The instruments discussed in this chapter are those required
by Title 14 of the Code of Federal Regulations (14 CFR)
part 91, and are organized into three groups: pitot-static
instruments, compass systems, and gyroscopic instruments.
The chapter concludes with a discussion of how to prefl ight
these systems for IFR fl ight. This chapter addresses additional
avionics systems such as Electronic Flight Information
Systems (EFIS), Ground Proximity Warning System
(GPWS), Terrain Awareness and Warning System (TAWS),
Traffi c Alert and Collision Avoidance System (TCAS),
Head Up Display (HUD), etc., that are increasingly being
incorporated into general aviation aircraft.
Pitot/Static Systems
Pitot pressure, or impact air pressure, is sensed through an
open-end tube pointed directly into the relative wind fl owing
around the aircraft. The pitot tube connects to pressure
operated fl ight instruments such as the ASI.
Static Pressure
Other instruments depend upon accurate sampling of the
ambient still air atmospheric pressure to determine the
height and speed of movement of the aircraft through the
air, both horizontally and vertically. This pressure, called
static pressure, is sampled at one or more locations outside
the aircraft. The pressure of the static air is sensed at a fl ush
port where the air is not disturbed. On some aircraft, air is
sampled by static ports on the side of the electrically heated
pitot-static head. Other aircraft pick up the static
pressure through fl ush ports on the side of the fuselage or
the vertical fi n. These ports are in locations proven by fl ight
tests to be in undisturbed air, and they are normally paired,
one on either side of the aircraft. This dual location prevents
lateral movement of the aircraft from giving erroneous static
pressure indications. The areas around the static ports may be
heated with electric heater elements to prevent ice forming
over the port and blocking the entry of the static air.
Three basic pressure-operated instruments are found in most
aircraft instrument panels. These are the sensitive altimeter,
ASI, and vertical speed indicator (VSI). All three receive
pressures sensed by the aircraft pitot-static system. The static
ports supply pressure to the ASI, altimeter, and VSI.
Blockage Considerations
The pitot tube is particularly sensitive to blockage especially
by icing. Even light icing can block the entry hole of the pitot
tube where ram air enters the system. This affects the ASI
and is the reason most airplanes are equipped with a pitot
heating system.
3-3
Figure 3-2. A Typical Pitot-Static System.
Indications of Pitot Tube Blockage
If the pitot tube becomes blocked, the ASI displays inaccurate
speeds. At the altitude where the pitot tube becomes blocked,
the ASI remains at the existing airspeed and doesn’t refl ect
actual changes in speed.
• At altitudes above where the pitot tube became
blocked, the ASI displays a higher-than-actual
airspeed increasing steadily as altitude increases.
• At lower altitudes, the ASI displays a lower-than-actual
airspeed decreasing steadily as altitude decreases.
Indications from Static Port Blockage
Many aircraft also have a heating system to protect the
static ports to ensure the entire pitot-static system is clear
of ice. If the static ports become blocked, the ASI would
still function but could produce inaccurate indications. At
the altitude where the blockage occurs, airspeed indications
would be normal.
• At altitudes above which the static ports became
blocked, the ASI displays a lower-than-actual airspeed
continually decreasing as altitude is increased.
• At lower altitudes, the ASI displays a higher-than-actual
airspeed increasing steadily as altitude decreases.
The trapped pressure in the static system causes the altimeter
to remain at the altitude where the blockage occurred. The
VSI remains at zero. On some aircraft, an alternate static
air source valve is used for emergencies. If
the alternate source is vented inside the airplane, where
static pressure is usually lower than outside static pressure,
selection of the alternate source may result in the following
erroneous instrument indications:
1. Altimeter reads higher than normal,
2. Indicated airspeed (IAS) reads greater than normal,
and
3. VSI momentarily shows a climb. Consult the Pilot’s
Operating Handbook/Airplane Flight Manual (POH/
AFM) to determine the amount of error.
Effects of Flight Conditions
The static ports are located in a position where the air at
their surface is as undisturbed as possible. But under some
fl ight conditions, particularly at a high angle of attack with
the landing gear and fl aps down, the air around the static
port may be disturbed to the extent that it can cause an error
in the indication of the altimeter and ASI. Because of the
importance of accuracy in these instruments, part of the
certifi cation tests for an aircraft is a check of position error
in the static system.
The POH/AFM contains any corrections that must be applied
to the airspeed for the various confi gurations of fl aps and
landing gear.
Pitot/Static Instruments
Sensitive Altimeter
A sensitive altimeter is an aneroid barometer that measures
the absolute pressure of the ambient air and displays it in
terms of feet or meters above a selected pressure level.
Principle of Operation
The sensitive element in a sensitive altimeter is a stack of
evacuated, corrugated bronze aneroid capsules.
The air pressure acting on these aneroids tries to compress
them against their natural springiness, which tries to expand
them. The result is that their thickness changes as the air
pressure changes. Stacking several aneroids increases the
dimension change as the pressure varies over the usable
range of the instrument.
Below 10,000 feet, a striped segment is visible. Above this
altitude, a mask begins to cover it, and above 15,000 feet,
all of the stripes are covered.
Another confi guration of the altimeter is the drum-type.
These instruments have only one pointer that
makes one revolution for every 1,000 feet. Each number
represents 100 feet and each mark represents 20 feet. A drum,
marked in thousands of feet, is geared to the mechanism that
drives the pointer. To read this type of altimeter, fi rst look at
3-4
Figure 3-3. Sensitive Altimeter Components.
Figure 3-4. Three-Pointer Altimeter. Figure 3-5. Drum-Type Altimeter.
the drum to get the thousands of feet, and then at the pointer
to get the feet and hundreds of feet.
A sensitive altimeter is one with an adjustable barometric scale
allowing the pilot to set the reference pressure from which the
altitude is measured. This scale is visible in a small window
called the Kollsman window. A knob on the instrument adjusts
the scale. The range of the scale is from 28.00" to 31.00"
inches of mercury (Hg), or 948 to 1,050 millibars.
Rotating the knob changes both the barometric scale and
the altimeter pointers in such a way that a change in the
barometric scale of 1" Hg changes the pointer indication
by 1,000 feet. This is the standard pressure lapse rate
below 5,000 feet. When the barometric scale is adjusted
to 29.92" Hg or 1,013.2 millibars, the pointers indicate the
pressure altitude. The pilot displays indicate altitude by
adjusting the barometric scale to the local altimeter setting.
The altimeter then indicates the height above the existing
sea level pressure.
Altimeter Errors
A sensitive altimeter is designed to indicate standard changes
from standard conditions, but most fl ying involves errors
caused by nonstandard conditions and the pilot must be able
to modify the indications to correct for these errors. There
are two types of errors: mechanical and inherent.
Mechanical
A prefl ight check to determine the condition of an altimeter
consists of setting the barometric scale to the local altimeter
setting. The altimeter should indicate the surveyed elevation
3-5
Figure 3-6. The loss of altitude experienced when fl ying into an area where the air is warmer (less dense) than standard.
of the airport. If the indication is off by more than 75 feet from
the surveyed elevation, the instrument should be referred
to a certifi cated instrument repair station for recalibration.
Differences between ambient temperature and/or pressure
causes an erroneous indication on the altimeter.
Inherent Altimeter Error
Figure 3-6 shows how nonstandard temperature affects an
altimeter. When the aircraft is fl ying in air that is warmer
than standard, the air is less dense and the pressure levels
are farther apart. When the aircraft is fl ying at an indicated
altitude of 5,000 feet, the pressure level for that altitude is
higher than it would be in air at standard temperature, and
the aircraft is higher than it would be if the air were cooler.
If the air is colder than standard, it is denser and the pressure
levels are closer together. When the aircraft is fl ying at an
indicated altitude of 5,000 feet, its true altitude is lower than
it would be if the air were warmer.
Cold Weather Altimeter Errors
A correctly calibrated pressure altimeter indicates true
altitude above mean sea level (MSL) when operating within
the International Standard Atmosphere (ISA) parameters of
pressure and temperature. Nonstandard pressure conditions are
corrected by applying the correct local area altimeter setting.
Temperature errors from ISA result in true altitude being
higher than indicated altitude whenever the temperature is
warmer than ISA and true altitude being lower than indicated
altitude whenever the temperature is colder than ISA.
True altitude variance under conditions of colder than ISA
temperatures poses the risk of inadequate obstacle clearance.
Under extremely cold conditions, pilots may need to add an
appropriate temperature correction determined from the chart
in Figure 3-7 to charted IFR altitudes to ensure terrain and
obstacle clearance with the following restrictions:
• Altitudes specifi cally assigned by Air Traffi c Control
(ATC), such as “maintain 5,000 feet” shall not be
corrected. Assigned altitudes may be rejected if the
pilot decides that low temperatures pose a risk of
inadequate terrain or obstacle clearance.
• If temperature corrections are applied to charted
IFR altitudes (such as procedure turn altitudes, fi nal
approach fi x crossing altitudes, etc.), the pilot must
advise ATC of the applied correction.
ICAO Cold Temperature Error Table
The cold temperature induced altimeter error may be
significant when considering obstacle clearances when
temperatures are well below standard. Pilots may wish to
increase their minimum terrain clearance altitudes with a
corresponding increase in ceiling from the normal minimum
when fl ying in extreme cold temperature conditions. Higher
altitudes may need to be selected when fl ying at low terrain
clearances. Most fl ight management systems (FMS) with
air data computers implement a capability to compensate
for cold temperature errors. Pilots fl ying with these systems
should ensure they are aware of the conditions under which
the system will automatically compensate. If compensation
is applied by the FMS or manually, ATC must be informed
that the aircraft is not fl ying the assigned altitude. Otherwise,
vertical separation from other aircraft may be reduced
creating a potentially hazardous situation. The table in
Figure 3-7, derived from International Civil Aviation
3-6
Figure 3-7. ICAO Cold Temperature Error Table.
Organization (ICAO) standard formulas, shows how much
error can exist when the temperature is extremely cold. To
use the table, fi nd the reported temperature in the left column,
and then read across the top row to the height above the
airport/reporting station. Subtract the airport elevation from
the altitude of the fi nal approach fi x (FAF). The intersection
of the column and row is the amount of possible error.
Example: The reported temperature is -10° Celsius and the
FAF is 500 feet above the airport elevation. The reported
current altimeter setting may place the aircraft as much as 50
feet below the altitude indicated by the altimeter.
When using the cold temperature error table, the altitude
error is proportional to both the height above the reporting
station elevation and the temperature at the reporting
station. For IFR approach procedures, the reporting station
elevation is assumed to be airport elevation. It is important
to understand that corrections are based upon the temperature
at the reporting station, not the temperature observed at the
aircraft’s current altitude and height above the reporting
station and not the charted IFR altitude.
To see how corrections are applied, note the following
example:
Airport Elevation 496 feet
Airport Temperature - 50° C
A charted IFR approach to the airport provides the following
data:
Minimum Procedure Turn Altitude 1,800 feet
Minimum FAF Crossing Altitude 1,200 feet
Straight-in Minimum Descent Altitude 800 feet
Circling MDA 1,000 feet
The Minimum Procedure Turn Altitude of 1,800 feet will
be used as an example to demonstrate determination of
the appropriate temperature correction. Typically, altitude
values are rounded up to the nearest 100-foot level. The
charted procedure turn altitude of 1,800 feet minus the airport
elevation of 500 feet equals 1,300 feet. The altitude difference
of 1,300 feet falls between the correction chart elevations of
1,000 feet and 1,500 feet. At the station temperature of -50°C,
the correction falls between 300 feet and 450 feet. Dividing
the difference in compensation values by the difference in
altitude above the airport gives the error value per foot.
In this case, 150 feet divided by 500 feet = 0.33 feet for each
additional foot of altitude above 1,000 feet. This provides a
correction of 300 feet for the fi rst 1,000 feet and an additional
value of 0.33 times 300 feet, or 99 feet, which is rounded to
100 feet. 300 feet + 100 feet = total temperature correction
of 400 feet. For the given conditions, correcting the charted
value of 1,800 feet above MSL (equal to a height above the
reporting station of 1,300 feet) requires the addition of 400
feet. Thus, when fl ying at an indicated altitude of 2,200 feet,
the aircraft is actually fl ying a true altitude of 1,800 feet.
Minimum Procedure Turn Altitude
1,800 feet charted = 2,200 feet corrected
Minimum FAF Crossing Altitude
1,200 feet charted = 1,500 feet corrected
Straight-in MDA
800 feet charted = 900 feet corrected
Circling MDA
1,000 feet charted = 1,200 feet corrected
Nonstandard Pressure on an Altimeter
Maintaining a current altimeter setting is critical because the
atmosphere pressure is not constant. That is, in one location
the pressure might be higher than the pressure just a short
distance away. Take an aircraft whose altimeter setting is set
to 29.92" of local pressure. As the aircraft moves to an area
of lower pressure (Point A to B in Figure 3-8) and the pilot
fails to readjust the altimeter setting (essentially calibrating
it to local pressure), then as the pressure decreases, the
indicated altitude will be lower. Adjusting the altimeter
3-7
Figure 3-8. Effects of Nonstandard Pressure on an Altimeter of an
Aircraft Flown into Air of Lower Than Standard Pressure (Air is
Less Dense).
settings compensates for this. When the altimeter shows an
indicated altitude of 5,000 feet, the true altitude at Point A
(the height above mean sea level) is only 3,500 feet at Point
B. The fact that the altitude indication is not always true lends
itself to the memory aid, “When fl ying from hot to cold or
from a high to a low, look out below.”
Altimeter Enhancements (Encoding)
It is not suffi cient in the airspace system for only the pilot
to have an indication of the aircraft’s altitude; the air traffi c
controller on the ground must also know the altitude of the
aircraft. To provide this information, the aircraft is typically
equipped with an encoding altimeter.
When the ATC transponder is set to Mode C, the encoding
altimeter supplies the transponder with a series of pulses
identifying the fl ight level (in increments of 100 feet) at
which the aircraft is fl ying. This series of pulses is transmitted
to the ground radar where they appear on the controller’s
scope as an alphanumeric display around the return for the
aircraft. The transponder allows the ground controller to
identify the aircraft and determine the pressure altitude at
which it is fl ying.
A computer inside the encoding altimeter measures the
pressure referenced from 29.92" Hg and delivers this data to
the transponder. When the pilot adjusts the barometric scale
to the local altimeter setting, the data sent to the transponder
is not affected. This is to ensure that all Mode C aircraft are
transmitting data referenced to a common pressure level. ATC
equipment adjusts the displayed altitudes to compensate for
local pressure differences allowing display of targets at correct
altitudes. 14 CFR part 91 requires the altitude transmitted by
the transponder to be within 125 feet of the altitude indicated
on the instrument used to maintain fl ight altitude.
Reduced Vertical Separation Minimum (RVSM)
Below 31,000 feet, a 1,000 foot separation is the minimum
required between usable fl ight levels. Flight levels (FLs)
generally start at 18,000 feet where the local pressure is
29.92" Hg or greater. All aircraft 18,000 feet and above use
a standard altimeter setting of 29.92" Hg, and the altitudes
are in reference to a standard hence termed FL. Between FL
180 and FL 290, the minimum altitude separation is 1,000
feet between aircraft. However, for fl ight above FL 290
(primarily due to aircraft equipage and reporting capability;
potential error) ATC applied the requirement of 2,000 feet of
separation. FL 290, an altitude appropriate for an eastbound
aircraft, would be followed by FL 310 for a westbound
aircraft, and so on to FL 410, or seven FLs available for fl ight.
With 1,000-foot separation, or a reduction of the vertical
separation between FL 290 and FL 410, an additional six
FLs become available. This results in normal fl ight level and
direction management being maintained from FL 180 through
FL 410. Hence the name is Reduced Vertical Separation
Minimum (RVSM). Because it is applied domestically, it is
called United States Domestic Reduced Vertical Separation
Minimum, or DRVSM.
However, there is a cost to participate in the DRVSM program
which relates to both aircraft equipage and pilot training. For
example, altimetry error must be reduced signifi cantly and
operators using RVSM must receive authorization from the
appropriate civil aviation authority. RVSM aircraft must
meet required altitude-keeping performance standards.
Additionally, operators must operate in accordance with
RVSM policies/procedures applicable to the airspace where
they are fl ying.
The aircraft must be equipped with at least one automatic
altitude control—
• Within a tolerance band of ±65 feet about an acquired
altitude when the aircraft is operated in straight-andlevel
fl ight.
• Within a tolerance band of ±130 feet under no
turbulent, conditions for aircraft for which application
for type certifi cation occurred on or before April 9,
1997 that are equipped with an automatic altitude
control system with fl ight management/performance
system inputs.
3-8
Figure 3-9. Increase in Aircraft Permitted Between FL 180 and
FL 410.
Figure 3-10. Rate of Climb or Descent in Thousands of Feet Per
Minute.
That aircraft must be equipped with an altitude alert system
that signals an alert when the altitude displayed to the fl ight
crew deviates from the selected altitude by more than (in most
cases) 200 feet. For each condition in the full RVSM fl ight
envelope, the largest combined absolute value for residual
static source error plus the avionics error may not exceed 200
feet. Aircraft with TCAS must have compatibility with RVSM
Operations. Figure 3-9 illustrates the increase in aircraft
permitted between FL 180 and FL 410. Most noteworthy,
however, is the economization that aircraft can take advantage
of by the higher FLs being available to more aircraft.
Vertical Speed Indicator (VSI)
The VSI in Figure 3-10 is also called a vertical velocity
indicator (VVI), and was formerly known as a rate-ofclimb
indicator. It is a rate-of-pressure change instrument
that gives an indication of any deviation from a constant
pressure level.
Inside the instrument case is an aneroid very much like the
one in an ASI. Both the inside of this aneroid and the inside
of the instrument case are vented to the static system, but
the case is vented through a calibrated orifi ce that causes
the pressure inside the case to change more slowly than
the pressure inside the aneroid. As the aircraft ascends, the
static pressure becomes lower. The pressure inside the case
compresses the aneroid, moving the pointer upward, showing
a climb and indicating the rate of ascent in number of feet
per minute (fpm).
When the aircraft levels off, the pressure no longer changes.
The pressure inside the case becomes equal to that inside
the aneroid, and the pointer returns to its horizontal, or
zero, position. When the aircraft descends, the static
pressure increases. The aneroid expands, moving the pointer
downward, indicating a descent.
The pointer indication in a VSI lags a few seconds behind
the actual change in pressure. However, it is more sensitive
than an altimeter and is useful in alerting the pilot of an
upward or downward trend, thereby helping maintain a
constant altitude.
Some of the more complex VSIs, called instantaneous vertical
speed indicators (IVSI), have two accelerometer-actuated air
pumps that sense an upward or downward pitch of the aircraft
and instantaneously create a pressure differential. By the time
the pressure caused by the pitch acceleration dissipates, the
altitude pressure change is effective.
Dynamic Pressure Type Instruments
Airspeed Indicator (ASI)
An ASI is a differential pressure gauge that measures the
dynamic pressure of the air through which the aircraft is
fl ying. Dynamic pressure is the difference in the ambient
static air pressure and the total, or ram, pressure caused by
the motion of the aircraft through the air. These two pressures
are taken from the pitot-static system.
3-9
Figure 3-11. Mechanism of an Airspeed Indicator.
Equivalent Airspeed (EAS)
EAS is CAS corrected for compression of the air inside the
pitot tube. EAS is the same as CAS in standard atmosphere
at sea level. As the airspeed and pressure altitude increase,
the CAS becomes higher than it should be, and a correction
for compression must be subtracted from the CAS.
True Airspeed (TAS)
TAS is CAS corrected for nonstandard pressure and
temperature. TAS and CAS are the same in standard
atmosphere at sea level. Under nonstandard conditions, TAS
is found by applying a correction for pressure altitude and
temperature to the CAS.
Some aircraft are equipped with true ASIs that have a
temperature-compensated aneroid bellows inside the
instrument case. This bellows modifi es the movement of
the rocking shaft inside the instrument case so the pointer
shows the actual TAS.
The TAS indicator provides both true and IAS. These
instruments have the conventional airspeed mechanism,
with an added subdial visible through cutouts in the regular
dial. A knob on the instrument allows the pilot to rotate the
subdial and align an indication of the outside air temperature
with the pressure altitude being fl own. This alignment causes
the instrument pointer to indicate the TAS on the subdial.
The mechanism of the ASI in Figure 3-11 consists of a thin,
corrugated phosphor bronze aneroid, or diaphragm, that
receives its pressure from the pitot tube. The instrument
case is sealed and connected to the static ports. As the
pitot pressure increases or the static pressure decreases, the
diaphragm expands. This dimensional change is measured by
a rocking shaft and a set of gears that drives a pointer across
the instrument dial. Most ASIs are calibrated in knots, or
nautical miles per hour; some instruments show statute miles
per hour, and some instruments show both.
Types of Airspeed
Just as there are several types of altitude, there are multiple
types of airspeed: Indicated Airspeed (IAS), Calibrated
Airspeed (CAS), Equivalent Airspeed (EAS), and True
Airspeed (TAS).
Indicated Airspeed (IAS)
IAS is shown on the dial of the instrument, uncorrected for
instrument or system errors.
Calibrated Airspeed (CAS)
CAS is the speed at which the aircraft is moving through
the air, which is found by correcting IAS for instrument
and position errors. The POH/AFM has a chart or graph to
correct IAS for these errors and provide the correct CAS for
the various fl ap and landing gear confi gurations.
3-10
Figure 3-12. A true airspeed indicator allows the pilot to correct
IAS for nonstandard temperature and pressure.
Figure 3-13. A Machmeter shows the ratio of the speed of sound to
the TAS the aircraft is fl ying.
Figure 3-14. A maximum allowable airspeed indicator has a movable
pointer that indicates the never-exceed speed, which changes with
altitude to avoid the onset of transonic shock waves.
Most high-speed aircraft are limited to a maximum Mach
number at which they can fl y. This is shown on a Machmeter
as a decimal fraction. For example, if the
Machmeter indicates .83 and the aircraft is fl ying at 30,000
feet where the speed of sound under standard conditions is
589.5 knots, the airspeed is 489.3 knots. The speed of sound
varies with the air temperature. If the aircraft were fl ying at
Mach .83 at 10,000 feet where the air is much warmer, its
airspeed would be 530 knots.
Maximum Allowable Airspeed
Some aircraft that fl y at high subsonic speeds are equipped
with maximum allowable ASIs like the one in Figure 3-14.
This instrument looks much like a standard air-speed indicator,
calibrated in knots, but has an additional pointer colored red,
checkered, or striped. The maximum airspeed pointer is
actuated by an aneroid, or altimeter mechanism, that moves
it to a lower value as air density decreases. By keeping the
airspeed pointer at a lower value than the maximum pointer,
the pilot avoids the onset of transonic shock waves.
Airspeed Color Codes
The dial of an ASI is color coded to alert the pilot, at a
glance, of the signifi cance of the speed at which the aircraft
is fl ying. These colors and their associated airspeeds are
shown in Figure 3-15.
Magnetism
The Earth is a huge magnet, spinning in space, surrounded
by a magnetic fi eld made up of invisible lines of fl ux. These
lines leave the surface at the magnetic north pole and reenter
at the magnetic South Pole.
Lines of magnetic fl ux have two important characteristics:
any magnet that is free to rotate will align with them, and
Mach Number
As an aircraft approaches the speed of sound, the air fl owing
over certain areas of its surface speeds up until it reaches the
speed of sound, and shock waves form. The IAS at which
these conditions occur changes with temperature. Therefore,
in this case, airspeed is not entirely adequate to warn the
pilot of the impending problems. Mach number is more
useful. Mach number is the ratio of the TAS of the aircraft
to the speed of sound in the same atmospheric conditions.
An aircraft fl ying at the speed of sound is fl ying at Mach
1.0. Some older mechanical Machmeters not driven from
an air data computer use an altitude aneroid inside the
instrument that converts pitot-static pressure into Mach
number. These systems assume that the temperature at any
altitude is standard; therefore, the indicated Mach number is
inaccurate whenever the temperature deviates from standard.
These systems are called indicated Machmeters. Modern
electronic Machmeters use information from an air data
computer system to correct for temperature errors. These
systems display true Mach number.
3-11
Figure 3-15. Color Codes for an Airspeed Indicator.
Figure 3-16. A Magnetic Compass. The vertical line is called the
lubber line.
an electrical current is induced into any conductor that cuts
across them. Most direction indicators installed in aircraft
make use of one of these two characteristics.
The Basic Aviation Magnetic Compass
One of the oldest and simplest instruments for indicating
direction is the magnetic compass. It is also one of the basic
instruments required by 14 CFR part 91 for both VFR and
IFR fl ight.
Magnetic Compass Overview
A magnet is a piece of material, usually a metal containing
iron, which attracts and holds lines of magnetic flux.
Regardless of size, every magnet has two poles: a north
pole and a south pole. When one magnet is placed in the
fi eld of another, the unlike poles attract each other and like
poles repel.
An aircraft magnetic compass, such as the one in Figure 3-16,
has two small magnets attached to a metal fl oat sealed inside a
bowl of clear compass fl uid similar to kerosene. A graduated
scale, called a card, is wrapped around the fl oat and viewed
through a glass window with a lubber line across it. The card
is marked with letters representing the cardinal directions,
north, east, south, and west, and a number for each 30°
between these letters. The fi nal “0” is omitted from these
directions; for example, 3 = 30°, 6 = 60°, and 33 = 330°.
There are long and short graduation marks between the letters
and numbers, with each long mark representing 10° and each
short mark representing 5°.
Magnetic Compass Construction
The fl oat and card assembly has a hardened steel pivot in its
center that rides inside a special, spring-loaded, hard-glass
jewel cup. The buoyancy of the fl oat takes most of the weight
off the pivot, and the fl uid damps the oscillation of the fl oat
and card. This jewel-and-pivot type mounting allows the fl oat
freedom to rotate and tilt up to approximately 18° angle of
bank. At steeper bank angles, the compass indications are
erratic and unpredictable.
The compass housing is entirely full of compass fl uid. To
prevent damage or leakage when the fl uid expands and
contracts with temperature changes, the rear of the compass
case is sealed with a fl exible diaphragm, or with a metal
bellows in some compasses.
Magnetic Compass Theory of Operations
The magnets align with the Earth’s magnetic fi eld and the
pilot reads the direction on the scale opposite the lubber line.
Note that in Figure 3-16, the pilot sees the compass card from
its backside. When the pilot is fl ying north as the compass
shows, east is to the pilot’s right, but on the card “33”, which
represents 330° (west of north), is to the right of north. The
reason for this apparent backward graduation is that the card
remains stationary, and the compass housing and the pilot turn
around it, always viewing the card from its backside.
3-12
Figure 3-17. Isogonic lines are lines of equal variation.
Figure 3-18. Utilization of a Compass Rose Aids Compensation
for Deviation Errors.
A compensator assembly mounted on the top or bottom of the
compass allows an aviation maintenance technician (AMT)
to create a magnetic fi eld inside the compass housing that
cancels the infl uence of local outside magnetic fi elds. This is
done to correct for deviation error. The compensator assembly
has two shafts whose ends have screwdriver slots accessible
from the front of the compass. Each shaft rotates one or two
small compensating magnets. The end of one shaft is marked
E-W, and its magnets affect the compass when the aircraft is
pointed east or west. The other shaft is marked N-S and its
magnets affect the compass when the aircraft is pointed north
or south.
Magnetic Compass Induced Errors
The magnetic compass is the simplest instrument in the
panel, but it is subject to a number of errors that must be
considered.
Variation
The Earth rotates about its geographic axis; maps and charts
are drawn using meridians of longitude that pass through the
geographic poles. Directions measured from the geographic
poles are called true directions. The north magnetic pole to
which the magnetic compass points is not collocated with
the geographic north pole, but is some 1,300 miles away;
directions measured from the magnetic poles are called
magnetic directions. In aerial navigation, the difference
between true and magnetic directions is called variation. This
same angular difference in surveying and land navigation is
called declination.
Figure 3-17 shows the isogonic lines that identify the number
of degrees of variation in their area. The line that passes near
Chicago is called the agonic line. Anywhere along this line
the two poles are aligned, and there is no variation. East of
this line, the magnetic pole is to the west of the geographic
pole and a correction must be applied to a compass indication
to get a true direction.
Flying in the Washington, D.C. area, for example, the variation
is 10° west. If the pilot wants to fl y a true course of south (180°),
the variation must be added to this resulting in a magnetic course
to fl y of 190°. Flying in the Los Angeles, CA area, the variation
is 14° east. To fl y a true course of 180° there, the pilot would
have to subtract the variation and fl y a magnetic course of 166°.
The variation error does not change with the heading of the
aircraft; it is the same anywhere along the isogonic line.
Deviation
The magnets in a compass align with any magnetic fi eld.
Local magnetic fi elds in an aircraft caused by electrical current
fl owing in the structure, in nearby wiring or any magnetized
part of the structure, confl ict with the Earth’s magnetic fi eld
and cause a compass error called deviation.
Deviation, unlike variation, is different on each heading, but it is
not affected by the geographic location. Variation error cannot
be reduced or changed, but deviation error can be minimized
when a pilot or AMT performs the maintenance task known
as “swinging the compass.”
Most airports have a compass rose, which is a series of lines
marked out on a taxiway or ramp at some location where there
is no magnetic interference. Lines, oriented to magnetic north,
are painted every 30°, as shown in Figure 3-18.
The pilot or AMT aligns the aircraft on each magnetic
heading and adjusts the compensating magnets to minimize
the difference between the compass indication and the actual
magnetic heading of the aircraft. Any error that cannot be
removed is recorded on a compass correction card, like the one
in Figure 3-19, and placed in a cardholder near the compass.
If the pilot wants to fl y a magnetic heading of 120° and the
3-13
Figure 3-19. A compass correction card shows the deviation
correction for any heading.
Figure 3-20. Northerly Turning Error.
aircraft is operating with the radios on, the pilot should fl y a
compass heading of 123°.
The corrections for variation and deviation must be applied
in the correct sequence and is shown below starting from the
true course desired.
Step 1: Determine the Magnetic Course
True Course (180°) ± Variation (+10°) = Magnetic Course (190°)
The Magnetic Course (190°) is steered if there is no deviation
error to be applied. The compass card must now be considered
for the compass course of 190°.
Step 2: Determine the Compass Course
Magnetic Course (190°, from step 1) ± Deviation (-2°, from
correction card) = Compass Course (188°)
NOTE: Intermediate magnetic courses between those listed on
the compass card need to be interpreted. Therefore, to steer a true
course of 180°, the pilot would follow a compass course of 188°.
To fi nd the true course that is being fl own when the compass
course is known:
Compass Course ± Deviation = Magnetic Course ± Variation
= True Course
Dip Errors
The lines of magnetic fl ux are considered to leave the Earth at
the magnetic north pole and enter at the magnetic South Pole. At
both locations the lines are perpendicular to the Earth’s surface.
At the magnetic equator, which is halfway between the poles,
the lines are parallel with the surface. The magnets in a compass
align with this fi eld, and near the poles they dip, or tilt, the fl oat
and card. The fl oat is balanced with a small dip-compensating
weight, so it stays relatively level when operating in the middle
latitudes of the northern hemisphere. This dip along with this
weight causes two very noticeable errors: northerly turning error
and acceleration error.
The pull of the vertical component of the Earth’s magnetic fi eld
causes northerly turning error, which is apparent on a heading
of north or south. When an aircraft fl ying on a heading of north
makes a turn toward east, the aircraft banks to the right, and the
compass card tilts to the right. The vertical component of the
Earth’s magnetic fi eld pulls the north-seeking end of the magnet
to the right, and the fl oat rotates, causing the card to rotate toward
west, the direction opposite the direction the turn is being made.
If the turn is made from north to west, the aircraft banks to the left
and the compass card tilts down on the left side. The magnetic
fi eld pulls on the end of the magnet that causes the card to rotate
toward east. This indication is again opposite to the direction
the turn is being made. The rule for this error is: when starting
3-14
Figure 3-21. The Effects of Acceleration Error.
a turn from a northerly heading, the compass indication lags
behind the turn.
When an aircraft is fl ying on a heading of south and begins
a turn toward east, the Earth’s magnetic fi eld pulls on the
end of the magnet that rotates the card toward east, the same
direction the turn is being made. If the turn is made from south
toward west, the magnetic pull starts the card rotating toward
west—again, in the same direction the turn is being made. The
rule for this error is: When starting a turn from a southerly
heading, the compass indication leads the turn.
In acceleration error, the dip-correction weight causes the end
of the fl oat and card marked N (the south-seeking end) to be
heavier than the opposite end. When the aircraft is fl ying at
a constant speed on a heading of east or west, the fl oat and
card is level. The effects of magnetic dip and the weight are
approximately equal. If the aircraft accelerates on a heading
of east , the inertia of the weight holds its end of
the fl oat back and the card rotates toward north. As soon as the
speed of the aircraft stabilizes, the card swings back to its east
indication. If, while fl ying on this easterly heading, the aircraft
decelerates, the inertia causes the weight to move ahead and the
card rotates toward south until the speed again stabilizes.
When fl ying on a heading of west, the same things happen.
Inertia from acceleration causes the weight to lag, and the
card rotates toward north. When the aircraft decelerates on a
heading of west, inertia causes the weight to move ahead and
the card rotates toward south.
Oscillation Error
Oscillation is a combination of all of the other errors, and it
results in the compass card swinging back and forth around
the heading being flown. When setting the gyroscopic
heading indicator to agree with the magnetic compass, use
the average indication between the swings.
The Vertical Card Magnetic Compass
The fl oating magnet type of compass not only has all the
errors just described, but also lends itself to confused reading.
It is easy to begin a turn in the wrong direction because its card
appears backward. East is on what the pilot would expect to be
the west side. The vertical card magnetic compass eliminates
some of the errors and confusion. The dial of this compass
is graduated with letters representing the cardinal directions,
numbers every 30°, and marks every 5°. The dial is rotated by
a set of gears from the shaft-mounted magnet, and the nose
of the symbolic airplane on the instrument glass represents
the lubber line for reading the heading of the aircraft from
the dial. Eddy currents induced into an aluminum-damping
cup damp oscillation of the magnet.
The Flux Gate Compass System
As mentioned earlier, the lines of fl ux in the Earth’s magnetic
fi eld have two basic characteristics: a magnet aligns with
these lines, and an electrical current is induced, or generated,
in any wire crossed by them.
3-15
Figure 3-22. Vertical Card Magnetic Compass.
Figure 3-23. The soft iron frame of the fl ux valve accepts the fl ux
from the Earth’s magnetic fi eld each time the current in the center
coil reverses. This fl ux causes current to fl ow in the three pickup
coils.
Figure 3-24. The current in each of the three pickup coils changes
with the heading of the aircraft.
Figure 3-25. Pictorial Navigation Indicator (HSI Top), Slaving
Control and Compensator Unit.
The fl ux gate compass that drives slaved gyros uses the
characteristic of current induction. The fl ux valve is a small,
segmented ring, like the one in Figure 3-23, made of soft
iron that readily accepts lines of magnetic fl ux. An electrical
coil is wound around each of the three legs to accept the
current induced in this ring by the Earth’s magnetic fi eld. A
coil wound around the iron spacer in the center of the frame
has 400-Hz alternating current (A.C.) fl owing through it.
During the times when this current reaches its peak, twice
during each cycle, there is so much magnetism produced by
this coil that the frame cannot accept the lines of fl ux from
the Earth’s fi eld.
But as the current reverses between the peaks, it demagnetizes
the frame so it can accept the fl ux from the Earth’s fi eld. As
this fl ux cuts across the windings in the three coils, it causes
current to fl ow in them. These three coils are connected in
such a way that the current fl owing in them changes as the
heading of the aircraft changes.
The three coils are connected to three similar but smaller coils
in a synchro inside the instrument case. The synchro rotates
the dial of a radio magnetic indicator (RMI) or a horizontal
situation indicator (HSI).
Remote Indicating Compass
Remote indicating compasses were developed to compensate
for the errors and limitations of the older type of heading
indicators. The two panel-mounted components of a typical
system are the pictorial navigation indicator and the slaving
control and compensator unit. The pictorial
navigation indicator is commonly referred to as a HSI.
3-16
Figure 3-26. Driven by signals from a fl ux valve, the compass card
in this RMI indicates the heading of the aircraft opposite the upper
center index mark. The green pointer is driven by the ADF.
The slaving control and compensator unit has a pushbutton
that provides a means of selecting either the “slaved gyro”
or “free gyro” mode. This unit also has a slaving meter
and two manual heading-drive buttons. The slaving meter
indicates the difference between the displayed heading and
the magnetic heading. A right defl ection indicates a clockwise
error of the compass card; a left deflection indicates a
counterclockwise error. Whenever the aircraft is in a turn
and the card rotates, the slaving meter shows a full defl ection
to one side or the other. When the system is in “free gyro”
mode, the compass card may be adjusted by depressing the
appropriate heading-drive button.
A separate unit, the magnetic slaving transmitter is mounted
remotely; usually in a wingtip to eliminate the possibility of
magnetic interference. It contains the fl ux valve, which is
the direction-sensing device of the system. A concentration
of lines of magnetic force, after being amplifi ed, becomes
a signal relayed to the heading indicator unit, which is also
remotely mounted. This signal operates a torque motor in
the heading indicator unit that processes the gyro unit until
it is aligned with the transmitter signal. The magnetic slaving
transmitter is connected electrically to the HSI.
There are a number of designs of the remote indicating
compass; therefore, only the basic features of the system are
covered here. Instrument pilots must become familiar with
the characteristics of the equipment in their aircraft.
As instrument panels become more crowded and the pilot’s
available scan time is reduced by a heavier fl ight deck
workload, instrument manufacturers have worked toward
combining instruments. One good example of this is the
RMI in Figure 3-26. The compass card is driven by signals
from the fl ux valve, and the two pointers are driven by an
automatic direction fi nder (ADF) and a very high frequency
omnidirectional range (VOR).
Gyroscopic Systems
Flight without reference to a visible horizon can be safely
accomplished by the use of gyroscopic instrument systems
and the two characteristics of gyroscopes, which are rigidity
and precession. These systems include attitude, heading,
and rate instruments, along with their power sources. These
instruments include a gyroscope (or gyro) that is a small wheel
with its weight concentrated around its periphery. When this
wheel is spun at high speed, it becomes rigid and resists tilting
or turning in any direction other than around its spin axis.
Attitude and heading instruments operate on the principle
of rigidity. For these instruments, the gyro remains rigid
in its case and the aircraft rotates about it. Rate indicators,
such as turn indicators and turn coordinators, operate on the
principle of precession. In this case, the gyro processes (or
rolls over) proportionate to the rate the aircraft rotates about
one or more of its axes.
Power Sources
Aircraft and instrument manufacturers have designed
redundancy in the fl ight instruments so that any single failure
will not deprive the pilot of the ability to safely conclude
the fl ight. Gyroscopic instruments are crucial for instrument
fl ight; therefore, they are powered by separate electrical or
pneumatic sources.
Pneumatic Systems
Pneumatic gyros are driven by a jet of air impinging on
buckets cut into the periphery of the wheel. On many aircraft
this stream of air is obtained by evacuating the instrument
case with a vacuum source and allowing fi ltered air to fl ow
into the case through a nozzle to spin the wheel.
Venturi Tube Systems
Aircraft that do not have a pneumatic pump to evacuate the
instrument case can use venturi tubes mounted on the outside
of the aircraft, similar to the system shown in Figure 3-27. Air
fl owing through the venturi tube speeds up in the narrowest
part and, according to Bernoulli’s principle, the pressure
drops. This location is connected to the instrument case by
a piece of tubing. The two attitude instruments operate on
approximately 4" Hg of suction; the turn-and-slip indicator
needs only 2" Hg, so a pressure-reducing needle valve is
used to decrease the suction. Air fl ows into the instruments
through fi lters built into the instrument cases. In this system,
ice can clog the venturi tube and stop the instruments when
they are most needed.
3-17
Figure 3-27. A venturi tube system that provides necessary vacuum
to operate key instruments.
Figure 3-28. Single-engine instrument vacuum system using a steel-vane wet-type vacuum pump.
Vacuum Pump Systems
Wet-Type Vacuum Pump
Steel-vane air pumps have been used for many years to
evacuate the instrument cases. The vanes in these pumps
are lubricated by a small amount of engine oil metered into
the pump and discharged with the air. In some aircraft the
discharge air is used to infl ate rubber deicer boots on the
wing and empennage leading edges. To keep the oil from
deteriorating the rubber boots, it must be removed with an
oil separator like the one in Figure 3-28.
The vacuum pump moves a greater volume of air than is
needed to supply the instruments with the suction needed,
so a suction-relief valve is installed in the inlet side of the
pump. This spring-loaded valve draws in just enough air to
maintain the required low pressure inside the instruments,
as is shown on the suction gauge in the instrument panel.
Filtered air enters the instrument cases from a central air
fi lter. As long as aircraft fl y at relatively low altitudes, enough
air is drawn into the instrument cases to spin the gyros at a
suffi ciently high speed.
Dry Air Vacuum Pump
As fl ight altitudes increase, the air is less dense and more air
must be forced through the instruments. Air pumps that do not
mix oil with the discharge air are used in high fl ying aircraft.
3-18
Figure 3-29. Twin-Engine Instrument Pressure System Using a Carbon-Vane Dry-Type Air Pump.
Steel vanes sliding in a steel housing need to be lubricated,
but vanes made of a special formulation of carbon sliding
inside carbon housing provide their own lubrication in a
microscopic amount as they wear.
Pressure Indicating Systems
Figure 3-29 is a diagram of the instrument pneumatic
system of a twin-engine general aviation airplane. Two dry
air pumps are used with fi lters in their inlet to fi lter out any
contaminants that could damage the fragile carbon vanes in
the pump. The discharge air from the pump fl ows through
a regulator, where excess air is bled off to maintain the
pressure in the system at the desired level. The regulated air
then fl ows through inline fi lters to remove any contamination
that could have been picked up from the pump, and from
there into a manifold check valve. If either engine should
become inoperative or either pump should fail, the check
valve isolates the inoperative system and the instruments are
driven by air from the operating system. After the air passes
through the instruments and drives the gyros, it is exhausted
from the case. The gyro pressure gauge measures the pressure
drop across the instruments.
Electrical Systems
Many general aviation aircraft that use pneumatic attitude
indicators use electric rate indicators and/or the reverse. Some
instruments identify their power source on their dial, but it
is extremely important that pilots consult the POH/AFM to
determine the power source of all instruments to know what
action to take in the event of an instrument failure. Direct
current (D.C.) electrical instruments are available in 14- or
28-volt models, depending upon the electrical system in
the aircraft. A.C. is used to operate some attitude gyros and
autopilots. Aircraft with only D.C. electrical systems can use
A.C. instruments via installation of a solid-state D.C. to A.C.
inverter, which changes 14 or 28 volts D.C. into three-phase
115-volt, 400-Hz A.C.
Gyroscopic Instruments
Attitude Indicators
The fi rst attitude instrument (AI) was originally referred to as
an artifi cial horizon, later as a gyro horizon; now it is more
properly called an attitude indicator. Its operating mechanism
is a small brass wheel with a vertical spin axis, spun at a high
speed by either a stream of air impinging on buckets cut into
its periphery, or by an electric motor. The gyro is mounted in
a double gimbal, which allows the aircraft to pitch and roll
about the gyro as it remains fi xed in space.
A horizon disk is attached to the gimbals so it remains in
the same plane as the gyro, and the aircraft pitches and
rolls about it. On early instruments, this was just a bar that
3-19
Figure 3-30. The dial of this attitude indicator has reference lines
to show pitch and roll.
represented the horizon, but now it is a disc with a line
representing the horizon and both pitch marks and bank-angle
lines. The top half of the instrument dial and horizon disc
is blue, representing the sky; and the bottom half is brown,
representing the ground. A bank index at the top of the
instrument shows the angle of bank marked on the banking
scale with lines that represent 10°, 20°, 30°, 45°, and 60°.
A small symbolic aircraft is mounted in the instrument case so it
appears to be fl ying relative to the horizon. A knob at the bottom
center of the instrument case raises or lowers the aircraft to
compensate for pitch trim changes as the airspeed changes. The
width of the wings of the symbolic aircraft and the dot in the center
of the wings represent a pitch change of approximately 2°.
For an AI to function properly, the gyro must remain
vertically upright while the aircraft rolls and pitches around
it. The bearings in these instruments have a minimum of
friction; however, even this small amount places a restraint
on the gyro producing precession and causing the gyro to tilt.
To minimize this tilting, an erection mechanism inside the
instrument case applies a force any time the gyro tilts from
its vertical position. This force acts in such a way to return
the spinning wheel to its upright position.
The older artifi cial horizons were limited in the amount of
pitch or roll they could tolerate, normally about 60° in pitch
and 100° in roll. After either of these limits was exceeded,
the gyro housing contacted the gimbals, applying such a
precessing force that the gyro tumbled. Because of this
limitation, these instruments had a caging mechanism that
locked the gyro in its vertical position during any maneuvers
that exceeded the instrument limits. Newer instruments do
not have these restrictive tumble limits; therefore, they do
not have a caging mechanism.
When an aircraft engine is fi rst started and pneumatic or
electric power is supplied to the instruments, the gyro is
not erect. A self-erecting mechanism inside the instrument
actuated by the force of gravity applies a precessing force,
causing the gyro to rise to its vertical position. This erection
can take as long as 5 minutes, but is normally done within
2 to 3 minutes.
Attitude indicators are free from most errors, but depending
upon the speed with which the erection system functions,
there may be a slight nose-up indication during a rapid
acceleration and a nose-down indication during a rapid
deceleration. There is also a possibility of a small bank angle
and pitch error after a 180° turn. These inherent errors are
small and correct themselves within a minute or so after
returning to straight-and-level fl ight.
Heading Indicators
A magnetic compass is a dependable instrument used as a
backup instrument. Although very reliable, it has so many
inherent errors that it has been supplemented with gyroscopic
heading indicators.
The gyro in a heading indicator is mounted in a double gimbal,
as in an attitude indicator, but its spin axis is horizontal
permitting sensing of rotation about the vertical axis of the
aircraft. Gyro heading indicators, with the exception of slaved
gyro indicators, are not north seeking, therefore they must
be manually set to the appropriate heading by referring to
a magnetic compass. Rigidity causes them to maintain this
heading indication, without the oscillation and other errors
inherent in a magnetic compass.
Older directional gyros use a drum-like card marked in the
same way as the magnetic compass card. The gyro and the
card remain rigid inside the case with the pilot viewing the
card from the back. This creates the possibility the pilot might
start a turn in the wrong direction similar to using a magnetic
compass. A knob on the front of the instrument, below the
dial, can be pushed in to engage the gimbals. This locks the
gimbals allowing the pilot to rotate the gyro and card until
the number opposite the lubber line agrees with the magnetic
compass. When the knob is pulled out, the gyro remains rigid
and the aircraft is free to turn around the card.
Directional gyros are almost all air-driven by evacuating
the case and allowing fi ltered air to fl ow into the case and
out through a nozzle, blowing against buckets cut in the
3-20
Figure 3-31. The heading indicator is not north seeking, but must
be set periodically (about every 15 minutes) to agree with the
magnetic compass.
Figure 3-32. Precession causes a force applied to a spinning
wheel to be felt 90° from the point of application in the direction
of rotation.
Figure 3-33. Turn-and-Slip Indicator.
periphery of the wheel. The Earth constantly rotates at 15°
per hour while the gyro is maintaining a position relative
to space, thus causing an apparent drift in the displayed
heading of 15° per hour. When using these instruments, it
is standard practice to compare the heading indicated on the
directional gyro with the magnetic compass at least every 15
minutes and to reset the heading as necessary to agree with
the magnetic compass.
Heading indicators like the one in Figure 3-31 work on the
same principle as the older horizontal card indicators, except
that the gyro drives a vertical dial that looks much like the
dial of a vertical card magnetic compass. The heading of the
aircraft is shown against the nose of the symbolic aircraft on
the instrument glass, which serves as the lubber line. A knob
in the front of the instrument may be pushed in and turned
to rotate the gyro and dial. The knob is spring loaded so it
disengages from the gimbals as soon as it is released. This
instrument should be checked about every 15 minutes to see
if it agrees with the magnetic compass.
Turn Indicators
Attitude and heading indicators function on the principle
of rigidity, but rate instruments such as the turn-andslip
indicator operate on precession. Precession is the
characteristic of a gyroscope that causes an applied force to
produce a movement, not at the point of application, but at
a point 90° from the point of application in the direction of
rotation.
Turn-and-Slip Indicator
The fi rst gyroscopic aircraft instrument was the turn indicator
in the needle and ball, or turn-and-bank indicator, which
has more recently been called a turn-and-slip indicator.
The inclinometer in the instrument is a black glass ball sealed
inside a curved glass tube that is partially fi lled with a liquid
for damping. This ball measures the relative strength of the
force of gravity and the force of inertia caused by a turn.
When the aircraft is fl ying straight-and-level, there is no
inertia acting on the ball, and it remains in the center of the
tube between two wires. In a turn made with a bank angle
that is too steep, the force of gravity is greater than the inertia
and the ball rolls down to the inside of the turn. If the turn is
made with too shallow a bank angle, the inertia is greater than
gravity and the ball rolls upward to the outside of the turn.
The inclinometer does not indicate the amount of bank, nor
does it indicate slip; it only indicates the relationship between
the angle of bank and the rate of yaw.
3-21
Figure 3-34. The rate gyro in both turn-and-slip indicator and turn
coordinator.
Figure 3-35. A turn coordinator senses rotation about both roll
and yaw axes.
The turn indicator is a small gyro spun either by air or by
an electric motor. The gyro is mounted in a single gimbal
with its spin axis parallel to the lateral axis of the aircraft
and the axis of the gimbal parallel with the longitudinal axis.
When the aircraft yaws, or rotates about its vertical axis, it
produces a force in the horizontal plane that, due to precession,
causes the gyro and its gimbal to rotate about the gimbal’s
axis. It is restrained in this rotation plane by a calibration
spring; it rolls over just enough to cause the pointer to defl ect
until it aligns with one of the doghouse-shaped marks on the
dial, when the aircraft is making a standard rate turn.
The dial of these instruments is marked “2 MIN TURN.” Some
turn-and-slip indicators used in faster aircraft are marked “4
MIN TURN.” In either instrument, a standard rate turn is being
made whenever the needle aligns with a doghouse.
Turn Coordinator
The major limitation of the older turn-and-slip indicator is that
it senses rotation only about the vertical axis of the aircraft. It
tells nothing of the rotation around the longitudinal axis, which
in normal fl ight occurs before the aircraft begins to turn.
A turn coordinator operates on precession, the same as the
turn indicator, but its gimbals frame is angled upward about
30° from the longitudinal axis of the aircraft.
This allows it to sense both roll and yaw. Therefore during
a turn, the indicator fi rst shows the rate of banking and once
stabilized, the turn rate. Some turn coordinator gyros are dualpowered
and can be driven by either air or electricity.
Rather than using a needle as an indicator, the gimbal moves
a dial that is the rear view of a symbolic aircraft. The bezel
of the instrument is marked to show wings-level fl ight and
bank angles for a standard rate turn.
The inclinometer, similar to the one in a turn-and-slip
indicator, is called a coordination ball, which shows the
relationship between the bank angle and the rate of yaw. The
turn is coordinated when the ball is in the center, between the
marks. The aircraft is skidding when the ball rolls toward the
outside of the turn and is slipping when it moves toward the
inside of the turn. A turn coordinator does not sense pitch.
This is indicated on some instruments by placing the words
“NO PITCH INFORMATION” on the dial.
3-22
Figure 3-36. The Kearfott Attitude Heading Reference System (AHRS) on the left incorporates a Monolithic Ring Laser Gyro (MRLG)
(center), which is housed in an Inertial Sensor Assembly (ISA) on the right.
Flight Support Systems
Attitude and Heading Reference System (AHRS)
As aircraft displays have transitioned to new technology,
the sensors that feed them have also undergone signifi cant
change. Traditional gyroscopic flight instruments have
been replaced by Attitude and Heading Reference Systems
(AHRS) improving reliability and thereby reducing cost and
maintenance.
The function of an AHRS is the same as gyroscopic systems;
that is, to determine which way is level and which way is north.
By knowing the initial heading the AHRS can determine both
the attitude and magnetic heading of the aircraft.
The genesis of this system was initiated by the development
of the ring-LASAR gyroscope developed by Kearfott located
in Little Falls, New Jersey. Their development
of the Ring-LASAR gyroscope in the 1960s/1970s was
in support of Department of Defense (DOD) programs to
include cruise missile technology. With the precision of
these gyroscopes, it became readily apparent that they could
be leveraged for multiple tasks and functions. Gyroscopic
miniaturization has become so common that solid-state
gyroscopes are found in products from robotics to toys.
Because the AHRS system replaces separate gyroscopes,
such as those associated with an attitude indicator, magnetic
heading indicator and turn indicator these individual systems
are no longer needed. As with many systems today, AHRS
itself had matured with time. Early AHRS systems used
expensive inertial sensors and fl ux valves. However, today the
AHRS for aviation and general aviation in particular are small
solid-state systems integrating a variety of technology such
as low cost inertial sensors, rate gyros, and magnetometers,
and have capability for satellite signal reception.
Air Data Computer (ADC)
An Air Data Computer (ADC) is an aircraft
computer that receives and processes pitot pressure, static
pressure, and temperature to calculate very precise altitude,
IAS, TAS, and air temperature. The ADC outputs this
information in a digital format that can be used by a variety
of aircraft systems including an EFIS. Modern ADCs
are small solid-state units. Increasingly, aircraft systems
such as autopilots, pressurization, and FMS utilize ADC
information for normal operations. NOTE: In most modern
general aviation systems, both the AHRS and ADC are
integrated within the electronic displays themselves thereby
reducing the number of units, reducing weight, and providing
simplifi cation for installation resulting in reduced costs.
Analog Pictorial Displays
Horizontal Situation Indicator (HSI)
The HSI is a direction indicator that uses the output from
a fl ux valve to drive the dial, which acts as the compass
card. This instrument, shown in Figure 3-37, combines the
magnetic compass with navigation signals and a glide slope.
This gives the pilot an indication of the location of the aircraft
with relationship to the chosen course.
3-23
Figure 3-38. Horizontal Situation Indicator (HSI).
Figure 3-37. Air Data Computer (Collins).
In Figure 3-38, the aircraft heading displayed on the rotating
azimuth card under the upper lubber line is North or 360°.
The course-indicating arrowhead shown is set to 020; the
tail indicates the reciprocal, 200°. The course deviation bar
operates with a VOR/Localizer (VOR/LOC) navigation
receiver to indicate left or right deviations from the course
selected with the course-indicating arrow, operating in the
same manner that the angular movement of a conventional
VOR/LOC needle indicates deviation from course.
The desired course is selected by rotating the courseindicating
arrow in relation to the azimuth card by means
of the course select knob. This gives the pilot a pictorial
presentation: the fi xed aircraft symbol and course deviation
bar display the aircraft relative to the selected course, as
though the pilot were above the aircraft looking down.
The TO/FROM indicator is a triangular pointer. When the
indicator points to the head of the course arrow, it shows
that the course selected, if properly intercepted and fl own,
takes the aircraft to the selected facility. When the indicator
points to the tail of the course arrow, it shows that the course
selected, if properly intercepted and fl own, takes the aircraft
directly away from the selected facility.
The glide slope deviation pointer indicates the relation of
the aircraft to the glide slope. When the pointer is below the
center position, the aircraft is above the glide slope, and an
increased rate of descent is required. In most installations,
the azimuth card is a remote indicating compass driven by
a fl uxgate; however, in few installations where a fl uxgate is
not installed, or in emergency operation, the heading must
be checked against the magnetic compass occasionally and
reset with the course select knob.
Attitude Direction Indicator (ADI)
Advances in attitude instrumentation combine the gyro
horizon with other instruments such as the HSI, thereby
reducing the number of separate instruments to which the
pilot must devote attention. The attitude direction indicator
(ADI) is an example of such technological advancement.
A fl ight director incorporates the ADI within its system,
which is further explained below (Flight Director System).
However, an ADI need not have command cues; however,
it is normally equipped with this feature.
Flight Director System (FDS)
A Flight Director System (FDS) combines many instruments
into one display that provides an easily interpreted
understanding of the aircraft’s fl ight path. The computed
solution furnishes the steering commands necessary to obtain
and hold a desired path.
Major components of an FDS include an ADI, also called
a Flight Director Indicator (FDI), an HSI, a mode selector,
and a fl ight director computer. It should be noted that a
fl ight director in use does not infer the aircraft is being
manipulated by the autopilot (coupled), but is providing
steering commands that the pilot (or the autopilot, if coupled)
follows.
Typical fl ight directors use one of two display systems for
steerage. The fi rst is a set of command bars, one horizontal
and one vertical. The command bars in this confi guration
are maintained in a centered position (much like a centered
glide slope). The second uses a miniature aircraft aligned to
a command cue.
A fl ight director displays steerage commands to the pilot on
the ADI. As previously mentioned, the fl ight director receives
its signals from one of various sources and provides that to the
ADI for steerage commands. The mode controller provides
signals through the ADI to drive the steering bars, e.g., the
3-24
Figure 3-39. A Typical Cue That a Pilot Would Follow.
Figure 3-40. Components of a Typical Flight Director System.
pilot fl ies the aircraft to place the delta symbol in the V of the
steering bars. “Command” indicators tell the pilot in which
direction and how much to change aircraft attitude to achieve
the desired result.
The computed command indications relieve the pilot of
many of the mental calculations required for instrument
fl ight. The yellow cue in the ADI provides all
steering commands to the pilot. It is driven by a computer that
receives information from the navigation systems, the ADC,
AHRS, and other sources of data. The computer processes this
information, providing the pilot with a single cue to follow.
Following the cue provides the pilot with the necessary threedimensional
fl ight trajectory to maintain the desired path.
One of the fi rst widely used fl ight directors was developed
by Sperry and was called the Sperry Three Axis Attitude
Reference System (STARS). Developed in the 1960s, it was
commonly found on both commercial and business aircraft
alike. STARS (with a modifi cation) and successive fl ight
directors were integrated with the autopilots and aircraft
providing a fully integrated fl ight system.
The flight director/autopilot system described below is
typical of installations in many general aviation aircraft.
The components of a typical fl ight director include the mode
controller, ADI, HSI, and annunciator panel. These units are
illustrated in Figure 3-40.
The pilot may choose from among many modes including
the HDG (heading) mode, the VOR/LOC (localizer tracking)
mode, or the AUTO Approach (APP) or G/S (automatic
capture and tracking of instrument landing system (ILS)
localizers and glide path) mode. The auto mode has a fully
automatic pitch selection computer that takes into account
aircraft performance and wind conditions, and operates once
the pilot has reached the ILS glide slope. More sophisticated
systems allow more fl ight director modes.
Integrated Flight Control System
The integrated fl ight control system integrates and merges
various systems into a system operated and controlled by one
principal component. Figure 3-41 illustrates key components
of the fl ight control system that was developed from the
onset as a fully integrated system comprised of the airframe,
autopilot, and fl ight director system. This trend of complete
integration, once seen only in large commercial aircraft, are
now becoming common in the general aviation fi eld.
Autopilot Systems
An autopilot is a mechanical means to control an aircraft
using electrical, hydraulic, or digital systems. Autopilots can
control three axes of the aircraft: roll, pitch, and yaw. Most
autopilots in general aviation control roll and pitch.
Autopilots also function using different methods. The fi rst
is position based. That is, the attitude gyro senses the degree
of difference from a position such as wings level, a change
in pitch, or a heading change.
3-25
Figure 3-41. The S-TEC/Meggit Corporation Integrated Autopilot Installed in the Cirrus.
Figure 3-42. An Autopilot by Century.
Determining whether a design is position based and/or rate
based lies primarily within the type of sensors used. In order
for an autopilot to possess the capability of controlling an
aircraft’s attitude (i.e., roll and pitch), that system must be
provided with constant information on the actual attitude
of that aircraft. This is accomplished by the use of several
different types of gyroscopic sensors. Some sensors are
designed to indicate the aircraft’s attitude in the form of
position in relation to the horizon, while others indicate rate
(position change over time).
Rate-based systems use the turn-and-bank sensor for the
autopilot system. The autopilot uses rate information on
two of the aircraft’s three axes: movement about the vertical
axis (heading change or yaw) and about the longitudinal
axis (roll). This combined information from a single sensor
is made possible by the 30° offset in the gyro’s axis to the
longitudinal axis.
Other systems use a combination of both position and ratebased
information to benefi t from the attributes of both systems
while newer autopilots are digital. Figure 3-42 illustrates an
autopilot by Century.
Figure 3-43 is a diagram layout of a rate-based autopilot by
S-Tec, which permits the purchaser to add modular capability
form basic wing leveling to increased capability.
Flight Management Systems (FMS)
In the mid-1970s, visionaries in the avionics industry such
as Hubert Naimer of Universal, and followed by others such
as Ed King, Jr., were looking to advance the technology of
aircraft navigation. As early as 1976, Naimer had a vision
of a “Master Navigation System” that would accept inputs
from a variety of different types of sensors on an aircraft
and automatically provide guidance throughout all phases
of fl ight.
At that time aircraft navigated over relatively short distances
with radio systems, principally VOR or ADF. For long-range
fl ight inertial navigation systems (INS), Omega, Doppler,
and Loran were in common use. Short-range radio systems
usually did not provide area navigation capability. Longrange
systems were only capable of en route point-to-point
navigation between manually entered waypoints described
as longitude and latitude coordinates, with typical systems
containing a limited number of waypoints.
3-26
Figure 3-43. A Diagram Layout of an Autopilot by S-Tec.
Figure 3-44. A Control Display Unit (CDU) Used to Control the
Flight Management System.
The laborious process of manually entering cryptic latitude
and longitude data for each fl ight waypoint created high
crew workloads and frequently resulted in incorrect data
entry. The requirement of a separate control panel for each
long-range system consumed precious fl ight deck space and
increased the complexity of interfacing the systems with
display instruments, fl ight directors, and autopilots.
The concept employed a master computer interfaced with all
of the navigation sensors on the aircraft. A common control
display unit (CDU) interfaced with the master computer would
provide the pilot with a single control point for all navigation
systems, thereby reducing the number of required fl ight deck
panels. Management of the various individual sensors would
be transferred from the pilot to the new computer.
Since navigation sensors rarely agree exactly about position,
Naimer believed that blending all available sensor position
data through a highly sophisticated, mathematical fi ltering
system would produce a more accurate aircraft position. He
called the process output the “Best Computed Position.” By
using all available sensors to keep track of position, the system
could readily provide area navigation capability. The master
computer, not the individual sensors, would be integrated into
the airplane, greatly reducing wiring complexity.
To solve the problems of manual waypoint entry, a preloaded
database of global navigation information would
be readily accessible by the pilot through the CDU. Using
such a system a pilot could quickly and accurately construct
a fl ight plan consisting of dozens of waypoints, avoiding
the tedious typing of data and the error potential of latitude/
longitude coordinates. Rather than simply navigating pointto-
point, the master system would be able to maneuver the
aircraft, permitting use of the system for terminal procedures
including departures, arrivals, and approaches. The system
would be able to automate any aspect of manual pilot
navigation of the aircraft. When the fi rst system, called the
UNS-1, was released by Universal in 1982, it was called a
fl ight management system (FMS).
3-27
An FMS uses an electronic database of worldwide
navigational data including navigation aids, airways and
intersections, Standard Instrument Departures (SIDs),
Standard Terminal Arrival Routes (STARs), and Instrument
Approach Procedures (IAPs) together with pilot input through
a CDU to create a fl ight plan. The FMS provides outputs to
several aircraft systems including desired track, bearing and
distance to the active waypoint, lateral course deviation and
related data to the fl ight guidance system for the HSI displays,
and roll steering command for the autopilot/fl ight director
system. This allows outputs from the FMS to command
the airplane where to go and when and how to turn. To
support adaptation to numerous aircraft types, an FMS is
usually capable of receiving and outputting both analog and
digital data and discrete information. Currently, electronic
navigation databases are updated every 28 days.
The introduction of the Global Positioning System (GPS) has
provided extremely precise position at low cost, making GPS
the dominant FMS navigation sensor today. Currently, typical
FMS installations require that air data and heading information
be available electronically from the aircraft. This limits FMS
usage in smaller aircraft, but emerging technologies allow this
data from increasingly smaller and less costly systems.
Some systems interface with a dedicated Distance Measuring
Equipment (DME) receiver channel under the control of the
FMS to provide an additional sensor. In these systems, the
FMS determines which DME sites should be interrogated
for distance information using aircraft position and the
navigation database to locate appropriate DME sites. The
FMS then compensates aircraft altitude and station altitude
with the aid of the database to determine the precise distance
to the station. With the distances from a number of sites the
FMS can compute a position nearly as accurately as GPS.
Aimer visualized three-dimensional aircraft control with
an FMS. Modern systems provide Vertical Navigation
(VNAV) as well as Lateral Navigation (LNAV) allowing
the pilot to create a vertical fl ight profi le synchronous with
the lateral fl ight plan. Unlike early systems, such as Inertial
Reference Systems (IRS) that were only suitable for en route
navigation, the modern FMS can guide an aircraft during
instrument approaches.
Today, an FMS provides not only real-time navigation
capability but typically interfaces with other aircraft systems
providing fuel management, control of cabin briefi ng and
display systems, display of uplinked text and graphic weather
data and air/ground data link communications.
Electronic Flight Instrument Systems
Modern technology has introduced into aviation a new
method of displaying fl ight instruments, such as electronic
fl ight instrument systems, integrated fl ight deck displays, and
others. For the purpose of the practical test standards, any
fl ight instrument display that utilizes LCD or picture tube like
displays is referred to as “electronic fl ight instrument display”
and/or a glass fl ight deck. In general aviation there is typically
a primary fl ight display (PFD) and a multi-function display
(MFD). Although both displays are in many cases identical,
the PFD provides the pilot instrumentation necessary for
fl ight to include altitude, airspeed, vertical velocity, attitude,
heading and trim and trend information.
Glass fl ight decks (a term coined to describe electronic fl ight
instrument systems) are becoming more widespread as cost
falls and dependability continually increases. These systems
provide many advantages such as being lighter, more reliable,
no moving parts to wear out, consuming less power, and
replacing numerous mechanical indicators with a single glass
display. Because the versatility offered by glass displays is
much greater than that offered by analog displays, the use
of such systems will only increase with time until analog
systems are eclipsed.
Primary Flight Display (PFD)
PFDs provide increased situational awareness to the pilot by
replacing the traditional six instruments used for instrument
fl ight with an easy-to-scan display that provides the horizon,
airspeed, altitude, vertical speed, trend, trim, rate of turn
among other key relevant indications. Examples of PFDs
are illustrated in Figure 3-45.
Synthetic Vision
Synthetic vision provides a realistic depiction of the aircraft
in relation to terrain and fl ight path. Systems such as those
produced by Chelton Flight Systems, Universal Flight
Systems, and others provide for depictions of terrain and
course. Figure 3-46 is an example of the Chelton Flight
System providing both 3-dimensional situational awareness
and a synthetic highway in the sky, representing the desired
fl ight path. Synthetic vision is used as a PFD, but provides
guidance in a more normal, outside reference format.
3-28
Figure 3-45. Two Primary Flight Displays (Avidyne on the Left and Garmin on the Right).
Figure 3-46. The benefi ts of realistic visualization imagery, as
illustrated by Synthetic Vision manufactured by Chelton Flight
Systems. The system provides the pilot a realistic, real-time, threedimensional
depiction of the aircraft and its relation to terrain
around it.
Multi-Function Display (MFD)
In addition to a PFD directly in front of the pilot, an MFD
that provides the display of information in addition to primary
fl ight information is used within the fl ight deck.
Information such as a moving map, approach charts, Terrain
Awareness Warning System, and weather depiction can all
be illustrated on the MFD. For additional redundancy both
the PFD and MFD can display all critical information that
the other normally presents thereby providing redundancy
(using a reversionary mode) not normally found in general
aviation fl ight decks.
Advanced Technology Systems
Automatic Dependent Surveillance—Broadcast
(ADS-B)
Although standards for Automatic Dependent Surveillance
(Broadcast) (ADS-B) are still under continuing development,
the concept is simple: aircraft broadcast a message on
a regular basis, which includes their position (such as
latitude, longitude and altitude), velocity, and possibly
other information. Other aircraft or systems can receive this
information for use in a wide variety of applications. The
key to ADS-B is GPS, which provides three-dimensional
position of the aircraft.
As an simplifi ed example, consider air-traffi c radar. The radar
measures the range and bearing of an aircraft. The bearing is
measured by the position of the rotating radar antenna when it
receives a reply to its interrogation from the aircraft, and the
range by the time it takes for the radar to receive the reply.
An ADS-B based system, on the other hand, would listen
for position reports broadcast by the aircraft.
These position reports are based on satellite navigation
systems. These transmissions include the transmitting
aircraft’s position, which the receiving aircraft processes into
usable pilot information. The accuracy of the system is now
determined by the accuracy of the navigation system, not
measurement errors. Furthermore the accuracy is unaffected
by the range to the aircraft as in the case of radar. With radar,
detecting aircraft speed changes require tracking the data and
changes can only be detected over a period of several position
updates. With ADS-B, speed changes are broadcast almost
instantaneously and received by properly equipped aircraft.
3-29
Figure 3-47. Example of a Multi-Function Display (MFD).
Figure 3-48. Aircraft equipped with Automatic Dependent Surveillance—Broadcast (ADS-B) continuously broadcast their identifi cation,
altitude, direction, and vertical trend. The transmitted signal carries signifi cant information for other aircraft and ground stations alike.
Other ADS-equipped aircraft receive this information and process it in a variety of ways. It is possible that in a saturated environment
(assuming all aircraft are ADS equipped), the systems can project tracks for their respective aircraft and retransmit to other aircraft
their projected tracks, thereby enhancing collision avoidance. At one time, there was an Automatic Dependent Surveillance—Addressed
(ADS-A) and that is explained in the Pilot’s Handbook of Aeronautical Knowledge.
3-30
Figure 3-49. An aircraft equipped with ADS will receive identifi cation, altitude in hundreds of feet (above or below using + or -), direction
of the traffi c, and aircraft descent or climb using an up or down arrow. The yellow target is an illustration of how a non-ADS equipped
aircraft would appear on an ADS-equipped aircraft’s display.
Figure 3-50. An aircraft equipped with ADS has the ability to upload and display weather.
Additionally, other information can be obtained by properly
equipped aircraft to include notices to airmen (NOTAM),
weather, etc. At the present time,
ADS-B is predominantly available along the east coast of
the United States where it is matured.
Safety Systems
Radio Altimeters
A radio altimeter, commonly referred to as a radar altimeter,
is a system used for accurately measuring and displaying the
height above the terrain directly beneath the aircraft. It sends
a signal to the ground and processes the timed information.
3-31
Figure 3-51. Components of a Radar Altimeter.
Figure 3-52. Coverage Provided by a Traffi c Information System.
Its primary application is to provide accurate absolute altitude
information to the pilot during approach and landing. In
advanced aircraft today, the radar altimeter also provides its
information to other onboard systems such as the autopilot
and fl ight directors while they are in the glide slope capture
mode below 200-300 feet above ground level (AGL).
A typical system consists of a receiver-transmitter (RT)
unit, antenna(s) for receiving and transmitting the signal,
and an indicator. Category II and III precision
approach procedures require the use of a radar altimeter and
specify the exact minimum height above the terrain as a
decision height (DH) or radio altitude (RA).
Traffi c Advisory Systems
Traffi c Information System
The Traffi c Information Service (TIS) is a ground-based
service providing information to the fl ight deck via data
link using the S-mode transponder and altitude encoder. TIS
improves the safety and effi ciency of “see and avoid” fl ight
through an automatic display that informs the pilot of nearby
traffi c. The display can show location, direction, altitude
and the climb/descent trend of other transponder-equipped
aircraft. TIS provides estimated position, altitude, altitude
trend, and ground track information for up to several aircraft
simultaneously within about 7 NM horizontally, 3,500 feet
above and 3,500 feet below the aircraft. This
data can be displayed on a variety of MFDs.
Figure 3-54 displays the pictorial concept of the traffi c
information system. Noteworthy is the requirement to have
Mode S and that the ground air traffi c station processes the
Mode S signal.
Traffi c Alert Systems
Traffi c alert systems receive transponder information from
nearby aircraft to help determine their relative position to the
equipped aircraft. They provide three-dimensional location
of other aircraft and are cost
effective alternatives to TCAS equipage for smaller aircraft.
Traffi c Avoidance Systems
Traffi c Alert and Collision Avoidance System (TCAS)
The TCAS is an airborne system developed by the FAA that
operates independently from the ground-based ATC system.
TCAS was designed to increase fl ight deck awareness of
proximate aircraft and to serve as a “last line of defense” for
the prevention of mid-air collisions.
There are two levels of TCAS systems. TCAS I was developed
to accommodate the general aviation (GA) community and
the regional airlines. This system issues traffi c advisories
(TAs) to assist pilots in visual acquisition of intruder aircraft.
TCAS I provides approximate bearing and relative altitude
of aircraft with a selectable range. It provides the pilot
with traffi c advisory (TA) alerting him or her to potentially
confl icting traffi c. The pilot then visually acquires the traffi c
and takes appropriate action for collision avoidance.
TCAS II is a more sophisticated system which provides the
same information of TCAS I. It also analyzes the projected
fl ight path of approaching aircraft and issues resolution
advisories (RAs) to the pilot to resolve potential mid-air
collisions. Additionally, if communicating with another
TCAS II equipped aircraft, the two systems coordinate the
resolution alerts provided to their respective fl ight crews.
3-32
Figure 3-53. Multi-Function Display (MFD).
Figure 3-54. Concept of the Traffi c Information System.
3-33
Figure 3-55. Theory of a Typical Alert System.
Figure 3-56. A Skywatch System.
Figure 3-57. Alert System by Avidyne (Ryan).
3-34
Figure 3-58. An example of a resolution advisory being provided
the pilot. In this case, the pilot is requested to climb, with 1,200
feet being the appropriate rate of ascent to avoid traffi c confl ict.
This visual indication plus the aural warning provide the pilot
with excellent traffi c awareness that augments see and avoid
practices.
Terrain Alerting Systems
Ground Proximity Warning System (GPWS)
An early application of technology to reduce CFIT was the
GPWS. In airline use since the early 1970s, GPWS uses the
radio altimeter, speed, and barometric altitude to determine the
aircraft’s position relative to the ground. The system uses this
information in determining aircraft clearance above the Earth
and provides limited predictability about aircraft position
relative to rising terrain. It does this based upon algorithms
within the system and developed by the manufacturer for
different airplanes or helicopters. However, in mountainous
areas the system is unable to provide predictive information
due to the unusual slope encountered.
This inability to provide predictive information was evidenced
in 1999 when a DH-7 crashed in South America. The crew
had a GPWS onboard, but the sudden rise of the terrain
rendered it ineffective; the crew continued unintentionally
into a mountain with steep terrain. Another incident involved
Secretary of Commerce Brown who, along with all on board,
was lost when the crew fl ew over rapidly rising terrain where
the GPWS capability is offset by terrain gradient. However,
the GPWS is tied into and considers landing gear status, fl ap
position, and ILS glide slope deviation to detect unsafe aircraft
operation with respect to terrain, excessive descent rate,
excessive closure rate to terrain, unsafe terrain clearance while
not in a landing confi guration, excessive deviation below an
ILS glide slope. It also provides advisory callouts.
Generally, the GPWS is tied into the hot bus bar of the electrical
system to prevent inadvertent switch off. This was demonstrated
in an accident involving a large four-engine turboprop airplane.
While on fi nal for landing with the landing gear inadvertently
up, the crew failed to heed the GPWS warning as the aircraft
crossed a large berm close to the threshold. In fact, the crew
attempted without success to shut the system down and attributed
the signal to a malfunction. Only after the mishap did the crew
realize the importance of the GPWS warning.
Terrain Awareness and Warning System (TAWS)
A TAWS uses GPS positioning and a database of terrain and
obstructions to provide true predictability of the upcoming
terrain and obstacles. The warnings it provides pilots are
both aural and visual, instructing the pilot to take specifi c
action. Because TAWS relies on GPS and a database of
terrain/obstacle information, predictability is based upon
aircraft location and projected location. The system is time
based and therefore compensates for the performance of the
aircraft and its speed.
Head-Up Display (HUD)
The HUD is a display system that provides a projection of
navigation and air data (airspeed in relation to approach
reference speed, altitude, left/right and up/down glide slope)
on a transparent screen between the pilot and the windshield.
The concept of a HUD is to diminish the shift between
looking at the instrument panel and outside. Virtually any
information desired can be displayed on the HUD if it is
available in the aircraft’s fl ight computer. The display for
the HUD can be projected on a separate panel near the
windscreen or as shown in Figure 3-60 on an eye piece. Other
information may be displayed, including a runway target in
relation to the nose of the aircraft, which allows the pilot to
see the information necessary to make the approach while
also being able to see out the windshield.
Required Navigation Instrument System
Inspection
Systems Prefl ight Procedures
Inspecting the instrument system requires a relatively small
part of the total time required for prefl ight activities, but its
importance cannot be overemphasized. Before any fl ight
involving aircraft control by instrument reference, the pilot
should check all instruments and their sources of power
for proper operation. NOTE: The following procedures are
appropriate for conventional aircraft instrument systems.
Aircraft equipped with electronic instrument systems utilize
different procedures.
3-35
Figure 3-59. A six-frame sequence illustrating the manner in which TAWS operates. A TAWS installation is aircraft specifi c and provides
warnings and cautions based upon time to potential impact with terrain rather than distance. The TAWS is illustrated in an upper left
window while aircrew view is provided out of the windscreen. illustrates the aircraft in relation to the outside terrain while and
illustrate the manner in which the TAWS system displays the terrain. is providing a caution of terrain to be traversed, while
provides an illustration of a warning with an aural and textural advisory (red) to pull up. also illustrates a pilot taking appropriate
action (climb in this case) while illustrates that a hazard is no longer a factor.
3-36
Figure 3-60. A Head-Up Display.
Before Engine Start
1. Walk-around inspection: Check the condition of all
antennas and check the pitot tube for the presence
of any obstructions and remove the cover. Check
the static ports to be sure they are free from dirt
and obstructions, and ensure there is nothing on the
structure near the ports that would disturb the air
fl owing over them.
2. Aircraft records: Confi rm that the altimeter and static
system have been checked and found within approved
limits within the past 24 calendar months. Check the
replacement date for the emergency locator transmitter
(ELT) batteries noted in the maintenance record, and
be sure they have been replaced within this time
interval.
3. Preflight paperwork: Check the Airport/Facility
Directory (A/FD) and all Notices to Airmen
(NOTAMs) for the condition and frequencies of all the
navigation aid (NAVAIDs) that are used on the fl ight.
Handbooks, en route charts, approach charts, computer
and fl ight log should be appropriate for the departure,
en route, destination, and alternate airports.
4. Radio equipment: Switches off.
5. Suction gauge: Proper markings as applicable if
electronic fl ight instrumentation is installed.
6. ASI: Proper reading, as applicable. If electronic
fl ight instrumentation is installed, check emergency
instrument.
7. Attitude indicator: Uncaged, if applicable. If electronic
fl ight instrumentation is installed, check emergency
system to include its battery as appropriate.
8. Altimeter: Set the current altimeter setting and ensure
that the pointers indicate the elevation of the airport.
9. VSI: Zero indication, as applicable (if electronic fl ight
instrumentation is installed).
10. Heading indicator: Uncaged, if applicable.
11. Turn coordinator: If applicable, miniature aircraft
level, ball approximately centered (level terrain).
12. Magnetic compass: Full of fl uid and the correction
card is in place and current.
13. Clock: Set to the correct time and running.
14. Engine instruments: Proper markings and readings,
as applicable if electronic fl ight instrumentation is
installed.
15. Deicing and anti-icing equipment: Check availability
and fl uid quantity.
16. Alternate static-source valve: Be sure it can be opened
if needed, and that it is fully closed.
3-37
17. Pitot tube heater: Check by watching the ammeter
when it is turned on, or by using the method specifi ed
in the POH/AFM.
After Engine Start
1. When the master switch is turned on, listen to the
gyros as they spin up. Any hesitation or unusual noises
should be investigated before fl ight.
2. Suction gauge or electrical indicators: Check the
source of power for the gyro instruments. The suction
developed should be appropriate for the instruments
in that particular aircraft. If the gyros are electrically
driven, check the generators and inverters for proper
operation.
3. Magnetic compass: Check the card for freedom of
movement and confirm the bowl is full of fluid.
Determine compass accuracy by comparing the
indicated heading against a known heading (runway
heading) while the airplane is stopped or taxiing
straight. Remote indicating compasses should also be
checked against known headings. Note the compass
card correction for the takeoff runway heading.
4. Heading indicator: Allow 5 minutes after starting
engines for the gyro to spin up. Before taxiing, or
while taxiing straight, set the heading indicator to
correspond with the magnetic compass heading. A
slaved gyrocompass should be checked for slaving
action and its indications compared with those of the
magnetic compass. If an electronic fl ight instrument
system is installed, consult the fl ight manual for proper
procedures.
5. Attitude indicator: Allow the same time as noted
above for gyros to spin up. If the horizon bar erects
to the horizontal position and remains at the correct
position for the attitude of the airplane, or if it begins
to vibrate after this attitude is reached and then slowly
stops vibrating altogether, the instrument is operating
properly. If an electronic fl ight instrument system
is installed, consult the flight manual for proper
procedures.
6. Altimeter: With the altimeter set to the current reported
altimeter setting, note any variation between the
known fi eld elevation and the altimeter indication. If
the indication is not within 75 feet of fi eld elevation,
the accuracy of the altimeter is questionable and
the problem should be referred to a repair station
for evaluation and possible correction. Because the
elevation of the ramp or hangar area might differ
signifi cantly from fi eld elevation, recheck when in
the run-up area if the error exceeds 75 feet. When
no altimeter setting is available, set the altimeter
to the published fi eld elevation during the prefl ight
instrument check.
7. VSI: The instrument should read zero. If it does not,
tap the panel gently. If an electronic fl ight instrument
system is installed, consult the fl ight manual for proper
procedures.
8. Engine instruments: Check for proper readings.
9. Radio equipment: Check for proper operation and set
as desired.
10. Deicing and anti-icing equipment: Check operation.
Taxiing and Takeoff
1. Turn coordinator: During taxi turns, check the
miniature aircraft for proper turn indications. The ball
or slip/skid should move freely. The ball or slip/skid
indicator should move opposite to the direction of
turns. The turn instrument should indicate the direction
of the turn. While taxiing straight, the miniature
aircraft (as appropriate) should be level.
2. Heading indicator: Before takeoff, recheck the heading
indicator. If the magnetic compass and deviation card
are accurate, the heading indicator should show the
known taxiway or runway direction when the airplane
is aligned with them (within 5°).
3. Attitude indicator: If the horizon bar fails to remain
in the horizontal position during straight taxiing, or
tips in excess of 5° during taxi turns, the instrument is
unreliable. Adjust the miniature aircraft with reference
to the horizon bar for the particular airplane while on
the ground. For some tricycle-gear airplanes, a slightly
nose-low attitude on the ground gives a level fl ight
attitude at normal cruising speed.
Engine Shut Down
When shutting down the engine, note any abnormal
instrument indications.
3-38
4-1
Introduction
Attitude instrument fl ying is defi ned as the control of an
aircraft’s spatial position by using instruments rather than
outside visual references. Today’s aircraft come equipped
with analog and/or digital instruments. Analog instrument
systems are mechanical and operate with numbers
representing directly measurable quantities, such as a watch
with a sweep second hand. In contrast, digital instrument
systems are electronic and operate with numbers expressed
in digits. Although more manufacturers are providing aircraft
with digital instrumentation, analog instruments remain more
prevalent. This section acquaints the pilot with the use of
analog fl ight instruments.
Airplane Attitude
Instrument Flying
Chapter 4, Section I
Using Analog Instrumentation
4-2
Figure 4-1. Control Instruments.
Any fl ight, regardless of the aircraft used or route fl own,
consists of basic maneuvers. In visual fl ight, aircraft attitude
is controlled by using certain reference points on the aircraft
with relation to the natural horizon. In instrument fl ight,
the aircraft attitude is controlled by reference to the fl ight
instruments. Proper interpretation of the fl ight instruments
provides essentially the same information that outside
references do in visual fl ight. Once the role of each instrument
in establishing and maintaining a desired aircraft attitude is
learned, a pilot is better equipped to control the aircraft in
emergency situations involving failure of one or more key
instruments.
Learning Methods
The two basic methods used for learning attitude instrument
fl ying are “control and performance” and “primary and
supporting.” Both methods utilize the same instruments
and responses for attitude control. They differ in their
reliance on the attitude indicator and interpretation of other
instruments.
Attitude Instrument Flying Using the Control and
Performance Method
Aircraft performance is achieved by controlling the aircraft
attitude and power. Aircraft attitude is the relationship
of both the aircraft’s pitch and roll axes in relation to the
Earth’s horizon. An aircraft is fl own in instrument fl ight by
controlling the attitude and power, as necessary, to produce
both controlled and stabilized fl ight without reference to a
visible horizon. This overall process is known as the control
and performance method of attitude instrument flying.
Starting with basic instrument maneuvers, this process can
be applied through the use of control, performance, and
navigation instruments, resulting in a smooth fl ight, from
takeoff to landing.
Control Instruments
The control instruments display immediate attitude and power
indications and are calibrated to permit those respective
adjustments in precise increments. In this discussion, the
term “power” is used in place of the more technically correct
term “thrust or drag relationship.” Control is determined
by reference to the attitude and power indicators. Power
indicators vary with aircraft and may include manifold
pressure, tachometers, fuel fl ow, etc.
Performance Instruments
The performance instruments indicate the aircraft’s actual
performance. Performance is determined by reference to the
altimeter, airspeed or vertical speed indicator (VSI), heading
indicator, and turn-and-slip indicator.
Navigation Instruments
The navigation instruments indicate the position of the aircraft
in relation to a selected navigation facility or fi x. This group
of instruments includes various types of course indicators,
range indicators, glide-slope indicators, and bearing pointers.
Newer aircraft with more technologically
advanced instrumentation provide blended information,
giving the pilot more accurate positional information.
Procedural Steps in Using Control and
Performance
1. Establish an attitude and power setting on the
control instruments that results in the desired
performance. Known or computed attitude changes
and approximated power settings helps to reduce the
pilot’s workload.
2. Trim (fine tune the control forces) until control
pressures are neutralized. Trimming for hands-off
fl ight is essential for smooth, precise aircraft control.
4-3
Figure 4-2. Performance Instruments.
Figure 4-3. Flight Panel Instrumentation.
It allows a pilot to attend to other fl ight deck duties
with minimum deviation from the desired attitude.
3. Cross-check the performance instruments to determine
if the established attitude or power setting is providing
the desired performance. The cross-check involves
both seeing and interpreting. If a deviation is noted,
determine the magnitude and direction of adjustment
required to achieve the desired performance.
4. Adjust the attitude and/or power setting on the control
instruments as necessary.
Aircraft Control During Instrument Flight
Attitude Control
Proper control of aircraft attitude is the result of proper use
of the attitude indicator, knowledge of when to change the
4-4
Figure 4-4. Pitch Instruments.
attitude, and then smoothly changing the attitude a precise
amount. The attitude reference provides an immediate, direct,
and corresponding indication of any change in aircraft pitch
or bank attitude.
Pitch Control
Changing the “pitch attitude” of the miniature aircraft or
fuselage dot by precise amounts in relation to the horizon
makes pitch changes. These changes are measured in degrees
or fractions thereof, or bar widths depending upon the type of
attitude reference. The amount of deviation from the desired
performance determines the magnitude of the correction.
Bank Control
Bank changes are made by changing the “bank attitude” or
bank pointers by precise amounts in relation to the bank scale.
The bank scale is normally graduated at 0°, 10°, 20°, 30°,
60°, and 90° and is located at the top or bottom of the attitude
reference. Normally, use a bank angle that approximates the
degrees to turn, not to exceed 30°.
