Chapter 1 — Part 7

each airplane is outlined in the pilot’s hand-

book and it is impcrativc that the specific tech-

nique be followed for successful recovery.

PITCH-UP

The term of “pitch-up” generally applies to

the static longitudinal instability encountered

by certain configurations at high angle of

attack. The condition of pitch-up is illustrated

by the graph of CM versus C, in figure 4.33.

Positive static longitudinal stability is evident

at low values of Cs by the negative slope of the

curve. At higher values of Cs the curve changes

to a positive slope and large positive pitching

moments are developed. This sort of in-

stability implies that an increase in angle of

attack produces nose up moments which tend

to bring about further increases in angle of

attack hence the term “pitch-up” is applied.

There are several items which may con-

tribute to a pitch-up tendency. Sweepback

of the wing planform can contribute unstable

moments when separation or stall occurs at

the tips first. The combination of sweepback

and taper alters the lift distribution to produce

high local lift coefficients and low energy

boundary layer near the tip. Thus, the tip

stall is an inherent tendency of such a plan-

form. In addition, if high local lift coefficients

exist near the tip, the tendency will be to incur

the shock induced separation first in these

areas. Generally, the wing will contribute

to pitch-up only when there is large sweepback.

Of course, the wing is not the only item con-

tributing to the longitudinal stability of the

airplane. Another item important as a source

of pitch-up is the downwash at the horizontal

tail. The contribution of the tail to stability

depends on the change in tail lift when the air-

plane is given a change in angle of attack.

Since the downwash at the tail reduces the

change in angle of attack at the tail, any in-

crease in downwash at the tail is destabilizing.

For certain low aspect ratio airplane configura-

tions, an increase in airplane angle of attack

may physically locate the horizontal tail in

NAWEPS DD-EDT-89

STABILITY AND CQNROL

the wing flow field where higher relative

downwash exists. Thus, a decrease in stability

would take place.

Certain changes in the flow field behind the

wing at high angles of attack can produce large

changes in the tail contribution to stability.

If the wing tips stall first, the vortices shift in-

board and increase the local downwash at the

tail for a given airplane C,. Also, the fusel~age

at high angle of attack can produce strong

cross flow separation vortices which increase

the local downwash for a horizontal tail placed

above the fuselage. Either one or a combiua-

tion of these downwash influences may provide

a large unstable contribution of the horizontal

tail.

The pitch-up instability is usually conlined

to the high angle of attack range and may be

a consequence of a configuration that otherwise

has very desirable flying qualities. In such a

case it would be necessary to provide some

automatic control function to prevent entry

into the pitch-up range or to provide synthetic

stability for the condition. Since the pitch-up

is usually a strong instability with a high1

rate of divergence, most pilots would not be

capable of contending with the condition. At

high 4, pitch-up would be of great danger in

that structural failure could easily result. At

low q, failing flight loads may not result but

the strong instability may preclude a successful

recovery from the ensuing motion of the, air-

plane.

EFFFCTS OF HIGH MACH NUMBFB

Certain stability problems are particular to

supersonic flight. While most of the problem

areas have been treated in particular in previous

discussion, it is worthwhile to review the

effects of supersonic flight on the various items

of stability.

The static longitudinal stability of an air-

plane increases during the transition from sub-

sonic to supersonic flight. Usually the prin-

cipal source of the change in stability is due to

the shift of the wing aerodynamic center with

NAVWEPS oo-s01-80

STABILITY AND CONTROL

Mach number. As a corollary of this increase

in stability is a decrease in controllability and

an increase in trim drag.

The static directional stability of an air-

plane decreases with Mach number in super-

sonic flight. The influence of the fuselage and

the decrease in vertical tail lift curve slope

bring about this condition.

The dynamic stability of the airplane

generally deteriorates with Mach number in

supersonic flight. Since a large part of the

damping depends on the tail surfaces, the

decrease in lift curve slope with Mach number

will account in part for the decrease in damp

ing. Of course, all principal motions of the

aircraft must have satisfactory damping and

if the damping is not available aerodynami-

cally it must be provided synthetically to

obtain satisfactory flying qualities. For many

high speed configurations the pitch and yaw

dampers, flight stabilization systems, etc.,

are basic necessities rather than luxuries.

Generally, flight at high Mach number will

cake place at high altitude hence the effect of

high altitude must be separated for study.

All of the basic aerodynamic damping is due

to moments created by pitching, rolling, or

yawing motion of the aircraft. These moments

are derived from the changes in angles of

attack on the tail surfaces with angular

rotation (see fig. 4.15). The very high true

airspeeds common to high altitude flight

reduce the angle of attack changes and reduce

the aerodynamic damping. In fact, the aero-

dynamic damping is proportional to &

similar to the proportion of true airspeed to

equivalent airspeed. Thus, at the altitude of

4O,C00 ft., the aerodynamic damping would

be reduced to one-half the sea level value and

at the altitude of 100,000 ft. the aerodynamic

damping would be reduced to one-tenth the

sea level value.

High dynamic pressures (high $I can be

common to flight at high Mach number and

adverse aeroelastic effects may be encountered.

If the aircraft surfaces, encounter significant

deflection when subject to load, the tendency

may be to lower the contribution to static

stability and reduce the damping contribution.

Thus, the problem of adequate stability of the

various airplane motions is aggravated.

PILOT INDUCED OSCILLATIONS

The pilot may purposely induce various

motions to the airplane by the action of the

controls. In additron, certain undesirable

motions may occur due to inadvertent action

on the controls. The most important con-

dition exists with the short period longitu-

dinal motion of the airplane where pilot-

control system response lag can produce an

unstable oscillation. The coupling possible

in the pilot-control system-airplane combi-

nation is most certainly capable of producing

damaging flight loads and loss of control of

the airplane.

When the normal human response lag and

control system lag are coupled with the air-

plane motion, inadvertent control reactions

by the pilot may furnish a negative damping

to the oscillatory motion and dynamic in-

stability exists. Since the short period motion

is of relatively high frequency, the amplitude

of the pitching oscillation can reach dangerous

proportions in an unbelievably short time.

When the pilot induced oscillation is en-

countered, the most effective solution is an

immediate release of the controls. Any at-

tempt to forcibly damp the oscillation simply

continues the excitation and amplifies the

oscillation. Freeing the controls removes

the unstable (but inadvertent) excitation and

allows the airplane to recover by virtue of

its inherent dynamic stability.

The pilot induced oscillation is most likely

under certain conditions, Most obvious is the

case of the pilot unfamiliar with the “feel”

of the airplane and likely to overcontrol or

have excessive response lag. High speed flight

at low. altitude (high 4) is most likely to

provide low stick-force gradients and periods

of oscillation which coincide with the pilot-

control system response lag. Also, the high 4

flight condition provides the aerodynamic

capability for failing flight loads during the

oscillation.

If a pilot induced oscillation is encountered

the pilot must rely on the inherent dynamic

stability of the airplane and immediately

release the controls. If the unstable excitation

is continued, dangerous oscillation amplitudes

will develop in a very short time.

ROLL COUPLING

The appearance of “inertia coupling” prob-

lems in modern airplanes was the natural result

of the progressive change in aerodynamic and

inertia characteristics to meet the demands of

high speed flight. Inertia coupling problems

were unexpected only when dynamic stability

analyses did not adequately account for the

rapid changes in aerodynamic and inertia

characteristics of airplane configurations. The

The term of “intertia coupling” is somewhat

misleading because the complete problem is

one of aerodynamic as well as inertia coupling.

“Coupling” results when some disturbance

about one airplane axis causes a disturbance

about another axis. An example of uncoupled

motion is the disturbance provided an airplane

when subjected to an elevator deflection. The

resulting motion is restricted to pitching

motion without disturbance in yaw or roll.

An example of, coupled motion could be the

disturbance provided an airplane when sub-

jected to rudder deflection. The ensuing mo-

tion can be some combination of yawing and

rolling motion. Hence, the rolling motion is

coupled with the yawing motion to define the

resulting motion. This sort of interaction

results from aerodynamic characteristics and is

termed “aerodynamic coupling.”

A separate type of coupling results from the

inertia characteristics of the airplane conligura-

tion. The inertia characteristics of the com-

plete airplane can be divided into the roll, yaw,

NAVWEPS OO-SOT-80

STABILITY AND CONTROL

and pitch inertia and each inertia is a measure

of the resistance to rolling, yawing, or pitching

acceleration of the airplane. The long,slender,

high-density fuselage with short, thin wings

produces a roll inertia which is quite small in

comparison to the pitch and yaw inertia.

These characteristics are typical of the modern

airplane configuration. The more conventional

low speed airplane may have a wingspan

greater than the fuselage length. This type of

configuration produces a relatively large roll

inertia. A comparison of these configurations

is shown in figure 4.34.

Inertia coupling can be illustrated by con-

sidering the mass of the airplane to be con-

centrated in two elements, one representing the

mass ahead of the c.g. and one representing the

mass behind the c.g. There are two principal

axis systems to consider: (1) the aerodynamic,

or wind axis is through the c.g. in the relative

wind direction, and (2) the inertia axis is

through the c.g. in the direction of the two

element masses. This axis system is illus-

trated in figure 4.34.

If the airplane shown in figure 4.34 were in

some flight condition where the inertia axis

and the aerodynamic axis are alined, no inertia

coupling would result from rolling motion.

However, if the inertia axis is inclined to the

aerodynamic axis, rotation about the aero-

dynamic axis will create centrifugal forces and

cause a pitching moment. In this case, a

rolling motion of the aircraft induces a pitch-

ing moment through the action of inertia

forces. This is “inertia coupling” and is

illustrated by part B of figure 4.34.

When the airplane is rotated about the

inertia axis no inertia coupling will exist but

aerodynamic coupling will be present. Part

C of figure 4.34 shows the airplane after rolling

90” about the inertia axis. The inclination

which was initially the angle of attack (a) is

now the angle of sideslip (-6). Also the

original zero sideslip has now become zero

angle of attack. The sideslip induced by this

90° displacement will affect the roll rate

NAVWEPS OD-3OT-80

STABILITY AND CONTROL

RELATIVELY

HIGH

ROLL

INERTIA

cc > -I

RELATIVELY

y$z>

0 A /7

MASS ROLL

MOTION

ROLL

MOTION

WSITIVE ANGLE

OF ATTACK.

ZERO SIDESLIP

FUSELAGE

fh SIDEFORCE

AERODYNAMIC

AXIS

FINITE SIDESLIP

u pgq ROLL

MOTION

Figure 4.34. Roll Coupling

depending on the nature of the dihedral effect

of the airplane.

It should be noted that initial inclination of

the inertia axis above the aerodynamic axis

will cause the inertia couple to provide adverse

yaw with rolling motion. If the inertia axis

were initially inclined below the aerodynamic

axis (as may happen at high 4 or negative load

factors), the roll induced inertia couple would

provide proverse yaw. Thus, roll coupling

may present a problem at both positive and

negative inclination of the inertia axis depend-

ing on the exact aerodynamic and inertia

characteristics of the configuration.

As a result of the aerodynamic and inertia

coupling, rolling motion can induce a great

variety of longitudinal, directional, and lateral

forces and moments. The actual motion of

the airplane is a result of a complex combina-

tion of the aerodynamic and inertia coupling.

Actually, all airplanes exhibit aerodynamic

and inertia coupling but of varying degrees.

The roll coupling causes no problem when the

moments resulting from the inertia couple are

easily counteracted by the aerodynamic re-

storing moments. The very short span, high

speed modern aircraft has the capability for

the high roll rates which cause large magni-

tudes of the inertia couple. The low aspect

ratio planform and flight at high Mach number

allow large inclination of the inertia axis with

respect to the aerodynamic axis and also add

to the magnitude of the inertia couple. In

addition, the aerodynamic restoring moments

deteriorate as a result of high Mach number

and angle of attack and can create the most

serious roll coupling conditions.

Since the roll coupling induces pitching and

yawing motion, the longitudinal and direc-

tional stability is important in determining the

overall characteristics of the coupled motion.

A stable airplane, when disturbed in pitch and

yaw, will return to equilibrium after a series

of oscillations. For each flight condition, the

airplane will have a coupled pitch-yaw fte-

quency between the uncoupled and separate

NAVWEPS 00-8OT-80

STABILITY AND CONTROL

pitch frequency and yaw frequency. Gen-

erally, the greater the static longitudinal and

directional stability, the higher will be the

coupled pitch-yaw frequency. When the air-

plane is subject to roiling motion, the inertia

couple disturbs the airplane in pitch and yaw

with each roll revolution and provides a dis-

turbing forcing function.’ If the airplane is

rolled at a rate equal to the coupled pitch-yaw

frequency, the oscillatory motion will either

diverge or stabilize at some maximum ampli-

tude depending on the airplane characteristics.

The longitudinal stability of the typical high

speed configuration is much greater than the

directional stability and results in a pitch fre-

quency higher than the yaw frequency. In-

creasing the directional stability by increasing

the vertical tail area, addition of ventral hns,

or use of stabilization systems will increase the

coupled pitch-yaw frequency and raise the roll

rate at which a possible divergent condition

could exist. Increasing directional stability

by the addition of ventral fins rather than by

addition to the vertical tail has an advantage

of not contributing to the positive dihedral

effect at low or negative angles of attack.

High dihedral effect makes higher roll rates

more easily attainable in roll motion where

proverse yaw occurs.

Since the uncoupled yawing frequency is

lower than the pitching frequency, a divergent

condition would lirst reach critical proportions

in yaw, closely followed by pitch. Of course,

whether the airplane motion becomes divergent

directionally or longitudinally is of academic

interest only.

There is one additional type of coupling

problem that is referred to as “autorotative

rolling.” A rolling airplane which has a high

positive dihedral effect may reach a large pro-

verse sideslip as a result of the inertia couple and

the rolling moment due to sideslip may exceed

that available from lateral control. In such

a case it would not be possible to stop the air-

plane from rolling although lateral control

was held full against the roll direction. The

design features which result in a large positive

dihedral effect are high sweepback, high wing

position, or large, high vertical tail, When

the inertia axis is inclined below the aero-

dynamic axis at low or negative angles of

attack, the roll induced inertia couple results

in proverse yaw.

Depending on the flight condition where the

roll coupling problem exists, four basic types

of airplane behavior are possible:

(1) Coupled motion stable but unacceptabk.

In this case the motion is stable but proves

unacceptable because of poor damping of the

motion. Poor damping would make it

dificult to track a target or the initial am-

plitudes of the motion may be great enough

to cause structural failure of loss of control.

(2) Coupled motion stable and acceptable.

The behavior of the airplane is stable and

adequately damped to allow acceptable

target tracking. The amplitudes of motion

are too slight to result in structural failure

or loss of control.

(3) Coupled motion divergent and unacceptable.

The rate of divergence is too rapid for the

pilot to recognize the condition and recover

prior’ to structural failure or complete loss

of control.

(4) Coupled. motion divergent but acceptable.

For such a condition the rate of divergence

is quite slow and considerable roll displace-

ment is necessary to produce a critical ampli-

tude. The condition can be recognized

easily in time to take corrective action.

There are available various means to cope

with the problem of roll coupling. The fol-

lowing items can be applied to control the

problem of roll coupling:

(ZZ) Increase directional stability.

(b) Reduce dihedral effect.

(c) M’ h 1‘ mnmze t e mc mation of the inertia

axis at normal flight conditions.

(d) Reduce undesirable aerodynamic

coupling.

(e) Limit roll rate, roll duration, and

angle of attack or load factor for performing

rolling maneuvers.

NAVWEPS DD-EOT-80

STABMTY AND CONTROL

The first four items can be effected,only during

design or by design changes. Some roll per-

formance restriction is inevitable since all of

the desirable characteristics are difficult to

obtain without serious compromise elsewhere

in the airplane design. The typical high

speed airplane will have some sort of roll pet-

formance limitation provided by flight restric-

tions or automatic control devices to prevent

reaching some critical condition from which

recovery is impossible. Any roll restriction

provided an airplane must be regarded as a

principal flight operating limitation since the

more severe motions can cause complete loss

of control and structural failure.

HELICOPTER STABILITY AND CONTROL

In discussing many of the problems of sta-

bility and control that occur in high speed

airplanes, one might be prone to believe that

the slow flying helicopter does not have any

such problems. Unfortunately, this is not

the case. Flying qualities that would be con-

sidered totally unsatisfactory by fixed-wing

standards ate normal for helicopters. Heli-

copter pilots are living evidence that an un-

stable aircraft ca. k‘.: controlled. Also, they

are evidence ~a. control without stability

requires constant attention and results in con-

siderable pilot fatigue.

“Inertia coupling” problems are relatively

new to fixed-wing aircraft but a similar effect

in the helicopter rotor has resulted in some

of its most important characteristics. This

aerodynamic-dynamic coupling effect is so im-

portant that it must be considered in discussing

both stability and control. The helicopter

derives both longitudinal and lateral control

by tilting the main rotor and thus producing

a pltchmg or rolling moment as indicated in

figure 4.35. The magnitude of the rotor thrust

the angle of tilt, and the height of the rotor

hub above the c.g. determine the control

moment produced. It should be noted that

low control effectiveness would result when

the rotor thrust is low. Some helicopters

NAWWEPS 00-8OT-80

STABILITY AND CONTROL

THRUST

A C.G.

ROTOR GYROSCOPIC ACTION

THESE FORCES PRODUCE THIS MOMEPdT AND DISPLACEMENT

Figure 4.35. Rotor Forces and Moments

employ an offset flapping hinge to increase the

control effectiveness by creating a centrifugal

force couple when the rotor is tilted. This is

shown in figure 4.35.

The rotor is tilted by taking advantage of

the gyroscopic effect of the rotor system. This

effect causes a rotating mass which is disturbed

about one axis to respond about another axis,

as shown in figure 4.35. A forward tilt to the

rotor is obtained by decreasing the pitch of the

blade when at the starboard position and in-

creasing the pitch of the blade when at the

port position. The lateral dissymmetry of

lift which results causes the rotor to tilt for-

ward by the gyroscopic effect.

A differential blade pitch change like this

is called a cyclic.pitch change since each blade

goes thr0ugh.a complete cycle of varying pitch

angles as it completes one revolution of rota-

tion about the hub. A cyclic pitch change is

accomplished by the pilot by the use of the

cyclic stick. The control arrangement is such

that the rotor tilts in the same direction that

the cyclic stick is deflected.

A variation in rotor thrust is accomplished

by increasing>sthe pitch of the blades simul-

taneously or collectively. This type of control

action is called “collective pitch” and is ac-

complished by the use of the collective pitch

stick. In operation, the cyclic stick is an-

alogous to the control stick of an airplane,

and the collective stick is analogous to the

throttle ~of an ,airplane.

There are several possibilities for longi-

tudinal control of a tandem-rotor helicopter.

A pitching moment can be produced by tilting

both rotors by a cyclic pitch change in each

rotor, by a differential collective pitch change

that increases the thrust on one rotor and de-

creases it on the other, or by some combination

of these methods. The two basic methods are

illustrated in figure 4.36. Obviously, a change

in fuselage attitude must accompany the dif-

ferential collective method of longitudinal

control.

NAVWEPS DO-80T-80

STABILITY AND CONTROL

Adequate pitch and lateral control effective-

ness are easy to obtain in the typical helicopter

and usually present no problems. The more

usual problem is an excess of control effective-

ness which results in an overly sensitive heli-

copter. The helicopter control specifications

attempt to assure satisfactory control charac-

teristics by requiring adequate margins of con-

trol travel and effectiveness without objection-

able sensitivity.

Directional control in a single rotor heli-

copter is obtained by a tail rotor (antitorque

rotor) since a conventional aerodynamic sur-

face would not be effective at low speeds or

hovering. The directional control require-

ments of the tail rotor on a typical shaft-driven

helicopter are quite demanding since it must

counteract the engine torque being supplied to

the main rotor as well as provide directional

control. Being a rotor in every respect, the

tail rotor requires some of the engine power to

generate its control forces. Unfortunately, the

maximum demands of the tail rotor occur at

conditions when engine power is also in great

demand. The most critical condition is while

hovering at maximum gross weight. The tail

rotor effectiveness is determined by the rotor

characteristics and the distance the tail rotor

is behind the c.g. The control specifications

require the helicopter to be able to turn in the

most critical direction at some specified rate

while hovering at maximum gross weight in a

specified wind condition. Also, it is required

that the helicopter have sufficient directional

control to fly sideways up to 30 knots, an

important requirement for plane guard duties.

The directional control requirements are

easily met by a tip-driven helicopter since the

directional control does not have to counter

the engine torque.

Directional control of a tandem-rotor heli-

copter is accomplished by differential cyclic

control of the main rotors. For a pedal turn

to the starboard, the forward rotor is tilted

to the starboard and the rear rotor is tilted to

port, creating a turning moment as shown in

NAVWEPS DD-80T-80

STABILITY AND CONTROL

TANDEM ROTOR LONGITUDINAL CONTROL

TANDEM ROTOR DIRECTIONAL CONTROL

AFT

ROTOR

*JR

F”&%iD

Fig&e 4.36. longitudinal and Directional Control

figure 4.36. The directional control require-

ments are easily met in a tandem-rotor heli-

copter because the engine torque from one

rotor is opposed by the torque of the other

rotor thereby eliminating one directional mo-

ment. Of course, some net unbalance of torque

may have to be overcome if the engine torque

on the two rotors is different.

When a tandem-rotor helicopter is rotated

rapidly about one of the rotors rather than

about the cg., the other rotor picks up

“translational lift” as a result of the velocity

due to rotation and an increase in rotor thrust

results. This causes pitch-up or pitch-down

depending on which rotor the helicopter is

being rotated about. Rotation about the

forward rotor, which is more common, re-

sults in pitch-down.

The overall stability of a helicopter results

from the individual stability contributions of

the various components just as in the case of

the fixed-wing airplane. The stability con-

tributions can be divided as follows:

(1) Rotor

(2) Fuselage

(3) Stabilizers

(4) Mechanical devices

The destabilizing contribution of the fuselage

and the stabilizing contribution of a stabilizing

surface are similar in effect to an airplane and

will not be discussed here. The principal

stability characteristics that make the heli-

copter different from an airplane are those of

the rotor.

Two types of stability are important in the

rotor: (1) angle of attack stability and (2)

velocity stability. In hovering flight the

relative wind velocity, angle of attack, and

lift on each blade of the rotor is the same. If

the rotor is displaced through some angle, no

changes in forces result. Therefore, the rotor

has neutral angle of attack stability when

hovering. However, in forward flight, an

increase in rotor angle of attack increases the

lift on the advancing blade more than on the

NAVWEPS 00-801-80

STABILITY AND CONTROL

retreating blade since the relative wind veloci-

ties are greater on the advancing blade. This

lateral dissymmetry of lift causes the rotor to

tilt back due to the gyroscopic effect of the

rotor, further increasing the rotor angle of

attack. Thus, the rotor is unstable with

changes in angle of attack at forward flight

speeds. Since the magnitude of the unstable

moment is affected by the magnitude of

the rotor thrust as well as the tilt of

the thrust force, a greater instability exists

for increases in angle of attack than for

decreases in angle of attack. In addition, the

instability is greater for increases in angle of

attack when the rotor thrust also increases.

If the rotor angle of attack is held constant

and the rotor is given a translational velocity,

a dissymmetry of lift results since the velocity

of the advancing blade is increased while the

velocity of the retreating blade is decreased.

This dissymmetry of lift causes the rotor to

tilt in a direction to oppose the change in

velocity due to the gyroscopic effect of the

rotor. Hence, the rotor has velocity stability.

A hovering helicopter exhibits some degree

of apparent stability by virtue of its velocity

stability although it has neutral angle of

attack stability. This type of hovering sta-

bility is analogous to the apparent lateral-

directional stability an airplane exhibits due

to dihedral effect. Additional hovering sta-

bility can be obtained by the use of mechanical

stabilizers such as th,e Bell stabilizer bar, by

the use of offset flapping hinges, or by syn-

thetic or artificial stabilization devices.

The total static stability of a helicopter is

determined by combining the stability con-

tributions of all the components. The usual

result for a typical helicopter is instability

with angle of attack and a variable velocity

stability which becomes neutral or unstable

at high speeds. Of course, the helicopter

could be made stable with angle of attack by

providing a large enough horizontal stabilizer.

Unfortunately, adverse effects at low speed or

NAVWEPS 00-801-80

STABILITY AND CONTROL

hovering and large trim moments upon entering

autorotation will limit the stabilizer size to

a relatively small surface. Usually the hori-

zontal stabilizer is used only to give the fuse-

lage the desired moment characteristics.

The angle of attack stability of a tandem-

rotor helicopter is adversely affected by the

downwash from the forward rotor reducing

the angle of attack and thrust of the rear

rotor. This reduction of thrust behind the

cg. causes the helicopter to pitch up to a

higher angle of attack, thereby adding to the

angle of attack instability.

As in the airplane, several oscillatory modes

of motion are characteristic of the dynamic

stability of a helicopter. The phugoid is the

most troublesome for the helicopter. The

phugoid mode is unstable in the majority of

helicopters which operate without the assist-

ance of artificial stabilization devices. The

dynamic instability of the helicopter is given

evidence by the flying qualities specification for

helicopters. These specifications essentially

limit the rate of divergence of the dynamic oscil-

lations for the ordinary helicopter. Although

this dynamic instability can be controlled, it

requires constant attention by the pilot and

results in pilot fatigue. The elimination of

the dynamic instability would contribute

greatly to improving the flying qualities of

the helicopter.

This dynamic instability characteristic is

particularly important if the helicopter is

expected to be used for instrument flight in

all-weather operations. In fact, a seriously

divergent phugoid mode would make instru-

ment flight impractical. For this reason, the

flying qualities specification requires that

helicopters with an instrument capability

exhibit varying degrees of stability or insta-

bility depending on the period of the oscilla-

tion. Long period oscillations (over 20 sec-

onds) must not double in amplitude in less

than 15 seconds whereas short period oscil-

lations (under 10 seconds) must damp to half

amplitude in two cycles.

The only immediate solution for the dynamic

instability is an attitude stabilization system

which is essentially an autopilot. Other

solutions to the dynamic instability problem

involve mechanical, aerodynamic, or elec-

tronic control feedback of pitch attitude,

pitch velocity, normal acceleration, or angle

of attack. The improvement of the heli-

copter’s stability is mandatory to fully utilize

its unique capability. As more of the heli-

copter problems are analyzed and studied, the

flying qualities of helicopters wiI1 improve

and be comparable to the fixed wing aircraft.

NAVWEPS 00-801-80

OPERATING STRENGTH LIMITATIONS

Chapter 5

OPERATING STRENGTH

LIMITATIONS

The weight of the structural components of

an aircraft is an extremely important factor in

the development of an efficient aircraft con-

figuration. In no other field of mechanical

design is there such necessary importance

assigned to structural weight. The efficient

aircraft and powerplant structure is the zenith

order to obtain the required service life from

his aircraft, the Naval Aviator must undet-

stand, appreciate, and observe the operating

strength limitations. Failure to do so will

incur excessive maintenance costs and a high

incidence of failure during the service life of

of highly reined rknimum weight design. in an aircraft.

NAVWEPS oo-EOT-80

OPERATING STRENGTH LIMITATIONS

GENERAL DEFINITIONS AND STRUC-

TURAL REQUIREMENTS

There are strength requirements which ate

common to all aircraft. In general, these re-

quirements can be separated into three particu-

lar areas. These are detailed in the following

discussion.

STATIC STRENGTH

The static strength requirement is the con-

sideration given to the effect of simple static

loads with none of the ramifications of the

repetition or cyclic variation of loads. An

important reference point in the static strength

requirement is the “limit load” condition.

When the aircraft is at the design conligura-

tion, there will be some maximum of load

which would be anticipated from the mission

requirement of the airplane. For example, a

fighter or attack type aircraft, at the design

configuration, may encounter a very peak load

factor of 7.5 in the accomplishment of its mis-

sion. Of course, such an aircraft may be sub-

ject to load factors of 3, 4, 5, 6, 1, etc., but no

more than 7.5 should be required to accom-

plish the mission. Thus, the limit load condi-

tion is the maximum of loads anticipated in

normal operation of the aircraft, Various

types of aircraft will have different limit load

factors according to the primary mission of

the aircraft. Typical values are tabulated

below:

Type of aircraft: hbi”< limi, hi,orror

Fighter or attack. 7.5

Trainer. 7.5

T ransport, patrol, antisubmarine. 3.0 or 2.5

Of course, these examples are quite general and

it is important to note that there may be varia-

tions according to specific mission require-

ments.

Since the limit load is the maximum of the

normally anticipated loads, the aircraft struc-

ture must withstand this load with no ill

effects. Specilicallv, the primary structure of

the aircraft should experience no objectionable

permanent deformation when subjected to the

limit load. In fact, the components must with-

stand this load with a positive margin. This

requirement implies that the aircraft should

withstand successfully the limit load and then

return to the original unstressed shape when

the load is removed. Obviously, if the air-

craft is subjected to some load which is in

excess of the limit load, the overstress may

incur an objectionable permanent deformation

of the primary structure and require replace-

ment of the damaged parts.

Many different flight and ground load condi-

tions must be considered to define the most

critical conditions for the structural com-

ponents. In addition to positive lift flight,

negative lift flight must be considered. Also,

the effect of flap and landing gear configura-

tion, gross weight, flight Mach,number, sym-

metry of loading, c.g. positions, etc., must be

studied to account for all possible sources of

critical loads. To verify the capability of the

structure, ground static tests are conducted

and flight demonstrations ate required.

To provide for the rare instances of flight

when a load greater than the limit is required

to prevent a disaster, an “ultimate factor of

safety” is provided. Experience has shown

that an ultimate factor of safety of 1.5 is suf-

ficient for piloted aircraft. Thus, the aircraft

must be capable of withstanding a load which

is 1.3 times the design limit load. The primary

structure of the aircraft must withstand the

“ultimate load” (1.5 times limit) without

failure. Of course, permanent deformation

may be expected with this “overstress” but

no actual failure of the major load-carrying

components should take place at ultimate load

Ground static tests are necessary to verify this

capability of the structure.

An appreciation of the static strength re-

quirements may be obtained by inspection of

the basic properties of a typical aircraft metal.

Figure 3.1 illustrates the typical static strength

properties of a metal sample by a plot of applied

stress versus resulting strain. At low values

3,26

NAVWEPS 00-EOT-80

OPERATING STRENGTH LIMITATIONS

STATIC STRENGTH OF TYPICAL AIRCRAFT METAL

ULTIMATE

STRENGTH -Cc

CYCLIC

STRESS

(PSI)

STRESSES APPLIED ABOVE

THIS POINT RESULT IN

OBJECTIONABLE PERMANENT

DEFORMATION

FAILURE

-I STRAIN

PERMANENT

(IN/IN)

SET II:

-I I--

FATIGUE STRENGTH OF TYPICAL AIRCRAFT METAL

HIGH CYCLIC STRESS

VERY FEW CYCLES REQUIRED

TO CAUSE FAILURE

MODERATE CYCLIC STRESS

RELATIVELY LARGE NUMBER OF

APPLICATIONS NECESSARY TO

CAUSE FAILURE LOW CYCLIC STRESS

ALMOST INFINITE CYCLES

TO CREATE FATIGUE FAILURE

NUMBER OF APPLICATIONS

TO CAUSE FATIGUE FAILURE

Egu,e 5.1. Strength Chomctorirfics

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OPERATING STRENGTH LIMITATIONS

of stress the plot of stress and strain is essen-

tially a straight line, i.e., the material in this

range is elastic. A stress applied in this range

incurs no permanent deformation and the ma-

terial returns to the original unstressed shape

when the stress is released. At higher values

of stress the plot of stress versus strain develops

a distinct curvature in the strain direction and

the material incurs disproportionate strains.

High levels of stress applied co the part and

then released produce a permanent deforma-

tion. Upon release of some high stress, the

metal snaps back-but not all the way. The

stress defining the limit of tolerable permanent

strain is the “yield stress” and stresses applied

above this point produce objectionable per-

manent deformation. The very highest stress

the material can withstand is the “ultimate

stress.” Noticeable permanent deformation

usually occurs in this range, but the material

does have the capability for withstanding one

application of the ultimate stress.

The relationship between the stress-strain

diagram and operating strength limits should

be obvious. If the aircraft is subjected to a

load greater than the limit, the yield stress

may be exceeded and objectionable permanent

deformation may result. If the aircraft is

subject to a load greater than the ultimate,

failure is imminent.

SERVICE LIFE

The various components of the aircraft and

powerplant structure must be capable of oper-

ating without failure or excessive deformation

throughout the intended service life. The

repetition of various service loads can produce

fatigue damage in the structure and special

attention must be given to prevent fatigue

failure within the service life, Also, the sus-

taining of various service loads can produce

creep damage and special attention must ‘be

given to prevent excessive deformation or

creep failure within the service life, This is

a particular feature of components which are

subjected to operation at high temperatures.

FATIGUE CONSIDERATIONS. The fa-

tigue strength requirement is the considera-

tion given the cumulative effect of repeated

or cyclic !oads during service. While there is

a vague relationship with the static strength,

repeated cyclic loads produce a completely

separate effect. If a cyclic, tensile stress is

applied to a metal sample, the part is subject

to a “fatigue” type loading. After a period

of time, the cyclic stressing will produce a

minute crack at some critical location in the

sample. With continued application of the

varying stress, the crack will enlarge and

propagate into the cross section. When the

crack has progressed sufficiently, the remaining

cross section is incapable of withstanding the

imposed stress and a sudden, final rupture

occurs. In this fashion, a metal can be failed

at stresses much lower than the static ultimate

strength.

Of course, the time necessary to produce

fatigue failure is related to the magnitude of

the cyclic stress. This relationship is typified

by the graph of figure 5.1. The fatigue

strength of a material can be demonstrated by

a plot of cyclic stress versus cycles of stress

required to produce fatigue failure. As might

be expected, a very high stress level requires

relatively few cycles to produce fatigue failure.

Moderate stress levels require a fairly large

number of cycles to produce failure and a very

low stress may require nearly an infinite num-

ber of cycles to produce failure. The very

certain implication is that the aircraft must

be capable of withstanding the gamut of

service loads without producing fatigue failure

of the primary structure.

For each mission type of aircraft there is

a probable spectrum of loads which the air-

craft will encounter. That is, various loads

will be encountered with a frequency particular

to the mission profile. The fighter or attack

type of aircraft usually experiences a pre-

dominance of maneuver loads while the trans-

port or patrol type usually encounters a pre-

dominance of gust loads. Since fatigue damage

SNOIlVlIWll HlOM3US ONllVU3dO

08-108-00 Sd3MAVN

NAVWEPS 00-8OT-80

OPERATING STRENGTH LIMITATIONS

is cti~n&zti~e during cyclic stressing, the useful

service life of the aircraft must be anticipated

to predict the gross effect of service loads.

Then, the primary structure is required to

sustain the typical load spectrum rhrough the

anticipated service life without the occurrence

of fatigue failure. To prove this capability

of the structure, various major components

must be subjected to an accelerated fatigue

test to verify the resistance to repeated loads.

The design of a highly stressed or long life

structure emphasizes the problems of fatigue.

Great care must be taken during design and

manufacture to minimize stress concentrations

which enhance fatigue. When the aircraft

enters service operation, care must be taken in

the maintenance of components to insure proper

adjustment, torquing, inspection, etc., as proper

maintenance is a necessity for achieving full

service life. Also, the structure must not be

subjected to a load spectrum more severe than

was considered in design or fatigue failures

may occur within the anticipated service life.

With this additional factor in mind, any pilot

should have all the more respect for the oper-

ating strength limits-recurring overstress

causes a high rate of fatigue damage.

There are many examples of the detrimental

effect of repeated overstress on service life.

One major automobile manufacturer adver-

tised his product as “guaranteed to provide

100,000 miles of normal driving without me-

chanical failure.” The little old lady from

Pasadena-the original owner of ALL used cars

-will probably best the guaranteed mileage

by many times. On the other hand, the hot-

rod artist and freeway Grand Prix contender

do not qualify for the guarantee since their

manner of operation could not be considered

normal. The typical modern automobile may

be capable of 60,000 to l~,OOO miles of normal

operation before an overhaul is necessary.

However, this same automobile may encounter

catastrophic failures in a few hundred miles if

operated continually at maximum torque in

low drive range. Obviously, there are similar

relationships for aircraft and powerplant

structures.

CREEP CONSIDERATIONS. By definition,

creep is the structural deformation which oc-

curs as a function of time. If a part is subjected

to a constant stress of sufficient magnitude, the

part will continue to develop plastic strain and

deform with time. Eventually, failure can

occur from the accumulation of creep damage.

Creep conditions are most critical at high

stress and high temperature since both factors

increase the rate of creep damage. Of course,

any structure subject to creep conditions should

not encounter excessive deformation or failure

within the anticipated service life.

The high operating temperatures of gas tur-

bine components furnish a critical environment

for creep conditions. The normal operating

temperatures and stresses of gas turbine com-

ponents create considerable problems in design

for service life. Thus, operating limitations

deserve very serious respect since excessive

engine speed or excessive turbine temperatures

will cause a large increase in the rate of creep

damage and lead to premature failure of com-

ponents. Gas turbines require high operating

temperatures to achieve high performance and

efficiency and short periods of excessive tem-

peratures can incur highly damaging creep

rates.

Airplane structures can be subject to high

temperatures due to aerodynamic heating at

high Mach numbers. Thus, very high speed

airplanes can be subject to operating limita-

tions due to creep conditions.

AFROELASTIC EFFECTS

The requirement for structural stiffness and

rigidity is the consideration given to the inter-

action of aerodynamic forces and deflections of

the structure. The aircraft and its components

must have sufficient stiffness to prevent or

minimize aeroelastic influences in the normal

flight range, Aileron reversal, divergence,

flutter, and vibration should not occur in the

range of flight speeds which will be normal

operation for the aircraft.

It is important to distinguish between

strength and stiffness. Strength is simply the

resistance to load while stiffness is the resist-

ance to deflection or deformation. While

strength and stiffness are related, it is necessary

to appreciate that adequate structural strength

does not automatically provide adequate stiff-

ness. Thus, special consideration is necessary

to provide the structural components with

specific stiffness characteristics to prevent un-

desirable aeroelastic effects during normal

operation.

An obvious solution to the apparent prob-

lems of static strength, fatigue strength,

stiffness and rigidity would be to build the

airplane like a product of an anvil works,

capable of withstanding all conceivable loads.

However, high performance airplane con-

figurations cannot be developed with inefi-

cient, lowly stressed structures. The effect of

additional weight is best illustrated by pre-

liminary design studies of a very long range,

high altitude bomber. In the preliminary

phases of design, each additional pound of

any weight would necessitate a 25-pound

increase in gross weight to maintain the same

performance. An increase in the weight of

any item produced a chain reaction-more

fuel, larger tanks, bigger engines, more fuel,

heavier landing gear, more fuel, etc. In the

competitive sense of design, no additional

structural weight can be tolerated to provide

more strength than is specified as necessary

for the design mission requirement.

AIRCRAFT LOADS AND OPERATING

LIMITATIONS

FLIGHT LOADS-MANEUVERS AND

GUSTS

The loads imposed on an aircraft in flight

are the result of maneuvers and gusts. The

maneuver loads may predominate in the

design of fighter airplanes while gust loads

may predominate in the design of the large

multiengine aircraft. The maneuver loads an

NAVWEPS 00-EOT-80

OPERATING STRENGTH LIMITATIONS

airplane may encounter depend in great part

on the mission type of the airplane. However,

the maximum maneuvering capability is of

interest because of the relationship with

strength limits.

The flight load factor is defined as the pro-

portion between airplane lift and weight,

where

n=L/W

n= load factor

L=lift, Ibs.

W= weight, Ibs.

MANEUVERING LOAD FACTORS. The

maximum lift attainable at any airspeed occurs

when the airplane is at CLmU. With the use

of the basic lift equation, this maximum lift

is expressed as:

Since maximum lift must be equal to the

weight at the stall speed,

If the effects of compressibility and viscosity

on Ch are neglected for simplification, the

maximum load factor attainable is determined

by the following relationship.

v.2

=(-) V*

Thus, if the airplane is flying at twice the

stall speed and the angle of attack is increased

to obtain maximum lift, a maximum load

factor of four will result. At three times the

stall speed, nine “g’s” would result; four

times the stall speed, sixteen g’s result; five

times the stall speed, twenty-five g’s result;

etc. Therefore, any airplane which has high

speed performance may have the capability of

high maneuvering load factors. The airplane

which is capable of flight speeds that are

NAVWEPS 00-801-80

OPERATING STRENGTH LIMITATIONS

many times the stall speed will require due

consideration of the operating strength limits.

The structural design of the aircraft must

consider the possibility of negative load factors

from maneuvers. Since the pilot cannot com-

fortably tolerate large prolonged negative “g”,

the aircraft need not be designed for negative

load factors as great as the positive load factors.

The effect of airplane gross weight during

maneuvers must be appreciated because of the

particular relation to flight operating strength

limitations. During flight, the pilot appre-

ciates the degree of a maneuver from the

inertia forces produced by various load factors;

the airplane structure senses the degree of a

maneuver principally by the airloads involved.

Thus, the pilot recognizes loadfactor while the

structure recognizes only load. To better

understand this relationship, consider an ex-

ample airplane whose basic configuration gross

weight is 20,000 lbs. At this basic configura-

tion assume a limit load factor for symmetrical

flight of 5.6 and an ultimate load factor of 8.4.

If the airplane is operated at any other con-

figuration, the load factor limits will be al-

tered. The following data illustrate this fact

by tabulating the load factors required to

produce identical airloads at various gross

weights.

Grass weight, Ibs. Limit load Ultimate

factor load factor

20,wO (basic). 5.60 8.40

30,003 (max. rakcoff). 3.73 5.60

13,333 (min. f”cl):. 8.40 12.60

As illustrated, at high gross weights above the

basic configuration weight, the limit and ulti-

mate load factors may be seriously reduced.

For the airplane shown, a 5-g maneuver im-

mediately after a high gross weight takeoff

could be very near the “disaster regime,”

especially if turbulence is associated with the

maneuver. In the same sense, this airplane

at very low operating weights below that of

the basic configuration would experience great-

ly increased limit and ultimate load factors.

Operation in this region of high load factors

at .low gross weight may create the impression

that the airplane has great excess strength

capability. This effect must be understood and

intelligently appreciated since it is not uncom-

mon to have a modern airplane configuration

with more than SO percent of its gross weight

as fuel.

GUST LOAD FACTORS. Gusts are asso-

ciated with the vertical and horizontal velocity

gradients in the atmosphere. A horizontal

gust produces a change in dynamic pressure on

the airplane but causes relatively small and

unimportant changes in flight load factor.

The more important gusts are the vertical gusts

which cause changes in angle of attack. This

process is illustrated in figure 5.2. The vec-

torial addition of the gust velocity to the air-

plane velocity causes the change in angle of

attack and change in lift. The change in angle

of attack at some flight condition causes a

change in the flight load factor. The incre-

ment change in load factor due to the vertical

gust can be determined from the following

equation:

where

An=change in load factor due to gust

m=lift curve slope, unit of C, per degree

of 01

o=altitude density ratio

W/S= wing loading, psf

V. = equivalent airspeed, knots

KU=equivalent sharp edged gust velocity

ft. per sec.

As an example, consider the case of an air-

plane with a lift curve slope m=O.OB and wing

loading, (W/S)=60 psf. If this airplane were

flying at sea level at 350 knots and encountered

an effective gust of 30 ft. per sec., the gust

would produce a load factor increment of 1.61.

This increment would be added to the flight

load factor of the airplane prior to the gust,

NAVWEPS OO-80T-80

OPERATING STRENGTH LIMITATIONS

CHANGE IN LIFT

AIRPLANE VELOCITY, V

GUST

VELOCITY RESULTANT VELOCITY

KU

figure 5.2. Effect of Vertical Gust

e.g., if in level flight before encountering the

gust, a final load factor of 1.0+1.61=2.61

would result. As a general requirement all

airplanes must be capable of withstanding an

approximate effective f30 ft. per sec. gust

when at maximum level flight speed for normal

rated power. Such a gust intensity has rela-

tively low frequency of occurrence in ordinary

flying operations.

The equation for gust load increment pro-

vides a basis for appreciating many of the

variables of flight. The gust load increment

varies directly with the equivalent sharp

edged gust velocity, KU, since this factor

effects the change in angle of attack.’ The

highest reasonable gust velocity that may be

anticipated is an actual vertical velocity, U,

of 50 ft. per sec. This value is tempered by

the fact that the airplane does not effectively

encounter the full effect because of the response

of the airplane and the gradient of the gust.

A gust factor, K (usually on the order of 0.6),

reduces the actual gust to the equivalent sharp

edged gust velocity, KU.

The properties of the airplane exert a power-

ful influence on the gust increment. The lift

curve slope, m, relates the sensitivity of the

airplane to changes in angle of attack. An

aircraft with a straight, high aspect ratio

wing would have a high lift curve slope and

would be quite sensitive to gusts. On the

other hand, the low aspect ratio, swept wing

airplane has a low lift curve slope and is com-

paratively less sensitive to turbulence. The

apparent effect of wing loading, W/S, is at

times misleading and is best understood by

considering a particular airplane encountering

a fixed gust condition at various gross weights.

If the airplane encounters the gust at lower

than ordinary gross weight, the accelerations

NAVWEPS 00-ROT-80

OPERATING STRENGTH LIMITATIONS

due to the gust condition are higher. This is

explained by the fact that essentially the

same lift change acts on the lighter mass.

The high accelerations and inertia forces

magnify the impression of the magnitude of

turbulence. If this same airplane encounters

the gust condition at higher than ordinary

gross weight, the accelerations due to the gust

condition are lower, i.e., the same lift change

acts on the gteatet mass. Since the pilot

primarily senses the degree of turbulence by

the resulting accelerations and inertia forces,

this effect can produce a very misleading

impression.

The effect of airspeed and altitude on the

gust load factor is important from the stand-

point of flying operations. The effect of alti-

tude is related by the term &, which would

related that an airplane flying at a given EAS

at 40,000 ft. (c=O.25) would experience a

gust load factor increment only one-half as

great as at sea level. This effect results be-

cause the true airspeed is twice as great and

only one-half the change in angle of attack

occurs for a given gust velocity. The effect of

airspeed is illustrated by the linear variation

of gust increment with equivalent airspeed.

Such a variation emphasizes the effect of gusts

at high flight speeds and the probability of

structural damage at excessive speeds in turbu-

lence.

The operation of any aircraft is subject to

specific operating strength limitations. A

single large overstress may cause structural

failure or damage severe enough to require

costly overhaul. Less severe overstress re-

peated for sufficient time will cause fatigue

cracking and require replacement of parts to

prevent subsequent failure. A combat airplane

need not be operated in a manner like the “little

old lady from Pasadena” driving to church on

Sunday but each aircraft type has strength

capability only specific to the mission require-

ment. Operating limitations must be given

due regard.

THE V-B OR V-g DIAGRAM

The operating flight strength limitations of

an airplane are presented in the form of a

V-‘-n or V-g diagram. This chart usually is

included in the aircraft flight handbook in the

section dealing with operating limitations.

A typical V-n diagram is shown in figure 5.3.

The V-n diagram presented in figure 5.3 is

intended to present the most important general

features of such a diagram and does not neces-

sarily represent the characteristics of any par-

ticular airplane. Each airplane type has its

own particular V-n diagram with specific V’s

and n’s,

The flight operating strength of an airplane

is presented on a graph whose horizontal scale

is airspeed (V) and vertical scale is load factor

(n). The presentation of the airplane strength

is contingent on four factors being known:

(I) the aircraft gross weight, (2) the configura-

tion of the aircraft (clean, external stores, flaps

and landing gear position, etc.), (3) symmetry

of loading (since a rolling pullout at high speed

can reduce the structural limits to approxi-

mately two-thirds of the symmetrical load

limits) and (4) the applicable altitude. A

change in any one of these four factors can

cause important changes in operating limits.

For the airplane shown, the positive limit

load factor is 7.5 and the, positive ultimate

load factor is Il.25 (7.5x1.5)- For negative

lift flight conditions the negative’limit load

factor is 3.0 and the negative ultimate load

factor is 4.5 (3.0x1.5). The limrt airspeed is

stated as 575 knots while the wing level stall

speed is apparently 100 knots.

Figure 5.4 provides supplementary informa-

tion to illustrate the significance of the V-n

diagram of figure 5.3. The lines of maximum

lift capa’bility are the first points of importance

on the’ V-n diagram. The subject aircraft is

capable bf developing no more than one posi-

tive “g” at 100 knots, the wing level stall speed

of the airplane. Since the maximum load

faztor varies with the square of the aitspeed,

LOAD

FACTOR,

IO-

I-.

-o-.

-I-

-2-

-3-

-4-

-5-

GROSS WEIGHT - 16.000 LBS

CLEAN CONFIGURATION

SEA LEVEL ALTITUDE

SYMMETRICAL LOADING

I ,.,/,,.~A~~‘~,, FACTOR i

~/POSlT,VE LlMlT LOAD FACTOR i

LIMIT LIMIT

AIRSPEED AIRSPEED

575 575 KNOTS KNOTS

INDICATED INDICATED AIRSPEED - KNOTS AIRSPEED - KNOTS

200 300 300 400 400 500 500 I 600 600

NEGATIVE LIMIT LOAO FACTOR

STALL

NEGATIVE ULTIMATE LOAD FACTOR

\

Figure 5.3. Flight Strength Diagram

STRUCTURAL FAILURE

IO-

7.5

7- UP

6- !

4- MAXIMUM

IN,,lCATF . . I -. . - - - -

NEGATIVE LIFT

CAPABILITY

Figure 5.4. Signikance o\ the V-n Diagram

the maximum positive lift capability of this

airplane is 4 “g” at 200 knots, 9 g at 300 knots,

16 g at 400 knots, etc. Any load factor above

this line is unavailable aerodynamically, i.e.,

the subject airplane cannot fly above the line of

maximum lift capability. Essentially the same

situation exists for negative lift flight with the

exception that the speed necessary to produce a

given negative load factor is higher than that

to produce the same positive load factor. Gen-

erally, the negative CL,., is less than the posi-

tive CL,., and the airplane may lack sufficient

control power to maneuver in this direction.

If the subject airplane is flown at a positive

load factor greater than the positive limit load’

factor of 7.5, structural damage will be possi-

ble. When the airplane is operated in this

region, objectionable permanent deformation

of the primary structure may take place and a

high rate of fatigue damage is incurred. Opera-

tion above the limit load factor must be

avoided in normal operation. If conditions of

extreme emergency require load factors above

the limit to prevent an immediate disaster, the

airplane should be capable of withstanding the

ultimate load factor without failure. The

same situation exists in negative lift flight

with the exception that the limit and ultimate

load factors are of smaller magnitude and the

negative limit load factor may not be the same

value at all airspeeds. At speeds above the

maximum level flight airspeed the negative

limit load factor may be of smaller magnitude.

The limit airspeed (or redline speed) is a de-

sign reference point for the airplane-the sub-

ject airplane is limited to 575 knots. If flight

is attempted beyond the limit airspeed struc-

tural,idamage or structural failure may result

from a variety of phenomena. The airplane in

flight above the limit airspeed may encounter:

(u) critical gust

(6) destructive flutter

(c) aileron reversal

(d) wing or surface divergence

(e) critical compressibility effects such as

stability and control problems,

damaging buffet, etc.

NAVWEPS 00-SOT-80

OPERATING STRENGTH LIMITATIONS

The occurrence of any one of these items could

cause structural damage or failure of the pri-

mary structure. A reasonable accounting of

these items is required during the design of an

airplane to prevent such occurrences in the re-

quired operating regions. The limit airspeed

of an airplane may be any value between termi-

nal dive speedand 1.2 times the maximum level

flight speed,depending on the aircraft type and

mission requirement. Whatever the resulting

limit airspeed happens to be, it deserves due

respect.

Thus, the airplane in flight is limited to a

regime of airspeeds and g’s which do not

exceed the limit (or redline) speed, do not

exceed the limit load factor, and cannot exceed

the maximum lift capability. The airplane

must be operated within this “envelope”

to prevent structural damage and ensure that

the anticipated service life of the airplane is

obtained. The pilot must appreciate the

V-n diagram as describing the allowable

combination of airspeeds and load factors for

safe operation. Any maneuver, gust, or gust

plus maneuver outside the structural envelope

can cause sttuctural damage and effectively

shorten the service.life, of the airplane.

There are two points of great importance on

the V-n diagram of figure 5.4. Point B is

the intersection of the negative limit load

factor and line of maximum negative lift

capability. Any airspeed greater than point

B provides a negative lift capability sufficient

to damage the airplane; any airspeed less

than point B does not provide negative lift

capability sufficient to damage the airplane

from excessive flight loads. Point A is the

intersection of the positive limit load factor

and the line of maximum, positive lift capa-

bility. The airspeed at this point is the

minimum airspeed at which the limit load

can be developed aerodynamically. Any air-

speed greater than pomt A provides a positive

lift capability sufficient to damage the air-

plane; any airspeed less than point A does

not provide Positive lift capability sufficient to

cause damage from excessive flight loads. The

usual term given to the speed at point A is the

“maneuver speed,” since consideration of

subsonic aerodynamics wouId predict mini-

mum usable turn radius to occur at this con-

dition. The maneuver speed is a valuable

reference point since an airplane operating

below this point cannot produce a damaging

positive flight load. Any combination of

maneuver and gust cannot create damage due

to excess airload when the airplane is below

the maneuver speed.

The maneuver speed can be computed from

the following equation:

where

VP= maneuver speed

V,= stall speed

n limit = limit load factor

Of course, the stall speed and limit load factor

must be appropriate for the airplane gross

weight. One notable fact is that this speed,

once properly computed, remains a constant

value if no significant change takes place in

the spanwise weight distribution. The ma-

neuver speed of the subject aircraft of figure

5.4. would be

v,= loo&3

= 274 knots

EFFECT OF HIGH SPEED FLIGHT

Many different factors may be of structural

importance in high speed flight. Any one or

combination of these factors may be encount-

ered if the airplane is operated beyond the

limit (or redline) airspeed.

At speeds beyond the limit speed the air-

plane may encounter a critical gust. This is

especially true of a high aspect ratio airplane

with a low limit load factor. Of course, this

NAVWEPS 00-801-80

OPERATING STRENGTH LIMITATIONS

is also an important consideration for an air-

plane with a high limit load factor if the gust

should be superimposed 00 a maneuver. Since

the gust Ioad factor increment varies directly

with airspeed and gust intensity, high airspeeds

must be avoided in turbulent conditions.

When it is impossible to avoid turbulent

conditions and the airplane must be subject to

gusts, the flight condition must be properly

controlled to minimize the effect of turbulence.

If possible, the airplane airspeed and power

should be adjusted prior to entry into turbu-

lence to provide a stabilized attitude. Ob-

viously, penetration of turbulence should not

be accomplished at an excess airspeed because

of possible structural damage. On the other

hand, an excessively low speed should not be

chosen to penetrate turbulence for the gusts

may cause stalling of the aircraft and difficulty

of control. To select a proper penetration

airspeed the speed should not be excessively

high or ‘low-the two extremes must be

tempered. The “maneuver” speed is an im-

portant reference point since it is the highest

speed that can be taken to alleviate stall due

to gust and the lowest speed at which limit

load factor can be develoPed aerodynamically.

The optimum penetration speed occurs at or

very near the maneuver speed.

Aileron rever& is a phenomenon particular

to high speed flight. When in flight at very

high dynamic pressures, the wing torsional

deflections which occur with aileron deflection

are considerable and cause noticeable change

in aileron effectiveness. The deflection of an

aileron on a rigid wing creates a change in lift

and produces a rolling moment. In addition

the deflection of the control surface creates a

twisting moment on the wing. When the

actual elastic wing is subject to this condition

at high dynamic pressures, the twisting mo-

ment produces measurable twisting deforma-

tions which affect the rolling performance of

the aircraft. Figure 5.5 illustrates this process

and the effect of airspeed on aileron effective-

ness. At some high dynamic pressure, the

NAVWEPS OO-SOT-80

OPERATING STRENGTH LIMITATIONS

-

RIGID WING ELASTIC WING

AILERON

EFFECTIVENESS 1.0

AILERON

C, ELASTIC REVERSAL

SPEED

C, RIGID

-0.

-3 w

EQUIVALENT AIRSPEED

DIVERGENCE

A+ LELASTIC

AXIS

Figure 5.5. Aeroelastic Effects (Sheet I of 2)

NAVWEPS 00-SOT-80

OPERATING STRENGTH LIMITATIONS

WING

ROOT’ /-TRAILING EDGE

9- LEADING EDGE

Figure 5.5. Aeroelastic Effects (Sheet 2 of 2)

NAVWEPS 00-ROT-80

OPERATING STRENGTH LIMITitTIONS

twisting deformation will be great enough to

nullify the effect on aileron deflection and the

aileron effectiveness will-be zero. Since speeds

above this point create rolling moments op-

posite to the direction controlled, this point

is termed the “aileron reversal speed.” Oper-

ation beyond the reversal speed would create

an obvious control difficulty. Also, the ex-

tremely large twisting moments which produce

loss of aileron effectiveness create large twist-

ing moments capable of structural damage.

In order to prevent loss of aileron effective-

ness at high airspeeds, the wing must have

high torsional stiffness. This may be a feature

difficult to accomplish in a wing of very thin

section and may favor the use of inboard ailer-

ons to reduce the twisted span length and

effectively increase torsional stiffness. The use

of spoilers for lateral control minimizes the

twisting moments and alleviates the reversal

problem.

Divergcm is another phenomenon common

to flight at high dynamic pressures. Like

aileron reversal, it is an effect due to the inter-

action of aerodynamic forces and elastic deflec-

tions of the structure. However, it differs

from aileron reversal in that it is a violent

instability which produces immediate failure.

Figure 5.5 illustrates the process of instability.

If the surface is above the divergence speed,

any disturbance precipitates this sequence.

Any change in lift takes place at the aerody-

namic center of the section. The change in

lift ahead of the elastic axis produces a twist-

ing moment and a consequent twisting deflec-

tion. The change in angle of attack creates

greater lift at the ac., greater twisting deflec-

tion, more lift, etc., until failure occurs.

At low flight speeds where the dynamic

pressure is low, the relationship between aero-

dynamic force buildup and torsional deflection

is ‘stable. However, the change in lift per

angle of attack is proportional to ‘vz but the

structural torsional stiffness of the wing re-

mains constant. This relationship implies

that at some high speed, the aerodynamic force

buildup may overpower the resisting torsional

stiffness and “divergence” will occur. The

divergence speed of the surfaces must be suf-

ficiently high that the airplane does not en-

counter this phenomenon within the normal

operating envelope. Sweepback, short span,

and high taper help raise the divergence speed.

F/titter involves aerodynamic forces, inertia

forces and the elastic properties of a surface.

The distribution of mass and stiffness in a

structure determine certain natural frequencies

and modes of vibration. If the structure is sub-

ject to a forcing frequency near these natural

frequencies, a resonant condition can result

with an unstable oscillation. The aircraft is

subject to many aerodynamic excitations while

in operation and the aerodynamic forces at

various speeds have characteristic properties

for rate of change of force and moment. The

aerodynamic forces may interact with the

structure in a fashion which may excite or

negatively damp the natural modes of the

structure and allow flutter. Flutter must not

occur within the normal flight operating en-

velope and the natural modes must be damped

if possible or designed to occur beyond the

limit speed. A’typical flutter mode is illus-

‘trated in figure 5.5.

Since the problem is one of high speed flight,

it is generally desirable to have ‘very high

natural frequencies and flutter speeds well

above the normal operating speeds. Any

change of stiffness or mass distribution will

alter the modes and frequencies and thus allow

a change in the flutter speeds. If the aircraft

is not properly maintained and excessive play

and flexibility exist, flutter could occur at

flight speeds below the limit airspeed.

Compres&ility pmblems may define the limit

airspeed for an airplane in terms of Mach num-

ber. The supersonic airplane may experience

a great decay of stability at some high Mach

number or encounter critical structural or

engine inlet temperatures due to aerodynamic

heating. The transonic airplane at an excessive

speed may encounter a variety of stability, con-

trol, or buffet problems associated with tran-

sonic flight. Since the equivalent airspeed for

a given Mach number decreases with altitude,

the magnitude of compressibility effects at

high altitude may be negligible for the tran-

sonic airplane. In this sense, the airplane may

not be able to fly at high enough dynamic

pressures within a certain range of Mach num-

bers to create any significant stability or

control problem.

The transonic airplane which is buffet lim-

ited requires due consideration of the effect of

load factor on the onset of buffet,. Since

critical Mach number decreases with lift coef-

ficient, the limit Mach number will decrease

with load factor. If the airplane is subject to

prolonged or repeated buffet for which it was

not designed, structural fatigue will be the

certain result.

The limit airspeed for each type aircraft is

set sufficiently high that full intended appli-

cation of the aircraft should be possible. Each

of the factors mentioned about the effect of

excess airspeed should provide due respect for

the limit airspeed.

LANDING AND GROUND LOADS

The most critical loads on the landing gear

occur at high gross weight and high rate of

descent at touchdown. Since the landing

gear has requirements of static strength and

fatigue strength similar to any other com-

ponent, overstress must be avoided to prevent

failure and derive the anticipated service life

from rhe components.

The most significant function of the landing

gear is to absorb the vertical energy of the air-

craft at touchdown. An aircraft at a given

weight and rate of descent at touchdown has

a certain kinetic energy which must be dis-

sipated in the shock absorbers of the landing

gear. If the energy were not absorbed at

touchdown, the aircraft would bounce along

similar to an automobile with faulty shock

absorbers. As the strut deflects on touchdown,

NAVWEPS 00-801-80

OPERATING STRENGTH LIMITATIONS

oil is forced through an orifice at high velocity

and the energy of the aircraft is absorbed. To

have an efficient strut the orifice size must be

controlled with a tapered pin to absorb the

energy with the most uniform force on the strut.

The vertical landing loads resulting at touch-

down can be simplified to an extent by assum-

ing the action of the strut to produce a uni-

formly accelerated motion of the aircraft. TV

landing load factor for touchdown at a consta

rate of descent can be expressed by the follow-

ing equation:

n= F/W

n = (ROD)’

where

a=landing load factor-the ratio of

the load in the strut, F, to the

weight, W

ROD=rate of descent, ft. per sec.

g= acceleration due to gravity

= 32 ft. per sec.’

S= effective stroke of the strut, ft.

As an example, assume that an aircraft touches

down at a constant rate of descent of 18 ft. per

sec. and the effective stroke of the strut is 18

inches (I.5 ft.). The landing load factor for

the condition would be 3.37; the average force

would be 3.37 times the weight of the aircraft.

(NOTE: there is no specific correlation between

the landing load factor and the indication of

a cockpit mounted flight accelerometer. The

response of the instrument, its mounting, and

the onset of landing loads usually prevent

direct correlation.)

This simplified equation points out two im-

portant facts. The effective stroke of the strut

should be large to minimize the loads since a

greater distance of travel reduces the force

necessary to do the work of arresting the ver-

tical descent of the aircraft. This should

NAVWEPS 00-EOT-80

OPERATING STRENGTH LIMITATIONS

emphasize the necessity of proper maintenance

of the struts. An additional fact illustrated is

that the landing load factor varies as the square

of the touchdown rate of descent. Therefore,

a 20 percent higher rate of descent increases

the landing load factor 44 percent. This fact

should emphasize the need for proper landing

technique to prevent a hard landing and over-

stress of the landing gear components and

associated structure.

The effect of landing gross weight is two-

fold. A higher gross weight at some landing

load factor produces a higher force in the

landing gear. The highe: gross weight re-

quires a higher approach speed and, if the same

glide path is used, a higher rate of descent

results. In addition to the principal vertical

loads on the landing gear, there are varied side

loads, wheel spin up and spring back loads,

etc., all of which tend to be more critical at

high gross weight, high touchdown ground

speed, and high rate of descent.

The function of the landing gear as a shock

absorbing device has an important application

when a forced landing must be accomplished

on an unprepared surface. If the terrain is

rough and the landing gear is not extended,

initial contact will be made with relatively

solid structure and whatever energy is ab-

sorbed will be accompanied by high vertical

accelerations. These high vertical accelera-

tions encountered with a gear-up landing on

an unprepared surface are the source of a very

incapacitating type injury-vertical compres-

sion fracture of the vertebrae. Unless some

peculiarity of the configuration makes it

inadvisable, it is generally recommended that

the landing gear be down for forced landing on

an unprepared surface. (NOTE: for those prone

to forget, it is also recommended that the gear

be down for landing on prepared surfaces.)

EFFECT OF OVERSTRESS ON SERVICE

LIFE

Accumulated periods of overstress can create

a very detrimental effect on the useful service

life of any structural component. This fact

is certain and irreversible. Thus, the opera-

tion of the airplane, powerplant, and various

systems must be limited to design values to

prevent failure or excessive maintenance costs

early in the anticipated service life. The

operating limitations presented in the hand-

book must be adhered to in a very strict

fashion.

In many cases of modern aircraft structures

it is very difficult to appreciate the effect of a

moderate overstress. This feature is due in

great part to the inherent strength of the

materials used in modern aircraft construction.

As a general airframe static strength require-

ment, the primary structure must not expe-

rience objectionable permanent deformation at

limit load or ~failure at 150 percent of limit

load (ultimate load is 1.5 times limit load).

To satisfy each part of the requirement, limit

load must not exceed the yield stress and ulti-

mate load must not exceed the ultimate stress

capability of the parts.

Many of the high strength materials used in

aircraft construction have stress-strain dia-

grams typical of figure 5.6. One feature of

these materials is that the yield point is at

some stress much greater than two-thirds of

the ultimate stress. Thus, the critical design

condition is the ultimate load. If 150 percent

of limit load corresponds to ultimate stress of

the material, 100 percent of limit load corre-

sponds to a stress much lower than the yield

stress. Because of the inherent properties of

the high strength material and the ultimate

factor of safety of 1.5, the limit load condition

is rarely the critical design point and usually

possesses a large positive margin of static

strength. This fact alone implies that the

structure must be grossly overstressed to pro-

duce damage easily vidble to the naked eye.

This lack of immediate visible damage with

“overstress” makes it quite diflicult to recog-

nize or appreciate the long range effect.

A reference point provided on the stress

strain diagram of figure 5.6 is a stress termed

STRESS,

PSI

100%

LIMIT LOAD

:NDURANCE

LIMIT

I- STRAIN, lN/,N

NAVWEPS 00-8OT-80

OPERATING STRENGTH LIMITATIONS

-

i\

ULTIMATE

STRENGTH

LIMIT LOAD

Figure 5.6. Typical Stress Strain Diagram for a High Strength Aluminum Alloy

the “endurance limit.” If the operating cyclic

stresses never exceed this “endurance limit” an

infinite (,or in some cases “near infinite”) num-

ber of cycles can be withstood without fatigue

failure. No significant fatigue damage accrues

from stresses below the endurance limit but

the value of this endurance limit is approxi-

mately 30 to 50 percent of the yield strength

for the light alloys used ia airkraft construc-

tion. The rate of fatigue damage caused by

stresses only &g/&y above the endurance limit

is insignificant. Even stresses near the limit

load do not cause a significant accumulation of

fatigue damage if the frequency of applicatibn

is reasonable and within the intended mission

requirement. However, stresses above the

limit load-and especially stresses well above

the limit load-create a very rapid rate of

fatigue damage.

inherent high yield strength and low ductility

of typical aircraft metals. These same over-

stresses cause high rate of fatigue damage and

create premature failure of parts in service.

The effect of accumulated overstress is rhe

formation and propagation of fatigue cracks.

While it is sure that fatigue crack always will

be formed before final failure of a part, accumu-

lated overstress is most severe and fatigue

provoking at the inevitable stress coticentra-

tions. Hence, disassembly and detailed’inspec-

tion is both costly and time-consuming. To

prevent in-service failures of a basically sound

structure, the part must be properly maintained

and operated within the design “envelope.”

Examples of in-service fatigue failures are

shown in figure 5.7.

A puzzling situation then exists. “Over-

stress” is difficult to recognize because of the

The operation of any aircraft and powerplant

must be conducted withm the operating limita-

tions prescribed in the flight handbook. No

hearsay or rumors can be substituted for chc

NAVWEPS 00-80T-80

OPERATING STRENGTH LIMITATIONS

ATTACHMENT FITTING FATiGUE FAILURES FATIGUE CRACKS IN STRUCTURAL SAMPLE

Figure 5.7. Examples of Fatigue Failures

NAVWEPS 00-801-80

OPi?RATlNG STRENGTH LIMITATIONS

accepted data presented in the aircraft hand-

book. All of the various static ‘strength,

operated past the specified time, speed, or

temperature limits without immediate appar-

service life, and aeroelastic effects must be ent damage. In each case. the cumulative

given proper respect. An airplane can be over- effect will tell at some later time when in-

stressed with the possibility that no immediate service failures occur and maintenance costs

damage is apparent. A powerplant may be increase.

SNOllVlIWll H13N3US ONllVU3dO

08-108-00 Sd3MAVN

NAVWEPS OD-8OT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

Chapter 6

APPLICATION OF AERODYNAMICS TO SPECIFBC PROW

OF FLYING

While the previous chapters have presented

the detailed parts of the general field of aero-

dynamics, there remain various problems of

flying which require the application of princi-

ples from many parts of aerodynamics. The

application of aerodynamics to these various

problems of flying will assist the Naval Aviator

in understanding these problems and develop-

ing good flying techniques.

PRIMARY CONTROL OF AIRSPEED AND

ALTITUDE

For the conditions of steady flight, the air-

plane must be in equilibrium. Equilibrium

will be achieved when there is no unbalance of

force’or moment acting on the airplane. If it is

assumed that the airplane is trimmed so that

no unbalance of pitching, yawing, or rolling

moments exists, the principal concern is for

NAVWE,PS OD-80T-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

the forces acting on the airplane, i.e., lift,

thrust, weight, and drag.

ANGLE OF ATTACK VERSUS AIRSPEED.

In order to achieve equilibrium in the vertical

direction, the net lift must equal the airplane

weight. This is a contingency of steady, level

flight or steady climbing and descending flight

when the flight path inclination is slight. A

refinement of the basic lift equation defines the

relationship of speed, weight, lift coefficient,

etc., for the condition of lift equal to weight.

V=17.2 y J TP

or

where

V=velocity, knots (TAX)

VE=equivalent airspeed, knots (EAS)

W=gross weight, lbs.

S= wing surface area, sq. ft.

W/S= wing loading, psf

g=altitudc density ratio

C,= lift coefficient

From this relationship it is appreciated that a

given configuration of airplane with a specific

wing loading, W/S, will achieve lift equal to

weight at particular combinations of velocity,

V, and lift coefficient, C,. In steady flight, each

equivalent airspeed demands a particular vaIue

of C,, and each value of C, demands a particular

equivalent airspeed to provide lift equal to

weight. Figure 6.1 illustrates a typical lift

curve for an airplane and shows the relation-

ship between C, and OL, angle of attack. For

this relationship, some specific value of a will

create a certain value of C, for any given aero-

dynamic configuration.

For the conditions of steady flight with

a given airplane, each angle of attack corre-

sponds to a specific airspeed. Each angle of

attack produces a specific value of CL and each

value of C, requires a specific value of equiva-

lent airspeed to provide lift equal to weight.

Hence, angle of attack is the primary .control of

airspeed in mad3 fright. If an airplane is es-

tablished in steady, level flight at a particular

airspeed, any increase in angle of attack will

result in some reduced airspeed common to the

increased C,. A decrease in angle of attack

will result in some increased airspeed com-

mon to the decreased CL. As a result of the

change in airspeed, the airplane may climb or

descend if there is no change in powet setting

but the change in airspeed was provided by

the change in angle of attack. The state of

the airplane during the change in speed will

be some transient condition between the

original and final steady state conditions.

Primary control of airspeed in steady flight

by angle of attack is an important principle.

With some configurations of airplanes, low

speed flight will bring about a low level of

longitudinal stick force stability and possi-

bility of low airplane static longitudi-

nal stability. In such a case, the “feel” for

airspeed will be light and may not furnish

a ready reference for easy control of the air-

plane. In addition, the high angles of attack

common to low speed flight are likely to pro-

vide large position errors to the airspeed indi-

cating system. Thus, proper control of air-

speed will be enhanced by good “attitude”

flying or-when the visual t;eference field is

poor-an angle of attack indicator.

RATE OF CLIMB AND .DESCENT. In

order for an airplane to achieve ‘equilibrium at

constant altitude, lift must be equal to weight

and thrust must be equal to drag. Steady,

level flight requires equilibrium in both the

vertical and horizontal directions. For the

case of climbing or descending flight condi-

tions, a component of weight is inclined along

the flight path direction and equilibrium is

achieved when thrust is not equal to the drag.

When the airplane is in a steady climb or

descent, the rate of climb is related by the

following expression:

NAVWEPS OO-80T-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

LIFT

COEFFICIENT

POWER

REQUIRED

AVAILABLE

HI?

\

FOR LIFT EQUAL

TO WEIGHT,

“= 17.2J$-

gs

ANGLE OF ATTACK

FOR A STEADY CLIMB,

ROC = 33,000

FPM POWER REQ’D-

POWER EXCESS

\ A 7- IER AVAILABLE

‘HIGH WILL ESTABLISH

PO’lh

w...-..

SNCY

..-- -_... - -.-.

LEVEL FLIGHT AT

--

I VELOCITY, KNOTS A 0 Figure 6.1. Primary Control of Airspeed and Altitude

NAVWEPS DO-ROT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

RC,,= 33,ooo pa;pr ( 1

where

RC=rate of climb, ft: per min.

Pn=propulsive power available, h.p.

Pr=power required for level flight, h.p.

W=gross weight, Ibs.

From this relationship it is appreciated that

the rate of climb in steady flight is a direct

function of the difference between power avail-

able and power required. If a given airplane

configuration is in lift-equal-to-weight flight

at some specific airspeed and altitude, there

is a specific power required to maintain these

conditions. If the power available from the

powerplant is adjusted LO equal the power

required, the rate of cl&b is zero (Pa--Pr=O).

This is illustrated in figure 6.1 where the power

available is ser equal to the power required at

velocity (A). If rhe airplane were in steady

level flight at velocity (A), an increase in

power available would create an excess of

power which will cause a rate of climb. Of

course, if the speed were allowed to increase

by a decreased angle of attack, the increased

power setting could simply maintain altitude

at some higher airspeed. However, if the

original aerodynamic conditions arc maintain-

ed, speed is maintained at (A) and an increased

power available results in a rate of climb.

Also, a decrease in power available at point (A)

will produce a deficiency in power and result

in a negative rate of climb (or a rate of descent).

For this reason, it is apparenr. that pomr

setting is the primary control of altitude in Jtcady

Bight. There is the direct correlation between

the excess power (Pa-P,>, and rhe airplane

rate of climb, RC.

FLYING TECHNIQUE, Since the condi-

tions of steady flight predominate during a

majority of all flying, the fundamentals of

flying technique are the principles of steady

flight:

(1) Angle of attack is the primary control

of airspeed.

(2) Power setting is the primary control

of altitude, i.e., rate of cl&b/descent.

With the exception of the transient conditions

of flight which occur during maneuvers and

acrobatics, the conditions of steady flight will

be applicable during such steady flight condi-

tions as cruise, climb, descent, takeoff, ap-

proach, landing, etc. A clear understanding

of these two principles will develop good, safe

flying techniques applicable to any sort of

airplane.

The primary control of airspeed during

steady flight conditions is the angle of attack.

However, changes in airspeed will necessitate

changes in power setting to maintain altitude

because of the variation of power required with

velocity. The primary control of altitude

(rate of climb/descent) is the power setting.

If an airplane is being flown at a particular

airspeed in level flight, an increase or decrease

in power setting will result in a rate of climb

or descent at this airspeed. While the angle

of attack must be maintained to hold airspeed

in steady flight, a change in power setting will

necessitate a change in nttitude;to.accommodate

the new flight path direction. These princi-

ples form the basis for “attitude” flying tech-

nique, i.e., “attitude plus, power equals per-

formance,” and provide .a background for

good instrument flying technique as well as

good flying technique for all ordinary flying

conditions.

One of the most important phases of flight

is the landing approach and it is during this

phase of flight that the principles of steady

flight are so applicable. If, during the landing

approach, it is realized that ithe airplane is

below the desired glide path, an increase in

nose up attitude will not insure that the

airplane will climb to the desired glide path.

In fact, an increase in nose-up attitude may

produce a greater race of descent and cause

the airplane co sink more below the desired

glide path. At a given airspeed, only an

increase in power setting can cause a rate of

climb (or lower rate of descent) and an in-

crease in nose up attitude without the appro-

priate power change only controls the airplane

to a lower speed.

REGION 0~ REVERSED COMMAND

The variation of power or thrust required

with velocity defines the power settings neces-

sary to maintain steady level flight at various

airspeeds. To simplify the situation, a gener-

ality could be,assumed that the airplane con-

figuration and. altitude define a variation of

power setting required (jet thrust required or

prop power required) versus velocity. This

general variation of required power setting

versus velocity is illustrated by’the first graph

of figure 6.2. This curve illustrates the fact

that at low speeds near the stall or minimum

control speed the power setting required for

steady level flight is quite high. However,

at low speeds, ant increase in speed reduces the

required power setting until some minimum

value is reached.at the conditions for maximum

endurance. Increased speed beyond the con-

ditions for maximum endurance will then

in&ease the power setting required for steady

level flight.

REGIONS OF NORMAL AND REVERSED

COMMAND. This .typical variation of re-

quired power setting with speed allows a

sort of terminology to be assigned to specific

regimes of velocity. Speeds greater than the

speed for maximum endurance require increas-

ingly greater power settings to achieve steady,

level flight. Since the normal command of

flight assumes a higher power setting will

achieve a greater speed, the regime of flight

speeds greater than the speed for minimum

required power setting is termed the “region

of normal command.” Obviously, parasite

drag or parasite power predominates in this

regime to produce the increased power setting

required with increased velocity. Of course,

the major items of airplane flight performance

take place in the region of normal command.

Flight speeds below the sperd for maximum

endurance produce required power settings

NAVWEPS DCI-ROT-RD

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYl,NG

which increase with a decrease in speed. Since

the increase in required power setting with

decreased velocity is contrary to the normal

command of flight, the regime of flight speeds

between the speed for minimum required

power setting and, the stall speed (or minimum

control speed) is termed the “region of re-

versed command. ” In this regime of flight,

a decrease in airspeed ‘must be accompanied

by an increased power setting in order to main-

tain steady flight. Obviously, induced drag

or induced power required predominates in

this regime to produce the increased power

setting required with decreased velocity. One

fact should be made clear about the region of

reversed command: flight in the “reversed”

region of command does not imply that a

decreased power setting will bring about a

higher airspeed or an increased power setting

will produce a lower airspeed. To be sure,

the primary control of airspeed is not the

power setting. Flight in the region of re-

versed command only implies that a higher

airspeed will repire a lower power setting and

a lower airspeed will require a higher power

setting to hold altitude.

Because of the variation of required power

setting throughout the range of flight speeds,

it is possible that one particular power setting

may be capable of achieving steady, level flight

at two different, airspeeds. As shown on the

first curve of figure 6.2, one given power setting

would meet the power requirements and allow

steady, level flight at both points 1 and 2. At

speeds lower than point 2, a deficiency of power 1

would exist and a rate of descent would be in-

curred. Similarly, at speeds greater than point

1, a deficiency of power would exist and the 1

airplane would descend. The speed range be-

tween points 1 and 2 would provide an excess

of power and climbing flight would be pro-

duccd

FEATURES OF FLIGHT IN THE NOR-

MAL AND REVERSED REGIONS OF COM-

MAND. The majority of all airplane flight is

conducted in the region of normal command,

Revised January 1965

NAVWEPS 00-6OT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

POWER

SETTING

REQUIREt

AND

AVAILABL

-

POWER

SETTING,

REOUIRED

AND

AVAILABLE

REGION OF

REVERSED COMMA

L- 0 I - -.

REGION OF=

NORMAL COMMAND

REQUIRED

SETTING

-SPEED FOR MINIMUM ,REQUlRED

POWER SETTING.ie MAX.ENDURANCE

I VELOCITY, KNOTS

REGION OF

REVERSED COMMAND

-I--

REGION OF

NORMAL COMMAND

VELOCITY, KNOTS

Figure 6.2. Region of Reversed Command

1 REOUIRED

-POWER

DEFICIENCY

Revised January 1965

e.g., cruise, climb, maneuvers, etc. The region

of reversed command is encountered primarily

in the low speed phases of flight during takeoff

and landing. Because of the extensive low

speed flight during carrier operations, the

Naval Aviator will be more familiar with the

region of reversed command than the ordinary

pilot.

The characteristics of flight in the region’of

normal command are illustrated at point A on

the second curve of figure 6.2. If the airplane

is established in steady, level flight at point A,

lift is equal to weight and the power available

is set equal to the power required. When the

airplane is disturbed to some airspeed slightly

greater than point ‘A, a power deficiency exists

and, wheq,:the &+la&is disturbed to some air-

speed slightly lower than point A, a power

excess exists. This relationship provides a

tendency for the airplane to return to the equili-

brium of point A and resume the original flight

condition following a disturbance. Also, the

static longitudinal stability of the airplane

tends to return the airplane to the original

trimmed CL and velocity corresponding to this

C,. The phugoid usually has most satisfactory

qualities at low values of C,. so the high speed

of the region ‘of normal command provides

little tendency of. the airplane’s, airspeed to

vary or wander abom.

With all factors considered, flight in Lhe

region of noi& command is characterized by

a relatively strong tendency of the airplane to

maintain the trim speed quite naturally. How-

ever, flight in the region of normal command

can lead to some unusual and erroneous impres-,

sions regarding proper flying technique. For

example, if the airplane is established at point

A in steady level flight, a controlled increase in

airspeed without a change in power setting

will create a deficiency of power and cause the

airplane to descend. Similarly, a controlled

decrease in airspeed without a change in power

setting will create an excess of power and cause

the airplane to climb. This fact, coupled with

Lhe transient motion of the airplane when the

NAVWEPS OD4OT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

angle of attack is changed rapidly, may lead to

the impression thal rate of climb and descent

can be controlled by changes in angle of attack.

While such is true in the region of normal com-

mand, for the conditions of stead’ flight, pri-

mary control of altitude remains the power

setting and the primary control of airspeed re-

mains the angle of attack. The impressions

and habits that can be developed in the region

of normal command can bring about disastrous

consequences in the region of reversed com-

mand

The characteristics of flight in the region of

reversed command are illustrated at point B

on the second curve of figure 6.2. If the air-

plane is established in steady, level flight at

point B, lift is equal to weight and the power

available is set equal to the. power required.

When the airplane is disturbed to some air-

speed slightly greater than point B, an excess

of power exists and, when the airplane is dis-

turbed to some airspeed slightly lower than

point B, a deficiency of power exists. This

relationship is basically unstable because the

variation of excess power to either side of

point B tends to magnify any original dis-

turbance. While the static longitudinal sta-

bility of the airplane tends to maintain the

original trimmed C, and airspeed correspond-

ing to that CL, the phugoid usually has the

least satisfactory qualities at the high values

of CL corresponding to low speed flight.

When all factors are considered, flight in the

region of reversed command is characterized

by a relatively weak tendency of the airplane

to maintain the trim speed naturally. In fact

it is likely that the airplane will exhibit no

inherent tendency to maintain the trim speed

in this regime of flight. For this reason, the

pilot inust give particular attention to precise

control of airspeed when operating in the low

flight speeds of the region of reversed command.

While flight in the region of normal com-

mand may create doubt as to the primary con-

trol of airspeed and altitude, operation in the

region of reversed command should leave little

‘:

-.-

* ,-.

. :,,.

_,: .-,A*

doubt about proper flying techniques. For

example, if the airplane is established at point

B in level flight, a controlled increase in air-

speed (by reducing angle of attack) without

change in power setting will create an excess

of power at the higher airspeed and cause the

airplane to climb. Also, a controlled decrease

in airspeed (by increasing angle of attack)

without a change of power setting will create

a deficiency of power at the lower airspeed

and cause the airplane to descend. This rela-

tionship should leave little doubt as to the

primary control of airspeed and altitude.

The transient conditions during the changes

in airspeed in the region of reversed command

are of interest from the standpoint of landing

flare characteristics. Suppose the airplane is

in steady flight at point B and the airplane

angle of attack is increased to correspond with

the value for the lower airspeed of point C (see

fig. 6.2). The airplane would not instanta-

neously dPvelop the lower speed and rate of

descent common to point C but would approach

the conditions of point C through some tran,

sient process depending on the airplane char.

acteristics. If the airplane characteristics are

low wing loading, high L/D, and high lift curve

slope, the increase in angle of attack at point B

will produce a transient motion in which

curvature of the flight path demonstrates a

definite flare. That is, the increase in angle

of attack creates a momentary rate of climb

(or reduction of rate of descent) which would

be accompanied by a gradual loss of airspeed.

Of course, the speed eventually decreases to

point C and the steady state rate of descent is

achieved. If the airplane characteristics are

high wing loading, low L/D, and low lift curve

slope, the increase in angle of attack at point B

may produce a transient motion in which the

airplane does not flare. That is, the increase

in angle of attack may produce such rapid re-

duction of airspeed and increase in rate of

descent that the airplane may be incapable of

a flaring flight path without an increase in

power setting. Such characteristics may neces-

NAVWEPS 00-807-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

sitate special landing techniques, particularly

in the case of a flameout landing.

Operation in the region of reversed command

does not imply that great control difficulty and

dangerous conditions will exist. However,

flight in the region of reversed command does

amplify any errors of basic flying technique.

Hence, proper flying technique and precise

control of the airplane are most necessary in

the region of reversed command.

THE ANGLE OF ATTACK INDICATOR

AND THE MIRROR LANDING SYSTEM

The usual errors during the takeoff and

landing phases of flight involve improper con-

trol of airspeed and altitude along some desired

flight path. Any errors of technique are ampli-

fied when an adequate visual reference is not

available to the pilot. It is necessary to

provide the pilot with as complete as possible

visual reference field to minimize or eliminate

any errors in perception and orientation. The

angle of attack indicator and the mirror land-

ing system assist the pilot during the phases of

takeoff and landing and allow more consistent,

precise control of the airplane.

THE A.NGLE OF ATTACK INDICATOR.

Many specific aerodynamic conditions exist at

particular angles of attack for the airplane.

Generally, the conditions of stall, landing ap-

proach, takeoff, range, endurance, etc., all

occur at specific values of lift coefficient and

specific airplane angles of attack. Thus, an

instrument to indicate or relate airplane angle

of attack would be a valuable reference to aid

the pilot.

When the airplane is at high angles of attack

it becomes difficult to provide accurate indica-

tion of airspeed because of the possibility of

large position errors. In fact, for low aspect

ratio airplane configurations at high angles of

attack, it is possible to provide indications of

angle of attack which are more accurate than

indications of airspeed. As a result, an angle

of attack indicator can be of greatest utility ar

the high angles of attack.

NAVWEPS 00-BOT-80

APPLICATION OF AERODYNAMKS

TO SPECIFIC PROBLEMS OF FLYl,NG

A particular advantage of an angle of attack

indicator is that the indicator is not directly

affected by gross weight, bank angle, load

factor, velocity, or density altitude. The

typical lift curve of figure 6.3 illustrates the

variation of lift coefficient, C,, with angle of

attack. a. When a particular aerodynamic

configuration is in subsonic flight, each angle

of attack produces a particular value of lift

coefficient. Of course, a point of special

interest on the lift curve is the maximum lift

coefficient, C,,,,. Angles of attack greater

than that for C,,,, produce a decrease in lift

coefficient and constitute the stalled condition

of flight. Since Cz,., occurs at a particular

angle of attack, any device to provide a stall

warning should be predicated on the function

of this critical angle of, attack. Under these

conditions, stall of the airplane may take place

at various airspeeds depending on gross weight,

load factor, etc., but always the same angle of

attack.

In order to reduce takeoff and landing dis-

tances and minimize arresting loads, takeoff

and landing wil! be accomplished at minimum

practical speeds. The takeoff and landing

speeds inust provide suficient margin above

the stall speed (or minimum control speed)

and are usually specified at some fixed per-

centages of the stall speed. As such, takeoff,

approach, and landing will be accomplished

at specific values of lift coefficient and, thus,

particular angles of attack. For example,

assume that point A on the lift curve is defined

as the proper aerodynamic condition for the

landing approach. This condition exists as

a particular lift coefficient and angle of attack

for a specific aerodynamic configuration.

When the airplane is flown in a steady flight

path at the prescribed angle of attack, the

resulting airspeed will be appropriate for the

airplane gross weight. Any variation in gross

weight will Simply alter the airspeed necessary

to provide suificient lift. The use of an angle

of attack indicator to maintain the recom-

mended angle of attack will insure that the

airplane is operated at the proper approach

speed-not too low or too high an airspeed.

In addition to the tise of the angle of attack

indicator during approach and landing, the

instrument may te used as a principal reference

during takeoff. The use of the angle of attack

indicator to assume the proper takeoff angle

of attack will prevent both over-rotation and

excess takeoff speed. Also, the angle of at-

tack indicator may be applicable to assist in

control of the airplane for conditions of range,

endurance, maneuvers, etc.

THE MIRROR LANDING SYSTEM. A

well planned, stabilized approach is a funda-

mental requirement for a good landing. How-

ever, one of the more difficult problems of

perception and orientation is the positioning

of the airplane along a proper flight path dur-

ing approach to landing. While various de-

vices are possible, the most successful form of

glide path indicator applicable to both field

and shipboard operations is the mirror landing

system. The function of the mirror landing

system is to provide the pilot with an accurate

visual reference for a selected flight path which

has the desired inclination and point of touch-

down. Utilization of the mirror system will

allow the pilot to position the airplane along

the desired glide path and touch down at the

desired point. When the proper glide path

inclination is set, the pilot can be assured that

the rate of descent will not be excessive and

a foundation is established for a successful

landing.

The combination of the angle of attack in-

dicator and the mirror landing system can

provide an excellent referetice for a landing

technique. The use of the angle of attack

indicator will provide the airplane with the

proper airspeed while the mirror system refer-

ence will provide the desired flight path.: When

shipboard operations are conducted without

the mirror system and angle of attack indicator,

the landing signal officer must provide the

immediate reference of airspeed and flight path.

The LSO must perceive gnd judge the angle of

LIFT

COEFFICIENT

CL MAX -

NAVWEPS OD-BOT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

7 A

- STALL

ANGLE OF ATTACK, 0

3s9

NAVWEPS DD-BOT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

attack (and, hence, airspeed) and the flight

path of the landing aircraft and signal correc-

tions to be made in order to achieve the desired

flight path and angle of attack. Because of

the field of orientation available to the LSO,

he is able to perceive the flight path and angle

of attack more accurately than the pilot with-

out an angle of attack indicator and mirror

landing system.

THE APPROACH AND LANDING

The specific techniques necessary during the

phase of approach and landing may vary con-

siderably between various types of airplanes

and various operations. However, regardless

of the airplane type or operation, there are

certain fundamental principles which will de-

fine the basic techniques of flying during ap-

proach and landing. The specific procedures

recommended for each airplane type must be

followed exactly to insure a consistent, safe

landing technique.

THE APPROACH. The approach must be

conductrd to provide a stabilized, steady flight

path to the intended point of touchdown. The

approach speed specified for an airplane must

provide sufficient margin above the stall speed

or minimum control speed to allow satisfactory

control and adequate maneuverability. On the

other hand, the approach speed must not be

greatly in excess of the touchdown speed or a

large reduction in speed would be necessary

prior to ground contact. Generally, the ap-

proach speed will be from 10 to 30 percent

above the stall speed depending on the air-

plane type and the particular operation.

During the approach, the pilot must attempt

to maintain a smooth flight path and prepare

for the touchdown. A smooth, steady ap-

proach to landing will minimize the transient

items of the flight path and provide the pilot

better opportunity to perceive and orientate

the airplane along the desired flight path.

Steep turns must be avoided at the low speeds

of the approach because of the increase in drag

and stall speed in the turn. Figure 6.4 illus-

trates the typical change in thrust required

caused by a steep turn. A steep turn may cause

the airplane to stall or the large increase in in-

duced drag may create an excessive rate of

descent. In either case, there may not be suf-

ficient altitude to effect recovery. If the .air-

plane is not properly lined up on the final ap-

proach, it is certainly preferable to take a

waveoff and go around rather than “press on

regardless” and attempt to salvage a decent

landing from a poor approach.

The proper coordination of the controls is

an absolute necessity during the approach. In

this sense, due respect must be given to the

primary control of airspeed and race of descent

for the conditions of the steady approach.

Thus, the proper angle of attack will produce

the desired approach airspeed; too low an

angle of attack will incur an excess speed while

an excessive angle of attack will produce a

deficiency of speed and may cause stall or con-

trol problems. Once the proper airspeed and

angle of attack are attained the primary control

of rate of descent during the steady approach

will be the power setting. For example, if it

is realized that the airplane is above the de-

sired glide path, a more nose-down attitude

without a decrease in power setting will result

in a gain in airspeed. On the other hand, if it

is realized that the airplane is below the desired

glide path, a more nose-up attitude without an

increase in power setting will simply allow the

airplane to fly more slowly and-in the region

of reversed command-eventually produce a

greater rate of descent. For the conditions of

steady flight, angle of attack is the primary

control of airspeed and power setting is the

primary control of rate of climb and descent.

This is especially true during the steady ap-

proach to landing. Of course, the ability of

the powerplant to produce rapid changes in

thrust will affect the specific technique to be

used. If the powerplant is not capable of pro-

ducing immediate controlled changes in thrust,

the operating technique must’ account for this

NAVWEPS OD-BOT-80

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

EFFECT OF STEEP TURNS ON THRUST REQ’D

THRUST

REO’D

LBS.

WING LEVEL FLIGHT

VARloUS

APPROACH PATHS

TYPICAL LIFT n

CURVES

LIFT

COEFFICIENT

ANGLE OF ATTACK, a

Figure 6.4. The Approach and Landing

NAVWEPS OD-BOT-BO

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYI’NG

deficiency. It is most desirable that the power-

plant be capable of effecting rapid changes in

thrust to allow precise control of the airplane

during approach.

The type of approach path is an important

factor since it affects the requirement of the

flare, the touchdown rate of descent, and-to

some extent-the ability to control the point

of touchdown. Approach path A of figure

6.4 depicts the steep, low power approach.

Such a flight path generally involves a low

power setting near idle conditions and a high

rate of descent. Precise control of the air-

plane is difficult and an excess airspeed usually

results from an approach path similar to A.

Waveoff may be difficult because of the re-

quired engine acceleration and the high rate

of descent. In addition, the steep approach

path with high rate of descent requires con-

siderable flare to reduce the rate of descent at

touchdown. This extreme flare requirement

will be di,fficult to execute with consistency

and will generally result in great variation

in the speed, rate of descent, and point of

touchdown.

Approach path C of figure 6.4 typifies the

long, shallow approach with too small an

inclination of the flight path.. Such a flight

path requires a relatively high power setting

and a deficiency of airspeed is a usual conse-

quence. This extreme of an approach path

is not desirable because it is difficult to control

the point of touchdown and the low speed

may allow the airplane to settle prematurely

short of the intended landing touchdown.

Some approach path between the extremes

of A and C must be selected, e.g., flight path

B. The desirable approach path must not

incur excessive speed and rate of descent or

require excessive flaring prior to touchdown.

Also, some moderate power setting must be

required which will allow accurate control

of the flight path and provide suitable waveoff

characteristics. The approach flight path

cannot be too shallow for excessive power

setting may be required and it may be difficult

to judge and control the point of touchdown.

The LSO, mirror landing system, and various

approach lighting systems will aid the pilot

in achieving the desired approach flight path.

THE LANDING FLARE AND TOUCH-

DOWN. The specific techniques of landing

flare and touchdown will vary considerably

between various types of airplanes. In fact,

for certain types of airplanes, a flare from a

properly executed approach may not be de-

sirable because of the possibility of certain

critical dynamic landing loads or because of

the necessity for a certain standard of tech-

nique when aerodynamic flare characteristics

are critical. The landing speed should be the

lowest practical speed above the stall or mini-

mum control speed to reduce landing distances

and arresting loads. Generally, the landing

speed will be from 5 to 25 percent above the

stall speed depending on the airplane type

and the particular operation.

The technique required for the landing will

be determined in great part by the aerodynamic

characteristics of the airplane. If the airplane

characteristics are low wing loading, high LID,

and relatively high lift curve slope, the airplane

usually will have good landing flare charac-

teristics. If the airplane characteristics are

high wing loading, low L/D, and relatively low

lift curve slope, the airplane may not possess

desirable flare characteristics and landing tech-

nique may require a minimum of flare to

touchdown. These extremes are illustrated by

the lift curves of figure 6.4.

In preparation for the landing, several factors

must be accounted for because of their effect

on landing distance, landing loads, and arrest-

ing loads. These factors are:

(1) Landing gross weight must be con-

sidered because of its effect on landing speed

and landing loads. Since the landing is

accomplished at a specific angle of attack or

margin above the stall speed, gross weight

will define the landing speed. In addition,

the gross weight is an important factor in

determining the landing distance and energy

dissipating requirements of the brakes.

There will be a maximum design landing

weight specified for each airpIane and this

limitation must be respected because of

critical landing loads, arresting loads, or

brake requirements. Of course, any air-

plane will have a limiting touchdown rate

of descent specified with the maximum land-

ing weight and the principal landing load

limitations will be defined by the combina-

tion. of gross weight and rate of descent at

touchdown.

(2) The surface winds must be considered

because of the large effect of a headwind or

tailwind OR the landing distance. In the

case of the crosswind, the component of

wind along the runway will be the effective

headwind or tailwind velocity. Also, the

crosswind component across the runway will

define certain requirements of lateral control

power. The airplane which exhibirs large

dihedral effect at high lift coefficients is

quite sensitive to crosswind and a limiting

crosswind component will be defined for the

configuration.

(3) Press.w~ dtitsde and tmpma~e will

affect the landing distance because of the

effect on the true airspeed for landing.

Thus, pressure altitude and temperature must

be considered to define the density altitude.

(4) The runway condition must be con-

sidered for its effect on landing distances.

Runway slope of ordinary values will ordi-

narily favor selection of a runway for a

favorable headwind at landing. The surface

condition of the runway will determine

braking effectiveness and ice or water on the

runway may produce a considerable increase

in the minimum landing distance.

Thus, preparation for the landing must in-

clude determination of the landing distance of

the airplane and comparison with the runway

length available. Use of the angle of attack

indicator and the mirror landing system will

assist the pilot in effecting touchdown at the

desired location with the proper airspeed. Of

NAVWEPS OD-BOT-BO

APPLICATION OF AERODYNAMICS

TO SPECIFIC PROBLEMS OF FLYING

course, the landing is not completed until the

airplane is slowed to turn off the runway.

Control of the airpIane must be maintained

after the touchdown and proper technique must

be used to decelerate the airplane.

TYPICAL ERRORS. There are many un-

desirable consequences when basic principles

and specific procedures are not followed during

the approach and landing. Some of the typical

errors involved in landing accidents are out-

lined in the following discussion.

The steep, low power approach leads to an

exce.rsive rate of descent and the possibility of a

hard landing. This is particularly the case

for the modern, low aspect ratio, swept wing

airplane configuration which incurs very large

induced drag at low speeds and does not have

very conventional flare characteristics. For

this type of airplane in a steep, low power

approach, an increased angle of attack without

a change of power setting may not cause a

reduction of rate of descent and may even in-

crease the rate of descent at touchdown. For

this reason, a moderate stabilized approach is

necessary and the principal changes in rate of

descent must be controlled by changes in power

setting and principal changes in airspeed must

be controlled by changes in angle of attack.

‘An excessive angle of attack during the ap-

proach and landing implies that the airplane is

being operated at too low an airspeed. Of

course, excessive angle of attack may cause the

airplane to stall or spin and the low altitude

may preclude recovery. Also, the low aspect

ratio configuration at an excessively low air-

speed will incur very high induced drag and

will necessitate a high power setting or other-

wise incur an excessive rate of descent. An

additional problem is created by an excessive

angle of attack for the airplane which exhibits

a large dihedral effect at high lift coefficients.

In this case, the airplane would be more sensi-

tive to crosswind.s and adequate lateral control

may not be available to effect a safe landing at

a critical value of crosswind.

MAVWEPS OO-BOLBO

APPLICA’IIOM OF AERODYNAMICS

10 SPECIFIC PROBLEMS OF FLYI~NG