Par 12-10 Page 12-604 proper bearing to station, audio identification and tone/beat frequency oscillator (BFO), correct operation in closed circuit (LOOP) and sense modes. Note the orientation of the selected station to the aircraft using an appropriate chart. Observ e the ADF relative bearing reading, and compare to the ch art. Slew the needle and observe how fast (or slowly) it returns to the reading.
ADF performance may be degraded by lightning activity, airframe chargi ng, ignition noise and atmospheric phenomena.
INSTRUMENT LANDING SYSTEMS (ILS). The ILS consist of several components, such as the localizer, glide slope, marker beacon, radio altimeter, and DME. Localizer and glide slope receivers and marker beacons will be discussed in this section.
Localizer receiver operates on one of 40 ILS channels within the frequency range of 108.10 to 111.95 MHz.
These signals provide course guidance to the pilot to the runway centerline through th e lateral displacement of the VOR/localizer (LOC) de viation indicator. The ground transmitter is sighted at the far end of the runway and provides a valid signal from a distance of 18 NM from the transmitter. The indication gives a full fly left/right deviation of 700 feet at the runway threshold. Identification of the transmitter is in International Morse Code and consists of a three letter identifier preceded by the Morse Code letter I (two dots). The localizer function is usually integral with the VOR system, and when maintenance is performed on the VOR unit, the localizer is also included. The accuracy of the system can be effectively evaluated through normal flight operations if evaluated during visual meteorological conditions. Any determination of airworthiness after reinstallation before instrument flight must be accomplished with ground test equipment.
The glide slope receiver operates on one of 40 channels within the frequency range 329.15 MHz, to
335.00 MHz. The glide slope transmitter is located between 750 feet and 1250 feet from the approach end of the runway and offset 250 to 650 feet. In the absence of questionable performance, periodic functional flight checks of the glide slope system would be an acceptable way to ensure continued system performance. The functional flight test must be conducted under visual flight rules (VFR) conditions. A failed or misleading system must be serviced by an appropriately-rated repair station. Ground test equipment can be used to verify glide slope operation.
Localizer/Glide Slope (LOC/GS) may have self test function, otherwise the proper ground test equipment must be used. Refer to manufacturer’s or aircraft instruction manual.
MARKER BEACON. Marker beacon receivers operate at 75 MHz and sense the audio signature of each of the three types of beacons. The marker beacon receiver is not tunable. The blue outer marker light illuminates when the receiver acquires a 75 MHz signal modulated with 400 Hz, an amber middle marker light for a 75 MHz signal modulated with 1300 Hz and, a white inner marker light for a 75 MHz signal modulated with 3000 Hz. The marker beacon system must be operationally evaluated in VFR when an ILS runway is available. The receiver sensitivity switch must be placed in LOW SENSE (the normal setting). Marker audio must be adequate. Ground test equipment must be used to verify marker beacon operation.
Marker beacon with self test feature, verify lamps, audio and lamp dimming.
LONG RANGE NAVIGATION (LORAN). The LORAN has been an effective alternative to Rho/Theta R-Nav systems.
Par 12-16 Page 12-605 Hyperbolic systems require waypoint designation in terms of latitude and longitude, unlike original R-Nav (distance navigation) systems, which define waypoints in terms of distance (Rho) and angle (Theta) from established VOR or Tacan fac ilities. Accuracy is better than the VOR/Tacan system but LORAN is more prone to problems with precipitation static. Proper bonding of aircraft structure and the use of high-quality static wicks will not only produce improved LORAN system performance, but can also benefit the very high frequency (VHF) navigati on and communications systems. This system has an automatic test equipment (ATE).
NOTE: Aircraft must be outside of hangar for LORAN to operate.
Normally self test check units, verification of position, and loading of flight plan will verify operation verification of proper flight manual supplements and operating handbooks on board, and proper software status can also be verified.
GLOBAL POSITIONING SYSTEM (GPS). The GPS is at the forefront of present generation navigation systems. This space-based navigation system is based on a 24-satellite system and is highly accurate (within 100 meters) for establishing position. The system is unaffected by weather and provides a world-wide common grid reference system.
Database updating and antenna maintenance are of primary concern to the GPS user.
NOTE: Aircraft must be outside of hangar for ground test of GPS.
AUTOPILOT SYSTEMS. Automatic Flight Control Systems (AFCS) are the most efficient managers of aircraft performance and control.
There are three kinds of autopilot; two axes, three axes, and three axes with coupled approach capability. Attention must be given to the disconnect switch operation, aural and visual alerts of automatic and intentional autopilot disconnects, override forces and mode annunciation, servo operation, rigging and bridle cable tension, and condition. In all cases the manufacturer’s inspection and maintenance instructions must be followed.
ALTIMETERS. Aircraft conducting operations in controlled airspace under instrument flight rule (IFR) are required to have their static system(s) and each altimeter instrument inspected and tested within the previous 24 calendar months. Frequent functional checks of all altimeters and automatic pressure altitude reporting systems are recommended.
Examine the altimeter face for evidence of needle scrapes or other damage. Check smoothness of operation, with particular attention to altimeter performance during decent.
Contact an appropriate air traffic facility for the pressure altitude displayed to the controller from your aircraft. Correct the reported altitude as needed, and compare to the reading on the altimeter instrument. The difference must not exceed 125 feet.
TRANSPONDERS. There are three modes (types) of transponders that can be used on various aircraft. Mode A provides a (non altitude-reporting) four-digit coded reply;
Mode C provides a code reply identical to Mode A with an altitude-reporting signal; and Mode S has the same capabilities as Mode A and Mode C and responds to traffic alert and collision avoidance system (TCAS)-Equipped Aircraft.
Ground ramp equipment must be used to demonstrate proper operation. Enough codes must be selected so that each switch
Par 12-16 Page 12-606 position is checked at least once. Low and high sensitivity operation must be checked. Identification operation must be checked. Altitude reporting mode must be demonstrated. Demonstrate that the transponder system does not interfere with other systems aboard the aircraft, and that other equipment does not interfere with transponder operation. Special consideration must be given to other pulse equipment, such as DME and weather radar.
All transponders must be tested every 24-calendar months, or during an annual inspection, if requested by the owner. The test must be conducted by an authorized avionics repair facility.
EMERGENCY LOCATOR TRANSMITTERS (ELT). The ELT must be evaluated in accordance with TSO-C91a, TSO-C126 for 406 MHz ELT’s, or later TSO’s issued for ELT’s.
ELT installations must be examined for potential operational problems at least once a year (section 91.207(d)). There have been numerous instances of interaction between ELT and other VHF installations. Antenna location should be as far as possible from other antennas to prevent efficiency losses.
Check ELT antenna installations in close proximity to other VHF antennas for suspected interference. Antenna patterns of previously installed VHF antennas could be measured after an ELT installation. Testing of an ELT must be performed within the first 5 minutes of an hour, and only three pulses of the transmitter should be activated. For example, a test could be conducted between 1:00 p.m.
and 1:05 p.m., with a maximum of three beeps being heard on a frequency of 121.5 MHz.
INSPECTION OF ELT. An inspection of the following must be accomplished by a properly certified person or repair station within 12- calendar months after the last inspection:
Proper Installation.
Remove all interconnections to the ELT unit and ELT antenna. Visually inspect and confirm proper seating of all connector pins.
Special attention should be given to coaxial center conductor pins which are prone to retracting into the connector housing.
Remove the ELT from the mount and inspect the mounting hardware for proper installation and security.
Reinstall the ELT into its mount and verify the proper direction for crash activation.
Reconnect all cables. They should have some slack at each end and should be properly secured to the airplane structure for support and protection.
Battery Corrosion. Gain access to the ELT battery and inspect. No corrosion should be detectable. Verify the ELT battery is approved and check its expiration date.
Operation of the Controls and Crash Sensor.
Activate the ELT using an applied force. Consult the ELT manufacturer’s instructions before activation. The direction for mounting and force activation is indicated on the ELT. A TSO-C91 ELT can be activated by using a quick rap with the palm. A TSO-C91a ELT can be activated by using a rapid forward (throwing) motion coupled by a rapid reversing action. Verify that the ELT can be activated using a watt meter, the airplane’s VHF radio communications receiver tuned to
121.5 MHz, or other means (see NOTE 1). Insure that the “G” switch has been reset if applicable.
For a Sufficient Signal Radiated From its Antenna.
Activate the ELT using the ON or ELT TEST switch.
A low-quality
Par 12-22 Page 12-607 AM broadcast radio recei ver should be used to determine if energy is being transmitted from the antenna. When the antenna of the AM broadcast radio receiver (tuning dial on any setting) is held about 6 inches from the activated ELT antenna, the ELT aural tone will be heard (see NOTE 2 and 3).
Verify That All Switches are Properly Labeled and Positioned.
Record the Inspection. Record the inspection in the aircraft maintenance records according to 14 CFR part 43, section 43.9. We suggest the following:
I inspected the Make/Model ______________ ELT system in this aircraft accord ing to applicable Aircraft and ELT manufacturer’s instru ctions and applicable FAA guidance and found that it meets the requirements of section 91.207(d).
Signed:
______________________ Certificate No. ______________________ Date:
______________________
NOTE 1: This is not a measured check; it only indicates that the G-switch is working.
NOTE 2: This is not a measured check; but it does provide confidence that the antenna is radiating with sufficient power to aid search and rescue. The signal may be weak even if it is picked up by an aircraft VHF receiver located at a considerable distance from the radiating ELT. Therefore, this check does not check the integrity of the ELT system or provide the same level of confidence as does the AM radio check.
NOTE 3: Because the ELT radiates on the emergency frequency, the Federal Communications Commission allows these tests only to be conducted within the first five minutes after any hour and is limited in three sweeps of the transmitter audio modulation.
FLIGHT DATA RECORDER. The flight data recorder is housed in a crush-proof container located near the tail section of the aircraft. The tape unit is fire resistant, and contains a radio transmitter to help crash investigators locate the unit under water. Inspection/Operational checks include:
Check special sticker on front of the flight data recorder for the date of th e next tape replacement, if applicable.
Remove recorder magazine and inspect tape for the following:
broken or torn tape,
proper feed of tape, and
all scribes were recording properly for approximately the last hour of flight.
Conditions for tape replacement (as applicable):
There is less than 20 hours remaining in the magazine as read on the tape remaining indicator.
Tape has run out.
Broken tape.
After hard landings and severe air turbulence have been encountered as reported by the pilots.
Par 12-22 Page 12-608 After the same tape has been in use 1 year (12 months), it must be replaced.
Ensure that a correlation test has been performed and then recorded in the aircraft records.
Refer to the specific equipment manufacturer’s manuals and procedures.
The state-of-the art Solid-State Flight Data Recorder (SSFDR) is a highly flexible model able to support a wide variety of aeronautical radio, incorporated (ARINC) configurations. It has a Built-In Test Equipment (BITE) that establishes and monitors the mission fitness of the hardware. BITE performs verification after storage (read after write) of flight data and status condition of the memory. These recorders have an underwater acoustic beacon mounted on its front panel which must be returned to their respective manufacturer’s for battery servicing. For maintenance information refer to the equipment or aircraft manufacture’s maintenance instruction manual.
COCKPIT VOICE RECORDERS (CVR). CVR’s are very similar to flight data recorders. They look nearly identical and operate in almost the same way. CVR’s monitors the last 30 minutes of flight deck conversations and radio communications. The flight deck conversations are recorded via the microphone monitor panel located on the flight deck. This panel is also used to test the system and erase the tape, if so desired. Before operating the erase CVR mode, consult the operational manual of the manufacturer for the CVR.
Playback is possible only after the recorder is removed from the aircraft.
Refer to the specific equipment manufacturer’s manuals and procedures.
The Solid State Cockpi t Voice Recorder system is composed of three essential components a solid state recorder, a control unit (remo te mic amplifier), and an area microphone. Also installed on one end of the recorder is an Under water Locator Beacon (ULB). The recorder accepts four separate audio inputs; pilot, copilot, public address/third crew member, and cockpit area microphone and where applicable, rotor speed input and flight data recorder synchronization tone input. For maintenance information refer to the equipment manufacturer’s maintenance manual.
WEATHER RADAR. Ground performance shall include antenna rotation, tilt, indicator brilliance, scan rotation, and indication of received echoes. It must be determined that no objectionable interference from other electrical/electronic equipment appears on the radar indicator, and that the radar system does not interfere with the operation of any of the aircraft’s communications or navigation systems.
CAUTION: Do not turn radar on within 15 feet of ground personnel, or containers holding flammable or explosive materials.
The radar should never operate during fueling operations. Do not operate radar system when beam may intercept larger metallic objects closer than 150 feet, as crystal damage might occur. Do not operate radar when cooling fans are inoperative. Refer to the specific Radar System equipment manufacturer’s manuals and procedures.
RADOME INSPECTION. Inspection of aircraft having weather radar installations should include a visual check of the radome surface for signs of surface damage,
Par 12-26 Page 12-609 (and 12-12) holes, cracks, chipping, and peeling of paint, etc.
Attach fittings and fastenings, neoprene erosion caps, and lightening strips, when installed, should also be inspected.
DATA BUS. Data Buses provide the physical and functional partitioning needed to enable different companies to design different avionics boxes to be able to communicate information to each other. It defines the framework for system(s) intergration. There are several types of data bus analyzers used to receive and review transmitted data or to transmit data to a bus user. Before using an analyzer, make sure that the bus language is compatible with the bus analyzer. For further information refer to ARINC specifications such as 429 Digital Information Transfer System, Mark 33 which offers simple and affordable answers t data communications on aircraft.
12-28.12-36. [RESERVED.]
Par 12-37 Page 12-610
SECTION 3. GROUND OPERATIONAL CHECKS FOR AVIONICS
EQUIPMENT (NON ELECTRICAL)
COMPASS SWING must be performed whenever a new compass is installed. The magnetic compass can be checked for accuracy by using a compass rose located on an airport and by using a hand held master compass. The check swing is normally effected by placing the aircraft on various magnetic headings and comparing the deviations with those on the deviation cards.
Refer to equipment or aircraft manufacture’s manual.
A compass swing must be performed on the following occasions:
When the accuracy of the compass is suspected.
After any cockpit modification or major replacement involving ferrous metal.
Whenever a compass has been subjected to a shock; for example, after a hard landing or turbulence.
After aircraft has passed through a severe electrical storm.
After lighting strike.
Whenever a change is made to the electrical system.
Whenever a change of cargo is likely to affect the compass.
When an aircraft operation is changed to a different geographic location (e.g., Miami, Florida to Fairbanks, Alaska) with a major change in magnetic deviation.
After aircraft has been parked on one heading for over a year.
Par 12-37 Page 12-611 When flux valves are replaced.
Compass Swing Procedures. The magnetic compass must be checked for accuracy in a location free of steel structures, underground pipes or cables, or equipment that produces magnetic fields.
The master compass is a reverse reading compass with a gun-sight arrangement mounted on top of it. With the aircraft facing North and the person in the cockpit running the engine(s) at 1000 rpm, a mechanic standing approximately 30 feet in front of the aircraft, facing South, “shoots” or aligns the master compass with the aircraft center line.
Using hand signals, the mechanic signals the person in the cockpit to make additional adjustments to align the aircraft with the master compass. Once aligned on the heading, the person in the cockpit runs the engine(s) to approximately 1,700 rpm to duplicate the aircraft’s magnetic field and then the person reads the compass.
NOTE: For conventional gear aircraft, the mechanic will have to position the magnetic compass in the straight and level position or mount the tail of the aircraft on a moveable dolly to simulate a straight and level cruise configuration.
If the aircraft compass is not in alignment with the magnetic North with the master compass, then the mechanic can correct the error by making small adjustments to the North-South brass adjustment screw with a nonmetallic screw driver. This screw driver can be made out of brass stock, or stainless steel welding rod. The aircraft should be positioned facing South and aligned with the
Par 12-37 Page 12-612 master compass. Using the same procedures, correct any error in the compass r eading using the check for errors on the East/West heading using the same procedures for the Nort h-South check, except the corrections should be made using the East-West correction brass screw.
Check the compass reading on all cardinal headings. Record the last reading and prepare a compass correction card. The maximum deviation (plus or minus) is 10 degrees on any one heading.
If the compass cannot be adjusted to meet the requirements, install another one.
NOTE: A common error that affects the compass’s accuracy is the mounting of a compass on or in the instrument panel using steel machine screws/nuts rather than brass hardware.
If the aircraft has an electrical system it is recommended that two complete compass checks be performed, one with minimum electrical equipment operating and the other with all electrical accessories on (e.g. radios, navigation radar, and lights). If the compass readings are not identical, then the mechanic should make up two separate compass correction cards. One with all the equipment on and one with the equipment off.
PNEUMATIC GYROS.
Venturi Systems. The early gyro instruments were all operated by air flowing out of a jet over buckets cut into the periphery of the gyro rotor. A venturi was mounted on the outside of the aircraft to produce a low pressure, or vacuum, which evacuated the instrument case, and air flowed into the instrument through a paper filter and then through a nozzle onto the rotor.
Venturi systems have the advantage of being extremely simple and requiring no power from the engine, nor from any of the other aircraft systems; but they do have the disadvantage of being susceptible to ice, and when they are most needed, they may become unusable.
There are two sizes of venturi tubes: those which produce four inches of suction are used to drive the attitude gyros, and smaller tubes, which produce two inches of suction, are used for the turn and slip indicator. Some installations use two of the larger venturi tubes connected in parallel to the two attitude gyros, and the turn and slip indicator is connected to one of these instruments with a needle valve between them. A suction gage is temporarily connected to the turn and slip indicator, and the aircraft is flown so the needle valve can be adjusted to the required suction at the instrument when the aircraft is operated at its cruise speed. (See figure 12-1.)
Vacuum Pump Systems. In order to overcome the major drawback of the venturi tube, that is, its susceptibility to ice, aircraft were equipped with engine driven vacuum pumps and the gyro instruments were driven by air pulled through the instrument by the suction produced by these pumps. A suction relief valve maintained the desired pressure (usually about four inches of mercury) on the attitude gyro instruments, and a needle valve between one of the attitude indicators and the turn and slip indicator restricted the airflow to maintain the desired 2 inches of suction in its case. Most of the early instruments used only paper filters in each of the instrument cases, but in some installations a central air filter was used to remove contaminants from the cabin air before it entered the instrument case.
The early vacuum pumps were vane-type pumps of what is called the wet
Par 12-38 Page 12-613
FIGURE 12-1. Venturi system for providing airflow through gyro instruments.
type-one with a cast iron housing and steel vanes.
Engine oil was metered into the pump to provide sealing, lubrication, and cooling, and then this oil, along with the air, was blown through an oil separator where the oil collected on baffles and was returned to the engine crankcase. The air was then exhausted overboard. Aircraft equipped with rubber deicer boots used this discharge air to inflate the boots. But before it could be used, this air was passed through a second stage of oil separation and then to the distributor valve and finally to the boots. (See figure 12-2.)
The airflow through the instruments is controlled by maintaining the suction in the instrument case at the desired level with a suction relief valve mounted between the pump and the instruments. This valve has a spring-loaded poppet that offsets to allow cabin air to enter the pump and maintain the correct negative pressure inside the instrument case.
The more modern vacuum pumps are of the dry type. These pumps use carbon vanes and do not require any lubrication, as the vanes provide their own lubrication as they wear away at a carefully predetermined rate. Other than the fact that they do not require an oil separator, systems using dry air pumps are quite similar to those using a wet pump. One slight difference, however, is in the need for keeping the inside of the pump perfectly clean. Any solid particles drawn into the system through the suction relief valve can damage one of the carbon vanes, and this can lead to destruction of the pump, as the particles
Par 12-38 Page 12-614
FIGURE 12-2. Instrument vacuum system using a wet-type vacuum pump.
broken off of one vane will damage all of the other vanes. To prevent particles entering the relief valve, its air inlet is covered with a filter, and this must be cleaned or replaced at the interval recommended by the aircraft manufacturer.
Positive Pressure Systems. Above about 18,000 feet there is not enough mass to the air drawn through the instruments to provide sufficient rotor speed, and, to remedy this problem, many aircraft that fly at high altitude use positive pressure systems to drive the gyros.
These systems use the same type of air pump as is used for vacuum systems, but th e discharged air from the pump is filtered and directed into the instrument case through the same fitting that receives the filtered air when the vacuum system is used. A filter is installed on the inlet of the pump, and then, before the air is directed into the instrument case, it is again filtered. A pressure regulator is located between the pump and the in-line filter to control the air pressure so only the correct amount is directed into the instrument case.
System Filters. The life of an air-driven gyro instrument is determined to a great extent by the cleanliness of the air that flows over the rotor. In vacuum systems, this air is taken from the cabin where there is usually a good deal of dust and very often tobacco
Par 12-38 Page 12-615 (and 12-18) smoke. Unless all of the solid contaminants are removed from the air before it enters the instrument, they will accumulate, usually in the rotor bearings, and slow the rotor. This cause s an inaccurate indication of the instrument and will definitely shorten its service life.
Dry air pumps are also subject to damage from ingested contaminants, and all of the filters in the system must be replaced on the schedule recommended by the aircraft manufacturer, and more often if the aircraft is operated under pa rticularly dusty conditions, especially if the occupants of the aircraft regularly smoke while flying. (See figures 12-3 and 12-4.)
FIGURE 12-3. Instrument vacuum system using a dry-type air pump.
FIGURE 12-4. Instrument pressure system using a dry-type air pump.
12-39.12-50. [RESERVED.]