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KC-10 Aircrew Training System 

Flying the KC-10

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LIST OF EFFECTIVE PAGES
KC-10 Aircrew Training System 

Flying the KC-10 

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W001-01 i 01 JUN 06 

TABLE OF CONTENTS 

1.0 INTRODUCTION .................................................................................................. 1-1 

2.0 AERODYNAMICS, DESIGN, AND FLIGHT CHARACTERISTICS ....................... 2-1 

3.0 MANEUVERS ....................................................................................................... 3-1 

4.0 ABNORMAL APPROACH - LANDING ................................................................. 4-1 

5.0 CONSERVATION ................................................................................................ 5-1 

APPENDIX

Definition of Symbols .................................................................................................... A-1 

Glossary of Terms ........................................................................................................ B-1 
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1.0 INTRODUCTION

This workbook is compiled from available design documents, the supplemental 

information workbook, Douglas flight test documentation, and aircrew inputs throughout 

the life of the KC-10. 

The information is presented as "information only" and is provided to those who want to 

know the "why" as well as the "what". 

For the Flight Engineer, the performance manual suggests that an "aft CG" improves 

fuel economy. The Aerodynamics Section of this workbook explains "Why", while the 

Conservation Section provides some proven ideas for accomplishing engine and fuel 

conservation. 

For the Pilot, in the discussion of longitudinal stability, you will see the effect of CG 

movement on aircraft performance. While an aft CG is best for fuel conservation, a 

forward CG provides a more stable platform for maintaining contact during receiver air 

refueling. This workbook tells you why. This knowledge could be the difference 

between getting your fuel onloads in turbulence and aborting the mission. 

For the Boom Operator, the cargo vs. fuel loading places restrictions on the amount of 

cargo that can be loaded with various fuel loads. This workbook explains why. 

Most of this information is available in other publications. However, it is not available 

within one cover, and very few crewmembers have access to that information at a 

moment’s notice. With that in mind, this workbook was assembled. 

The information contained within this workbook is accurate at time of publication; 

however, any conflict that may arise between data in this workbook and the flight 

manual must be resolved in favor of the flight manual or other governing directives. 

Section Two presents an expanded look at the basic aerodynamics of the KC-10 and 

explains in simple English why it performs as it does. 

Section Three covers practice maneuvers that are accomplished in the simulator and 

not in the aircraft. It amplifies the established procedures for accomplishing these 

maneuvers. 

Section Four presents accepted techniques for accomplishing a single engine approach 

and landing. Also included are methods for dealing with dual hydraulic failures, 

landings with partial flap and slat configurations, landing with the stabilizer inoperative 

and flight operations with a severely damaged radome and/or suspect airspeed 

indications. 
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Section Five completes the workbook and covers conservation of the two most 

expensive areas of KC-10 operation: Engines and Fuel. 

The Appendix contains the definitions of common symbols used in aerodynamics and 

used throughout the flight manual. A glossary of terms is also provided. 

Fly Safe!!!!! 
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2.0 AERODYNAMICS, DESIGN, AND FLIGHT CHARACTERISTICS 

2.1 INTRODUCTION

The purpose of this section is to provide you with an understanding of the flight 

characteristics of the KC-10A. It is practically impossible to discuss flight characteristics 

without questions arising as to why the airplane does what it does or won't do. These 

questions most likely concern the aerodynamics involved and can only be answered 

through an understanding of aerodynamic principles. 

At the risk of "preaching to the choir," this section contains some basic aerodynamic 

information and definitions related to the design and, therefore, the flying characteristics 

of the KC-10A. It is included as reference material only in case a question arises 

regarding some aerodynamic concept and you just can't remember the principles 

involved. 
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2.2 AERODYNAMIC PERFORMANCE 

2.2.1 GENERAL

The KC-10A is a big airplane. How Big? Well, in terms of dimensions Figure 2-1 will 

give you an idea, but dimensions are only one aspect of "bigness". The KC-10A 

operates at heavy weights, up to 590,000 pounds, and has a wide operational weight 

range. During a mission, the weight may vary between around 260,000 pounds and 

590,000 pounds. A good "round number" weight is 350,000 pounds. The allowable 

Center of Gravity (CG) ranges from 8% Mean Aerodynamic Chord (MAC) to 33%. 

Large weight and CG variations may take place rather suddenly in the case of a heavy 

airborne fuel onload or offload. 

In conjunction with the ranges in weight during a mission are the associated changes in 

1G angle of attack and stall speeds. For instance, during a mission, the 1G-stall speed 

may vary from 178 knots (590,000 pounds, flaps and slats retracted) to 90 knots 

(260,000 pounds, landing configuration). Nearly a 100-knot change in stall speed 

during a single sortie! Granted, that is the extreme case; however, 70 knots is a fairly 

typical stall speed range for a given mission. 

Another aspect of "bigness" is inertia. It takes a greater force, applied sooner, to 

change a KC-10A from its projected flight path to the required flight path than in a lighter 

aircraft for any given speed. What this means to the Pilot is that maneuvers must be 

anticipated. What it means to the aircraft manufacturer is that the displacement of 

control surfaces sufficient to accomplish the maneuver is beyond the physical capability 

of the Pilot, and some type of augmentation is required. The KC-10A controls are 

covered later in the text. 

Big airplanes have many motors. Big airplanes have big motors. The KC-10A has 

three, and each is capable of generating 52,500 pounds of thrust. Among other things, 

what this means to the Pilot is that if an engine fails, particularly a wing engine, 

asymmetrical forces will upset the stability of the aircraft. The Pilot must counter these 

forces in order to maintain control. Asymmetrical thrust is thoroughly discussed in a 

later segment. 

The size of the KC-10A impacts all aspects of aerodynamic design. The following 

section will hopefully provide an understanding and appreciation for the flight 

characteristics of this versatile aircraft. 
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VERTICAL CLEARANCE 

 NOMINAL CLEARANCE 

MAXIMUM RAMP WEIGHT

NOMINAL CENTER 

GRAVITY 

MINIMUM CLEARANCE 

CRITICAL WEIGHT AND 

CENTER GRAVITY 

MAXIMUM CLEARANCE 

CRITICAL WEIGHT 

AND CENTER GRAVITY 

 FT. IN. METERS FT. IN. METERS FT. IN. METERS 

A 27-2 8.28 27-1 8.25 28-1 8.56 

B 6-1 1.85 5-10 1.78 6-10 2.08 

C 57-7 17.55 57-2 17.42 58-7 17.86 

D 10-9 3.28 10-8 3.25 11-10 3.61 

E 9-8 2.95 9-7 2.92 10-6 3.20 

F 7-9 2.36 7-9 2.36 8-5 2.57 

G 2-11 0.89 2-10 0.86 3-7 1.09 

H 14-6 4.42 14-4 4.37 16-3 4.95 

I 23-10 7.26 23-5 7.14 24-10 7.57 

J 36-10 11.23 36-7 11.15 37-8 11.48 

Figure 2-1. Airplane Dimensions and Vertical Clearance Table. 

165 FEET 4 INCHES

71 FEET 2 INCHES

34 FEET

8 INCHES

G

D

F

A

H I

J

27 FEET

11 INCHES

72 FEET 8 INCHES

181 FEET 7 INCHES

B

C

E

40 FEET

10 INCHES

82 FEET 8 INCHES

62 FEET

3 INCHES
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Why big? Basically, for economy and flexibility. For example: In 1980 it took 32 KC135 tankers, a large percentage of which had to be pre-positioned overseas, and 2 EC135 command and control aircraft to move 12 F-4E's to Egypt. In 1982, eight KC-10's 

did the same job, all staged from the States, and they carried 175,000 pounds of cargo 

to boot. 

Some other impressive statistics: The KC-10A can deliver 200,000 pounds of fuel to a 

receiver 2,200 miles from home base, and return. It can carry 169,400 pounds of cargo 

4,370 air miles unrefueled. With its receiver air refueling capability, the flexibility of the 

airplane is virtually limitless. 

2.2.2 THE WING

2.2.2.1 THE WING COMPROMISE

Since the characteristics of the wing dictate a majority of the flying characteristics and 

performance envelope of any airplane, the foundation for understanding how to fly the 

KC-10A is understanding the wing. 

GIVEN: Design a wing for a large, heavy airplane that will facilitate economical high 

subsonic cruise Mach, greater than .8, and which will still provide a sufficient coefficient 

of lift at low speeds so that high gross weight landings and takeoffs may be 

accomplished from nominal distance runways. The wing should have low coefficients of 

drag for economy, no undesirable high subsonic Mach (up to .95) characteristics or 

dangerous pre-stall or post-stall tendencies. The wing must be structurally sound to 

withstand bending moments produced by a high gross weight aircraft and be capable of 

carrying a large amount of fuel within its structure. 

Now, there are specific wing design features for each one of these requirements. 

Unfortunately, designing a wing to maximize one attribute generally minimizes all the 

others. The wing that ultimately satisfies all the requirements is a compromise. 

2.2.2.2 THE ASPECT RATIO COMPROMISE

Very basically, aspect ratio is the ratio of the length of the wing to its width (see Figure 

2-2). For a given angle of attack, the greater the aspect ratio (the longer and thinner the 

wing), the less the drag generated by lift (induced drag). Why? During generation of lift 

by the wing, the flow of air from the high-pressure lower surface to the low-pressure 

upper surface occurs at the wing tips. 

The wing tip vortex created by this flow induces a transverse movement of air across 

the wing surfaces: inboard on the upper surface and outboard on the lower surface (see 

Figure 2-3). These diverging flows meet at the trailing edge of the wing and their 

interaction creates a swirling downward motion or downwash. The effect of this 

downwash is to create a geometric angle of attack greater than the lift producing angle 

of attack, and thereby creating additional drag. 
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Figure 2-2. Aspect Ratio Compromise

Length

Width

Calculated

at various

points.

The Aspect Ratio Compromise

Ratio of Length vs. Width
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Figure 2-3. Wing Tip Vortex.
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Think of the wingtip vortex as a pulling force acting on the lower wing surface airflow, 

and a pushing force acting on the upper surface airflow. The further from the wingtip, 

the weaker the force becomes; so the length of the wing becomes a player. Also, the 

narrower the wing, the shorter the distance the airflow across it has to travel, 

consequently reducing the time that the pull/push force is acting upon it. So, the longer 

and narrower the wing, the smaller the adverse effects in relation to the total area of the 

wing. 

Obviously, there are engineering problems associated with creating a structurally sound 

long thin wing on a heavy jet transport; such a wing is not generally compatible with 

high speeds. So, there must be an aspect ratio compromise. 

2.2.2.3 AIRFOIL COMPROMISE AND WING SWEEP

The subject of critical Mach number arises when discussing wing sweep and wing airfoil 

on jet aircraft. At some speed, the air passing over the wing experiences the first 

evidence of local sonic (Mach 1.0) flow. This speed is termed the Critical Mach 

Number. Any increase above the Critical Mach Number changes the sonic flow to a 

supersonic flow and creates a shock wave. 

The supersonic air changing to subsonic as it reaches the trailing edge of the wing 

causes the shock wave. Airflow separation induced by shock wave formation increases 

drag, decreases lift, and affects the pitching moment. 

2.2.2.3.1 HIGH SPEED FLIGHT

The maximum operating airspeed in the KC-10A is represented by the symbol VMO. 

VMO is 350 knots at sea level, increasing with altitude to a maximum of 368 knots at 

25,000 feet. This is a structural limit, and although the aircraft exhibits normal 

aerodynamic characteristics above this speed, any attempt to maneuver may overstress 

aircraft components. VMO is depicted as a barber pole symbol on the airspeed indicator, 

and when exceeded, an audible warning recognizes it. 

Critical Mach is also a limiting speed, but is a limit imposed by aerodynamic rather than 

by structural considerations. The Critical Mach in the KC-10A is .88 Mach and is given 

the symbol MMO. As speed reaches Critical Mach, the shock wave formation causes a 

recognizable buffet. The shock wave forms first at the wing root because of the 

exaggerated chord, and as airflow disturbance behind the shock wave causes loss of lift 

on the inboard portion of the wing, the nose will drop. This phenomena is known as 

"Mach Tuck." 

If speed is allowed to increase, the shock wave forms on the remainder of the wing and 

counters the nose down tendency. Above approximately .89 Mach, the roll due to yaw 

reverses and becomes negative as the shock wave and associated drag become 

stronger on the leading wing edge. 
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Lateral control at speeds above Critical Mach remain normal because of the effectiveness of the spoilers and the fact that the inboard aileron location is not susceptible to 

wing warp due to the rigidity of the wing at this point. The .95 Mach design diving speed 

of the KC-10A is a demonstrated speed and the aircraft is normally controllable at this 

speed. 

The most likely occurrence of a speed in excess of Critical Mach would be during 

recovery from an upset of some sort. The aircraft accelerates quite rapidly in a dive. 

The inherent dangers in exceeding the Critical Mach during a dive result from the 

possibility of overstressing components during the pullout. 

2.2.2.3.2 AIRFOIL COMPROMISE

The ideal high speed wing would be a flat plate since the airflow does not accelerate 

across a flat airfoil and the critical Mach would be Mach 1.0. Obviously, such a wing 

presents other problems. It would be structurally weak, and takeoff and landing speeds 

would be abnormally high due to the poor lift qualities. 

The thickness of an airfoil is dictated by the strength required, the low speed airfoil 

qualities, and the accommodation of the high lift devices that are necessary for 

reasonable takeoff and landing weights. The KC-10A airfoil is the "peaky" design where 

the distribution of lift induces a high value at the leading edge, but the velocity is made 

to fall very rapidly in order to delay the onset of supersonic flow to a higher indicated 

Mach number (see Figure 2-4). The peaky airfoil provides for a higher cruise Mach 

number by delaying the formation of the drag producing shock wave; however, there is 

a drag penalty when cruising at less than maximum range Mach. See Figure 2-5 which 

is an extract from the KC-10A Performance Manual. 
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Figure 2-4. Airfoil Compromise. 
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Figure 2-5. Range Penalty. 
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2.2.2.3.3 SWEEP COMPROMISE

The reason for sweeping a wing is to delay the compressibility induced shock wave to a 

higher indicated Mach number, i.e., increase the critical Mach speed. Sweeping the 

wing changes the airflow velocity vectors, and the wing behaves as though it is flying 

slower than it actually is. The speed of the free stream air ahead of the wing, 

represented by a vector perpendicular to the leading edge, determines the velocity of 

the air over the wing. The greater the sweep, the shorter the vector becomes for a 

given free stream velocity (see Figure 2-6). 

The KC-10A wing is swept 35°, and at this sweep angle the critical Mach number is 

increased about 12% and the peak drag Mach number is increased about 24% versus if 

the wing were not swept. 

Another advantage of the swept wing is that it increases the static directional stability 

(weather cocking) and lateral stability. This phenomena is covered in a later section. 

There are three primary disadvantages to the swept wing that are of interest to the Pilot: 

The low speed maximum coefficient of lift is reduced, about 19% for 35° of sweep. 

What this means is that the swept wing requires a higher angle of attack to produce the 

same lift. The tips of a swept wing tend to stall before the rest of the wing, and wing 

taper further aggravates this tendency. Taper ratio is the ratio of wing root chord to 

wingtip chord (see Figure 2-2). The KC-10A has a taper ratio of .26, mainly for 

structural reasons. Given a swept and tapered wing, the tips will stall first and pitch the 

airplane up because the center of lift then moves forward of the center of gravity. 

This rather undesirable condition must be countered by changing the angle of incidence 

of the wingtip in relation to the rest of the wing. The engineering warp required to 

accomplish the task is called "washout". The use of sweepback also reduces the 

effectiveness of the trailing edge control devices (ailerons) and lift devices (flaps). The 

lift effectiveness of a single slotted flap in a 35° swept wing is reduced by more than 

50% versus if the wing were not swept. 

The shape of the entire wing includes the changing or section airfoil, sweep, taper and 

aspect ratio, and is called the wing planform. The primary impact of the planform on 

flight characteristics is in the realm of high speed and low speed flight. The KC-10A 

wing has been designed to maximize the aerodynamic advantages of planform and 

minimize the disadvantages. 

Where aerodynamic penalties still exist, such as aileron ineffectiveness and low speed 

lift characteristics, additional devices have been incorporated into the wing to alleviate 

the deficiencies. These devices are spoilers, slats, and flaps. 
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Figure 2-6. Effects of Sweepback.
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2.2.2.3.3.1 SPOILERS

The spoilers on the KC-10A wing are used to augment lateral control, and as drag 

devices (see Figure 2-7). Spoiler operation is covered in another section. 

2.2.2.3.3.2 SLATS

The leading edge slats increase the low speed lift qualities of the wing by changing the 

effective chord. The leading edge slats on the KC-10A are divided into eight sections 

on each wing: two inboard of the wing engines and six outboard (see Figure 2-8). The 

slats are extended for takeoff and landing. The outboard slats extend to greater 

deflection angles than the inboard slats. There are two normal settings for the slats and 

a single handle in the cockpit controls all slats. The settings are: takeoff - inboard 15°, 

outboard 25.4°; landing - inboard 19°, outboard 30°. 

2.2.2.3.3.3 FLAPS

The flap system consists of inboard and outboard double-slotted flaps on each wing 

(see Figure 2-9). The movable vanes that form the slots are incorporated to provide 

maximum chord extension and thereby increase the lift over drag with the flaps 

extended. The flap handle in the cockpit controls flap position. The flaps are used as 

lift devices for takeoff - up to 25°, and lift/drag devices for landing - up to a maximum 

setting of 50°. 

2.2.2.4 WING FUEL

One last subject before leaving the wing discussion is the subject of wing fuel. The fuel 

tanks in the wing have a capacity of 146,800 pounds contained in three tanks located as 

shown in Figure 2-10. When the main body of the aircraft is heavy with cargo and/or 

fuel contained in the body tanks, the wing tanks must remain full and fuel should be 

burned from the body tanks until body weight is reduced. This restriction is due to the 

bending moment of the wing at the root. 

As an example: think of the wing as a floating bridge. Place a heavy weight at the 

center, and the bridge will bend at the center; the amount of bend is dependent upon 

the weight. Now if weight can be added over the span of the bridge, the angle of the 

bend at the center will be reduced. The reason then for retaining fuel in the wing, is to 

reduce the bending moment at the wing root and consequently increase the fatigue life 

and safety margin of the wing. 

What does it all mean to you, the pilot? Well, bear with the lesson a bit longer and it will 

all come together, in the segment on "Flight Characteristics." 
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Figure 2-7. Aileron and Spoilers.

AILERON LOCATION

SPOILER LOCATION

INBOARD

OUTBOARD

W0016-07
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Figure 2-8. Slats.
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Figure 2-9. Flaps.

FLAP-VANE RELATIONSHIP

FLAP LOCATION

W0016-09
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Figure 2-10. Fuel.
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2.3 AIRCRAFT CONTROL

2.3.1 LATERAL CONTROL

Conventional ailerons and wing spoilers provide lateral control of the KC-10A. The case 

for spoilers evolves from the fact that the KC-10A is a high-speed aircraft with a large 

wing. Aileron deflection of sufficient magnitude to give desired roll rates, twists the wing 

at high speeds. What happens is the downward movement of the aileron on the up 

going wing creates a twisting moment, which actually reduces the angle of attack of the 

wing. This reduction in angle of attack counteracts the rolling moment caused by the 

aileron deflection. At high speeds, this phenomena may actually cause a roll in the 

opposite direction and is termed "aileron reversal." 

Two other problems with ailerons are that large ailerons can cause adverse yaw due to 

drag, and ailerons lose their effectiveness at high speeds due to the effects of 

compressibility. 

The roll rate of the KC-10A is excellent at all speeds up to 40° per second. In order to 

achieve these roll rates in spite of aileron ineffectiveness several things must be done. 

First of all, there are two sets of ailerons. Both sets work at low speeds, but only one 

set works at high speeds. The high-speed aileron is smaller, and is placed inboard to 

take advantage of the stiffness of the wing at that point, and reduce the percentage of 

the total span subjected to the twisting moment. 

The outboard aileron works at low speeds to augment the reduced effectiveness of the 

inboard aileron at these slower speeds. 

The location of the spoilers is shown in Figure 2-7. The spoilers operate in conjunction 

with the ailerons at all speeds to increase roll rates. They open on the side of the 

upgoing aileron to spoil lift and drive the wing down. The spoilers are also used 

symmetrically as speedbrakes to produce drag when required and to destroy lift on the 

ground after touchdown or during an abort. 

The automatic operation of the speedbrakes on the ground will be covered during 

academic training. Spoiler operation on the KC-10A is of the differential type; meaning: 

when deployed by the pilot in the air to produce drag (speedbrake operation), they still 

operate in conjunction with the ailerons to increase the rolling moment. 

The speedbrake setting becomes the new reference point for spoiler movement, 

extending further on the downgoing wing and retracting on the upgoing wing in 

response to the roll command. 
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Don't let the terminology throw you. The actual panels are the same, but are referred to 

during flight as "speedbrakes" when used by the Pilot as drag devices, and "spoilers" 

when actuated in conjunction with the ailerons for roll. These same panels are actuated 

on the ground to destroy lift during landing or an aborted takeoff. In this role, they are 

termed "Ground Spoilers." 

2.3.2 LONGITUDINAL AND DIRECTIONAL CONTROL

The KC-10A empennage is comprised of the vertical stabilizer, rudder, horizontal 

stabilizer and elevator, which is nothing new in that arrangement. Of course, there is an 

engine stuck in the lower section of the vertical stabilizer, and there are some 

aerodynamic idiosyncrasies. 

The wide speed range, CG, and weight range of the KC-10A have already been 

discussed. These considerations all impact the design of the empennage. Another 

player in the design of the tail section, specifically the vertical stabilizer and rudder, is 

the possibility of an asymmetrical thrust situation. 

2.3.3 HORIZONTAL STABILIZER AND ELEVATOR

The KC-10A has a jackscrew actuated movable stabilizer which is controlled through 

pilot operated trim switches or the autopilot. The stabilizer is hinged at the aft spar and 

moves through a 16.5° arc as measured at the leading edge. The stabilizer is adjusted 

as necessary to balance the force required (tail down or tail up) as the center of lift 

changes with the angle of attack, changes in flap/slat configurations, or imbalances 

caused by CG movement (see Figure 2-11). 

If the horizontal stabilizer were fixed, the balancing forces would have to be provided by 

the elevator alone and a drag penalty would be imposed. In some extreme cases (such 

as CG at forward or aft limit,) full deflection of the elevator may be required to offset the 

pitching moment, and there would be a decided lack of additional elevator authority in 

that direction (see Figure 2-12). 

With a movable stabilizer, the fore and aft stick forces are trimmed to zero by moving 

the stabilizer, leaving the elevator free to operate from a neutral position in a 

conventional manner as commanded by the pilot. The limits of elevator movement are 

20 degrees up and 20° down. 

The KC-10A horizontal stabilizer is swept and tapered for the same reasons as wing 

sweep and taper. The sweep is 35°, and taper ratio is .37. 

2.3.4 VERTICAL STABILIZER AND RUDDER

The size of the vertical stabilizer and rudder is a compromise dictated by stability 

requirements. Stability is covered later. The rudder size and effectiveness is further 

impacted by the maximum turning moment possible during an asymmetrical thrust 

situation. 
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Figure 2-11. C.G. Change - Stabilizer Operative.

DOWNWARD

BALANCING

WEIGHT FORCE

FORWARD C.G.

LIFT

UPWARD

BALANCING

FORCE

WEIGHT

AFT C.G.

LIFT

AFT SPAR HINGE.
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Figure 2-12. Elevator - Stabilizer Inoperative. 

DOWNWARD

BALANCING

WEIGHT FORCE

FORWARD C.G.

LIFT

WEIGHT

AFT C.G.

LIFT

UPWARD

BALANCING

FORCE
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Figure 2-13. Vertical Stabilizer and Rudder.
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The KC-10A rudder is double-hinged to increase its effectiveness (see Figure 2-13). 

The effect of the double hinge arrangement is to modify the chord and camber of the 

rudder as it is displaced in order to make a more efficient airfoil. The rudder is subject 

to stall the same as any airfoil, and on the KC-10A or any aircraft subject to asymmetric 

thrust, a large rudder deflection (angle of attack) is required to offset the yawing 

moment caused by failure of a wing engine. The KC-10A vertical stabilizer has a 40° 

sweep and a taper ratio of .35. 

2.3.5 CONTROL SURFACE ACTUATION

All control surfaces on the KC-10A (see Figure 2-14) are hydraulically powered in order 

to provide constant control pressures throughout the operating envelope of the aircraft, 

and to provide the necessary control at high speeds where the required deflection of the 

control surfaces is beyond the physical capability of the Pilot. 

There are two dominant drawbacks associated with powered controls. One, the 

feedback or feel to the Pilot does not change with speed. It takes the same amount of 

stick force to fully displace the controls at maximum speed that it does at minimum 

speed. This results in overcontrol, and in extreme cases, overstress and failure. The 

second problem is, of course, that if the power fails, there is no control. 

Both of these problems have been engineered out of the KC-10A. In the first instance, 

artificial feel has been added to the elevator. This feel or force felt by the Pilot 

increases with speed. Stabilizer trim rates are automatically reduced as speed 

increases above 250 knots. One set of ailerons, the larger outboard ailerons, lock out 

mechanically, leaving only the smaller inboard ailerons at the higher speeds. 

The differential spoilers blow back as air load increases with speed, thereby reducing 

their effectiveness. The rudder has been left alone because the Pilot, as a matter of 

course, does not induce large rudder deflections. 

Now, if the controls rely exclusively on hydraulics, and they do in the KC-10A, what 

happens when there is a hydraulic failure? The redundancy features engineered into 

the KC-10A render the possibility of complete control loss statistically remote (however 

it has occurred once). Spring devices add artificial feel of Rudder air loads at the 

Rudder Pedals 

Without going into too much detail at this point, suffice to say that there are three selfcontained hydraulic systems. Each control has multiple actuators divided among the 

hydraulic systems so that even with complete loss of fluid in two systems, approximately 

1/3 of the control about each axis remains. The hydraulic systems are also tied 

together with a series of pumps that allow the hydraulic systems to pressurize each 

other as long as at least one system has at least one of its two engine driven pumps 

operating. There are backup systems available in the form of electrical pumps, and 

even a pump operated by electric power generated with ram air. Hydraulic systems and 

flight controls will be covered in detail in academics. 
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Figure 2-14. Flight Control Surfaces.

RUDDERS

ELEVATORS

ALL SPEED AILERON

STABILIZER

FULL SPAN SLATS

SPOILERS

LOW SPEED

AILERON

FLAPS

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2.4 AIRCRAFT STABILITY

2.4.1 DEFINITIONS

2.4.1.1 STATIC STABILITY 

The initial tendency to return to equilibrium following some disturbance from equilibrium. 

Static stability is either: positive, the tendency to return to equilibrium; negative, the 

tendency to continue in the direction of the disturbance; or neutral, neither the tendency 

to return to equilibrium nor continue in the direction of the disturbance. 

Static stability is only concerned with the TENDENCY to return to equilibrium and does 

not take into account the dynamic stability. 

2.4.1.2 DYNAMIC STABILITY 

The time history or duration/amplitude of any resulting motion following a disturbance 

from equilibrium. Dynamic stability is graphically depicted in Figure 2-15. 

Chart A illustrates the non-oscillatory mode of direct return to equilibrium, and Chart 

B depicts direct divergence by a non-cyclic increase of amplitude with time. Pure 

neutral stability is shown in Chart C. The original disturbance creates a displacement, 

which remains constant. 

The oscillatory modes shown are all statically stable since the oscillations do return to 

neutral. What happens after they return to neutral is determined by their dynamic 

stability. Chart D shows a damped oscillation, dynamically positive, which remains at 

neutral once it arrives there. The undamped oscillation, Chart E, continues to return to 

neutral and also continues to oscillate at the same amplitude and frequency. The 

negative dynamic oscillation depicted in Chart F continues to return to neutral, but each 

time it passes through neutral, energy is applied to the motion such that the amplitude 

continues to increase with time. 

The Pilot Induced Oscillation (PIO) or porpoise maneuver would be an example of 

negative dynamic stability. The pilot adds control force too late and the force is not 

applied until the oscillation passes through neutral; the "Porpoise does it on purpose." 

Too much stability, as well as too little, is quite undesirable. The greater the stability, 

the less the controllability because of the increased force required displacing or 

maneuvering the aircraft. On the other hand, the effect of reduced stability is 

exaggerated displacement following a control input and lack of tendency to return to 

equilibrium. 
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Figure 2-15. Dynamic Stability.
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2.4.2 LONGITUDINAL STABILITY 

An understanding of KC-10A static longitudinal stability is most important to the crew 

simply because the stability of the aircraft is dependent upon the Center of Gravity 

(CG), and the crew has control over the location of the CG. The KC-10A CG is easily 

shifted by the Flight Engineer by moving fuel from tank to tank, varying the fuel burn 

sequence, or by the Boom Operator during cargo loading and unloading operations. 

To understand the role CG plays in the stability of the KC-10A, it is necessary to discuss 

the contribution to stability of the wing and horizontal stabilizer. The result of any 

changes in lift on an airfoil takes place at the Aerodynamic Center (AC) of the airfoil 

(see Figure 2-16). The Center of Gravity (CG) is the pitching axis of the aircraft. It is 

apparent then that a change in lift acting on the AC will create a pitching moment about 

the CG. 

For any given change in lift, the magnitude of the pitching moment depends upon the 

distance of the CG from the AC. CG ahead, positive lift causes a pitch down; CG 

behind, pitch is up. The magnitude of the pitching moment for a given CG depends 

upon the change in lift, and direction of the moment is determined by the fore/aft 

relationship of the AC and CG. 

When the CG is forward of the AC, a disturbance which increases lift will create a nose 

down pitching moment that results in a decreased angle of attack. This reduces lift 

which tends to return the aircraft to equilibrium. If the CG is aft of the AC, any increase 

in lift will pitch the nose up, increasing angle of attack, increasing lift, pitching the nose 

further up results in a continuous divergence from equilibrium. 

The Aerodynamic Center (AC) of the KC-10A wing is located one quarter of the way 

back from the leading edge of the Mean Aerodynamic Chord (MAC); in other words, 

25% MAC. The AC does move about a bit during low and high speeds, but for practical 

purposes it is considered to be at 25% MAC. 

The allowable CG range of the KC-10A varies with weight (see Figure 2-17), but for 

purposes of discussion here, the limits are considered to be 8% MAC - forward limit, 

and 29% MAC - aft limit (see the wing depiction in Figure 2-18). 

This is interesting. If the AC is at 25% MAC and the CG is at the aft limit of 29% MAC, 

the KC-10A is longitudinally unstable! Well, not the whole KC-10A, just the wing - and 

thereby hangs a tale, er.. tail. 

The horizontal stabilizer is needed to counter the pitching moments created when the 

wing is longitudinally unstable due to aft CG, and to reinforce the pitching moments 

resulting from lift disturbances when the wing is longitudinally stable and the stabilizer is 

in trim. It is also there, along with the elevator, to balance forces acting on the wing 

during constant attitude flight (stabilizer in trim) and to give the Pilot the ability to change 

and sustain pitching moments required amending the flight path (elevator movement). 
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The contribution of the horizontal stabilizer to longitudinal stability is depicted graphically 

in Figure 2-18. 

Figure 2-16. Pitching Moment.
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As far as the flight characteristics of the KC-10A are concerned, here is what is 

important to understand: when the CG is forward of the AC, the horizontal stabilizer is 

trimmed to a negative angle of attack to create down force (negative lift) to counter the 

nose down pitching moment created by the wing. As the CG moves aft, less and less 

down force is required, until the lift on the horizontal stabilizer reaches zero when the 

CG and the AC converge at 25% MAC. As the CG moves aft of the AC, the horizontal 

stabilizer must create positive lift to counter the nose up pitching moment of the wing. 

Maximum longitudinal stability is achieved at the most forward CG limit where the wing 

is most stable and the contribution of the horizontal stabilizer is maximum. At this CG 

position, large elevator forces are required to change the aircraft deck angle because of 

opposing forces. For example: the Pilot pushes forward on the yoke - lift decreases on 

the wing creating a nose up pitching moment, and negative lift increases on the 

horizontal stabilizer which pulls the tail down. 

On the other hand, the further aft the CG, the less the wing and the horizontal stabilizer 

contribute to stability, and very small elevator corrections result in large pitch changes. 

The aircraft is very hard to keep in trim, and Pilot induced oscillations are common. So 

why not fly with the CG at the forward limit? Very simply, "drag." Large horizontal 

stabilizer downloads must be offset by increases in wing lift, and both downloads 

(negative lift) on the stabilizer and positive lift on the wing create additional drag. 

Another reason is that large elevator forces are required to change pitch in a forward 

CG (maximum stability) situation. In practice, the KC-10A CG is kept near the aft limit 

to reduce drag, and the Pilots accept the pitch sensitivity. During critical maneuvers 

such as receiver refueling during turbulence, it might be prudent to onload fuel in a 

manner to shift the CG forward for increased stability. It can always be shifted aft after 

the refueling is complete by transferring fuel. 
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Figure 2-17. Center of Gravity in % MAC. 

Forward limit 

For demonstrated controllability. 

Aircraft trim critical 
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Figure 2-18. Longitudinal Stability Vs. C.G.

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2.4.3 LATERAL AND DIRECTIONAL STABILITY

Lateral and directional stability are practically inseparable but are discussed separately; 

initially for simplicity, and then combined in the discussion of spiral and oscillatory 

stability. 

2.4.4 LATERAL STABILITY

Positive lateral stability is demonstrated when lateral displacement from "wings level" 

flight (about the longitudinal axis) produces sideslip, and sideslip produces a rolling 

moment which returns the aircraft to wings level flight. The two main contributors to 

positive lateral stability in the KC-10A are wing dihedral and wing sweep (see Figure 2-

19). 

Generally, the aircraft handles better with relatively weak lateral stability. Strong rolling 

tendencies in a deliberate sideslip, such as during a crosswind takeoff or landing, tend 

to be a bit upsetting. Pun intended! 

2.4.5 DIRECTIONAL STABILITY

It is important when discussing directional stability to define the difference between yaw 

and sideslip. The salient difference is that during a sideslip, the vertical fin has a 

positive angle of attack equal to the sideslip angle. This creates a yawing moment 

which weathercocks the aircraft into the relative wind (reduces the vertical fin angle of 

attack and the sideslip to zero). So by way of definition, we can say that the sideslip 

angle is the angle between the relative wind and the longitudinal axis. 

The yaw angle relates the displacement of the longitudinal axis from some reference 

azimuth, and is basically a time history motion. An aircraft is considered to have static 

directional stability if a positive sideslip results in a positive yaw. The positive reference 

in all cases being clockwise about the respective axis (see Figure 2-20). 

It follows that a negative sideslip would result in a negative yaw in this representative 

statically stable aircraft. The vertical stabilizer is the main contributor to directional 

stability. Again, a compromise in the size of the vertical stabilizer is necessary. If it is 

too large, there is not only a drag penalty, but as we are about to learn, the aircraft will 

exhibit undesirable spiral divergence. 

2.4.6 SPIRAL STABILITY

The aircraft is considered to have spiral instability if a sideslip condition creates a 

divergence resulting ultimately in a spiral dive. What happens is that the sideslip, acting 

on the large vertical stabilizer, rapidly weathercocks (yaws) the aircraft into the relative 

wind. The lift on the outer wing increases due to yaw creating another sideslip, creating 

another yaw, and so on. The inner wing continues to drop, vertical lift is lost, and a 

spiral dive is the result. 
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Figure 2-19. Contribution of Components to Lateral Stability.
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Figure 2-20. Airplane Reference Axis.
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The KC-10A is spirally unstable over much of the takeoff envelope. This spiral 

instability does not create any problems for the Pilot since turns during takeoff are not 

normally accomplished, and even if a turn is initiated; the spiral divergence is quite slow 

and easily controllable. The Pilot should be aware that after initiating a turn in the 

takeoff configuration, the bank would tend to increase if the controls are released.

2.4.7 OSCILLATORY STABILITY

Oscillatory motion is a yawing/rolling oscillation more commonly known as "Dutch Roll". 

A yaw disturbance causes the aircraft to roll (dihedral and sweep). The vertical fin 

opposes the yaw since the yaw disturbance is moving the longitudinal axis away from 

the relative wind. If the fin and rudder are large enough, the motion is damped out 

(oscillatory stability). If the vertical fin and rudder are too small, the yaw and roll get 

progressively larger (oscillatory instability). 

2.4.8 RELATIONSHIP OF OSCILLATORY AND SPIRAL STABILITY

Since wing dihedral and sweep basically determine the lateral stability, and the vertical 

fin and rudder dictate the directional stability, it should be apparent that the dominance 

of the former decreases oscillatory stability, and dominance of the latter decreases 

spiral stability. Since an increase in oscillatory stability decreases spiral stability, and 

vice versa, a wing design and rudder design compromise is required to prevent 

dominant stability in one regime and instability in the other. 

2.4.9 HIGH ALTITUDE FLIGHT

The aerodynamic penalties incurred with flight at high altitude are accepted as a small 

price to pay when compared to the benefits of speed and economy. The penalties are: 

reduced aerodynamic damping, reduced stability, reduced maneuverability, and reduced speed range. 

The reduction in aerodynamic damping is a direct result of the decrease in air density. 

The smaller the number of air molecules that exist to be pushed out of the way following 

a disturbance from equilibrium, the smaller the opposing or damping force. 

Reduced stability results from the increase in the velocity of the aircraft at high altitude 

relative to a given equivalent airspeed. For any equivalent airspeed, the dynamic 

pressure or airflow acting on the airfoil is the same, but the airfoil angle of incidence 

created following a disturbance is less at high altitude because of the increase in 

velocity vector. 

Recall that the angle of incidence of the airfoil in disturbance is part of the equilibrium 

restoring force. The predominant loser is oscillatory stability, which deteriorates quite 

rapidly with altitude (see "Dutch Roll" Recovery - Section 3.3). As a matter of fact, spiral 

stability, which always opposes oscillatory stability, actually increases with altitude. 
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The decline of maneuverability at altitude is due to the general deterioration in damping 

and stability. The aircraft feels less stable, and it is. It will diverge further for a given 

disturbing force and will take longer to settle following a disturbance. 

The KC-10A is subject to the same deterioration in oscillatory stability, maneuverability, 

and reduction in speed range at high altitude as any swept wing transport. The least 

stable flight condition, neutral oscillatory stability, occurs at the upper altitude limit of the 

KC-10A, about 42,000 feet. 

The only technique to remember in the case of reduced maneuverability at high altitude 

is to be gentle with the aircraft. Make control movements slowly and gently, and 

recover from divergences in the same way. Also important is the fact that either one of 

the two autopilots will quite effectively maneuver the aircraft at all altitudes. At least one 

yaw damper must be operational before an autopilot can be engaged. 

The speed range becomes more restricted at altitude because the stall speed increases 

and the maximum equivalent airspeed decreases due to Mach effect. Since a wing 

always stalls at the same incidence, and the wing is only interested in EAS, the stall will 

occur at a fairly constant EAS. 

The stall speed increases with altitude for two reasons. First, EAS for a given IAS 

decreases with altitude. This difference can be as much as 16 knots in the KC-10A. 

Second, the Mach number effect on the wing can actually increase the EAS where the 

wing will stall, and this difference can amount to another 10 knots or so. On the highspeed side, it has already been explained how the Mach number increases at a given 

EAS as altitude increases. 

The matter of the reduced speed margin opens the door to a discussion of speed 

stability at altitude, which further leads to the subject of why the airplane is operated at 

high altitude in the first place. The reason is quite simple: efficiency! There is a charted 

altitude and speed for a given gross weight where air miles traveled per pound of fuel 

(specific range) is maximum. These charts are contained in the Performance Manual 

and are used during flight planning. 

The altitude is called OPTIMUM ALTITUDE and the speed is called MAXIMUM RANGE 

MACH. An increase in specific range with altitude can be attributed to three factors: 

One, an increase in altitude provides a greater true airspeed for the same amount of 

thrust. Two, an increase in altitude (below the tropopause) produces lower engine inlet 

temperature that reduces specific fuel consumption. Three, an increase in altitude 

requires increased engine RPM for the same thrust, and the CF6 engine is most fuel 

efficient at about 96% RPM N1. 

Any altitude higher than optimum requires the engine to operate at an RPM higher than 

its most efficient RPM to maintain the desired Mach. Speed stability becomes a factor 

because at a given altitude, the desired Mach for fuel economy is slower than 

achievable Mach numbers because of drag penalties associated with higher Mach. The 
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aircraft is more speed stable at higher Mach numbers because it takes more thrust to 

cruise at the higher speed. What this means, of course, is that at higher speeds if a 

sudden gust slows the airplane down, the thrust available will exceed thrust required by 

a greater amount than would exist if a gust of equal magnitude affected the airplane in a 

lower thrust/speed configuration. That greater amount of thrust, therefore, makes it 

easier for the airplane to regain its pre-gust cruise speed; thus it is more speed stable. 

Reduced speed stability at altitude is accepted for economic reasons. It means that 

speed excursions must be countered by changes in thrust. 

2.5 FLIGHT CHARACTERISTICS

The KC-10A does not possess any aerodynamic "bad habits". There are, however, 

characteristics unique to the KC-10A that are encountered during some flight regimes. 

You, as the pilot, should understand these characteristics. 

2.5.1 STALL

Stalls and stall recovery procedures will be briefed and practiced during simulator 

training. A documented occurrence of a DC-10 stall may give you an idea why stall 

recognition and proper recovery procedures are important. The stall occurred when the 

airspeed was allowed to bleed off during climb. The aircraft lost 4000 feet before the 

crew realized they were in a stall, and another 6000 feet during the recovery. 

The Flight Manual contains several terms related to stall speeds. They are presented 

here and defined: 

● MINIMUM STALL SPEED - This is the FAA recognized KC-10A speed. It 

is a speed demonstrated by reducing the airspeed at a rate of one knot 

per second until the stall occurs. The symbol for this speed is Vs and is 

the basis for charted maneuvering speeds for the KC-10A. 

● 1G STALL SPEED - A theoretical speed at which the KC-10A will stall in 

1G flight at a constant airspeed. It is included in the Flight Manual for 

information only. This is the stall speed that takeoff performance is based 

on. 

● CRUISE BUFFET SPEEDS - Flight test speeds where high speed and low 

speed buffet first occurs as boundary layer separation begins. 

An understanding of the KC-10A prestall, stall, and post-stall characteristics can 

perhaps best be gained by describing them as they occur during the maneuver. 

Notification of the approach to a stall is provided by a perceptible buffet as airflow 

separation begins. When the flaps are in the landing configuration, there is a flap buffet, 

which will mask the prestall buffet. 
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Below 15,000′, approach to stall is usually annunciated first by stick shaker. Between 

15000’ and 25,000’ stick shaker may be preceded by low speed buffet. Above 25,000′, 

stall can occur prior to stick shaker with buffet serving as the only stall warning. 

If the slats are up, the stall warning system will automatically extend all the slats 

outboard of the engine nacelles at this time (Auto-Slat Extension,) but only if the speed 

is below maximum slat extension speed. 

Extension of the slats improves the lift on the outboard portion of the wing for better 

lateral control. Auto-slat airspeed limits are 270 knots IAS, or .55 Mach because the 

outboard portions extend to 13.5° or approximately 50% of the takeoff position. Stall 

speeds in excess of these limits only occur at high altitude or during a grossly 

accelerated (high G) situation. If the proper recovery procedure is not initiated at stick 

shaker speed, the buffet will increase in intensity as the full stall approaches. 

If any yaw is present during the approach to the stall, and there probably will be some, 

one wing will lose lift before the other and will drop. This roll is easily controlled if 

outboard slats are extended, and rather hard to control if they are not. 

The stall will be accompanied by a nose down pitch as the inboard section of the wing 

stalls. The only deterrent to a "hands off" recovery at this point would be if the pilot 

were holding backpressure on the yoke, or if the stabilizer were trimmed to an extreme 

nose up position. 

Recovery from stalls in the takeoff and landing configuration can be accomplished with 

very little altitude loss. On the other hand, a clean wing stall at high altitude requires a 

substantial altitude loss to get the wing flying again. This is due to the limited thrust 

available, the inability to use slats due to airspeed limitations, and the relationship of 

deck angle to relative wind. The latter probably requires a short explanation. 

If a stall is allowed to occur, the aircraft begins to sink, which in effect increases the 

angle of attack without an increase in deck angle (see Figure 2-21). This phenomena 

occurs during any stall, but results in greater altitude loss during clean wing stalls at 

altitude. It also demonstrates the need for positive recovery to minimize altitude loss. 

Let's recap with some representative figures for a clean wing stall: Initial buffet occurs 

10 to 12 knots above automatic stall warning, stick shaker and auto-slat extension 

(speed permitting) occur at about 20% above stall, and at onset of stall, the nose will 

pitch down. 

A bit of food for thought. The probability of approaching stall speed in the KC-10A is 

remote, but is most likely to occur as a result of inadvertent early slat retraction after 

takeoff or failure to adjust orbit speeds for gross weight. An orbit is the aircraft loitering 

at altitude as a tanker awaiting a receiver. It is similar to a holding pattern. 
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Figure 2-21. The Lack of Relationship Between Attitude and Incidence.

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2.5.2 APPROACH AND LANDING

As the KC-10A is configured for approach and landing, the L/D max changes from 17 in 

the clean configuration to less than half of that in the approach configuration (see 

Appendix B, Glossary of Terms, L/D max). Large increases in thrust and deck angle 

are required as the configuration changes and speed is reduced. 

The speed stability decreases with the increase in drag and decrease in speed, and is 

approximately neutral at flare speed with a constant throttle setting. These aerodynamic characteristics are not unexpected, given the low speed curves of a swept wing 

and the wing chord changes that alter the lift characteristics of the wing as flaps and 

slats are extended. 

The Pilot should expect exaggerated deck angles during maneuvering and approach. 

For instance: to maintain altitude at maneuvering speed with maneuvering flaps and 

slats (flaps 22°, slats in takeoff range,) a deck angle of 7° is required. To maintain a 3° 

glideslope with landing flaps and slats (flaps 35°, slats in landing range,) expect a deck 

angle of 5°. Again, don't be confused by the term "deck angle". It's the pitch angle, 

graphically depicted on the attitude indicator as the relationship between the horizon 

and the aircraft symbol. Because of the reduction in speed stability, any reduction in 

speed from desired should be countered by a prompt increase in thrust. 

2.6 ASYMMETRIC THRUST

2.6.1 GENERAL

The possibility of an engine failure must be considered in the design of a multi-engine 

aircraft. The loss of thrust not only affects the overall performance, but in the case of a 

wing engine failure, the resultant asymmetrical thrust presents a directional control 

problem due to the induced yaw. The magnitude of the yawing moment resulting from 

an asymmetrical thrust situation depends upon the thrust of the remaining wing engine 

and its distance from the yaw axis of the aircraft. 

The wing engines of the KC-10A are located quite close in, which is good. However, 

the amount of thrust the engine can produce is substantial: 52,500 pounds. A graphic 

presentation of an asymmetrical thrust moment is shown in Figure 2-22. 
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Figure 2-22. Requirements for Directional Control.
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The yaw and resultant roll due to yaw during the asymmetrical thrust situation is 

countered by using the rudder to regain directional control, and ailerons as required to 

maintain wings level. Since the effectiveness (lift or side force) of any given rudder is 

dependent upon equivalent airspeed, it follows that there exists some speed below 

which there will be insufficient rudder force available to counteract the yaw. 

This speed is called Flight Minimum Control Speed, and is given the symbol "VMCA". 

There is a related speed, which is used in takeoff planning called Ground Minimum 

Control Speed, and its symbol is "VMCG". VMCG takes into account the effectiveness of 

the nose wheel steering to counteract the yaw should a wing engine fail during takeoff 

roll. 

By definition, VMCA is the minimum control speed in the air to maintain directional 

control, with one wing engine inoperative, full thrust on the good wing engine, full 

rudder, and a maximum bank angle of 5 degrees with the failed engine on the high 

wing. If the airspeed deteriorates to VMCA, power must be reduced on the good engine, 

or a descent initiated, to increase airspeed in order to prevent an out of control situation. 

VMCG is the minimum speed on the ground at which a wing engine can fail and 

directional control be maintained with rudder and nose-wheel steering without exceeding a lateral deviation of more than 25 feet. Use of VMCG in takeoff planning will be 

covered in academic training. 

Due to the design of the KC-10A, VMCA is not a factor unless one of the rudder 

segments (upper or lower) is inoperative. This is because VMCA is below the clean wing 

stall speed of the aircraft and theoretically cannot be attained. If the upper or lower 

rudder is inoperative, then the approach speed will be limited by VMCA. The appropriate 

VMCA speeds are stated in the three amplified checklist procedures for "Dual Hydraulic 

System Failure" in Section II of the KC-10A Flight Manual. 

The Pilot reaction to a wing engine failure requires immediate application of rudder to 

counteract the yaw, and ailerons as required to achieve 5° of bank or less. The rudder 

application must be timely, but the yaw is not so violent that there is insufficient time to 

think about which rudder pedal to push. 

The yaw will displace the slip/skid ball in the direction of the good engine, which is also 

the direction in which to apply rudder. 

The KC-10A has plenty of rudder available and it isn't necessary to slam the rudder in. 

Once the yaw has started, give the rudder a judicious push; it will become apparent 

immediately if the wrong rudder correction is being applied. Apply sufficient rudder to 

counter the yaw, then trim the pedal pressure out with the rudder trim. 
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The aircraft is properly trimmed when the slip/skid ball is centered and the control wheel 

is neutral. Insufficient rudder, or too much rudder, requires use of the ailerons to 

maintain directional control resulting in a sideslip as evidenced by the displaced slip/skid 

ball. This is not dangerous, but it does result in increased drag due to the deflected 

ailerons and spoilers. 

The KC-10A has adequate thrust at all weights with one engine inoperative. The Pilot 

must keep in mind that during flight with asymmetric thrust, any change in airspeed or 

power will change the yaw moment and the rudder will need to be re-trimmed. A failure 

of the tail engine (#2) will have a slight affect on the pitch axis of the airplane. The 

thrust vector of this engine tends to push the nose down and, as a result of an engine 

failure, the nose will rise slightly but can easily be controlled by the Pilot. 

There is one flight performance consideration associated with engine failure that should 

be covered here. If the engine fails at three-engine optimum altitude, the loss of thrust 

available will require a descent to or below two engine optimum. Refer to KC-10A "Two 

Engine Operation" in Section IIA of the KC-10A Flight Manual and the Aircraft's 

Performance Manual. 

An engine failure after passing takeoff decision speed is the most critical of all engine 

failure situations. Takeoff decision speed, known as V1, will be covered in academic 

training. The pilot is faced with reduced performance during a critical phase of flight, 

and if it is a wing engine, an asymmetric thrust situation. It is most important to be 

MENTALLY PREPARED for engine failure during takeoff. The simulator mission will 

adequately prepare you to cope with the "engine failure after V1" procedure. 
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2.6.2 SUMMARY

1. Asymmetric thrust in flight causes: 

a. Yaw. 

 b. Roll due to yaw. 

 c. A spiral dive due to roll and resultant reduction of vertical lift 

 component. 

2. Proper corrective action requires: 

 a. Rudder to counteract the yaw. 

 b. Center the control wheel. 

 c. Trim out the pedal pressures. 

 d. Do not reduce airspeed below VMCA (VMCA not limiting with both 

 rudders operative). 

The autopilot does a fine job of flying the aircraft with a failed engine but only if the 

aircraft is trimmed properly. The autopilot is not to be used for engine out go-around. 
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3.0 MANEUVERS

Due to the excellent fidelity inherent in the KC-10 simulators, many maneuvers that are 

performed inflight in other aircraft are restricted to the simulator in the KC-10 program. 

Although steep turns may be accomplished inflight with direct IP supervision, that 

maneuver is also discussed here for convenience. 

3.1 ACCELERATION/DECELERATION

Acceleration/deceleration maneuvers will provide you with knowledge of required 

control forces, power requirements and ADI presentations that produce the desired 

paths. This maneuver will be completed with the FGS OFF and is not to be memorized, 

but should be reviewed prior to the mission. The instructor will call for each item as it is 

to be accomplished: 

• Entry level will be 12,000 feet, speed 350 KIAS. Airspeed bugs set for landing. 

• Set throttles to idle, establish 750 fpm descent, and decrease speed to 250 KIAS. 

• When reaching 250 KIAS, extend slats. With throttles still in idle, continue 

descent to 7000 feet. 

• When reaching 7000 feet, extend speed brakes (note tendency to pitch up) and 

maintain 250 KIAS. With throttles still in idle, continue descent to 5000 feet. 

• Level at 5000 feet, decrease speed to 210 KIAS, retract speed brakes, and then 

extend 22° flaps. Continue speed decrease to minimum maneuver speed and 

make 180° turn to the right. 

• When rolled out, descend 1000 feet at minimum maneuver speed and level off. 

Select GA on the thrust computer. 

• Extend gear. Turn left 180°, descend at 500 fpm to 3000 feet with 22° flap 

maneuver speed. 

• When rolled out, set flaps to 35°/EXT, establish 700 fpm descent at approach 

speed until reaching 1000 feet. At 1000 feet call "Go-around power, flaps 22°," 

Set power and adjust pitch attitude to maintain salmon bug airspeed setting. 

(35/EXT VTH approximates V2 plus 10) (Use 15° pitch as a maximum.) Retract 

gear at positive rate of climb. 

• At 3000 feet, lower the nose approximately 5°, retract flaps at 22°/EXT VMM

(approximates VFR), continue acceleration and retract slats at 0°/EXT VMM

(approximates VSR). 
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• Select CL on the thrust computer, reduce power to CL, and continue acceleration 

to 250 KIAS while maintaining the adjusted pitch angle. 

• Maintain 250 KIAS as a pitch reference until 10,000 feet, then establish 1000 fpm 

climb and accelerate to 330 KIAS. Level at 14,000 feet. 

• When leveled off, move throttles to idle, speed brakes full, and decelerate to 250 

KIAS. Retract speed brakes and use power as required to maintain 250 KIAS 

and level flight. 

3.2 STEEP TURNS

Steep turns are practiced to augment the Pilot's feel of the airplane and to assist in 

instrument crosscheck development. The maneuver is practiced in level flight with 45°

of bank with 275 KIAS recommended. The recommended parameters for this maneuver are; a maximum of 450,000 lbs gross weight, AR orbit speed or 250 KIAS which 

ever is greater, an altitude of between 10,000 feet AGL and FL 250 in VMC conditions, 

and required altitude block of 2,000 feet. Minimum speed is 180 KIAS or cruise buffet 

onset (clean configuration at 55E bank) whichever is greater. 

NOTE 

At 450,000 pounds, GW FL 200 cruise buffet onset boundary 

speed (55-degree bank) is 277 KIAS.

Prior to beginning the steep turn, stabilize all parameters (power, speed, heading, and 

altitude). When stabilized, start a bank in the desired direction toward 45°. On entering 

the turn, there is a tendency for the airplane to nose down but this is easily 

compensated for with elevator force. As the turn continues, a slight increase in power 

(about 500 pph fuel flow) will be required to hold speed. Use the ADI as reference, 

crosschecking with the altimeter and the vertical speed indicator. Pitch attitude varies 

from 2 1/2° ANU in level flight to approximately 4 1/2° to 5° in a stabilized 45° bank. 

With an INS ADI, precession error is nonexistent. 

Start the rollout within 15° of the intended heading. Power will have to be reduced if 

added during the turn. 

Steep turn target parameters are airspeed ± 5 knots, bank angle ± 5°, and altitude ± 100 

feet. 

NOTE 

At the 250 knot point on the airspeed indicator, the dial 

divisions change from 2 knot to 10 knot increments.
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COMMON ERRORS--STEEP TURNS

• Entering turn with improper roll rate. 

• Rapid roll rate causing altitude loss on entry. 

• Slow roll rate causing altitude gain on entry. 

• Erratic use of control column to maintain altitude. 

• Slow instrument scan, fixation on one instrument. 

• Slowness to notice need for or make required thrust changes to control airspeed. 

• Not maintaining constant bank angle during the exercise. Bank angle should not 

be permitted to gradually decrease without taking corrective action to reinstate 

the 45° bank angle. 

3.3 DUTCH ROLL

Dutch roll is a common characteristic of swept wing aircraft. In this aircraft however, 

Dutch roll will only be experienced when all four yaw dampers are inoperative or turned 

off. 

There are several ways of recovering from Dutch roll. The aileron technique is 

described here. This method positively ensures that the proper control inputs will be 

applied to stop the rolling motion. If Dutch roll is encountered and the yaw dampers are 

turned off, rather than inoperative, simply turn on the damper(s). The KC-10 will 

recover from the roll in approximately one cycle. If dutch roll is encountered without the 

availability of yaw dampers, use the following technique: 

Freeze the rudder pedals and give one firm correction on the aileron control against the 

up-coming wing. Do not hold the correction too long. Apply and remove force. This will 

immediately stop the rolling tendency and normal flight may resume. 

COMMON ERRORS--DUTCH ROLL

• Attempting to correct the condition by applying control forces too hastily. 

• Holding corrective input too long--thus prolonging the dutch roll. 

• Using rudder and aggravating the yaw. 

3.4 HIGH SINK

The purpose of this maneuver is to demonstrate the ability of the KC-10 to recover from 

an undesirable high sink rate condition during landing approach when the proper 

recover procedure is followed. The following configuration will be used: 

Flaps 50°

Gear Down 

Maintain Threshold speed 

Throttles idle 
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When the above configuration is established, trim the airplane to maintain airspeed with 

power in idle. The airplane will assume a high sink rate. After this rate is established 

(approximately 1000 ft. altitude loss), position all throttles to the approximate go-around 

N1, and rotate to the missed approach attitude necessary to maintain the threshold 

speed. Maintain landing configuration, then note the altitude loss at level flight and 

climb is possible (while maintaining speed). Due to excellent engine control response, 

the altitude loss (approximately 300-400 feet) will not be excessive. However, this loss 

could put the aircraft below the safe obstacle clearance height during an approach. 

NOTE

Increasing lift is the only way to control or stop an 

unacceptable rate of descent. Attempting a landing from a 

high rate of descent is extremely hazardous due to the 

airplane's large mass and inertia effect, and engine 

acceleration time.

3.5 STALLS

The following discussion includes indications and warning of an impending aircraft stall, 

along with proper recovery procedures, during all phases of flight. Stalls will not be 

practiced in the aircraft; however, a section on practice approach to stall is included 

since this is a mandatory simulator-training item for all pilot courses. A table listing 

normal control column forces during various maneuvers is also included in this section. 

3.5.1 STALL CHARACTERISTICS

In the high lift configuration with slats extended, stall characteristics are excellent. 

Notification of the approach to stall is provided by light buffet which starts approximately 

10-12 knots above stall warning speed and by the control column stick shaker in turn. 

At low flap settings, the stick shaker will be closely followed by a perceptible buffet, 

which increases in intensity. At high flap settings, the prestall buffet will be almost 

completely masked by flap buffet. Because the stall buffet is partially or completely 

masked by the flap buffet, the stick shaker provides a more definite warning. Stability 

and controllability are excellent as the stall is approached. If the speed is further 

reduced, some rolling tendency may be noticed. A nose down pitch that will increase 

with the deceleration will accompany this. Relaxing the pull force and allowing the pitch 

down attitude to accelerate the aircraft to normal flight accomplish recovery. 

When in the clean configuration and at or below the slat placard IAS/MACH, the same 

signal that actuates the stick shaker will simultaneously extend the outboard slats 13.5° 

(approximately halfway to the takeoff position) and illuminate the amber SLAT RESET 

light (the slat DISAGREE light will also come on). The action, which causes the 

outboard slats to extend, is referred to as Automatic Slat Extension (ASE). If the proper 

recovery procedure is not initiated at ASE, the buffet will increase in intensity as the full 

stall approaches. The ASE will prevent outboard wing stall and maintain essentially the 
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same stall characteristics as in the 0°/EXT configuration. Moving the slats to the 0°/EXT 

position provides 15° extension of the inboard slats and 25.4° extension of the outboard 

slats. The following table shows the reduction in stall speeds between the 0°/RET and 

0°/EXT flap/slat positions for the noted configuration: 

GROSS WEIGHT DELTA STALLS 

X1000 LBS SPEED (KIAS) 

 350-450 18-22 

NOTE

If the slats and flaps are extended after stall recovery, the 

subsequent clean up of flaps and slats will not retract the 

automatically deployed slats. Only the pushing of the SLAT 

RESET button will retract the ASE mechanism and 

extinguish the SLAT RESET light. This may be done while 

the flaps and slats are extended.

3.5.2 KC-10 STALL CONSIDERATIONS

There are many ways the KC-10 can be placed in a regime of flight, which may result in 

an inadvertent approach to a stall, and aircrews are frequently exposed to many 

potentially dangerous situations. Such situations include cell formation join-ups, air 

refueling activities and traffic pattern operations. 

An excessive closure rate during cell join-up or failure to monitor airspeed and 

configuration during air refueling orbit and rendezvous, particularly during turns, could 

result in an unobserved speed reduction, leading to an inadvertent approach to a stall. 

Likewise, failure to monitor airspeed and configuration while operating at high gross 

weights during orbit or slow speed air refueling, and dependency on the autothrottles 

during traffic pattern operations with autothrottles off could place the crew in situations 

conducive to an approach to a stall. 

Aircrew situational awareness cannot be over emphasized. Crew members must be 

aware of the possible flight regimes in which the aircraft will be operated that are 

conducive to an approach to a stall, recognize the deteriorating performance 

characteristics and take the appropriate action. Training will emphasize the conditions 

leading to the approach to a stall, recognition, and recovery procedures. 
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3.5.3 PRACTICE APPROACH TO STALLS ENTRY AND RECOVERY TECHNIQUE

The approach to the stall is a training maneuver and will be performed only in the 

simulator. The procedures will not be practiced in the aircraft. 

During training for the "real world", the approach to stall indications should be stressed. 

Emphasis should be placed on a rapid decrease of airspeed below bug setting, rapid 

decrease of climb rate during takeoff or go-around, or a rapid increase in sink rate 

during an approach. Proper actions will greatly improve aircraft performance during 

stall recoveries resulting from decreasing performance during takeoff into windshear, 

approach or go-around; microburst; or inadvertent initial stall buffet or stick shaker. 

NOTE 

There have been reported instances of early retraction of the 

slats during departures. The slats were inadvertently 

retracted with the flaps as flap retraction speeds were 

reached. Even with the airspeed increase during this 

configuration change, the clean configuration may be 

achieved as much as 60 to 70 knots below the 0°/RET 

minimum maneuvering speed. In all known departure cases, 

the stick shaker and automatic slat extension systems have 

operated properly to alert the crew of the situation and there 

have been no serious consequences; however, in such an 

event, the clean configuration approach to stall procedure 

should be executed.

Although entry procedures should not be emphasized during the approach to the stall, 

certain actions, which are discussed under the appropriate entry and recovery 

procedures, will aid you in establishing the entry parameters. Power settings used 

during the recoveries should be the maximum thrust for the existing conditions, i.e., 

Maximum Continuous Thrust above 14,000 feet MSL, and Go-Around power below 

14,000 feet. 

3.5.3.1 CLEAN CONFIGURATION 

Retard the throttles to idle and maintain heading and altitude. As the airspeed 

decreases, attempt to maintain altitude (do not descend) by trimming the stabilizer to 

relieve backpressure. 

When the airspeed has decreased to stick shaker value, the stick shaker will actuate 

and the auto slats will extend, providing the airspeed is not in excess of approximately 

MACH .55 or 270 KIAS. The amber SLAT RESET and slat DISAGREE lights will 

illuminate. Upon recognizing the approaching stall (stick shaker), apply maximum 

available thrust; adjust pitch as required to minimize altitude loss; extend the slats below 

MACH/IAS slat limit speed; and accelerate to the appropriate airspeed (minimum safe 
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speed for the existing configuration). As the slats extend, the slat DISAGREE light will 

go out and the blue slat TAKEOFF light will remain illuminated. 

Avoid unnecessary loss of altitude during the recovery. Accelerate to the slat retraction 

speed (approximates 0°/EXT VMM) and retract the slats by moving the flap/slat handle to 

the 0°/RET position and pressing the SLAT RESET light. Regain any altitude lost by 

climbing at the slat retract speed or its practical equivalent, or the 0°/EXT minimum 

maneuvering speed, until reaching an appropriate altitude. Remember that at heavy 

gross weights and higher altitudes it is possible to exceed the slat MACH limit when 

accelerating to the slat retract speed. When at the desired altitude, continue to 

accelerate to at least the 0°/RET minimum maneuvering speed. Refer to Figure 3-1, 

Approach to Stall - Clean With ASE. 

NOTE 

If the stick shaker actuates and auto slat extension occurs 

due to inadvertent slat retraction at insufficient airspeed, 

thereby placing the airplane in a stalled condition, 

immediately extend the slats to the 0°/EXT position and 

recover as previously described. 

If the stick shaker actuates when the airplane is at high altitude and airspeed above the 

slat limiting speed of approximately MACH .55 or 270 KIAS (auto slats inhibited). Do 

not extend the slats during the recovery. Apply maximum available thrust, adjust the 

pitch as required to minimize altitude loss and accelerate to an appropriate airspeed 

(minimum safe speed for the existing configuration). 

During recoveries at altitudes where performance is limited, accept an altitude loss, 

exceeding 1,000 feet, while accelerating to the 0°/RET minimum maneuvering speed. 

Adjust the pitch attitude as necessary and return to the desired altitude while 

maintaining this airspeed. Refer to Figures 3-2, Approach to Stall - Clean Without ASE. 
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Figure 3-1. Approach to Stall – Clean with ASE.

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Figure 3-2. Approach to Stall - Clean without ASE.

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3.5.3.2 TAKEOFF, APPROACH, LANDING OR GO-AROUND CONFIGURATION 

ENTRY AND RECOVERY

Establish the appropriate configuration for the phase of flight; disengage the 

autothrottles; set the airspeed reference bugs; and select G/A on the thrust computer. 

Reduce airspeed and set power to approximately 60% N1 for takeoff or 80% N1 for 

approach, landing or go-around. Trim to relieve backpressure; maintain altitude until 

approaching the stall. 

When the airspeed decreases to shaker value, the stick shaker will actuate. Upon 

recognizing the approach to the stall, apply maximum available thrust (G/A) while 

simultaneously adjusting pitch to minimize the altitude loss or to provide obstacle 

clearance. Maintain the existing flap/slat and gear configuration and adjust pitch to 

accelerate to the salmon bug +5 (approximates V2), which is the minimum maneuvering 

speed for the existing configuration. After attaining the minimum maneuvering speed, 

adjust the configuration as desired. A safe recovery to the entry altitude may be 

performed at the salmon bug speed and G/A power; however, if ground contact is 

imminent, apply thrust up to the throttle mechanical stops. Refer to Figure 3-3, Approach To Stall - Takeoff Configuration, and Figure 3-4, Approach to Stall - Approach, 

Landing or Go-Around Configuration. 

In summary, an inadvertent approach to a stall in the KC-10, as in most other aircraft, 

can usually be attributed to one or more of the following factors: inattention; inadequate 

situational awareness, resulting in excessively low airspeed; incorrect configuration for 

the actual airspeed; or actual wing loading, which is a function of bank angle. The KC10 gross weight range, mission flexibility, and aircrew workload collectively challenge 

each member to "fly the jet" and avoid low speed, high wing load flight. 
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 Figure 3-3. Approach to Stall – Takeoff Configuration 

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Figure 3-4. Approach to Stall - Approach, Landing, or Go-Around Configuration.

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COMMON ERRORS--STALL RECOVERY

• Not applying maximum available thrust. 

• Failing to adjust pitch as required minimizing altitude loss or providing obstacle 

clearance. 

• Abruptly raising the nose during recovery, reactivating the stall warning by 

increased G loading. 

• Failing to adjust configuration after accelerating to the minimum maneuvering 

speed for the existing configuration. 

• Not coordinating the rollout and decreasing pitch during the recovery from 

approaches to stalls in turns. 

• Using rudder instead of ailerons to control lateral stability (rudder should not be 

used excessively - this could result in abrupt roll off). 
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3.6 CONTROL COLUMN FORCES

The following table presents the control column forces you can expect when performing 

the listed maneuvers. These forces are easily handled until a stabilizer trim change can 

be made. Note the force reduction with the tail (#2) engine inoperative. 

 CONTROL 

COLUMN FORCE 

CONFIGURATION THRUST PULL PUSH 

TAKEOFF · ROTATION 

· CLIMB AT V2+10 ALL ENGINE 

· CLIMB AT V2 WING ENG INOP

· CLIMB AT V2 TAIL ENG INOP 

· FLAP RETRACTION (15°/EXT0°/EXT) 

· SLAT RETRACTION 

T.O. 

T.O. 

T.O. 

T.O. 

T.O. 

T.O. 

10-35 

0-5 

10-15 

5-10 

20-35 

5-15 

CRUISE · FULL-SPEED BRAKE 

EXTENSION 

· POWER APPLICATION 

T.F.L.F.*

IDLE-T.O.

 2-10 

0-4 

APPROACH/ 

LANDING 

· FLAP EXTENSION (0°/RET22°/EXT) 

· LANDING GEAR EXTENSION 

· FLAP EXTENSION (22°/EXT50°/EXT) 

· POWER APPLICATION 

(50°/EXT) 

· FLARE 

IDLE 

T.F.L.F.*

IDLE 

IDLE-T.O.

10-40 

5-10 

20-30 

0-5 

5-15 

* Thrust For Level Flight 

Figure 3-5. Longitudinal Control During Trim Changes 

(Approximate Control Column Forces in Pounds). 
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4.0 ABNORMAL APPROACH - LANDING 

This section addresses malfunctions that have significant impact on the approach and 

landing phase. The most serious situations involve loss of two engines or two or more 

hydraulic systems. Flight control malfunctions and inaccurate airspeed indications are 

also included. 

4.1 SINGLE ENGINE OPERATION

One of the most critical situations you may encounter in the KC-10 is the loss of two 

engines, more commonly referred to as single engine operation. Safe recovery will 

require the utmost in pilot skill and crew coordination. The following information will 

guide you through to a successful conclusion. 

4.1.1 GENERAL 

The point at which a second engine fails dictates what the Pilot does and the urgency of 

immediate action. If an engine is shut down, the pilot should be prepared for the 

eventuality of the loss of another engine. Steps should be taken to prevent a "crunch" 

caused by crew not being prepared if a second engine fails. Following an engine 

shutdown, if time and circumstances permit, brief the single engine operation checklist. 

Specifically cover the crew actions if the second engine should fail during the approach 

phase; power settings, flap/slat configuration, speed, etc. Refer to Figure 5-16 in T.O. 

1C-10(K)A-1, Single-Engine VFR Approach and Landing. If you are on final approach, 

configured with gear down, flaps and slats at 35/EXT + 10 (plus wind additives) when a 

second engine fails, then refer to Figure 4-4 and Paragraph 4.1.5. 

4.1.2 PRIOR TO REACHING AIRFIELD 

Assuming that the engine losses occur prior to reaching the airport, the following 

procedures are recommended: (Assuming Driftdown is not required) 

• Initiate the single engine operation checklist unless on final approach, flaps and 

slats extended and gear down. 

• Dump fuel to the lowest practical weight. Consider the maximum weight for 

single engine go-around (Figure 4-1) and the maximum normal landing weight for 

center gear retracted (400,000 pounds). 

• Reduce electrical loads to a minimum. 

• Ensure that both hydraulic motor pumps and both auxiliary pumps are ON 

(electrical loads permitting). APU is available for pneumatics and electrics. 

However, range and go-around capability must be considered. 

• Maintain a clean configuration. 

• Maintain at least minimum maneuver speed for the clean configuration. 

• Select GA on the thrust computer (14,000 feet and below). 

• Carry inoperative engine throttles along to eliminate an unsafe gear indication 

due to throttle position. 
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• Below 10,000 feet or when entering the traffic pattern, shut off all bleed air 

demands except engine and wing anti-ice (when required). 

• Consider destination options. The closest runway may not be the best choice for 

landing a single engine aircraft if weather, approach availability, and conditions of 

the aircraft suggest choosing another location. 

• Use a downwind and base leg extended enough to provide a minimum of a 7-

mile final. Maintain a clean configuration and at least minimum maneuver speed 

for the configuration. 

Maintain 0°/RET minimum maneuver until established on straight-in approach, 1,500 

feet AGL minimum. Starting descent, select 0°/EXT and reduce to single engine 

approach speed. Single engine approach speed is a charted speed based upon 1.5VS

for 0°/EXT. At heavier Gross Weights the Pilot will have to aggressively manage his 

power to achieve the desired single engine approach speed by 1,000 feet above airfield 

elevation. Wind additives are NOT added to this speed. 

This no-flap approach will necessitate higher than normal airspeed and pitch attitude. 

Plan on an increased descent rate. 

One thousand feet above the airfield elevation is the commit point for a landing or goaround. When this point is reached, and the glidepath and speed are satisfactory, 

continue the approach for landing while extending the landing gear. 

NOTE

Do not extend the landing gear before 1000 feet above airfield 

elevation since the landing gear drag is approximately three times 

that of the slats, and there may be insufficient thrust to maintain 

the desired flight profile if this configuration is selected too early. 

If rudder trim is used, zero the trim and hold the required rudder force to compensate. If 

the trim is not removed the aircraft will yaw when the power is reduced to idle. Adjust 

thrust on the available engine to arrive at the threshold with a minimum airspeed of 

single engine threshold speed (plus wind additives). When the threshold is crossed, the 

throttle is slowly moved to idle while flaring slightly through touchdown. 

After nose gear is on runway, raise all three reverse levers.. Do not delay thrust on 

available engine because of inability to raise the other reverse levers beyond reverse 

idle. Spoilers will not deploy until reverse lever actuation. 

4.1.3 MISSED APPROACH 

The decision to go-around must be made at prior to 1000 feet AGL (see figure 4-2). If 

go around is required, accomplish the following: 

• Set Go-Around power. 
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• Retract the slats, continue descent and accelerate to 0°/RET minimum 

maneuvering speed on the glidepath (ILS glideslope or maintain approximately 

600-800 fpm if visual). 

• Only when the airplane is clean, and minimum maneuver speed is attained, 

should the climb back to altitude begin. 

A single engine missed approach should not be attempted if the gear is down, 

regardless of altitude. If a missed approach is unavoidable under these conditions, the 

gear must be up prior to 1000 feet AGL and any additional demands on the hydraulic 

system should be delayed until the gear lights are out; this includes slat retraction. 

A single engine missed approach will not be made under any of the following conditions: 

1. Below 1000 feet above airport elevation. 

2. Weight, pressure altitude, and temperature exceed charted capability. 

(Figure 4-1) 

3. Airspeed is below single engine approach speed. 

4. Gear is extended. 

4.1.4 HEIGHT LOSS DURING A SINGLE ENGINE MISSED APPROACH 

A computed height loss during a single-engine missed approach initiated at 1000 feet 

with the proper approach speed, a loss of approximately 500 feet will occur prior to 

climb. A missed approach at 1000 feet (with a speed below the recommended) will 

result in a greater altitude loss. In this case, the airplane could contact the ground prior 

to the runway. 

WARNING

The single engine missed approach performance is based 

on having the landing gear up and locked at the 1000 feet 

commit point. 
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SINGLE ENGINE LANDING DO’S AND DON’TS

• DON’T - Extend slats too early on final. 

• DON’T - Over boost engine. 

• DO - Adjust pattern for winds (the suggested 7 mile final is for no-wind 

conditions). 

• DON’T - Fail to disconnect ATS for deceleration to VTH. 

• DO - Slowly reduce thrust and establish a steady speed trend, recognizing the 

advantage of being a little fast at the threshold rather than slow. 

• DON’T - Allow airspeed to decay below VTH plus wind additives on final 

approach. 

• DON’T - Go below glideslope during deceleration to threshold speed. 

• DON’T - Bleed airspeed from the 1000 foot point, producing floating and a long 

landing, and a potential uneven rate of airspeed bleed. 

• DON’T - Try to hold off the airplane during landing. 

• DO - Get all 3 reverse levers to reverse idle for auto spoiler deployment. 

SINGLE ENGINE MISSED APPROACH CAPABILITY

NOTE 

•Ice Protection OFF. 

•Reduce weight by 1.6% with APU operating.

A 

L 

T

AMBIENT TEMPERATURE °C 

-54 -20 10 15 20 25 30 32 34 36 38 40 42 44 46 48 50

GROSS WEIGHT (X 1,000 LB) 

SL 458 461 464 464 465 465 461 451 442 435 425 418 410 401 395 387 380

2 441 443 446 446 447 447 428 420 411 405 397 390 382 374 369 

4 421 425 427 428 428 416 397 391 383 377 369 362 356 

6 402 404 407 407 402 387 370 365 357 350 343 

8 381 384 385 385 372 360 344 338 331 

10 360 362 365 358 346 333 318 

Figure 4-1. Maximum Recommended Gross Weight 

for Single Engine Missed Approach.
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For purposes of illustration: GW = 336,000 lbs. 

0/RET

VMM

VAPP VTH

SINGLE ENGINE MISSED APPROACH

ABOVE 1,000' AGL

APPLY GO-AROUND POWER

RETRACT SLATS

ACCELERATE TO V AT 700-800 FPM

DESCENT

MM

BEGIN CLIMB AT 0/RET V 

MAINTAIN

0

MM

VMM

0/RET

VMM

VAPP

VTH

W0013-03

BEGIN CLIMB AT 0°RET/VMM

MAINTAIN VMM

ABOVE 1,000’ AGL 

APPLY GO-AROUND POWER 

RETRACT SLATS 

ACCELERATE TO VMM AT 700-800 

FPM DESCENT 

VTH 

VTH 

0/RET

VMM VAPP 

VAPP 0/RET

VMM

Figure 4-2. Single Engine Missed Approach. 
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4.1.5 DURING FINAL APPROACH

If a second engine failure occurs during final approach, after the gear is extended, the 

focus of attention must be on completing the approach to a successful single engine 

landing. Since two engine approach procedures call for increasing the 22°/EXT 

Minimum Maneuvering and 35°/EXT Threshold speeds (VTH) by 10 knots until landing is 

assured, the aircraft is carrying an airspeed pad in the event of an additional engine 

failure. Approach speed should not exceed charted 35°/EXT VTH + 20 knots. This 

speed approximates 22°/EXT VTH. 

Upon failure of the second engine, immediately increase thrust on the remaining engine, 

as required, not to exceed go-around power and retract the flaps to 22°/EXT regardless 

of airspeed. Maintain the increased approach speed of 35°/EXT VTH + 10 plus normal 

wind additives that were being maintained on final prior to the second engine failure. 

Maintain this speed to the threshold. Adjust power as required to maintain the 

glideslope. 

Prompt pilot response to the loss of a second engine should result in no deviations from 

glideslope at gross weights up to the maximum normal landing weight with center gear 

extended (436,000 pounds). 

Common errors during the loss of a second engine on final are late detection of the loss 

and allowing heading and altitude to vary excessively due to surprise or concentration 

on the second loss. The arbitrary selection of go-around power at very low gross 

weights will make heading and glide path control difficult. 
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Figure 4-3. Loss of Second Engine on Final Approach.

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4.2 HYDRAULIC SYSTEM FAILURE

Loss of two or more hydraulic systems also makes safe recovery especially critical 

since all of the primary and secondary flight controls require hydraulic power to operate. 

Built-in safety features, discussed below, prevent loss of all flight controls in all but one 

instance. That condition, complete hydraulic power loss, is discussed after dual 

hydraulic failure. 

4.2.1 CONTROLLABILITY WITH DUAL HYDRAULIC FAILURE

A lot of thought went into the design of the KC-10 hydraulic system redundancy, 

especially when it comes to control of the airplane during a dual hydraulic system loss. 

The airplane can be safely flown with only one hydraulic system, but there is very little 

margin for error. Part of your responsibilities as an aircrew member is to become 

mentally prepared to handle any emergency. What follows are some ideas, which may 

be termed, "How to prepare for, and cope with, a dual hydraulic system failure in the 

KC-10." 

1. Become familiar with the general hydraulic systems arrangement as affects flight 

controls--see Figures 4-4 and 4-5. 

A. Any dual hydraulic system loss reduces lateral control about 66% and 

longitudinal/directional control about 50%. 

B. A 1-2 system loss renders flaps, upper rudder and autopilot stab 

inoperative (do not use autopilot or auto-throttles with dual hydraulic 

failure). 

C. A 1-3 system loss renders slats and alternate stab trim inoperative. Only 

manual gear lowering, and accumulator brake pressure is available. 

D. A 2-3-system loss renders the lower rudder and alternate stab trim 

inoperative. 

E. The checklist for a 1-3 system loss will have you turn off anti-skid to 

preserve what braking capability exists. 

2. Carefully follow the steps of the Flight Manual emergency procedure for dual 

hydraulic system failure. Knowledge of what systems are affected aids in 

understanding what the checklist steps are designed to accomplish. 

3. Protect the functioning system; reduce demand on the system while attempting 

to extend slats and/or flaps. Remember that the use of AR pumps for fuel dump 

is prohibited with a dual hydraulic system failure. 
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4. Determine VMCA; the checklist will tell you which rudder(s) may be affected. 

When in doubt, check the surface position indicator. 

5. Evaluate your situation; is the emergency compounded by a loss of thrust, what 

is the most likely flap/slat configuration for landing (the present configuration plus 

what you can expect to get). Review the landing profile (contained in Sect. 5 of 

the Flight Manual) with the crew. 
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Establish the final landing configuration prior to beginning the approach. Don't get in a 

“crunch” situation wherein you are still trying to extend flaps and/or slats and running the 

checklists as you cross the threshold.EFFECTS ON CONTROLLABILITY

SYSTEM SYSTEM 1 REMAINS SYSTEM 2 REMAINS SYSTEM 3 REMAINS 

AILERONS LEFT INBOARD AND 

RIGHT OUTBOARD 

INOP. LEFT 

OUTBOARD AVAILABLE 

AFTER FLAP OR SLAT 

EXTENSION OR AFTER 

MAIN GEAR 

ALTERNATE 

EXTENSION LEVER IS 

RAISED. 

RIGHT INBOARD 

INOPERATIVE. 

BOTH OUTBOARD 

AVAILABLE AFTER 

FLAP EXTENSION 

OR MAIN GEAR 

ALTERNATE 

EXTENSION LEVER 

IS RAISED. 

LEFT OUTBOARD 

INOPERATIVE. RIGHT 

OUTBOARD AVAILABLE 

AFTER SLATS OR FLAPS 

ARE EXTENDED, OR 

WHEN ALTERNATE 

LANDING GEAR LEVER 

IS RAISED. 

ELEVATORS LEFT INBOARD 

INOPERATIVE 

RIGHT INBOARD 

INOPERATIVE 

LEFT AND RIGHT 

OUTBOARD 

INOPERATIVE 

SPOILERS EFFECT REDUCED 

ONLY PANELS 2 AND 4 

LEFT AND RIGHT 

OPERATIVE. PILOTS 

CONTROL WHEEL MAY 

NOT BE CENTERED 

LATERALLY, DUE TO 

SPOILER PANEL 

FLOAT. 

EFFECT REDUCED. 

ONLY PANELS 1 

AND 5 LEFT AND 

RIGHT OPERATIVE. 

PILOTS CONTROL 

WHEEL MAY NOT BE 

CENTERED 

LATERALLY, DUE TO 

SPOILER PANEL 

FLOAT. 

EFFECT REDUCED 

ONLY PANEL 3 LEFT 

AND RIGHT OPERATIVE.

STAB TRIM REDUCED TO ONE 

HALF NORMAL RATE 

REDUCED TO ONE 

HALF NORMAL RATE

REDUCED TO ONE HALF 

NORMAL RATE. USE 

SPARINGLY. 

RUDDER UPPER RUDDER ONLY. UPPER RUDDER, 

LOWER RUDDER 

THROUGH RUDDER 

STANDBY POWER 

LOWER RUDDER 

THROUGH RUDDER 

STANDBY POWER 

VMCA 172 KIAS (INCREASES 

WITH DECREASING 

GROSS WEIGHT) 

NORMAL 157 KIAS 

MAX X-WIND 15 KNOTS (0 IF 

RUDDER IS INOP) 

NORMAL 15 KNOTS (0 IF RUDDER 

IS INOP) 

FLAPS WILL EXTEND WILL EXTEND INOPERATIVE, WILL 

REMAIN IN LAST 

POSITION 
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SLATS WILL EXTEND INOPERATIVE, WILL 

REMAIN IN LAST 

POSITION 

WILL EXTEND 

LANDING 

GEAR 

ALTERNATE 

EXTENSION, NO 

RETRACTION 

ALTERNATE 

EXTENSION, NO 

RETRACTION 

ALTERNATE EXTENSION 

RETRACTION POSSIBLE

NOSE WHEEL 

STEERING 

LEFT NORMAL RIGHT 

25 DEGREES 

INOPERATIVE RIGHT NORMAL, LEFT 

25 DEGREES 

BRAKES NORMAL, WITH ANTISKID SYSTEM 1, 

ACCUMULATORS 

SYSTEM 2 

MANUAL BRAKING 

ONLY THROUGH 

ACCUMULATORS. 

(SEE WARNING 

CONCERNING ANTISKID INOP) 

NORMAL, WITH ANTISKID SYSTEM 2, 

ACCUMULATORS 

SYSTEM 1 

PROBABLE 

LANDING 

CONFIG. 

35°/EXT 22° RET/MANUAL 

BRAKING. 

(OBSERVE 

WARNING - ANTISKID INOP) 

0°/EXT 

Figure 4-4. Effect on Controllability. 
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NOTE

Reduce gross weight as low as practical. This will reduce 

stopping distances and flaps/slats, if available, are more 

likely to extend at lower airspeeds. 

7. Increase lateral control during low altitude maneuvering by unlocking the 

outboard ailerons as soon as practical. This is especially important during a 1-3 

or 2-3 system failure where only one inboard aileron is operative. Outboard 

ailerons are unlocked by extending flaps or slats, or by raising the alternate gear 

extension lever. 

8. Set the bug speeds for the landing configuration. Compute the stopping distance 

and land at the nearest suitable airport. 

9. During approach with only a single hydraulic system operating and two engines 

inoperative, use the normal single engine operation landing procedure. 

In all cases, the following NOTES apply:

• Gear should be extended with alternate gear extension lever 

and the landing gear handle left in DOWN position. 

• Do not extend the flaps more than 35° to minimize effects of 

possible spoiler float. 

• Avoid using more than one hydraulic powered control at a time 

to ensure adequate hydraulic pressure. 

• Do not use either autopilot. 

• Have a plan for go-around or missed approach (configuration, turn 

capability, etc.) and brief it. 
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4.2.2 CONTROLLABILITY WITH ALL HYDRAULIC FAILURE

After the DC-10 accident at Sioux City, Iowa on 19 July 1989, the U.S. National 

Transportation Safety Board directed a simulator reenactment of the events leading up 

to the crash. Flight recorder data was used to replicate the accident aircraft dynamics 

resulting from the uncontained failure of the aircraft's number 2 engine and damage 

occasioned by the explosion, including the loss of fluid in all three hydraulic systems. 

The purpose of this simulator study was to determine if DC-10 flight crews could be 

taught to control the airplane and land safely with no hydraulic power available to 

actuate the flight controls. The result of this study showed that such a maneuver 

involved many unknown variables, and the degree of uncontrollability during the 

approach and landing rendered a simulator training exercise virtually impossible. 

However, the results of these simulator studies did provide some advice, which may be 

helpful to flight crews in the extremely unlikely event they are faced with a similar 

situation. 

Douglas offered enhancements for DC/KC-10 hydraulic systems, which preserve some 

longitudinal and lateral control even with catastrophic in-flight damage to all three 

hydraulic systems in the tail area such as occurred to the Sioux City accident aircraft. 

The enhancements consist of three separate installations -- an electrically operated 

shutoff valve in the supply line and a check valve in the return line of the number three 

hydraulic system, a sensor switch in the number three hydraulic reservoir, and the 

addition of an annunciator light in the cockpit to alert the crew if the shutoff valve has 

been activated. Hydraulic pressure in system no. 3 will be preserved forward of the 

valve to provide the crew with longitudinal control by stabilizer trim input at half rate; 

with lateral control through right inboard, right outboard and left inboard aileron input 

coupled with some spoiler deflection; with nose wheel steering; and with slats but no 

flaps. Simulator tests of the Sioux City accident configuration and flight tests of a DC-10 

powered only with hydraulic system no. 3 forward of the shutoff valve have enabled 

Douglas to develop procedures and suggested techniques for flying the aircraft with the 

enhancement. 

Section 1 discusses the techniques and suggestions developed for use with a complete 

hydraulic power loss such as occurred near Sioux City. Section 2 discusses the specific 

suggestions, which apply to flying the aircraft with the enhancement operative. 

4.2.2.1 AIRCRAFT CONTROL WITH COMPLETE HYDRAULIC POWER LOSS

Without hydraulic power, the ailerons will tend to float trailing edge up over a period of 

time, but there is little observable float on the elevators and spoilers, and none on the 

rudder. The only remaining means of control to the crew is from the operating wing 

engines. The application of asymmetric power to the wing engines changes the 

heading, hence roll angle, and increasing or decreasing power equally on both wing 

engines has a limited effect on the pitch attitude. It is important to note that the pilot has 

no direct control of airspeed because airspeed is primarily determined by pitch trim 

configuration. With this very limited means of control available, the following techniques 
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and suggestions are offered for the specific case of a number 2 engine failure and loss 

of all hydraulic power: 

4.2.2.1.1 REGAIN LEVEL FLIGHT

Immediately after the failure is apparent, attempt to keep or regain control of the aircraft 

by conventional control inputs for as long as the hydraulic power lasts. The autopilot 

and autothrottles should be disconnected and every effort made to level the wings and 

maintain pitch for level flight. 

4.2.2.1.2 ASSESS THE AIRCRAFT DAMAGE

Assess the aircraft damage by whatever means possible. Weather and other 

operational considerations permitting, such as day VFR conditions, consider calling for 

a chase aircraft to advise of the extent of damage. Visual examination from the inside 

of the aircraft may afford some degree of damage assessment. 

4.2.2.1.3 USE OF ASYMMETRIC THRUST TO MAINTAIN DIRECTIONAL 

CONTROL

The aircraft may have a tendency to roll in one direction due to damage in the tail area 

as in the Sioux City accident. Higher thrust on the right wing engine will roll the aircraft 

to the left and vice versa. To maintain wings level in the accident simulations the right 

throttle had to be maintained considerably ahead of the left. It was discovered that to 

minimize the rate of descent, the high throttle must be kept at or near Maximum 

Continuous Thrust (MCT) and the low throttle must be maintained as high as possible 

without causing undesired pitch or roll. 

4.2.2.1.4 PHUGOID MODE

With no pitch control, the aircraft will naturally tend to oscillate in the pitch axis. The 

nature of the pitch oscillations must be understood before discussing aircraft control. If 

the pitch attitude is upset with the aircraft stable in roll and yaw, the airspeed will 

change depending on the direction of the pitch change. As the aircraft pitch attitude 

increases (aircraft nose up) the airspeed will decrease, resulting in less lift on the wing 

and horizontal stabilizer. At the point where there is insufficient lift on the wing/stabilizer 

combination to maintain the pitch attitude, the nose begins to fall. As it does, the 

airspeed increases causing lift to increase, to the point where the pitch begins to 

increase again. This cycle repeats itself over a period of time. This long period pitch 

oscillation is called the “phugoid mode”, a characteristic inherent in all aircraft designs. 

The degree of “damping”, or the time it takes for the pitch oscillations to subside varies 

with aircraft design. On the DC-10, this pitch oscillation is eventually damped to low 

amplitude due to the inherent stability of the aircraft. 
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4.2.2.1.5 USE OF THRUST FOR PITCH CONTROL

Because the wing engines are mounted below the center of gravity of the aircraft, a net 

increase in thrust will tend to pitch the nose up and a decrease in thrust will pitch the 

nose down. Consequently, to control pitch oscillations with wing engine thrust changes, 

it is necessary to apply power just as the nose is coming down, and to retard power as 

soon as the nose is rising. (If asymmetric thrust is necessary to maintain wings level, 

power changes must be made while maintaining the power asymmetry). Judicious and 

aggressive use of power changes (i.e. jam accelerations followed by rapid throttle 

chops) is instrumental in helping to dampen the phugoid oscillations. However, care 

must be taken to avoid aggravating the situation by applying power out-of-phase 

thereby increasing the amplitude of the pitch oscillations. 

4.2.2.1.6 CHOICE OF LANDING SITE

As the aircraft returns to some degree of control using the above techniques, 

consideration should be given to a landing site. Many factors need to be taken into 

consideration, such as runway length and width, navigation aids, meteorological 

conditions, terrain, populated areas, crash-fire-rescue capabilities, and most important, 

the degree of aircraft control. Given some confidence in the structural integrity of the 

aircraft and the degree of aircraft control, consideration should be given to remaining 

airborne to reach a more desirable landing site versus a hasty landing in the immediate 

vicinity. If a distant landing site is selected, determine the aircraft's capability to 

maintain altitude enroute to the landing area. Use discretion before extending the 

landing gear because of the irreversible effect on increased airplane drag, hence 

decreased range. In addition, consider the fuel remaining and its distribution before 

dumping fuel to reduce the gross weight. 

4.2.2.1.7 FUEL TRANSFER

The fuel distribution may enable transfer of fuel both fore and aft for C.G. control and 

laterally for roll control. Lateral fuel transfer to level the wings is desirable because it 

allows both wing engines to be at the same throttle setting if asymmetric power was 

required to maintain constant heading. Matched throttles allow full concentration on 

pitch and facilitate heading changes simply by changing the left and right engine power 

simultaneously with one hand on both throttles. Longitudinal fuel transfer from the 

number 2 tank (or aux tanks) to tanks 1 and 3 moves the center of gravity aft, which 

effectively reduces the trim speed without changing the configuration. A reduction in 

trim speed is very important because it tends to reduce the landing speed thus 

enhancing stopping capability. 

4.2.2.1.8 ASSESSMENT OF AIRCRAFT CONTROL

Continually assess the controllability of the aircraft and remain alert for any further 

degradation of control. Attempt small turns and climb or descend to learn the aircraft 
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response to power inputs. Consider attempting a practice landing approach at a lower 

altitude by maneuvering to a predetermined heading at a specific altitude. 

4.2.2.1.9 LANDING GEAR EXTENSION

Gear extension will cause a nose up pitching tendency, thus upsetting the airplane in 

the pitch axis and changing the natural phugoid oscillation. Consequently, when the 

landing site has been determined, and before lowering the gear, allow sufficient time for 

the pitch recovery. Once the initial pitch upset is controlled, the aircraft will tend to be 

more stable due to the increased drag from the gear. Use the alternate gear extension 

lever to free-fall the nose and main gears for all DC-10 models. Give consideration to 

extending the center gear on aircraft so equipped. The extension of the center gear is 

optional when below the weight limit for landing with the center gear retracted. 

However, use of the center gear may provide some additional braking during the 

landing roll, and absorb some energy on impact. 

4.2.2.1.10 APPROACH

When preparing for the approach, note the capability to maintain altitude during the 

initial approach. To maintain as low a rate of descent as possible, plan for a long flat 

approach. The initial part of the approach should incorporate the use of all approach 

aids available at the landing site, as well as heavy reliance on visual cues. If the 

runway has no ILS, the VOR or ADF guidance may aid in alignment at the initial stage 

of the approach. Radar vectors may also be helpful at the early stage, particularly in 

cases of reduced visibility. Attempt to reach the vicinity of the outer marker (or 5 to 7 

miles from touchdown) at approach altitude with the gear extended, aligned with the 

runway, and with the pitch controlled as well as possible. Keep in mind that there is 

more thrust available at lower altitudes. Jam accelerations followed by rapid throttle 

chops may be necessary to control pitch without generating additional pitch oscillations 

as a sustained thrust input would do. If asymmetric thrust is required to keep the wings 

level, a rolling tendency will occur if both throttles are retarded to idle without 

maintaining the thrust asymmetry. Also be aware of any tendency for the engines to 

accelerate asymmetrically. This is more likely to occur when the throttles are moved 

from idle than from approach power settings. 

The final phase of the approach from the outer marker (or equivalent distance) to the 

runway should be accomplished mainly with visual cues, supported by instrument 

indications as a reference. If an ILS is available, the raw data may be used to 

determine the start of descent and to compare altitude to the glide slope. The approach 

descent profile should be slightly below the glide slope (approximately 1⁄2 dot) but no 

higher than 2 1⁄2 to 3 degrees. A higher glide path means higher descent close to the 

runway, which will be difficult to arrest. An approach slightly below the glide path should 

reduce the sink rate and will also require higher approach thrust, which places the 

engine power in a range to provide quicker response to power changes. 
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The use of the Flight Director (FD) commands while far out on the approach may be 

helpful to determine direction and relative magnitude of control (power) input, but 

insufficient control exists to fly the FD commands at lower altitudes. Consequently, it 

may be beneficial to bias the command bars from view to facilitate reference to the 

attitude data on the FD. 

4.2.2.1.11 GROUND EFFECT

Before discussing the landing, a review of the phenomenon known as “ground effect” 

will be useful. As an aircraft comes in close proximity to the ground, a slight increase in 

lift and decrease in drag occurs at an altitude beginning about 1⁄2 the wing span. 

Another tendency in ground effect, generally not as obvious, may be a nose down pitch 

moment as the aircraft enters ground effect, which is a function of sink rate and 

configuration. The DC-10 begins to enter ground effect at about 100 feet A.G.L. and the 

effect increases exponentially as altitude decreases. Without elevators, the only means 

of controlling the ground effect pitch change is with a sharp power advance followed by 

a throttle retard. This is a matter of judgment and is mentioned as a necessary step in 

reducing what could be an excessively high sink rate at touchdown. Obviously, an 

approach above the glideslope with a rapid sink rate coupled with ground effect could 

result in a very hard touch down. 

4.2.2.1.12 FINAL APPROACH

If the final approach is stable from about 500 feet A.G.L., a landing should be 

attempted, but extreme vigilance is required from all crewmembers. The pilot may find 

it useful for the pilot not flying to call radio altitude, sink rate, and sink rate increasing or 

decreasing trends from the outer marker (or equivalent distance) inbound. As the 

aircraft nears the runway, sink rate and sink rate trend are sufficient, and in ground 

effect sink rate trend only. The pilot not flying and the Flight Engineer are in a good 

position to judge pitching tendencies by watching the horizon in relation to the 

glareshield. This can be critical information when approaching touchdown. 

Because the ground effect may increase the rate of descent near the ground, it is 

apparent that the touch down aiming point must be moved to compensate for this 

tendency. Simulator trials have shown that if the aiming point is moved toward the far 

end of the runway, the touch down will have a better chance of occurring in the normal 

first third of the runway. 

In an unstable approach when the aircraft is in a phugoid oscillation, three things can 

happen: (a) the aircraft can touch down on the pitch down phase of the phugoid, which 

means a hard impact when coupled with a possible pitch down due to ground effect; (b) 

the aircraft can enter the pitch up phase of the phugoid during final approach and not 

touch down at all (in which case a go-around should be attempted; or, (c) it can touch 

down somewhere in between the two extremes. The likelihood of touching down 

smoothly is highly unlikely. Consequently, it is recommended to attempt to reduce the 

rate of descent before touch down as much as possible. 
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4.2.2.1.13 GO-AROUND CAPABILITY

If the final approach is not stable nearing touchdown, a go-around may be attempted by 

advancing the power and allowing the aircraft to pitch up. Remember that the power is 

not controlling speed, so only enough thrust should be used to initiate a climb at 

moderate pitch attitude. Too much pitch could cause the stick shaker to activate and/or 

the aircraft may stall. Keep in mind that adding power if a power differential was 

required to keep the wings level, will necessitate that the power differential be 

maintained to avoid the initiation of roll or a heading change during the go-around. 

Attempt to level off at a safe altitude and reinitiate the approach as before. If the 

approach is unstable and insufficient fuel exists for another attempted landing, or for 

other operational reasons a landing is necessary, it may be advisable to continue the 

approach and trade the original aiming point on the runway for a more stable and 

controlled touchdown in close proximity to the runway. In this sense it is possible to 

choose the landing site, or maintain a controlled stable descent and touch down, but it is 

extremely difficult to successfully accomplish both simultaneously. 

4.2.2.1.14 LANDING

When entering ground effect with the intention of landing, be aware of, and be quick to 

respond to, the necessity to add power to keep the nose from falling. Remember that 

the speed at touchdown will be a function of the phase of the phugoid oscillation, and 

could be well over 200 knots. It is most important to increase thrust, to raise the nose if 

necessary thereby decreasing the sink rate -- even if this results in an increase in touch 

down speed. Maneuver the power as necessary up to maximum thrust to reduce the 

sink rate to an acceptable value. 

If the final approach has been stable and the sink rate has been slowed, the aircraft will 

touch down at a pitch attitude approaching 10o. If the aircraft skips back into the air, 

attempt to ride out the skip or apply power for a go-around, depending on the degree 

and duration of the skip. If there is no skip or the aircraft returns to the runway 

immediately, apply full antiskid braking to forcefully bring the nose down. (Brake 

accumulators will allow full braking even with the loss of all hydraulic fluid.) The wing 

engines should be placed in maximum reverse immediately after nose wheel touch 

down while maintaining full braking. Manually modulate the brakes only as necessary 

for steering, but do not pump the brakes under any circumstances, as this will deplete 

the brake accumulators more rapidly. Steering will not be available from the normal 

means, but only as a last resort, asymmetric reverse thrust could be used. 

Adherence to these recommendations and techniques is by no means a guarantee of a 

successful landing. However, these suggestions represent the best information 

available from the studies and empirical simulator results into the nature of flight without 

hydraulic pressure powering the flight controls. 
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4.2.2.2 AIRCRAFT CONTROL WITH HYDRAULIC SYSTEM ENHANCEMENT

When the hydraulic shut off valve closes in response to a hydraulic quantity loss in 

system 3, hydraulic fluid is preserved for all controls using system 3 forward of, but not 

including the elevators. This provides longitudinal control by means of stabilizer trim 

with one trim motor (half rate) and lateral control by the inboard ailerons as well as 

aileron trim. The following will address the flyability of the DC-10 with the enhanced 

hydraulic system, and provide flight crews with the recommendations, techniques, and 

advice gleaned from simulator and actual flight test evaluations of the enhancement. 

The information presented regarding the flight characteristics of the DC-10 without 

hydraulic control, and the recommendations and techniques for flying and landing in this 

condition remains valid and is basically applicable to flying the aircraft with the 

enhanced hydraulic system, with the exception of the control that the enhancement 

provides in the pitch and roll axis. (If the ailerons were not trimmed and the stabilizer 

was not used to control pitch, the condition would be identical to the Sioux City aircraft.) 

It is this control that will be discussed below. 

4.2.2.2.1 REGAIN AIRCRAFT CONTROL

If damage similar to the Sioux City accident ever occurs again, attempt to gain/recover 

aircraft control using conventional aileron input for lateral control, and a combination of 

power and stabilizer trim for pitch control. It is recommended that the autopilot and 

autothrottles be disengaged and a wings level attitude held with aileron control. Aileron 

trim may now be used to trim out the aileron control wheel forces due to the rolling 

moment (if any) due to aircraft damage. It should be noted that without rudder control 

(i.e., no fluid in hydraulic system 2 for the 3-2 nonreversible motor pump) there will be a 

slight unbalanced side force that may be easily controlled by a very small bank angle. 

The relative ease of establishing roll control, and the use of symmetric throttles now 

allows full attention to pitch control. As in the case with no hydraulics, the throttles are 

used in the same manner to control pitch (that is an aggressive advance just as the 

nose is coming down, or a rapid retard just as the nose begins to rise). But now there is 

additional pitch control available through the horizontal stabilizer. Even at half rate 

(about 1/4o/second) sufficient trim exists to bring the aircraft under control with the use 

of a combination of thrust and pitch trim. The most important aspect of the stabilizer 

trim is the reduction in trim speed down to and including approach speed. Controlling 

the speed is the most significant factor in accomplishing a successful landing. 

4.2.2.2.2 STABILIZER TRIM

When trimming the stabilizer to make pitch changes, either the wheel trim switches or 

the suitcase handles may be used. The wheel trim switches must be held momentarily 

to assure stabilizer movement in either direction. When approaching the desired pitch 

attitude, it may be necessary to reverse the trim input to stop the motion at the required 

level. After the first few attempts at trimming, it becomes apparent that some 
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anticipation and input reversal is required. The suitcase handles may also be used to 

trim the stabilizer by placing both handles firmly to at least the first soft detent position 

(or beyond if necessary). However, this may be distracting because it requires the Pilot 

to alternate his right hand between the throttle and the suitcase handles. 

4.2.2.2.3 SLAT EXTENSION

The initial approach should be flown to the vicinity of the outer marker (or 5-7 miles out) 

in the same manner as previously discussed. During maneuvering in the terminal area, 

the slats may be extended to allow a reduction in airspeed. Slat extension may require 

airspeed to be reduced because the slats are now powered by a single hydraulic 

system. The airspeed should be reduced to the 0/RET Minimum Maneuvering speed 

and the slat handle placed to 0/EXT. If the slats do not extend in the normal time of 

about 10 seconds, it will be necessary to follow published procedures and show to the 

1.3VS speed for 0o/RET. Slowly decelerate to the 1.3VS speed and observe slat 

extension. There will be a slight nose down pitching moment as the slats extend which 

is easily controllable by stabilizer trim. The speed may be reduced to the 0o/EXT 

maneuvering speed once the slats are extended. 

4.2.2.2.4 LANDING GEAR EXTENSION

No later than the outer marker (or equivalent distance), the landing gear should be 

extended. Allow sufficient time to accomplish the procedure as outlined in the dual 

hydraulic failure – SYS 1 & 2 INOP AND SYS 3 ELEV OFF LIGHT ON checklist. The 

alternate gear extension lever should be raised, and the gear monitored for normal free 

fall indications for all model DC-10 aircraft. Free fall of the center gear, on DC-10 

aircraft so equipped, should be accomplished if the gross weight requires extension. 

Observe the normal extension time of about 13 seconds for the main gear and 10 

seconds for the center gear. There will be a pitch up tendency as the gear extends and 

a noticeable drag increase. The nose up pitching moment is easily controlled by the 

stabilizer trim, and the drag increases are controlled by thrust application. When all 

required gears have been extended, the normal landing gear lever should be placed 

down and the alternate gear handle stowed. 

4.2.2.2.5 APPROACH SPEED AND DESCENT

Once in the vicinity of the outer marker with the slats extended and the landing gear 

down, the speed should be reduced from maneuvering speed 0o/EXT to an approach 

speed of 1.3VS 0o/EXT. Speed reduction should be accomplished with a throttle retard 

and stabilizer trim used to maintain the flight path angle. Passing the outer marker (or 

equivalent), the flight path angle may be adjusted to establish an approach path on or 

slightly below a normal glideslope. As previously discussed, a low rate of descent will 

make the final approach and landing more manageable. The pitch attitude on final will 

be about 10oANU and the descent rate about 500 feet per minute. 
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4.2.2.2.6 PREPARATION FOR TOUCHDOWN

Adjust the touchdown aim point toward the far end of the runway as before, and 

continue to fly the thrust for speed control and stabilizer trim to maintain the desired 

pitch attitude/flight path angle. Smooth deliberate throttle adjustments for speed control 

while far out on the approach will make the task of trimming the stabilizer easier due to 

the slower than normal rate of trim. Begin the transition from stabilizer trim to thrust for 

flight path angle control on the final part of the approach (about 500 feet AGL). 

Aggressive power application (i.e., rapid accelerations followed by immediate throttle 

chops) will allow the small changes in pitch attitude necessary to maintain the 

touchdown aiming point without significantly changing the speed, assuming the 

approach is stable in the pitch axis. Throttle adjustments may need to be more 

aggressive as the airplane enters ground effect. Once again, there should be no 

attempt to accomplish a smooth landing, but simply reduce the sink rate as much as 

practical without ballooning or skipping. 

4.2.2.2.7 STOPPING

Upon touchdown, apply full brake pedal deflection and, as the nose comes to the 

runway, initiate reverse thrust on the wing engines. Continue to hold full brake 

deflection and monitor brake system #2 pressure (powered by hydraulic system #3) 

noting antiskid operation. With the enhanced hydraulics, spoiler panels 3 on each wing 

will be powered and should be deployed at main gear touchdown. Nose wheel steering 

is available and should be used for directional control until the aircraft slows to taxi 

speed. 

The above recommendations represent our suggested means of accomplishing a 

controlled approach and safe landing and are based on the conditions that existed 

during the simulator and flight tests referenced above. Because of the multitude of 

unknown variables that could accompany any given set of actual conditions in any 

future incident, the above procedures may not be optimal for all conditions. However, 

they do provide a good foundation to use as guidance in a different set of 

circumstances. 
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SYSTEMS REMAINING WITH LOSS OF SYSTEMS 1 AND 2 

WITH HYD SYS 3 OFF LIGHT ON

AILERONS LEFT AND RIGHT INBOARD - RIGHT OUTBOARD WHEN SLATS ARE 

EXTENDED OR THE ALTERNATE GEAR EXTENSION LEVER IS 

RAISED. 

ELEVATORS INOPERATIVE - PITCH CONTROL AVAILABLE THROUGH THE RIGHT 

TRIM MOTOR 

SPOILERS NUMBER THREE SPOILER PANEL ON EACH WING 

PITCH TRIM RIGHT TRIM MOTOR 

SLATS OPERATIVE 

NOSE WHEEL 

STEERING 

25° LEFT - 68° RIGHT 

BRAKES NUMBER 2 BRAKE SYSTEM WITH ANTI-SKID - ACCUMULATORS 

SYSTEM ONE 

LANDING 

GEAR 

ALTERNATE GEAR EXTENSION 

Figure 4-5. Catastrophic Failure - Systems Remaining. 
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4.3 STABILIZER INOPERATIVE

If the stabilizer becomes inoperative during flight, high sink rates and steep body angles 

must be avoided, particularly during approach and landing. Improper techniques could 

result in insufficient elevator control being available if large changes in pitch are 

required at low airspeed. Control of descent and pitch must be accomplished by 

coordinated use of thrust and elevator. 

Remember that the autopilot, alternate trim, normal trim and suitcase handles all use 

separate means to energize the trim motors. Exhaust all methods of activating the trim 

before accepting the fact it is inoperative. 

If the flaps are set at 50° and the airplane is trimmed, land with 50° flaps, otherwise use 

35° due to the reduced nose-up force required and the increased elevator effectiveness 

at the higher approach speed. To ensure sufficient elevator for flare, the 35° flap 

threshold speed may have to be increased. The new threshold speed becomes the 

base speed for all normal additives used in determining approach speed. Increases in 

threshold speed depend upon the stabilizer angle when the stabilizer failed and the CG. 

The threshold speed additive can be determined from the chart located in the checklist 

in Section IIA of the Dash One. 

If the stabilizer failure position requires a high force at the control wheel, this force can 

be reduced by manually slewing the elevator feel system to the minimum reference 

airspeed and moving the CG toward the aft limit. 

NOTE

If the stabilizer fails when the airplane is in normal climb, cruise, or 

descent, the required offset pressure will be small until speed is 

reduced when maneuvering for landing. In such a case, as speed 

diminishes, the required backpressure to offset the mis-trim will 

increase as much as 30 lbs nose up. This will be particularly true if 

the stabilizer failed near zero trim units when at cruise, and will apply 

even after slewing the elevator feel system to minimum reference 

speed at the beginning of the approach. The copilot may assist in 

relieving the elevator force.

Engaging an autopilot in CWS (Autopilot Lever moved to CWS or CWS Switch 

depressed) will reduce the control force required to counteract an out-of-trim condition. 

Use of CWS requires accepting the risk that excessive control inputs may cause the 

autopilot to disengage, resulting in unexpected control forces required to maintain pitch 

attitude. 
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NOTE

In the event of a missed approach, the go-around speed should be 

the approach speed as computed in the previous paragraphs but no 

less than the computed 35°/EXT flap/slat threshold speed plus speed 

increment, to ensure elevator effectiveness. 

COMMON ERRORS-STABILIZER INOPERATIVE

• Excessive sink rate developed on final. 

• Trying to hold off airplane during flare. 

• Failing to stabilize at proper approach speed. 

• Reducing speed and thrust too soon. 

• Failing to call for check of chart speed to obtain required additive. 

14.4 ASYMMETRIC FLAPS OR NO FLAP LANDING WITH SLATS

This procedure is required when flaps will not extend or flap asymmetry occurs on initial 

extension. If asymmetry is evident when selecting flaps, immediately retract the flaps to 

zero and effect the approach and landing in the 0°/EXT configuration. If asymmetry 

occurs in a setting greater than the 22° flap selection, do not retract flaps to zero but 

only to the point where asymmetry occurred. Modify the landing procedure and use the 

threshold speed shown for that configuration. If flap symmetry cannot be obtained, find 

the position that provides the most symmetrical configuration and use the speeds 

associated with the least (in degrees) flap position indicated. The following procedures 

should be used for approach and landing with 0°/EXT configuration: 

• Consider landing distance (Figures 4-12 and 4-13) and, if required, dump fuel to 

reduce gross weight. Obtain threshold speed for 0°/EXT from the Performance 

Manual. 

• Plan a slightly wider pattern and a longer final. 

• Extend slats as the pattern (vector) is entered and maintain minimum maneuver 

speed for 0°/EXT. 

• When turning base (final vector heading), bleed off 10 KIAS in level flight. 

• When stabilized on the final approach heading, further bleed off airspeed to the 

approach (1.3VS) speed for 0°/EXT. 

NOTE

In the 0°/EXT configuration, the fast/slow indicator is 

referenced to 1.5VS and will be indicating slow during the 

approach. Do not use autothrottles.
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• Fly a normal glide slope angle. Do not attempt to fly a flat approach. Anticipate 

a higher than normal (800-1000 fpm) rate of descent due to the faster than 

normal speeds involved. Extend the “normal” visual aim point further down the 

runway to compensate for the increased pitch attitude on final approach. 

• As the threshold is attained, reduce sink rate slightly (do not hold off), retard the 

throttles to idle, and fly to a positive touchdown. 

• Lower the nose positively to the runway to reduce lift and increase braking 

capability. Use immediate reverse thrust for assistance in stopping and to 

provide auto-spoiler deployment. See ASYMMETRIC FLAPS OR NO FLAP 

LANDING WITH SLATS profile (Figure 4-6). 

NOTE 

At threshold speed in 0°/EXT configuration, the body angle 

(with a 2.75° glidepath) approximates 8° ANU. The boom 

will contact the runway at a 13° body angle with the struts 

extended, and 11° with the struts compressed.
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For purposes of illustration, the speeds and landing distances are for a 350,000-pound 

airplane. 

Figure 4-6. Asymmetric Flaps or No Flap Landing with Slats. 
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For purposes of illustration, the speeds and landing distances are for a 350,000-pound 

airplane. (See Figure 4-6). 

Figure 4-7. No Flap Landing With Slats Speed Bugs.

The fast/slow indicator in this configuration is referenced to 0°/EXT VMM (1.5VS) and will 

be reading slow during the approach. The fourth bug may be set at VMCA if either rudder 

is inoperative. 

4.4.1 STOPPING DISTANCE

Stopping distance is computed using the Performance Manual and is based upon full 

spoilers with anti-skid braking on all gear, but does not take into account reverse thrust. 

0°/EXT Flare (2200) + Stopping (3600) = 5800 feet 

Landing distances in this case are not appreciably affected by a 1-2 dual hydraulic 

system loss since anti-skid braking is available. 

COMMON ERRORS - NO FLAP LANDING WITH SLATS

• Failing to disconnect ATS for deceleration to VTH. 

• Ducking under the required glide path to compensate for the increased 

pitch attitude. 

• Over flaring with resultant balloon maneuver. 

• Holding the aircraft off instead of flying it onto the runway with a slight 

flare. 

0/EX

V

0

TH1

RET

15V

0°/RET 

VMM1.5V

0°/EXT 

VMM1.5VS 

0°/EXT 

VTH1.3VS 
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4.5 ASYMMETRIC SLATS OR NO SLAT LANDING WITH FLAPS

Whenever slat asymmetry occurs, always attempt to regain symmetry by returning to 

the last position where the slats were symmetrical. If, on initial slat selection, the slats 

will not extend or if asymmetry is evident, place the flap/slat handles in the UP/RET 

detent and split the handles, leaving the slat handle in the RET detent. 

If slat asymmetry occurs when flaps are at a greater setting than 22° (landing slats), do 

NOT retract the flaps and slats completely. Return to the 22° flap position and affect the 

landing using the 22°/EXT configuration and associated speeds regardless of slat 

symmetry. 

The NO SLAT WITH FLAPS procedure is most demanding. The technique is as 

follows: 

• Place the flap handle in the zero detent and leave the slat handle in the RET 

position (split the handles). 

• Consider landing distance (Figure 4-13) and reduce gross weight by fuel 

dumping, if required. 

• Obtain the minimum maneuver speed for 22°/RET and threshold speed for 

22°/RET and 35°/RET from the checklist. 

• Plan a slightly wider pattern (vector), which will provide a longer final for speed 

stabilization. 

• When entering the pattern, select flaps 22°, lower the gear, and maintain 

minimum maneuver speed for 22°/RET. 

• When turning base (final vector), set flaps to 35° or 22°. Bleed off 10 knots in 

level flight if using flaps 35°. 

• When stabilized on the final approach heading (course) and still in level flight, 

further reduce KIAS to approach speed for 35°/RET or 22°/RET. 

• Fly a normal glideslope angle (700-900 fpm rate of descent). Do not attempt to 

flatten approach or drag airplane in. Move the “normal” visual aim point closer to 

the threshold to compensate for the decreased pitch attitude. 

• The airplane will be in a flat or nose-down attitude. When the threshold is 

crossed, use a slight flare to diminish the sink rate and raise the nose of the 

aircraft to at least a level attitude. Reduce thrust. Avoid holding off the airplane 

at touchdown. 

• Immediately use reverse thrust while holding forward pressure on the control 

column to reduce lift and increase braking capability. 

Spoilers will deploy automatically upon touchdown with 35° flap and upon selection of 

reverse thrust if 22° flaps are used. For ASYMMETRIC SLATS OR NO SLAT 

LANDING WITH FLAPS PROFILE, See Figure 4-8. 
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Figure 4-8. Asymmetric Slats or No Slat Landing with Flaps 
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For purposes of illustration, the speeds and landing distances shown are for a 350,000-

pound airplane. At this weight, 35°/RET VTH is only two knots below 35° flap placard. 

Landing with 50°/RET is not considered due to objectionable body angle and absence 

of charted data. 

22 /RET/TH = 191

V

0

22 /RET = 199 TH S 1.35 V

V

0

MM S 1.4 V

0 /RET = 225

V

0

MM S 1.5 V

W0013-09

35 /RET/TH = 183

V

0

TH S 1.35 V 22 /RET

V

0

MM S 1.4 V

0 /RET

V

0

MM S 1.5 V

W0013-10

 Figure 4-9. 22oRET/Landing Figure 4-10. 35o/RET/Landing 

4.5.1 STOPPING DISTANCES 

The difference in landing distance (flare distance and stopping distance) between 

22°/RET and 35°/RET at 350,000 pounds is about 1,000 feet with all other factors being 

equal. These distances are extracted from the Performance Manual and are based 

upon full spoilers with anti-skid braking on all gear, but do not take into account reverse 

thrust. 

22°/RET Flare (2150) + Stopping (4750) = 6,900 feet 

35°/RET Flare (1950) + Stopping (3900) = 5,850 feet 

Landing distances nearly double if the situation is complicated by a 1-3 dual hydraulic 

system failure even when taking into account reverse thrust. This is due to manual 

braking with accumulator pressure. See the landing distance chart in the 1-3 DUAL 

HYDRAULIC SYSTEM FAILURE procedures. 

Note the small difference between 22°/RET 

VMM and VTH. During approach, the fast/slow is 

referenced 22°/RET VTH (1.35VS). 

You can see that the VTH for 35°/RET is not 

much less than VTH for 22°/RET. During 

approach, the fast/slow is referenced to 

35°/RET (1.35VS). 
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COMMON ERRORS - NO SLAT LANDING WITH FLAPS

• Failing to disconnect ATS when landing with 22o/Ret. 

• Constantly going above the required glide path to compensate for the 

decreased pitch attitude. 

• Over flaring with resultant balloon maneuver. 

• Under flaring with resultant nose gear first landing. 

4.6 NO FLAP/NO SLAT LANDING

If multiple malfunctions occur that prevent the use of both flaps and slats for landing, or 

due to a LOSS OF WING POD DROGUE use the following procedure: 

NOTE 

Review DUAL HYDRAULIC SYSTEM FAILURE to 

determine effect on airplane systems and controllability (If 

applicable.)

• Reduce gross weight by dumping fuel to the lowest practical level. Obtain 

minimum maneuvering and threshold speed for 0°/RET from the Performance 

Manual. Limiting tire speed may be a factor. 

• Plan a wide pattern, if visual, or use a longer than normal intercept for ILS 

Localizer. 

• Always extend the gear with the alternate gear extension while level, IAS 230 

KTS or less. This will unlock the outboard ailerons. Lowering the gear will aid in 

speed stabilization. 

• Maintain minimum maneuvering speed throughout the pattern (vectors) until the 

base turn begins. Bleed off 10 knots on base leg (final vector) in level flight. 

• When stabilized on the final approach heading (course), further decelerate while 

in level flight until approach speed for 0°/RET is achieved. A final approach that 

provides a level portion (for speed stabilization) must be used. Deceleration will 

be slow and acceleration rapid. At approximately 600 feet AGL, begin reducing 

speed to the threshold speed. 

NOTE

In the 0°/RET configuration, the fast/slow indicator is 

referenced to 1.5VS and will be indicating slow during the 

approach. Do not use autothrottles.

• Fly a normal glideslope angle. Do not fly a flat approach or attempt to drag in 

the airplane. This complicates finding the correct target touchdown point. 

• When the threshold is attained, an increase in control effectiveness due to 

ground effect will be apparent, especially in elevator control. 
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• Reduce the throttles to idle and use a slight flare to diminish the sink rate. Do not 

hold the airplane off. Flaring will result in float and waste of runway. 

On touchdown, take positive action to lower the nosegear to the runway. This will 

destroy lift and increase braking effectiveness. 

Use reverse thrust immediately. Reverse thrust is most effective at high speeds and will 

provide auto-spoiler deployment. For NO FLAP/NO SLAT LANDING profile, see Figure 

4-12. 

For purposes of illustration, the speeds and landing distances shown are for a 350,000 

pound airplane. 

Figure 4-11. No Flap/No Slat Speed Bugs

The fast/slow indicator in this configuration is referenced to 0°/RET VMM (1.5VS) and will 

be indicating slow during the approach. 

The third bug may be set to VMCA if either rudder is inoperative. 

RET

15V 0°/RET 

VMM1.5VS 0°/RET 

VTH1.35VS
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STOPPING DISTANCE

Stopping distance is computed using the Performance Manual and is based upon full 

spoilers with anti-skid braking on all gear, but does not take into account reverse thrust. 

0°/RET Flare (2750) + Stopping (5500) = 8250 feet 

COMMON ERRORS--NO FLAP/NO SLAT LANDINGS

• Failing to disconnect the Autothrottles. 

• Failing to stabilize speed prior to beginning glideslope. 

• Over controlling longitudinally with resulting speed changes. 

• Flying a flat approach. 

• Abandoning the stabilized glideslope when the runway is near. 

• Failing to move throttles to idle when the runway is near. 

• Flaring and attempting to hold the airplane off rather than flying it on to the 

runway. 
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Figure 4-12. No Flap/No Slat Landing. 
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Figure 4-13. Landing Distance Comparison Dry Runway for 

Training Purposes Only - Not FAA Approved, (Chart extracted 

from the DC-10 Flight Manual). 
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4.7 FLIGHT OPERATION WITH SEVERELY DAMAGED RADOME AND/OR 

SUSPECT AIRSPEED INDICATION

Loss of a major portion of the nose radome on the KC-10 may disrupt the airflow over 

the pitot probes, thereby affecting the measured total pressure. Should this occur, 

those functions dependent on total pressure from the pitot probes (such as indicated 

airspeed and MACH number, overspeed warning, certain autopilot functions, elevator 

load feel and flap limiter programming, etc') may become erroneous or inoperative. 

Damage or loss of a radome will be accompanied by an increase in noise level, a 

relatively small loss in airplane handling characteristics, and a possible change in stall 

or buffet onset speeds due to radome damage. 

Hail encounters, in addition to causing radome and/or pitot tube damage, may also 

cause damage to the wing leading edges, thereby resulting in an increase in airplane 

stall speeds and a degradation of handling characteristics at, or near, stall. (Certain 

types of damage can cause airplane stall prior to stall warning.) Further, the angle-ofattack probes may be damaged, thereby invalidating stall warning. Terminal area and 

final approach speeds contain an arbitrary margin to account for the degree of wing 

leading edge damage nominally occurring due to hail encounter. The final approach 

speeds should be taken into consideration when determining landing field length 

requirements. 

Noise and low frequency buffeting that is distinguishable from stall buffet in that it does 

not change greatly with changes in speed or attitude usually accompany loss of a major 

portion of the nose radome. Once the problem is identified and airspeed appears 

suspect, no difficulties should be encountered as long as pitch attitude is used as the 

primary speed reference by use of the attitude director indicator, and indicated airspeed 

readings are disregarded. 

4.7.1 BLOCKED OR FROZEN PITOT-STATIC SYSTEM 

Unreliable airspeed indications may result from blocking or freezing of the pitot system. 

When the ram air inlet to the pitot head is blocked, pressure in the probe is released 

through the drain holes and the airspeed will drop to zero. If the ram air inlet and the 

probe drain holes are both blocked, trapped pressure within the system reacts 

unpredictably. The pressure may increase through expansion, decrease through 

contraction or remain constant. In each case, the airspeed indications and associated 

air data parameters would be abnormal. This could mean increasing indicated airspeed 

in climb, decreasing indicated speeds in descent or unpredictable indicated speeds in 

cruise. 

Although extremely unlikely, icing of BOTH left and right side fuselage static ports will 

affect the airspeed and the altimeter systems. The pilot and copilot have independent 

pitot-static systems, each with its own power source for anti-icing so that the probability 

of a malfunction occurring in both systems simultaneously is extremely remote. The 

failure of one system, while easier to recognize, could present a problem in determining 
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the valid system. However, the standby system, which is also independent of Pilot and 

Copilot systems, allows comparison for fault isolation purposes. 

Information for pitch attitude flying is contained in the Flight Manual Section 2A, and 

may be used any time airspeed indications are suspect. Linear interpolation for the 

target pitch attitude corresponding to the appropriate gross weight is permissible. 

NOTE 

Groundspeed as displayed on the EHSI will continue to be 

accurate during IAS/TAS malfunctions. Groundspeed may 

be of some value during low altitude operations if winds are 

known to be relatively light. INS wind and TAS readout will 

NOT be accurate.

See figures 4-14 and 4-15 for primary airplane systems that a blocked or frozen Pitot 

and/or static system will affect. Charts are located in the Flight Instruments Section of 

the Flight Manual. 
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Figure 4-14. Pitot/Static System 

6 0 0

5

7

2 9,

0

STANDBY

ALTIMETER

Q

TRANSDUCER

PILOT'S

PITOT

P

AUXILIARY

PITOT

P

1 3

COPILOT'S

PITOT

P 2

1

1301

2299 7 3

2

ALT

5 6 4

8

9 100

FEET

MILLIBARS

IN HG

400

350

300

250 200180160

140

120

100

80

0

KT

60

IAS

STANDBY

AIRSPEED

INDICATOR

CAB ALT

0

0 1

2

3

4

6 5

7

8

9

10

1 1000

FT

2

3

4

5

6

7

8

BARO

PSI

DIFF PRESS

9 10

15

20

25

30

BARO

970 1000 1050 PRESS mb

CABIN

DIFFERENTIAL

PRESSURE

INDICATOR

CABIN

PRESSURE

CONTROLLER

AVIONIC FAN

CONTROL

PRESSURE SWITCH

STATIC AIR

NORM

ALT

STATIC AIR

NORM

ALT

COPILOT'S

INSTRUMENT

PANEL

PILOT'S

INSTRUMENT

PANEL

STATIC SELECTOR

VALVE

(TYPICAL 2 PLACES) 

LEGEND

PILOT'S STATIC

COPILOT'S STATIC

ALTERNATE STATIC

PILOT'S PITOT 2-POSITION

VALVE (AIR)

SA1-206B

MECHANICALLY

ACTUATED

COPILOT'S PITOT

AUXILIARY PITOT

CENTRAL

AIR DATA

COMPUTER-1

CENTRAL

AIR DATA

COMPUTER-2
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Figure 4-15. Central Air Data Computer Pitot Static Interface 

SA1-152D
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THIS PAGE INTENTIONALLY LEFT BLANK 
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5.0 CONSERVATION

The objective of this section is to provide you with relevant information pertaining to 

engine and equipment conservation and, as an adjunct, the conservation of fuel. The 

reasons for the necessity of conservation are threefold: economy, the protection of 

scarce resources, and the ability to be able to rely on the equipment, especially the 

engines, to produce 100% when the chips are down. 

You must instill in yourself a philosophy of conservation, an awareness of the results of 

actions that may be detrimental to equipment life or fuel economy. Preach and practice 

conservation. 

5.1 ENGINES AND FUEL

Your unit has already established procedures and techniques for fuel conservation and 

they will not be repeated here. It is your responsibility to become an expert on these 

directives and demand compliance. 

So, once the established fuel conservation directives are followed to the letter, the 

maximum amount of fuel is being saved, right? Wrong. Specific Fuel Consumption 

(SFC) is directly related to engine deterioration and engine deterioration is directly 

related to the manner in which they are operated. Minimizing the deterioration saves 

fuel and obviously leads to increased engine life. 

What will be covered here are characteristics of the CF6-50 engine that relate directly to 

engine conservation through operational techniques. 

5.2 ENGINE START

Extended start times, increased starting EGT, and the potential for hot/hung starts is 

directly related to pneumatic pressure available to the starter. Slow starts, hung starts, 

and hot starts are detrimental to engine hot section life. It is important to note that the 

starter function does not end when initial lightoff occurs. The starter must continue to 

provide sustained follow-through torque during the lightoff phase to assist engine 

acceleration toward idle. This requirement emphasizes the need for adequate 

pneumatic supply. GE recommends a minimum of 30 PSI for cold weather starts and 

35 PSI will give better assurance of a satisfactory start. The more pressure the better 

for any start regardless of the temperature. 

A vertical thermal gradient affects the CF6-50 after shutdown when the hot air rises and 

is contained by the nacelle. The lower portion of the engine cools more quickly than the 

upper portion. This temperature differential causes a very slight "bowing" of the engine 

rotor and case. If the engine is restarted while this condition exists (typically following a 

short turnaround), an increase in the engine vibration may be detected. The "bowing" 

may cause rubbing and consequently wear between static and rotating components. It 

may even prevent N1 rotation during the initial start sequence. 
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5.2.1 RECOMMENDED STARTING PRACTICES 

1. Start one engine at a time if operational requirements permit. 

2. Operate the APU and pneumatic system in a manner that will provide 

optimum pneumatic pressure for start; i.e., all bleeds off in accordance 

with tech order procedures and only use the standby mode of the APU as 

a last resort. 

3. Comply with engine cooling recommendations prior to shutdown (see 

engine shutdown in this section). 

4. Maximize the amount of time between engine shutdown and engine start. 

After starting a warm engine, operate it at low power for several minutes 

for thermal stabilization. This minimizes transitory vibration stresses and 

the probability of rubbing between static and rotating components. 

Preventing rubs reduces EGT and SFC deterioration. 

5.3 TAXI

The primary cause of engine deterioration during taxi operations is the susceptibility to 

FOD or the sandblasting effect of contaminant ingestion. A characteristic of the CF6-50 

engine is the vortex that occurs between the lip of the wing engine inlet and the ground. 

This vortex may begin to form at idle thrust if there is no wind and the aircraft is static. 

When the vortex is present, contaminants will be sucked into the engine. The formation 

and intensity of this vortex is dependent upon the relationship between airplane velocity 

and engine RPM (see Figure 5-1). Power settings as low as 40% N1 create a vortex 

with sufficient suction to lift one-inch stones into the inlet. Forward taxi speed of 10 

KIAS is sufficient to eliminate the vortex. 

5.3.1 RECOMMENDED TAXI PRACTICES 

1. If taxi is to be accomplished on contaminated taxiways (sand, slush, ice, 

etc.) or if the taxiway width results in the wing engines extending beyond 

the taxiway, restrict the use of more than idle thrust on the wing engines. 

Start number two engine and use it to accelerate to, and maintain, taxi 

speed. Observe the cautions relating to center engine at more than idle 

thrust in the terminal area or in the proximity of other aircraft. 

2. Adhere to proper taxi speeds and practices. 
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5.4 TAKEOFF

Engines are particularly susceptible to deterioration during takeoff. The low speed, high 

thrust conditions for wing engine vortex are present during the initial portion of the 

takeoff roll and a conscious effort is required to prevent vortex induced FOD. 

Target N1 should be captured by 80 KIAS. A slight N1 overshoot and then a drop after 

throttle set may be observed. This is a normal characteristic of high by-pass engines 

and should not be countered with a throttle adjustment. Rapid throttle movement 

toward takeoff thrust will cause excessive N1 and EGT overshoot. These overshoots 

are detrimental to engine life and are caused by uneven heating between engine 

components effecting clearances and flow areas (see Figure 5-2). Reduced thrust 

takeoff results in reduced fuel flow, N1, N2, and EGT. A 1% thrust reduction represents 

approximately a 1% fuel flow reduction. However, since the reduced thrust is 

maintained for a longer period than full thrust, the net fuel used may not change significantly. The major benefit of reduced thrust takeoffs is the reduction of engine EGT 

and SFC deterioration rate. 

Compressor erosion also reduces EGT/SFC margins as previously discussed. Rolling 

takeoff technique inhibits formation of ground to inlet vortices minimizing foreign object 

particle ingestion, a major cause of erosion. At 30 KIAS vortex formation is prevented 

even at maximum power settings. 

Bleed extraction also effects fuel flow and EGT. When bleed air is extracted, the N2

RPM must be increased. This requires increased fuel flow to the combustor, which 

results in increased EGT. Typical pneumatic bleed extraction at takeoff power results in 

a fuel flow increase of approximately one percent with a resultant EGT increase of 

approximately 7° C. 

5.4.1 RECOMMENDED TAKEOFF PRACTICES 

1. When practical, takeoff with zero bleed until after first power reduction. 

2. Rolling takeoffs are preferred. If a static takeoff must be made, keep in 

mind the susceptibility of the wing engines to FOD due to vortex at high 

power settings below 30 KIAS. 

3. Advance throttles slowly as aircraft gains speed. Approximate takeoff 

thrust before engaging autothrottles to prevent excessive N1 overshoot. 

4. Do not adjust throttles after 80 KIAS. A decrease in N1 as the speed 

increases is normal and is caused by "Ram Drag" and increased TAT. 

5. Do not engage the autothrottles after takeoff if the TRC is in takeoff. An 

overboost can occur. 
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Figure 5-1. Aircraft Taxi Speed (Knots).

Figure 5-2. Time - Minutes.

W0017-01

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5.5 INFLIGHT ENGINE OPERATION, GENERAL

Rapid throttle movement with a heat soaked N2 rotor may cause turbine blade tip to 

turbine shroud rubbing with subsequent SFC deterioration. GE recommends that rapid 

throttle movements be avoided any time the engine has been operated at an EGT 

above 800° C. (EGT during climb normally exceeds 800.) Generally speaking, slow, 

smooth throttle movements lead to longer engine life and fuel economy due to 

prolonged SFC values. Minimizing bleed extraction during flight operations reduces fuel 

flow and EGT for the same reasons stated under the takeoff discussion. Derated climbs 

save engines. Consideration should be given to using MCR thrust for a portion of the 

climb anytime the gross weight is low enough to allow a derated takeoff. Generally, 

derated climbs are less fuel efficient in the short term than CL thrust climbs; however, 

the difference is some cases is negligible. For instance, derating climb power to 10,000 

feet at light gross weights. 

5.5.1 RECOMMENDED CLIMB PRACTICES 

1. Achieve climb thrust and climb configuration as soon as flight manual 

procedures and safety considerations permit. 

2. Adhere to the flight manual climb profile. 

3. Anticipate thrust reductions/applications and make smooth, slow throttle 

movements. During climb, fan speed increases with altitude. When 

autothrottles are not engaged, the thrust setting should be refined about 

every 5000 feet to prevent exceeding climb N1. 

4. Eliminate unnecessary bleed extraction. 

5. Consider use of derated (MCR) climb for the initial portion of the climb 

during light gross weight takeoffs. 

5.5.2 RECOMMENDED CRUISE/AIR REFUELING PROCEDURES 

1. Minimize bleed extraction through use of single pack operation when 

circumstances permit and avoid unnecessary use of anti-ice. (If a pack is 

not to be used, closing the ram air door by manually positioning the pack 

temperature valve to full hot will reduce drag and save fuel.) 

2. Use manual throttle control in turbulence, during large speed changes in 

SPD mode or anytime autothrottle excursions exceed 5% N1. 

3. Practice breakaway procedures in the aircraft only when absolutely 

necessary. 
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5.5.3 RECOMMENDED DESCENT PROCEDURES 

1. Accomplish descents in idle thrust with autothrottles off. 

2. Avoid rapid thrust reduction at top of descent. 

3. Anticipate intermediate leveloffs with slow, smooth throttle applications. 

5.5.4 TRAFFIC PATTERN RECOMMENDATIONS 

1. Large throttle excursions should be avoided. Autothrottle go-around is 

particularly tough on the engines and should be avoided if at all possible. 

5.6 LANDINGS

Moderate use of reverse thrust during normal landings provides optimum balance of 

costs considering fuel consumption, engine deterioration rate, and aircraft brake wear. 

A Lufthansa study revealed that a reverse thrust N1 RPM of 50%-70% results in the 

greatest cost savings. There have been numerous problems associated with reverse 

thrust at low speeds during landing roll including contaminant precipitated malfunctions 

of the fan air valve and thrust reverser shuttle valve along with FOD caused blade 

erosion. 

5.6.1 RECOMMENDED LANDING PRACTICES 

1. Use reduced reverse thrust if conditions permit. 

2. Insure that engines are in reverse idle by 60 KIAS, and forward idle by 

turn off speed. 

5.6.2 RECOMMENDED ENGINE SHUTDOWN PRACTICES 

1. The engine cool down procedures are outlined in the flight manual. 

5.7 AVIONICS COOLING

As a general rule, do not operate the avionic equipment when it is not needed. There is 

no reason to have the black boxes powered up when conducting operations that require 

only ground service bus power such as loading and unloading cargo. Heat sensitive 

avionics equipment receives power from AC bus 1 and AC bus 3. Using the applicable 

Ext Pwr Avail or APU Pwr Avail switch on the ground service bus panel to connect 

power to the ground service busses causes the AC bus 1 and AC bus 3 to remain 

unpowered. The ground service bus Ext Pwr Avail and APU Pwr Avail switches are 

located at the service bus Ext Pwr Avail and APU Pwr Avail switches are located at the 

external power receptacle and at the galley. If the avionic equipment is to be operated 

on ground, comply with flight manual cooling procedures. 
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5.8 TIRES AND BRAKES

See taxi procedures. 

5.9 FLIGHT CONTROLS

During the flight control check, avoid rapid inputs and control reversal. "Slamming" the 

control surfaces from stop to stop is particularly detrimental to the hinges and attach 

points. 

5.10 MISCELLANEOUS

If you don't need it, don't use it. Ignition, strobes, etc. Component life is directly related 

to component use, think and practice conservation. 

Material in this section has been extracted from, DAC "Know Your DC-10" letters, GE 

"Operations engineering bulletins", GE "Customer Rep. tips", DAC "All operator letters" 

and GE "Items of interest" from product support engineering. 
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APPENDIX A

DEFINITION OF SYMBOLS
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DEFINITION OF SYMBOLS

 

Symbol Definition Symbol Definition 

 

ICAO International Civil Aviation RCR Runway Condition Reading 

Organization 

IMN Indicated Mach Number R/D Rate of Descent 

in Inches RL Runway Length 

ISA International Standard rpm Revolution per Minute 

 Atmosphere 

LB, Lb Pounds(s) RSC Runway Surface Condition 

SDR Stopping Distance Ratio Sec Seconds 

SL Sea Level V2 Climbout Speed 

Std Standard Δδ Increment of Weight, Drag, 

Airspeed, etc. and Air 

Pressure Ratio (P/Po) 

TAS True Airspeed--Equivalent δ f Wing Flap Deflection 

airspeed corrected for density 

TAT Total Air Temperature ρ (Rho) Air Density, Slugs Per 

 Cubic Foot 

Temp. Temperature σ (Sigma) Air Density Ratio (P/Po) 

TD Touchdown θ (Theta) Temperature Ratio (T/To) 

TO Takeoff < Less Than 

TOGW Takeoff Gross Weight > Greater Than 

TSLW Tire Speed Limiting Weight ≤ Less Than or Equal To 

V Velocity ≥ Greater Than or Equal To 

VA Design Maneuvering Speed AERODYNAMICS AND

PERFORMANCE

VC Maximum Cruise Speed ASE Automatic slat extension 

VD Design Diving Speed CG Center of Gravity 

 
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Symbol Definition Symbol Definition 

 .. 

VLD Airspeed for Maximum Glide CP Center of pressure (of lift) 

 Ratio 

 DLE Drooped leading edge 

VLO Liftoff Speed 

VMBE Maximum Brake Energy Speed FGS Flight Guidance System 

VMCA Minimum Control Speed With k Temperature recovery factor 

One Wing Engine Inoperative 

In-Flight L/D Lift to drag ratio 

VMCG Minimum Control Speed With NAM/LB Nautical air miles per 

 One Wing Engine Inoperative pound of fuel 

 On-Ground 

VMM Minimum Maneuver Speed SAT Static Air Temperature 

 

VMO Maximum Operating Limit TAT Total Air Temperature (ram 

 Speed air temperature) 

VMU Minimum Unstick Speed 

VR Rotation Speed 

VSMIN Minimum Stalling Speed 

VSIG 1g Stalling Speed 

VTH Threshold Speed 

V1 Critical Engine Failure Speed 
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APPENDIX B

GLOSSARY OF TERMS
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GLOSSARY OF TERMS

Aerodynamic Center: The point along the chord where all the changes in lift take 

place. 

AOA: Angle of attack is the angle between a reference line in a 

body and the line of the relative airflow. 

Angle Of Incidence: The acute angle formed by the chord line of the wing and 

the longitudinal axis of the airplane. It is the angle at which 

the wing is attached to the fuselage. Since it was 

determined by the airplane design, it is a fixed angle and 

cannot be changed by the Pilot. Angle of incidence should 

not be confused with angle of attack. 

Asymmetrical Thrust: An imbalance in thrust forces acting through one of the 

aircraft axis. Generally associated with failure of a wing 

engine on a multi-engine aircraft. 

Aspect Ratio: The proportion of the span and average chord of an airfoil. 

Calibrated Airspeed: Indicated airspeed corrected for instrument and position 

error. In the KC-10, the Pilot and Copilot's airspeed 

indicators read calibrated airspeed. 

Center of Gravity: The balance point of the aircraft. 

Compressibility: The changes in air density and pressure that take place 

during high-speed flight. 

Critical Mach: For any given airfoil, the highest speed possible without the 

airflow across the airfoil going supersonic. 

Deck Angle: Synonymous with pitch attitude. This term is used in the 

KC-10 to denote pitch angle and should not be confused 

with angle of attack. The deck angle is the relationship of 

the longitudinal axis of the airplane with the horizon. 

Dihedral: The angle that the wing or horizontal stabilizer is displaced 

from horizontal when in level flight. 

Dutch Roll: A yawing and rolling oscillation. 

Dynamic Stability: Behavior of the airplane as a function of time following a 

disturbance from equilibrium. 

EGT: Exhaust gas temperature.

Equivalent Airspeed: Calibrated airspeed corrected for compressibility. 
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F/F: Fuel flow. 

Flow Separation: When the boundary layer flow no longer adheres to the 

airfoil surface. 

Indicated Airspeed: Airspeed read on the airspeed indicator. In the KC-10 the 

indicated airspeed read on the Pilot's and Copilot's airspeed 

indicators is the same as calibrated airspeed. 

iH: Angle of incidence between the horizontal stabilizer chord 

and the longitudinal axis of the airplane. 

Induced Drag: The drag generated as a result of lift. 

iW: Angle of incidence is between the chord line of the wing and 

the longitudinal axis of the airplane. 

L/D ("L over D"): Ratio of drag to lift for any given airfoil. 

L/D Max: For a given airfoil, the point on the L/D curve where the 

most lift for the least drag is obtained. Clean wing L/D max 

for the KC-10 is 17. 

Lateral Stability: Stability of the airplane about the lateral axis. 

Limit Load Factor: The load factor which corresponds to 100 percent of design 

limit. That load factor which an airplane can withstand 

without permanent structural deformation. 

Longitudinal Stability: Stability of the airplane about the longitudinal axis. 

Manufacturer’s Empty Weight: Weight of the airplane, including the structure, power plant, 

furnishings, systems and other items of equipment that are 

considered an integral part of a particular configuration. It is 

essentially a "dry" weight, including only those fluids, which 

are considered in a closed system such as the hydraulic 

system. 

Maximum Fuel Unbalance The maximum permissible lateral unbalance caused by 

Weight: uneven fuel distribution. 

Maximum Gross Weight: The maximum weight authorized by the Flight Manual for an 

airplane and its contents. 

Maximum Landing Weight: The maximum airplane weight permitted for a touchdown 

with a rate of sink 10 feet per second. This may also be 

further restricted by performance requirements. 
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Maximum Ramp Weight: That weight above maximum takeoff weight, which may 

include additional fuel for ground maneuvering, taxi, and 

run-up. 

Maximum Takeoff Weight: The maximum airplane weight permitted for takeoff under 

any condition but which may be further restricted by 

performance requirements. 

Mean Aerodynamic Chord 

(MAC): The chord drawn through the centroid or geographical 

center of a planform. 

Normal Load Factor: The ratio of lift to airplane gross weight, or the acceleration 

perpendicular to the flight path in "g" units. 

N1 and N2: Fan (and) high pressure rotor speeds expressed as rpm (a 

percent of reference rpm value). 

Operating Empty Weight: The manufacturer's empty weight of the airplane plus the 

weight of the oil, water, undisposable fuel, crew, crew 

baggage, passenger equipment and emergency equipment. 

Peaky: A descriptive word applied to an airfoil that has a high value 

of lift at the leading edge. The KC-10 has a "peaky" wing. 

Planform: The overall shape and size of a wing or stabilizer. 

Profile Drag: Synonymous with parasite drag. The drag generated by 

friction and form. 

Section: A cross-section of an airfoil at a given point. The overall 

aerodynamic quality of an airfoil is a sum of the section 

qualities. 

SFC: Specific fuel consumption. The amount of fuel consumed in 

pounds per hour divided by the thrust of the engine in 

pounds. 

Shock Waves: A line in the flow field about the airplane across which an 

instantaneous reduction of velocity occurs, together with 

associated instantaneous changes in density, pressure, and 

temperature. 

Sideslip: Displacement of the airplane centerline about the vertical 

axis in reference to the relative wind. 

Speedbrakes: Panels which are deployed to destroy lift and/or add drag. 

Slats: High lift devices applied to the leading edge of a wing. 
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Specific Consumption: Amount of fuel consumed by an engine for a given amount 

of thrust. 

Spiral Stability: The resistance of an airplane to lateral and directional 

deviation from a desired flight path following a disturbance. 

Spoilers: Synonymous with speedbrakes. 

Sweep: The angle formed by the wing or stabilizer and the 

longitudinal axis of the aircraft. 

Taper Ratio: The ratio of the width of the root of the wing or stabilizer to 

the width of its tip. 

Ultimate Load Factor: The maximum load factor an airplane can withstand without 

suffering structural failure. This is never less than 1.5 times 

the design limit load factor. 

Washout: The change in the angle of incidence of an airfoil relative to 

the distance from the root. 

Weathercock: The tendency of an airplane to align itself longitudinally 

about its vertical axis with the relative wind. 

Yaw: Displacement of the airplane centerline from some 

reference azimuth. 

Zero Fuel Weight: The maximum weight authorized for an airplane with 

payload, with no disposable fuel. 
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