ENGLISH TRANSLATION OF A SOVIET MANUAL ENTITLED AIRCRAFT MIG-21F-13, TECHNICAL DESCRIPTION, BOOK I, FLIGHT CHARACTERISTICS

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP80T00246A030200200001-3
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 193
• Document Creation Date: December 23, 2016
• Document Release Date: August 8, 2013
• Sequence Number: 1
• Publication Date: April 1, 1963
• Content Type: MISC

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RET3ORT? I
:RAL ,NT.ELLIGeNCE GEVSY,
?? - LLA
This material contains information affecting the National Defense of the United States within the meaning of t
18 U.S.0 Secs. 793 and 794, the transmission or revelation of which in any manner to an unauthorized per
COUNTRY USSR
REPORT
SUBJECT English Translation of a Soviet DATE DISTR.
Manual Entitled Aircraft MIG-21F-13,
Technical Description, Book I, NO. PAGES
Flight Characteristics
DATE OF
INFO.
PLACE &
DATE ACQ.
THIS IS UNEVALUATED *FORMATION. SOURCE GRADINGS ARE DEFINITIVE. APPRAISAL OF CONTENT IS TENTATIVE.
REFERENCES
April 1963
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English translation of a Soviet manual onhe 50X1-HUM
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MIG-21F-13 aircraft, containing 188 pages
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EXCIAMED FR(50X1-HUM!
DOWNGRA
DECLASSIFICATION
STATE
I ARMY'
I NAVY I I AIR
I I NSA
I DIA I I ?'50X1-HUM !
NIC-: SAC
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(Note: Washington distribution indicated by "X"; Field distribution by "#".)
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AIRCRAFT MIG-21F-13
TECHNICAL DESCRIPTION
Book I
Flight Characteristics
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THE MIG-21F-13 TECHNICAL DESCRIPTION
First Book
?light Characteristics
Chapter 1
GENERAL DATA
.The jet aircraft MiG-21F-13 is a tactical supersonic fighter.
.The MiG-21F-13 aircraft is a single-engine fighter, with one turbojet
engine, with delta wing and controllable stabilizer; it is armed with two
self-guiding rockets *13 and cannon NR-30.
The basic flight characteristics of the MiG-21F-13 are given in Table
1.
Table 1
Flight Characteristics of the Aircraft
1 Maximum speed in.kmihr 2125 (at an altitude of 12.5 to
-18.5 km)
2 Static ceiling in meters 19,000 (at M=1.85)
3 Time to climb to practical ceiling
in minutes up to 6
4 Time of climbing to static ceiling
in minutes:
a) with afterburner turned on at
the moment of takeoff (without
turning during climb)
b) With afterburner turiled on at
an altitude of 8000 m (with
180? turn during climb)
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13.5
16.9
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5 Maximum tactical range at an
altitude of 11,000 m, in km (at
fuel density = 0.83 gr/cm3)
without suspension tank 1400
with suspension tank 1670
6 Maximum tactical duration at an
altitude 11,000 m in hr-min (at
fuel density = 0.83 gr/cm3)
without suspension tank 1 hr 13 min
with suspension tank 2 hr 3 min
7 Takeoff run in meters with
afterburner 800
8 Length of landing run! in meters
with brake parachute 900
without brake parachute 1200 to 1800
9 Takeoff speed in km/hr 315 to 330
10 Landing speed in km/hr 260 to 270
Remarks: Ranges and flight durations are given for a single aircraft with
reserves of 77 of the initial fuel load, allowance for seven minutes/of (
engine operation on the ground prior to takeoff, and with the external tank
being jettisoned when empty,
The strength of the MiG-21F-13 is calculated in accordance with the
following data:
, Table 2
Structural Limitations of the Aircraft
Maximum operational load factor
Maximum indicated speed in km/hr
2
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Maximum Mach nr of flight 2.35
Maximum head pressure in kg/m2 7500
The suspension tank and its attachment is figured by the following
limits (when (when flying with filled and empty suspension tank).
Maximum-permissible indicated
speed km/hr 1000
Maximum permissible Mach nr 1.8
Maximum head pressure in kg/m2 4830
Maximum operational load-factor 6
Table 3
Basic Weight and Centering Characteristics
of the MiG-21F-13 Aircraft
Initial weight kg 7370
Landing weight (minimum at 77. fuel
supply in main tanks less K-13
rockets and cannon shells) kg 5217
Weight of fuel kg (at gamma = .83
gm/cm3) 2080
Practical center of gravity travel
in % MAC 31 to 35
Total center of gravity travel in
% MAC 31.3 to 36.3
General view of the MiG-21F-13 aircraft is shown in Fig 1, 2, 3.
Table 4
Geometric Data of the MiG-21F-13
Wing area 23 m2
Wing span.. 7.150
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Mean aerodynamic chord of the wing
Angle of wing sweep-back along leading edge
Angle of dihedral "V" of the wing
Area (overall) of two ailerons
Maximum angle of deflection of ailerons
(perpendicular to the axis of rotation)
Length of trim tab (only on left
Width of trim tab (perpendicular
edge)
Flap area (2)
aileron) /
to trailing
4.002 '
/57
deg ///
-2 deg
1.18 m2
?20 deg
0.4
0.01 m /
1.87
m2
Angle of flap deflection for takeoff and .
landing (from the streamline position) 24.5 deg
f / /
Total area of two forward speed brake panels 0.76 m2
I
Area of rear speed brake panel 0.47
1.
Maximum angle of deviation of two forward
braking panels 25 deg
Maximum angle of deviation of one rear
braking panel
Area of horizontal empennage
Maximum angles of stabilizer deflection:
?Nose of stabilizer deflected upwards
Nose of stabilizer deflected downwards
Vertical stabilizer area
Maximum angle of deviation of rudder (perpen-
dicular to the axis of rotation)
Length of aircraft (i
Without pitot tube
With pitot tube
4
line of flight)
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40 deg
3.94 m2
7.5 deg
16.5 deg
4.45 m2
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13.46
15.76
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Height of aircraft (when parked with
non-compressed shock absorbers) 4.10 m
Width between landing gear wheels 2.69 m
Longitudinal base of landing gear 4.81 in
Standing angle of aircraft (with uncompressed
shock absorber) 0?16
Landing angle of aircraft (with shock
absorbers uncompressed) 14?03
Chapter 2
FLIGHT CHARACTERISTICS OF THE AIRCRAFT
The flight characteristics of the MiG-21F-13 aircraft include: maxi-
mum speeds at flight altitudes, ranges and flight durations, stability
qualities, controllability and maneuverability as well as takeoff and
landing qualities.
All data, with the exception of stability qualities, controllability
and maneuverability are given for standard atmospheric conditions.
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1 Speeds and Flight Altitudes
The MiG-21F713 aircraft has a wide range of cruising speeds: from'
minimum V of the instrument = 215 km/hr to supersonic speed of 2125 km/hr
at altitudes of more than 12,300*.
Maximum cruising speed is attained during accelerations, when the
engine is working on afterburner, at Altitudes ranging from 12300m to 18500m.
? _ 5.
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The aircraft, under power, has the ability of continuing acceleration'
to greater speeds.
The following maximum permissible operational speeds/Mach nrs of '
flight are established for, the MiG-21F-13 aircraft both with and without /
the K-13 rockets:
a) With K-13 rockets suspended and without them (with pylons)
, H=0-5000m - indicated speed Vlnd=1100 km/hr
(according to the wide arrow)
H=5000-12,300m - indicated speed Vind=1200 km/hr
(according to the wide arrow)
H=12,306m and above - Mach nr=2.CY
b) With non-controllable missiles S-5M or S-5K (suspended instead of
K-13 rockets)
H=0-13,500m - indicated speed Vind=1200 km/hr
(acrording to the wide arrow)
H=13,500m and over - Mach nr=1.8
c) With suspended fuel tanks (filled and empty), with K-13 rockets
or with missiles 1S-5M or S-5K, as well as without rockets and
missiles (with suspension tanks only)
H=0-12,000m - indicated speed Vind=1000 km/hr
(according to the wide arrow)
H=12,000 and over - Mach nr=1.6 Fig 4 and 5
Best operational speed for level flight and maneuver Vind=350 km/hr.
For the MiG-21F-13 aircraft with K-13 rockets the static ceiling at
full afterburning, is best at M1.8 to 1.86 and ,m 19,000m, and without K-13
rockets = 19,500m. For the MiG-21F-13 aircraft with K-13 rockets and with
suspension tank the static ceiling equals 17,500m at M 1.6. The reduction
in ceiling of 1500m is due to the additional drag of the suspension tank
and a reduction of Mach number in climb to M=1.6.
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Practical ceiling of the aircraft with two missiles S-5M or S-5K
without suspension tank is 17 to 17.5 km depending on the fuel supply.
MiG-21F-13 aircraft possesses greater vertical climb velocities,
the value of which at small altitudes with afterburner equals 130 to
140 m/sec and at maximum power without afterburner, 70 to 80 m/sec.
Depending upon the climbing profile and the mode Of engine opera-
tion the static ceiling can be reached in various tines and with the.
consumption of various quantities of fuel. One of the schedules con-
cerning minimum fuel consumption in climbing to static ceiling recom-
mends the following sequence:
- Takeoff and climb to 8000 in and at maximum non-afterburning power
at true speed of 925 to 930 km/hr.
- At an altitude of 8000 in, the afterburner is cut in and the
aircraft climbs to an Altitude of 12,000 to 13,000 in, with an accelera-
tion up to M=1.1 to 1.35, simultaneously with a 1800 turn. After this
the acceleration to M=1.8 to 1.85 and climb to the practical ceiling at
a constant M=1.8 to 1.85 is accomplished.
The vertical velocities of the aircraft MiG-21F-13, Mach numbers in
climb, and time of climb are shown. in Fig 6, 7, and 8.
For the MiG-21F-13 aircraft with afterburning during climb, the time
to climb to 5000 m = 2 min, 10,000 in = 3.2 min, 15)000 in = 5.5 min. When
using maximum power to an altitude of 8000 in the time of climb to each
indicated altitudes increases, for example for 10,000 in to 6.6 min. The
,
time is indicated from the moment of takeoff.
Remarks: The time for takeoff run and acceleration to a climbing speed
at maximum power equals 1.5 min and with afterburning equals 1 min.
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The flight profile and operating mode of the engine when flying with
suspended tank remain unchanged, but the time to climb to altitude with
suspended tank increases by approximately 0.5 min for an altitude of 5000m,.
and correspondingly by 2.5 min for an altitude of 10,000m. To increase
the rate of climb (to reduce the time of climb) the suspension tank can be
released after the fuel from it has been consumed. With this available
fuel system the fuel from the suspension tank is consumed completely when
reaching an altitude of 11,000 to 12,000m.
In Fig 9 and 10 are given approximate practical consumptions of fuel
for various altitude climbing schedules.
2 Range and Flight Duration
The MiG-21F-13, as well as any other supersonic aircraft, possesses
various characteristics with respect to range and flight duration depending
upon the speed and flight altitude.
When strictly adhering to the recommended flight profiles from the
point of takeoff to reaching the ceiling, the amount of fuel consumed in
the attainment of the practical ceiling of the MiG-21F-13 aircraft is 800 .
to 900 liters and the flight duration of the aircraft to the point of the
ceiling is about 6 min.
The time of climbing to static ceiling can be reduced by 0.5 min with
a simultaneous fuel saving of up to 100 liters through a preliminary accel-
eration to a speed of 2100 km/hr and then followed by deceleration of the
aircraft during climb.
Maximum range and flight duration of the MIG-21F-13 aircraft is
attained at an altitude of 11,000m at a cruising speed Of flight. I
table 5 are given data concerning the flight conditions for maximum range
and flight duration.
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Table 5
14IG-21F713 AIRCRAFT
Practical Range and Flight Duration of the Aircraft at an Altitude of 11,000m
with Reserve Fuel Equal to 7% of the Basic Amount
(At specific weight of fuel=0.83 g/cm3)
a) With rockets K-13
Initial gross weight of aircraft - 7370 kg
Total fuel supply - 2080 kg
Supply of fuel for horizontal flight - 1330 kg
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Flight
Speed of flight
Rev of low
km-consump
Hourly
Range
Flight duration
condition
km/hr
pressure en-
of fuel
fuel
km
(hrs-min)
Instr
True
gine rotor
in %
kg/km
consump
kg/hr
Hor
Pract
Hor
-Pract
1
2
3
4
5
6
7
8
9
10
Max range
,
520
925
91
1.12
1040
1190
1400
1-17
1-37
?
Max flight
duration
440
795
89
1.20
955
1110
1320
1-23 -
1-43
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11} Without rockets (with pylons AHY)
Initial gross weight
Total fuel supply
Supply of fuel for horizontal flight
- 7215 kg'
- 2080 kg
1410 kg
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Flight
condition
?
Speed of flight
km/hr
Rev of low
pressure en-
gine rotorkg/km
inl
km-consump
,of fuel
Hourly
fuel
consump
kg/hr ?
Range
km
Flight duration
(hrs-min)
Instr
True
_ _ ___
Hor
Pract
Hor
Pract
1
2
3
4
5
6
7
8
9
10
Max range
520
925
87
1.01
937
1400
1580
1-31
1-49
Max flight
d 09
440
795
85
1.10
885
1280
1460
1-36
1-54
c) With K-13 and suspension tank, electable after consumption
Gross flight weight - 7840 kg
Total fuel supply - 2480 kg
Fuel supply for horizontal flight - 1680 kg
1
2
3
4
5
6
7
8
9
10
Max range
520
925
92
1.17
1080
1440
1670
1-34
1-56
Max flight
duration
440
795
90
1.26
1000
1840
1570
1-41
2-03
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Table 5a
Practical Range and Duration of Aircraft at an Altitude
of 10,000m with S5M or S5K Missiles with 77 Fuel Remaining
(At a specific weight of fuel gamma = 0..83 gm/cm3)
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Flight
condition
Speed of flight
km/hr
Rev of low
pressure en-
gine rotor
in %
km-consump
of fuel
kg/km
Hourly
fuel
consump
kg/hr
Range
km
Flight duration
(hrs-min)
Instr
True
Hor
Pract
Hot
Pract
1
2
3
4
5 6
7
8
9
10
a) With two non-controllable reaction missiles S-5M or S-5K
Initial gross weight - 7417 kg
Total fuel supply - 2080 kg
Fuel supply for horizontal flight - 1397 kg
1
2
3
4
5
6
7
8
9
10
Max range
550
910
88.8
1.20
1090
1160
1320
1-17
1-35
Max flight
duration
440
745
88.4
1.42
1060
985
1155
1-19
1-37
b) With two non-controllable reaction missiles S-5M or S-5K and suspension tank
Initial gross weight - 7877 kg
Total fuel supply -i 2480 kg
'Fuel supply for horizontal flight - 1722 kg
1
2
3
4
5
6
7
8
9
10
Max range
550
910
90
1.32
1200
1300
1490
1-26
1-46
Max flight
duration
440
745
89.2
1.53
1140
1130
1320
1-30
1-50
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In calculating the range and flight duration of the aircraft, the
following factors were considered:
-a) Fuel consumption during the operation of engine on the ground
(starting and testing of engines and taxiing) for a period of
7 min = 60 kg I
?b) Fuel consumption,, distance and time for takeoff and climb and
during gliding in, conformity with tables 6, 7, and 8
c) Fuel consumption When flying in circle before landing for a
period of 4 min 80 kg
d) Non-consumed supply of fuel = 30 kg
e) 77, of fuel supply of the basic supply at a specific weight
gamma = 0.83 g/cm = 145 kg.
In the table is given, the value of range and flight duration without
launching K-13 rockets during flight. With consideration of the launching
of K-13 rockets during the, horizontal flight the range and duration of the
horizontal section increased by 5%.
Fuel consumption, time, and distance during takeoff and climb are
taken at maximum operating
12
condition of the engine.
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Table 6
Fuel Consumption, Time, and Distance for Takeoff and Climb at
Maximum Condition of Engine of MiG-21F-13 Aircraft
..._
Flight
Altitude
in
Meters
- Without K-13 Rockets
(With Pylons)
V = 930 km/hr
With K-13 rockets
V = 930 km/hr
With K-13 rockets
and susp tank
y = 930 km/hr
Fuel
consump
kg
Time
min
Distance
km
Fuel
consump
kg
Time
min
Distance
km
Fuel
consump
kg
Time
min
Distance
km
1
2
3
4
5
6
78
.
105
220
305
380
415
2.0
4.3
6.5
8.4
10.0
10
10
45
75
110
130
1,000
-5,000
-,_
10,000
11,000
70
165
225
,
265
,
285
1.3
3.1
4.1
5.4
6.2
5
30
50
65
85
85
195 ?
260
325
360
1.5
3.4
5.0
7.3
8.6
5
35 -
60
-90
110
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Table 7
Fuel Consumption, Time, and Distance
during Gliding of MiG-21F-13 Aircraft
For all variants of external suspensions
including
Time
min
suspension tank)
Flight altitude
m
Fuel Consumption
kg
Distance
km
5,000
20
3.0
35
8,000
40
5.0
60
10,000
60
7.0
85
11)000
75
8.0
100
17_,500
120
135
13.0
14.0
165
180
19,000
Remarks:
1 Gliding from all altitudes is done at Vins = 500-550 km/hr
2 Engine control lever on small gas
3 Brake panels in the retracted position
Table 8
Fuel Consumption, Time, and Distance during Takeoff and
Climb at Maximum Operation of the MiG-21F-13 Aircraft
Flight
Altitude
in
Meters
With S-SM or S-51?. missiles
and without suspension tank
V = 900 km/hr
With suspension tank and
S-SM or S-SR missiles
V = 830 km/hr
Fuel
cons
kg
Time
min
Distance
km
Fuel
cons
kg
Time
min
Distance
km
11000
85
1.35
10
90
1.5
10
5,000
180
3.0
35
'
210
3.7
40
10 000
_ A
320
7.0
90
390
9.3
115
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3 Takeoff and Landing Characteristics of the Aircraft
For the MiG-21F-13 aircraft without suspension tank the speed of
breaking from the ground with deflected flaps equals 315 to 320 km/hr
(under standard atmospheric conditions in the absence of wind) at an
angle of attack at takeoff of 9 to 100.
To assure this speed of break away of the aircraft the beginning,
of lifting the nose wheel should be realized at a speed of not more than
210 to 220 km/hr. The time of climb at the takeoff angle of attack with
afterburning should be 3 to 4 seconds, and at maximum power should be 5
to 6 seconds.
When the nose wheel is lifted off at higher speed or during slow
attainment of the takeoff angle of attack the speed of separation of the
MiG-21F-13 aircraft increases.
At normal rate of separation, V = 315 to 330 km/hr, the length of '
the takeoff run of the aircraft is: L takeoff run = 1150 to 1300m. at
/
maximum power; L takeoff run = 850 to 900 with afterburning.
With suspended tank and a rate of separation V = 330 to 340 km/hr,
the length of the takeoff run is: L takeoff run = 1250 to 1350m at
maximum power; L takeoff run = 900 to 1000m with afterburning.
In figure 11 is given an approximate dependence of the takeoff speed
and length of takeoff run upon the angle of attack during takeoff.
In table 9 are given approximate fuel consumptions andtimes necessary
for takeoff and for acceleration of the aircraft from the moment of lift?''
off to the beginning of. climb.
The length of the landing run after touching ground depends upon:the
landing speed of the aircraft, timely application of the brakes
15
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three wheels, and upon the state of the surface of the landing strip, as
well as upon timely deployment of the brake parachute.
Table 9
Engine Operation
Afterburning
Maximum
Time of takeoff run (sec)
17
25
Fuel consumption on takeoff run (kg)
55
26
Time of acceleration up to V climb =
1000 km/hr (sec)
32
50
Fuel consumption during acceleration
to V climb = 1000 km/hr (kg)
110
60
Data concerning the length of landing run is given for a dry runway
strip. The landing speed of MiG-21F-13 aircraft is equal to 260 to 270
km/hr at an angle of attack of 9 to 10?.
The length of the landing run of the MiG-21F-13 aircraft without brake
parachute, but with braking of all three wheels, equals 1200 to 1300 m.
A reduction in landing run is attained to a large extent by applying.
the brake parachute.
With released brake parachute the length of the landing run of the
MiG-21F-13 aircraft is reduced to 900 m. We must remember that, even when
releasing the parachute at the moment the aircraft touches ground, the time
between releasing the parachute and its inflation is 2 to 2.5 sec during
which time the aircraft covers a distance on the ground of 150 to 200 m
actually without the braking effect of the parachute.
16
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The landing characteristics are listed
in the following table:
Without brake With brake
parachute
Landing weight (kg)
5480
/parachute
5480
Angle of flap deflection.
24.5?
24.5?
/
Landing speed (at alpha =
9 to 100) (km/hr)
260 to 270
260 to 270
Length of landing run (on
dry runway) (m)
1200 to 1300
900
Length of landing distance
(from an altitude H = 25m)
(m)
2300 to 2400
1600
In figure 12 are given approximate dependence of the landing speed
and length of landing run upon the angle of attack of the aircraft during
landing.
Takeoff and landing distances are shown in figure 13 and 14.
Stability and Controllability
In this chapter are given the basic characteristics for longitudinal
and lateral stability and controllability of the MiG-21F-13 aircraft,
derived from flight tests.
The basic criterion of ongitudinal stability of the aircraft appears
1
to be the degree of longitudinal static stability in accordance with .over
1
load. On the MiG-21F-13 aircraft the centering was selected in such a
way that under subsonic speed conditions the amount of longitudinal static
1
stability appears to be optimum for light fighters. It constitutes on an
average 3 to 57. MAC. The change in neutral centering of the aircraft by
17
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the Mach number of flight, as given in figure 15, shows an increase in
the amount of stability when changing over to supersonic speeds. When
the engine is cut off the amount of longitudinal stability rises by ,1 to
27 MAC. When launching K-13 rockets the stability is practically un-
changed.
Longitudinal controllability of the aircraft is determined by the
following parameters, felt directly by the pilot during flight, - overload
of the aircraft, shifting of control stick, and by the forces affecting
it. In figure 16 is given the required angle of deflection ft of the
stabilizer for a 1 g overload versus the Mach number of flight for several
flight altitudes. The relation of the angle of deflection of the stabili-
zer to the magnitude of overload, 1):: 4? , for all Mach numbers of
flight in the case of the MIG-21F-13 aircraft is linear, which appears to
be a positive quality of the aircraft. An automatic mechanism for changing
the gear ratio between the control stick and the stabilizer (ARC-KV) is
installed in the longitudinal control system to provide good control char-
acteristics - magnitude of stick movement (Xe) and stick force (Rn) required
to produce a 1 g overload.
In figure 13 are given dependence of Xnu upon the Mach number and
flight altitude, obtained during flight test. Mathematically .the_charac-...
ay ..'?
teris tics Xnu and Pnu,are determined by'fOrmulasA --- /
1
. .
where: n) =. the gear ratio between the Movement of the handle and the
.deflection of stabilizer; Px = gradient of,change ih forces on the control
handle by its movement'.
18
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?
The valueS)03 and Px? which appear to be the characteristics of
the ARU-ZV mechanism mounted on the aircraft do change-in relation to the
speed and flight altitude as shown in figures 18 and 19.
The balancing curves - dependence of the angles of stabilizer deflec-
tion, movement of control handle, forces on the handle in recti-linear
flight (for a 1 g overload) upon the M nr of flight is given in figures 20,
21, and 22.
Balancing curves have a smooth character of change in accordance with
the flight speed. In the) range of near-sonic speeds there is a small in-
stability with respect to, speed, which does not deteriorate the technique
--of aircraft piloting. The balance of the MiG-21F-13 aircraft is such that
the trimmer mechanism produces the effect of zero stick force when the stick
is in the neutral position at flight speeds V ins = 750 km/hr under climbing
conditions after takeoff
at altitudes ranging from 0 to 3000 m.
The nature of changes in the forces against the control handle during
a change in flight velocity depends considerably upon the magnitude of
stabilizer "manipulation." "Manipulation" refers to the change in the angle
1
of deflection of the stabilizer when shifting the rod of the operational
mechanism of the ARU-ZU epparatus on the control handle and when the mechan-
ism ARU is switched over
from the larger to the smaller arm.
By changing the magnitude of stabilizer "manipulation" it is possible
to balance the control forces by the air flow, attaining thereat a reduction
in instability with respect to speed to a minimum.
In the chapter dealing in leveling and balancing of the aircraft is
thoroughly discussed the
stabilizer "manipulation
19
method of changing the balancing on account of
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The flight with 0 forces against the control handle is achieved by
installing in the control system a trimmer effect mechanism, the opera-
tional range of which offers the possibility of such flight at practically
all speeds and altitudes.
The extension of landing gears, ming flaps, brake flaps, change in
operating conditions of engine, and the mounting of a suspension fuel tank
have practically no effect on the change in stability and controllability
characteristics.
The basic characteristic of lateral stability appears to be the degree
of directional stability. On supersonic aircraft the degree of directional
stability decreases, beginning with ,a Mach number and at subsonic flight
velocities corresponds to the standard requirement;and in the range of Mach
number = 1.2 to 1.6, it rises. When flying with suspended fuel tank the
degree of directional stability corresponds to the standard requirement
at Mach 1.6.
The characteristic of Lateral control is determined by the magnitude
of angular banking speed(W) produced by one degree deflection of
ailerons, and by the magnitude of forces acting against the control handle
(Rzwx) during the creation of an angular banking speed of 1 rad/sec as
shown in figures 23, 24, and 25.
And so on the MiG-21F-13,aircraft the system of irreversible booster
control of ailerons has a charging mechanism; and the value RzWx can be
f,
determined: ; ig) if; ? ,rr f- Am 7
a-7-77.-1'4264/
x
The characteristic of the charging mechanism R,Xr is non-linear; also
non-linear is the gear ratio from the handle to the ailerons Xr8x (see
20
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figure 26). The non-linear
and the deflection of the
connections between the movement of the handle
ailerons on the aircraft are for the purpose of
reducing the extreme effectiveness of ailerons, at greater instrument speeds
(in a small range of aileron deflection angles).
Maneuverability
Maximum overloads of the aircraft (in the plane of symmetry) are
determined in the range of subsonic Mach numbers.
Maximum lift is produced by the aircraft at angles of attack smaller
than critical, and at sUpeIrsonic speeds--with sufficient longitudinal
stability and stabilizer effectiveness. In the range of Mach numbers from
minimum to. Mach lA or 1.4 when, during overload conditions, the aircraft
approaches a discontinuous angleof attack, there is felt a warning vibra-
tion in the aircraft. In figure 27 is shown a change in Maximum-oriented
I '
overloads of the aircraft in the plane of its symmetry by the Mach number
for a number of altitudes.I
Maximum negative overloads are limited by-the condition of continuous
feeding of fuel into the engine. The fuel feeding system is intended for
a 15 sec delivery for the engine under negative load conditions.
One of the important features, characterizing the maneuverability of
the aircraft, is the ability of the aircraft to rapidly change the speed
of flight. The characteristics of aircraft acceleration at various alti-
tudes, time of acceleration, fuel consumption and length of path during
.accelerations are given in figures 28, 29, 30, and 31.
Braking characteristics of the aircraft with cutoff engine are shown
in figures 32 and 33. In figure 34 is given the change in maximum -overload
21
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of the aircraft, originating during the release of brake flaps versus Mach
number of flight.
On the MiG-21F-13 aircraft it is quite simple with small deflections
of control surfaces to execute an acrobatic figure, whereby all control
surfaces retain their effectiveness at any evolution of the aircraft. ,The
nature of spins is shown' in figures 35, 36, and 37.
22
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Chapter 3
WEIGHT AND CENTERING CHARACTERISTICS
OF THE AIRCRAFT
In the chapter concerning weight and centering characteristics are
given weight and centering characteristics of the MiG-21F-13 aircraft
with K-13 rockets--a variant with normal load; with K-13 rockets and.sus-
pended fuel tank, with non-controllable missiles S-SM or S-5K, mounted
instead of K-13 rockets, .with non-controllable S-SM or S-5K missiles and
with suspended fuel tank.
A calculation of the aircraft's centering is done in coordinate axes
"X" and ")t". Axis "X" appears to be the longitudinal axis of the aircraft
and is situated in the plane of structural horizontal. The axis "X" is
perpendicular to axis "X"i and passes in vertical plane, situated at 600 mm
toward the tail of the aircraft from bulkhead nr 16. The intersection of
axes "V' and "Y" is the starting point for the reading of CG coordinates
of the aircraft and its components.
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The coordinates of the units, arranged in the direction of the air-
craft's tail along the axis "X" and upwards along the axis 'T" have the plub
sign, and toward the nose, of the aircraft and below they have the minus sign. ?
The weight characteristics with respect to centering relative to the
centering of axes "X" and "Y" of the aircraft at takeoff with normal load,
and with normal load plus suspended fueltank and non-controllable missiles
S-SM or S-5K are given in table 10.
The centering of the aircraft is given in %MAC of the wing relative
to the point where the MAC (average aerodynamic shord of the wing) intersects
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??;
the wing leading edge. The IC" coordinates of the CG of the aircraft' in.
direction of axes "X" in %MAC are determined in the following manner:
. _
?
= r,. r-e
? .
where: alpha X = 1.302 m ? distance from the intersection point of the
.wing leading edge and the MAC to the centering axes "Y" (see dis-
positionrif MAC on figure 38)
X = distance of CG of the aircraft from the origin of the coordi-
nates, which are taken with their own sign
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bA = 4.002 m = average aerodynamic wing chord (MAC).
Weight and centering characteristics of the aircraft with normal load
and with K-13 rockets plus suspended fuel tank are given in table 11: with
non-controllable missiles S-5M or S-5K and with non-controllable missiles
S-5M or S-5K plus suspended fuel tank are given in table 12.
Centering of the aircraft in flight changes depending upon the consump-
tion of fuel from the tanks and upon the operating conditions of the engine.'
Consumption of fuel from the tanks during non-afterburning engine
operation takes place in the following order:
1. From tanks 1, 2, 3, 4 5, 6 until a special float valve opens the
connection to the suspended tank.
2. From rear section of suspended tank.
3. From forward section of suspended tank.
4. From rear wing tanks.
5. From forward wing tanks. ?
6. From first group of tanks.
7. From third group of tanks.
8. From second group of tanks.
Small changes in' aircraftcentering during flight, depending on fuel
consumption, occur when operating the engine With afterburner. The order
24
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of consuming the fuel in flight with afterburning operation of the engine
differs from that. during non-afterburning operation by combined consumption
of fuel from wing tanks and partial consumption from fuselage tanks.
The change in aircraft centering during flight in the variant with nor-
mal load due to fuel consumption for afterburning are given in figure 39.
The change in aircraft centering in flight for the variant with K-13
rockets and with suspended fuel tank due to the consumption of fuel for
afterburning isgiven in figure 40.
The change in aircraft centering in flight with non-controllable
missiles S-5M or S-5K due to fuel consumption for afterburning is given in
figure 41.
The change in aircraft centering in flight with non-controllable
missiles S-SM or S-5K and suspended fuel tank due to fuel consumption for
afterburning is given in figure 42.
A change in aircraft centering as a result of a change in weight or
as a result of any other change or modification is permitted within limits
of plus or minus 0.57.MAC of the maximum forward and maximum rear balancing
limits mentioned in tables and shown in figures.
25
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Table 10
WEIGHT AND CENTERING DATA ON THE
111G-21F-13 AIRCRAFT
Name
RX
kg.m
X
in
R
kg
Y
m
Aircraft with normal load at
takeoff (landing gear out)
47
0.006
7370
0.053
Empty aircraft
1020
0.205
4980
-0.021
Normal load
-973
-0.407
2390
0.207
Normal load
-973
-0.407
2390
0.207
Pilot with parachute /
-319
-3.19
100
0.39
Rockets, K-13 (2 units)/
25
0.165
154
-0.52
Rounds, NR-30 (60 unitr)
-100
-1.80
56
0
Fuel in Tank One
-482
-2.41
200
0.38
Fuel in Tank Two
-798
-1.33
600
0.35
Fuel in Tank Three
-51
-0.22
232
0.19
Fuel in Tank Four
90
0.52
174
0.34
Fuel in Tank Five
232
1.16
200
0.40
Fuel in Tank Six
384
1.92
200
0.40
Fuel in Rear wing tanks
205
1.18
174
-0.03
Fuel in forward wing tanks
-159
-0.53
300
-0.03
Retraction of landing gear
-196
26
-
P C-R-R-T
RY
kg.m-
390
-104
494
494
39
-80
0
76
210
44
59
80
80
-9
278
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Name
RX X R Y RY
kg.m m kg m kg.m
Installation of suspension
fuel tank V = 500 liters -132 -0.28 470 -0.96 -452
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Suspension tank -7 -0.146 46 -0.98
Beam of of suspension tank -11 -0.485 24 -0.65 -15
Fuel in suspension tank
V = 490 liters -114 -0.285 400 -0.98 -392
Mounting of non-controllable
missiles S-5M 61 0.22 277 -0.46 -127
Beams 17 0.39 43 -0.23 -10
Missiles S-SM (32 units) 20 0.158 124 -0.504 -62
Pods 24 0.22 110 -0.504 -55
Remarks: 1. Weight and centering with non-controllable missiles S-SM
are no different than with non-controllable missiles S-SR.
2. When mounting non-controllable missiles S-SM instead of
K-13 rockets the rocket installation is removed from the
aircraft.
Mounting of K-13 rockets
46
0.19
237
-0.45
-106
Beams
17
0.39
42
-0.23
-10
Starting device, APU-13u
4
0.107
41
-0.40
-16
Rockets, K-13 (2 units)
25
0.165
154
-0.525
-80
27
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Table 11
TABLE OF CENTERING VARIANTS OF THE
MIG-21F-13 AIRCRAFT WITH K-13 ROCKETS
Loading Conditions
Acft Acft Acft Acft Acft Acft
with with with with w/o with
normal 50% max max fuel ext
load fuel, fwd aft and fuel
w/o e.g. /e.g. ammo tank.,
ammo V=500
liters
Weight of aircraft kg
c.g. locationjgear down %MAC
on X axis (gear up 7.M.la
Empty aircraft weight kg
Useful load kg
Pilot with chute kg
K-13 Rockets kg
Ammo (NR-30) kg
Total fuel weight* kg
Tank #1 (V=241 1) kg
Tank #2 (V=720 1) kg j.
Tank #3 (V=277 1) kg
Tank #4 (V=208 1) kg
Tank #5 (V=241 1) kg
1
Tank #6 (V=241 1) kg :
Fwd wing tanks (V=360 1) kg
Aft wing tanks (V=210 1) kg
External tank (V=490 1) kg,
7370
32.7
32.1
4980
2390
100
154
56
2080
200
600
232
174
200
200
300
174
-
6120
36,0
35.2
4980
1140
100
-
- ,
1040/
i
16
308
192
159
183
182
-
7046
32.1
31.4
4980
,2066
/ 100
/ 154
56
1756
180
560
192
159
183
182
300
-
6086
36.3
35.5
4980
1106
100
-
1006
-
290
'192
159 0
183/
182'.
-
5080
36.1
35.1
4980
100
100
-
.
-
-
-
7840
32.3
31.7
4980
2860
100 -
154
/ 56
2480
200
600/
232'
174
200
200
300
174
400
Wt of external tank installa-
tion without fuel kg i 70
' -*Specific weight of fuel .2)) = .83 gm/cm3 throughout.
28
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Table 12
TABLE OF CENTERING VARIANTS OF THE
MIG-21F-13 AIRCRAFT WITH S-5M MISSILES
Loading Conditions
Acft Acft Acft ACft Acft Acft
with with with with w/o with
normal 507 max max fuel ext
load fuel, fwd aft and fuel
w/o c.g c.g. ammo tank,
IIMMO V=500
liters
Weight of aircraft kg
7410
6190
7086
6156
5150
7880
c.g. locationjgear down %MAC
32.8
36.1
32.2
36.4
36.2
32.4
on X axis (gear up 714AC
32.2
35.3
31.5
35.6
35.2
31.8
Empty aircraft weight kg
5050
5050
5050
5050
5050
5050
Useful load kg
2360
1140
2036
1106
100
2830
Pilot with chute kg
100
100
100
100
100
100
S-5M missiles kg
124
-
124
-
124
Ammo (NR-30) kg
56
-
56
56
Total fuel weight* kg
2080
1040
1756
1006
2480
Tank #1 (V=241 1) kg
200
16
180
-
-
200
Tank #2 (V=720 1) kg
600
308
560
290
-
600
Tank #3 (V=277 1) kg
232
192
192
192
232
Tank #4 (V-208 1) kg
174
159
159
159
174
Tank #5 (V=241 1) kg
200
183
183
183
-
200
Tank #6 (V=241 1) kg
200
182
182
182
-
200
Fwd wing tanks (V=360 1) kg
300
-
300
-
-
300
Aft wing tanks (V=210 1) kg
174
-
-
174
External tank (V=490 1) kg
-
400
Wt of external tank installa-
tion without fuel kg
70
*Specific weight of fuel = .83 gm/ m3 throughout.
29
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Chapter 4
ADDITIONAL DATA ON FLIGHT CHARACTERISTICS
I. Method of determining flight speed and altitude
On board the aircraft are installed instruments for determining speed
of flight relative to the air flow and barometric altitude of flight. For
this purpose an air pressure receiver, DUAS-8M, is mounted in the nose sec;
tion of the aircraft fuselage.
The location of the receiver on board the aircraft has been selected
with consideration that distortions in velocity and flight altitude indi-
cations on the cockpit instruments should be at a minimum.
Flight altitude is determined by instrument indications with intro-
duction into same of instrumental and aerodynamic corrections.
The geometric altitude, i.e., flight altitude over the surfade of the
earth is determined as follows: Hg = R + delta Hz where: H = baro-
metric altitude and delta Hz = local altitude variation over or under
relative to sea level. Delta Hz takes into consideration a correction
for barometric altitude in. connection with the difference in actual atmos-
pheric pressure on the ground and standard pressure of 760 mm of mercury.
For example, at an actual pressure on the ground of 740 mm Hg, the
magnitude of the correction delta Hy determined by a table of atmospheric
standard's equals +220m, and at a pressure of 780mm Hg the correction
delta Hz is -220m.
The magnitude of barometric altitude H is determined by formula:
Hpr 4-L H instr +Ha +.2.1 HB Hzzp where: Hpr = actual
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altitude in accordance with instruments indication in the cockpit;.AH instr
= instrumental correction to altitude indicator obtained after callibration
of instrument; LI Ha = aerodynamic correction to the altitude indicator
(altimeter);/\ HB = altitude correction to altimeter.
The total of aerodynamic and wave corrections (ZAHa +AHB) depending
upon the number M ind and the flight altitude are given in figure 43.
- Remarks:
In figure 43 along the axis of the abscissa, M ind number is taken
with the introduction of an instrumental correction pm instr, which is
obtained during the calibration of the M nr indicator.
Hlag correction for the lag in altimeter indications.
The magnitude of correction for the lag in altimeter indication depends
upon the rate of flight altitude change and upon the volume and length of
wiring of the system for flight altitude measurements, PVD.
In the presence on the aircraft of board instruments only, i.e., in
the absence of special recorders, the value of the correction for lagging
is small.
The true rate of flight V (relative to the flow of air) is determined
Vi
by V = where: V; = indicated speed, the speed.which would be
indicated by the speed indicator, if it would not have any errors in the
instrument and would be situated at an altitude, of H = 0, i.e., under con-
ditions for which it has been calibrated.'
ZA4k
= relative air density.
The magnitude of indicated speed Vi is determined by formula:
Vi = Vpr +ci V instr +S-Va +S Vb Vcom trc Vlag
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0 ?
instrumental speed (according to the wide arrow);411V instr = instrumental
correction to speed indicaLr, obtainable during the calibration of instru-
ment;Pra = aerodynamic correction of speed indicator, taking ipto consid-
eration the distortion of static pressure at the point of installation of
the air pressure indicator;6Vb = wave correction, taking into considera-
tion the distortion of static pressure at the point of installation of the
air pressure receiver, due
to the effect of density jumps; the total of
aerodynamic and wave corrections (pa +ciVb) depending upon Mach number
and flight altitude is given in figure 44;Plag ' correction for the lag
of the speed indicator in time, practically equals 0; Pcom ''' correction
(
for the difference in compressibility of air at given altitude and on zero
altitude (for which instrument calibration was made).
The dependence of the
correction? Vcom upon air compressibility upon .
the indicated speed, ViVercor +ccVa +crib, is given in figure 73.
where: VPrc0r = Vpr +cr V instr
We want to point out that the correction 07 instr is always negative.
This leads to the fact tha the speed indicator will always indicate a
somewhat delayed speed. This explains the frequent misunderstandings when
the pilot is convinced thati? he attained in flight a greater speed than in
actuality.
One of the highly important parameters characterizing the speed of
flight is the Mach number, indicating the ratio of true -flight speed to
V
the speed of sound, i.e., M = ?
V a
The value M = ? = 1.0 appears to be the boundary of the zone of sub-
sonic speeds (M smaller than 1.0) and supersonic velocities (at M greater '
than 1.0).
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The true Mach number of flight is determined by formula:
V
M = - = +1M
Mpr instr
a TMa + (Mb +6141ag
by instrument in the cockpit:cc M
-instr instrumental correction to Mach
number indicator:cc-Ma = aerodynamic correction to Mach number indicator;
where: Myr = Mach number
mb
= wave correction to Mach number indicator.
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The dependence of (.Na +crMb) upon the number Mpr (with consideration
of instrumental corrections) of flight is shown inligure 45.
It is evident from this figure that for Mach numbers greater than
Mpr = 1.08 the air pressure, receiver has no' aerodynamic or wave correction,
i.e., for flight Mach numbers of more than Myr = 1.08 the speed indicators,
indicators of Mach number and flight altitude give indications without die-
tortions.
S-mlag
= correction for lag to Mach number indicator.
Analogous to the correctionilag the correction Miag is also small.
P
In addition it is also possible to employ the following ratio of
connections between the Mach number and the speed of flight: V = 72.2M1[
where T = absolute temperature of outer air. T = 273? + t? where t? is
the temperature in degrees Centigrade.
For practical utilization when changing instrumental speed into .rue
speed and into Mach number of flight or when changing from true speed or
Mach number of flight into instrumental speed are given graphs Vyr = f(M)
and V = f(V;H) in figure 74 and 75. On these graphs of V values should
pr pr
be taken with the introduction of an instrumental correction eiVinstr?
2 Bringing the Static Ceiling to Standard Conditions
The ceiling of the aircraft depends to a large extent upon the tempera-
ture of the outside air,.because the thrust of the engine decreases with
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increase in temperature of outside air and increases during its reduction.
This leads to the fact that the static ceiling will depend upon the
temperature of the outside air, and the higher the temperature the lower
will be the static ceiling.
Evaluation of the obtained ceiling during afterburning operation of
the engine and its conformity to the ceiling under standard conditions is
given in figure 46, where the dependence of the static ceiling value is
shown in the relationship to the actual temperature of the outside air.
. I
The static ceiling value is also affected by the flight weight of the
aircraft, therefore when
evaluating the obtained ceiling at given actual
temperature of the air it is necessary to bring in a correction for the
weight.
If the flight weight of the aircraft of the aircraft at the ceiling
differs from normal weigbt, then for each 100 kg the actual weight exceeds
the normal one the value
of the obtained altitude would be decreased by
100m, and for each 100 kg of reduction
tude would be increased by 100m.
in actual weight, the obtained alti-.
Taking into consideration the production deviations in the role of
aircraft manufacturing and a tolerance for engine thrust, a static ceiling
-tolerance has been established which reduces the static ceiling by 2% of
the standard. In figure
46 the lower boundary of static ceiling is shown
in relation to the actual outer temperature of air.
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Chapter 5
LEVELING AND BALANCING OF THE AIRCRAFT
I Ground Leveling and Balancing
Technical control over the position of outer contours of the aircraft
and over dimensions of control systems is realized with the aid of a
leveling arrangement.
When making flight tests and during flight operations some control
dimensions can be changed. All these changes should be reflected in the
leveling system. As has been established by many years of practice, a part
of the leveling can be done during the process of flight operations, but
only on certain selected aircraft. Consequently the leveling system is
divided into two parts: a basic leveling system for each aircraft and an
addition to it for selective aircraft.
Measurement by the leveling system should be carried out in accordance
to methodical instructions (see Appendix 1 to Chapter 5).
Measurements by the leveling system and bythe addition to the leveling
system are also carried out in the event repairs were made on the aircraft
or when, in the process of flying, any kind of abnormalities in the behavior
of the aircraft are detected ( e.g., unusual, above normal turns, banks; etc.).
Given below are methodical instructions on the conduct of ground
leveling and balancing of each aircraft in addition to the measurement
carried out with this instrument for selective aircraft (see Appendix 1 to
Chapter 5).
The leveling system of the MiG-21F-13 aircraft is shown in figures 48-61).
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2 Balancing of Aircraft in Flight
Balancing of the aircraft in flight is realized in accordance with a
program of flight tests in conformity with methodical instructions.
Methodical instructions on the balancing of an aircraft in flight are given
1
in.the Appendix 2 to Chapter 5.
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Appendix 1
METHODICAL INSTRUCTIONS ON LEVELING AND GROUND BALANCING
General Data
1.- The leveling system dexcribed in this book and the auxilary graphs
(Figures 47-61) represent a list of control and assembly dimensions for
the aircraft airframe.
The balancing system is filled in during the assembly of the aircraft
at the aircraft manufacturing plant and appears to be the only document
giving the actual state of the aircraft after assembly and in its further
exploitation.
' The centering system records all changes in the adjustment of the con-
trol system and observed changes in the process of aircraft operation.
The filled-in form of the leveling system is then attached to the cer-
tificate of the aircraft.
Before flights begin the pilot should acquaint himself with the aitual
data concerned with aircraft control recorded in the leveling system.
2. Leveling of the aircraft and control systems is done for each
aircraft in the entire process of leveling systems.
3. For selective aircraft are also carried additional measurements
using the attachment to the leveling system for selective aircraft.
4. In figures 62-72 are given additional data pertaining to ground
leveling of aircraft. This also constitutes a reference material on the
characteristics of the Control systems.
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I. General Leveling of Object
5. Leveling of aircraft is done without fuel, pilot, and special
loads with equipment installed onboard the aircraft.
6. When leveling, the aircraft is placed on support (horses) at bulk-
heads of the fuselage numbers 2 and 28 (weights to the nose of the fuselage
are not suspended). See figure 47.
The aircraft retains stable position on the supports even during the
application on the horizontal empennage of a load of 100 kg. In..addition,
are placed protective supports over the wing (at a distance of 2 meters from
the axis of the aircraft) and under the fuselage along bulkhead number 35.
7. The axis of the main part of the fuselage is placed horizontally
over reference points number I left and 2 left.
In the lateral respect the aircraft is placed horizontally over the
reference points of the wing "8" lower left and "8" lower right (Figure 48).
Allowance for the difference along these points was set up at ?0.5 mm.
8. All leveling points are plotted in the figures up to the point of
general leveling. The points on the wing and horizontal empennage are plotted
only below and on the fin only to the left.
9. After filling in the cards with differences of the type "a-b" and
"b-a" etc., the designations are engraved on the surface.
Remarks:
1. Leveling points number 3, 6, 22, and 18m are used when leveling
selective aircraft.
2. Leveling points 6a, 40, 41, 46, 30, and 27 are used, for unit
leveling.
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10. Measurements after mounting the wing are carried out for the
right and left panels by three sections.
When determining the difference in excess of the leveling points, it
is necessary to take into consideration the local contour deflections from
the chart showing wing contour measurements, i.e., note in the tables the
dimensions with addition (or subtraction) of local deflections. In the
presence of local deflections along the reference points in the leveling
system we make "remarks," -- "point is higher than the contour" or "point
is lower than the contour."
After filling in the boxes in the leveling system with respect to
differences in exceeding points of the right and left wing, for all points
the measurement values of the left wing are subtracted from the measure-
ment values of the right wing, i.e., "right" orileft." If "left" is
greater than "right," then the difference is written in the table with
the minus sight. The wedging angle of the right wing will be greater
than that of the left one in the case where the sign Of measurement differ-
ences "right-left" is as follows:
a) For points 8-9, 8-10, 12-184 16-17 -- minus sign:'
b) For points 7-9, 11L13, 15-17 -7 plus sign.'
In figure 49 is given the measurement arrangement along the wing.
Rudder
11. Measurements on the wedging of the' rudder, figure 49, are carried
out with respect to the axis of the main part of the fuselage (according
to reference points 4-5).
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Rudder tilting measureMents with respect to height are made perpen-
dicularly to the axes, passing over points 4-5.
Stabilizer
12. Stabilizer measurements are carried out when testing the control.
(Mounting of stabilizer in lateral ratio is checked only on selective
aircraft.)
Fuselage
13. Fuselage measurements are not carried out for each individual
aircraft, but only on selective aircraft.
Brake Flaps
14. Measurements of angles of deflection of brake flaps (Figure 49)
are carried out at maximum flap deflection with connected hydraulic system
along the linear dimension according to points 33 and 34.
An angle of flap deflection of 24? corresponds to a nominal dimension
M ... 368 mm.
Flap Blades
16. "Blades" (plates) mounted on the rear edge of the flaps (Figure
49) are intended for flight adjustment of lateral stability of the aircraft
during deflection of the hydraulic system in the process of delivery tests
and are not deflected during operation.
Flap Gap
17. Flap gap is measured (Figure 49) by means of a ruler along
points 33 and 34 at connected hydraulic system with flaps in retracted
. position.
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When measuring the gap, a force of 2 kg is applied to the trailing
edge of the flap.
Bar of Air Pressure Receiver
18. Measurements on the installation of the APR bar (Figure 49)
are carried out with respect to any given axes of the leveling device in
vertical and horizontal planes.
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Ventral Fin
19. Measurement on the installation of ventral fin (Figure 50) /is
done relative to the leveling axes over points 4-5.
Suspension Tank
/
20. Measurements with respect to the installation 'of the suspension
tank (Figure 50) are made from the axes of the main part of the fuselage
(on the 'side) and relative to the axes of symmetry of the main part of the
fuselage (in plan) according to leveling points 56a and 57a.
21. .leveling the pylon under the suspension tank (Figure 50) is
carried out over points 61, 62, 63 and 64 relative to the axes of symmetry
of the main part of the fuselage.
Pylon of Special Suspension on the Wing
22. Leveling of pylon (Figure 50) is done with respect to the line
of the leveling device parallel to the axes of the main part of the fuse-
lage according to points 56, 57, 58, and 59.
Cone
/
23: done measurements (Figure 51) are carried out for three/positions--
cone retracted, and cone extended in two and three positions-- in addition
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is checked the beginning of cone outlet by the flashing of a signal light.
Testing cone installation with respect to the rim should be carried
out in accordance with a special pattern; a schematic drawing of the
pattern is shown in figure 52 where:
ek = distance from leading edge of fuselage to tip of cone in its
first position (retracted).
ek2 = distance from leading edge of fuselage to tip of cone in ex-
tended position.
ek3 = distance from leading edge of fuselage to tip of cone in third
position (extreme extended).
The basic surface should be the plane of the forward edge and outer
surface of the envelope.
Fuselage Flaps
24. Measuring the angles of deflection of fuselage flaps (front and
rear) figure 51 is done in accordance with linear dimensions by points
53 and 56.
Nominal dimensions according to points 53 and 56 correspond to angles
of deflection of the flaps: 200 for forward ones and 350 for rear ones...
Landing Gear
25. On each aircraft are measured the sizes of the base and wheels
of the landing gear (Figure 48).
II. Testing Control Systems
26. Testing of control systems is carried out at the assembly plant.
in conformity with tolerances established for :the Assembly in the leveling
system. When carrying out the receiver-delivering program at a series
manufacturing plant it is permitted to readjust the control systemyith
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expansion of the range of permissible installation dimensions, listed in
the leveling system.
A. Longitudinal Control
Mounting of Stabilizer
27. The leveling point on the stabilizer is point 55.
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point in the stabilizer insert -- point 54 is applied on the left, as well!
as on the right stabilizer insert when mounting each half of it in accor-/
dance with leveling points,19 and 20 (Figure 53).
The dimension (L) in this case should be equal to 0 (i.e., points
54 and 55 on each side of the insert and stabilizer are combined). This
corresponds to a height differential from the axes of the leveling device
along points (19) and (20)' i.e., the dimension b-a = 13mm (here "b" = dis-
tance from the axes of the' leveler to point 20, "a" = distance to point 19).
/e
Remarks:
At such a control arrangement, the"shears" of left and right halves
!
of the stabilizer are equal to zero with a tolerance of ?1mm along points
19 and 20 or along points 54 and 55 (dimension D = 0 ?1mm).
The distance D between Points 54 and 55 when checking the neutral
position is measured in the projection of the vertical for the purpose
of eliminating the effect of slot widths.
Stabilizer Shears
Stabilizer "shear" is: defined as the angle of deflection (in degrees
in accordance with the angle gauge or by the linear dimension "D", or by
the difference "b-a") of the right half of the stabilizer during the mount-
ing of its left half under, a zero angle (see figure 53). When assembling
the aircraft at the plant ithe "shears" of stabilizer halves should be equal
to zero (allowance ?1mm). !
!
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28. Measuring the "shears" of left and right halves of the stabili-
zer are carried out with connected hydraulic system. Using the control
stick it is necessary to place the left half of stabilizer along points
54 and 55 in zero position, D = 0. Measurement along points 54 and 55
of the dimension D of the right half of the stabilizer follows next.
In the case of presence of local contour deviations at measured
points, the measurement should be corrected by a value of initial stages
between stabilizer and insert.
Neutral Position-of Randle
29. Checking the neutral of the control handle is done (see figure
53) in the following manner: with connected hydraulic system set control
handle of the left half of the stabilizer by pulling it toward yourself
from forward position according to points 54 and 55 by the dimension "left"
0 mm, ARU should in this case be on the longer arm, the trimmer effect
mechanism should be in neutral.
To measure the position of the handle from the instrument panel
along the distance "V, the position of point "T" on the handle is fixed
by means of a special collar.
"Adjustment" of Stabilizer when SwitChing
over ARU from Greater to Smaller Arm
" 30. The check is. made with the "trimmer effect" mechanism in neutral
position with hydraulic system on (see figure 53).
By pulling the control handle toward yourself from the forward posi-
tion set left stabilizer according to points 54 and 55 by a dimension
left =-0 tlmm on the larger ARU arm.
Fix handle in this position. ,
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Switch oyer ARU from larger to smaller arm. On smaller arm of the
ARU measure the dimension "K" left; when doing the adjustment the nose
of the stabilizer should be lowered, i.e., point 55 will be lower than
point 54. The dimension D, left = 18mm (up to points 54-55) corresponds
to an angle of deflection of the stabilizer of 2?40' with nose down.
A change in the "lead-away" value of the stabilizer is utilized to
change the nature of the balancing curve.
- If at greater balancing speeds (Vis = 900 to 1000 km/hr) pulling
forces appear greater than the permissible ones by 3 to 4 kg, then the
"lead-away" should be increased (lower nose of stabilizer even more down).
When changing the lead-away of the stabilizer (change in position
of the ARU rod along the vertical) it is necessary to adjust the loading
mechanism to zero forces on the larger arm, with the stabilizer fixed
by the controltandle in accordance with the actual position "DSR." In
figure 54 is given a schematic drawing of the kinematic connection from
the stabilizer to the ARU.
A change in "lead-away" of the stabilizer along the dimension D by
?
1 mm corresponds to a change in control handle forces of 2 to 3 kg at a
speed Vins = 900 to 1000 km/hr.
Neutral Position of the Trimmer Effect Mechanism
31. A check of the neutral position is made (Figure 53) in accor-
dance with the dimension OR determinable by the dimensions D1 and D2 of
DI +
the stabilizer, as a mean value D2. Measuring the values Di and
2
D2 is done in the following manner: ARU is placed on the larger arm.
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The trimmer effect mechanism is placed in neutral position by the
signal of a flashing-on green light designating trimmer effect in neu-
tral position, by shifting the switch on the control handle toward
yourself. Before this is executed the light is put out by shifting the
switch away from yourself.
We measure the distance D left between points 54 and 55 after the
handle is turned into neutral position from position D1 and from posi-
tion D2'
The turning of the handle into neutral position is done by. 'first
tilting it into extreme position, slowly at a rate of not more than 100mm
of handle movement per 10 sec.
In this case the handle will stop at any distance from neutral posi-
tion, thanks to the friction of the control system.
The dimensions Di and D2 are placed in a formula with their own
signs: Minus if point 55 is lower than'point 54 (stabilizer nose is
down);plus if point 55 is above point 54 (stabilizer nose upwards).
When balancing the aircraft, a change in neutral position of the .
trimmer effect mechanism is permitted. The change in neutral position
of the trimmer effect mechanism is applied to assure balancing of the
aircraft in flight at instrument speed of Vins = 750 ?100 km/hr.
At a balancing rate of less than Vins = 750 ?100 km/hr it is necessary
to change the DSR by raising the nose of the stabilizer upwards (according
to points 54 and 55), and at a speed greater than Vins = 750 ?100 km/hr
in direction of lowering the stabilizer nose downward. Adjustment of the
DSR is done after the adjustment of the stabilizer lead-away.
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To a dimension of DSR equaling plus 10 mm corresponds an angle of
stabilizer deflection (perpendicular to the axes of rotation) of 1030'
nose upwards. In figure 55 is given the relationship between the dimen-
sion D and stabilizer angle' for small angles.
Balancing the aircraft by the forces at a given lead-away D 18mm
will be assured at a Vim, .1. 750 ? km/hr at an altitude H 0 to 3 km.
ARU under these conditions
will be situated in an intermediate position.
An increase in DSR (raising the stabilizer nose according to points
54 and 55 upwards) by 1 mm increases the balancing speed by approximately
10 km/hr.
The friction force and gap in the control system on the greater arm
when the trimmer effect mechanism is in neutral position is determined
by the difference between D and D2 along points 54 and 55 of the left
half of the stabilizer.
The difference between D1-D2 = 10 mm corresponds to a friction force,
applied to the handle, on the larger ARU arm equaling 0.6 kg at a 0 gap
condition. Allowances forD1 and D2 are not set up and allowance is only
made for their difference, i.e., D1-D2.
Maximum Movements of the Handle and
Maximum Deflections bf Stabilizer
32. On every aircraft are measured only the maximum movements of
the handle and maximum angles of deflection of the stabilizer on the large
and small arm of the ARU.
Measuring the movement of the handle is done by points listed in the
schematic drawing shown in figure 53 and 56 during its total deflection
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toward yourself and away from yourself by the difference in the dimension
"K-M" and "N-K," point "T"-conditional point of applying the force of the
pilot's hand. Position of point "T" on the handle when measuring the
movement ofthe handle is fixed by a special collar..
Measurements are carried out with hydraulic system connected with
normal pressure and with trimmer effect mechanism in neutral position.
The fixing and control of the ARU position is done by the indicator in
the cockpit.
The angles of stabilizer deflection are measured by an angle gauge
perpendicularly to the axes of rotation. The angle gauge is placed in
zero position Along points 19 and 20 on each half of the stabilizer.
The reading of angles foreach half is done from zero (Figure 53):
When it is necessary to change the angle of inclination from perpen-
dicular into a direction parallel to the flow it is necessary to use the
following ratio: yoin direction of flow = 0.559
to the axis of rotation.
Clearance along the Movement
in the Control System
'P-
st
perpendicular
?
33. The clearance on the handle in longitudinal direction is measured
(Figure 53) when the hydraulic system is disconnected and the loading
mechanism on the small ARU arm is off. The clearance is measured in the
chain running from the handle to the booster. The magnitude of clearance
allowance includes the gap of the booster slide-valve. The booster isfixed
automatically (thanks to a hydraulic lock). To the handle is applied a
force of 2 kg pushing the.handle forward and the movement of the handle is.
measured, when pulling the handle toward yourself we also measure the
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movement of the handle. The sum of handle movements away from yourself and
toward yourself appears to be the gap. The measurement is repeated twice
but in a different sequence.l
The mean value is inserted in a table.
B. Lateral Control
Neutral Position of Handle
34. Neutral position of the handle (i.e., at a force applied to it
equaling zero (P = 0) should be in the plane of symmetry of the aircraft.'
The hydraulic system isIon, pressure normal. The shifting of the
handle into neutral position'during operation is controlled by the dis-
tance "from left" end of cabin (from reference point 21) to point T on .
the handle, measured when the, airplane is leveled at the assembly plant
and the handle is fixed according to a balance (see figure 57).
lAileron Shears
35. The "shears" of ailerons are the sum of angles of deflection
of the left and right ailerons (in various directions) during neutral
position of the handle. When, the aircraft leaves the assembly plant the
shears of the ailerons should be zero. The shears of ailerons are used
for lateral balancing of the aircraft when carrying out the acceptance-
delivery program.
The shears of ailerons (see figure 57) are measured in the following
order: - hydraulic system "on"; '
- handle in neutral position according to the leveling dimension
,"from the left."
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It is unnecessary to fix the handle because it is held in neutral
position by the friction force and by the force of the loading mechanism.
We measure the.linear dimensions "N" according to points 38 and 38a
along the left and right ailerons.
We determine the magnitudes of the shears of ailerons by the number
of measurements of "N" (at a deflection of ailerons in various directions
The direction of shears is determined by the right aileron using terms--
"right aileron up," "right aileron down," at a condition when points 38
and 38a of the left aileron coincide.
In the process of changing the shears of the aileron for the purpose
of balancing the neutral position of the handle does not change.
When measuring the shears in the case of presence of local deviations
of the contours at the measured points, the measurements should be correc-
ted by a value of the initial stage between wing and aileron.
The shears of ailerons are used to eliminate the banking of the air-
craft during flight with boosters on. The rules of applying aileron shears
,when balancing the aircraft with hydraulic system on are:
a)' during left banking of the aircraft it is necessary, during neu-
tral position of the handle (by the dimension "form the left") by adjusting
the control cables, to deflect the left aileron downward and the right one
upwards.
b) during right banking of the aircraft left aileron up and right one
down.
The magnitude of aileron deviation along points 38-38a is evaluated
by the following ratios:/ During banking, moving the handle 5 mm from neu-
tral position (force against the handle approximately 0.3 to 0.4 kg) each
1
aileron should be deflected by 5 mm along points 38-38a.
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Maximum Movements of the. Handle and Maximum
Angles of Aileron Deflection
36. Maximum movements of the handle are checked with hydraulic system
connected. The movement of the handle is measured by a ruler from neutral
position (by the dimension "from the left") along point "T" on the handle
(see figure 57 and 58).
37. The angles of aileron deflection are measured by an angle gauge
with connected and disconnected hydraulic system perpendicularly to the
axis of aileron rotation (Figure 57).
Zero position of the angle gauge is set up for each aileron by com-
bining points 38-38a of the left and right sides (see figure 57).
Non-Linear Mechanisms
38. On each aircraft is made (figure 57) a check of the non-linearity
of the kinematic bond between the handle and the ailerons with connected
hydraulic system and one position of the handle.
To this condition corresponds a maximum coefficient of non-linearity.
Measurements are conducted simultaneously with the measurements described
in paragraph 36.
When entering into the table the angles of aileron deviations for a
given handle movement from neutral, the obtained angle of deviation should
be deducted from the value of the angle corresponding to the neutral posi-
tion (with the existance,of shears). Measurements are carried out for both
left and right ailerons for handle deflections both left and right ?50 mm.
A check of the nonlinearity of the kinematic coupling between the lever and
aileron for the entire range of handle movements is made according to the
additional instructions.
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Knife (Plate) of Left Aileron
39. The knifed the left aileron serves for transverse balancing
of the aircraft.
When the aircraft leaves the assembly plant of'a series manufactur-
ing plant and before the first flight, the angle Of bending of the knife
should be zero (see figure 57). The deflection of the knife is further
used to assure lateral balancing of the aircraft with deflected aileron
boosters.
The rule for bending the knife of the left aileron to assure lateral
balancing with engaged hydraulic system is as follows:
e) If the aircraft banks to the left it is necessary to bend the
knife of the left aileron upwards.
b) If the aircraft banks to the right the knife of the left aileron
is bent downwards.
The bending of the knife by 1 mm against the banking reduces the
control handle force by 4 to 8 kg at Vind = 1000 to 1100 km/hr.
C. Path Control
Maximum Angles of Deflection of the/Rudder
and the Movement of Pedals'
40. Angles of deflection Of the rudder (Figure 57) is measured with
an angle gauge perpendicularly to the axis of rotation. Zero position
of the angle gauge, corresponds to a coordination of points 47 on the rudder
and 27 on the fin.
Movement of the pedals is measured frim zero position of
using a ruler.
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Knife of Rudder
41. After assembling, the angle of bending of knife of the rudder should
be equal to zero (see figure 57).
The bending of the knife is used to assure path balancing of the
aircraft with respect to force.
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Control of knife bending is carried out along two sections, lower and
upper, using the following rule: if the aircraft turns to the right, the
knife of the rudder is bent to the right ( in direction of flight); during
left turn--to left side.
For the turning of the aircraft (leadout of "pellet" of EUP-6):
a) up to subsonic speeds - bend knife in lower sections;
b) atsupersonic speeds in upper section.
In the center section .the knife should not be bent.
If the aircraft turns such that the leadout of the "pellet" is one
diameter, the knife should be bent approximately 0.5 mm.
III. Measuring the Forces on the Handle and in Pedals
A. Longitudinal Control
Forces on the control Handle from
the Loading Mechanism
42. The forces acting against the handle are measured on each air-
craft when the ARU is on the small aim, when shifting toward yourself and
away from yourself but only in direct movement.
The forces (Figure 59) are measured after a control check has been
made, i.e., by the established values of Duvod and Dsr.
The measurements are made under the following conditions:
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The forces against the handle are measured with a special dynamometer
(of the "TOKAR" type).!
When the handle is shifted away from yourself or toward yourself
the axis of the dynamometer should be parallel to the axis of symmetry /
of the fuselage and perpendicular to the axis of the handle. The appli-
/
cation of dynamometer forces to the handle should be at point
The movements of the handle are measured with the aid of a special
ruler installed in the cabin on a bracket. The reading of movements is
done along the arc of the ruler with its radius of 605 mm.
The beginning of reading the handle movements is assumed to be the
position of the handle, at a zero angle of deflection of the stabilizer,
i.e., at a stabilizer position (left) ?by the dimension D = 0 (along points
54 and 55).
The measurements are made when the handle is shifted from position
D = 0 into position toward yourself and away from yourself by 5 mm less
than the maximum movements.!
Selection of handle movement less than maximum is made for the purpose
of eliminating inaccurate measurements of forces on account of handle
resistance.
The forces are measured during direct movement of the handle, i.e.,
when it moves from neutral toward yourself, then from neutral away from
yourself. In the force measurements will be included the magnitudes of /
friction forces in the control system and ?8h change in forces on account
of the friction forces of the loading mechanism.
The forces during the movements from extreme positions into neutral
(reverse movement) are not measured.
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The forces are measured only in movement, i.e., other forces are
not measured.
The measurements are carried out at neutral position of the trimmer
effect.
The mechanism of the trimmer effect is placed in neutral position
by the movement of the switch on the handle toward yourself, the signal
light is put out by the movement of the switch away from yourself. The
measurements are carried out at normal pressure of the hydraulic system.
The measurements are made at a temperature of the outer air within
limits of the +10? to +20?C.
The measurements are carried out on the aircraft with fuselage hatches
closed.
Prior to carrying out the measurements are made several movements
of the stabilizer.
Number of points to be checked during measurements:
In the range of a rigid spring of the loading mechanism during the
movement toward yourself - 6 to 8 points, away from yourself also 6 to 8
points..
In the range of a soft spring during the movement toward yourself
10 to 15 points, away from yourself - 8 to 10 points.
The ARU is set on the small arm by the cabin indicator.
The forces are measured in the following manner:
a) Deflect handle into forward position and then, due to friction,
lower it into neutral position with a speed of not more than 100 mm of
handle movement per 10 sec, so as to assure the attainment of the neutral
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position exactly. (This leads to the elimination of lags between the
handle and loading mechanism.)
b) Measure the forces along the movement of the handle when shifted
from this position into the direction away from yourself.
c) Having deflected
the handle into position toward yourself,place
it in neutral position, analogous to point "A".
d) Measure forces during movement toward yourself from this posi-
tion. The force values by the measurements should be adjusted by the /
available sample, shown in figure 60.
In the table should be, entered the maximum forces appearing during
movement toward yourself and away from yourself.
After plotting the force points on the handle along its movement
and after drawing a line along these points,
in forces in the range of
The gradient of force
AR
the gain in the movement of the handle, i.e., Rx =
AK
The gradient of force change Rx is determined for the smai (rigid)
spring;in this case the gradient is determined conditionally by the forces
at points A, F, and A2E (see figure 60).
The values of the force gradients Rx should be determined during
the movement of the handle away from yourself by: RN ...AR RE
X KE _ KA2
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check the gradient of changes
handle movement to points E-F.
changes is called the ratio of force gain to
and toward yourself by: RX
R RF
6K KF__EA1
On the graph of the leveling scheme the nominal forces are plotted
with consideration of the frictionforce and ?8% allowance for forces from
the loading mechanism. Differences in forces allow for nominal values' of
leadout and Dsr of the stabilizer.
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Range of Operation of the "Trimmer Effect" Mechanism
43. The range of the operation of the trimmer effect mechanism,
Figure 59, is checked under the following conditions: Hydraulic system
connected; ARU in two position--larger and smaller arm.
Movement of the handle is measured by a ruler in the cabin from the
position of the handle at neutral position of the trimmer effect mechanism.
To measure the range of operation of the trimmer effect mechanism
it is necessary to fix the trimmer effect mechanism in neutral position
(by the flashing of green light) using switch during movement toward
yourself; shift handle into extreme position away from yourself, applying
pressure to the trimmer effect switch; measure maximum movement of handle;
repeat measurement with handle deflected toward yourself.
B. Lateral Control
Forces Against Aileron Control Handle
44. The forces against the handle in relation to the movement of the
handle are measured only on selective aircraft.
On each aircraft are checked the nominal values of forces acting
against the handle from the loading mechanism with an allowance of ?8%
for changein forces on account of the loading mechanism and friction
forces only at extreme deflections of the handle (see figure 59).
The measurements are made With a special dynamometer (of the "TOKAR"
type).
The application,of dynamometer-forceg'should be at point "r! on the
control handle (see figure 58).
The axis of the dynamometer should be perpendicular to the axis of
the handle.
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When measuring the forces the handle should not be.moved within
4
5 mm of the extreme position to avoid inaccuracy when measuring/because
of the handle stop.
Measurements are made with connected and disconnected hydraulic
system.
Before measuring it is necessary to move the handle several times
into extreme position.
Measurements are made at an outside temperature ranging from +100
to +20?C.
The forces Are measured for a smooth deflection of dynamometer con-
trol handle to the left and right.
Measurement of forces
In the obtained force
during reverse movement is not made.
values (and in the nominal values) are included
(: friction force values of control systems and ?87. for change in forces on
1
account of the friction forces in the loading mechanism). When the.hy-
draulic system is off the forces against the handle increase sharply due
?
I
to the higher friction forces of the boosters.
I
C. Path Control
Forces Against the Pedals
45. The forces against the pedals (Figure 59) are measured by means
of a special dynamometer with pedals in extreme positions, but not nearer
than 10 mm to the point of rest during the movement of left pedal away
from yourielf and toward yourself. The force of the dynamometer is applied
at the point of conditional application of the pilot's foot force, i.e., .
at a distance of 200 mm from the axis of rotation of the pedal.
58
In Figure 61 is given a schematic drawing of pedal movement.
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Supplement to Methodical Instructions
on Leveling and Ground Balancing
for Selective Aircraft
General Data
Measurements by the "supplement" on leveling system is carried out
only on selective aircraft or in the case abnormalities in the behavior
of the aircraft are revealed in the process of carrying out the accep-
tance-delivery program (banking, unusual turns, etc.).
1
Prior to carrying out leveling by the "supplement" the aircraft
should be leveled in confprmity with methodical instructions previously
explained.
In figures 62-72 is given supplementary material necessary for the
checking of leveling data and the control systems of selective aircraft.
I. General Leveling of an'Aircraft
1.41118.
1. The fixing of lateral "V" Of the wing is checked by points 18a-8a.
When looking in plan view the bend of the wing relative to the axes Of the
main part of the fuselage is checked in accordance with points T.1 and
T.2 by dimensions "A left," and "A right."
The curving of the wing relative to the axes of the tail section of .
the fuselage is checked along points T.3 and T.18a, by dimensions "B right"
and "B left" (see figure 62). .
A nominal dimension E 22 mm corresponds to an angle of lateral V 2?.
Landing Gear
2. We measure the angle,of landing gear collapse by the dimension
E (Figure 62)..
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Fuselage
3. When checking the curving of the fuselage are measured the dis-
placements of point 3 relative to the axis of the main part of the fuselage
(the fuselage is located along points T.1 and T.2) by the dimension "L".
When viewed in the plan the displacement of the axis of the tail section
relative to the axis of the main section of the fuselage is measured by
the dimension "V" between the axis of the main part over points T.4 and
T.5 and point 6 on the tail section of the fuselage (Figure 62).
The twist of the fuselage is checked by sections. Measurement of
fuselage twist is done when it is placed over points I left and 2.1eft
along the length and over points 8a left and 8a right in width, by the
differences between the measurements on the left and right halves of the
fuselage (Figures 62 and 63).!
(:= Stabilizer
4. The mounting of the stabilizer and its curviture, relative to
the main part of the fuselage, is checked along point T.2 on the main
part of the fuselage and along point T.22 on the stabilizer by the dimen-
sion "R" (see figure 62).
Checking the lateral "V" of the left and right halves of stabilizer
is done along points 20-22 by setting the stabilizer by the dimension
"v-a." A check of the stabilizer at a height relative to the axes of
the main part is made by the dimension "a" by placing the stabilizer at
a zero angle by the difference in "b-a" between points 19 and 20 (see
figure 63). .1
To the nominal dimensione u22 ' Upo = 412 corresponds the lateral
angle Of the stabilizer "V 0."
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Flaps
5. The flap slots are measured by the dimension "n" (Figure 63)
during flap deflection at the maximum angle with connected hydraulic
system for left and right flaps.
Brake Flaps
6. Measurements of the angles of deflection of the brake flaps
(Figure 62) are made during maximum deflection of brake flaps with hy-
draulic system on: (a) Forward (left and right brake flap) by the linear
.dimensions'UShch" over points 57-57; (b) rear one by the dimension "Kshch"
over points 51-51.
To the nominal dimension Kshch = 446 corresponds an angle of deflection
of the forward brake flaps of 25?.
To a nominal dimension Kshch = 672 corresponds an angle of deflection
of the rear brake flap of 40?.
Wing Fence
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7. A check of the fence installation relative to the axes of the
main part of the fuselage (along points 4 and 5) (see figure 63) is made
by measuring the distance from the axis of the fuselage to the nose of
the fence "Gpgr" and by the distance of the nose of the fence-and the
tail of the fence "Gngr - G,gr" for the right and left fence.
r
II. Checking the Control System
A. Longitudinal Control
8. A check of the angles of stabilizer deflection. The stabilizer
angles are measured in the same way as in paragraph 32 "Methodical Instruc-
tions on Leveling for Left and Right Halves of Stabilizer at Various Posi-
tions of the ARU rod" (see figure 64).
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In figure 18 is shown a change in the stabilizer angles by the move-
ment of the handle in, relation to the instrument speed and flight altitude.
In figure 65 is shown the arm change of the ARU-3V automat in rela-
tion to the instrument speed, and flight altitude.
A change in arm, of the ARU-3V with flight altitude (altitude correc-
tion)?takes place automatically to an altitude of 10,000 in. At an altitude -
of 10,000 m the ARU-3V changes into the larger arm which at altitude of
10,000 m and over remains constant at all flight velocities.
In figure 64 are given complete characteristics of the forces acting
against the handle of the stabilizer mechanism when the ARU is in working
position and the trimmer effect mechanism is in neutral position, and the
booster is on.
The forces affecting the handle during its movement as shown in figure
19 are given without consideration of friction in the control system..
The angles of stabilizer deflection when ARU is in positions corre-
sponding to altitudes of 5 km, 7.5 and 10 kin, are checked at .a speed of
Vins
1100 km/hr (Figure 18).
When doing measurements we write into the tables of the leveling
system (Figure 64) the actual speed values, Vis, at the beginning of
ARU operation and Vins at the end of ARU operation (corresponding to EL:
displacement) and with ARU in position of small and larger arms.
For the given speeds have been established tolerances.
In an analogous manner are checked the altitudes at the beginning
of the correction after the connection of the ARU automat.
The actual altitude values at the beginning of corrections and the
end of corrections are entered into the very same table.
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In addition to angle measurement, is also measured the dimension D
(along points 54 and 55) of the nominal dimensions. The norm for the D
dimension is not given.
The instrumental speed (rated speed) and the altitude when measuring
stabilizer deflection angles is produced by RPU-3 instrument. For this
purpose one RPU-3 instrument is connected to the dynamic aperture DUAS-8M;
(RPU-3) is connected to the static aperture DUAS-8M.
The pressure in the dynamic line of the ARU sensing element controls
the indicated speed instrument (according to the wide arrow )--Vins;and the
pressure in the static line controls the altitude indicator--H5.
It is necessary to set the altimeter dial on the altitude value which
is shown by the instrument in standard atmosphere at a barometric pressure
.,uthile the aircraft is parked (this assures the realization of measurements
in accordance with standard conditions).
To do this it is necessary: (a) to tightenthe fastening clamp and
set the pressure in the little window at 760 mm Hg; (b) to consider the
, instrument errors of the speed indicator and altimeter.
When setting the necessary gear ratio, the values Vins and Hins must
- be given accurately, starting with small values of Vins and Mins.
Remark: The control gear ratio is the ratio of increase in handle
movement to'the increase in stabilizer movement, i.e.
X
The relative gear ratio Ko. of the control is the ratio of values
7
6 at given altitude and flight speed to the value-4L- at H = 0 and
y
/
Vins = 0. In Figure 60 is shown the value R0( depending
, 0,V-.0
upon the altitude indicated by flight speed le instrument.
/
Warning: at a constant Vins it is necessary to keep in mind that
exceeding a Vins of 1250/km/hr may lead to a stoppage of the ARU automat.
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B. Lateral Control
Non-Linear Aileron Mechanisms
9. A check of the non-linearity of kinematic coupling between'control
handle and ailerons is also done as described in paragraph 38, "Methodical
Information," but only in forward and reverse movement of the handle and
within 20 mm of the neutral handle position (see figure 67). On figure 67
is shown the angles of aileron deflection for movement of the handle to
the right and left with boosters connected for the entire range of handle
movement.
Play of the Control Handle
10. The play of the aileron control handle is measured with hydraulic
? system off, see figure 67. When measuring the amount of play the left
aileron should be secured in the neutral position. The handle should
then be moved sideways with a pressure of 2 kg and the position of the
handle should be noted, then the handle should be moved toward the oppo-
site side with pressure of 2 kg and this position of the handle should be
noted. The full displacement of the handle in this range is the amount
of play. Then repeat handle movement measurement again under the very
same effects but in reverse sequence.
In the square of figure 67 write in the mean value.
III. Measuring the Forces Acting against the Handle and Pedals
A. Longitudinal Control
Forces against the Control Handle
from the Charging Mechanism
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11. The forces against the handle are measured in the same way as
described in paragraph 42, "Methodical Instructions," for three positions
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of the ARU rod on the small, large, and medium arm, during forward and
reverse movements of the handle (see figure. 64).
The number of points to be measured with the fixed spring of charg-
ing mechanism for movement toward yourself is 6 to 8 and away from yourself
is 6 to 8 points; and with the flexible spring for movement toward yourself
it is 10 to 15 points, and away from yourself it is 8 to 10 points.
If, after measurement of forces at each condition of Vins and
there is a noticeable discrepancy of points, it is necessary to measure
again.
The values of force measurements are plotted for forward and reverse
movements under various conditions using the same designations as in figure
68.
When measuring the forces the handle should not be moved within 5 mm
of extreme position in order to avoid distortions due to the point of
handle rest.
The allowable force value includes an allowance of ?87. for spring
forces of the charging mechanism and for friction.
The forces during forward and reverse movements differ by the value'
of double friction force.
The control characteristic is considered satisfactory if the force
measuring points are situated within the tolerances, and in addition an
allowance for the force gradient should be maintaine'd.
Determining the Force Change Gradient on
the Handle from the Stabilizer
12. The force change gradient in accordance with movement of the
. /
handle is determined only on the small ARU arm during forward movement.
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The force change gradient according to the fixed spring is checked
on every aircraft as mentioned in paragraph 42 "Methodical Instructions."
After plotting a force dependence for the handle in accordance with
its movement and after drawing a line along these points, as is shown in
figure 69, it is necessary to check the force change gradient for the
hollow sections (according to points F--toward yourself, and E--away from
8
yourself) Rx X RE2-RE =
XE2-XE
The gradient Rx of the hollow sections is checked only for selective
____------
aircraft.
Measuring the Friction Forces
of Longitudinal Control,
13. The friction forces are determined after plotting on the graph,
figure 64, the force measurement results on the handle in accordance with
its movement from the charging mechanism in forward and reverse movements.
The friction forces are determined with the handle in neutral posi-
tion and in extreme positions on the greater and smaller ARU arms (H = Oi
Vins = 0 and V ins 1100 km/hr).
The friction force in the control system is determined from the
ratios of forces acting during forward and reverse movements, e.g., if
the force against the handle during forward movement equals 22 kg and
during reverse movement of the handle only 20 kg, then the value of the
22-20
friction force will be equal to T = --T--.= 1 kg.
In figure 68 is shown a picture of plotting the results of measure-
ments during direct and reverse movements of the handle.
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Play (According to Forces) on the
Longitudinal Control Handle
1
14. Play measurements are carried out in two positions of the ARU
on the greater and smaller/arms. The play of the stabilizer handle is
determined by figure 64,plotted in accordance with measurement results.
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The play should be cheated between the center lines of actual force
measurements regarding forward and reverse movements of the handle,(handle
"toward yourself" andlaway from yourself") at P 0 as mentioned in figure
68.
B. Lateral Control
Forces against Aileron Control Handle
?
15. The forces against the aileron handle are measured both during
right and left movements and during forward and reverse movements (Figure
67).
The number of points during measurement is 6 to 8 during left and /
right movements in the range of action of rigid spring of the charging
mechanism and 8 to 10 during left and right movement in the range of ac-
tion of flexible spring of charging mechanism. / /
In figure 70 is given a sample of plotting force measurement results
on the handle for its right and left movements. The value of the force..
should be plotted for the forward and reverse movements with various con-
ditional signs.
When measuring handle forces the handle should not be moved within
5 mm of the extreme position in order to avoid inaccuracy due to the
handle stops.
67
47'
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v R V T
The force allowance value includes allowances for the spring of the
charging mechanism equaling ?87. and the friction force. The forces during
forward and reverse movements differ by the value of double friction
force. The control characteristic is considered satisfactory, if the force
measurement points are within allowances; and in addition it is necessary
to maintain an allowance for the gradient of forces in accordance with
the rigid spring.
In figure 71 is given a theoretical dependence of the change in forces
on the handle upon its right and left movements. The forces are given
without consideration of the friction force effect in the system of aileron
control.
Measuring Friction Forces of Lateral Control
16. The friction forces are determined after plotting on the graph
(see figure 67) the force measurement results on the handle from the'charg-
ing mechanism during forward and reverse movements.
Friction forces are determined during neutral position of the handle
and with the handle in extreme positions. ' The friction force in the '
aileron control system is determined from ratios of forces, acting during
forward and reverse movements, e.g., if the force against the handle during
forward movement equals 6 kg and during reverse movement 4 kg, then the
6-4
friction force value equals T = 2 = 1 kg.
Play (in Accordance to Forces) on
the Aileron Control Handle
17. The force play on the aileron control handle is measured from
graphs (figure 67) as distances between the center lines of actual force
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measurements for the movement of the handle during forward and reverse
movements (handle "to the left" and handle "to the right") at 1) 0, as
it is shown in figure 70.
/
Forces against the Pedals
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18. The forces acting against the pedals are measured for each
aircraft (paragraph 45).
Determining the Gradient of Force Change
on the Aileron Control Handle
19. The gradients of force change on the aileron control handle are
determined along the section of the rigid spring during forward movement
of the handle to the right and left (figure 67).
/
In a similar way are determined the forces acting against the handle/.
of longitudinal control (paragraph 12).
C. Directional Control.
Non-Linear Rudder Mechanism
20. A check of the non-linearity of kinematic coupling between pedals
and the rudder is done during forward and reverse movements of the pedals
within 20 mm from neutral position (see figure 72).
In figure 72 is given the change in the angles of rudder deflection
to the right and left for the entire range of pedal movements.
Methodical Instructions on Balancing the Aircraft in Flight
General Data
1. The present methodical instructions on balancing the aircraft in
flight have been compiled in conformity with the methodical instructions
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governing the leveling and ground balancing of aircraft, described in
appendix 1 to chapter 5 of this report.
2. Methodical instructions contain no allowances. All allowances
are listed only in the leveling system.
3. The sequence of flights for aircraft balancing is given below.
I. Instructions for Flight Crew
A. Pre-Flight Control Test
Before flying for aircraft balancing the pilot should become acquainted
with the leveling system, and as he makes a pre-flight inspection of the
aircraft in conformity with instructions governing the operation he should
check: (a) arrangement of ARU, which should be on the greater arm; (b)
arrangement of trimmer effect mechanism in neutral position; (c) shears of
stabilizer (difference in setting angles of right and left halves of sta-
bilizer) according to lines 54 and 55 from the left (see figure 53) should
be 0 ?1 mm when combining lines 54 and 55 of the left stabilizer half; (d)
the correctness of the position of the rear stabilizer edges, rudder,
ailerons, and flaps. The position of the rear edges should be as follows:
The knife on the trailing edge of stabilizer tilted upward by 40; knives
of trailing edges of rudder and ailerons should be: before the first flight
during acceptance-delivery tests at the series manufacturing plant in zero
position. In subsequent flights they are controlled in conformity with
the flight results for balancing.
B. Longitudinal Balancing of Aircraft in Flight
Ihe purpose of this task.is to balance the aircraft with the trimmer
effect mechanism during climb to an altitude, pf:3 000 m (in rectilinear
climb without overloads) during maximum operational condition of the engine
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by instrument speed showing 750 t100 km/hr (by the wide arrow) with re-
tracted landing gear, flaps, brake flaps, and without suspension tank.
Make sure that the balancing mechanism of the trimmer effect is set
in the neutral position according to the signal lamp prior to takeoff.
The order of carrying out the balancing--the balancing should be
carried out in rectilinear flight after reaching an instrument speed of
750 ?100 km/hr, and then not using the trimmer effect mechanism speeding .
up to an instrument speed of 1000 km/hr he must check the nature of
changes in forces affecting the handle.
At speeds greater than the balancing speed (Vis = 900 to 1000 km/hr)
pulling forces against the control handle of up to 3 to 4 kg are permitted.
In figure 20a is given a typical graph of a balancing curve -P = f(V)) i.e.,
a change in handle forces in relation to the speed of flight according to
instruments when climbing to an altitude from H = 0 to H = 3,000 m.
Remarks: If to secure longitudinal balancing of the aircraft at an
instrument speed of 750 t100 km/hr the pilot has to utilize the trimmer
effect mechanism and the illumination of the signal light has not corre-
sponded with the balancing speed, then the pilot should determine the
balancing speed at which the signal light does go on.
If after the balancing flight it is necessary to adjust the trimmer
effect mechanism or to adjust the "lead-out" of the stabilizer, then it
is necessary to make a control check of the balancing i9' the next flight.
C. Lateral Balancing of Aircraft in Flight
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When doing lateral balancing evaluate the lateral balancing of the
aircraft With aileron boosters on in accordance with an instrument speed
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showing 1000 ?50 km/hr. If for the purpose of balancing aircraft banking
the movement of the handle, in this case does not exceed 1/4 of the total
movement of the handle (which corresponds to an aileron deflection angle
of 2.5?) then it is possible to begin checking the lateral balancing of
the aircraft.
Remark: If to balance banking we need more than 1/4 of handle move-1
ment, then after the flight it is necessary to adjust the "shears" of the:
ailerons.
The purpose of the miss
ion is to assure lateral balancing of the
aircraft over the entire range of speed operations.
Balancing is assured when the boosters of ailerons (adjusters of
shears) are in "on" position, and when aileron boosters are "off",/by
tilting the knife on the trailing edge of the left aileron.
Remark: When checking
lateral balancing the aircraft should fly
without slip. If in heading ratio the aircraft isEtill not balanced, it
is allowed to eliminate slip by tilting the rudder pedals.
Balancing when Aileron Boosters are Off
The purpose of the mission is to assure balancing of the aircraft
up. to_an instrument speed of 1000 ?50 km/hr at an altitude of 2000 to
2,500 m. The order of carrying out the balancing at an altitude of 4000
to 5000 in at instrument speed of 600 km/hr cut off the aileron boosters
and then in slanting descent from that altitude to a height of 2000 63
1
2,500 m at maximum engine operation accelerate to an instrument speed of
1000 ?50 km/hr, then decelerate to an instrument speed of 750 km/hr after
which the aileron boosters are connected again.
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Up to an instrument speed of 800 km/hr is permitted a slight banking,
balanced by the deflection of the handle with a force of not more than
4 kg, and as the speed increases these forces can be smoothly increased
to change the sign but at the end of the acceleration to an instrument
speed of 1000 ?50 km/hr (Mach number ,.. 0.89 to 0.92) should not accede
15 kg.
In the event the forces against the control handle reach a value of
15 to 20 kg at lower speeds, it is necessary to discontinue the accelera-
tion, to decelerate and again cut in the aileron boosters.
Fix the speed, flight altitude, direction of bank, approximate value
of the force against the handle and handle movement, necessary to elimi-
nate bank.
Remark: If after the flight it was necessary to unbend the knife
on the aileron, then in the next flight it is necessary to check the-
lateral balancing with aileron boosters shut off.
Balancing with Aileron Boosters On
The puspose of the mission is to check the balancing of the aircraft
in acceleration to a maximum permissible Mach number at an altitude of
13)000 in.
Special attention should be devoted to lateral balancing during
acceleration to a maximum permissible Mach number. In case the aircraft
is not balanced at any given condition, the pilot should fix the instru-
mentspeed or Mach number, flight altitude, direction of banking, approxi-
mate force values and movements of the handle, necessary to eliminate
bank.
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At near-sonic Mach numbers (0.89 to 0.92) and up to maximum permissi-
ble Mach numbers (or at an instrument speed of more than 1000 km/hr and up
to Vins maximum) is permitted a smooth increase in control handle deflec-
tion of the aileron handle in order to eliminate bank up to 1/4 of the
movement.
Remark: If after the flight it was necessary to adjust shears of
_ -
the ailerons, then in the next flight it is necessary to check the lateral
balancing with aileron boosters on.
D. Directional Balancing of Aircraft in Flight
The purpose of the mission is to assure directional balancing of the
aircraft over the entire range of operational speeds.
The balancing is assured by bending the knife of the rudder.
For directional balancing it is necessary to pay attention to the
behavior of the aircraft when flying Up to maximum permissible instrument
speeds and maximum permissible Mach number.
In case of non-balancing of the aircraft at any given flight condition,
the pilot should fix instrument speed or Mach number, flight altitude,
direction and magnitude of deflection of the pellet of the EUP-56 indica-
tor with free pedals and the
to eliminate turn.
At an instrument flight
the wide arrow) and at a Mach
approximate force value on the pedals necessary
speed of more than 1000 km/hr,(according to
number of more than 0.89 to 0.92 is permitted
a smooth lead-out of the pellet of the EUP-56 indicator by ?1 diameter with
free pedals in rectilinear flight without overload.
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For a lead-out of the pellet of up to ?1,0 diameter the pilot should
have the ability to balance_the,turn by rudder pedal deflections operating
under acceptable forces.- Sharp spontaneous aircraft turns are not per-
mitted.- The accuracy in determining the position of the pellet in various
flights is 0.5 diameter.
Remark: If after the flight balancing was done by bending the knife
on the rudder, the next flight it is necessary to.check directional balan-
cing..
E. Filling out the Leveling System
The pilot together with the LIS engineer should record in the flight
log book the speed values of longitudinal and lateral balancing.
Remark: After adjusting balancing data the pilot, prior to the next
flight, should be acquainted with the introduced changes in control data.
II. Instructions for the Technical Crew
Post-Flight Control Adjustment
Post flight aircraft control adjustment is done on the ground pro-
vided the adjustment data applied before the first flight have not provided.
the necessary conditions for balancing in conformity with chapter "A" of
the present methodical instructions.
A. Longitudinal Control
If the pilot in the flight has not used the trimmer effect mechanism
and the balancing was found to be proper, then the test of the balancing
is considered complete; and if he did use it, then after landing it is?
necessary to adjust the neutral position of the trimmer effect mechanism.
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To do this it is necessary to adjust the flashing of the signal light
during neutral position of the trimmer effect mechanism, fixed by the pilot
when flying for balancing.
If the pilot reports that at speeds greater than the balancing speed
(Vins = 900 to 1000 km/hr) pulling forces appear of more than 4 kg, then
it is necessary to increase the lead-out of the stabilizer (to increase the
lowering of stabilizer nose). The lead-out of the stabilizer is adjusted
at Der = 0 (along points 54 and 55).
Adjustment of the lead-out of the stabilizer is done by dimensioned
change of the ARU by changing the length of the cable running to the ARU
-
from the pilot and the length of cables running from the ARU to the booster.
To increase the lead-out of the stabilizer (to lower stabilizer nose)
it is necessary to tilt the ARU mechanism in clockwise direction (as viewed
(17 ? from the left side). The lead-out of the stabilizer from the greater arm
of the ARU to the smaller one must be checked with stabilizer control
handle arrested and away from Der = 0 (combining points 54 and 55 of the
left stabilizer half).
The adjustment is made in accordance with methodical instructions
listed in appendix 1 to chapter 5, paragraph 30.
During ground adjustment of longitudinal control it is necessary to
keep in mind the dimension Der 10 mm (stabilizer nose along points 54 and
55 tilted upwards) corresponds with the angle of stabilizer deflection
(perpendicular to the axis of rotation) 1?30 nose upwards (see figure 55).
The increase (lifting stabilizer nose upwards) of Der (along points
54 and 55) by 1 mm raises the
76
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An increase (lowering stabilizer nose) (or deduction) of the lead-out
of the stabilizer when switching over the ARU from larger arm to the smaller
one by 1 mm decreases (increases) the traction (compression) forces at
speeds Vins = 900 to 1000 km/hr by 2 to 3 kg.
Remark: The knives on the trailing edge of the stabilizer are not
a means for longitudinal balancing of the aircraft, but they serve to
assure minimum hinge moments of the stabilizer.
The unbending of the knives of the stabilizer is prohibited.
B. Lateral Control
For Aileron Booster Deflections
Lateral balancing of the aircraft with aileron boosters deflected
is assured by unbending the knife on the left aileron. If according to
the pilot's report during flight with deflected boosters the forces against
the handle necessary for balancing aircraft bank during acceleration to
Vins 1000 ?50 km/hr and the deceleration to Vim, = 750 km/hr, were more
than 15 kg, then it is necessary to unbend the knife of the left aileron.
For left bank the knife of left aileron is tilted upwards.
For right bank the .knife of left aileron is tilted downwards.
Adjustment is donw in conformity with paragraph 35 of the leveling
scheme.
The bending of the knife of the left aileron in opposite direction
- of the bank by 1 mm removes from the control handle a force equaling 4 to
8 kg at Vins = 1000 to 1100 km/hr.
With Aileron Boosters Turned On
Lateral balancing with connected boosters is assured by shears of
ailerons. If according to pilot's report when flying with connected
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aileron boosters a lateral unbalancing of the aircraft is observed, then
1
it is necessary to adjust the shears of the ailerons.
During left bank it is necessary (at neutral position of handle), by
adjusting the control cables, \to deflect the left aileron downwards and
the right one upwards.
At a right bank--left aileron up and right down.
In the process of changing the shears of the ailerons the neutral
position of the handle should not be changed.
The adjustment should be
carried out in conformity with methodical
instructions governing the leveling and balancing.
The magnitude of the necessary deviation of shears of ailerons along
points 38 and 38a is evaluated, by the following ratio: during banking
for equal 5 mm movements of the handle from neutral (the force against
the handle in this case equals approximately 0.4 kg) each aileron should
be deflected along points 38 and 38a by ?5 mm.
C. Path Control
Path unbalancing of the aircraft is eliminated by tilting the knife
on the trailing edge of the rudder.
When tilting the knife it is necessary to use the following rule:
For right turn of the aircraft in flight the knife of the rudder should
be tilted to the right (in direction of flight); for left it should be
to the left. When the aircraft is making a turn at subsonic speeds the
knife of the rudder is tilted in the lower section; at supersonic speeds
it is in upper section.
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In the center section the knife is not tilted. If when making a
turn we evaluate the lead-out of the pellet to be 1 diameter, the knife
should be bent approximately 0.5 mm.
Adjustment is made in conformity with methodical instructions
governing leveling and balancing.
Filling out the Leveling Documents by the Mechanic
The aircriftzathaAic,_after-completing ground Adjustments concern-
ing the control, to assure the balancing of theAircraft, has to fill out
the leveling -diagram..,'
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