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angle of the surface must be sufficiently great
to prevent stall and subsequent loss of effec-
tiveness at ordinary sideslip angles. The high
Mach numbers of supersonic flight produces a
decrease in lift curve slope with the consequent
reduction in tail contribution to stability. In
order to have sufficient directional stability at
high Mach numbers, the typical supersonic
configuration will exhibit relatively large
vertical tail surfaces.
The flow field in which the vertical tail
operates is affected by the othei components
of the airplane as well as powe; effects. The
dynamic pressure at the vertical tail could
depend on the slipstream of a propeller or the
boundary layer of the fuselage. Also, the
local flow direction at the vertical tail is in-
fluenced by the wing wake, fuselage crossflow,
induced flow of the horizontal tail, or the
direction of slipstream from a propeller. Each
of these factors must be considered as possibly
affecting the contribution of the vertical tail
to directional stability.
The contribution of the wing tb %tatic direc-
tional stability is tisually small: The swept
wing provides a stable contribution’depending
on the amount of sweepback but the contribu-
tion is relatively weak when compared with
other components. :. The contribution of the fuselage and nacelles
is of primary importance since these compo-
nents furnish rhe greatest destabilizing in-
fluence. The contribution of the fuselage and
nacelles is similar to the longitudinal case
with the exception that there is no large in-
fluence of the induced flow field of the wing.
The subsonic center of pressure of the fuselage
will be located at or forward of the quarter-
length point and, since the airplane c.g. is
usually considerably aft of this point, the
fuselage contribution will be destabilizing.
However, at large angles of sideslip the large
destabilizing contribution of the fuselage di-
minishes which is some relief to the problem
of maintaining directional stability at large
displacements. The supersonic pressure,. dis-
tribution on the body provides a relatively
NAVWEPS OO-ROLRO
STARIUTY AND CONTROL
greater aerodynamic force and, generally, a
continued destabilizing influence.
Figure 4.23 illustrates a typical buildup of
the directional stability of an airplane by
separating the contribution of the fuselage
and tail. As shown by the graph of C. versus
6, the contribution of the fuselage is de-
stabilizing but the instability decreases at
large sideslip angles. Tbe contribution of the
vertical tail alone is highly stabilizing up to
the point where the surface begins to stall.
The contribution of the vertical tail must be
large enough so that the complete airplane
(wing-fuselage-tail combination) exhibits the
required degree of stability.
The dorsal fin has a powerful effect on pre-
serving the directional stability at large angles
of sideslip wliich would produce stall of the
vertical tail. The addition of a dorsal fin to
the airplane will allay the decay of directional
stability at high sideslip in two ways. The
least obvious but most important effect is a
large increase in the fuselage stability at large
sideslip angles. In addition, the effective
aspect rario of the vertical tail is reduced
which increases the stall angle for the surface.
By this twofold effect, the addition of the
dorsal fin is a v useful’ device.
Poluer effects on static directional stability
are similar to the power effects on static
longitudinal stability. The direct effects are
confined to the normal force at the propeller
plane or the jet inlet and, of course, are de-
stabilizing when the propeller or inlet is
located ahead of the c.g. The indirect effects
of power induced velocities and flow dirkccion
changes at the vertical tail are quite significant
for the propeller driven airplane and can pro-
duce large directional trim changes. As in
the lontitudinal case, the indirect effects are
negligible for the jet powered airplane.
The contribution of the direct and indirect
power effects to static directional stability is
greatest for the propeller powered airplane
and usually slight for the jet powered airplane.
In either case, the general effect of power is
287 | 304 | 304 | 00-80T-80.pdf |
NAVWEPS oO-801-80
STABILITY AND CONTROL
CONTRIBUTION OF VERTICALTAIL
CHANGE IN
TAIL LIFT
TYPICAL DIRECTIONAL STABILITY
BUILD-UP
AIRPLANE WITH
DORSAL FIN
STALL ,-ADDED
Figure 4.23. Contribution of Components to Directional Stability
288 | 305 | 305 | 00-80T-80.pdf |
NAVWEPS Oe8OT-80
STABILITY AND CONTROL
EFFECT OF RUDDER FLOAT ON STATIC
DIRECTIONAL STABILITY
t
\ RUDDER-FIXED
RUDDER-FREE
RUDDER FLOAT
-e ANGLE
t
SIDESLIP ANGLE, p
EFFECT OF ANGLE OF ATTACK
HIGH ANGLE
OF ATTACK
w
SIDESLIP ANGLE, fla
EFFECT OF MACH NUMBER
A
SIDESLIP ANGLE, p
Figure 4.24. Factors Affecting Direcfional Stability
289 | 306 | 306 | 00-80T-80.pdf |
NAVWRPS DD-807-80
STABILITY AND CONTROL
destabilizing and the greatest contribution
will occur at high power and low dynamic
pressure as during a waveoff.
As in the case of longitudinal static stability,
freeing the controls will reduce the effective-
ness of the tail and alter the stability. While
the rudder must be balanced to reduce control
pedal forces, the rudder will tend to float or
streamline and reduce the contribution of the
vertical tail to static directional stability. The
floating tendency is greatest at large angles of
sideslip where large angles of attack for the
vertical tail tend to decrease aerodynamic bal-
ante. Figure 4.24 illustrates the difference be-
tween rudder-fixed and rudder-free static di-
rectional stability.
CRITICAL CONDITIONS. The most criti-
cal conditions of,staric directional stability are
usually the combination of several separate
effects. The combination which produces the
most critical condition is much dependent upon
the type and mission of the airplane. In addi-
tion, there exists a coupling of lateral and di-
rectional effects such that the required degree
of static directional stability may be deter-
mined by some of these coupled conditions.
Center of gravity position has a relatively
negligible effect on static directional stability.
The usual range of c.g. position on any air-
plane is set by the Jinits of long&d&a/ stability
and control. Within this limiting range of
c.g. position, no significant changes take place
in the contribution of the vertical tail, fuselage,
nacelles, etc. Hence, the static directional
stability is essentially unaffected by the varia-
tion of c.g. position within the longitudinal
limits.
When the airplane is at a high angle of a$tack
a decrease in static directional stability can be
anticipated. As shown by the second chart of
figure 4.24, a high angle of attack reduces the
stable slope of the curve of C,, versus 8, The
decrease in static directional stability is due in
great part to the reduction in the contribution
of the vertica1 tail. At high angles of attack,
the effectiveness of the vertical tail is reduced
because of increase in the fuselage boundary
layer at the vertical tail location. The decay of
dir&ctional stability with angle of attack is
most significant for the low aspect ratjo air-
plane with sweepback since this configuration
requires such high angles of attack to achieve
high lifr coefficients. Such decay in directional
stability can have a profound effect on the re-
sponse of the airplane to adverse yaw and spin
characteristics.
High Mach ntrmbers of supersonic flight reduce
the contribution of the vertical tail to direc-
tional stability because of the reduction of lift
cnrve slope with Mach number. The third
chart of figure 4.24 illustrates the typical decay
of directional stability with Mach number. To
produce the required directional stability at
high Mach numbers, a viziy large vertical tail
area may be necessary. Ventral fins may be
added as an additional contribution to direc-
tional stability but landing clearance require-
ments may limir their size or require the fins to
be retractable.
Hence, the most critical demands of static
directional stability will occur from some
combination of the following effects:
(1) high angle of sideslip
(2) high power at low airspeed
(3) high angle of attack
(4) high Mach number
The propeller powered airplane may have such
considerable power effects that the critical
conditions may occur at low speed while the
effect of high Mach numbers may produce the
critical conditions for the typical supersonic
airplane. In addition, the coupling of lateral
and directional effects may require prescribed
degrees of directional stability.
DIRECTIONAL CONTROL
In addition to directional stability, the air-
plane must have adequate directional control
to coordinate turns, balance power effects,
create sideslip, balance unsymmetrical power,
etc. The principal source of directional con-
trol is the rudder and the rudder must be | 307 | 307 | 00-80T-80.pdf |
capable of producing sufhcient yawing moment
for the critical conditions of flight.
The effect of rudder deflection is to produce
a yawing moment coefficient according to
control deflection and produce equilibrium at
some angle of sideslip. For small deflections
of the rudder, there is no change in stability
but a change in equilibrium. Figure 4.25
shows the effect of rudder deflection on yawing
moment coefficient curves with the change in
equilibrium sideslip angle.
If the airplane exhibits static directional
stability with rudder lixed, each angle of side-
slip requires a particular deflection of the
rudder to achieve equilibrium. Rudder-free
directional stability will exist when the float
angle of the rudder is less than the rudder
deflection required for equilibrium. However
at high angles of sideslip, the floating tend-
ency of the rudder increases. This is illus-
trated by the second chart of figure 4.25 where
the line of rudder float angle shows a sharp
increase at large values of sideslip. If the
floating angle of the rudder catches up with
the required rudder angle, the, rudder pedal
force will decrease to zero and rudder lock will
occur. Sideslip angles beyond this point pro-
duce a floating angle greater than the required
rudder deflection and the rudder tends to float
to the limit of deflection.
Rudder lock is accompanied by a reversal of
pedal force and rudder-free instability will
exist. The dorsal fin is a useful addition in
this case since it will improve the directional
stability at high angles of sideslip. The re-
sulting increase in stability requires larger
deflections of the rudder to achieve equilibrium
at high sideslip and the tendency for rudder
lock is reduced.
Rudder-free directional stability is appre-
ciated by the pilot as the rudder pedal force to
maintain a given sideslip. If the rudder pedal
force gradient is too low near zero sideslip, it
will be difficult to maintain zero sideslip dur-
ing various maneuvers. The airplane should
NAVWEPS 00-SOT-80
STABIUTY AND CONTROL
have a stable rudder pedal feel through the
available range of sideslip.
DIRECTIONAL CONTROL REQUIRE-
MENTS. The control power of the rudder
must be adequate to contend with the many
unsymmetrical conditions of flight. Gener-
ally, there are five conditions of flight which
provide the most criticalrequirements of di-
‘rectional control power. The type and mission
of the airplane will decide which of these
conditions is most important.
ADVERSE YAW. When an airplane is
rolled into a turn yawing moments are pro-
duced which require rudder deflection to main-
tain zero sideslip, i.e., coordinate the turn.
The usual source of adverse yawing moment is
illustrated in figure 4.26. When the airplane
shown is subject to a roll to the left, the down-
going port wing will experience a new relative
wind and an increase in angle of attack. The
inclination of the lift vector produces a com-
ponent force forward on the downgoing wing.
The upgoing starboard wing has its lift in-
clined with a component force aft. The re-
sulting yawing moment due to rolling motion
is in a direction opposite to the roll and is
hence “adverse yaw.” The yaw due to roll is
primarily a function of the wing lift coefficient
and is greatest at high C,.
In addition to the yaw due to rolling motion
there will be a yawing moment contribution
due to control surface deflection. Conventional
ailerons usually contribute an adverse yaw
while spoilers may contribute a favorable or
“proverse” yaw. The high wing airplane
with a large vertical tail may encounter an
influence from inboard ailerons. Such a con-
figuration may induce flow directions at the
vertical tail to cause proverse yaw.
Since adverse yaw will be greatest at high
C, and full deflection of the ailerons, coordi-
nating steep turns at low speed may produce
a critical requirement for rudder control power.
SPIN RECOVERY. In the majority of air-
planes, the rudder is the principal control for
spin recovery. Powerful control of sideslip at
291 | 308 | 308 | 00-80T-80.pdf |
NAVWEPS 00-807-80
STABILITY AND CONTROL
EFFECT OF RUDDER DEFLECTION ON
EOlJlLlSRlUM SIDESLIP ANGLE
RUDDER DEFLECTION
RUDDER LOCK.
RUDDER DEFLECTION
FLOAT ANGLE
/ SIDESLIP ANGLE, p +
EFFECT DF RUDDER LOCK ON PEDAL FORCE
RUDDER LOCK
+P
w
---
DORSAL FIN ADDED
Figure 4.25. Directional Control
192 | 309 | 309 | 00-80T-80.pdf |
ADVERSE YAW DUE TO ROLL
FORCE FORWARD
DOWNGOING
PORT WING
,IRPLANE.IN ROLL TO LEFT
NAVWEPS 00-8OT-30
STABILITY AND CONTROL
\FOR SAKE OF CLARITY. /
SLIPSTREAM SWIRL ON THE PROPFLLER POWERED AIRPLANE
YAWING MOMENT COEFFICIENT
FROM ASYMMETRICAL /
THRUST
YAWING MOMENT DUE TO ASYMMETRICAL THRUST
I
I
w
EQUIVALENT AIRSPEED, KNOTS
Figure 4.26. Requirements for Directional Control
293 | 310 | 310 | 00-80T-80.pdf |
NAVWEPS 00-8OT-80
STABILITY AND CONTROL
high angles of attack is required to effect re-
covery during a spin. Since the effectiveness
of the vertical tail is reduced at large angles of
attack, the directional control power neces-
sary for spin recovery may produce a critical
requirement of rudder power.
SLIPSTREAM ROTATION. A critical di-
rectional control requirement may exist when
the propeller powered airplane is at high
power and low airspeed. As shown in figure
4.26, the single rotation propeller induces
a slipstream swirl which causes a change in
flow direction at the vertical tail. The rudder
must furnish sufficient control power to balance
this condition and achieve zero sideslip.
CROSSWIND TAKEOFF AND LANDING.
Since the airplane must make a true path down
the runway, a crosswind during takeoff or
landing will require that the airplane be.con-
trolled in a sideslip. The rudder must have
sufficient control power to create the required
sideslip for the expected crosswinds.
ASYMMETRICAL POWER. The design
of a multiengine airplane must account for the
possibility of an engine failure at low airspeed.
The unbalance of thrust from a condition of
unsymmetrical power produces a yawing mo-
ment dependent upon the thrust unbalance
and the lever arm of the force. The deflection
of the rudder will create a side force on the tail
and contribute a yawing moment to balance
the yawing moment due to the unbalance of
thrust. Since the yawing moment coefficient
from the unbalance of thrust will be greatest
at low speed, the critical requirement will be
at a low speed with the one critical engine
out and the remaining engines at maximum
power. Figure 4.26 compares the yawing
moment coeflicient for maximum rudder deflec-
tion with the yawing moment coefficient for
the unbalance of thrust. The intersection of
the two lines,determines the minimum speed
for directional control, i.e., the lowest speed
at which the rudder control moment can equal
the moment of unbalanced thrust, It is usually
specified that the minimum directional control
speed be no greater than 1.2 times the stall
294
Revised January 1965
speed of the airplane in the lightest practical
takeoff configuration. This will provide ade-
quate directional control for the remaining
conditions of flight.
Once defined, the minimum directional con-
trol speed is not a function of weight, altitude,
etc., but is simply the equivalent airspeed (or
dynamic pressure). to produce a required yaw-
ing moment with the maximum rudder deflec-
tion. If the airplane is operated in the critical
unbalance of power below the minimum con
trol speed, the airplane will yaw uncontrolla-
bly into the inoperative engine. In order to
regain directional control below the minimum
speed certain alternatives exist: reduce power
on the operating engines or sacrifice altitude
for airspeed. Neither alternative is satisfac-
tory if the airplane is in a marginal condition
of powered flight so due respect must be given
to the minimum control speed.
Due to the side force on the vertical tail, a
slight bank is necessary to prevent turning
flight at zero sideslip. The inoperative engine
will be raised and the inclined wing lift will
provide a component of force to balance the 1
side force on the tail.
In each of the critical conditions of required
directional control, high directional stability
is desirable as it will reduce the displacement
of the aircraft from any disturbing influence.
Of course, directional control must he sufficient
to attain zero sideslip. The critical control
requirement for the multiengine airplane is
the condition of asymmetrical power since
spinning is not common to this type of airplane.
The single engine propeller airplane may have
either the spin recovery or the slipstream rota-
tion as a critical design condition. The single
engine jet airplane may have a variety of
critical items but the spin recovery require-
ment usually predominates.
LATERAL STABILITY AND CONTROL
LATERAL STABILITY
The static lateral stability of an airplane
involves consideration of rolling moments due | 311 | 311 | 00-80T-80.pdf |
to sideslip. If an airplane has favorable rolling
moment due to sideslip, a lateral displacement
from wing level flight produces sideslip and
the sideslip creates rolling moments tending
to return the airplane to wing level flight.
By this action, static lateral stability will be
evident. Of course, a sideslip will produce
yawing moments depending on the nature of
the static directional stability but the consid-
rations of static lateral stability will involve
only the ‘relationship of rolling moments and
sideslip.
DEFINITIONS. The axis system of an
airplane defines a positive rolling, L, as a
moment about the longitudinal axis which
tends to rotate the right wing down. As in
other aerodynamic considerations, it is con-
venient to consider rolling moments in the
coefficient form so that lateral stability can
be evaluated independent of weight, altitude,
speeds, etc. The rolling moment, L, is defined
in the coeflicient form by the following equa-
tion :
or
L=C,qSb
*
+I
0
where
L=rolling moment, ft.-lbs., positive to
the right
4 = dynamic pressure, psf.
S=wing area, sq. ft.
b = wingspan, ft.
C,=rolling moment coeflicient, positive
to the right
The angle of sideslip, 8, has been defined
previously as the angle between the airplane
centerline and the relative wind and is positive
when the relative wind is to the right of the
centerline.
The static lateral stability of an airplane can
be illustrated by a graph of rolling moment
coefficient, Cl, versus sideslip angle, 8, such
as shown in figure 4.27. When the airplane
is subject to a positive sideslip angle, lateral
stability will be evident if a negative rolling
NAVWEPS 00-8OT-80
STABILITY AND COI’ITROL
moment coefficient results. Thus, when the
relative wind comes from the right (+-a>,
a rolling moment to the left (-Cl> should be
created which tends to roll the airplane to
the left. Lateral stability will exist when
the curve of C1 versus p has a negative slope
and the degree of stability will be a function
of the slope of this curve. If the slope of the
curve is zero, neutral lateral stability exists;
if the slope is positive lateral instability is
present.
It is desirable to have lateral stability or
favorable roll due to sideslip. However, the
required magnitude of lateral stability is deter-
mined by many factors. Excessive roll due to
sideslip complicates crosswind takeoff and
landing and may lead to undesirable oscil-
latory coupling with the directional motion of
the airplane. In addition, a high lateral sta-
bility may combine with adverse yaw to hinder
rolling performance. Generally, favorable han-
dling qualities are obtained with a relatively
light-or weak positive-lateral stability.
CONTRIBUTION OF THE AIRPLANE
COMPONENTS. In order to appreciate the
development of lateral stability in an airplane,
each of the contribution components must be
inspected. Of course, there will be interference
between the components which will alter the
contribution to stability of each component on
the airplane.
The principal surface contributing to the
lateral stability of an airplane is the wing. The
effect of the geometric dihedral of a wing is a
powerful contribution to lateral stability. As
shown in figure 4.28, a wing with dihedral will
develop stable rolling moments with sideslip.
If the relative wind comes from the side, the
wing into the wind is subject to an increase in
angle of attack and develops an increase in lift.
The wing away from the wind. is subject to a
decrease in angle of attack and develops a de-
crease in lift. The changes in lift effect a rolling
moment tending to raise the windward wing
hence dihedral contributes a stable roll due to
sideslip.
295 | 312 | 312 | 00-80T-80.pdf |
NAVWEPS DD-8OT-80
STABILITY AND CONTROL
RELATIVE WIND
+L, ROLLING MOMENT
ROLLING MOMENT COEFFICIENT
UNSTABLE 7,
-I
TABLE ROLL DUE
TO SIDESLIP
SIDESLIP ANGLE, /3
NEUTRAL
Figure 4.27. Static Lateral Stability
296 | 313 | 313 | 00-80T-80.pdf |
NAVWEPS CID-8OT-80
STABILITY AND CONTROL
EFFECT OF DlilEDRAL
EFFECTIVE INCREASE IN
--SE IN
LIFT DUE TO SIDESLIP
EFFECT OF SWEEPBACK
R~~;~~~p
CONTRIBUTION OF VERTICAL TAIL
SIDESLIP CONTRIBUTES
ROLLING MOMENT
Figure 4.28. Contribution of Components to Lateral Stability
297 | 314 | 314 | 00-80T-80.pdf |
NAVWEPS OO-BOT-80
STABILITY AND CONTROL
Since wing dihedral is so powerful in pro-
ducing lateral stability it is taken as a common
denominator of the lateral stability contribu-
tion of all other components. Generally, the
contribution of wing position, flaps, power,
etc., is expressed as an equivalent amount of
“effective dihedral” or “dihedral effect.”
The contribution of the fadage alone is
usually quite small depending on the location
of the resultant aerodynamic side force on the
fuselage. However, the effect of the wing-
fuselage-tail combination is significant since
the vertical placement of the wing on the fuse-
lage can greatly affect the stability of the com-
bination. A wing located at the mid wing
position will generally exhibit a dihedral effect
no different from that of the wing alone. A
low wing location on the fuselage may con-
tribute an effect equivalent to 3’ or 4’ of nega-
tive dihedral while a high wing location may
contribute a positive dihedral of 2’ or 3’. The
magnitude of dihedral effect contributed by
vertical position of the wing is large and may
necessitate a noticeable dihedral angle for the
low wing configuration.
The contribution of wccpback to dihedral ef-
fect is important because of the nature of the
contribution. As shown in figure 4.28, the
swept wing in a sideslip has the wing into
wind operating with an effective decrease in
sweepback while the wing out of the wind
is operating with an effective increase in
sweepback. If the wing is at a positive lift
coefficient, the wing into the wind has less
sweep and an increase in lift and the wing out
of the wind has more sweep and a decrease in
lift. In this manner the swept back wing
would contribute a positive dihedral effect and
the swept forward wing would contribute a
negative dihedral effect.
The unusual nature of the contribution of
sweepback to dihedral effect is that the con-
tribution is proportional to the wing lift
coefficient as well as the angle of sweepback.
It should be clear that the swept wing at zero
lift will provide no roll due to sideslip since
there is no wing lift to change. Thus, the
dihedral effect due to sweepback is zero at zero
lift and increases directly with wing lift
coefficient. When the demands of high speed
flight require a large amount of sweepback, the
resulting configuration may have an excessive-
ly high dihedral effect at low speeds (high CL)
while the dihedral effect may be satisfactory
in normal flight (low or medium C,).
The vertical tail of modern configurations
can provide a sign&ant-and, at times, un-
desirable-contribution to the effective dihe-
dral. If the vertical tail is large, the side force
produced by sideslip may produce a noticeable
rolling moment as well as the important yaw-
ing moment contribution. Such an effect is
usually small for the conventional airplane
configuration but the modern high speed
airplane configuration induces this effect to a
great magnitude. It is difficult then to obtain
a large vertical tail contribution to directional
stability without incurring an additional con-
tribution to dihedral effect.
The amount of effective dihedral necessary
to produce satisfactory flying qualities varies
greatly with the type and purpose of the air-
plane. Generally, the effective dihedral should
not be too great since high roll due to side-
slip can create certain problems. Excessive
dihedral effect can lead to “Dutch roll,”
difficult rudder coordination in rolling maneu-
vers, or place extreme demands for lateral
control power during crosswind takeoff and
landing. Of course, the effective dihedral
should not be negative during the predominat-
ing conditions of flight, e.g., cruise, high
speed, etc. If the airplane demonstrates satis-
factory dihedral effect for these conditions of
flight, certain exceptions can be considered
when the airplane is in the takeoff and landing
configuration. Since the effects of flaps and
power are destablizing and reduce the dihedral
effect, a certain amount of negative dihedral
effect may be possible due to these sources.
The deflection of flaps causes the inboard
sections of the wing to become relatively more
298 | 315 | 315 | 00-80T-80.pdf |
effective and these sections have a small
spanwise moment arm. Therefore, the changes
in wing lift due to sideslip occur closer in-
board and the dihedral effect is reduced. The
effect of power on dihedral effect is negligible
for the jet airplane but considerable for the
propeller driven airplane. The propeller slip-
stream at high power and low airspeed makes
the inboard wing sections much more effective
and reduces the dihedral effect. The reduction
in dihedral effect is most critical when the
flap and power effects are combined, e.g., the
propeller driven airplane in the power approach
or waveoff.
With certain exceptions during the condi-
tions of landing and takeoff, the dihedral
effect or lateral stability should be positive
but light. The problems created by excessive
dihedral effect are considerable and difficult
to contend with. Lateral stability will be
evident to a pilot by stick forces and displace-
ments required to maintain sideslip. Positive
stick force stability will be evident by stick
forces required in the direction of the controlled
sideslip.
LATERAL DYNAMIC EFFECTS
Previous discussion has separated the lateral
and directional response of the airplane to
sideslip. This separation is convenient for
detailed study of each the airplane static
lateral stability and the airplane static direc-
tional stability. However, when the airplane
in free flight is placed in a sideslip, the lateral
and directional response will be coupled, i.e.,
simultaneously the airplane produces rolling
moment due to sideslip and yawing moment
due to sideslip. Thus, the lateral dynamic
motion of the airplane in free flight must
consider the coupling or interaction of the
lateral and directional effects.
The principal effects which determine the
lateral dynamic characteristics of an airplane
are :
(1) Rolling moment due to sideslip or
dihedral effect (lateral stability).
NAVWEPS, OO-ROT-80
STABILITY AND CONTROL
(2) Yawing moment due to sideslip or
static directional stability.
(3) Yawing moment due to rolling veloc-
ity or the adverse (or proverse) yaw.
(4) Rolling moment due to yawing ve-
locity-a cross effect similar to (3). If the
aircraft has a yawing motion to the right,
the left wing will move forward faster and
momentarily develop more lift than the
right and cause a rolling moment to the
right.
(3) Aerodynamic side force due to side-
slip.
(6) Rolling moment due to rolling ve-
locity or damping in roll.
(7) Yawing moment due yawing velocity
or damping in yaw.
(8) The moments of inertia of the air-
plane about the roll and yaw axes.
The complex interaction of these effects pro-
duces three possible types of motion of the
airplane: (a) a directional divergence, (b)
a spiral divergence, and (c) an oscillatory
mode termed Dutch roll.
Directional divergence is a condition which
cannot be tolerated. If the reaction to a small
initial sideslip is such as to create moments
which tend to increase the sideslip, directional
divergence will exist. The sideslip would in-
crease until the airplane is broadside to the
wind or structural failure occurs. Of course,
increasing the static directional stability re-
duces the tendency for directional divergence.
Spiral divergence will exist when the static
directional stability is very large when com-
pared with the dihedral effect. The character
of spiral divergence is by no means violent,
The airplane, when disturbed from the equilib-
rium of level flight, begins a slow spiral which
gradually increases to a spiral dive. When a
small sideslip is introduced, the strong direc-
tional stability tends to restore the nose into
the wind while the relatively weak dihedral
effect lags in restoring the airplane laterally,
In the usual case, the rate of divergence in the
299 | 316 | 316 | 00-80T-80.pdf |
NAVWEPS DGROT-50
STABBITY AND CONTROL
spiral motion is so gradual that the pilot can
control the tendency without difficulty.
Dutch roll is a coupled lateral-directional
oscillation which is usually dynamically stable
but is objectionable because of the oscillatory
nature. The damping of this oscillatory mode
may be weak or strong depending on the prop-
erties of the airplane. The response of the air-
plane to a disturbance from equilibrium is a
combined rolling-yawing oscillation in which
the rolling motion is phased to precede the
yawing motion. Such a motion is quite unde-
sirable because of the great havoc it would
create with a bomb, rocket, or gun platform.
Generally, Dutch roll will occur when the
dihedral effect is large when compared to static
directional stability. Unfortunately, Dutch
roll will exist for relative magnitudes of dihe-
dral effect and static directional stability be-
tween the limiting conditions for directional
divergence and spiral divergence. When the
dihedral effect is large in comparison with
static directional stability, the Dutch roll
motion has weak damping and is objectionable.
When the static directional stability is strong
in comparison with the dihedral effect, the
Dutch roll motion has such heavy damping
that it is not objectionable. However, these
qualities tend toward spiral divergence.
The choice is then the least of three evils.
Directional divergence cannot be tolerated,
Dutch roll is objectionable, and spiral diver-
gence is tolerable if the rate of divergence is
low. For this reason the dihedral effect should
be no more than that required for satisfactory
lateral stability. If the static directional sta-
bility is made adequate to prevent objection-
able Dutch roll, this will automatically be
sufficient to prevent directional divergence,
Since the more important handling qualities
are a result of high static directional stability
and minimum necessary dihedral effect, most
airplanes demonstrate a mild spiral tendency.
As previously mentioned, a weak spiral tend-
ency is of little concern to the pilot and cer-
tainly preferable to Dutch roll.
The contribution of sweepback to the lateral
dynamics of an airplane is significant. Since
the dihedral effect from sweepback is a function
of lift coefficient, the dynamic characteristics
may vary throughout the flight speed range.
When the swept wing airplane is at low C,, the
dihedral effect is small and the spiral tendency
may be apparent. When the swept wing air-
plane is at high C,, the dihedral effect is in-
creased and the Dutch Roll oscillatory tendency
is increased.
An additional oscillatory mode is possible
in the lateral dynamic effects with the rudder
free and the mode is termed a “snaking” oscil-
lation. This yawing oscillation is greatly
affected by the aerodynamic balance of the
rudder and requires careful consideration in
design to prevent light or unstable damping
of the oscillation.
CONTROL IN ROLL
The lateral control of an airplane is ac-
complished by producing differential lift on
the wings. The rolling, moment created by
the differential lift can be used to accelerate
the airplane to some rolling motion or control
the airplane in a sideslip by opposing dihedral
effect. The differential lift for control in
roll is usually obtained by some type of ailerons
or spoilers.
ROLLING MOTION OF AN AIRPLANE.
/ When an airplane is given a rolling motion in
flight, the wing tips move in a helical path
through the air. As shown in figure 4.29, a
rolling velocity to the right gives the right
wing tip a downward velocity component and
the left wing tip an upward velocity com-
ponent. By inspection of the motion of the
left wing tip, the velocity of the tip due to
roll combines with the airplane flight path
velocity to define the resultants motion. The
resulting angle between the flight path vector
and the resultant path of the tip is the helix
angle of roll. From the trigonometry of small
angles, the helix angle of roll can be defined as: | 317 | 317 | 00-80T-80.pdf |
Roll helix angle=&; (radians)
where
p=rate of roll, radians per second
6=wing span, ft.
V=airplane flight velocity, ft. per sec.
and, one radian=S7.3 degrees
pb Generally, the maximum values of rVobtained
by control in roll are approximately 0.1 to 0.07.
The helix angle of roll, $i, is, actually a com-
mon denominator of rolling performance.
The deflection of the lateral control surfaces
creates the differential lift and the rolling
moment to accelerate the airplane in roll. The
roll rate increases until an equal and opposite
moment is created by the resistance to rolling
motion or “damping in roll.” The second
illustration of figure 4.29 defines the source
of the damping in roll. When the airplane
is given a rolling velocity to the right, the
downgoing wing experiences an increase in
angle of attack due to the helix angle of roll.
Of course, the upgoing wing experiences a
decrease in angle of attack. In flight at angles
of attack less than that for maximum lift, the
downgoing wing experiences an increase in
lift and the upgoing wing experiences a de-
crease in lift and a rolling moment is developed
which opposes the rolling motion. Thus, the
steady state rolling motion occurs when the
damping moment equals the control moment.
The response of the airplane to aileron deflec-
tion is shown by the time history diagram of
figure 4.29. When the airplane is restrained
so that pure rolling motion is obtained, the
initial response to an aileron deflection is a
steady increase in roll rate. As the roll rate
increases so does the damping moment and the
roll acceleration decreases. Finally, the
damping moment approaches the control mo-
ment and a steady state roll rate is achieved.
NAVWEPS 00-6OT-60
STABILITY AND CONTROL
If the airplane is unrestrained and sideslip is
allowed, the affect of the directional stability
and dihedral effect can be appreciated. The
conventional airplane will develop adverse
yawing moments due to aileron deflection and
rolling motio6. Adverse yaw tends to produce
yawing displacements and sideslip but this is
resisted by the directional stability of the air-
plane. If adverse yaw produces sideslip, di-
hedral effect creates a rolling moment opposing
the roll and tends to reduce the roll rate. The
typical transient motions (A) and (B) of the
time history diagram of figure 4.29 show that
high directional stability with low dihedral
effect is the preferable combination. Such a
combination provides an airplane which has
no extreme requirement of coordinating aileron
and rudder in order to achieve satisfactory
rolling performance. While the coupled mo-
tion of the airplane in roll is important,
further discussion of lateral control will be
directed to pure uncoupled rolling performance.
ROLLING PERFORMANCE. The required
rolling performance of an airplane is generally
specified as certain necessary values of the roll 1
helix angle, &I$ However, in certain condi-
tions of flight, it may be more appropriate to
specify minimum times for the airplane to
accelerate through a given angle of roll.
Usually, the maximum value of 2% should be
on the order of 0.10. Of course, fighters and
attack airplanes have a more specific require-
ment for high rolling performance and 0.09
Pb may be considered a minimum necessary 2v.
Patrol, transport, and bomberairplaneshaveless
requirement for high rolling performance and a
Pb 2-V of 0.07 may be adequate for these types.
The ailerons or spoilers must be powerful
Pb enough to provide the required rV’ While
the size and effectiveness of the lateral control
devices is important, consideration must be
301
Revised January 1965 | 318 | 318 | 00-80T-80.pdf |
NAVWEPS OO-80T-80
STABILITY AND CONTROL
HELIX ANGLE OF ROLL
IP VELOCITY,$
RCUING VELOCITY, P
TIP VELOCITY WE TO ROLL
RESULTANT PATH
( RADIANS 1
DAMPING IN ROLL
STARBOARD WING
AIRPLANE RESPONSE TO AILERON DEFLECTION
PIRPLANE RESTRAINED
TO ROLLING MOTION ONLY
------(A) HIGH DlRECTlCNAL STABILITY
m DIHEDRAL EFFECT
AIRPLANE UNRESTRAINED
\ AND FREE TO SIDESLIP
\ (RUDDER FIXED) ( B ) LOW DIRECTIONAL STABIUTY
.---A HIGH MHEDRAL EFFECT
w TIME, SECONDS
Figure 4.29. Rolling Performance
302 | 319 | 319 | 00-80T-80.pdf |
given to the airplane size. For geometrically
similar airplanes, a certain deflection of the
I!!. ailerons will produce a fixed value of zlr mde-
pendent of the airplane size. However, the
roll rate of the geometrically similar airplanes
at a given speed will vary inversely with the
span, b.
If
Pb -
~-constant
p=(constant) 7
( )
Thus, the smaller airplane will have an ad-
vantage in roll rate or in time to accelerate
through a prescribed angle of roll. For ex-
ample, a one-half scale airplane will develop
twice the rate of roll of the full scale airplane.
This relationship points to the favor of the
small, short span airplane for achieving high
roll performance.
An important variable affecting the rate of
roll is the true airspeed or flight velocity, V.
If a certain deflection of the ailerons creates a
Pb specific value of -7 the rate of roll varies 2V
directly with the true airspeed. Thus, if the
roll helix angle is held constant, the rate of
roll at a particular true airspeed will not be
affected by altitude. The linear variation of
roll rate with airspeed points out the fact that
high roll rates will require high airspeeds.
The low roll rates at low airspeeds are simply
a consequence of the low flight speed and this
condition may provide a critical lateral con-
trol requirement for satisfactory handling
qualities.
Figure 4.30 illustrates the typical rolling
paformance of a low speed airplane. When
the ailerons are at full deflection, the maximum
roll helix angle is obtained. The rate of roll
increases linearly with speed until the control
forces increase to limit of pilot effort and full
control deflection cannot be maintained. Past
NAVWEPS OO-BOT-BO
STABILIJY AND CONTROL
the critical speed, with some limited amount
of force applied by the pilot (usually the limit
of lateral force is assumed to be 30 lbs.), the
Pb ailerons cannot be held at full deflection, ~~
drops, and rate of roll decreases. In this exam-
ple, the rolling performance at high speeds is
limited by the ability of the pilot to maintain
full deflection of the controls. In an effort to
reduce the aileron hinge moments and control
forces, extensive application is made of aerody-
namic balance and various tab devices. How-
ever, 100 percent aerodynamic balance is not
always feasible or practical but a sufficient
Pb value of - must be maintained at high speeds. ZV
Rather than developing an extensive weight
lifting program mandatory for all Naval
Aviators, mechanical assistance in lateral con-
trol can be provided. If a power boost is
provided for the lateral control system, the
rolling performance of the airplane may be
extended to higher speeds since pilot effort
will not be a limiting factor. The effect of a
power boost is denoted by the dashed line
extensions of figure 4.30. A full powered,
irreversible lateral control system is common
for high speed airplanes. In the power oper-
ated system there is no immediate limit to the
deflection of the control surfaces and none of
the aberrations in hinge moments due to com-
pressibility are fed back to the pilot. Control
forces are provided by the stick centering
lateral bungee or spring.
A problem particular to the high speed is
due to the interaction of aerodynamic forces
and the elastic deflections of the wing in
torsion. The deflection of ailerons creates
twisting moments on the wing which can cause
significant torsional deflections of the wing.
At the low dynamic pressures of low flight
speeds, the twisting moments and twisting
deflections are too small to be of importance.
However, at high dynamic pressures, the
deflection of an aileron creates significant
303 | 320 | 320 | 00-80T-80.pdf |
NAVWEPS 00-807-80
STABILITY AND CONTROL
P,
RAl ^,
r%“LL
O/SEC.
0
,/ <EiECT
0 OF ADDED
POWER
BOOST
V. KNOTS
4 ROLL .lD
HELIX
ANGLE
pb
TT
-
V, KNOTS
A
AILERON ----- DEFLECTlON
8,
V.KNOTS
SPEED CORRESPONDING
TO LIMIT OF PILOT EFFORT
TO MAINTAIN MAXIMUM DEFLECTION
A (3 z 3 is
1.0
0 c 5 ELASTIC WING TWISTING
REVERSAL
Figure 4.30. Control in Roll
304 | 321 | 321 | 00-80T-80.pdf |
twisting deflections which reduce the effec-
tiveness of the aileron, e.g., downward deflec-
tion of an aileron creates a nose down twist of
the wing which reduces the rolling moment
due to aileron deflection. At very high speeds,
the torsional deflection of the wing may be
so great than a rolling moment is created
opposite to the direction controlled and “aile-
ron reversal” occurs. Prior to the speed for
aileron reversal, a serious loss of roll helix
angle may be encountered. The effect of this
aeroelastic phenomenon on rolling perform-
ance is illustrated in figure 4.30.
To counter the undesirable inceractiuo be-
tween aerodynamic forces and wing torsional
deflections, the trailing edge ailerons may be
moved inboard to reduce the portion of the
span subjected to twisting moments. Of
course, the short span, highly tapered wing
planform is favorable for providing relatively
high stiffness. In addition, various configura-
tions of spoilers may be capabIe of producing
the required rolling performance without the
development of large twisting moments.
CRITICAL REQUIREMENTS, The critical
conditions for requiring adequate lateral con-
trol power may occur at either high speed or
low speed depending on the airplane configura-
tion and intended use. In transonic and super-
sonic flight, compressibility effects tend to
reduce the effectiveness of lateral control de-
vices to produce required roll helix angles.
These effects are most significant when com-
bined with a loss of control effectiveness due to
aeroelastic effects. Airplanes designed for
high speed flight must maintain suflicient
lateral control effectiveness at the design dive
speed and this is usually the predominating
requirement.
During landing and takeoff, the airplane
must have adequate lateral control power to
contend with the ordinary conditions of flight.
The lateral controls must be capable of achiev-
ing required roll helix angles and acceleration
through prescribed roll dispIacements. Also,
the airplane must be capable of being con-
NAVWEPS OO-UOT-80
STABILITY AND CONTROL
trolled in a sideslip to accomplish crosswind
takeoff and landing. The lateral control dur-
ing crosswind takeoff and landing is a par-
ticular problem when the dihedral effect is
high. Since the sweepback contributes a large
dihedral effect at high lift coefficients, the
problem is most important for the airplane
with considerable sweepback. The limiting
crosswind components must be given due re-
spect especially when the airplane is at low
gross weight. At low gross weight the speci-
fied takeoff and landing speeds will be low and
the controlled angle of sideslip will be largest
for a given crosswind velocity.
MISCELLANEOUS STABILITY PROBLEMS
There are several general problems of flying
which involve certain principles of stability as
well as specific areas of longitudinal, direc-
tional and lateral stability. Various condi-
tions of flight will exist in which certain
problems of stability (or instability) are un-
avoidable for some reason or another. any
of the following items deserve consideration
because of the possible unsafe condition of flight
and the contribution to an aircraft accident.
LANDING GEAR CONFIGURATIONS
There are three general configurations for the
aircraft landing gear: the tricycle, bicycle, and
“conventional” tail wheel arrangement. At
low rolling speeds where the airplane aerody-
namic forces are negligible, the “control-fixed”
static stability of each of these configurations
is determined by the side force characteristics
of the tires and is not a significant problem.
The instability which allows ground loops
in an aircraft with a conventional tail wheel
landing gear is quite basic and can be appre-
ciated from the illustration of figure 4.31. Cen-
trifugal force produced by a turn must be
balanced and the aircraft placed in equilibrium.
The greatest side force is produced at the main
wheels but to achieve equilibrium with the | 322 | 322 | 00-80T-80.pdf |
NAVWEPS oo-SOT-80
STABILITY AND CONTROL
-
“CONVENTIONAL’
TAIL WHEEL
CONFIGURATION
SIDE FORCE ON
MAIN WHEELS
CENTRIFUGAL FORCE
TRICYCLE
\\ CONFIGURATION
--BALANCING
NOSE WHEEL
SIDE FORCE
CENTRIFUGAL FORCE
BICYCLE CONFIGURATION
FORCE
Figure 4.31. Landing Gear Configurations | 323 | 323 | 00-80T-80.pdf |
center of gravity aft of the main wheels a bal-
ancing load on the tail wheel must be produced
toward the center of turn. When the tail
wheel is free to swivel, the equilibrium of the
turn requires a control force opposite to the
direction of turn-i.e.. control force insta-
bility. The inherent stability problem exists
because the center of gravity is aft of the point
where the main side forces are developed. This
condition is analogous to the case of static
longitudinal stability with the center of
gravity aft of the neutral point.
The conventional tail wheel configuration
has this basic instability or ground loop tend-
ency which must be stabilized by the pilot.
At high rolling speeds where aerodynamic
forces are significant, the aerodynamic direc-
tional stability of the airplane resists the
ground looping tendency. The most likely
times for a ground loop exist when rolling
speeds are not high enough to provide a con-
tribution of the aerodyhamic forces. When the
tail wheel is free to swivel or when the normal
force on the tail wheel is small, lack of pilot
attention can allow the ground loop to take
place.
The tricycle landing gear configuration has
an inherent stability d,ue to the relative posi-
tion of the main wheels and the center of
gravity. Centrifugal force produced by a
turn is balanced by the side force on the main
wheels and a side force on the nose wheel in
the direction of turn. Note that the freeing
the nose wheel to swivel produces moments
which bring the aircraft out of the turn. Thus,
the tricycle configuration has a basic stability
which.is given evidence by control displace-
ment and a wheel side force in the direction
of turn. Because of the contrast in stability,
the tricycle configuration is much less difficult
to maneuver than the tail wheel configuration
and does not provide an inherent ground loop
tendency. However, a steerable nose wheel
is usually necessary to provide satisfactory
maneuvering capabilities.
NAVWEPS DD-BDT-80
STABILITY AND CONTROL
The bicycle configuration of landing gear
has stability characteristics more like the
automobile. If directional control is ac-
complished with the front wheels operated
by power controls, no stability problem exists
at low speeds. A problem can exist when the
airplane is at high speeds because of a distribu-
tion of normal force being different from the
ordinary static weight distribution. If the
airplane is held onto the runway at speeds
well above the normal takeoff and landing
speeds, the front wheels carry a greater than
ordinary amount of normal force and a tend-
ency for instability exists. However, at these
same high speeds the rudder is quite powerful
and the condition is usually well within
control.
The basically stable nature of the tricycle
and bicycle landing gear configurations is best
appreciated by the ease of control and ground
maneuvering of the airplane. Operation of
a conventional tail wheel configuration after
considerable experience with tricycle cohfigu-
rations requires careful consideration af the
stability that must be furnished by the pilot
during ground maneuvering.
SPINS AND PROBLEMS OP SPIN
RECOVERY
The motion of an airplane in a spin can
involve many complex aerodynamic and in-
ertia forces and moments. However, there are
certain fundamental relationships regarding
spins and spin recoveries with which all
aviators should be familiar. The spin differs
from a spiral dive in that the spin always
involves flight at high angle of attack while
the spiral dive involves a spiral motion of
the airplane at relatively low angle of attack.
The stall characteristics and stability of
the airplane at high lift coefficients are im-
portant in the initial tendencies of the airplane.
As previously mentioned, it is desirable to
have the wing initiate stall at the root first
rather than tip first. Such a stall pattern
prevents the undesirable rolling moments at
high lift coeGients, provides suitable stall | 324 | 324 | 00-80T-80.pdf |
_~. ,,
. | 325 | 325 | 00-80T-80.pdf |
warning, and preserves lateral control effec-
tiveness at high angles of attack. Also, the
airplane must maintain positive static longi-
tudinal stability at high lift coe&ients and
should demonstrate satisfactory stall recovery
characteristics.
In order to visualize the principal effects of
an airplane entering a spin, suppose the air-
plane is subjected to the rolling and yawing
velocities shown in figure 4.32. The yawing
velocity to the right tends to produce higher
local velocities on the left wing than on the
right wing. The rolling velocity tends to
increase the angle of attack for the downgoing
right wing (a,) and. decrease the angle of
attack for the upgoing left wing (al). At
airplane angles of attack below the stall this
relationship produces roll due to yaw, damping
in roll, etc., and some related motion of the
airplane in unstalled flight. However, at
angles of attack above the stall, important
changes take place in the aerodynamic char-
acteristics.
Figure 4.32 illustrates the aerodynamic
characteristics typical of a conventional air-
plane configuration, i.e., moderate or high
aspect ratio and little-if any-sweepback.
Ifs this airplane is provided a rolling displace-
ment when at some angle of attack above
the stall, the upgoing wing experiences a
decrease in angle of attack with a correspond-
ing increase in C, and decrease in C,,. In other
words, the upgoing wing becomes less stalled.
Similarly, the downgoing wing experiences
an increase in angle of attack with a corre-
sponding decrease in CL and increase in CD. Es-
sentially, the downgoing wing becomes more
stalled. Thus, the rolling motion is aided
rather than resisted and a yawing moment is
produced in the direction of roll. At angles
of attack below stall the rolling motion is
resisted by damping in roll and adverse yaw
is usually present. At angles of attack above
the stall, the damping in roll is negative and
a rolling motion produces a rolling moment
in the direction of the roll. This negative
NAVWEPS OO-BOY-BO
STABIUTY AND CoMml
damping in roll is generally referred to as
“autorotation.”
When the conventional airplane is stalk4
and some rolling-yawing displacement takes
place, the resulting autotiotation rolling mo-
ments and yawing moments start the airplane
into a self-sustaining rolling-yawing motion.
The autorotation rolling and yawing tenden-
cies of the airplane at high angles of attack
are the principal prospin moments of the
conventional airplane configuration and these
tendencies accelerate the airplane into the
spin until some limiting condition exists.
The stabilized spin is not necessaray a simple
steady vertical spiral but may involve some
coupled unsteady oscillatory motion.
An important characteristic of the mote
conventional airplane configuration is that the
spin shows a predominating contribution of
the autorotation tendency. Generally, the
conventional configuration has a spin motion
which is primarily rolling with moderate yaw.
High directional stability is favorable since it
will limit or minimize the yaw displacement
of the spinning airplane.
The fundamental requirement of the spin is
that the airplane be placed at an excessive
angle of attack to produce the autorotation
rolling and yawing tendencies. Generally
speaking, the conventional airplane must be
stalled .before a spin can take place. This
relationship establishes a fundamental p&r-
ciple of recovery-the airplane must be un-
stalled by decreasing the wing angle of attack.
The most dfective procedure for the conven-
tional configuration is to use opposite rudder
to stop the sideslip, then lower the angle of
attack with the elevators. With sufficient
rudder power this procedure will produce a
positive recovery with a minimum loss of
altitude. Care should be taken during pullout
from the ensuing dive to prevent excessive
angle of attack and entry into another spin.
It should be appreciated that a spin is always
a possible corollary of a stall and the self-
sustaining motion of a spin will take place at | 326 | 326 | 00-80T-80.pdf |
NAVWEPS OO-BOT-80
STABILITY AND CONTROL
YAWING
VELOCITY
ROLLING
VELOCITY
AERODYNAMIC CHARACTERISTICS TYPICAL OF AERODYNAMIC CHARACTERISTICS TYPICAL OF
A CONVENTIONAL CONFIGURATION A CONVENTIONAL CONFIGURATION
I--- I---
STALL STALL
CL CL
AND AND
CD CD
I 0, ANGLE OF ATTACK QL OR
t TYPICAL Q’ . ..I,.” cncrn lr H “lU” a-LL” A CD
AERODYNAMIC CHARACTERISTICS
co NFIGURATION
I
a, ANGLE OF ATTACK OL aR
Figure 4.32. Spin Characteristics
310 | 327 | 327 | 00-80T-80.pdf |
excessive angles of attack. Of course, a low
speed airplane could be: designed to be spin-
proof by making it stallproof. By limiting
the amount of control deflection, the airplane
may not have the longitudinal control power
to trim to maximum lift angle of attack. Such
a provision may be possible for certain light
planes and commercial aircraft but would
create an unrealistic and impractical limita-
tion on the utility of a military airplane.
The modern high speed airplane configura-
tion is typified by low aspect ratio, swept wing
planforms with relatively large yaw and pitch
inertia. The aerodynamic characteristics of
such a configuration are shown in figure 4.32.
The lift curve (C, versus U) is quite shallow at
high angles of attack and maximum lift is not
clearly defined. When this type of airplane is
provided a rolling motion at high angles of
attack, relatively small changes in C, take
place. When this effect is combined with the
relatively short span of this type airplane, it is
apparent that the wing autorotation contribu-
tion will be quite weak and will not be a pre-
dominating pro-spin moment. The relatively
large changes in drag coefficient with rolling
motion imply .a predominance of yaw for the
spin of the high speed airplane configuration.
Actually, various other factors contribute
to the predominating yaw tendency for the
spin of the modern airplane configuration.
The static directional stability deteriorates at
high angles of attack and may be so weak that
extemely large yaw displacements result. In
certain instances, very high angles of attack
may bring such a decay in directional stability
that a “slice” or extreme yaw displacement
takes place before a true spin is apparent. At
these high angles of attack, the adverse yaw
due to roll and aileron deflection can be very
strong and create large yaw displacements of
the airplane prior to realizing a stall.
The aircraft with the relatively large, long
fuselage can exhibit a significant moment con-
tribution from the fuselage alone. The cross
flow pattern on the fuselage at high angles of
NAWWEPS DO-BOT-BO
STABILITY AND CONTROL
attack is capable of producing pro-spin mo-
ments of considerable magnitude which con-
tribute to the self-sustaining nature of the
spin. Also, the large distributed mass of the
fuselage in rolling-yawing rotation contributes
to inertia moments which flatten the spin and
place the aircraft at extreme angles of attack.
The spin recovery of the modern high speed
airplane involves principles which are similar
to those of the spin recovery of the conven-
tional airplane. However, the nature of the
spin for the modern configuration may involve
specific differences in technique necessary to
reduce the sideslip and angle of attack. The
use of opposite rudder to control the sideslip
and effect recovery will depend on the effective-
ness of the rudder when the airplane is in the
spin. At high positive angles of attack and
high sideslip the rudder effectiveness may be
reduced and additional anti-spin moments must
be provided for rapid recovery. The deflection
of ailerons into the spin reduces the autorota-
tion rolling moment and can produce adverse
yaw to aid the rudder yawing moment in
effecting recovery.
There may be many other specific differences
in the technique necessary to effect spin re-
covery . The effectiveness of the rudder during
recovery may be altered by the position of
elevators or horizontal tail. Generally, full
aft stick may be necessary during the initial
phase of recovery to increase the effectiveness
of the rudder. The use of power during the
spin recovery of a propeller powered airplane
may or may not aid recovery depending on the
specific airplane and the particular nature of
the slipstream effects. The use of power during
the spin recovery of a jet powered airplane
induces no significant or helpful flow but does
offer the possibility of a severe compressor
stall and adverse gyroscopic moments. Since
the airplane is at high angle of attack and
sideslip, the flow at the inlet may be very
poor and the staI1 limits considerably reduced.
These items serve to point out possible dif-
ferences in technique required for various con-
figurations. The spin recovery specific for
31.1 | 328 | 328 | 00-80T-80.pdf |
NAVWEPS woT-80
STABILITY AND CONTROL
-UNSTABLE
I
v
w
CL
PITCH-UP
NEUTRAL
SEPARATION OR
STALLTIP FIRST
RD SHIFT OF VORTEX
INCREASE IN LOCAL
DDWNWAM AT TAIL
: :
4b FUSELAGE CROSS-
FLOW SEPARATION
VORTICES INCREASE
LOCAL DOWNWASH AT TAIL
Figure 4.33. Pitch-up
312 | 329 | 329 | 00-80T-80.pdf |
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 | 330 | 330 | 00-80T-80.pdf |
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
314 | 331 | 331 | 00-80T-80.pdf |
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
315 | 332 | 332 | 00-80T-80.pdf |
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
316 | 333 | 333 | 00-80T-80.pdf |
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
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335 | 335 | 00-80T-80.pdf |
|
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
319 | 336 | 336 | 00-80T-80.pdf |
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
320 | 337 | 337 | 00-80T-80.pdf |
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
321 | 338 | 338 | 00-80T-80.pdf |
NAVWEPS DD-80T-80
STABILITY AND CONTROL
TANDEM ROTOR LONGITUDINAL CONTROL
TANDEM ROTOR DIRECTIONAL CONTROL
AFT
ROTOR
*JR
9
F”&%iD
Fig&e 4.36. longitudinal and Directional Control
322 | 339 | 339 | 00-80T-80.pdf |
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
323 | 340 | 340 | 00-80T-80.pdf |
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.
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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.
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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
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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
Q
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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NAVWEPS oo-8oT-80
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
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SNOIlVlIWll HlOM3US ONllVU3dO
08-108-00 Sd3MAVN | 346 | 346 | 00-80T-80.pdf |
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.
330 | 347 | 347 | 00-80T-80.pdf |
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
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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,
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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 | 350 | 350 | 00-80T-80.pdf |
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,
334 | 351 | 351 | 00-80T-80.pdf |
LOAD
FACTOR,
n
12-
II-
IO-
9-
B-
7-
6-
5-
4-
3
2-
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 | 352 | 352 | 00-80T-80.pdf |
STRUCTURAL FAILURE
12-
II-
IO-
9-
a-
7.5
7- UP
6- !
5-
4- MAXIMUM
IN,,lCATF . . I -. . - - - -
500
NEGATIVE LIFT
CAPABILITY
Figure 5.4. Signikance o\ the V-n Diagram | 353 | 353 | 00-80T-80.pdf |
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 | 354 | 354 | 00-80T-80.pdf |
355 | 355 | 00-80T-80.pdf |
|
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
339 | 356 | 356 | 00-80T-80.pdf |
NAVWEPS OO-SOT-80
OPERATING STRENGTH LIMITATIONS
-
RIGID WING ELASTIC WING
A
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)
340 | 357 | 357 | 00-80T-80.pdf |
NAVWEPS 00-SOT-80
OPERATING STRENGTH LIMITATIONS
WING
ROOT’ /-TRAILING EDGE
9- LEADING EDGE
Figure 5.5. Aeroelastic Effects (Sheet 2 of 2)
341 | 358 | 358 | 00-80T-80.pdf |
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
342 | 359 | 359 | 00-80T-80.pdf |
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)’
w
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
343 | 360 | 360 | 00-80T-80.pdf |
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
344 | 361 | 361 | 00-80T-80.pdf |
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
345 | 362 | 362 | 00-80T-80.pdf |
NAVWEPS 00-80T-80
OPERATING STRENGTH LIMITATIONS
ATTACHMENT FITTING FATiGUE FAILURES FATIGUE CRACKS IN STRUCTURAL SAMPLE
Figure 5.7. Examples of Fatigue Failures
346 | 363 | 363 | 00-80T-80.pdf |
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.
347 | 364 | 364 | 00-80T-80.pdf |
SNOllVlIWll H13N3US ONllVU3dO
08-108-00 Sd3MAVN | 365 | 365 | 00-80T-80.pdf |
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
349 | 366 | 366 | 00-80T-80.pdf |
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:
350 | 367 | 367 | 00-80T-80.pdf |
NAVWEPS OO-80T-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
LIFT
COEFFICIENT
CL
POWER
REQUIRED
0
AVAILABLE
HI?
\
FOR LIFT EQUAL
TO WEIGHT,
“= 17.2J$-
gs
a
ANGLE OF ATTACK
FOR A STEADY CLIMB,
t
ROC = 33,000
FPM POWER REQ’D-
POWER EXCESS
\ A 7- IER AVAILABLE
‘HIGH WILL ESTABLISH
PO’lh
w...-..
SNCY
..-- -_... - -.-.
LEVEL FLIGHT AT
--
t
I VELOCITY, KNOTS A 0 Figure 6.1. Primary Control of Airspeed and Altitude
351 | 368 | 368 | 00-80T-80.pdf |
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-
352 | 369 | 369 | 00-80T-80.pdf |
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,
353
Revised January 1965 | 370 | 370 | 00-80T-80.pdf |
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
354
Revised January 1965 | 371 | 371 | 00-80T-80.pdf |
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 | 372 | 372 | 00-80T-80.pdf |
‘:
-.-
* ,-.
. :,,.
_,: .-,A* | 373 | 373 | 00-80T-80.pdf |
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.
357 | 374 | 374 | 00-80T-80.pdf |
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
358 | 375 | 375 | 00-80T-80.pdf |
LIFT
COEFFICIENT
CL
CL MAX -
NAVWEPS OD-BOT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
7 A
- STALL
ANGLE OF ATTACK, 0
3s9 | 376 | 376 | 00-80T-80.pdf |
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
360 | 377 | 377 | 00-80T-80.pdf |
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
CL
ANGLE OF ATTACK, a
Figure 6.4. The Approach and Landing | 378 | 378 | 00-80T-80.pdf |
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
362 | 379 | 379 | 00-80T-80.pdf |
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. | 380 | 380 | 00-80T-80.pdf |
MAVWEPS OO-BOLBO
APPLICA’IIOM OF AERODYNAMICS
10 SPECIFIC PROBLEMS OF FLYI~NG | 381 | 381 | 00-80T-80.pdf |
Excess airspeed at landing is just as undesira-
ble as a deficiency of airspeed. An cxccssivc
airspeed at landing will produce an undesirable
increase in landing distance and the energy to
be dissipated by the brakes for the field landing
or excessive arresting loads for theshipboard
landing. In addition, the excess airspeed is a
corollary of too low an angle of attack and the
airplane may contact the deck or runway nose
wheel first and cause damage to the nose wheel
or begin a porpoising of the airplane. During
a flare to landing, any excess speed will be
difficult to dissipate due to the reduction of
drag due to ground effect. Thus, if the air-
plane is held off with excess airspeed the air-
plane will “float” with the consequence of a
barrier engagement, barricade engagement,
bolter, or considerable runway distance used
before touchdown.
A fundamental requirement for a good land-
ing is a well planned and executed approach.
The possibility of errors during the landing
process is minimized when the airplane is
brought to the point of touchdown with the
proper glide path and airspeed. With the
proper approach, there is no need for drastic
changes in the flight path, angle of attack, or
power setting to accomplish touchdown at the
intended point on the deck or runway. Late
corrections to line up with the deck or diving
for the deck are common errors which eventu-
ally result in landing accidents. Accurate
control of airspeed and glide path are ab-
solutely necessary and the LSO, angle of attack
indicator, and the mirror landing system pro-
vide great assistance in accurate control of the
airplane.
THE TAKEOFF
As in the case of landing, the specific tech-
niques necessary may vary greatly between
various types of airplanes and various oper-
ations but certain fundamental principles will
be common to all airplanes and all operations.
The specific procedures recommended for each
airplane type must be followed exactly to
MAVWEPS OD-BOT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
insure a consistent, safe takeoff flying tech-
nique.
TAKEOFF SPEED AND DISTANCE. The
takeoff speed of any airplane is some mini-
mum practical airspeed which allows sufficient
margin above stall and provides satisfactory
control and initial rate of climb.. Depending
on the airplane characteristics, the takeoff
speed will be some value 5 to 25 percent above
the stall or minimum control speed. As such,
the takeoff will be accomplished at a certain
value of lift coefficient and angle of attack
specific to each airplane configuration. As
a result, the takeoff airspeed (FAX or CM) of
any specific airplane configuration is a function
of the gross weight at takeoff. Too low an
airspeed at takeoff may cause stall, lack of
adequate control, or poor initial climb per-
formance. An excess of speed at takeoff may
provide better control and initial rate of climb
but the higher speed requires additional dis-
tance and may provide critical conditions for
the tires.
The takeoff distance of an airplane is affected
by many different factors other than technique
and, prior to takeoff, the takeoff distance
must be determined and compared with the
runway length available. The principal factors
affecting the takeoff distance are as follows:
(1) The gross weight of the airplane has
a considerable effect on takeoff distance be-
cause it affects both takeoff speed and ac-
celeration during takeoff roll.
(2) The surface winrls must be considered
because of the powerful effect of a headwind
or tailwind on the takeoff distance. In the
case of the crosswind, the component of
wind along the runway will be the effective
headwind or tailwind velocity. In addi-
tion, the component of wind across the run-
way will define certain requirements of lateral
control power and the limiting compo-
nent wind must not be exceeded.
(3) Pressure altitude and temperature can
cause a large effect on takeoff distance, es-
pecially in the case of the turbine powered | 382 | 382 | 00-80T-80.pdf |
NAVWEPS OD-SOT-80
APPLICATION OF AERODYNAMICS,
TO SPECIFIC PROBLEMS OF FLYING
airplane. Density altitude will determine
the true airspeed at takeoff and can affect
the takeoff acceleration by altering the
powerplant thrust. The effect of tempeta-
ture alone is important in the case of the
turbine powered aircraft since inlet air tem-
perature will affect powerplant thrust. Ic
should be noted that a typical turbojet ait-
plane may he approximately twice as sensi-
tive to density altitude and five to ten times
as sensitive to temperature as a representa-
cive reciprocating engine powered airplane.
(4) Specific humidity must be accounted
for in the case of the reciprocating engine
powered airplane. A high water vapor
content in the air will cause a definite reduc-
tion in takeoff power and takeoff acceler-
ation.
(5) The runluay condition will deserve con-
sideration when the takeoff acceleration is
basically low. The runway slope must be
compared carefully with the surface winds
because ordinary values of runway slope will
usually favor choice of the runway with
headwind and upslope rather than down-
slope and tailwind. The surface condition
of the runway has little bearing on takeoff
distance as long as the runway is a hard
surface.
Each .of these factors must be accounted
for and the takeoff distance properly com-
puted for the existing conditions. Since
obstacle clearance distance is generally a
function of the same factors which affect
takeoff distance, the obstacle clearance dis-
tance is usually related as some proportion
of the takeoff distance. Of course, the take-
off and obstacle clearance distances related
by the handbook data will be obtained by
the techniques and procedures outlined in the
handbook.
TYPICAL ERRORS. The takeoff distance
of an airplane should be computed for each
takeoff. A most inexcusable error would be to
attempt takeoff from a runway of insufficient
length. Familiarity with the airplane hand-
book performance data and proper accounting
of weight, wind, altitude, temperature, etc.,
are necessary parts of flying. Conditions of
high gross weight, high pressure altitude and
temperature, and unfavorable winds create the
extreme requirements of runway length, espe-
cially for the turbine powered airplane. Under
these conditions, use of the handbook data is
mandatory and no guesswork can be tolerated.
One typical.etror of takeoff technique is the
premature or excess pitch rotation of the air-
plane. Premzture or excm pitch rotation of the
airplane may seriously reduce the takeoff accel-
eration and increase the takeoff distance. In
addition, when the airplane is placed at an
excessive angle of attack during takeoff, the
airplane may become airborne at too low a
speed and the result may be a stall, lack of ade-
quate control (especially in a crosswind), or
poor initial climb performance. In fact. there
are certain low aspect ratio configurations of
airplanes which, at an excessive angle of ar-
tack, will not fly out of ground effect. Thus,
over-rotation of the airplane ,during takeoff
may hinder takeoff acceleration or the.initial
climb. It is quite typical for an airplane to be
placed at an excess angle of attack and become
airborne prematurely then settle back fo rhe
runway. When the proper angle of attack is
assumed, the airplane simply accelerates to the
takeoff speed and becomes airborne wirh suf-
ficient initial rate of climb. In this sense, the
appropriate rotation and takeoff speeds or an
angle of attack indicator must be used.
If the airplane is subject to a sudden pull-up
or Jteep tzzra after becoming airborne, rhe,resulr
may be a stall, spin, or reduction in initial rate
of climb. The increased angle of attack may
exceed the critical angle of attack or the in-
crease in induced drag may be quite large. For
this reason, any clearing turns made immedi-
ately after takeoff or deck launch must be slight
and well within the capabilities of the air-
plane.
366 | 383 | 383 | 00-80T-80.pdf |
In order to obviate some of the problems
of a deficiency of airspeed at takeoff, usual
result can be an excess of airspeed at takeoff.
The principal effect of an BXCBJS takeoff air@pssd
is the greater takeoff distance which results.
The general effect is that each 1 percent excess
takeoff velocity incurs approximately 2 per-
cent additional takeoff distance. Thus, excess
speed must be compared with the additional
runway required to produce the higher speed.
In addition, the aircraft tires may be subject
to critical loads when the airplane is at very
high rolling speeds and speeds in excess of a
basically high takeoff speed may produce
damage or failure of the tires.
As with the conditions of landing, excess
velocity or deficiency of velocity at takeoff
is undesirable. The proper takeoff speeds and
angle of attack must be utilized to assure
satisfactory takeoff performance.
GUSTS AND WIND SHEAR
The variation of wind velocity and direction
throughout the atmosphere is important be-
cause of its effect on the aerodynamic forces
and moments on an airplane. As the airplane
traverses this variation of wind velocity and
direction during flight, the changes in airflow
direction and velocity create changes in the
aerodynamic forces and moments and produce
a response of the airplane. The variation of
airflow velocity along a given direction exists
with shear parallel to the flow direction.
Hence, the velocity gradients are often re-
ferred to as the wind “shear.”
The effect of the vertical gust has important
effects on the airplane at high speed because
of the possibility of damaging flight loads.
The mechanism of vertical gust is illustrated
in figure 6.5 where the vertical gust velocity
is added vectorially to the flight velocity to
produce some resultant velocity. The principal
effect of the vertical gust is to produce a change
in airplane angle of attack, e.g., a positive
(up) gust causes an increase in angle of attack
NAVWEPS DD-BOT-BD
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
while a negative (down) gust causes a de-
crease in angle of attack. Of course, a change
in angle of attack will effect a change in lift
and, if some critical combination of high gust
intensity and high flight speed is encountered,
the change in lift may be large enough to
cause structural damage.
At low flight speeds during approach, land-
ing, and takeoff, the effect of the vertical gust
is due to the same mechanism of the change
in angle of attack. However, at these low
flight speeds, the problem is one of possible
incipient stalling and sinking rather than
overstress. When the airplane is at high
angle of attack, a further increase in angle of
attack due to a gust may exceed the critical
angle of attack and cause an incipient stalling
of the airplane. Also, a decrease in angle
of attack due to a gust will cause a loss of lift
and allow the airplane to sink. For this
reason, any deficiency of airspeed will be quite
critical when operating in gusty conditions.
The effect of the hori<oonral gust differs from
the effect of the vertical gust in that the im-
mediate effect is a change of airspeed rather
than a change in angle of attack. In this
sense, the horizontal gust is of little conse-
quence in the major airplane airloads and
strength limitations. Of greater significance
is the response of the airplane to horizontal
gusts and wind shear when operating at low
flight speeds. The possible conditions in
which an airplane may encounter horizontal
gusts and wind shear are illustrated in figure
6.5. As the airplane traverses a shear of wind
direction, a change in headwind component
will exist. Also, a climbing or descending
airplane may traverse a shear of wind velocity,
i.e., a wind profile in which the wind velocity
varies with altitude.
The response of an airplane is much de-
pendent upon the airplane characteristics but
certain basic effects are common to all ait-
planes. Suppose that an airplane is estab-
lished in steady, level flight with lift equal to
weight, thrust equal to drag, and trimmedso
367 | 384 | 384 | 00-80T-80.pdf |
NAVWEPS OG-8OT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
EFFECT OF VERTICAL GUST
CHANGE IN ANGLE OF ATTACK
OF WIND
-1
VERTICAL VARll
TRANSIENT
CONDITION FROM
WIND SHEAR
STEADY
LEVEL LIFT
FLIGHT
OR HORIZONTAL GUST 1 LIFT I
Figure 6.5. Effect of Wind Shear
368 | 385 | 385 | 00-80T-80.pdf |
there is no unbalance of pitching, yawing, or
rolling moment. If the airplane traverses a
sharp wind shear equivalent to a horizontal
gust, the resulting change in airspeed will
disturb such an equilibrium. For example, if
the airplane encounrers a sharp horizontal
gust which reduces the airspeed 20percent,the
new airspeed (80 percent of the original value)
produces lift and drag at the same angle of
attack which are 64 percent of the original
value. The change in these aerodynamic forces
would cause the airplane to accelerate in the
direction of resultant unbalance of force.
That is, the airplane would accelerate down
and forward until a new equilibrium is
achieved. In addition, there would be a
change in pitching moment which would
produce a response of the airplane in pitch.
The response of the airplane to a horizontal
gust will differ according to the gust gradient
and airplane characccristics. Gmcrally, if the
airplane encounters a sharp wind shear which
reduces the airspeed, the airplane tends to sink
and incur a loss of altitude ‘before equilibrium
conditions are achieved. Similarly, if the
airplane encounters a sharp wind shear which
increases the airspeed, the airplane tends to
float and incur a gain of altitude before equilib-
rium conditions are achieved.
Significant vertical and horizontal gusts may
be due to the terrain or atmospheric conditions.
The proximity’of an unstable front or thunder-
storm activity’in the vicinity of the airfield is
likely to create significant wind shear and gust
activity at low altitude. During gusty condi-
tions every effort must be made for precise con-
trol of airspeed and flight path and any changes
due to gusts must be corrected by proper con-
trol action. Under extreme gusts conditions,
it may be advisable to utilize approach, land-
ing, and takeoff speeds slightly greater than
normal to provide margin for adequate control.
POWER-OFF GLIDE PERFORMANCE
The gliding performance of an airplane is of
special interest for the single-engine airplane
NAVWEPS O&ROT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
in the case of powerplant failure or malfunc-
tion. When a powerplant failure or malfunc-
tion occurs, it is usually of interest to obtain a
gliding flight path which results in the mini-
mum glide angle. The minimum glide angle
will produce the greatest proportion of glide
distance to altitude loss and will result in
maximum glide range or minimum expendi-
ture of altitude for a specific glide distance.
GLIDE ANGLE AND LIFT-DRAG RATIO.
In the study of climb performance, the forces
acting on the airplane in a steady climb (or
glide) produce the following relationship:
where
Y=: angle of climb, degrees
T-thrust, lbs.
D-drag, lbs.
W=: lbs.
In the case of power-off glide performance, the
thrust, T, is zero and the relationship reduces
to:
D sin y= -- W
By this relationship it is evident that the mini-
mum angle of glide-or minimum negative
climb angle-is obtained at the aerodynamic
conditions which incur the minimum total
drag. Since the airplane lift is essentially equal
to the weight, the minimum angle of glide
will be obtained when the airplane is operated
at maximum lift-drag ratio, (L/D)ma,. When
the angle of glide is relatively small, the ratio
of glide distance to glide altitude is numeri-
cally equal to the airplane lift-drag ratio.
glide ratio= glide distance, ft.
glide altitude, ft.
glide ratio = (L/D)
Figure 6.6 illustrates the forces acting on the
airplane in a power-off glide. The equilibrium
of the steady glide is obtained when the sum-
mation of forces in the vertical and horizontal
directions is equal to zeta.
369 | 386 | 386 | 00-80T-80.pdf |
NAVWEPS 00-801-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
+Y
I
l--J-@$!?e
DRAG
\ SIN 7 = - j WEIGHT
LIFT-OR&G
RATIO
L4l
i
GLIDE RATIO * L/o
r )M
-CLEAN CONFIGURATION
!
A <LANDING CONFIGURATION
LIFT COEFFICIENT, CL
RATE OF
DESCENT,
FPM
CLEAN CONFIGURATION
POWER OFF
VELOCITY, KNOTS
Figure 6.6. Glide Performance
370 | 387 | 387 | 00-80T-80.pdf |
NAVWEPS W-ROT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
speed will not cause any significant reduction
of glide ratio. This is fortunate and allows the
specifying of convenient glide speeds which
will be appropriate for a range of gross weights
at which power-off gliding may be encoun-
tered, e.g., small quantities of fuel remaining.
An attempt to stretch a glide by flying at
speeds above or below the best glide speed will
prove futile. As shown by the illustration of
figure 6.6, any C, above or below the optimum
will produce a lift-drag ratio less than the
maximum. If the airplane angle of attack is
increased above the value for (L/D),,, a tran-
sient reduction in rate of descent will take place
but this process must be reserved for the land-
ing phase. Eventually, the steady-state condi-
tions would be achieved and the increased
angle of attack would incur a lower airspeed
and a reduction in (L/D) and glide ratio.
The effect of gross weight on glide performance
may be difficult to appreciate. Since (L/D)-
of a given airplane configuration will occur at a
specific value of C,, the gross weight of the air-
plane will not affect the glide ratio if the air-
plane is operated at the optimum C,. Thus,
two airplanes of identical aerodynamic con-
figuration but different gross weight could
glide the same distance from the same altitude.
Of course, this fact would be true only if both
airplanes are flown at the specific C, to produce
(L/D),,. The principal difference would be
that the heavier airplane must fly at a higher
airspeed to support the greater weight at the
optimum C,. In addition, the heavier airplane
flying at the greater speed along the same flight
path would develop a greater rate of descent.
The relationship which exists between gross
weight and velocity for a particular C, is as
follows:
In order to obtain maximum glide ratio, the
airplane must be operated at the angle of at-
tack and lift coefficient which provide maxi-
mum lift-drag ratio. The illustration of figure
6.6 depicts a variation of lift-drag ratio, L/D,
with lift coeficient, C,, for a typical airplane
in the clean and landing configurations. Note
that (LID),,, for each configuration will occur
at a specihc value of lift coefficient and, hence,
a specific angle of attack. Thus, the maximum
glide performance of a given airplane configu-
ration will be unaffected by gross weight and
altitude when the airplane is operated at
(L/D),az. Of course, an exception occurs at
very high altitudes where compressibility ef-
fects may alter the aerodynamic characteristics.
The highest value of (L/D) will occur with the
airplane in the clean configuration. As the
airplane is changed to the landing configura-
tion, the added parasite drag reduces (L/D)_nz
and the C, which produces (L/D),, will be in-
creased. Thus, the best glide speed for the
landingconhguration generallywill be lessthan
the best glidespeed .for theclean configuration.
The power-off glide performance may be
appreciated also by the graph of rate of descent
versus velocity shown in figure 6.6. When a
straight line is drawn from the origin tangent
to the curve, a point is located which produces
the maximum proportion of velocity to rate
of descent. Obviously, this condition provides
maximum glide ratio. Since the rate of descent
is proportional to the power required, the
points of tangency define the aerodynamic
condition of (L/D)m.z.
FACTORS AFFECTING GLIDE PER-
FORMANCE. In order to obtain the mini-
mum glide angle through the air, the airplane
must be operated at (L/D)mor. The subsonic
(LIDL of a given airplane configuration will
occur at a specific value of lift coefficient and
angle of attack. However, as can be noted
from the curves of figure 6.6, small deviations
from the optimum C, will not cause a drastic
reduction of (L/D) and glide ratio. In fact, a
5 percent deviation in speed from the best glide
-4 VT- w, VI w, (constant C,>
where
VI= best glide speed corresponding to
some original gross weight. WI
V,=best glide speed corresponding to
some new gross weight, IV2 | 388 | 388 | 00-80T-80.pdf |
NAVWEPS OD-ROT-80
APPLICATION OF AERODYNAMICS
TO ‘SPECIFIC PROBLEMS OF FLYl,NG
As a result of this relationship, a IO percent
increase in gross weight would require a 5 per-
cent increase in glide speed to maintain
(L/D),,. While small. variations in gross
weight may produce a measurable change in
best glide speed, the airplane can tolerate small
deviations from the optimum C, without signif-
icant change in (L/D) and glide ratio. For this
reason, a standard, single value of glide speed
may be specified for a small range of gross
weights at which glide performance can be of
importance. A gross weight which is con-
siderably different from the normal range will
require a modification of best glide speed to
maintain the maximum glide ratio.
The effect a! &it.& on glide performance is
insignificant if there is no change in (L/D),.,.
Generally, the glide performance of the major-
ity of airplanes is subsonic and there is no
noticeable variation of (L/D),, with altitude.
Any specific airplane configuration at a partic-
ular gross weight will require a specific value
of dynamic pressure to sustain flight at the
C, for (L/D),,. Thus, the airplane will have
a best glide speed which is a specific value of
equivalent airspeed (EAS) independent of
altitude. For convenience and simplicity, this
best glide speed is specified as a specific value
of indicated airspeed (IAS) and compressibility
and position errors are neglected. The prin-
cipal effect of altitude is that at high altitude
the true airspeed (TAX) and rate of descent
along the optimum glide path are increased
above the low altitude conditions. However,
if WD),.z is maintained, the glide angle and
glide ratio are identical to the low altitllde
conditions.
The effect of configura+~n has been noted pre-
viously in that the addition of parasite drag by
flaps, landing gear, speed brakes, external
stores, etc. will reduce the maximum lift-drag
ratio and cause a reduction of glide ratio. In
the case where glide distance is of great im-
portance, the airplane must be maintained in
the clean configuration and flown at (L/D),=,
The eficct aj wind on gliding performance is
similar to the effect of wind on cruising range.
That is, a headwind will always reduce the
glide range and a tailwind will always increase
the glide range. The maximum glide range
of the airplane in still air will be obtained by
flight at (L/D),,,,. However, when a wind is
present, the optimum gliding conditions may
not be accomplished by operation at (L/D)ma.
For example, when a headwind is present,
the optimum glide speed will be increased to
obtain a maximum proportion of ground dis-
tance to altitude. In this sense, the increased
glide speed helps to minimize the detrimental
effect of the headwind. In the case of a tail-
wind, the optimum glide speed will be reduced
to maximize the benefit of the tailwind. For
ordinary wind conditions, maintaining the
glide speed best for zero wind conditions will
suffice and the loss or gain in glide distance
must be accepted. However, when the wind
conditions are extreme and the wind velocity
is large in comparison with the glide speed,
e.g., wind velocity greater than 25 percent of
the glide speed, changes in the glide speed must
be made to obtain maximum possible ground
distance.
THE FLAMEOUT PATTERN. In the case
of failure of the powerplant, every effort
should be made to establish a well-planned,
stabilized approach if a suitable landing area
is available. Generally a 360’ overhead ap-
proach is specified with the approach begin-
ning from the “high key” point of the flameout
pattern. The function of a standardized
pattern is to provide a flight path well within
the capabilities of the airplane and the abilities
of the pilot to judge and control the flight
path. The flight handbook will generally
specify the particulars of the flameout pattern
such as the altitude at the high key, glide
speeds, use of flaps, etc. Of course, the par-
ticulars of the flameout pattern will be de-
termined by the aerodynamic characteristics
of the airplane. A principal factor is the
272 | 389 | 389 | 00-80T-80.pdf |
effect of glide ratio, or (L/D),,, on the alti-
tude required at the high key point at the ~ _
beginning of the flameour pattern. The air-
plane with a low value of (L/D),,* will require
a high altitude at the high key point.
The most favorable situation during a
flameout would be for the airplane to in posi-
tion to arrive over the intended landing area
the altitude for the high key point. In this
case, the standard flameout pattern could be
utilized. If the airplane does not have s&i-
cient glide range to arrive at the landing
area with the altitude for the high key point,
it is desirable to fit the approach into the
lower portions of the standard flameout ap-
proach. If it is not possible to arrive at the
intended landing area with sufficient altitude
to “play” the approach, serious considera-
tion should be given to ejection while suffi-
cient altitude remains. Deviations from a
well-planned approach such as the standard
flameout pattern may allow gross errors in
judgment. A typical error of a non-standard
or poorly executed flameout approach is the
use of excessive angles of bank in turns to
correct the approach. Because of the great
increase ‘in induced drag at large angles of
bank, excessive rates of descent will be incurred
and there will be further deviations from a
desirable flight path.
The power-off gliding characteristics of the
airplane can be simulated in power on flight by
certain combinations of engine power setting
and position of the speed brake or dive Rap.
This will allow the pilot to become familiar
with the power-off glide performance and the
flameopt landing pattern. In addition, the
simulated flameout pattern is useful during a
precautionary landing when the powerplant is
malfunctioning and there is the possibility of
an actual flarneout.
The final approach and landing flare will be
particularly critical for the airplane which has*
a low glide ratio but a high best glide speed.
These airplane characteristics are typical of the
modern configuration of airplane which has
NAVWEPS OD-ROT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
low aspect ratio, sweepback, and high wing
loading. Since these airplane characteristics
also produce marginal flare capability in power-
off flight, great care should be taken to follow
the procedure recommended for the specific
airplane.
As an example of the power-off glide per-
formance of an airplane with low aspect ratio,
sweepback, and high wing loading, a best
glide speed of 220 knots and a glide ratio of 6
may be typical. In such a case, the rate of
descent during the glide at low altitude would
be on the order of 3,700 FPM. Any deviations
from the recommended landing technique can-
not be tolerated because of the possibility of an
excessive rate of descent. Either premature
flare or delayed flare may allow the airplane to
touch down at a rate of descent which would
cause structural failure. Because of the mar-
ginal flare characteristics in power-off flight,
the best glide speed recommended for the land-
ing configuration may be well above the speed
corresponding to the exact maximum lift-drag
ratio. The greater speed reduces induced drag
and provides a greater margin for a successful
power-off landing flare.
In the extreme case, the power-off glide and
landing flare characteristics may be very criti-
cal for certain airplane configurations. Thus,
a well-planned standard flameout pattern and
precise flying technique are necessary and, if
very suitable conditions are not available, the
recommended alternative is simple: eject!
EFFECT OF ICE AND FROST ON AIRPLANE
PERFORMANCE
Without exception, the formation of ice or
frost on the surfaces of an airplane will cause a
detrimental effect on aerodynamic performance.
The ice or frost formation on the airplane sur-
faces will alter the aerodynamic contours and
affect the nature of the boundary layer. Of
course, the most important surface of the air-
plane is the wing and the formation of ice or
frost can create significant changes in the aero-
dynamic characteristics.
373 | 390 | 390 | 00-80T-80.pdf |
NAVWEPS DO-80T-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
EDGE ICE FOR_“ATlON
UPPER SURFACE FROST
A
BASIC SMOOTH WING
WING WITH FROST
LIFT
COEFFICIENT WITH ICE
CL
t
ANGLE OF ATTACK, a
Figure 6.7. Effect of ice and Frost
374 | 391 | 391 | 00-80T-80.pdf |
A large formation of ice on the leading edge
of the wing can produce large changes in the
local contours and severe local pressure gra-
dients. The extreme surface roughness common
to some forms of ice will cause high surface
friction and a considerable reduction of bound-
ary layer energy. As a result of these effects,
the ice formation can produce considerable in-
crease in drag and a large reduction in maxi-
mum lift coefficient. Thus, the ice formation
will cause an increase in power required and
stall speed. In addition, the added weight of
the ice formation on the airplane will provide
an undesirable effect. Because of the detri-
mental effects of ice formation, recommended
anti-icing procedures must be followed to
preserve the airplane performance.
The effect of frost is perhaps more subtle
than the effect of ice formation on the aero-
dynamic characteristics of the wing. The ac-
cumulation of a hard coat of frost on the wing
upper surface will provide a surface texture of
considerable roughness. While the basic shape
and aerodynamic contour is unchanged, the
increase in surface roughness increases skin-
friction and reduces the kinetic energy of the
boundary layer. As a result, there will be an
increase in drag but, of course, the magnitude
of drag increase will not compare with the
considerable increase due to a severe ice forma-
tion. The reduction of boundary layer kinetic
energy will cause incipient stalling of the wing,
i.e., separation will occur at angles of attack
and lift coefficients lower than for the clean,
smooth wing. While the reduction in C,,,,
due to frost formation ordinarily is not as great
as that due to ice formation, it is usually un-
expected because it may be thought that large
changes in the aerodynamic shape (such as
due to ice) are necessary to reduce CL,az. How-
ever, the kinetic energy of the boundary layer
is an important factor influencing separation
of the airflow and this energy is reduced by an
increase in surface roughness.
The general effects of ice and frost formation
NAVWEPS OD-BOT-80
APP,LlCATlON OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
on the lift characteristics is typified by the il-
lustration of figure 6.7.
The effect of ice or frost on takeoff and land-
ing performance is of great importance. The
effects are so detrimental to the landing and
takeoff that no effort should be spared to keep
the airplane as free as possible from any ac-
cumulation of ice or frost. If any ice remains
on the airplane as the landing phase approaches
it must be appreciated that the ice formation
will have reduced CL,., and incurred an increase
in stall speed. Thus, the landing speed will be
greater. When this effect is coupled with the
possibility of poor braking action during the
landing roll, a critical situation can exist. It is
obvious that .great effort must be made to
prevent the accumulation of ice during flight.
In no circumstances should a formation of ice
or frost be allowed to remain on the airplane
wing surfaces prior to takeoff. The undesir-
able effects of ice are obvious but, as previously
mentioned, the effects of frost are more subtle.
If a heavy coat of hard frost exists on the wing
upper surface, a typical reduction in CL,..
would cause a 5 to 10 percent increase in the
airplane stall speed. Because of this magnitude
of effect, the effect of frost on takeoff per-
formance may not be realized until too late.
The takeoff speed of an airplane is generally
some speed 5 to 25 percent greater than the
stall speed, hence the takeoff lift coefficient
will be value from 90 to 65 percent of C1,,..,
Thus, it is possible that the airplane with frost
cannot become airborne at the specified take-
off speed because of premature stalling. Even
if the airplane with frost were to become air-
borne at the specified takeoff speed, the air-
plane could have insufficient margin of air-
speed above stall and turbulence, gusts, turning
flight could produce incipient or con plete
stalling of the airplane.
The increase in drag during takeoff roll due
to frost or ice is not considerable and there
will not be any significant effect on the initial
acceleration during takeoff. Thus, the effect
of frost or ice will be most apparent during the
375 | 392 | 392 | 00-80T-80.pdf |
NAVWEPS DD-8OT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
later portions of takeoff if the airplane is un-
able to become airborne or if insufficient margin
above stall speed prevents successful initial
climb, In no circumstances should a formation
of ice or frost be allowed to remain on the air-
plane wing surfaces prior to takeoff.
ENGINE FAILURE ON THE MULTIENGINE
AIRPLANE
In the case of the single-engine airplane,
power-plant failure leaves only the alternatives
of effecting a successful power-off landing or
abandoning the airplane. In the case of the
multiengine airplane, the failure of a power-
plant does nor necessarily constitute a disaster
since flight may be continued with the remain-
ing powerplants functioning. However, the
performance of the multiengine airplane with
a powerplant inoperative may be critical for
certain conditions of flight and specific tech-
niques and procedures must be observed to
obtain adequate performance.
The effect of a powerplant failure on the
multiengine turbojet airplane is illustrated by
the first chart of figure 6.8 with the variation
of required and available thrust with velocity.
If half of the airplane powerplants are inoper-
ative, e.g., single-engine operation of a twin-
engine airplane, the maximum thrust available
at each velocity is reduced to half that avail-
able prior to the engine failure. The variation
of thrust required with velocity may be
affected by the failure of a powerplant in that
there may be significant increases in drag if
specific procedures are not followed. The
inoperative powerplant may contribute addi-
tional drag and the pilot must insure that the
additional drag is held to a minimum. In the
case of the propeller powered airplane, the
propeller must be feathered, cowl flaps closed.
etc., as the increased drag will detract con-
siderably from the performance.
The principal effects of the reduced available
thrust are pointed out by the illustration of
figure 6.6. Of course, the lower available
thrust will reduce the maximum level flight
speed but of greater importance is the reduc-
tion in excess thrust. Since the acceleration
and climb performance is a function of the
excess thrust and power, the failure of a power-
plant will be most immediately appreciated in
this area of performance. As illustrated in
figure 6.8, loss of one-half the maximum avail-
able thrust will reduce the excess thrust to less
than half the original value. Since some
thrust is required to sustain flight, the excess
which remains to accelerate and climb the
airplane may be greatly reduced. The most
critical conditions will exist when various
factors combine to produce a minimum of
excess thrust or power when engine failure
occurs. Thus, critical conditions will be com-
mon to high gross weight and high density altitude
(and high temperatures in the case of the
turbine powered airplane) as each of these
factors will reduce the excess thrust at any
specific flight condition.
The asymmetrical power condition which
results when a powerplant fails can provide
critical control requirements. First consid-
eration is due the yawing moment produced
by the asymmetrical power condition. Ade-
quate directional control will be available
only when the airplane speed is greater than
the minimum directional control speed. Thus,
the pilot must insure that the flight speed never
falls below the minimum directional control
speed because the application of maximum
power on the functioning powerplants will
produce an uncontrollable yaw if adequate
directional control is unavailable. A second
consideration which is due the propeller
powered airplane involves the rolling moments
caused by the slipstream velocity. Asym-
metrical power on the propeller airplane will
create a dissymmetry of the slipstream veloc-
ities on the wing and create rolling moments
which must be controlled. These slipstream
induced rolling moments will be greatest at
high power and low velocity and the pilot
must be sure of adequate lateral control,
especially for the crosswind landing.
376 | 393 | 393 | 00-80T-80.pdf |
The effect of an engine failure on the remain-
ing range and endurance is specific to the air-
plane type and configuration. If an engine
fails during optimum cruise of the turbojet
airplane, the airplane must descend and experi-
ence a loss of range. Since the turbojet air-
plane is generally overpowered at (L/D),,,
a loss of a powerplant will not cause a signi-
ficant change in maximum endurance. If an
engine fails during cruise of a reciprocating
powered airplane, there will be a significant
loss of range only if the maximum range condi-
tion cannot be sustained with the remaining
powerplants operating within the cruise power
rating. If a power greater than the maximum
cruise rating is necessary to sustain cruise, the
specific fuel consumption increases and causes
a reduction of range. Essentially the same
relationship exists regarding maximum endur-
ance of the reciprocating powered airplane.
When critical conditions exist due to failure
of a powerplant, the pilot must appreciate the
reduced excess thrust and operate the airplane
within specific limitations. If the engine-out
performance of the airplane is marginal, the
pilot must be aware of the very detrimental
effect of steep turns.. Due to the increased load
factor in a coordinated turn, there will be an
increase in stall speed and-of greater import-
ance to engine-out performance-an increase
in induced drag. The following table illus-
trates the effect of bank angle on stall speed
and induced drag.
TABLE 6.1
Bank mglc, 6, dcgrccs Load factor
0 0
0.2 0.8
0.7 3.1
1.7 7.2
3.2 13.3
5.0 21.7
7.5 33.3
10.5 4% 0
14.3 70.4
IS. 9 loo. 0
41.4 303.0
NAVWEPS OD-BOT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
The previous table of values illustrates the
fact that coordinated turns with less than 15”
of bank .cause no appreciable effect on stall
speed or induced drag. However, note that 30”
of bank will increase the induced drag by 33.3
percent. Under critical conditions, such an in-
crease in induced drag (and, hence, total drag)
would be prohibitive causing the airplane to
descend rather than climb. The second graph
of figure 6.7 illustrates the case where the steep
turn causes such a large increase in required
thrust that a deficiency of thrust exists. When-
ever engine failure produces critical perform-
ance conditions it is wise to limit all turns to
II0 of bank wherever possible.
Another factor to consider in turning flight
is the effect of sideslip. If the turn is not coor-
dinated to hold sideslip to a minimum, addi-
tional drag will be incurred due to the sideslip.
The use of the flaps and landing gear can
greatly affect the performance of the multi-
engine airplane when a powerplant is inopera-
tive. Since the extension of the landing gear
and flaps increases the parasite drag, maximum
performance of the airplane will be obtained
with airplane in the clean configuration. In
certain critical conditions, the extension of the
landing gear and full flaps may create a defi-
ciency of thrust at any speed and commit the
airplane to descend. This condition is illus-
trated by the second graph of figure 6.8. Thus,
judicious use of the flaps and landing gear is
necessary in the case of an engine failure.
In the case of engine failure immediately
after takeoff, it is important to maintain air-
speed in excess of the minimum directional con-
trol speed and accelerate to the best climb
speed. After the engine failure, it will be fa-
vorable to climb only as necenary to clear obstacles
until the airplane reaches the best climb speed. Of
course, the landing gear should be retracted as
soon as the airplane is airborne to reduce para-
site drag and, in the case of the propeller pow-
ered airplane, it is imperative that the wind
milling propeller be feathered. The flaps should
be retracted only as rapidly as the increase in
377 | 394 | 394 | 00-80T-80.pdf |
NAVWEPS 00-801-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
THRUST
REQ’D
AND
AVAILABLE
LBS.
THRUST AVAJLABLE WJTH ALL
ENGINES OPERATING
THRUST AVAILABLE WITH
NES OPERATING
THRUST
REO’U
AND
AVAILABLE
LB.%
I VELOCITY, KNOTS
THRUST REP’D THRUST REO’D CLEAN
LANDING CONFIGURATION CONFIGURATION
WING LEVEL FLIGHT , BUT TURNING FLIGHT
WING LEVEL FLIGHT
AVAILABLE DUE TO
ENGINE FAILURE
4
VELOCITY, KNOTS
Figure 6.8. Engine Failure on Multi-engine Aircraft
378 | 395 | 395 | 00-80T-80.pdf |
airspeed will allow. If full flap deflection is
utilized for takeoff it is important to recall that
the last 50 percent of flap deflection creates
more than half the total drag increase but less
than half the total change in CL,-. Thus, for
some configurations of airplanes, a greater re-
duction in drag may be accomplished by partial
retraction of the flaps rather than retraction of
the landing gear. Also, it is important that no
steep turns be attempted because of the unde-
sirable increase in induced drag.
During the landing with an engine inopera-
tive, the same fundamental precautions must
be observed as during takeoff, i.e., minimum
directional control speed must be maintained
(or exceeded), no steep turns should be at-
tempted, and the extension of the flaps and
landing gear must be well planned. In the case
of’a critical power condition it may be neces-
sary to delay the extension of the landing gear
and full flaps until a successful landing is as-
sured. If a waveoff is necessary, maximum per-
formance will be obtained cleaning up the air-
plane and accelerating to the best climb speed
before attempting any gain in altitude.
At all times during flight with an engine
inoperative, the pilot must utilize the proper
techniques for control of airspeed and altitude,
e.g., for the conditions of steady flight, angle
of attack is the primary control of airspeed
and excess power is the primary control of
rate of climb. For example, if during approach
to landing the extension of full flaps and
landing gear creates a deficiency of power at
all speeds, the airplane will be committed to
descend. If the approach is not properly
planned and the airplane sinks below the
desired glide path, an increase in angle of
attack will only allow the airplane to fly more
slowly and descend more rapidly. An attempt
to hold altitude by increased angle of attack
when a power deficiency exists only causes a
continued loss of airspeed. Proper procedures
and technique are an absolute necessity for
safe flight when an engine failure occurs.
NAVWEPS OO-BOT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
GROUND EFFECT
When an airplane in flight nears the ground
(or water) surface, a change occurs in the
three dimensional flow pattern because the
local airflow cannot have a vertical component
at the ground plane. Thus, the ground plane
will furnish a restriction to the flow and alter
the wing upwash, downwash, and tip vortices.
These general effects due to the presence of
the ground plane are referred to as “ground
effect. ‘*
AERODYNAMIC INFLUENCE OF
GROUND EFFECT. While the aerodynamic
characteristics of the tail and fuselage are
altered by ground effects, the principal effects
due to proximity of the ground plane are the
changes in the aerodynamic characteristics
of the wing. As the wing encounters ground
effect and is maintained at a constant lift
coefficient, there is a reduction in the upwash,
downwash, and the tip vortices. These effects
are illustrated by the sketches of figure 6.9.
As a result of the reduced tip vortices, the wing
in the presence of ground effect will behave as
if it were of a greater aspect ratio. In other
words, the induced velocities due to the tip
(or trailing) vortices will be reduced and the
wing will incur smaller values of induced
drag coefficient, C,<, and induced angle of
attack, OL;, for any specific lift coefhcient, C,.
In order for ground effect to be of a signifi-
cant magnitude, the wing must be quite close
to the ground plane. Figure 6.9 illustrates
one of the direct results of ground effect by
the variation of induced drag coefficient with
wing height above the ground plane for a
representative unswept wing at constant lift
coefficient. Notice that the wing must be
quite close to the ground for a noticeable
reduction in induced drag. When the wing
is at a height equal to the span (h/b=l.O),
the reduction in induced drag is only 1.4
percent. However, when the wing is at a
height equal to one-fourth the span (b/b=
0.25), the reduction in induced drag is 23.5
379 | 396 | 396 | 00-80T-80.pdf |
NAVWEPS 00-802-80
APPLICATION OF AERODYNAMICS
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AIRPLANE OUT OF
GROUND EFFECT TIP VORTEX
/
REDUCED DOWNWASH
AND UPWASH- REDUCED t-- SFkAN’ --I
PERCENT
REDUCTION
IN
INDUCED
DRAG
COEFFICIENT
LIFT
COEFFICIEN’
CL
40
CL CONSTANT
30
20
IO
0
r
1
RATIO OF WING HEIGHT TO SPAN, h/b
AIRPLANE IN
THRUST
REQ’D
LBS.
AIRPLANE OUT OF
GROUND EFFECT
w - w
ANGLE OF ATTACK, 0 VELOCITY, KNOTS
Figure 6.9. Ground Effect
380
AIRPLANE OUT OF /
(-,’ < AIRPLANE IN
GROUND EFFECT | 397 | 397 | 00-80T-80.pdf |
percent and, when the wing is at a height
equal to one-tenth the span (h/b=O.l), the
reduction in induced drag is 47.6 percent.
Thus, a large reduction in induced drag will
take place only when the wing is very close
to the ground. Because of this variation,
ground effect is most usually recognized during
the liftoff of takeoff or prior to touchdown on
landing.
The reduction of the tip or trailing vortices
due to ground effect alters the spanwise lift
distribution and reduces the induced angle of
attack. In this case, the wing will require
a lower angle of attack in ground effect to
produce the same lift coefficient. This effect
is illustrated by the lift curves of figure 6.9
which show that the airplane in ground effect
will develop a greater slope of the lift curve.
For the wing in ground effect, a lower angle of
attack is necessary to produce the same lift
coefficient or, if a constant angle of attack is
maintained, an increase in lift coefficient will
result.
Figure 6.9 illustrates the manner in which
ground effect will alter the curve of thrust re-
quired versus velocity. Since induced drag
predominates at low speeds, the reduction of
induced drag due to ground effect will cause
the most significant reduction of thrust re-
quired (parasite plus induced drag) only at
low speeds. At high speeds where parasite
drag predominates, the induced drag is but
a small part of the total drag and ground
effect causes no significant change in thrust re-
quired. Because ground effect involves the
induced effects of airplane when in close prox-
imity to the ground, its effects are of greatest
concern during the takeoff and landing. Ordi-
narily, these are the only phases of flight in
which the airplane would be in close proximity
to the ground.
GROUND EFFECT ON SPECIFIC FLIGHT
CONDITIONS. The overall influence of
ground effect is best realized by assuming that
the airplane descends into ground effect while
maintaining a constant lift coefficient and,
NAVWEPS CKLBOT-BO
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
thus, a constant dynamic pressure and equiva-
lent airspeed. As the airplane descends into
ground effect, the following. effects will take
place:
(1) Because of the reduced induced angle
of attack and change in lift distribution, a
smaller wing angle of attack will be required
to produce the same lift coefficient. If a
constant pitch attitude is maintained as
ground effect is encountered, an increase in
lift coefficient will be incurred.
(2) The reduction in induced flow due to
ground effect causes a significant reduction
in induced drag but causes no direct effect on
parasite drag. As a result of the reduction
in induced drag, the thrust required at low
speeds will be reduced.
(3) The reduction in downwash due to
ground effect will produce a change in longi-
tudinal stability and trim. Generally, the
reduction in downwash at the horizontal
tail increases the contribution to static longi-
tudinal stability. In addition, the reduction
of downwash at the tail usually requires
a greater up elevator to trim the airplane at
a specific lift coefficient. For the conven-
tional airplane configuration, encountering
ground effect will produce a nose-down
change in pitching moment. Of course, the
increase in stability and trim change associ-
ated with ground effect provide a critical re-
quirement of adequate longitudinal control
power for landing and takeoff.
(4) Due to the change in upwash, down-
wash, and tip vortices, there will be a change
in position error of the airspeed system, as-
sociated with ground effect. In the majority
of cases, ground effect will cause an increase
in the local pressure at the static source and
produce a lower indication of airspeed and
altitude.
During the landing pha~c of flight, the effect
of proximity to the ground plane must be
understood and appreciated. If the airplane
is brought into ground effect with a constant
angle of attack, the airplane will experience
3&l | 398 | 398 | 00-80T-80.pdf |
NAVWEPS OD-8OT-80
APPLICATION OF AERODYNAMICS
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an increase in lift coeflicient and reduction in
thrust required. Hence, a “floating” sensa-
tion may be experienced. Because of the re-
duced drag and power-off deceleration in
ground effect, any excess speed at the point of
flare may incur a considerable “float” distance.
As the airplane nears the point of touchdown
on the approach, ground effect will be most
realized at altitudes less than the wing span.
An exact appreciation of the ground effect may
be obtained during a PcZd approach with the
mirror landing system furnishing an exact
reference of the flight path. During the final
phases of the field approach as the airplane
nears the ground plane, a reduced power
setting is necessary or the reduced thrust re-
quired would allow the airplane to climb
above the desired glide path. During ship-
board operations, ground effect will be delayed
until the airplane passes the edge of the deck
and the reduction in power setting that is
common to field operations should not be
encountered. Thus, a habit pattern should
not be formed during field landings which
would prove dangerous during carrier oper-
ations.
An additional factor to consider is the aero-
dynamic drag of the airplane during the land-
ing roll. Because of the reduced induced drag
when in ground effect, aerodynamic braking
will be of greatest significance only when
partial stalling of the wing can be accom-
plished. The reduced drag when in ground
effect accounts for the fact that the brakes
are the most effective source of deceleration
for the majority of airplane configurations.
During the takeoff pharc of flight ground
effect produces some important relationships.
Of course, the airplane leaving ground effect
encounters just the reverse of the airplane
entering ground effect, i.e., the airplane leaving
ground effect will (1) require an increase in
angle of attack to maintain the same lift
coefficient, (2) experience an increase in in-
duced drag and thrust required, (3) experience a
decrease in stability and a nose-up change in
moment, and (4) usually a reduction in static
source pressure and increase in indicated air-
speed. These general effects should point out
the possible danger in attempting takeoff
prior to achieving the recommended takeoff
speed. Due to the reduced drag in ground
effect the airplane may seem capable of takeoff
below the recommended speed. However, as
the airplane rises out of ground effect with a
deficiency of speed, the greater induced drag
may produce marginal initial climb perform-
ance. In the extreme conditions such as high
gross weight, high density altitude, and high
temperature, a deficiency of airspeed at takeoff
may permit the airplane to become airborne
but be incapable of flying’out of ground effect.
In this case, the airplane may become airborne
initially with a deficiency of speed, but later
settle back to the runway. It is imperative
that no attempt be made to force the airplane
to become airborne with a deficiency of speed;
the recommended takeoff speed is necessary to
provide adequate initial climb performance.
In fact, ground effect can be used to advantage
if no obstacles exist by using the reduced drag
to improve initial acceleration.
The results of the airplane leaving ground
effect can be most easily realized during the
deck launch of a heavily loaded airplane. As
the airplane moves forward and passes over the
edge of the deck, whatever ground effect exists
will be lost immediately. Thus, proper rota-
tion of the airplane will be necessary to main-
tain the same lift coefficient and the increase
in induced drag must be expected.
The rotor of the helicopter experiences a
similar restraint of induced flow when in prox-
imity to the ground plane. Since the induced
rotor power required will predominate at low
flight speeds, ground effect will produce a con-
siderable effect on the power required at low
speeds. During hovering and flight at low
speeds, the elevation of the rotor above the
ground plane will be an important factor de-
termining the power required for flight.
3882 | 399 | 399 | 00-80T-80.pdf |
The range sf the reciprocating powered air-
plane can be augmented by the use of ground
effect. When the airplane is close to the
ground or water surface the reduction of in-
duced drag increases the maximum lift-drag
ratio and causes a corresponding increase in
range. Of course, the airplane must be quite
close to the surface to obtain a noticeable in-
crease in (L/D),., and range. The difficulty in
holding the airplane at the precise altitude
without contacting the ground or water will
preclude the use of ground effect during ordi-
nary flying operations. The use of ground
effect to extend range should be reserved as
a final measure in case of emergency. Because
of the very detrimental effect of low altitude
on the range of the turbojet, ground effect will
not be of a particular advantage in an attempt
to augment range.
The most outstanding examples of the use
of ground effect are shown in the cases of multi-
engine airplanes with some engines inoperative.
When the power loss is quite severe, the air-
plane may not be capable of sustaining altitude
and will descend. As ground effect is en-
countered, the reduced power required may
allow the airplane to sustain flight at extremely
low altitude with the remaining powerplants
functioning. In ground effect, the recipro-
cating powered airplane will encounter a
greater (L/D),, which occurs at a lower air-
speed and power required and the increase in
range may be quite important during emer-
gency conditions.
INTERFERENCE BETWEEN AIRPLANES IN
FLIGHT
During formation flying and inflight refuel-
ing, airplanes in proximity to one another will
produce a mutual interference of the flow pat-
terns and alter the aerodynamic characteristics
of each airplane. The principal effects of this
interference must be appreciated since certain
factors due to the mutual interference may
enhance the possibility of a collision.
NAVWEPS D&ROT-R0
APPLICATION OF AERODYNAMICS
TO SPECIFIC ‘PROBLEMS OF FLYING
One example of interference between air-
planes in flight is shown first in figure 6.10 with
the effect of lateral separation of two airplanes
flying in line abreast. A plane of symmetry
would exist halfway between two identical air-
planes and would furnish a boundary of flow
across which there would be no lateral com-
ponents of flow. As the two airplane wing
tips are in proximity, the effect is to reduce the
strength of the tip or trailing vortices and re-
duce the induced velocities in the vicinity of
wing tip. Thus, each airplane will experience
a local increase in the lift distribution as the
tip vortices are reduced and a rolling moment is
developed which tends to roll each airplane
away from the other. This disturbance may
provide the possibility of collision if other air-
planes are in the vicinity and there is delay in
control correction or overcontrol. If the wing
tips are displaced in a fore-and-aft direction,
the same effect exists but generally it is of a
lower magnitude.
The magnitude of the interference effect due
to lateral separation of the wing tips depends
on the proximity of the wi.ig tips and the ex-
tent of induced Pov;. This implies that the
interference v-r 1 e grealest when the tips
are very close AL-L the airplanes are operating
at high lift coefficients. An interesting ramifi-
cation of this effect is that several airplanes in
line abreast with the wing tips quite close will
experience a reduction in induced drag.
An indirect form of interference can be en-
countered from the vortex system created by a
preceding airplane along the intended flight
path. The vortex sheet rolls up a considerable
distance behind an airplane and creates consid-
erable turbulence for any closely following air-
plane. This wake can prove troublesome if air-
planes taking off and landing are not provided
adequate separation. The rolled-up vortex
sheet will be strongest when the preceding air-
planes is large, high gross weight, and operat-
ing at high lift coefhcients. At times this tur-
bulence may be falsely attributed to propwash
or jetwash.
383 | 400 | 400 | 00-80T-80.pdf |
NAVWBPS OO-BOT-BO
APPLICATION OF AERODYNAMlCS
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LIFT DISTRIBUTION
LBWFT OF SPAN
TIP VORTEX
PLANE OF
SYMMETRY
REDUCED CHANGE IN
LIFT DISTRIBUTION
-
--
%SH
\’
DOCASH
11
TRIM CHANGE
figure 6.10. Interference 8etween Airplanes in Flight | 401 | 401 | 00-80T-80.pdf |
Another important form of direct inter-
ference is common when the two airplanes are
in a trail position and stepped down. As shown
in figure 6.10, the single airplane in flight de-
velops upwash ahead of the wing and down-
wash behind and any restriction accorded the
flow can alter the distribution and magnitude of
the upwash and downwash. When the trailing
airplane is in close proximity aft and below the
leading airplane a mutual interference takes
place betweetrthe two airplanes. The leading
airplane above will experience an effect which
would be somewhat similar to encountermg
ground effect, i.e., a reduction in induced drag,
a reduction in downwash at the tail, and a
change in pitching moment nose down. The
trailing airplane below will experience an effect
which is generally the opposite of the airplane
above. In other words, the airplane below
will experience an increase in induced drag, an
increase in downwash at the tail, and a change
in pitching moment nose up. Thus, when
the airplanes are in close proximity, a definite
collision possibility exists because of the trim
change experienced by each airplane. The
magnitude of the trim change is greatest
when the airplanes are operating at high lift
coefficients, e.g., low speed flight, and when
the airplanes are in close proximity.
In formation flying, this sort of interference
must be appreciated and anticipated. In cross-
ing under another airplane, care must be
taken to anticipate the trim change and
adequate clearance must be maintained, other-
wise a collision may result. The pilot of the
leading aircraft will know of the presence of
the trailing airplane by the trim change
experienced. Obviously, some anticipation is
necessary and adequate separation is necessary
to prevent a disturbing magnitude of the
trim change. In a close diamond formation
the leader will be able to “feel” the presence
of the slot man even though the airplane is
not within view. Obviously, the slot man
will have a difficult job during formation
maneuvers because of the unstable trim changes
NAVWEPS OO-ROT-80
APPLICATION OF AERODYNAMICS
TO SPECIFIC PROBLEMS OF FLYING
and greater power changes required to hold
position.
A common collision problem is the case of
an airplane with a malfunctioning landing
gear. If another”airpIane is called to inspect
the malfunctioning landing gear, great care
must be taken to maintain adequate separation
and preserve orientation. Many instances
such as this have resulted in a collision when
the pilo: of the trailing airplane became dis-
oriented and did not maintain adequate sepa-
ration.
During inflight refueling, essentially the
same problems of interference exist. AS the
receiver approaches the tanker from behind
and below, the receiver will encounter the
downwash from the tanker and require a
slight, gradual increase in power and pitch
attitude to continue approach to the receiving
position. While.‘the .receiver may not be
visible to the pilot ‘of the tanker, he will
anticipate the receiver coming into position
by the slight reduttion in power required and
nose down changein pitching moment. Ade-
quate clearance and, proper position must be
maintained by the pilot of the receiver for a
collision possibility is enhanced by the rela-
tive positions of the airplanes. A hazardous
condition exists if the pilot of the receiver
has excessive speed and runs under the tanker
in close proximity.* ‘The trim change expe-
rienced by both airphines may be large and
unexpected and it may be difficult to avoid a
collision.
In addition to the forms of interference
previously mentioned, there exists the possi-
bility of strong interference between airplanes
in supersonic flight. In this case, the shock
waves from one airplane may strongly affect
the pressure distribution and rolling, yawing,
and pitching moments of an adjacent air-
Pl ane. It is difficult to express general rela-
tionships of the effect except that magnitude
of the effects will be greatest when in close
proximity at low altitude and high 4. General-
ly, the trailing airplane will be most affected. | 402 | 402 | 00-80T-80.pdf |
403 | 403 | 00-80T-80.pdf |
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