Instrument Flying Handbook

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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. [Figure 3-1] 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. [Figure 3-2] 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. [Figure 3-3] 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. [Figure 3-4] Another confi guration of the altimeter is the drum-type. [Figure 3-5] 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

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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.” [Figure 3-8] 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).

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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. [Figure 3-12] 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. [Figure 3-13] 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.

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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. [Figure 3-20] 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 [Figure 3-21], 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. [Figure 3-22] 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. [Figure 3-24] 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. [Figure 3-25] 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.

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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°. [Figure 3-30]

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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. [Figure 3-32] 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. [Figure 3-33] 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. [Figure 3-34] 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. [Figure 3-34] 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. [Figure 3-35] 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. [Figure 3-36] 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.

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Air Data Computer (ADC) An Air Data Computer (ADC) [Figure 3-37] 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 [Figure 3-39] 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).

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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). [Figure 3-44] 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.

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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. [Figure 3-47] 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. [Figure 3-48] 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. [Figures 3-49 and 3-50] 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. [Figure 3-51] 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. [Figure 3-52] This data can be displayed on a variety of MFDs. [Figure 3-53] 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 [Figures 3-55, 3-56, and 3-57] 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. [Figure 3-58]

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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. [Figure 3-59] 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. [Figure 4-1] 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. [Figure 4-2] 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. [Figure 4-3] 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°.