Par 12-18 Page 12-7 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.
12-19. ALTIMETERS. Aircraft conducting operations in controlled airspace under 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.
a. The tests required must be performed by:
(1) The manufacturer of the aircraft on which the tests and inspection are to be performed.
(2) A certificated repair station (CRS) properly equipped to perform those functions and holding:
(a) An instrument rating Class I.
(b) A limited instrument rating appropriate to the make and model (M/M) of appliance to be tested.
(c) A certified/qualified mechanic with an airframe rating (static system tests and inspections only). Any adjustments shall be accomplished only by an instrument shop certified/qualified person using proper test equipment and adequate reference to the manufacturer’s maintenance manuals. The altimeter correlation adjustment shall not be adjusted in the field. Changing this adjustment will nullify the correspondence between the basic test equipment calibration standards and the altimeter. It will also nullify correspondence between the encoding altimeter and its encoding digitizer or the associated blind encoder.
b. Examine the altimeter face for evidence of needle scrapes or other damage. Check smoothness of operation, with particular attention to altimeter performance during decent.
c. 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 ft.
12-20. 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.
a. Ground ramp equipment must be used to demonstrate proper operation. Enough codes must be selected so that each switch 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
Par 12-20 Page 12-8 transponder operation. Special consideration must be given to other pulse equipment, such as DME and weather radar.
b. 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.
12-21. AUTOMATIC DEPENDENT SURVEILLANCE–BROADCAST (ADS-B). ADS-B is a system that uses positioning data, such as from GPS, downlinked from aircraft avionics for use in air traffic control (ATC) and by other ADS-B In-equipped aircraft. ADS-B avionics operate on the 1090 MHz or 978 MHz frequencies. In use, ADS-B avionics consists of two operational capabilities, ADS-B “Out” and ADS-B “In”. ADS-B Out is the capability to periodically broadcast information from the aircraft, such as identification, current position, altitude, and velocity, through an onboard transmitter. ADS-B In is the capability to receive ADS-B data uplinked from FAA ground radios such as Flight Information Service-Broadcast (FIS-B) and Traffic Information Service-Broadcast (TIS-B) or from other ADS-B Out transmitting aircraft.
12-22. EMERGENCY LOCATOR TRANSMITTERS (ELT). The ELT must be evaluated in accordance with TSO-C91a, TSO-C126 for 406 MHz ELTs, or later TSOs issued for ELTs. ELT installations must be examined for potential operational problems at least once a year (§ 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. Tests should be conducted during the first 5 minutes after any hour. If operational tests must be made outside of this timeframe, they should be coordinated with the nearest FAA Control Tower or Flight Service Station (FSS). Tests should be no longer than three audible sweeps.
12-23. 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:
a. Proper Installation.
(1) 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.
(2) Remove the ELT from the mount and inspect the mounting hardware for proper installation and security.
(3) 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.
b. 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.
Par 12-23 Page 12-9
c. 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.
d. For a Sufficient Signal Radiated From its Antenna. Activate the ELT using the ON or ELT TEST switch. A low-quality AM broadcast radio receiver 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).
e. Ensure Proper Labeling and Positioning. Verify that all switches are properly labeled and positioned.
f. Record the Inspection. Record the inspection in the aircraft maintenance records according to 14 CFR part 43, § 43.9. We suggest the following:
I inspected the Make/Model ELT system in this aircraft according to applicable Aircraft and ELT manufacturer’s instructions and applicable FAA guidance and found that it meets the requirements of section (§) 91.207(d).
Signed: ______________________ C ertificate 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 (FCC) allows these tests only to be conducted within the first 5 minutes after any hour and is limited in three sweeps of the transmitter audio modulation.
12-24. FLIGHT DATA RECORDER (FDR). The FDR 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 underwater. Inspection/Operational checks include:
Par 12-24 Page 12-10
a. Check special sticker on front of the FDR for the date of the next tape replacement, if applicable.
b. Remove recorder magazine and inspect tape for the following:
(1) Broken or torn tape,
(2) Proper feed of tape, and
(3) All scribes were recording properly for approximately the last hour of flight.
c. Conditions for tape replacement (as applicable):
(1) There is less than 20 hours remaining in the magazine as read on the tape remaining indicator.
(2) Tape has run out.
(3) Broken tape.
(4) After hard landings and severe air turbulence have been encountered as reported by the pilots. After the same tape has been in use 1 year (12 months), it must be replaced.
(5) Ensure that a correlation test has been performed and then recorded in the aircraft records.
d. Refer to the specific equipment manufacturer’s manuals and procedures.
e. The state-of-the-art Solid State Flight Data Recorder (SSFDR) is a highly flexible model able to support a wide variety of 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 manufacturers for battery servicing. For maintenance information, refer to the equipment or aircraft manufacturer’s maintenance instruction manual.
12-25. COCKPIT VOICE RECORDERS (CVR). CVRs are very similar to FDRs. They look nearly identical and operate in almost the same way. CVRs 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.
a. Playback is possible only after the recorder is removed from the aircraft.
b. Refer to the specific equipment manufacturer’s manuals and procedures.
c. The Solid State Cockpit Voice Recorder system is composed of three essential components a solid state recorder, a control unit (remote mic amplifier), and an area microphone.
Par 12-26 Page 12-11 Also installed on one end of the recorder is an Underwater Locator Beacon (ULB). The recorder accepts four separate audio inputs: pilot, copilot, public address/third crewmember, and cockpit area microphone and where applicable, rotor speed input and FDR synchronization tone input.
For maintenance information, refer to the equipment manufacturer’s maintenance manual.
12-26. 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 ft 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 ft, 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.
12-27. RADOME INSPECTION. Inspection of aircraft having weather radar installations should include a visual check of the radome surface for signs of surface damage, holes, cracks, chipping, and peeling of paint, etc. Attach fittings and fastenings, neoprene erosion caps, and lightening strips, when installed, should also be inspected.
12-28. DATABUS. Databuses 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) integration. There are several types of databus 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 (DITS), Mark 33 which offers simple and affordable answers to data communications on aircraft.
12-29. ELECTRIC COMPATIBILITY. When replacing an instrument with one which provides additional functions or when adding new instruments, check the following electrical (where applicable) for compatibility:
a. Voltage (AC/DC).
b. Voltage polarity (DC).
c. Voltage phase(s) (AC).
d. Frequency (AC).
e. Grounding (AC/DC).
Par 12-29 Page 12-12
f. System impedance matching.
g. Compatibility with system to which connected.
12-30.–12-36. [RESERVED.]
Par 12-37 Page 12-13 SECTION 3. GROUND OPERATIONAL CHECKS FOR AVIONICS EQUIPMENT (NON-ELECTRICAL)
12-37. COMPASS SWING. This must be performed whenever any ferrous component of the system (i.e., flux valve compensator, or Standby Compass) is installed, removed, repaired, or a new compass is installed. The magnetic compass can be checked for accuracy by using a compass rose located on an airport. The compass swing is normally affected by placing the aircraft on various magnetic headings and comparing the deviations with those on the deviation cards. Refer to 14 CFR part 23, § 23.2500 and part 91, § 91.205(b), and the equipment or aircraft manufacturer’s manual.
a. A compass swing must be performed on the following occasions:
(1) When the accuracy of the compass is suspected.
(2) After any cockpit modification or major replacement involving ferrous metal.
(3) Whenever a compass has been subjected to a shock; for example, after a hard landing or turbulence.
(4) After aircraft has passed through a severe electrical storm.
(5) After lightning strike.
(6) Whenever a change is made to the electrical system.
(7) Whenever a change of cargo is likely to affect the compass.
(8) When an aircraft operation is changed to a different geographic location with a major change in magnetic deviation (e.g., from Miami, Florida to Fairbanks, Alaska).
(9) After aircraft has been parked on one heading for over a year.
(10) When flux valves are replaced.
b. Precautions.
(1) 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.
(2) Personnel engaged in the compensation of the compass shall remove all magnetic or ferrous material from their possession.
(3) Use only nonmagnetic tools when adjusting the compass.
(4) Position the aircraft at least 100 yards from any metal object.
(5) All equipment in the aircraft having any magnetic effect on the compass must be secured in the position occupied in normal flight.
Par 12-37 Page 12-14
c. Compass Swing Procedures.
(1) Have the aircraft taxied to the NORTH (0°) radial on the compass rose. Use a hairline sight compass (a reverse reading compass with a gun sight arrangement mounted on top of it) to place the aircraft in the general vicinity. With the aircraft facing North and the person in the cockpit running the engine(s) at 1,000 rpm, a mechanic, standing approximately 30 ft in front of the aircraft and 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: (1) 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. (2) Remember the hairline sight compass is only intended to be used as a general piece of test equipment.
(2) If the aircraft compass is not in alignment with the magnetic North of the master compass, correct the error by making small adjustments to the North-South brass adjustment screw with a nonmetallic screwdriver (made out of brass stock, or stainless steel welding rod).
Adjust the N-S compensator screw until the compass reads North (0°). Turn the aircraft until it is aligned with the East-West, pointing East. Adjust the E-W compensator screw until it reads 90°.
Continue by turning the aircraft South 180° and adjust the N-S screw to remove one-half of the South’s heading error. This will throw the North off, but the total North-South should be divided equally between the two headings. Turn the aircraft until it is heading West 270°, and adjust the E-W screw on the compensator to remove one-half of the West error. This should divide equally the total E-W error. The engine(s) should be running.
(3) With the aircraft heading West, start your calibration card here and record the magnetic heading of 270° and the compass reading with the avionics/electrical systems on then off. Turn the aircraft to align with each of the lines on the compass rose and record the compass reading every 30°. There should be not more than a plus or minus 10° difference between any of the compass’ heading and the magnetic heading of the aircraft.
(4) If the compass cannot be adjusted to meet the requirements, install another one.
NOTE: A common error that affects the compass’ accuracy is the mounting of a compass or instruments on or in the instrument panel using steel machine screws/nuts rather than brass hardware, magnetized control yoke, structural tubing, and improperly routed electrical wiring, which can cause unreasonable compass error.
(5) If the aircraft has an electrical system, two complete compass checks should 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
Par 12-37 Page 12-15 identical, the mechanic should make up two separate compass correction cards, one with all the equipment on and one with the equipment off.
(6) When the compass is satisfactorily swung, fill out the calibration card properly and put it in the holder in full view for the pilot’s reference.
d. Standby (wet) Compass. Adjustment and compensation of the Standby Compass may also be accomplished by using the “compass swing” method.
12-38. PNEUMATIC GYROS.
a. 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.
(1) 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.
(2) There are two sizes of venturi tubes: those which produce 4 inches of suction are used to drive the attitude gyros, and smaller tubes, which produce 2 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.)
FIGURE 12-1. VENTURI SYSTEM FOR PROVIDING AIRFLOW THROUGH GYRO
INSTRUMENTS
Par 12-38 Page 12-16
b. 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 4 inches of mercury (inHg)) 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.
(1) The early vacuum pumps were vane-type pumps of what is called the wet 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.)
FIGURE 12-2. INSTRUMENT VACUUM SYSTEM USING A WET-TYPE VACUUM PUMP
Par 12-38 Page 12-17
(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.
(3) 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 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.
c. Positive Pressure Systems. Above about 18,000 ft, 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 the 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.
d. 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 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 causes 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 particularly dusty conditions, especially if the occupants of the aircraft regularly smoke while flying. (See Figures 12-3 and 12-4.)
Par 12-38 Page 12-18 FIGURE 12-3. INSTRUMENT VACUUM SYSTEM USING A DRY-TYPE AIR PUMP F IGURE 12-4. INSTRUMENT PRESSURE SYSTEM USING A DRY-TYPE AIR PUMP 12 -39.–12-50. [RESERVED.]