https://ntrs.nasa.gov/search.jsp?R=19730001300 2020-04-28T23:25:46+00:00Z
_S
THE SIMULATION OF A JUMBO JET
TRANSPORT AIRCRAFT
VOLUME II: MODELING DATA
D6 -30643
Prepared
C. Rodney Hanke
by
and Donald
R. Nordwall
THE BOEING C_PANY
Wichita
Division Wichita,
September
Kansas
1970
for
NATIONAL
Ames Research
AERONAUTICS
Center
AND SPACE ADMINISTRATION
Moffett
Field,
California
'j
PREFACE
This report summarizes all work conducted by The Boeing
Company
under
Task II of Contract
for the Simulation
Aeronautics
of Advanced
Hanke
Flying
Company
Project
Qualities
dynamics
Leader
Division
Organization.
in Everett,
The National
Monitor
Sciences
Division.
was Mr. C. Roduey
Stability,
Technical
by Mr. Robert A Curnutt
Staff
"Design
Technical
of the Simulation
of the Wichita
was provided
Aircraft".
and Space Administration
was John Dusterberry
The Boeing
NAS2-552_,
Control
and
assistance
of the 7_7 Aero-
Washington.
D6-30643
__'fJ_'_
REV LT'R:
E-3033
R|
ISECT
IN O. Vo i. II
IPAGE
i
TABLE
OF CONT_TTS
SECTION
INTRODUCTION
1.0
AIRPLAI_E
FLIGHT
DESCRIPTION
CO_fROLS
DISCUSSION
OF
1.!
SYST_IS
THE
1.2
SIMULATION
1.3
LIFT
FORCE
COEFFICIET{T
2.0
DRAG
FORCE
COEFFICIENT
3.0
PITCHING
MOMI2_T
ROLLING
YAWING
SIDE
MOM_ENT
MOMENT
FORCE
LANDING
SD_UI_TION
APPENDIX
5.0
COEFFICIENT
CONTROL
LIFT
PROPULSION
COEFFICIENT
6.0
7.0
CONTROL
DIRECTIONAL
HIGH
4.0
COEFFICIENT
LONGITUDINAL
LATERAL
COEFFICIENT
SYSTEM
8.0
SYSTEM
CONTROL
9.0
SYSTEM
10.O
SYSTEM
ll.O
SYSTEM
GEAR
12.0
SYSTEM
z3.o
z4.o
Ct-]ECKOUT
B - BUFFET
CHARACTERISTICS
16.0
NO.
REV LT'R:
E-3033
RI
I SECT
Vol.
II
I PAGE _-3o6_3
_ ;_J.
SECTIO[_
APPenDIX
C - AUTOTHROTTLE
APPk_,FDIX E - P_-_SED
APPENDIX
REV LT'R:
E-3033
R_
F - AIRPLANE
SIMUI_T!0N
RESPONSE
17.0
DATA
TO CONTROL
19.0
INIr0TS
ISECT
20.0
NO.
PAGE
_V
!.0
P_
INTROIECTI_[
m
General
Arra_4_ement
Summary
of
Areas
Limits
and Placards
Center
of
Moments
Dimensions
1.1-3
i.I-_
Limits
1.1-5
of Inertia
1.1-6,
1.3-_
Equations
Sign
9
1.2-7
System
Systems
Flight
1.3-5
1.3-6
Convention
Maximum
Control
Maximum
Values
2.0
and
Gravity
Hydraulic
Axes
1.1-2
Surface
Deflections
and Rates
1.3-7
1.3-8
LIFT FORCE _IEFFICIENT
Effect
of
Angle
of
Attack
on Basic
CL
Effect
of
Angle
of
Attack
and
Number
Effect
of
Flaps
on
Aeroelastic
EfTect
(ACL)=w.a
on
Mach
2.0-7
on Basic
2.0-8
CL.
2.0-9
R . Oe
(ACL_=..ap..
2.0-10
0o
Effect
of napso,. A(,Jc./d=)
Aeroelastic
on
Effect
A(dCL/_
2.0-11
2.0-12
)
Effect
of
2.0-13
Effect
of
2.0-1_
Effect
of
Flaps
Effect
of
Normal
Effect
of Stabilizer
on
dCa/dn
Load
2.0-_
=
Factar
2.0-16
O
!
I
REV 5YM
D
[PAOE"
_4-
_.>.
@.?000
%
2.0
J
LIFT
FORCE
(Cont'd)
PAGE
m
Effect
of
Inboard
Effect
of
Outboard
Effectiveness
Effect
of Spoilers
Effect
of
Mach
(6 aud
Number
Effect
7)
2.0-22
on Spoilers
2.0-23
on Lift
Force
or 5,
Aeroelastic
Effect
on
Lift
to
Spoilers
Aeroelastic
Effect
(9,
on
lO.
Lift
to
Spoilers
(12
of
Outboard
Aileron
Landing
REV
2.0-21
(8
Effect
2.0-20
Spoilers
Spoilers
Due
•
- Spoilers
to
Due
2.o-19
Elevators
Factor
of
Due
2.o-18
Elevators
Effect
Aeroelastic
¢
COEFFICIENT
Gear
or
Coefficient
7 and
6)
2.0-24
Force
Coefficient
II or
2,
Force
Coefficient
3,
2.0-25
4)
i)
2.0-26
2.0-27
Effectiveness
2.0--28
Factor
Effect
of
Landing
Gear
Effect
of
Mach
Ground
Effect
Ground
Effect
Effect
of
Flaps
on
2.0-33
Effect
of
}'laps
on
2.o-:_
Number
Height
2.0-29
on
ACtS_A
Factor.
of Attack
for
Stick
A_le
of
for
Initial
SYM
E
2.o-)I -
K_Z
2.0-32
An_le
Attack
2.o-_,o
M
Shaker
Actuation
2.o-35
2.0-36
Buffet
a_a_N_
NO;Vol.
I! '
.;.
'-
.::;-
2.0 L_
FOR_
COEFFICIENT
PAGE
(Cont'd)
n
Certification
Buffet
Stall Speeds
Boundary
u:
2.O-38
_LMAx.
and
3.0 DRAG FORCE
2.0-37
COEFFICIENT
3.0-5
Effect
of Angle of Attack
on Basic
CD
Effect
of Angle
of Attack
on Stabilizer
(Flaps Up, i,
Effect
of Angle of Attack
on Stabilizer
(Flaps 20, 25, 30)
Effect
of Math
Number,
M
_
Effect
of Mach Number,
M
Effect
of Spol]ers
Effect
of Angle
Effect
of Spoilers
Effect
of Mach
Gro_md
Effect
Effect
of Landing
Effect
of Mach Number
Ground
Effect
5, i0)
3.0-6
3.0-7
.7
3.o-8
.7
3.o-9
_
3.O-lO
of Attack
3.o-n
on S_oilers
(6 and 7)
Number
3.O-le
3.o-13
on Spoilers
Lateral
Control
Factor
, FD
3.0-I_
Gear
Height
3.0-15
on _CD___
3.0-16
M
Fr_ctor, K_E
3.0-17
Ground Effect
3.0-18
Effect
of Sideslip
Effect
of
Flaps
Effect
o_
Flaps on
on
and Rudder
3.o-19
3.0-20
F(_Co)_W._RmO'IFLAPFAILURE
3.0-21
A _dCo/d_
_n
C
.,L
% %.'_
REV S'fM
E
_i'O_IN_
NO Vol.
I!
o6-3o6
-AG,,
_iL-
3
4.0 PITCHING
MOMENT
PAGE
OOEFFICIE_T
Effect
of Angle
of Attack
on Basic
Cm_5(Flaps
Up, i, 5, I0)
Z,.o..8
Effect
of Angle
of Attack
on Basic
CmA_(Flaps
20, 25, 30)
_.0-9
Effect
of Angle
of Attack
and Mach
Effect
of Flaps
on
Aeroelastic
Effect
Number
Cm. _
(ACre _)_R
_.o-1o
11.o-.ll
• O"
(AC_._)_p.
on
on Basic
= Oe
z,.o-_
z,.o..zS
Aeroelastic
Effect
,,(dCm,,/d,()
on
z,.o-zp
Effect
of
O_
Effect
of
Effect
of Flaps
Effect
of Normal
z,.o-_
Effectiveness
on
.o-17
dCm.zs/dnz
.o-18
Load Factor
Factor
- Stabilizer
Aeroelastic
Effect
on Stabilizer
Aeroelastic
Effect
on Inboard
Aeroelastic
Effect
on Outboard
Effectiveness
Factor
Effect
of Spoilers
Effect
of
Spoilers
Effect
of
Mach
Aeroelastic
Aeroelastic
Due to
Aeroelastic
Effect
Spoilers
Effect
Elevator
Effectiveness
Elevator
Effectiveness
_.o-2o
1_.o-21
_.o-22
_.o-2k
Number
Due _o Spoilers
Effectiveness
_.o-z9
- Spoilers
(6
Effect
and Elevators
and
on
or
5,
10,
Moment
7 and
on Pitching
(9,
_.0-26
Spoilers
on Pitching
(8
_.O-2p
7)
11 or
on Pitching
Coefficient
_.o.27
6)
Moment
2,
3,
Moment
Coefficient
_)
k.o-28
Coefficient
O
!
Due to
Spoilers
(12
0:' 1)
.o-29
O
e[
D6-30___M3
REV SYM
NO. :Vol. II
PaGZ V|_L
6-?000
•4.0
PIT_TING
MC_',_
Effect
of
I_oard
Effect
of
Outboard
Ground
Effect
Effect
of
Effect
of Mach
Ground
Effect
Effect
of
COEFFICIE_
PAGE
(Cont'd)
m
4.o-30
Aileron
4.o-31
Aileron
Lateral
Landing
Control
Factor,
4 .o-32
Fm
4.o-33
Gear
_.o-34
Number
on
ACm.z56_
a
4.o-35
Sideslip
and
4.0-36
Rudder
_
D/fect
of
F:aps
5 ..0_ROLLLNG
MOMENT
Effect
of
Effect
of Mach
Sideslip
Effect
Effect
of Roll
Aeroe!astlc
toRoll
to
5- 0-7
dCi/d
d Cl/d
5.0-8
_
5.0-9
_
dCi/d
Sideslip
Factor
5.0-lO
_
5.O-ll
_ FI_
5.0-L9
Rate
on Ro_llng
Moment
Coefficient
on
Moment
Coefficient
Rate
Yaw
Rate
Effect
Yaw
on
on
Effect
Aeroelastic
Due
on
Effect
Effect
of
4. o-38
,A,L,,Z
COEFFICIENT
Number
Ground
Effect
(dCm._/d_)]FLA,
Sideslip
Aeroelastic
Due
[A
on
4.0-37
Rolling
Rate
f
Effectiveness
CJ
Factor
- Spoilers
us
C)
REV SYM
E
_#A#'_
No. Vol.
P_GE iX
II
5-0
ROLLING
MOMENT
(Cont'd)
COEFFICIENT
PAGE
era.rammeD
Effect
of Spellers
Effect
of _ach
Aeroelastic
Due
to
Effect
to
Due
to
of
Inboard
Effect
of
Mach
Aeroelastic
Effect
of
Due
Effect
of
Aeroelastic
t-0
Effect
E Moment
3,
4)
_.O_0
Coefficient
1)
5.O_I
Ailerons
5.O-22
5.o 3
on
Inboard
Rollin
Ailerons
E Moment
Coefficient
- Outboard
Ailerons
5.o 6
Aileron
on
Rolling
5.o 7
Moment
Coefficient
5.o 8
Ailerons
Lateral
Control
Factor,
5.o 9
F_
Rudders
5.0-30
Effect
Aeroelastic
Due
Factor
Outboard
Effect
2,
Coefficient
Ailerons
Outboard
Ground
DUe
on
Effect
to
11 or
Rollin
or
Moment
- Xnboard
Number
Inboard
Aeroelastic
5.o-19
Aileron
Effect
Effectiveness
Coefficient
5)
lO0
(12
Factor
Effect
to
on
Spoilers
Moment
Rolling
(9,
Effect
Effectiveness
Due
on
Spoilers
Spoilers
Rolling
(8 or
Effect
Aeroelastic
on
on
Spoilers
Aeroelastic
Due
Number
Upper
on
Rolling
Moment
Coefficient
5.0 r
Rudder
Effect
on
to
Lower
of
Asymmetric
Rolling
Moment
Coefficient
Rudder
5.oa2
Inboard
Flap
Failure
For
a
Flap
REV $YM
D
Extension
5.0_
.o. Vol.
II
I PAGI[ D6-306__3
X
6-?OOO
_.0 ROLLING
Effect
of
Flap
MOMENT
COEFFICIENT
Asymmetric
Inboard
of
Flap
Effect
Effect
For
Outboard
Failure
For
5.O-3_
Asymmetric
Outboard
Flap
Failure
For
5.o-36
Retraction
Asymmetric
L.E.
I, 2, 3, _, 5 or
Effect
Flap
Extension
of
7,
8 or
19,
5.o-_
23, 2_, 25, 26
5.o-_
E1
COEFFICIENT
6.0-6
Effect
of
Ground
Effect
Effect
of
Effect
of
Roll
Effect
of
Taw Rate
Sideslip
Sideslip
Factor,
6.0-7
FN_
6.0-8
Effectiveness
Rate
6.0-9
6.o-1o
Factor
Effect
of
Spoilers
Effect
of
Mach
Ground
Effect
Effect
of
Inboard
Effect
of
Outboard
Effectiveness
Segments
L.E. Flap Segments
20,
MOMENT
Flap
22,
of Asymmetric
6.0 YAWING
t
Failure
5.O-3_
Asynnnetric
of
Flap
REV SYM
Flap
Retraction
Effect
6,
PAGE
m
(Cont'd)
6.O-ll
- Spoiler8
6.o-Ia
Number
on
Lateral
6.o-i_
Spoilers
Control
Factor,
6.O-I_
FN
Aileron
Factor
6.o-15
Aileron
-
Upper
6.O-16
Rudder
6.O-17
BOfJNO
j
D6-306__3
.o. Vol. II
PAG,. X A
->I-?O00
,6.0 YAWING
MOMENT
COEFFICIENT
PAGE
(Cont 'd)
u
Effectiveness
Effect
Factor
of
Effect
to
Upper
Aeroelastic
Due
Effect
Effect
of
Asymmetric
6.0-20
Yawing
Coefficient
Rudder
6.O-21
Inboard
Flap
Failure
For
6.0-22
Asymmetric
Inboard
Flap
Failure
For
6.O-e_
Asymmetric
Outboard
Flap
Failure
For
6.0-24
of
Asymmetric
Outboard
Flap
Failure
For
Retraction
of
6.0-25
Asymmetric
L.E.
1, 2, 3, 4, 5 or 2,
of
Asymmetric
6, 7, 8 or
%
Moment
Extension
Flap
7.0
Coefficient
Retraction
Flap
Effect
Moment
Extension
of
Effect
Yawing
on
Lower
Flap
Effect
Rudder
Rudder
to
of
Effect
on
Effect
Flap
Lower
Rudder8
Aeroelastic
Due
-
SIDE
of
Ground
Effect
Effect
of
Roll
Effect
of
Yaw
Segments
23, 24, 25, 26
L.E.
Flap
6.0-26
Segments
6.o-2?
19, 20, ?-I
FORCE
Effect
Flap
COEFFICIIRT
7.o-5
Sideslip
Sideslip
Factor,
7.0-6
Fy_
Rate
7.0-7
7.0-8
Rate
i
REV SYM
0
_P'O,L_"JN_
' J.o. Vol.
II
IP,,G, xD,
-.>@-?ooo
7.o s_
F_cE
Aeroelastic
PAGE
COmrF!CU_T
Effect
on Side Force
Coefficient
Due to Taw Rate
Effectiveness
7.0-9
Factor
7.o-11
Effect
of Spoilers
Effect
of Mach Number
Effect
of Rudders
Aeroelastic
Effect
7.o-Io
- Spoilers
on Spoilers
7.0-12
7.0-13
on Side Force
Coefficient
7.o-m_
Due to Upper Rudder
Aeroelastic
Effect
on Side Force
Coefficient
Due to Lower Rudder
Effect
Flap
Effect
of Asymmetric
Effect
Inboard
Flap Failure
For
7.o-16
Extension
Of Asymmetric
Flap
7.o-15
Inboard
_
Failure
l_r
7.o-17
Retraction
of
Asymmetric
Outboard
Flap
Failure
l_r
7.o-18
Flap Extensiou
Effect
of Asymmetric
Outboard
Flap
Failure
For
7.o-19
Flap Retraction
Effect
1, 2,
Effect
of Asymmetric
L.E. Flap Se_nent_
3, 1_, 5 or 22,
of Asymmetric
6, 7, 8 or 19, 20,
23,
2_,
25,
26
7.0-20
L.E. Flap "Se_nents
21
7.0-21
O
D
O
REV SYM
D
BO_='JAS'O
D6-306_3
,o. Vol. II
PA_E
xlli
9@-7000
8.0
LONGI_IDINAL
Elevator
and Stick Force
Forward
Feel
Quadrant
Unit
SYSTEM
System
PAGE
8.l-3
Schematic
8.l-_
Travel
8.l-5
Torque
Mechanical
8.l-6
Advantage
Column
Force
Gradient
at qc = 0
8.z-7
Column
Force
Gradient
qc Limit
8.1-8
Inboard
Elevator
Outboard
Inboard
Elevator
Force
8.2_3
Blowdown
8.2-4
8.2-5
Float Angles
Trim Rate
LATERAL
Lateral
8.2-2
Float Angles
Elevator
Stabilizer
9.0
Blowdown
Elevator
Outboard
CONTROL
Control
8.3-2
SYST_4
System
Schematic
9.1-2a
9.l-3
Due to Spring and Cam Mechanism
Lateral
Trim
9.l-4
Aileron
- Wheel Program
9.2-3
Intermediate
Flight
Spoilers
Inboard
Outboard
Aileron
9.2-4
Program
9-2-5
- Wheel Program
Program
at Combined
Lateral
- Speed Brake Program
Aileron
Aileron
Blowdown
Blowdown
Control
- Speed
Brakes
9.2-6
9.2-7
9.3-2
9.3-3
NO. Vol. II
iPAGED6-3o643
xi,J
REV LTR:
R
Outboard
Spoiler
Spoiler
E-3033
CONTROL
I
9.0
LATERAL
CONTROL
SYSTEM
(Cont'd)
Spoiler
Blowdown
(Panel
Spoiler
Blowdown
(Panels
2,
3,
Spoiler
Blowdown
(Panels
5,
6 or
Inboard
Aileron
Outboard
10.0
Float
Aileron
Force
Due
to
4 or 9,
10,
ll)
9.3-5
7, 8)
9.3-6
9.3-7
and
9.3-8
SYSTEM
System
Spring
9.3-4
Angles
CONTROL
Control
12)
Angles
Float
DIRECTIONAL
Directional
1 or
PAGE
Schematic
Cam
lO. 1-2a
Mechanism
i0.1-3
Rudder
Trim
Rudder
Pedal
Rudder
Travel
Rudder
Hinge
Moments
- Segments
Deflected
Together
10.2-2
Rudder
Hinge
Moments
- Segments
Deflected
Separately
i0.2-3
Rudder
Hinge
Moments
- Effect
747
Yaw
lO.l-4
Motion
10.i-5
Limits
Damper
lO.1-6
Transfer
II.0
HIGH
LIFT
High
Lift
System
Flap
Screw
12.0
PROPULSION
of
Mach
Function
and
Number
Block
1o.2-4
Diagram
lO.3-2
SYSTEM
Schematic
Travel
SYST_4
Engine
Pressure
Thrust
Lever
EPR
Due
EPR
Correction
Ratio
and
to Power
Simulation
Power
Lever
Lever
Angle
Due
to
Speed
Schematic
Lq.l-2
Angle
].2.1-3
12.1-4
and
Altitude
le.l-5
NO. Vol.
REV LTR:
_-3033
RI
±s.
iPAGED6-3_._3
xv
12.0 PROPULSION
SYSTEM
Engine
Transient
Engine
Thrust
(Cont'd)
Characteristics
Simulation
12.1-6 - -17
Schematic
].2.2-2
Forward
Net Thrust
12.2-3, -4
Reverse
Net Thrust
12.2-5 - -8
Engine
Gage Display
Exhaust
Schematic
12.3-2, -3
Gas Temperature
Low Speed Compressor
Windmilling
Drag
Aerodynamic
Effects
12.3-5
Ratio Speed
12.3-6
12.4-2
Due to Thrust
Reversers
12.5-2
13.0 LANDING GEAR
Strut
Compression
Strut Force
Schematic
13.0-2
Schematic
13.0-3
Wheel
Side Force
Schematic
13.0-4
Wheel
Drag Force
Schematic
13.0-5
0leo Damping
13.0-6
Landing
13.0-7
Gear Air Curve
I SECT
IPAGE
_vi
V
REFERENCES :
1.
Boeing Document D6-20423 "Aerodynamic Data for the 747 Flight
Simulator", Revision D, _rch
ll, 1970.
2.
Boeing Document D6-30833-I "Flight Performance
Tests for
Validation of the 7_7 Flight Simulator",
June 3, 1970.
e
Boeing Document D6-13302 "JTgD-3 Turbofan Engine Operating
Parameters and Performance Data for 7_7 Airplane Flight
Simulator Design", Revision F, September 9, 1969.
4.
Boeing Document D6-13703
December 30, 1969.
5-
Boeing Document D6-30643 "The Simulation of a Large Jet Transport
Aircraft Vol. l: Mathematical Model", 30 September 1970.
6.
Boeing
1968.
Document
D6-13645,
"Boeing
747 Flight
Manual",
"747 IEFCS System Description",
April
93
-2
REV LT'R:
E-3033
R I
ISECT
IT
x_Ji
I3_
PAGE
Vol.
_
I
NO.
i. 0
INTRODUCTION
The Boeing
matical
models
of the Boeing
Company
747 on the NASA Flight
Simulator
The work performed
2.
Generalized
3-
The form
4.
Nomenclature
the Government
document
systems,
istics
Sections
propulsion
ll, 12
of the simulation
REV LT_:
E'3033
R T
simulator
design.
and a short
(FSAA).
Volume
data.
I in-
in the simulation.
data are to be
chooses
envelope,
occurs
section
a general
of the simulation.
the airplane
rolling
sections
to treat
moment,
describe
The final
contains
description
character-
ya_Ning moment,
the control
and side
character-
modes.
of the high
section
first.
The follow-
aerodynamic
operating
the characteristics
the
of the 747 aerodynamic
This introductory
roll and yaw in the various
gear.
These
Company
as a summary
discussion
present
The next three
and landing
used
15, 1971, whichever
its flight
moment,
and 13 describe
system
rights
September
has been prepared
drag, pitching
for pitch,
Aircraft
to NASA.
until the Boeing
of the document
coefficients.
and approximations
only limited
of the 747 including
lift,
For Advanced
for the report.
or until
data for use in flight
force
used
within
ing six sections
and characteristics
into two volumes.
of the data furnished
II contains
of the control
qualities
mathe-
under the contract.
equations
Volume
a description
is divided
Center with
of:
data asnon-proprietary
istics:
report
i.
This
Research
the flying
a description
retained
NASA-Ames
and data to simulate
The contractual
cludes
provided
contains
llft system,
the results
checkout.
SECT
i__JAIr_
I PAGE
NO. V_o_.3_
1.0-1
"
"
The appendices
characteristics,
contain
autothrottle,
a portion
autopilot
of the 7_7 Flight
and revised
simulation
__='J,_l#_
REV LTR:
E'3033
R I
I SECT
Manual,
buffet
data.
INO.
IPAGE
D6-_O64_
1.o-]a
i. i
AIRPLANE
The
DESCRIPTION
Boeing
designed
the
to
flaps
and
flaps
landing
the
side
with
_obtained
lateral
between
which
panels
the
on
drawing
showing
1.1-2.
A
wing
also
down
from
1.1-3.
only
two
four-wheel
trucks
on
and
are
the
of
and
each
spoiler
segments.
pertinent
and
dimensions
a_rplane
operating
inboard
The
aileron
aileron
five'spoller
general
necessary
for
and
In
Directional
dimensions
limits
is
speedbrakes
panel.
A
to be
stabilizer.
outboard
as
aft
each
control
an
wing.
on
rotation
a movable
an
slightly
trucks
pitch
panels,
sixth
rudder
controls
mounted
symmetrically
inboard
of areas
The
spoiler
and
The
which
and
Krueger
cambered
Longitudinal
flaps,
operate
the most
summary
five
The
unslotted.
between
seBments
obtain
standard
trucks
center
To
triple-slotted
flaps.
wing
trucks.
outboard
flaps
these
the
has
variable
are
of
system
of
elevator
with
obtained
page
pairs
and
is
sho_aon
four
with
a pair
allows
inboard
each
page
two
flaps
transport
airports.
wing
edge
are
four-wheel
employs
control
on
of
control
operates
conjunction
Is
the
nacelle
equalizing
travel
the
leading
Krueger
mounted
limited
through
inboard
consists
A load
between
type
intercontinental
international
characteristics
inboard
of body
four-fanjet
existing
Krueger
the
gear
wing.
midway
The
and
the
large
from
speed
of
while
a pair
a very
low
outboard
slotted
main
is
operate
necessary
trailing
of
747
arrangement
is
on
page
simulation
placards
are
1.1-4.
i
,
REV SYM
BO_rJNO
-3o6_ 3
INo. Vol.
JPAOE
lI
i.i-i
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t
/
i
AOflNO
.D6-306h.3_
NO.
VOI.
II
I
PAOE I .I-
P_
I
SUMMARY
OF AREAS
AND DIMENSIONS
VALUE
ITEM
DIMENSION
m
m
Wing
Area
Wing Mean Aerodynamic
Chord (MAC)
Wing Span
Wheel
(b)
Ft. 2
27.31
Ft.
195.68
Ft.
78.96
88.96
Ft.
Base
Wing Gear
Body Gear
Wheel
55oo
(S)
_Q
Tread
Wing Gear
Body Gear
_te
Ft.
Effective Engine
.Moment Arms
Inboard
39.6
Ytz
_te
Ft.
(ground)
Outboard
69.4
Yzo
Fte
(air)
[7.3
Note
The transition
engine
main
between
pitching
landing
arms, ZKZ
gear
For 0 • R _ I,
__
where.
and
the ground
and
compression
Zao
=
z_oa,_
ZaX
= ZaX._a
_Z_o
=
ZEO&ROUNO
_Z_I
=
ZE_e.ouNo-
-!
-- _T
E
-
(ground)
and air values
ZEo,
is a function
ratio,
+
•
Ft.
Ft.
for the
effective
of the averaged
_ .
_'_za0
_'AZEI
ZEo_t_
_
_EXA,_
=-5.@
Y_in Landing
Gear
1.9
Oleo
FT.
FT.
Compression
(inches).
C
SEE SECTION 19
FOR REVISED DATA
REV SYM
D
_:__-/':.f/i:.,,_
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DATE
THE
.,
PAOE
1.2
FLIGHT
All
CONTROL
primary
hydraulic
Each
a
flight
control
actuators
segment
single
of
the
hydraulic
system
driven
conventional
by
by
schematic
to
the
All
elevators,
type
of the
showing
the
unrestricted
control
input
is
on any
_r_itudinal
is
the
shown
ailerons
by
and
four
on
overtravel
to
the
There
the
is
event
irreversible
hydraulic
is
systems.
driven
ground
spoilers
are
individually
A hydraulic
hydraulic
page
systems
1.2-7.
capability
to allow
remaining
surfaces
no manual
reversion
total
by
various
systems.
of
motion
by
actuators
hydraulic
input
in
flight
cylinder
inoperative.
surface
and
powered
The
actuators
have
powered
independent
actuator
four
actuators
capability
four
distribution
surface
actuator
are
rudders
simgle
control
given
by
combinations.
three
control
surfaces
supplied
"dual-tandem"
supplied
1.2.1
SYS_ES
hydraulic
when
power
is
a
lost.
Controls
The
longitudi_al
control
system
and
a trlmmable
stabilizer.
consists
of
four
elevator
se_ntl
Elevators
Each
of
the
dual-tandem
powered
actuator
_e
four
elevator
actuators.
by
by
inboard
two
one
segments
*Each
hydraulic
hydraulic
elevators
are
is
inboard
systems
system
and
as
controlled
independently
elevator
each
shown
from
actuator
outboard
on
the
powered
page
aft
by
is
elevator
1.2-7.
quadrant
while
O
I
REV SYM
D
BOXJNO
.o.
pasz
Vol.
II _
1.2-1
lt-'7000
1.2
(Cont'd)
the
outboard
elevators
the opposite
any segment
the
hinge
inboard
moment
is
by a control
surface.
will
cable
If hydraulic
trail
system
to
power
to
at an angle
where
zero.
is us_
of a hydraulic"_
setting
is powered
elevator
feel system
consists
bilizer
slaved
is lost the segment
An artificial
which
are
to program
' spring
and a mechanical
by hydraulic
number
i as a backup
normal
feel forces.
system
to number
modulated
centering
numbers
3-
the feel forces
with
spring.
sta-
The system
2 and 3 normally
Any one
with
system will
provide
Stabilizer
The trimmable
driving
stabilizer
a single
both motors
is actuated
jackscrew.
The power
available
is the same, but the trim rate with
half that with both motors
control
by two hydraulic
which
varies
operating.
the rate
with
motors
one or
one motor
Each motor
for both
motors
has
is one-
a rate
operating
from
-L
0.5
deg./sec,
Simultaneous
callyby
at low speeds
control
the thumb
by the control
signal.
of both
switch
motors
on each
stand levers
The autopilot
to 0.2 deg./sec,
which
system will
at high
is obtained
control
wheel
override
operate
speed.
either
electri-
or mechanically
any electrical
either
input
hydraulic
motor.
Lateral
Controls
The lateral
and outboard
control
system
ailerons
comprises
a combination
and spoilers which
also
of inboard
can be used as
o
D6-306_3
REV SYM
D
Bo ,.,vo
vol.n
IP_,_E
1.2 2
1.2
speedbrakes.
(Cont'd)in
Two dual-tandem
the wheel
wells
and right wing
central
drive
lateral
control
control
independent
control
actuators
each of the pilot
central
cable
surfaces.
is provided
and co-pilot
actuators
systems to the left
Pilot
input
by a cable
control
located
to the
system
from
wheels.
Ailerons
Both
inboard
actuators.
and outboard
ailerons
are actuated
by dual-tandem
_'ne inboard
ailerons
operate
in all flight
tions, but the outboard
ailerons
operate
with
A lockout
mechanism
on the outboard
aileron
will
is actuated
flap follow-up
actuators
lic systems
which
to neutral
to any aileron
trailat
down
only.
by a switch
positions
the outboard
up.
When
both hydrau-
are inoperative
the surface
flaps
surface
an angle where
electrically
linkage,
with
flaps
condi-
the hinge
moment
is zero.
Spoilers
There are six spoiler panels
with
lateral
board panel)
isactuated
to prevent
when
control
which
and speedbrake
the panel from
system
floating
and one
only.
which
from
are modulated
(the most
The five
moment
handle
valve
angle
modulated
two "mixer boxes"
speedbrake
in-
Each panel
has a check
up to a zero hinge
is inoperative.
from the pilot's
five which
speedbrake
actuator
on each wing are controlled
sum the inputs
inputs
is an unmodulated
by a single hydraulic
the hydraulic
panels
on each wing,
which
and the central
G
I
0
!
REV SYM
D
AD'Of/AIrO
@-?000
1.2
control
(Cont'd)
actuators.
panels
The
up
or
down
speedbrake
functions.
will
the
detent
spoiler
position
panels
solenoid
Further
handle
only
the
on
box
other
mixer
Wheel
forces
will
travel
limits
at
the
31
41
full
is
reached,
6 and
movement
to
handle
to
the
91
i0 which
and
will
81
full
the
raise
the
controlled
angle.
directly
position
remaining
Before
handle
or blowdown
detent
detent
are
speedbrake
deflection
settimg.
two
"inflight"
controlled
ground
spoiler
into
or blowdown
moving
are
the
speedbrake
divided
deflection
the
any
is
7 which
to
move
spoilers
5 and
to
valvel
ground
inputs
speedbrake
panels
mixer
"inflight"
of
the
control
the
operation
spoiler
one
raise
within
Moving
raise
through
Lateral
will
by
a
two
amgle.
is
panels
possible
through
the
box.
-
cam
an
i. 2.3
provided
arrangement.
electric
the
on
are
wheel
the
servo
angle
being
and
by
shifting
The
control
actuated
by
a
available
changer
with
the
ratio
two
obtained
servo
spring
loaded
rotating
this
foll_er
and
mechanism
the
zero
wheel
force
motor
is
operated
by
with
datumm to
a switch
Controls
is provided
rudder
motor
a simple
stand.
directional
each
is
desired.
control
Directional
The
Trim
with
a
by
system
consists
a dual-tandem
"q" Programmed
fr_
full
comparator
changers
pedal
circuit
disagree
of two
actuator.
gear
ratio
travel.
to
beyond
changer
monitor
segments,
Rudder
Each
the
rudder
which
rudder
their
system
llmitlmg
limits
has
a
the
ratio
operation.
If
tolerance
i
o
REV SYM
D
:
D_AL'JNO
.o.
Vol.
II
PAG, 1.2-4
_7_
1.2
limits,
a warning
(Ce,nt'd) certain
c_ditions
rudder
available
_aller
angles
_e
rudder
light in the cockpit
such as one hydraulic
_ill be limited
than the ratio
pedal
rudder
knob
changer
between
airspeed.
the feel
of airspeed
A series
stand.
force
_c
due to the ratio
feel
rudder
actuator.
add to the inputs
loaded
system.
required
and the actuator
the feel
trim
authority
changer
function.
Rudder
from the pedals
to
.
inputs
follower
However,
to obtain
a given
the ratio
input.
unit with
type yaw damper and turn coordinator
into each
_here
capability
Xhis is done by installing
system
are
off operation
by a spring
force
trim is obtained by rotating
on the control
_ene
allaws.
to the lateral
varies the pedal
angle with
changer
Rudder
changer
similar
system
by actuat_
forces are programmed
and cam arrangement
the ratio
is activated.
the trim
is also a function
system
from
these
or trim but will
is incorporated
systems
will
not fe_d back
.i
through
the control
system.
1.2.4
The flap
system
flaps as shown
The
leading
powered
consists
edge flaps
retracted
motor.
edge flaps
and trailing
edge
on page 1.1-2.
by a separate
electric
of leading
_e
comprise
air motor
leading
or extended)
four
sets per wi_,
or, as a backup_
edge
flaps
each
by a separate
are two-positioned
and are programmed
set being
to operate
(ful_
in conjunction
0
!
D
q[i
I
REV SYM
D
_P'd_'lJ_VO
D6-30643
.o. Vol. II.
p,_E 1.2"5
___
f
@-70@0
1.2
(Cont'd)
with
the trailimg
flap
sets
extend
(6,
edge flaps.
7,
8),
fully when
flaps i.
As the inboard
4, 5), (9,
tube extending
is driven
appropriate
torque
power.
an inboard
trailing
shutoff
tion difference
flaps exceeds
flaps asymmetry
D
flaps
161,
flaps
extend
extend
edge
Each
During
trailing
a predetermined
powered
system
right and left
amount.
operates
one power
consists
gears
trailing
extension
3p
drive
edge flap
of the
motor
or retractionj
automatically
edge flaps
of
to a torque
by the turning
monitor
trailing
5,
fully.
has an electric
edge flap
asymmetry
to
(i_ 2;
system
wings.
Each drive
to the inboard
by
through
actuators
edge flaps
drive
21)
to flaps
flap sets
coupled
tube.
between the
Each
edge
(19, 20,
edge flaps
are actuated
along both
screw
monitor
15,
leading
by another.
laterally
for backup
hydraulic
edge
(see page 1.2-71
by two ball
(l_,
leading
(22, 23, 2&, 25, 261 extend
trailing edge
motor
A comprising
trailing
the remaining
the two outboard
a hydraulic
131 /
trailing
lO) / (17; 18),
The two inboard
REV SVM
12,
the outboard
group B comprising
system,
(3_I,
Group
when
causes
the posi-
inboard
trailing
e_e
An outboard
trailing
edge
similarly.
BO_='jNO
.o. Vol.Ii
P,aE 1.2-6
@-?o0o
J
O
Q
<
RFV SYM E
.o+.++
i.o.
+oL!_
13
DISCUSSION
An
OF THE
aircraft
resulting
SIMULATION
in
motion
is
from
thrust
and
and
aerodynamic
the
coefficients
the
key
to
loads.
of
acted
upon
gravity
These
the
a realistic
by
external
effects,
force
airplane
and
landing
moment
equations
description
of
forces
and
gear
forces,
components
of motion,
the
moments
comprise
which
aircraft's
are
flight
characteristics.
Aircraft
flight
solution
of
simulation,
these
representation
allow
This
the
document
created
of
by
mity
of
to
geometry
extensions.
The
data
contained
the
in
axes
from
y-axls
the
This
plane
xs-axis
body
through
loads
on
axes
the
lles
in
as
that
by
in
center
the
of
the
the
the
an
of
necessary
They
on
page
and
and
may
be
and
gear
about
Stability
rotated
that
is,
the
x -axis
s
relative
wind
and
the
body
plane
of
symmetry
axes
about
attack;
(x-z)
.
proxi-
computed
are
functions
number,
ground
1.3-4.
z-axes
moments
Mach
deflections
are
to
realism.
forces
gravity,
sections
x-
accurate
airspeed,
control
shown
real-time
sufficient
airplane.
following
angle
with
altitude,
as
as
characteristics
the
such
Zs)
well
describe
rates,
Ys'
continuous,
as
simulator
including
determined
also
and
which
in the
(Xs,
requires
motion,
data
changes,
flap
stability
the
ro%ation
and
differ
"fly"
variables,
of
systems
contains
attack,
and
those
aerodynamic
several
angle
equations
of
pilot
then,
lies
y-axls.
of
the
D
|
O
•
REV
/
SYM
,_,a,_rjN_
I_5-3_3
INO. Vol.
IPAOE
II
1.3-1
9@-7OOO
l.B
airplane
(Cont'd)
by
and is thus rotated
sideslip
system
angle,_.
effects
While
as dimensionless
terms
(stability
of each important
functions,
the appropriate
motion.
method
axes
used
dynamic
to solve
forces
systems
reference
line
This allows
flow
with
for any axis
A second
moments
E-3033
R1
the
in which
here
or through
to
of
the
the aero-
into body axes.
Note
is to the fuselage
the angle _
inertias,
transformations.
the aerodynamic
of the
be resolved
1.B-5 presents
are transformed
any inertia
must
F.R.L."
Ixx , lyy, Izz ,
Note also that
the
= o_ F.R.L. + 20) is used only in
data curves
- it is not used
transformations.
due to thrust.
engine
is the simulation
Flight
moment-arm
testing
of forces
has shown
representation
ISEcT
and
that a simple
accounts
BOEJNO
REV LT,R:
showing
are broken
of the equations
equations
use of body axis
item of importance
effective
and moments
chart on Page
wing angle of attack (_W.D.P.
conjunction
which
the presentation
for solution
(FRL) body axes,
and Ixz without
derivatives)
axis transformation
the direct
in this
coefficients
simplifies
the dynamic
and moments
that the appropriate
axes
the forces
The schematic
measured
wind
parameter.
the use of stability
aerodynamic
zs away from the relative
Forces and moments
are presented
down into separate
about
for thrust.
,
D6-S06_S
INO.
Vol.
II
IPAGE z.3-a
1.3
At present,
(Cont'd)
2.5 ° from
the thrust
the fuselage
inward by 2 ° .
vector may be considered
reference
The corresponding
line,
to be inclined
with each
thrust
effects
Tx = Teng. #l + Teng. #2 + Teng. #3
engine
up
canted
are as follows:
+ Teng. #4
Ty = .0349 (T#I + T#2 - T#3 - T#4)
Tz = -.O436
T
X
LT = .O_36 NT
_ =(%_+_#_)"
%o+ (%_+_#3
)"_
_T= (T#I
+ _#4)"_o +(_- %3).%
The effective
The thrust
moment
reverser
effects
coefficients
crements
Cm
engine moment
2.0-1,
are found
on the lift,
are presented
are to be added
on pages
arm values
3.0-1 and 4.0-1
1.1-3.
drag and pitching
in Section
to the equations
on page
12.5.
These
in-
for CL, CD and
respectively.
e.g.
Sign
conventions
for the controls
are shown on page
listed
on page
maximum
control
a list of maximum
E'3033
RI
deflections
and a summary
of nomenclature
values
is given
begins
on page
B_'_=TA#4_'
REV LTR:
coefficients
are
1.3-7.
To aid in scaling,
1.3-8,
1.3-6, while
and aerodynamic
on page
1.3-9.
l N O.
D6-30____3
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747
D6-306153
Vol. II
PAGE
1.3- 6,
•
?001
MAXIMUM
CONTROL
SURFACE
}_a_cimum
Control
Surface
STmbol
DEFLECTION
AND
Normal
(Full
.
Displacement
Rate
(deg)
RATES
Operation
Boost)
One Hydraulic.
System
Failure
(deg/sec)_
Rate(deg/sec)
Elevators
37
Inboard
-23
+17
+17
-23
Outboard
Stabilizer
Pilot's
Thumb
Switch
Control
Stand
Levers
+20
-20
26
40
2?down
-25
Spoilers
Panels
1,2,3,4
down
up
35 up
45 up
+15
45
down
22
55
up
45up
down
1
_p
20
(Speedbrakes
!°.
.I
J
I
L
9,10,11,1R
6,7
37up
{+3
-32
_utboard
Panels
down
up
down
+
-IO0.5
.
5,8
30
26
30
•_ F.R.I..
Ailerons
Inboard
Panels
down
37up
37 down
25
only)
Rudder
Upper
I_R u
Lower.
_RL
Note
The
deflection
flight
40
40
rates
conditions.
characteristics
0
Data
For
NASA
simulation
place
The
_ne
are
74.7
ofindividual
shown
actual
presented
Flight
used
above
a
are
average
non-linear
in
values
applicable
deflection
D6-13336_
'Tlight
rate
Control
to
all
response
Systems
Simulator".
control
control
_heel
rate
rate
limit
of
i00
deg/sec
in
limits.
Db- 30o,. _,
-
"_r,-,_
TT"
PAGE 1-3-7
f
Parameter
MAXIMUM
VALUES
Maximum
Value
Units
30
ftlsec
cry
25
ftlsec 2
120
aL
ee
ft/sec
50
e
deg/sec
2
30
deg/see
2
60
deg/sec
&
75
deg/sec
+
30
deg/sec
40
deg/see
30
des/see
+
+
Altitude
_
ft
Rate
of
Climb
15000
ft/min
Rate
of
Descent
15000
ft/mln
0-40
I_4AC
NOTE : .
of
2
deg/sec 2
130
Center
2
_X
Gravity
For
Range
complete
D6-13585,
center of gravity
"M_ass Properties
information
- 7_7 Flight
consult
Simulator".
O
8
g.
Q
REV SYM D
_P'a_'_'P_VO
,
D6-30_ 3
I .o.
Vol.
JpAOE |.5-8
II
6-?OOO
NOMENCLATURE
SYMBOL
DEFINITION
Acceleration
along
b
Wing spa. (_).
Co
Airplane
drag
the
"i"
axis
, where
CO
coefficient.
a
_s
CL
positive
drag force
Airplane
lift coefficient.
a positive
Cv
C!
Cm
lift
acts
force
the negative
_t."
--LIFT
acts along
the negative
side
force
coefficient.Cy-__.,
positive
side
force
acts
Airplane
rolling
axis, X$
.
Airplane
pitching
C_
Airplane
moment
the
coefficient
axis.
I_$ axis.
where
positive
"_i
about
the
a
axis.
reference
m _Jm.
_sb
yawing
@15 "
axis,
along
X$
, where
%s
Airplane
moment
axis,_s " Cm "
C_n
along
about
the reference
M
moment
C n
coefficient
--
coefficient
about
the
reference
N
%5b
Airplane
center
of
gravity
the wing mean aerodynamic
I
m
Wing meanaerodynamlc
Rudder pedal
pedal
pull
s
Control
moment
column
force
as
a
fraction
of
chord.
chord
(ft).
positive
for a left
rudder
{llb).
force
Control
-
force,
position
force,
positive
for
_ colu_m
.
(Ib).
wheel
force,
positive
for a clockwise
wheel
(ib),
o
%
,
REV SYM D
ROXJMM_I.o.
D6-30_3
Vol. II
IPAsz 1.3-9
DEFINITION
HM
=
Hinge
moment
=
Pressure
altitude
=
Airplane
mass
axes
L,M,N
=
Rolling,
system.
number.
Mach
I_z
=
A'rplane
rl z
normal
-Z
=
moment
pitching
axes
=
(ft).
of
inertia
about
the
reference
j (slug-ft2).
i,
M
(lb-ft).
and
load
yawing
factor
5 C.O_,(X-X.351NO(
moments
along
+.043f.T-IZdu
about
the
a reference
z-axis.
_ ,=, ll.d,n_ r"-'-*" -e-__-__,_
m9
p:i_jr
=
Roll,
pitch
system
=
Dynamic
and
yaw
=
pressure
Impact
pressure
a reference
axes
(lb/ft2).
9,c"
P','o',',,L-
vt"
a95.4
(ib/ft2).
Pm'.'r,c
=
s[(,
S
=
Wing area
T
=
_tne
V
=
1'rue airspeed
=
Calibrated
v_
about
(radians/sec).
= _ pV t .
c_c
rates
,]
(ecZ).
thrust (lb).
(ft/sec).
-_rspeed
(knots).
v_.= ,,,,_.,o_../'(_1,_,,.._.,,.,..I)"'-,
=
Equivalent
airspeed
(knotm).
0
I
P
O
9£
REV SYM
D
Bd_'EIN_
NO
VOI.
II
D6-306__3'
,,Ao1.3-I06-7OOO
SYMBOL
DEFINITION
=
Airplane
=
Axial
gross
weight
force, side
reference
force,
=
Body axes system
XSj I.JS_T-S
=
Stability
×w,
=
Wind axes system
YL,
=
Effective
inboard
Yto
=
Effective
outboard
Z£x
=
Effective
inboard
Z[o
=
Effective
outboard
c_, 0_F.R.L.
=
Airplane
2w
axes
angle
reference
_W.D.R
=
Airplane
plane
I_A z
=
=
--
described
system
in Section
described
described
engine
in Section
a
engine
arm
yawing
(ft).
arm (ft).
pitching
arm
engine pitching
of attack
1.3.
1.3.
arm
(ft).
(ft).
relative
to the
fuselage
relative
to the wing
(degrees).
of attack
angle
design
ratio.
Aileron
deflection
Inboard
aileron
Control
=
Elevator
=
Inboard
=
Outboard
=
Flap
(degrees).
angle
deflection
aileron
deflection
column
deflection
movement
(de&Tees).
deflection
elevator
elevator
(degrees).
angle
deflection
deflection
angle
(degrees).
angle
angle,
(degrees).
positive
for rearwaz_
(degrees).
angle
angle
(de&Tees).
(degrees).
Setting
,
REV SYU D
along
1.3.
in Section
yawing
engine
sideslip
pressure
column
_V..z
force
(degrees).
Outboard
_o
l_r.
Air
line
angle
Airplane
g
and normal
system.
X, tJ,Z
_w,
(ib).
_'_'l'JN_'
D6-3iO i3
JNo. Vol.
JP*OE 1.3-11
e-7_
SYMBOL
DEFINITION
Rudder
pedal deflection
rudder
deflection
(degrees).
_R
Rudder
deflection
angle
_aL
Lower
rudder
deflection
angle
(degrees).
Upper
rudder
deflection
angle
(degrees).
_p
Spoiler
deflection
Control
wheel
clockwise
reference
@
angle
line
for positive
(degrees).
angle,
movement
stabilizer
positive
(degrees).
deflection
wheel
Horizontal
_KR.L.
angle,
positive
for a
(degrees).
angle
relative
to the fuselage
(degrees).
Airplane
Euler
pitch
Airplane
Euler
roll
Airplane
Euler
yaw angle.
P
Air mass density
0
Air density
angle.
angle.
(slugs/ft3).
ratio
SUBS_
F.R.L.
=
Fuselage
W._P.
=
Wing
As used
with
reference
line
(any body
water
line).
design plane.
VE and M placards:--
Ft
=
Operational
LO
=
Landing
gear operating
LF.
=
Landing
gear extended
-
flaps
extended
placard.
placard.
gear cannot be extended
to X_
............
placard.
above
(Note:
The
the LO limit,
landing
but fli_t
limits is possible).
;
-
o .....
O
,
REV SYM
D
BO_"J,
AIfO
-306_ 3-
I NO. Vol.
IY
@-7O00
MO
=
Maximum
OF
=
Design
The
this
time
derivative
document
by
operation
the
dot
operating
dive
flight
normally
derivative
limit.
placard.
denoted
by
d_
dt
is
replaced
d(_.___
. (-)
notation,
dt
in
A.O
.i
D
t
D
REV 5YM
.
D
AI'_'_TMM_
_-3_9
INO. Vol.
II
IPA_E 1.3-13
6-7000
2.0
LIFT
The
FORCE
COEFFICIENT
dimensionless
given
in
terms
aerodynamic
of
its
lift
significant
force
coefficient
components
by
is
the
equation
below.
At
C_.O.P.)
a given
• O(w._p "
+'
--dCL
" _a
dnt
+
Koc "de"
+
+
•
",_F.tu
+
K,_.
A C_. Seolu[_
+
ACL om-_
_umo_s
bCL
+ [_Ct
c,Aoo_=
dCL
.&s_=
u
F_
.
+
_UjRt]
.
K_.
de,.
"_-o
_
r
where,
C
L IL_SI¢
=
Basic
lift
_L_.=O*,
in
retracted.
For
on
page
plotted
&CL)e(w.o.R.Oe--
coefficient
Change
due
2.0-7
on
page
in basic
free
•
low
for
air
and
speed,
For
flaps
the
rigid
with
the
CL_$1&
up,
airplane
at
landinz
gear
is plotted
CL_SI
&
is
2.0-8.
lift
to aeroelasticity.
coefficient
For
low
at
O_w._.p.-
0 e
speed,(AC_1,_._.,
Oe
D
t
m
I
D
,
REV 5YM
B
n_l_
].o.
Vol.
]P*"z2.0-1
II
2.0
is
plotted
on
page
2.0- _.
(Cant'd)
is
plotted
on
page
2.0-I0.
i_r
For flapsup,
%
elastic
effect
on
the
rigid
airplane
basic
lift
is
%0_/
plotted
= Change
change
on
'%_Vl
page
= Change
rate.
page
2.0-IZ.
in basic
lift
of
of
on
angle
attack•
in basic
llft
Change
to
rate
_C_
coefficient
and
are
=
due
is
of
plotted
2.0- Q.
where
dCk__.nt
coefficient
plotted
in basic
on
lift
the
center
page
to
due
of
pitch
gravity
factor,
2.0-14,
coefficient
to
due
aeroelastlc
dn_
inertia
relief
caused
by
normal
is
plotted
load
factor,
_IZ
"
dCL
For
low
speed,
-----
on
page
2.0-15.
dnt
For
flaps
up,
dC&.
is
plotted
on
2.0-1&.
page
dn I
Change
in
basic
lift
coefficient
due
to
change
in
d_
stabilizer
angle
from
_F._.t
w 0 b "
-----dC_ is plotted
a_
on
REV SYM
page
2.0-I'I.
The
effectiveness
_O_l_O
factor
i _,
for
Vol.
the
II
I'*@"2.0-2
9I-?ooO
%
2.0
stabilizer
elevators),
is
_
on
plotted
4.0-19.
page
(Cont'd)
(and
_Change
in basic
inboard
lift
elevator
coefficient
angle
from
due
_%_
=O
_
to
change
. dC_
in
is
d6_
plotted
on
= Change
normal
T_OuLtR%:
Change
coefficient
angle
2.0-..
is
+
lift
"spoilers"
extended
while
used
"speed
panels
to
determined
page
and
1.2-?.
ground
the
(rigged)
faired
change
in
out-
is plotted
on
one
due
wing
elevator
_osition.
It should
symmetrically
be
are
extended
to
spoiler
noted
used
or
that
for
lateral
spoilers
are
brakes."
(_CL1s,)e 6
coefficient
to
BCL
coefficient
deflection.
control,
free
from
speedbrake
where
_£_0"
from
stick
2°
in basic
for
due
_
system
deflection
AC_
lift
elevator
page
The
2.0-18.
in basic
board
on
page
due
45".
from
is
to
the
deflecting
in basic
the
lift,
operating
spoiler
The
operating
spoiler
panels
the
hydraulic
systems
schematic
(ACLsIp)4_S
spoilers
change
on
is
page
plotted
2.0-_land
for
are
on
spoilers
page
2.0-_
!
D6-3o6_B
REV SYM
]l)
-_
l._t
Vol.
2.0-3
II
->&.'/O00
2.0
¸
(Cont
respectively.
°d)
(Kis@)
a is
The
plotted
effect,
asp
The
2.O-23.
spoiler
effectiveness
on page
_
2.0-20.
asp
_-o
The
is
erred,
aeroelastic
-_actor,
_ach
plotted
nulber
on page
(L-_-__ is plctt_
%
on pages
effect
_CLouviOAa
b _- Change
2.O-2_,
2.0-25,
factor,
F_
E
in basic
lift
and 2.0-26.
is obtained
coefficient
The
from
ground
page
5.0-Zq.
due to _utboard
A I i., _u R, OI_I5
aileron
deflection.
aCL O#o A-
° . ac,.Ao.
LEI:_ N_bRIGI4T
ocmK_pW _u.tRowS
where
_CLA
due to
deflecting
the
is
0
opposite
plotted
one
outboard
on page
effectiveness
5.0-Z6.
is the change
outboard
aileron
The
I_l;Ao
factor,
effect
li_t
aileron
2_ e or
_CL.Ao
1_ e.
outboard
aileron
is plotted
factor,
coefficient
up to
down to
2.0-27.
The ground
in basic
on page
_a_a
is obtained
from page _.O-Z_.
_CL-
LANbiNiP
=
Change
in basic
lift
coefficient
due
to
main
and
nose
OE&IL
landing
where
number
gear
AC_l_is
effect,
2.0-_0.
o
is
8
REV SYM
extenmion.
plotted
plotted
on page
¢Ca_^R.)pt/_L,_t)_s
Theianding
on page
gear
2.0.-29.
The
plotted
effectiveness
factor,
_c,_
on page
ICGIIMt,
2.0-28.
->_-?ooo
f,
•j
2.0
.
ACL
aROUNd
-Cnange
in
basic
llft
coefficient
due
to
groun_
effect.
F..FFF.CT
(Cont'
d )
AC
ACt.RAp
FAiLU_[
=
-
_CLr=F
where
effect
height
Change
in
or
is
basic
For
of
from
both
symmetric
hC
L FLAP
KGED
lift
CLG L
page
is
,
2.0-
the
flap
inboard
ed
on
due
to
flap
an
which
inboard
position
_r
or
both
_ne
=.
plot
coefficient
outboard
outboard
ground
page
2.0-_I.
extension
s_#mmetric
flaps
flap
occurs.
failure,
FAILURE.
(dCL_
A \ d_, / FLAP FA=LUr,_.. ¢X w.o.P.
+
[(
AC L
where
_
on
plot%ed
,factor,
retraction
failure
K G_.."
e
=
L eP_OLIND
EFFECT
_aR
]
FLAP _AmLUMt
=O"
is
the
ci_ar_e_
tn
no
s yTn.metrlc
basic
= 00
lift
Ooard
!
-_
coefficient
,°.
2[
at
outboard
0(w.D.p.
flap
due
[(AC,
failure.
in- °
L)(XW.O.p.=O-]FLAPFAILURF,.
(dc.
i .....
is
plotted
the
change
of
symmetric
on
page
in
basic
2.0-3_.
lift
inboard
Ak--_---/R.Apr_ILURZ-0(W=.p.
coefficient
o___routbt_rd
due
flap
_o
failure
.
is
the
effect
on
_e
i
rigi4
airplene
basic
A
llft
coefficient
A, uMe
_._e
above
data
tor
limited)
is
inboard
inboard
flap
failed
monitor
limited
curve
slope.
I ,
•
also
applicable
or
and
extension
outboard
_ne
or
for
flap
opposite
_sy_etric
(moni-
failure,
inb.oard
retraction
e,_],,
flap
st
one
Dhe
position,
C_
O
i
REV SYM
NO. Vol. I!
,:,AGE2.0-5
96-7000
2.0
AC"
(Cont'd)
L
FLAP
computed
Lhat
I
for
The
angle
of attack
angle
for
2.0-35.
The
2.0-36.
Certification
the
inboard
fw_ctions
!
of
is
FAILURE
of
stick
attack
to
inboard
flap
flap
posltion
flap
in
shaker
for
Oe
this
_o
c_,.ca-
posiblon.
should
De
CL
Note
used
for
all
docunen%.
actuatior,
i_litial
added
is
plotted
o_
oa-Se
buffet
is
plosted
on
plosced
on
p%_,'e 2.C,-[_
p_e
$
stall
speeds
are
The
t
initial
page
buffet
boundary
and
trimmed
CLMAx"
values
are
plotued
on
2.0-38.
l
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E
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O(W.D.p "
"_" A CO SPOSLZ_
"J-
AC.OLANOING
GEAR
4"
Z_,C.OGROUN 0
EFFECT
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"J- llC
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airplane
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at
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on page
4"fans, i,
FAILUP. _
gear
3.0-5.
_., i0, 20,
2>
".0.
ana
dCO
A CO RUDOZR. _
coeYficient
rcuracted.
K
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2hanEe
in b:,,niu dr'-/ ,:oefficier, t due
co
change
in
d_
•"a :
....bLlizer
ant;le
from
'_F.R.L.
=
dCo
O*
is
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d_
o:-: oa6e
_.0-(- ant
p_r,e ;.0- v.
o
ion] M
-->-_v
_n
coefl'ieienU
_age
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NOT
REV SYM
-
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Tr,T_,;]I_A_'r','lrr_,.Tn_
3.0-8
Lerms
of
at
_'_ci, numoer,
and
_.0-9.
the
lift
CL
force
_
M.
[CO] H
is given
ig
oy the
coefficient
plotted
first
equation
on
2.0-i.
B_-gAM_
_o
Vol.
II
i_A_E DG-306h3
3.0-1
3.0
(Cont
ACDspoIL
'd )
=Change
in drag
coefficient
due
to
spoiler
or
speedbrake
deflection.
_Cospo,.zRs
=
Z_Cos p [I
+ Fo" KssL]
'
+.OS-Fo-K_• 45n
Z
"
_"r_" "
OPCR/mNG
"_,ImOII.E,R
ACDsll
where
spoiler
D
is
the
in
drag
or speedbrakedeflection
,...,
where
change
coefficient
in
free
coefficient
"°"
o
is
at _w.D.R-4"
spoiler
_ne!s.
due
change
the
to
to
air.
•-
(_CD_ID)_I_I_._4
due
Cl
in
drag
deflecting
the
operatlug
(_Co_)
_w°.'*"isplotted
for
spoilers
on
page
3.0-10
spoilers
on page
3.0-1E
and
.
is
d
tabulated
{AC°_I
for
is
the
ground
rate
of
d_
change
with
angle
of
attack
of
drag
coefficient
due
to
deflecting
theoperatin_
spoile_
p=els,d IAC=_)
i.
do(
plotted
for
for
ground
effect,
spoilers
on page
spoilers
on page
(C°sP)M
3.0-II
3.0-I_
is plotted
)
ground
effect
on
page
is
plotted
spoiler
3.0-14
on
panels.
The
page
control
ground
2.0-51
_
is
The
.
page
tabulated
Mach
number
3.0-13
The
on
lateral
•
and
is
.
effect
.
the
Fo
factor,
I"I is
height
the
, is
factor,
number
spoiler
panel
plotted
of
K_
operati_
deflection
o
I
IQ
o
D6-3061_3
REV
SYM
B
JP'47_='IAYd_
'
NO,
PA_
VO1.
II
3.0-2
@-?oo@
angle
3.0
(C_nt
(degrees).
'd)
ACo
L_NO,Nm
=Change
in drag
coefficient
due to main
and nose
landing
gear extension.
:
AC-DLAND|N6
KBF.A
R " _COe,
F.A.m_ " (CD6F..N_)M
F°.,.,.,)..o
where
/ICoo_m _
is plotted
on page 3.0-1S.
number
effect,
ICDeam_) M
_ is plotted
The
on page
_ch
3.0-I_,.
(C'D,ba,_) M. o
The landing
on page
plotted
_,C_
GROUND
gear effectiveness
:Change
factor,
Kr,_a_R1
is
2.0-Z8.
in drag coefficient
rJ'FEC'r
due to ground
effect.
A
ACD
r:_OUNO
EFFECT
where
_CDe _
is plotted
effect height
ACD
Sl
DF...%LI p
"-Change
factor,
P
is plotted
On page
where
A_.,OR u
coefficient
the lover
due
rudder
are obtained
and
to
"
A_..OK L
deflection
respectively.
The
ground
3.0-17.
of sideslip,
_.
3.0-I_.
due to rudder
ACDR
from page
.
on page
due to angle
---Change in drag coefficient
ACDRuoOF_R s
3.0-18
is plotted
in drag coefficient
ACDslo[_U
AC. o auoDa_,
on page
A
KE_,
deflection.
U "1" ACDRI.
are
of
the
the
_DR
changes
upper
u and
in
drag
rudder
and
_CDR
L
3.0-_.
D
CK
REVSYM
D6-30643
B
I
6-7OOO
L
"-.0
Z_CDFLAP
FAILURE.
Cont'd)
finds
0o_::-:
.
{dCo]
_ _-_!
+
[( )
•,,."_<_e
A.CD
dr_,
]
=O°
FLAP
CC"['<;_c-'-:.,.
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IP..I?.-6]
R_v_S_D
EI_
?.4.b1_ CURUUT'I
COEFFICIENT
?_1"7
3.E,."/O
GROUND
_.PR
EFFECT
D6-306_:
VoW. IT _
APR
INK
TD
461
_
OAT_
DRAG
CHEC_ FO_T
:_++_
PA_E
K I NSM/kI_
C,.-R4
3./.,.'7O
THE
BOEING
COMPANY
"S.O- _8
_e_
ALpANEN(
15s_
TN[::
BO£1NG
COMPANY
E
:!i:i
ODE_RD
TD
4_1
C_lltl
THE
BOEING
(_OMPANY
::,
;:::if
:_.!:ii:l• :!
i
J
4.0
PITCHING
The
MOMENT
COEFFICIENT
dimensionless
given
in
aerodynamic
terms
of
its
pitching
significant
moment
components
coeft_cient
by
the
is
.-.
equati_r
b elow.
At
a
given
_W.O.P.)
where,
=
Cm.__
Basic
pitching
moment
airplane
at
landing
gear
For
speed,C
low
_p._=
coefficient
O"
, in
retracted,
m_
for
and
is
with
plotted
free
the
on
the
rigid
air,
C.G.
pages
=
with
25%
4.0-
the
M.A.C.
6
and
_
4.D-9.
mAgl_
I
w_.
= Change
-4D"
For
fi_ps
in basic
O<_m._.-O
°
uD,Cm_w_a
pitching
due
to
is plotted
moment
on
coefficient
aeroelasticity.
For
,
REV SYM B
__MM_
page
4.0-_.
at
low
speed,
D6-306 3
I,o. Vol.
| PAoz 4.0-i
II
96.7000
4.0
(A,Clll.z$)_w.ap.,,o
(Cont' d)
For
flaps
Change
the
up,
is
(_Cm_s)_
in
basic
speed,
moment
f_aps
plotted
on
in
is
moment
from
is
is basic
of
airplane
on
to
basic
For
page
25%
of
_Cm._
ted
M.A.C.
defined
pitching
angle
p_ot
low
4.0-13.
on
coefficient
in
page
due
The
to
total
Section
moment
4e
0-,..
center
of
lift
2.0.
coefficient
due
to
rate
attack.
dC.__._m =
d_
where
due
slope.
plotted
is
pitching
change
riBid
curve
__
variation
Change
the
coefficient
/
coefficient,CL,
of
is
m Oe
coefficient
up,
= Change
=
on
_(_CmC_.
gravity
dGn/'_c-'_
4.0-11.
on page
moment
effect
\d_
CL'(C_.n_
W._P.
pitching
aeroelastic
pitching
For
plotted
4.0- It.
page
_(d_._._r.s
*
K_
and
the
plotted
on
• dCm'zs'
d_
center
of
gray-fry
factor,
da
are
=
Change
pitch
in
basic
rate.
Change
im
aeroelastic
factor,
n_
pitching
d_m._
pitching
inertia
•
For
coefficient
plotted
moment
relief
low
4.0-1S.
moment
is
.-V-
basic
page
on
page
by
due
normal
_
to
4.0-|6.
coefficient
caused
speed,
due
to
load
is
plotted
is
plotted
dn_
on paBe
o
4.0-I_.
For
flaps
up,
_
dn_.
S
D
r
I
REV SYM
_rJMM_IN_-VoI.
D6-306_3
II
l.,,_z 4.0-2
96-700C
on
4.0
page
4.0-
t_.
(cant 'd)
_x"
_'_._L_.'-
Change
in basic
pitching
change
in stabilizer
moment
coefficient
due
to
F
C_n_._
is
effectiveness
_O(
_.
_n._._=
I
_X"
d_n.IS
._
Change
is basic
change
in inboard
in basic
change
in
d_._$
Change
spoiler
the
page
2°
in basic
stabilizer
The
(and
coefficient
from
page
4.0-_.
moment
elevators),
on
page
st'ick
free
from
the
deflection.
to
I
.
_
due
to
_.Om_o
4.0-_.
faired
brake
_E_O
from
(rigged)
moment
due
coefficient
angle,
pitching
speed
4.0-ZO.
angle
elevator
system
page
-_0 o
4.0-_.
on
pitching
plotted
_.eL_
moment
plotted
is ÷
or
for
elevator
is
deflection
on
on
outboard
normal
plotted
pitching
is
Change
The
from
factor
is plotted
dCm_$
angle
.
_
elevator
l_OSitlon.
coefficient
due
to
•
where
(/_C_._m
moment
spoiler
for
.._ ) _
coefficient
panels
spoilers
to
and
is
due
45 ° .
ground
the
to
change
in basic
deflecting
the
(_Crn._S_.._
spoilers
on
operating
is
page
pitching
plotted
4.0-24and
n
I
REV SYM
D
_O_'JNO
JNo. Vol.D6"30_3zl
IpA_K
4.0-3
..>
6-?OO0
4.0
4.0-25
(Cont
'd)
respectively.
factor,
is
The
(KS_)m
plotted
-M---_sp
Mi
The
The
plotted
on
4.0-27,
effect
factor,
is
Fm
page
is
plotted
pages
page
4.0-23.
aeroelastic
Fro@ E
on page
effectiveness
on
4.0-26.
ground
where
plotted
is
on
spoiler
effect,
4.0-28,
is
and
obtained
4.0-32.
The
4.O-29.
as
follon:
ground
a
effect
height
factor,
K_
is
plotted
on
page
:'.o-31.
A Cm.2S=NBOARD
---_ Change
in basic
pitching
moment
coefficient
due
to
A t t.E.I_NS
inboard
aileron
deflection.
pitching
moment
inboard
aileron
on
4.0-30.
page
coefficient
uP
to
KSA
X
aileron
deflection
is
due
20 e.
The
factor,
(C. m.z_a._l
K&A_
"2"S'_"
&, Nt i,
| ILOII,,IS
O_I_' _--
to
lO
deflecting
one
(_Cm.2sAx)_ois
inboard
plotted
aileron
to
be
obtained
and
is
plotted
effectivenesll
for
on
the
page
inboard
up
_.O-ZL
ill
'_
plotted
on
Fm@E
ACre
._.SOU.rtOARD
AI I.. IE RO1_1$
o
Change
page
is
in
basic
4.O-30.
obtained
pitching
_'_,_
The
from
,
ground
page
moment
4 rmU
effect
factor,
4.O-_.
coefficient
due
to
!
P
REV SYM
.i D
,'AO- 4.0--4
6-7O00
4.0
outboard
aileron
deflection.
(Cont'd)
_Cm._s_o-_.
where
_Cm._o
moment
coefficient
aileron
down
up
outboard
plotted
-
on
25 ° or
aileron
page
is
basic
nose
deflecting
is
The
from
pitching
landing
_
The
Mach
number
is
plotted
on
factor,
Change
in
ground
effect.
is
outboard
on
aileron
page
4.0-_|.
_o
effect
"
is
factor,
4.0-52.
coefficient
due
to
extension.
( Cm.ts a f..s,a) o,a
(Cm.,,,,_0,.o
plotted
on
page
4.0-311.
(Cm._s_)_/(Cmz$_M.o
4.0-_MF.
_(CnlzAR _
basic
pitching
factor,
ground
moment
effect,
page
one
plotted
page
gear
basic
outboard
.- Kc_t,- _Cm.z_-
KCm._t
ness
in
opposite
5.0-_&_
Change
where
the
Fm_
effectiveness
is obtained
L_Cm._0,_
to
change
_Cm.1_Ao
_
and
the
due
Fm_
main
/%Cm.zsr,,am.._ =
is
to
to 15 ° .
The
Y-..m._,..,_.,_
1<.8_o" ACm._.
pitching
The
is
landing
plotted
moment
on
I
gear
page
coefficient
effective2.0-211.
due
to
ifll_'t"
/XCm._61m.._
where
_C
= I_%.-
m,.._Gl_
_,Cm._.
is
plotted
on
page
4.0-_.
o
I
im
o
REV SYM
,D
D6-30643
_'JMM_
.o. Vol.
p,oe 4.0-5
II
->6-7000
1
• ..,..,
AC
.
!
\
m._P¢_ _IDE_LIP
Cm.as S,OESL,p
"
"
"
"
:-.'-3&.
.
AC m.z5 RUDOERS
Z_Crn.ZSRuoOERS
....
Z_CrlI.ZSRu
=
".
ACrnlsR
ACltl.t-,.5
FLAP
FAILURE
=
u +
Z_Cm.ZSRL
_'-
._ o
ITI.Z_Ru
= •
_._
....
4..
.
,
ACm._SRL
:,_re
t:-.e
cY._i_L_ec, _n
_,
.,,_!
"-," ,_ :: .... _ L 0-:_o.
ITI.ISRI_
-
.
. . ..
•
"
•
.
...
• ?L_:
-
,, _-
.......
'.'-'n"j
_;" ; :,th
c,/tbo:_rd.flaps
.
'
""1','_'_ °" 7" " :"
"
i
AC'I'FI'z5
FLAP
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APR
INK
ODEGARE} 5-3.Gc)
TD 461C-_I
THE BOEING
_
COMPANY
_,e_
4-.0-_
14-_ *"'"'
......
REV.
F.
5.0
ROLLING
MOMENT
COEFFICIENT
The dimensionless
aerodynamic
rolling
given in terms of its significant
moment
components
coefficient
is
by the equation
below.
At a given
C!
D(W._R
_
,
dC!
._
+
dCl
dF
+
ACI
d_
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ACI
SPOILZRS
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dP
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AILERONS
"1"
ACI
RUBDEJ_
"1" _C
1 FLAP
AILERONS
-I"
FAILURE
AC 1 L__ FAILURE_
where,
-- Rolling
moment
The complete
coefficient
expression
due to angle
dCl
for
of sideslip, _.
is given
as
d#
follows :
where
(d___
_d_
is the basic
moment coefficient
speed,
(d__
rate of
due
to
is plotted
\d_l
oo
angle
of
on page
of
sideslip.
rolling
5.O-7.
For
For flaps
5.0-8.
kd_/
plotted
change
/
up,
,.
"('_,),. o
on pa_e
5.0-9.
?JOT _'_
o
REV SYM • D
BOIJA_O
D6-30643
NO. VOI. II
b
6-7OOO
5.0
The aeroelastic
effect,
:(CAO)t
is
t_,.)..,
(Cont
' d)
plotted
on
page
5.0-10.
The
effect
of
main
and
nose
_ea_
extension, _"_-2,.---,,_is givenby:
The landing
plotted
on page
__F20_)
where
gear
2.0-28.
The
is obtained
as
the ground
effect
plotted
factor,
w Rolling
factor,
ground
K@LaA t is
effect
factor,
follows:
sideslip
factor,
F_
,
is
on page 5.0-11.
The ground effect height
B
_@_
t
is plotted on page 2.0-_I.
moment
the stability
for
effectiveness
dC_
coefficient
axis, X_.
is given
due to roll
The
complete
rate about
expression
as follows:
d_
The
k-_J
aeroelastic
page
--Rolling
effect,
on
5.0-13.
moment
coefficient
the stability
for
is plotted
___Ci
axis,
is given
_
as
due
.
to
yaw
rate
The complete
about
expression
follows:
o
o
U%
REV SYM D
lPAGE D6-3o6_3
5.0-2
9@-TOO0
5.0
(Conttd)
dr
where
IdC1__
is
aeroelastic
(C.i.p_Moo
_dP/
plotted
effect,
on
page
{_l_IM
5.0-|_.
The
, is
plotted
on
page
5.0-I_.
to
spoiler
deflection.
_Cm. P) _.o
CJL spm-u_s
"
Rolling
moment
coefficient
due
o,:n._.s
_lm,O
where
due
_
I_CiS4w)4_
to
sp_-o
is
deflecting
the
(AC,JSp_4S
is
the
rolling
operating
plotted
moment
spoiler
on
page
effectiveness
on
page
5.0-J&.
is
plotted
ground
effect
R/sP
factor,
The
pages
Math
5.0-1_,
factor,
panels
P1
_
number
5.0-20,
45 °.
is
plotted
|_P)I_I'%M
t
'CJ,--%-_IWtlO. _IS
%"
and
i
to
The
'
_
effect,
._
coefficient
5.0-17.
(_
spoiler
on
\
ILILPt
5.0-ti.
is
The
as
obtained
follows:
[
rjt, e =
where
is
the
plotted
factor,
Rolling
ground
on
K
moment
-]
| . FjL. K.:
effect
page
GZ
,
lateral
5.0-Z9.
is
The
plotted
coefficient
control
ground
on
due
factor,
page
to
effect
F L
height
2.0-_|°
inboard
aileron
kt_.NONIi
deflectlon.
REV SYM D
B_'l'J,_V_
NO. vol. II
lFAGZ5.0--3
D6-3o6_3
->6*7OOO
5.0
(Cont'd)
where
due
I_C_X)_O
is
to
deflecting
opposite
inboard
plotted
on
one
effectiveness
inboard
aileron
page
rolling
the
aileron
down
5.0-Z3.
The
to
Mach
up
coefficient
to
20 ° .
inboard
20 ° or
the
_&
aileron
factor,
K_AXw
The
moment
number
effect,
is
plotted
(CIAz)_
on
is
page
5.0-Za.
plotted
on
{C._._,z)_. o'
page
5.0-,4.
plotted
is
AC*I
Ou_OARD
AILERONIb
on
The
page
obtained
-----Rolling
aeroelastic
5.0-_.
from
moment
The
page
' (R_._. _)
effect
ground
is
p
_JAz
effect
factor,
5.0-Z_.
coefficient
due
to
outboard
aileron
deflection.
_A
where
0
is
deflecting
one
opposite
outboard
plotted
ness
on
aeroelastic
5.0-28.
from
page
ground
The
is
up
down
5.0-Z7.
K_Ao,
moment
aileron
aileron
effect,
The
rolling
outboard
page
factor,-
the
to
effect
to
25"
15 °.
outboard
plotted
_
coefficient
on
_
factor,
is
due
or
the
GC.,IA
0
aileron
page
is
effective-
The
5.0-_.
plotted
F_G
E
_
tO
on
page
is
obtained
5.0-Z9.
n
!
Q
i
REV SYM D
2.
Bd_'EINd_
• _ 3
D6-30o
J .o. Vo I • I I
-IP A_[
5.0-4
6,.'/0oo
5.0
AC
( Cont
i_OD
m
Rolling
moment
coefficient
to
rudder
deflection.
' d)
where
(ACI,Jz
s
coefficients
rudder
and
5.0-30.
and
the
The
the
lower
plotted
the
page
FAlUJRIr
=
where
ACI
inboard
extension
A_e_m_
plotted
rudder
rudder
and
,Ns'o Fa,_Re
flap
or
an
failure.
on
page
K_R
K_R
and
(CIML-) M
The
, for
(c,.j..o
are plotted
rudder,
u
L , are
respectively.
due
or
to
an
on
asymmetric
outboard
the
rolling
asymmetric
A_9",NWO
flap
(monitor
limited)
_'_,l._m[
is plotted
failure.
moment
for
on
pages
flap
5.0-35
and
respectively.
_,_,
on pages
for
5.0-_
flap
and
extension
5.0-3@
or
retraction
is
respectively.
t
REV SYM
upper
respectively.
is
retraction
the
factor,
factor,
6.0-18
5.0-3Z
inboard
to
plotted
effectiveness
coefficient
due
are
_
F,..)M.o
the lower
page
of
respectively.
(C'_u)M
and
the :rolling
deflection
effectiveness
6.0-17
limited)
coefficient
5.0-3@
upper
and
full
rudder
(A_i_a)
moment
(monitor
to
and
rudder
5.0-51
Rolling
due
effects,
upper
are
(AC, IRL)"
lower
on pages
aeroelastic
FlAP
and
moment
(AC.JIRu)___
AC.j_
due
. -3o6 3
B
I
6-70O0
5.0
Rolling
(Cont'd)
edge
pages
moment
flap
coefficient
failure.
5.0-37
and
ACILa.
due
_sLu_
to
asymmetric
leading
is plotted
on
5.0-38.
P
REV SYM
b
NO
DG-306__3
Vol.
l±
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6-?000
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THI: BOt::IN_ COMPANY
TD
.61
C-R4
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D6-306_3
Vol. II
PAG
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5.O-Z_
REV. D
RICHARDSON
FOSTER
REYISED
LOW
BECK
ODF..GARD
AD
461
C-IR4
ROLLING
EFFECT
THE
MOMENT
OF
BOEING
COEFFICIENT
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COMPANY
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LOW
LOW
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DATI
2"I770]EFFECT
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,
MOMENT
OF
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MOMENT
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FOR
THE
FLAP
BOEING
COEFFIEIENT
OUTI_ARD
EXTENSION
COMPANY
FLAP
f
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LOW
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FLAP
BOEING
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COEFFICIENT
OUTBOARD
RETRACTION
COMPANY
FLAP FAILURE
7,+7
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1ol.
IIdK
AD
461
KINS_A,N
C-R4
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THE
BOEING
COMPANY
TT
PAO,:
5.0- _7
I,(-IE %"',._:_.';.'.,_,REV.
D
THE
BOEING
COMPANY
6.0
YAWING
The
MOMENT
COEFFICIENT
dimensionless
given
in
terms
aerodynamic
of
its
yawing
significant
moment
coefficient
components
by
the
is
equation
below.
At
a given
O(w.mp ")
+ A Cn _,L_s
•4- ACn
+ _¢_ o_
+ aCn z.1_a_
Ru.oru,,_
+ ACn._.t.
,_RL
,Aiu.w_z'_
"K"
where,
=
dJ
Yawing
The
moment
complete
coefficient
expression
due
for
to
angle
_Cn
is
of
sideslip,
given
_.
as
d_
follows
where
moment
:
C_d__)is
the
coefficient
basic
due
rate
to
of
angle
change
of
of yawing
sideslip.
For
lowspeed,
fdC__
isplotted
onpage
6.O-6.
me
_dP/
aeroelastic
effect,
C_P_M
is plotted
on
page
(c.,),.o'
6.0-6.
as
NOT
I
Ii':NASA
follows
The
ground
effect
factor,
_n_t
is obtained
:
SI__!UI_%'!
z0_,
O
R EV SYM B
NO,
D6-3o6h.3
Vol.
II
PAaz 6.0-i
s-7_
6.0
(Cont'd)
Fn_s
where
_
the
plotted
Fn_,"
ground
on
effect
page
sideslip
6.0-'I.
K6E Is
factor,
•
.:.3
The
is
factor,
ground
plotted
Fn_
effect
on
page
•
is
height
dC v
2.0-_|.
is
ap
obtained
from
Yawing
page
moment
7.0-I.
coefficient
sideslip
angle.
is
as
The
due
to
complete
rate
of
expression
change
of
for
_p
given
follows:
fJ_%
where
I_--_J
aeroelastic_;ll
is
plotted
effect,
on
page
_
6.0-B.
: is
The
plotted
on
page
6.0-8.
Yawing
moment
stability
coefficient
axis,
X_.
due
to
(_Cn
rollrate
about
is
plotted
yaw
rate
the
on
page
about
the
6.0-9.
Yawing
moment
stability
dC_
coefficient
axis,
is
given
Z s
as
due
The
to
complete
expression
for
follows:
dP
where
"
%_aCn_L_
--
and
the
center
of
gravity
factor,
K e
,
_dr/
REV SYM A"
_J_,_
NO. _oi.
11
jp*_ -3o64_3
6.0-2
.._
l-
6.0
are
(Cont'd)
plotted
on
(Cn_)M
page
is
(Cn )
Yawing
6.0-IO.
plotted
The
on
page
aeroelastic
effect,
6.0-_.
'
moment
coefficient
due
to spoiler
deflection.
%_muu_
where
due
_Cn_4%
to
deflecting
on
spoiler
page
plotted
is
6.0-Ii.
on
The
page
as
the
factor,
_
- Yawing
The
ground
KGZ
moment
6.0-I_'.
(K_n
, is
effect,
ground
to
plotted
_
effect
, is
factor,
FI%GE
t
follows:
B
on
number
panels
on page
factor,
6.0-1B.
coefficient
spoiler
plotted
Mach
Fn.K
plotted
moment
operating
effectiveness
obtained
where
_Cn_
the
yawing
is
FnG,-
is
the
{_Cn_4_
45 °.
The
is
]
effect
lateral
page
6.0-14.
,
plotted
is
coefficient
control
The
on
due
ground
page
to
factor,
effect
_n
i
height
2.0-_I.
inboard
aileron
deflection.
where
(ACI%AX)_
is. the'yawing
moment--coefflclent.:-
o
due
I
to
deflecting
one
inboard
aileron
up
to
20 ° or
I
Q
REV SYM A
_'JMM_
INO. vo i.
Ii
I PA°'6" 0-3
9-
6.0
°
(Cont' d)
is plotted
on
page
effectiveness
5.0-=.
6.0-#S.
K_A
factor,
The
Mach
The
number
z j
inboard
is
effect,
aileron
plotted
on
page
(C'n&z)_sis
plotted
(c.,,)...
on page
is
AC'n
OUTIIOARO
6.0-15.
obtained
==Yawing
The
from
moment
ground
page
effect
factor,
Fn6E
6.0-14_.
coefficient
due
to
outboard
aileron
AILIrNOkrl
deflection.
_'CnourrmoAR
D =_--_'_
AII-ERON_
K_o
.
&C, nA
yawi_L_
moment
° - Fnel
[
1.1Pr ARO Ittret4T
OU'flIO_lID
A|iJ[NONm
GC.rtAO
where
to
is
deflecting
opposite
one
on
page
effectiveness
15"0--_"
The
from
AC
n Ruootws
page
ACnRuooE_
moment
upper
aileron
aileron
down
6.0-I_.
ground
The
K_A
factor,
effect
coefficient
up
to
to
the
15 °. G_hAois
outboard
O _
25 ° or
due
is
factor,
aileron
plotted
FnO
on
L , is
page
obtained
6.0-_.
-----Yawing moment
where
outboard
outboard
plotted
the
coefficient
due
to
rudder
= K_Ru'IACnRu_'_q"
(AC_rt.u)lt$
and
coefficients
rudder
and
the
K_ R -{ACn
I_.nRL)I_
due
deflection.
to
lower
are
full
the
*_
yawing
deflection
rudder
_
of
the
respectively.
Q
!
REV SYM
D
BO,
E'INO
I .o. Vol.
II
[PA=Z 6.0--4
6-700
(" )
6.0
CrlRu
(Cont'd)
(,,:o.,.).
_
6.0-19.
The
and
the
are
plotted
upper
lower
rudder
on
The
on
effectiveness
6.6-|7_
, for
page
factor,
factor,
and
aeroelastic
)M
(CnRL
•plotted
effectiveness
page
respectively.
and
rudder
are
on
H_Ro'
_R
page
L )
6.0-18
effects,
the
upper
rudder
and
the
olower
(Crt._M.o
rudder,
are
plotted
on
It
should
respectively.
deflection
with
rudder
one
_Ap
FA_U._=
=
Yawing
be
is
obtained
FA,U.JR¢ == ACn
A_n
'NB_
coefficient
due
flap
extension
or
6.0-Z5
ACnoov_D
L.[. FAI I.IL,
IRE.
flap
6.0-ZA
D
to
outboard
an
page
to
an
flap
is
the
retraction
is
rudder
appropriate
by
1.12.
?.0-3.
asymmetric
(monitor
failure.
OR
yawing
asymmetric
ACn
for
6.O-El
the
,NSO ra,,u,m.
failure.
ACnourm
moment
(monitor
limited)
INS'O V_ILUR_
plotted
o a_=ua,ae
for
on pages
flap
6.0-EE
respectively.
on pages
moment
failure.
and
due
FA,I.URI_
_ILURa
plotted
--Yawing
or
page
multiplied
from
coefficient
ACnFLAP
inboard
A_'I
that
should
inboard
is
noted
contribution
limited)
and
be
inoperative,
moment
where
and
6.0-ZO
rudder
A_yRu_
ACn
page
for
6.0-Z@
coefficient
_Cn
flap
L I"
extension
and
due
FAILAJR
to
6.0-Z5
or
retraction
respectively.
asymmetric
leading
is
on
plotted
edge
pages
6.0-ZT.
!
REV SYM
B
_D6-306M3
,,AOE 6.0-5
6-7000
Ilrl
.Trr
2"-
:2:
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MOMENT
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J_R
OF
COEFFICIENT
51DESLIP
i
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INI_ ODE6ARO
AD
4111
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THE
BOEING
COMPANY
C-R4
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® .....
ALIANI'N[
l_ik
REV. D
REVISED
YAWING
MOMENT
COEFFICIENT
747
LOW
GROUND
EFFECT
THE
SIDESLIP
BO£1NG
FAL-"TOR,
Fne
COMPANY
6.0-7
PAGE
TD
461
C_R4
R_V, D
I
YAWING
MOMENT
EFFECT
THE
"° "" :""
BOEING
COEFFICIENT
OF
,_
COMPANY
K-E _"3:.',.,';.'_, REV. O
RICHARDSON
CURNUTT
YAWING
MOMENT
EFFECT
ODF..GARD
AD
46|
C-R4
THE
OF
BOEING
74-7
COEFFICIENT
ROLL
RATE
COMPANY
_o_
® ........lg&k.
ALliAN[N[
REV. D
OF
THE
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'D6-306_ 3
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_,.0-
_7
7.0
SIDE FORCE
COEFFICIENT
The dimensionless
terms
aerodynamic
of its significant
At a given
O(W._p.
side force
components
coefficient
is given
by the equation
in
below.
•
C¥
"I" AC.y,A=Oll.[m
-t- dC,,y. Iosb
d_
ZV
+
+
+[_AC'yI'L_p
ACYRuooFJ_S
dC_y. r,_b
d---F ZV
r-all.UmF.
-W-I" AC, ya..E. FA,LUR:_
where,
= Side
force coefficient
The complete
of sideslip, _.
due to angle
expression
for
_Cy_y
is
given
as
d_
follows:
(_C-YI is the basic
where ___j
rate
coefficient
of sideslip.
{dc
l is
_--_--/
effect.
effect
due to angle
plotted
(Cy_)_
of main
on page
of change
7.0-5.
, is plotted
and nose gear
of side
For
force
low speed,
The aeroelastic
on page
7.0-5.
extension,
A
The
cl(_-_ul_o,_
is given by:
(.dCy_
HOT !1'_NASA
%.%
REV SYM
B
=
-O.OO&
P_
ou._.F.
S_,'ZT!_TION
j
_-30643
,o Vol. YI
PA_e
7-O-I
.
6-70_
7.0
The landing
( Cont ' d)
plotted
on page
The ground
where
gear effectiveness
effect
on page
factor, K_
dCy.
psb
•
-- Side force
K_
is
2.0-_.
the ground
plotted
factor,
factor,
effect
Fy_
sideslip
7.0-@.
is obtained
as follows:
factor, _
, is
The ground
is plotted
coefficient
effect
on page 2.0-51
due to roll
height
.
rate about
the
ZV
dCv
axis, x$ • d-_
stability
dCv.
r'$b
zv
=
Side
force coefficient
axis,
stability
is plotted
7.0-7.
due to yaw rate about
.
Z$
on page
The
complete
the
.expression
for
dCy
as follows:
is given
=
LdCY'._
where
(cve)..o
is plotted
on paEe
7.0-8.
The
U --I
aeroelastic
effect,
, is
plotted
on page
7.0-9.
ACYs_,_e._
----Side force
where
coefficient
(ACre)4
the
deflecting
(ACYm-)
4s
6 is
due to spoiler
the
operating
is plotted
side
force
spoiler
on page
deflection.
coefficient
to
panels
7.0-II
due
to
_'.
.
The spoiler
D
REV SYM
B
NO. Vol. Yl
IpAo, D6-306_43
7.0-2
->6-7000
7.0
effectiveness
(Cont'd)
7.0-10.
factor,
The Hach
plotted
on page
number
=
Side
force
, is plotted
effect,
7.0-IE.
Fne e , is obtained
ACYRuooe_
(Kst_D) y
(Cy_)_
The ground
from page
coefficient
on page
, is
effect
factor,
6.0-14.
due to rudder
deflection.
YRU M-O
where
(ACYRu) _
coefficients
rudder
and
(ACYRk)
are
=
due to full deflection
and the lower
rudder
the side
force
of the upper
respectively.
(Z_CYRu) zS
and (ACy,
L)m areplotted
onpage7.0-,3.The
upper
rudder effectiveness
factor,
lower
rudder effectiveness
factor,
on pages
elastic
_CYF_P
FAlaUnE =
6.0-17
and 6.0-18
effects,
upper
rudder
pages
7.0-14
Side force
limited)
where
and
coefficient
inboard
ACy
i
retraction
REV SYM D
is
plotted
the
• are
plotted
The
for
the
on
respectively.
(monitor
flap failure.
is the side force
(monitor
limited)
coefficient
izboard
I'AILURF.. for flap extension
on
aero-
FYRL) M-O
are plotted
due to an asymmetric
|NB'0 PAtLURt
&CYINB'O
, and
FYRL)M
rudder,
or outboard
due to an asymmetric
failure.
and
FY*u)M.O
the lower
7.0-15
K_R L
respectively.
(CYRu)M
and
P_Ru
pages
7.0-1_
az_
B_JAW_
flap
or
7.0-17
Vol.
._E
II
7.0-3
6-7OOO
"7.0
respectively.
(Cont'd)
ACy
LL.
FAII-UI_
_CYotrr_D
FAILUP4L
is plotted
on pages
= Side
force
for flap extension
7.0-18
coefficient
flap failure.
_CYL_.
and 7.0-19
or retraction
respectively.
due to asymmetric
leadingedge
FAI,,,_ is plotted
on pages
7.0-_0
And ?.O-Zl.
m
D
REV SYM
B
_B'_'_J_If_
No.Vol.
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7.0--I_
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BIDE
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V4V
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PAGE
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THE BOEING
COMPANY
'1'.0-2.1
8.0
LONGITUDINAL
A
general
in
A
the
diagram
of the
to
the
program
is
Control
the
longitudinal
on Pages
of
particular
jacent
8.1
description
force
each
SYSTEM
Introduction
block
stick
CONTROL
block
1.2-1
simulated
shown
Column
Force
and
Elevator
travel
in
the
747
The
column
travel
in
the
FSAA
was
scaled
column
FSAA
in
the
and
in Volume
control
8.1-3
.
the
page
on
is presented
system
The
data
numbers
I.
and
for
ad-
block.
column
the
1.2-2
elevator
found
The
of
and
on Page
can be
system
FSAA
up by
resulted
Deflection
is 12.67 ° pull
is + ll ° .
1.15
(=
and
The
-12.5 ° push.
column
12.67/]_I) so
in maximum
column
in
deflection
that
the
maximum
simulated
7_7.
The
feel
torque,
unit
pressure,
used
in
determining
the
feel
unit
is
Pf - ___Fs
. (-12o.9) lb/in
_e
2
where
dFs
= [-.0025.q
c + Fs]
>i mc
d6e
Pf
=
Feel
unit
pressure,
Ib/in 2
dFs
=
Control
"Fs"
=
Column
force gradient
at q_ =
This is plotted
on Page 8.5-7
"FQC"
=
column
force
gradient,
ib/deg
0.
Column
force gradient
qa limit.
This is plotted
on Page-8.1-8
D6-3o643
BO_'JN_
REV LTR:
E-3033
RI
ISECT
I N O.
Vo i • I I
IPAGE _,.I - '
The stick force due to the column mass unbalance
FSmass
unbalance
The FSAA
plus
=
control
a constant
nz_
loader
B +
_column-5.4! ._
was programmed
gradient
for zero
control
loader preload
plotter
are shown on Page 8.1-9.
were
subtracted
digital
computer.
incremental
computed
(computed
force
stick force.
recorded
The
by an x-y
The preload
characteristics
stick
on the NASA
The control loader
stick force
(-.275)lb
with a breakout
characteristics
from the computed
is:
force
was
commanded
stick force
with an
minus
preload
stick force).
BOfl/VO
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8.'2
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Control
8.2.1
Elevator
Limits
The
maximum
p_ges
8.2-2
S_'sta_
- Boost
inboard
and
On
and
8.2-3
and
Off
outboard
for
full
elevator
limits
arid half
boost
surface
with
are
sho_n
on
operation,
respectively.
At. inboard
off
wi!l
or
outboard
trail
at
the
elevator
float
angles
shown
on
both
pages
hydraulic
systems
8.2-4
8.2-5,
and
respectively.
8.2.2
Elevotor
Ri_ging
The
elevator
is
The
outboard
elevator
angle
up
to
the
downrigged
blowdown
at +2 ° from
angle
angle
is
equal
of
the
the
to the
outboard
faired
position.
inboard
elevator
elevator.
O
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O
FOSTE.R
HOLTZNER
LONGITUDINAL
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461
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THE
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FLOAT
ANGLES
COMPANY
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REV. D
ALlAN(N(
LONGITUDINAL
ELEVATOR
THE
BOEING
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FLOAT
COMPANY
ANGLE5
8.3
Stabilizer
A b!ocX
8.3-1.
Page
The
8.3-2
S_rste_._
v
diagram
stabilizer
shows
Trim
the
of
trim
_ilot
_
the
rate
and
si;_ulated
trim
is programmed
sys_e:_
by
is
impact
sho_
on Page
_ressure_
qc"
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COMPUTER
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PILOT
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--_l_. PILOT
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POSITION
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LIMIT
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•
BY COMPUTER
STABILIZER
TRIM
PROGRAM
TRIM
m4_p_='fN_
REV LT,R.
E-3033
RI
ISFCT
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m_mmu_
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I
Vol. Ii
'IPAGE S.'-'
I-IOLTZNER
_ER
PAGE
THI:
BOEING
COMPANY
B.3-?_
R_V. O
9.0
LATERAL
CONTROL
A general
duction
block
SYSTEM
description
of the system
on Pages 1.2-2,
diagram
1.2-3
The data
be found on the page
requirements
an equivalent
spoiler
panel.
rather
9.1
Control
Wheel Force
9.1.1
Control
Wheel
Fw
=
were
and
rather
aileron
was a function
reduced
A
block
is
can
The comby applying
than to each
lockout
program
of flap screw
of time,
I.
system
to the block.
for the simulation
than a function
control
for each particular
number adjacent
The outboard
in the Intro-
and in Volume
lateral
rate limit to the wheel
in the simulation
9.2-8,
and 1.2-4
of the NASA simulated
shown on Page 9.1-2a.
puter
is presented
as shown
utilized
travel,
on Page
Page
9.2-4.
An_le
Force
Fw$
f
F'wF _
where,
Fw
=
Control
wheel
Fw 5
=
wheel
moment
Control
Friction
positive
for a clockwise
(ib).
wheel
mechanism
=
force,
force
(lb).
force
due to the spring and cam
This
is plotted
opposing
control
on page 9.1-3.
wheel
motion
(= 2.0 Ib).
,m'OEfMMG
i
REV LTR:
E'3033
RT
I SECT
N O. Vo _.
z7
i
I
b5-3o643
PAGE
9.l-1
9.1.2
Control
_w
Wheel Angle
"=
6warn,.
+
"35Fw
whe_
Control
Sw
limits
wheel
(degrees).
are '6 w : ± 88 degrees.
Reference
control
not include
input
angle
wheel
the effect
to the aileron
The control
wheel
"
angle
(degrees).
of cable
stretch
and spoiler
This
does
and la the
programs
plotted
on
pages 9.2-3 and 9.2-5 respectivel_.
.35
9.1.3
Iateral
:
Cable
stretch
factor
(deg/lb).
Trim
Zero
force datum
This
is plotted
the wheel
wheel
force
limits.
The nominal
wheel
wheel
angle
due to trim
on page 9.1-4.
datum
trim rate
The trim shifts
but does not charge
The trim limits
IE-3033
R1
the
are ± 6.27units.
is 2.5 degrees
of control
per second.
_I'_EJAV'_
REV LT'P-,:
(degrees).
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CONTROL
TRIM,
APR
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461
ODE_ARD
Z.IO'70
THE
BOEING
PA_E
COMPANY
9.1-4-
C-R4
®
.........
REV.
F...
9.2
Aileron-Spoiler-Wheel
9.2.1
Iateral
Control
The inboard
Program
On_
aileron
wheel
program
is plotted
For fully unlocked
outboard
program
on pa6e 9.2- 3 .
is plotted
is available
Outboard
aileron
extension
about
at all times
ailerons,
unlocking
After
is available.
Outboard
flap
retraction
jackscrew
locking
takes about
aileron
authority
travel
this
time,
aileron
outboard
is available.
aileron
at outboard
flap
and full unlocking
full outboard
locking
wheel
authority
After
time,
For
intermediate
angle
(degrees).
takes
authority
at outboard
travel
this
jackscrew
aileron
is started
to O_ of jackscrew
15 seconds.
aileron
5, I0, 20, 25 and 30.
is started
to 09 of jackscrew
15 seconds.
the outboard
Full
for flaps
on page 9.2-3.
and full
no outboard
outboard
aileron
authority
"
o " aOmu, "
where,
=
Outboard
aileron
aileron
mechanical
limits
The outboard
are 25 e T.E.
up and 15 e
T.E. down.
=
Intermediate
outboard
is a function
of flap
aileron
screw
gain factor.
travel
This
and is_plotted
on Page 9.2-4.
O
!
D6-30643
REV SYM
D
_t_'J_
.o. Vol.
II
PAOE 9.2-I
9-
_AORzF.
=
Reference
from
In
the
fully
The
9.2.2
NASA
intermediate
the
spoiler
progran_
Page
ailerons
are
locked
page
9.2-_.
commanded
9.2-3.
out
when
flaps
are
speed
and
Operation
remains
normal
program
brake
the
same
page
9.2-7.
used
Inflight
speed
as
handle
for
as
ground
Spoiler
speed
either
position
speed
is plotted
of
brakes
all
on
brake
speed
brake
handle
positions,
9.2-6.
page
the
speed
speed
brake
3,
4,
brakes
7,
cannot
or
control
5,
remaining
6 and
on_y,
extended
lateral
6,
7,
(intermediate
The
panels
brakes
fully
no
Pan@is
positions).
only
as
inflight
programwith
on
Cally
Brake
on
Operation
position).
D
program
wheel
Brake
operate
SYM
(degrees)
wheel
is plotted
Speed
plotted
are
angle
positions.
brake
REV
With
wheel
For
The
-
outboard
program
Control
aileron
Speed
@
simulation
spoiler-wheel
handle
9.2.3
the aileron
aileron
retracted.
Lateral
The
outboard
fully
and
(ground
be
8,
panels
which
inputs
9,
are
inflight
speed
speed
Ii,
brake
operate
retracted
i0
I, 2,
modulated.
is
12
handle
symmetriThese
depending
surfaces
on
the
handle.
Ba,
C=',/'N_
n6-3o6_3
NO. VOI.
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9.2-2
IT
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nun
DATE
cHr_K
APR
APR
PAGE
THE
AD
,461
C-R4
BOEING
COMPANY
_',9.-
FLIG_'T
U
THI:
TD
4&l
C-R4
BOEING
COMPANY
SEE SECTION
FOR REVISED
19
DATA
PIOOIWEER
LATERAL
KUPCI5
KUF_.IS
L---
REVIS£D
5POLLER
PRO_RAH
LATERAL
CONTROL-
THE BOEING
CONTROL.
AT
747
C.X:)MBINED
SPEED
COMPANY
BRN_F..._
Vol.
iI
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TD
461
C_lt4
THE
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COMPANY
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MODEL
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n
9.3
Control
Surface
The maximum
Limits and Float
inboard
aileron
and maximum
for full boost and half boost
9.3-3
respectively.
Angles
outboard
are plotted
The maximum
spoiler
aileron
on pages
angles
angles
9.3-2
and
are plotted
on
pa_es 9.3-_, 9.3-5 and 9.3-6.
Any inboard
systems
9.3-7
faired
REV SYM
D
or outboard
off, will
aileron
trail at the float
and 9.3-8 respectively.
position,
surface
boost
with both hydraulic
angles
The spoilers
plotted
are held
off, due to the hold down
BOfJ_
on pages
in a
check
VO
valves.
NO. Vol. II
j PA_EI_5-30_3
9.3-1
HOL'I"ZNER
FOSTER
INI_)/_RD
AILERON
THE
BOEING
E$LOWDOWN
COMPANY
, •
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FOSTER
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BOEING
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pages
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inoperative.
coefficient
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fo_
10.2-3;'
any
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flight
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c..
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rudders
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12.2
Engine
Thrust
A block
Page
The
diagram
of
the
engine
thrust
simulation
is
shown
12.2-2.
reverse
thrust
were
a function
This
data
was
data
as
function
12.2-7
a
and
of
characteristics
ambient
based
12.2-8
pressure
on a V_ch
of
V_ch
used
and
number
number
is
of
in
the
simulation
EPR,
Page
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Reverse
sho%_
on
12.2-5
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Pages
12.2-6
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THE
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12.3
Engine
Cockpit
Display
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12.3-2
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Parameters
for engine display
diagram
lation
is shown on Page 12.3-
scaled
up by a factor
ature
is shown
for the NASA engine
3.
The compressor
of 8.8 to obtain
in the simulation.
accuracy
parameters
The table
and the data points used
gauge
simu-
RPM was
the exhaust
gas temper-
on Page 12.3-4
in determining
on Page
shows
the
the scale
factor.
BOEI_
REV LTR:
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R1
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N O . vol . II
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Windmillin 6 Dra 6
The windmilling
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as a term in the engine
were
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equations.
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THE BOEING COMPANY
.
I_,.q-_.
_
12.5
Thrust
Reverser
The thrust
moment
Effects
reverser
coefficients
approximations
computer
used
storage
are recommended
approximately
on Aerod_uamic
effects
Coefficients
on the lift,
are presented
on Page
in the simulation
and j_stified
because
drag and pitching
12.5-2
were made
.
The
to conserve
the thrust
to be in the Idle reverse position
reversers
by
60 knots.
D6-30643
REV LTR:
1_-3033
RT
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12.5-1
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LANDING
GEAR
A general
included
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in Volume
of the landing
to Volume
Block
NASA
and landing
I of this report.
gear equations
gear
A detailed
is contained
equations
is
derivation
in the Appendix
I.
diagrams
pression
of system
showing
and force
simulation
Refer to Volume
the method
and wheel
of calculating
side and drag
are shown on Pages
I for the landing
13.0-2
forces
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strut
com-
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13.0-5.
gear nomenclature.
D6-306h_3
REV LTR:
_-3033
lq I
I SECT
IPAGE
13.0-1
CG
DISTANCE
FROM CG
TO LEFT MAIN GEAR
XL2 = -7+.01
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DISTANCE
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C
XL1 = 77+ .01 (CG-25)
DISTANCE
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14.1.1
COCKPIT
INSTH_vIENTS
The cockpit
the piloted
checkout
checks were made
of the following
Rate-of-climb-indicator
3.
Airspeed
indicator
4.
Attitude
indicator
5.
Turn and bank
6.
_ch
7.
Flap position
8.
Stabilizer
readings
smoothly
tabulated
during
parison
between
largest
discrepancy.
to a fraction
specified
Rudimentary
indicator
position
with
the computer
the computer values.
during
the simulation
with
values
the four engine
checks
pitch
attitude
The differences
and comparing
comparisons
checkout
values.
The instruevaluation.
instrument
test.
attitude
readings
The com-
shows
is difficult
are within
were
and all of the
and piloted
and cockpit
and indicated
on Page
These
climb performance
However 3 the pitch
of a degree.
in "hold"
the computed
the dynamic
of digital
the computed
in the table
indicator
by putting
in good agreement
A comparison
were
that
indicator
of times throughout
operated
confidence.
to assure
meter
conducted
were
with
checked
instruments:
2.
made a number
statically
could be conducted
Altimeter
the instrument
ments
were
1.
The checks were
instruments
instruments
the
to read
the tolerances
14.2-10.
NO. Vol. I_
I SECT
i PAGE
I_.I ' 2.
The stabilizer
using
stabilizer
of the aerodynamic
dicator
referenced
indicator
to the fuselage
data are in terms
and stabilizer
D6-30833-1
position
information
is in stabilizer
of_FR
used
in the FSAA _s
reference
programmed
line, _RL"
All
The 747 stabilizer
L.
in the flight
"pilot's units", _L .
P
manual
in-
and
The conversion
is
given by:
= 3° "_FRL
•
-
__J_
I SECT
IN O.
_6_3o_I13
Vol.
IPAGE _.,
-3
I_. I, 2
A_,_0SPHERE MODEL
The atmosphere
_as
checked
for a number
of V e for a number
values
model3
presented
of altitude
of altitudes
and airspeed
were input
of M, V c (VI) , V, g . and qc were
table on Page
in Volume
I of this
conditions.
to the computer.
checked
report 3
Values
Output
_rith the data in the
14.1-5.
l SECT
le
2.
REV SYM
Standard Day, No
Pilot's Airspeed
Wind
Vl is equal
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¥
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14.1.3
kW_GINE CHARACTERISTICS
Static
by comparing
same
forward
and reverse
the computer
condition.
output
The comparisons
thrust
characteristics
data with a manual
are shown
were
checked
calculation
in the examples
of the
on Pages
14.1-7
and -8.
The transient
and -10.
_ere
mates
A
sea level
commanded.
the engine
engine
static
The resulting
transient
characteristics
condition
are shown
on Pages
was set up and throttle
transient
characteristics
data shown in Section
14.1-9
changes
for EPR approxi-
12.
__fMM_
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NO.
I
Vol. II
D6-306k3
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14.1.4
LONGI_n_JDINAL TR_.I
The simulated
of weight,
eB,
c.g., altitude,
and EPR/F n were
Comparisons
on Pages
results
are plotted
results
are tabulated
-18.
and flight
RT
values
of_,
test results.
14.1-12
to flight
of configuration
changes,
computed
landing
and -13 and
test and simulator
gear up and do}m
and the results
are tabulated
14.1-22.
of ground
effect on trim is tabulated
effect
on Page
curves on Pages
14.1-12.
2.0-31,
-32,
-18 and _.0-35.
,_'O_'IN_
E-_033
Computed
variations
thru -21.
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REV LTR:
at several
on Pages
Comparisons
on Pages 14.1-19
up and down, were
The effect
3.0-17,
trimmed
to Boeing simulator
14.1-14-thru
The effects
and speed brakes
vas statically
speed, and flap positions.
compared
to simulator
plotted
on Page
airplane
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14.1.7
FLIGHT
14.1.7.1
Force
CONTROLS
and Displacement
The force and displacement
shown on Pages
14.1-52
control
in the FSAA
loader
cozpared
to the force
and 14.1-53.
cab.
Wheel
and rudder forces
These
to the force
were made
to verify
initially
A condition
e relationship
tabulated
comparisons
14.1-5_,
55.
placement
The data
also
obtained
x-y plotter.
Page
and _col
force
on Page 10.1-3.
in the computer
loader
forces
and a modified
to augment
control
the constant
loader.
Tests
to check
the _col
on Page 14.1-56.
obtained
to a position
were obtained.
are tabulated
to
for two c.g. positions.
from the column control
The results
the
and cab column force.
cab _ras also
column _ras deflected
of stick force, _e,
The rudder
Boeing and NASA simulations
is plotted
data in the FSAA
The control
were
between
from
in the computer.
was set up in the computer
and the resulting
obtained
in the cab should be
set in the column
the computed
and rudder are
characteristics
were not calculated
was input to the control
stick force gradient
force
on Page 9.1-3.
The column force was calculated
force increment
for the wheel
curves were
The wheel
characteristics
in the cab should be compared
curves
are on Pages
Column
for this
force
loader
on Page
stick force
through
14.1-55
and dis-
condition.
and the computed
Column
The
values
and _col
a calibrated
and plotted
on
14.1-57.
stabilizer
plotted
Similar
comp_ted
trade
condition.
stick
force data was obtained
The results
are tabulated
from the elevator
on Page 14.1-54
and
on Page 14. l- 58.
_-3o6_3
REV LTR:
E-3033
RI
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PAGE
I _..I - go
_neuvering
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force data was obtained
during
the piloted
checkout.
4
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14.1.7.2
RIGGING
The lateral
table on Page
results
14.1-60
for various
control
rigging was
shows a comparison
combinations
of wheel
che eked at high
of Boeing
speed.
and NASA
and speed brake
The
simulation
inputs.
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IA.I.7.3
BLOWDOWN
The
limit
value
blowdown
for
rudder
simulation
checkout.
data
for
point
spoiler
A
is a
comparison
between
Boeing
and
was
verified
at
elevator
blowdown
spoiler
different
stabilizer
and
and
the
times
trade
lateral
the
ratio
changer
throughout
test
provided
rigging
check
the
a
provided
data.
further
accelerations
elevator
The
elevator
blowdown
for
with
of
and
check
full
on
blowdown
control
initial
_as
a
comparison
inputs.
The
tabulation
accelerations
and
control
of
initial
on Page
surface
14.1-62
positions
IIASA simulations.
D6-30643
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14.1.8
LA_,TDINGGEAR
The landing
to settle
gear was
on the ground
Dynamic
responses
were
initial
conditions.
static checked by allowing
and verifying
made by releasing
of the computer
the landing
simulation
gear
frame time of 39 milliseconds.
the response
to a _h
initial
The increase
in a smoother
than
lecting
the limit
to have
the same frequency
limitations
condition
compression.
from _h and _
and the overall
was run at a computational
The time histories
in computation
response
of strut
the airplane
In consideration
simulation,
times.
the amount
the airplane
on Page
for two different
speed,
the "normal"
frame
frame
time
14.1-64
computation
to 13 ms, results
time of 39 ms.
cycle, the response with the 39 ms frame
and slightly
increased
shows
damping
Neg-
time appears
over the 13 ms
re sponse.
The time histories
initial
the _
condition.
The comments
the frame
of slowing
of the limit
realistic
though
cycle increased
to the Boeing
a limit
was high
enough
by the pilot.
RT
response
to a A_
are applicable
to
14.1-66.
as the computation
inthe
limit
speed
The amplidecreased.
gear deflections
moving
in the computed
of the computed
speed 3 or increasing
on Page
due to landing
test pilot
cycle existed
time the amplitude
E-3033
the response
response
the computation
time, are shown in the responses
The aircraft
REV LTR:
on the _h
shows
responses.
The effects
tude
on Page 14.1-65
base
response
felt
simulation.
with
Even
a 39 ms frame
cycle was low and the frequency
(3 cps) that no objectionable
characteristics
_'0_'_
I SECT
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PAGE
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I PAGE _6-306__3
Ii.I.. I- GG
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14.2
PILOTED
CHECKOUT
The piloted
simulation
project
as well
pilot,
3 sessions.
1.
as quantitative
flew the motion
He qualitatively
Airplane
simulator
evaluated
handling
Dutch
b.
Spiral mode
c.
Short period
d.
Phugoid
e.
Roll
f.
Climb performance
g.
Flap
h.
Speed brakes
Jack Waddell 3 Boeing
for a total
747
of 5 hours
the following
in
characteristics:
mode
mode
rate
extension
and retraction
response
3.
Ground
effect
4.
Control
5-
Take off (3 and 4 engine)
6.
Landing
7.
stall
8.
Air minimum
9-
Buffet
and landing
of the
roll mode
Engine
The ground
assessment
characteristics
a.
Stick
an overall
data.
2.
10.
approach
checkout provided
forces
control
speed
shaker
effect
characteristics
evaluation.
The ground
were modified
effect
during
data on Pages
I
the
2.0-31_
-32,
z_-3o643
INo. vol. zz
REV LTR:
E-3033
R I
SECT
IPAGE
_.
2. - I
3.0-17,
The
-18,
_.0-35
were
•7 times
the
pitching
•3 times
the
drag
increment
•9
the
lift
increment
above
times
changes
additional
Boeing
revisions
the
felt
location
simulator
due
to
were
in
the
roll
above
6 feet
the
the
pilot
the
stick
747
in Section
19.
shaker
c.g.
(i0
simulator
location
in
in the
feet)
drive
comparable
to
factors:
were
substantiated
simulation.
and
pilot.
The
frequency
His
by
overs_l
resulting
characteristics
comments
sub-
simulation.
resulted
than
following
amplitude
of the
the
inputs
data
Boeing
satisfaction
of
the
increment
effect
included
greater
motion
the
on
validity
Rapid
pilot
testing
and
by
moment
ground
are
to
modified
the
buffet
tailored
stantiated
of
to
piloted
The
were
and
The
reduced
equations.
the
acceleration
airplane.
w_s
the
above
lateral
to an
The
airplane.
c.g.
pilots
felt
lateral
is attributed
to
the
vertical
effective
pilot
The
which
distance
the
resulting
acceleration
aircraft
flexibility.
Quantitative
pages
contain
the
data
results
were
of
the
obtained
piloted
with
NASA
pilots.
The
following
tests.
_O.E'INO
I SECT
I N O.
I
D6-30643
Vol. II
PAGE i_.2-2
I$. 2.1
TAKEOFF
Two
verify
the
istics,
takeoff
thrust
efficient
takeoff
time
while
Boeing
the
thrust.
from
.
Boeing
gross
The
ground
weights
were
simulation
effect
inertia
be
compare
added
to
average
the
curves
Velocity
results
obtain
in
effects
rotation
and
E'3033
RT
drag
attitudes_
indirectly
a
and
lift-off
to
character-
rolling
checked
stop
_ratch.
shown
in the
table
started
with
V = 15
kt
the
the
data
two,
NASA
started
the
data.
acceleration.
on Pages
histories
versus
on Pages
The
14.2-7
from
time
to
This
time
the
on
co-
by
timing
= 0
and
-8
(t = 0)
is
Boeing
.
NASA
equal
to
The
takeoff
and
from
acceleration
!_.2-8
the
and
determined
between
i$.2-$
t = 0)
ta/_eoff
0
15
thrust
to
kt
w_s
15 knots
divided
determined
.
I%_SA takeoff
from
Page
accelerate
average
and
data
14.2-7
time
were
Comparisons
(at
at V
speeds
NASA
are
test
shown
are
on
Pages
plotted
14.2-5
with
and
the
.
__JAr_
REV LTR:
performed
of basic
taxi
are
are
takeoff
To
Time
_6
and
to
simulations
takeoff
the
rate,
indicator
simulated
by
acceleration.
lapse
airspeed
EASA
must
different
acceleration.
the
and
at
of friction
The
by
takeoffs
ISECT
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14.2.2
CLI_B PERFOE_NCE
The purpose
simulation.
of this test was
The pilot
started
to verify
the test near
was to fly the air speed and EPR schedule
Page l_.2-10.
At the check altitudes,
and data was obtained
The comparison
on Page
14.2-10
this task
from the cockpit
between
is to have
sea level
prescribed
the computer
instruments
Boeing and I_ASA simulator
and plotted
on Page 14.2-Ii.
the airplane
the climb performance
stabilized
and his task
for this
condition,
was put in "hold"
and the computer.
results
is tabulated
The difficulty
when putting
in performing
the computer
to
"hold" .
_6-3o6;,3
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R T
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14.2.3
IICFLIGHT ACCELERATION
The acceleration
with
maximum
Speed
continuous
tween Boeing
and NASA
starting
on Page 14.2-14.
result
simulations
Y_ch
and altitude
high
altitude
The differences
for the higher
is shown
The NASA
idle thrust,
to the maximum
on Page
The trend
appeared
idle EPR was made in the NASA
after
brakes
results
be-
and
are plotted
data are the
test utilized
a revised
the idle thrust
of the results_
correct.
simulation
speed
and Boeing
affected
speed.
14.2-13.
with
The Boeing
at M = .65
The comparison
simulation
in the NASA
to EPR which
number.
by starting
from the cockpit.
in idle EPR.
correction
and Mach
performed
check was performed
at M = .85.
of a difference
ms
and accelerating
obtained
The deceleration
idle power,
DECELERATIOE
condition
thrust
and time data were
_
at
accounting
The adjustment
the checkout
to the
data were
obtained.
,
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14.2.4
STZADY TU_S
This
test
required
to hold
airplane
for
coordinated
to
in
dition
was
values
of
the
when
is
going
plotting
order
stabilized
flight
to
the
angle 3 or
load
level
check
flight
were
results
rolled
of
results_
the
A
was
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calculated
The
put
in
value
into
and
are
The
data
having
and
a
a
the
allowed
the
the
concomputed
tabulated
are
-18.
was
with
plotted,
The
stabilized
of _Iz = 1/cos_
elevator
steady
was
When
results
14.2-17
and
trimmed
airplane
"hold"
14.2-16.
in
pilot
airspeed.
The
on Pages
difficulty
then
angle.
on Page
force
and
a constant
computer
stick
The
to maintain
3 the
the
factor.
bank
test
"hold".
to
a specified
and _e
a result
conducted
and
simulation
with
scatter
at
Fs, @B'
Boeing
along
a bank
straight
turn
descend
was
data
condition
used
in
data.
D6-3o643
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BOEING
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14.2.5
LONGI_JDINAL
STATIC
The airplane
_as trimmed
out changing
speed using
a rate
the stabilizer
the elevator.
the airplane
column _ras then released
speed.
The results
for straight
or throttle,
and level
the pilot
The thrust discrepancy
of climb or descent.
corded when
STABILITY
Comp_ted
values
_as stabilized
OfSe
on Page
changed
_as allowed
Withthe
to generate
, Fs, ands B were re-
at the prescribed
to see if the airplane
are plotted
slowly
flight.
_uld
speed.
return
The
to the trim
14.2-20.
D6-3o643
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E-3033
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14.2.6
STEADY
SIDESLIPS
After a trimmed
applied
rudder
to hold
flight
condition
a steady sideslip.
trol and bank angle was _sed to maintain
When
the airplane
were
recorded.
was stabilized,
Comparisons
test are shown on Pages 14.2-22
A combination
a straight
computed
bet_een
NASA
was set up3 the pilot
values
of lateral
and level
of_,
and Boeing
E-3033
RI
con-
flight
_R , _w'
simulations
path.
and_B
and flight
and -23.
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D6-306_3
Vol.
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SIDESLIP
COMPANY
14.2.7
ROLL RATES
The pilot
and then
pedals
rolled
trimmed
into a 30 ° bank.
were not used.
input was
rapidly
in the opposite
a stop watch
were
obtained
required
applied
and held until
thru
through
flight
time histories.
simulator
14.2-25
data
a wheel
banked
to 30 °
the 60 ° was obtained
roll rate and magnitude
of wheel
The average
the change in bank angle,
The NASA
and level
_,_s established,
the airplane
The time to bank
in the table on Page
14.2-26
the bank angle
from the computed
by dividing
for straight
The yaw damper was on and the rudder
and the maximum
to bank 60 ° .
results
After
direction.
with
was determined
the airplane
input
roll rate
60 °, by the time
is compared
and to the plotted
to Boeing
data
on Pages
-28.
,/
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ROLL
i.
RATE
CAPABILITY
FL AI::_
IO
_< ;:_:1.
"74"7
I
]
A,,.
:1:_-306_3
'
A_r
"
ODE_ARD
'
5-14-'/0
Vol.
II
PAGE
THE
BOEING
COMPANY
I_._.-;_(=
I
+,
+ .
i_£i ':::i'iW
L_
ROLL
II
INK
AD
4_1
RATE CAPABILITY
I=LAPS 30
THE BOEING COMPANY
14.2.8
AIR IIIND_4
Minimum
CONTROL
control
at _¢hich an airplane
SPEED
speed
in the air is defined
can maintain
failed_
rudder at full
high).
The results
deflection,
from
straight
flight
and bank angle
the NASA test are plotted
_th
at
as the lowest
a critical
5 degrees
on Page
speed
engine
(dead engine
14.2-30.
D6-3o643
14_IIP_IP_I_'_
INO.
VOW"
Z
REV LTR:
E'3033
R|
I SECT
IPAGE ,'-t-._- 2._
15.0
APPE_,_IX A - FLIGHT _I_UAL
Selected
airplane
will
flight
allow
performance
Part
pages
manual
are
the simulation
criteria
from the performance
contained
in this
to be operated
and certification
section
section.
in compliance
requirements
of the FAA approved
This
information
_rith the appropriate
of FAR Part
25 and
36.
D6-3o643
I NO. Vo1. z I
I
_r_'LL"J'_I/_
REV LTR:
E-_033
RI
I SECT
I
PAGE
_.o-
I
.,i/c-,,,_A
APP.,#_,,E_.m,//,,L='.a7
FLIGHTNANUAL
/
--_
'
--
JL-__
_
......
THIS AIRPLANE MUST BE OPERATED IN COMPLIANCE WITH THE
PRESCRIBED CERTIFICATE
LIMITATIONS
IN SECTION 1 HEREIN
.,,00rov,,_,_.,-C;__N"_
Chief, Aircraft Engineering Division,
FAA, WESTERN REGION
Date
PUBLISHED
BY
THE
BOEING
COMPANY,
December
COMMERCIAL
30,
1969
AIRPLANE
GROUP,
RENTON,
WASHINGTON
U.S.A.
D6-306__3 "
NO. Vol. II
-
PERIFOlqMA_(Z
AIRPLANE
SE
C T_
Q_
FLIGHT
_-
MANUAL
P ERgO
RMAN
CE
OENERA_
The information
in this section is presented for compliance with the appropriate
performance
criteria and certification
requirements
of FAR Part 25 and Part 36.
STARDARD
PERFORMANCE
All performance
CC_DITI_S
in this
section
ratings
on the following:
i.
Approved engine
sory losses.
2.
Full temperature
accountability
within the operational
limits,
landing distance, which is based on standard d_7 temperatures.
Trailing edge flaps positions as follows:
3.
thrust
is based
less installation
losses,
Trailin_
Takeoff
Transition
Enroute
Approach
Flap
and acces-
except
for
FJ.aI_O
i0, 20
I,
5
0
20
Setting
Lancling
25, 30
Leading edge devices
position.
Y_
Ed_e
airbleed
FACTORS
in the appropriate
position
for trailing
edge
flap
A_'_'_.'C"-'A_G
F_RFORMA_CE
Details of the variable factors affecting performance
are given under Performance
configuration,
but certain assumptions relating to all Performance
data are as
follows:
ICING PROTECTION
The effect
of anti-iclng
systems
operation
is shown
on applicable
charts.
NJ_IDITY
Humidity has no appreciable
effect on the thrust
has not been considered in the perforaance
data.
of the engines;
therefore,
it
WIND
The wind velocity used in calculations is factored
relevant operating regulations.
All charts should
reported wind components.
to assure compliance with the
be entered with actual tower..........
FAA
+ APPROVED
12"30-69
---
ISECT
_6
IPAGE
m
......
,,
AIRPLANE
FLIGHT
MANUAL
G ERAL
A_SPE_S
All
airspeed
All
Indicated
Position
and Mach
Airspeeds
error
values
in this manual
are based
is the instrument
on normal
indication
assume
a zero
instrument
static
source
position
error
error.
error.
due to location
of static
ports.
Equivalent
Airspeed. EAS - Airspeed indicator reading, as installed in the airplane, corrected for static source position error and compressibility.
Ca_ibrate¢ A_rsoeed. CAS - Airspeed indicator reading,
plane, corrected for static source position error.
as installed
in the air-
;_zlicate_ Airspeed. IAS - Airspeed indicator reading,
plane, uncorrected
for static source position error.
as installed
True Mach
in the airplane,
N_ber.
for static
source
M - Mac_meter
position
reading,
as installed
in the air-
corrected
error.
Critical Enzine Failure Soeed. VI - The speed at which, when an engine failure is
recognized,
the distance to continue the takeoff to a height of 35 feet will not
exceed the usable takeoff distance; or, the distance to bring the airplane to a
full stop will not exceed the accelerate-stop
distance available.
VI must not
be less than the Ground Minimum Control Speed, VMCG, or greater than the rotation
speed,
VR, or greater
Engine
Failure
for actual
than
Soeed Ratio.
runway
dimensions
the maximum
VI/VR
brake
- The ratio
and conditions,
Maximum B_e
Enerav Speed. V_
stop can be accomplished within
energy
speed, VMBE.
of the engine
to the rotation
failure
speed,
speed,
V1,
VR.
- The maximum speed on the ground from which
the energy capabilities
of the brakes.
_'_I'AL"JAI"_I'!
NO.
a
I_-
+
FAA APPROVED
12-30-69
ISECT
IPAGE t .o- @
AIRPLANE
FLIGHT
MANUAL
GENERAL
AIRSPEEDS
Rotation
(Continued )
Speed.
VR - The speed
at which
Takeoff Safe_.y Soeed. V_ - The scheduled
foot height wlth one engine inoperative.
rotation
target
is initiated
speed
during
to be attained
the takeoff.
at the 35
Air Minimum Control speed, VMCA - The minimum flight speed at which the airplane
Is controllable
with a maximum of 5 ° bank when the critical engine suddenly becomes inoperative
with the remaining engines at takeoff thrust.
Ground Minimum Cont_l
Speed, _MCG - The minimum speed on the ground at which the
takeoff can be continued, utilizing aerodynamic
controls alone, when the critical
engine suddenly becomes inoperative with the remaining engines at takeoff thrust.
Landln_ Reference Speed, VREF - The minimum speed at the 50 foot height in a norreal landing.
Thls speed is equal to 1.3 times the stall speed In the landing
configuration.
Desl_n Maneuverin_
available aileron,
Soeed. VA - The maximum
rudder or elevator will
speed at which
not overstress
application
of full
the airplane.
6-3
+
FAA APPROVED
_rJAY_
12-30-69
I SECT
I NO.
IP AGE
Vol.
I_'.0-
II5-
AIRPLANE
FLIGHT
MANUAL
GENERAL
(Continued)
TEMPERA_JRE
ISA - International
Standard
Aviation Organizatlon.
Air Temperature
Atmosphere,
- the
as accepted
free air
static
by the International
(ambient)
Civil
OAT
- Outside
temperature.
SAT
- Static Air Temperature
- outside air (ambient) temperature
as computed by
the Air Data Computer and presented on the Static Air Temperature
indicator.
TAT
- Total Air Temperature
- static air temperature
plus a_llabatic compression
(ram) rise as indicated ca the Total Air Temperature
indicator.
WIND
Wi_d Velocity
- The actual wind velocity at a 50 foot height reported frmm the
tower and corrected by the wind component chart to a headwlnd or tailwlnd component parallel to the flight path.
HEIGHT
9rQss Height - The geometric height attained at any point in the takeoff flight
path using gross climb performance.
Gross height is used for calculating
actual
pressure altitudes at which obstacle clearance procedures and wing fl_p retraction are initiated, and level-off-height
scheduled.
_et Height - The geometric height attained at any point in the takeoff flight
path using net climb performance.
Net height is used to determine the net flight
path which _st
clear any obstacles by at least 35 feet to comply with the regular ions.
IC ING
Icing Conditions
- Icing may develop when visible moisture such as
wet snow is ]_resent with Static Air Temperature
below 8"C (46"F).
fog, rain,
or
TAKEOFF DATA
Balanced
distance
Field L_h
- The condition
equal to the accelerate-stop
where Vl is selected
distance.
to make
Unbalanced
F_eld Length - The condition where Vl is selected
distance and accelerate-stop
distance unequal.
the t_off
to make
the takeoff
t
+
FAA APPROVED
12-30-69
I
I
SECT
I
I No.
PAGE IS.O-
6
AIRPLANE
FUGHT
MANUAL
GENERAL
GRADIenT
Gross
OF CLIMB
Grsd/_D$
- The demonstrated
ratio,
expressed
as a percentage,
of:
Chaa_e i:n Height
Horizontal Dis tauce Traveled
The gradients shown on the charts are true
true, not pressure, rates of climb.
N_t Gradlen_ - The demonstrated
by regulation.
BUFFET
_SET
gross
gradients,
i.e.,
graY,dent reduced
they
are based
by the increment
on
as requlred
CHARACTERISTICS
The buffet boundary is a basic characteris_;ic
of the airplane that is defined by
angle of attack and Mach numbers.
Below aj_proximately 0.85 M, the buffet is related to the wing maximum llft capability
(therefore
to true airplane stall).
In the range above O. 85 M, buffet is rela_i
to the growth of shock waves on
the wing.
At any flight condition, it is possible to
high-speed,
and maneuvering
margins before
Maneuvering
Capability
chart. )
determine the altitude, low-speed,
buffet onset occurs.
(See Cruise
C _ 0 S S W I N D
and Landing)
V A L U E S
The maximum demonstrated
height.
This component
(Takeoff
crosswind componer it is 30 knots reported
is not considered
_ _ be limiting.
wind
at 50 foot
For performance
scheduling,
the corresponding
crosswind
the full head_ "J.nd component may be used provided
component does
not exceed 30 knots.
MINIMUM
L
CONTRO
SPEED
_[
The Air Minim_
Control Speeds (VMCA), and
(VMCG), of this airplane are shown on a chs
Speeds charts where applicable.
+
FAA A/qIROVED
the Ground Minim_
Control Speeds
rt in this Section, and on the takeoff
NA_'_"J,/VLW
12-30-69
that
I SECT
I N O.
IPAGE
V_o_.SO_
_5,o-
7
P_'O_E
80A ING
AIRPLANE
PERFORMANCE
FLIGHT
C OND_TIONS
MANUAL
AND
PROCEDURES
Takeoff field length performance
shown in this section accounts for 115 percent of
the all-engines
operating distance, or the total distance considering an engine
failure recognition at VI, whichever is greater.
These distances are based on a
smooth, dry, hard-surfaced
runway.
The appropriate
was used.
airplane
conflgurati_,
outlined
under
Performance
Conflgurations,
The conditions and procedures used in establishing
the performance
data in this
manual are presented under each phase of operation.
Procedures are guidance
material only.
TAKEOFF
Conditions
Prior to takeoff, a review was m_de of stabilizer and flap settings, takeoff
speeds, and that sufficient field length w_s available for the gross weight and
ambient conditions.
Corrections were applied, when necessary, for significantly
altered ambient conditions or loading.
Procedures
Thrust was set to i.i EPR prior to brake release, or as the airplane was aligned
with the runway.
With no wind, or adirect
headwind, EPR was advanced prior to
brake release.
Thrust levers were adjusted as necessary to obtain target EPR
values by approximately
40 to 80knots
on the takeoff roll.
Rudder
pedals
were
used
for directional
control
through
the nose
wheel
and rudder.
Rotation to takeoff attitude was initiated at VR.
A speed not less than V2 was
obtained at a height of 35 feet.
The landing gear was retracted after a positive
rate of climb was established.
A smooth positive rotation was used to the initial climb attitude (approximately
13 to 22 degrees depending upon gross weight and thrust available).
Minor attitude variations were made after liftoff to achieve the initial climb speed.
Engine failure results in approximately
2 to 2-1/2 degrees lower attitude than
normal climb.
NOTE: With center of gravity at or near the aft limit, avoid sudden brake release
and maintain forward pressure on the control column
knots to increase nose wheel steering effectiveness.
BO_L'JAk'O
+
FAA
to approximately
APPROVED
2-20-70
I SECT
m NO.
80
D6-3000000000
Vol.
IPAGE
II-
8
AIRPLANE
PERFORMANCE
R_FUSED
FLIGHT
C 0_DIT_0NS
TAKEOFF
(Anti-Skld
MANUAL
_ND
FROCEDUR_S
Or4.)
C ondit ions
Calculated accelerate-stop
distances account
reaction times, plus arbitrary time delays.
Reverse
thrust
for demonstrated
recognition
and
was not used.
Procedures
When an engine failure occurred,
recognized prior to VI.
the takeoff
If the takeoff was refused for any reason,
was accomplished
as rapidly as possible:
Wheel Brakes - MAXI_
BRAKING
All Thrust Levers - IDLE
Speed Brakes - UP
CLIMB-OUT
was refused
prior
when the failure
to VI, the following
was
procedure
APPLIED
(B or _ Enalnes)
C audit ions
Climb gradient and obstacle cleaxm_ce flight path performance
most critical engine inoperative at VI.
is based
on the
procedures
Takeoff flap setting and V2 speed were maintained
for initiation of flap retraction.
Flaps were
section.
retracted
according
to the Flap
to at least
Retraction
Enroute procedures
were followed after climbing
_y
elevation,
or after all takeoff flight path
Speed
the height
Schedule,
selected
in this
to at least 1500 feet above
obstacles had been cleared.
run-
+
FAA
APPROVED
2-20-70
I SECT
IPAGE
R
AIRPLANE
pER_QRMANCE
FLIGHT
CONDITIONS
MANUAL
AND
pROCEDUBES
OBSTA_C_CE
Conditions
With all engines operating, a speed not greater than V2 + i0 knots was maintained
until either the scheduled flap retraction height, or the minimum gross height
for obstacle clearance (whichever was lower), was reached.
When engine failure occured Drlor to V2, and the takeoff weight was obstacle
limited, V2 was maintained
up to the gross height required for obstacle clearance.
If an engine failure occured afte____!r
V2, speed at engine failure (V2 + i0 knots
maximum) was maintained
up to the gross height required for obstacle clearance.
Procedures
When the height selected for initiation of flap retraction was limited due to
dlstant-obstacle
considerations,
the procedure was to initiate flap retraction
and accelerate
to final takeoff climb speed while maintaining
constant altitude
andinitial
takeoff thrust setting.
Final
takeoff
climb was
continued
to 1500 feet above
minimum gross height required for obstacle
and maxlmum continuous thrust.
clearance,
runway
elevation,
at final
takeoff
or to the
climb
speed
The airplane should be levelled off, and flaps retracted at the selected
level-off height, only if the limiting obstacle is beyond the Third Segment and an engine failure has occurred.
The height selected for initiation of flap retraction may be limited by
available performance
as described under Takeoff Flight Path, this section.
Vertical clearance of either close-in or distant obstacles in the intended
flight path must be established
clearance charts.
by reference
to the appropriate
obstacle
+
FAA
APPROVED
12-30-69
ISECT
IPAGF
'0
PERFORMANCE
AIRPLANE
PERFORMANCE
LANDING
FLIGHT
CONDITI0_
MANUAL
AND
pRO_PUR_
FIELD LENGTH
All landing field lengths shown in this section are based ca stamdard day temperatures on a smooth, level, hard-surfaced
rumw_y.
Dry landing field lengths are
demonstrated
landing distances, from a 50 foot height at VREFa divided by a factor
0.6.
Wet landing field lengths are determined by multiplying
the dry landing
field length by a factor of 1.15.
proc ecl_res
Approach and landing were made with landing gear down, fl_ps in landing position,
thrust reduced to flight-idle ca all engines before touch-down,
and automatically
set to ground-idle on all engines after touch-down.
When the landing was made with anti-skld operating, full speed brakes
wheel braking were applied 2 seconds or less after touch-down.
and maximum
When the landing was made with anti-skid inoperative,
speed brakes were raised
immediately upon touch-down and stes_Iy, light bra_
was used if gross weight
was approximately
500,000 lb (006,800 kg. ) or less.
Steeply, light to moderate
braking was used if landing weight was over 500,000 lbs (26,800
k_. ). Brakes
were modulated as necessary to prevent skidding.
NOTE:
Landing
field lengths
+
FAA APPROVED
03-09-70
are not based
ca use of reverse
thrust.
R
PERFORMANCE
AIRPLANE
FLAP
RETRACTION
FLIGHT
SPEED
MANUAL
SC
H_D[_LE
Maximum level-off heights, Third Segment distances and Final Segment climb performance shown in this m_nual are based upon retracting the wing flaps during Third
Segment acceleration using the schedule below.
This schedule is recommended
for
all normal flap retraction operations.
During
acceleration,
select
flap positions
TAKEOFF
Initiation
V2
V2
V2
V2
Spee¢,
+
+
+
+
FLAPS
at the following
_0SITION
Knots
Final Segment
speeds:
i0 OR _0
Select
20
40
60
80
initiation
Flap Position
l0
5
1
0
Climb
Speed:
V2 + 80 Knots.
When flaps are being retracted at a constant altitude due to engine failure, with
a critical final segment obstacle, begin climbing when V2 + 80 knots is achieved,
maintaining
takeoff thrust setting untll flaps are completely retracted.
+
FAA APPROVED
BOIJMM_
5-ii-70
I SECT
I
N O.
R
Vo i • II
i PAGE D6-306
B
_:_.o1,2..
PF_Rv,_CE
AIRPLANE
PERFORMANCE
FLIGHT
MANUAL
CONFIGURATION
The airplane configuration associated with the performance data in this manual
is shown below.
Performance conditions not shown below are on the appropriate
charts.
nm s
TAKEOFF
Takeoff on all engines to
VI, and three engines subsequent
to VI.
Takeoff setting.
Noted on charts.
Down
Takeoff
engines
on operating
(3).
Same as takeoff.
Down
Same as ist segment.
Same as takeoff.
Up
3nI SE T
Same as ist segment.
Takeoff setting to
flaps up, according
to schedule.
Up
FINAL TAKEOFF
CLIMB
Maximum continuous on
operating engines (3).
Up
Up
_VROUTE
Maximum continuous on operating engines (3 Or 2).
Up
Up
APPROACH
Takeoff on operating
engines
(3).
Approach setting.
Noted
on chart.
Up
LANDING CLIMB
Maximum
seconds
Landing setting.
Noted on chart.
Down
LANDING
Ground idle on all engines
after touchdown.
Landing setting.
Noted on chart.
Down
ist SE_
_d
CLIMB
SE(RENT CLD_
CLIMB
available in 8
on all engines.
Anti-skid is on for takeoff and lauding with up to two brakes deactivated,
unless anti-skid inoperative braking performance is used.
thrust operations.
R
All three Air Conditioning packs are on for final takeoff climb.
For enroute
climb the number of the Air Conditioning packs on is noted on the charts.
R
One or three Air Conditioning
Wing anti-icing
is off during
Packs are on for all takeoff
all flaps-extended
operations.
+
F AF ROVED
NO.
6-5-70
IsEcT
Vol.
II
IPAGE D6-306
I .o- 3
AIRPLANE
NQ_E
FLIGHT
MANUAL
R
CHARACTERISTICS
No determination
has been made by the Federal Aviation Administratlcm
that the
noise levels in this manual are, or ahould be, acceptable or unacceptable
for
operation at, into, or out of, an airport.
The noise levels tabulated below
part 36 certification
tests:
are the result
Kilograms
Position
_eozr
3.5 nautical miles
710,000
673,000
564,000
322,056
305,273
255,830
The
traded noise
lO
20
30
level
Aviation
Regnlations,
NOISE LEVEL (EPNdB)
n_s
Pounds
of Federal
_
rAm_m_
0.35 nautlcal mile
1_5.o
lOl.9
114.5
......
lOl. 9
1.0 nautical mile
----113.6
is 112.0 EPNdB
Normal all engines takeoff procedures were used, with a climb-out at V2 . i0 knots
at Takeoff Thrust with no thrust cut-back.
The landing
approach was made at VP_F
+ i0 knots.
! i
I
-tFAA
iNo. vol.
APPROVED
03-09-70
I SECT
'IPAGE L_-.o- _A-
"J
AIRPLANE
FLIGHT
MANUAL
PERFORMANCE
I ENVIRONMENT
ENVELOPE
5o
FAA
APPROVED
12-50-
69
0o_'#_
ISECT
I NO.
voz._'3°_-'3"rr
IPAGE _'.o-
,_"
AIRPLANE
FLIGHT
MANUAL
PERFORMANCE
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COMPONENT
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_
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30
5O
KNOTS
BOEINO
12-30-69
-+-+-f_
__ii-t
-20
APPROVED
_..i....i...._...._
' .....
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:..
""ii'"i"4"'
FAA
_"'_'"i'"4
.-..'.--4..._....!.-.4.
_""_"";,'"_'"!
ISECT
I
I NO"
;:;;-30643
.,_k. _T
IpAGE
-it:,
AIRPLANE
FLIGHT MANUAL
PERFORMANCE
40-8OKNOTS
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-_
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OAT --DEGREES
C
!
FAA APPROVED
6-50-70
I SECT
o.
PAGE
t_'. o - V-/
:
AIRPLANE
FLIGHT
MANUAL
PERFORMANCE
I TAKEOFF
EPR
GO.ROUND
O_
<
o
.,
_-4
9
-I0
0
I0
20
TAT ml)EGR£E$
FAA
APPROVED
6-
5-70
30
40
50
60
C
,,,,t.H,_ INo. _ _°__3
I secT
IpAOE _.o-
,8
AIRPLANE
FLIGHT
MANUAL
ifiiii!!iii
,II'_L"JAF_
FAA
APPROVED
12-30-69
I
SECT
NO.
VOlo i±
.I PAGE _-3o6___3
15".0 -19
AIRPLANE
FAA
APPROVED
2.-30-m
FLIGHT
MANUAL
'dl'_'_mJAV'a_'
I SECT
I NO.
IPAGE
vD6"3ol.
°_I
_s.o - 2.0
AIRPLANE
FLIGHT
MANUAL
PIE RFOAMANCE
RECOMMENDED
USE
SET
OF
FOR
ALL
STABILIZER
GRAVITY
TAKEOFF
FDR
TAKEOFF I
FLAP
SETTINGS
TAKEOFF
DETERMINED
BY
OR GRAPHICAL
METHOD,
WITH
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FLAPS.
CENTER
CALCULATED
GEAR
DOWN
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CENTER
|A4
II_ID
5-11-70
OF
GRAVITY_
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D6-30___3
FAA
APPROVED
12-30
-69
I SECT
IPAGE _s-.o"
"?_'_
_ "+'_'_
"
:._:i::!
:i:;: _: .,..L_+
AIRPLANE
FLIGHT
MANUAL
PERFORMANCE
I STALL
NOTE•
APPLICABLE
,: _i,_!i
FOR
i_ii
TAKEOFF
:!!!l:
AND
SPEEDS
LANDING
i,i:_i
ALTITUDES
tii!'!
ONLY.
:._'!i __..':_-:i:_,.!: i_!il
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;550
400
450
!::::!Ii77
'500
GROSS
FAA
APPROVED
12-_69
!i
550
WEIGHT--
!';:!
600
I000
: :17!7
650
700
750
LB.
I SECT
IP_(:;E_.o-
z._-
_RPLANE
FAA
APPROVED
12-30-69
FLIGHT
MANUAL
:_p,O_rJA#'4_
CT
NO.
PAGE
v01.
__'.o-z-_
,,
P]_:_OR4ANCZ
AIRPLANE
MAXIMUM
NORMAL
ALLOWABLE
TAKEOFF
CHART READ_G
WEIGHT
FLIGHT
MANUAL
TAKEOFF
WEIGHT
ANALYSIS
PROCEDUI_
The following steps w_ll be adequate for most takeoff situations.
_
dotted
gulde-lines
on typical charts and the Illustrative
Exmmples are for guidance
purposes.
I.
Select most
2.
Runway Len_h
Co_rectlo_
(A_I Engine@) cha_t_ - Enter with actual
available.
Make slope and wind corrections.
Read correcte_ _7
e
4.
5.
e
probable
takeoff
flap
setting.
l_v
Length and V1 A_lu_DSs
char_ available on both the Run_
Length, and
Make slope and wind corrections from the
the two corrected distances intersect on
read VI/VR , add corrected _umw_7 length.
runway length
length.
Enter with actual runway leugth
Accelerate-Stop
Distance
scales.
respective
reference lines.
Where
the "web" portion of the chart,
M_xlm_
T_keoff Weight. Fie_d Length Limits c_art - Enter with the lesser of
the corrected run_
lengths from Step 2 or Step 3, airport pressure
altitude,
and temperature.
Read gross weight.
MaYim_
Tak_ff
Weight, C14mb T.Imlts chart - Enter
altitude and temperature.
R_a_ _ross weight.
IF 200 MPH TIRES ARE INSTAI/_,
USE:
Mmximum Takeoff Wei_t,
T_._e Speed chart
and temperature.
Make wind correction.
NOTE:
If obstacles
are present,
proceed
with
with
airport
- Enter with airport
Bead _ro_s weight.
Steps
7 through
pressure
pressure
altitude
i0.
- _3061 3
.+
FAA
APPROVEO
1 2-_O-_O
ISECT
IPAGE
o-z6
- _
P]_FO_C_
AIRPLANE
MAXIMQM
NOBMAL
A_OWAB_E
TAKEOFF
CHART HEADING
e
,
PBOCEDURE
TAKEOFF
MANUAL
WEIGHT
ANALYSIS
(Continued)
Takeoff Climb 9hart - Enter with leas_____t
gross weight determined by Steps _, 5,
6, structural limits, or operational requirements,
then proceed to temperature
and airport pressure altitude.
Read Second Segment _radient.
Correct the
gradient for wind, if applicable,
using Gradient Corrections
chart.
Obstacle C_eara_ce (Close-in
from end of takeoff distance
ceed vertically
to zero-wind
(a)
(b)
,
WEIGHT
FLIGHT
Obstacles) chart - Enter with obstacle distance
required.
Correct the distance for wind.
Progradient available, from Step 7.
If this exceeds obstacle height, obstacle is cleared.
If obstacle height is no___t
exceeded, follow dashed field-length-trade
guide lines to obstacle height.
Read _radient require d for obstacle
clearance.
This gradient accounts for the shorter takeoff distances
required when gross weight is reduced.
(See SLOPED RUNWAY EFFECT ON
OBSTACLE CI_ICE
).
Obstacle Clearance (Distant Obstacles ) chart - Enter with obstacle distance
from end of takeoff distance required.
Proceed vertically to wind-corrected
gradient available, from Step 7.
(a)
(b)
If this exceeds obstacle height, obstacle is cleared.
However, now drop
back down to obstacle height, then move left (horizontally)
to windcorrected gradient available.
Read _ross height.
If obstacle is not cleared at wind-corrected
gradient available, follow
dashed field-length-trade
lines to the obstacle height.
Bead gros_
height and gradlen._ required for obstacle clearange.
Correct the gradient
required to zero-wind gradient, on the Gradient Corrections
chart.
(See
OBSTACLE IN THIRD SE_T,
and OBSTACLE IN FINAL SEGMENT. )
+
FAA
APPROVED
12-30-69
ISECT
IPAGE tS". o - Z7
P_e0R_J_CE
AIRPLANE
MAXIMUM
N(M_AL
AL_QWA_LE
TAKEOFF
CHART READING
10.
ll.
L_e
FLIGHT
WEIGH_
P_
MANUAL
T_EOFF
WEIGHT
ANALYSIS
(Continued)
Takeoff C_imb 9ha_t - Enter gradient scale with the largest zero-wlnd gradient
required for obstacle clearance as determined
from Step 8 or 9. Proceed to
airport pressure altitude, and temperature.
Read _ross weight.
Takeoff Soee_
chart - Enter with airport pressure altitude, temperature,
and
the _
gross weight determined by Steps h, 5, 6, i0, structural
limits, or
operating requirements.
Read VR, a_ V2.
Read VI at the V_/yR ratio from
e_.
The resulting VI must exceed the minimum V1 (VMCG), (See V1 Less
than VMCG), and the resulting VR must exceed the VMCA limit.
(See VR Less
than VMCA Limlt).
Max_
Brake Energy L_m_t S_eed chart - Enter with airport pressure altitude,
temperature,
and gross weight used in Step ll.
Correct for _
slope, and
wind.
Read VM_E.
This speed must exceed VI of Step i0.
(See VI Greater
than VM_).
13.
Gross Height - Pressure Alti_d_
Co_ver_ion
_hart0 - Enter with gross height
obtained f r_n Step 9. Correct for temperature
and altitude.
Re_d _ressure
_ltitude increment.
Add this value to the airport pressure altitude.
Acceleration and flap retraction
should not be scheduled below this altimeter
reading.
lb.
Determine flap retraction
SPEED SCHEDULE.
..
FAA
and final
climb
speeds
from the FLAP RETRACTION
Lo-306 3
NO. Vol. Ti
Ar_'_"J,_V_
APPROVED
12-30-69
I SECT
i
_r
PAGE
4
I_'. o - ?__,
AIRPLANE
MAXIMUM
TAKEOFF
ALLOWA_E
WEIGHT
ANALYSIS
FLIGHT
MANUAL
_A_OFF
WITH
IMPROVED
CLIMB
W_IGH
T
PERFOI_CE
The following steps describe the procedure for finding the maximum allowable takeoff weight when using Improved Climb Performance.
Improved Obstacle Clearance
wlth speed increase is considered under a separate paragraph heading.
The term,
"normal", refers to the weights and speeds based on no increase in takeoff speeds.
1.
Steps
1 through
6 of the Normal
Takeoff
Analysis
are unchanged.
e
Maxim_
Takeoff Weight and
the weight portion of this
gross weights.
Follow the
section, read _ross weight
climb limited gross weights
weight and speed increase.
e
Takeoff Spe@dS chart - Enter this chart, as in Normal Analysis, with pressure
altitude and temperature,
and with the gross weight determined
from Step 2
above.
Read VR. V2. and Vl for the VI/VR ratio Ip Step _ of Normal Analysis.
e
Maximt_n Takeoff Weig_
a_ Speeds (Imoix_ged C_$m b Performance ) chart - Enter
the speed correction portion with the VR, V2 and V1 speeds determined
in Step
3 above.
Follow the guidelines to the speed increase determined
in Step 2
above.
Read Corrected VR, %"2. and VI.
V1 must be less than V_H_E. (See VI
Greater than VMSE).
5.
Determine
above.
flap-retraction
Soeeds {Imorovcd Climb Perfo_mapce ) 9har% - Enter
chart with normal field length and climb limited
guidelines until they intersect.
At the interand speed increase.
Repeat, using tire speed and
if 200 MPH tires are installed.
Use lesser
speeds based
on the V2
speed
derived
m m,,A, l
FAA APPROVED
12-30-69
ISECT
from
NO.
Step
Vol.
IPAGE
ii
_s.o-z.9
3
AIRPLANE
MAXIMUM
SELECTIC_
A_OWAB_E
FLIGHT
MANUAL
TAKEOFF
.WEIGHT
OF V_
When the calculated or scheduled takeoff weight is not field-length
limited, V1
may be raised or lowered to suit operating conditions within the restrictions
imposed by the available runway length, VMCG, VR, and VMHE.
Even when takeoff
weight is field-length limited, a small reduction in V1 is available if the
3-engine field length requirement
is shorter than the all-englne requirement.
This is generally true for this airplane.
The following additional Steps are required
to determine the V1 limits:
in the Normal
Takeoff
Weight
Analysis
Maximum Takeoff Weight (Field l_n_th Limits ) cha_t - Enter with airport
pressure altitude, temperature,
and actual scheduled takeoff weight.
_ad
corrected runway length reauired.
le
,
Runway Length and V1 A_lustments
chart - Enter, in upper right portion of the
chart, with the corrected runway length obtained from Step 1. Where this
runway length intersects the slope and wind corrected takeoff distance available line as determined in Step B of the Normal Analysis procedure, read minim_
V1/VR.
Where this runway length intersects the slope and wind corrected
accelerate-st0p
distance line as determined in Step 3 of the Normal Analysis
procedure,
r_ad maximum V1/VB.
If either intersect occurs off the limits
values shown on the chart.
3.
,
of the chart,
use minimum
or maximum
Takeoff Speed chart - At the actual scheduled takeoff weight, and the VI/VR
ratios determined In Step 2 above, read minimum V1 and maximum _ri.
Choose a suitable, single value of VI between the limits determined in Step 3
above.
The speed selected must be greater than VMCG, and less than the VR
and V_
determined in Steps 11 and ]2 of the Normal Amalysis procedure.
_I'_'_L'JMM_
+
FAA APPROVED
12-30-69
ISECT
_6-306h3
I NO. i VO_ . _ _
IPAGE
0- o
AIRPLANE
MAXIMUM
ALLOWABLE
FLIGHT
MANUAL
TAKEOFF
WEIGHT
Vl LESS
Usually
this
will
only
occur
when
the takeoff
weight
is not
field-length
limited:
Ts/_eoff Speeds
chart - Read Vl at the field-length
limited
gross weight
obtained
from Step 4 of Normal Analysis
procedure,
and at the V1/VR ratio obtained
from
Step 3 of Normal Analysis.
Use a V1 between
VMCG and the maximum
V1 so determined,
but not exceeding
VR.
V_G_TERTHANVMBE
If the minimum
VI, as determined
exceeds
the brs/_e energy
speed,
Maximum
Reduce
weight.
VR
Brake
gross
Energy
Limit
weight
as
by Steps i through
3 under SELECTION
OF VI,
the takeoff
weight
will have to be reduced:
Sp@_d
indicated
chart
- Determine
and redetermine
A
VR,
VI,
V1
as defined
and
V2
for
on the
the
still
chart.
lower
_STHANVMCAL_T
On this airplane,
VMCG is greater
than 1.05 VMCA; therefore,
mini_mAm VR is equal
to VMCG.
This implies
that the available
runway
length _,st be sufficient
to
permit
a V1 equal to VR.
This condition
will be satisfied
if the slope - and
wind-corrected
accelerate-stop
distance
exceeds
4500 feet on the Runway Length
and VI Adjustments
chart.
OPERATION
WITH
ONE
AIR
CONDITIONING
PACK
R
ON
Performance
limited
takeoff
weights
and takeoff
climb gradients
available
can be
increased
by scheduling
takeoff with only one air conditioning
pack on.
Performance
increments
are shown on the Takeoff
With i A/C Pack On chart;
these increments
are applied
to the weights
and gradients
obtained
in steps 4, 5 and 7 of the
Normal
Takeoff
Weight
Analysis
procedure.
The gradient
increment
_,St be subtracted
from the gradient
r_quir_d
for use in step l0 to find an obstacle
limited
weight.
÷
FAA APPROVED
ROI/NO
IN O.
D_-306_3
vol. II
IPAGE
_s.o-3_
6-5-70
ISECT
AIRPLANE
MAXIMUM
ALLOWABLE
sumPz
FLIGHT
_A_EQFF
MANUAL
WEIGHT
mmWAX
Wet or icy runway conditions have the same effect on takeoff field length requirements as an inoperative
anti-skid system; namely, reduced braking effectiveness.
This makes the one engine inoperative condition critical.
The operator must review the existing runw_
conditions,
decide how much additional
stopping distance is desired, then proceed with the followir_ steps to determine
the changes in the field length limited takeoff weight, and V1 speed:
i.
Omit
Step
2 of the Normal
Weight
Analysis
procedure.
_unwav Length and V1 Ad_lus_ments char_ - Determine V1/VR as usual, except that
the accelerate-stop
distance must be reduced by an arbitrary amount equal to
the desired extra stopping margin before making corrections
for slope and
wind.
At the intersection of the corrected takeoff distance available and
corrected accelerate-stop
distance lines in the "web" portion of the chart,
read correct_ d _v
ler4Eth and Vl_ R.
.
3.
Maxlmtm2 Takeoff weight chart - Enter the appropriate
runway length from Step 2 above.
Rea_ _ross weight.
NOTE:
chart
with
the corrected
If the V1, determined by using the VI/VR ratio from Step 2 above, is
greater than the minlmun VI as determined under SELECTION OF VI, no reduction in field length limited takeoff weight is necessary.
In fact, the
use of minimum V1 in this case will give a greater stopping distance margin.
+
FAA
Takeoff
APPROVED
12-30-69
I
NO.
Vol.
I±
PAGE
D6-3o6_.3
I _'. o -'Z, ?_
AIRPLANE
MAXIMUM
ALLOWABLE
FLIGHT
MANUAL
TAKEOFF
W_
IG_T
EXAMPLES
Example
lines
ILLUSTRATIVE
Given:
shown
on the
EXILE
Airport
charts
do not
necessarily
reflect
these
examples.
1
Conditions
Runway
Length
Run
Slope : O.5% (Uphill)
Available
= 12,000
ft.
Airport
Elevation
(Pressure
Altitude ) = 2000
Obstacles:
120 ft. high at 3400 ft.
500
A_mospheric
Reported
ft.
high
at 25,000
ft.
ft.
C ondltions
Wind
(at
50 ft. ) = i0
kt.
(Headwlnd)
OAT = 20°C (68°F)
Airplane
Conditions
T_keoff
flap position
I0
This example
illustrates
the general
use of the charts,
and
clearance
charts
in particular,
and is restricted
to normal
tion.
See
See
example
5 for improved
chart reading
procedure
From Runway
12,070
ft.
le
1
.
4e
e
e
Length
Corrections
From Runway
Length
and
ft. and VI/VR = .970
obstacle
clearance
for chart titles
(All Engines)
V1 Adjustments
chart,
with speed
and methods
chart,
increase
of use.
corrected
corrected
the obstacle
speed opera-
runway
method.
runway
length
length
= L_,570
From Maximum
Takeoff
Weight
the corrected
runway
length
weight
= 683,000
lb.
(Field Length
Limits)
chart,
using the lesser
of
from Steps 1 and 2, field length
limited
takeoff
From Maximum
Takeoff
= 649,000
lb.
(Climb
Weight
Limits)
chart,
climb
limited
takeoff
=
weight
From Maxi_im
Takeoff
Weight
(Tire Speed Limits)
chart,
200 _H
tire speed
limited
takeoff
weight
= 743,000 lb.
Maximum
Takeoff
weight
= 649,000
lb.
From Takeoff
Climb chart, using field length
limited
second segment
gross gradient
at 400 ft. = 2.30%
weight
Using
for
Gradient
Corrections
chart,
gradient
corrected
of
wind
683,000
lb,
= 2.37_
(Continued)
+
FAA APPROVED
2-20-70
BOIJN_
I
SECT
NO.
I
Vol.
II
_-30643
PAGE
15". 0 - E'_
PERFORI_J_
AIRPLANE
MAXIMUM
FLIGHT
ALLOWABLE
E
MANUAL
TAKEOFF
WEIGHT
EXA/_LES
ILLUSTRATIVE
e
1
(Continued)
From Close-in Obstacle Clearance chart, second segment gross gradient required
for close-ln obstacle clearance = 4.50%.
Obstacle is not cleared with gradlent available (uncorrected for wind).
Change in reference zero - 5100 3400 = 1700 ft. due to the decrease in field length required for the reduced
takeoff weight.
In this distance, reference
due to runway slope.
The obstacle
the required
.
zero shifts
1700 x .005 = 8.5 ft. (vertical)
height is now 128.5 ft. at 5100 ft. from reference
gradient is 3.08_.
From Distance Obstacle
minimum gross height.
Clearance
chart
R
determine
gross
zero and
gra_lient required
and
Gradient capability
(from Step 6) = 3.08% (corrected for wind = 3.17%)
Corrected obstacle distance = 25,000 + 1700 = 26,700 ft.
Required gross gradient = 2.89_ (less than gradient available).
Distance from reference zero to reach obstacle height of 500 ft. is
23,300 ft.
Minimum gross height for flap retraction is 770 ft.
1
I0.
lie
From Takeoff Climb chart,
gradient of 3.08%.
obstacle
limited
weight
= 638,500
lb. at a gross
From Distance Obstacle Clearance chart, maximum level-off height is 940 ft.
Therefore extended final segment climb is unnecessary
in this case.
From Gross Height - Pressure Altitude Conversion chart, the pressure altitude
increment for a gross height of 770 ft. is 750 ft. Therefore, the m_nlrmlm
pressure altitude for level-off and flap retraction is 2750 ft.
12.
From Takeoff Speeds chart for Vl_R = .970 and
VR = 152.5 kt.
V1 = 147.5 kt.
13.
Determine
flap retraction
final climb
speeds.
__='f#IM_
+
FAA APPROVED
and
638,500 lb.
V2 = 159 kt.
5-11-70
ISECT
INO.
IPAGE
D6"3061_3
Vol. II
o-
R
AIRPLANE
FAA
APPROVED
12-30-6S
FLIGHT
MANUAL
AIRPLANE
i
FLIGHT
MANUAL
.+ ++_
, X.T_+c
m +-..+
V'_.
.÷_..
-.++
.÷++.o.,..+.
FAA
APPROVED
1-20-70
_o_..,,,_
lNo._k_o_3
ISECT
IPAGE ,s.o-'_eo
_RPLANE
FAA
APPROVED
1-20--70
FLIGHT
MANUAL
_,_
ISECT
iNo.
_15".
o_
PAGE
0-_7
_RPLANE
0
0
W
m
I'-
-I
FAA
APPROVED
IZ-30
-69
FLIGHT
MANUAL
FAA APPROVED
O0_rl_I/'O
ISECT
lNO. _o_.3('_Z
_3
IPAGE,_.o-_
AIRPLANE
FAA
APPROVED
17.-30-6S
FLIGHT
MANUAL
AIRPLAN
lr
FLIGHT
MANUAL
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.+++.
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FAA
APPROVED
IZ-30-_
oo*'-N ,I
IsEcT
,,,+
+_+++ ++++
,_
i+-++'+
+;++
+.pp.+
NO.vo:.._:3'_6_-3]:z
IPAOE _-. o- W
AIRPLANE
FAA
APPROVED
IZ-50-_
FLIGHT
MANUAL
AIRPLANE
FLIGHT
MANUAL
I-
FAA
APPROVED
5-11-70
__L'J_b'_
i SECT
]NO.
D6-3064_3
I PAGE vol.
zz
_.o--LI3
_,, -
,UFII_Ld,NI,:
FLIGHT
MANUAL
'i
!
FAA
APPROVED
12-30-69
__L'JAV"I_
iSECT
INO.
D6-30643
Voi. II
IPAGE I_'.O
- V'-/
FAA
APPROVEO
12-30-69
_XI_VOINO"
_-3o643vol.
T_
ISECT
IPAGE,_.a-@G-
18
!
FAA
APPROVED
12-30- 69
I SECT
• Vol. II
I NO
PAGED6-_30643
t_-o-_6
I
AIRPLANE
FLIGHT
MANUAL
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AIRPLANE
APPROACH
AND
FLIGHT
MANUAL
LANDING
Charts
on the following
pages present approach
and landing
gradients,
maxia_un
landing
weights
as limited
by approach
and landing
performance,
landing
field
length
requirements,
and landing
weights
for the maximum
brake energy
at which
wheel
thermal
plugs will remain
intact.
The speed schedules
shown on the Approach
and Landing
Climb charts
are those at
which
the gradients
were calculated
in accordance
with FAR 25.12_1(d) and 25.119,
and have no operational
significance.
The minimum
landing
approach
speeds are
shown on the Landing
Field Length
and Speed charts.
The ICE CORRECTION
on the charts accounts
for maximum
probable
performance
effects
of ice remaining
on the airplane
surfaces without
_uti-ice
protection.
The correction
applies
when operating
in icing conditions
during
any part of the flight,
unless
the forecast
temperature
at the destination
airport
is high enough
(above
8°C or 46@F) to ensure that the ice will melt off prior to approach
and landing.
The Maxlmum
Landing
Weight,
Climb Limits,
can be increased
for one air conditioning pack on operation.
This information
is included
to permit
a showing
of compliance
with FAR 25.1001
(c) at the increased
allowable
takeoff
weight.
R
This information
may also be used to schedule
higher
allowable
landing
weights
provided
approach
and landing
procedures
are modified
to have only one air conditioning
pack on prior to the point where a go-around
might be initiated.
See
Normal
Procedures.
R
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16.0
APPE}[DIX B - BUFFET
The buffet
frequency
in the table on Page 16.0-_-.
each flap detent and
From
Buffet
For illustration,
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range
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These conditions
compressibility
this investigation
Average
CHARACTERISTICS
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17.0
APPENDIX
C - AUTOTHROTITZ
The block diagram
block
diagram
for use in
This data is provided
on Page 17.0-2
small perturbation
for information
is a simplified
747 control
purposes
and ms
autothrottle
simulation.
not incorporated
in the simulation.
.
B_IIIM_
REV LTR:
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1 19_
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PAGE
I"-/. o - 2.
18.0
APPENDIX
This
flight
director,
aircraft.
D - AUTOPILOT
section
contains
yaw dsm_er,
and autothrottle
This data is provided
pilot and autothrottle
were
a description
of the automatic
systems
for information
not incorporated
pilot,
for the 7_7
purposes.
The auto-
in the simulation.
D6 - 30643 - RESTRICTED USE - See Notice on Cover
])6-3o643
,,,,F,_
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!
|NO.
REV LT'R:
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D6-30643
Vol. II
IsEct:,s IPAGE'
Page
F.
G.
II.
III.
PITCH AUTOPILOT
i.
Engage
2.
F_nual
3.
Turbulence
4.
Co,_d
YAW DAMPER
Synchronization
h5
........................
, ....
.............................................
45
.........................................
.........................................
45
45a
...
b7
a.
Altitude
Hold
b.
Altitude
Select
c.
Vertical
Speed
d.
IAS Hold
e.
_ch
f.
ILS Mode .......................................
52
g.
Land
.....................................
53
TRIM UNIT
61
AUTOMATIC
FACH TRIM
.........................................
47
................................
.................................
51
52
.....................................
52
Hold
Mode
?$°
47
......................................
STABILIZER
SYSTEM
.................................
...........................
(Section Deleted)
SYSTEM
69
............................................
A.
GENERAL
.................................................
B.
SYSTEM DESCRIPTION
Io
Dutch
2.
Band
3.
Dutch
4.
5.
Turn
6.
Yaw Da_per
7.
Self Test
......................................
Roll Damping
Pass Filter
Roll Damper
Signals
Gain
Circuit
70
70
.........................
................................
Coordination
Easy On-Off
69
•..
70
.............................
.................................
.
and Confidence
Test
Servo
71
.................
...................
71
71
................................
Electro-Hydraulic
70
...
72
O
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i PAGE
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6
6-7'_,_
Page
8.
Performance
9-
Tolerance
i0
IV.
•
Failure
AUTOTHROTTLE
A.
GENERAL
B.
SYSTEM
.................................................
..................................................
Mode
SYSTEM
Design Features
DESCRIPTION
2.
Accelerometer,
3.
Computation
_.
Throttle
5-
ADI Signal
6.
Airspeed
7.
Disengage
9.
Test
73
eeeooeeDeeeeeeeeeeeeeeeeee@eeoe
........................................................
Command
Flare
72
................................................
I.
e
72
.............................................
and Airspeed
Error
Attitude,
Signals
.........................
and Elevator
Signals
..............
................................................
Control
and Limits
................................
.................................................
Error
Warning
......................................
..................................................
Ioaei•e•eooeoeee•eeeooeee•ooo•eeeo,oeooeoeeoeoeoeeoeee
83
83
83.,,.
83
87
87
87
88
88
88
88
.......................................................
88
O
!
1)6 - 30643 - RESTRICTED
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p
REV SYM (_
BOEINd_
|8
No Vol. II
iPAGE D6_30643
7
fr-7000
SUMMARY
This document
provides
a Aascription
_f _he _
_
Flight
Director,
Yaw Damper,
and Autothrottle
Systems
for
the
747 aircraft.
It contains
block
dia_rs_m,
_i_t_r_al_,
xnd _ables
_o _n-l_e
_ration
of e_ch
of these
systems as they are planned for i_ti_
in the aircraft.
Two autopilots
give the flight crew a choice of fail-safe systems, each of
which can provide all manual and path modes except LAND.
Both autopilot
channels are used to give "fail passive" control at low altitudes during the
automatic landing sequence.
Two separate dual-channel
automatic pilot pitch
trim systems and two yaw dampers give a high level of system integrity.
0
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ip,_,E _-3o643
8
LIST OF ILLUSTRATIONS
Figure
Titl_____e
I
Autopilot/Flight
la
747 Cockpit
Flight
2
Mode
Panel
2a
Control
3
Preset
Course
4
Flight
Controller
4a
Control
4b
Flight
4c
Typical
4d
ILLS Deviation
Warning
5
Basic
Autopilot
6
Force Detent
7
INS Mode Sequence
8
True
9
Radio
Select
Director
Controls
BITE Test
Force
747 Autopilot/Flight
lla
Roll
12
747 Lateral
13
LAND
Wheel
Flight
15
19
................
19a
20a
21b
Servo
Mechanization
Characteristics
21d
..... 23
........
.............................
Pitch
Steering
Control
Block
15
Attitude
16
Pitch
17
Altitude
18
Glide
19
X_ml-Channel
20
LAND
Mode Sequence
21
LAND
Mode Interlocks
22
Elevator
23
Autopilot
23a
Block
Wheel
39
•. .........•.................
Roll Axis
Computer
Block Diagram
40
...............
......................
Diagram
....................
Autopilot
Steering
and Mach
41
........
System
42
42a
43
• 44
...................
Hold Options
25
33
................................
Mode - Pitch
14a
46
...... 46a
- Path Nodes
...........................
48
Mode Gain Schedule
..........................
49
Select
Mode
Slope Mode
Control
...................................
Gain Schedule
Pitch
Stabilizer
Landing
System
54
........
..... : ...............................
55
• 56
....................................
57
.................................
Schematic
Gearing
Trim
59
............................
Elevator-Autopilot
to Column
50
...........................
Axis Automatic
System
Elevator
Diagram,
Automatic
Control
Director
Mode Monitoring
Control
l_a
...........................
Gain Program
ii
Elevator
System
Gain Program
Control
...............
................................
Gain Program
Hold
14
......................................
Time-Based
Control
13a
......................................
i0
Path
Transducer
and Servo Operating
Altitude
..............
.......................................
Steering,
Airspeed
12
.................................
Mode Annunciator
Attitude
and Indicators
Mode Select Panel
Switching
Lateral
........................
.......................................
Wheel Steering,
Wheel
System
Mode
60
..................
60a
...............................
................................
60b
62
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Titl___Re
25
Deleted
........................................................
26
Deleted
.................................................
27
Yaw Damper
28
747 Yaw Da_per
29
Mathematical
Pictorial
Diagram
66
...... 67
..................................
and Turn Coordinator
in Block
Model of the 747 Yaw Damper
Diagram
and Turn
Form
74
..... 74
Coordinator
Systems
........................................................
76
30
Deleted
.......................................................
77
31
Frequency
32
Phase Angle
33
747 Yaw Damper
34
35
Response
of Yaw Damper
Plot of Yaw Damper
System Root
Bandpass
Bandpass
Locus
Filter
..............
78
Filter-. ...............
79
..............................
80
Tolerance
Effects on Landing Approach Condition Flaps 33°
Down ........................................................
Tolerance
Effects
on HaximumDynamic
Autothrottle
Speed
Control
37
Autothrottle
Controls
38
Autothrottle
System
System
Pressure
...... . 82
.............................
84
...........................
85
and Monitors
Implementation
Condition
81
Block
Diagram
..............
86
LIST OF TABLES
Table
i
Autopilot/Flight
Director
Mode Chart
- Pitch
2
Autopilot/Flight
Director
Mode Chart
- Roll
3
Flight
4
5
Summary
6
Flight
V
Autopilot
Mode Annunciator
of Autopilot
SummarTof
Flight
Director
Chart
Roll Axis
Director
PitchAxis
Pitch Axis
_II
Mode
Mode
..... •.......
17
........... ... 18
•................................
21
................................
37
Axis
38
..........................
....................................
45b
..........................................
45c
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p^GE I0
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.I$ o
I.
747 INTEGRATED
A.
Ab_OPILOT
AND FLIGHT DIRECTOR
GENERAL
The Sperr_ SPZ-I integrated autopilot flight director system was developed for the Boeing 747.
The autopilot is used for path or manual con•-r_l via the _
_rh_el _d
turn ku_rb:* Flight director commands are
provided and may be used to monitor autopilot operation, or to be
tracked with the autopilot manual mode, or to be tracked manually. '_The basic 747 is equipped with a dual autopilot syst_
and a triple
flight director system.
Each pitch or roll computer contains all of
the circuitry necessary to compute the respective pitch or roll autopilot and flight director co,ands.
Each mutopilot or flight director
channel consists of two computers:
one roll, and one pitch.
The general autopilot flight director systa_ layout and interconnects
used on the basic Boeing 747 are shown in Figure I. A dual set of
sensors provide
flight director
navigational
computers.
and performance
signals
to the autopilot
Only autopilot computers of channel A and B are connected to hydraulic
servos.
Computer C is used only for computing flight director co_ands
and cannot be switched in as an autopilot.
The number i sensor group
serves both the A and C computers while sensor group 2 supplies information to the B computers.
Instrument switching allows either the pilot or co-pilot to select any
one of the three flight director computer signals for display on his
ADI.
_2
The autopilot can be used as a single-channel
system in the navigation, cruise, and manual modes Of operation.
Either channel A or B
can be selected for these single-channel
modes.
For automatic landing only, fail passive operation is obtained with
the dual channel system.
Both A and B autopilot channels must be
engaged for this mode of operation.
The cockpit location of the controls and indicators for the autopilot/
Flight Director, as well as the other systems
of the IEFCS, are shown
in Figure IA.
*Maneuvering
control for certain airlines will be via control
wheel steering instead of turn and pitch knob.
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B.
SYST_2_ INTRODUCTION
All components associated with each flight axis of each channel are
packaged in separate computer units.
The elements which comprise a
basic autopilot/flight
director system are:
1.
Mode
I)
One mode
2)
One flight
3)
Two flight mode
4)
One monitor
5)
Three
accessory
6)
Three
roll computers
7)
Three
pitch
8)
One automatic
Select
selector
panel
controller
or optionally
3 force
sensors.
annunciators
and logic
unit
boxes
computers
stabilizer
trim unit
Panel
The mode select panel contains the switches and logic for mode
selection and control of all the autopilot/flight
director computer
channels.
Figure 2 shows the mode select panel.
The controls
shown shaded in are for optional modes.
The engage switches are solenoid held in MAN and COMMAND with
locking provisions at OFF.
Each engage switch controls both the
pitch and roll computer associated with that channel.
Either
channel A or B maybe
engaged in Manual or Command by choice of
engage switches.
Once one engage switch has been placed into the
Nanual or Co_nd
position, the other switch is locked off and
cannot be moved from the OFF position except when LAND has been
selected and a monitor check has been satisfactorily
completed.
The course select switch may be placed in either the course 1 or
course 2 position without regard to which channel of the autopilot
has been selected.
It is solenoid held in both positions.
The
switch drops to the dual position when the LAND mode has been selected.
Course I position feeds the No. 1 VOR/LOC receiver output
and No. 1 course error signal to all three A/P - F/D roll computers.
and course 2 does likewise for the No. 2 VOR/LOC'and
No. 2 course
error.
In the dual position, LOC No. I and Course error No. I are
fed to computers A and C while LOC No. 2 and course error No. 2
are fed to the B computer.
Figure 3 is a pictorial diagram of the
above switching functions and indicates the isolation and independence achieved for single and dual channel modes of autopilot
operation.
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Heading select is accomplished for all three channels through a
single set knob which positions two synchros fed from two separate
magnetic heading reference units to maintain heading isolation.
The lateral navigation modes (INS, HDG,VOR/LOC),as well as the
landing approach control modesILS and LAND, are selected by the
main modeselector switch. This is a simple rotary type switch
without holding or centering devices. A second selector switch,
o,_.
also rotary type, allows selection of the Turbulence mode, V/S
(vertical speed), IAS hold, or Machhold. This second switch is
solenoid held in each modeselect position.
The V/S and IAS modes
may be selected alone or during the arm phases of ALT. Select, ILS,
and LAND. Whenused during the armed phases of another mode, the
switch will drop to OFFwhen the modegoes into the capture phase.
The switch will not hold in Turb. if the autopilot is not engaged.
Also, if the switch is movedto Turb while the altitude select is
engaged, the altitude select switch will drop to OFF. Further,
if Turb. is engaged and ALT hold is selected, the Turb. switch will
drop
to OFF.
The Back Beam switch is designed to mechanically
preclude accidental
turn on and is solenoid held in the ON position.
It will hold only
if VOR/LOC is selected on the mode selector switch.
Also available
on the mode select panel are the ON/OFF switches and pitch trim controls for flight directors.
The mode selection and mode compatibility
interlocks related to the
position of the switches on theM ode Select Panel are shown in Table 3
(Pitch channel) and Table 2 (Roll channel) for both the autopilot and
the Flight Director.
2.
Flight
Controller
The Flight Controller which contains the autopilct pitch wheels
and turn knob is shown in Figure 4. The pitch wheel and turn knob
provide attitude commands proportional to their respective
rotations.
The outputs for channels A and B of the autcpilot are electrically independent.
The potentiometers
used with the pitch wheels
are unclutched
electrically from the wheels during all pitch path
modes.
For customers who order the control wheel steering option "the flight
controller is not used.
Instead three force transducers
shown in
Figure _a are provided to maneuver the aircraft via the normal
control column and control wheel with autopilot engaged.
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3.
Flight
Mode Annunciation
Dual flight
ector modes
annunciator
mode annunciators are provided on the 747. Flight dirand autopilot modes are displayed side by side on each
panel as shown in Figure 4a.
The basic modes
o
o
o
o
o
annunciated
are.
ALT SEL (Altitude Select)
NAV
(Navigation)
GS
(Glide Slope)
FLARE
GO AROUND
?$o
In each mode the annunciator displays an amber light when the particular mode is armed and switches to a green light when the mode
is initiated.
Table 3 shows each autopilot and Flight Director
mode as it will be displayed on the mode annunciator.
Also displayed on the annunciator panel
throttle warning and disengage lights.
are the autopilot
and auto-
A press-to-test
feature is included in the flight mode annunciator. All amber lights are tested by depressing the left hand section of
the panel.
All green lights as well as the red warning lights
of the autopilot and autothrottle are tested by depressing
the right
hand section of the panel.
4.
MonitOr
,and
Logic
Unit
The monitor and logic unit is a separate package which contains
much of the autopilot/Flight
Director engage interlock logic, dualchannel monitoring logic, mode annunciation
logic, and the warning
light circuits.
The physical and electrical arrangement
of the
monitor and logic unit provides complete channel isolation.
5. ,Accessory
Boxes
The accessory boxes are Boeing supplied units which provide the
switching and interconnect functions necessary to interface the
autopilot and flight director systems to other airplane systems.
6.
Pitch
and Roll
Comuuters
The A/P - F/D pitch and roll computers include the computing and
logic circuitry necessary to receive data from the aircraft sensors
and produce autopilot _ervo and flight director commands for all
modes of operation.
O
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7.
Automatic
Stabilizer
Trim Unit
The automatic stabilizer trim unit (ASTU) contains the computational
circuitry necessary to provide automatic stabilizer trim whenever
the autopilot is engaged.
The ASTU contains two independent
self
monitored trim channels.
One channel provides trim operation
while the other is in standby.
Automatic transfer is provided to
the standby channel in the event that the trim monitor detects a
malfunction.
:,.
8.
Self-Test
and Maintenance
Self-test
of the pitch
Monitorin_
and roll computers,
the automatic
stabilizer
trim unit and the monitor and logic unit is performed wi th a go/no-go
readout by means of the Built-In-Test-Equipment
(BITE).
The BITE switches and lights, located on the front panels of each of
the above units, permit the rapid isolation of a faulty unit, while
the complete system is installed on the airplane.
A typical BITE test is shown in Figure 4b.
The BITE sets up input
slgnals and sensors to adjust two or more signal path gains in the
unit under test (DIE) as required to achieve a null summation of the
signal path outputs, if all circuits are normal.
The null condition
is sensed by BITE null detection logic.
Simultaneously,
specific
portions of the unit logic are addressed by BITE.
The status of the
unit logic circuits and the output of the BITE signal null detector
are combined in BITE logic to produce a go/no-go output to the BITE
readout lights.
Maintenance
monitoring has been included in the system as a basic
feature to help in isolating the cause of autopilot warning or disengagement during dual-channel autopilot operation.
The maintenance
monitoring circuits and readouts are included in the
Monitor and Logic Unit.
The readout is located in the top portion of
the Monitor and Logic Unit fron_
panel, above the BITE readout lights.
The maintenance
monitor readout consists of four latching indicators
which trip to indicate that autopilot warning or disengagement
occurred
for one of the following reasons:
i.
2.
3.
4.
Power
Power
Pitch
Roll
loss to channel A (single or dual channel operation)
loss to channel B (single or dual channeY operation)
channel camout monitoring
trip (dual channel operation)
channel camout monitoring trip (dual channel operation)
The above faults are monitored and displayed
latched indicators, until these are manually
permanently
reset.
by the
o
I)6 - 30643 - RESI1UCTEDUSE- See Notice on Cover
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Aa'd_EJAV'd_
18
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6.306__3.
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21a
_7000
BITE
SIGNAL
INPUTS
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NORMAL
SIGNAL
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SOLID
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$-'/000
.
ILS Deviation
Warning
System
(Optional)
The Deviation Warning System is a dual system, monitoring the
Captain's and First Officer's navigation receiver outputs.
(See
Figure 4d)
The system warns the pilots when the outputs of the navigation
receivers exceed 20 ua from the localizer beam centerline or 75 ua
from the Glide Slope beam centerline with a delay time of 2.2
_°
seconds.
The warning system is in operation when the autopilot is
engaged in the ILS or LAND mode and the radio altitude is below
500 feet altitude.
From 500 feet to 200 feet the system provides
a warning if the receiver output exceeds either GS or LOC thresholds.
If the nav. receiver signal causing the warning is reduced below
the detection threshold, the warning light will be turned off.
However, below 20Ofeet
altitude the system becomes latching and
once the warning system has been tripped the warning light will
remain on even though the signal error has been reduced below the
threshold.
Below Flare altitude (53 feet) the warning system
monitors
only
LOC deviation
signal
errors.
The two monitor systems are independent.
However, to provide
greater redundancy the deviation warning signals are cross fed
such that if one monitor is tripped it will also switch on the
warning light driver stage of the second monitor, thus illuminating
all warning lamps associated with the ILS deviation warning system.
Confidence tests of the system can be performed either on the
ground or in flight.
The confidence tests are pilot activated and
monitored..
The tests are activated by proper mode selection and
engagement of the A/P and F/D and the deflection of the VHF NAV
test switch.
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C.
DESCRIPTION
OF THE AUTOPILOT
ACTUATOR
SYSTEM
When the autopilot is engaged, autopilot commands are coupled into the
primary flight control system via parallel servo actuators.
Thus,
the control wheel and column, as well as the control surfaces, move in
response to autopilot commands.
The lateral control system of the 747 utilizes a oair of hydraulic
_o
central control actuators which control the hydraulic-powered
surface
actuators.
These central control actuators accept ccm_uands from either
the autopilot or the manual control system.
During all manual or autopilot operation, the two central control actt_torsare
slaved together
through a cross link.
The 747 lateral autopilot servo system is mechanized as shown in Figure
5. The pitch axis autopilot servo mechanization
is similar in concept
to the lateral system.
The mechanization
of the servo makes the system fail safe for all
single-channel
operation and fail passive for all dual-channel
(LAND)
operation.
I.
Single-Channel
Operation
Single-channel
autopilot operation is used in all modes except
LAND.
In single-channel
operation, either the A channel or B
channel
autopilot maybe
selected.
The autopilot drives an autopilot servo actuator integral with the central control package.
The
servo actuator output displacements
are proportional
to the autopilol
co_m_nd signals.
The autopilot actuator drives the manual controls
via a force detent.
When either autopilot is engaged, both central control
actuator.
packages
are driven
from
the engaged
autopilot
Autopilot authority in the lateral axis is stroke limited to the
equivalent of twenty-five degrees of control wheel displacement.
In the pitch axis, authority is limited by reacting the force
detent against the manual feel pressure system.
The pilot can overpower the autopilot at any time by applying approx.
imately fourteen pounds at the control wheel in the lateral axis
and about twenty-seven pounds in the longitudinal
axis.
g6 - 30643 - RESTRICTED
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PAGED6-3O643
22
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2.
Engage
Synchronization
Synchronization
loops
are provided
to
eliminate
engage
trans-
b_en _
_l_t
_w _li_en_aged, _he autopilot actuator is by_
draulically
de-ener_zed
a_d caged to _he null position.
The
sTncbronizer loop around the servo amplifier holds the servo
amplifier output near null.
As soon as the autopilot is engage@,
hydraulic pressure builds up first in the force detent mechanism
moving the autopilot actuator, which is still not pressurized,
to a position matching that of the control wheel.
Since the
autopilot LVDT is active, the voltage resulting from a change
in autopilot actuator position will be fed back to the servo
amplifier and synchronized within 0.25 sec.
Additional loops
pitch autopilot,
3.
_Dual-ChannF!
are included in the roll autopilot and the
to synchronize
the attitude con_ands.
O_eration
During dual-channel operation, both autopilot actuators are
coupled to the Jaanual controls via their respective force detents
as shown in Figure 5. The force detents are mechanized to give
the characteristics
shown in Figure 6a.
To reach equilibriwa, the forces applied to the cross link must
sum to zero.
The forces applied to the cross link are from the
two force detents, feel system and control system friction.
Since
the initial force gradients of the detents are steep, disagreement between the two autopilot actuator positions of more than
approximately
0._ aileron will cause one or both of the detentw
to reach its maximum force level.
If one channel fails hardover, the second channel and the feel
system will keep airplane controls in the trim position.
If the
second channel should command in the same direction as the hardover, the surfaces will correctly
as illustrated
in Figure 6b.
respond
to the second
Except
in the case of rather precise agre_ent
channels, the resultant dual-channel
autopilot
lesser of the two autopilot commands.
Several
can be concluded:
..
D
channel
between autopilot
command is the
pertinent points
.
(a)
A hardover command results
gible surfaces deflection.
in a passive
(b)
If the two autopilot commands
zero and the airplane remains
failure
are opposing,
in trim.
with
negli-
the output
is
D6. 30643 - RESTRICTED
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D.
FLIGHT
During normal operation,
the airplane will track the
autopilot command baying the lesser value.
If one autopilot fails passive, the resultant output is nearly zero.
Thus, the dual-channel servo actuator system provides true
"Fail Passive" operation for use in the _D
mode.
IKRECTOR
CP_ATION
AND MCDK5
The flight director and autopilot computation
paths become separate
Just prior to the autopilot path integrator as shown in the roll axi%and pitch axis computer block diagrams,
Conmand and rate limits are
incorporated
into the separate autopilot and flight director computer
circuits.
These limits are switched as a function of the mode selected
and the associated
submodes.
The Flight Director system characteristics are sunmarized in Table 5 on page
38 for the roll axis and
Table 6 on page
45a for the pitch axis.
As shown in the autopilot-Flight
Director mode charts (Tables i and 2
on pages 17 and 18), the flight director computes and displays the
navigational and vertical path data even when the autopilot engage
switches are in the OFF position.
When the autopilot is off, the
pilots can fly the displayed flight director commands using the primary
flight controls.
When the autopilot is engaged in MANUAL, the flight
director commands can be followed using the pitch wheel and turn knob.
All modes of the autopilot except turbulence are also provided for the
flight director.
There are two control modes which are exclusively
flight director modes.
These are go-around and back beam (optional).
Go-around provides a
wings leveling conmand in the ]_teral axis and a fixed pitch attitude
climb command in the pitch axis.
The go-around mode is initiated by
the pilot operating either of two go-around switches located on the
inboard throttle levers.
Operating a go-around switch will cause the
autopilot engage switch to drop to the OFF position from either MANUAL
or COMMAND.
The back beam mode is initiated by selecting the VOR/LOC mode, then
placing the solenoid held back beam switch in ON.
In this mode, the
flight director provides localizer back beam intercept, capture and
track conmands which can be flown by the pilot using either the turn
knob with the autopilot in MANUAL, or the primary flight controls with
the autopilot off.
During back beam, any one of several pitch control
modes can be chosen to provide flight director pitch commands.
D6 - 30643 - RESTRICTED
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E.
RCLL AUTOPILOY
The lateral autopilot computer block diagram is shown in Figure Ii
on Page 42. Gains, transfer functions, special gain programs,
mode engage and switching logic are summarized in Table 4 and
Figures 8, 9 and i0 on Page 37and following.
Finally, a block diagram of the lateral flight control system is shown in Figure 12 on
Page
43 .
The roll attitude and rate loops (see Figure
all modes of operation.
These loops provide
ing and spiral mode stabilization.
ii) are basic for
both roll mode damp-
The rate signals are sensed by the gyros installed in each roll
autopilot computer.
The roll attitude signals are from the Inertial
Navigation
System.
Fixed attitude and rate gains of
C
_AIL
= 3.2Deg
Deg
and
_,L
@t
_ 3_6
Deg
are employed
DeglSec
When engaged in MANUAL, the roll autpilot responds to bank commands
inserted via the turn knob.
When zero bank is commanded
(turn knob in
detent), wings leveling occurs after which the autopilot holds airplane
heading.
When engaged in COMMAND, the pilot has the option of control
by any of the following modes:
Heading Select, VOR/Localizer,
INS,
ILS, or LAND.
Prior to the autopilot engagement,
dual channel mode, synchronization
pilot engage transients.
i.
_Igage
either in the single channel or the
loops operate to eliminate auto-
Sync.hronization
In addition to the servo amplifier output synchronization
mentioned
earlier (See page 24), a second loop (co_uand synchronizer)
holds
the servo amplifier input near zero by synchronizing
the attitude
con_uand, prior to autopilot engagement in any single channel mode.
A thi_dloop, used only for the dual channel mode, synchronizes
the attitude co_and
by nulling the servo amplifier input through
feedback to the lateral path integrator of the channel yet to
be engaged.
This loop operates only after both autopilot switches
are in the command position and until the second autopilot channel
becomes engaged.
During the operation of this loop, the previously
mentioned command synchronizer loop is inhibited.
2.
Manual
The manual mode is engaged by placing the autopilot engage switch
in the MANUAL position.
The autopilot cannot be engaged with the
D6 - 30643 - RESTRICTED
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turn knob out of detent.
When the turn knob is in its center detent position, wings leveling occurs and the autopilot flys to
hold heading.
The heading reference is established by the Magnetic
_eadlngReTerence
_nlt (RHRU).
The turn
knob produces
a bank angle
cummana _i_nal
tD
.turn
knab Atsplassmnt.
_
ttumkn_b
output
is generated
by
a shaped
pot with a dead-zone in the center equivalent to ± 16 °
of knob rotation.
The maximum range of rotation of the turn
.s.
knob is ±128 ° which ccnnands ±30 e bank angle.
Maximum roll
rate for the Manual Mode is ±4°/sec.
Provisions are available
for a 7 degrees/second
rate limit.
When the magnetic heading clutch in the MHRU is engaged, the
heading error synchro is clutched to a magnetic heading repeater
and provides a heading error signal.
When the turn knob is
in detent, the clutch automatically
engages.
Proportional
and integral heading error signals produce the
bank angle ccmmandnecessary
to maintain the airplaneheading.
The integral of heading error reduces heading errors to zero
in the presence of thrust assymmetry or other lateral control
system m is-trim.
Both gains are scheduled as a functibn of
true airspeed to maintain consistent system performance
throughout the flight regime.
Certain airline customers will have control wheel steering
incorporated
instead of a turn knob.
The control wheel steering
option will be usable at all times when the autopilot is engaged
in Manual Mode.
In addition, when the autopilot is engaged
in "Command" roll control wheel steering is available in the
arm phase of lateral path modes such as LOC, VOR, or INS.
Thus the control wheel steering will be available to establish
the intercept angle desired prior to the capture maneuver
on these modes.
A special switch is included on the Mode
Select Panel as shown in Figure 2a.
This switch allows either
CWS or Heading Select to be used for the above mentioned
arm phases.
The roll control wheel steering block diagram is shown in
figure fla. Force on the control wheel commands roll rate via
an integral path. The integrator is bypassed with a displacement or "boost" path to minimize the velocity error or overshoot
which results when force is abruptly removed from the wheel.
Constant CWS gains are used over the speed range of the 747
so that light feel forces are present at all times in roll.
Two electronic detents are used.
The lower detent activates
the CWS mode and is equivalent to the turn knob detent.
The
higher value detent is used in the autopilot to disconnect
path modes and drop the engage switch from comm_nd tO manual.
This action overrides these modes and provides the CWS function.
O
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3.
Turbulence
Mode
The turbulence mode may be engaged at any time
the nay. mode selector is on ILS or LAND.
If the autopilot
mode is engaged,
except when
engage switch is in COMMAND when the turbulence
the switch will revert from COMMAND to MANUAL.
The system configuration in turbulence is identical with the
MANUAL configuration
except that the gains are reduced by
approximately
one half and the heading hold signal is removed.'$_o
Command
a.
Heading
Select Mode
The heading select mode allows
to fly on a desired heading.
the pilot
to use the autopilot
The desired heading is selected by means of the heading knob
on the Mode Select Panel.
The heading select mode is
engaged by placing the Nay. mode select switch in KDGand
positioning
the autopilot engage switch in COMMAND.
The command signal is the Heading Selector error (instantaneous
heading of the airplane minus the selected heading).
The
gain is scheduled as a function of true airspeed to malntain
consistent system performance throughout the flight regime.
The attitude command limit for the
±30" with the option for change to
500 feet/sec.
The roll rate limit
to iS deg/sec as a function of the
error.
b.
Localizer
Use
Heading Select mode is
_I0 ° for TAS above
is variable frum 1.Sdeg/sec
amount of heading select
Mode
of this mode requires
(i)
Tune
(2)
Dial in the runway
on the mode select
(3)
Position
in the localizer
the mode
the following
pilot
procedures:
receivers.
heading
panel.
with
selectswitch
the course
selectors
in VO_/LOC.
Dial in the desired localizer beam intercept
displayed on the Heading Select window.
(5)
!
O
Position
the Automatic
pilot
engage
switch
heading
in COMMAND.
The autopilot is in the Heading Select mode until
localizer capture sensor operates ahd the capture
initiated.
the
mode
is
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After capture, the system switches to the localizer
when the on-course logic is satisfied.
on-course
The basic damping
obtained by suing
signal in the localizer mode is ground
drift, angle with course error.
The use of ground
heading
is contingent
upon receipt
mode
heading,
of a drift
angle valid signal (DAV) from the INS.
If this signal is lost,
I
the roll computer automatically
reverts to the use of derived beam
I
rate, washed out heading, and lagged roll for localizer mode damping_." !
The beam displacement
and integral parameters
are the same for the drif_
angle valid and non-valid conditions.
The localizer syste, has three submodes of operation; namely: capture,
on-course, and on-course approach.
The circuit implementation
of the
syste, is such that it will automatically
switch to the proper submode
configuration
when predetermined
requirements
are satisfied.
(I)
LOC Capture
The iocalizer capture is initiated when a summed combination
of intercept angle and derived beam rate becomes equal to or
less than the instantaneous
beam error or when beam error is
less than one degree.
The exact
and gains are shown in Table 4.
capture
and on-course
logic
When the drift angle valid signal is present, the displacement
co,_nd
is the localizer beam error and damping is provided
by the derived ground headi n_ signal.
If the drift a r_le
valid signal is lost, the ground heading signal is removed
and
derived beam rate and course error are substituted.
(2)
LOC On-Course
The LOC on-course submode is initiated when
conditions of bank angle, beam displacement
satisfied, as su_narized in Table 4.
The localizer
configuration,
predetermined
and beam rate
are
on-course configuration
is similar to the capture
with the following modifications :
a signal proportional
to the integral of beam error is introduced to reduce the steady state error in presence of thrust
asymmetry and lateral mistrim conditions,
a thirty-three
second time constant high pass filter in the
ground heading and course error path washes out steady state
errors due to INS and course error signals offsets, as well as
heading errors due to crosswinds in the drift angle non-valid
condition,
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a signal proportional to lagged roll is added for increased
damping in the drift angle non-valid condition.
(3)
-
LOC On-Course A_proanh
autopilot
-.approach _
md_t, ches
at _
from the on-course
feet
ot' al__
sub._e
to
oa-course
Mechanization
of this submode is similar to that of the on-course
submode whe_ the drift angle is valid.
When the drift angle is
not valid, the washed out course error signal is removed to improve wind shear performance.
The beam displacement
and integral
gains are scheduled linearly with radio altitude to compensate
for beam convergence while the damping parameter gains are increased to improve close-in performance.
The submode gains, command
in Table 4.
C@
limits
and engage
logic
are summarized
VOR Mode
The procedure for the pilot to engage this mode is identical to that
of the localizer except that the VOR frequency has to be selected
rather than the LOC frequency.
The VOR mode has three submodes;
namely, capture, on-course, and over the station.
The basic damping
parameter is ground heading.
If the drift angle valid signal is lost,
the INS drift angle signal is removed, leaving the course heading error
as _he system damping signal.
The displacement
con_and is the beam
error for both drift angle valid and non-valid conditions.
System
gains are scheduled as a function of TAS to maintain good performance
throughout the flight regime.
The system gains, limits and mode initiation logic are summarized in Table 4.
(1) _pture
When
the VCR mode
the capture
mode which
established
Panel.
The
is first
selected
and the aircraft
is outside
threshold, the autopilot is in the heading select
steers the airplane to the desired intercept angle
with the heading select control on the Mode Select
capture sensor is armed.
The variable engage point logic used for VOR capture is shown
in Table. 4. For intercept angles between 90 degrees and i0
degrees, the capture starts at beam error between 1.8 and 0.2
dots.
The greater the intercept angle, the earlier the capture
maneuver is initiated.
A _4 degrees
course cut limit
is provided.
O
et
m
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2)
On-Course
The on-course submode is initiated when the bank
than 3 degrees and the course error is less than
angle is less
15 degrees.
The beam displacement
gain is reduced to one-half the value
used @_r1_
t_a_ture. A beam integral _igr_l is introduced
to reduce the beam error in presence of thrust a_try
and
lateral control surface mistrims.
A 2OO-second washout of the
ground heading signal is also introduced to eliminate INS
o,$.
drift angle and course error static offsets.
In the drift
angle non-valid condition, this washout improves the system
performance
under cross wind conditions.
Maximum position and rate commands are limited to + I0 degrees
and + 1.5 degrees/second
respectively
during the _R on-course
submode.
(3)
Over
the Station
The VOR over-the-station
sensor initiates this submode upon
detection of beam rates higher than 0.5 degrees/second.
The
system automatically
reverts to the on-course mode, after
passing the station, when the beam rate signal has decreased
below 0.5 degrees/second
for 20 seconds.
During the time that the airplane is over the station, the VOR
beam signal is removed and the autopilot command signal is
either ground heading for the drift angle valid condition, or
course error for drift angle non-valid condition.
If the pilot
desires to make a course change while over the station, he may
dial in the change in course setting and the system will track
outbound on the new radial.
d.
INS Node
The autopilot
circle routes
maybe
used to capture and track any of the great
that have been progra_ned
into the INS computer.
The INS mode is armed by placing the Nav. mode selector switch of
the autopilot mode select panel in the INS position.
Figure 7 shows
a typical control sequence.
If the aircraft is further than 7.5
nautical miles off the desired great circle course, the autopilot is
in the heading select mode, and steers the airplane to the desired
intercept angle established with the heading select knob.
In order
for the Heading Select mode to operate on those airplanes with
remote set preselect heading where the heading error signal is
developed in the HSI, the INS-RADIO switch must be left in RADIO
position until the INS capture point is reached.
When the 7.5 mile
point is reached, the INS mode is initiated.
Cross-track
deviation and track angle error outputs
used to compute the desired steering command.
O
of the INS are
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Control
a.
be
is similar
to VOR control
with
the following
exceptions
:
Since tP_ INS provides the equivalent of a non-convergent
beam, the distance from destination
does not effect system
performance.
The damping signal, track angle error, is not susceptible
_ress-_rlnd _f._c_
which w_'_l_ result in lateral displacement
from the desired
course.
?$
(1)
°
INS Capture
The capture maneuver is automatically
initiated at 7.5 nautical
miles from the desired INS course.
The bank angle command is
limited to 30 degrees and bank rate is limited to 4 degrees per
second during capture.
The system gains are summarized in
Table 4. A 25-degree course cut limiter is employed.
(2)
INS On-Course
The capture is complete and on-course control is initiated
when the cross-track deviation is reduced below 1,070 feet and
track angle is below 3 degrees.
The sa_e gains are employed
in the autopilot as during capture.
The on-course sensor
_wnshifts
the bank and bank rate limiters to I0 degrees and
1.5 degrees per second respectively.
(3)
Waypoint
switching
Automatic switching from one great circle course to another
is provided.
On the INS system Control and Display Unit,
the "Auto-Manual"
switch must be set to the "Auto" position
to use this feature.
If this is not done, the autopilot
will overfly the waypoint and continue on the extension
of the great circle.
If the "Auto" sequence is used,
switching of the autopilot to the next route occurs a
short distance before the waypoint is reached.
(See Table 4)
It will occur at 3.5 nautical miles for cases where the angular
change between successive courses is small.
Restoration
of the capture bank limits is provided automatically.
When the on-course conditions are again satisfied,
the reduced
bank and bank rate limits are automatically
reinstated.
ee
ILS Mode
The ILS mode is identical to the Localizer Mode for the roll autopilot.
Glideslope control is armed by this mode and the autopilot continues
to fly toward the glideslope beam on either pitch attitude, vertical
speed, altitude or IAS hold (optional mode) until a predetermined
glideslope
signal level is reached.
(See pitch axis system description
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Land Mode
The autopilot
dual-channel
features dual
LAND mode provides
the Boeing 747 with
automatic approach and landing system.
ILS and flare coupling.
The LAND mode is properly
are satisfied:
selected
when
a
the following
switch
fail-passive
The LAND mode
prerequisites
I)
Nay. Mode Selector
2)
Both course
3)
Heading select control set for the desired
angle with the localizer
4)
VHF/NAV
5)
Both autopilot
controls
receivers
in LAND
set to the runway
set to the proper
engage
switches
heading
,_o
intercept
localizer
frequency
in the command
position.
Upon selection of this mode, single-channel
operation, identical to that for ILS mode, is initiated.
Dual-channel
operation
does not begin until after the autopilot is on Loc. approach,
glide slope capture is completed, and the airplane is less
than I,500 feet altitude above the surface.
The second channel
synchronizes
to the controlling channel until it is engaged;
at this point, the flare computer is armed and equalization
and monitoring of both channels begins.
(1)
Equalization
Equalization is accomplished bytaking
the difference of
the central control actuator and the autopilot actuator
LVDT signals and feeding this signal back to the autopilot
path integrator.
The output position of the central
be equal to the output position of
having the least value (see Section
having the greater command will be
central control actuator output by
the A and B autopilot commands.
control actuator will
the autopilot actuator
C).
Thus, the channel
different from the
the difference between
The channel with the lower command signal receives no
equalization signal since the difference between commanded
and output position is zero, while the channel having the
higher signal receives an equalization.that
tends to
reduce its output signal to match that of the'controlling
channel.
It should be noted that each channel has an
independent equalization
system and that there are no
cross ties between channels.
The equalization signal in the lateral axis is limited,
so that ramp faults of relatively
low values may be detected.
The equalization
signal is also gain scheduled as a
function of radio altitude from i to 0.5.
Thus, equalization is decreased as the airplane approaches touchdown.
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(2)
Monitoring
The monitoring system on the 747 dual-channel
autopilot
acts on %he meaB_Ted diff_L_,,_e between _he output positio_
_f each autopilo% actuator and the output position of the
central camtrol actua%or associated with that autopilot
actuator.
A block diagram of the monitoring
system is?s.
shown in Figure 13.
The two actuator position sensors
used in the equalization
circuit also feed the monitor.
Each channel has a pair of monitorsone for pitch and
one for roll.
A failure is indicated by the monitor if an
autopilot channel disagrees with the central control
actuator output by a given amount for a set length of
time.
These values are presently set at Six degrees and
w o seconds in the roll channel.
A steady red cockpit
warning light is initiated any time a monitor is tripped.
Thus, early warning of any potential problems is brought
quickly to the pilot's attention without disconnecting
the autopilots.
Should the warning light be activated
by noise or radio beam masking, it is not latched to the
ON condition; and thus, goes out at the conclusion of
these temporary disturbances.
' i
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FQ
PITCI{ AUTOFILOT
A summary
contained
of Autopilot/Flight
in Tables 6 a_d 7.
Director
pitch
axis characteristics
is
The 747 autopilot _Fera_s
through a parallel servo system.
Thus all
actuator response due to a command from the autopilot is also fed back
to the control column.
The feedback path is the 1_
elevator
control system.
At the frequencies which the autopilot system normally
operates, the control column can be assumed to be directly proportional
to the elevator surface votion.
Figure 23a shows the linearized transfer
functions for the Au_pilot
servc a_d elevator power control unit.
The
normal gain between
_column
and O elev. is .478 degrees of column
per degree of elevator displacement about surface neutral.
The full
column to elevator curve is shown in Figure 23b.
Figure 14 is the block diagram of the basic pitch autopilot control
system.
Pitch attitude and rate loops are basic for pitch modes of
operation.
These loops provide both short and long period stabilization
and damping.
The rate signals are sensed by gyros installed in each
autopilot computer.
System (INS).
Fixed
Attitude
attitude
signals are from Inertial Navi_tion
and rate gains of ___
= _ _ _eg and
2.2 Deg/Sec are employed for flaps
flight.
Rate
damping is increased from 2.2 to 3.5 for flaps down flight.
The pitch
rate is passed through a i/(.is+l) lag for higher frequency suppression,
a i/(.O5s ÷ i) lag for structural mode decoupling and a washout
ts/(ts + I) where t = 2 for flaps up and to provide roll compensation.
The compensation
signal (1-cos _) is gain scheduled as a function of
airspeed.
When the autopilot is on, the automatic trim system maintains pitch
trim of the airplane.
This is true for all autopilct modes except
turbulence.
When this mode is engaged, pitch trim is not active.
When engaged in MANUAL, the pitch autopilot responds to commands inserted
via the pitch knob.
When engaged in CO_AND,
the pilot has the option of
control by any of the following modes:
ALT. HOLD, ALT. SELECT, IAS HOLD,
V/S (Vertical Speed Control), MACH HOLD, ILS, or LAND.
Engage
l.
S_nchronization
Before the autopilot is engaged, the output of the servo amplifier
is fed back to the path integrator at a high gain.
This loop maintains the output of the servo amplifier at zero in order to obtain
transient-free
elevator when the A/P is engaged.
At engagement,
the
synchronizing
path is opened.
The integrator hold circuit tracks
airplane attitude which serves as the reference for attitude hold,
the initial mode of the A/P.
Q
J.
Manual
The MANUAL mode is engaged by placing the autopilot engage switch
in the MANUAL position.
Synchronization
is provided so there is
no attitude transient when the mode is engaged.
a.
u,
n
PITCH WHEEL:
If the pilot desires to change the airplane
pitch attitude, the pitch wheel on the flight controller
is used. The pitch wheel produces an attitude command
proportional
to wheel displacement.
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CONTROL WHEEL STEERING (CWS):
(CWS is alternative
to
Pitch Wheel.)
CWS enables the pilot, by applying a force
to the column to insert a command signal into the A/P to
change the airplane attitude.
The force signal is processed as in Figure 14a.
Force on the control column
commands pitch rate via an integral path, the gain of
which is programmed with True Airspeed to give uniform
performance over the flight regime.
When the force is
removed, the path integrator is synchronized
to the
airplane attitude at the time of force release.
When the A/P is in MANUAL or in COMMAND and not in any path
mode, the A/P will be in the CWS mode.
If in COMMAND and in a
path mode and the high detent force (19 pounds) is exceeded,
the path mode drops off and the A/P drops to MANUAL except
when in Altitude Hold.
When in CWS and the deadzone (low
detent) is exceeded, the automatic trim is inhibited.
Whenever
an attitude of + 25 degrees is reached and the pilot applies a
force in the direction to increase this attitude, the rate
path is inhibited (MANEUVER LIMIT CONTROL).
Be
Turbulence
In a turbulent environment TURB. may be engaged on the Turb/
Speed mode select switch when the NAVmode
switch is not in
ILS or LAND.
With TURB engaged, the autopilot is automatically
switched from COMMAND to MANUAL and the flight director pitch
trim control becomes effective.
The pitch attitude and pitch
rate gains are reduced by one-half.
The automatic stabilizer
trim is off in the turbulence mode.
O
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ao%
r.AS
J_G/FTmr=c
.032
t
eEGmGG
mS
ATTITUOE
1
OEWDL_
.13
lb/Av
ATTITUOE
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MAtH ERROR
2S
12114SX.154 IIL0$S÷
ID[O
FCN OF DEG
ALT HOlD
ATTnlR)(
AIRSPEED NOLD
VERTICAL
IFIE|D
I¢OLD
OI[G/OEG_I¢
I&J
14LTUOLD
PROGRAM
IN/OEO
JIO
AJ_lr _
GRADIENT
MACH HOI.D
UNITE
.059
lIAR SENSITIVITY
IAII I¢OLD
REMARKS
OEGnCT
I
OEG/OEG
95
OEGAIN:H
!
OEG/Or_
ATIlllJOE
• 16"_ t
PATH FILTI[R Am_
LRIITER AT tltll:
TO ALL
ImX)ES I_OWN
3)
1#s.1
IWA&
VIEWncJ_
IPEED ERROR
ATTITUDE
mEG_r_-c
.05
CAPTURE
1
I¢Ot_
I1
DEG/DEG
s}
SE
I+1
20S
laOS+ !
.o0_5
_aN
ALTITU_
ERROR GAIN
OEG/IrT
all
HOLD
ALTITUOE PATE
,05
DEG
-092
FT;SE¢
I
ATTIll_E
OEG/OEG
I.SS
I.SS¢' I
m
I
14S+ 1
¢Nmml
FLARE
6AIN
BEAM ERROR
k4
PATH DAMPING
I_
.21
DEG/FT/SEC
28
DEG/IDEG
r.,Jmme
AND NOLO
mNK RATE
ERROR
2.5
_+
.m
OEG/FT/SEC
DEG/SEC
FT/SEC
FLAME RATIO
Wk
.n
DAWING
.lip
DEG/SEC
.t2
.2
DEGSEC/FT
IITCH RATE
DEG/I)EG/SEC
IqTCH
ATTITUOE
DEG/DIEG
WUP
I .lesS+ 1 108+I
.--I-.35+I
laNK RATE
INTEGRAL
GOAR(XJIR)
SINK RATE llAII
12 FT/SEC
_.J._
35+I
GAIN COMMON TO
AUTmqLOT
I
_U
10_
(108+ !)(.IS+
,
1
1)
4S+I
12
EiAII
DEG
_
m
1)
AUTOMATIC IrrAIMLIZER TRIM IS INHIBITED DURING THESE IdOOES.
SEE DIAGRAM 1 FOIl CASGAIN PROGRAM.
:ll
4)
THE GAiN PATH NICU.IOEIi THE .1__
PATH FILTER AS WELL AS AN S DEG (k: AMPLITUDE LIMIT.
WITH THE RADIO ALTIMETE R, VALID SIGNAL PRESENTGAINS ARE PROGRAMMED AS PER DIAGRAM 2 WiTH RADIO ALTITUDE.
OTHERWISE, THEY ARE PROGRAMMED WiTH THE TIME BASE PROGRAM SHOWN IN DIAGRAM 4.
l)
SEE DIAGRAM 3 _
@)
IN FOR GLIDE SLOPE CAPTURE, TRACK; FLARE; GO AROUNO ONLY,
.
TAS GAIN PROGRAM.
j)6
- 30643
-RESTRICTED USE - See Notice on Cover....
._
%,
REV SYM
Fligh_
Table
6
Director
Fitch
BgE/AY_
Axis
NO.
I_'
VOI.
II
i_AO_ D6-3o6_3
h5
6-7000
F
,¢d'"
AUTOPILOT
PITCH
AXIS
GAIN
PITCH
:
'
AUTOPILOT
MODE
PARAMETER
SYMBOL
ATTITUDE
ATTITUDE
_l/e
SUBMOOE
UNITS
VALUE
OTHER
SHAPING
OEG/DEG
HOCO
(.oss* Q
_'.Z
GLIDE
SLOPE
AND
DAMPING
6el6"
(.05S
FLARE
FLAPS
UP
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OEG
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F LAPS
COMP
PITCH
Co/(
KNOB
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1).,
(.IS+
II(2S
(.lS+
li
+
li
i.05S
÷ I)
BODE
LOSS
BAND
NOT
AT
,¢
MID-
INCLUDED.
9.5
DOWN
---
8c/STROKE
*
._.--.Z.%-_
DEG/SEC
LIFT
REMARKS
PROGRAM
18.9
---
(lOS+
_
._s_
OEG/STROKE
---
,05S
1) (.05S+
I1
CAB 2 )
30°$
* 1
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IS
KNOB
THE
CAN
WITHOUT
AMOUNT
BE MOVED
REMOVAL
OF
HAND.
TURBU-
1)
ALL
ABOVE
_4/_¢
----
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THE
LIFT
---
COMP
REMAINS
UNCHANGED
LENCE
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ABOVE
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i CONTROL
WHEEL
I)
CWS
STEERING
CWS
(OPTIONAL)
GAIN
ALT
GAIN
8¢/
LB
STICK
FORCE
INTEGRAL
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RCE
---
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T_S
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LB
ALT
HOLD
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B
--
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1
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A he
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HOLD
RATE
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---
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---
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----
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FILTER
LIMITER
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TIVE
TO
IS ALTERNA-
PITCH
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AND
AT
COMMON
MODES
AIRSPEED
MODE
+ 1
PATH
DISPLACEMENT
lAB
SEC
THIS
CAB2)
31
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INTEGRAL
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ALT
= .35
0
---
8°_k:
TO
ALL
SHOWN
DEG/KT
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HOLD
INTEGRAL
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RAL
NO
MACH
MACH
HOLD
GAIN
V
GAIN
DISPLACEMENT
PROGRAM
HOLD
0c/MACH
HOLD
Oc/JMACH
----
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120
NO.
INTEGRAL
(OPTIONALI
MACH
INTEGRAL
.05
DISPLACEMENT
_¢IMACH
CAB
VERTICAL
VERTICAL
8c/_t_
---
VERTICAL
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I_
8c/A
hf
2
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RATE
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TED
TO
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2
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d.
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ALTITUDE
ALTITUDE
SELECT
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AND
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CAPTURE
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SEL
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GLIDE
DEG/FT
DEG
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C_
MAY
fi
DISPLACEMENT
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BEAM
INTEGRAL
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0c/J(
CONTROL
INTEGRATOR
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IN
FOR
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BIAS
ONLY
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I
FT/SEC
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RATE
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AS ABOVE
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CAPTURE
LIMIT
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4)
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OR
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FOR
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EE
TIME
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OF
VALID
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785.1
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oO
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THIS
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CONTROL
PATH
TO
THE
PATH
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RATE
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---
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ALSO
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FEEDS
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ERROR
S_NK
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---
.64
INTEGRAL
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MEMORY
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2,
SEE
THE
4)
WITH
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DIAGRAM
GAIN
THE
30
DEG/SEC
I
PATH
RADIO
FOR
TRIM
CAS
INCLUDES
GAIN
THE
ALTIMETER,
WITH
IE
FOR
THE
TAB
iS
INHIBITED
THESE
MOOES.
I)6 - 30643- RESTRICTED
USE- SeeNotice
onCove_
PROGRAM
_.
VALID
TIME
GAIN
DURING
PATH
SIGNAL
FILTER
PRESENT
5)
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SEE
FrGURE
6)
7)
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AT 3.0 DEGREES
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CAM
OUT
LEVEL(_
DEGREES
ELEVATOR
DIFFERENCE.
AS WELL
GAINS
AS AN
ARE
S DEG
0¢ AMPLITUDE
PROGRAMMED
AS
PER
BASE
PROGRAM.
PROGRAM.
LIMIT.
FIGURE
IS
WITH
RADIO
ALTITUDE,
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THEY
ARE
; .......................
T
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LIGHT'DELAY
I
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..............
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D6-3o61_3
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Vol.
NO..EI_
_8
PAGE
II
14'-l-
4
_:) (9_
"_
6-7000
4.
Command
When engaged in CGMMAND and no other pitch mode is selected, the
A/P is in pitch attitude hold and the pitch knob is operative.
a.
Altit_e
L
I
_Id
The altitude Hold mode holds the airplane at the altitude
,
existing when the mode is engaged.
This mode can be engaged, in
either MANUAL or COMMAND.
If the mode is engaged with the
airplane climbing or descending
at a reasonable rate, the air-
l
[
I
I
plane
[
returns
to and holds
the engage
altitude.
I
The Air Data Computer provides the reference signal
mode.
This altitude error signal is provided
after
a mechanically nulled synchro to the altitude shaft
engage.
Aircraft altitude rate (or vertical speed)
sensed and provided
The block
included
diagram
by the Air Data
and gains
as part of Figure
for this
clutching
at mode
is also
I
]
I
I
Computer.
for the Altitude
Hold mode
l
i
I
are
[
15.
I
The altitude error signal commands attitude changes required to
maneuver the airplane toward zero altitude error.
Altitude
rate is added to improve damping.
Integral control on the
altitude error removes standoffs.
The gains used are programmed
down at high speeds to about I/3 of the.low speeds.
(See
i
[
I
I
I
Figure
I
16.)
-
I
be
Altitude
Select
"
[
[
The Altitude
Select
mode allows
the pilot
to select
a desired
flight altitude.
If the selected altitude is more than 1200
feet from actual altitude, the pilot also selects the desired
mode of climb or descent.
With these selections made and the
mode engaged, the autopilot maneuvers the airplane to smoothly capture and hold the selected altitude.
This mode is particularly useful when a number of successive altitude changes
]
[
I
I
I
I
I
are required.
I
I
Altitude
Select has three
(See Figures
(1)
Arm
submodes.,
arm,
capture,
and track.
I
15 and 1 7. )
-
I
I
"
I
I
In the arm submode, the pilot selects some other pitch
mode of the autopilot,
such as vertical speed, IAS hold,
or pitch knob, to command the aircraft to climb or descend
toward the desired altitude.
In this submode, the altitude select switch is engaged and the capture logic is
I
I
I
I
I
armed.
I
D6 - 30643
REV SYM A
- RESTRICTED USE - See Notice on Cover
Bd_'_"J_V_
/ 8
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I
Vol.
II
r6-3o6__3
PAGE 47
J
6.7000
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(2)
C_t_e
Capture is initiated when the aircraft' s altitude approaches the selected altitude to within 1200 feet, and the altitude error (in feet) is less than X5 times the altitude
rate (in feet/sec.) as sensed by the coincidence
detector.
The detector uses course altitude and rate data from th_CADC to determine the switching point.
When the detector
is satisfied, it switches off the previously used climb
or descent mode and initiates capture.
During capture,
the fine altitude data from the CADC and altitude rate
information are summed as indicated by the following control equation to cc,,,and the aircraft to maneuver and
capture the selected altitude-
ae,
where:
_u
-
G. (.ooc_b,
+.oq r,)
is the gain program with
(unity at low speed)
n ec is pitch
attitude
A k is altitude
error
is rate of climb
command
computed
airspeed
in degrees
in feet
in feet per second
This produces a flare toward the selected attitude which
is rate limited to reduce "g',forces if initiated from
steep climbs or descents.
(3)
Track
The track phase is initiated when the airplane approaches
within lOO feet of the selected attitude, with a rate not
exceeding 5 feet/seconds.
The control equation is as
•follows:
• @c
c.
Vertical
-
G_, (.029_k
. .0'3
K)
with
the units
above.
defined
as
Speed
The vertical speed mode is selectedby
the Turb/Speed
switch on
the Mode Select Panel.
Priorto
selection, the vertical speed
wheel on the Mode Select Panel is synchronized
to aircraft
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vertical speed.
It is driven by an electro-mechanical
servo
follow-up to accomplish this repeating function.
Upon engagement of this mode, the motor of this follow-up de-activates
to provide the reference required to generate a vertical speed
error signal.
Should the pilot desire to operate at a different
vertical speed, the command wheel is moved to establish a new
between
the
_u_i_s
of
operation
which
are
+4000
ft/minute
to
-80oo t/ nu e.
I
The block
diagrLm and gains
used for the vertical
speed mode ar_ _"
shown in Figure
15.
An acceleration
l_lter
provides
smooth
system response to con_and changes.
Gain scheduling is included
to allow good performance over the speed range of the 747.
IAS Hold
de
This mode gives
capability.
the pilot
automatic
Indicated
Airspeed
hold
A clutched synchro on the computed _rspeed shaft of the CADC
provides an IAS error signal which is the difference between
actual IAS and the IAS at the time of mode engage.
The error
signal con_ands pitch attitude required to hold the reference
airspeed.
The block diagram and gains used are shown in
Figure 15.
.
Mach
Hold
The Mach Hold mode provides the capability to automatically
hold
the airplane Mach existing when the mode is engaged.
The mode
is selected by the Turb/Speed
select switch on the Mode Select
Panel.
The Mach error signal is generated in the CADC by a
clutched synchro.
The error signal commands the airplane
pitch attitude required to zero the Mach error signal.
The
block diagram and gains are shown in Figure 15.
f.
ILS Mode
The pitch axis control system for dual channel automatic landing is
shown in Figure 19. The ILS mode uses part of one channel,
that
is, glide slope capture and control.
Upon selection of ILS, glide
slope capture is armed and the glide slope indicator light on the
mode annunciator
panel is amber.
During the glide slope arm phase,
a glide slope intercept path from above or below the glide path
can be flown on ALT. HOLD, ALT. SELECT, MANUAL, MACH
IAS, or V/S.
When the glide slope receiver is within _ 30 mv (¢.i_ °) of beam
center, the vertical beam sensor is tripped and glide slope capture
begins.
A sink rate error signal, barometric
altitude rate plus
a 12.5 fps bias, is switched into the path integrator for I0 second
to produce a pitch angle
the glide slope.
approximately
equal
to that needed
to fly
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'-1000
Normal acceleration,
through a washout to eliminate steady state
accelerometer
outputs and filtered through a iO second lag to
pseudo-integrate,
is switched in for beam damping.
During the
i0 second pitch down, beam error is synchronized
by a limited
integrator at the integrator' s high rate.
At the end of the IO
seconds, the sink rate error signal is switched off and glide
slope control begins.
Beam error minus the synchronized
signal stored on the rate-limited
integrator is switched in, with this stored value bleeding down
at the low rate of the integrator to give an easing on of the signal.
On glide slope, beam error is programmed with radio altitude
to provide a nearly constant elevator deflection per foot of beam
error.
The gain programmer is shown in Figure 18.
In the event of
a radio altimeter failure, a time-based gain program is also pro-
vlded. (Figure18-2).
g.
Land Mode
The LAND Mode is shown in Figure 19.
It is selected on the NAV
mode selector when dual-channel,
fail passive localizer, glide
slope, and flare is planned.
Sequencing during the LAND mode is
shown in Figure 20.
Upon selecting of the LAND mode and with one channel in
C_AND,
the autopilot is in single channel operation.
The
mode is the same as ILS except that the autopilot warning
light flashes amber and the autopilot will automatically
disengage at 150 ft. if the second channel is not engaged.
The second
channel
is locked
in OFF until
the automatic
con-
fidence test of the dual-channel
camout monitor is completed.
This takes about I second after LA_
is selected and the
first channel is engaged in COMMAND.
Upon moving the second
channel engage switch to C_iAND,
the flashing amber warning
light is extinguished and the second channel is armed.
The
systemwill
operate in this configuration
until the dual
channel engage interlock logic is satisfied.
The autopilot remains in single channel
of the following conditions are met:
operation
until
all
LOC on-course
Rad. Altitude less than 1,500
Glide Slope control
Rad. Altitude valid
ft.
When these conditions are met, dual-channel
operation begins
and is annunciated after a three second delay by an amber
A/P FLARE ARM light on the Mode Annunciator
Panel.
The nonlatching autopilot camout monitor gives the pilot a stea_v
red warning light on the Mode Annunciator
Panel if a
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6-7000
camout
exists
(Page 61) in either
Interlocks for the LAND mode
Figure 21.
(2)
channel
are shown
for two seconds.
in flow diagram,
Synchronization
When in IAND with both channel engage switches in CO_4AND,
but still flying single channel, the second channel actuator
is in the caged position and its servo amplifier output is
being nulled by the path integrator by means of the synchronization loop. At A/P FLARE ARM, the hydraulic pressure
'_"
builds up first in the force detent mechanism moving the
autopilot actuator to a position matching that of the elevator.
The-LVDT signal generated when the autopilot actuator
is moved into position is fed back to the servo amplifier, and
is hulled out by the synchronization
loop.
The actuator is
then pressurized and the synchronization
loop is opened.
Since
the servo amp was held to zero, the actuator does not move until
it is commanded from the autopilot.
(3)
Glide
Slope
The glide
slope control
of ILS. Figure
slope control.
19 shows
law for each channel
all switches
is the same as that
in position
for glide
During dual-channel operation, equalization
is in effect, without signal intertie between channels.
An equalization
con_and
proportional
to the difference,
if any, between the autopilot
actuator and the elevator is fed back to the path integrator.
The equalizer signal to the path integrator is limited.
The
equalization gain _s programmed down as a function of altitude
since the requirement for equalization
to reduce null offsets
is likewise reduced.
(4) Flare
At the flare point, the FLARE
changes from _mber to green.
light on the mode annunciator
The camout monitor remains active.
Since the gain programmer
is zero, beam error and the gain
programmed portion of equalization
are not applied during
flare.
However, a small fixed gain equalization
signal remains on throughout flare.
The flare
he
law:
_+
+2;
ER=h
--_@a
commands an airplane descent rate, linearly decreasing with
altitude, from the descent rate at flare to 2 feet/second.
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9
The descent rate is derived from the radio altitude
and vertical accelerometer
signals.
The descent rate
error is su_ed
with filtered acceleration
for damping
and then passed through direct and integral paths to
produce an attitude co_a_i.
A synchronizer
loop mainrains
the command to zero
until
flare.
The flare command is limited
now down pitch attitude.
C5)
Fail-Passive
to 7.5° nose up and .75°'_-
Actuators
The autopilot actuators are located on the two inboard
elevator control packages.
Figures 22 and 23 show the
autopilot elevator control configuration.
The feel computer
provides
a centering
force
for all auto-
pilot and manual control commands.
Feel force is programmed as a function of dynamic pressure and stabilizer position and provides authority limitation to the autopilot.
The autopilot actuators are tied to the main valve inputs
through force limited detents.
In single-channel
operation
only one detent is engaged and both main valves follow the
engaged autopilot actuator.
Figure 5 shows a schematic of the detents and servosystem
mechanization
for theroll
axis.
Pitch mechanization
is
similar.
In dual-channel
operation, both autopilot actuatios are powered and the detents are engaged.
If the
autopilot actuator positions disagree, the elevators
will follow the command nearer zero.
The detent monitors
associated with each autopilot channel measure the disagreement of the autopilot actuators with the output.
A
disagreement of 6.0 ° elevator will give a camout and a
consequent pilot warning light after a one-second
delay.
Dual-channel
elevator authority is double the singlechannel authority or approximately
+i0, -30 degrees at final
approach speed.
The detent authori_y remains constant
at 27 pounds stick force, maximum.
At maximum q conditions
elevator single-channel authority diminishes to + 1.5
degrees elevator.
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4.
AUTOPILOT
AUTOMATIC
STABILIZER
TRIM
The autopilot is provided with two separate stabilizer trim systems.
During single-channel
cruise operation of the autopilot, the A trim
system is utilized with the "A autopilot"
_
the B trim system is
utilized with the '_ autopilot".
During dual-channel
operation of
the autopilot, both trim systems are capable of operation.
One of
the two is _tilized and _he ohher is armed and in a standby state.
Suitable monitoring and logic is included to effect an automatic
transfer to the standby trim, should a malfunction occur in the
active channel of trim.
The auto-trim unit drives the stabilizer to reduce steady state
elevator displacement
from neutral when the autopilot is engaged.
This reduces to a low level the transient which occurs when autopilot
modes
is disengaged.
Auto-trim
except Turbulence.
is obtained
during
all autopilot
The stabilizer has no direct connection with the primary control
system.
Thus, motions of the stabilizer show up on the column only
as the amount of elevator surface which must be held on the column to
maintain airplane trim. When the autopilot is engaged the autopilot
commands the amount of elevator needed to maintain the desired flight
path and airplane trim. When the autopilot is co_m_nding an elevator
position greater than the trim threshold (Figure 24) the trim system
will drive the stabilizer until the elevator required to maintain
the desired flight path is reduced below .185 degrees of elevator.
The effect of automatic trim as seen on the column is a smooth return
of the column to near neutral as the trim system operates.
The rate
of stabilizer
trim is inversely proportional
to the impact pressure
(_).
Thus, the column rate of return is similarly reduced.
Autopilot Stabilizer Trim Unit (ASTU) channel is shown in Figure 24.
The arm and control circuits are identical.
Trim is effected by the
presence of discretes from both arm and control when there is hydraulic
pressure to operate the brake pressure switch.
When the elevator
exceeds the trim threshold for five seconds, the stabilizer is driven
until the elevator decreases to 0.185 degrees.
The trim rate and
thresholds are variable with feel pressure as indicated in Figures 24b
and 24c.
.2
The trim warning monitor is activated 8.5 seconds after
passive failure of the arm or control circuits or brake
A warning is also obtained for an out-of-trim
condi$ion
12 seconds.
Dual-Channel
Autopilot
Node
an active or
pressure switch.
sustained for
(Autoland)
One auto-trim channel is engaged in this condition; and in the event
of failure, this channel is automatically
disengaged and the other
standby channel is engaged by the changeover
switches in the Boeing
accessory box.
Thus, "fail operational"
trim is provided during
automatic landing.
Should a camout occur, the ASTU is inhibited from
moving the stabilizer.
D6 - 30643 - RESTRICTEDUSE - See Notice on Cover
D6-3o643
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p_,E 61
II
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II.
MACH TRIM SYSTEM
.iS
(SYSTEM
NOT INSTALLED
o
ON 747 AIRPLANE)
Pages 63 - 68 deleted.
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;3o643..
NO. Vol.
PAOC
63
II
96.7000
III. YAW DAMPER
A.
SYST_
GENERAL
The 747 directional control system comprises dual rudders, each
independently
powered by a dual tandem hydraulic actuator.
Dual
redundant yaw damping is provided on the 747 by incorporating
an
independent
electro-hydraulic
augmentation
system with each rudder
segment.
iS.
Yaw damping signals are connected to the rudder actuators in a series
fashion such that the pilot's rudder pedals are not displaced by
yaw damper commands.
This feature permits the yaw dampers to be
operative at all times through take-off, cruise, and landing without
interferring
with normal pilot rudder control.
A preflight cockpit operated confidence test is provided to check
each yaw damper system prior to departure from the ramp area.
Disengage switches mounted on the pilot's overhead panel are provided
to enable the flight crew to shut off either yaw damper system should
a malfunction
occur.
Figure 27 is a pictorial diagram of the lower
rudder yaw damper system.
In addition to providing additional damping of basic airframe lateral
directional
oscillations
(dutch roll), the 747 yaw damper system has
an added feature designed to improve airplane response to turning
maneuvers in flap down flight conditions.
This system is called the
,turn coordinator" and deflects rudders proportional
to roll rate
in a nturn coordinatingU
sense, thereby improving roll control
re sports e.
Due to a favorable phase relationship,
the turn coordinator
has the added benefit of further improving basic dutch roll
beyond that available from the yaw damping mode alone.
feature
damping
The .turn coordinator" system is used at flaps down condition only.
An .easy on/off" circuit is implemented
in the "turn coordinator"
command path to eliminate transient rudder kicks resulting from the
flap switching should the aircraft not be at zero roll attitude.
The design objective is to provide additional dutch roll damping and
turn coordination
with a system which provides no potentially
hazardous failure conditions.
(See document D6-13647, IEFCS Failure
Analysis ).
USE- SeeNoticeon Cover
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BOEINO
I _
,_o. vol.
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69
"
B@
SYST_
DESCRIPTION
Figure 28 is the block diagram of the 747 dutch roll damper
turn coordination system.
Its
mathematical
model is shown
Figure 29.
I@
and
in
_Dumb
Dutch roll damping Is provided by a yaw rate (_) signal.
At flaps down condition, additional bank angle signal is
provided to increase system damping.
Yaw rate is sensed by
a rate gyro, mounted in the yaw damper chassis.
Bank angle
is obtained from the Inertial Navigation
System (INS).
These signals are independently
demodulated.
2.
Band
Pass Filter
At flaps up condition, the yaw rate signal passes through a
band pass filter into the servo amplifier.
The band pass
filter is composed of R-C components and operational amplifiers.
The transfer function of the filter can be expressed in
Laplace form as the following:
2.72S
(2.72S*l)
(.272S,l)
At flaps down condition, roll attitude signal passes through
a similar band pass filter and is su_aedwith
the filtered
yaw rate signal into the servo amplifier.
This additional
filtered roll attitude signal provides additional dutch roll
d
iug.
The functions
(i)
(2)
(3)
of the band pass
are:
To washout the steady yaw rate and roll attitude
and to eliminate null offset of sensors.
To reduce high frequency signal
minimize possible coupling with
Dutch Roll
Damper
At flaps down
gains are
2.5
respectively._,
of the band
so as to
modes.
pass filter
are
C_in
condition,
_
_
amplitudes
structural
the yaw rate
(degree)
(degrees/sec)
lOAd1
and
and the roll
.076
attitude
_R (degree)
_ (degree)
. RFSTRIL"IED USE - See Neti¢e on Cover
,
REV SYM
signals
To provide dutch roll damping signals with minimum phase
shift at dutch roll frequencies,
in order to achieve
optimum damping.
The Bode and phase angle plots
shown in Figures 31 and 32.
3.
filter
_'_-/N_
D6-306]+3
I NO. Vol.
|_
IP^GE
70
II
6-7000
At flaps
_R
up condition , the yaw rate
(de_ree)
.
root
loci plot.
Turn
Coo_a_on
For detail
gain is 1.25
information,
refer
to Figure
33
m
A shaping circuit is used to derive a roll rate signal from
the roll angle input.
The rate circuit used has a transfer
function
.is
of (.is +i) z
which
yields
a derived
roll
rate
-'_"
signal
from roll angle at turn entry frequencies, but cuts off at
frequencies
above 1.4 cps in order to minimize the effects
of system noise.
The turn coordinator system
conditions only.
Switching
The system
gain
is 0.693
operates at flaps down flight
is accomplished
by a flaps switch.
_R
(de_ree)
_ (degree/sec)
.
Easy On-Off
Circuit
An easy on-off circuit is implemented
in the filtered roll
attitude signals path to eliminate transient rudder kicks
when the roll attitude signal is turned off or on by the
flap actuated switches.
.
Yaw Damper
Electro-Hydraulic
The servo valve
amplifier
Servo
accepts
dutch
roll damper,
turn
coordinator,
and servo feedback inputs and provides an output
to the electro-hydraulic
transfer valve.
The transfer valve
controls motion of the yaw damper actuator which is linked
to the main power control unit valve via a summing link.
The electro-hydraulic
transfer valve is supplied system
pressure via a solenoid operated shut-off valve when the yaw
damper is engaged.
The yaw damper servo actua@or is selfcentered by caging springs when the system is de-energized.
This preserves the integrity of manual commands when the yaw dampm
is not energized.
The maximum rudder rate which the yaw damper can command
controlled by the area of the servo actuator orifice and
the area of the piston and is + 15 deg/sec at no load.
maximum rudder displacement is controlled by the summing
stops and is ÷ 3.6 degrees.
is
by
The
lever
D6 - 30643_-_RESTRICTED
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_L'JMM_
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,
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71
There are two feedback paths
path directly feeds back yaw
feedback is required so that
follow com_a_i inputs.
The
actuator position via a
in the yaw damper.
The first
damper actuator position.
This
yaw damper servo output will
second path feeds back yaw damper
8
in parallel with the
normal
position feedback.
Whe_ the yaw damper system is
disengaged,
the servo amplifier autput is fed back through
this lag to provide synchronization.
.I$
.
Self-Test
and Confidence
°
Test
The yaw damper computer provides self-test circuitry for the
two categories of fault isolation testingLine Replaceable
Unit test (electronic components)
and system test (electronic
components plus the actuator loop).
The self-tests are performed by positioning
the test switch and momentarily
depressing
the press-to-test
switch on the yaw damper front panel.
The
monitor light on the front panel will indicate the test results
(go or no go).
Confidence Test is a pilot actuated and monitored system test.
Channel confidence tests can be initiated singularly or simultaneously by use of the cockpit mounted test switches.
By
observing the upper and lower rudder surfaces position indicator located on the pilot's instrument panel, the operating
status of the systems can be assessed.
The confidence test and self-test
test initiation in flight.
are interlocked
to prevent
Separate confidence tests are provided for the dutch
damping and the turn coordinator functions.
@
roll
Performance
Analog computer simulation and digital computer root locus
analysis were conducted to investigate yaw damper system
performance.
Results indicate that the system performs well
at all flight conditions.
The lowest augmented airplane
dutch roll damping ratio is 0.30 at the landing approach
33 degrees flaps down condition.
System performance
for
various flight conditions are summarized in root loci plots
in Figure 33.
.
Tolerance
Phase shifts a_ gain variations affect system performance.
Root locus analysis were conducted to investigate
the effect
of tolerances on performance.
The Bode plot and phase angle
plot for the nominal and the maximum time constants of the rudder to sensors input transfer function are shown in Figures 31
and 32 respectively.
The two flight conditions, which describe
the lower and upper bound of the unaugmenteddutch
roll frequency
D6 - _N&A_.
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-3o643
->72
6-7000
ware investigated.. They are the landing approach 33 degrees
flaps down and maximum dynamic pressure conditions where the
dutch roll frequencies are 0.12 cps and 0.20 cps respectively.
The yaw damper system performance
for the above two conditions
is summarized in Figures 34 and 35.
If the system components
are within the design tolerances, the system performances
as described by the envelope in the figures is satisfactory.
The upper and lower bounds of the envelope are the gain tolerance limits while the left and right bounds are the phase
angle tolerance limits.
i0.
Failure
Mode Design
$
o
Features
At flaps down condition, three sensed signals are used in the
yaw damper.
The yaw rate, roll angles, and derived roll rate
signals each provide an increment of dutch roll damping.
If
failures occur in any of these signal chains such that one or
more of the signals is not driving the rudder, the remaining
signals will continue to provide some dutch roll damping.
All
three of these signals are washed out and thus cannot command
a continuous steady-state rudder position.
At flaps up condition, roll angle and derived roll rate
signals are removed by the easy on-off circuit and the delayed
flaps energized switch.
Monitoring
circuitry is implemented
to detect failure of this flap actuated switching.
A red
light is illuminated on the yaw damper control panel if flap
switching failure occurs.
Any type of malfunction causing a no-signal condition or a
sustained large signal will not cause more than a modest
transient airplane response.
Control excursions are positively
limited by hydraulic system velocity and force limits, plus
displacement
stops effective at all speeds.
Q
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IV.
AUTOTHROTTLE
A.
SYSTEM
GENERAL
The 7h7 is equipped with a single-channel
autothrottle
system.
This
system is designed to capture and hold a selected indicated airspeed
during terminal area maneuvering,
and approach and landing flight
regimes by automatically positioning
the throttles.
It may be used
for indicated airspeeds up to 400 kts.
Figures 36 and 37 pictorially
describe the system.
Salient features of the autothrottle
are:
_Clutches in the mechanical drive to the throttles
enable the pilot to override the action of the system
at any time.
I.
o
.
.
e
Be
The system is operable
or while the Autopilot
during manual
is engaged.
flight
control
The system limits the rate of change of commanded
airspeed so that throttle motions occur smoothly.
Thrust changes following pitch maneuvering
are minimized, and the magnitude of transient thrust changes
arising from wind gusts are limited byagust
filter.
A throttle retardfunction,
interlocked with the
Autopilot Land Node, is provided to automatically
retard the throttles during an autopilot flare.
SYSTEM DESCRIPTION
A block diagram of the autothrottle
system is shown
more detailed description
is provided below.
i.
Co_Inand and Airspeed
Error
inFigure
38.
A
Signals
The system engage switch and the speed
on the AFCS Mode Select Panel.
control
knob are situated
The Autothrottle computer receives an airspeed error control
signal from the Captain's airspeed indicator.
Whenever a
difference
exists between the pilot-selected
airspeed command,
indicated in digital form on the AFCSMode
Select Panel, and the
actual captain's airspeed, throttle action to reduce" the error
is commanded.
The servo-motor
is geared to the throttle
levers via clutches which allow the pilot to easily override
the system.
I)6 - 30643 - RESTRICTED
USE- SeeNoticeon Cover
k_
REV SYM
B_EJN_
i
D6-306_43
No. Vol. II
83
/
REV SYM
==_,.,,= .o. _#.o_3
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PA_E
84
->6-70oo
0
0
D6 - 30643
REV
SYM
- RESTRICTEDUSE - See Notice on Cover
Bd_EIN_I_
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ac
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-
.
Accelerometert
Attitudet
and Elevator
The longitudinal accelerometer,
which
the computer, provides rate of change
Signals
is an integral part of
of speed information.
The pitch attitude input, obtained from the Inertial Navigation
System, cancels the attitude component of the accelerometer
output.
IS
.
o
Computation
The computer utilizes transformer coupling for the A.C. input
signals.
A differential amplifier is used for the D.C. radio
altimeter signal to provide isolation and noise rejection.
All A.C. signals are demodulated
prior to shaping, filtering,
and error level detection.
The processed D.C. signals are
summed and modulated for use in the A.C. power amplifier which
drives the servo-motor.
Prior to engagement, the system is
synchronized for inputs other than airspeed error.
See block
diagram Figure 38.
The basic airspeed error input signal is processed through an
acceleration limiting circuit.
This circuit assymmetrically
limits the airspeed commandrate.
The rate limited command
signal is passed through an asymmetric gain program which reduces
the gain for overspeed errors larger than 2 kt to .25 of its
value.
This gain reduction for large overspeed errors compensates
for the fast deceleration
of the engine and the higher
authority of the throttle in the aft direction.
The IAE signal
is summed with the compensated
longitudinal
accelerometer
signal and passed through the gust filter to suppress the effects
of air turbulence on system activity.
.2
h@
Throttle
Control
and Limits
Control of the throttle levers is accomplished
by a proportional
plus integral servo configuration.
The servo proportional
response isobtained
by integrating the tachometer output.
Switches are provided to limit forward and aft throttle motion.
The forward limit position is set to avoid exceedin_ the maximum
allowable engine pressure ratio or temperature;
the aft position
closely corresponds to the flight idle thrust value.
Whenever
I
D6 - 30643 - RESTRICTEDUSE - See Noticeon Cover
{,
REV SYM
18
,
D6-306_3
i' .......
87
the throttles are driven to either limit position, the
integrator is put into a "hold" condition to prevent it
from computing an erroneous throttle position.
The system
remains engaged and will drive the throttles out of the
limit position when the appropriate
signal is developed.
5.
ADI Sisnal
The demodulated
output from the airspeed indicator, filtered
by a B-second lag, is supplied to fast-slow indicators on the
two Attitude Director Indicators in the cockpit.
6.
Airspeed
Error
, o
Warnin6
When the Autothrottle
is engaged, an airspeed error greater
than ten knots causes the amber Flight Mode Annunicator
light
to illuminate.
Disenga6e
We
A pilot can disengage
following:
.
the system by means
(a)
The engage
switch
(b)
Disconnect
switches
(c)
Go-around
switches
on the AFCS
mode
on throttle
on throttle
of any of the
selector
levers
levers
panel.
i and _.
2 and 3.
Flare
In conjunction with the Land Mode of the Autopilot,
the system
provides autou_tic throttle retard during the flare maneuver.
The conditions necessary to activate this function are that
the Autopilot must have been armed for flare, the Autopilot
flare must have commenced, and the radio altimeter signal
must be below the trigger altitude.
The logic requirements
prevent inadvertent operation of the retard function.
e
Test
Incorporated
in the computer is an automatic test arrangement
which can isolate a failure to the computer or.servo-motor
without requiring the use of supplementary
ground-test
equipment.
This test is accomplished by means of a rotary switch and a pushbutton switch located on the front panel of the computer.
The
rotary switch is used to select the unit to be tested (i.e., the
computer or the servo-motor) and the push-button
initiates
the test.
The rotary switch will remain in the selected
r_
P
D6 - 30643
(3
4C
REV SYM
- RESTRICTEDUSE - See Notice on Cover
Bd_'EINd_
|_
N0
PAGE
VOI.
II
.D6-306h3
.->-
88
6-7OOO
position until it is manually returned to the "OFF" position.
The Autothrottle
system cannot be engaged, and a steady red
warning light shows on the front panel (and in the cockpit)
when the switch is not in the "OFF" position.
The test results are indicated by lights located on the computer's
front panel.
A "test-in-progress"
light is illuminated
on
initiation of a test; and a "go" light is illuminated upon
successful completion.
A failure is indicated by the "test-in _"
progress" light being extinguished without a "go" indication.
The equipment functionally tests the entire Autothrottle
system
in approximately
one minute.
The testing is performed by
injecting test signals into the various computer inputs while
monitoring both the input and output of the feedback integrator
in the servo loop.
If each of these monitoring
devices yields
the proper indication, then the test step is completed and the
next step is performed.
If each test step is successful, the
program will proceed to the last step and yield a "go" indication is withheld.
C3
D6 - 30643 - RESTRICTED
USE - See Notice on Cover
C3
qC
°o
REV SYM
Bd_'_='J, AV_
IB
Vol. II
i _o.
pA_,E_-3o6h3
89
_b-6-700O
19.0
APPE_[DIX E - REVISED
S]]_ULATION DATA
The data in this section
recommends
incorporating
contains
revisions
that Boeing
in the NASA simulation.
D6-3o6_3
|
AD'_='_FAb'_
REV LTR:
E-3033
R I
I SECT
IN O. Vol.
I
PAGE
II
I?.O-
I
SUMMARY
OF
AREAS
ITEM
Wing
AND
DIMENSIONS
VALUE
Area
DIMENSION
Ft .2
(S)
Wing Mean Aerodynamic
Chord (MAC)
Wing
Wheel
Span
195.68
(b)
Ft.
Base
Wing
Body
Wheel
Ft.
Gear
Gear
78.96
88.96
Ft.
Ft.
36.16
12.5
Ft.
P_.
39.6
Ft.
Tread
Wing
Body
Gear
Gear
Effective
Engine
Moment
Arms
Inboard
•
YEz
Za z .
{'q_
(air)
(ground
Ft.
Ft.
)
Outboard
Yzo
ZE o
Note
The
transition
engine
where
and
REV SYM
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pitching
between
the
arm,s, ZEz
ground
and
ZZo,
and
air
is a
values
function
for
the
effective
of
the
averaged
CURNUTT
FOSTF..R
ODEGARD
TD
4_1 C-R4
GROUND
EFFECT
THE
HEI6HT
BOEING
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COMPANY
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LIFT
COEFFICIENT
GROUND
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EFFECT
APR
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BOEING
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Xg,o_Lt
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APPENDIX
F -AIRPLANE
Time histories
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CONDITION
PAGE
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