resulting in two voltage pulses as shown in Fig.93b, for
,ea(41 cycle of variation in the voltage Ux. It is these pulses of U7 that are fed to
/ the plates (5) and (6) of the cathode-ray tube. In this way, during the time of its
*otion along the =axis, the electron ray is twice thrown into the direction yy? as
istwen in Fig.94a. The distance between the obtained image and the straight line is
avoporti.onal to the time T between the pulses, and to the altitude H.
Br using another method of controlling the electron ray it is possible, instead
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of a rectilinear displacement of the light spot along the horizontal axis, to obtain
a circular action. In this case the initial mark and the mark corresponding to the
measured altitude take the form of peaks or pips on a circle (Fig.94b). The existing
models of high-altitude radio altimeters have an accuracy of measurement of the order
of 0.25% 20 m.
? Radio Semi-Compasses and Radio Compasses
The methods of determining the position of an aircraft based on the use of ra-
dio semi-compasses and radio compasses are very useful to aviation in many cases,
particularly in high latitudes, where the use of magnetic compasses is difficult or
impossible.
The radio semi-compass and the radio compass allow the direction of the axis of
the aircraft to be determined with respect to a ground radio station. The radio
camas* differs from the radio semi-compass in that the latter automatically indi-
cates the direction of the ground radio station, while the former requires the in-
tervention of the navigator.
By using a radio semi-compass the angle n between the direction of flight
toward a given radio station and the longitudinal axis of the aircraft can be de-
termined.
By means of the radio semi-compass and a magnetic compass, we may determine the
angle of true radio bearing flY 0, i.e., the angle between the meridian at the posi-
tion of the aircraft and the direction to the radio station (Fig.95). By connecting
ail points of the earth's surface at which the angle of true radio bearing has one
and the same value, we obtain the position line, i.e., the line of possible posi-
tions of the aircraft. The radio semi-compass does not indicate in exactly what po-
:sition on this line the aircraft is located. In order to determine the aircraft po-
;
Aition auxiliary methods must be used; for example, the true radio bearing of two
4mnd-transmitters-may be found and in this way two position lines may be obtained,
.?
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whose point of intersection determines the position of the aircraft. The diagram of
the radio semi-compas is given in Fig.96. The instrument has two antennas, one
loop and one open vertical antenna. The relation between the strength of reception
Fig.95 - Reading of the Radio Compass
P - Ground radio station; ity- Course
angle; fl ? Bearing
tor
.141.10.0=???
???????
4411
Fig.96 - Summation of Signals from
Loop and Vertical Antennas
by the loop and the angle y between the plane of the loop and the direction to the
radio station is given in Fig.97. The signal becomes loudest when the plane of the
?loop is directed toward the transmitter, and is equal to zero at two dia.:metrically
opposite positions of the loop, 900 and 270?, when the plane of the loop is perpen-
Aicular to the direction of the radio station. In order to determine which of the
positions of the loop is the correct one the open antenna is used.
The high-frequency signal voltages from both antennas are combined in the com-
Jaime circuit (Fig.96). By choosing the mutual inductance M, the voltage of the open
*antenna (a constant quantity) is mads equal to the maximum intensity of the voltage
the loOp-Con the coincidence of its loop with the direction to the radio trans-
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mitte ). By the aid of capacitor C, the phases of the voltage from the vertical an-
ltage from the loop are made to coincide when its plane coincidei
Fl
.57RINeirW
Or.ferePriew
270* 150
Diagram of the Directivity of the Loop
with the direction to the transmitter. This coincidence is possible only in one of
the two positions of the
loop at which audibility is zero, since in the diametrical-
ly opposite positions the sign (phase) of
the voltage changes (Fig.98).
Fig.98 - Obtaining the Voltage
Difference U1 - U2 for the Indicator
of the Radio Semi-Compass
;directions is not the only possible one.
By adding the voltages of both an-
tennas we obtain a total characteristic
from which it will be seen that at y 120?
the total voltage is zero.
Only in this position of the loop
will the signals from the transmitter on
the ground fail to be heard in the ear,-
phones of the receiver. At Y = 3600, the
toU.l voltage reaches its maximum.
This method of determining the two
Very often the following procedure is used
tinstead: If the positions of the ends of the loop are interchanged, then the sign
f=2,
_jot' the voltage coming frosithe loop antenna is reversed; in this case, the total
I voltg? is equal to CVat an angle Of 'Y 0, instead of at ia IN?.
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n the radio semi-compass the ends of the loop are periodically switched over
,;.9..that:the voltage U impressed on the indicator consiStP of the difference between
the two voltages U1 and U2,- depends on the angle y of rotation of the loop
(Cf.Fig,.98), and has a positive value at angles from 0 to 900 and at angles from
,270 to 360P? and a-negative value at, angles from 90 to 2709.
At a voltage Of U u 0, the pointer of the instrument is at the midpoint of the
scale. When U is positive, the pointer deviates to the right, and at negative value
of U, to the left; in the former position of the loop, when y 900, the reduction.
:in the angle, i.e., the rotation.of the leftward' rotation of the loop toward the
Asero position, leads to a rightward deflection of the instrument pointer. In the
Asecond position, when y is 270? such a rotation leads instead to a leftward deflec-
tion of the pointer. Consequently, the first position will be the correct one, and
the radio station is actually located to the right of the perpendicular to the plane
of the loop. In order to determine the line of possible positions of the aircraft,
the loop antenna is used, since in this ease the minimum of reception will be sharp-
er. To find the correct position, both antennas are used.
By the aid of the radio compass the same problems are solved as by means of the
radio semi-compass, except that in that case the correct position of the loop of the
. radio compass is automatically found by means of the following system, which sets
' the loop in a certain way with respect to the radio station.
The automatic radio compass consists of a radio receiver which automatically
,,..,locates the loop antenna, an automatically rotating loop antenna, an open nondirec-
_tional antenna, a bearing indicator, a control board, and a relay box. In tuning
the receiver to the wave of the radio station, the output signal from the radio re-
ceiver is supplied to the mechanism that rotates the loop, ?forcing it to turn this
--:loop until it coincides with the direction toward the radio station* After setting
-,the loop in the direction or the radio station, signals in the receiver disappear
4and the rotations of the loop stops. The loop is connected with a synchronous elec-
162
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trio transmission from the indicators of the radio compass, whose pointers change
their position when the loop it rotated and stop together with it showing the bear-
ing of the radio station.
If the aircraft continues changing its course until the pointer of the indica-
tor is at zero, and then keeps the pointers of the indicator at zero, its flight in
the direction of the radio station will be insured.
Figure 99 shows the block diagram of the radio compass.
Fi 99 - Electric Block Diagram of the Radio Compass
In the loop antenna, the incoming electromagnetic oscillations set up an emf
'whose value is smaller the closer the plane of the loop is to a position normal to
the direction of the incoming waves. At the moment when the plane of the loop is
.perpendicular to the direction of the wave, the emf of the loop is eaval to zero.
On passage through the zero position, the amf of the loop changes its phase by 180P
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with respect to the emf of the nondirectional antenna. In addition, the emf of the
loop has a 900 lag relative to the emf of the open antenna.
? The omf of the loop, after amplification in the high-frequency amplifier stage,
is fed to the phase shifter where it is
given an additional phase shift by 900.
Now the ere of the loop may either coin-
cide with the enf of the nondirectional
antenna or (when this loop passes to the
zero position) it will have a pha dif-
fering by 1800 from the emf of the anten-
na. The balance modulator on which the
emf of the frame is impressed, is designed
so as to indicate its phase and, conse-
quently, also the direction of any devia-
tion from zero position.
The low-frequency (60-100 cycles)
oscillator feeds two generated voltages
With opposite phases to the grid of the
modulator tube, a balanced modulator, to
put of first half of signal from balancing
which the emf of the loop is also fed.
Fig.100 - Diagram of Voltage in the
Channel of the Balancing Modulator?
- Signal of loop after a 90? phase
shift; b - Modulator to low-frequency
voltage in balancing modulator; c - Out-
..thodulator; d Output of second hall of
signal from balancing modulator; e - Sig-
the channel of the balanced modulator,
nal from nondirectional antenna; f- Re-
The graph (Fig.100a) shows on the
sultant signal after combination of mod-
left and right the variation of voltage
-ulated signal and signal from nondirec-
in the loop when it is deflected to the
left or right of the neutral position,
his direction is fed to the grid of the modulator tube (a double triode).
-,Figure .100b shows the Variation of the voltage from the low-frequency oscillator
STAT
Figure 100 is a diagram of the voltage in
tiona1 antenna
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on one of the grids of the tube. When this voltage is positive, a h gh-frequency
current corresponding to the voltage on the grid, passes through the plate circuit
of this section of the tube, an showii in .Fig.lOOc.
At this time, the voltage on the second grid of the modulator tube is negative,
and there is no current flow in the plate circuit of the second section or the tube.
When the low-frequency voltage on the first grid becomes negative, the current in
-11f1fIrPteflIt trittPlff Mgr rtrrtritfirtIFIPPOtIft tior1ir1PPOrstrtflf
Fig.101 - The Radio Compass Set
Directional antenna; 2 - Nondirectional antenna; 3 - Radio receiver; 4 - Re-
lay box: 5 - Bearing indicator; 6 - Bearing indicator, pilot; 7 Control panels;
-8 Inverter; 9 - Headphone
the plate circuit of this section is interrupted. The voltage from the second
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grid then becomes positive, and a high frequency current flows in the plate circuit
of the second section, as shown by Fig.100d. The current of this section of the
tube, in the winding 1' (cf. Fig.99) of the transformer T, has a direction opposite
to the plate current of the first section of the tube. For this reason, the high-
frequency current in Fig.100c is shown as opposite in pease to the , h-frequency
voltage coming from the loop antenna (cf. Fig.100a).
In the second y winding of the transformer T the unmodulatcd oscillations of
the same frequency coming from the nondirectional antenna. across the nrimary wind-
ing 1" are added to the modulated high-frequency oscillations from the loop. The
current in this winding is shown in Fig.100e.
The result of the combination of the oscillation shown in Fig.10Ga,c d, and e
is shown in Fir.100f. These are modulated oscillations with the modulation phase
of their amplitude being determined by the direction of the deviation of the loop
from the zero position.
The total high-frequency voltage is amplified in the receiver, after which it
is detected and is again amplified at low frequency. At the output of the receiver
a low-frequency voltage, corresponding to the law of modulation of the input signal,
is Obtained.
e
14 J544JrJ
? deflection of the loop.
At zero position of the loop, there is no high-frequency voltage. In this
there is also no low-frequency voltage at the output of the receiver.
The voltage taken from the output of the receiver is amplified in a magnetic
power amplifier and is-fed to one of the phase windings of a two-phase asynchronous
motor. The second phase winding is fed by alternating current from an inverter.
The motOr rotates the loop until it is set in zero position, after which the low-
frequency voltage at the receiver output disappears and the motion stops. When the
loop is deflected toward the Other side, the low-frequency voltage changes its
phase by 190?, as a result of which the direction of rotation of the motor is
case
this voltage likewise depends on the direction of the
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raver5ed, and the motor again returns the loop to the r4ero position.
? The radio compass is f d with direct current (tubt filaments, control signal,
? etc.) from the aircraft 27-volt system while the alternating current feed (supply-
ing the circuit of the radio receiver, motor, etc.) is supplied by a 115-volt
400-cyc1e inverter.
There is a control panel for the remote control of the c gure 101ompass.
shows the compl t set of the radio compass.
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CHAPTER V
GYROSCOPIC INSTRUMENTS
Section 1
MEMENTARY THEORY OF THE GYROSCOPE
The Concent or the Gvroscone
In technology, a-flywheel (rotor) 1 (Fig.102) rapidly rotating about its axis
of symmetry, held by one or two likewise movable rings (frames) 2, clledgiMbals,
is known as a gyroscope.
Fig.102 - Gyroscope with Three
Degrees of Freedom
1 - Rotor; 2 - Pram,.
cfi
Fig.103 - Angles Determining the Di-
Yr
rection of the Rotor Axis in Space
U - Azimuth; 0 - Altitude; Q - Velocity
of the natural rotation of the gyroscope
Depending on the structure of the suspension, the number of degrees of freedom
of the gyroscope varies. With two movable frames the gyroscope has three degrees
f freedom, since the rotor may rotate about three mutuAlly perpendicular axes xx,
yy. zz. If one frame is made immdbile, the gyroscope will have two degrees of
'freedom.
The axis xx of rotation of the rotor is called the principal axis and the
,rotation of the rotor about this axis is known as the natural rotation of the gyro-
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?
t.
scope. The iuibai i1 ow the rotor axis, to occupy any position -in space determined
by the azimuth 0 and the elevation .0 (Fig.103). If the moments of the forc8 of
, gravity-With respect to all three,gyrpscope axes re equal to zero, the gyroscope
- iO called astatic. This is possible is two
, 1. When all three axes of ,the gyros of xx, yy, and zz,intersect in a single
point which remains fixed when the gyroscope aor nd coincides with the center of
gravity of the gyroscope;
2. When the center of gravity of the rotor lies on the xx axis of natural ro-
tation;
the center of gravity of the system rotor - inner ring is located on the
zz axis of this ring; and the center of gravity of the system - inner rinF
:outer ring is located on the yy axis of the outer ring.
not equal to zero, then such a gyroscope is called a gyro pendulum.
A gyroscope is known as a free gyroscope if no external forces or narents act
on
In considering g,yrosCopc phenonna we are dealirc with the rotary motions of
body which are characterized by the direction of the axis of rotation, the sense
the rotation, and the agu1ar velocity of rotation Q., i.e., they may be charac-
--vector Q to be placed in such a way that, when looking from the point of the vector,
,system. The lensth of the vector Q
- to the magnitude of this velocity.
The value of the 'angular velocity is measured in 1/sec.
Lolutions per ninute NI or per second ns is knoi, then
- ( -
--I from the equation
,
the quantity Q may be fourxt
? - - - ? - - " ' ???'"-z=-
4
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For approximate calculation, the equation
used.
The rotarymotion is due to a moment of force or a couple.
The noent of the force is calculated by the formula
where F is the force;
R is the arm of the force;
a is the angle between the direction of the force and the axis of rotation.
If the force acts at an angle of 900 to the axis, then
The moment of force is represented likewise in the form of a vector placed along
the axis of rotation and directed in such a way that, When looking from its ti
the force applied to the body tends to turn the body counterclockwise. The kinetic
energy (vis viva) of a rotating solid body is expressed by the formula
rat 1 Inv* ?
2 ... 2
but since / mr2 f r2dm ,,Tx it follows that
??
2 n2
4
where m is the mass of an elementary pirticle of the given body
V is the velocity in az/sec;
12 is the angular velocity of natural rotation;
r -is the -distance of the given particle from the axis of rotation;
six ?is the moment of-inertia of the body with respect to its axis of rotation,
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in gm-mt- se-6
6
? Knowing the mass of the body and its form we may determine the moment of iner-
tia of the body with respect to any axis. If the body i. of complex form, it i
divided into simpler bodies, and their moments of inertia with respect to one and
the same axis are found. Then, by summation, the moment of inertia of the whole
body?is found with respect to the same axis. In cases where the form of the body
is so complex that the moments of inertia cannot be calculated by this method, the
moment of inertia of the body is found experimentally.
The moment of inertia of a rapidly rotating symmetrical eyroscope with respect
to the proper axis will be d noted in what follows by the letter C, and with respect
to the other two axes by A.
If a rotating body is exposed to the effects of several forces, the sum of the
moments of which, with respect to the axis of rotation, is equal to zero, then the
body will move without angular acceleration.
if the sum of the moments applied to the body is not equal to zero, then the
body will move with an angular velocity which is measured by the increment of angu-
lar velocity of rotation in unit time.
The displacement, velocity, and acceleration of a body, relative to a fixed
system of ccordinates are called absolute; and those relative to a moving system of
coordinates are called relative. The motion of - moving system of coordinates
with respect to a fixed system is called transport motion and is characterized by
the transport velocity. Vt and by the acceleration at.
The absolute velocity V of a certain point A which is in complex motion is
4. egmalto.the geometric sum,of the relative velocity Vo and to the transport veloc-
ity Vt
By analogy, the absolute acceleration 7; m 30 .4-?
at.
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If a transport motion is rotary, then the rotary acceleration ac, -which is
o termed the Coriolis acceleration, must be added to this acceleration:
111???
+ a
Let us assume that Oxy is a certain fixed system of coordinates (Fig.104) and
Oxiy, is a moving system of coordinates,
the system of coordinates Oxlyi being in
uniform rotation at the angular veloci-
ty about the point 0. Then the angu-
lar displacement (41 of this system in the
time T will be
Fig.104 - Determination of Absolute
Acceleration at Rotary Transport
Motion
Oxy - Fixed system of coordinates;
?ova. - Moving system of coordinates;
- Angular velocity of rotation of
moving system of coordinates; r - Dis-
tance of point A from origin of coor-
dinates; V - Absolute velocity of
Let us assume that, in turn, the point A
on the axis -Oxi is '44''ln^A4
nect__
to this
axis by
the velocity
- Relative velocity of
- Transport velocity of
point A
port velocity Vt Yr o
he velocities
= dr
'0 dt 2
where r is the distance between point A
and the origin of coordinates. The tra-
jectory of the point A, performing a coal--;
plex motion, lel) be a spiral. The abso-
lute velocity V of the motion of the
point A represents the geometric SUM of
the relative velocity Vo and the trans-
of this point
&WM
+r
The projection of the absolute velocity onto the fixed system of coordinates may
be found from the equations
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In order to find the projection of the absolute' acceleration onto the axes x
and y) we use. the derivatives of these equations with respect to time
$incew2r = a, (transport acceleration) we find the equation for the instant of
reading t, when the axes Oxi and Ox coincide: = 0;
We Obtain the absolute accelerhtion a as the geomeLric the
zum of
three
erations
The acceleration 2 Vow is called the rotary or Coriolis acceleration ac. The
'direction of this acceleration is obtained by rotating the vector of relative ve-
locity Vo (in our case it is directed along the axis 04 through 900 in the direc-
tion of the trapsport rotation w. In the example we are discussing the rotary
acceleration is directed along the axis Oy.
If the vector of relative velocity Vo is not perpendicular to the vector of
the transport angular acceleration w and forms the angle a with it (Fig.105) then,
by resolving this velocity into a parallel and a perpendicular component with re-
spect to the,aris-of transport rotation we find that the rotary acceleration is
due only to the perpendicular component of this velocity Vo sin a. The value of
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The direction of the acceleratio. a0 is d termined by turning the component .110 sin a
through 900 in the direction of the transport rotation.
Thus, on considering the complex motion of a body with a transport rotary
motion it is necessary to take account of all three accelerations ao, at, ac.
Princioal Pro
rties of a Ra di Rotating Gyroscope with Three Degrees of Freedom
Since the solution of the prOblem of the motion of the gyroscope with respect
.to a fixed point under the action of assigned external forces is very complex,let
us consider instead the inverse problem:
Let us assign the motion and find that
cud forces or moments cause that motion, and
whether this is the result of the action
21,;(aisila
10;siirict
of external forces or moments or takes
place as a result of inertia.
Fig.105 - Rotary Acceleration of the
Let the rotor of a gyroscope with
'Body in the General Case of Motion
three degrees of freedom rotate about its.
Anglo between the direction of rel-
natural axis xx at an angular velocity Cl
ative Velocity and the vector of
(cf. Fig.102).
,transport angular velocity
At the same time the axis Ox and,
-1,41t41 it., the whole system or coordinates Oxz, connected with the gyroscope, rotates
'with the angular velocity 14 about the axis Oy (Fig.10.6a).
The relative motion of the gyroscope is characterized by the velocity Q
the'transpOrtmotion by the velocity W., if 4P
The relative linear velocity of any point B of the rotor (Fig.1060'is equal
Ito',ur-QT -Iiirre-r is-the distance-of this point from the as of rotationi
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Kr.
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a'
For points ly
by the xpresion
on the periphery of the rotors the velocity Vo is determined
where R is the radius of the rotor.
21
if
a
C
sls
Q
vo
8
?
Po
/ha
fig.106 - Precessional Motion of the Gyroscope
Q- Velocity of relative motions of the gyroscope; Transport velocity
of the gyroscope; R Radius of rotor; Vo - Relative linear velocity of
any point of the rotor; mi, m2, m5 m, s
- Mases of points lying on the
4
rotor rim; B - Arbitrary point on the rotor body
The direction of the velocity Vo is shown in Fig.106b.
The angular acceleration generated by the motion of the gyroscope under consid-
eration is caused by the component of this velocity along the axis Oz, i.e., Voz
Vo sin (4
12 r sin a.
The value of the angular acceleration is determimd by the relation
Vo w sin a
For all points of a !:kody located in the region I, the acceleration is directed
pward while for points located in region II, it is directed downward (with the
175
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ta?
selected directions of Q and to).
At the points A&, these velocities are_zero[Voi while at the points CC the
accelerationa reach their maiinum value, Since Voz V0aqQR.
It is clear from this that the forces producing the rotational acceleration
and causing the motion of the rotor about the axis Oy with the angular velopityw
'form a couple acting with respect to the axis Oz, about which, by hypothesis- there
is no motion.
The vector of the moment-M of this couple is directed from right to left.
' Thus, at a transport motion of the rapidly rotating gyroscope about the axis Oy,
with an angular velocity w, due to the couple acting about the axis Oz, the points m
of the rotor will have an acceleration of the direction shown in Fig.106s. Since
these accelerations are not the result of the action of internal forces (the reac-
tions of internal connections) but appear as a result of the interaction of the
components of the relative velocity Vo of the gyroscope, directed along the ZZ axis
;(the velocity Voz), and of the transport Velocity w, it follows that the gyroscope
'motion under consideration, about the axis Oy, is the result of the action of ex-
.:.ternal forces or moments. This motion is termed precessional, and its velocity is
'called velocity of precession. The precessional motion of a gyroscope, at first
glance, seems to be paradoxical,- since the external moment M acts about one axis of
'the gyroscope while this motion takes place about another axis of the gyroscope.
It will be easy to convince oneself from the following example that this contradic-
-tion is an apparent one.
A body being displaced in a horizontal plane in a vacuum at the initial
Istant of time with the velocity V, under the action of the force of gravity begins
Hto vary its trajectory of motion (which will curve downward along a parablol). .In
this case, the direction of the velocity vector of the potion of the body will like.-
wise vary, and this variation will be slower, the higher the initial velocity of the
, body.
Pt.
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'17)
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Let us return to the roscope rotor. Assume that a gyroscope (Fig.107) with a
fixed, point 0 corresponding to the center of gravity, is subjected to the moment of
Ahe couple M. It may be assumed that forces directed upward act on the points of
the rotor in the region I, and forces di-
rected downward on those in the region U.
The point A of the rotor is displaced
with the velocity Voz (the linear rela-
tive velocity of this point Vo = QR). Un-
der the action of the downward forces, the
trajectory of the point A must likewise
curve downward, and the projection of its
velocity vector Voz onto the plane x0z
likewise occupies the position Vozi in-
clined to the plane of the horizon. Con-
sequently, the point of the periphery A
rotates and moves into the position D.
This means that the whole rotor, together
Fig.107 - Trajectory of an Arbitrary
with it, has rotated about the yy axis in
Point of a Gyroscope Rotor in Pre-
such a way that its axis now occupies the
cessional Motion.
position zizi (Fig.107b).
Thus, under the action of the moment of the external forces 14, acting about the
,zz axis of the gyroscope rotating at high relative velocity Vo =QR about the xx
-axie, the gyroscope will begin to precess about the yy axis (at an angular velocity
Let us find the relation between theuccent of inertia of the rotor Co the an-
gular velocity of its proper rotation CI, the angular velocity of precession w and
-I the external moment IL The direction of the rotational acceleration is shown in
4ig,106 and the value of the force applied to the point D of the gyroscope which
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The moment Md, produced by the force P is found from the ?qwttion Md
where 1 is the arm of the moment equal to the distance of the point D from the ss
axis, and consequently 1 r sin a.
since r sin a y.
All the particles of the gyroscope produce a total moment equal to
2 Q tdatatiyi in2y4 ? *4
are the masses of the particles of the gyroscope;
...yn are corresponding coordinates of these points;
SW'2 is the moment of inertia of the gyroscope with respect to the
plane Ome.
Since the gyroscope rotor usually takes the form of a body of revolution
(roughly in the form of a cylinder), its moments of inertia with respect to the
planes Ozz and Oxy are equal, i.e.,. A, while the moment of inertia C
xz?xy
with respect to its axis of rotation is C . 2A . 2 m12. In this case, the ex-
ternal moment is -
,
!since CiZ 0 11 (kinetic moment of the gyroscope). The expression (V.5) is tailed
ithe equation of precession.
On rapid rotation of the rotor, the vector of the kinetic moment of the gyro-
tscoPe, 11, practically coincides with the vector of the angular velocity Qs, and at
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+OA
, the =me time with the proper axis of the gyroscope.
The direction of the precession of the gyroscope under the action of the ex7-
ternal moment may be found by the following method: Together the precessional mo-
tion of the gyroscope, its vector of kinetic motion rotates, tending by the shortest
?
path to coincide with the vector of the external moment M (cf.Fig.106).
Let Us consider th6 ease where the angle between the vectors Q and w, does not
equal_, i.e., when for example, the axis of transport rotation tai makes the an-
gle 4 with the axis xx of natural rotation 'of the gyroscope (Fig.108).
Let us resolve the vector w into the
direction of the rotor axis and the nor-
mal to' it. The vector w2= w cos a coin-
cides in direction with the vector of
natural rotation Q, while the vector
w1 a W sin a represents the angular ve-
locity of precession about the axis Cy;
consequently, now in the equation M
the quantity w should be substituted by
wimw sin am to sin (A),so that the equation
of precession in the general ease can now
cc
Fig.108 - Motion of the Gyroscope under
the influence of the Moment of External
Forces
be written in the form
M = Mw sin (42) (V.6)
It follows from eq (V.6) that:
1) The precession caused by the action of the external moment M takes place at
, constant velocity w, and at the instant of cessation of the action of the external
moment, the precession of the gyroscope likewise ceases.
-1 2) If the kinetic moment H of the rotor is large, while the moment of external
tortes 14 is small, then the velocity of precession ha M mill be vary small; it
_
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therefore be considered that the gyroscope keeps the direction of its proper
axis in space practically constant. The higher the value of H C 94 the more ac-
curately will the constant direction of the axis of the gyroscope rotor be main-
tained. For this reason, an effort is made to increase this moment as much as pos-
sible in gyroscopic instruments, thus increasing the inertia of the gyroscope.
3) Under the action of the external moment 114 the gyroscope precesses in the
direction perpendicular to the plane of the couple. Consequently, the external mo-
ment is balanced by some moment, which is obviously the moment of the inertia
forces. Indeed, the moment that luipedes the rotation of the gyroscope in the di-
rection of the action of external forces is basically the moment of the inertia
forces due to the rotational accelerations ac. This moment is equal to the exter-
nal moment, but with reversed sign. It is termed the gyroscopic moment*.
The direction of the vector of the gyroscopic moment Mg may be found by rotat-
ing that component of the vector w which is perpendicular to the rotor axis about
, this axis in the direction of its rotation through 9e. To determine the direction
of the gyroscopic moment, we ordinarily use the following rule: The gyroscopic mo-
merit
-
ment ME created when the gyroscope is rotated at the instantaneous angular velocity
w is equal to Mg Hw sin (Hw). This moment is perpendicular to the plane in
which the vectors H. andw are located, and so directed as to tend to bring the vee-
In this case, we neglect the centripetal accelerations due to rotation at the
, !angular velocity wt. The forces of inertia of this acceleration form a couple which
ml
_j acts as though it tended to reduce the angle Q. The external moment producing
-t the centripetal accelerations must be algegraically added to the moment producing
the rotational accelerations.
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'
1,Cc.
7yy,.
'tor toa (Fig.106a). If the friction in the, bearings of the frames in the astatic free
aroscope were assumed equal to zeros, then such a gyroscope would maintain its posi-
tion at rest, and in that case the direction of its axis would remain fixed in space,
.
regardless of the rotation of the rotor.
In reality, such an ideal suspension without friction is not feasible, and the
character or the motion of a gyroscope with a fixed rotating rotor would be differ-
ent., It is dell= trated in mechanics that even if ne external moments at all act on
the arescepe? the rotor axis will still remain fixed in space or will describe a
circular or elliptical cone, according to the initial conditions, and that the per-
led of that cone will be
I
ry,
(v.8)
where H is the kinetic moment of the gyroscope;
Hy, Jz are the moments of inertia w,,ith respect to the Oy and Oz axes of the
system of gimbals together with the rotor.
The motion of the liv-ica of the rotor
takes place without the action of an ex-
ternal moment, i.e., by virtue of inertia.
In theoretical mechanics this is known as
regular precession, and in technology as
nutation. This motion is executed at
high frequency, corresponding roughly to
Fig.109 - Regular Precession of the the velocity Q of the natural rotation of
Gyroscope the rotor and is manifested in the form
of motion of the axis of symmetery of the
rotor along a cone about the vector of kinetic moment (Fig.109). As a result of
-- friction in the bearings of the frames, the nutational oscillations are rapidly
- damped while the precession remains.
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Returning now to the question of the inertialess nature of the precessional mo-
tion of the gyroscope, the following must be remarked;
As already stated, the precessional motion of the gyroscope,is. ediately in-
terrupted at the instant at which the cauoative moment ceases to act.
The motion due to inertia, however, takes place in the form of hutationalos-
0-nations, which are so small in the rapidly rotating gyroscope that they cannot be
detected by the naked eye. The above-enumerated properties of the astatic gyroscope
(with three degrees of freedom) and particularly its peculiar stabil3ty in the se-
lected direction, i.e., the property of slowly varying its initial posItion under
the action of external moments, 'which is inherent tothe gyroscope to a greater ex,-
tent than: to any other body, opened excellent prospects for the use of the gyre-
scope in technology.
At the present time, gyroscopes are being used as systems for maintaining a
-fixed direction with respect to the vert-cal, the meridian, the fixed stars, etc.
which allow us to determine the angles of inclination with respect to the plane Of
the horizon, the course of an aircraft or a ship: etc. and also as systems for sta-
bilizing a given object, such as an aircraft, in a definite position (automatic pi-
lot).
Figure 110 shows the utilization of the free gyroscope for observing the rota-
tion of the earth.
If no external forces or moments act on the axis of a free gyroscope, it will
maintain a constant position in space. Owing to the diurnal rotation of the earth,
the position of this axis with respect to the meridian will vary. To an observer
4
on the earth, moving together with it, it will seem that the axis of the gyroscope
jialso rotates and describes, during one rotation of the earth, a complete cone whose
f axis is parallel to the ,axis of rotation of the earth. The observer sees the ap-
.,,?
Declassified
parent motion of the gyroscope.
Thus, even if it were possible to make an ideal free gyroscope, it could not
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4..,????
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eOlo
4
be used for determining a direction that was constant with respect to the earth,
since its axis maintains its direction constant with respect to space. For example,
if the axis of such a free gyroscope at any point of the earth were directed along
the vertical, then at all other points of the earth it would indic te the former di-
rection in space, i.e., a direction different from that of the local vertical,
a I veir.
0 to p *A rit 111040f:Vi eir\f
1
4141, --Crt'iarree
i
219
Fig.110 - Apparent Motion of the Free Gyroscope
In addition, if such a gyroscope, under the influence of any forces or moments,
'were to deviate from the initial position., then it would no longer return to it but
:mould instead tend to maintain its new position. Such a gyroscope does not possess
sdlectivity in the choice of its direction withrespect to the selected direction
and is ."indifferent" in the same may to any position with respect to the earth.
The deviation of the gyroscope from the initial position may take place under,.
,.the influence of the moments .of inertia forces which arise in the presence of all
types of acceleration, on displacements of the object, friction, residual unbalance,
4.El_letc. The rate of this deviation (precession), if kinetic moment of the gyro-
, i
scope H CCL is properly selected, is Mmall. For this reason, in.cases.mhen a
-idefinite direction must be maintained for'alimited and short period of. time
1
. , .10 min),,It.15:.entirely possible to use an ordinary astatic gyroscope, prodded
:it,is properly set at the initial moment
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war?
In cases where the gyroscope is being operated for a long period, or when it is
impossible to guarantee the accuracy of its setting at the initial moment, it is
necessary, even during short-period operation, to give the gyroscope axis a selectiv-
ity with respect to the earth, i.e., the power of assuming only a single selected
?direction.
Two methods are used for giving selectivity to gyroscopes with three degrees of
:freedom:
The first method is based on the displacement of the center of gravity of the
.system away ,framthe fixed point. For example in order to give selectivity with
..'respect to the vertical, the center of gravity in the system is shifted along the
rotor axie (Fig.111), i.e., a gyroscopic pendulum is used.
Fig.111 - Giving a Gyroscope, with Three Degrees of Freedom, Selectivity of
Displacement of its Center of Gravity.
The restoring moments, i.e., the moments restoring the axis of the rotor to the
selected direction, ate obtained by a very simple method; however, in doing this,
the dimensions of the instruments are increased and in addition, the oscillatory mo-
,
:tions of the proper axis of the gyroscope, generated in such instruments, require
MI6 use of damping devices which considerably complicate the design.
For this reason, a second method is used in aviation, based on the utilization
of-special:mechanisms (correcting devices) which return the proper axis of the gyro-
le4
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?;
o
*cope to the selected direction, and retain it at this direction.
The operation of the correction device is based on the following two pri ci-
ple
1. Duringthe deviation of the gyroscope, the correcting went, i.e., themo?
ment produced. by the correcting device, is applied to the gyroscope in such a waj
that it caused the rotor axis to move in the ame plane in which the deflection
tosok place, but in the opposite direction.
Consequently, the moment that eliminates the deviation of the rotor axis must
act in the plane perpendicular to this deviation.
2. The fact of the deviation of the gyroscopc from the selected direction is
deteroined by comparing the position of the gyrosco e with the position of any
other device possessing selectivity with respect to the direction selected.
This device is sometimes termed asensitive correcting element. The selection
of a sensitive correcting element depends on the fixed direction the gyroscope must
maintain. For example, if the gyroscope must be set along the vertical direction
(gyro vertical), then the sensitive element is made in the form of a pendulum.
The action of the correcting device is dependent on the comparison of the po-
sition of the gyroscope and the sensitive element.
? As a result, a system is obtained that co-mbines selectivity (the sensitive
correcting element) and high inertia (a gyroscope with a very low rate of preces-
sion).
Owing to this brief and random deviations of the sensitive element of correc-
tion to not noticeably affect the position of the axis of a gyroscope with a low
rate of precession. At the same time, the precession of the gyroscope that is' due
to moments of friction unbalances of the gyroscope, etc. is eliminate 4 in time by
means of the correcting device.
The diagrams of gyroscopic instruments with various correcting devices as well
as the characteristics of such corrective devices are described in Sections 3 and 4
:1.7- ?
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4 '
????41,..0 ,T?
?
I
c's
scope
to the selected direction, and retain it at thin direction.
The operation of the correction device is based on the following two princi-
ples;
1. During the deviation of the gyroscope, the correcting =meat i.e., the ro-
mtnt produced by the correcting device is applied to the gyroscope in such a way
? that it causes the rotor axis to move in the same plane in which the deflection
took place, but in the opposite direction.
Consequently, the moment that eliminates the deviation of the rotor axis must
act in the plane perpendicular to this deviation.
2. The fact of the deviation of the gyroscope from the selected direction is
determined by comparing the position of the gyroscope with the position of any
other device possessing selectivity with respect to the direction selected.
This device is sometimes termed asensitive correcting element. The selection
of a sensitive correcting element depends on the fixed direction the gyroscope must
maintain. For example if the gyroscope must be set along the vertical direction
(gyro vertical), then the sensitive element is made in the form of a pendulum.
711.4 A.fttion nr thn ftnnl=nntina rinvino in depnntintt nn -------- nannarinnn nr tha
eition of the gyroscope and the sensitive element.
As a result, a system is obtained that combines selectivity (the sensitive
correcting element) and high inertia (a gyroscope with a very low rate of preces-
sion).
Owing to this, brief and random deviations of the sensitive element of correc-
tion to not noticeably affect the position of the axis of a gyroscope with a low
rate of precession. At the same time, the precession of the gyroscope that is due
to moments of friction, unbalances of the gyroscope, etc. is eliminated in time by
means of the correcting device.
The diagrams of gyroscopic instraments with various correcting devices as well
' as the characteristics of such corrective devices, are described in Sections 3 and 4
STAT
? - -
185
.1
A
11
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`7 4
'
,of Chapter V.
? osco.e with Two Dec:a of Freedom.
By attaching an outer frame (Fig.102) to a gyroscope with three degrees of
reedom, i.e., by deprifin,g it of one degree of freedom (displacement about the
?iyy axial here $ 22 Op 0 " Op 0 " 0) we. geta gyroscope with two degrees of freedom
(Fig.112a).
Let us apply to the gyroscope the moment Ms acting about the zz axis.
? If the gyroscope had three degrees of freedom, then the external moment would
--be balanced by the gyroscopic mOment., and ..a precessional moment would appear about
the axis yy, in accordance with the equation
??-????,..
4 6._
Or,
Itzg
H
Since in our ease a . 0, there is no gyroscopic moment with respect to the
Motion undor
its ? *Won oj fordo
a)
t.
'St
ArrlAkiPsincr
Fig.112 ?Gyroscope with Two Degrees of Freedom
zz axis, Consequently, under the influence of the external moment /42, a gyroscope
with two degrees of freedom, in spite of the existence of a natural rotation will
behave like any other body, i.e., it will rotate with acceleration about the zz
ax-
is.
tra ?
Let us now suppose that no external moments about the zz axis act on a gyro-
-
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k,
4
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fi
42
Section 2.
scope with two degrees of freedom, i.e., that 1.1,21 0, and that the whole ini rument
is installed on a moving base, which rotates about the YY aria at the velocity ba.
In this case the gyroscope will, as it were recover its lost degree of freedom.
411 this ease, the external moment coinciding in direction with the angular
velocity of rotation of the base, is applied to the gyroscope.
A precessional moment about the zz axis now arises and tends to bring the vec-
tors H and Hy into coincidence by the shortest route.
The zz axis of rotation of the rotor tends to become parallel to the axis of
rotation or the whole system so that the direction of rotation of the rotor a will
coincide with the direction of w.
Such a displacement of the gyroscope leads to a variation in the mutual posi-
tion of the axis of rotation of the movinz base and the proper axis of the rotor
(Fig.1121); the angle a between these axes varies with the rotation of the gyro-
scope in the range from7to O.
If we prevent the rotation of the gyroscope frame, the gyroscopic moment
= H w sin (90 - a) - H 4) cos a will act on the bearings, preventing the
frame of the gyroscope from rotating. At a constant kinetic moment of the gyro-
scope, this moment is proportional to the angular velocity of rotation of the mov-
ing base and to the sine of the angle between the axis of the rotor and the axis of
rotation. This peculiarity of a gyroscope with two degrees of freedom is utilized
,- in instruments measuring the angular velocity of rotation.
MAIN PARTS AND EIENMETS OF THE GYROSCOPE
In their structure, gyroscopic instruments are subdivided into gyroscopes with
,n
4."-- three degrees of freedom, and high-speed ;gyroscopes with two degrees of freedom.
Most gyroscopes consist of the following main units:
11-base-or-botr of- the instrument;
J
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) rotor;
:3) bearings;
4) gyroscope gimbals;
) rotor motor, if the instrument ts electric;
) correcting devices;
7) arresting devices;
8) power source.
Let us consider in more detail the rotors and bearings, which are necessary
elements of any gyroscop c irstrument.
The rate of precession the to the action of the moments of external forces
(friction in the bearings, in the contacts of the servosystem, etc.) depends on the
magnitude of the kinetic moment H.
To reduce the rate of precession, the moment of inertia must be increased by
increasing the weight of the rotor, and the distribution of the principal mass must
;be as far as possible from the axis of rotation* or the rotational speed of the ro-
tor must be increased.
The possibility of increasing the rotational speed of the rotor is limited by
the wear on the bearings, which at high rotor speeds (100000 - 25,000 rpm) operate
under severe conditions. For gyroscopic instruments, special types of ball bearings
jare used, which absorb axial and radial loads with equal facility, and have minimum
clearanmas to avoid *pairment of balance. The rotor bearings which are usually
callokimain bearings, operate under more Severe conditions than the bearings of the
4 4
gimbals since the spindle resting on them rotates at a velocity many times that of
41i
Ithe outer and inner frames.
For this reason, the rotor bearings have a shorter service life than the frame
,bearings.
-A The moments of friction generated at the main bearings and frame bearings have,
different effects on the gyroscope. The moments of friction in the main bearings
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.:::;;;;''.?
,
? ' ? 4
affect only the power expended to rotate the rotor, while the friction in
ings of the suspension leads to the appearance of a precessional moment of the gyro-
The displacement of the center of gravity of the gyroscope with respect to the
various axes likewise has a different effect on the behavior of the gyroscope. The
displacement of the center of gravity along the yy axis (cf.Fig.102) does not pro-
duce a moment. Its displacement along the ez axis by the quantity ? a causes the
appearance of a moment i Pa acting with respect to the xx axis and absorbed by the ,
bearings of the outer gyroscope franc.
On displacement of the center of gravity of the gyroscope along the xx axis by
the amount of the clearance i c, a moment i Pe is produced relative to the ze axis,
Causing a precessional moment with respect to the yy axis at the angular velocity
Consequently, a clearance in the bearings of the gimbals is allowable, but in
the main bearings it must be reduced to the lowest possible value.
All the gyroscope bearings must be precisely manufactured and. must
'''----iaervice life. In addition the friction in the suspension bearings must
las possible, and there =ust be no axixi clearances in the main bearings.
The material of which the bearing is made must be very hard and
Shich-15 steel is used for the rigs.
According to the design of the inner frame, rotors are subdivided into open
;
----i and closed types.
44 f
In closed-type rotors, the inner frame consists
1
l'reitor-rotaterw -To maintain the rotor in continuous rotation at a constant angular
-
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velocity IQ, either a pneumatic or electric gyromotor is used.
The PneuPlatie ar?1140tor (Fig.113) consiete of the rotor (2) with a rim on which
a number of grooves are milled. Blasts of air from the two nozzles (1) attached to
1 _J.
the,she.u.? V4 which the inner frame consists, strike these grooves. The air brings
the rotor into rotation. Such gyromotors were very widely used a few years ago and
are still it in pneumatic gyro instruments. At the present time, however, electric
aTtmotors operating on direct or alternating current are widely used. In both
cases, they are inverted type electric motors, that is, the stator is located inside
the rotor; this gives an increased moment of inertia and consequently also an in-
creased kinetic moment.
Direct-current gyromotors are convenient in that they can be directly connected
o the aircraft DC system and have only two supply leads. A shortcoming of DC gyroe
;motors:is the rapid wear of the friction parts (collector and brushes), sparking,
f.
etc.
Alternating-current gyromotors (Fig.114) are more often used; they require no
iS lead to the rotor, having a short-circuited squirrel-cage type winding and no col-
lector.
Alternating-current gyromotors require a special three-phase generator.
Pendulum, induction, electromagnetic, or other correctors are used to stabil-
ize the rotor axis with respect to the earth. The feed is through the bearings to
the stator winding. The current-input devices must not produce an appreciable ad-
ditional moment causing precession of the gyroscope, and must not limit the dis-
placment of the gyroscope with respect to the axes of suspension.
When the range of displacements of the gyroscope is limitedoo-called moment-
053 filamentst.i.e.? spiral springs with a very low moment, may be used.
Supply leads are often designed according to the diagram shown .
-4 Sueli leads do not restrict rotation, but they can be used only with a two-wire feed
;system in form of :axial contacts at both ends of the rotor shaft. When
190
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isplac nt is I tited, several axial contacts may be used.
60-cal1ed comb contacts are often used in three-A=1e Potors (Fi 115 b).
The pneumatic instruments receive their feed' front vacuum or pressure pumps,
Fig.113 - Pneumatic Gyromoter
1 - Nozzle; 2 and 3 - Gyryscope
frames; 4 - Rotor
-.Venturi tubes, etc.
Electric gyroscopes are fed either directly from the aircraft DC system, or
!across a converter yielding higher frequency three-phase current (at about 500
cycles).
The accuracy of the readings of gyro instrumentsdepends to a considerable
iex-
tent on the relationbetween the moments Of friction in the bearings of the frames
"'--L-iand the moments of the Correcting devices, The moment of the correction devices
!must be 3 to 5 times as great as the sum of the moments of friction. To reduce the;
Jrietiam-in-the-bearingsl-vibratingt.rookingt or rotating bearingsvae-vell-as
Fig.114 - Alternating Current Gyromotor
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elastic bearing, are used.
Since the rotor of a aromotor,rotat at verY high speed, great care must be
used in its manufacture.
The rotor, and thus the weight of the gyro unit, is exactly balanced, since
any unbalance leads to' ?additional load on the bearing and causes an error in the
awilk0011111RMISMIX11???
IIdil 11
\ nk . ? f . . ?
'WW1', ,
`? ? =ris =
',Aoff.fit
NV te
441
'?) l,/17/ ?Vo
Fig.115 - Current Leads to a Gyromotor through the Bearings
a - axial lead;. b Comb lead
1 - Frame of gyroscope suspension; 2 - Shaft; 3 - Insulation; 4 - Contact;
5 - Fixed contact plates
readings of the instrument. The rotor must be made of a material of high specific
gravity which is uniform and strong, since considerable stresses arise in the body
of the rotor when it rotates at high speed. Brass, steel, and bronze are usually
employed.
If the circular velocity V of the rotor, the radius R and the density of the
t material- p are- known, then the stress in the circular rim of the rotor-is found
192
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from the .conditions of equilibrium between the centrifugal force F mi2/11 and the
?41,1aptie tensile forces P QS. An element of the rim with an infinitesimal central
angle dais separated. The rrss of this element is mr pSRda (Fig.116).
On projecting all forces in the direction of the force,?, we obtain the equa-
tion of equilibrium in the form
=-.12/-r)cos ( 90 - ?17
sin Pd2.
(v.10)
? In view of the smallness of the angle da, we may take the quantityALinstead
2
of sinAS- . By substituting the values
of P and F in eq.(V.10)we get
Fig.116 - Tension Stresses in the
Rotor Body
S - Area of rim cross section; F -
Tensile force; F Centrifugal force;
V Linear velocity; 6- Stress on
?rotor rim.
Ui 3 \
then
'whence
niv:
=== pS V' 112
',Sti=da c;Scix,
Since the material of the rotor,
and consequently the assigned allowable
stress cb is known, it follows that the
maximum allowable circular velocity is
(1.12)
In a rotor, having the form of a solid disk of a radius R and a constant
thickness, rotating at a velocity V, the maxi= tensile stresses will be concen-
trated at the center of the disk.
The?streases cr2- at the outer periphery are equal to cy2
193
?
4
2 where
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is the Poissoncoefficient; the stress at the center of the disk is
(V.13)
If the rotor has a round opening of a radius r at the center, then the modal=
tensile stresses 03 at the inner periphery are equal to
where a e-E.
In all cases considered, the value of the stress will be less than =P I/2
(since the coefficient for 072 is less than 1), and only for a thin ring, when
R r* will the stress be cl :t e oV2.
This calculation does not allow for the strain in the body of the rotor created
during stamping of the shaft nor for the varying thickness of the rotor.
In connection with the non-uniformity of the material and the inaccuracy of
.manufacture, the rotor may prove to be unbalanced both statistically and dynamical-
ly. Static unbalance- is expressed in the failure of the center of gravity to coin-
cide with the axis of rotation of the rotor, and to its being displaced from it by
:a certain distance t (Fig.11
) ?
On rotation of a statically unbalanced rotor, the centrifugal force F =
is generated, creating a pulsating load.
To eliminate this shortcoming, static balancing of rotors is used, detect' 4
the unbalance on a special installation, and then drilling holes in the body of the
,i rotor on the heavier side.
Dynamic unbalance is expressed in the production of a moment (couple of forces)
--, by the centrifugal forces that appear on rapid rotation, even though the center of
gravity of the body lies on the axis of rotation. Such unbalance ray be obtained,
194
STAT
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Vt,
for example, as a result of a skewing .of the rotor shaft during the stamping, as a
result of nonuniform material of the rotor, etc. Rotors are first subjected to
static balancingl.followed by dynamic balancing.
After static balancing, the rotor :ie
fixed in the frame and set in rapid rota-
tion. If vibration occurs this indicates
'dynamic unbalance.
Plastilline or wax are used for.bal-
ancing the rotor (Fig.118), after which
holes are drilled at places opposite the
point of application of the wax; the wax
F1g.117 - Static Unbalance of Rotor
is then removed. For balancing it is necessary that the weight f the metal re-
Loved by drilling satisfies the equation C,r1 = MI where P is the weight of the
plastilline and r, 1, R, L are the dimensions shown in Fi .118. The static and dy-
namic unbalance are tested on special instrurents.
-1
-1 the variation. of the parameters (p, T, etc.) of the atmosphere with height have -a
?J great influence on the operation.
j
-1 If the instrument operates on compressed air, the form of the blades of the
-iminiature turbine and the guide nozzles (the body of the rotor may be considered a
Prilod hole
Fig.118 Dradc Balancing of Rotor
As already stated, the increase of
the kinetic moment of the rotor is ob-
tained by increasing the moment of iner-
tia and the rate of rotation.
The increase in the rate of rotation
is limited mainly by the quAlity of the
ball bearings.
It must be remarked that, in a pneu-
matic instrument operating in a vacuum
195
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Lminiature turbine) exert a great influence on the rate of rotation of the rotor. Ir
the electric feed not only the power supplied is extremely important, but also .Ue.
t method of manufacturing the electric motor that turns the rotor, which affects the
rotor efficiency. In designing a new gyroscopic instrument, the designer in most
oases does not-ealculate-the gyromotor, but is guided by the standards in forte at
the plant.
In designing gyro instruments it is necessary to aflow for the ventialtion
losses, i.e., losses due to friction between the rotor and the air, on which as!
'much as 80-95% of all the energy required by the gyroscope may be expended.
To reduce these !Losses, the rotor is given a lenticular streamlined shape, pol-;-
ished, and placed in a special chamber whose inner surface is likewise polished.
Fig.119 - Rotors of Gyroscopic Instruments
? Figure 119 shows the most. common type of rotors. The rotors shown in Fig.119
Fig.119 a, b are charaCteristic for pneumatic low-power gyroscopes. The rotors in
--Fig.119 bandc are sometimes used in low-power electric instruments. The rotor in
r
-1Fig.119 d is characteristic of electric gyroscopes rotating at high speed.
4 8
50-
The dimensions and characteristics of rotors of a few gyro instruments are
given in Tables 3 and 4.
The rotors of. pneumatic gyroscopes are made of aluminum-nickel bronze with a
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The rotors of electric, gyro cop 5 are nude of high-grade steel with a specific
gravity of 7.8 - 8.2.
In designing a gyroscopic instrument :the specifications for it and the neces-
sary accuracy of operation of the instrument, together with its dimensions, are
'prescribed.
Table 3
Characteristics of hotors of Gyro Instruments
Model of Weight Moment of
Angular Kinetic I.nstraMent Type
Rotor in grams Inertia in rpm Velocity Moment
gp-cm-sec2 1/sec
1 700 3.5 20 000 2094 7 330 Anschuetz V-02-50
2 330 56 20 000 2094 117 000 Anschuetz three-rotor
aro
Table 4
Dimensions of Rotors
Model
1
81 1,038
62 2,355
26
35
60 1,40
130 1,122
55
55
35
? 51,5
15
40
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Section
THE TUEN-AND-B4NE IN CATOR
When an aircraft makes a turn, the magnetic compass is deflected under the in-
fluence of the vertical component of the magnetic field of the earth to such an ex-
tent that, in some cases, the instrument completely fails to show the turn or even
shows a turn in the opposite direction (turning error). For this reason, the mag-
._retic_compass is not used in. aircraft turns and is replaced by a turn indicator,
which is a compound instrument consisting of a gyroscope with two. degrees. of free-
dom and a pendulum bank indicator. The turn-and-bank indicator shows the rotation
of the aircralt.about-the vertical axis as: well as its sideslip.
The turn indicator (Fig.)20) is a glasS tube curved along a circle with a .radius
Fig.120 The Bank of an Aircraft
- Operation of the turn-and-bank indicator in a regular turn;
- Operation of the turn-and-bank indicator in a turn and bank.
Apparent hank
-ipf R and filled with liquid; the tube contains a heavy ball moving freely inside it.
--iThe behavior of the ball is analogous to the behavior of a pendulum of the length R
,and massm-9-twhereGis the weight of the ball andgthe acceleration of gravity.
411111111111ININIssi.
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When , the tube is inclined, the ball slides downward, and the line connecting the
center of gravity of the ball with the center of gravity of the crcua,ference of the
tube coincides with the direction of the vertical. In a regular turn this line co-
incides with the direction of the apparent vortical, i.e., in a regular turn the
ball will occupy the central position in the tube (Fig.120 a).
During a turn, the forces of gravity 0 . mg and the centrifugal force F ?mVia
are applied to the ball (Fig.120 b). The displacement of the ball from the middle
of the tube at an angle of sideslip'' 1, is equal to S Y1 i.e., the value of
360
f
he tube is filled with liquid to damp the oscillations of the ball that take place
!,t)?
_lunder the effect of accelerations. The damping of the ball is stronger the smaller
--tthe Clearance between it and the inner wail of the tube (which is taken as about
5-mm)1 and the higher the viscosity of the liquid. Toluene is often used as a
Fig.121 - Diagram of Turn Indicator
- Body; 2 Pin; 3 - Port for admission of air; 4 - Rotor;
Frame1-6 - Spring; 7 - Damper; 8 - Pointer; 9. .Bankrindicator;
10 - Adjusting screw.
the displacement is proportional to the angle of slip and to the radius of the tube.
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id,r'11ping liquid.
Figure 121 given 4 schematic diagram of the turn indicator in coioination with
the bank indicator. The sensitive element of the instrument is a gyroscope with two
degrees of freedom (rotation about the
and zz axes shown in Fig.112 a).
When the gyroscope is inclined about the longitudinal and lateral axes of the
aircraft? no precessional motions occur. The rotation of the aircraft with respect
to the yy axis causes a gyroscopic moment to appear and the frame of the gyroscope
to rotate about the xx axis in accordance with the equation 1%, = w cos a. The
moment Y7.1, is balanced by the moment developed by the steel spring (6) connecting
the gyroscope frame with the body of the instrument. As a result of the action of
these two moments, the instrument pointer (8)j deflected.
The moment developed by the spring is equal to Y51) Z: c5, where c is the moment
developed by the spring on a 10 rotation of the frame, i.e., the coefficient of
rigidity of the spring; and 6 is the angle of rotation of the gyroscope frame about
the xx axis. The gyroscopic moment is
Mr= ?12 w sin a= ho sin it,
where J is the principal moment of inertia of the gyroscope;
Q is the angular velocity of rotation of the rotor;
w is the angular velocity of rotation of the aircraft;
a is the angle between the axis of rotation of the aircraft and the rotor
axis.
n the equilibrium position,
J9 ? sin Ile sin a?(3
??!::
or
F
?
?1 but
/slit coa(4
cos ma/ +1,104sinye.
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In practice the angle 6 of inclination of the frame does not exceed 150; we may
therefore write
cos(?)* cos I 4- sin Tvw
Since, in this case,
we may write
and
coct it a sin 1 est
.1441(corg +a i)
asi
t .1?10 ibis ;
(v.15)
The angle 6 of rotation of the gyroscope frame about the xx axis depends not
only on the angular velocity of rotation of the aircraft, but also on the kinetic
moment H of the gyroscope, the angle of bank y and the elastic properties of the
.spring c. Since some of these values, for example the angle y, may vary during
flight and these variations are not taken into account by the instrument, the turn
indicator shows only the direction of the turn but not the angular velocity of the ?
iaircraft.
In order to make the notion of the pointer smooth without sharp fluctuations,
A the frame of the gyroscope (5) is connected by a special tie rod with the damper (7),
consisting of a cylinder rigidly connected to the body of the instrument, and a
, Iplungerconnected with the gyroscope frame (5),
! The wall of the cylinder has a capillary opening connecting the inner cavity of
the cylinder with the atmosphere and covered by the adjusting screw (10). On sharp
rotations of the framel-i,e., .on rapid displacement of the plunger of the damper,
,
'the,..air,cannet.pass through the capillary, and a deceleration force, counteracting
ithe_oscillations. of the gyroscope frame and, consequently, also those of the pointer
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f the instrument, is produced. The deg of dampingle rgulated by varying the
area of the croon
tion of the capillary openine
ns of the adjusting
crew (10)? whose position may be changed by turning it head, which la located on
ho outside of the instrument body. This method is not the only way of p od ci
damping force.
The gyroecope rotor may be actuated both by a pneumatic or an electric moth
Recently, instrument
g
h an o1ectrc gyromotor have cone into wide use, and coin-
pound electric instrumento are being ued snore and more, Whlch, in a .nle ease,
cobino a turn indicator, 'a gyro horizon, and a hank ndicator.
Figure 121 shows the diagram of the pneumatic turn indicator. The rotor rotates
in two radial ball bearings which are lubricated through .spectal opening in the
instilment body. The distance between the centers in the ball bearings is regu-
lated by means of special screws placed on the 1rtricnt body. The frame, together
with the rotor, Is carefully baaanced.
The pointer (8) is attached to one axis with a yoke, provided with a counter-
:weight for balancing the entire unit.
Since the instrument ,does not measure the rate or turn cf4.eq.(V.15)j, its
stale has only three divisions: the center one coy esponding to a zero rate of
turn, and two extreme divisions (without figures).
On the face aide of the instrument the bank indicator, consisting of a pendu-
lum, is mounted!
The position of the pointer of the turn indicator and the ball of the bank in-
cater, during various evelutions of the aircraft, are given in fig.122.
In rectilinear flight and during a regular turn, the ball of the bank indicator
occupies the central potation. The pilot decides from the position of the bank-
indicator ball whether the aircraft is sideslipping in flight.. from the position
of the pointer of the turn indicator, he judges the turns. The pointer occupies
the central position in rectilinear flight, ie. deflected to the right on a right
202
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turn and to the left on a left turn.
Principal Characteristics,of the Pneumatic Turn Indicator
Diameter of case. 80rrim
Length of case 115 mm
Moment of inertia of rotor 0.5 gm-om-8ec2
Rigidity of spring 160 gm,-cm/rad a 2.8 gm-cm/deg
Rate of rotation of rotor
(at mortal vacutm.equal to 50 mm Hg) 7000 - 8000 rPal
?o
Angle of stagnation -
Deflection of pointer at turning
rate of 60 per sec (360P per min) 26 - 300
Radius of curvature of bank indicator
tube
Diameter of ball
140 rn
Weight of of instrument 650 gm
Since the instrument has only one numbered division (at zero), the error of the
instrument, and in particular, the scale error, is tested at the zero division.
The scale error must not exceed ? 1?.
.The pointer oscillations at vibrations of the instrument at a frequency of
40 cycles and an amplitude of 0.15 mm must not exceed ?10.
Figure 123 gives the diagram of the electric turn indicator. The rotor of the
gyroscope is rotated by means of a DC electric motor with parallel excitation. The
rotor speed is 8000 - 10,000 rpm. The gyroscopic moment is balanced by a spiral
springs and oscillation of the pointer is prevented by an air damper of the plunger
type (as in the pneumatic turn indicator).
203
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-r
WI left wing
si4leei,Ap
on r.ight
F4022 - Folnitr tialing# 6f 1)1rn Indicator at Varlet10 roPitIonf! of Aircraft
ttUnr flirht
-1Aft turn
- tnrn
Fig.123 Diagram of
1etri Thrr inctioator
2 Oxroseope unit 3 Damper; 4 rointer; vat
2014
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Section h
GYRO-RECORDE8S
Gyroscopes with two degrees of freedom are used not only In turn indicators but
also in gyro-recorders, i.e.
Figure 124
in Instr.
(Nits for recording angular velocities.
ives a schematic diagram of the gyroscopic part of the instrument.
The instrument records the value and direction of the angular velocity of the
aircraft with respect to its three axes.
Each of the gyroscopes registers the
value and direction of one of the compo-
nents of this velocity. The gyroscopes
are therefore arranged, respectively, along
the three aircraft axes.
As a result of the action of an ex-
ternal mo=ent, proportional to one of the
components of the angular velocity of the
aircraft, the frame of the corresponding
gyroscope is rotated (just as in the turn
indicator).
For recording various angular veloc-ities, the set of the gyro-recorder in-
cludes interchangeable frames, springs, and dampers.
The gyroscope rotors consist of DC motors (AC cromotors may also be used). To
ensure constancy of the angular velocity of the flywheel rotor, a centrifugal gov-
Fig.124 - Diagram of Gyro-Recorder
ernor is used. The record is made in most cases on a paper chart by perforations
- or with special inks. The paper is moved by an electric motor. The time marker is
--!most often provided in the for m of an electromagnet which receives an impulse from
an ,electric clock.
,
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Section 5
GM WHIZ=
.tr
The gyro horizon is designed for determining the position of the aircraft with
respect to the plane of the horizon. The instrument permits holding the aircraft in
' horizontal flight even when the natural horizon is invisible, or indicates the de-
gree of bank and pitch on turns and changes of altitude.
The artificial horizon is a gyroscopic instrument with a pneumatic or electric
drive for the rotor. Depending on the drive used in the rotor, a gyro horizon is
called pneumatic or electric. The principal part of the instrument is a gyroscope
with three degrees of freedom, which, as already stated, is indifferent to the po-
sition in which it is installed and may be used for determining an assigned direc-
tion. As is commonly known, a prerequisite for using an astatic gyroscope as a de-
vice for maintaining a given position constant, is to give it selectivity.
One of the most widely used methods of giving a gyroscopic instrument the neces-
sary selectivity is the introduction of a radial correction*.
The action of the radial corrector is based on the following principle:
1. When the axis of a gyroscope deviates from the desired direction, a correc-
tion moment is applied to the gyroscope in such a way as to cause it to move in the
same plane as the deviation in question but in the opposite direction. mother
words, the deviation of the gyroscopic unit is eliminated, and the gyroscope is re-
turned to its initial position along the shortest radial direction. This is accom-
-plished by rotating the plane of action by the correcting moment through 00 with
respect to the plane in which the deviation of the gyroscope occurs.
* The tem r4radial correction" was introduced by B.V.Bulgakov? who performed the
itundameetal..inveatizaktions on &Tom:Tic deviclgi eTaPPed wiA.41 4.401,_1701,91-1.4.71
! vics40._
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.The fact of the deviation and its value is established by compari
.. ? . . ? ? ' '
on with a
certain 'element possessing selectivity, with respect to the position selected.. This
;-element is termed the sensitive element or the correction system. For example, when:
selectivity is is
4,14
4T* the vertical rct1on
UNA 1 I. 4.4.L i434:4 pendultaut used;
? -4,1ve selectivity in a horizontal plane, magnets are often employed, etc.
'In addition to the radial corrector, Other systems of correction are also used
which the return ? of the gyroscope to the selected position may also take place
-
over a distance that is not necessarily the shortest one.
The question is very often asked as to how the necessity of using a gyroscope
is to be explained. Would it not be better to use only a sensitive element?
The point is that a gyroscope with a radial or other for:: of correction is a
system that possesses not only selectivity but also stability in the selected direc-
tion (the precessional motion of the gyroscope takes place at a very low speed,
?;while a sensitive element, for example, a pendulum,
-tion of disturbing forces).
The gyroscope used in the gyro horizon is Liven
reacts very
rapidly to the ac-
selectivity with respect to the
'7vertical by use of some system of correction, making this gyroscope a gyro vertical
Figure 125 gives a schematic diagram of the pneumatic gyro horizon. In any po-
Sitions of the aircraft, the principal axis of the instrument I - I (the axis of
.-)natural rotation) maintains its vertical direction, so that the angle between the
-plane of rotation of the rotor and the longitudinal axis of the aircraft, xx is
-equal to the absolute pitch 9, while the angle between this plane and the lateral
- wooris of the aircraft, zz, is equal to the absolute bank y (Fig.126).
correction with pendulum slide valves is used in the pneumatic gyro hori-
? The instrument can operate either on vacuum or on pressure.
The air driving the gyroscope rotor passes through the bearings and channels in
-ithe outer frame (8) (cf.Fig.125), through the bearings and ducts in the inner
z
trame (2), which consists of a closed shell, and in two jets, issuing from the
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-
grooves of the rotor (I) nettii
rotation.
The inner frame (2) in the lower part ends in the hollow cylinder with four win-'
dome (openings) (5) for the discharge of *ir, which are covered by the pendulum
shutters (4). When the shell is in the vertical position,? corresponding to the
vertical position of the rotor axis each shutter covers half of the opening of the
Fig.125 - Diagram of Pneumatic Gyro Horizon
1 - Rotor; 2 - Inner frame (shell);. 3 - Correction chamber; 4 - Pendulum
slide gate; 5 - Opening of correction chamber; 6 - Silhouette of aircraft;
7 - Gear transmission; 8 - Outer frame
vindow, and the air is ejected in four equal jets through these openings. The re-
-j active pressures of the jets, in this ease are mutually balanced, no external mo-
(1--1 seas are applied to the gyroscope, and its axis maintains its initial position
(Fig.127,a)*
-The-slide valves-are arranged-in pairs on the spindles, one of which is parallel
to-the-xx sods, the other to the as axis.
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4?"
under the influence of any external iments, the gyroscope begins to process
and deviate from the vertical in any direction, for le with respect to the
sz axis, while the slide valves remain in the vertical position then the pair of
, opening, or the correction at chamber (Wwill now no longer be closed to the mole
Fig.126 -Measurement of Angle of Bank and Angle of Pitch by Means of the
Gyro Horizon
extent, i.e., one window will be open wider, while the opposite window will be
closed more. The equilibrium of the reactive forces perpendicular to the rotor axis
will now be disturbed.
Fig.127 - Pendulum Pneumatic Correction for Gyro Horizon
- Axis of rotor is vertical; b is of rotor deviates from the vertical
The resultant reactive force will provide a correction moment which will act
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perpendicular to the direction of the gyroscope deviation.
Precession will now take place and return the axis of the gyroscope to the ver-
tical position (Fig.127 b).
The sensitive element of the correction is a pendulum installed on the aircraft,
, which, under the action of accelerations, produces oscillations, causing the pre-
cession of the gyroscope. However,since the rate of precession is very small, the
axis of the gyroscope will hardly deviate from the vertical. All gyro verticals
operate on this principle, including gyro horizons. The most widely used designs
for gyro horizons are the diving gyro horizon AGP and the electric gyro horizon of
the ACK-4713 type.
filltigt:LIDASEJW.21.191.1.29.11
The sensitive element of a gyro vertical is an astatic gyroscope whose rotor is
actuated by their jets (cf,Fig.125). The axis of rotation of the rotor, which is
made integral with the rotor, rests on two ball bearings installed in the shell (2)
of the rotor. The shell, together with the rotor, is able to rotate with respect to
the outer frame (8) which, together with the shell, is able to rotate freely in the
?body of the instrument at any angle. The shaft of the rotor is vertical. The axes
?of the outer and inner gimbals lie in the horizontal plane. In the AGP the axis of
rotation of the outer frame coincides with the lateral axis of the aircraft (in some
,designs of the gyro vertical, the axis of the inner frame coincides instead with the
-longitudinal axis of the aircraft).
-1
('C-Ispecial orifices and from there enters the body of the instrument through the ori-
!flees (5)P which are covered by pendulum shutters.
The diving gyro horizon can operate on vacuum or on pressure.
From the upper part of the shell, the air enters the correction chamber through
The body and frames are cast of silusdn.
2---The--rotor-is-a- solid 'brass ring with a patent of inertia of the order of about -
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1..03 gm-cm-sec and the rotor speed is 14,000 - 15,000 rpm.
The rotor rests on radial ball bearings which are lubricated by special inserts
saturated with oil. The same bearings are installed in the frames. The clearance
in the rotor spindle is selected by means of calibrated inserts, placed under the
upper spring bearing.
A spring bearing is used to avoid possible compression of the aluminum frame by
? the steel axis of the rotor when the outside temperature rises, since the coeffi-
? cient of temperature expansion of the frame is considerably greater than that of the
rotor spindle. The lower bearings forms a tight fit.
The clearance in the suspension bearings are maintained by tightening the cen-
ters. To reduce the escape of air through the bearings, special air gaskets are in-
stalled in the form of condensing rings or German silver. The front face of the in-
Fig.128 Beading the Angle of Bank from the Position of the Aircraft Silhouette
Left: Horizon; Right: Instrumental horizontal line
strument case carries a screen with a luminous line representing the line-of the
&0-J1y3rizon. Through a vertical slit in the screen passes the spindle of a movable in-
!tidex (the aircraft silhouette) placed between the screen and the glass of the instru-
serving_as indicator of the position of the horizon.
56- aircraft_ ballig5 at the angle y, the index, pin of the horizon line like-
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tilts at the angle y. If the-aircraft silhouette is attached directly to the spin-
..,Hdle of .the gyroscope unit, then, when the aircraft banks at the angle yo the air-
craft silhouette, together with the gyro unit, will remain fixed in space and indi-
cate the bank at an angle y on the opposite side with respect to the horizon.
To ensure correct reading, of the bank of the aircraft, the silhouette of the
aircraft is connected with the inner frame over a pair of gear wheels with a trans-,
mission ratio of 1:1 (in some instruments,a different design of transmission mech-
anism is used). When the aircraft banks at an angle the aircraft silhouette on
the scale of the gyro horizon rotates through an artgle y in the same direction
? (Fig.128), while its position with respect to the index of the horizon line will
correspond to the true angle of bank of the aircraft.
The right side of the instrument carries a scale attached to the outer frame,
which is in the form of a ring with graduations within the 15Ttlits of ..1t9e; on this
scale the angle of pitch is read off. Since, at high angles of pitch, the aircraft
silhouette goes outside the face of the instrument the scale is painted in two
'
colors: blue for climb and brown for dive.
The diving gyro horizon allows the execution of aerobatics. To prevent the
, spindles of the frame and rotor from coinciding at banks over 75 - 800, stops are
placed above and below on the rotor shell.
. The Errors of the GYM Horizon
1) Errors due to Aircraft Accelerations. The accelerations produced in recti-
linear flight, as. well as those during turns, cause a deviation of the shutter
415-- elides of the correction and a precession of the gyroscope axis, which deflects the
4C- rotor axis from the vertical. Because of the low rate of precession, eqyal on the
&0-1 average to 6?/Min, the gyroscope is unable to deviate significantly from the verti-
cal if the action of the force of inertia deflecting the shutter is brief. In
Y,---reaseek-where-the-aotion of the inertia fore* is prolonged, the dofloctions of the
_1
scop4, &xis ray reach considerable values.
212
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'
The magnitude of the deviation of the gyro rotor axis depends not only on the
time of flight under acceleration but also on the nature of the acceleration. For
,example, at constant longitudinal acceleration, when one of the windows of the cor-
,
:reetion ohaMber is completely open, and WhAn the increment of the flying speed is
A V 360 km/hr, the error in the position of the rotor axis may reach about 8? If,
_l however, the value or the acceleration at which the shutters are .completely opened
As variable, then this error may be greater than 100.
If the flying speed is increasing, the direction of the acceleration will coin-
cide with the direction of the velocity; in this case, the shutters will deviate in
.the opposite direction (backward). The rotor will deviate forward, away from the
in such a way that the instrument indicates a climb of the aircraft although
A.t is actually flying horizontally. On reduction of the flying speed the gyro hori-
zon will show a descent.
Considerable errors in the reading of the instruments also arise in the ease of
a turn. After prolonged horizontal acceleration during a turn, the lateral pair of
poemdulum slides assume a position on the apparent vertical. As a result, the pre-
- cession of the gyroscope takes place in the direction of the apparent vertical.
Since the lateral plane of the aircraft is rotated about the vertical during a
turn, at a velocity equal to the angular velocity of the turn, the direction of pre-
cession of the gyroscope in space will also vary continuously, causing the position
--of the gyroscope, with respect to the vertical, to vary also. As an ultimate re-
--413U, the axis of the rotor will deviate from the vertical forward in flight, and
'inside of the bank. The angle of deflection of the rotor axis from the vertical
f
during a turn may reach 4 - 50 or more. To reduce this error, the gyro unit of the
gyro horizon is arranged in such a way that, at a vertical position of the slides
and zero position of the pointer, the upper end of the rotor axis will be deflected
from the vertical forward in the direction of flight by 2 - 3?.
During a turn, this longitudinal inclination of the rotor axis changes to a lat-
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inclination
'will completely compensate the precession of the gyroscope while for all other val.,
of the angular velocity, the compensation will be only partial. The angular ve-
locity of turning wc of the aircraft, at Which the turning error is completely com-
pensated, is determined by the relation
where wc is the angular velocity of turning;
too the angular velocity of precession;
00 the angle of inclination of the rotor axis.
2) The errors caused by the flying speed of the aircraft, which appear during
rectilinear flight as a result of the curvature of the earth's surface (while the
'aircraft rotates about its lateral axis zz), as well as the errors resulting from
the diurnal rotation of the earth, have a negligible value, not exceeding 2 - 4",
-'and are usually disregarded.
3) The instrument errors of the gyro horizon due to friction in the suspension
:bearings, friction in the shutter pins, vibrations, and unbalance, are reduced to a
-Atli= by selecting a suitable design and by efficient work in the manufacture and
For example, to reduce the moments of friction in the bearings of the frames
J
-lacting in 4, direction opposite to that of the relative rotation, bearings of the
4-2-
Ivibrating? rocking, rotating, or elastic type may be used.
To reduce the influence of vibration on the gyroscope it is necessary that for ,
each frame, the shaft of one bearing and he bearing of another act as supports;
this causes the moments of friction in the bearings under vibration to have oppo-
site directions so that not the ma but the difference of the moments of friction
in the bearings will, act on the gyroscope.
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?
? , , ?
?
: ?
he ACIE-47B itleetric Combinatior?ftto Ipriz
,The altitude-dependence of the rotor. 'speed is one shortco nmig of pneumatic gyro
Alorizons? that is, the dependence of rotor speed on the variation in atmospheric
,
pressure. Electric instruments fed from the aircraft electrical system do not have ,
-1 this drawback.
The ACK -47B gyro horizon consists of a combination of three instruments mounted :
in a single ease:
1) a gyro horizon in the form of a gyroscope with three degrees of freedom, in-
? ?dicating the position of the aircraft with respect to the plane of the horizon;
2)?urn indicator in the form of a gyroscope with two degrees of freedom,
dicating the rotation of the aircraft about its vertical axis;
3) a bank indicator, consisting of a pendulum indicating the sideslip of th
air raft in reeitlinear flight and during turns.
Figure 129 shows the readings of the gyro 'horizon during various evolutions
- the aircraft.
Figure 130 gimes the kinematic diagram of the gyro horizon.
The gyro unit: consists of a rotor installed in gimbals.
The axis of rotation of the rotor is inclined forward in the direction of flight
--Iky an angle of 2?. This reduces the errors of the rro ye
The axis of rotation of the outer frame of the gldbals
t
42
--eral axis of the aircraft, while the axis of the inner frame is parallel to the
,
t .
,
71onsitudinal axis of the aircraft.
,
-i Pitch and, bank scales are attached to the face side of the outer frame; the air,?
1 ,
--Icraft silhouette attached to the gimbals serves as the horizon indicator; the index:
?
--of the horizon line is attached to the instrument case. ,
!ALA
-4 Figure 131 shows the design of the ACK-47B gyro horizon. The unit (1), and the
fa- '
_ -*mil' of the gyromotor (2) form the gimbal*. The gyromotor is a three-phase electric
4, .
, _
' motor with a speed of 20,000 rpm. The stator is attached to the top of the instru-
1,6 t ?
t
STAT
4"
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The stitor block is made of dynameAron? carrying, in twelve slots, the three+
two-pole winding in star connection 4 The rotor block, which is likewise made
of dynamo iron, has a short-circuit squirrel-cage winding, east of aluminum alloy.
,The rotor turns in radial bearings with brass separators. The bearings are lubri-
cated by a felt wick soaked with VW oil.
Radial bearings are also used in the gimbals. The gyromotor and correction de-
vice are supplied with current across the gimbals, using contact rings and brushes.
.The block of the contact rings consists of three polished silver rings with insu-
lated bushings
used instead of a wrist-pin bearing. The contact brushes consist of
flat rings, to the ends of which platinum-
iridium alloy wires are soldered, which
cake contact with the silver contact rings.
A brush pressure of 1 - 1.5 gm ensures re-
liable contact.
The position of the rotor axis of the
gyro horizon is determined and maintained
by the aid of a correcting device consist-
ing of a liquid switch and two solenoids
placed in the gyromotor body (Fig.132).
Each solenoid has two windings lo-
cated to the right and left of its geomet7
ric center. An armature (core) is placed,
inside the winding and can be displaced
yJ
1
WOW WOW S -
Fig.130 - Kinematic Diagram
of the Gyro Horizon
4? 1 - Axis of outer frame of gimbals;
42J 2 - Axis of inner frame of gimbals;
of the armature, when it is displaced witti respect to the center of the solenoid.
The magnitude of the moment depends on the magnitude of the displacement of the ar,-
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The correction switch (Fig.133) consists of a copper vessel filled with an elec-
trically conducting liquid. Into the upper part of the vessel, which is made of an
,insulating material, four red copper contacts are stamped. The assembled and ad-
justed switch is a self-contained unit which needs no further adjustment during oper-
ation or repair of the instrument. It is categorically forbidden to test the switch
with direct current* since this will put the switch out of order.
The static balancing of the gyro unit is performed by centering the cores in
the solenoids in such a manner that produces no moments with respect to the axes of
? the gyroscope gimbals. Each of the soleneids has two windings. The windings of the
solenoids and the switch are connected with the switch of the correcting device ac-
cording to Fig.134.
The correcting device operates in the following way:
Any deviation of the rotor axis from its original vertical position, relative
to the lateral axis of the aircraft, results in an inclination of the switch with
respect to the horizontal plane, causing a change in the wetted area of the contact
surfaces along the longitudinal axis of the instrument. The forward contact (with
respect to direction of flight) is immersed deeper into the liquid than the rear
contact. These contacts are connected in the circuit of the windings of the sole-
poid4 placed parallel to the lateral axis of the instrument. when the equality of
1
_Ahe resistances of the circuits of the solenoid windings is disturbed, a redistri-
ii
....,?Ibution of the currents flowing in the solenoid windings occurs, and the armature is
42_4
_Idisplaced toward the winding through which the greater current flows. In its new
__Ipositiont the solenoid armature produces a moment acting on the gyroscope and caus-
46_,_
?4ing its rotor axis to return to its original position.
.4&1
4
A rotation of the rotor axis about the longitudinal axis of the instrument
causes a change in the liquid coverage of the conductors of the switch contacts a-
Oco"'"' 5
--liong the lateral axis,
J?f?.
In the original calibration of the gyro horizon at the factory, the caging mech.-
4
218
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cu
(D
0
0
-0
-0
(D
(D
(T-
CU
CD
0
0
? ?
0
33
0
co
, 4:7N ?
ee4
, ? t*.:1.1
? , ?
L I L 1, .4
rilipoietioitoiwiAl
ali is ippri i ...de
;:.
41 4017
1 r 4
. / i
i ,
/ i
IIL P ?
As
a 1r ?
A
'S 417
, p6a. Ana ? . ? owc 1
)
?1111WAst."*.s ?."710,67 ..*1".t:::".
.4. ""1"411404-0 A
....."
' iirmazzaw '?ar,:=NmssrakoriEdimaalsoits=can..,01101:200.
7 ' ellaillh-- ,,,,,r ....1Pdmo-UllAt4,61100:40.-:' ? ..=',:, ...... ? '411"...;.0... ,:. 0, ... roily** i 1
mee.
Fig.131 - Design of the AGK-47B.
1 - Gimbal unit of gyro horizon; 2 - Gyromotor
of gyro horizon; 3 - Switch of correction de-
vice" 4 - Solenoids; 5 - Leads for gyromotor; '
6 - Turn indicator; 7 - Gyromotor of turn indi-
cator; 8 - Bank indicator; 9 - Caging mach-7
anism; 10 - Case; 11 - Rear coveri
-
? .
''"'
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Fig.132 - Circuit Diagram of AC-47B
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used for rapid elimination of errors in the readings after deviation
,from the vertical, and also for preventing damage in shipment (cf. Fig.131). This
'device consists of three cams, push rods,
qoalmas a stop, working and return springs,
And a signal blinker.
When the caging knob is moved, the
force developed by the working spring, is
transmitted through the push rod and cam
?
Fig.134 - Wiring Diagram of the Correcting Device
- Central position; b - Rotation of the gyro-horizon axis about the lateral
axis of the aircraft; c - Rotation of the gyro-horizon axis about the vertical
axis of the aircraft.
A Boundary of the liquid meniscus
Solemoid of longitudinal correction; 2 ?- Solenoid of lateral correction;
- Switch; 4 - Switch contact; 5 - Iimulating bushing; 6 - Longitudinal axis
of instrument; 7 - Late.' axis of instrument
'.---gleated on the spring axle ofthe gimbals, to the gyroscope unit.
- The-profile of the CSSI is cut in a logarithmic spiral, thus causing the force
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?
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-
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, ? ?. ? . ?
, ? , ?
1?f
acting on the east to create a moment about the outer axis of the gimbals. Under the
,action of this moment, the
unit begins to precess until th axis of rotation of
:the gyroscope coincides with the outer axis of the gimbals. In this position, the
?gyroscope loses one degree of freedom, its! gyroscopic moment disappears, and the
frame, under the pressure of the plunger, rotates freely about the outer a..Itis of the
'gimbals until the plunger engages the cam slot. As soon as the plunger enters the
cam slot and fixes the gyro unit with re-
spect to the outer axis of the instrument,
the arrest of the gyro unit relative to
the inner axis of the gimbals begins.
The turn indicator, which is designed
tached to the rear flange of the instru-
ment ease. The axis of rotation of the
rotor is parallel to the longitudinal axis
Fig.135 - The ACK 478 Turn Indicator
1 - Gyromotor; 2 - Spring balancing
the gyroscopic moment; 3.- Damper;
- Pointer; 5 - Frame; 6 - Leads;
- Bearings; 8 - Adjusting lever;
9,- Damper lever
scope, which arises when the aircraft
turns, is balanced by the spring (2). To
extinguish the oscillations of the system
the air damper (3) is installed. The de-
Which covers the orifice by which the in-
ner cavity of the damper communicates with
he atmosphere (as in the ordinary turn indicator).
30 ? - . ?
? , The instrument is fed by a PAG-IF type converter.
-instruments (for instance two gyro horizons or one gyro horizon and a DGMK -2 remote-
,
40'
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? The gyro horizon causes considerable distortion in the character of the magnetic
field in, the space surrounding it; therefore magnetic compasses must not be installed
closer than 200 mm to the AGE-47B gyro horizon.
Total reading error of the AGE-47B gYie horizon:
in rectilinear flight,. not more than 10
after coming out of a turn with a 209 bank
at a flying speed of 400 km/hr not more than 29
The principal characteristics do not depend on the flying height.
Electric gyro horizons have come into wider use. There is a great variety in
the designs of these instruments.
In principle their operation does not differ from that of the ACE-47B. Only
dee4gn f the gyro unit and the method of correction differ.
The primary advantage of the air radial correction is its simplicity. This ex,
plains the attempts to introduce it in electric gyro instruments.
A substantial shortcoming of the air radial correction is the fact that its ef-
_,ficiency depends on the flying height. When the rotor is fed by air this drawback
is not too noticeable, since the rotor speed is reduced simultaneously with a reduc-
tion in effectiveness of the correction. When the instrument is electrically polo-
.,ered the speed of the rotor does not depend on the flying height and the lowering of.
the effectiveness of the correction makes it too sluggish at high elevations.
Consequently, the transition to electric rotors whose speed can be considerably
Adgher than that of pneumatic rotors and is practically independent of the flying
requires the development of electrical methods of correction.
Induction correction may serve as an example of electric correction.
Fipre 156 a gives the induction diagram of the electric correction of one of
--be frames of a gyro vertical,. The signal pickup consists of the contact roller (1),
frame being corrected (2) and the contact plate (3)0 connected
element of the correction and divided into two parts by an inS11--
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in the coinciding position?,the.contact roller is in
the center of the insulating gap. When it is in disagreement the roller is dis-
placed from the center.
system of two induction coils (4) ad (5) is used as pickup for the correc-
in these coils engage the conductor disk (6), connected with the
axis of rotation of the corresponding
frame. One of the coils constantly carries
the voltage of one of the phases of a
three-phase alternating-current line
6) ?
Fig.136 - Diagram of Electric Inductive
Correction
- Schematic diagram; b - Circuit
diagram; 1,11 - Axes of gimbals;
1 - Contact roller; 2 - Frame to be
velP
Fig.137 - Diagram of Electric Correc-
tion with Mercury Switches
I and Il - Axes of gimbals;
1 - Ampoul ? 2 - Mercury; 3 and 4 -
114 virlA ng
of induction motor; 5
Shot-
corrected; 3 Contact plate; 4, 5 - circuited rotor of induction motor.
Induction coils; 6 - Conductor disk
(cf. Fig,136 b) The second coil is connected across the signal pickup to one of
he other two phases of the line, depending on which half of the contact plate makes
--contat with the contact roller. In this may, as soon as the contact roller touches
the insulating gap i.e., when the signal pickup is in the coinciding state current
.6?1
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4
low through only one of the coils, and there will be no moment on the conduc-
Uhen the roller Passes to one side of the Plate, current will flow through
both-of the
There is a 1200 phase shift between the alternating magnetic
fluxes produced by the coils. On the con-
ductor disk, as a result of the interaction
of the eddy currents and the magnetic
fluxes, which are phase-shifted with re-
spect to each other, a moment is produced,
constant in magnitude and having a sign
depending on the side of the plate on
which the roller is located. The same
type of klorrection can be obtained by mer-
cury switches (Fig.137). The signal pick-
ups in this eases is an ampoule with a
droplet of mercury having three contacts
and so located that the plane of its rock-
ing coincides with the plane of rocking of
the frame to be corrected. When the posi-
tion of the gyroscope axis and the pendu-
lum coincide, the mercury droplet is in
? the center, while when they are in dis-
agreement it is in one of the extreme po-
sitions. In the latter case, the mercury
droplet connects two contacts, thus actu-
ating one of the extreme contacts. The
moment pickup is an induction motor with
a short-circuited rotor mounted to the
-Zone of insensi-
&xis of rotation of the corresponding
Fig.138 -Characteristics of Correa-
tion Device
- Correcting moment; 0 - Angle of
discrepancy in signal pickup; 0 - Sec-
tor of proportional part; l - Moment
of correction of constant part of
characteristic; V - Steepness of
225
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diagonally opposite) frame to be corrected. One of the phases of the motor is
always supplied by voltage from one of the phases of the three-phase feed line
When two contacts are closed by the mercury, the second phase of the motor is
turned on, causing a rotary field there. The direction of this field depends on
which of th, extreme contacts is closed by the mercury switch. In this way, a moment
of either sign and of constant magnitude, is developed on the rotor.
Gyro-verticals include widely varying devices for performing the correction;
these may be divided into three groups according to their characteristics;
1. Proportional correction (Fig.13$ a).
2. Constant correction (Fig.138 b).
3. Mixed correction (Fig.138 c).
Section 6
WIESE GTROSCOPES
The disadvantages of magnetic compasses and the impossibility of installing
gyroscopic compasses on an aircraft, owing to their large size, has led to the cre-
ation of so-called directional gyro turn indicators (DGT)? gyremagnetic compass-
es (GMK) and, finally, to the creation of remote-reading gyromagnetic and induction
?-compasses.
THE DIRECTIOVAL GIRO TURN INDICATOR
As with gyro horizons, there are both pneumatic and electric directional gyro
urn indicators in existence. The pneumatic type consists of an astatic gyroscope
with three degrees offreedom, mounted in gimbals. To maintain its perpendicular
position, the special flanges (2) are provided between the principal axis of the
roscope and the azte of the outer frame of the rotor (1) of the instrument. The
iair jet from the rotor, striking against these flanges, produces a correcting mo-
'Since the &static free gyroscope has no directional force, the axis of the gyro-
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;
scope, in flight', begins to deviate more and more from the initial direction mainly
under the influence of friction in the suspension bearings. The instrument readings
will be correct for 10-15 minutes. A correction of the instrument is then neoessarb
Fig.139 - Rotor of a Pneumatic Directional Gyro Turn Indicator
1 - Rotor; 2 - Flanges; 3 - Nozzles; 4 - Tubes conducting air to the noz-
zles 3; 0 = Angular velocity of rotor; F - Air jet spinning the rotor;
p - Pressure. of jet on flange mused by the moment lip and the precession
tzld is made manually by the pilot.
The tern gyro semicompass is also used for such an instrument and is explained
by the fact that this instrument cannot completely replace the compass; the gyro
sesicompass provides only a possibility of holding the course with an accuracy with-
in 2 - 3? for a limite'd time interval (10.- 15 sdn).
Figure 140 gives the diagram of a pneumatic gyro semicompass. The rotor is
----driven by an air jet, issuing from two nozzles, one on each side of the rotor. The
:moment of inertia of the rotor is about 0.0 gs-cm-stic2 and its speed is 10,000 to
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rpm. The rotor rests on ball bearings mounted in the inner frame. The outer,
and inner frames likewise rest on ball bearings. The scale of the instrument (2) i
uter frame (3) of the suspension. The course line from which the
readinks are taken is placed on the glass mounted in the instrument case.
TRI-4
13
\14
Fig.140 - The Gyro Unit of the Directional Gyro Turn Indicator
1 Rotor; 2 - Gard; 3 - Outer frame; 4 - Inner frame; 5 - Bearing; 6 - Small
cog wheel; 7 - Caging knob; 8 - Frame cog wheel; 9 - Caging lever; 10 - Noz-
zle; 11 Caging fork; 12 - Caging bushing; 13 - Balancing stud and nut;
14 - Spring plunger
The stop lever (9) is attached to the outer frame and is actuated by the easing
knob (7). Mb= this knob is pushed "inn, the yoke (11) engages the conical cut-out
the face of the cog wheel Oh in this case the yoke is set in the horizontal di-
rection and lifts the bushing (12), turning the lever (9) upward. The lever (9)
comes to rest on the inner frame and does not allow it to be displaced. At the same
ime the cogwheel (6), engages the cog wheel (8), so that turning the knob (7) will
,
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!
rotate the entire gyroscopic unit is rotated by the required angle.
0 precessional
motion? occurs when this is done, since the inner frame is fixed by the
,that is, the gyroscope is_deprived of one degree of freedom. When the
uouttl, the gyro unit is released .art_d the instrument begins to operate.
In most oases, after setting the aircraft on the required course, the pilot sets
the gyro card mounted to the inner frame, to zero division, and then only observes
the deviations. This relieves him of the necessity for remembering the flight-path
angle
stop levers
knob is pulled
Errors of Directima.14=Turn Ingcators
Error due to Diurnal Rotation of the Earth. The principal axis of the free
gyroscope maintains a constant direction in space; in this ease, owing to the rota-
tion of the earth, an apparent motion of the gyroscope about the xx and zz axes oc-
-curs (Fig.141). The angular velocity of the apparent motion depends on the geo-
graphic latitude and longitude of the
place and on the angle of inclination of
the principal axis to the horizon; the
.P1
Fig.141 - Apparent Motion of the Rotor
of the Semicompass with Respect to
the Earth
Fig. 142 - Compensation of the Earthts
Rotation by Displacement of the Center
of Gravity of the Rotor
angular velocity is equal in value to the vertical component of the velocity of ro-
tation of the earth. For example, for Leningrad, the velocity of the apparent pre-
cession of the gyroscope due to the earthts rotation is equal to lribour.
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The gyro unit of the semicompa'ss is balanced so as to eliminate, as far as
,sible the apparent motion of the free gYreseoPe about its axis This is aeeesi-
plished by displacing the center of gravity of the rotor by a certain value 1 with
'resPect to the center ,of the rotor (Fig.142) under the action of the weight P. If
,the magnitude of the moment M = Pi produced by this weight is so selected that the
'precession due to this moment is equal in magnitude but opposite in direction to the
apparent motion, there will be no errors in the instrument readings This motion
however, can be completely compensated only for the given latitude and at a constant
rotor speed, since the rate of precession, due to the fact that M P/, is deter-
mined by the relation
w is the rate of precession;
we the velocity of rotation of the earth.
2. Error due to Variation in Latitude. This error appears in flights at lati-
tudes different from that for which the error due to the earth's rotation had been
eliminated by balancing. For example, a semicompass balanced for Leningrad will de-
viate by 2.e/hour in flights to the Crimea.
3. Errors due? to Variation in Longitude and Latitude (without allowing for the
-?.rotation of the earth). In flights along a parallel, the position of the gyro axis
in space will vary by the value
--4.f the influence of the earth's rotation is neglected (Fig.143). For example, in a,
41'1144 along the 600 parallel of latitude, the readings of the semi compass will be
Il?561 too low for a distance of 100 km eaetward.
A Variation in latitude causes a.varistion in the vertical component of the.
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, ?
angular velocity of the earth's rotation. For this reason, if a semicompass is bal-
anced for a certain latitude w, there will be an error in its readings due to uzbal-
co at all other latitudes. This error is relatively small (of the order of about
2.4Vhour).
4. Error due to Frietion in the Fftrings. This errOl, 1TIAV 1?44A-eh 3 ? 50 in
15 mixt, which necessitates frequent corrections of the semicompass readings with
another instrument.
Fig.143 - Error Due to Variati4m in Longitude
Radius of earth; r - Radius of parallel; p ? Latitude of the place;
North Pole;-A, 8 - Two points at the same latitude; X1X2, Longitudes
of these points; A - Angle between the meridians of the points A and B;
AB r (X) - X2).
5 Errors due to Unbalance of the Suspension.
6. Cardanic Error. The axis of the outer frame to which the card is mounted
is attached to the instrument case, parallel to the vertical axis of the aircraft.
lam the aircraft is inclined, this axis and the gyro card are inclined together
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Sty..
,with the instrument, while the position co
40 to a geometric error in course read
Princi
the rotor axis remains constant. This
g? called the Cardanic error.
Characteristics of a DirectionalTurn Indicator
Moment of inertia of rotor 0.8 gm-cmr-sec2
Eoment of friction in bearings 0.3 gm-am;
Speed of rotor 12,000 rpm
Maximum deviation in 15 min 3 - 5*
Weight 1.5 kg
Allowable pitch and bank not over 45?
Like the gyro horizon, a gyro semicompass can operate either on pneumatic or
electric power.
Recently electric gyro semicompasses have came into wider use.
The gyro semicompass is used as the sensitive element of an automatic course
indicator, which is one of the principal units of an automatic pilot.
9n2RMESt1.9.--913,22
The readings of the gyro semicompass must be periodically corrected by the
readings of a magnetic compass, selecting for this a time when the magnetic compass
has the smallest errors. In the gyramagnetic compass this correction is performed
automatically and continuously and converts the instrument from a semicompass into a
COMpa$8.
Two types of correction by a magnetic compass are used.
1. Remote control when a remote magnetic compass is used for the correction.
2 Direct control, when a magnetic systez located on the gyroscope is used.
The former variant of the correction gives readings of higher accuracy, but the
correction device, in this case is bulkier. The latter variant gives the most com-
et sise, from the design point of view. The most successful design of a gyromag-
, .
lietic compass, constructed -according to the latter method, is the GME-2 gyromagnetic
232
, 444.i
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k'4
!?"
s proposed in 1932 by the soviet designers D.A.Braslavekiy, M.G.Eltkind and
t?Kachkacht ya,n.
etic C os
Figure 144 gives a schematic diagram of the GMK-2. The GMK-2 gyromagnetic com-
pass consists of a static gyroscope with three degrees of freedom, with its princi-
4P
? Fig.144 - Schematic Diagram of the Gyromagnetic Compass
1.- Rotor; 2 - Rotor shell; 3 - Frame; 4 - Card; 5 - Magnet; 6 - Eccentric
slide; 7 - Eozzle; 8 - Tubes; 9 - Pneumatic relay; 10 - Aneroid; 11 - Slide
of magnetic correction; 12 - Orifice of correction chamber; 13 - Slide of
pendulum correction; 14 - Rotor, driven by nozzle; 15 - Air duct for rotor;
16 - Shaft
pal axis in the horizontal plane. The rotor (1) is enclosed in the shell (2), which
isths inner ring of a Cardanic suspension. The shell is attached to the center
-
233
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A
777" ; ?
?-Xrame (3) whose axis of rotation coincides with the normal axis of the,aircr ft, yy.
The outer fraMe carries a card graduated from 0 to 3600
The shell (2) of the rotor carri s the moving magnetic system (5) with the
.ec-
centric slide (6) attached to its axle, and covering the nozzle (7) of the magnetic
.corrector. The receiving nozzles are connected with the pneumatic relay (9) which
As also installed on the rotor shell.,
The pneumatic relay consists of an aneroid (10) whose inner part is connected
H'with one of the nozzles; it transmits to its outer side the pressure from a second
nozzle. The center of the box is connected over the shaft (16) with the slides(11),.
.
'which cover the jet slide (12) and are located on the upper and lower parts of the
'chamber. In the neutral position of the aneroid, when. the pressure difference is.
equal to zero, the slide covers the upper and lower orifice.
The two magnetic slides (13)0 which half-cover two side windows when the rotor
---sbaft is horizontal, are also attached to the rotor shell.
The instrument is pneumatic, and its feed system is entirely analogous to the
feed system of the gyro horizon.
After the rotor is started, a considerable part of the air enters directly into
the correction Chamber and only an insignificant part of it passes through the gys-
-Aem of nozzles into the pneumatic relay. The magnetic correction operates in the
following y:
if the axis of the rotor is parallel, to the magnetic Meridian i.e., parallel
to the eads of the magnetic system, the eccentric slide (6) covers the nozzles, the
--membmane occuplesHa neutral position, the actions of the air jets entering from the
-*wand lower orifices balance each other, and no new external moments are applied
to the gyroscope.,
If the rote!' Shaft deflates from the morthseuth direction, the nozzles of the
poumatic relay will be covered differently, the pressure difference on the _aneroid
will deform it and cause -a displacement of the slide valve; in this case one of the:
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-orifices (1.2) of the nozzle will be covered more and the' other les, the air jets
will no longer balance each other'andra moment will be applied to the gyroscop
causing its precession in a direction opposite to the original deviation of the rotor
shaft. The magnetic correction, operating in the same way as in the gyro horizon,
maintains the horizontal direction of the rotor axis.
In this instrument, a new version of the pneumatic radial correction device is
:used in which the signal pickup is an air collector with two nozzles covered by ex-
centric slide valves, while the moment pickup is an air chamber with two pairs of
?:windows, one of which provides a correction in the vertical plane (pendulum) the
other a correction in azimuth (magnetic).
rrors of the GK-2 G' .mar etic Co ass
1. Deviations created by the influence of the gyromagnetic force of the air-
craft on the magnetic part of the correc-
tion device; the deviation is reduced by
the aid of a deviation instrument.
Asstsivor
Figure 145 shows its principle of operat
Am44w4s
tion and its appearance.
The deviation instrument consists of
four permanent magnets placed in the alumr-
,inum ease of the instrument in the shape
or a cross, with the south poles at the
ends. One pair of magnets is parallel to.
Fig 145 - Principle of Operation of
the longitudinal axis of the aircraft, the
the Deviation Instrumgnt
other to the lateral axis. The magnetic
field of these magnets is closed by a screen of soft iron placed under the magnets.
The screen may be displaced in two mutually perpendicular directions. Accord-
ing to the location of the screen, the value and direction of the stray field will
vary and will compensate the semicircular deviation of the instrument.
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tion in the magnetic and corredtion system; does not exceed 1?.
The banking deviation and rotational error are the same as in any magnetic
instrument. The Cardanic error is the same as in the gyro sericompas13.
4. Errors due to eddy currents; these arise in the rotor when it rotates in the
magnetic field and interacts with the magnets, causing their deviation. To reduce
the value of the currents, the rotor is milled perpendicular to its axis.
Principal Characteristics or the GyromagnitlisSman
Moment of inertia of rotor 0.64 gm-cm-sec2
Moment of friction in suspension 0.3 gm-cm
Stagnation
Setting to meridian accurate to 30
Rotor speed 12,000 rpr
Rate of precession 10 - 15 rpm
gm-cm
Reactive moment of correction 2.5
Weight about 2 kg
Period of rotation of rotor due to inertia 16 - 23 in
Deviation on inclination not more than 50 in 2 min
Restota_.eticaes
Remote-reading gyromagnetic compasses are being more widely used in recent
times (Bib1.12).
Figure 146 gives a schematic diagram of one of these remote-reading gyromagnet-
ic compasses.
The DOM remotereading gyromagnetic compass is a gyro semicompass with read-
ings corrected by a'type POI magnetic compass.
The full set of the instrument contains:
a. a gyro semicompass;
b, a type POI magnetic transmitter;
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an atplifier;
d. two indicators;
e. a junction box;
f. a knob for coincidence;
g. a transformer.
On the axis of the outer frame of the free gyroscope (1) is attached the type
PDK ring potentiometer (2) with two diametrically opposite points a direct-current
voltage of 2? v is applied. Three brushes 120? apart, slide on the potentiometer.
The position of the brushes with respect to the conductors determine the values of
the potential of each brush. Through the contact ring (9) and the current collect-
ing springs the brushes are connected simultaneously with the potentiometer of the
Gyro wait
- Schematic Diagram of the,Rete-Reading Gyromagnetic Compass DG4-3
1 - Outer frame of free gyroscope; 2 - Ring potentiometer; 3 - Magnetic pickup
111K-45; 4 - Indicators; 5 - Amplifier; 6 - Electric motor; 7 Brushes; 8 - Re-
ducer; 9 - Contact rings; 10 - kdjusting button.
etic pickup (3) and the two indicators (4).
If the angles between the brushes and the conductors of the gyro unit, and be--
tween the brushes and conductors of the magnetic pickup, respectively, differ by
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.-then the potential difference across the Conductors ,of the transmitter is equal
,However, if the difference between these angles is not equal to 900, a poten-
tial difference is produced across the conductors of the transmitter and is deliv-
ered to the input of the amplifier (5). The voltage tapped, from the secondary wind,
ing of the output transformer or the amplifier is fed to the control winding of the
..electric motor (6) of the gYro unit which rotates the potentiometer brushes of the
gyro unit in the direction in which the angle of discrepancy will be reduced.
Thus, the readings of the indicator depend on the position of the brushes with
respect to the potentiometer conductors of the magnetic pickup. In this case the
Indicators show the compass course with an error of not over -2?.
A discrepancy between the position of the brushes of the potentiometer of the
gyro unit and that of the magnet of the pickup may be due to either of two causes:
a. Precession of the gyroscope in azimuth.
Since the gyroscope is not corrected in azimuth it may preeess under the action
Of the forces of friction and of a certain unbalance of the inner frame, which is
-always present. This precession of the gyroscope is compensated by the correspond-
Sing movement of the brushes or the potentiometer of the gyro unit. Since the rate
--Of precession of the gyroscope, under no conditions, exceeds 1?/min, the precession
.of the gyroscope in azimuth cannot cause discrepancy in the system. In addition,
-this rate of deflection of the brushes also assures, regardless of the fluctuations
-Of the magnetic system of the PM pickup that the indicators will show the mean
--.compass course.
-b. Shifting of the card of the magnetic pickup.
The dynamic errors of the magnetic system of the PDX, which arise under high
;load factors of the ,aircraft cause a displacement of the card of the magnetic pick-
VP, leading to diScrepancies in the system. *In this case the electric motor (6) of
the gyro unit rotates the potentiometer brushes in a direction diminishing the dis-
crepancy so that the readings of the indicator show an error. However, since the
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high overloads, especially those created during turns, are uswoly of brief duration
and since the gyroscope during this time, maintains its position in space practical-
ly while the rate of deflection of the brushes is small (3?/nm), the er-
ror each minute of turn will not exceed 2.50 on the average.
The eoccesei-v
iazrom 4 tothorant in
the nmsnetic compass and the necessity for a
precise determination of the aircraft course in solving a number of problems, force
us to seek new solutions for determining the aircraft course. One such solution is
the creation of instruments based on the direct mtasur ent of the earth's magnetic
field, i.e., the creation of induction compasses.
emote-Readin Induction C
sees
The remote-reading induction co ss consists of:
1. an induction pickup;
2. several of indicators (including one master indicator);
3. an amplifier;
4 an inverter;
5. a caging mechanism.
The induction pickup is a principal part of the instrument. It is placed in
the aircraft wing or in the tail of the fuselage to isolate it frog the interference
of the ferromagnetic masses of the electric currents.
The sensitive element of the pickup consists of three pairs of rods (Fig.147)
made of 'a material with a high magnetic permeability (permalloy). Each pair of rods
is provided with two windings, one of which serves for the magnetic excitation of
.t,he rods and the other for indicating the change in direction of the rod, with re-
,ispeet of the earth's magnetic field.
The excitation windings ? are connected in series and are fed by 500-cycle alter-
'rating current
. The alternating magnetic fluxed produced by these currents in the rods of each
pair have opposite directions. For this reason, they do not induce an se in the
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.44,44ilmr?
ry windings but merely
40VINi.e.;
a change in the sagnetic perxieability of
he
Because of the periodic variation in the magnetic permeability, the horizontal
onent of the earth's magnetic field produces pulsating magnetic fluxee in the
rods, leading to the appearance of an electromotive forte in the secondary winding
at ermmvogrme
C "4WD
ourantraw
Cht roff UliSilliff
A eisogrAfr $y
_01
gintievitrows
CORIZINT
I
es
4;
ovmvir
0A490142WW5
MaNare
Fig.147 Diagram of Sensitive Element of the Induction Pickup
of a magnitude which is a function of the arrangement of the rods with respect to
the magnetic meridian
The emf reaches its greatest value in the winding of the rod located parallel
to the lines of the field (north-south). No emf appears in the windings perpendicu-
lar to the field (east-vest) since their loops are not cut by the magnetic field.
To each course of the aircraft there corresponds a definite ratio of the emf
of the three windings. This fact is utilized for measuring the aircraft course.
The sensitive element of the pickup is mounted on a gyroscope with three de-
grees of freedom. The body axis of the gyroscope is located in the vertical plane.
The gyroscope maintains the horizontal position of the sensitive element and,
:on any change in course, does not interfere with the rotation of the element togeth-
er with the aircraft about the vertical axis.
Figure 148 gives a diagram for computing the course by such a compass.
240
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The main indicator (4) contains a selsyn (2), which receives the end signals
from the sensitive element.
The secondary winding of the induction pickup feeds the stator winding of the
selsyn (2). In the winding of the selsyn:rotor an emf is thus induced by the resul-
tant magnetic flux of all three stator windings. Consequently, the resultant emf of
the rotor minding will have a different angular direction, depending on the ratio
between the electromotive forces of the pickup windings (i.e., depending on the
course). Since the emf of the selsyn rotor is small, it is amplified in the vacuum-
TO "WC Mgt Y
warairolf
-
Fig.148 7 Circuit Diagram of Remote-Rdi ng Induction Compass
1 - Pickup; 2 - SelsYn; Magnegyn ; 4 7 Hain indicator; 5 - Ar*iliary indi-
cator; 6 - Reducer; 7 - Amplifier; 8 .7.! Asynchronaus motor.
- , 3
, tube amplifier(7) from where the amplified signals are fed to one of the windings ,
,
,
J
of the two-phase synchronous motor (8) of the mechanism of the course computer. 1
-,
11
The secondary winding of the motor is fed by alternating current of the same
4
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?
,to the secondary minding is eons
?r,1
ca.
The electric motor (g) over the reducer (6), rotates the pointer of the main
,indicator (4) and the rotor of the selsyn (2). When the rotor of the selsyn has ro-
tated until its winding is perpendicular to the resultant magnetic flux of the sta-
tor, the emf in the selsyn disappears, the electric motor (8) stops, and the pointer
of the instrument indicates the course of the aircraft.
Every change of course is accompanied by a change in direction of the magnetic
flux; in this case the electric motor is again excited and shifts the rotor and the
instrument pointer to a new position,corresponding to the new course.
The course computer mechanism, while spinning the rotor of the selsyn tube,
also spins the rotor of the magnesyn (3) which transmits the readings to the auxil-
iary indicator (5).
The auxiliary indicators have a receiving magnesyn with a pointer on the rotor.
Since the transmitter of the compass has no moving sensitive elements to cause
inertia errors and
rarnetto
from the ferromagnetic masses producing deviations; it
will give readings with a high degree of accuracy; this has led to its widespread
use in medium and heavy aircraft.
The absence of a Cardanic error is a great advantage of the induction compass,
With the sensitive element stabilized in the plane of the horizon.
In designing a remote-reading induction compass, it must be remembered that the
, 'very accurate gyro vertical used in it must be installed at the center of gravity of
/-
the aircraft (to ensure adequate accuracy of the readings) while the magnetic
;
Aion pickup must be placed in a location where the compass deviations are negligible.
1111 practice, the magnetic deviations are not at a minimum at the center of gravity
Of the aircraft.
A.drawback which limits the usefulness of the remote-reading induction compass
LI its complexity and bulkiness.
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In flight, an aircraft may rotate simultaneously with respect to all three axes.
Since the aircraft rotates with respect to its center of gravity, it is neces-
sary for maintaining the longitudinal equilibrium that the sum of the moments act-
on the aircraft with respect to the zz axis (cf. Fig.1), are equal to zero. The
-moments turning the aircraft with respect to the zz axis may be due to aerodynamic
forces, to forces of gravity, etc.
If the longitudinal equilibrium of the aircraft is disturbed, the elevator must
be deflected to restore it.
The equilibrium of the aircraft with respect to the yy axis is ensured if the
SUR of the momsnts acting in the plane of symmetry is equal to zero. This is called
course equilibrium The disturbance of this equilibrium may be due to a number of
-'reasons: distortion of the planform of one half of the wing, deflection of the rud-
:der, unequal thrust of the propellers in multi-engine aircraft, etc.
A disturbance of eqmilibrilmi results in turns of the aircraft to the right or
left, with a simultaneous dip in the direction of the turn. This inclination is the
result of the fact that the outside wing (with respect to the turn) traverses a
,longer path during the same period of time; consequently, its speed, and lift will
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?????.The'eqUiLibrimm?Of the aircraft with to the kxaxis is called lateral
. . . . ? ,
. . . . . . . .
? , ? ? ? . ? ? ? . . . ? ? . ? . . ? ? ? ? - ? ? ? ? ? . .
and is: maintained primarily by strict symmetry of the Alreraft With re
The reaction of the propeller-engine group, deflecting the aircraft toward the
side opposite the sense of rotation of the propeller, is absorbed by aileron tabs
(small movable control surfaces circumscribed in the dimensions of the aileron).
To restore the lateral equilibrium, both ailerons and tabs are used.
While banking, the aircraft tends to turn toward the side of the bank.
this reason, in banks and turns the pilot always moves rudder and ailerons simulta-
neously.
The ability of the aircraft to maintain a constant flight attitude without in-
tervention by the pilot as well as its ability, during a short disturbance of equili-
"brium, to restore this attitude rapidly, is called the stability of the aircraft.
The stabty of the aircraft with respect to the xx and yy axes is callwilat-
eral stability, while the stability with respect to the mc axis is known as trans-
verse stability and the stability with respect to the yy axis as directional stabil-
The additional concepts of dynamic stability and stability in roll are also
An aircraft is called dynamically stable if, without interference by the pilot,
- :the Nor* *table the aircraft* the harder it is for the pilot to chance its attitude,
4 '
, ' ?,?-? ??- - ? ,5
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?Iev
To facilitate the work of the pilot in controlling the aircraft, automatic sta-
bilizers known as autopilots are used.
Almost all present day automatic pilots have gyroscopic units
Automatic pilots with gyroscopes having three degrees of freedom, are very im-
portant for flying without landmarks or at zero visibility.
The automatic pilot is an automatic regulator designed to hold the aircraft at
a predetermined attitude without intervention by the pilot. It is therefore neces-
sary that the parameters characterizing the given attitude remain constant or, at
least, that the deviations are as small as possible. For example, in rectilinear
horizontal flight, the mean speed, altitude, and course must remain constant, and
the amplitude of their variations must be small. In addition the oscillations of
the aircraft about all three of its axes must also be eliminated.
Certain parameters characterizing the state of equilibrium of the aircraft rel-
ative to its axes may vary during flight (centering, weight, propeller thrust, etc);
the automatic pilot must, without intervention by the pilot, hold the aircraft at
the assigned attitude (or more exactly, at an attitude differing only slightly from
? the assigned one).
The automatic pilot must execute all principal maneuvers, i.e., right and left
turns, in horizontal flight as well as in climbing and descending.
The automatic pilot must be reliable, independent in its characteristics of the
surrounding medium, and simple in operation.
Since the simultaneous stabilization of the four principal parameters of
flight; altitude, speed, course, bank, is possible only by the aid of four control
elements, i.e., throttle control, elevator, rudder, and ailerons it follows that an
automatic pilot which completely frees the pilot from controlling the aircraft must
',consist of four automatic devices, for the throttle control of the aircraft engine,
,the elevator, the rudder, and the ailerons. In most modern autopilots this problem
has not been fully solved, and the only types of autopilots available at present are
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those manipulating the control surfaces of the aircraft.
Figure 149 shows schematically the main parts of the automatic portion of the
automatic pilot.
The sensitive elements consisting of a gyroscope with three degrees of freedom,
picks up any variation in the aircraft parameter regulated by it.
cow rwoz
crow air
rrmfa ?awe
$ONS/ rve4F
emsov
forweiroo Ts
4f4,
"weave
smear
or0A 01V- l/P
ON/ r
Same ihvir
100%/741404
awn&
nuusoomcd
Fig.149 - Principal Parts of the Automatic Pilot
At present the sensitive element for maintaining the selected direction in an
,automatic pilot, usually is a gyroscopic instrument in the form of a gyroscope with
three degrees of freedom, which is less subject to the action of gusts than other
- ;instruments.
The correction in these instruments does not basically differ from the correc-
tions used in gyro horizons, directional gyros, etc. Nembrane instruments operated
177 air pickoffs are used for stabilizing the flying speed and altitude.
The intermediate mechanisms receive the impulses from the sensitive elements,
4,44,
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amplify and compound them where nacos ary(if one control surface is controlled by
severalsensitive elements), and control the operation of the servo unit which de-
flects the corresponding follow-up unit (the control surface).
Intermediate mechanisms consist of:
a. a sensing element: the pickup;
b. an amplifying element;
t. a compounding element;
d. a control element.
The power sources supply the necessary energy to all remaining elements. The
most widely used are pneumatic, hydraulic, and electric energy. Electric energy is
coming into increasingly wider use.
The follow-up unit coordinates the angle of control-surface deflection with the
value of the deflection of the parameter being stabilized.
The indicator units allow the pilot to follow the operation of the automatic
pilot and its components. The automatic pilot is turned on and off by control knobs.
The sevo unit deflects the control surface of the aircraft by means of an ex,-
ternal source of energy. According to the type of energy used, these are divided
into pneumatic, hydraulic, electric, and mechanical. The pneumatic units, usually
operating on a pressure of 1.5 - 5 kg/cm2, have the smoothest action. Of all power-
consuming elements of the? automatic pilot, the servo units consume the most power.
The source of energy used generally is a pneumatic or hydraulic pump. Accord-
ing to their purpose, these pumps are divided into groups serving: a) the sensitive
.elements; b),the Servo units; c) the intermediate mechanisms.
? Section 2
PRINCIPLW OF CONTROL
Any automatic pilot operates on one of the following control principles:
?1. Direct Control: a Direct?Acting Automatic Device. The sensitive elements
of such automatic devices take the form of pendulums or free gyroscopes. Pendulum
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automatic pilots have not come into wide use, since all or them operate at consider-
ble lag and cause flutter of the aircraft (yawing).
Gyroscopic automatic pilots of this type also are not in wide use, because of
the need for high-power gyroscopes.
Fre:Jen-day automatic pilots do not use direct control principles.
2. Indirect Control without Follow-Up. This type of control leads to undamped
oscillations of the control surfaces and
Fig.150 - Principle of Indirect Con-
trol with Follow-Up
1 - Gyroscope; 2 - Tie rod; 3 - Valve
tie rod; 4 - Valve; 5 -Servo unit;
6 - Piston rod 7 - Control surface;
8 - Follow-up clutch; a, b, c, and
the aircraft itself (auto-oscillations).
This method is not used on aircraft.
3. Indirect Control with Rigid Follow-
up (Fig.150). An automatic pilot with
such a control consists of the free gyro-
scope (1), whose frame is connected over
the tie rods (2) and (3) with the valve (4),
controlling the operation of the piston
servo unit (5); the rod (6) of the unit (5)
is connected over a tie rod with the air,-
craft control surface (7). The lever (8)
connects the valve (4) with the upper end
of the piston rod of the servo unit
(follow-up). In this arrangement, the po-
sit ion of the valve depends on the position
of the gyroscope and the position of the
control surface.
On any deviation from the assigned
flight attitude the valve is displaced,
d - Arms of clutch; y Angoe of devi-
lor example, upward; in this ease, the
ation of aircraft.
piston of the servo unit moves downward,
deflecting the control surface to the left. At
the same time the piston rod of the
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.servo unit moves the valve downward. The motion of the piston and the deflection of
the control surface is interrupted, as soon as the valve returns to the neutral po-
sition.
Automatic pilots operating on 'this principle do not always completely restore
'the assigned state of operation.- For example, in aircraft* equipped with a longitudi-
nal automatic device, on any change in centering due, for example, to the displace-
pent of a passenger in the cabin, which will cause a nose-down (or nose-up) of the
aircraft, the automatic pilot will defleat.the elevator upward (or downward) and the
equilibrium of the aircraft would be restored at the instant when the moment pro-
duced by the control surface balances the moment produced by the displacement of the
center of gravity, i.e., at some new .position of the aircraft. The angle at which
the aircraft still deviates from its original position is called the residnal error
of control. For return to the original flight attitude, the pilot must apply correc-
tions manually in accordance with his instrument readings.
4. Indirect Control with a Nonrigid Follow-Up, also Termed Isodromic.
In this ease, the valve (4) is connected with the piston rod of the servo u-
nit (5) over the damper (9), called a cataract (Fig.151)whose degree of damping is
regulated by the cock (10). If the cock (10) is closed, the follow-up is converted
:into a rigid follow-up. The end of the lover (8) carries the restraining spring (i1).
The action of the follow-up of the servo unit on the valve depends on the rate*
of relative displacement of the aircraft and the aircraft control surfaces. In the
,-ease of brief disturbances,this 'automatic pilot operates like an automatic device
,
with a rigid follow-up.
Or prolonged disturbances-, such as due to a change in the thrust of the right
--impeller, the aircraft begins to rotate (in this case to the right) while the valve
.forces the rudder to deflect to the left As in the preceding ease, the aircraft.
Autsume a new position, different from itslnitial.position, at Which the moment of
tthe deflected control surface compensates the moment of asymmetry of the thrust. We
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obtain a new position of equilibrium of the aircraft but not of the automatic de?.
since the spring (ll) in this case will be stretched and will displace the pis-
ton of the cataract (9) and the valve (4)
II upuards for which reason the angle of de-
flection of the directional rudder to the
a
Fig.151 - Indirect Regulation with
Nonrigid Follow-Up
1 - scope; 2 - Tie rod; 3 - Valve
tie rod; 4 - Valve; 5 - Servo unit;
6 - Piston rod of unit; 7 - Control
pensate the moment due to the asymmetry of
surface; 8 - Follow-up knob; 9 - Cata-
the thrust, and all elements of the auto-
matic device, except the piston rod of the
rudder machine will take neutral positions.
left will increase still more. The ecidli-
brium of the aircraft with respect to the
yy axis will be disturbed, and the aircraft
will begin to return to its course.
This process will continue until the
spring and the valve are established in
the neutral position corresponding to the
assigned course. As a result, the air-
craft will now be stabilized on this
course. In this case, the rudder is de-
flected leftward by an angle such that the
moment from the rudder will exactly como-
ract; 10 - Cock; 11 - Spring; - An-
gle of deflection of aircraft.
Section 3
STRUCTURE= OPERATION OF AUTOMATIC PIMA
1.....1tAlErgillealledkANIssagallat
The operation of all three automatic devices of the LP-45 autopilot, longitudi-
lateral, and directional, is based on a single principle. The sensitive ele-
ment which consists of an astatic gyroscope with three degrees of freedom, actuates
a servo rudder over a valve, by means of a special pneumatic relay. This unit then
250
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